Far Part25

Part 25: Airworthiness Standards: Transport Category Part 25

ASA 1

25

PART 25

AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY

AIRPLANES

S

PECIAL

F

EDERAL

A

VIATION

R

EGULATIONS

SFAR No. 13SFAR No. 109

Subpart A — General

Sec.25.1 Applicability.25.2 Special retroactive requirements.25.3 Special provisions for ETOPS type design

approvals.25.5 Incorporations by reference.

Subpart B — Flight

G

ENERAL

25.21 Proof of compliance.25.23 Load distribution limits.25.25 Weight limits.25.27 Center of gravity limits.25.29 Empty weight and corresponding center

of gravity.25.31 Removable ballast.25.33 Propeller speed and pitch limits.

P

ERFORMANCE

25.101 General.25.103 Stall speed.25.105 Takeoff.25.107 Takeoff speeds.25.109 Accelerate-stop distance.25.111 Takeoff path.25.113 Takeoff distance and takeoff run.25.115 Takeoff flight path.25.117 Climb: general.25.119 Landing climb: All-engines-operating.25.121 Climb: One-engine-inoperative.25.123 En route flight paths.25.125 Landing.

C

ONTROLLABILITY

AND

M

ANEUVERABILITY

25.143 General.25.145 Longitudinal control.25.147 Directional and lateral control.25.149 Minimum control speed.

T

RIM

25.161 Trim.

S

TABILITY

25.171 General.25.173 Static longitudinal stability.

25.175 Demonstration of static longitudinal stability.

25.177 Static lateral-directional stability.25.181 Dynamic stability.

S

TALLS

25.201 Stall demonstration.25.203 Stall characteristics.25.207 Stall warning.

G

ROUND

AND

W

ATER

H

ANDLING

C

HARACTERISTICS

25.231 Longitudinal stability and control.25.233 Directional stability and control.25.235 Taxiing condition.25.237 Wind velocities.25.239 Spray characteristics, control, and

stability on water.

M

ISCELLANEOUS

F

LIGHT

R

EQUIREMENTS

25.251 Vibration and buffeting.25.253 High-speed characteristics.25.255 Out-of-trim characteristics.

Subpart C — Structure

G

ENERAL

25.301 Loads.25.303 Factor of safety.25.305 Strength and deformation.25.307 Proof of structure.

F

LIGHT

L

OADS

25.321 General.

F

LIGHT

M

ANEUVER

AND

G

UST

C

ONDITIONS

25.331 Symmetric maneuvering conditions.25.333 Flight maneuvering envelope.25.335 Design airspeeds.25.337 Limit maneuvering load factors.25.341 Gust and turbulence loads.25.343 Design fuel and oil loads.25.345 High lift devices.25.349 Rolling conditions.25.351 Yaw maneuver conditions.

S

UPPLEMENTARY

C

ONDITIONS

25.361 Engine torque.25.363 Side load on engine and auxiliary power

unit mounts.25.365 Pressurized compartment loads.25.367 Unsymmetrical loads due to engine

failure.25.371 Gyroscopic loads.25.373 Speed control devices.

Part 25 Federal Aviation Regulations

2 ASA

C

ONTROL

S

URFACE

AND

S

YSTEM

L

OADS

25.391 Control surface loads: general.25.393 Loads parallel to hinge line.25.395 Control system.25.397 Control system loads.25.399 Dual control system.25.405 Secondary control system.25.407 Trim tab effects.25.409 Tabs.25.415 Ground gust conditions.25.427 Unsymmetrical loads.25.445 Auxiliary aerodynamic surfaces.25.457 Wing flaps.25.459 Special devices.

G

ROUND

L

OADS

25.471 General.25.473 Landing load conditions and

assumptions.25.477 Landing gear arrangement.25.479 Level landing conditions.25.481 Tail-down landing conditions.25.483 One-gear landing conditions.25.485 Side load conditions.25.487 Rebound landing condition.25.489 Ground handling conditions.25.491 Taxi, takeoff and landing roll.25.493 Braked roll conditions.25.495 Turning.25.497 Tail-wheel yawing.25.499 Nose-wheel yaw and steering.25.503 Pivoting.25.507 Reversed braking.25.509 Towing loads.25.511 Ground load: unsymmetrical loads on

multiple-wheel units.25.519 Jacking and tie-down provisions.

W

ATER

L

OADS

25.521 General.25.523 Design weights and center of gravity

positions.25.525 Application of loads.25.527 Hull and main float load factors.25.529 Hull and main float landing conditions.25.531 Hull and main float takeoff condition.25.533 Hull and main float bottom pressures.25.535 Auxiliary float loads.25.537 Seawing loads.

E

MERGENCY

L

ANDING

C

ONDITIONS

25.561 General.25.562 Emergency landing dynamic conditions.25.563 Structural ditching provisions.

F

ATIGUE

E

VALUATION

25.571 Damage-tolerance and fatigue evaluation of structure.

L

IGHTNING

P

ROTECTION

25.581 Lightning protection.

Subpart D — Design and Construction

G

ENERAL

25.601 General.25.603 Materials.25.605 Fabrication methods.25.607 Fasteners.25.609 Protection of structure.25.611 Accessibility provisions.25.613 Material strength properties and material

design values.25.619 Special factors.25.621 Casting factors.25.623 Bearing factors.25.625 Fitting factors.25.629 Aeroelastic stability requirements.25.631 Bird strike damage.

C

ONTROL

S

URFACES

25.651 Proof of strength.25.655 Installation.25.657 Hinges.

C

ONTROL

S

YSTEMS

25.671 General.25.672 Stability augmentation and automatic

and power-operated systems.25.675 Stops.25.677 Trim systems.25.679 Control system gust locks.25.681 Limit load static tests.25.683 Operation tests.25.685 Control system details.25.689 Cable systems.25.693 Joints.25.697 Lift and drag devices, controls.25.699 Lift and drag device indicator.25.701 Flap and slat interconnection.25.703 Takeoff warning system.

L

ANDING

G

EAR

25.721 General.25.723 Shock absorption tests.25.725 [Reserved]25.727 [Reserved]25.729 Retracting mechanism.25.731 Wheels.25.733 Tires.25.735 Brakes and braking systems.25.737 Skis.


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25

F

LOATS

AND

H

ULLS

25.751 Main float buoyancy.25.753 Main float design.25.755 Hulls.

P

ERSONNEL

AND

C

ARGO

A

CCOMMODATIONS

25.771 Pilot compartment.25.772 Pilot compartment doors.25.773 Pilot compartment view.25.775 Windshields and windows.25.777 Cockpit controls.25.779 Motion and effect of cockpit controls.25.781 Cockpit control knob shape.25.783 Fuselage doors.25.785 Seats, berths, safety belts, and

harnesses.25.787 Stowage compartments.25.789 Retention of items of mass in passenger

and crew compartments and galleys.25.791 Passenger information signs and

placards.25.793 Floor surfaces.25.795 Security considerations.

E

MERGENCY

P

ROVISIONS

25.801 Ditching.25.803 Emergency evacuation.25.807 Emergency exits.25.809 Emergency exit arrangement.25.810 Emergency egress assist means and

escape routes.25.811 Emergency exit marking.25.812 Emergency lighting.25.813 Emergency exit access.25.815 Width of aisle.25.817 Maximum number of seats abreast.25.819 Lower deck surface compartments

(including galleys).25.820 Lavatory doors.

V

ENTILATION

AND

H

EATING

25.831 Ventilation.25.832 Cabin ozone concentration.25.833 Combustion heating systems.

P

RESSURIZATION

25.841 Pressurized cabins.25.843 Tests for pressurized cabins.

F

IRE

P

ROTECTION

25.851 Fire extinguishers.25.853 Compartment interiors.25.854 Lavatory fire protection.25.855 Cargo or baggage compartments.25.856 Thermal/Acoustic insulation materials.25.857 Cargo compartment classification.25.858 Cargo or baggage compartment smoke

or fire detection systems.

25.859 Combustion heater fire protection.25.863 Flammable fluid fire protection.25.865 Fire protection of flight controls, engine

mounts, and other flight structure.25.867 Fire protection: other components.25.869 Fire protection: systems.

M

ISCELLANEOUS

25.871 Leveling means.25.875 Reinforcement near propellers.25.899 Electrical bonding and protection against

static electricity.

Subpart E — Powerplant

G

ENERAL

25.901 Installation.25.903 Engines.25.904 Automatic takeoff thrust control system

(ATTCS).25.905 Propellers.25.907 Propeller vibration and fatigue.25.925 Propeller clearance.25.929 Propeller deicing.25.933 Reversing systems.25.934 Turbojet engine thrust reverser

system tests.25.937 Turbopropeller-drag limiting systems.25.939 Turbine engine operating characteristics.25.941 Inlet, engine, and exhaust compatibility.25.943 Negative acceleration.25.945 Thrust or power augmentation system.

F

UEL

S

YSTEM

25.951 General.25.952 Fuel system analysis and test.25.953 Fuel system independence.25.954 Fuel system lightning protection.25.955 Fuel flow.25.957 Flow between interconnected tanks.25.959 Unusable fuel supply.25.961 Fuel system hot weather operation.25.963 Fuel tanks: general.25.965 Fuel tank tests.25.967 Fuel tank installations.25.969 Fuel tank expansion space.25.971 Fuel tank sump.25.973 Fuel tank filler connection.25.975 Fuel tank vents and carburetor

vapor vents.25.977 Fuel tank outlet.25.979 Pressure fueling system.25.981 Fuel tank explosion prevention.

F

UEL

S

YSTEM

C

OMPONENTS

25.991 Fuel pumps.25.993 Fuel system lines and fittings.25.994 Fuel system components.25.995 Fuel valves.

Part 25 Federal Aviation Regulations

4 ASA

25.997 Fuel strainer or filter.25.999 Fuel system drains.25.1001 Fuel jettisoning system.

O

IL

S

YSTEM

25.1011 General.25.1013 Oil tanks.25.1015 Oil tank tests.25.1017 Oil lines and fittings.25.1019 Oil strainer or filter.25.1021 Oil system drains.25.1023 Oil radiators.25.1025 Oil valves.25.1027 Propeller feathering system.

C

OOLING

25.1041 General.25.1043 Cooling tests.25.1045 Cooling test procedures.

I

NDUCTION

S

YSTEM

25.1091 Air induction.25.1093 Induction system icing protection.25.1101 Carburetor air preheater design.25.1103 Induction system ducts and air

duct systems.25.1105 Induction system screens.25.1107 Inter-coolers and after-coolers.

E

XHAUST

S

YSTEM

25.1121 General.25.1123 Exhaust piping.25.1125 Exhaust heat exchangers.25.1127 Exhaust driven turbo-superchargers.

P

OWERPLANT

C

ONTROLS

AND

A

CCESSORIES

25.1141 Powerplant controls: general.25.1142 Auxiliary power unit controls.25.1143 Engine controls.25.1145 Ignition switches.25.1147 Mixture controls.25.1149 Propeller speed and pitch controls.25.1153 Propeller feathering controls.25.1155 Reverse thrust and propeller pitch

settings below the flight regime.25.1157 Carburetor air temperature controls.25.1159 Supercharger controls.25.1161 Fuel jettisoning system controls.25.1163 Powerplant accessories.25.1165 Engine ignition systems.25.1167 Accessory gearboxes.

P

OWERPLANT

F

IRE

P

ROTECTION

25.1181 Designated fire zones; regions included.

25.1182 Nacelle areas behind firewalls, and engine pod attaching structures containing flammable fluid lines.

25.1183 Flammable fluid-carrying components.25.1185 Flammable fluids.25.1187 Drainage and ventilation of fire zones.25.1189 Shutoff means.25.1191 Firewalls.25.1192 Engine accessory section diaphragm.25.1193 Cowling and nacelle skin.25.1195 Fire extinguishing systems.25.1197 Fire extinguishing agents.25.1199 Extinguishing agent containers.25.1201 Fire extinguishing system materials.25.1203 Fire detector system.25.1207 Compliance.

Subpart F — Equipment

G

ENERAL

25.1301 Function and installation.25.1303 Flight and navigation instruments.25.1305 Powerplant instruments.25.1307 Miscellaneous equipment.25.1309 Equipment, systems, and installations.25.1310 Power source capacity and distribution.25.1316 System lightning protection.25.1317 High-Intensity Radiated Fields (HIRF)

Protection.

I

NSTRUMENTS

: I

NSTALLATION

25.1321 Arrangement and visibility.25.1322 Warning, caution, and advisory lights.25.1323 Airspeed indicating system.25.1325 Static pressure systems.25.1326 Pitot heat indication systems.25.1327 Magnetic direction indicator.25.1329 Flight guidance system.25.1331 Instruments using a power supply.25.1333 Instrument systems.25.1335 Flight director systems.25.1337 Powerplant instruments.

E

LECTRICAL

S

YSTEMS

AND

E

QUIPMENT

25.1351 General.25.1353 Electrical equipment and installations.25.1355 Distribution system.25.1357 Circuit protective devices.25.1360 Precautions against injury.25.1362 Electrical supplies for emergency

conditions.25.1363 Electrical system tests.25.1365 Electrical appliances, motors, and

transformers.

L

IGHTS

25.1381 Instrument lights.25.1383 Landing lights.25.1385 Position light system installation.25.1387 Position light system dihedral angles.25.1389 Position light distribution and intensities.


ASA 5

25

25.1391 Minimum intensities in the horizontal plane of forward and rear position lights.

25.1393 Minimum intensities in any vertical plane of forward and rear position lights.

25.1395 Maximum intensities in overlapping beams of forward and rear position lights.

25.1397 Color specifications.25.1399 Riding light.25.1401 Anticollision light system.25.1403 Wing icing detection lights.

SAFETY EQUIPMENT

25.1411 General.25.1415 Ditching equipment.25.1419 Ice protection.25.1421 Megaphones.25.1423 Public address system.

MISCELLANEOUS EQUIPMENT

25.1431 Electronic equipment.25.1433 Vacuum systems.25.1435 Hydraulic systems.25.1438 Pressurization and pneumatic systems.25.1439 Protective breathing equipment.25.1441 Oxygen equipment and supply.25.1443 Minimum mass flow of supplemental

oxygen.25.1445 Equipment standards for the oxygen

distributing system.25.1447 Equipment standards for oxygen

dispensing units.25.1449 Means for determining use of oxygen.25.1450 Chemical oxygen generators.25.1453 Protection of oxygen equipment

from rupture.25.1455 Draining of fluids subject to freezing.25.1457 Cockpit voice recorders.25.1459 Flight data recorders.25.1461 Equipment containing high energy

rotors.

Subpart G —Operating Limitations and Information

25.1501 General.

OPERATING LIMITATIONS

25.1503 Airspeed limitations: general.25.1505 Maximum operating limit speed.25.1507 Maneuvering speed.25.1511 Flap extended speed.25.1513 Minimum control speed.25.1515 Landing gear speeds.25.1516 Other speed limitations.25.1517 Rough air speed, VRA.

25.1519 Weight, center of gravity, and weight distribution.

25.1521 Powerplant limitations.25.1522 Auxiliary power unit limitations.25.1523 Minimum flight crew.25.1525 Kinds of operation.25.1527 Ambient air temperature and operating

altitude.25.1529 Instructions for Continued

Airworthiness.25.1531 Maneuvering flight load factors.25.1533 Additional operating limitations.25.1535 ETOPS approval.

MARKINGS AND PLACARDS

25.1541 General.25.1543 Instrument markings: general.25.1545 Airspeed limitation information.25.1547 Magnetic direction indicator.25.1549 Powerplant and auxiliary power

unit instruments.25.1551 Oil quantity indication.25.1553 Fuel quantity indicator.25.1555 Control markings.25.1557 Miscellaneous markings and placards.25.1561 Safety equipment.25.1563 Airspeed placard.

AIRPLANE FLIGHT MANUAL

25.1581 General.25.1583 Operating limitations.25.1585 Operating procedures.25.1587 Performance information.

Subpart H — Electrical Wiring Interconnection Systems (EWIS)

25.1701 Definition.25.1703 Function and installation: EWIS.25.1705 Systems and functions: EWIS.25.1707 System separation: EWIS.25.1709 System safety: EWIS.25.1711 Component identification: EWIS.25.1713 Fire protection: EWIS.25.1715 Electrical bonding and protection

against static electricity: EWIS.25.1717 Circuit protective devices: EWIS.25.1719 Accessibility provisions: EWIS.25.1721 Protection of EWIS.25.1723 Flammable fluid fire protection: EWIS.25.1725 Powerplants: EWIS.25.1727 Flammable fluid shutoff means: EWIS.25.1729 Instructions for Continued

Airworthiness: EWIS.25.1731 Powerplant and APU fire detector

system: EWIS.25.1733 Fire detector systems, general: EWIS.

SFAR No. 13 to Part 25 Federal Aviation Regulations

6 ASA

APPENDICES TO PART 25Appendix A to Part 25Appendix B to Part 25Appendix C to Part 25Appendix D to Part 25Appendix E to Part 25Appendix F to Part 25Appendix G to Part 25 —

Continuous Gust Design CriteriaAppendix H to Part 25 —

Instructions for Continued AirworthinessAppendix I to Part 25 —

Installation of an Automatic Takeoff Thrust Control System (ATTCS)

Appendix J to Part 25 — Emergency EvacuationAppendix K to Part 25 — Extended Operations

(ETOPS)Appendix L to Part 25—HIRF Environments and

Equipment HIRF Test LevelsAppendix M to Part 25—Fuel Tank System

Flammability Reduction MeansAppendix N to Part 25—Fuel Tank Flammability

Exposure and Reliability Analysis

Authority: 49 U.S.C 106(g), 40113, 44701, 44702 and44704.

Source: Docket No. 5066, 29 FR 18291, Dec. 24, 1964,unless otherwise noted.

SFAR NO. 13 TO PART 251. Applicability. Contrary provisions of the Civil

Air Regulations regarding certification notwith-standing,1 this regulation shall provide the basisfor approval by the Administrator of modificationsof individual Douglas DC-3 and Lockheed L-18airplanes subsequent to the effective date of thisregulation.1 It is not intended to waive compliance with suchairworthiness requirements as are included in theoperating parts of the Civil Air Regulations forspecific types of operation.

2. General modifications. Except as modified insections 3 and 4 of this regulation, an applicantfor approval of modifications to a DC-3 or L-18 air-plane which result in changes in design or inchanges to approved limitations shall show thatthe modifications were accomplished in accor-dance with the rules of either Part 4a or Part 4b ineffect on September 1, 1953, which are applica-ble to the modification being made: Provided,That an applicant may elect to accomplish a mod-ification in accordance with the rules of Part 4b ineffect on the date of application for the modifica-tion in lieu of Part 4a or Part 4b as in effect on

September 1, 1953: And provided further, Thateach specific modification must be accomplishedin accordance with all of the provisions containedin the elected rules relating to the particular modi-fication.

3. Specific conditions for approval. An appli-cant for any approval of the following specificchanges shall comply with section 2 of this regu-lation as modified by the applicable provisions ofthis section.

(a) Increase in take-off power limitation —1,200to 1,350 horsepower. The engine take-off powerlimitation for the airplane may be increased tomore than 1,200 horsepower but not to more than1,350 horsepower per engine if the increase inpower does not adversely affect the flight charac-teristics of the airplane.

(b) Increase in take-off power limitation to morethan 1,350 horsepower. The engine take-offpower limitation for the airplane may be increasedto more than 1,350 horsepower per engine if com-pliance is shown with the flight characteristics andground handling requirements of Part 4b.

(c) Installation of engines of not more than1,830 cubic inches displacement and not having acertificated take-off rating of more than 1,350horsepower. Engines of not more than 1,830 cu-bic inches displacement and not having a certifi-cated take-off rating of more than 1,350 horse-power which necessitate a major modification ofredesign of the engine installation may be in-stalled, if the engine fire prevention and fire pro-tection are equivalent to that on the prior engineinstallation.

(d) Installation of engines of more than 1,830cubic inches displacement or having certificatedtake-off rating of more than 1,350 horsepower.Engines of more than 1,830 cubic inches dis-placement or having certificated take-off rating ofmore than 1,350 horsepower may be installed ifcompliance is shown with the engine installationrequirements of Part 4b: Provided, That where lit-eral compliance with the engine installation re-quirements of Part 4b is extremely difficult to ac-complish and would not contribute materially tothe objective sought, and the Administrator findsthat the experience with the DC-3 or L-18 air-planes justifies it, he is authorized to accept suchmeasures of compliance as he finds will effec-tively accomplish the basic objective.

4. Establishment of new maximum certificatedweights. An applicant for approval of new maxi-mum certificated weights shall apply for anamendment of the airworthiness certificate of theairplane and shall show that the weights soughthave been established, and the appropriate man-ual material obtained, as provided in this section.

Part 25: Airworthiness Standards: Transport Category SFAR No. 109 to Part 25

ASA 7

25

Note: Transport category performance require-ments result in the establishment of maximumcertificated weights for various altitudes.

(a) Weights — 25,200 to 26,900 for the DC-3and 18,500 to 19,500 for the L-18. New maximumcertificated weights of more than 25,200 but notmore than 26,900 pounds for DC-3 and more than18,500 but not more than 19,500 pounds for L-18airplanes may be established in accordance withthe transport category performance requirementsof either Part 4a or Part 4b, if the airplane at thenew maximum weights can meet the structural re-quirements of the elected part.

(b) Weights of more than 26,900 for the DC-3and 19,500 for the L-18. New maximum certifi-cated weights of more than 26,900 pounds forDC-3 and 19,500 pounds for L-18 airplanes shallbe established in accordance with the structuralperformance, flight characteristics, and groundhandling requirements of Part 4b: Provided, Thatwhere literal compliance with the structural re-quirements of Part 4b is extremely difficult to ac-complish and would not contribute materially tothe objective sought, and the Administrator findsthat the experience with the DC-3 or L-18 air-planes justifies it, he is authorized to accept suchmeasures of compliance as he finds will effec-tively accomplish the basic objective.

(c) Airplane flight manual-performance operat-ing information. An approved airplane flight man-ual shall be provided for each DC-3 and L-18 air-plane which has had new maximum certificatedweights established under this section. The air-plane flight manual shall contain the applicableperformance information prescribed in that part ofthe regulations under which the new certificatedweights were established and such additional in-formation as may be necessary to enable the ap-plication of the take-off, en route, and landing lim-itations prescribed for transport category air-planes in the operating parts of the Civil AirRegulations.

(d) Performance operating limitations. Each air-plane for which new maximum certificated weightsare established in accordance with paragraphs (a)or (b) of this section shall be considered a trans-port category airplane for the purpose of comply-ing with the performance operating limitations ap-plicable to the operations in which it is utilized.

5. Reference. Unless otherwise provided, allreferences in this regulation to Part 4a and Part4b are those parts of the Civil Air Regulations ineffect on September 1, 1953.

This regulation supersedes Special Civil AirRegulation SR-398 and shall remain effective un-til superseded or rescinded by the Board.

[19 FR 5039, Aug. 11, 1954. Redesignated at 29 FR19099, Dec. 30, 1964]

SFAR NO. 109 TO PART 251. Applicability. Contrary provisions of 14

CFR parts 21, 25, and 119 of this chapter notwith-standing, an applicant is entitled to an amendedtype certificate or supplemental type certificate inthe transport category, if the applicant complieswith all applicable provisions of this SFAR.

OPERATIONS

2. General.(a) The passenger capacity may not exceed 60.

If more than 60 passenger seats are installed,then:

(1) If the extra seats are not suitable for occu-pancy during taxi, takeoff and landing, each extraseat must be clearly marked (e.g., a placard onthe top of an armrest, or a placard sewn into thetop of the back cushion) that the seat is not to beoccupied during taxi, takeoff and landing.

(2) If the extra seats are suitable for occupancyduring taxi, takeoff and landing (i.e., meet all thestrength and passenger injury criteria in part 25),then a note must be included in the LimitationsSection of the Airplane Flight Manual that thereare extra seats installed but that the number ofpassengers on the airplane must not exceed 60.Additionally, there must be a placard installed ad-jacent to each door that can be used as a passen-ger boarding door that states that the maximumpassenger capacity is 60. The placard must beclearly legible to passengers entering the air-plane.

(b) For airplanes outfitted with interior doorsunder paragraph 10 of this SFAR, the airplaneflight manual (AFM) must include an appropriatelimitation that the airplane must be staffed with atleast the following number of flight attendants whomeet the requirements of 14 CFR 91.533(b):

(1) The number of flight attendants required by§91.533(a)(1) and (2) of this chapter, and

(2) At least one flight attendant if the airplanemodel was originally certified for 75 passengersor more.

(c) The AFM must include appropriate limita-tion(s) to require a preflight passenger briefingdescribing the appropriate functions to be per-formed by the passengers and the relevant fea-tures of the airplane to ensure the safety of thepassengers and crew.

(d) The airplane may not be offered for com-mon carriage or operated for hire. The operatinglimitations section of the AFM must be revised toprohibit any operations involving the carriage ofpersons or property for compensation or hire. Theoperators may receive remuneration to the extentconsistent with parts 125 and 91, subpart F, ofthis chapter.

(e) A placard stating that “Operations involvingthe carriage of persons or property for compensa-


8 ASA

tion or hire are prohibited,” must be located in thearea of the Airworthiness Certificate holder at theentrance to the flightdeck.

(f) For passenger capacities of 45 to 60 pas-sengers, analysis must be submitted that demon-strates that the airplane can be evacuated in lessthan 90 seconds under the conditions specified in§25.803 and Appendix J to part 25.

(g) In order for any airplane certified under thisSFAR to be placed in part 135 or part 121 opera-tions, the airplane must be brought back into fullcompliance with the applicable operational part.

EQUIPMENT AND DESIGN

3. General. Unless otherwise noted, compli-ance is required with the applicable certificationbasis for the airplane. Some provisions of thisSFAR impose alternative requirements to certainairworthiness standards that do not apply to air-planes certificated to earlier standards. Those air-planes with an earlier certification basis are notrequired to comply with those alternative require-ments.

4. Occupant Protection.(a) Firm Handhold. In lieu of the requirements

of §25.785(j), there must be means provided toenable persons to steady themselves in moder-ately rough air while occupying aisles that arealong the cabin sidewall, or where practicable,bordered by seats (seat backs providing a 25-pound minimum breakaway force are an accept-able means of compliance).

(b) Injury criteria for multiple occupancy side-facing seats. The following requirements are onlyapplicable to airplanes that are subject to§25.562.

(1) Existing Criteria. All injury protection criteriaof §25.562(c)(1) through (c)(6) apply to the occu-pants of side-facing seating. The Head Injury Cri-terion (HIC) assessments are only required forhead contact with the seat and/or adjacent struc-tures.

(2) Body-to-Body Contact. Contact betweenthe head, pelvis, torso or shoulder area of one An-thropomorphic Test Dummy (ATD) with the head,pelvis, torso or shoulder area of the ATD in the ad-jacent seat is not allowed during the tests con-ducted in accordance with §25.562(b)(1) and(b)(2). Contact during rebound is allowed.

(3) Thoracic Trauma. If the torso of an ATD atthe forward-most seat place impacts the seatand/or adjacent structure during testing, compli-ance with the Thoracic Trauma Index (TTI) injurycriterion must be substantiated by dynamic test orby rational analysis based on previous test(s) of asimilar seat installation. TTI data must be ac-quired with a Side Impact Dummy (SID), as de-fined by 49 CFR part 572, subpart F, or an equiv-alent ATD or a more appropriate ATD and must be

processed as defined in Federal Motor VehicleSafety Standards (FMVSS) part 571.214, sectionS6.13.5 (49 CFR 571.214). The TTI must be lessthan 85, as defined in 49 CFR part 572, subpart F.Torso contact during rebound is acceptable andneed not be measured.

(4) Pelvis. If the pelvis of an ATD at any seatplace impacts seat and/or adjacent structure dur-ing testing, pelvic lateral acceleration injury crite-ria must be substantiated by dynamic test or byrational analysis based on previous test(s) of asimilar seat installation. Pelvic lateral accelerationmay not exceed 130g. Pelvic acceleration datamust be processed as defined in FMVSS part571.214, section S6.13.5 (49 CFR 571.214).

(5) Body-to-Wall/Furnishing Contact. If the seatis installed aft of a structure—such as an interiorwall or furnishing that may contact the pelvis, up-per arm, chest, or head of an occupant seatednext to the structure—the structure or a conserva-tive representation of the structure and its stiff-ness must be included in the tests. It is recom-mended, but not required, that the contact surfaceof the actual structure be covered with at least twoinches of energy absorbing protective padding(foam or equivalent) such as Ensolite.

(6) Shoulder Strap Loads. Where upper torsostraps (shoulder straps) are used for sofa occu-pants, the tension loads in individual straps maynot exceed 1,750 pounds. If dual straps are usedfor restraining the upper torso, the total strap ten-sion loads may not exceed 2,000 pounds.

(7) Occupant Retention. All side-facing seatsrequire end closures or other means to preventthe ATD’s pelvis from translating beyond the endof the seat at any time during testing.

(8) Test Parameters.(i) All seat positions need to be occupied by

ATDs for the longitudinal tests.(ii) A minimum of one longitudinal test, con-

ducted in accordance with the conditions speci-fied in §25.562(b)(2), is required to assess the in-jury criteria as follows. Note that if a seat is in-stalled aft of structure (such as an interior wall orfurnishing) that does not have a homogeneoussurface, an additional test or tests may be re-quired to demonstrate that the injury criteria aremet for the area which an occupant could contact.For example, different yaw angles could result indifferent injury considerations and may requireseparate tests to evaluate.

(A) For configurations without structure (suchas a wall or bulkhead) installed directly forward ofthe forward seat place, Hybrid II ATDs or equiva-lent must be in all seat places.

(B) For configurations with structure (such as awall or bulkhead) installed directly forward of theforward seat place, a side impact dummy orequivalent ATD or more appropriate ATD must be

Part 25: Airworthiness Standards: Transport Category SFAR No. 109 to Part 25

ASA 9

25

in the forward seat place and a Hybrid II ATD orequivalent must be in all other seat places.

(C) The test may be conducted with or withoutdeformed floor.

(D) The test must be conducted with either noyaw or 10 degrees yaw for evaluating occupant in-jury. Deviating from the no yaw condition may notresult in the critical area of contact not being eval-uated. The upper torso restraint straps, where in-stalled, must remain on the occupant’s shoulderduring the impact condition of §25.562(b)(2).

(c) For the vertical test, conducted in accor-dance with the conditions specified in§25.562(b)(1), Hybrid II ATDs or equivalent mustbe used in all seat positions.

5. Direct View. In lieu of the requirements of§25.785(h)(2), to the extent practical without com-promising proximity to a required floor level emer-gency exit, the majority of installed flight attendantseats must be located to face the cabin area forwhich the flight attendant is responsible.

6. Passenger Information Signs. Compliancewith §25.791 is required except that for§25.791(a), when smoking is to be prohibited, no-tification to the passengers may be provided by asingle placard so stating, to be conspicuously lo-cated inside the passenger compartment, easilyvisible to all persons entering the cabin in the im-mediate vicinity of each passenger entry door.

7. Distance Between Exits. For an airplanethat is required to comply with §25.807(f)(4), in ef-fect as of July 24, 1989, which has more than onepassenger emergency exit on each side of the fu-selage, no passenger emergency exit may bemore than 60 feet from any adjacent passengeremergency exit on the same side of the samedeck of the fuselage, as measured parallel to theairplane’s longitudinal axis between the nearestexit edges, unless the following conditions aremet:

(a) Each passenger seat must be locatedwithin 30 feet from the nearest exit on each sideof the fuselage, as measured parallel to the air-plane’s longitudinal axis, between the nearest exitedge and the front of the seat bottom cushion.

(b) The number of passenger seats located be-tween two adjacent pairs of emergency exits(commonly referred to as a passenger zone) orbetween a pair of exits and a bulkhead or a com-partment door (commonly referred to as a “dead-end zone”), may not exceed the following:

(1) For zones between two pairs of exits, 50percent of the combined rated capacity of the twopairs of emergency exits.

(2) For zones between one pair of exits and abulkhead, 40 percent of the rated capacity of thepair of emergency exits.

(c) The total number of passenger seats in theairplane may not exceed 33 percent of the maxi-mum seating capacity for the airplane model us-

ing the exit ratings listed in §25.807(g) for theoriginal certified exits or the maximum allowableafter modification when exits are deactivated,whichever is less.

(d) A distance of more than 60 feet betweenadjacent passenger emergency exits on the sameside of the same deck of the fuselage, as mea-sured parallel to the airplane’s longitudinal axisbetween the nearest exit edges, is allowed onlyonce on each side of the fuselage.

8. Emergency Exit Signs. In lieu of the re-quirements of §25.811(d)(1) and (2) a single signat each exit may be installed provided:

(a) The sign can be read from the aisle while di-rectly facing the exit, and

(b) The sign can be read from the aisle adja-cent to the passenger seat that is farthest fromthe exit and that does not have an interveningbulkhead/divider or exit.

9. Emergency Lighting.(a) Exit Signs. In lieu of the requirements of

§25.812(b)(1), for airplanes that have a passen-ger seating configuration, excluding pilot seats, of19 seats or less, the emergency exit signs re-quired by §25.811(d)(1), (2), and (3) must havered letters at least 1-inch high on a white back-ground at least 2 inches high. These signs may beinternally electrically illuminated, or self illumi-nated by other than electrical means, with an ini-tial brightness of at least 160 microlamberts. Thecolor may be reversed in the case of a sign that isself-illuminated by other than electrical means.

(b) Floor Proximity Escape Path Marking. Inlieu of the requirements of §25.812(e)(1), forcabin seating compartments that do not have themain cabin aisle entering and exiting the compart-ment, the following are applicable:

(1) After a passenger leaves any passengerseat in the compartment, he/she must be able toexit the compartment to the main cabin aisle us-ing only markings and visual features not morethat 4 feet above the cabin floor, and

(2) Proceed to the exits using the marking sys-tem necessary to accomplish the actions in§25.812(e)(1) and (e)(2).

(c) Transverse Separation of the Fuselage. Inthe event of a transverse separation of the fuse-lage, compliance must be shown with §25.812(l)except as follows:

(1) For each airplane type originally type certif-icated with a maximum passenger seating capac-ity of 9 or less, not more than 50 percent of allelectrically illuminated emergency lights requiredby §25.812 may be rendered inoperative in addi-tion to the lights that are directly damaged by theseparation.

(2) For each airplane type originally type certif-icated with a maximum passenger seating capac-ity of 10 to 19, not more than 33 percent of allelectrically illuminated emergency lights required


10 ASA

by §25.812 may be rendered inoperative in addi-tion to the lights that are directly damaged by theseparation.

10. Interior doors. In lieu of the requirementsof §25.813(e), interior doors may be installed be-tween passenger seats and exits, provided thefollowing requirements are met.

(a) Each door between any passenger seat,occupiable for taxi, takeoff, and landing, and anyemergency exit must have a means to signal tothe flightcrew, at the flightdeck, that the door is inthe open position for taxi, takeoff and landing.

(b) Appropriate procedures/limitations must beestablished to ensure that any such door is in theopen configuration for takeoff and landing.

(c) Each door between any passenger seat andany exit must have dual means to retain it in theopen position, each of which is capable of react-ing the inertia loads specified in §25.561.

(d) Doors installed across a longitudinal aislemust translate laterally to open and close, e.g.,pocket doors.

(e) Each door between any passenger seatand any exit must be frangible in either direction.

(f) Each door between any passenger seat andany exit must be operable from either side, and ifa locking mechanism is installed, it must be capa-ble of being unlocked from either side without theuse of special tools.

11. Width of Aisle. Compliance is requiredwith §25.815, except that aisle width may be re-duced to 0 inches between passenger seats dur-ing in-flight operations only, provided that the ap-plicant demonstrates that all areas of the cabinare easily accessible by a crew member in theevent of an emergency (e.g., in-flight fire, decom-pression). Additionally, instructions must be pro-vided at each passenger seat for restoring theaisle width required by §25.815. Procedures mustbe established and documented in the AFM to en-sure that the required aisle widths are providedduring taxi, takeoff, and landing.

12. Materials for Compartment Interiors.Compliance is required with the applicable provi-sions of §25.853, except that compliance with Ap-pendix F, parts IV and V, to part 25, need not bedemonstrated if it can be shown by test or a com-bination of test and analysis that the maximumtime for evacuation of all occupants does not ex-ceed 45 seconds under the conditions specified inAppendix J to part 25.

13. Fire Detection. For airplanes with a typecertificated passenger capacity of 20 or more,there must be means that meet the requirementsof §25.858(a) through (d) to signal the flightcrewin the event of a fire in any isolated room not oc-cupiable for taxi, takeoff and landing, which canbe closed off from the rest of the cabin by a door.The indication must identify the compartmentwhere the fire is located. This does not apply to

lavatories, which continue to be governed by§25.854.

14. Cooktops. Each cooktop must be de-signed and installed to minimize any potentialthreat to the airplane, passengers, and crew.Compliance with this requirement must be foundin accordance with the following criteria:

(a) Means, such as conspicuous burner-on in-dicators, physical barriers, or handholds, must beinstalled to minimize the potential for inadvertentpersonnel contact with hot surfaces of both thecooktop and cookware. Conditions of turbulencemust be considered.

(b) Sufficient design means must be includedto restrain cookware while in place on the cook-top, as well as representative contents, e.g., soup,sauces, etc., from the effects of flight loads andturbulence. Restraints must be provided to pre-clude hazardous movement of cookware and con-tents. These restraints must accommodate anycookware that is identified for use with the cook-top. Restraints must be designed to be easily uti-lized and effective in service. The cookware re-straint system should also be designed so that itwill not be easily disabled, thus rendering it unus-able. Placarding must be installed which prohibitsthe use of cookware that cannot be accommo-dated by the restraint system.

(c) Placarding must be installed which prohibitsthe use of cooktops (i.e., power on any burner)during taxi, takeoff, and landing.

(d) Means must be provided to address thepossibility of a fire occurring on or in the immedi-ate vicinity of the cooktop. Two acceptable meansof complying with this requirement are as follows:

(1) Placarding must be installed that prohibitsany burner from being powered when the cooktopis unattended. (Note: This would prohibit a singleperson from cooking on the cooktop and intermit-tently serving food to passengers while anyburner is powered.) A fire detector must be in-stalled in the vicinity of the cooktop which pro-vides an audible warning in the passenger cabin,and a fire extinguisher of appropriate size and ex-tinguishing agent must be installed in the immedi-ate vicinity of the cooktop. Access to the extin-guisher may not be blocked by a fire on or aroundthe cooktop.

(2) An automatic, thermally activated fire sup-pression system must be installed to extinguish afire at the cooktop and immediately adjacent sur-faces. The agent used in the system must be anapproved total flooding agent suitable for use inan occupied area. The fire suppression systemmust have a manual override. The automatic acti-vation of the fire suppression system must alsoautomatically shut off power to the cooktop.

(e) The surfaces of the galley surrounding thecooktop which would be exposed to a fire on thecooktop surface or in cookware on the cooktop

Part 25: Airworthiness Standards: Transport Category

ASA 11

25

must be constructed of materials that comply withthe flammability requirements of part III of Appen-dix F to part 25. This requirement is in addition tothe flammability requirements typically required ofthe materials in these galley surfaces. During theselection of these materials, consideration mustalso be given to ensure that the flammability char-acteristics of the materials will not be adverselyaffected by the use of cleaning agents and uten-sils used to remove cooking stains.

(f) The cooktop must be ventilated with a sys-tem independent of the airplane cabin and cargoventilation system. Procedures and time intervalsmust be established to inspect and clean or re-place the ventilation system to prevent a fire haz-ard from the accumulation of flammable oils andbe included in the instructions for continued air-worthiness. The ventilation system ducting mustbe protected by a flame arrestor. [Note: The appli-cant may find additional useful information in So-ciety of Automotive Engineers, Aerospace Rec-ommended Practice 85, Rev. E, entitled “Air Con-ditioning Systems for Subsonic Airplanes,” datedAugust 1, 1991.]

(g) Means must be provided to contain spilledfoods or fluids in a manner that will prevent thecreation of a slipping hazard to occupants and willnot lead to the loss of structural strength due toairplane corrosion.

(h) Cooktop installations must provide ade-quate space for the user to immediately escape ahazardous cooktop condition.

(i) A means to shut off power to the cooktopmust be provided at the galley containing thecooktop and in the cockpit. If additional switchesare introduced in the cockpit, revisions to smokeor fire emergency procedures of the AFM will berequired.

(j) If the cooktop is required to have a lid to en-close the cooktop there must be a means to auto-matically shut off power to the cooktop when thelid is closed.

15. Hand-Held Fire Extinguishers.(a) For airplanes that were originally type certif-

icated with more than 60 passengers, the numberof hand-held fire extinguishers must be thegreater of—

(1) That provided in accordance with the re-quirements of §25.851, or

(2) A number equal to the number of originallytype certificated exit pairs, regardless of whetherthe exits are deactivated for the proposed config-uration.

(b) Extinguishers must be evenly distributedthroughout the cabin. These extinguishers are inaddition to those required by paragraph 14 of thisSFAR, unless it can be shown that the cooktopwas installed in the immediate vicinity of the origi-nal exits.

16. Security. The requirements of §25.795 arenot applicable to airplanes approved in accor-dance with this SFAR.

[Docket No. FAA–2007–28250, SFAR No. 109; 74 FR21541, May 8, 2009]

§25.1 Federal Aviation Regulations

12 ASA

Subpart A — General

§25.1 Applicability.(a) This part prescribes airworthiness standards

for the issue of type certificates, and changes tothose certificates, for transport category airplanes.

(b) Each person who applies under Part 21 forsuch a certificate or change must show compli-ance with the applicable requirements in this part.

§25.2 Special retroactive requirements.The following special retroactive requirements

are applicable to an airplane for which the regula-tions referenced in the type certificate predate thesections specified below —

(a) Irrespective of the date of application, eachapplicant for a supplemental type certificate (or anamendment to a type certificate) involving an in-crease in passenger seating capacity to a totalgreater than that for which the airplane has beentype certificated must show that the airplane con-cerned meets the requirements of:

(1) Sections 25.721(d), 25.783(g), 25.785(c),25.803(c) (2) through (9), 25.803 (d) and (e),25.807 (a), (c), and (d), 25.809 (f) and (h), 25.811,25.812, 25.813 (a), (b), and (c), 25.815, 25.817,25.853 (a) and (b), 25.855(a), 25.993(f), and25.1359(c) in effect on October 24, 1967, and

(2) Sections 25.803(b) and 25.803(c)(1) in ef-fect on April 23, 1969.

(b) Irrespective of the date of application, eachapplicant for a supplemental type certificate (or anamendment to a type certificate) for an airplanemanufactured after October 16, 1987, must showthat the airplane meets the requirements of§25.807(c)(7) in effect on July 24, 1989.

(c) Compliance with subsequent revisions tothe sections specified in paragraph (a) or (b) ofthis section may be elected or may be required inaccordance with §21.101(a) of this chapter.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–72, 55 FR 29773, July 20, 1990;Amdt. No. 25–99, 65 FR 36266, June 7, 2000; Amdt. 25–99, 66 FR 56989, Nov. 14, 2001]

§25.3 Special provisions for ETOPS type design approvals.(a) Applicability. This section applies to an ap-

plicant for ETOPS type design approval of an air-plane:

(1) That has an existing type certificate on Feb-ruary 15, 2007; or

(2) For which an application for an original typecertificate was submitted before February 15,2007.

(b) Airplanes with two engines.(1) For ETOPS type design approval of an air-

plane up to and including 180 minutes, an appli-

cant must comply with §25.1535, except that itneed not comply with the following provisions ofAppendix K, K25.1.4, of this part:

(i) K25.1.4(a), fuel system pressure and flowrequirements;

(ii) K25.1.4(a)(3), low fuel alerting; and(iii) K25.1.4(c), engine oil tank design.(2) For ETOPS type design approval of an air-

plane beyond 180 minutes an applicant mustcomply with §25.1535.

(c) Airplanes with more than two engines.An applicant for ETOPS type design approvalmust comply with §25.1535 for an airplane manu-factured on or after February 17, 2015, exceptthat, for an airplane configured for a three personflight crew, the applicant need not comply with Ap-pendix K, K25.1.4(a)(3), of this part, low fuel alert-ing.

[Docket No. FAA–2002–6717, 72 FR 1873, Jan. 16,2007]

§25.5 Incorporations by reference.(a) The materials listed in this section are incor-

porated by reference in the corresponding sec-tions noted. These incorporations by referencewere approved by the Director of the Federal Reg-ister in accordance with 5 U.S.C. 552(a) and 1CFR part 51. These materials are incorporated asthey exist on the date of the approval, and noticeof any change in these materials will be publishedin the Federal Register. The materials are avail-able for purchase at the corresponding addressesnoted below, and all are available for inspection atthe National Archives and Records Administration(NARA), and at FAA, Transport Airplane Director-ate, Aircraft Certification Service, 1601 Lind Ave-nue SW, Renton, Washington 98057-3356. For in-formation on the availability of this material atNARA, call 202-741-6030, or go to:

http://www.archives.gov/federal_register/code_of_federal_regulations/ibr_locations.html

(b) The following materials are available forpurchase from the following address: The Na-tional Technical Information Services (NTIS),Springfield, Virginia 22166.

(1) Fuel Tank Flammability AssessmentMethod User’s Manual, dated May 2008, docu-ment number DOT/FAA/AR-05/8, IBR approvedfor §25.981 and Appendix N. It can also be ob-tained at the following Web site:

http://www.fire.tc.faa.gov/systems/fueltank/FTFAM.stm

(2) [Reserved]

[Docket No. FAA–2005–22997, 73 FR 42494, July 21,2008]

Part 25: Airworthiness Standards: Transport Category §25.25

ASA 13

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Subpart B—Flight

GENERAL

§25.21 Proof of compliance.(a) Each requirement of this subpart must be

met at each appropriate combination of weightand center of gravity within the range of loadingconditions for which certification is requested.This must be shown—

(1) By tests upon an airplane of the type forwhich certification is requested, or by calculationsbased on, and equal in accuracy to, the results oftesting; and

(2) By systematic investigation of each proba-ble combination of weight and center of gravity, ifcompliance cannot be reasonably inferred fromcombinations investigated.

(b) [Reserved](c) The controllability, stability, trim, and stalling

characteristics of the airplane must be shown foreach altitude up to the maximum expected in op-eration.

(d) Parameters critical for the test being con-ducted, such as weight, loading (center of gravityand inertia), airspeed, power, and wind, must bemaintained within acceptable tolerances of thecritical values during flight testing.

(e) If compliance with the flight characteristicsrequirements is dependent upon a stability aug-mentation system or upon any other automatic orpower-operated system, compliance must beshown with §§25.671 and 25.672.

(f) In meeting the requirements of §§25.105(d),25.125, 25.233, and 25.237, the wind velocitymust be measured at a height of 10 meters abovethe surface, or corrected for the difference be-tween the height at which the wind velocity is mea-sured and the 10-meter height.

(g) The requirements of this subpart associ-ated with icing conditions apply only if the appli-cant is seeking certification for flight in icing condi-tions.

(1) Each requirement of this subpart, except§§25.121(a), 25.123(c), 25.143(b)(1) and (b)(2),25.149, 25.201(c)(2), 25.207(c) and (d), 25.239,and 25.251(b) through (e), must be met in icingconditions. Compliance must be shown using theice accretions defined in appendix C, assumingnormal operation of the airplane and its ice pro-tection system in accordance with the operatinglimitations and operating procedures establishedby the applicant and provided in the AirplaneFlight Manual.

(2) No changes in the load distribution limits of§25.23, the weight limits of §25.25 (except wherelimited by performance requirements of this sub-part), and the center of gravity limits of §25.27,

from those for non-icing conditions, are allowedfor flight in icing conditions or with ice accretion.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5671, April 8, 1970;Amdt. 25–42, 43 FR 2320, Jan. 16, 1978; Amdt. 25–72,55 FR 29774, July 20, 1990; Amdt. 25–121, 72 FR44665, Aug. 8, 2007]

§25.23 Load distribution limits.(a) Ranges of weights and centers of gravity

within which the airplane may be safely operatedmust be established. If a weight and center ofgravity combination is allowable only within cer-tain load distribution limits (such as spanwise)that could be inadvertently exceeded, these limitsand the corresponding weight and center of grav-ity combinations must be established.

(b) The load distribution limits may not ex-ceed—

(1) The selected limits;(2) The limits at which the structure is proven; or(3) The limits at which compliance with each

applicable flight requirement of this subpart isshown.

§25.25 Weight limits.(a) Maximum weights. Maximum weights corre-

sponding to the airplane operating conditions(such as ramp, ground or water taxi, takeoff, enroute, and landing), environmental conditions(such as altitude and temperature), and loadingconditions (such as zero fuel weight, center ofgravity position and weight distribution) must beestablished so that they are not more than—

(1) The highest weight selected by the appli-cant for the particular conditions; or

(2) The highest weight at which compliancewith each applicable structural loading and flightrequirement is shown, except that for airplanesequipped with standby power rocket engines themaximum weight must not be more than the high-est weight established in accordance with Appen-dix E of this part; or

(3) The highest weight at which compliance isshown with the certification requirements of Part36 of this chapter.

(b) Minimum weight. The minimum weight (thelowest weight at which compliance with each ap-plicable requirement of this part is shown) mustbe established so that it is not less than—

(1) The lowest weight selected by the applicant;(2) The design minimum weight (the lowest

weight at which compliance with each structuralloading condition of this part is shown); or

(3) The lowest weight at which compliance witheach applicable flight requirement is shown.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5671, April 8, 1970;Amdt. 25–63, 53 FR 16365, May 6, 1988]


14 ASA

§25.27 Center of gravity limits.The extreme forward and the extreme aft cen-

ter of gravity limitations must be established foreach practicably separable operating condition.No such limit may lie beyond—

(a) The extremes selected by the applicant;(b) The extremes within which the structure is

proven; or(c) The extremes within which compliance with

each applicable flight requirement is shown.

§25.29 Empty weight and corresponding center of gravity.(a) The empty weight and corresponding center

of gravity must be determined by weighing the air-plane with—

(1) Fixed ballast;(2) Unusable fuel determined under §25.959;

and(3) Full operating fluids, including—(i) Oil;(ii) Hydraulic fluid; and(iii) Other fluids required for normal operation of

airplane systems, except potable water, lavatoryprecharge water, and fluids intended for injectionin the engine.

(b) The condition of the airplane at the time ofdetermining empty weight must be one that is welldefined and can be easily repeated.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–42, 43 FR 2320, Jan. 16, 1978;Amdt. 25–72, 55 FR 29774, July 20, 1990]

§25.31 Removable ballast.Removable ballast may be used on showing

compliance with the flight requirements of thissubpart.

§25.33 Propeller speed and pitch limits.(a) The propeller speed and pitch must be lim-

ited to values that will ensure—(1) Safe operation under normal operating con-

ditions; and(2) Compliance with the performance require-

ments of §§25.101 through 25.125.(b) There must be a propeller speed limiting

means at the governor. It must limit the maximumpossible governed engine speed to a value notexceeding the maximum allowable r.p.m.

(c) The means used to limit the low pitch posi-tion of the propeller blades must be set so that theengine does not exceed 103 percent of the maxi-mum allowable engine rpm or 99 percent of anapproved maximum overspeed, whichever isgreater, with—

(1) The propeller blades at the low pitch limitand governor inoperative;

(2) The airplane stationary under standard at-mospheric conditions with no wind; and

(3) The engines operating at the takeoff mani-fold pressure limit for reciprocating engine pow-ered airplanes or the maximum takeoff torquelimit for turbopropeller engine-powered airplanes.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–57, 49 FR 6848, Feb. 23, 1984;Amdt. 25–72, 55 FR 29774, July 20, 1990]

PERFORMANCE

§25.101 General.(a) Unless otherwise prescribed, airplanes

must meet the applicable performance require-ments of this subpart for ambient atmosphericconditions and still air.

(b) The performance, as affected by enginepower or thrust, must be based on the followingrelative humidities;

(1) For turbine engine powered airplanes, a rel-ative humidity of—

(i) 80 percent, at and below standard tempera-tures; and

(ii) 34 percent, at and above standard tempera-tures plus 50°F.

Between these two temperatures, the relativehumidity must vary linearly.

(2) For reciprocating engine powered air-planes, a relative humidity of 80 percent in a stan-dard atmosphere. Engine power corrections forvapor pressure must be made in accordance withthe following table:

Altitude H (ft.)

Vapor pressure e (In. Hg.)

Specific humidity w (Lb. moisture per lb. dry air)

Density ratio ρ/σ = 0.0023769

0 0.403 0.00849 0.99508

1,000 .354 .00773 .96672

2,000 .311 .00703 .93895

3,000 .272 .00638 .91178

4,000 .238 .00578 .88514

5,000 .207 .00523 .85910

6,000 .1805 .00472 .83361

7,000 .1566 .00425 .80870

8,000 .1356 .00382 .78434

9,000 .1172 .00343 .76053

10,000 .1010 .00307 .73722

15,000 .0463 .001710 .62868

20,000 .01978 .000896 .53263

25,000 .00778 .000436 .44806


ASA 15

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(c) The performance must correspond to thepropulsive thrust available under the particularambient atmospheric conditions, the particularflight condition, and the relative humidity specifiedin paragraph (b) of this section. The available pro-pulsive thrust must correspond to engine power orthrust, not exceeding the approved power orthrust less—

(1) Installation losses; and(2) The power or equivalent thrust absorbed by

the accessories and services appropriate to theparticular ambient atmospheric conditions andthe particular flight condition.

(d) Unless otherwise prescribed, the applicantmust select the takeoff, en route, approach, andlanding configurations for the airplane.

(e) The airplane configurations may vary withweight, altitude, and temperature, to the extentthey are compatible with the operating proce-dures required by paragraph (f) of this section.

(f) Unless otherwise prescribed, in determiningthe accelerate-stop distances, takeoff flight paths,takeoff distances, and landing distances, changesin the airplane’s configuration, speed, power, andthrust, must be made in accordance with proce-dures established by the applicant for operation inservice.

(g) Procedures for the execution of balkedlandings and missed approaches associated withthe conditions prescribed in §§25.119 and25.121(d) must be established.

(h) The procedures established under para-graphs (f) and (g) of this section must—

(1) Be able to be consistently executed in ser-vice by crews of average skill;

(2) Use methods or devices that are safe andreliable; and

(3) Include allowance for any time delays, in theexecution of the procedures, that may reasonablybe expected in service.

(i) The accelerate-stop and landing distancesprescribed in §§25.109 and 25.125, respectively,must be determined with all the airplane wheelbrake assemblies at the fully worn limit of their al-lowable wear range.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–38, 41 FR 55466, Dec. 20, 1976;Amdt. 25–92, 63 FR 8318, Feb. 18, 1998]

§25.103 Stall speed.(a) The reference stall speed, VSR, is a cali-

brated airspeed defined by the applicant. VSRmay not be less than a 1-g stall speed. VSR is ex-pressed as:

where:VCLMAX = Calibrated airspeed obtained when theload factor-corrected lift coefficient

is first a maximum during the maneuver prescribed in paragraph (c) of this section. In addition, when the maneuver is limited by a device that abruptly pushes the nose down at a selected angle of attack (e.g., a stick pusher), VCLMAX may not be less than the speed existing at the instant the device operates;

nZW = Load factor normal to the flight path at VCLMAX

W = Airplane gross weight;S = Aerodynamic reference wing area; andq = Dynamic pressure.

(b) VCLMAX is determined with:(1) Engines idling, or, if that resultant thrust

causes an appreciable decrease in stall speed,not more than zero thrust at the stall speed;

(2) Propeller pitch controls (if applicable) in thetakeoff position;

(3) The airplane in other respects (such asflaps, landing gear, and ice accretions) in the con-dition existing in the test or performance standardin which VSR is being used;

(4) The weight used when VSR is being used asa factor to determine compliance with a requiredperformance standard;

(5) The center of gravity position that results inthe highest value of reference stall speed; and

(6) The airplane trimmed for straight flight at aspeed selected by the applicant, but not less than1.13 VSR and not greater than 1.3 VSR.

(c) Starting from the stabilized trim condition,apply the longitudinal control to decelerate the air-plane so that the speed reduction does not ex-ceed one knot per second.

(d) In addition to the requirements of paragraph(a) of this section, when a device that abruptlypushes the nose down at a selected angle of at-tack (e.g., a stick pusher) is installed, the refer-ence stall speed, VSR, may not be less than 2knots or 2 percent, whichever is greater, abovethe speed at which the device operates.

[FAA–2002–13902, 67 FR 70825, Nov. 26, 2002; asamended by Amdt. 25–121, 72 FR 44665, Aug. 8, 2007]

§25.105 Takeoff.(a) The takeoff speeds prescribed by §25.107,

the accelerate-stop distance prescribed by§25.109, the takeoff path prescribed by §25.111,the takeoff distance and takeoff run prescribed by§25.113, and the net takeoff flight path prescribed

VSR

VCLMAX

nZW

-----------------≥

nZWWqS

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


16 ASA

by §25.115, must be determined in the selectedconfiguration for takeoff at each weight, altitude,and ambient temperature within the operationallimits selected by the applicant—

(1) In non-icing conditions; and(2) In icing conditions, if in the configuration of

§25.121(b) with the takeoff ice accretion definedin appendix C:

(i) The stall speed at maximum takeoff weightexceeds that in non-icing conditions by more thanthe greater of 3 knots CAS or 3 percent of VSR; or

(ii) The degradation of the gradient of climb de-termined in accordance with §25.121(b) is greaterthan one-half of the applicable actual-to-net take-off flight path gradient reduction defined in§25.115(b).

(b) No takeoff made to determine the data re-quired by this section may require exceptional pi-loting skill or alertness.

(c) The takeoff data must be based on—(1) In the case of land planes and amphibians:(i) Smooth, dry and wet, hard-surfaced run-

ways; and(ii) At the option of the applicant, grooved or po-

rous friction course wet, hard-surfaced runways.(d) The takeoff data must include, within the es-

tablished operational limits of the airplane, the fol-lowing operational correction factors:

(1) Not more than 50 percent of nominal windcomponents along the takeoff path opposite tothe direction of takeoff, and not less than 150 per-cent of nominal wind components along the take-off path in the direction of takeoff.

(2) Effective runway gradients.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–92, 63 FR 8318, Feb. 18, 1998;Amdt. 25–121, 72 FR 44665, Aug. 8, 2007]

§25.107 Takeoff speeds.(a) V1 must be established in relation to VEF as

follows:(1) VEF is the calibrated airspeed at which the

critical engine is assumed to fail. VEF must be se-lected by the applicant, but may not be less thanVMCG determined under §25.149(e).

(2) V1, in terms of calibrated airspeed, is se-lected by the applicant; however, V1 may not beless than VEF plus the speed gained with criticalengine inoperative during the time interval be-tween the instant at which the critical engine isfailed, and the instant at which the pilot recog-nizes and reacts to the engine failure, as indicatedby the pilot’s initiation of the first action (e.g., ap-plying brakes, reducing thrust, deploying speedbrakes) to stop the airplane during accelerate-stop tests.

(b) V2MIN, in terms of calibrated airspeed, maynot be less than—

(1) 1.13 VSR for—

(i) Two-engine and three-engine turbopropellerand reciprocating engine powered airplanes; and

(ii) Turbojet powered airplanes without provi-sions for obtaining a significant reduction in theone-engine-inoperative power-on stall speed;

(2) 1.08 VSR for—(i) Turbopropeller and reciprocating engine

powered airplanes with more than three engines;and

(ii) Turbojet powered airplanes with provisionsfor obtaining a significant reduction in the one-en-gine-inoperative power-on stall speed; and

(3) 1.10 times VMC established under §25.149.(c) V2, in terms of calibrated airspeed, must be

selected by the applicant to provide at least thegradient of climb required by §25.121(b) but maynot be less than—

(1) V2MIN;(2) VR plus the speed increment attained (in ac-

cordance with §25.111(c)(2)) before reaching aheight of 35 feet above the takeoff surface; and

(3) A speed that provides the maneuvering ca-pability specified in §25.143(h).

(d) VMU is the calibrated airspeed at and abovewhich the airplane can safely lift off the ground,and continue the takeoff. VMU speeds must be se-lected by the applicant throughout the range ofthrust-to-weight ratios to be certificated. Thesespeeds may be established from free air data ifthese data are verified by ground takeoff tests.

(e) VR, in terms of calibrated airspeed, must beselected in accordance with the conditions ofparagraphs (e) (1) through (4) of this section:

(1) VR may not be less than—(i) V1(ii) 105 percent of VMC;(iii) The speed (determined in accordance with

§25.111(c)(2)) that allows reaching V2 beforereaching a height of 35 feet above the takeoff sur-face; or

(iv) A speed that, if the airplane is rotated at itsmaximum practicable rate, will result in a VLOF ofnot less than 110 percent of VMU in the all-en-gines-operating condition and not less than 105percent of VMU determined at the thrust-to-weightratio corresponding to the one-engine-inoperativecondition.

(2) For any given set of conditions (such asweight, configuration, and temperature), a singlevalue of VR, obtained in accordance with thisparagraph, must be used to show compliancewith both the one-engine-inoperative and the all-engines-operating takeoff provisions.

(3) It must be shown that the one-engine-inop-erative takeoff distance, using a rotation speed of5 knots less than VR established in accordancewith paragraphs (e)(1) and (2) of this section,does not exceed the corresponding one-engine-inoperative takeoff distance using the established


ASA 17

25

VR. The takeoff distances must be determined inaccordance with §25.113(a)(1).

(4) Reasonably expected variations in servicefrom the established takeoff procedures for theoperation of the airplane (such as over-rotation ofthe airplane and out-of-trim conditions) may notresult in unsafe flight characteristics or in markedincreases in the scheduled takeoff distances es-tablished in accordance with §25.113(a).

(f) VLOF is the calibrated airspeed at which theairplane first becomes airborne.

(g) VFTO, in terms of calibrated airspeed, mustbe selected by the applicant to provide at least thegradient of climb required by §25.121(c), but maynot be less than—

(1) 1.18 VSR; and(2) A speed that provides the maneuvering ca-

pability specified in §25.143(h).(h) In determining the takeoff speeds V1, VR,

and V2 for flight in icing conditions, the values ofVMCG, VMC, and VMU determined for non-icingconditions may be used.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–38, 41 FR 55466, Dec. 20, 1976;Amdt. 25–42, 43 FR 2320, Jan. 16, 1978; Amdt. 25–92,63 FR 8318, Feb. 18, 1998; Amdt. 25–94, 63 FR 8848,Feb. 23, 1998; Amdt. 25–108, 67 FR 70826, Nov. 26,2002; Amdt. 25–121, 72 FR 44665, Aug. 8, 2007]

§25.109 Accelerate-stop distance.(a) The accelerate-stop distance on a dry run-

way is the greater of the following distances:(1) The sum of the distances necessary to—(i) Accelerate the airplane from a standing start

with all engines operating to VEF for takeoff from adry runway;

(ii) Allow the airplane to accelerate from VEF tothe highest speed reached during the rejectedtakeoff, assuming the critical engine fails at VEFand the pilot takes the first action to reject thetakeoff at the V1 for takeoff from a dry runway;and

(iii) Come to a full stop on a dry runway fromthe speed reached as prescribed in paragraph(a)(1)(ii) of this section; plus

(iv) A distance equivalent to 2 seconds at theV1 for takeoff from a dry runway.

(2) The sum of the distances necessary to—(i) Accelerate the airplane from a standing start

with all engines operating to the highest speedreached during the rejected takeoff, assuming thepilot takes the first action to reject the takeoff atthe V1 for takeoff from a dry runway; and

(ii) With all engines still operating, come to a fullstop on dry runway from the speed reached asprescribed in paragraph (a)(2)(i) of this section;plus

(iii) A distance equivalent to 2 seconds at theV1 for takeoff from a dry runway.

(b) The accelerate-stop distance on a wet run-way is the greater of the following distances:

(1) The accelerate-stop distance on a dry run-way determined in accordance with paragraph (a)of this section; or

(2) The accelerate-stop distance determined inaccordance with paragraph (a) of this section, ex-cept that the runway is wet and the correspondingwet runway values of VEF and V1 are used. In de-termining the wet runway accelerate-stop dis-tance, the stopping force from the wheel brakesmay never exceed:

(i) The wheel brakes stopping force determinedin meeting the requirements of §25.101(i) andparagraph (a) of this section; and

(ii) The force resulting from the wet runwaybraking coefficient of friction determined in accor-dance with paragraphs (c) or (d) of this section, asapplicable, taking into account the distribution ofthe normal load between braked and unbrakedwheels at the most adverse center-of-gravity posi-tion approved for takeoff.

(c) The wet runway braking coefficient of fric-tion for a smooth wet runway is defined as a curveof friction coefficient versus ground speed andmust be computed as follows:


18 ASA

(1) The maximum tire-to-ground wet runwaybraking coefficient of friction is defined as:

Where—

Tire Pressure = maximum airplane operating tire pressure (psi);

µt/gMAX = maximum tire-to-ground braking coefficient;

V = airplane true ground speed (knots); andLinear interpolation may be used for tire pres-sures other than those listed.

(2) The maximum tire-to-ground wet runwaybraking coefficient of friction must be adjusted totake into account the efficiency of the anti-skidsystem on a wet runway. Anti-skid system opera-tion must be demonstrated by flight testing on asmooth wet runway, and its efficiency must be de-termined. Unless a specific anti-skid system effi-ciency is determined from a quantitative analysisof the flight testing on a smooth wet runway, themaximum tire-to-ground wet runway braking coef-ficient of friction determined in paragraph (c)(1) ofthis section must be multiplied by the efficiencyvalue associated with the type of anti-skid systeminstalled on the airplane:

(d) At the option of the applicant, a higher wetrunway braking coefficient of friction may be usedfor runway surfaces that have been grooved ortreated with a porous friction course material. Forgrooved and porous friction course runways, thewet runway braking coefficient of friction is de-fined as either:

(1) 70 percent of the dry runway braking coeffi-cient of friction used to determine the dry runwayaccelerate-stop distance; or

(2) The wet runway braking coefficient definedin paragraph (c) of this section, except that a spe-cific anti-skid system efficiency, if determined, isappropriate for a grooved or porous frictioncourse wet runway, and the maximum tire-to-ground wet runway braking coefficient of friction isdefined as:

Tire Pressure (psi) Maximum Braking Coefficient (tire-to-ground)

50

100

200

300

µ t g⁄ MAX0.0350

V100---------

3– 0.306

V100---------

20.851

V100---------

– 0.883+ +=

µ t g⁄ MAX0.0437

V100---------

3– 0.320

V100---------

20.805

V100---------

– 0.804+ +=

µ t g⁄ MAX0.0331

V100---------

3– 0.252

V100---------

20.658

V100---------

– 0.692+ +=

µ t g⁄ MAX0.0401

V100---------

3– 0.263

V100---------

20.611

V100---------

– 0.614+ +=

Type of anti-skid system Efficiency value

On-Off 0.30

Quasi-Modulating 0.50

Full Modulating 0.80


ASA 19

25

Where—

Tire Pressure = maximum airplane operating tire pressure (psi);

µt/gMAX = maximum tire-to-ground braking coefficient;

V = airplane true ground speed (knots); andLinear interpolation may be used for tire pres-sures other than those listed.

(e) Except as provided in paragraph (f)(1) ofthis section, means other than wheel brakes maybe used to determine the accelerate-stop dis-tance if that means—

(1) Is safe and reliable;(2) Is used so that consistent results can be ex-

pected under normal operating conditions; and(3) Is such that exceptional skill is not required

to control the airplane.(f) The effects of available reverse thrust—(1) Shall not be included as an additional

means of deceleration when determining the ac-celerate-stop distance on a dry runway; and

(2) May be included as an additional means ofdeceleration using recommended reverse thrustprocedures when determining the accelerate-stopdistance on a wet runway, provided the require-ments of paragraph (e) of this section are met.

(g) The landing gear must remain extendedthroughout the accelerate-stop distance.

(h) If the accelerate-stop distance includes astopway with surface characteristics substantiallydifferent from those on the runway, the takeoffdata must include operational correction factorsfor the accelerate-stop distance. The correctionfactors must account for the particular surfacecharacteristics of the stopway and the variationsin these characteristics with seasonal weatherconditions (such as temperature, rain, snow, andice) within the established operational limits.

(i) A flight test demonstration of the maximumbrake kinetic energy accelerate-stop distancemust be conducted with not more than 10 percent

of the allowable brake wear range remaining oneach of the airplane wheel brakes.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–42, 43 FR 2321, Jan. 16, 1978;Amdt. 25–92, 63 FR 8318, Feb. 18, 1998]

§25.111 Takeoff path.(a) The takeoff path extends from a standing

start to a point in the takeoff at which the airplaneis 1,500 feet above the takeoff surface, or at whichthe transition from the takeoff to the en route con-figuration is completed and VFTO is reached,whichever point is higher. In addition—

(1) The takeoff path must be based on the pro-cedures prescribed in §25.101(f);

(2) The airplane must be accelerated on theground to VEF, at which point the critical enginemust be made inoperative and remain inoperativefor the rest of the takeoff; and

(3) After reaching VEF, the airplane must be ac-celerated to V2.

(b) During the acceleration to speed V2, thenose gear may be raised off the ground at aspeed not less than VR. However, landing gear re-traction may not be begun until the airplane is air-borne.

(c) During the takeoff path determination in ac-cordance with paragraphs (a) and (b) of this sec-tion—

(1) The slope of the airborne part of the takeoffpath must be positive at each point;

(2) The airplane must reach V2 before it is 35feet above the takeoff surface and must continueat a speed as close as practical to, but not lessthan V2, until it is 400 feet above the takeoff sur-face;

(3) At each point along the takeoff path, start-ing at the point at which the airplane reaches 400feet above the takeoff surface, the available gradi-ent of climb may not be less than—

(i) 1.2 percent for two-engine airplanes;(ii) 1.5 percent for three-engine airplanes; and(iii) 1.7 percent for four-engine airplanes.

Tire Pressure (psi) Maximum Braking Coefficient (tire-to-ground)

50

100

200

300

µ t g⁄ MAX0.1470

V100-------

5

= 1.050V

100-------

4

– 2.673V

100-------

3

2.683V

100-------

2

– 0.403V

100-------

0.859+ + +

µ t g⁄ MAX0.1106

V100-------

5

= 0.813V

100-------

4

– 2.130V

100-------

3

2.200V

100-------

2

– 0.317V

100-------

0.807+ + +

µ t g⁄ MAX0.0498

V100-------

5

= 0.398V

100-------

4

– 1.140V

100-------

3

1.285V

100-------

2

– 0.140V

100-------

0.701+ + +

µ t g⁄ MAX0.0314

V100-------

5

= 0.247V

100-------

4

– 0.703V

100-------

3

0.779V

100-------

2

– 0.00954V

100-------

0.614+ + +


20 ASA

(4) The airplane configuration may not bechanged, except for gear retraction and automaticpropeller feathering, and no change in power orthrust that requires action by the pilot may bemade until the airplane is 400 feet above the take-off surface; and

(5) If §25.105(a)(2) requires the takeoff path tobe determined for flight in icing conditions, the air-borne part of the takeoff must be based on the air-plane drag:

(i) With the takeoff ice accretion defined in ap-pendix C, from a height of 35 feet above the take-off surface up to the point where the airplane is400 feet above the takeoff surface; and

(ii) With the final takeoff ice accretion defined inappendix C, from the point where the airplane is400 feet above the takeoff surface to the end ofthe takeoff path.

(4) Except for gear retraction and propellerfeathering, the airplane configuration may not bechanged, and no change in power or thrust thatrequires action by the pilot may be made, until theairplane is 400 feet above the takeoff surface.

(d) The takeoff path must be determined by acontinuous demonstrated takeoff or by synthesisfrom segments. If the takeoff path is determinedby the segmental method—

(1) The segments must be clearly defined andmust be related to the distinct changes in the con-figuration, power or thrust, and speed;

(2) The weight of the airplane, the configura-tion, and the power or thrust must be constantthroughout each segment and must correspondto the most critical condition prevailing in the seg-ment;

(3) The flight path must be based on the air-plane’s performance without ground effect; and

(4) The takeoff path data must be checked bycontinuous demonstrated takeoffs up to the pointat which the airplane is out of ground effect andits speed is stabilized, to ensure that the path isconservative relative to the continuous path.

The airplane is considered to be out of the groundeffect when it reaches a height equal to its wingspan.

(e) For airplanes equipped with standby powerrocket engines, the takeoff path may be deter-mined in accordance with section II of Appendix E.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–6, 30 FR 8468, July 2, 1965;Amdt. 25–42, 43 FR 2321, Jan. 16, 1978; Amdt. 25–54,45 FR 60172, Sept. 11, 1980; Amdt. 25–72, 55 FR29774, July 20, 1990; Amdt. 25–94, 63 FR 8848, Feb.23, 1998; Amdt. 25–108, 67 FR 70826, Nov. 26, 2002;Amdt. 25–121, 72 FR 44666, Aug. 8, 2007]

§25.113 Takeoff distance and takeoff run.(a) Takeoff distance on a dry runway is the

greater of—(1) The horizontal distance along the takeoff

path from the start of the takeoff to the point atwhich the airplane is 35 feet above the takeoffsurface, determined under §25.111 for a dry run-way; or

(2) 115 percent of the horizontal distance alongthe takeoff path, with all engines operating, fromthe start of the takeoff to the point at which the air-plane is 35 feet above the takeoff surface, as de-termined by a procedure consistent with §25.111.

(b) Takeoff distance on a wet runway is thegreater of—

(1) The takeoff distance on a dry runway deter-mined in accordance with paragraph (a) of thissection; or

(2) The horizontal distance along the takeoffpath from the start of the takeoff to the point atwhich the airplane is 15 feet above the takeoffsurface, achieved in a manner consistent with theachievement of V2 before reaching 35 feet abovethe takeoff surface, determined under §25.111 fora wet runway.

(c) If the takeoff distance does not include aclearway, the takeoff run is equal to the takeoffdistance. If the takeoff distance includes a clear-way—

(1) The takeoff run on a dry runway is thegreater of—

(i) The horizontal distance along the takeoffpath from the start of the takeoff to a point equi-distant between the point at which VLOF isreached and the point at which the airplane is 35feet above the takeoff surface, as determined un-der §25.111 for a dry runway; or

(ii) 115 percent of the horizontal distance alongthe takeoff path, with all engines operating, fromthe start of the takeoff to a point equidistant be-tween the point at which VLOF is reached and thepoint at which the airplane is 35 feet above thetakeoff surface, determined by a procedure con-sistent with §25.111.

(2) The takeoff run on a wet runway is thegreater of—

(i) The horizontal distance along the takeoffpath from the start of the takeoff to the point atwhich the airplane is 15 feet above the takeoffsurface, achieved in a manner consistent with theachievement of V2 before reaching 35 feet abovethe takeoff surface, as determined under §25.111for a wet runway; or

(ii) 115 percent of the horizontal distance alongthe takeoff path, with all engines operating, fromthe start of the takeoff to a point equidistant be-tween the point at which VLOF is reached and thepoint at which the airplane is 35 feet above the


ASA 21

25

takeoff surface, determined by a procedure con-sistent with §25.111.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5671, April 8, 1970;Amdt. 25–92, 63 FR 8320, Feb. 18, 1998]

§25.115 Takeoff flight path.(a) The takeoff flight path shall be considered to

begin 35 feet above the takeoff surface at the endof the takeoff distance determined in accordancewith §25.113(a) or (b), as appropriate for the run-way surface condition.

(b) The net takeoff flight path data must be de-termined so that they represent the actual takeoffflight paths (determined in accordance with§25.111 and with paragraph (a) of this section) re-duced at each point by a gradient of climb equalto—

(1) 0.8 percent for two-engine airplanes;(2) 0.9 percent for three-engine airplanes; and(3) 1.0 percent for four-engine airplanes.(c) The prescribed reduction in climb gradient

may be applied as an equivalent reduction in ac-celeration along that part of the takeoff flight pathat which the airplane is accelerated in level flight.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–92, 63 FR 8320, Feb. 18, 1998]

§25.117 Climb: general.Compliance with the requirements of §§25.119

and 25.121 must be shown at each weight, alti-tude, and ambient temperature within the opera-tional limits established for the airplane and withthe most unfavorable center of gravity for eachconfiguration.

§25.119 Landing climb: All-engines-operating.In the landing configuration, the steady gradi-

ent of climb may not be less than 3.2 percent, withthe engines at the power or thrust that is available8 seconds after initiation of movement of thepower or thrust controls from the minimum flightidle to the go-around power or thrust setting—

(a) In non-icing conditions, with a climb speedof VREF determined in accordance with§25.125(b)(2)(i); and

(b) In icing conditions with the landing ice ac-cretion defined in appendix C, and with a climbspeed of VREF determined in accordance with§25.125(b)(2)(ii).

[Docket No. FAA–2005–22840, 72 FR 44666, Aug. 8,2007]

§25.121 Climb: One-engine-inoperative.(a) Takeoff; landing gear extended. In the criti-

cal takeoff configuration existing along the flightpath (between the points at which the airplane

reaches VLOF and at which the landing gear isfully retracted) and in the configuration used in§25.111 but without ground effect, the steady gra-dient of climb must be positive for two-engine air-planes, and not less than 0.3 percent for three-en-gine airplanes or 0.5 percent for four-engine air-planes, at VLOF and with—

(1) The critical engine inoperative and the re-maining engines at the power or thrust availablewhen retraction of the landing gear is begun in ac-cordance with §25.111 unless there is a more crit-ical power operating condition existing later alongthe flight path but before the point at which thelanding gear is fully retracted; and

(2) The weight equal to the weight existingwhen retraction of the landing gear is begun, de-termined under §25.111.

(b) Takeoff; landing gear retracted. In the take-off configuration existing at the point of the flightpath at which the landing gear is fully retracted,and in the configuration used in §25.111 but with-out ground effect:

(1) The steady gradient of climb may not beless than 2.4 percent for two-engine airplanes,2.7 percent for three-engine airplanes, and 3.0percent for four-engine airplanes, at V2 with:

(i) The critical engine inoperative, the remainingengines at the takeoff power or thrust available atthe time the landing gear is fully retracted, deter-mined under §25.111, unless there is a more crit-ical power operating condition existing later alongthe flight path but before the point where the air-plane reaches a height of 400 feet above the take-off surface; and

(ii) The weight equal to the weight existingwhen the airplane’s landing gear is fully retracted,determined under §25.111.

(2) The requirements of paragraph (b)(1) of thissection must be met:

(i) In non-icing conditions; and(ii) In icing conditions with the takeoff ice accre-

tion defined in appendix C, if in the configurationof §25.121(b) with the takeoff ice accretion:

(A) The stall speed at maximum takeoff weightexceeds that in non-icing conditions by more thanthe greater of 3 knots CAS or 3 percent of VSR; or

(B) The degradation of the gradient of climb de-termined in accordance with §25.121(b) is greaterthan one-half of the applicable actual-to-net take-off flight path gradient reduction defined in§25.115(b).

(c) Final takeoff. In the en route configuration atthe end of the takeoff path determined in accor-dance with §25.111:

(1) The steady gradient of climb may not beless than 1.2 percent for two-engine airplanes,1.5 percent for three-engine airplanes, and 1.7percent for four-engine airplanes, at VFTO with—


22 ASA

(i) The critical engine inoperative and the re-maining engines at the available maximum contin-uous power or thrust; and

(ii) The weight equal to the weight existing atthe end of the takeoff path, determined under§25.111.

(2) The requirements of paragraph (c)(1) of thissection must be met:

(i) In non-icing conditions; and(ii) In icing conditions with the final takeoff ice

accretion defined in appendix C, if in the configu-ration of §25.121(b) with the takeoff ice accretion:

(A) The stall speed at maximum takeoff weightexceeds that in non-icing conditions by more thanthe greater of 3 knots CAS or 3 percent of VSR; or

(B) The degradation of the gradient of climb de-termined in accordance with §25.121(b) is greaterthan one-half of the applicable actual-to-net take-off flight path gradient reduction defined in§25.115(b).

(d) Approach. In a configuration correspondingto the normal all-engines-operating procedure inwhich VSR for this configuration does not exceed110 percent of the VSR for the related all-engines-operating landing configuration:

(1) The steady gradient of climb may not beless than 2.1 percent for two-engine airplanes,2.4 percent for three-engine airplanes, and 2.7percent for four-engine airplanes, with—

(i) The critical engine inoperative, the remainingengines at the go-around power or thrust setting;

(ii) The maximum landing weight;(iii) A climb speed established in connection

with normal landing procedures, but not exceed-ing 1.4 VSR; and

(iv) Landing gear retracted.(2) The requirements of paragraph (d)(1) of this

section must be met:(i) In non-icing conditions; and(ii) In icing conditions with the approach ice ac-

cretion defined in appendix C. The climb speedselected for non-icing conditions may be used ifthe climb speed for icing conditions, computed inaccordance with paragraph (d)(1)(iii) of this sec-tion, does not exceed that for non-icing conditionsby more than the greater of 3 knots CAS or 3 per-cent.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–84, 60 FR 30749, June 9, 1995;Amdt. 25–108, 67 FR 70826, Nov. 26, 2002; Amdt. 25–121, 72 FR 44666, Aug. 8, 2007]

§25.123 En route flight paths.(a) For the en route configuration, the flight

paths prescribed in paragraph (b) and (c) of thissection must be determined at each weight, alti-tude, and ambient temperature, within the operat-ing limits established for the airplane. The varia-tion of weight along the flight path, accounting for

the progressive consumption of fuel and oil by theoperating engines, may be included in the compu-tation. The flight paths must be determined at aspeed not less than VFTO, with—

(1) The most unfavorable center of gravity;(2) The critical engines inoperative;(3) The remaining engines at the available

maximum continuous power or thrust; and(4)The means for controlling the engine-cooling

air supply in the position that provides adequatecooling in the hot-day condition.

(b) The one-engine-inoperative net flight pathdata must represent the actual climb performancediminished by a gradient of climb of 1.1 percentfor two-engine airplanes, 1.4 percent for three-en-gine airplanes, and 1.6 percent for four-engineairplanes—

(1) In non-icing conditions; and(2) In icing conditions with the en route ice ac-

cretion defined in appendix C, if:(i) A speed of 1.18 VSR with the en route ice ac-

cretion exceeds the en route speed selected fornon-icing conditions by more than the greater of 3knots CAS or 3 percent of VSR; or

(ii) The degradation of the gradient of climb isgreater than one-half of the applicable actual-to-net flight path reduction defined in paragraph (b)of this section.

(c) For three- or four-engine airplanes, thetwo-engine-inoperative net flight path data mustrepresent the actual climb performance dimin-ished by a gradient of climb of 0.3 percent forthree-engine airplanes and 0.5 percent for four-engine airplanes.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–121, 72 FR 44666, Aug. 8, 2007]

§25.125 Landing.(a) The horizontal distance necessary to land

and to come to a complete stop (or to a speed ofapproximately 3 knots for water landings) from apoint 50 feet above the landing surface must bedetermined (for standard temperatures, at eachweight, altitude, and wind within the operationallimits established by the applicant for the air-plane):

(1) In non-icing conditions; and(2) In icing conditions with the landing ice ac-

cretion defined in appendix C if VREF for icing con-ditions exceeds VREF for non-icing conditions bymore than 5 knots CAS at the maximum landingweight.

(b) In determining the distance in paragraph (a)of this section:

(1) The airplane must be in the landing configu-ration.

(2) A stabilized approach, with a calibrated air-speed of not less than VREF, must be maintaineddown to the 50-foot height.


ASA 23

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(i) In non-icing conditions, VREF may not beless than:

(A) 1.23 VSR0;(B) VMCL established under §25.149(f); and(C) A speed that provides the maneuvering ca-

pability specified in §25.143(h).(ii) In icing conditions, VREF may not be less

than:(A) The speed determined in paragraph

(b)(2)(i) of this section;(B) 1.23 VSR0 with the landing ice accretion de-

fined in appendix C if that speed exceeds VREF fornon-icing conditions by more than 5 knots CAS;and

(C) A speed that provides the maneuvering ca-pability specified in §25.143(h) with the landingice accretion defined in appendix C.

(3) Changes in configuration, power or thrust,and speed, must be made in accordance with theestablished procedures for service operation.

(4) The landing must be made without exces-sive vertical acceleration, tendency to bounce,nose over, ground loop, porpoise, or water loop.

(5) The landings may not require exceptionalpiloting skill or alertness.

(c) For landplanes and amphibians, the landingdistance on land must be determined on a level,smooth, dry, hard-surfaced runway. In addition—

(1) The pressures on the wheel braking sys-tems may not exceed those specified by the brakemanufacturer;

(2) The brakes may not be used so as to causeexcessive wear of brakes or tires; and

(3) Means other than wheel brakes may beused if that means—

(i) Is safe and reliable;(ii) Is used so that consistent results can be ex-

pected in service; and(iii) Is such that exceptional skill is not required

to control the airplane.(d) For seaplanes and amphibians, the landing

distance on water must be determined on smoothwater.

(e) For skiplanes, the landing distance on snowmust be determined on smooth, dry, snow.

(f) The landing distance data must include cor-rection factors for not more than 50 percent of thenominal wind components along the landing pathopposite to the direction of landing, and not lessthan 150 percent of the nominal wind componentsalong the landing path in the direction of landing.

(g) If any device is used that depends on theoperation of any engine, and if the landing dis-tance would be noticeably increased when a land-ing is made with that engine inoperative, the land-ing distance must be determined with that engineinoperative unless the use of compensatingmeans will result in a landing distance not morethan that with each engine operating.


CONTROLLABILITY AND MANEUVERABILITY

§25.143 General.(a) The airplane must be safely controllable

and maneuverable during—(1) Takeoff;(2) Climb;(3) Level flight;(4) Descent; and(5) Landing.(b) It must be possible to make a smooth tran-

sition from one flight condition to any other flightcondition without exceptional piloting skill, alert-ness, or strength, and without danger of exceed-ing the airplane limit-load factor under any proba-ble operating conditions, including—

(1) The sudden failure of the critical engine;(2) For airplanes with three or more engines,

the sudden failure of the second critical enginewhen the airplane is in the en route, approach, orlanding configuration and is trimmed with the crit-ical engine inoperative; and

(3) Configuration changes, including deploy-ment or retraction of deceleration devices.

(c) The airplane must be shown to be safelycontrollable and maneuverable with the critical iceaccretion appropriate to the phase of flight de-fined in appendix C, and with the critical engineinoperative and its propeller (if applicable) in theminimum drag position:

(1) At the minimum V2 for takeoff;(2) During an approach and go-around; and(3) During an approach and landing.(d) The following table prescribes, for conven-

tional wheel type controls, the maximum controlforces permitted during the testing required byparagraphs (a) and (c) of this section:

(e) Approved operating procedures or conven-tional operating practices must be followed whendemonstrating compliance with the control forcelimitations for short term application that are pre-scribed in paragraph (d) of this section. The air-

Force, in pounds, applied to the control wheel or rudder pedals

Pitch Roll Yaw

For short term application for pitch and roll control — two hands available for control

75 50 —

For short term application for pitch and roll control — one hand available for control

50 25 —

For short term application for yaw control

— — 150

For long term application 10 5 20


24 ASA

plane must be in trim, or as near to being in trimas practical, in the preceding steady flight condi-tion. For the takeoff condition, the airplane mustbe trimmed according to the approved operatingprocedures.

(f) When demonstrating compliance with thecontrol force limitations for long term applicationthat are prescribed in paragraph (d) of this sec-tion, the airplane must be in trim, or as near to be-ing in trim as practical.

(g) When maneuvering at a constant airspeedor Mach number (up to VFC/MFC), the stick forcesand the gradient of the stick force versus maneu-vering load factor must lie within satisfactory lim-

its. The stick forces must not be so great as tomake excessive demands on the pilot’s strengthwhen maneuvering the airplane, and must not beso low that the airplane can easily be over-stressed inadvertently. Changes of gradient thatoccur with changes of load factor must not causeundue difficulty in maintaining control of the air-plane, and local gradients must not be so low asto result in a danger of overcontrolling.

(h) The maneuvering capabilities in a constant-speed coordinated turn at forward center of grav-ity, as specified in the following table, must be freeof stall warning or other characteristics that mightinterfere with normal maneuvering:

1 A combination of weight, altitude, and temperature (WAT) such that the thrust or power setting produces the mini-mum climb gradient specified in §25.121 for the flight condition.2 Airspeed approved for all-engines-operating initial climb.3 That thrust or power setting which, in the event of failure of the critical engine and without any crew action to adjustthe thrust or power of the remaining engines, would result in the thrust or power specified for the takeoff condition atV2, or any lesser thrust or power setting that is used for all-engines-operating initial climb procedures.

(i) When demonstrating compliance with§25.143 in icing conditions—

(1) Controllability must be demonstrated withthe ice accretion defined in appendix C that ismost critical for the particular flight phase;

(2) It must be shown that a push force is re-quired throughout a pushover maneuver down toa zero g load factor, or the lowest load factor ob-tainable if limited by elevator power or other de-sign characteristic of the flight control system. Itmust be possible to promptly recover from themaneuver without exceeding a pull control forceof 50 pounds; and

(3) Any changes in force that the pilot must ap-ply to the pitch control to maintain speed with in-creasing sideslip angle must be steadily increas-ing with no force reversals, unless the change incontrol force is gradual and easily controllable bythe pilot without using exceptional piloting skill,alertness, or strength.

(j) For flight in icing conditions before the iceprotection system has been activated and is per-forming its intended function, it must be demon-strated in flight with the ice accretion defined inappendix C, part II(e) of this part that:

(1) The airplane is controllable in a pull-up ma-neuver up to 1.5 g load factor; and

(2) There is no pitch control force reversal dur-ing a pushover maneuver down to 0.5 g load fac-tor.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–42, 43 FR 2321, Jan. 16, 1978;Amdt. 25–84, 60 FR 30749, June 9, 1995; Amdt. 25–108, 67 FR 70826, Nov. 26, 2002; Amdt. 25–121, 72 FR44667, Aug. 8, 2007; Amdt. 25–129, 74 FR 38339, Aug.3, 2009]

§25.145 Longitudinal control.(a) It must be possible, at any point between

the trim speed prescribed in §25.103(b)(6) andstall identification (as defined in §25.201(d)), topitch the nose downward so that the accelerationto this selected trim speed is prompt with—

(1) The airplane trimmed at the trim speed pre-scribed in §25.103(b)(6).

(2) The landing gear extended;(3) The wing flaps (i) retracted and (ii) ex-

tended; and(4) Power (i) off and (ii) at maximum continuous

power on the engines.(b) With the landing gear extended, no change

in trim control, or exertion of more than 50 poundscontrol force (representative of the maximumshort term force that can be applied readily byone hand) may be required for the following ma-neuvers:

Configuration Speed Maneuvering bank angle in a coordinated turn

Thrust / power setting

Takeoff V2 30° Asymmetric WAT-Limited.1

Takeoff 2V2 + XX 40° All-engines-operating climb.3

En route VFTO 40° Asymmetric WAT-Limited.1

Landing VREF 40° Symmetric for -3° flight path angle.


ASA 25

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(1) With power off, flaps retracted, and the air-plane trimmed at 1.3 VSR1, extend the flaps asrapidly as possible while maintaining the airspeedat approximately 30 percent above the referencestall speed existing at each instant throughout themaneuver.

(2) Repeat paragraph (b)(1) except initially ex-tend the flaps and then retract them as rapidly aspossible.

(3) Repeat paragraph (b)(2), except at the go-around power or thrust setting.

(4) With power off, flaps retracted, and the air-plane trimmed at 1.3 VSR1, rapidly set go-aroundpower or thrust while maintaining the same air-speed.

(5) Repeat paragraph (b)(4) except with flapsextended.

(6) With power off, flaps extended, and the air-plane trimmed at 1.3 VSR1, obtain and maintainairspeeds between VSW and either 1.6 VSR1 orVFE, whichever is lower.

(c) It must be possible, without exceptional pi-loting skill, to prevent loss of altitude when com-plete retraction of the high-lift devices from anyposition is begun during steady, straight, levelflight at 1.08 VSR1 for propeller powered air-planes, or 1.13 VSR1 for turbojet powered air-planes, with—

(1) Simultaneous movement of the power orthrust controls to the go-around power or thrustsetting;

(2) The landing gear extended; and(3) The critical combinations of landing weights

and altitudes.(d) If gated high-lift device control positions are

provided, paragraph (c) of this section applies toretractions of the high-lift devices from any posi-tion from the maximum landing position to the firstgated position, between gated positions, and fromthe last gated position to the fully retracted posi-tion. The requirements of paragraph (c) of thissection also apply to retractions from each ap-proval landing position to the control position(s)associated with the high-lift device configura-tion(s) used to establish the go-around proce-dure(s) from that landing position. In addition, thefirst gated control position from the maximumlanding position must correspond with a configu-ration of the high-lift devices used to establish ago-around procedure from a landing configura-tion. Each gated control position must require aseparate and distinct motion of the control to passthrough the gated position and must have fea-tures to prevent inadvertent movement of the con-trol through the gated position. It must only bepossible to make this separate and distinct motiononce the control has reached the gated position.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5671, April 8, 1970;Amdt. 25–72, 55 FR 29774, July 20, 1990; Amdt. 25–84,60 FR 30749, June 9, 1995; Amdt. 25–98, 64 FR 6164,Feb. 8, 1999; Amdt. 25–108, 67 FR 70827, Nov. 26,2002]

§25.147 Directional and lateral control.(a) Directional control; general. It must be pos-

sible, with the wings level, to yaw into the opera-tive engine and to safely make a reasonably sud-den change in heading of up to 15 degrees in thedirection of the critical inoperative engine. Thismust be shown at 1.3 VSR1 for heading changesup to 15 degrees (except that the heading changeat which the rudder pedal force is 150 poundsneed not be exceeded), and with—

(1) The critical engine inoperative and its pro-peller in the minimum drag position;

(2) The power required for level flight at 1.3VSR1, but not more than maximum continuouspower;

(3) The most unfavorable center of gravity;(4) Landing gear retracted;(5) Flaps in the approach position; and(6) Maximum landing weight.(b) Directional control; airplanes with four or

more engines. Airplanes with four or more en-gines must meet the requirements of paragraph(a) of this section except that—

(1) The two critical engines must be inoperativewith their propellers (if applicable) in the minimumdrag position;

(2) [Reserved](3) The flaps must be in the most favorable

climb position.(c) Lateral control; general. It must be possible

to make 20° banked turns, with and against the in-operative engine, from steady flight at a speedequal to 1.3 VSR1, with—

(1) The critical engine inoperative and its pro-peller (if applicable) in the minimum drag position;

(2) The remaining engines at maximum contin-uous power;

(3) The most unfavorable center of gravity;(4) Landing gear (i) retracted and (ii) extended;(5) Flaps in the most favorable climb position;

and(6) Maximum takeoff weight.(d) Lateral control; airplanes with four or more

engines. Airplanes with four or more enginesmust be able to make 20° banked turns, with andagainst the inoperative engines, from steady flightat a speed equal to 1.3 VSR1, with maximum con-tinuous power, and with the airplane in the config-uration prescribed by paragraph (b) of this sec-tion.

(e) Lateral control; all engines operating. Withthe engines operating, roll response must allownormal maneuvers (such as recovery from upsets


26 ASA

produced by gusts and the initiation of evasivemaneuvers). There must be enough excess lat-eral control in sideslips (up to sideslip angles thatmight be required in normal operation), to allow alimited amount of maneuvering and to correct forgusts. Lateral control must be enough at anyspeed up to VFC/ MFC to provide a peak roll ratenecessary for safety, without excessive controlforces or travel.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–42, 43 FR 2321, Jan. 16, 1978;Amdt. 25–72, 55 FR 29774, July 20, 1990; Amdt. 25–108; Amdt. 25–108, 67 FR 70827, Nov. 26, 2002]

§25.149 Minimum control speed.(a) In establishing the minimum control speeds

required by this section, the method used to simu-late critical engine failure must represent the mostcritical mode of powerplant failure with respect tocontrollability expected in service.

(b) VMC is the calibrated airspeed at which,when the critical engine is suddenly made inoper-ative, it is possible to maintain control of the air-plane with that engine still inoperative and main-tain straight flight with an angle of bank of notmore than 5 degrees.

(c) VMC may not exceed 1.13 VSR with—(1) Maximum available takeoff power or thrust

on the engines;(2) The most unfavorable center of gravity;(3) The airplane trimmed for takeoff;(4) The maximum sea level takeoff weight (or

any lesser weight necessary to show VMC);(5) The airplane in the most critical takeoff con-

figuration existing along the flight path after theairplane becomes airborne, except with the land-ing gear retracted;

(6) The airplane airborne and the ground effectnegligible; and

(7) If applicable, the propeller of the inoperativeengine—

(i) Windmilling;(ii) In the most probable position for the specific

design of the propeller control; or(iii) Feathered, if the airplane has an automatic

feathering device acceptable for showing compli-ance with the climb requirements of §25.121.

(d) The rudder forces required to maintain con-trol at VMC may not exceed 150 pounds nor may itbe necessary to reduce power or thrust of the op-erative engines. During recovery, the airplanemay not assume any dangerous attitude or re-quire exceptional piloting skill, alertness, orstrength to prevent a heading change of morethan 20 degrees.

(e) VMCG, the minimum control speed on theground, is the calibrated airspeed during the take-off run at which, when the critical engine is sud-denly made inoperative, it is possible to maintain

control of the airplane using the rudder controlalone (without the use of nosewheel steering), aslimited by 150 pounds of force, and the lateralcontrol to the extent of keeping the wings level toenable the takeoff to be safely continued usingnormal piloting skill. In the determination of VMCG,assuming that the path of the airplane accelerat-ing with all engines operating is along the center-line of the runway, its path from the point at whichthe critical engine is made inoperative to the pointat which recovery to a direction parallel to thecenterline is completed may not deviate morethan 30 feet laterally from the centerline at anypoint. VMCG must be established with—

(1) The airplane in each takeoff configurationor, at the option of the applicant, in the most criti-cal takeoff configuration;

(2) Maximum available takeoff power or thruston the operating engines;

(3) The most unfavorable center of gravity;(4) The airplane trimmed for takeoff; and(5) The most unfavorable weight in the range of

takeoff weights.(f) VMCL, the minimum control speed during ap-

proach and landing with all engines operating, isthe calibrated airspeed at which, when the criticalengine is suddenly made inoperative, it is possi-ble to maintain control of the airplane with that en-gine still inoperative, and maintain straight flightwith an angle of bank of not more than 5 degrees.VMCL must be established with—

(1) The airplane in the most critical configuration(or, at the option of the applicant, each configura-tion) for approach and landing with all engines op-erating;

(2) The most unfavorable center of gravity;(3) The airplane trimmed for approach with all

engines operating;(4) The most favorable weight, or, at the option

of the applicant, as a function of weight;(5) For propeller airplanes, the propeller of the

inoperative engine in the position it achieves with-out pilot action, assuming the engine fails while atthe power or thrust necessary to maintain a threedegree approach path angle; and

(6) Go-around power or thrust setting on theoperating engine(s).

(g) For airplanes with three or more engines,VMCL-2, the minimum control speed during ap-proach and landing with one critical engine inop-erative, is the calibrated airspeed at which, whena second critical engine is suddenly made inoper-ative, it is possible to maintain control of the air-plane with both engines still inoperative, andmaintain straight flight with an angle of bank of notmore than 5 degrees. VMCL-2 must be establishedwith—

(1) The airplane in the most critical configura-tion (or, at the option of the applicant, each config-


ASA 27

25

uration) for approach and landing with one criticalengine inoperative;

(2) The most unfavorable center of gravity;(3) The airplane trimmed for approach with one

critical engine inoperative;(4) The most unfavorable weight, or, at the op-

tion of the applicant, as a function of weight;(5) For propeller airplanes, the propeller of the

more critical inoperative engine in the position itachieves without pilot action, assuming the en-gine fails while at the power or thrust necessary tomaintain a three degree approach path angle,and the propeller of the other inoperative enginefeathered;

(6) The power or thrust on the operating en-gine(s) necessary to maintain an approach pathangle of three degrees when one critical engine isinoperative; and

(7) The power or thrust on the operating en-gine(s) rapidly changed, immediately after thesecond critical engine is made inoperative, fromthe power or thrust prescribed in paragraph (g)(6)of this section to—

(i) Minimum power or thrust; and(ii) Go-around power or thrust setting.(h) In demonstration of VMCL and VMCL-2 —(1) The rudder force may not exceed 150

pounds;(2) The airplane may not exhibit hazardous

flight characteristics or require exceptional pilot-ing skill, alertness, or strength;

(3) Lateral control must be sufficient to roll theairplane, from an initial condition of steady flight,through an angle of 20 degrees in the directionnecessary to initiate a turn away from the inoper-ative engine(s), in not more than 5 seconds; and

(4) For propeller airplanes, hazardous flightcharacteristics must not be exhibited due to anypropeller position achieved when the engine failsor during any likely subsequent movements of theengine or propeller controls.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–42, 43 FR 2321, Jan. 16, 1978;Amdt. 25–72, 55 FR 29774, July 20, 1990; 55 FR 37607,Sept. 12, 1990; Amdt. 25–84, 60 FR 30750, June 9,1995; Amdt. 25–108, 67 FR 70827, Nov. 26, 2002]

TRIM

§25.161 Trim.(a) General. Each airplane must meet the trim

requirements of this section after being trimmed,and without further pressure upon, or movementof, either the primary controls or their correspond-ing trim controls by the pilot or the automatic pilot.

(b) Lateral and directional trim. The airplanemust maintain lateral and directional trim with themost adverse lateral displacement of the center ofgravity within the relevant operating limitations,during normally expected conditions of operation

(including operation at any speed from 1.3 VSR1to VMO/MMO).

(c) Longitudinal trim. The airplane must main-tain longitudinal trim during—

(1) A climb with maximum continuous power ata speed not more than 1.3 VSR1, with the landinggear retracted, and the flaps (i) retracted and (ii)in the takeoff position;

(2) A glide with power off at a speed not morethan 1.3 VSR1, with the landing gear extended, thewing flaps (i) retracted and (ii) extended, the mostunfavorable center of gravity position approved forlanding with the maximum landing weight, andwith the most unfavorable center of gravity posi-tion approved for landing regardless of weight;and

(3) Level flight at any speed from 1.3 VSR1, toVMO/MMO, with the landing gear and flaps re-tracted, and from 1.3 VSR1 to VLE with the landinggear extended.

(d) Longitudinal, directional, and lateral trim.The airplane must maintain longitudinal, direc-tional, and lateral trim (and for the lateral trim, theangle of bank may not exceed five degrees) at 1.3VSR1 during climbing flight with—

(1) The critical engine inoperative;(2) The remaining engines at maximum contin-

uous power; and(3) The landing gear and flaps retracted.(e) Airplanes with four or more engines. Each

airplane with four or more engines must maintaintrim in rectilinear flight—

(1) At the climb speed, configuration, andpower required by §25.123(a) for the purpose ofestablishing the rate of climb;

(2) With the most unfavorable center of gravityposition; and

(3) At the weight at which the two-engine-inop-erative climb is equal to at least 0.013 VSR0

2 at analtitude of 5,000 feet.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5671, April 8, 1970;Amdt. 25–38, 41 FR 55466, Dec. 20, 1976; Amdt. 25–108, 67 FR 70827, Nov. 26, 2002]

STABILITY

§25.171 General.The airplane must be longitudinally, direction-

ally, and laterally stable in accordance with theprovisions of §§25.173 through 25.177. In addi-tion, suitable stability and control feel (static sta-bility) is required in any condition normally en-countered in service, if flight tests show it is nec-essary for safe operation.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–7, 30 FR 13117, Oct. 15, 1965]


28 ASA

§25.173 Static longitudinal stability.Under the conditions specified in §25.175, the

characteristics of the elevator control forces (in-cluding friction) must be as follows:

(a) A pull must be required to obtain and main-tain speeds below the specified trim speed, and apush must be required to obtain and maintainspeeds above the specified trim speed. This mustbe shown at any speed that can be obtained ex-cept speeds higher than the landing gear or wingflap operating limit speeds or VFC/MFC, whicheveris appropriate, or lower than the minimum speedfor steady unstalled flight.

(b) The airspeed must return to within 10 per-cent of the original trim speed for the climb, ap-proach, and landing conditions specified in§25.175 (a), (c), and (d), and must return to within7.5 percent of the original trim speed for the cruis-ing condition specified in §25.175(b), when thecontrol force is slowly released from any speedwithin the range specified in paragraph (a) of thissection.

(c) The average gradient of the stable slope ofthe stick force versus speed curve may not beless than 1 pound for each 6 knots.

(d) Within the free return speed range specifiedin paragraph (b) of this section, it is permissiblefor the airplane, without control forces, to stabilizeon speeds above or below the desired trimspeeds if exceptional attention on the part of thepilot is not required to return to and maintain thedesired trim speed and altitude.


§25.175 Demonstration of static longitudinal stability.Static longitudinal stability must be shown as

follows:(a) Climb. The stick force curve must have a

stable slope at speeds between 85 and 115 per-cent of the speed at which the airplane—

(1) Is trimmed, with—(i) Wing flaps retracted;(ii) Landing gear retracted;(iii) Maximum takeoff weight; and(iv) 75 percent of maximum continuous power

for reciprocating engines or the maximum poweror thrust selected by the applicant as an operatinglimitation for use during climb for turbine engines;and

(2) Is trimmed at the speed for best rate-of-climb except that the speed need not be less than1.3 VS R1.

(b) Cruise. Static longitudinal stability must beshown in the cruise condition as follows:

(1) With the landing gear retracted at highspeed, the stick force curve must have a stableslope at all speeds within a range which is the

greater of 15 percent of the trim speed plus theresulting free return speed range, or 50 knots plusthe resulting free return speed range, above andbelow the trim speed (except that the speedrange need not include speeds less than 1.3VS R1, nor speeds greater than VFC/MFC, norspeeds that require a stick force of more than 50pounds), with —

(i) The wing flaps retracted;(ii) The center of gravity in the most adverse

position (see §25.27);(iii) The most critical weight between the maxi-

mum takeoff and maximum landing weights;(iv) 75 percent of maximum continuous power

for reciprocating engines or for turbine engines,the maximum cruising power selected by the ap-plicant as an operating limitation (see §25.1521),except that the power need not exceed that re-quired at VMO/MMO; and

(v) The airplane trimmed for level flight withthe power required in paragraph (b)(1)(iv) of thissection.

(2) With the landing gear retracted at lowspeed, the stick force curve must have a stableslope at all speeds within a range which is thegreater of 15 percent of the trim speed plus theresulting free return speed range, or 50 knots plusthe resulting free return speed range, above andbelow the trim speed (except that the speedrange need not include speeds less than 1.3VS R1, nor speeds greater than the minimumspeed of the applicable speed range prescribed inparagraph (b)(1), nor speeds that require a stickforce of more than 50 pounds), with—

(i) Wing flaps, center of gravity position, andweight as specified in paragraph (b)(1) of thissection;

(ii) Power required for level flight at a speedequal to (VMO + 1.3 VSR1) /2; and

(iii) The airplane trimmed for level flight with thepower required in paragraph (b)(2)(ii) of this sec-tion.

(3) With the landing gear extended, the stickforce curve must have a stable slope at all speedswithin a range which is the greater of 15 percentof the trim speed plus the resulting free returnspeed range, or 50 knots plus the resulting freereturn speed range, above and below the trimspeed (except that the speed range need not in-clude speeds less than 1.3 VS R1, nor speedsgreater than VLE, nor speeds that require a stickforce of more than 50 pounds), with—

(i) Wing flap, center of gravity position, andweight as specified in paragraph (b)(1) of thissection;

(ii) 75 percent of maximum continuous powerfor reciprocating engines or, for turbine engines,the maximum cruising power selected by the ap-plicant as an operating limitation, except that the


ASA 29

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power need not exceed that required for levelflight at VLE; and

(iii) The aircraft trimmed for level flight with thepower required in paragraph (b)(3)(ii) of this sec-tion.

(c) Approach. The stick force curve must havea stable slope at speeds between VS W and 1.7VSR1, with—

(1) Wing flaps in the approach position;(2) Landing gear retracted;(3) Maximum landing weight; and(4) The airplane trimmed at 1.3 VS R1 with

enough power to maintain level flight at thisspeed.

(d) Landing. The stick force curve must have astable slope, and the stick force may not exceed80 pounds, at speeds between VS W and 1.7 VSR0with—

(1) Wing flaps in the landing position;(2) Landing gear extended;(3) Maximum landing weight;(4) Power or thrust off on the engines; and(5) The airplane trimmed at 1.3 VSR0 with

power or thrust off.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–7, 30 FR 13117, Oct. 15, 1965;Amdt. 25–108, 67 FR 70827, Nov. 26, 2002]

§25.177 Static lateral-directional stability.(a) [Reserved](b) [Reserved](c) In straight, steady sideslips, the aileron and

rudder control movements and forces must besubstantially proportional to the angle of sideslipin a stable sense; and the factor of proportionalitymust lie between limits found necessary for safeoperation throughout the range of sideslip anglesappropriate to the operation of the airplane. Atgreater angles, up to the angle at which full rudderis used or a rudder force of 180 pounds is ob-tained, the rudder pedal forces may not reverse;and increased rudder deflection must be neededfor increased angles of sideslip. Compliance withthis paragraph must be demonstrated for all land-ing gear and flap positions and symmetricalpower conditions at speeds from 1.13 VSR1 toVFE, VLE, or VFC/MFC, as appropriate.

(d) The rudder gradients must meet the re-quirements of paragraph (c) at speeds betweenVMO/MMO and VFC/MFC except that the dihedraleffect (aileron deflection opposite the correspond-ing rudder input) may be negative provided the di-vergence is gradual, easily recognized, and easilycontrolled by the pilot.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–72, 55 FR 29774, July 20, 1990;55 FR 37607, Sept. 12, 1990; Amdt. 25–108, 67 FR70827, Nov. 26, 2002]

§25.181 Dynamic stability.(a) Any short period oscillation, not including

combined lateral-directional oscillations, occur-ring between 1.13 VSR and maximum allowablespeed appropriate to the configuration of the air-plane must be heavily damped with the primarycontrols—

(1) Free; and(2) In a fixed position.(b) Any combined lateral-directional oscilla-

tions (“Dutch roll”) occurring between 1.13 VSRand maximum allowable speed appropriate to theconfiguration of the airplane must be positivelydamped with controls free, and must be controlla-ble with normal use of the primary controls with-out requiring exceptional pilot skill.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–42, 43 FR 2322, Jan. 16, 1978;Amdt. 25–72, 55 FR 29775, July 20, 1990; 55 FR 37607,Sept. 12, 1990; Amdt. 25–108, 67 FR 70827, Nov. 26,2002]

STALLS

§25.201 Stall demonstration.(a) Stalls must be shown in straight flight and in

30 degree banked turns with—(1) Power off; and(2) The power necessary to maintain level flight

at 1.5 VSR1 (where VSR1 corresponds to the refer-ence stall speed at maximum landing weight withflaps in the approach position and the landinggear retracted).

(b) In each condition required by paragraph (a)of this section, it must be possible to meet the ap-plicable requirements of §25.203 with—

(1) Flaps, landing gear, and deceleration de-vices in any likely combination of positions ap-proved for operation;

(2) Representative weights within the range forwhich certification is requested;

(3) The most adverse center of gravity for re-covery; and

(4) The airplane trimmed for straight flight atthe speed prescribed in §25.103(b)(6).

(c) The following procedures must be used toshow compliance with §25.203;

(1) Starting at a speed sufficiently above thestalling speed to ensure that a steady rate ofspeed reduction can be established, apply thelongitudinal control so that the speed reductiondoes not exceed one knot per second until the air-plane is stalled.

(2) In addition, for turning flight stalls, apply thelongitudinal control to achieve airspeed decelera-tion rates up to 3 knots per second.

(3) As soon as the airplane is stalled, recoverby normal recovery techniques.


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(d) The airplane is considered stalled when thebehavior of the airplane gives the pilot a clear anddistinctive indication of an acceptable nature thatthe airplane is stalled. Acceptable indications of astall, occurring either individually or in combina-tion, are—

(1) A nose-down pitch that cannot be readily ar-rested;

(2) Buffeting, of a magnitude and severity thatis a strong and effective deterrent to further speedreduction; or

(3) The pitch control reaches the aft stop andno further increase in pitch attitude occurs whenthe control is held full aft for a short time beforerecovery is initiated.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–38, 41 FR 55466, Dec. 20, 1976;Amdt. 25–42, 43 FR 2322, Jan. 16, 1978; Amdt. 25–84,60 FR 30750, June 9, 1995; Amdt. 25–108, 67 FR70827, Nov. 26, 2002]

§25.203 Stall characteristics.(a) It must be possible to produce and to cor-

rect roll and yaw by unreversed use of the aileronand rudder controls, up to the time the airplane isstalled. No abnormal nose-up pitching may occur.The longitudinal control force must be positive upto and throughout the stall. In addition, it must bepossible to promptly prevent stalling and to re-cover from a stall by normal use of the controls.

(b) For level wing stalls, the roll occurring be-tween the stall and the completion of the recoverymay not exceed approximately 20 degrees.

(c) For turning flight stalls, the action of the air-plane after the stall may not be so violent or ex-treme as to make it difficult, with normal pilotingskill, to effect a prompt recovery and to regaincontrol of the airplane. The maximum bank anglethat occurs during the recovery may not exceed—

(1) Approximately 60 degrees in the original di-rection of the turn, or 30 degrees in the oppositedirection, for deceleration rates up to 1 knot persecond; and

(2) Approximately 90 degrees in the original di-rection of the turn, or 60 degrees in the oppositedirection, for deceleration rates in excess of 1knot per second.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–84, 60 FR 30750, June 9, 1995]

§25.207 Stall warning.(a) Stall warning with sufficient margin to pre-

vent inadvertent stalling with the flaps and landinggear in any normal position must be clear and dis-tinctive to the pilot in straight and turning flight.

(b) The warning must be furnished eitherthrough the inherent aerodynamic qualities of theairplane or by a device that will give clearly distin-guishable indications under expected conditions

of flight. However, a visual stall warning devicethat requires the attention of the crew within thecockpit is not acceptable by itself. If a warning de-vice is used, it must provide a warning in each ofthe airplane configurations prescribed in para-graph (a) of this section at the speed prescribed inparagraphs (c) and (d) of this section. Except forshowing compliance with the stall warning marginprescribed in paragraph (h)(3)(ii) of this section,stall warning for flight in icing conditions must beprovided by the same means as stall warning forflight in non-icing conditions.

(c) When the speed is reduced at rates not ex-ceeding one knot per second, stall warning mustbegin, in each normal configuration, at a speed,VSW, exceeding the speed at which the stall isidentified in accordance with §25.201(d) by notless than five knots or five percent CAS, which-ever is greater. Once initiated, stall warning mustcontinue until the angle of attack is reduced to ap-proximately that at which stall warning began.

(d) In addition to the requirement of paragraph(c) of this section, when the speed is reduced atrates not exceeding one knot per second, instraight flight with engines idling and at the cen-ter-of-gravity position specified in §25.103(b)(5),VS W, in each normal configuration, must exceedVSR by not less than three knots or three percentCAS, whichever is greater.

(e) In icing conditions, the stall warning marginin straight and turning flight must be sufficient toallow the pilot to prevent stalling (as defined in§25.201(d)) when the pilot starts a recovery ma-neuver not less than three seconds after the on-set of stall warning. When demonstrating compli-ance with this paragraph, the pilot must performthe recovery maneuver in the same way as for theairplane in non-icing conditions. Compliance withthis requirement must be demonstrated in flightwith the speed reduced at rates not exceedingone knot per second, with—

(1) The more critical of the takeoff ice and finaltakeoff ice accretions defined in appendix C foreach configuration used in the takeoff phase offlight;

(2) The en route ice accretion defined in appen-dix C for the en route configuration;

(3) The holding ice accretion defined in appen-dix C for the holding configuration(s);

(4) The approach ice accretion defined in ap-pendix C for the approach configuration(s); and

(5) The landing ice accretion defined in appen-dix C for the landing and go-around configura-tion(s).

(f) The stall warning margin must be sufficientin both non-icing and icing conditions to allow thepilot to prevent stalling when the pilot starts a re-covery maneuver not less than one second afterthe onset of stall warning in slow-down turns withat least 1.5 g load factor normal to the flight path


ASA 31

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and airspeed deceleration rates of at least 2 knotsper second. When demonstrating compliancewith this paragraph for icing conditions, the pilotmust perform the recovery maneuver in the sameway as for the airplane in non-icing conditions.Compliance with this requirement must be dem-onstrated in flight with—

(1) The flaps and landing gear in any normalposition;

(2) The airplane trimmed for straight flight at aspeed of 1.3 VSR; and

(3) The power or thrust necessary to maintainlevel flight at 1.3 VSR.

(g) Stall warning must also be provided in eachabnormal configuration of the high lift devices thatis likely to be used in flight following system fail-ures (including all configurations covered by Air-plane Flight Manual procedures).

(h) For flight in icing conditions before the iceprotection system has been activated and is per-forming its intended function, with the ice accre-tion defined in appendix C, part II(e) of this part,the stall warning margin in straight and turningflight must be sufficient to allow the pilot to pre-vent stalling without encountering any adverseflight characteristics when:

(1) The speed is reduced at rates not exceed-ing one knot per second;

(2) The pilot performs the recovery maneuverin the same way as for flight in non-icing condi-tions; and

(3) The recovery maneuver is started no earlierthan:

(i) One second after the onset of stall warning ifstall warning is provided by the same means asfor flight in non-icing conditions; or

(ii) Three seconds after the onset of stall warn-ing if stall warning is provided by a differentmeans than for flight in non-icing conditions.

(i) In showing compliance with paragraph (h) ofthis section, if stall warning is provided by a differ-ent means in icing conditions than for non-icingconditions, compliance with §25.203 must beshown using the accretion defined in appendix C,part II(e) of this part. Compliance with this re-quirement must be shown using the demonstra-tion prescribed by §25.201, except that the decel-eration rates of §25.201(c)(2) need not be dem-onstrated.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–7, 30 FR 13118, Oct. 15, 1965;Amdt. 25–42, 43 FR 2322, Jan. 16, 1978; Amdt. 25–108,67 FR 70827, Nov. 26, 2002; Amdt. 25–121, 72 FR44668, Aug. 8, 2007; Amdt. 25–129, 74 FR 38339, Aug.3, 2009]

GROUND AND WATER HANDLING CHARACTERISTICS

§25.231 Longitudinal stability and control.(a) Landplanes may have no uncontrollable

tendency to nose over in any reasonably ex-pected operating condition or when rebound oc-curs during landing or takeoff. In addition—

(1) Wheel brakes must operate smoothly andmay not cause any undue tendency to nose over;and

(2) If a tail-wheel landing gear is used, it mustbe possible, during the takeoff ground run on con-crete, to maintain any attitude up to thrust linelevel, at 75 percent of VSR1.

(b) For seaplanes and amphibians, the mostadverse water conditions safe for takeoff, taxiing,and landing, must be established.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–108, 67 FR 70828, Nov. 26,2002]

§25.233 Directional stability and control.(a) There may be no uncontrollable ground-

looping tendency in 90° cross winds, up to a windvelocity of 20 knots or 0.2 VSR0, whichever isgreater, except that the wind velocity need not ex-ceed 25 knots at any speed at which the airplanemay be expected to be operated on the ground.This may be shown while establishing the 90°cross component of wind velocity required by§25.237.

(b) Landplanes must be satisfactorily controlla-ble, without exceptional piloting skill or alertness,in power-off landings at normal landing speed,without using brakes or engine power to maintaina straight path. This may be shown during power-off landings made in conjunction with other tests.

(c) The airplane must have adequate direc-tional control during taxiing. This may be shownduring taxiing prior to takeoffs made in conjunc-tion with other tests.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5671, April 8, 1970;Amdt. 25–42, 43 FR 2322, Jan. 16, 1978; Amdt. 25–94,63 FR 8848, Feb. 23, 1998; Amdt. 25–108, 67 FR 70828,Nov. 26, 2002]


32 ASA

§25.235 Taxiing condition.The shock absorbing mechanism may not

damage the structure of the airplane when the air-plane is taxied on the roughest ground that mayreasonably be expected in normal operation.

§25.237 Wind velocities.(a) For land planes and amphibians, the follow-

ing applies:(1) A 90-degree cross component of wind ve-

locity, demonstrated to be safe for takeoff andlanding, must be established for dry runways andmust be at least 20 knots or 0.2 VSR0, whicheveris greater, except that it need not exceed 25knots.

(2) The crosswind component for takeoff estab-lished without ice accretions is valid in icing condi-tions.

(3) The landing crosswind component must beestablished for:

(i) Non-icing conditions, and(ii) Icing conditions with the landing ice accre-

tion defined in appendix C.(b) For seaplanes and amphibians, the follow-

ing applies:(1) A 90-degree cross component of wind ve-

locity, up to which takeoff and landing is safe un-der all water conditions that may reasonably beexpected in normal operation, must be estab-lished and must be at least 20 knots or 0.2 VSR0,whichever is greater, except that it need not ex-ceed 25 knots.

(2) A wind velocity, for which taxiing is safe inany direction under all water conditions that mayreasonably be expected in normal operation,must be established and must be at least 20 knotsor 0.2 VSR0, whichever is greater, except that itneed not exceed 25 knots.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–42, 43 FR 2322, Jan. 16, 1978;Amdt. 25–108, 67 FR 70828, Nov. 26, 2002; Amdt. 25–121, 72 FR 44668, Aug. 8, 2007]

§25.239 Spray characteristics, control, and stability on water.(a) For seaplanes and amphibians, during

take-off, taxiing, and landing, and in the condi-tions set forth in paragraph (b) of this section,there may be no—

(1) Spray characteristics that would impair thepilot’s view, cause damage, or result in the takingin of an undue quantity of water;

(2) Dangerously uncontrollable porpoising,bounding, or swinging tendency; or

(3) Immersion of auxiliary floats or sponsons,wing tips, propeller blades, or other parts not de-signed to withstand the resulting water loads.

(b) Compliance with the requirements of para-graph (a) of this section must be shown—

(1) In water conditions, from smooth to themost adverse condition established in accor-dance with §25.231;

(2) In wind and cross-wind velocities, water cur-rents, and associated waves and swells that mayreasonably be expected in operation on water;

(3) At speeds that may reasonably be expectedin operation on water;

(4) With sudden failure of the critical engine atany time while on water; and

(5) At each weight and center of gravity posi-tion, relevant to each operating condition, withinthe range of loading conditions for which certifica-tion is requested.

(c) In the water conditions of paragraph (b) ofthis section, and in the corresponding wind condi-tions, the seaplane or amphibian must be able todrift for five minutes with engines inoperative,aided, if necessary, by a sea anchor.

MISCELLANEOUS FLIGHT REQUIREMENTS

§25.251 Vibration and buffeting.(a) The airplane must be demonstrated in flight

to be free from any vibration and buffeting thatwould prevent continued safe flight in any likelyoperating condition.

(b) Each part of the airplane must be demon-strated in flight to be free from excessive vibrationunder any appropriate speed and power condi-tions up to VDF/MDF. The maximum speeds shownmust be used in establishing the operating limita-tions of the airplane in accordance with §25.1505.

(c) Except as provided in paragraph (d) of thissection, there may be no buffeting condition, innormal flight, including configuration changesduring cruise, severe enough to interfere with thecontrol of the airplane, to cause excessive fatigueto the crew, or to cause structural damage. Stallwarning buffeting within these limits is allowable.

(d) There may be no perceptible buffeting con-dition in the cruise configuration in straight flight atany speed up to VMO/MMO, except that stall warn-ing buffeting is allowable.

(e) For an airplane with MD greater than .6 orwith a maximum operating altitude greater than25,000 feet, the positive maneuvering load factorsat which the onset of perceptible buffeting occursmust be determined with the airplane in the cruiseconfiguration for the ranges of airspeed or Machnumber, weight, and altitude for which the air-plane is to be certificated. The envelopes of loadfactor, speed, altitude, and weight must provide asufficient range of speeds and load factors fornormal operations. Probable inadvertent excur-sions beyond the boundaries of the buffet onsetenvelopes may not result in unsafe conditions.


ASA 33

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[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5671, April 8, 1970;Amdt. 25–72, 55 FR 29775, July 20, 1990; Amdt. 25–77,57 FR 28949, June 29, 1992]

§25.253 High-speed characteristics.(a) Speed increase and recovery characteris-

tics. The following speed increase and recoverycharacteristics must be met:

(1) Operating conditions and characteristicslikely to cause inadvertent speed increases (in-cluding upsets in pitch and roll) must be simulatedwith the airplane trimmed at any likely cruisespeed up to VMO/MMO. These conditions andcharacteristics include gust upsets, inadvertentcontrol movements, low stick force gradient in re-lation to control friction, passenger movement,leveling off from climb, and descent from Mach toairspeed limit altitudes.

(2) Allowing for pilot reaction time after effectiveinherent or artificial speed warning occurs, it mustbe shown that the airplane can be recovered to anormal attitude and its speed reduced toVMO/MMO, without—

(i) Exceptional piloting strength or skill;(ii) Exceeding VD/MD, VDF/MDF, or the structural

limitations; and(iii) Buffeting that would impair the pilot’s ability

to read the instruments or control the airplane forrecovery.

(3) With the airplane trimmed at any speed upto VMO/MMO, there must be no reversal of the re-sponse to control input about any axis at anyspeed up to VDF/MDF. Any tendency to pitch, roll,or yaw must be mild and readily controllable, us-ing normal piloting techniques. When the airplaneis trimmed at VMO/MMO, the slope of the elevatorcontrol force versus speed curve need not be sta-ble at speeds greater than VFC/MFC, but theremust be a push force at all speeds up to VDF/MDFand there must be no sudden or excessive reduc-tion of elevator control force as VDF/MDF isreached.

(b) Maximum speed for stability characteristics.VFC/MFC. VFC/MFC is the maximum speed atwhich the requirements of §§25.143(g),25.147(E), 25.175(b)(1), 25.177, and 25.181 mustbe met with flaps and landing gear retracted. Ex-cept as noted in §25.253(c), VFC/MFC may not beless than a speed midway between VMO/MMO andVDF/MDF, except that for altitudes where Machnumber is the limiting factor, MFC need not ex-ceed the Mach number at which effective speedwarning occurs.

(c) Maximum speed for stability characteristicsin icing conditions. The maximum speed for stabil-ity characteristics with the ice accretions definedin appendix C, at which the requirements of§§25.143(g), 25.147(e), 25.175(b)(1), 25.177,and 25.181 must be met, is the lower of:

(1) 300 knots CAS;(2) VFC; or(3) A speed at which it is demonstrated that the

airframe will be free of ice accretion due to the ef-fects of increased dynamic pressure.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5671, April 8, 1970;Amdt. 25–54, 45 FR 60172, Sept. 11, 1980; Amdt. 25–72, 55 FR 29775, July 20, 1990; Amdt. 25–84, 60 FR30750, June 9, 1995; Amdt. 25–121, 72 FR 44668, Aug.8, 2007]

§25.255 Out-of-trim characteristics.(a) From an initial condition with the airplane

trimmed at cruise speeds up to VMO/MMO, the air-plane must have satisfactory maneuvering stabil-ity and controllability with the degree of out-of-trimin both the airplane nose-up and nose-down di-rections, which results from the greater of—

(1) A three-second movement of the longitudi-nal trim system at its normal rate for the particularflight condition with no aerodynamic load (or anequivalent degree of trim for airplanes that do nothave a power-operated trim system), except aslimited by stops in the trim system, includingthose required by §25.655(b) for adjustable stabi-lizers; or

(2) The maximum mistrim that can be sus-tained by the autopilot while maintaining levelflight in the high speed cruising condition.

(b) In the out-of-trim condition specified inparagraph (a) of this section, when the normal ac-celeration is varied from +1 g to the positive andnegative values specified in paragraph (c) of thissection—

(1) The stick force vs. g curve must have a pos-itive slope at any speed up to and includingVFC/MFC; and

(2) At speeds between VFC/MFC and VDF/MDFthe direction of the primary longitudinal controlforce may not reverse.

(c) Except as provided in paragraphs (d) and(e) of this section, compliance with the provisionsof paragraph (a) of this section must be demon-strated in flight over the acceleration range—

(1) -1 g to +2.5 g; or(2) 0 g to 2.0 g, and extrapolating by an accept-

able method to -1 g and +2.5 g.(d) If the procedure set forth in paragraph (c)(2)

of this section is used to demonstrate complianceand marginal conditions exist during flight testwith regard to reversal of primary longitudinalcontrol force, flight tests must be accomplishedfrom the normal acceleration at which a marginalcondition is found to exist to the applicable limitspecified in paragraph (b)(1) of this section.

(e) During flight tests required by paragraph (a)of this section, the limit maneuvering load factorsprescribed in §§25.333(b) and 25.337, and themaneuvering load factors associated with proba-


34 ASA

ble inadvertent excursions beyond the boundariesof the buffet onset envelopes determined under§25.251(e), need not be exceeded. In addition,the entry speeds for flight test demonstrations atnormal acceleration values less than 1 g must belimited to the extent necessary to accomplish arecovery without exceeding VDF/MDF.

(f) In the out-of-trim condition specified in para-graph (a) of this section, it must be possible froman overspeed condition at VDF/MDF to produce atleast 1.5 g for recovery by applying not more than125 pounds of longitudinal control force using ei-ther the primary longitudinal control alone or theprimary longitudinal control and the longitudinaltrim system. If the longitudinal trim is used to as-sist in producing the required load factor, it mustbe shown at VDF/MDF that the longitudinal trimcan be actuated in the airplane nose-up directionwith the primary surface loaded to correspond tothe least of the following airplane nose-up controlforces:

(1) The maximum control forces expected inservice as specified in §§25.301 and 25.397.

(2) The control force required to produce 1.5 g.(3) The control force corresponding to buffeting

or other phenomena of such intensity that it is astrong deterrent to further application of primarylongitudinal control force.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–42, 43 FR 2322, Jan. 16, 1978]


ASA 35

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Subpart C — Structure

GENERAL

§25.301 Loads.(a) Strength requirements are specified in

terms of limit loads (the maximum loads to be ex-pected in service) and ultimate loads (limit loadsmultiplied by prescribed factors of safety). Unlessotherwise provided, prescribed loads are limitloads.

(b) Unless otherwise provided, the specifiedair, ground, and water loads must be placed inequilibrium with inertia forces, considering eachitem of mass in the airplane. These loads must bedistributed to conservatively approximate orclosely represent actual conditions. Methodsused to determine load intensities and distributionmust be validated by flight load measurement un-less the methods used for determining thoseloading conditions are shown to be reliable.

(c) If deflections under load would significantlychange the distribution of external or internalloads, this redistribution must be taken into ac-count.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5672, April 8, 1970]

§25.303 Factor of safety.Unless otherwise specified, a factor of safety of

1.5 must be applied to the prescribed limit loadwhich are considered external loads on the struc-ture. When a loading condition is prescribed interms of ultimate loads, a factor of safety need notbe applied unless otherwise specified.


§25.305 Strength and deformation.(a) The structure must be able to support limit

loads without detrimental permanent deformation.At any load up to limit loads, the deformation maynot interfere with safe operation.

(b) The structure must be able to support ulti-mate loads without failure for at least 3 seconds.However, when proof of strength is shown by dy-namic tests simulating actual load conditions, the3-second limit does not apply. Static tests con-ducted to ultimate load must include the ultimatedeflections and ultimate deformation induced bythe loading. When analytical methods are used toshow compliance with the ultimate load strengthrequirements, it must be shown that—

(1) The effects of deformation are not significant;(2) The deformations involved are fully ac-

counted for in the analysis; or(3) The methods and assumptions used are suf-

ficient to cover the effects of these deformations.

(c) Where structural flexibility is such that anyrate of load application likely to occur in the oper-ating conditions might produce transient stressesappreciably higher than those corresponding tostatic loads, the effects of this rate of applicationmust be considered.

(d) [Reserved](e) The airplane must be designed to withstand

any vibration and buffeting that might occur in anylikely operating condition up to VD/MD, includingstall and probable inadvertent excursions beyondthe boundaries of the buffet onset envelope. Thismust be shown by analysis, flight tests, or othertests found necessary by the Administrator.

(f) Unless shown to be extremely improbable,the airplane must be designed to withstand anyforced structural vibration resulting from any fail-ure, malfunction or adverse condition in the flightcontrol system. These must be considered limitloads and must be investigated at airspeeds up toVC/MC.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5672, April 8, 1970;Amdt. 25–54, 45 FR 60172, Sept. 11, 1980; Amdt. 25–77, 57 FR 28949, June 29, 1992; Amdt. 25–86, 61 FR5220, Feb. 9, 1996]

§25.307 Proof of structure.(a) Compliance with the strength and deforma-

tion requirements of this subpart must be shownfor each critical loading condition. Structural anal-ysis may be used only if the structure conforms tothat for which experience has shown this methodto be reliable. The Administrator may require ulti-mate load tests in cases where limit load testsmay be inadequate.

(b) – (c) [Reserved](d) When static or dynamic tests are used to

show compliance with the requirements of§25.305(b) for flight structures, appropriate mate-rial correction factors must be applied to the testresults, unless the structure, or part thereof, beingtested has features such that a number of ele-ments contribute to the total strength of the struc-ture and the failure of one element results in theredistribution of the load through alternate loadpaths.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5672, April 8, 1970;Amdt. 25–54, 45 FR 60172, Sept. 11, 1980; Amdt. 25–72, 55 FR 29775, July 20, 1990]

FLIGHT LOADS

§25.321 General.(a) Flight load factors represent the ratio of the

aerodynamic force component (acting normal tothe assumed longitudinal axis of the airplane) tothe weight of the airplane. A positive load factor is


36 ASA

one in which the aerodynamic force acts upwardwith respect to the airplane.

(b) Considering compressibility effects at eachspeed, compliance with the flight load require-ments of this subpart must be shown—

(1) At each critical altitude within the range ofaltitudes selected by the applicant;

(2) At each weight from the design minimumweight to the design maximum weight appropriateto each particular flight load condition; and

(3) For each required altitude and weight, forany practicable distribution of disposable loadwithin the operating limitations recorded in theAirplane Flight Manual.

(c) Enough points on and within the boundariesof the design envelope must be investigated toensure that the maximum load for each part of theairplane structure is obtained.

(d) The significant forces acting on the airplanemust be placed in equilibrium in a rational or con-servative manner. The linear inertia forces mustbe considered in equilibrium with the thrust andall aerodynamic loads, while the angular (pitch-ing) inertia forces must be considered in equilib-rium with thrust and all aerodynamic moments, in-cluding moments due to loads on componentssuch as tail surfaces and nacelles. Critical thrustvalues in the range from zero to maximum contin-uous thrust must be considered.


FLIGHT MANEUVER AND GUST CONDITIONS

§25.331 Symmetric maneuvering conditions.(a) Procedure. For the analysis of the maneu-

vering flight conditions specified in paragraphs(b) and (c) of this section, the following provisionsapply:

(1) Where sudden displacement of a control isspecified, the assumed rate of control surface dis-placement may not be less than the rate thatcould be applied by the pilot through the controlsystem.

(2) In determining elevator angles and chord-wise load distribution in the maneuvering condi-tions of paragraphs (b) and (c) of this section, theeffect of corresponding pitching velocities must betaken into account. The in-trim and out-of-trimflight conditions specified in §25.255 must be con-sidered.

(b) Maneuvering balanced conditions. Assum-ing the airplane to be in equilibrium with zero pitch-ing acceleration, the maneuvering conditions Athrough I on the maneuvering envelope in§25.333(b) must be investigated.

(c) Pitch maneuver conditions. The conditionsspecified in paragraphs (c)(1) and (2) of this sec-

tion must be investigated. The movement of thepitch control surfaces may be adjusted to take intoaccount limitations imposed by the maximum piloteffort specified by §25.397(b), control systemstops and any indirect effect imposed by limita-tions in the output side of the control system (forexample, stalling torque or maximum rate obtain-able by a power control system.)

(1) Maximum pitch control displacement at VA.The airplane is assumed to be flying in steadylevel flight (point A1, §25.333(b)) and the cockpitpitch control is suddenly moved to obtain extremenose up pitching acceleration. In defining the tailload, the response of the airplane must be takeninto account. Airplane loads that occur subse-quent to the time when normal acceleration at thec.g. exceeds the positive limit maneuvering loadfactor (at point A2 in §25.333(b)), or the resultingtailplane normal load reaches its maximum,whichever occurs first, need not be considered.

(2) Specified control displacement. A checkedmaneuver, based on a rational pitching controlmotion vs. time profile, must be established inwhich the design limit load factor specified in§25.337 will not be exceeded. Unless lesser val-ues cannot be exceeded, the airplane responsemust result in pitching accelerations not less thanthe following:

(i) A positive pitching acceleration (nose up) isassumed to be reached concurrently with the air-plane load factor of 1.0 (Points A1 to D1,§25.333(b)). The positive acceleration must beequal to at least

where—

n is the positive load factor at the speed under consideration, and V is the airplane equivalent speed in knots.

(ii) A negative pitching acceleration (nosedown) is assumed to be reached concurrentlywith the positive maneuvering load factor (pointsA2 to D2, §25.333(b)). This negative pitching ac-celeration must be equal to at least

where—

n is the positive load factor at the speed under consideration; and V is the airplane equivalent speed in knots.

39nv

--------- n 1.5–( ) Radians sec.2⁄( ),

26n–v

------------ n 1.5–( ) Radians sec.2⁄( ),


ASA 37

25

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5672, April 8, 1970;Amdt. 25–46, 43 FR 50594, Oct. 30, 1978; 43 FR 52495,Nov. 13, 1978; 43 FR 54082, Nov. 20, 1978; Amdt. 25–72, 55 FR 29775, July 20, 1990; 55 FR 37607, Sept. 12,1990; Amdt. 25–86, 61 FR 5220, Feb. 9, 1996; Amdt.25–91, 62 FR 40704, July 29, 1997]

§25.333 Flight maneuvering envelope.(a) General. The strength requirements must

be met at each combination of airspeed and loadfactor on and within the boundaries of the repre-sentative maneuvering envelope (V-n diagram) ofparagraph (b) of this section. This envelope mustalso be used in determining the airplane struc-tural operating limitations as specified in§25.1501.

(b) Maneuvering envelope.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; as amended by Amdt. 25–86, 61 FR 5220, Feb. 9, 1996]

§25.335 Design airspeeds.The selected design airspeeds are equivalent

airspeeds (EAS). Estimated values of VS0 andVS1 must be conservative.

(a) Design cruising speed, VC. For VC, the fol-lowing apply:

(1) The minimum value of VC must be suffi-ciently greater than VB to provide for inadvertentspeed increases likely to occur as a result of se-vere atmospheric turbulence.

(2) Except as provided in §25.335(d)(2), VCmay not be less than VB + 1.32 UREF (with UREFas specified in §25.341(a)(5)(i)). However VCneed not exceed the maximum speed in levelflight at maximum continuous power for the corre-sponding altitude.

(3) At altitudes where VD is limited by Machnumber, VC may be limited to a selected Machnumber.

(b) Design dive speed, VD. VD must be selectedso that VC/MC is not greater than 0.8 VD/MD, or sothat the minimum speed margin between VC/MCand VD/MD is the greater of the following values:

(1) From an initial condition of stabilized flight atVC/MC, the airplane is upset, flown for 20 secondsalong a flight path 7.5° below the initial path, andthen pulled up at a load factor of 1.5 g (0.5 g ac-celeration increment). The speed increase occur-ring in this maneuver may be calculated if reliableor conservative aerodynamic data is used. Poweras specified in §25.175(b)(1)(iv) is assumed untilthe pullup is initiated, at which time power reduc-

tion and the use of pilot controlled drag devices may be assumed;

D, D2

D1

E

VD

V C“EQUIVALENT” AIR SPEED

FLAPS DOWN

LOA

D F

AC

TOR

, n

+ CN MAX

+ CN MAX

FLAPS UP3

2

1

0

-1

- CN MAX

FLAPS UP

H F

A2

A1

AV

A

VF

VS

1

I


38 ASA

(2) The minimum speed margin must beenough to provide for atmospheric variations(such as horizontal gusts, and penetration of jetstreams and cold fronts) and for instrument errorsand airframe production variations. These factorsmay be considered on a probability basis. Themargin at altitude where MC is limited by com-pressibility effects must not be less than 0.07Munless a lower margin is determined using a ratio-nal analysis that includes the effects of any auto-matic systems. In any case, the margin may notbe reduced to less than 0.05M.

(c) Design maneuvering speed VA. For VA, thefollowing apply:

(1) VA may not be less than VS1 √ n where—(i) n is the limit positive maneuvering load factor

at VC; and(ii) VS1 is the stalling speed with flaps retracted.(2) VA and VS must be evaluated at the design

weight and altitude under consideration.(3) VA need not be more than VC or the speed

at which the positive CNmax curve intersects thepositive maneuver load factor line, whichever isless.

(d) Design speed for maximum gust intensity,VB.

(1) VB may not be less than

where—

VS1 = the 1-g stalling speed based on CNAmax with the flaps retracted at the particular weight under consideration;

VC = design cruise speed (knots equivalent airspeed);

Uref = the reference gust velocity (feet per second equivalent airspeed) from §25.341(a)(5)(i);

w = average wing loading (pounds per square foot) at the particular weight under consideration.

ρ = density of air (slugs/ft3);c = mean geometric chord of the wing (feet);g = acceleration due to gravity (ft/sec2);a = slope of the airplane normal force coefficient

curve, CNA per radian;

(2) At altitudes where VC is limited by Machnumber—

(i) VB may be chosen to provide an optimummargin between low and high speed buffet bound-aries; and,

(ii) VB need not be greater than VC.(e) Design flap speeds, VF. For VF, the following

apply:(1) The design flap speed for each flap position

(established in accordance with §25.697(a)) mustbe sufficiently greater than the operating speedrecommended for the corresponding stage offlight (including balked landings) to allow for prob-able variations in control of airspeed and for tran-sition from one flap position to another.

(2) If an automatic flap positioning or load limit-ing device is used, the speeds and correspondingflap positions programmed or allowed by the de-vice may be used.

(3) VF may not be less than—(i) 1.6 VS1 with the flaps in takeoff position at

maximum takeoff weight;(ii) 1.8 VS1 with the flaps in approach position

at maximum landing weight, and(iii) 1.8 VS0 with the flaps in landing position at

maximum landing weight.(f) Design drag device speeds, VDD. The se-

lected design speed for each drag device must besufficiently greater than the speed recommendedfor the operation of the device to allow for proba-ble variations in speed control. For drag devicesintended for use in high speed descents, VDD maynot be less than VD. When an automatic drag de-vice positioning or load limiting means is used,the speeds and corresponding drag device posi-tions programmed or allowed by the automaticmeans must be used for design.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5672, April 8, 1970;Amdt. 25–86, 61 FR 5220, Feb. 9, 1996; 25–91, 62 FR40704, July 29, 1997]

§25.337 Limit maneuvering load factors.(a) Except where limited by maximum (static)

lift coefficients, the airplane is assumed to be sub-jected to symmetrical maneuvers resulting in thelimit maneuvering load factors prescribed in thissection. Pitching velocities appropriate to the cor-responding pull-up and steady turn maneuversmust be taken into account.

(b) The positive limit maneuvering load factor“n” for any speed up to VN may not be less than2.1+24,000/ (W +10,000) except that “n” may notbe less than 2.5 and need not be greater than3.8 — where “W” is the design maximum takeoffweight.

(c) The negative limit maneuvering load fac-tor—

(1) May not be less than –1.0 at speeds up toVC; and

V S1 1KgUref V ca

498w----------------------------+

1 2⁄

Kg.88µ

5.3 µ+-----------------=

µ 2wρcag-------------=


ASA 39

25

(2) Must vary linearly with speed from the valueat VC to zero at VD.

(d) Maneuvering load factors lower than thosespecified in this section may be used if the air-plane has design features that make it impossibleto exceed these values in flight.


§25.341 Gust and turbulence loads.(a) Discrete Gust Design Criteria. The airplane

is assumed to be subjected to symmetrical verti-cal and lateral gusts in level flight. Limit gust loadsmust be determined in accordance with the provi-sions:

(1) Loads on each part of the structure must bedetermined by dynamic analysis. The analysismust take into account unsteady aerodynamiccharacteristics and all significant structural de-grees of freedom including rigid body motions.

(2) The shape of the gust must be:

for 0 ≤ s ≤ 2H

where—

s = distance penetrated into the gust (feet);Uds = the design gust velocity in equivalent

airspeed specified in paragraph (a)(4) of this section; and

H = the gust gradient which is the distance (feet) parallel to the airplane’s flight path for the gust to reach its peak velocity.

(3) A sufficient number of gust gradient dis-tances in the range 30 feet to 350 feet must be in-vestigated to find the critical response for eachload quantity.

(4) The design gust velocity must be:

where—

Uref = the reference gust velocity in equivalent airspeed defined in paragraph (a)(5) of this section.

Fg = the flight profile alleviation factor defined in paragraph (a)(6) of this section.

(5) The following reference gust velocities apply:(i) At the airplane design speed VC: Positive

and negative gusts with reference gust velocitiesof 56.0 ft/sec EAS must be considered at sealevel. The reference gust velocity may be reduced

linearly from 56.0 ft/sec EAS at sea level to 44.0ft/sec EAS at 15000 feet. The reference gust ve-locity may be further reduced linearly from 44.0ft/sec EAS at 15000 feet to 26.0 ft/sec EAS at50000 feet.

(ii) At the airplane design speed VD: The refer-ence gust velocity must be 0.5 times the value ob-tained under §25.341(a)(5)(i).

(6) The flight profile alleviation factor, Fg, mustbe increased linearly from the sea level value to avalue of 1.0 at the maximum operating altitude de-fined in §25.1527. At sea level, the flight profile al-leviation factor is determined by the followingequation:

Zmo = Maximum operating altitude defined in §25.1527.

(7) When a stability augmentation system is in-cluded in the analysis, the effect of any significantsystem nonlinearities should be accounted forwhen deriving limit loads from limit gust condi-tions.

(b) Continuous Gust Design Criteria. The dy-namic response of the airplane to vertical and lat-eral continuous turbulence must be taken into ac-count. The continuous gust design criteria of Ap-pendix G of this part must be used to establish thedynamic response unless more rational criteriaare shown.

[Docket No. 27902, 61 FR 5221, Feb. 9, 1996; asamended by Amdt. 25–86, 61 FR 9533, March 8, 1996]

§25.343 Design fuel and oil loads.(a) The disposable load combinations must in-

clude each fuel and oil load in the range from zerofuel and oil to the selected maximum fuel and oilload. A structural reserve fuel condition, not ex-ceeding 45 minutes of fuel under the operatingconditions in §25.1001(e) and (f), as applicable,may be selected.

(b) If a structural reserve fuel condition is se-lected, it must be used as the minimum fuel

UUds

2--------- 1 Cos

πsH-----

–=

Uds Uref FgH

350⁄

1 6⁄=

Fg 0.5 Fgz

Fgm+

=

Where:

Fgz 1Zmo

250000------------------–=

Fgm R2Tan πR1 4⁄( )=

R1Maximum Landing Weight

Maximum Take-off Weight------------------------------------------------------------------=

R2Maximum Zero Fuel Weight

Maximum Take-off Weight---------------------------------------------------------------------=


40 ASA

weight condition for showing compliance with theflight load requirements as prescribed in this sub-part. In addition—

(1) The structure must be designed for a condi-tion of zero fuel and oil in the wing at limit loadscorresponding to—

(i) A maneuvering load factor of +2.25; and(ii) The gust conditions of §25.341(a) but as-

suming 85% of the design velocities prescribed in§25.341(a)(4).

(2) Fatigue evaluation of the structure must ac-count for any increase in operating stresses re-sulting from the design condition of paragraph(b)(1) of this section; and

(3) The flutter, deformation, and vibration re-quirements must also be met with zero fuel.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–18, 33 FR 12226, Aug. 30, 1968;Amdt. 25–72, 55 FR 37607, Sept. 12, 1990; Amdt. 25–86, 61 FR 5221, Feb. 9, 1996]

§25.345 High lift devices.(a) If wing flaps are to be used during takeoff,

approach, or landing, at the design flap speedsestablished for these stages of flight under§25.335(e) and with the wing flaps in the corre-sponding positions, the airplane is assumed to besubjected to symmetrical maneuvers and gusts.The resulting limit loads must correspond to theconditions determined as follows:

(1) Maneuvering to a positive limit load factor of2.0; and

(2) Positive and negative gusts of 25 ft/sec EASacting normal to the flight path in level flight. Gustloads resulting on each part of the structure mustbe determined by rational analysis. The analysismust take into account the unsteady aerodynamiccharacteristics and rigid body motions of the air-craft. The shape of the gust must be as describedin §25.341(a)(2) except that—

Uds = 25 ft/sec EAS;H = 12.5 c; andc = mean geometric chord of the wing (feet).

(b) The airplane must be designed for the con-ditions prescribed in paragraph (a) of this section,except that the airplane load factor need not ex-ceed 1.0, taking into account, as separate condi-tions, the effects of—

(1) Propeller slipstream corresponding to maxi-mum continuous power at the design flap speedsVF, and with takeoff power at not less than 1.4times the stalling speed for the particular flap po-sition and associated maximum weight; and

(2) A head-on gust of 25 feet per second veloc-ity (EAS).

(c) If flaps or other high lift devices are to beused in en route conditions, and with flaps in theappropriate position at speeds up to the flap de-

sign speed chosen for these conditions, the air-plane is assumed to be subjected to symmetricalmaneuvers and gusts within the range deter-mined by—

(1) Maneuvering to a positive limit load factoras prescribed in §25.337(b); and

(2) The discrete vertical gust criteria in§25.341(a).

(d) The airplane must be designed for a ma-neuvering load factor of 1.5g at the maximumtake-off weight with the wing-flaps and similarhigh lift devices in the landing configurations.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–46, 43 FR 50595, Oct. 30, 1978;Amdt. 25–72, 55 FR 37607, Sept. 17, 1990; Amdt. 25–86, 61 FR 5221, Feb. 9, 1996; Amdt. 25–91, 62 FR40704, July 29, 1997]

§25.349 Rolling conditions.The airplane must be designed for loads result-

ing from the rolling conditions specified in para-graphs (a) and (b) of this section. Unbalancedaerodynamic moments about the center of gravitymust be reacted in a rational or conservativemanner, considering the principal masses fur-nishing the reacting inertia forces.

(a) Maneuvering. The following conditions,speeds, and aileron deflections (except as the de-flections may be limited by pilot effort) must beconsidered in combination with an airplane loadfactor of zero and of two-thirds of the positive ma-neuvering factor used in design. In determiningthe required aileron deflections, the torsional flex-ibility of the wing must be considered in accor-dance with §25.301(b):

(1) Conditions corresponding to steady rollingvelocities must be investigated. In addition, condi-tions corresponding to maximum angular acceler-ation must be investigated for airplanes with en-gines or other weight concentrations outboard ofthe fuselage. For the angular acceleration condi-tions, zero rolling velocity may be assumed in theabsence of a rational time history investigation ofthe maneuver.

(2) At VA, a sudden deflection of the aileron tothe stop is assumed.

(3) At VC, the aileron deflection must be that re-quired to produce a rate of roll not less than thatobtained in paragraph (a)(2) of this section.

(4) At VD, the aileron deflection must be that re-quired to produce a rate of roll not less than one-third of that in paragraph (a)(2) of this section.

(b) Unsymmetrical gusts. The airplane is as-sumed to be subjected to unsymmetrical verticalgusts in level flight. The resulting limit loads mustbe determined from either the wing maximum air-load derived directly from §25.341(a), or the wingmaximum airload derived indirectly from the verti-cal load factor calculated from §25.341(a). It mustbe assumed that 100 percent of the wing air load


ASA 41

25

acts on one side of the airplane and 80 percent ofthe wing air load acts on the other side.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5672, April 8, 1970;Amdt. 25–86, 61 FR 5222, Feb. 9, 1996; Amdt. 25–94,63 FR 8848, Feb. 23, 1998]

§25.351 Yaw maneuver conditions.The airplane must be designed for loads result-

ing from the yaw maneuver conditions specified inparagraphs (a) through (d) of this section atspeeds from VMC to VD. Unbalanced aerodynamicmoments about the center of gravity must be re-acted in a rational or conservative manner consid-ering the airplane inertia forces. In computing thetail loads the yawing velocity may be assumed tobe zero.

(a) With the airplane in unaccelerated flight atzero yaw, it is assumed that the cockpit ruddercontrol is suddenly displaced to achieve the re-sulting rudder deflection, as limited by:

(1) The control system on control surfacestops; or

(2) A limit pilot force of 300 pounds from VMC toVA and 200 pounds from VC/MC to VD/MD, with alinear variation between VA and VC/MC.

(b) With the cockpit rudder control deflected soas always to maintain the maximum rudder de-flection available within the limitations specified inparagraph (a) of this section, it is assumed thatthe airplane yaws to the overswing sideslip angle.

(c) With the airplane yawed to the static equilib-rium sideslip angle, it is assumed that the cockpitrudder control is held so as to achieve the maxi-mum rudder deflection available within the limita-tions specified in paragraph (a) of this section.

(d) With the airplane yawed to the static equilib-rium sideslip angle of paragraph (c) of this sec-tion, it is assumed that the cockpit rudder controlis suddenly returned to neutral.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5672, April 8, 1970;Amdt. 25–46, 43 FR 50595, Oct. 30, 1978; Amdt. 25–72,55 FR 29775, July 20, 1990; 55 FR 37607, Sept. 12,1990; 55 FR 41415, Oct. 11, 1990; Amdt. 25–86, 61 FR5222, Feb. 9, 1996; Amdt. 25–91, 62 FR 40704, July 29,1997]

SUPPLEMENTARY CONDITIONS

§25.361 Engine torque.(a) Each engine mount and its supporting

structure must be designed for the effects of—(1) A limit engine torque corresponding to take-

off power and propeller speed acting simulta-neously with 75 percent of the limit loads fromflight condition A of §25.333(b);

(2) A limit torque corresponding to the maxi-mum continuous power and propeller speed, act-

ing simultaneously with the limit loads from flightcondition A of §25.333(b); and

(3) For turbopropeller installations, in additionto the conditions specified in paragraphs (a)(1)and (2) of this section, a limit engine torque corre-sponding to takeoff power and propeller speed,multiplied by a factor accounting for propeller con-trol system malfunction, including quick feather-ing, acting simultaneously with 1g level flightloads. In the absence of a rational analysis, a fac-tor of 1.6 must be used.

(b) For turbine engine installations, the enginemounts and supporting structure must be de-signed to withstand each of the following:

(1) A limit engine torque load imposed by sud-den engine stoppage due to malfunction or struc-tural failure (such as compressor jamming).

(2) A limit engine torque load imposed by themaximum acceleration of the engine.

(c) The limit engine torque to be considered un-der paragraph (a) of this section must be obtainedby multiplying mean torque for the specifiedpower and speed by a factor of—

(1) 1.25 for turbopropeller installations;(2) 1.33 for reciprocating engines with five or

more cylinders; or(3) Two, three, or four, for engines with four,

three, or two cylinders, respectively.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5672, April 8, 1970;Amdt. 25–46, 43 FR 50595, Oct. 30, 1978; Amdt. 25–72,55 FR 29776, July 20, 1990]

§25.363 Side load on engine and auxiliary power unit mounts.(a) Each engine and auxiliary power unit mount

and its supporting structure must be designed fora limit load factor in lateral direction, for the sideload on the engine and auxiliary power unitmount, at least equal to the maximum load factorobtained in the yawing conditions but not lessthan—

(1) 1.33; or(2) One-third of the limit load factor for flight

condition A as prescribed in §25.333(b).(b) The side load prescribed in paragraph (a) of

this section may be assumed to be independentof other flight conditions.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5672, April 8, 1970;Amdt. 25–91, 61 FR 40704, July 29, 1997]

§25.365 Pressurized compartment loads.For airplanes with one or more pressurized

compartments the following apply:(a) The airplane structure must be strong

enough to withstand the flight loads combined


42 ASA

with pressure differential loads from zero up to themaximum relief valve setting.

(b) The external pressure distribution in flight,and stress concentrations and fatigue effectsmust be accounted for.

(c) If landings may be made with the compart-ment pressurized, landing loads must be com-bined with pressure differential loads from zero upto the maximum allowed during landing.

(d) The airplane structure must be designed tobe able to withstand the pressure differentialloads corresponding to the maximum relief valvesetting multiplied by a factor of 1.33 for airplanesto be approved for operation to 45,000 feet or by afactor of 1.67 for airplanes to be approved for op-eration above 45,000 feet, omitting other loads.

(e) Any structure, component or part, inside oroutside a pressurized compartment, the failure ofwhich could interfere with continued safe flight andlanding, must be designed to withstand the effectsof a sudden release of pressure through an open-ing in any compartment at any operating altituderesulting from each of the following conditions:

(1) The penetration of the compartment by aportion of an engine following an engine disinte-gration;

(2) Any opening in any pressurized compart-ment up to the size Ho in square feet; however,small compartments may be combined with anadjacent pressurized compartment and both con-sidered as a single compartment for openingsthat cannot reasonably be expected to be con-fined to the small compartment. The size Ho mustbe computed by the following formula:

Ho = PAS

where,

Ho = Maximum opening in square feet, need not exceed 20 square feet.

AS = Maximum cross-sectional area of the pressurized shell normal to the longitudinal axis, in square feet; and

(3) The maximum opening caused by airplaneor equipment failures not shown to be extremelyimprobable.

(f) In complying with paragraph (e) of this sec-tion, the fail-safe features of the design may beconsidered in determining the probability of failureor penetration and probable size of openings, pro-vided that possible improper operation of closuredevices and inadvertent door openings are alsoconsidered. Furthermore, the resulting differentialpressure loads must be combined in a rationaland conservative manner with 1-g level flight

loads and any loads arising from emergency de-pressurization conditions. These loads may beconsidered as ultimate conditions; however, anydeformations associated with these conditionsmust not interfere with continued safe flight andlanding. The pressure relief provided by intercom-partment venting may also be considered.

(g) Bulkheads, floors, and partitions in pressur-ized compartments for occupants must be de-signed to withstand the conditions specified inparagraph (e) of this section. In addition, reason-able design precautions must be taken to mini-mize the probability of parts becoming detachedand injuring occupants while in their seats.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–54, 45 FR 60172, Sept. 11, 1980;Amdt. 25–71, 55 FR 13477, April 10, 1990; Amdt. 25–72,55 FR 29776, July 20, 1990; Amdt. 25–87, 61 FR 28695,June 5, 1996]

§25.367 Unsymmetrical loads due to engine failure.(a) The airplane must be designed for the un-

symmetrical loads resulting from the failure of thecritical engine. Turbopropeller airplanes must bedesigned for the following conditions in combina-tion with a single malfunction of the propeller draglimiting system, considering the probable pilotcorrective action on the flight controls:

(1) At speeds between VMC and VD, the loadsresulting from power failure because of fuel flowinterruption are considered to be limit loads.

(2) At speeds between VMC and VC, the loadsresulting from the disconnection of the enginecompressor from the turbine or from loss of theturbine blades are considered to be ultimateloads.

(3) The time history of the thrust decay anddrag build-up occurring as a result of the pre-scribed engine failures must be substantiated bytest or other data applicable to the particular en-gine-propeller combination.

(4) The timing and magnitude of the probablepilot corrective action must be conservatively esti-mated, considering the characteristics of the par-ticular engine-propeller-airplane combination.

(b) Pilot corrective action may be assumed tobe initiated at the time maximum yawing velocityis reached, but not earlier than two seconds afterthe engine failure. The magnitude of the correc-tive action may be based on the control forcesspecified in §25.397(b) except that lower forcesmay be assumed where it is shown by analysisor test that these forces can control the yaw androll resulting from the prescribed engine failureconditions.

PAS

6240------------ .024+=


ASA 43

25

§25.371 Gyroscopic loads.The structure supporting any engine or auxil-

iary power unit must be designed for the loads in-cluding the gyroscopic loads arising from the con-ditions specified in §§25.331, 25.341(a), 25.349,25.351, 25.473, 25.479, and 25.481, with the en-gine or auxiliary power unit at maximum rpm ap-propriate to the condition. For the purposes ofcompliance with this section, the pitch maneuverin §25.331(c)(1) must be carried out until the pos-itive limit maneuvering load factor (point A2 in§25.333(b)) is reached.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–86, 61 FR 5222, Feb. 9, 1996;Amdt. 25–91, 62 FR 40704, July 29, 1997]

§25.373 Speed control devices.If speed control devices (such as spoilers and

drag flaps) are installed for use in en route condi-tions—

(a) The airplane must be designed for the sym-metrical maneuvers prescribed in §25.333 and§25.337, the yawing maneuvers prescribed in§25.351, and the vertical and later gust conditionsprescribed in §25.341(a), at each setting and themaximum speed associated with that setting; and

(b) If the device has automatic operating orload limiting features, the airplane must be de-signed for the maneuver and gust conditions pre-scribed in paragraph (a) of this section, at thespeeds and corresponding device positions thatthe mechanism allows.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–72, 55 FR 29776, July 20, 1990;Amdt. 25–86, 61 FR 5222, Feb. 9, 1996]

CONTROL SURFACE AND SYSTEM LOADS

§25.391 Control surface loads: general.The control surfaces must be designed for the

limit loads resulting from the flight conditions in§§25.331, 25.341(a), 25.349 and 25.351 and theground gust conditions in §25.415, consideringthe requirements for—

(a) Loads parallel to hinge line, in §25.393;(b) Pilot effort effects, in §25.397;(c) Trim tab effects, in §25.407;(d) Unsymmetrical loads, in §25.427; and(e) Auxiliary aerodynamic surfaces, in §25.445.


§25.393 Loads parallel to hinge line.(a) Control surfaces and supporting hinge

brackets must be designed for inertia loads actingparallel to the hinge line.

(b) In the absence of more rational data, the in-ertia loads may be assumed to be equal to KW,where—

(1) K = 24 for vertical surfaces;(2) K = 12 for horizontal surfaces; and(3) W = weight of the movable surfaces.

§25.395 Control system.(a) Longitudinal, lateral, directional, and drag

control system and their supporting structuresmust be designed for loads corresponding to 125percent of the computed hinge moments of themovable control surface in the conditions pre-scribed in §25.391.

(b) The system limit loads, except the loads re-sulting from ground gusts, need not exceed theloads that can be produced by the pilot (or pilots)and by automatic or power devices operating thecontrols.

(c) The loads must not be less than those re-sulting from application of the minimum forcesprescribed in §25.397(c).


§25.397 Control system loads.(a) General. The maximum and minimum pilot

forces, specified in paragraph (c) of this section,are assumed to act at the appropriate controlgrips or pads (in a manner simulating flight condi-tions) and to be reacted at the attachment of thecontrol system to the control surface horn.

(b) Pilot effort effects. In the control surfaceflight loading condition, the air loads on movablesurfaces and the corresponding deflections neednot exceed those that would result in flight fromthe application of any pilot force within the rangesspecified in paragraph (c) of this section. Two-thirds of the maximum values specified for the ai-leron and elevator may be used if control surfacehinge moments are based on reliable data. In ap-plying this criterion, the effects of servo mecha-nisms, tabs, and automatic pilot systems, must beconsidered.

(c) Limit pilot forces and torques. The limit pilotforces and torques are as follows:

1The critical parts of the aileron control systemmust be designed for a single tangential force with

ControlMaximum forces or torques

Minimum forces or torques

Aileron:StickWheel1

100 lbs.80 D in.-lbs2

40 lbs.40 D in.-lbs.

Elevator:StickWheel (symmetrical)Wheel (unsymmetrical)3

250 lbs.300 lbs.—

100 lbs.100 lbs.100 lbs.

Rudder 300 lbs. 130 lbs.


44 ASA

a limit value equal to 1.25 times the couple forcedetermined from these criteria.

2 D = wheel diameter (inches).3 The unsymmetrical forces must be applied at

one of the normal handgrip points on the periph-ery of the control wheel.

[Docket 5066, 29 FR 18291, Dec. 24, 1964; as amendedby Amdt. 25–38, 41 FR 55466, Dec. 20, 1976; Amdt. 25–72, 55 FR 29776, July 20, 1990]

§25.399 Dual control system.(a) Each dual control system must be designed

for the pilots operating in opposition, using individ-ual pilot forces not less than—

(1) 0.75 times those obtained under §25.395; or(2) The minimum forces specified in §25.397(c).(b) The control system must be designed for

pilot forces applied in the same direction, using in-dividual pilot forces not less than 0.75 times thoseobtained under §25.395.

§25.405 Secondary control system.Secondary controls, such as wheel brake,

spoiler, and tab controls, must be designed for themaximum forces that a pilot is likely to apply tothose controls. The following values may be used:

Pilot Control Force Limits (Secondary Controls)

*Limited to flap, tab, stabilizer, spoiler, andlanding gear operation controls.

§25.407 Trim tab effects.The effects of trim tabs on the control surface

design conditions must be accounted for onlywhere the surface loads are limited by maximumpilot effort. In these cases, the tabs are consid-ered to be deflected in the direction that would as-sist the pilot, and the deflections are—

(a) For elevator trim tabs, those required to trimthe airplane at any point within the positive por-tion of the pertinent flight envelope in §25.333(b),except as limited by the stops; and

(b) For aileron and rudder trim tabs, those re-quired to trim the airplane in the critical unsym-metrical power and loading conditions, with ap-propriate allowance for rigging tolerances.

§25.409 Tabs.(a) Trim tabs. Trim tabs must be designed to

withstand loads arising from all likely combina-tions of tab setting, primary control position, andairplane speed (obtainable without exceeding theflight load conditions prescribed for the airplaneas a whole), when the effect of the tab is opposedby pilot effort forces up to those specified in§25.397(b).

(b) Balancing tabs. Balancing tabs must be de-signed for deflections consistent with the primarycontrol surface loading conditions.

(c) Servo tabs. Servo tabs must be designedfor deflections consistent with the primary controlsurface loading conditions obtainable within thepilot maneuvering effort, considering possible op-position from the trim tabs.

§25.415 Ground gust conditions.(a) The control system must be designed as fol-

lows for control surface loads due to ground gustsand taxiing downwind:

(1) The control system between the stops near-est the surfaces and the cockpit controls must bedesigned for loads corresponding to the limithinge moments H of paragraph (a)(2) of this sec-tion. These loads need not exceed—

(i) The loads corresponding to the maximumpilot loads in §25.397(c) for each pilot alone; or

(ii) 0.75 times these maximum loads for eachpilot when the pilot forces are applied in the samedirection.

(2) The control system stops nearest the sur-faces, the control system locks, and the parts ofthe systems (if any) between these stops andlocks and the control surface horns, must be de-signed for limit hinge moments H, in foot pounds,obtained from the formula, H = .0034KV2cS,where —

V = 65 (wind speed in knots)K = limit hinge moment factor for ground gusts

derived in paragraph (b) of this section.c = mean chord of the control surface aft of the

hinge line (ft);S = area of the control surface aft of the hinge line

(sq ft).

Control Limit pilot forces

Miscellaneous:* Crank, wheel or lever

but not less than 50 lbs. nor more than 150 lbs. (R = radius). (Applicable to any angle within 20° of plane of control).

Twist 133 in.-lbs.

Push-pull To be chosen by applicant

1 R+3

----------- 50 lbs.×


ASA 45

25

(b) The limit hinge moment factor K for groundgusts must be derived as follows:

1 A positive value of K indicates a moment tend-ing to depress the surface, while a negative valueof K indicates a moment tending to raise the sur-face.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–72, 55 FR 29776, July 20, 1990;Amdt. 25–91, 62 FR 40705, July 29, 1997]

§25.427 Unsymmetrical loads.(a) In designing the airplane for lateral gust,

yaw maneuver and roll maneuver conditions, ac-count must be taken of unsymmetrical loads onthe empennage arising from effects such as slip-stream and aerodynamic interference with thewing, vertical fin and other aerodynamic surfaces.

(b) The horizontal tail must be assumed to besubjected to unsymmetrical loading conditionsdetermined as follows:

(1) 100 percent of the maximum loading fromthe symmetrical maneuver conditions of §25.331and the vertical gust conditions of §25.341(a) act-ing separately on the surface on one side of theplane of symmetry; and

(2) 80 percent of these loadings acting on theother side.

(c) For empennage arrangements where thehorizontal tail surfaces have dihedral anglesgreater than plus or minus 10 degrees, or aresupported by the vertical tail surfaces, the sur-faces and the supporting structure must be de-signed for gust velocities specified in §25.341(a)acting in any orientation at right angles to theflight path.

(d) Unsymmetrical loading on the empennagearising from buffet conditions of §25.305(e) mustbe taken into account.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5673, April 8, 1970Amdt. 25–86, 61 FR 5222, Feb. 9, 1996]

§25.445 Auxiliary aerodynamic surfaces.(a) When significant, the aerodynamic influ-

ence between auxiliary aerodynamic surfaces,such as outboard fins and winglets, and their sup-porting aerodynamic surfaces, must be taken into

account for all loading conditions including pitch,roll, and yaw maneuvers, and gusts as specifiedin §25.341(a) acting at any orientation at right an-gles to the flight path.

(b) To provide for unsymmetrical loading whenoutboard fins extend above and below the hori-zontal surface, the critical vertical surface loading(load per unit area) determined under §25.391must also be applied as follows:

(1) 100 percent to the area of the vertical sur-faces above (or below) the horizontal surface.

(2) 80 percent to the area below (or above) thehorizontal surface.


§25.457 Wing flaps.Wing flaps, their operating mechanisms, and

their supporting structures must be designed forcritical loads occurring in the conditions pre-scribed in §25.345, accounting for the loads oc-curring during transition from one flap positionand airspeed to another.

§25.459 Special devices.The loading for special devices using aerody-

namic surfaces (such as slots, slats and spoilers)must be determined from test data.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–72, 55 FR 29776, July 20, 1990]

GROUND LOADS

§25.471 General.(a) Loads and equilibrium. For limit ground

loads—(1) Limit ground loads obtained under this sub-

part are considered to be external forces appliedto the airplane structure; and

(2) In each specified ground load condition, theexternal loads must be placed in equilibrium withthe linear and angular inertia loads in a rational orconservative manner.

(b) Critical centers of gravity. The critical cen-ters of gravity within the range for which certifica-tion is requested must be selected so that themaximum design loads are obtained in each land-ing gear element. Fore and aft, vertical, and lat-eral airplane centers of gravity must be consid-ered. Lateral displacements of the c.g. from theairplane centerline which would result in maingear loads not greater than 103 percent of thecritical design load for symmetrical loading condi-tions may be selected without considering the ef-fects of these lateral c.g. displacements on theloading of the main gear elements, or on the air-plane structure provided—

(1) The lateral displacement of the c.g. resultsfrom random passenger or cargo disposition

Surface K Position of controls

(a) Aileron 0.75 Control column locked or lashed in mid-position

(b) do 11± 0.50 Ailerons at full throw

(c) Elevator 11± 0.75 (c) Elevator full down

(d) do 11± 0.75 (d) Elevator full up

(e) Rudder 0.75 (e) Rudder in neutral

(f) do 0.75 (f) Rudder at full throw


46 ASA

within the fuselage or from random unsymmetri-cal fuel loading or fuel usage; and

(2) Appropriate loading instructions for randomdisposable loads are included under the provi-sions of §25.1583(c)(1) to ensure that the lateraldisplacement of the center of gravity is main-tained within these limits.

(c) Landing gear dimension data. Figure 1 ofAppendix A contains the basic landing gear di-mension data.


§25.473 Landing load conditions and assumptions.(a) For the landing conditions specified in

§25.479 to §25.485 the airplane is assumed tocontact the ground—

(1) In the attitudes defined in §25.479 and§25.481;

(2) With a limit descent velocity of 10 fps at thedesign landing weight (the maximum weight forlanding conditions at the maximum descent veloc-ity); and

(3) With a limit descent velocity of 6 fps at thedesign takeoff weight (the maximum weight forlanding conditions at a reduced descent velocity).

(4) The prescribed descent velocities may bemodified if it is shown that the airplane has designfeatures that make it impossible to develop thesevelocities.

(b) Airplane lift, not exceeding airplane weight,may be assumed unless the presence of systemsor procedures significantly affects the lift.

(c) The method of analysis of airplane andlanding gear loads must take into account at leastthe following elements:

(1) Landing gear dynamic characteristics.(2) Spin-up and springback.(3) Rigid body response.(4) Structural dynamic response of the air-

frame, if significant.(d) The landing gear dynamic characteristics

must be validated by tests as defined in§25.723(a).

(e) The coefficient of friction between the tiresand the ground may be established by consider-ing the effects of skidding velocity and tire pres-sure. However, this coefficient of friction need notbe more than 0.8.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5673, April 8, 1970;Amdt. 25–91, 62 FR 40705, July 29, 1997; Amdt. 25–103, 66 FR 27394, May 16, 2001]

§25.477 Landing gear arrangement.Sections 25.479 through 25.485 apply to air-

planes with conventional arrangements of main

and nose gears, or main and tail gears, when nor-mal operating techniques are used.

§25.479 Level landing conditions.(a) In the level attitude, the airplane is assumed

to contact the ground at forward velocity compo-nents, ranging from VL1 to 1.25 VL2 parallel to theground under the conditions prescribed in§25.473 with—

(1) VL1 equal to VS0 (TAS) at the appropriatelanding weight and in standard sea level condi-tions; and

(2) VL2 equal to VS0 (TAS) at the appropriatelanding weight and altitudes in a hot day tempera-ture of 41 degrees F above standard.

(3) The effects of increased contact speedmust be investigated if approval of downwindlandings exceeding 10 knots is requested.

(b) For the level landing attitude for airplaneswith tail wheels, the conditions specified in thissection must be investigated with the airplanehorizontal reference line horizontal in accordancewith Figure 2 of Appendix A of this part.

(c) For the level landing attitude for airplaneswith nose wheels, shown in Figure 2 of AppendixA of this part, the conditions specified in this sec-tion must be investigated assuming the followingattitudes:

(1) An attitude in which the main wheels are as-sumed to contact the ground with the nose wheeljust clear of the ground; and

(2) If reasonably attainable at the specified de-scent and forward velocities, an attitude in whichthe nose and main wheels are assumed to con-tact the ground simultaneously.

(d) In addition to the loading conditions pre-scribed in paragraph (a) of this section, but withmaximum vertical ground reactions calculatedfrom paragraph (a), the following apply:

(1) The landing gear and directly affected at-taching structure must be designed for the maxi-mum vertical ground reaction combined with anaft acting drag component of not less than 25% ofthis maximum vertical ground reaction.

(2) The most severe combination of loads thatare likely to arise during a lateral drift landingmust be taken into account. In absence of a morerational analysis of this condition, the followingmust be investigated:

(i) A vertical load equal to 75% of the maximumground reaction of §25.473 must be considered incombination with a drag and side load of 40% and25% respectively of that vertical load.

(ii) The shock absorber and tire deflectionsmust be assumed to be 75% of the deflection cor-responding to the maximum ground reaction of§25.473(a)(2). This load case need not be consid-ered in combination with flat tires.


ASA 47

25

(3) The combination of vertical and drag com-ponents is considered to be acting at the wheelaxle centerline.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5673, April 8, 1970;Amdt. 25–91, 62 FR 40705, July 29, 1997; Amdt. 25–91,62 FR 45481, Aug. 27, 1997]

§25.481 Tail-down landing conditions.(a) In the tail-down attitude, the airplane is as-

sumed to contact the ground at forward velocitycomponents, ranging from VL1 to VL2 parallel tothe ground under the conditions prescribed in§25.473 with—

(1) VL1 equal to VS0 (TAS) at the appropriatelanding weight and in standard sea level condi-tions; and

(2) VL2 equal to VS0 (TAS) at the appropriatelanding weight and altitudes in a hot day tempera-ture of 41 degrees F above standard.

(3) The combination of vertical and drag com-ponents considered to be acting at the mainwheel axle centerline.

(b) For the tail-down landing condition for air-planes with tail wheels, the main and tail wheelsare assumed to contact the ground simulta-neously, in accordance with figure 3 of AppendixA. Ground reaction conditions on the tail wheelare assumed to act—

(1) Vertically; and(2) Up and aft through the axle at 45 degrees to

the ground line.(c) For the tail-down landing condition for air-

planes with nose wheels, the airplane is assumedto be at an attitude corresponding to either thestalling angle or the maximum angle allowingclearance with the ground by each part of the air-plane other than the main wheels, in accordancewith figure 3 of Appendix A, whichever is less.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–91, 62 FR 40705, July 29, 1997;Amdt. 25–94, 63 FR 8848, Feb. 23, 1998]

§25.483 One-gear landing conditions.For the one-gear landing conditions, the air-

plane is assumed to be in the level attitude and tocontact the ground on one main landing gear, inaccordance with Figure 4 of Appendix A of thispart. In this attitude—

(a) The ground reactions must be the same asthose obtained on that side under §25.479(d)(1);and

(b) Each unbalanced external load must be re-acted by airplane inertia in a rational or conserva-tive manner.


§25.485 Side load conditions.In addition to §25.479(d)(2) the following condi-

tions must be considered:(a) For the side load condition, the airplane is

assumed to be in the level attitude with only themain wheels contacting the ground, in accor-dance with figure 5 of Appendix A.

(b) Side loads of 0.8 of the vertical reaction (onone side) acting inward and 0.6 of the vertical re-action (on the other side) acting outward must becombined with one-half of the maximum verticalground reactions obtained in the level landingconditions. These loads are assumed to be ap-plied at the ground contact point and to be re-sisted by the inertia of the airplane. The dragloads may be assumed to be zero.


§25.487 Rebound landing condition.(a) The landing gear and its supporting struc-

ture must be investigated for the loads occurringduring rebound of the airplane from the landingsurface.

(b) With the landing gear fully extended and notin contact with the ground, a load factor of 20.0must act on the unsprung weights of the landinggear. This load factor must act in the direction ofmotion of the unsprung weights as they reachtheir limiting positions in extending with relation tothe sprung parts of the landing gear.

§25.489 Ground handling conditions.Unless otherwise prescribed, the landing gear

and airplane structure must be investigated forthe conditions in §§25.491 through 25.509 withthe airplane at the design ramp weight (the maxi-mum weight for ground handling conditions). Nowing lift may be considered. The shock absorbersand tires may be assumed to be in their static po-sition.


§25.491 Taxi, takeoff and landing roll.Within the range of appropriate ground speeds

and approved weights, the airplane structure andlanding gear are assumed to be subjected toloads not less than those obtained when the air-craft is operating over the roughest ground thatmay reasonably be expected in normal operation.


§25.493 Braked roll conditions.(a) An airplane with a tail wheel is assumed to

be in the level attitude with the load on the mainwheels, in accordance with figure 6 of Appendix


48 ASA

A. The limit vertical load factor is 1.2 at the designlanding weight and 1.0 at the design ramp weight.A drag reaction equal to the vertical reaction mul-tiplied by a coefficient of friction of 0.8, must becombined with the vertical ground reaction andapplied at the ground contact point.

(b) For an airplane with a nose wheel the limitvertical load factor is 1.2 at the design landingweight, and 1.0 at the design ramp weight. A dragreaction equal to the vertical reaction, multipliedby a coefficient of friction of 0.8, must be com-bined with the vertical reaction and applied at theground contact point of each wheel with brakes.The following two attitudes, in accordance withfigure 6 of Appendix A, must be considered:

(1) The level attitude with the wheels contactingthe ground and the loads distributed between themain and nose gear. Zero pitching acceleration isassumed.

(2) The level attitude with only the main gearcontacting the ground and with the pitching mo-ment resisted by angular acceleration.

(c) A drag reaction lower than that prescribedin this section may be used if it is substantiatedthat an effective drag force of 0.8 times the verti-cal reaction cannot be attained under any likelyloading condition.

(d) An airplane equipped with a nose gearmust be designed to withstand the loads arisingfrom the dynamic pitching motion of the airplanedue to sudden application of maximum brakingforce. The airplane is considered to be at designtakeoff weight with the nose and main gears incontact with the ground, and with a steady-statevertical load factor of 1.0. The steady-state nosegear reaction must be combined with the maxi-mum incremental nose gear vertical reactioncaused by the sudden application of maximumbraking force as described in paragraphs (b) and(c) of this section.

(e) In the absence of a more rational analysis,the nose gear vertical reaction prescribed in para-graph (d) of this section must be calculated ac-cording to the following formula:

Where:

VN = Nose gear vertical reaction.WT = Design takeoff weight.A = Horizontal distance between the c.g. of the

airplane and the nose wheel.B = Horizontal distance between the c.g. of the

airplane and the line joining the centers of the main wheels.

E = Vertical height of the c.g. of the airplane above the ground in the 1.0 g static condition.

µ = Coefficient of friction of 0.80.f = Dynamic response factor; 2.0 is to be used

unless a lower factor is substantiated. In the absence of other information, the dynamic response factor f may be defined by the equation:

Where:

is the effective critical damping ratio of the rigid body pitching mode about the main landing gear effective ground contact point.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5673, April 8, 1970;Amdt. 25–97, 63 FR 29072, May 27, 1998]

§25.495 Turning.In the static position, in accordance with figure

7 of Appendix A, the airplane is assumed to exe-cute a steady turn by nose gear steering, or byapplication of sufficient differential power, so thatthe limit load factors applied at the center of grav-ity are 1.0 vertically and 0.5 laterally. The sideground reaction of each wheel must be 0.5 of thevertical reaction.

§25.497 Tail-wheel yawing.(a) A vertical ground reaction equal to the static

load on the tail wheel, in combination with a sidecomponent of equal magnitude, is assumed.

(b) If there is a swivel, the tail wheel is as-sumed to be swiveled 90° to the airplane longitu-dinal axis with the resultant load passing throughthe axle.

(c) If there is a lock, steering device, or shimmydamper the tail wheel is also assumed to be in thetrailing position with the side load acting at theground contact point.

§25.499 Nose-wheel yaw and steering.(a) A vertical load factor of 1.0 at the airplane

center of gravity, and a side component at thenose wheel ground contact equal to 0.8 of the ver-tical ground reaction at that point are assumed.

(b) With the airplane assumed to be in staticequilibrium with the loads resulting from the useof brakes on one side of the main landing gear,the nose gear, its attaching structure, and the fu-selage structure forward of the center of gravitymust be designed for the following loads:

(1) A vertical load factor at the center of gravityof 1.0.

VN

WT

A B+-------------- B

fµAEA B µE+ +----------------------------+=

f 1 expπ ξ–

1 ξ2–

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

+=

ξ


ASA 49

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(2) A forward acting load at the airplane centerof gravity of 0.8 times the vertical load on onemain gear.

(3) Side and vertical loads at the ground con-tact point on the nose gear that are required forstatic equilibrium.

(4) A side load factor at the airplane center ofgravity of zero.

(c) If the loads prescribed in paragraph (b) ofthis section result in a nose gear side load higherthan 0.8 times the vertical nose gear load, the de-sign nose gear side load may be limited to 0.8times the vertical load, with unbalanced yawingmoments assumed to be resisted by airplane in-ertia forces.

(d) For other than the nose gear, its attachingstructure, and the forward fuselage structure, theloading conditions are those prescribed in para-graph (b) of this section, except that—

(1) A lower drag reaction may be used if an ef-fective drag force of 0.8 times the vertical reactioncannot be reached under any likely loading condi-tion; and

(2) The forward acting load at the center ofgravity need not exceed the maximum drag reac-tion on one main gear, determined in accordancewith §25.493(b).

(e) With the airplane at design ramp weight, andthe nose gear in any steerable position, the com-bined application of full normal steering torque andvertical force equal to 1.33 times the maximumstatic reaction on the nose gear must be consid-ered in designing the nose gear, its attachingstructure, and the forward fuselage structure.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5673, April 8, 1970;Amdt. 25–46, 43 FR 50595, Oct. 30, 1978; Amdt. 25–91,62 FR 40705, July 29, 1997]

§25.503 Pivoting.(a) The airplane is assumed to pivot about one

side of the main gear with the brakes on that sidelocked. The limit vertical load factor must be 1.0and the coefficient of friction 0.8.

(b) The airplane is assumed to be in static equi-librium, with the loads being applied at the groundcontact points, in accordance with figure 8 of Ap-pendix A.

§25.507 Reversed braking.(a) The airplane must be in a three point static

ground attitude. Horizontal reactions parallel tothe ground and directed forward must be appliedat the ground contact point of each wheel withbrakes. The limit loads must be equal to 0.55times the vertical load at each wheel or to the loaddeveloped by 1.2 times the nominal maximumstatic brake torque, whichever is less.

(b) For airplanes with nose wheels, the pitchingmoment must be balanced by rotational inertia.

(c) For airplanes with tail wheels, the resultantof the ground reactions must pass through thecenter of gravity of the airplane.

§25.509 Towing loads.(a) The towing loads specified in paragraph (d)

of this section must be considered separately.These loads must be applied at the towing fittingsand must act parallel to the ground. In addition—

(1) A vertical load factor equal to 1.0 must beconsidered acting at the center of gravity;

(2) The shock struts and tires must be in theirstatic positions; and

(3) With WT as the design ramp weight, thetowing load, FTOW, is—

(i) 0.3 WT for WT less than 30,000 pounds;(ii) (6 WT + 450,000)/7 for WT between 30,000

and 100,000 pounds; and(iii) 0.15 WT for WT over 100,000 pounds.(b) For towing points not on the landing gear

but near the plane of symmetry of the airplane,the drag and side tow load components specifiedfor the auxiliary gear apply. For towing points lo-cated outboard of the main gear, the drag andside tow load components specified for the maingear apply. Where the specified angle of swivelcannot be reached, the maximum obtainable an-gle must be used.

(c) The towing loads specified in paragraph (d)of this section must be reacted as follows:

(1) The side component of the towing load atthe main gear must be reacted by a side force atthe static ground line of the wheel to which theload is applied.

(2) The towing loads at the auxiliary gear andthe drag components of the towing loads at themain gear must be reacted as follows:

(i) A reaction with a maximum value equal tothe vertical reaction must be applied at the axle ofthe wheel to which the load is applied. Enoughairplane inertia to achieve equilibrium must be ap-plied.

(ii) The loads must be reacted by airplane inertia.


50 ASA

(d) The prescribed towing loads are as follows:

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; as amended by Amdt. 25–23, 35 FR 5673, April 8, 1970]

§25.511 Ground load: unsymmetrical loads on multiple-wheel units.(a) General. Multiple-wheel landing gear units

are assumed to be subjected to the limit groundloads prescribed in this subpart under paragraphs(b) through (f) of this section. In addition—

(1) A tandem strut gear arrangement is a multi-ple-wheel unit; and

(2) In determining the total load on a gear unitwith respect to the provisions of paragraphs (b)through (f) of this section, the transverse shift inthe load centroid, due to unsymmetrical load dis-tribution on the wheels, may be neglected.

(b) Distribution of limit loads to wheels; tires in-flated. The distribution of the limit loads amongthe wheels of the landing gear must be estab-lished for each landing, taxiing, and ground han-dling condition, taking into account the effects ofthe following factors:

(1) The number of wheels and their physical ar-rangements. For truck type landing gear units, theeffects of any seesaw motion of the truck duringthe landing impact must be considered in deter-mining the maximum design loads for the fore andaft wheel pairs.

(2) Any differentials in tire diameters resultingfrom a combination of manufacturing tolerances,tire growth, and tire wear. A maximum tire-diame-ter differential equal to 2/3 of the most unfavorablecombination of diameter variations that is ob-tained when taking into account manufacturingtolerances, tire growth, and tire wear, may be as-sumed.

(3) Any unequal tire inflation pressure, assum-ing the maximum variation to be ±5 percent of thenominal tire inflation pressure.

(4) A runway crown of zero and a runway crownhaving a convex upward shape that may be ap-proximated by a slope of 11⁄2 percent with the hor-

izontal. Runway crown effects must be consid-ered with the nose gear unit on either slope of thecrown.

(5) The airplane attitude.(6) Any structural deflections.(c) Deflated tires. The effect of deflated tires on

the structure must be considered with respect tothe loading conditions specified in paragraphs (d)through (f) of this section, taking into account thephysical arrangement of the gear components. Inaddition—

(1) The deflation of any one tire for each multi-ple wheel landing gear unit, and the deflation ofany two critical tires for each landing gear unit us-ing four or more wheels per unit, must be consid-ered; and

(2) The ground reactions must be applied to thewheels with inflated tires except that, for multiple-wheel gear units with more than one shock strut,a rational distribution of the ground reactions be-tween the deflated and inflated tires, accountingfor the differences in shock strut extensions re-sulting from a deflated tire, may be used.

(d) Landing conditions. For one and for two de-flated tires, the applied load to each gear unit isassumed to be 60 percent and 50 percent, respec-tively, of the limit load applied to each gear foreach of the prescribed landing conditions. How-ever, for the drift landing condition of §25.485, 100percent of the vertical load must be applied.

(e) Taxiing and ground handling conditions. Forone and for two deflated tires—

(1) The applied side or drag load factor, or bothfactors, at the center of gravity must be the mostcritical value up to 50 percent and 40 percent, re-spectively, of the limit side or drag load factors, orboth factors, corresponding to the most severecondition resulting from consideration of the pre-scribed taxiing and ground handling conditions;

Tow point PositionLoad

Magnitude No. Direction

Main gear 0.75 FTow per main gear unit

1234

Forward, parallel to drag axisForward, at 30° to drag axisAft, parallel to drag axisAft, at 30° to drag axis

Auxiliary gear Swiveled forward

Swiveled aft

Swiveled 45° from forward

Swiveled 45° from aft

1.0 FTow

do

0.5 FTow

do

56789

101112

ForwardAftForwardAftForward, in plane of wheelAft, in plane of wheelForward, in plane of wheelAft, in plane of wheel


ASA 51

25

(2) For the braked roll conditions of §25.493 (a)and (b)(2), the drag loads on each inflated tire maynot be less than those at each tire for the symmet-rical load distribution with no deflated tires;

(3) The vertical load factor at the center of grav-ity must be 60 percent and 50 percent, respec-tively, of the factor with no deflated tires, exceptthat it may not be less than 1g; and

(4) Pivoting need not be considered.(f) Towing conditions. For one and for two de-

flated tires, the towing load, FTOW, must be 60percent and 50 percent, respectively, of the loadprescribed.

§25.519 Jacking and tie-down provisions.(a) General. The airplane must be designed to

withstand the limit load conditions resulting fromthe static ground load conditions of paragraph (b)of this section and, if applicable, paragraph (c) ofthis section at the most critical combinations ofairplane weight and center of gravity. The maxi-mum allowable load at each jack pad must bespecified.

(b) Jacking. The airplane must have provisionsfor jacking and must withstand the following limitloads when the airplane is supported on jacks—

(1) For jacking by the landing gear at the maxi-mum ramp weight of the airplane, the airplanestructure must be designed for a vertical load of1.33 times the vertical static reaction at each jack-ing point acting singly and in combination with ahorizontal load of 0.33 times the vertical static re-action applied in any direction.

(2) For jacking by other airplane structure atmaximum approved jacking weight:

(i) The airplane structure must be designed fora vertical load of 1.33 times the vertical reactionat each jacking point acting singly and in combi-nation with a horizontal load of 0.33 times the ver-tical static reaction applied in any direction.

(ii) The jacking pads and local structure mustbe designed for a vertical load of 2.0 times thevertical static reaction at each jacking point, act-ing singly and in combination with a horizontalload of 0.33 times the vertical static reaction ap-plied in any direction.

(c) Tie-down. If tie-down points are provided,the main tie-down points and local structure mustwithstand the limit loads resulting from a 65-knothorizontal wind from any direction.

[Docket No. 26129, 59 FR 22102, April 28, 1994]

WATER LOADS

§25.521 General.(a) Seaplanes must be designed for the water

loads developed during takeoff and landing, withthe seaplane in any attitude likely to occur in nor-mal operation, and at the appropriate forward and

sinking velocities under the most severe sea con-ditions likely to be encountered.

(b) Unless a more rational analysis of the waterloads is made, or the standards in ANC-3 areused, §§25.523 through 25.537 apply.

(c) The requirements of this section and§§25.523 through 25.537 apply also to amphibians.

§25.523 Design weights and center of gravity positions.(a) Design weights. The water load require-

ments must be met at each operating weight up tothe design landing weight except that, for thetakeoff condition prescribed in §25.531, the de-sign water takeoff weight (the maximum weight forwater taxi and takeoff run) must be used.

(b) Center of gravity positions. The critical cen-ters of gravity within the limits for which certifica-tion is requested must be considered to reachmaximum design loads for each part of the sea-plane structure.


§25.525 Application of loads.(a) Unless otherwise prescribed, the seaplane

as a whole is assumed to be subjected to theloads corresponding to the load factors specifiedin §25.527.

(b) In applying the loads resulting from the loadfactors prescribed in §25.527, the loads may bedistributed over the hull or main float bottom (inorder to avoid excessive local shear loads andbending moments at the location of water load ap-plication) using pressures not less than those pre-scribed in §25.533(b).

(c) For twin float seaplanes, each float must betreated as an equivalent hull on a fictitious sea-plane with a weight equal to one-half the weight ofthe twin float seaplane.

(d) Except in the takeoff condition of §25.531,the aerodynamic lift on the seaplane during theimpact is assumed to be 2⁄3 of the weight of theseaplane.

§25.527 Hull and main float load factors.(a) Water reaction load factors nW must be

computed in the following manner:(1) For the step landing case

nwC1VS0 2

Tan2 3⁄ β

W1 3⁄

----------------------------------------=


52 ASA

(2) For the bow and stern landing cases

(b) The following values are used:(1) nW = water reaction load factor (that is, the

water reaction divided by seaplane weight).(2) C1 = empirical seaplane operations factor

equal to 0.012 (except that this factor may not beless than that necessary to obtain the minimumvalue of step load factor of 2.33).

(3) VS0 = seaplane stalling speed in knots withflaps extended in the appropriate landing positionand with no slipstream effect.

(4) β = angle of dead rise at the longitudinalstation at which the load factor is being deter-mined in accordance with figure 1 of Appendix B.

(5) W = seaplane design landing weight inpounds.

(6) K1 = empirical hull station weighing factor,in accordance with figure 2 of Appendix B.

(7) rx = ratio of distance, measured parallel tohull reference axis, from the center of gravity ofthe seaplane to the hull longitudinal station atwhich the load factor is being computed to the ra-dius of gyration in pitch of the seaplane, the hullreference axis being a straight line, in the planeof symmetry, tangential to the keel at the mainstep.

(c) For a twin float seaplane, because of theeffect of flexibility of the attachment of the floatsto the seaplane, the factor K1 may be reduced atthe bow and stern to 0.8 of the value shown in fig-ure 2 of Appendix B. This reduction applies onlyto the design of the carrythrough and seaplanestructure.


§25.529 Hull and main float landing conditions.(a) Symmetrical step, bow, and stern landing.

For symmetrical step, bow, and stern landings,the limit water reaction load factors are thosecomputed under §25.527. In addition—

(1) For symmetrical step landings, the resultantwater load must be applied at the keel, throughthe center of gravity, and must be directed per-pendicularly to the keel line;

(2) For symmetrical bow landings, the resultantwater load must be applied at the keel, one-fifth ofthe longitudinal distance from the bow to the step,and must be directed perpendicularly to the keelline; and

(3) For symmetrical stern landings, the result-ant water load must be applied at the keel, at apoint 85 percent of the longitudinal distance from

the step to the stern post, and must be directedperpendicularly to the keel line.

(b) Unsymmetrical landing for hull and singlefloat seaplanes. Unsymmetrical step, bow, andstern landing conditions must be investigated. Inaddition—

(1) The loading for each condition consists ofan upward component and a side componentequal, respectively, to 0.75 and 0.25 tan β timesthe resultant load in the corresponding symmetri-cal landing condition; and

(2) The point of application and direction of theupward component of the load is the same as thatin the symmetrical condition, and the point of ap-plication of the side component is at the samelongitudinal station as the upward component butis directed inward perpendicularly to the plane ofsymmetry at a point midway between the keel andchine lines.

(c) Unsymmetrical landing; twin float sea-planes. The unsymmetrical loading consists of anupward load at the step of each float of 0.75 and aside load of 0.25 tan β at one float times the steplanding load reached under §25.527. The sideload is directed inboard, perpendicularly to theplane of symmetry midway between the keel andchine lines of the float, at the same longitudinalstation as the upward load.

§25.531 Hull and main float takeoff condition.For the wing and its attachment to the hull or

main float—(a) The aerodynamic wing lift is assumed to be

zero; and(b) A downward inertia load, corresponding to

a load factor computed from the following formula,must be applied:

where—

n = inertia load factor;CTO = empirical seaplane operations factor equal

to 0.004;VS1 = seaplane stalling speed (knots) at the

design takeoff weight with the flaps extended in the appropriate takeoff position;

β = angle of dead rise at the main step (degrees); and

W = design water takeoff weight in pounds.


nw

C1VS02

Tan2 3⁄ β

W1 3⁄

----------------------------------------K1

1 rx2+

2 3⁄-----------------------------×=

nC TO VS1 2

tan2 3⁄ β

W1 3⁄

--------------------------------------=


ASA 53

25

§25.533 Hull and main float bottom pressures.(a) General. The hull and main float structure,

including frames and bulkheads, stringers, andbottom plating, must be designed under this sec-tion.

(b) Local pressures. For the design of the bot-tom plating and stringers and their attachments tothe supporting structure, the following pressuredistributions must be applied:

(1) For an unflared bottom, the pressure at thechine is 0.75 times the pressure at the keel, andthe pressures between the keel and chine vary lin-early, in accordance with figure 3 of Appendix B.The pressure at the keel (psi) is computed as fol-lows:

where—

Pk = pressure (p.s.i.) at the keel;C2 = 0.00213;K2 = hull station weighing factor, in accordance

with figure 2 of Appendix B;VS1 = seaplane stalling speed (Knots) at the de-

sign water takeoff weight with flaps extended in the appropriate takeoff position; and

βk = angle of dead rise at keel, in accordance with figure 1 of Appendix B.

(2) For a flared bottom, the pressure at the be-ginning of the flare is the same as that for an un-flared bottom, and the pressure between thechine and the beginning of the flare varies lin-early, in accordance with figure 3 of Appendix B.The pressure distribution is the same as that pre-scribed in paragraph (b)(1) of this section for anunflared bottom except that the pressure at thechine is computed as follows:

where—

Pch = pressure (p.s.i.) at the chine;C3 = 0.0016;K2 = hull station weighing factor, in accordance

with figure 2 of Appendix B;VS1 = seaplane stalling speed at the design water

takeoff weight with flaps extended in the appropriate takeoff position; and

β = angle of dead rise at appropriate station.

The area over which these pressures are appliedmust simulate pressures occurring during high lo-

calized impacts on the hull or float, but need notextend over an area that would induce criticalstresses in the frames or in the overall structure.

(c) Distributed pressures. For the design of theframes, keel, and chine structure, the followingpressure distributions apply:

(1) Symmetrical pressures are computed asfollows:

where—

P = pressure (p.s.i.);C4 = 0.078 C1 (with C1 computed under §25.527);K2 = hull station weighing factor, determined in

accordance with figure 2 of Appendix B;VS0 = seaplane stalling speed (Knots) with

landing flaps extended in the appropriate position and with no slipstream effect; and

VS0 = seaplane stalling speed with landing flaps extended in the appropriate position and with no slipstream effect; and

β = angle of dead rise at appropriate station.

(2) The unsymmetrical pressure distributionconsists of the pressures prescribed in paragraph(c)(1) of this section on one side of the hull ormain float centerline and one-half of that pressureon the other side of the hull or main float center-line, in accordance with figure 3 of Appendix B.

These pressures are uniform and must be appliedsimultaneously over the entire hull or main floatbottom. The loads obtained must be carried intothe sidewall structure of the hull proper, but neednot be transmitted in a fore and aft direction asshear and bending loads.


§25.535 Auxiliary float loads.(a) General. Auxiliary floats and their attach-

ments and supporting structures must be de-signed for the conditions prescribed in this sec-tion. In the cases specified in paragraphs (b)through (e) of this section, the prescribed waterloads may be distributed over the float bottom toavoid excessive local loads, using bottom pres-sures not less than those prescribed in paragraph(g) of this section.

(b) Step loading. The resultant water load mustbe applied in the plane of symmetry of the float ata point three-fourths of the distance from the bowto the step and must be perpendicular to the keel.The resultant limit load is computed as follows,except that the value of L need not exceed three

Pk C2

K 2 VS1 2

βktan--------------------×=

Pch C3

K 2 VS1 2

βtan--------------------×=

P C4

K 2 VS0 2

βtan---------------------×=


54 ASA

times the weight of the displaced water when thefloat is completely submerged:

where—

L = limit load (lbs.);C5 = 0.0053;VS0 = seaplane stalling speed (knots) with landing

flaps extended in the appropriate position and with no slipstream effect;

W = seaplane design landing weight in pounds;βS = angle of dead rise at a station 3⁄4 of the

distance from the bow to the step, but need not be less than 15 degrees; and

ry = ratio of the lateral distance between the center of gravity and the plane of symmetry of the float to the radius of gyration in roll.

(c) Bow loading. The resultant limit load mustbe applied in the plane of symmetry of the float ata point one-fourth of the distance from the bow tothe step and must be perpendicular to the tangentto the keel line at that point. The magnitude of theresultant load is that specified in paragraph (b) ofthis section.

(d) Unsymmetrical step loading. The resultantwater load consists of a component equal to 0.75times the load specified in paragraph (a) of thissection and a side component equal to 3.25 tan βtimes the load specified in paragraph (b) of thissection. The side load must be applied perpendic-ularly to the plane of symmetry of the float at apoint midway between the keel and the chine.

(e) Unsymmetrical bow loading. The resultantwater load consists of a component equal to 0.75times the load specified in paragraph (b) of thissection and a side component equal to 0.25 tan βtimes the load specified in paragraph (c) of thissection. The side load must be applied perpendic-ularly to the plane of symmetry at a point midwaybetween the keel and the chine.

(f) Immersed float condition. The resultant loadmust be applied at the centroid of the cross sec-tion of the float at a point one-third of the distancefrom the bow to the step. The limit load compo-nents are as follows:

where—

ρ = mass density of water (slugs/ft.2);V = volume of float (ft.2);Cx = coefficient of drag force, equal to 0.133;Cy = coefficient of side force, equal to 0.106;K = 0.8, except that lower values may be used if it

is shown that the floats are incapable of submerging at a speed of 0.8 VS0 in normal operations;

VS0 = seaplane stalling speed (knots) with landing flaps extended in the appropriate position and with no slipstream effect; and

g = acceleration due to gravity (ft./sec.2).

(g) Float bottom pressures. The float bottompressures must be established under §25.533,except that the value of K2 in the formulae may betaken as 1.0. The angle of dead rise to be used indetermining the float bottom pressures is set forthin paragraph (b) of this section.


§25.537 Seawing loads.Seawing design loads must be based on appli-

cable test data.

EMERGENCY LANDING CONDITIONS

§25.561 General.(a) The airplane, although it may be damaged

in emergency landing conditions on land or water,must be designed as prescribed in this section toprotect each occupant under those conditions.

(b) The structure must be designed to giveeach occupant every reasonable chance of es-caping serious injury in a minor crash landingwhen—

(1) Proper use is made of seats, belts, and allother safety design provisions;

(2) The wheels are retracted (where applica-ble); and

(3) The occupant experiences the following ulti-mate inertia forces acting separately relative tothe surrounding structure:

(i) Upward, 3.0g(ii) Forward, 9.0g(iii) Sideward, 3.0g on the airframe; and 4.0g

on the seats and their attachments.(iv) Downward, 6.0g(v) Rearward, 1.5g(c) For equipment, cargo in the passenger

compartments and any other large masses, thefollowing apply:

(1) Except as provided in paragraph (c)(2) ofthis section, these items must be positioned sothat if they break loose they will be unlikely to:

(i) Cause direct injury to occupants;

LC5VS0 2 W

2 3⁄

tan2 3⁄ βs

1 ry2+

2 3⁄-----------------------------------------------------=

vertical Vρg S0=

aft Cxρ2V

2 3⁄KV S0( )2

=

side Cyρ2V

2 3⁄KV S0( )2

=


ASA 55

25

(ii) Penetrate fuel tanks or lines or cause fire orexplosion hazard by damage to adjacent sys-tems; or

(iii) Nullify any of the escape facilities providedfor use after an emergency landing.

(2) When such positioning is not practical (e.g.fuselage mounted engines or auxiliary powerunits) each such item of mass shall be restrainedunder all loads up to those specified in paragraph(b)(3) of this section. The local attachments forthese items should be designed to withstand 1.33times the specified loads if these items are sub-ject to severe wear and tear through frequent re-moval (e.g. quick change interior items).

(d) Seats and items of mass (and their support-ing structure) must not deform under any loads upto those specified in paragraph (b)(3) of this sec-tion in any manner that would impede subsequentrapid evacuation of occupants.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5673, April 8, 1970;Amdt. 25–64, 53 FR 17646, May 17, 1988; Amdt. 25–91,62 FR 40706, July 29, 1997]

§25.562 Emergency landing dynamic conditions.(a) The seat and restraint system in the air-

plane must be designed as prescribed in this sec-tion to protect each occupant during an emer-gency landing condition when—

(1) Proper use is made of seats, safety belts,and shoulder harnesses provided for in the de-sign; and

(2) The occupant is exposed to loads resultingfrom the conditions prescribed in this section.

(b) Each seat type design approved for crew orpassenger occupancy during takeoff and landingmust successfully complete dynamic tests or bedemonstrated by rational analysis based on dy-namic tests of a similar type seat, in accordancewith each of the following emergency landing con-ditions. The tests must be conducted with an oc-cupant simulated by a 170-pound anthropomor-phic test dummy, as defined by 49 CFR Part 572,Subpart B, or its equivalent, sitting in the normalupright position.

(1) A change in downward vertical velocity (∆v)of not less than 35 feet per second, with the air-plane’s longitudinal axis canted downward 30 de-grees with respect to the horizontal plane andwith the wings level. Peak floor deceleration mustoccur in not more than 0.08 seconds after impactand must reach a minimum of 14g.

(2) A change in forward longitudinal velocity(∆v) of not less than 44 feet per second, with theairplane’s longitudinal axis horizontal and yawed10 degrees either right or left, whichever wouldcause the greatest likelihood of the upper torsorestraint system (where installed) moving off the

occupant’s shoulder, and with the wings level.Peak floor deceleration must occur in not morethan 0.09 seconds after impact and must reach aminimum of 16g. Where floor rails or floor fittingsare used to attach the seating devices to the testfixture, the rails or fittings must be misaligned withrespect to the adjacent set of rails or fittings by atleast 10 degrees vertically (i.e., out of Parallel)with one rolled 10 degrees.

(c) The following performance measures mustnot be exceeded during the dynamic tests con-ducted in accordance with paragraph (b) of thissection:

(1) Where upper torso straps are used for crew-members, tension loads in individual straps mustnot exceed 1,750 pounds. If dual straps are usedfor restraining the upper torso, the total strap ten-sion loads must not exceed 2,000 pounds.

(2) The maximum compressive load measuredbetween the pelvis and the lumbar column of theanthropomorphic dummy must not exceed 1,500pounds.

(3) The upper torso restraint straps (where in-stalled) must remain on the occupant’s shoulderduring the impact.

(4) The lap safety belt must remain on the oc-cupant’s pelvis during the impact.

(5) Each occupant must be protected from seri-ous head injury under the conditions prescribed inparagraph (b) of this section. Where head contactwith seats or other structure can occur, protectionmust be provided so that the head impact doesnot exceed a Head Injury Criterion (HIC) of 1,000units. The level of HIC is defined by the equation:

Where:

t1 is the initial integration time,t2 is the final integration time, anda(t) is the total acceleration vs. time curve for the

head strike, and where(t) is in seconds, and (a) is in units of gravity (g).

(6) Where leg injuries may result from contactwith seats or other structure, protection must beprovided to prevent axially compressive loads ex-ceeding 2,250 pounds in each femur.

(7) The seat must remain attached at all pointsof attachment, although the structure may haveyielded.

(8) Seats must not yield under the tests speci-fied in paragraphs (b)(1) and (b)(2) of this sectionto the extent they would impede rapid evacuationof the airplane occupants.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–64, 53 FR 17646, May 17, 1988]

HIC t2 t1–( ) 1t2 t1–( )

------------------- a t( ) tdt1

t2∫2.5

max

=


56 ASA

§25.563 Structural ditching provisions.Structural strength considerations of ditching

provisions must be in accordance with §25.801(e).

FATIGUE EVALUATION

§25.571 Damage-tolerance and fatigue evaluation of structure.(a) General. An evaluation of the strength, de-

tail design, and fabrication must show that cata-strophic failure due to fatigue, corrosion, manu-facturing defects, or accidental damage, will beavoided throughout the operational life of the air-plane. This evaluation must be conducted in ac-cordance with the provisions of paragraphs (b)and (e) of this section, except as specified inparagraph (c) of this section, for each part of thestructure that could contribute to a catastrophicfailure (such as wing, empennage, control sur-faces and their systems, the fuselage, enginemounting, landing gear, and their related primaryattachments). For turbojet powered airplanes,those parts that could contribute to a catastrophicfailure must also be evaluated under paragraph(d) of this section. In addition, the following apply:

(1) Each evaluation required by this sectionmust include—

(i) The typical loading spectra, temperatures,and humidities expected in service;

(ii) The identification of principal structural ele-ments and detail design points, the failure ofwhich could cause catastrophic failure of the air-plane; and

(iii) An analysis, supported by test evidence, ofthe principal structural elements and detail de-sign points identified in paragraph (a)(1)(ii) of thissection.

(2) The service history of airplanes of similarstructural design, taking due account of differ-ences in operating conditions and procedures,may be used in the evaluations required by thissection.

(3) Based on the evaluations required by thissection, inspections or other procedures must beestablished, as necessary, to prevent catastrophicfailure, and must be included in the AirworthinessLimitations Section of the Instructions for Contin-ued Airworthiness required by §25.1529. Inspec-tion thresholds for the following types of structuremust be established based on crack growth anal-yses and/or tests, assuming the structure con-tains an initial flaw of the maximum probable sizethat could exist as a result of manufacturing orservice-induced damage:

(i) Single load path structure, and(ii) Multiple load path “fail-safe” structure and

crack arrest “fail-safe” structure, where it cannotbe demonstrated that load path failure, partial fail-ure, or crack arrest will be detected and repairedduring normal maintenance, inspection, or opera-

tion of an airplane prior to failure of the remainingstructure.

(b) Damage-tolerance evaluation. The evalua-tion must include a determination of the probablelocations and modes of damage due to fatigue,corrosion, or accidental damage. Repeated loadand static analyses supported by test evidenceand (if available) service experience must also beincorporated in the evaluation. Special consider-ation for widespread fatigue damage must be in-cluded where the design is such that this type ofdamage could occur. It must be demonstratedwith sufficient full-scale fatigue test evidence thatwidespread fatigue damage will not occur withinthe design service goal of the airplane. The typecertificate may be issued prior to completion offull-scale fatigue testing, provided the Administra-tor has approved a plan for completing the re-quired tests, and the airworthiness limitationssection of the instructions for continued airworthi-ness required by §25.1529 of this part specifiesthat no airplane may be operated beyond a num-ber of cycles equal to 1⁄2 the number of cycles ac-cumulated on the fatigue test article, until suchtesting is completed. The extent of damage for re-sidual strength evaluation at any time within theoperational life of the airplane must be consistentwith the initial detectability and subsequentgrowth under repeated loads. The residualstrength evaluation must show that the remainingstructure is able to withstand loads (consideredas static ultimate loads) corresponding to the fol-lowing conditions:

(1) The limit symmetrical maneuvering condi-tions specified in §25.337 at all speeds up to VCand in §25.345.

(2) The limit gust conditions specified in§25.341 at the specified speeds up to VC and in§25.345.

(3) The limit rolling conditions specified in§25.349 and the limit unsymmetrical conditionsspecified in §§25.367 and 25.427 (a) through (c),at speeds up to VC.

(4) The limit yaw maneuvering conditionsspecified in §25.351(a) at the specified speedsup to VC.

(5) For pressurized cabins, the following condi-tions:

(i) The normal operating differential pressurecombined with the expected external aerody-namic pressures applied simultaneously with theflight loading conditions specified in paragraphs(b) (1) through (4) of this section, if they have asignificant effect.

(ii) The maximum value of normal operating dif-ferential pressure (including the expected externalaerodynamic pressures during 1 g level flight)multiplied by a factor of 1.15, omitting other loads.


ASA 57

25

(6) For landing gear and directly-affected air-frame structure, the limit ground loading condi-tions specified in §§25.473, 25.491, and 25.493.

If significant changes in structural stiffness or ge-ometry, or both, follow from a structural failure, orpartial failure, the effect on damage tolerancemust be further investigated.

(c) Fatigue (safe-life) evaluation. Compliancewith the damage-tolerance requirements of para-graph (b) of this section is not required if the appli-cant establishes that their application for particu-lar structure is impractical. This structure must beshown by analysis, supported by test evidence, tobe able to withstand the repeated loads of vari-able magnitude expected during its service lifewithout detectable cracks. Appropriate safe-lifescatter factors must be applied.

(d) Sonic fatigue strength. It must be shown byanalysis, supported by test evidence, or by theservice history of airplanes of similar structuraldesign and sonic excitation environment, that— (1) Sonic fatigue cracks are not probable in anypart of the flight structure subject to sonic excita-tion; or

(2) Catastrophic failure caused by sonic cracksis not probable assuming that the loads pre-scribed in paragraph (b) of this section are ap-plied to all areas affected by those cracks.

(e) Damage-tolerance (discrete source) evalu-ation. The airplane must be capable of success-fully completing a flight during which likely struc-tural damage occurs as a result of—

(1) Impact with a 4-pound bird when the veloc-ity of the airplane relative to the bird along the air-plane’s flight path is equal to VC at sea level or0.85VC at 8,000 feet, whichever is more critical;

(2) Uncontained fan blade impact;(3) Uncontained engine failure; or(4) Uncontained high energy rotating machin-

ery failure.

The damaged structure must be able to withstandthe static loads (considered as ultimate loads)which are reasonably expected to occur on theflight. Dynamic effects on these static loads neednot be considered. Corrective action to be taken bythe pilot following the incident, such as limiting ma-neuvers, avoiding turbulence, and reducing speed,must be considered. If significant changes in struc-tural stiffness or geometry, or both, follow from astructural failure or partial failure, the effect ondamage tolerance must be further investigated.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–45, 43 FR 46242, Oct. 5, 1978;Amdt. 25–54, 45 FR 60173, Sept. 11, 1980; Amdt. 25–72, 55 FR 29776, July 20, 1990; Amdt. 25–86, 61 FR5222, Feb. 9, 1996; Amdt. 25–96, 63 FR 15714, March31, 1998; Amdt. 25–92, 63 FR 23338, April 28, 1998]

LIGHTNING PROTECTION

§25.581 Lightning protection.(a) The airplane must be protected against cat-

astrophic effects from lightning.(b) For metallic components, compliance with

paragraph (a) of this section may be shown by—(1) Bonding the components properly to the air-

frame; or(2) Designing the components so that a strike

will not endanger the airplane.(c) For nonmetallic components, compliance

with paragraph (a) of this section may be shownby—

(1) Designing the components to minimize theeffect of a strike; or

(2) Incorporating acceptable means of divertingthe resulting electrical current so as not to endan-ger the airplane.



58 ASA

Subpart D —Design and Construction

GENERAL

§25.601 General.The airplane may not have design features or

details that experience has shown to be hazard-ous or unreliable. The suitability of each question-able design detail and part must be establishedby tests.

§25.603 Materials.The suitability and durability of materials used

for parts, the failure of which could adversely af-fect safety, must—

(a) Be established on the basis of experienceor tests;

(b) Conform to approved specifications (suchas industry or military specifications, or TechnicalStandard Orders) that ensure their having thestrength and other properties assumed in the de-sign data; and

(c) Take into account the effects of environmen-tal conditions, such as temperature and humidity,expected in service.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–38, 41 FR 55466, Dec. 20 1976;Amdt. 25–46, 43 FR 50595, Oct. 30, 1978]

§25.605 Fabrication methods.(a) The methods of fabrication used must pro-

duce a consistently sound structure. If a fabrica-tion process (such as gluing, spot welding, or heattreating) requires close control to reach this ob-jective, the process must be performed under anapproved process specification.

(b) Each new aircraft fabrication method mustbe substantiated by a test program.


§25.607 Fasteners.(a) Each removable bolt, screw, nut, pin, or

other removable fastener must incorporate twoseparate locking devices if—

(1) Its loss could preclude continued flight andlanding within the design limitations of the air-plane using normal pilot skill and strength; or

(2) Its loss could result in reduction in pitch,yaw, or roll control capability or response belowthat required by Subpart B of this chapter.

(b) The fasteners specified in paragraph (a) ofthis section and their locking devices may not beadversely affected by the environmental condi-tions associated with the particular installation.

(c) No self-locking nut may be used on any boltsubject to rotation in operation unless a nonfriction

locking device is used in addition to the self-lockingdevice.


§25.609 Protection of structure.Each part of the structure must—(a) Be suitably protected against deterioration

or loss of strength in service due to any cause,including—

(1) Weathering;(2) Corrosion; and(3) Abrasion; and(b) Have provisions for ventilation and drainage

where necessary for protection.

§25.611 Accessibility provisions.(a) Means must be provided to allow inspection

(including inspection of principal structural ele-ments and control systems), replacement of partsnormally requiring replacement, adjustment, andlubrication as necessary for continued airworthi-ness. The inspection means for each item mustbe practicable for the inspection interval for theitem. Nondestructive inspection aids may be usedto inspect structural elements where it is impracti-cable to provide means for direct visual inspectionif it is shown that the inspection is effective andthe inspection procedures are specified in themaintenance manual required by §25.1529.

(b) EWIS must meet the accessibility require-ments of §25.1719.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5674, April 8, 1970;Amdt. 25–123, 72 FR 63404, Nov. 8, 2007]

§25.613 Material strength properties and material design values.(a) Material strength properties must be based

on enough tests of material meeting approvedspecifications to establish design values on a sta-tistical basis.

(b) Material design values must be chosen tominimize the probability of structural failures dueto material variability. Except as provided in para-graphs (e) and (f) of this section, compliance mustbe shown by selecting material design valueswhich assure material strength with the followingprobability:

(1) Where applied loads are eventually distrib-uted through a single member within an assem-bly, the failure of which would result in loss ofstructural integrity of the component, 99 percentprobability with 95 percent confidence.

(2) For redundant structure, in which the failureof individual elements would result in appliedloads being safely distributed to other load carry-ing members, 90 percent probability with 95 per-cent confidence.


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(c) The effects of environmental conditions,such as temperature and moisture, on materialdesign values used in an essential component orstructure must be considered where these effectsare significant within the airplane operating enve-lope.

(d) [Reserved](e) Greater material design values may be

used if a “premium selection” of the material ismade in which a specimen of each individual itemis tested before use to determine that the actualstrength properties of that particular item willequal or exceed those used in design.

(f) Other material design values may be used ifapproved by the Administrator.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–46, 43 FR 50595, Oct. 30, 1978;Amdt. 25–72, 55 FR 29776, July 20, 1990; Amdt. 25–112, 68 FR 46430, Aug. 5, 2003]

§25.619 Special factors.The factor of safety prescribed in §25.303 must

be multiplied by the highest pertinent special fac-tor of safety prescribed in §§25.621 through25.625 for each part of the structure whosestrength is—

(a) Uncertain;(b) Likely to deteriorate in service before nor-

mal replacement; or(c) Subject to appreciable variability because of

uncertainties in manufacturing processes or in-spection methods.


§25.621 Casting factors.(a) General. The factors, tests, and inspections

specified in paragraphs (b) through (d) of this sec-tion must be applied in addition to those neces-sary to establish foundry quality control. The in-spections must meet approved specifications.Paragraphs (c) and (d) of this section apply to anystructural castings except castings that are pres-sure tested as parts of hydraulic or other fluid sys-tems and do not support structural loads.

(b) Bearing stresses and surfaces. The castingfactors specified in paragraphs (c) and (d) of thissection—

(1) Need not exceed 1.25 with respect to bear-ing stresses regardless of the method of inspec-tion used; and

(2) Need not be used with respect to the bear-ing surfaces of a part whose bearing factor islarger than the applicable casting factor.

(c) Critical castings. For each casting whosefailure would preclude continued safe flight andlanding of the airplane or result in serious injury tooccupants, the following apply:

(1) Each critical casting must—(i) Have a casting factor of not less than 1.25;

and(ii) Receive 100 percent inspection by visual,

radiographic, and magnetic particle or penetrantinspection methods or approved equivalent non-destructive inspection methods.

(2) For each critical casting with a casting fac-tor less than 1.50, three sample castings must bestatic tested and shown to meet—

(i) The strength requirements of §25.305 at anultimate load corresponding to a casting factor of1.25; and

(ii) The deformation requirements of §25.305 ata load of 1.15 times the limit load.

(3) Examples of these castings are structuralattachment fittings, parts of flight control systems,control surface hinges and balance weight attach-ments, seat, berth, safety belt, and fuel and oiltank supports and attachments, and cabin pres-sure valves.

(d) Noncritical castings. For each casting otherthan those specified in paragraph (c) of this sec-tion, the following apply:

(1) Except as provided in paragraphs (d)(2) and(3) of this section, the casting factors and corre-sponding inspections must meet the following ta-ble:

(2) The percentage of castings inspected bynonvisual methods may be reduced below thatspecified in paragraph (d)(1) of this section whenan approved quality control procedure is estab-lished.

(3) For castings procured to a specification thatguarantees the mechanical properties of the ma-terial in the casting and provides for demonstra-tion of these properties by test of coupons cutfrom the castings on a sampling basis—

(i) A casting factor of 1.0 may be used; and(ii) The castings must be inspected as provided

in paragraph (d)(1) of this section for casting fac-tors of “1.25 through 1.50” and tested under para-graph (c)(2) of this section.

§25.623 Bearing factors.(a) Except as provided in paragraph (b) of this

section, each part that has clearance (free fit),

Casting factor Inspection

2.0 or more 100 percent visual

Less than 2.0 but more than 1.5

100 percent visual, and magnetic particle or penetrant or equivalent nondestructive inspection methods

1.25 through 1.50 100 percent visual, magnetic particle or penetrant, and radiographic, or approved equivalent nondestructive inspection methods


60 ASA

and that is subject to pounding or vibration, musthave a bearing factor large enough to provide forthe effects of normal relative motion.

(b) No bearing factor need be used for a partfor which any larger special factor is prescribed.

§25.625 Fitting factors.For each fitting (a part or terminal used to join

one structural member to another), the followingapply:

(a) For each fitting whose strength is notproven by limit and ultimate load tests in which ac-tual stress conditions are simulated in the fittingand surrounding structures, a fitting factor of atleast 1.15 must be applied to each part of—

(1) The fitting;(2) The means of attachment; and(3) The bearing on the joined members.(b) No fitting factor need be used—(1) For joints made under approved practices

and based on comprehensive test data (such ascontinuous joints in metal plating, welded joints,and scarf joints in wood); or

(2) With respect to any bearing surface forwhich a larger special factor is used.

(c) For each integral fitting, the part must betreated as a fitting up to the point at which thesection properties become typical of the member.

(d) For each seat, berth, safety belt, and har-ness, the fitting factor specified in §25.785(f)(3)applies.


§25.629 Aeroelastic stability requirements.(a) General. The aeroelastic stability evalua-

tions required under this section include flutter, di-vergence, control reversal and any undue loss ofstability and control as a result of structural defor-mation. The aeroelastic evaluation must includewhirl modes associated with any propeller or ro-tating device that contributes significant dynamicforces. Compliance with this section must beshown by analyses, wind tunnel tests, ground vi-bration tests, flight tests, or other means foundnecessary by the Administrator.

(b) Aeroelastic stability envelopes. The air-plane must be designed to be free from aeroelas-tic instability for all configurations and design con-ditions within the aeroelastic stability envelopesas follows:

(1) For normal conditions without failures, mal-functions, or adverse conditions, all combinationsof altitudes and speeds encompassed by theVD/MD versus altitude envelope enlarged at allpoints by an increase of 15 percent in equivalentairspeed at both constant Mach number and con-

stant altitude. In addition, a proper margin of sta-bility must exist at all speeds up to VD/MD and,there must be no large and rapid reduction in sta-bility as VD/MD is approached. The enlarged en-velope may be limited to Mach 1.0 when MD isless than 1.0 at all design altitudes, and

(2) For the conditions described in §25.629(d)below, for all approved altitudes, any airspeed upto the greater airspeed defined by;

(i) The VD/MD envelope determined by§25.335(b); or

(ii) An altitude-airspeed envelope defined by a15-percent increase in equivalent airspeed aboveVC at constant altitude, from sea level to the alti-tude of the intersection of 1.15 VC with the exten-sion of the constant cruise Mach number line, MC,then a linear variation in equivalent airspeed toMC+.05 at the altitude of the lowest VC/MC inter-section; then, at higher altitudes, up to the maxi-mum flight altitude, the boundary defined by a .05Mach increase in MC at constant altitude.

(c) Balance weights. If concentrated balanceweights are used, their effectiveness and strength,including supporting structure, must be substanti-ated.

(d) Failures, malfunctions, and adverse condi-tions. The failures, malfunctions, and adverseconditions which must be considered in showingcompliance with this section are:

(1) Any critical fuel loading conditions, notshown to be extremely improbable, which may re-sult from mismanagement of fuel.

(2) Any single failure in any flutter damper sys-tem.

(3) For airplanes not approved for operation inicing conditions, the maximum likely ice accumula-tion expected as a result of an inadvertent encoun-ter.

(4) Failure of any single element of the struc-ture supporting any engine, independentlymounted propeller shaft, large auxiliary powerunit, or large externally mounted aerodynamicbody (such as an external fuel tank).

(5) For airplanes with engines that have propel-lers or large rotating devices capable of significantdynamic forces, any single failure of the enginestructure that would reduce the rigidity of the rota-tional axis.

(6) The absence of aerodynamic or gyroscopicforces resulting from the most adverse combina-tion of feathered propellers or other rotating de-vices capable of significant dynamic forces. In ad-dition, the effect of a single feathered propeller orrotating device must be coupled with the failuresof paragraphs (d)(4) and (d)(5) of this section.

(7) Any single propeller or rotating device capa-ble of significant dynamic forces rotating at thehighest likely overspeed.

(8) Any damage or failure condition, required orselected for investigation by §25.571. The single


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structural failures described in paragraphs (d)(4)and (d)(5) of this section need not be consideredin showing compliance with this section if:

(i) The structural element could not fail due todiscrete source damage resulting from the condi-tions described in §25.571(e), and

(ii) A damage tolerance investigation in accor-dance with §25.571(b) shows that the maximumextent of damage assumed for the purpose of re-sidual strength evaluation does not involve com-plete failure of the structural element.

(9) Any damage, failure, or malfunction consid-ered under §§25.631, 25.671, 25.672, and25.1309.

(10) Any other combination of failures, malfunc-tions, or adverse conditions not shown to be ex-tremely improbable.

(e) Flight flutter testing. Full scale flight fluttertests at speeds up to VDF/MDF must be conductedfor new type designs and for modifications to atype design unless the modifications have beenshown to have an insignificant effect on theaeroelastic stability. These tests must demon-strate that the airplane has a proper margin ofdamping at all speeds up to VDF/MDF, and thatthere is no large and rapid reduction in dampingas VDF/MDF, is approached. If a failure, malfunc-tion, or adverse condition is simulated duringflight test in showing compliance with paragraph(d) of this section, the maximum speed investi-gated need not exceed VFC/MFC if it is shown, bycorrelation of the flight test data with other testdata or analyses, that the airplane is free fromany aeroelastic instability at all speeds within thealtitude-airspeed envelope described in para-graph (b)(2) of this section.

[Docket No. 26007, 57 FR 28949, June 29, 1992]

§25.631 Bird strike damage.The empennage structure must be designed to

assure capability of continued safe flight and land-ing of the airplane after impact with an 8-poundbird when the velocity of the airplane (relative tothe bird along the airplane’s flight path) is equal toVC at sea level, selected under §25.335(a). Com-pliance with this section by provision of redundantstructure and protected location of control systemelements or protective devices such as splitterplates or energy absorbing material is acceptable.Where compliance is shown by analysis, tests, orboth, use of data on airplanes having similarstructural design is acceptable.


CONTROL SURFACES

§25.651 Proof of strength.(a) Limit load tests of control surfaces are re-

quired. These tests must include the horn or fittingto which the control system is attached.

(b) Compliance with the special factors require-ments of §§25.619 through 25.625 and 25.657 forcontrol surface hinges must be shown by analysisor individual load tests.

§25.655 Installation.(a) Movable tail surfaces must be installed so

that there is no interference between any surfaceswhen one is held in its extreme position and theothers are operated through their full angularmovement.

(b) If an adjustable stabilizer is used, it musthave stops that will limit its range of travel to themaximum for which the airplane is shown to meetthe trim requirements of §25.161.

§25.657 Hinges.(a) For control surface hinges, including ball,

roller, and self-lubricated bearing hinges, the ap-proved rating of the bearing may not be ex-ceeded. For nonstandard bearing hinge configu-rations, the rating must be established on the ba-sis of experience or tests and, in the absence of arational investigation, a factor of safety of not lessthan 6.67 must be used with respect to the ulti-mate bearing strength of the softest material usedas a bearing.

(b) Hinges must have enough strength and ri-gidity for loads parallel to the hinge line.


CONTROL SYSTEMS

§25.671 General.(a) Each control and control system must oper-

ate with the ease, smoothness, and positivenessappropriate to its function.

(b) Each element of each flight control systemmust be designed, or distinctively and perma-nently marked, to minimize the probability of in-correct assembly that could result in the malfunc-tioning of the system.

(c) The airplane must be shown by analysis,tests, or both, to be capable of continued safeflight and landing after any of the following failuresor jamming in the flight control system and sur-faces (including trim, lift, drag, and feel systems),within the normal flight envelope, without requir-ing exceptional piloting skill or strength. Probablemalfunctions must have only minor effects oncontrol system operation and must be capable ofbeing readily counteracted by the pilot.


62 ASA

(1) Any single failure, excluding jamming (forexample, disconnection or failure of mechanicalelements, or structural failure of hydraulic compo-nents, such as actuators, control spool housing,and valves).

(2) Any combination of failures not shown to beextremely improbable, excluding jamming (for ex-ample, dual electrical or hydraulic system failures,or any single failure in combination with any prob-able hydraulic or electrical failure).

(3) Any jam in a control position normally en-countered during takeoff, climb, cruise, normalturns, descent, and landing unless the jam isshown to be extremely improbable, or can be alle-viated. A runaway of a flight control to an adverseposition and jam must be accounted for if suchrunaway and subsequent jamming is not ex-tremely improbable.

(d) The airplane must be designed so that it iscontrollable if all engines fail. Compliance withthis requirement may be shown by analysis wherethat method has been shown to be reliable.


§25.672 Stability augmentation and automatic and power-operated systems.If the functioning of stability augmentation or

other automatic or power-operated systems isnecessary to show compliance with the flightcharacteristics requirements of this part, suchsystems must comply with §25.671 and the fol-lowing:

(a) A warning which is clearly distinguishable tothe pilot under expected flight conditions withoutrequiring his attention must be provided for anyfailure in the stability augmentation system or inany other automatic or power-operated systemwhich could result in an unsafe condition if thepilot were not aware of the failure. Warning sys-tems must not activate the control systems.

(b) The design of the stability augmentationsystem or of any other automatic or power-oper-ated system must permit initial counteraction offailures of the type specified in §25.671(c) withoutrequiring exceptional pilot skill or strength, by ei-ther the deactivation of the system, or a failed por-tion thereof, or by overriding the failure by move-ment of the flight controls in the normal sense.

(c) It must be shown that after any single failureof the stability augmentation system or any otherautomatic or power-operated system—

(1) The airplane is safely controllable when thefailure or malfunction occurs at any speed or alti-tude within the approved operating limitations thatis critical for the type of failure being considered;

(2) The controllability and maneuverability re-quirements of this part are met within a practical

operational flight envelope (for example, speed,altitude, normal acceleration, and airplane config-urations) which is described in the Airplane FlightManual; and

(3) The trim, stability, and stall characteristicsare not impaired below a level needed to permitcontinued safe flight and landing.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5675 April 8, 1970]

§25.675 Stops.(a) Each control system must have stops that

positively limit the range of motion of each movableaerodynamic surface controlled by the system.

(b) Each stop must be located so that wear,slackness, or take-up adjustments will not ad-versely affect the control characteristics of the air-plane because of a change in the range of surfacetravel.

(c) Each stop must be able to withstand anyloads corresponding to the design conditions forthe control system.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–38, 41 FR 55466, Dec. 20, 1976]

§25.677 Trim systems.(a) Trim controls must be designed to prevent

inadvertent or abrupt operation and to operate inthe plane, and with the sense of motion, of the air-plane.

(b) There must be means adjacent to the trimcontrol to indicate the direction of the controlmovement relative to the airplane motion. In addi-tion, there must be clearly visible means to indi-cate the position of the trim device with respect tothe range of adjustment.

(c) Trim control systems must be designed toprevent creeping in flight. Trim tab controls mustbe irreversible unless the tab is appropriately bal-anced and shown to be free from flutter.

(d) If an irreversible tab control system is used,the part from the tab to the attachment of the irre-versible unit to the airplane structure must consistof a rigid connection.


§25.679 Control system gust locks.(a) There must be a device to prevent damage

to the control surfaces (including tabs), and to thecontrol system, from gusts striking the airplanewhile it is on the ground or water. If the device,when engaged, prevents normal operation of thecontrol surfaces by the pilot, it must—

(1) Automatically disengage when the pilot op-erates the primary flight controls in a normal man-ner; or


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(2) Limit the operation of the airplane so thatthe pilot receives unmistakable warning at thestart of takeoff.

(b) The device must have means to precludethe possibility of it becoming inadvertently en-gaged in flight.

§25.681 Limit load static tests.(a) Compliance with the limit load requirements

of this Part must be shown by tests in which—(1) The direction of the test loads produces the

most severe loading in the control system; and(2) Each fitting, pulley, and bracket used in at-

taching the system to the main structure is in-cluded.

(b) Compliance must be shown (by analyses orindividual load tests) with the special factor re-quirements for control system joints subject to an-gular motion.

§25.683 Operation tests.It must be shown by operation tests that when

portions of the control system subject to pilot ef-fort loads are loaded to 80 percent of the limit loadspecified for the system and the powered portionsof the control system are loaded to the maximumload expected in normal operation, the system isfree from—

(a) Jamming;(b) Excessive friction; and(c) Excessive deflection.


§25.685 Control system details.(a) Each detail of each control system must be

designed and installed to prevent jamming, chaf-ing, and interference from cargo, passengers,loose objects, or the freezing of moisture.

(b) There must be means in the cockpit to pre-vent the entry of foreign objects into places wherethey would jam the system.

(c) There must be means to prevent the slap-ping of cables or tubes against other parts.

(d) Sections 25.689 and 25.693 apply to cablesystems and joints.


§25.689 Cable systems.(a) Each cable, cable fitting, turnbuckle, splice,

and pulley must be approved. In addition—(1) No cable smaller than 1⁄8 inch in diameter

may be used in the aileron, elevator, or ruddersystems; and

(2) Each cable system must be designed sothat there will be no hazardous change in cable

tension throughout the range of travel under oper-ating conditions and temperature variations.

(b) Each kind and size of pulley must corre-spond to the cable with which it is used. Pulleysand sprockets must have closely fitted guards toprevent the cables and chains from being dis-placed or fouled. Each pulley must lie in the planepassing through the cable so that the cable doesnot rub against the pulley flange.

(c) Fairleads must be installed so that they donot cause a change in cable direction of morethan three degrees.

(d) Clevis pins subject to load or motion and re-tained only by cotter pins may not be used in thecontrol system.

(e) Turnbuckles must be attached to parts hav-ing angular motion in a manner that will positivelyprevent binding throughout the range of travel.

(f) There must be provisions for visual inspectionof fairleads, pulleys, terminals, and turnbuckles.

§25.693 Joints.Control system joints (in push-pull systems)

that are subject to angular motion, except those inball and roller bearing systems, must have a spe-cial factor of safety of not less than 3.33 with re-spect to the ultimate bearing strength of the soft-est material used as a bearing. This factor may bereduced to 2.0 for joints in cable control systems.For ball or roller bearings, the approved ratingsmay not be exceeded.


§25.697 Lift and drag devices, controls.(a) Each lift device control must be designed so

that the pilots can place the device in any takeoff,en route, approach, or landing position estab-lished under §25.101(d). Lift and drag devicesmust maintain the selected positions, except formovement produced by an automatic positioningor load limiting device, without further attention bythe pilots.

(b) Each lift and drag device control must bedesigned and located to make inadvertent opera-tion improbable. Lift and drag devices intended forground operation only must have means to pre-vent the inadvertent operation of their controls inflight if that operation could be hazardous.

(c) The rate of motion of the surfaces in re-sponse to the operation of the control and thecharacteristics of the automatic positioning orload limiting device must give satisfactory flightand performance characteristics under steady orchanging conditions of airspeed, engine power,and airplane attitude.

(d) The lift device control must be designed toretract the surfaces from the fully extended posi-


64 ASA

tion, during steady flight at maximum continuousengine power at any speed below VF +9.0 (knots).

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5675, April 8, 1970;Amdt. 25–46, 43 FR 50595, Oct. 30, 1978; Amdt. 25–57,49 FR 6848, Feb. 23, 1984]

§25.699 Lift and drag device indicator.(a) There must be means to indicate to the

pilots the position of each lift or drag device hav-ing a separate control in the cockpit to adjust itsposition. In addition, an indication of unsymmetri-cal operation or other malfunction in the lift ordrag device systems must be provided when suchindication is necessary to enable the pilots to pre-vent or counteract an unsafe flight or ground con-dition, considering the effects on flight character-istics and performance.

(b) There must be means to indicate to thepilots the takeoff, en route, approach, and landinglift device positions.

(c) If any extension of the lift and drag devicesbeyond the landing position is possible, the con-trols must be clearly marked to identify this rangeof extension.


§25.701 Flap and slat interconnection.(a) Unless the airplane has safe flight charac-

teristics with the flaps or slats retracted on oneside and extended on the other, the motion offlaps or slats on opposite sides of the plane ofsymmetry must be synchronized by a mechanicalinterconnection or approved equivalent means.

(b) If a wing flap or slat interconnection orequivalent means is used, it must be designed toaccount for the applicable unsymmetrical loads,including those resulting from flight with the en-gines on one side of the plane of symmetry inop-erative and the remaining engines at takeoffpower.

(c) For airplanes with flaps or slats that are notsubjected to slipstream conditions, the structuremust be designed for the loads imposed when thewing flaps or slats on one side are carrying themost severe load occurring in the prescribed sym-metrical conditions and those on the other sideare carrying not more than 80 percent of thatload.

(d) The interconnection must be designed forthe loads resulting when interconnected flap orslat surfaces on one side of the plane of symme-try are jammed and immovable while the surfaceson the other side are free to move and the fullpower of the surface actuating system is applied.


§25.703 Takeoff warning system.A takeoff warning system must be installed and

must meet the following requirements:(a) The system must provide to the pilots an au-

ral warning that is automatically activated duringthe initial portion of the takeoff roll if the airplaneis in a configuration, including any of the following,that would not allow a safe takeoff:

(1) The wing flaps or leading edge devices arenot within the approved range of takeoff positions.

(2) Wing spoilers (except lateral control spoil-ers meeting the requirements of §25.671), speedbrakes, or longitudinal trim devices are in a posi-tion that would not allow a safe takeoff.

(b) The warning required by paragraph (a) ofthis section must continue until—

(1) The configuration is changed to allow a safetakeoff;

(2) Action is taken by the pilot to terminate thetakeoff 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 oftakeoff weights, altitudes, and temperatures forwhich certification is requested.


LANDING GEAR

§25.721 General.(a) The main landing gear system must be de-

signed so that if it fails due to overloads duringtakeoff and landing (assuming the overloads toact in the upward and aft directions), the failuremode is not likely to cause—

(1) For airplanes that have passenger seatingconfiguration, excluding pilots seats, of nine seatsor less, the spillage of enough fuel from any fuelsystem in the fuselage to constitute a fire hazard;and

(2) For airplanes that have a passenger seatingconfiguration, excluding pilots seats, of 10 seatsor more, the spillage of enough fuel from any partof the fuel system to constitute a fire hazard.

(b) Each airplane that has a passenger seatingconfiguration excluding pilots seats, of 10 seats ormore must be designed so that with the airplaneunder control it can be landed on a paved runwaywith any one or more landing gear legs not ex-tended without sustaining a structural componentfailure that is likely to cause the spillage ofenough fuel to constitute a fire hazard.

(c) Compliance with the provisions of this sec-tion may be shown by analysis or tests, or both.



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§25.723 Shock absorption tests.(a) The analytical representation of the landing

gear dynamic characteristics that is used in deter-mining the landing loads must be validated by en-ergy absorption tests. A range of tests must beconducted to ensure that the analytical represen-tation is valid for the design conditions specified in§25.473.

(1) The configurations subjected to energy ab-sorption tests at limit design conditions must in-clude at least the design landing weight or the de-sign takeoff weight, whichever produces thegreater value of landing impact energy.

(2) The test attitude of the landing gear unit andthe application of appropriate drag loads duringthe test must simulate the airplane landing condi-tions in a manner consistent with the developmentof rational or conservative limit loads.

(b) The landing gear may not fail in a test, dem-onstrating its reserve energy absorption capacity,simulating a descent velocity of 12 f.p.s. at designlanding weight, assuming airplane lift not greaterthan airplane weight acting during the landing im-pact.

(c) In lieu of the tests prescribed in this section,changes in previously approved design weightsand minor changes in design may be substanti-ated by analyses based on previous tests con-ducted on the same basic landing gear systemthat has similar energy absorption characteristics.

[Docket No. FAA–1999–5835, 66 FR 27394, May 16,2001]

§25.725 [Reserved]

§25.727 [Reserved]

§25.729 Retracting mechanism.(a) General. For airplanes with retractable land-

ing gear, the following apply:(1) The landing gear retracting mechanism,

wheel well doors, and supporting structure, mustbe designed for—

(i) The loads occurring in the flight conditionswhen the gear is in the retracted position,

(ii) The combination of friction loads, inertialoads, brake torque loads, air loads, and gyro-scopic loads resulting from the wheels rotating ata peripheral speed equal to 1.3 VS (with the flapsin takeoff position at design takeoff weight), oc-curring during retraction and extension at any air-speed up to 1.6 VS1 (with the flaps in the ap-proach 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 deceleratethe airplane in flight at this speed, the landinggear, the retracting mechanism, and the airplanestructure (including wheel well doors) must be de-

signed to withstand the flight loads occurring withthe landing gear in the extended position at anyspeed up to 0.67 VC.

(3) Landing gear doors, their operating mecha-nism, and their supporting structures must be de-signed for the yawing maneuvers prescribed forthe airplane in addition to the conditions of air-speed and load factor prescribed in paragraphs(a)(1) and (2) of this section.

(b) Landing gear lock. There must be positivemeans to keep the landing gear extended, in flightand on the ground.

(c) Emergency operation. There must be anemergency means for extending the landing gearin the event of—

(1) Any reasonably probable failure in the nor-mal retraction system; or

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

(d) Operation test. The proper functioning ofthe retracting mechanism must be shown by oper-ation tests.

(e) Position indicator and warning device. If aretractable landing gear is used, there must be alanding gear position indicator (as well as neces-sary switches to actuate the indicator) or othermeans to inform the pilot that the gear is securedin the extended (or retracted) position. Thismeans must be designed as follows:

(1) If switches are used, they must be locatedand coupled to the landing gear mechanical sys-tems in a manner that prevents an erroneous indi-cation of “down and locked” if the landing gear isnot in a fully extended position, or of “up andlocked” if the landing gear is not in the fully re-tracted position. The switches may be locatedwhere they are operated by the actual landinggear locking latch or device.

(2) The flightcrew must be given an aural warn-ing that functions continuously, or is periodicallyrepeated, if a landing is attempted when the land-ing gear is not locked down.

(3) The warning must be given in sufficient timeto allow the landing gear to be locked down or ago-around to be made.

(4) There must be a manual shut-off meansreadily available to the flightcrew for the warningrequired by paragraph (e)(2) of this section suchthat it could be operated instinctively, inadvert-ently, or by habitual reflexive action.

(5) The system used to generate the auralwarning must be designed to eliminate false or in-appropriate alerts.

(6) Failures of systems used to inhibit the landinggear aural warning, that would prevent the warningsystem from operating, must be improbable.

(f) Protection of equipment in wheel wells.Equipment that is essential to safe operation ofthe airplane and that is located in wheel wellsmust be protected from the damaging effects of—


66 ASA

(1) A bursting tire, unless it is shown that a tirecannot burst from overheat; and

(2) A loose tire tread, unless it is shown that aloose tire tread cannot cause damage.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5676, April 8, 1970;Amdt. 25–42, 43 FR 2323, Jan. 16, 1978; Amdt. 25–72,55 FR 29777, July 20, 1990; Amdt. 25–75, 56 FR 63762,Dec. 5, 1991]

§25.731 Wheels.(a) Each main and nose wheel must be ap-

proved.(b) The maximum static load rating of each

wheel may not be less than the correspondingstatic 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 radiallimit load determined under the applicable groundload requirements of this part.

(d) Overpressure burst prevention. Meansmust be provided in each wheel to prevent wheelfailure and tire burst that may result from exces-sive pressurization of the wheel and tire assem-bly.

(e) Braked wheels. Each braked wheel mustmeet the applicable requirements of §25.735.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–72, 55 FR 29777, July 20, 1990;Amdt. 25–107, 67 FR 20420, April 24, 2002]

§25.733 Tires.(a) When a landing gear axle is fitted with a sin-

gle wheel and tire assembly, the wheel must befitted with a suitable tire of proper fit with a speedrating approved by the Administrator that is notexceeded under critical conditions and with a loadrating approved by the Administrator that is notexceeded under—

(1) The loads on the main wheel tire, corre-sponding to the most critical combination of air-plane weight (up to maximum weight) and centerof gravity position, and

(2) The loads corresponding to the ground re-actions in paragraph (b) of this section, on thenose wheel tire, except as provided in paragraphs(b)(2) and (b)(3) of this section.

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

(1) The static ground reaction for the tire corre-sponding to the most critical combination of air-plane weight (up to maximum ramp weight) andcenter of gravity position with a force of 1.0g act-ing downward at the center of gravity. This loadmay not exceed the load rating of the tire.

(2) The ground reaction of the tire correspond-ing to the most critical combination of airplane

weight (up to maximum landing weight) and cen-ter of gravity position combined with forces of 1.0gdownward and 0.31g forward acting at the centerof gravity. The reactions in this case must be dis-tributed to the nose and main wheels by the prin-ciples of statics with a drag reaction equal to 0.31times the vertical load at each wheel with brakescapable of producing this ground reaction. Thisnose tire load may not exceed 1.5 times the loadrating of the tire.

(3) The ground reaction of the tire correspond-ing to the most critical combination of airplaneweight (up to maximum ramp weight) and centerof gravity position combined with forces of 1.0gdownward and 0.20g forward acting at the centerof gravity. The reactions in this case must be dis-tributed to the nose and main wheels by the prin-ciples of statics with a drag reaction equal to 0.20times the vertical load at each wheel with brakescapable of producing this ground reaction. Thisnose tire load may not exceed 1.5 times the loadrating of the tire.

(c) When a landing gear axle is fitted with morethan one wheel and tire assembly, such as dual ordual-tandem, each wheel must be fitted with asuitable tire of proper fit with a speed rating ap-proved by the Administrator that is not exceededunder critical conditions, and with a load ratingapproved by the Administrator that is not ex-ceeded by—

(1) The loads on each main wheel tire, corre-sponding to the most critical combination of air-plane weight (up to maximum weight) and centerof gravity position, when multiplied by a factor of1.07; and

(2) Loads specified in paragraphs (a)(2), (b)(1),(b)(2), and (b)(3) of this section on each nosewheel tire.

(d) Each tire installed on a retractable landinggear system must, at the maximum size of the tiretype expected in service, have a clearance to sur-rounding structure and systems that is adequateto prevent unintended contact between the tireand any part of the structure or systems.

(e) For an airplane with a maximum certificatedtakeoff weight of more than 75,000 pounds, tiresmounted on braked wheels must be inflated withdry nitrogen or other gases shown to be inert sothat the gas mixture in the tire does not containoxygen in excess of 5 percent by volume, unless itcan be shown that the tire liner material will notproduce a volatile gas when heated or that meansare provided to prevent tire temperatures fromreaching unsafe levels.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–48, 44 FR 68752, Nov. 29, 1979;Amdt. 25–72, 55 FR 29777, July 20, 1990; Amdt. 25–78,58 FR 11781, Feb. 26, 1993]


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§25.735 Brakes and braking systems.(a) Approval. Each assembly consisting of a

wheel(s) and brake(s) must be approved.(b) Brake system capability. The brake system,

associated systems and components must be de-signed and constructed so that:

(1) If any electrical, pneumatic, hydraulic, ormechanical connecting or transmitting elementfails, or if any single source of hydraulic or otherbrake operating energy supply is lost, it is possi-ble to bring the airplane to rest with a braked rollstopping distance of not more than two times thatobtained in determining the landing distance asprescribed in §25.125.

(2) Fluid lost from a brake hydraulic system fol-lowing a failure in, or in the vicinity of, the brakesis insufficient to cause or support a hazardous fireon the ground or in flight.

(c) Brake controls. The brake controls must bedesigned and constructed so that:

(1) Excessive control force is not required fortheir operation.

(2) If an automatic braking system is installed,means are provided to:

(i) Arm and disarm the system, and(ii) Allow the pilot(s) to override the system by

use of manual braking.(d) Parking brake. The airplane must have a

parking brake control that, when selected on, will,without further attention, prevent the airplanefrom rolling on a dry and level paved runway whenthe most adverse combination of maximum thruston one engine and up to maximum ground idlethrust on any, or all, other engine(s) is applied.The control must be suitably located or be ade-quately protected to prevent inadvertent opera-tion. There must be indication in the cockpit whenthe parking brake is not fully released.

(e) Antiskid system. If an antiskid system is in-stalled:

(1) It must operate satisfactorily over the rangeof expected runway conditions, without externaladjustment.

(2) It must, at all times, have priority over theautomatic braking system, if installed.

(f) Kinetic energy capacity—(1) Design landing stop. The design landing

stop is an operational landing stop at maximumlanding weight. The design landing stop brake ki-netic energy absorption requirement of eachwheel, brake, and tire assembly must be deter-mined. It must be substantiated by dynamometertesting that the wheel, brake and tire assembly iscapable of absorbing not less than this level of ki-netic energy throughout the defined wear range ofthe brake. The energy absorption rate derivedfrom the airplane manufacturer’s braking require-ments must be achieved. The mean decelerationmust not be less than 10 fps2.

(2) Maximum kinetic energy accelerate-stop.The maximum kinetic energy accelerate-stop is arejected takeoff for the most critical combinationof airplane takeoff weight and speed. The acceler-ate-stop brake kinetic energy absorption require-ment of each wheel, brake, and tire assemblymust be determined. It must be substantiated bydynamometer testing that the wheel, brake, andtire assembly is capable of absorbing not lessthan this level of kinetic energy throughout the de-fined wear range of the brake. The energy ab-sorption rate derived from the airplane manufac-turer’s braking requirements must be achieved.The mean deceleration must not be less than 6fps2.

(3) Most severe landing stop. The most severelanding stop is a stop at the most critical combina-tion of airplane landing weight and speed. Themost severe landing stop brake kinetic energy ab-sorption requirement of each wheel, brake, andtire assembly must be determined. It must be sub-stantiated by dynamometer testing that, at the de-clared fully worn limit(s) of the brake heat sink, thewheel, brake and tire assembly is capable of ab-sorbing not less than this level of kinetic energy.The most severe landing stop need not be consid-ered for extremely improbable failure conditionsor if the maximum kinetic energy accelerate-stopenergy is more severe.

(g) In the landing case, the minimum speed rat-ing 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 determi-nation of kinetic energy in accordance with para-graph (f) of this section, assuming that the testprocedures for wheel-brake assemblies involve aspecified rate of deceleration, and, therefore, forthe same amount of kinetic energy, the rate of en-ergy absorption (the power absorbing ability ofthe brake) varies inversely with the initial speed.

(h) Stored energy systems. An indication to theflightcrew of the usable stored energy must beprovided if a stored energy system is used toshow compliance with paragraph (b)(1) of thissection. The available stored energy must be suf-ficient for:

(1) At least 6 full applications of the brakeswhen an antiskid system is not operating; and

(2) Bringing the airplane to a complete stopwhen an antiskid system is operating, under allrunway surface conditions for which the airplaneis certificated.

(i) Brake wear indicators. Means must be pro-vided for each brake assembly to indicate whenthe heat sink is worn to the permissible limit. Themeans must be reliable and readily visible.

(j) Overtemperature burst prevention. Meansmust be provided in each braked wheel to preventa wheel failure, a tire burst, or both, that may re-sult from elevated brake temperatures. Addition-


68 ASA

ally, all wheels must meet the requirements of§25.731(d).

(k) Compatibility. Compatibility of the wheeland brake assemblies with the airplane and itssystems must be substantiated.

[Docket No. FAA–1999–6063, 67 FR 20420, April 24,2002; as amended by Amdt. 25–108, 67 FR 70828, Nov.26, 2002]

§25.737 Skis.Each ski must be approved. The maximum limit

load rating of each ski must equal or exceed themaximum limit load determined under the appli-cable ground load requirements of this part.

FLOATS AND HULLS

§25.751 Main float buoyancy.Each main float must have—(a) A buoyancy of 80 percent in excess of that

required to support the maximum weight of theseaplane or amphibian in fresh water; and

(b) Not less than five watertight compartmentsapproximately equal in volume.

§25.753 Main float design.Each main float must be approved and must

meet the requirements of §25.521.

§25.755 Hulls.(a) Each hull must have enough watertight

compartments so that, with any two adjacentcompartments flooded, the buoyancy of the hulland auxiliary floats (and wheel tires, if used) pro-vides a margin of positive stability great enough tominimize the probability of capsizing in rough,fresh water.

(b) Bulkheads with watertight doors may beused for communication between compartments.

PERSONNEL AND CARGO ACCOMMODATIONS

§25.771 Pilot compartment.(a) Each pilot compartment and its equipment

must allow the minimum flight crew (establishedunder §25.1523) to perform their duties withoutunreasonable concentration or fatigue.

(b) The primary controls listed in §25.779(a),excluding cables and control rods, must be lo-cated with respect to the propellers so that nomember of the minimum flight crew (establishedunder §25.1523), or part of the controls, lies in theregion between the plane of rotation of any in-board propeller and the surface generated by aline passing through the center of the propellerhub making an angle of five degrees forward or aftof the plane of rotation of the propeller.

(c) If provision is made for a second pilot, theairplane must be controllable with equal safetyfrom either pilot seat.

(d) The pilot compartment must be constructedso that, when flying in rain or snow, it will not leakin a manner that will distract the crew or harm thestructure.

(e) Vibration and noise characteristics of cock-pit equipment may not interfere with safe opera-tion of the airplane.


§25.772 Pilot compartment doors.For an airplane that has a lockable door in-

stalled between the pilot compartment and thepassenger compartment:

(a) For airplanes with a maximum passengerseating configuration of more than 20 seats, theemergency exit configuration must be designedso that neither crewmembers nor passengers re-quire use of the flightdeck door in order to reachthe emergency exits provided for them; and

(b) Means must be provided to enable flightcrewmembers to directly enter the passengercompartment from the pilot compartment if thecockpit door becomes jammed.

(c) There must be an emergency means to en-able a flight attendant to enter the pilot compart-ment in the event that the flightcrew becomes in-capacitated.

[Docket No. 24344, 55 FR 29777, July 20, 1990; asamended by Amdt. 25–106, 67 FR 2127, Jan. 15, 2002]

§25.773 Pilot compartment view.(a) Nonprecipitation conditions. For nonprecipi-

tation 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 safelyperform any maneuvers within the operating limi-tations of the airplane, including taxiing takeoff,approach, and landing.

(2) Each pilot compartment must be free ofglare and reflection that could interfere with thenormal duties of the minimum flight crew (estab-lished under §25.1523). This must be shown inday and night flight tests under nonprecipitationconditions.

(b) Precipitation conditions. For precipitationconditions, the following apply:

(1) The airplane must have a means to main-tain a clear portion of the windshield, during pre-cipitation conditions, sufficient for both pilots tohave a sufficiently extensive view along the flightpath in normal flight attitudes of the airplane. Thismeans must be designed to function, without con-tinuous attention on the part of the crew, in—

(i) Heavy rain at speeds up to 1.5 VSR1 with liftand drag devices retracted; and


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(ii) The icing conditions specified in §25.1419 ifcertification for flight in icing conditions is re-quested.

(2) The first pilot must have—(i) A window that is openable under the condi-

tions prescribed in paragraph (b)(1) of this sectionwhen the cabin is not pressurized, provides theview specified in that paragraph, and gives suffi-cient protection from the elements against impair-ment of the pilot’s vision; or

(ii) An alternate means to maintain a clear viewunder the conditions specified in paragraph (b)(1)of this section, considering the probable damagedue to a severe hail encounter.

(c) Internal windshield and window fogging.The airplane must have a means to prevent fog-ging of the internal portions of the windshield andwindow panels over an area which would providethe visibility specified in paragraph (a) of this sec-tion under all internal and external ambient condi-tions, including precipitation conditions, in whichthe airplane is intended to be operated.

(d) Fixed markers or other guides must be in-stalled at each pilot station to enable the pilots toposition themselves in their seats for an optimumcombination of outside visibility and instrumentscan. If lighted markers or guides are used theymust comply with the requirements specified in§25.1381.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5676, April 8, 1970;Amdt. 25–46, 43 FR 50595, Oct. 30, 1978; Amdt. 25–72,55 FR 29778, July 20, 1990; Amdt. 25–108, 67 FR70828, Nov. 26, 2002; Amdt. 25–121, 72 FR 44669, Aug.8, 2007]

§25.775 Windshields and windows.(a) Internal panes must be made of nonsplin-

tering material.(b) Windshield panes directly in front of the

pilots in the normal conduct of their duties, andthe supporting structures for these panes, mustwithstand, without penetration, the impact of afour-pound bird when the velocity of the airplane(relative to the bird along the airplane’s flight path)is equal to the value of VC, at sea level, selectedunder §25.335(a).

(c) Unless it can be shown by analysis or teststhat the probability of occurrence of a criticalwindshield fragmentation condition is of a low or-der, the airplane must have a means to minimizethe danger to the pilots from flying windshieldfragments due to bird impact. This must be shownfor each transparent pane in the cockpit that—

(1) Appears in the front view of the airplane;(2) Is inclined 15 degrees or more to the longi-

tudinal axis of the airplane; and(3) Has any part of the pane located where its

fragmentation will constitute a hazard to thepilots.

d) The design of windshields and windows inpressurized airplanes must be based on factorspeculiar to high altitude operation, including theeffects of continuous and cyclic pressurizationloadings, the inherent characteristics of the mate-rial used, and the effects of temperatures andtemperature differentials. The windshield and win-dow panels must be capable of withstanding themaximum cabin pressure differential loads com-bined with critical aerodynamic pressure and tem-perature effects after any single failure in the in-stallation or associated systems. It may be as-sumed that, after a single failure that is obvious tothe flight crew (established under §25.1523), thecabin pressure differential is reduced from themaximum, in accordance with appropriate operat-ing limitations, to allow continued safe flight of theairplane with a cabin pressure altitude of not morethan 15,000 feet.

(e) The windshield panels in front of the pilotsmust be arranged so that, assuming the loss of vi-sion through any one panel, one or more panelsremain available for use by a pilot seated at a pilotstation to permit continued safe flight and landing.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5676, April 8, 1970;Amdt. 25–38, 41 FR 55466, Dec. 20, 1976]

§25.777 Cockpit controls.(a) Each cockpit control must be located to pro-

vide convenient operation and to prevent confu-sion and inadvertent operation.

(b) The direction of movement of cockpit con-trols must meet the requirements of §25.779.Wherever practicable, the sense of motion involvedin the operation of other controls must correspondto the sense of the effect of the operation upon theairplane or upon the part operated. Controls of avariable nature using a rotary motion must moveclockwise from the off position, through an increas-ing range, to the full on position.

(c) The controls must be located and arranged,with respect to the pilots’ seats, so that there isfull and unrestricted movement of each controlwithout interference from the cockpit structure orthe clothing of the minimum flight crew (estab-lished under §25.1523) when any member of thisflight crew, from 5'2" to 6'3" in height, is seatedwith the seat belt and shoulder harness (if pro-vided) fastened.

(d) Identical powerplant controls for each en-gine must be located to prevent confusion as tothe engines they control.

(e) Wing flap controls and other auxiliary lift de-vice controls must be located on top of the pedes-tal, aft of the throttles, centrally or to the right ofthe pedestal centerline, and not less than 10inches aft of the landing gear control.


70 ASA

(f) The landing gear control must be locatedforward of the throttles and must be operable byeach pilot when seated with seat belt and shoul-der harness (if provided) fastened.

(g) Control knobs must be shaped in accor-dance with §25.781. In addition, the knobs mustbe of the same color, and this color must contrastwith the color of control knobs for other purposesand the surrounding cockpit.

(h) If a flight engineer is required as part of theminimum flight crew (established under §25.1523),the airplane must have a flight engineer station lo-cated and arranged so that the flight crewmemberscan perform their functions efficiently and withoutinterfering with each other.


§25.779 Motion and effect of cockpit controls.Cockpit controls must be designed so that they

operate in accordance with the following move-ment and actuation:

(a) Aerodynamic controls:(1) Primary.

(2) Secondary.

(b) Powerplant and auxiliary controls:(1) Powerplant.

(2) Auxiliary.


Controls Motion and effect

Aileron Right (clockwise) for right wing down

Elevator Rearward for nose up

Rudder Right pedal forward for nose right


Flaps (or auxiliary lift devices)

Forward for flaps up; rearward for flaps down

Trim tabs (or equivalent)

Rotate to produce similar rotation of the airplane about an axis parallel to the axis of the control


Power or thrust Forward to increase forward thrust and rearward to increase rearward thrust

Propellers Forward to increase rpm

Mixture Forward or upward for rich

Carburetor air heat

Forward or upward for cold

Supercharger Forward or upward for low blower For turbosuperchargers, forward, upward, or clockwise, to increase pressure


Landing gear Down to extend


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§25.781 Cockpit control knob shape.Cockpit control knobs must conform to the gen-

eral shapes (but not necessarily the exact sizes orspecific proportions) in the following figure:

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; as amended by Amdt. 25–72, 55 FR 29779, July 20, 1990]

§25.783 Fuselage doors.(a) General. This section applies to fuselage

doors, which includes all doors, hatches, open-able windows, access panels, covers, etc., on theexterior of the fuselage that do not require the useof tools to open or close. This also applies to eachdoor or hatch through a pressure bulkhead, in-

cluding any bulkhead that is specifically designedto function as a secondary bulkhead under theprescribed failure conditions of part 25. Thesedoors must meet the requirements of this section,taking into account both pressurized and unpres-surized flight, and must be designed as follows:

FLAP CONTROL KNOB LANDING GEAR CONTROL KNOB

MIXTURE CONTROL KNOB SUPERCHARGER CONTROL KNOB

POWER OR THRUST KNOB PROPELLER CONTROL KNOB


72 ASA

(1) Each door must have means to safeguardagainst opening in flight as a result of mechanicalfailure, or failure of any single structural element.

(2) Each door that could be a hazard if it un-latches must be designed so that unlatching dur-ing pressurized and unpressurized flight from thefully closed, latched, and locked condition is ex-tremely improbable. This must be shown by safetyanalysis.

(3) Each element of each door operating sys-tem must be designed or, where impracticable,distinctively and permanently marked, to minimizethe probability of incorrect assembly and adjust-ment that could result in a malfunction.

(4) All sources of power that could initiate un-locking or unlatching of any door must be auto-matically isolated from the latching and lockingsystems prior to flight and it must not be possibleto restore power to the door during flight.

(5) Each removable bolt, screw, nut, pin, orother removable fastener must meet the lockingrequirements of §25.607.

(6) Certain doors, as specified by §25.807(h),must also meet the applicable requirements of§§25.809 through 25.812 for emergency exits.

(b) Opening by persons. There must be ameans to safeguard each door against openingduring flight due to inadvertent action by persons.In addition, design precautions must be taken tominimize the possibility for a person to open adoor intentionally during flight. If these precau-tions include the use of auxiliary devices, thosedevices and their controlling systems must be de-signed so that—

(1) No single failure will prevent more than oneexit from being opened; and

(2) Failures that would prevent opening of theexit after landing are improbable.

(c) Pressurization prevention means. Theremust be a provision to prevent pressurization ofthe airplane to an unsafe level if any door subjectto pressurization is not fully closed, latched, andlocked.

(1) The provision must be designed to functionafter any single failure, or after any combination offailures not shown to be extremely improbable.

(2) Doors that meet the conditions described inparagraph (h) of this section are not required tohave a dedicated pressurization preventionmeans if, from every possible position of the door,it will remain open to the extent that it preventspressurization or safely close and latch as pres-surization takes place. This must also be shownwith any single failure and malfunction, exceptthat—

(i) With failures or malfunctions in the latchingmechanism, it need not latch after closing; and

(ii) With jamming as a result of mechanical fail-ure or blocking debris, the door need not closeand latch if it can be shown that the pressurization

loads on the jammed door or mechanism wouldnot result in an unsafe condition.

(d) Latching and locking. The latching andlocking mechanisms must be designed as follows:

(1) There must be a provision to latch eachdoor.

(2) The latches and their operating mechanismmust be designed so that, under all airplane flightand ground loading conditions, with the doorlatched, there is no force or torque tending to un-latch the latches. In addition, the latching systemmust include a means to secure the latches in thelatched position. This means must be indepen-dent of the locking system.

(3) Each door subject to pressurization, and forwhich the initial opening movement is not inward,must—

(i) Have an individual lock for each latch;(ii) Have the lock located as close as practica-

ble to the latch; and(iii) Be designed so that, during pressurized

flight, no single failure in the locking system wouldprevent the locks from restraining the latches nec-essary to secure the door.

(4) Each door for which the initial openingmovement is inward, and unlatching of the doorcould result in a hazard, must have a lockingmeans to prevent the latches from becoming dis-engaged. The locking means must ensure suffi-cient latching to prevent opening of the door evenwith a single failure of the latching mechanism.

(5) It must not be possible to position the lock inthe locked position if the latch and the latchingmechanism are not in the latched position.

(6) It must not be possible to unlatch thelatches with the locks in the locked position. Locksmust be designed to withstand the limit loads re-sulting from—

(i) The maximum operator effort when thelatches are operated manually;

(ii) The powered latch actuators, if installed;and

(iii) The relative motion between the latch andthe structural counterpart.

(7) Each door for which unlatching would notresult in a hazard is not required to have a lockingmechanism meeting the requirements of para-graphs (d)(3) through (d)(6) of this section.

(e) Warning, caution, and advisory indica-tions. Doors must be provided with the followingindications:

(1) There must be a positive means to indicateat each door operator’s station that all requiredoperations to close, latch, and lock the door(s)have been completed.

(2) There must be a positive means clearly vis-ible from each operator station for any door thatcould be a hazard if unlatched to indicate if thedoor is not fully closed, latched, and locked.


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(3) There must be a visual means on the flightdeck to signal the pilots if any door is not fullyclosed, latched, and locked. The means must bedesigned such that any failure or combination offailures that would result in an erroneous closed,latched, and locked indication is improbable for—

(i) Each door that is subject to pressurizationand for which the initial opening movement is notinward; or

(ii) Each door that could be a hazard if un-latched.

(4) There must be an aural warning to the pilotsprior to or during the initial portion of takeoff roll ifany door is not fully closed, latched, and locked,and its opening would prevent a safe takeoff andreturn to landing.

(f) Visual inspection provision. Each door forwhich unlatching of the door could be a hazardmust have a provision for direct visual inspectionto determine, without ambiguity, if the door is fullyclosed, latched, and locked. The provision mustbe permanent and discernible under operationallighting conditions, or by means of a flashlight orequivalent light source.

(g) Certain maintenance doors, removableemergency exits, and access panels. Somedoors not normally opened except for mainte-nance purposes or emergency evacuation andsome access panels need not comply with certainparagraphs of this section as follows:

(1) Access panels that are not subject to cabinpressurization and would not be a hazard if openduring flight need not comply with paragraphs (a)through (f) of this section, but must have a meansto prevent inadvertent opening during flight.

(2) Inward-opening removable emergency exitsthat are not normally removed, except for mainte-nance purposes or emergency evacuation, andflight deck-openable windows need not complywith paragraphs (c) and (f) of this section.

(3) Maintenance doors that meet the conditionsof paragraph (h) of this section, and for which aplacard is provided limiting use to maintenanceaccess, need not comply with paragraphs (c) and(f) of this section.

(h) Doors that are not a hazard. For the pur-poses of this section, a door is considered not tobe a hazard in the unlatched condition duringflight, provided it can be shown to meet all of thefollowing conditions:

(1) Doors in pressurized compartments wouldremain in the fully closed position if not restrainedby the latches when subject to a pressure greaterthan 1/2 psi. Opening by persons, either inadvert-ently or intentionally, need not be considered inmaking this determination.

(2) The door would remain inside the airplaneor remain attached to the airplane if it opens ei-ther in pressurized or unpressurized portions ofthe flight. This determination must include the

consideration of inadvertent and intentional open-ing by persons during either pressurized or un-pressurized portions of the flight.

(3) The disengagement of the latches duringflight would not allow depressurization of thecabin to an unsafe level. This safety assessmentmust include the physiological effects on the oc-cupants.

(4) The open door during flight would not createaerodynamic interference that could precludesafe flight and landing.

(5) The airplane would meet the structural de-sign requirements with the door open. This as-sessment must include the aeroelastic stability re-quirements of §25.629, as well as the strength re-quirements of subpart C of this part.

(6) The unlatching or opening of the door mustnot preclude safe flight and landing as a result ofinteraction with other systems or structures.


§25.785 Seats, berths, safety belts, and harnesses.(a) A seat (or berth for a nonambulant person)

must be provided for each occupant who hasreached his or her second birthday.

(b) Each seat, berth, safety belt, harness, andadjacent part of the airplane at each station des-ignated as occupiable during takeoff and landingmust be designed so that a person making properuse of these facilities will not suffer serious injuryin an emergency landing as a result of the inertiaforces specified in §§25.561 and 25.562.

(c) Each seat or berth must be approved.(d) Each occupant of a seat that makes more

than an 18-degree angle with the vertical planecontaining the airplane centerline must be pro-tected from head injury by a safety belt and an en-ergy absorbing rest that will support the arms,shoulders, head, and spine, or by a safety beltand shoulder harness that will prevent the headfrom contacting any injurious object. Each occu-pant of any other seat must be protected fromhead injury by a safety belt and, as appropriate tothe type, location, and angle of facing of eachseat, by one or more of the following:

(1) A shoulder harness that will prevent thehead from contacting any injurious object.

(2) The elimination of any injurious object withinstriking radius of the head.

(3) An energy absorbing rest that will supportthe arms, shoulders, head, and spine.

(e) Each berth must be designed so that theforward part has a padded end board, canvas dia-phragm, or equivalent means, that can withstandthe static load reaction of the occupant when sub-jected to the forward inertia force specified in§25.561. Berths must be free from corners and


74 ASA

protuberances likely to cause injury to a personoccupying the berth during emergency condi-tions.

(f) Each seat or berth, and its supporting struc-ture, and each safety belt or harness and its an-chorage must be designed for an occupant weightof 170 pounds, considering the maximum loadfactors, inertia forces, and reactions among theoccupant, seat, safety belt, and harness for eachrelevant flight and ground load condition (includ-ing the emergency landing conditions prescribedin §25.561). In addition—

(1) The structural analysis and testing of theseats, berths, and their supporting structures maybe determined by assuming that the critical loadin the forward, sideward, downward, upward, andrearward directions (as determined from the pre-scribed flight, ground, and emergency landingconditions) acts separately or using selectedcombinations of loads if the required strength ineach specified direction is substantiated. The for-ward load factor need not be applied to safetybelts for berths.

(2) Each pilot seat must be designed for the re-actions resulting from the application of the pilotforces prescribed in §25.395.

(3) The inertia forces specified in §25.561 mustbe multiplied by a factor of 1.33 (instead of the fit-ting factor prescribed in §25.625) in determiningthe strength of the attachment of each seat to thestructure and each belt or harness to the seat orstructure.

(g) Each seat at a flight deck station must havea restraint system consisting of a combined safetybelt and shoulder harness with a single-point re-lease that permits the flight deck occupant, whenseated with the restraint system fastened, to per-form all of the occupant’s necessary flight deckfunctions. There must be a means to secure eachcombined restraint system when not in use to pre-vent interference with the operation of the air-plane and with rapid egress in an emergency.

(h) Each seat located in the passenger com-partment and designated for use during takeoffand landing by a flight attendant required by theoperating rules of this chapter must be:

(1) Near a required floor level emergency exit,except that another location is acceptable if theemergency egress of passengers would be en-hanced with that location. A flight attendant seatmust be located adjacent to each Type A or Bemergency exit. Other flight attendant seats mustbe evenly distributed among the required floor-level emergency exits to the extent feasible.

(2) To the extent possible, without compromis-ing proximity to a required floor level emergencyexit, located to provide a direct view of the cabinarea for which the flight attendant is responsible.

(3) Positioned so that the seat will not interferewith the use of a passageway or exit when theseat is not in use.

(4) Located to minimize the probability that oc-cupants would suffer injury by being struck byitems dislodged from service areas, stowagecompartments, or service equipment.

(5) Either forward or rearward facing with anenergy absorbing rest that is designed to supportthe arms, shoulders, head, and spine.

(6) Equipped with a restraint system consistingof a combined safety belt and shoulder harnessunit with a single point release. There must bemeans to secure each restraint system when notin use to prevent interference with rapid egress inan emergency.

(i) Each safety belt must be equipped with ametal to metal latching device.

(j) If the seat backs do not provide a firm hand-hold, there must be a handgrip or rail along eachaisle to enable persons to steady themselveswhile using the aisles in moderately rough air.

(k) Each projecting object that would injure per-sons seated or moving about the airplane in nor-mal flight must be padded.

(l) Each forward observer’s seat required by theoperating rules must be shown to be suitable foruse in conducting the necessary enroute inspec-tion.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–72, 55 FR 29780, July 20, 1990;Amdt. 25–88, 61 FR 57956, Nov. 8, 1996]

§25.787 Stowage compartments.(a) Each compartment for the stowage of

cargo, baggage, carry-on articles, and equipment(such as life rafts), and any other stowage com-partment must be designed for its placarded max-imum weight of contents and for the critical loaddistribution at the appropriate maximum load fac-tors corresponding to the specified flight andground load conditions, and to the emergencylanding conditions of §25.561(b), except that theforces specified in the emergency landing condi-tions need not be applied to compartments lo-cated below, or forward, of all occupants in the air-plane. If the airplane has a passenger seatingconfiguration, excluding pilots seats, of 10 seatsor more, each stowage compartment in the pas-senger cabin, except for underseat and overheadcompartments for passenger convenience, mustbe completely enclosed.

(b) There must be a means to prevent the con-tents in the compartments from becoming a haz-ard by shifting, under the loads specified in para-graph (a) of this section. For stowage compart-ments in the passenger and crew cabin, if themeans used is a latched door, the design must


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take into consideration the wear and deteriorationexpected in service.

(c) If cargo compartment lamps are installed,each lamp must be installed so as to prevent con-tact between lamp bulb and cargo.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–32, 37 FR 3969, Feb. 24, 1972;Amdt. 25–38, 41 FR 55466, Dec. 20, 1976; Amdt. 25–51, 45 FR 7755, Feb. 4, 1980]

§25.789 Retention of items of mass in passenger and crew compartments and galleys.(a) Means must be provided to prevent each

item of mass (that is part of the airplane type de-sign) in a passenger or crew compartment or gal-ley from becoming a hazard by shifting under theappropriate maximum load factors correspondingto the specified flight and ground load conditions,and to the emergency landing conditions of§25.561(b).

(b) Each interphone restraint system must bedesigned so that when subjected to the load fac-tors specified in §25.561(b)(3), the interphone willremain in its stowed position.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–32, 37 FR 3969, Feb. 24, 1972;Amdt. 25–46, 43 FR 50596, Oct. 30, 1978]

§25.791 Passenger information signs and placards.(a) If smoking is to be prohibited, there must be

at least one placard so stating that is legible toeach person seated in the cabin. If smoking is tobe allowed, and if the crew compartment is sepa-rated from the passenger compartment, theremust be at least one sign notifying when smokingis prohibited. Signs which notify when smoking isprohibited must be operable by a member of theflightcrew and, when illuminated, must be legibleunder all probable conditions of cabin illuminationto each person seated in the cabin.

(b) Signs that notify when seat belts should befastened and that are installed to comply with theoperating rules of this chapter must be operableby a member of the flightcrew and, when illumi-nated, must be legible under all probable condi-tions of cabin illumination to each person seatedin the cabin.

(c) A placard must be located on or adjacent tothe door of each receptacle used for the disposalof flammable waste materials to indicate that useof the receptacle for disposal of cigarettes, etc., isprohibited.

(d) Lavatories must have “No Smoking” or “NoSmoking in Lavatory” placards conspicuously lo-cated on or adjacent to each side of the entrydoor.

(e) Symbols that clearly express the intent ofthe sign or placard may be used in lieu of letters.


§25.793 Floor surfaces.The floor surface of all areas which are likely to

become wet in service must have slip resistantproperties.


§25.795 Security considerations.(a) Protection of flightcrew compartment. If

a flightdeck door is required by operating rules:(1) The bulkhead, door, and any other accessi-

ble boundary separating the flightcrew compart-ment from occupied areas must be designed toresist forcible intrusion by unauthorized personsand be capable of withstanding impacts of 300joules (221.3 foot pounds).

(2) The bulkhead, door, and any other accessi-ble boundary separating the flightcrew compart-ment from occupied areas must be designed toresist a constant 250 pound (1,113 Newtons) ten-sile load on accessible handholds, including thedoorknob or handle.

(3) The bulkhead, door, and any other bound-ary separating the flightcrew compartment fromany occupied areas must be designed to resistpenetration by small arms fire and fragmentationdevices to a level equivalent to level IIIa of the Na-tional Institute of Justice (NIJ) Standard 0101.04.

(b) Airplanes with a maximum certificated pas-senger seating capacity of more than 60 personsor a maximum certificated takeoff gross weight ofover 100,000 pounds (45,359 Kilograms) must bedesigned to limit the effects of an explosive or in-cendiary device as follows:

(1) Flightdeck smoke protection. Means mustbe provided to limit entry of smoke, fumes, andnoxious gases into the flightdeck.

(2) Passenger cabin smoke protection. Meansmust be provided to prevent passenger incapaci-tation in the cabin resulting from smoke, fumes,and noxious gases as represented by the initialcombined volumetric concentrations of 0.59%carbon monoxide and 1.23% carbon dioxide.

(3) Cargo compartment fire suppression. Anextinguishing agent must be capable of suppress-ing a fire. All cargo-compartment fire suppressionsystems must be designed to withstand the fol-lowing effects, including support structure dis-placements or adjacent materials displacingagainst the distribution system:

(i) Impact or damage from a 0.5-inch diameteraluminum sphere traveling at 430 feet per second(131.1 meters per second);


76 ASA

(ii) A 15-pound per square-inch (103.4 kPa)pressure load if the projected surface area of thecomponent is greater than 4 square feet. Any sin-gle dimension greater than 4 feet (1.22 meters)may be assumed to be 4 feet (1.22 meters) inlength; and

(iii) A 6-inch (0.152 meters) displacement, ex-cept where limited by the fuselage contour, from asingle point force applied anywhere along the dis-tribution system where relative movement be-tween the system and its attachment can occur.

(iv) Paragraphs (b)(3)(i) through (iii) of this sec-tion do not apply to components that are redun-dant and separated in accordance with paragraph(c)(2) of this section or are installed remotely fromthe cargo compartment.

(c) An airplane with a maximum certificatedpassenger seating capacity of more than 60 per-sons or a maximum certificated takeoff grossweight of over 100,000 pounds (45,359 Kilo-grams) must comply with the following:

(1) Least risk bomb location. An airplane mustbe designed with a designated location where abomb or other explosive device could be placed tobest protect flight-critical structures and systemsfrom damage in the case of detonation.

(2) Survivability of systems.(i) Except where impracticable, redundant air-

plane systems necessary for continued safe flightand landing must be physically separated, at aminimum, by an amount equal to a sphere of di-ameter

(where H0 is defined under §25.365(e)(2) of thispart and D need not exceed 5.05 feet (1.54meters)). The sphere is applied everywhere withinthe fuselage—limited by the forward bulkheadand the aft bulkhead of the passenger cabin andcargo compartment beyond which only one-halfthe sphere is applied.

(ii) Where compliance with paragraph (c)(2)(i)of this section is impracticable, other design pre-cautions must be taken to maximize the surviv-ability of those systems.

(3) Interior design to facilitate searches. Designfeatures must be incorporated that will deter con-cealment or promote discovery of weapons, ex-plosives, or other objects from a simple inspectionin the following areas of the airplane cabin:

(i) Areas above the overhead bins must be de-signed to prevent objects from being hidden fromview in a simple search from the aisle. Designsthat prevent concealment of objects with volumes20 cubic inches and greater satisfy this require-ment.

(ii) Toilets must be designed to prevent the pas-sage of solid objects greater than 2.0 inches in di-ameter.

(iii) Life preservers or their storage locationsmust be designed so that tampering is evident.

(d) Exceptions. Airplanes used solely to trans-port cargo only need to meet the requirements ofparagraphs (b)(1), (b)(3), and (c)(2) of this sec-tion.

(e) Material Incorporated by Reference. Youmust use National Institute of Justice (NIJ) Stan-dard 0101.04, Ballistic Resistance of PersonalBody Armor, June 2001, Revision A, to establishballistic resistance as required by paragraph(a)(3) of this section.

(1) The Director of the Federal Register ap-proved the incorporation by reference of this doc-ument under 5 U.S.C. 552(a) and 1 CFR part 51.

(2) You may review copies of NIJ Standard0101.04 at the:

(i) FAA Transport Airplane Directorate, 1601Lind Avenue, SW., Renton, Washington 98055;

(ii) National Institute of Justice (NIJ),http://www.ojp.usdoj.gov/nij, telephone (202) 307-2942; or

(iii) National Archives and Records Administra-tion (NARA). For information on the availability ofthis material at NARA go to:http://www.archives.gov/federal_register/code_of_federal_regulations/ibr_locations.html or call (202) 741-6030.

(3) You may obtain copies of NIJ Standard0101.04 from the National Criminal Justice Refer-ence Service, P.O. Box 6000, Rockville, MD20849-6000, telephone (800) 851-3420.

[Docket No. FAA–2006–26722, 73 FR 63879, Oct. 28,2008]

EMERGENCY PROVISIONS

§25.801 Ditching.(a) If certification with ditching provisions is re-

quested, the airplane must meet the requirementsof this section and §§25.807(e), 25.1411, and25.1415(a).

(b) Each practicable design measure, compati-ble with the general characteristics of the air-plane, must be taken to minimize the probabilitythat in an emergency landing on water, the behav-ior of the airplane would cause immediate injuryto the occupants or would make it impossible forthem to escape.

(c) The probable behavior of the airplane in awater landing must be investigated by model testsor by comparison with airplanes of similar config-uration for which the ditching characteristics areknown. Scoops, flaps, projections, and any otherfactor likely to affect the hydrodynamic character-istics of the airplane, must be considered.

(d) It must be shown that, under reasonablyprobable water conditions, the flotation time andtrim of the airplane will allow the occupants toleave the airplane and enter the liferafts required

D 2 HO π⁄( )=


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by §25.1415. If compliance with this provision isshown by buoyancy and trim computations, ap-propriate allowances must be made for probablestructural damage and leakage. If the airplanehas fuel tanks (with fuel jettisoning provisions)that can reasonably be expected to withstand aditching without leakage, the jettisonable volumeof fuel may be considered as buoyancy volume.

(e) Unless the effects of the collapse of exter-nal doors and windows are accounted for in theinvestigation of the probable behavior of the air-plane in a water landing (as prescribed in para-graphs (c) and (d) of this section), the externaldoors and windows must be designed to with-stand the probable maximum local pressures.


§25.803 Emergency evacuation.(a) Each crew and passenger area must have

emergency means to allow rapid evacuation incrash landings, with the landing gear extended aswell as with the landing gear retracted, consider-ing the possibility of the airplane being on fire.

(b) [Reserved](c) For airplanes having a seating capacity of

more than 44 passengers, it must be shown thatthe maximum seating capacity, including the num-ber of crewmembers required by the operatingrules for which certification is requested, can beevacuated from the airplane to the ground undersimulated emergency conditions within 90 sec-onds. Compliance with this requirement must beshown by actual demonstration using the test cri-teria outlined in Appendix J of this part unless theAdministrator finds that a combination of analysisand testing will provide data equivalent to thatwhich would be obtained by actual demonstration.

(d) – (e) [Reserved]

[Docket No. 24344, 55 FR 29781, July 20, 1990]

§25.807 Emergency exits.(a) Type. For the purpose of this part, the types

of exits are defined as follows:(1) Type I. This type is a floor-level exit with a

rectangular opening of not less than 24 incheswide by 48 inches high, with corner radii notgreater than eight inches.

(2) Type II. This type is a rectangular opening ofnot less than 20 inches wide by 44 inches high,with corner radii not greater than seven inches.Type II exits must be floor-level exits unless lo-cated over the wing, in which case they must nothave a step-up inside the airplane of more than10 inches nor a step-down outside the airplane ofmore than 17 inches.

(3) Type III. This type is a rectangular openingof not less than 20 inches wide by 36 inches highwith corner radii not greater than seven inches,

and with a step-up inside the airplane of not morethan 20 inches. If the exit is located over the wing,the step-down outside the airplane may not ex-ceed 27 inches.

(4) Type IV. This type is a rectangular openingof not less than 19 inches wide by 26 inches high,with corner radii not greater than 6.3 inches, lo-cated over the wing, with a step-up inside the air-plane of not more than 29 inches and a step-downoutside the airplane of not more than 36 inches.

(5) Ventral. This type is an exit from the passen-ger compartment through the pressure shell andthe bottom fuselage skin. The dimensions andphysical configuration of this type of exit must al-low at least the same rate of egress as a Type Iexit with the airplane in the normal ground atti-tude, with landing gear extended.

(6) Tailcone. This type is an aft exit from thepassenger compartment through the pressureshell and through an openable cone of the fuse-lage aft of the pressure shell. The means of open-ing the tailcone must be simple and obvious andmust employ a single operation.

(7) Type A. This type is a floor-level exit with arectangular opening of not less than 42 incheswide by 72 inches high, with corner radii notgreater than seven inches.

(8) Type B. This type is a floor-level exit with arectangular opening of not less than 32 incheswide by 72 inches high, with corner radii notgreater than six inches.

(9) Type C. This type is a floor-level exit with arectangular opening of not less than 30 incheswide by 48 inches high, with corner radii notgreater than 10 inches.

(b) Step down distance. Step down distance,as used in this section, means the actual distancebetween the bottom of the required opening and ausable foot hold, extending out from the fuselage,that is large enough to be effective withoutsearching by sight or feel.

(c) Over-sized exits. Openings larger thanthose specified in this section, whether or not ofrectangular shape, may be used if the specifiedrectangular opening can be inscribed within theopening and the base of the inscribed rectangularopening meets the specified step-up and step-down heights.

(d) Asymmetry. Exits of an exit pair need not bediametrically opposite each other nor of the samesize; however, the number of passenger seatspermitted under paragraph (g) of this section isbased on the smaller of the two exits.

(e) Uniformity. Exits must be distributed as uni-formly as practical, taking into account passengerseat distribution.

(f) Location.(1) Each required passenger emergency exit

must be accessible to the passengers and lo-


78 ASA

cated where it will afford the most effective meansof passenger evacuation.

(2) If only one floor-level exit per side is pre-scribed, and the airplane does not have a tail-cone or ventral emergency exit, the floor-level ex-its must be in the rearward part of the passengercompartment unless another location affords amore effective means of passenger evacuation.

(3) If more than one floor-level exit per side isprescribed, and the airplane does not have acombination cargo and passenger configuration,at least one floor-level exit must be located ineach side near each end of the cabin.

(4) For an airplane that is required to havemore than one passenger emergency exit foreach side of the fuselage, no passenger emer-gency exit shall be more than 60 feet from any ad-jacent passenger emergency exit on the sameside of the same deck of the fuselage, as mea-sured parallel to the airplane’s longitudinal axisbetween the nearest exit edges.

(g) Type and number required. The maximumnumber of passenger seats permitted depends onthe type and number of exits installed in each sideof the fuselage. Except as further restricted inparagraphs (g)(1) through (g)(9) of this section,the maximum number of passenger seats permit-ted for each exit of a specific type installed in eachside of the fuselage is as follows:

Type A 110

Type B 75

Type C 55

Type I 45

Type II 40

Type III 35

Type IV 9

(1) For a passenger seating configuration of 1to 9 seats, there must be at least one Type IV orlarger overwing exit in each side of the fuselageor, if overwing exits are not provided, at least oneexit in each side that meets the minimum dimen-sions of a Type III exit.

(2) For a passenger seating configuration ofmore than 9 seats, each exit must be a Type III orlarger exit.

(3) For a passenger seating configuration of 10to 19 seats, there must be at least one Type III orlarger exit in each side of the fuselage.

(4) For a passenger seating configuration of 20to 40 seats, there must be at least two exits, oneof which must be a Type II or larger exit, in eachside of the fuselage.

(5) For a passenger seating configuration of 41to 110 seats, there must be at least two exits, oneof which must be a Type I or larger exit, in eachside of the fuselage.

(6) For a passenger seating configuration ofmore than 110 seats, the emergency exits in eachside of the fuselage must include at least two TypeI or larger exits.

(7) The combined maximum number of pas-senger seats permitted for all Type III exits is 70,and the combined maximum number of passen-ger seats permitted for two Type III exits in eachside of the fuselage that are separated by fewerthan three passenger seat rows is 65.

(8) If a Type A, Type B, or Type C exit is in-stalled, there must be at least two Type C or largerexits in each side of the fuselage.

(9) If a passenger ventral or tail cone exit is in-stalled and that exit provides at least the samerate of egress as a Type III exit with the airplane inthe most adverse exit opening condition thatwould result from the collapse of one or more legsof the landing gear, an increase in the passengerseating configuration is permitted as follows:

(i) For a ventral exit, 12 additional passengerseats.

(ii) For a tail cone exit incorporating a floor levelopening of not less than 20 inches wide by 60inches high, with corner radii not greater thanseven inches, in the pressure shell and incorpo-rating an approved assist means in accordancewith §25.810(a), 25 additional passenger seats.

(iii) For a tail cone exit incorporating an openingin the pressure shell which is at least equivalent toa Type III emergency exit with respect to dimen-sions, step-up and step-down distance, and withthe top of the opening not less than 56 inchesfrom the passenger compartment floor, 15 addi-tional passenger seats.

(h) Other exits. The following exits also mustmeet the applicable emergency exit requirementsof §§25.809 through 25.812, and must be readilyaccessible:

(1) Each emergency exit in the passenger com-partment in excess of the minimum number of re-quired emergency exits.

(2) Any other floor-level door or exit that is ac-cessible from the passenger compartment and isas large or larger than a Type II exit, but less than46 inches wide.

(3) Any other ventral or tail cone passengerexit.

(i) Ditching emergency exits for passengers.Whether or not ditching certification is requested,ditching emergency exits must be provided in ac-cordance with the following requirements, unlessthe emergency exits required by paragraph (g) ofthis section already meet them:

(1) For airplanes that have a passenger seatingconfiguration of nine or fewer seats, excludingpilot seats, one exit above the waterline in eachside of the airplane, meeting at least the dimen-sions of a Type IV exit.


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(2) For airplanes that have a passenger seatingconfiguration of 10 of more seats, excluding pilotseats, one exit above the waterline in a side of theairplane, meeting at least the dimensions of aType III exit for each unit (or part of a unit) of 35passenger seats, but no less than two such exitsin the passenger cabin, with one on each side ofthe airplane. The passenger seat/exit ratio may beincreased through the use of larger exits, or othermeans, provided it is shown that the evacuationcapability during ditching has been improved ac-cordingly.

(3) If it is impractical to locate side exits abovethe waterline, the side exits must be replaced byan equal number of readily accessible overheadhatches of not less than the dimensions of a TypeIII exit, except that for airplanes with a passengerconfiguration of 35 or fewer seats, excluding pilotseats, the two required Type III side exits need bereplaced by only one overhead hatch.

(j) Flightcrew emergency exits. For airplanes inwhich the proximity of passenger emergency exitsto the flightcrew area does not offer a convenientand readily accessible means of evacuation of theflightcrew, and for all airplanes having a passengerseating capacity greater than 20, flightcrew exitsshall be located in the flightcrew area. Such exitsshall be of sufficient size and so located as to per-mit rapid evacuation by the crew. One exit shall beprovided on each side of the airplane; or, alterna-tively, a top hatch shall be provided. Each exit mustencompass an unobstructed rectangular openingof at least 19 by 20 inches unless satisfactory exitutility can be demonstrated by a typical crewmem-ber.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–72, 55 FR 29781, July 20, 1990;Amdt. 25–88, 61 FR 57957, Nov. 8, 1996; Amdt. 25–88,62 FR 1817, Jan. 13, 1997; Amdt. 25–94, 63 FR 8848,Feb. 23, 1998; Amdt. 25–94, 63 FR 12862, March 16,1998; Amdt. 25–114, 69 FR 24501, May 3, 2004]

§25.809 Emergency exit arrangement.(a) Each emergency exit, including each flight-

crew emergency exit, must be a moveable door orhatch in the external walls of the fuselage, allow-ing an unobstructed opening to the outside. In ad-dition, each emergency exit must have means topermit viewing of the conditions outside the exitwhen the exit is closed. The viewing means maybe on or adjacent to the exit provided no obstruc-tions exist between the exit and the viewingmeans. Means must also be provided to permitviewing of the likely areas of evacuee ground con-tact. The likely areas of evacuee ground contactmust be viewable during all lighting conditionswith the landing gear extended as well as in allconditions of landing gear collapse.

(b) Each emergency exit must be openablefrom the inside and the outside except that sliding

window emergency exits in the flight crew areaneed not be openable from the outside if other ap-proved exits are convenient and readily accessi-ble to the flight crew area. Each emergency exitmust be capable of being opened, when there isno fuselage deformation—

(1) With the airplane in the normal ground atti-tude and in each of the attitudes corresponding tocollapse of one or more legs of the landing gear;and

(2) Within 10 seconds measured from the timewhen the opening means is actuated to the timewhen the exit is fully opened.

(3) Even though persons may be crowdedagainst the door on the inside of the airplane.

(c) The means of opening emergency exitsmust be simple and obvious; may not require ex-ceptional effort; and must be arranged andmarked so that it can be readily located and oper-ated, even in darkness. Internal exit-openingmeans involving sequence operations (such asoperation of two handles or latches, or the releaseof safety catches) may be used for flightcrewemergency exits if it can be reasonably estab-lished that these means are simple and obviousto crewmembers trained in their use.

(d) If a single power-boost or single power-oper-ated system is the primary system for operatingmore than one exit in an emergency, each exitmust be capable of meeting the requirements ofparagraph (b) of this section in the event of failureof the primary system. Manual operation of the exit(after failure of the primary system) is acceptable.

(e) Each emergency exit must be shown bytests, or by a combination of analysis and tests, tomeet the requirements of paragraphs (b) and (c)of this section.

(f) Each door must be located where personsusing them will not be endangered by the propel-lers when appropriate operating procedures areused.

(g) There must be provisions to minimize theprobability of jamming of the emergency exits re-sulting from fuselage deformation in a minorcrash landing.

(h) When required by the operating rules forany large passenger-carrying turbojet-poweredairplane, each ventral exit and tailcone exit mustbe—

(1) Designed and constructed so that it cannotbe opened during flight; and

(2) Marked with a placard readable from a dis-tance of 30 inches and installed at a conspicuouslocation near the means of opening the exit, stat-ing that the exit has been designed and con-structed so that it cannot be opened during flight.

(i) Each emergency exit must have a means toretain the exit in the open position, once the exit isopened in an emergency. The means must not re-quire separate action to engage when the exit is


80 ASA

opened, and must require positive action to disen-gage.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–15, 32 FR 13264, Sept. 20, 1967;Amdt. 25–32, 37 FR 3970, Feb. 24, 1972; Amdt. 25–34,37 FR 25355, Nov. 30, 1972; Amdt. 25–46, 43 FR50597, Oct. 30, 1978; Amdt. 25–47, 44 FR 61325, Oct.25, 1979; Amdt. 25–72, 55 FR 29782, July 20, 1990;Amdt. 25–114, 69 FR 24501, May 3, 2004; Amdt. 25–116, 69 FR 62788, Oct. 27, 2004]

§25.810 Emergency egress assist means and escape routes.(a) Each non over-wing Type A, Type B or Type

C exit, and any other non over-wing landplaneemergency exit more than 6 feet from the groundwith the airplane on the ground and the landinggear extended, must have an approved means toassist the occupants in descending to the ground.

(1) The assisting means for each passengeremergency exit must be a self-supporting slide orequivalent; and, in the case of a Type A or Type Bexits, it must be capable of carrying simulta-neously two parallel lines of evacuees. In addition,the assisting means must be designed to meetthe following requirements—

(i) It must be automatically deployed and de-ployment must begin during the interval betweenthe time the exit opening means is actuated frominside the airplane and the time the exit is fullyopened. However, each passenger emergencyexit which is also a passenger entrance door or aservice door must be provided with means to pre-vent deployment of the assisting means when it isopened from either the inside or the outside undernonemergency conditions for normal use.

(ii) Except for assisting means installed at TypeC exits, it must be automatically erected within 6seconds after deployment is begun. Assistingmeans installed at Type C exits must be automati-cally erected within 10 seconds from the time theopening means of the exit is actuated.

(iii) It must be of such length after full deploy-ment that the lower end is self-supporting on theground and provides safe evacuation of occu-pants to the ground after collapse of one or morelegs of the landing gear.

(iv) It must have the capability, in 25-knot windsdirected from the most critical angle, to deployand, with the assistance of only one person, to re-main usable after full deployment to evacuate oc-cupants safely to the ground.

(v) For each system installation (mockup or air-plane installed), five consecutive deployment andinflation tests must be conducted (per exit) with-out failure, and at least three tests of each suchfive-test series must be conducted using a singlerepresentative sample of the device. The sampledevices must be deployed and inflated by the sys-tem’s primary means after being subjected to the

inertia forces specified in §25.561(b). If any partof the system fails or does not function properlyduring the required tests, the cause of the failureor malfunction must be corrected by positivemeans and after that, the full series of five con-secutive deployment and inflation tests must beconducted without failure.

(2) The assisting means for flightcrew emer-gency exits may be a rope or any other meansdemonstrated to be suitable for the purpose. If theassisting means is a rope, or an approved deviceequivalent to a rope, it must be—

(i) Attached to the fuselage structure at orabove the top of the emergency exit opening, or,for a device at a pilot’s emergency exit window, atanother approved location if the stowed device, orits attachment, would reduce the pilot’s view inflight;

(ii) Able (with its attachment) to withstand a400-pound static load.

(b) Assist means from the cabin to the wing arerequired for each Type A or Type B exit locatedabove the wing and having a stepdown unless theexit without an assist-means can be shown tohave a rate of passenger egress at least equal tothat of the same type of nonoverwing exit. If anassist means is required, it must be automaticallydeployed and automatically erected concurrentwith the opening of the exit. In the case of assistmeans installed at Type C exits, it must be self-supporting within 10 seconds from the time theopening means of the exits is actuated. For allother exit types, it must be self-supporting 6 sec-onds after deployment is begun.

(c) An escape route must be established fromeach overwing emergency exit, and (except forflap surfaces suitable as slides) covered with aslip resistant surface. Except where a means forchanneling the flow of evacuees is provided—

(1) The escape route from each Type A or TypeB passenger emergency exit, or any common es-cape route from two Type III passenger emer-gency exits, must be at least 42 inches wide; thatfrom any other passenger emergency exit mustbe at least 24 inches wide; and

(2) The escape route surface must have a re-flectance of at least 80 percent, and must be de-fined by markings with a surface-to-marking con-trast ratio of at least 5:1.

(d) Means must be provided to assist evacueesto reach the ground for all Type C exits locatedover the wing and, if the place on the airplanestructure at which the escape route required inparagraph (c) of this section terminates is morethan 6 feet from the ground with the airplane onthe ground and the landing gear extended, for allother exit types.

(1) If the escape route is over the flap, theheight of the terminal edge must be measured


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with the flap in the takeoff or landing position,whichever is higher from the ground.

(2) The assisting means must be usable andself-supporting with one or more landing gearlegs collapsed and under a 25-knot wind directedfrom the most critical angle.

(3) The assisting means provided for each es-cape route leading from a Type A or B emergencyexit must be capable of carrying simultaneouslytwo parallel lines of evacuees; and, the assistingmeans leading from any other exit type must becapable of carrying as many parallel lines of evac-uees as there are required escape routes.

(4) The assisting means provided for each es-cape route leading from a Type C exit must be au-tomatically erected within 10 seconds from thetime the opening means of the exit is actuated,and that provided for the escape route leadingfrom any other exit type must be automaticallyerected within 10 seconds after actuation of theerection system.

(e) If an integral stair is installed in a passengerentry door that is qualified as a passenger emer-gency exit, the stair must be designed so that, un-der the following conditions, the effectiveness ofpassenger emergency egress will not be im-paired:

(1) The door, integral stair, and operatingmechanism have been subjected to the inertiaforces specified in §25.561(b)(3), acting sepa-rately relative to the surrounding structure.

(2) The airplane is in the normal ground attitudeand in each of the attitudes corresponding to col-lapse of one or more legs of the landing gear.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–72, 55 FR 29782, July 20, 1990;Amdt. 25–88, 61 FR 57958, Nov. 8, 1996; Amdt. 25–88,62 FR 1817, Jan. 13, 1997; Amdt. 25–114, 69 FR 24501,May 3, 2004]

§25.811 Emergency exit marking.(a) Each passenger emergency exit, its means

of access, and its means of opening must be con-spicuously marked.

(b) The identity and location of each passengeremergency exit must be recognizable from a dis-tance equal to the width of the cabin.

(c) Means must be provided to assist the occu-pants in locating the exits in conditions of densesmoke.

(d) The location of each passenger emergencyexit must be indicated by a sign visible to occu-pants approaching along the main passengeraisle (or aisles). There must be—

(1) A passenger emergency exit locator signabove the aisle (or aisles) near each passengeremergency exit, or at another overhead location ifit is more practical because of low headroom, ex-cept that one sign may serve more than one exit ifeach exit can be seen readily from the sign;

(2) A passenger emergency exit marking signnext to each passenger emergency exit, exceptthat one sign may serve two such exits if they bothcan be seen readily from the sign; and

(3) A sign on each bulkhead or divider that pre-vents fore and aft vision along the passengercabin to indicate emergency exits beyond and ob-scured by the bulkhead or divider, except that ifthis is not possible the sign may be placed at an-other appropriate location.

(e) The location of the operating handle and in-structions for opening exits from the inside of theairplane must be shown in the following manner—

(1) Each passenger emergency exit must have,on or near the exit, a marking that is readablefrom a distance of 30 inches.

(2) Each Type A, Type B, Type C or Type I pas-senger emergency exit operating handle must—

(i) Be self-illuminated with an initial brightnessof at least 160 microlamberts; or

(ii) Be conspicuously located and well illumi-nated by the emergency lighting even in condi-tions of occupant crowding at the exit.

(3) [Reserved](4) Each Type A, Type B, Type C, Type I, or Type

II passenger emergency exit with a locking mech-anism released by rotary motion of the handlemust be marked—

(i) With a red arrow, with a shaft at least three-fourths of an inch wide and a head twice the widthof the shaft, extending along at least 70 degreesof arc at a radius approximately equal to three-fourths of the handle length.

(ii) So that the centerline of the exit handle iswithin ±1 inch of the projected point of the arrowwhen the handle has reached full travel and hasreleased the locking mechanism, and

(iii) With the word “open” in red letters 1 inchhigh, placed horizontally near the head of the ar-row.

(f) Each emergency exit that is required to beopenable from the outside, and its means ofopening, must be marked on the outside of theairplane. In addition, the following apply:

(1) The outside marking for each passengeremergency exit in the side of the fuselage mustinclude a 2-inch colored band outlining the exit.

(2) Each outside marking including the band,must have color contrast to be readily distinguish-able from the surrounding fuselage surface. Thecontrast must be such that if the reflectance of thedarker color is 15 percent or less, the reflectanceof the lighter color must be at least 45 percent.“Reflectance” is the ratio of the luminous flux re-flected by a body to the luminous flux it receives.When the reflectance of the darker color isgreater than 15 percent, at least a 30-percent dif-ference between its reflectance and the reflec-tance of the lighter color must be provided.


82 ASA

(3) In the case of exists other than those in theside of the fuselage, such as ventral or tail coneexists, the external means of opening, includinginstructions if applicable, must be conspicuouslymarked in red, or bright chrome yellow if the back-ground color is such that red is inconspicuous.When the opening means is located on only oneside of the fuselage, a conspicuous marking tothat effect must be provided on the other side.

(g) Each sign required by paragraph (d) of thissection may use the word “exit” in its legend inplace of the term “emergency exit.”

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–15, 32 FR 13264, Sept. 20, 1967;Amdt. 25–32, 37 FR 3970, Feb. 24, 1972; Amdt. 25–46,43 FR 50597, Oct. 30, 1978; 43 FR 52495, Nov. 13,1978; Amdt. 25–79, 58 FR 45229, Aug. 26, 1993; Amdt.25–88, 61 FR 57958, Nov. 8, 1996]

§25.812 Emergency lighting.(a) An emergency lighting system, independent

of the main lighting system, must be installed.However, the sources of general cabin illumina-tion may be common to both the emergency andthe main lighting systems if the power supply tothe emergency lighting system is independent ofthe power supply to the main lighting system. Theemergency lighting system must include:

(1) Illuminated emergency exit marking and lo-cating signs, sources of general cabin illumina-tion, interior lighting in emergency exit areas, andfloor proximity escape path marking.

(2) Exterior emergency lighting.(b) Emergency exit signs—(1) For airplanes that have a passenger seating

configuration, excluding pilot seats, of 10 seats ormore must meet the following requirements:

(i) Each passenger emergency exit locatorsign required by §25.811(d)(1) and each passen-ger emergency exit marking sign required by§25.811(d)(2) must have red letters at least 11⁄2inches high on an illuminated white background,and must have an area of at least 21 squareinches excluding the letters. The lighted back-ground-to-letter contrast must be at least 10:1.The letter height to stroke-width ratio may not bemore than 7:1 nor less than 6:1. These signsmust be internally electrically illuminated with abackground brightness of at least 25 foot-lam-berts and a high-to-low background contrast nogreater than 3:1.

(ii) Each passenger emergency exit sign re-quired by §25.811(d)(3) must have red letters atleast 11⁄2 inches high on a white background hav-ing an area of at least 21 square inches excludingthe letters. These signs must be internally electri-cally illuminated or self-illuminated by other thanelectrical means and must have an initial bright-ness of at least 400 microlamberts. The colors

may be reversed in the case of a sign that is self-illuminated by other than electrical means.

(2) For airplanes that have a passenger seatingconfiguration, excluding pilot seats, of nine seatsor less, that are required by §25.811(d) (1), (2),and (3) must have red letters at least 1 inch highon a white background at least 2 inches high.These signs may be internally electrically illumi-nated, or self-illuminated by other than electricalmeans, with an initial brightness of at least 160microlamberts. The colors may be reversed in thecase of a sign that is self-illuminated by other thanelectrical means.

(c) General illumination in the passenger cabinmust be provided so that when measured alongthe centerline of main passenger aisle(s), andcross aisle(s) between main aisles, at seat arm-rest height and at 40-inch intervals, the average il-lumination is not less than 0.05 foot-candle andthe illumination at each 40-inch interval is not lessthan 0.01 foot-candle. A main passenger aisle(s)is considered to extend along the fuselage fromthe most forward passenger emergency exit orcabin occupant seat, whichever is farther forward,to the most rearward passenger emergency exitor cabin occupant seat, whichever is farther aft.

(d) The floor of the passageway leading to eachfloor-level passenger emergency exit, betweenthe main aisles and the exit openings, must beprovided with illumination that is not less than0.02 foot-candle measured along a line that iswithin 6 inches of and parallel to the floor and iscentered on the passenger evacuation path.

(e) Floor proximity emergency escape pathmarking must provide emergency evacuationguidance for passengers when all sources of illu-mination more than 4 feet above the cabin aislefloor are totally obscured. In the dark of the night,the floor proximity emergency escape path mark-ing must enable each passenger to—

(1) After leaving the passenger seat, visuallyidentify the emergency escape path along thecabin aisle floor to the first exits or pair of exits for-ward and aft of the seat; and

(2) Readily identify each exit from the emer-gency escape path by reference only to markingsand visual features not more than 4 feet above thecabin floor.

(f) Except for subsystems provided in accor-dance with paragraph (h) of this section that serveno more than one assist means, are independentof the airplane’s main emergency lighting system,and are automatically activated when the assistmeans is erected, the emergency lighting systemmust be designed as follows.

(1) The lights must be operable manually fromthe flight crew station and from a point in the pas-senger compartment that is readily accessible toa normal flight attendant seat.


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(2) There must be a flight crew warning lightwhich illuminates when power is on in the airplaneand the emergency lighting control device is notarmed.

(3) The cockpit control device must have an“on,” “off,” and “armed” position so that whenarmed in the cockpit or turned on at either thecockpit or flight attendant station the lights will ei-ther light or remain lighted upon interruption (ex-cept an interruption caused by a transverse verti-cal separation of the fuselage during crash land-ing) of the airplane’s normal electric power. Theremust be a means to safeguard against inadvert-ent operation of the control device from the“armed” or “on” positions.

(g) Exterior emergency lighting must be pro-vided as follows:

(1) At each overwing emergency exit the illumi-nation must be—

(i) Not less than 0.03 foot-candle (measurednormal to the direction of the incident light) on a 2-square-foot area where an evacuee is likely tomake his first step outside the cabin;

(ii) Not less than 0.05 foot-candle (measurednormal to the direction of the incident light) for aminimum width of 42 inches for a Type A overwingemergency exit and two feet for all other overwingemergency exits along the 30 percent of the slip-resistant portion of the escape route required in§25.810(c) that is farthest from the exit; and

(iii) Not less than 0.03 foot-candle on theground surface with the landing gear extended(measured normal to the direction of the incidentlight) where an evacuee using the established es-cape route would normally make first contact withthe ground.

(2) At each non-overwing emergency exit notrequired by §25.810(a) to have descent assistmeans the illumination must be not less than 0.03foot-candle (measured normal to the direction ofthe incident light) on the ground surface with thelanding gear extended where an evacuee is likelyto make first contact with the ground outside thecabin.

(h) The means required in §25.810(a)(1) and(d) to assist the occupants in descending to theground must be illuminated so that the erectedassist means is visible from the airplane.

(1) If the assist means is illuminated by exterioremergency lighting, it must provide illumination ofnot less than 0.03 foot-candle (measured normalto the direction of the incident light) at the groundend of the erected assist means where an evac-uee using the established escape route wouldnormally make first contact with the ground, withthe airplane in each of the attitudes correspond-ing to the collapse of one or more legs of the land-ing gear.

(2) If the emergency lighting subsystem illumi-nating the assist means serves no other assist

means, is independent of the airplane’s mainemergency lighting system, and is automaticallyactivated when the assist means is erected, thelighting provisions—

(i) May not be adversely affected by stowage;and

(ii) Must provide illumination of not less than0.03 foot-candle (measured normal to the direc-tion of incident light) at the ground and of theerected assist means where an evacuee wouldnormally make first contact with the ground, withthe airplane in each of the attitudes correspond-ing to the collapse of one or more legs of the land-ing gear.

(i) The energy supply to each emergency light-ing unit must provide the required level of illumi-nation for at least 10 minutes at the critical ambi-ent conditions after emergency landing.

(j) If storage batteries are used as the energysupply for the emergency lighting system, theymay be recharged from the airplane’s main elec-tric power system: Provided, That, the chargingcircuit is designed to preclude inadvertent batterydischarge into charging circuit faults.

(k) Components of the emergency lighting sys-tem, including batteries, wiring relays, lamps, andswitches must be capable of normal operation af-ter having been subjected to the inertia forceslisted in §25.561(b).

(l) The emergency lighting system must be de-signed so that after any single transverse verticalseparation of the fuselage during crash landing—

(1) Not more than 25 percent of all electricallyilluminated emergency lights required by this sec-tion are rendered inoperative, in addition to thelights that are directly damaged by the separation;

(2) Each electrically illuminated exit sign re-quired under §25.811(d)(2) remains operative ex-clusive of those that are directly damaged by theseparation; and

(3) At least one required exterior emergencylight for each side of the airplane remains opera-tive exclusive of those that are directly damagedby the separation.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–15, 32 FR 13265, Sept. 20, 1967;Amdt. 25–28, 36 FR 16899, Aug. 26, 1971; Amdt. 25–32,37 FR 3971, Feb. 24, 1972; Amdt. 25–46, 43 FR 50597,Oct. 30, 1978; Amdt. 25–58, 49 FR 43186, Oct. 26,1984; Amdt. 25–88, 61 FR 57958, Nov. 8, 1996; Amdt.25–116, 69 FR 62788, Oct. 27, 2004; Amdt. 25–128, 74FR 25645, May 29, 2009]


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§25.813 Emergency exit access.Each required emergency exit must be acces-

sible to the passengers and located where it willafford an effective means of evacuation. Emer-gency exit distribution must be as uniform aspractical, taking passenger distribution into ac-count; however, the size and location of exits onboth sides of the cabin need not be symmetrical.If only one floor level exit per side is prescribed,and the airplane does not have a tail cone or ven-tral emergency exit, the floor level exit must be inthe rearward part of the passenger compartment,unless another location affords a more effectivemeans of passenger evacuation. Where morethan one floor level exit per side is prescribed, atleast one floor level exit per side must be locatednear each end of the cabin, except that this provi-sion does not apply to combination cargo/passen-ger configurations. In addition—

(a) There must be a passageway leading fromthe nearest main aisle to each Type A, Type B,Type C, Type I, or Type II emergency exit and be-tween individual passenger areas. Each passage-way leading to a Type A or Type B exit must be un-obstructed and at least 36 inches wide. Passage-ways between individual passenger areas andthose leading to Type I, Type II, or Type C emer-gency exits must be unobstructed and at least 20inches wide. Unless there are two or more mainaisles, each Type A or B exit must be located sothat there is passenger flow along the main aisleto that exit from both the forward and aft direc-tions. If two or more main aisles are provided,there must be unobstructed cross-aisles at least20 inches wide between main aisles. There mustbe—

(1) A cross-aisle which leads directly to eachpassageway between the nearest main aisle anda Type A or B exit; and

(2) A cross-aisle which leads to the immediatevicinity of each passageway between the nearestmain aisle and a Type I, Type II, or Type III exit; ex-cept that when two Type III exits are located withinthree passenger rows of each other, a singlecross-aisle may be used if it leads to the vicinitybetween the passageways from the nearest mainaisle to each exit.

(b) Adequate space to allow crewmember(s) toassist in the evacuation of passengers must beprovided as follows:

(1) Each assist space must be a rectangle onthe floor, of sufficient size to enable a crewmem-ber, standing erect, to effectively assist evacuees.The assist space must not reduce the unob-structed width of the passageway below that re-quired for the exit.

(2) For each Type A or B exit, assist space mustbe provided at each side of the exit regardless of

whether an assist means is required by§25.810(a).

(3) For each Type C, I or II exit installed in anairplane with seating for more than 80 passen-gers, an assist space must be provided at oneside of the passageway regardless of whether anassist means is required by §25.810(a).

(4) For each Type C, I or II exit, an assist spacemust be provided at one side of the passageway ifan assist means is required by §25.810(a).

(5) For any tailcone exit that qualifies for 25 ad-ditional passenger seats under the provisions of§25.807(g)(9)(ii), an assist space must be pro-vided, if an assist means is required by§25.810(a).

(6) There must be a handle, or handles, at eachassist space, located to enable the crewmemberto steady himself or herself:

(i) While manually activating the assist means(where applicable) and,

(ii) While assisting passengers during an evac-uation.

(c) The following must be provided for eachType III or Type IV exit—

(1) There must be access from the nearest toeach exit. In addition, for each Type III exit in anairplane that has a passenger seating configura-tion of 60 or more—

(i) Except as provided in paragraph (c)(1)(ii),the access must be provided by an unobstructedpassageway that is at least 10 inches in width forinterior arrangements in which the adjacent seatrows on the exit side of the aisle contain no morethan two seats, or 20 inches in width for interiorarrangements in which those rows contain threeseats. The width of the passageway must be mea-sured with adjacent seats adjusted to their mostadverse position. The centerline of the requiredpassageway width must not be displaced morethan 5 inches horizontally from that of the exit.

(ii) In lieu of one 10- or 20-inch passageway,there may be two passageways, between seatrows only, that must be at least 6 inches in widthand lead to an unobstructed space adjacent toeach exit. (Adjacent exits must not share a com-mon passageway.) The width of the passagewaysmust be measured with adjacent seats adjustedto their most adverse position. The unobstructedspace adjacent to the exit must extend verticallyfrom the floor to the ceiling (or bottom of sidewallstowage bins), inboard from the exit for a distancenot less than the width of the narrowest passen-ger seat installed on the airplane, and from theforward edge of the forward passageway to the aftedge of the aft passageway. The exit openingmust be totally within the fore and aft bounds ofthe unobstructed space.

(2) In addition to the access—(i) For airplanes that have a passenger seating

configuration of 20 or more, the projected opening


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of the exit provided must be obstructed and theremust be no interference in opening the exit byseats, berths, or other protrusions (including anyseatback in the most adverse position) for a dis-tance from that exit not less than the width of thenarrowest passenger seat installed on the air-plane.

(ii) For airplanes that have a passenger seatingconfiguration of 19 or fewer, there may be minorobstructions in this region, if there are compen-sating factors to maintain the effectiveness of theexit.

(3) For each Type III exit, regardless of the pas-senger capacity of the airplane in which it is in-stalled, there must be placards that—

(i) Are readable by all persons seated adjacentto and facing a passageway to the exit;

(ii) Accurately state or illustrate that propermethod of opening the exit including the use ofhandholds; and

(iii) If the exit is a removable hatch, state theweight of the hatch and indicate an appropriate lo-cation to place the hatch after removal.

(d) If it is necessary to pass through a pas-sageway between passenger compartments toreach any required emergency exit from any seatin the passenger cabin, the passageway must beunobstructed. However, curtains may be used ifthey allow free entry through the passageway.

(e) No door may be installed between any pas-senger seat that is occupiable for takeoff andlanding and any passenger emergency exit, suchthat the door crosses any egress path (includingaisles, crossaisles and passageways).

(f) If it is necessary to pass through a doorwayseparating any crewmember seat (except thoseseats on the flightdeck), occupiable for takeoffand landing, from any emergency exit, the doormust have a means to latch it in the open position.The latching means must be able to withstand theloads imposed upon it when the door is subjectedto the ultimate inertia forces, relative to the sur-rounding structure, listed in §25.561(b).

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–1, 30 FR 3204, March 9, 1965;Amdt. 25–15, 32 FR 13265, Sept. 20, 1967; Amdt. 25–32, 37 FR 3971, Feb. 24, 1972; Amdt. 25–46, 43 FR50597, Oct. 30, 1978; Amdt. 25–72, 55 FR 29783, July20, 1990; Amdt. 25–76, 57 FR 19244, May 4, 1992;Amdt. 25–76, 57 FR 29120, June 30, 1992; Amdt. 25–88, 61 FR 57958, Nov. 8, 1996; Amdt. 25–116, 69 FR62788, Oct. 27, 2004; Amdt. 25–128, 74 FR 25645, May29, 2009]

§25.815 Width of aisle.The passenger aisle width at any point be-

tween seats must equal or exceed the values inthe following table:

1A narrower width not less than 9 inches maybe approved when substantiated by tests foundnecessary by the Administrator.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–15, 32 FR 13265, Sept. 20, 1967;Amdt. 25–38, 41 FR 55466, Dec. 20, 1976]

§25.817 Maximum number of seats abreast.On airplanes having only one passenger aisle,

no more than three seats abreast may be placedon each side of the aisle in any one row.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–15, 32 FR 13265, Sept. 20, 1967]

§25.819 Lower deck service compartments (including galleys).For airplanes with a service compartment lo-

cated below the main deck, which may be occu-pied during taxi or flight but not during takeoff orlanding, the following apply:

(a) There must be at least two emergencyevacuation routes, one at each end of each lowerdeck service compartment or two having suffi-cient separation within each compartment, whichcould be used by each occupant of the lower deckservice compartment to rapidly evacuate to themain deck under normal and emergency lightingconditions. The routes must provide for the evacu-ation of incapacitated persons, with assistance.The use of the evacuation routes may not be de-pendent on any powered device. The routes mustbe designed to minimize the possibility of block-age which might result from fire, mechanical orstructural failure, or persons standing on top of oragainst the escape routes. In the event the air-plane’s main power system or compartment mainlighting system should fail, emergency illumina-tion for each lower deck service compartmentmust be automatically provided.

(b) There must be a means for two-way voicecommunication between the flight deck and eachlower deck service compartment, which remainsavailable following loss of normal electrical powergenerating system.

Passenger seating capacity

Minimum passenger aisle width (inches)

Less than 25 in. from floor

25 in. and more from floor

10 or less 112 15

11 through 19 12 20

20 or more 15 20


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(c) There must be an aural emergency alarmsystem, audible during normal and emergencyconditions, to enable crewmembers on the flightdeck and at each required floor level emergencyexit to alert occupants of each lower deck servicecompartment of an emergency situation.

(d) There must be a means, readily detectableby occupants of each lower deck service com-partment, that indicates when seat belts shouldbe fastened.

(e) If a public address system is installed in theairplane, speakers must be provided in eachlower deck service compartment.

(f) For each occupant permitted in a lower deckservice compartment, there must be a forward oraft facing seat which meets the requirements of§25.785(d), and must be able to withstand maxi-mum flight loads when occupied.

(g) For each powered lift system installed be-tween a lower deck service compartment and themain deck for the carriage of persons or equip-ment, or both, the system must meet the followingrequirements:

(1) Each lift control switch outside the lift, ex-cept emergency stop buttons, must be designedto prevent the activation of the life if the lift door, orthe hatch required by paragraph (g)(3) of this sec-tion, or both are open.

(2) An emergency stop button, that when acti-vated will immediately stop the lift, must be in-stalled within the lift and at each entrance to the lift.

(3) There must be a hatch capable of beingused for evacuating persons from the lift that isopenable from inside and outside the lift withouttools, with the lift in any position.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–53, 45 FR 41593, June 19, 1980;45 FR 43154, June 26, 1980; Amdt. 25–110, 68 FR36883, June 19, 2003]

§25.820 Lavatory doors.All lavatory doors must be designed to pre-

clude anyone from becoming trapped inside thelavatory. If a locking mechanism is installed, itmust be capable of being unlocked from the out-side without the aid of special tools.


VENTILATION AND HEATING

§25.831 Ventilation.(a) Under normal operating conditions and in

the event of any probable failure conditions of anysystem which would adversely affect the ventilat-ing air, the ventilation system must be designed toprovide a sufficient amount of uncontaminated airto enable the crewmembers to perform their du-ties without undue discomfort or fatigue and toprovide reasonable passenger comfort. For nor-

mal operating conditions, the ventilation systemmust be designed to provide each occupant withan airflow containing at least 0.55 pounds of freshair per minute.

(b) Crew and passenger compartment air mustbe free from harmful or hazardous concentrationsof gases or vapors. In meeting this requirement,the following apply:

(1) Carbon monoxide concentrations in excessof 1 part in 20,000 parts of air are considered haz-ardous. For test purposes, any acceptable carbonmonoxide detection method may be used.

(2) Carbon dioxide concentration during flightmust be shown not to exceed 0.5 percent by vol-ume (sea level equivalent) in compartments nor-mally occupied by passengers or crewmembers.

(c) There must be provisions made to ensurethat the conditions prescribed in paragraph (b) ofthis section are met after reasonably probable fail-ures or malfunctioning of the ventilating, heating,pressurization, or other systems and equipment.

(d) If accumulation of hazardous quantities ofsmoke in the cockpit area is reasonably probable,smoke evacuation must be readily accomplished,starting with full pressurization and without de-pressurizing beyond safe limits.

(e) Except as provided in paragraph (f) of thissection, means must be provided to enable the oc-cupants of the following compartments and areasto control the temperature and quantity of ventilat-ing air supplied to their compartment or area inde-pendently of the temperature and quantity of airsupplied to other compartments and areas:

(1) The flight crew compartment.(2) Crewmember compartments and areas

other than the flight crew compartment unless thecrewmember compartment or area is ventilatedby air interchange with other compartments orareas under all operating conditions.

(f) Means to enable the flight crew to controlthe temperature and quantity of ventilating airsupplied to the flight crew compartment indepen-dently of the temperature and quantity of ventilat-ing air supplied to other compartments are not re-quired if all of the following conditions are met:

(1) The total volume of the flight crew and pas-senger compartments is 800 cubic feet or less.

(2) The air inlets and passages for air to flowbetween flight crew and passenger compart-ments are arranged to provide compartment tem-peratures within 5 degrees F of each other andadequate ventilation to occupants in both com-partments.

(3) The temperature and ventilation controlsare accessible to the flight crew.


ASA 87

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(g) The exposure time at any given temperaturemust not exceed the values shown in the followinggraph after any improbable failure condition.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; as amended by Amdt. 25–41, 42 FR 36970, July 18, 1977; Amdt. 25–87, 61 FR 28695, June 5, 1996; Amdt. 25–89, 61 FR 63956, Dec. 2, 1996]

§25.832 Cabin ozone concentration.(a) The airplane cabin ozone concentration

during flight must be shown not to exceed—(1) 0.25 parts per million by volume, sea level

equivalent, at any time above flight level 320; and(2) 0.1 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 levelequivalent” refers to conditions of 25°C and 760millimeters of mercury pressure.

(c) Compliance with this section must beshown by analysis or tests based on airplane op-erational procedures and performance limitations,that demonstrate that either—

(1) The airplane cannot be operated at an alti-tude which would result in cabin ozone concentra-tions exceeding the limits prescribed by para-graph (a) of this section; or

(2) The airplane ventilation system, includingany ozone control equipment, will maintain cabinozone concentrations at or below the limits pre-scribed by paragraph (a) of this section.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–50, 45 FR 3883, Jan. 1, 1980;Amdt. 25–56, 47 FR 58489, Dec. 30, 1982; Amdt. 25–94, 63 FR 8848, Feb. 23, 1998]

§25.833 Combustion heating systems.Combustion heaters must be approved.


PRESSURIZATION

§25.841 Pressurized cabins.(a) Pressurized cabins and compartments to

be occupied must be equipped to provide a cabinpressure altitude of not more than 8,000 feet atthe maximum operating altitude of the airplaneunder normal operating conditions.

(1) If certification for operation above 25,000feet is requested, the airplane must be designedso that occupants will not be exposed to cabinpressure altitudes in excess of 15,000 feet afterany probable failure condition in the pressuriza-tion system.

(2) The airplane must be designed so that oc-cupants will not be exposed to a cabin pressurealtitude that exceeds the following after decom-pression from any failure condition not shown tobe extremely improbable:

(i) Twenty-five thousand (25,000) feet for morethan 2 minutes; or

(ii) Forty thousand (40,000) feet for any duration.(3) Fuselage structure, engine and system fail-

ures are to be considered in evaluating the cabindecompression.

TEMPERATURE

TIME — MINUTESTIME — TEMPERATURE RELATIONSHIP

Humidity < 2700 N/m2 (27 mbar) Vapor Pressure

0 50 100 150 200 25075

100

125

150F

Degrees

C

65-60-55-50-45-40-35-30-25-

90


88 ASA

(b) Pressurized cabins must have at least thefollowing valves, controls, and indicators for con-trolling cabin pressure:

(1) Two pressure relief valves to automaticallylimit the positive pressure differential to a prede-termined value at the maximum rate of flow deliv-ered by the pressure source. The combined ca-pacity of the relief valves must be large enough sothat the failure of any one valve would not causean appreciable rise in the pressure differential.The pressure differential is positive when the in-ternal pressure is greater than the external.

(2) Two reverse pressure differential reliefvalves (or their equivalents) to automatically pre-vent a negative pressure differential that woulddamage the structure. One valve is enough, how-ever, if it is of a design that reasonably precludesits malfunctioning.

(3) A means by which the pressure differentialcan be rapidly equalized.

(4) An automatic or manual regulator for con-trolling the intake or exhaust airflow, or both, formaintaining the required internal pressures andairflow rates.

(5) Instruments at the pilot or flight engineerstation to show the pressure differential, the cabinpressure altitude, and the rate of change of thecabin pressure altitude.

(6) Warning indication at the pilot or flight engi-neer station to indicate when the safe or presetpressure differential and cabin pressure altitudelimits are exceeded. Appropriate warning mark-ings on the cabin pressure differential indicatormeet the warning requirement for pressure differ-ential limits and an aural or visual signal (in addi-tion to cabin altitude indicating means) meets thewarning requirement for cabin pressure altitudelimits if it warns the flight crew when the cabinpressure altitude exceeds 10,000 feet.

(7) A warning placard at the pilot or flight engi-neer station if the structure is not designed forpressure differentials up to the maximum reliefvalve setting in combination with landing loads.

(8) The pressure sensors necessary to meetthe requirements of paragraphs (b)(5) and (b)(6)of this section and §25.1447(c), must be locatedand the sensing system designed so that, in theevent of loss of cabin pressure in any passengeror crew compartment (including upper and lowerlobe galleys), the warning and automatic presen-tation devices, required by those provisions, willbe actuated without any delay that would signifi-cantly increase the hazards resulting from de-compression.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–38, 41 FR 55466, Dec. 20, 1976;Amdt. 25–87, 61 FR 28696, June 5, 1996]

§25.843 Tests for pressurized cabins.(a) Strength test. The complete pressurized

cabin, including doors, windows, and valves, mustbe tested as a pressure vessel for the pressuredifferential specified in §25.365(d).

(b) Functional tests. The following functionaltests must be performed:

(1) Tests of the functioning and capacity of thepositive and negative pressure differential valves,and of the emergency release valve, to stimulatethe effects of closed regulator valves.

(2) Tests of the pressurization system to showproper functioning under each possible conditionof pressure, temperature, and moisture, up to themaximum altitude for which certification is re-quested.

(3) Flight tests, to show the performance of thepressure supply, pressure and flow regulators, in-dicators, and warning signals, in steady andstepped climbs and descents at rates correspond-ing to the maximum attainable within the operatinglimitations of the airplane, up to the maximum alti-tude for which certification is requested.

(4) Tests of each door and emergency exit, toshow that they operate properly after being sub-jected to the flight tests prescribed in paragraph(b)(3) of this section.

FIRE PROTECTION

§25.851 Fire extinguishers.(a) Hand fire extinguishers.(1) The following minimum number of hand fire

extinguishers must be conveniently located andevenly distributed in passenger compartments:

(2) At least one hand fire extinguisher must beconveniently located in the pilot compartment.

(3) At least one readily accessible hand fire ex-tinguisher must be available for use in each ClassA or Class B cargo or baggage compartment andin each Class E cargo or baggage compartmentthat is accessible to crewmembers in flight.

(4) At least one hand fire extinguisher must belocated in, or readily accessible for use in, eachgalley located above or below the passengercompartment.

(5) Each hand fire extinguisher must be ap-proved.

Passenger capacity No. 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


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(6) At least one of the required fire extinguish-ers located in the passenger compartment of anairplane with a passenger capacity of at least 31and not more than 60, and at least two of the fireextinguishers located in the passenger compart-ment of an airplane with a passenger capacity of61 or more must contain Halon 1211 (bromochlo-rodifluoromethane CBrC1F2), or equivalent, asthe extinguishing agent. The type of extinguishingagent used in any other extinguisher required bythis section must be appropriate for the kinds offires likely to occur where used.

(7) The quantity of extinguishing agent used ineach extinguisher required by this section mustbe appropriate for the kinds of fires likely to occurwhere used.

(8) Each extinguisher intended for use in a per-sonnel compartment must be designed to mini-mize the hazard of toxic gas concentration.

(b) Built-in fire extinguishers. If a built-in fire ex-tinguisher is provided—

(1) Each built-in fire extinguishing system mustbe installed so that—

(i) No extinguishing agent likely to enter per-sonnel compartments will be hazardous to the oc-cupants; and

(ii) No discharge of the extinguisher can causestructural damage.

(2) The capacity of each required built-in fireextinguishing system must be adequate for anyfire likely to occur in the compartment whereused, considering the volume of the compartmentand the ventilation rate.


§25.853 Compartment interiors.For each compartment occupied by the crew or

passengers, the following apply:(a) Materials (including finishes or decorative

surfaces applied to the materials) must meet theapplicable test criteria prescribed in Part I of Ap-pendix F of this part, or other approved equivalentmethods, regardless of the passenger capacity ofthe airplane.

(b) [Reserved](c) In addition to meeting the requirements of

paragraph (a) of this section, seat cushions, ex-cept those on flight crewmember seats, mustmeet the test requirements of part II of AppendixF of this part, or other equivalent methods, re-gardless of the passenger capacity of the air-plane.

(d) Except as provided in paragraph (e) of thissection, the following interior components of air-planes with passenger capacities of 20 or moremust also meet the test requirements of parts IVand V of Appendix F of this part, or other ap-proved equivalent method, in addition to the flam-

mability requirements prescribed in paragraph (a)of this section:

(1) Interior ceiling and wall panels, other thanlighting lenses and windows;

(2) Partitions, other than transparent panelsneeded to enhance cabin safety;

(3) Galley structure, including exposed sur-faces of stowed carts and standard containersand the cavity walls that are exposed when a fullcomplement of such carts or containers is not car-ried; and

(4) Large cabinets and cabin stowage compart-ments, other than underseat stowage compart-ments for stowing small items such as magazinesand maps.

(e) The interiors of compartments, such as pilotcompartments, galleys, lavatories, crew rest quar-ters, cabinets and stowage compartments, neednot meet the standards of paragraph (d) of thissection, provided the interiors of such compart-ments are isolated from the main passengercabin by doors or equivalent means that wouldnormally be closed during an emergency landingcondition.

(f) Smoking is not allowed in lavatories. Ifsmoking is allowed in any area occupied by thecrew or passengers, an adequate number of self-contained, removable ashtrays must be providedin designated smoking sections for all seated oc-cupants.

(g) Regardless of whether smoking is allowedin any other part of the airplane, lavatories musthave self-contained, removable ashtrays locatedconspicuously on or near the entry side of eachlavatory door, except that one ashtray may servemore than one lavatory door if the ashtray can beseen readily from the cabin side of each lavatoryserved.

(h) Each receptacle used for the disposal offlammable waste material must be fully enclosed,constructed of at least fire resistant materials, andmust contain fires likely to occur in it under normaluse. The capability of the receptacle to containthose fires under all probable conditions of wear,misalignment, and ventilation expected in servicemust be demonstrated by test.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–72, 55 FR 29783, July 20, 1990;Amdt. 25–83, 60 FR 6623, Feb. 2, 1995; Amdt. 25–116,69 FR 62788, Oct. 27, 2004]

§25.854 Lavatory fire protection.For airplanes with a passenger capacity of 20

or more:(a) Each lavatory must be equipped with a

smoke detector system or equivalent that pro-vides a warning light in the cockpit, or provides awarning light or audible warning in the passenger


90 ASA

cabin that would be readily detected by a flight at-tendant; and

(b) Each lavatory must be equipped with abuilt-in fire extinguisher for each disposal recepta-cle for towels, paper, or waste, located within thelavatory. The extinguisher must be designed todischarge automatically into each disposal recep-tacle upon occurrence of a fire in that receptacle.


§25.855 Cargo or baggage compartments.For each cargo or baggage compartment, 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 havea liner, and the liner must be separate from (butmay be attached to) the airplane structure.

(c) Ceiling and sidewall liner panels of Class Ccompartments must meet the test requirements ofpart III of Appendix F of this part or other ap-proved equivalent methods.

(d) All other materials used in the constructionof the cargo or baggage compartment must meetthe applicable test criteria prescribed in part I ofAppendix F of this part or other approved equiva-lent methods.

(e) No compartment may contain any controls,lines, equipment, or accessories whose damageor failure would affect safe operation, unlessthose items are protected so that—

(1) They cannot be damaged by the movementof cargo in the compartment, and

(2) Their breakage or failure will not create afire hazard.

(f) There must be means to prevent cargo orbaggage from interfering with the functioning ofthe fire protective features of the compartment.

(g) Sources of heat within the compartmentmust be shielded and insulated to prevent ignitingthe cargo or baggage.

(h) Flight tests must be conducted to showcompliance with the provisions of §25.857 con-cerning—

(1) Compartment accessibility,(2) The entries of hazardous quantities of

smoke or extinguishing agent into compartmentsoccupied by the crew or passengers, and

(3) The dissipation of the extinguishing agent inClass C compartments.

(i) During the above tests, it must be shownthat no inadvertent operation of smoke or fire de-tectors in any compartment would occur as a re-sult of fire contained in any other compartment,either during or after extinguishment, unless the

extinguishing system floods each such compart-ment simultaneously.

(j) Cargo or baggage compartment electricalwiring interconnection system components mustmeet the requirements of §25.1721.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–72, 55 FR 29784, July 20, 1990;Amdt. 25–93, 63 FR 8048, Feb. 17, 1998; Amdt. 25–116,69 FR 62789, Oct. 27, 2004; Amdt. 25–123, 72 FR63405, Nov. 8, 2007]

§ 25.856 Thermal/Acoustic insulation materials.(a) Thermal/acoustic insulation material in-

stalled in the fuselage must meet the flame prop-agation test requirements of part VI of Appendix Fto this part, or other approved equivalent test re-quirements. This requirement does not apply to“small parts,” as defined in part I of Appendix F ofthis part.

(b) For airplanes with a passenger capacity of20 or greater, thermal/acoustic insulation materi-als (including the means of fastening the materi-als to the fuselage) installed in the lower half ofthe airplane fuselage must meet the flame pene-tration resistance test requirements of part VII ofAppendix F to this part, or other approved equiva-lent test requirements. This requirement does notapply to thermal/acoustic insulation installationsthat the FAA finds would not contribute to fire pen-etration resistance.

[Docket No. FAA–2000–7909, 68 FR 45059, July 31,2003]

§25.857 Cargo compartment classification.(a) Class A. A Class A cargo or baggage com-

partment is one in which—(1) The presence of a fire would be easily dis-

covered by a crewmember while at his station;and

(2) Each part of the compartment is easily ac-cessible in flight.

(b) Class B. A Class B cargo or baggage com-partment is one in which—

(1) There is sufficient access in flight to enablea crewmember to effectively reach any part of thecompartment with the contents of a hand fire ex-tinguisher;

(2) When the access provisions are beingused, no hazardous quantity of smoke, flames, orextinguishing agent, will enter any compartmentoccupied by the crew or passengers;

(3) There is a separate approved smoke detec-tor or fire detector system to give warning at thepilot or flight engineer station.

(c) Class C A Class C cargo or baggage com-partment is one not meeting the requirements foreither a Class A or B compartment but in which—


ASA 91

25

(1) There is a separate approved smoke detec-tor or fire detector system to give warning at thepilot or flight engineer station;

(2) There is an approved built-in fire extinguish-ing or suppression system controllable from thecockpit;

(3) There are means to exclude hazardous quan-tities of smoke, flames, or extinguishing agent, fromany compartment occupied by the crew or passen-gers;

(4) There are means to control ventilation anddrafts within the compartment so that the extin-guishing agent used can control any fire that maystart within the compartment.

(d) [Reserved](e) Class E. A Class E cargo compartment is

one on airplanes used only for the carriage ofcargo and in which—

(1) [Reserved](2) There is a separate approved smoke or fire

detector system to give warning at the pilot orflight engineer station;

(3) There are means to shut off the ventilatingairflow to, or within, the compartment, and thecontrols for these means are accessible to theflight crew in the crew compartment;

(4) There are means to exclude hazardousquantities of smoke, flames, or noxious gases,from the flight crew compartment; and

(5) The required crew emergency exits are ac-cessible under any cargo loading condition.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended 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]

§25.858 Cargo or baggage compartment smoke or fire detection systems.If certification with cargo or baggage compart-

ment smoke or fire detection provisions is re-quested, the following must be met for each cargoor baggage compartment with those provisions:

(a) The detection system must provide a visualindication to the flight crew within one minute afterthe start of a fire.

(b) The system must be capable of detecting afire at a temperature significantly below that atwhich the structural integrity of the airplane issubstantially decreased.

(c) There must be means to allow the crew tocheck in flight, the functioning of each fire detec-tor circuit.

(d) The effectiveness of the detection systemmust be shown for all approved operating configu-rations and conditions.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–54, 45 FR 60173, Sept. 11, 1980;Amdt. 25–93, 63 FR 8048, Feb. 17, 1998]

§25.859 Combustion heater fire protection.(a) Combustion heater fire zones. The following

combustion heater fire zones must be protectedfrom fire in accordance with the applicable provi-sions of §§25.1181 through 25.1191 and§§25.1195 through 25.1203;

(1) The region surrounding the heater, if this re-gion contains any flammable fluid system compo-nents (excluding the heater fuel system), thatcould—

(i) Be damaged by heater malfunctioning; or(ii) Allow flammable fluids or vapors to reach

the heater in case of leakage.(2) The region surrounding the heater, if the

heater fuel system has fittings that, if they leaked,would allow fuel or vapors to enter this region.

(3) The part of the ventilating air passage thatsurrounds the combustion chamber. However, nofire extinguishment is required in cabin ventilatingair passages.

(b) Ventilating air ducts. Each ventilating airduct passing through any fire zone must be fire-proof. In addition—

(1) Unless isolation is provided by fireproofvalves or by equally effective means, the ventilat-ing air duct downstream of each heater must befireproof for a distance great enough to ensurethat any fire originating in the heater can be con-tained in the duct; and

(2) Each part of any ventilating duct passingthrough any region having a flammable fluid sys-tem must be constructed or isolated from that sys-tem so that the malfunctioning of any componentof that system cannot introduce flammable fluidsor vapors into the ventilating airstream.

(c) Combustion air ducts. Each combustion airduct must be fireproof for a distance great enoughto prevent damage from backfiring or reverseflame propagation. In addition—

(1) No combustion air duct may have a com-mon opening with the ventilating airstream unlessflames from backfires or reverse burning cannotenter the ventilating airstream under any operat-ing condition, including reverse flow or malfunc-tioning of the heater or its associated compo-nents; and

(2) No combustion air duct may restrict theprompt relief of any backfire that, if so restricted,could cause heater failure.

(d) Heater controls; general. Provision must bemade to prevent the hazardous accumulation ofwater or ice on or in any heater control compo-nent, control system tubing, or safety control.

(e) Heater safety controls. For each combus-tion heater there must be the following safety con-trol means:

(1) Means independent of the components pro-vided for the normal continuous control of air tem-


92 ASA

perature, airflow, and fuel flow must be provided,for each heater, to automatically shut off the igni-tion and fuel supply to that heater at a point re-mote from that heater when any of the followingoccurs:

(i) The heat exchanger temperature exceedssafe limits.

(ii) The ventilating air temperature exceeds safelimits.

(iii) The combustion airflow becomes inade-quate for safe operation.

(iv) The ventilating airflow becomes inadequatefor safe operation.

(2) The means of complying with paragraph(e)(1) of this section for any individual heatermust—

(i) Be independent of components serving anyother heater whose heat output is essential forsafe operation; and

(ii) Keep the heater off until restarted by thecrew.

(3) There must be means to warn the crewwhen any heater whose heat output is essentialfor safe operation has been shut off by the auto-matic means prescribed in paragraph (e)(1) ofthis section.

(f) Air intakes. Each combustion and ventilatingair intake must be located so that no flammablefluids or vapors can enter the heater system un-der any operating condition—

(1) During normal operation; or(2) As a result of the malfunctioning of any

other component.(g) Heater exhaust. Heater exhaust systems

must meet the provisions of §§25.1121 and25.1123. In addition, there must be provisions in thedesign of the heater exhaust system to safely expelthe products of combustion to prevent the occur-rence of—

(1) Fuel leakage from the exhaust to surround-ing compartments;

(2) Exhaust gas impingement on surroundingequipment or structure;

(3) Ignition of flammable fluids by the exhaust,if the exhaust is in a compartment containingflammable fluid lines; and

(4) Restriction by the exhaust of the prompt re-lief of backfires that, if so restricted, could causeheater failure.

(h) Heater fuel systems. Each heater fuel sys-tem must meet each powerplant fuel system re-quirement affecting safe heater operation. Eachheater fuel system component within the ventilat-ing airstream must be protected by shrouds sothat no leakage from those components can enterthe ventilating airstream.

(i) Drains. There must be means to safely drainfuel that might accumulate within the combustionchamber or the heat exchanger. In addition—

(1) Each part of any drain that operates at hightemperatures must be protected in the samemanner as heater exhausts; and

(2) Each drain must be protected from hazard-ous ice accumulation under any operating condi-tion.

[Docket No. 5066, 29 FR 18291, Dec. 24 1964; asamended by Amdt. 25–11, 32 FR 6912, May 5, 1967;Amdt. 25–23, 35 FR 5676, April 8, 1970]

§25.863 Flammable fluid fire protection.(a) In each area where flammable fluids or va-

pors might escape by leakage of a fluid system,there must be means to minimize the probabilityof ignition of the fluids and vapors, and the result-ant hazards if ignition does occur.

(b) Compliance with paragraph (a) of this sec-tion must be shown by analysis or tests, and thefollowing factors must be considered:

(1) Possible sources and paths of fluid leakage,and means of detecting leakage.

(2) Flammability characteristics of fluids, in-cluding effects of any combustible or absorbingmaterials.

(3) Possible ignition sources, including electri-cal faults, overheating of equipment, and malfunc-tioning of protective devices.

(4) Means available for controlling or extin-guishing 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 criti-cal to safety of flight to withstand fire and heat.

(c) If action by the flight crew is required to pre-vent or counteract a fluid fire (e.g., equipmentshutdown or actuation of a fire extinguisher) quickacting means must be provided to alert the crew.

(d) Each area where flammable fluids or vaporsmight escape by leakage of a fluid system mustbe identified and defined.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5676, April 8, 1970;Amdt. 25–46, 43 FR 50597, Oct. 30, 1978]

§25.865 Fire protection of flight controls, engine mounts, and other flight structure.Essential flight controls, engine mounts, and

other flight structures located in designated firezones or in adjacent areas which would be sub-jected to the effects of fire in the fire zone must beconstructed of fireproof material or shielded sothat they are capable of withstanding the effectsof fire.



ASA 93

25

§25.867 Fire protection: other components.(a) Surfaces to the rear of the nacelles, within

one nacelle diameter of the nacelle centerline,must be at least fire-resistant.

(b) Paragraph (a) of this section does not applyto tail surfaces to the rear of the nacelles thatcould not be readily affected by heat, flames, orsparks coming from a designated fire zone or en-gine compartment of any nacelle.


§25.869 Fire protection: systems.(a) Electrical system components:(1) Components of the electrical system must

meet the applicable fire and smoke protection re-quirements of §§25.831(c) and 25.863.

(2) Equipment that is located in designated firezones and is used during emergency proceduresmust be at least fire resistant.

(3) EWIS components must meet the require-ments of §25.1713.

(b) Each vacuum air system line and fitting onthe discharge side of the pump that might containflammable vapors or fluids must meet the require-ments of §25.1183 if the line or fitting is in a des-ignated fire zone. Other vacuum air systems com-ponents in designated fire zones must be at leastfire resistant.

(c) Oxygen equipment and lines must—(1) Not be located in any designated fire zone,(2) Be protected from heat that may be gener-

ated in, or escape from, any designated fire zone,and

(3) Be installed so that escaping oxygen cannotcause ignition of grease, fluid, or vapor accumula-tions that are present in normal operation or as aresult of failure or malfunction of any system.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–72, 55 FR 29784, July 20, 1990;Docket Nos. FAA–2001–9634, FAA–2001–9633, FAA–2001–9638, FAA–2001–9637; Amdt. 25–113; 69 FR12529, March 16, 2004]

MISCELLANEOUS

§25.871 Leveling means.There must be means for determining when the

airplane is in a level position on the ground.


§25.875 Reinforcement near propellers.(a) Each part of the airplane near the propeller

tips must be strong and stiff enough to withstandthe effects of the induced vibration and of icethrown from the propeller.

(b) No window may be near the propeller tipsunless it can withstand the most severe ice impactlikely to occur.

§25.899 Electrical bonding and protection against static electricity.(a) Electrical bonding and protection against

static electricity must be designed to minimize ac-cumulation of electrostatic charge that wouldcause—

(1) Human injury from electrical shock,(2) Ignition of flammable vapors, or(3) Interference with installed electrical/elec-

tronic equipment.(b) Compliance with paragraph (a) of this sec-

tion may be shown by—(1) Bonding the components properly to the air-

frame; or(2) Incorporating other acceptable means to

dissipate the static charge so as not to endangerthe airplane, personnel, or operation of the in-stalled electrical/electronic systems.

[Docket No. FAA–2004–18379, 72 FR 63405, Nov. 8,2007]


94 ASA

Subpart E — Powerplant

GENERAL

§25.901 Installation.(a) For the purpose of this part, the airplane

powerplant installation includes each componentthat—

(1) Is necessary for propulsion;(2) Affects the control of the major propulsive

units; or(3) Affects the safety of the major propulsive

units between normal inspections or overhauls.(b) For each powerplant—(1) The installation must comply with—(i) The installation instructions provided under

§§33.5 and 35.3 of this chapter; and(ii) The applicable provisions of this subpart;(2) The components of the installation must be

constructed, arranged, and installed so as to en-sure their continued safe operation between nor-mal inspections or overhauls;

(3) The installation must be accessible for nec-essary inspections and maintenance; and

(4) The major components of the installationmust be electrically bonded to the other parts ofthe airplane.

(c) For each powerplant and auxiliary powerunit installation, it must be established that no sin-gle failure or malfunction or probable combinationof failures will jeopardize the safe operation of theairplane except that the failure of structural ele-ments need not be considered if the probability ofsuch failure is extremely remote.

(d) Each auxiliary power unit installation mustmeet the applicable provisions of this subpart.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5676, April 8, 1970;Amdt. 25–40, 42 FR 15042, March 17, 1977; Amdt. 25–46, 43 FR 50597, Oct. 30, 1978; Amdt. 25–126, 73 FR63345, Oct. 24, 2008]

§25.903 Engines.(a) Engine type certificate.(1) Each engine must have a type certificate

and must meet the applicable requirements ofpart 34 of this chapter.

(2) Each turbine engine must comply with oneof the following:

(i) Sections 33.76, 33.77 and 33.78 of thischapter in effect on December 13, 2000, or assubsequently amended; or

(ii) Sections 33.77 and 33.78 of this chapter ineffect on April 30, 1998, or as subsequentlyamended before December 13, 2000; or

(iii) Comply with §33.77 of this chapter in effecton October 31, 1974, or as subsequentlyamended prior to April 30, 1998, unless that en-gine’s foreign object ingestion service history hasresulted in an unsafe condition; or

(iv) Be shown to have a foreign object ingestionservice history in similar installation locationswhich has not resulted in any unsafe condition.

(b) Engine isolation. The powerplants must bearranged and isolated from each other to allowoperation, in at least one configuration, so that thefailure or malfunction of any engine, or of any sys-tem that can affect the engine, will not—

(1) Prevent the continued safe operation of theremaining engines; or

(2) Require immediate action by any crewmem-ber for continued safe operation.

(c) Control of engine rotation. There must bemeans for stopping the rotation of any engine in-dividually in flight, except that, for turbine engineinstallations, the means for stopping the rotationof any engine need be provided only where con-tinued rotation could jeopardize the safety of theairplane. Each component of the stopping systemon the engine side of the firewall that might be ex-posed to fire must be at least fire-resistant. If hy-draulic propeller feathering systems are used forthis purpose, the feathering lines must be at leastfire resistant under the operating conditions thatmay be expected to exist during feathering.

(d) Turbine engine installations. For turbine en-gine installations—

(1) Design precautions must be taken to mini-mize the hazards to the airplane in the event of anengine rotor failure or of a fire originating withinthe engine which burns through the engine case.

(2) The powerplant systems associated withengine control devices, systems, and instrumen-tation, must be designed to give reasonable as-surance that those engine operating limitationsthat adversely affect turbine rotor structural integ-rity will not be exceeded in service.

(e) Restart capability.(1) Means to restart any engine in flight must

be provided.(2) An altitude and airspeed envelope must be

established for in-flight engine restarting, andeach engine must have a restart capability withinthat envelope.

(3) For turbine engine powered airplanes, if theminimum windmilling speed of the engines, fol-lowing the inflight shutdown of all engines, is in-sufficient to provide the necessary electricalpower for engine ignition, a power source inde-pendent of the engine-driven electrical powergenerating system must be provided to permit in-flight engine ignition for restarting.

(f) Auxiliary Power Unit. Each auxiliary powerunit must be approved or meet the requirementsof the category for its intended use.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5676, April 8, 1970;Amdt. 25–40, 42 FR 15042, March 17, 1977; Amdt. 25–57, 49 FR 6848, Feb. 23, 1984; Amdt. 25–72, 55 FR29784, July 20, 1990; Amdt. 25–73, 55 FR 32861, Aug.


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10, 1990; Amdt. 25–94, 63 FR 8848, Feb. 23, 1998;Amdt. 25–95, 63 FR 14798, March 26, 1998; Amdt. 25–100, 65 FR 55854, Sept. 14, 2000]

§25.904 Automatic takeoff thrust control system (ATTCS).Each applicant seeking approval for installation

of an engine power control system that automati-cally resets the power or thrust on the operatingengine(s) when any engine fails during the takeoffmust comply with the requirements of Appendix Iof this part.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–62, 52 FR 43156, Nov. 9, 1987]

§25.905 Propellers.(a) Each propeller must have a type certificate.(b) Engine power and propeller shaft rotational

speed may not exceed the limits for which the pro-peller is certificated.

(c) The propeller blade pitch control systemmust meet the requirements of §§35.21, 35.23,35.42 and 35.43 of this chapter.

(d) Design precautions must be taken to mini-mize the hazards to the airplane in the event apropeller blade fails or is released by a hub failure.The hazards which must be considered includedamage to structure and vital systems due to im-pact of a failed or released blade and the unbal-ance created by such failure or release.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–54, 45 FR 60173, Sept. 11, 1980;Amdt. 25–57, 49 FR 6848, Feb. 23, 1984; Amdt. 25–72,55 FR 29784, July 20, 1990; Amdt. 25–126, 73 FR63345, Oct. 24, 2008]

§25.907 Propeller vibration and fatigue.This section does not apply to fixed-pitch wood

propellers of conventional design.(a) The applicant must determine the magni-

tude of the propeller vibration stresses or loads,including any stress peaks and resonant condi-tions, throughout the operational envelope of theairplane by either:

(1) Measurement of stresses or loads throughdirect testing or analysis based on direct testing ofthe propeller on the airplane and engine installa-tion for which approval is sought; or

(2) Comparison of the propeller to similar pro-pellers installed on similar airplane installationsfor which these measurements have been made.

(b) The applicant must demonstrate by tests,analysis based on tests, or previous experienceon similar designs that the propeller does not ex-perience harmful effects of flutter throughout theoperational envelope of the airplane.

(c) The applicant must perform an evaluation ofthe propeller to show that failure due to fatigue willbe avoided throughout the operational life of the

propeller using the fatigue and structural data ob-tained in accordance with part 35 of this chapterand the vibration data obtained from compliancewith paragraph (a) of this section. For the purposeof this paragraph, the propeller includes the hub,blades, blade retention component and any otherpropeller component whose failure due to fatiguecould be catastrophic to the airplane. This evalua-tion must include:

(1) The intended loading spectra including allreasonably foreseeable propeller vibration andcyclic load patterns, identified emergency condi-tions, allowable overspeeds and overtorques, andthe effects of temperatures and humidity ex-pected in service.

(2) The effects of airplane and propeller operat-ing and airworthiness limitations.

[Docket No. FAA–2007–27310, 73 FR 63345, Oct. 24,2008]

§25.925 Propeller clearance.Unless smaller clearances are substantiated,

propeller clearances with the airplane at maxi-mum weight, with the most adverse center ofgravity, and with the propeller in the most adversepitch position, may not be less than the following:

(a) Ground clearance. There must be a clear-ance of at least seven inches (for each airplanewith nose wheel landing gear) or nine inches (foreach airplane with tail wheel landing gear) be-tween each propeller and the ground with thelanding gear statically deflected and in the leveltakeoff, or taxiing attitude, whichever is most criti-cal. In addition, there must be positive clearancebetween the propeller and the ground when in thelevel takeoff attitude with the critical tire(s) com-pletely deflated and the corresponding landinggear strut bottomed.

(b) Water clearance. There must be a clearanceof at least 18 inches between each propeller andthe water, unless compliance with §25.239(a) canbe shown with a lesser clearance.

(c) Structural clearance. There must be—(1) At least one inch radial clearance between

the blade tips and the airplane structure, plus anyadditional radial clearance necessary to preventharmful vibration;

(2) At least one-half inch longitudinal clearancebetween the propeller blades or cuffs and station-ary parts of the airplane; and

(3) Positive clearance between other rotatingparts of the propeller or spinner and stationaryparts of the airplane.


§25.929 Propeller deicing.(a) For airplanes intended for use where icing

may be expected, there must be a means to pre-


96 ASA

vent or remove hazardous ice accumulation onpropellers or on accessories where ice accumula-tion would jeopardize engine performance.

(b) If combustible fluid is used for propeller de-icing, §§25.1181 through 25.1185 and 25.1189apply.

§25.933 Reversing systems.(a) For turbojet reversing systems—(1) Each system intended for ground operation

only must be designed so that during any reversalin flight the engine will produce no more thanflight idle thrust. In addition, it must be shown byanalysis or test, or both, that—

(i) Each operable reverser can be restored tothe forward thrust position; and

(ii) The airplane is capable of continued safeflight and landing under any possible position ofthe thrust reverser.

(2) Each system intended for inflight use mustbe designed so that no unsafe condition will resultduring normal operation of the system, or fromany failure (or reasonably likely combination offailures) of the reversing system, under any antic-ipated condition of operation of the airplane in-cluding ground operation. Failure of structural ele-ments need not be considered if the probability ofthis kind of failure is extremely remote.

(3) Each system must have means to prevent theengine from producing more than idle thrust whenthe reversing system malfunctions, except that itmay produce any greater forward thrust that isshown to allow directional control to be maintained,with aerodynamic means alone, under the mostcritical reversing condition expected in operation.

(b) For propeller reversing systems—(1) Each system intended for ground operation

only must be designed so that no single failure (orreasonably likely combination of failures) or mal-function of the system will result in unwanted re-verse thrust under any expected operating condi-tion. Failure of structural elements need not beconsidered if this kind of failure is extremely re-mote.

(2) Compliance with this section may be shownby failure analysis or testing, or both, for propellersystems that allow propeller blades to move fromthe flight low-pitch position to a position that issubstantially less than that at the normal flightlow-pitch position. The analysis may include or besupported by the analysis made to show compli-ance with the requirements of §35.21 of this chap-ter for the propeller and associated installationcomponents.


§25.934 Turbojet engine thrust reverser system tests.Thrust reversers installed on turbojet engines

must meet the requirements of §33.97 of thischapter.


§25.937 Turbopropeller-drag limiting systems.Turbopropeller power airplane propeller-drag

limiting systems must be designed so that no sin-gle failure or malfunction of any of the systemsduring normal or emergency operation results inpropeller drag in excess of that for which the air-plane was designed under §25.367. Failure ofstructural elements of the drag limiting systemsneed not be considered if the probability of thiskind of failure is extremely remote.

§25.939 Turbine engine operating characteristics.(a) Turbine engine operating characteristics

must be investigated in flight to determine that noadverse characteristics (such as stall, surge, orflameout) are present, to a hazardous degree,during normal and emergency operation withinthe range of operating limitations of the airplaneand of the engine.

(b) [Reserved](c) The turbine engine air inlet system may not,

as a result of air flow distortion during normal op-eration, cause vibration harmful to the engine.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–11, 32 FR 6912, May 5, 1967;Amdt. 25–40, 42 FR 15043, March 17, 1977]

§25.941 Inlet, engine, and exhaust compatibility.For airplanes using variable inlet or exhaust

system geometry, or both—(a) The system comprised of the inlet, engine

(including thrust augmentation systems, if incor-porated), and exhaust must be shown to functionproperly under all operating conditions for whichapproval is sought, including all engine rotatingspeeds and power settings, and engine inlet andexhaust configurations;

(b) The dynamic effects of the operation ofthese (including consideration of probable mal-functions) upon the aerodynamic control of theairplane may not result in any condition that wouldrequire exceptional skill, alertness, or strength onthe part of the pilot to avoid exceeding an opera-tional or structural limitation of the airplane; and

(c) In showing compliance with paragraph (b)of this section, the pilot strength required may notexceed the limits set forth in §25.143(d), subject


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to the conditions set forth in paragraphs (e) and (f)of §25.143.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–38, 41 FR 55467, Dec. 20, 1976;Amdt. 25–121, 72 FR 44669, Aug. 8, 2007]

§25.943 Negative acceleration.No hazardous malfunction of an engine, an

auxiliary power unit approved for use in flight, orany component or system associated with thepowerplant or auxiliary power unit may occurwhen the airplane is operated at the negative ac-celerations within the flight envelopes prescribedin §25.333. This must be shown for the greatestduration expected for the acceleration.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–40, 42 FR 15043, March 17,1977]

§25.945 Thrust or power augmentation system.(a) General. Each fluid injection system must

provide a flow of fluid at the rate and pressure es-tablished for proper engine functioning undereach intended operating condition. If the fluid canfreeze, fluid freezing may not damage the air-plane or adversely affect airplane performance.

(b) Fluid tanks. Each augmentation systemfluid tank must meet the following requirements:

(1) Each tank must be able to withstand withoutfailure the vibration, inertia, fluid, and structuralloads that it may be subject to in operation.

(2) The tanks as mounted in the airplane mustbe able to withstand without failure or leakage aninternal pressure 1.5 times the maximum operat-ing pressure.

(3) If a vent is provided, the venting must be ef-fective under all normal flight conditions.

(4) [Reserved](c) Augmentation system drains must be de-

signed and located in accordance with §25.1455if—

(1) The augmentation system fluid is subject tofreezing; and

(2) The fluid may be drained in flight or duringground operation.

(d) The augmentation liquid tank capacity avail-able for the use of each engine must be largeenough to allow operation of the airplane underthe approved procedures for the use of liquid-aug-mented power. The computation of liquid con-sumption must be based on the maximum ap-proved rate appropriate for the desired engineoutput and must include the effect of temperatureon engine performance as well as any other fac-tors that might vary the amount of liquid required.

(e) This section does not apply to fuel injectionsystems.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–40, 42 FR 15043, March 17,1977; Amdt. 25–72, 55 FR 29785, July 20, 1990]

FUEL SYSTEM

§25.951 General.(a) Each fuel system must be constructed and

arranged to ensure a flow of fuel at a rate andpressure established for proper engine and auxil-iary power unit functioning under each likely oper-ating condition, including any maneuver for whichcertification is requested and during which the en-gine or auxiliary power unit is permitted to be inoperation.

(b) Each fuel system must be arranged so thatany air which is introduced into the system will notresult in—

(1) Power interruption for more than 20 sec-onds for reciprocating engines; or

(2) Flameout for turbine engines.(c) Each fuel system for a turbine engine must

be capable of sustained operation throughout itsflow and pressure range with fuel initially satu-rated with water at 80°F and having 0.75cc of freewater per gallon added and cooled to the mostcritical condition for icing likely to be encounteredin operation.

(d) Each fuel system for a turbine engine pow-ered airplane must meet the applicable fuel vent-ing requirements of part 34 of this chapter.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5677, April 8, 1970;Amdt. 25–36, 39 FR 35460, Oct. 1, 1974; Amdt. 25–38,41 FR 55467, Dec. 20, 1976; Amdt. 25–73, 55 FR32861, Aug. 10, 1990]

§25.952 Fuel system analysis and test.(a) Proper fuel system functioning under all

probable operating conditions must be shown byanalysis and those tests found necessary by theAdministrator. Tests, if required, must be madeusing the airplane fuel system or a test article thatreproduces the operating characteristics of theportion of the fuel system to be tested.

(b) The likely failure of any heat exchanger us-ing fuel as one of its fluids may not result in a haz-ardous condition.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–40, 42 FR 15043, Mar. 17, 1977]

§25.953 Fuel system independence.Each fuel system must meet the requirements

of §25.903(b) by—(a) Allowing the supply of fuel to each engine

through a system independent of each part of thesystem supplying fuel to any other engine; or

(b) Any other acceptable method.


98 ASA

§25.954 Fuel system lightning protection.The fuel system must be designed and ar-

ranged to prevent the ignition of fuel vapor withinthe system by—

(a) Direct lightning strikes to areas having ahigh probability of stroke attachment;

(b) Swept lightning strokes to areas whereswept strokes are highly probable; and

(c) Corona and streamering at fuel vent outlets.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–14, 32 FR 11629, Aug. 11, 1967]

§25.955 Fuel flow.(a) Each fuel system must provide at least 100

percent of the fuel flow required under each in-tended operating condition and maneuver. Com-pliance must be shown as follows:

(1) Fuel must be delivered to each engine at apressure within the limits specified in the enginetype certificate.

(2) The quantity of fuel in the tank may not ex-ceed the amount established as the unusable fuelsupply for that tank under the requirements of§25.959 plus that necessary to show compliancewith this section.

(3) Each main pump must be used that is nec-essary for each operating condition and attitudefor which compliance with this section is shown,and the appropriate emergency pump must besubstituted for each main pump so used.

(4) If there is a fuel flowmeter, it must beblocked and the fuel must flow through the meteror its bypass.

(b) If an engine can be supplied with fuel frommore than one tank, the fuel system must—

(1) For each reciprocating engine, supply thefull fuel pressure to that engine in not more than20 seconds after switching to any other fuel tankcontaining usable fuel when engine malfunction-ing becomes apparent due to the depletion of thefuel supply in any tank from which the engine canbe fed; and

(2) For each turbine engine, in addition to hav-ing appropriate manual switching capability, bedesigned to prevent interruption of fuel flow to thatengine, without attention by the flight crew, whenany tank supplying fuel to that engine is depletedof usable fuel during normal operation, and anyother tank, that normally supplies fuel to that en-gine alone, contains usable fuel.


§25.957 Flow between interconnected tanks.If fuel can be pumped from one tank to another

in flight, the fuel tank vents and the fuel transfer

system must be designed so that no structuraldamage to the tanks can occur because of over-filling.

§25.959 Unusable fuel supply.The unusable fuel quantity for each fuel tank

and its fuel system components must be estab-lished at not less than the quantity at which thefirst evidence of engine malfunction occurs underthe most adverse fuel feed condition for all in-tended operations and flight maneuvers involvingfuel feeding from that tank. Fuel system compo-nent failures need not be considered.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5677, April 8, 1970;Amdt. 25–40, 42 FR 15043, March 17, 1977]

§25.961 Fuel system hot weather operation.(a) The fuel system must perform satisfactorily

in hot weather operation. This must be shown byshowing that the fuel system from the tank outletsto each engine is pressurized, under all intendedoperations, so as to prevent vapor formation, ormust be shown by climbing from the altitude of theairport elected by the applicant to the maximumaltitude established as an operating limitation un-der §25.1527. If a climb test is elected, there maybe no evidence of vapor lock or other malfunction-ing during the climb test conducted under the fol-lowing conditions:

(1) For reciprocating engine powered air-planes, the engines must operate at maximumcontinuous power, except that takeoff power mustbe used for the altitudes from 1,000 feet below thecritical altitude through the critical altitude. Thetime interval during which takeoff power is usedmay not be less than the takeoff time limitation.

(2) For turbine engine powered airplanes, theengines must operate at takeoff power for thetime interval selected for showing the takeoff flightpath, and at maximum continuous power for therest of the climb.

(3) The weight of the airplane must be theweight with full fuel tanks, minimum crew, and theballast necessary to maintain the center of gravitywithin allowable limits.

(4) The climb airspeed may not exceed—(i) For reciprocating engine powered airplanes,

the maximum airspeed established for climbingfrom takeoff to the maximum operating altitudewith the airplane in the following configuration:

(A) Landing gear retracted.(B) Wing flaps in the most favorable position.(C) Cowl flaps (or other means of controlling

the engine cooling supply) in the position that pro-vides adequate cooling in the hot-day condition.

(D) Engine operating within the maximum con-tinuous power limitations.


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(E) Maximum takeoff weight; and(ii) For turbine engine powered airplanes, the

maximum airspeed established for climbing fromtakeoff to the maximum operating altitude.

(5) The fuel temperature must be at least110°F.

(b) The test prescribed in paragraph (a) of thissection may be performed in flight or on theground under closely simulated flight conditions. Ifa flight test is performed in weather cold enoughto interfere with the proper conduct of the test, thefuel tank surfaces, fuel lines, and other fuel sys-tem parts subject to cold air must be insulated tosimulate, insofar as practicable, flight in hotweather.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–11, 32 FR 6912, May 5, 1967;Amdt. 25–57, 49 FR 6848, Feb. 23, 1984]

§25.963 Fuel tanks: general.(a) Each fuel tank must be able to withstand,

without failure, the vibration, inertia, fluid, andstructural loads that it may be subjected to in op-eration.

(b) Flexible fuel tank liners must be approvedor must be shown to be suitable for the particularapplication.

(c) Integral fuel tanks must have facilities for in-terior inspection and repair.

(d) Fuel tanks within the fuselage contour mustbe able to resist rupture and to retain fuel, underthe inertia forces prescribed for the emergencylanding conditions in §25.561. In addition, thesetanks must be in a protected position so that ex-posure of the tanks to scraping action with theground is unlikely.

(e) Fuel tank access covers must comply withthe following criteria in order to avoid loss of haz-ardous quantities of fuel:

(1) All covers located in an area where experi-ence or analysis indicates a strike is likely mustbe shown by analysis or tests to minimize pene-tration and deformation by tire fragments, low en-ergy engine debris, or other likely debris.

(2) All covers must be fire resistant as definedin part 1 of this chapter.

(f) For pressurized fuel tanks, a means with fail-safe features must be provided to prevent thebuildup of an excessive pressure difference be-tween the inside and the outside of the tank.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–40, 42 FR 15043, March 17,1977; Amdt. 25–69, 54 FR 40354, Sept. 29, 1989]

§25.965 Fuel tank tests.(a) It must be shown by tests that the fuel

tanks, as mounted in the airplane, can withstand,without failure or leakage, the more critical of thepressures resulting from the conditions specified

in paragraphs (a)(1) and (2) of this section. In ad-dition, it must be shown by either analysis or tests,that tank surfaces subjected to more critical pres-sures resulting from the condition of paragraphs(a)(3) and (4) of this section, are able to withstandthe following pressures:

(1) An internal pressure of 3.5 psi.(2) 125 percent of the maximum air pressure

developed in the tank from ram effect.(3) Fluid pressures developed during maximum

limit accelerations, and deflections, of the air-plane with a full tank.

(4) Fluid pressures developed during the mostadverse combination of airplane roll and fuel load.

(b) Each metallic tank with large unsupportedor unstiffened flat surfaces, whose failure or defor-mation could cause fuel leakage, must be able towithstand the following test, or its equivalent, with-out leakage or excessive deformation of the tankwalls:

(1) Each complete tank assembly and its sup-ports must be vibration tested while mounted tosimulate the actual installation.

(2) Except as specified in paragraph (b)(4) ofthis section, the tank assembly must be vibratedfor 25 hours at an amplitude of not less than 1⁄32of an inch (unless another amplitude is substanti-ated) while 2⁄3 filled with water or other suitabletest fluid.

(3) The test frequency of vibration must be asfollows:

(i) If no frequency of vibration resulting fromany r.p.m. within the normal operating range ofengine speeds is critical, the test frequency of vi-bration must be 2,000 cycles per minute.

(ii) If only one frequency of vibration resultingfrom any r.p.m. within the normal operating rangeof engine speeds is critical, that frequency of vi-bration must be the test frequency.

(iii) If more than one frequency of vibration re-sulting from any r.p.m. within the normal operat-ing range of engine speeds is critical, the mostcritical of these frequencies must be the test fre-quency.

(4) Under paragraphs (b)(3)(ii) and (iii) of thissection, the time of test must be adjusted to ac-complish the same number of vibration cyclesthat would be accomplished in 25 hours at thefrequency specified in paragraph (b)(3)(i) of thissection.

(5) During the test, the tank assembly must berocked at the rate of 16 to 20 complete cycles perminute, through an angle of 15° on both sides ofthe horizontal (30 total), about the most criticalaxis, for 25 hours. If motion about more than oneaxis is likely to be critical, the tank must be rockedabout each critical axis for 121⁄2 hours.

(c) Except where satisfactory operating experi-ence with a similar tank in a similar installation isshown, nonmetallic tanks must withstand the test


100 ASA

specified in paragraph (b)(5) of this section, withfuel at a temperature of 110°F. During this test, arepresentative specimen of the tank must be in-stalled in a supporting structure simulating the in-stallation in the airplane.

(d) For pressurized fuel tanks, it must be shownby analysis or tests that the fuel tanks can with-stand the maximum pressure likely to occur onthe ground or in flight.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–11, 32 FR 6913, May 5, 1967;Amdt. 25–40, 42 FR 15043, March 17, 1977]

§25.967 Fuel tank installations.(a) Each fuel tank must be supported so that

tank loads (resulting from the weight of the fuel inthe tanks) are not concentrated on unsupportedtank surfaces. In addition—

(1) There must be pads, if necessary, to pre-vent chafing between the tank and its supports;

(2) Padding must be nonabsorbent or treated toprevent the absorption of fluids;

(3) If a flexible tank liner is used, it must be sup-ported so that it is not required to withstand fluidloads; and

(4) Each interior surface of the tank compart-ment must be smooth and free of projections thatcould cause wear of the liner unless—

(i) Provisions are made for protection of theliner at these points; or

(ii) The construction of the liner itself providesthat protection.

(b) Spaces adjacent to tank surfaces must beventilated to avoid fume accumulation due to mi-nor leakage. If the tank is in a sealed compart-ment, ventilation may be limited to drain holeslarge enough to prevent excessive pressure re-sulting from altitude changes.

(c) The location of each tank must meet the re-quirements of §25.1185(a).

(d) No engine nacelle skin immediately behinda major air outlet from the engine compartmentmay act as the wall of an integral tank.

(e) Each fuel tank must be isolated from per-sonnel compartments by a fumeproof and fuel-proof enclosure.

§25.969 Fuel tank expansion space.Each fuel tank must have an expansion space

of not less than 2 percent of the tank capacity. Itmust be impossible to fill the expansion space in-advertently with the airplane in the normal groundattitude. For pressure fueling systems, compli-ance with this section may be shown with themeans provided to comply with §25.979(b).


§25.971 Fuel tank sump.(a) Each fuel tank must have a sump with an ef-

fective capacity, in the normal ground attitude, ofnot less than the greater of 0.10 percent of thetank capacity or one-sixteenth of a gallon unlessoperating limitations are established to ensurethat the accumulation of water in service will notexceed the sump capacity.

(b) Each fuel tank must allow drainage of anyhazardous quantity of water from any part of thetank to its sump with the airplane in the ground at-titude.

(c) Each fuel tank sump must have an accessi-ble drain that—

(1) Allows complete drainage of the sump onthe ground;

(2) Discharges clear of each part of the air-plane; and

(3) Has manual or automatic means for positivelocking in the closed position.

§25.973 Fuel tank filler connection.Each fuel tank filler connection must prevent

the entrance of fuel into any part of the airplaneother than the tank itself. In addition—

(a) [Reserved](b) Each recessed filler connection that can re-

tain any appreciable quantity of fuel must have adrain that discharges clear of each part of the air-plane;

(c) Each filler cap must provide a fuel-tightseal; and

(d) Each fuel filling point, except pressure fuel-ing connection points, must have a provision forelectrically bonding the airplane to ground fuelingequipment.


§25.975 Fuel tank vents and carburetor vapor vents.(a) Fuel tank vents. Each fuel tank must be

vented from the top part of the expansion spaceso that venting is effective under any normal flightcondition. In addition—

(1) Each vent must be arranged to avoid stop-page by dirt or ice formation;

(2) The vent arrangement must prevent siphon-ing of fuel during normal operation;

(3) The venting capacity and vent pressure lev-els must maintain acceptable differences of pres-sure between the interior and exterior of the tank,during—

(i) Normal flight operation;(ii) Maximum rate of ascent and descent; and(iii) Refueling and defueling (where applicable);(4) Airspaces of tanks with interconnected out-

lets must be interconnected;


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(5) There may be no point in any vent linewhere moisture can accumulate with the airplanein the ground attitude or the level flight attitude,unless drainage is provided; and

(6) No vent or drainage provision may end atany point—

(i) Where the discharge of fuel from the ventoutlet would constitute a fire hazard; or

(ii) From which fumes could enter personnelcompartments.

(b) Carburetor vapor vents. Each carburetorwith vapor elimination connections must have avent line to lead vapors back to one of the fueltanks. In addition—

(1) Each vent system must have means toavoid stoppage by ice; and

(2) If there is more than one fuel tank, and it isnecessary to use the tanks in a definite se-quence, each vapor vent return line must leadback to the fuel tank used for takeoff and landing.

§25.977 Fuel tank outlet.(a) There must be a fuel strainer for the fuel

tank outlet or for the booster pump. This strainermust—

(1) For reciprocating engine powered air-planes, have 8 to 16 meshes per inch; and

(2) For turbine engine powered airplanes, pre-vent the passage of any object that could restrictfuel flow or damage any fuel system component.

(b) [Reserved](c) The clear area of each fuel tank outlet

strainer must be at least five times the area of theoutlet line.

(d) The diameter of each strainer must be atleast that of the fuel tank outlet.

(e) Each finger strainer must be accessible forinspection and cleaning.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–11, 32 FR 6913, May 5, 1967;Amdt. 25–36, 39 FR 35460, Oct. 1, 1974]

§25.979 Pressure fueling system.For pressure fueling systems, the following

apply:(a) Each pressure fueling system fuel manifold

connection must have means to prevent the es-cape of hazardous quantities of fuel from the sys-tem if the fuel entry valve fails.

(b) An automatic shutoff means must be pro-vided to prevent the quantity of fuel in each tankfrom exceeding the maximum quantity approvedfor that tank. This means must—

(1) Allow checking for proper shutoff operationbefore each fueling of the tank; and

(2) Provide indication at each fueling station offailure of the shutoff means to stop the fuel flow atthe maximum quantity approved for that tank.

(c) A means must be provided to prevent dam-age to the fuel system in the event of failure of theautomatic shutoff means prescribed in paragraph(b) of this section.

(d) The airplane pressure fueling system (notincluding fuel tanks and fuel tank vents) mustwithstand an ultimate load that is 2.0 times theload arising from the maximum pressures, includ-ing surge, that is likely to occur during fueling. Themaximum surge pressure must be establishedwith any combination of tank valves being eitherintentionally or inadvertently closed.

(e) The airplane defueling system (not includ-ing fuel tanks and fuel tank vents) must withstandan ultimate load that is 2.0 times the load arisingfrom the maximum permissible defueling pres-sure (positive or negative) at the airplane fuelingconnection.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–11, 32 FR 6913, May 5, 1967;Amdt. 25–38, 41 FR 55467, Dec. 20, 1976; Amdt. 25–72, 55 FR 29785, July 20, 1990]

§25.981 Fuel tank explosion prevention.(a) No ignition source may be present at each

point in the fuel tank or fuel tank system wherecatastrophic failure could occur due to ignition offuel or vapors. This must be shown by:

(1) Determining the highest temperature allow-ing a safe margin below the lowest expected au-toignition temperature of the fuel in the fuel tanks.

(2) Demonstrating that no temperature at eachplace inside each fuel tank where fuel ignition ispossible will exceed the temperature determinedunder paragraph (a)(1) of this section. This mustbe verified under all probable operating, failure,and malfunction conditions of each componentwhose operation, failure, or malfunction could in-crease the temperature inside the tank.

(3) Demonstrating that an ignition source couldnot result from each single failure, from each sin-gle failure in combination with each latent failurecondition not shown to be extremely remote, andfrom all combinations of failures not shown to beextremely improbable. The effects of manufactur-ing variability, aging, wear, corrosion, and likelydamage must be considered.

(b) Except as provided in paragraphs (b)(2)and (c) of this section, no fuel tank Fleet AverageFlammability Exposure on an airplane may ex-ceed three percent of the Flammability ExposureEvaluation Time (FEET) as defined in Appendix Nof this part, or that of a fuel tank within the wing ofthe airplane model being evaluated, whichever isgreater. If the wing is not a conventional unheatedaluminum wing, the analysis must be based on anassumed Equivalent Conventional Unheated Alu-minum Wing Tank.


102 ASA

(1) Fleet Average Flammability Exposure is de-termined in accordance with Appendix N of thispart. The assessment must be done in accor-dance with the methods and procedures set forthin the Fuel Tank Flammability AssessmentMethod User’s Manual, dated May 2008, docu-ment number DOT/FAA/AR-05/8 (incorporated byreference, see §25.5).

(2) Any fuel tank other than a main fuel tank onan airplane must meet the flammability exposurecriteria of Appendix M to this part if any portion ofthe tank is located within the fuselage contour.

(3) As used in this paragraph,(i) Equivalent Conventional Unheated Alumi-

num Wing Tank is an integral tank in an unheatedsemi-monocoque aluminum wing of a subsonicairplane that is equivalent in aerodynamic perfor-mance, structural capability, fuel tank capacityand tank configuration to the designed wing.

(ii) Fleet Average Flammability Exposure is de-fined in Appendix N to this part and means thepercentage of time each fuel tank ullage is flam-mable for a fleet of an airplane type operatingover the range of flight lengths.

(iii) Main Fuel Tank means a fuel tank thatfeeds fuel directly into one or more engines andholds required fuel reserves continually through-out each flight.

(c) Paragraph (b) of this section does not applyto a fuel tank if means are provided to mitigate theeffects of an ignition of fuel vapors within that fueltank such that no damage caused by an ignitionwill prevent continued safe flight and landing.

(d) Critical design configuration control limita-tions (CDCCL), inspections, or other proceduresmust be established, as necessary, to prevent de-velopment of ignition sources within the fuel tanksystem pursuant to paragraph (a) of this section,to prevent increasing the flammability exposure ofthe tanks above that permitted under paragraph(b) of this section, and to prevent degradation ofthe performance and reliability of any means pro-vided according to paragraphs (a) or (c) of thissection. These CDCCL, inspections, and proce-dures must be included in the Airworthiness Limi-tations section of the instructions for continuedairworthiness required by §25.1529. Visiblemeans of identifying critical features of the designmust be placed in areas of the airplane whereforeseeable maintenance actions, repairs, or al-terations may compromise the critical design con-figuration control limitations (e.g., color-coding ofwire to identify separation limitation). These visi-ble means must also be identified as CDCCL.

[Docket No. FAA–1999–6411, 66 FR 23130, May 7,2001; as amended by Amdt. 25–125, 73 FR 42494, July21, 2008]

FUEL SYSTEM COMPONENTS

§25.991 Fuel pumps.(a) Main pumps. Each fuel pump required for

proper engine operation, or required to meet thefuel system requirements of this subpart (otherthan those in paragraph (b) of this section, is amain pump. For each main pump, provision mustbe made to allow the bypass of each positive dis-placement fuel pump other than a fuel injectionpump (a pump that supplies the proper flow andpressure for fuel injection when the injection is notaccomplished in a carburetor) approved as part ofthe engine.

(b) Emergency pumps. There must be emer-gency pumps or another main pump to feed eachengine immediately after failure of any main pump(other than a fuel injection pump approved as partof the engine).

§25.993 Fuel system lines and fittings.(a) Each fuel line must be installed and sup-

ported to prevent excessive vibration and to with-stand loads due to fuel pressure and acceleratedflight conditions.

(b) Each fuel line connected to components ofthe airplane between which relative motion couldexist must have provisions for flexibility.

(c) Each flexible connection in fuel lines thatmay be under pressure and subjected to axialloading must use flexible hose assemblies.

(d) Flexible hose must be approved or must beshown to be suitable for the particular application.

(e) No flexible hose that might be adversely af-fected by exposure to high temperatures may beused where excessive temperatures will exist dur-ing operation or after engine shut-down.

(f) Each fuel line within the fuselage must be de-signed and installed to allow a reasonable degreeof deformation and stretching without leakage.


§25.994 Fuel system components.Fuel system components in an engine nacelle

or in the fuselage must be protected from damagewhich could result in spillage of enough fuel toconstitute a fire hazard as a result of a wheels-uplanding on a paved runway.


§25.995 Fuel valves.In addition to the requirements of §25.1189 for

shutoff means, each fuel valve must—(a) [Reserved](b) Be supported so that no loads resulting

from their operation or from accelerated flight


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conditions are transmitted to the lines attached tothe valve.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964;amended by Amdt. 25–40, 42 FR 15043, Mar. 17, 1977]

§25.997 Fuel strainer or filter.There must be a fuel strainer or filter between

the fuel tank outlet and the inlet of either the fuelmetering device or an engine driven positive dis-placement pump, whichever is nearer the fueltank outlet. This fuel strainer or filter must—

(a) Be accessible for draining and cleaning andmust incorporate a screen or element which iseasily removable;

(b) Have a sediment trap and drain except thatit need not have a drain if the strainer or filter iseasily removable for drain purposes;

(c) Be mounted so that its weight is not sup-ported by the connecting lines or by the inlet oroutlet connections of the strainer or filter itself, un-less adequate strength margins under all loadingconditions are provided in the lines and connec-tions; and

(d) Have the capacity (with respect to operatinglimitations established for the engine) to ensurethat engine fuel system functioning is not im-paired, with the fuel contaminated to a degree(with respect to particle size and density) that isgreater than that established for the engine in Part33 of this chapter.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–36, 39 FR 35460, Oct. 1, 1974;Amdt. 25–57, 49 FR 6848, Feb. 23, 1984]

§25.999 Fuel system drains.(a) Drainage of the fuel system must be accom-

plished by the use of fuel strainer and fuel tanksump drains.

(b) Each drain required by paragraph (a) of thissection must—

(1) Discharge clear of all parts of the airplane;(2) Have manual or automatic means for posi-

tive locking in the closed position; and(3) Have a drain valve—(i) That is readily accessible and which can be

easily opened and closed; and(ii) That is either located or protected to prevent

fuel spillage in the event of a landing with landinggear retracted.


§25.1001 Fuel jettisoning system.(a) A fuel jettisoning system must be installed

on each airplane unless it is shown that the air-plane meets the climb requirements of §§25.119and 25.121(d) at maximum takeoff weight, lessthe actual or computed weight of fuel necessary

for a 15-minute flight comprised of a takeoff, go-around, and landing at the airport of departurewith the airplane configuration, speed, power, andthrust the same as that used in meeting the appli-cable takeoff, approach, and landing climb perfor-mance requirements of this part.

(b) If a fuel jettisoning system is required itmust be capable of jettisoning enough fuel within15 minutes, starting with the weight given in para-graph (a) of this section, to enable the airplane tomeet the climb requirements of §§25.119 and25.121(d), assuming that the fuel is jettisoned un-der the conditions, except weight, found least fa-vorable during the flight tests prescribed in para-graph (c) of this section.

(c) Fuel jettisoning must be demonstrated be-ginning at maximum takeoff weight with flaps andlanding gear up and in—

(1) A power-off glide at 1.3 VSR1;(2) A climb at the one-engine inoperative best

rate-of-climb speed, with the critical engine inop-erative and the remaining engines at maximumcontinuous power; and

(3) Level flight at 1.3 VSR1; if the results of thetests in the conditions specified in paragraphs (c)(1) and (2) of this section show that this conditioncould be critical.

(d) During the flight tests prescribed in para-graph (c) of this section, it must be shown that—

(1) The fuel jettisoning system and its operationare free from fire hazard;

(2) The fuel discharges clear of any part of theairplane;

(3) Fuel or fumes do not enter any parts of theairplane; and

(4) The jettisoning operation does not ad-versely affect the controllability of the airplane.

(e) For reciprocating engine powered air-planes, means must be provided to prevent jetti-soning the fuel in the tanks used for takeoff andlanding below the level allowing 45 minutes flightat 75 percent maximum continuous power. How-ever, if there is an auxiliary control independent ofthe main jettisoning control, the system may bedesigned to jettison the remaining fuel by meansof the auxiliary jettisoning control.

(f) For turbine engine powered airplanes,means must be provided to prevent jettisoning thefuel in the tanks used for takeoff and landing be-low the level allowing climb from sea level to10,000 feet and thereafter allowing 45 minutescruise at a speed for maximum range. However, ifthere is an auxiliary control independent of themain jettisoning control, the system may be de-signed to jettison the remaining fuel by means ofthe auxiliary jettisoning control.

(g) The fuel jettisoning valve must be designedto allow flight personnel to close the valve duringany part of the jettisoning operation.


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(h) Unless it is shown that using any means (in-cluding flaps, slots, and slats) for changing the air-flow across or around the wings does not ad-versely affect fuel jettisoning, there must be aplacard, adjacent to the jettisoning control, towarn flight crewmembers against jettisoning fuelwhile the means that change the airflow are beingused.

(i) The fuel jettisoning system must be de-signed so that any reasonably probable singlemalfunction in the system will not result in a haz-ardous condition due to unsymmetrical jettisoningof, or inability to jettison, fuel.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–18, 33 FR 12226, Aug. 30, 1968;Amdt. 25–57, 49 FR 6848, Feb. 23, 1984; Amdt. 25–108,67 FR 70828, Nov. 26, 2002]

OIL SYSTEM

§25.1011 General.(a) Each engine must have an independent oil

system that can supply it with an appropriatequantity of oil at a temperature not above that safefor continuous operation.

(b) The usable oil capacity may not be lessthan the product of the endurance of the airplaneunder critical operating conditions and the ap-proved maximum allowable oil consumption of theengine under the same conditions, plus a suitablemargin to ensure system circulation. Instead of arational analysis of airplane range for the purposeof computing oil requirements for reciprocatingengine powered airplanes, the following fuel/oilratios may be used:

(1) For airplanes without a reserve oil or oiltransfer system, a fuel/oil ratio of 30:1 by volume.

(2) For airplanes with either a reserve oil or oiltransfer system, a fuel/oil ratio of 40:1 by volume.

(c) Fuel/oil ratios higher than those prescribedin paragraphs (b) (1) and (2) of this section maybe used if substantiated by data on actual engineoil consumption.

§25.1013 Oil tanks.(a) Installation. Each oil tank installation must

meet the requirements of §25.967.(b) Expansion space. Oil tank expansion space

must be provided as follows:(1) Each oil tank used with a reciprocating en-

gine must have an expansion space of not lessthan the greater of 10 percent of the tank capacityor 0.5 gallon, and each oil tank used with a tur-bine engine must have an expansion space of notless than 10 percent of the tank capacity.

(2) Each reserve oil tank not directly connectedto any engine may have an expansion space ofnot less than two percent of the tank capacity.

(3) It must be impossible to fill the expansionspace inadvertently with the airplane in the nor-mal ground attitude.

(c) Filler connection. Each recessed oil tankfiller connection that can retain any appreciablequantity of oil must have a drain that dischargesclear of each part of the airplane. In addition, eachoil tank filler cap must provide an oil-tight seal.

(d) Vent. Oil tanks must be vented as follows:(1) Each oil tank must be vented from the top

part of the expansion space so that venting is ef-fective under any normal flight condition.

(2) Oil tank vents must be arranged so thatcondensed water vapor that might freeze and ob-struct the line cannot accumulate at any point.

(e) Outlet. There must be means to prevent en-trance into the tank itself, or into the tank outlet, ofany object that might obstruct the flow of oilthrough the system. No oil tank outlet may be en-closed by any screen or guard that would reducethe flow of oil below a safe value at any operatingtemperature. There must be a shutoff valve at theoutlet of each oil tank used with a turbine engine,unless the external portion of the oil system (in-cluding the oil tank supports) is fireproof.

(f) Flexible oil tank liners. Each flexible oil tankliner must be approved or must be shown to besuitable for the particular application.

[Docket No. 5066, 29 FR 18291, Dec. 24; as amendedby Amdt. 25–19, 33 FR 15410, Oct. 17, 1968; Amdt. 25–23, 35 FR 5677, April 8, 1970; Amdt. 25–36, 39 FR35460, Oct. 1, 1974; Amdt. 25–57, 49 FR 6848, Feb. 23,1984; Amdt. 25–72, 55 FR 29785, July 20, 1990]

§25.1015 Oil tank tests.Each oil tank must be designed and installed

so that—(a) It can withstand, without failure, each vibra-

tion, inertia, and fluid load that it may be sub-jected to in operation; and

(b) It meets the provisions of §25.965, except—(1) The test pressure—(i) For pressurized tanks used with a turbine

engine, may not be less than 5 p.s.i. plus the max-imum operating pressure of the tank instead ofthe pressure specified in §25.965(a); and

(ii) For all other tanks may not be less than 5p.s.i. instead of the pressure specified in§25.965(a); and

(2) The test fluid must be oil at 250°F instead ofthe fluid specified in §25.965(c).


§25.1017 Oil lines and fittings.(a) Each oil line must meet the requirements of

§25.993 and each oil line and fitting in any desig-nated fire zone must meet the requirements of§25.1183.


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(b) Breather lines must be arranged so that—(1) Condensed water vapor that might freeze

and obstruct the line cannot accumulate at anypoint;

(2) The breather discharge does not constitutea fire hazard if foaming occurs or causes emittedoil to strike the pilot’s windshield; and

(3) The breather does not discharge into theengine air induction system.

§25.1019 Oil strainer or filter.(a) Each turbine engine installation must incor-

porate an oil strainer or filter through which all ofthe engine oil flows and which meets the followingrequirements:

(1) Each oil strainer or filter that has a bypassmust be constructed and installed so that oil willflow at the normal rate through the rest of the sys-tem with the strainer or filter completely blocked.

(2) The oil strainer or filter must have the ca-pacity (with respect to operating limitations estab-lished for the engine) to ensure that engine oilsystem functioning is not impaired when the oil iscontaminated to a degree (with respect to particlesize and density) that is greater than that estab-lished for the engine under Part 33 of this chapter.

(3) The oil strainer or filter, unless it is installedat an oil tank outlet, must incorporate an indicatorthat will indicate contamination before it reachesthe capacity established in accordance with para-graph (a)(2) of this section.

(4) The bypass of a strainer or filter must beconstructed and installed so that the release ofcollected contaminants is minimized by appropri-ate location of the bypass to ensure that collectedcontaminants are not in the bypass flow path.

(5) An oil strainer or filter that has no bypass,except one that is installed at an oil tank outlet,must have a means to connect it to the warningsystem required in §25.1305(c)(7).

(b) Each oil strainer or filter in a powerplant in-stallation using reciprocating engines must beconstructed and installed so that oil will flow at thenormal rate through the rest of the system withthe strainer or filter element completely blocked.


§25.1021 Oil system drains.A drain (or drains) must be provided to allow

safe drainage of the oil system. Each drainmust—

(a) Be accessible; and(b) Have manual or automatic means for posi-

tive locking in the closed position.


§25.1023 Oil radiators.(a) Each oil radiator must be able to withstand,

without failure, any vibration, inertia, and oil pres-sure load to which it would be subjected in opera-tion.

(b) Each oil radiator air duct must be located sothat, in case of fire, flames coming from normalopenings of the engine nacelle cannot impinge di-rectly upon the radiator.

§25.1025 Oil valves.(a) Each oil shutoff must meet the require-

ments of §25.1189.(b) The closing of oil shutoff means may not

prevent propeller feathering.(c) Each oil valve must have positive stops or

suitable index provisions in the “on” and “off” posi-tions and must be supported so that no loads re-sulting from its operation or from acceleratedflight conditions are transmitted to the lines at-tached to the valve.

§25.1027 Propeller feathering system.(a) If the propeller feathering system depends

on engine oil, there must be means to trap anamount of oil in the tank if the supply becomesdepleted due to failure of any part of the lubricat-ing system other than the tank itself.

(b) The amount of trapped oil must be enoughto accomplish the feathering operation and mustbe available only to the feathering pump.

(c) The ability of the system to accomplishfeathering with the trapped oil must be shown.This may be done on the ground using an auxil-iary source of oil for lubricating the engine duringoperation.

(d) Provision must be made to prevent sludgeor other foreign matter from affecting the safe op-eration of the propeller feathering system.


COOLING

§25.1041 General.The powerplant and auxiliary power unit cooling

provisions must be able to maintain the tempera-tures of powerplant components, engine fluids,and auxiliary power unit components and fluidswithin the temperature limits established for thesecomponents and fluids, under ground, water, andflight operating conditions, and after normal en-gine or auxiliary power unit shutdown, or both.


§25.1043 Cooling tests.(a) General. Compliance with §25.1041 must

be shown by tests, under critical ground, water,


106 ASA

and flight operating conditions. For these tests,the following apply:

(1) If the tests are conducted under conditionsdeviating from the maximum ambient atmo-spheric temperature, the recorded powerplanttemperatures must be corrected under para-graphs (c) and (d) of this section.

(2) No corrected temperatures determined un-der paragraph (a)(1) of this section may exceedestablished limits.

(3) For reciprocating engines, the fuel usedduring the cooling tests must be the minimumgrade approved for the engines, and the mixturesettings must be those normally used in the flightstages for which the cooling tests are conducted.The test procedures must be as prescribed in§25.1045.

(b) Maximum ambient atmospheric tempera-ture. A maximum ambient atmospheric tempera-ture corresponding to sea level conditions of atleast 100 degrees F must be established. The as-sumed temperature lapse rate is 3.6 degrees Fper thousand feet of altitude above sea level untila temperature of -69.7 degrees F is reached,above which altitude the temperature is consid-ered constant at - 69.7 degrees F. However, forwinterization installations, the applicant may se-lect a maximum ambient atmospheric tempera-ture corresponding to sea level conditions of lessthan 100 degrees F.

(c) Correction factor (except cylinder barrels).Unless a more rational correction applies, temper-atures of engine fluids and powerplant compo-nents (except cylinder barrels) for which tempera-ture limits are established, must be corrected byadding to them the difference between the maxi-mum ambient atmospheric temperature and thetemperature of the ambient air at the time of thefirst occurrence of the maximum component orfluid temperature recorded during the cooling test.

(d) Correction factor for cylinder barrel temper-atures. Unless a more rational correction applies,cylinder barrel temperatures must be corrected byadding to them 0.7 times the difference betweenthe maximum ambient atmospheric temperatureand the temperature of the ambient air at the timeof the first occurrence of the maximum cylinderbarrel temperature recorded during the coolingtest.


§25.1045 Cooling test procedures.(a) Compliance with §25.1041 must be shown

for the takeoff, climb, en route, and landing stagesof flight that correspond to the applicable perfor-mance requirements. The cooling tests must beconducted with the airplane in the configuration,and operating under the conditions, that are criti-

cal relative to cooling during each stage of flight.For the cooling tests, a temperature is “stabilized”when its rate of change is less than two degrees Fper minute.

(b) Temperatures must be stabilized under theconditions from which entry is made into eachstage of flight being investigated, unless the entrycondition normally is not one during which com-ponent and the engine fluid temperatures wouldstabilize (in which case, operation through the fullentry condition must be conducted before entryinto the stage of flight being investigated in orderto allow temperatures to reach their natural levelsat the time of entry). The takeoff cooling test mustbe preceded by a period during which the power-plant component and engine fluid temperaturesare stabilized with the engines at ground idle.

(c) Cooling tests for each stage of flight mustbe continued until—

(1) The component and engine fluid tempera-tures stabilize;

(2) The stage of flight is completed; or(3) An operating limitation is reached.(d) For reciprocating engine powered air-

planes, it may be assumed, for cooling test pur-poses, that the takeoff stage of flight is completewhen the airplane reaches an altitude of 1,500feet above the takeoff surface or reaches a pointin the takeoff where the transition from the takeoffto the en route configuration is completed and aspeed is reached at which compliance with§25.121(c) is shown, whichever point is at ahigher altitude. The airplane must be in the follow-ing configuration:

(1) Landing gear retracted.(2) Wing flaps in the most favorable position.(3) Cowl flaps (or other means of controlling the

engine cooling supply) in the position that pro-vides adequate cooling in the hot-day condition.

(4) Critical engine inoperative and its propellerstopped.

(5) Remaining engines at the maximum contin-uous power available for the altitude.

(e) For hull seaplanes and amphibians, coolingmust be shown during taxiing downwind for 10minutes, at five knots above step speed.


INDUCTION SYSTEM

§25.1091 Air induction.(a) The air induction system for each engine

and auxiliary power unit must supply—(1) The air required by that engine and auxiliary

power unit under each operating condition forwhich certification is requested; and

(2) The air for proper fuel metering and mixturedistribution with the induction system valves inany position.


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(b) Each reciprocating engine must have an al-ternate air source that prevents the entry of rain,ice, or any other foreign matter.

(c) Air intakes may not open within the cowling,unless—

(1) That part of the cowling is isolated from theengine accessory section by means of a fireproofdiaphragm; or

(2) For reciprocating engines, there are meansto prevent the emergence of backfire flames.

(d) For turbine engine powered airplanes andairplanes incorporating auxiliary power units—

(1) There must be means to prevent hazardousquantities of fuel leakage or overflow from drains,vents, or other components of flammable fluidsystems from entering the engine or auxiliarypower unit intake system; and

(2) The airplane must be designed to preventwater or slush on the runway, taxiway, or other air-port operating surfaces from being directed intothe engine or auxiliary power unit air inlet ducts inhazardous quantities, and the air inlet ducts mustbe located or protected so as to minimize the in-gestion of foreign matter during takeoff, landing,and taxiing.

(e) If the engine induction system containsparts or components that could be damaged byforeign objects entering the air inlet, it must beshown by tests or, if appropriate, by analysis thatthe induction system design can withstand theforeign object ingestion test conditions of§§33.76, 33.77 and 33.78(a)(1) of this chapterwithout failure of parts or components that couldcreate a hazard.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–38, 41 FR 55467, Dec. 20, 1976;Amdt. 25–40, 42 FR 15043, March 17, 1977; Amdt. 25–57, 49 FR 6849, Feb. 23, 1984; Amdt. 25–100, 65 FR55854, Sept. 14, 2000]

§25.1093 Induction system icing protection.(a) Reciprocating engines. Each reciprocating

engine air induction system must have means toprevent and eliminate icing. Unless this is done byother means, it must be shown that, in air free ofvisible moisture at a temperature of 30°F, eachairplane with altitude engines using—

(1) Conventional venturi carburetors have apreheater that can provide a heat rise of 120°Fwith the engine at 60 percent of maximum contin-uous power; or

(2) Carburetors tending to reduce the probabil-ity of ice formation has a preheater that can pro-vide a heat rise of 100°F with the engine at 60percent of maximum continuous power.

(b) Turbine engines.(1) Each turbine engine must operate through-

out the flight power range of the engine (including

idling), without the accumulation of ice on the en-gine, inlet system components, or airframe com-ponents that would adversely affect engine opera-tion or cause a serious loss of power or thrust—

(i) Under the icing conditions specified in Ap-pendix C, and

(ii) In falling and blowing snow within the limita-tions established for the airplane for such opera-tion.

(2) Each turbine engine must idle for 30 min-utes on the ground, with the air bleed available forengine icing protection at its critical condition,without adverse effect, in an atmosphere that is ata temperature between 15° and 30°F (between-9° and -1°C) and has a liquid water content notless than 0.3 grams per cubic meter in the form ofdrops having a mean effective diameter not lessthan 20 microns, followed by momentary opera-tion at takeoff power or thrust. During the 30 min-utes of idle operation, the engine may be run upperiodically to a moderate power or thrust settingin a manner acceptable to the Administrator.

(c) Supercharged reciprocating engines. Foreach engine having a supercharger to pressurizethe air before it enters the carburetor, the heat risein the air caused by that supercharging at any alti-tude may be utilized in determining compliancewith paragraph (a) of this section if the heat riseutilized is that which will be available, automati-cally, for the applicable altitude and operating con-dition because of supercharging.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–38, 41 FR 55467, Dec. 20, 1976;Amdt. 25–40, 42 FR 15043, March 17, 1977; Amdt. 25–57, 49 FR 6849, Feb. 23, 1984; Amdt. 25–72, 55 FR29785, July 20, 1990]

§25.1101 Carburetor air preheater design.Each carburetor air preheater must be de-

signed and constructed to—(a) Ensure ventilation of the preheater when

the engine is operated in cold air;(b) Allow inspection of the exhaust manifold

parts that it surrounds; and(c) Allow inspection of critical parts of the pre-

heater itself.

§25.1103 Induction system ducts and air duct systems.(a) Each induction system duct upstream of the

first stage of the engine supercharger and of theauxiliary power unit compressor must have adrain to prevent the hazardous accumulation offuel and moisture in the ground attitude. No drainmay discharge where it might cause a fire hazard.

(b) Each induction system duct must be—


108 ASA

(1) Strong enough to prevent induction systemfailures resulting from normal backfire conditions;and

(2) Fire-resistant if it is in any fire zone forwhich a fire-extinguishing system is required, ex-cept that ducts for auxiliary power units must befireproof within the auxiliary power unit fire zone.

(c) Each duct connected to components be-tween which relative motion could exist must havemeans for flexibility.

(d) For turbine engine and auxiliary power unitbleed air duct systems, no hazard may result if aduct failure occurs at any point between the airduct source and the airplane unit served by the air.

(e) Each auxiliary power unit induction systemduct must be fireproof for a sufficient distance up-stream of the auxiliary power unit compartment toprevent hot gas reverse flow from burning throughauxiliary power unit ducts and entering any othercompartment or area of the airplane in which ahazard would be created resulting from the entryof hot gases. The materials used to form the re-mainder of the induction system duct and plenumchamber of the auxiliary power unit must be capa-ble of resisting the maximum heat conditionslikely to occur.

(f) Each auxiliary power unit induction systemduct must be constructed of materials that will notabsorb or trap hazardous quantities of flammablefluids that could be ignited in the event of a surgeor reverse flow condition.


§25.1105 Induction system screens.If induction system screens are used—(a) Each screen must be upstream of the car-

buretor;(b) No screen may be in any part of the induc-

tion system that is the only passage throughwhich air can reach the engine, unless it can bedeiced by heated air;

(c) No screen may be deiced by alcohol alone;and

(d) It must be impossible for fuel to strike anyscreen.

§25.1107 Inter-coolers and after-coolers.Each inter-cooler and after-cooler must be

able to withstand any vibration, inertia, and airpressure load to which it would be subjected inoperation.

EXHAUST SYSTEM

§25.1121 General.For powerplant and auxiliary power unit instal-

lations the following apply:

(a) Each exhaust system must ensure safe dis-posal of exhaust gases without fire hazard or car-bon monoxide contamination in any personnelcompartment. For test purposes, any acceptablecarbon monoxide detection method may be usedto show the absence of carbon monoxide.

(b) Each exhaust system part with a surfacehot enough to ignite flammable fluids or vaporsmust be located or shielded so that leakage fromany system carrying flammable fluids or vaporswill not result in a fire caused by impingement ofthe fluids or vapors on any part of the exhaustsystem including shields for the exhaust system.

(c) Each component that hot exhaust gasescould strike, or that could be subjected to hightemperatures from exhaust system parts, must befireproof. All exhaust system components must beseparated by fireproof shields from adjacent partsof the airplane that are outside the engine andauxiliary power unit compartments.

(d) No exhaust gases may discharge so as tocause a fire hazard with respect to any flammablefluid vent or drain.

(e) No exhaust gases may discharge wherethey will cause a glare seriously affecting pilot vi-sion at night.

(f) Each exhaust system component must beventilated to prevent points of excessively hightemperature.

(g) Each exhaust shroud must be ventilated orinsulated to avoid, during normal operation, atemperature high enough to ignite any flammablefluids or vapors external to the shroud.


§25.1123 Exhaust piping.For powerplant and auxiliary power unit instal-

lations, the following apply:(a) Exhaust piping must be heat and corrosion

resistant, and must have provisions to prevent fail-ure due to expansion by operating temperatures.

(b) Piping must be supported to withstand anyvibration and inertia loads to which it would besubjected in operation; and

(c) Piping connected to components betweenwhich relative motion could exist must havemeans for flexibility.


§25.1125 Exhaust heat exchangers.For reciprocating engine powered airplanes,

the following apply:(a) Each exhaust heat exchanger must be con-

structed and installed to withstand each vibration,


ASA 109

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inertia, and other load to which it would be sub-jected in operation. In addition—

(1) Each exchanger must be suitable for contin-ued operation at high temperatures and resistantto corrosion from exhaust gases;

(2) There must be means for the inspection ofthe critical parts of each exchanger;

(3) Each exchanger must have cooling provi-sions wherever it is subject to contact with ex-haust gases; and

(4) No exhaust heat exchanger or muff mayhave any stagnant areas or liquid traps that wouldincrease the probability of ignition of flammablefluids or vapors that might be present in case ofthe failure or malfunction of components carryingflammable fluids.

(b) If an exhaust heat exchanger is used forheating ventilating air—

(1) There must be a secondary heat exchangerbetween the primary exhaust gas heat exchangerand the ventilating air system; or

(2) Other means must be used to preclude theharmful contamination of the ventilating air.


§25.1127 Exhaust driven turbo-superchargers.(a) Each exhaust driven turbo-supercharger

must be approved or shown to be suitable for theparticular application. It must be installed andsupported to ensure safe operation between nor-mal inspections and overhauls. In addition, theremust be provisions for expansion and flexibilitybetween exhaust conduits and the turbine.

(b) There must be provisions for lubricating theturbine and for cooling turbine parts where tem-peratures are critical.

(c) If the normal turbo-supercharger controlsystem malfunctions, the turbine speed may notexceed its maximum allowable value. Except forthe waste gate operating components, the com-ponents provided for meeting this requirementmust be independent of the normal turbo-super-charger controls.

POWERPLANT CONTROLS AND ACCESSORIES

§25.1141 Powerplant controls: general.Each powerplant control must be located, ar-

ranged, and designed under §§25.777 through25.781 and marked under §25.1555. In addition, itmust meet the following requirements:

(a) Each control must be located so that it can-not be inadvertently operated by persons enter-ing, leaving, or moving normally in, the cockpit.

(b) Each flexible control must be approved ormust be shown to be suitable for the particular ap-plication.

(c) Each control must have sufficient strengthand rigidity to withstand operating loads withoutfailure and without excessive deflection.

(d) Each control must be able to maintain anyset position without constant attention by flightcrewmembers and without creep due to controlloads or vibration.

(e) The portion of each powerplant control lo-cated in a designated fire zone that is required tobe operated in the event of fire must be at leastfire resistant.

(f) Powerplant valve controls located in thecockpit must have—

(1) For manual valves, positive stops or in thecase of fuel valves suitable index provisions, inthe open and closed position; and

(2) For power-assisted valves, a means to indi-cate to the flight crew when the valve—

(i) Is in the fully open or fully closed position; or(ii) Is moving between the fully open and fully

closed position.


§25.1142 Auxiliary power unit controls.Means must be provided on the flight deck for

starting, stopping, and emergency shutdown ofeach installed auxiliary power unit.


§25.1143 Engine controls.(a) There must be a separate power or thrust

control for each engine.(b) Power and thrust controls must be arranged

to allow—(1) Separate control of each engine; and(2) Simultaneous control of all engines.(c) Each power and thrust control must provide

a positive and immediately responsive means ofcontrolling its engine.

(d) For each fluid injection (other than fuel) sys-tem and its controls not provided and approved aspart of the engine, the applicant must show thatthe flow of the injection fluid is adequately con-trolled.

(e) If a power or thrust control incorporates afuel shutoff feature, the control must have ameans to prevent the inadvertent movement ofthe control into the shutoff position. The meansmust—

(1) Have a positive lock or stop at the idle posi-tion; and

(2) Require a separate and distinct operation toplace the control in the shutoff position.


110 ASA

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5677, April 8, 1970;Amdt. 25–38, 41 FR 55467, Dec. 20, 1976; Amdt. 25–57, 49 FR 6849, Feb. 23, 1984]

§25.1145 Ignition switches.(a) Ignition switches must control each engine

ignition circuit on each engine.(b) There must be means to quickly shut off all

ignition by the grouping of switches or by a masterignition control.

(c) Each group of ignition switches, except igni-tion switches for turbine engines for which contin-uous ignition is not required, and each master ig-nition control must have a means to prevent its in-advertent operation.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–40, 42 FR 15044 March 17, 1977]

§25.1147 Mixture controls.(a) If there are mixture controls, each engine

must have a separate control. The controls mustbe grouped and arranged to allow—

(1) Separate control of each engine; and(2) Simultaneous control of all engines.(b) Each intermediate position of the mixture

controls that corresponds to a normal operatingsetting must be identifiable by feel and sight.

(c) The mixture controls must be accessible toboth pilots. However, if there is a separate flightengineer station with a control panel, the controlsneed be accessible only to the flight engineer.

§25.1149 Propeller speed and pitch controls.(a) There must be a separate propeller speed

and pitch control for each propeller.(b) The controls must be grouped and arranged

to allow—(1) Separate control of each propeller; and(2) Simultaneous control of all propellers.(c) The controls must allow synchronization of

all propellers.(d) The propeller speed and pitch controls must

be to the right of, and at least one inch below, thepilot’s throttle controls.

§25.1153 Propeller feathering controls.(a) There must be a separate propeller feather-

ing control for each propeller. The control musthave means to prevent its inadvertent operation.

(b) If feathering is accomplished by movementof the propeller pitch or speed control lever, theremust be means to prevent the inadvertent move-ment of this lever to the feathering position duringnormal operation.


§25.1155 Reverse thrust and propeller pitch settings below the flight regime.Each control for reverse thrust and for propeller

pitch settings below the flight regime must havemeans to prevent its inadvertent operation. Themeans must have a positive lock or stop at theflight idle position and must require a separateand distinct operation by the crew to displace thecontrol from the flight regime (forward thrust re-gime for turbojet powered airplanes).


§25.1157 Carburetor air temperature controls.There must be a separate carburetor air tem-

perature control for each engine.

§25.1159 Supercharger controls.Each supercharger control must be accessible

to the pilots or, if there is a separate flight engineerstation with a control panel, to the flight engineer.

§25.1161 Fuel jettisoning system controls.Each fuel jettisoning system control must have

guards to prevent inadvertent operation. No con-trol may be near any fire extinguisher control orother control used to combat fire.

§25.1163 Powerplant accessories.(a) Each engine mounted accessory must—(1) Be approved for mounting on the engine in-

volved;(2) Use the provisions on the engine for mount-

ing; and(3) Be sealed to prevent contamination of the

engine oil system and the accessory system.(b) Electrical equipment subject to arcing or

sparking must be installed to minimize the proba-bility of contact with any flammable fluids or va-pors that might be present in a free state.

(c) If continued rotation of an engine-drivencabin supercharger or of any remote accessorydriven by the engine is hazardous if malfunction-ing occurs, there must be means to prevent rota-tion without interfering with the continued opera-tion of the engine.


§25.1165 Engine ignition systems.(a) Each battery ignition system must be sup-

plemented by a generator that is automaticallyavailable as an alternate source of electrical en-ergy to allow continued engine operation if anybattery becomes depleted.


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(b) The capacity of batteries and generatorsmust be large enough to meet the simultaneousdemands of the engine ignition system and thegreatest demands of any electrical system com-ponents that draw electrical energy from thesame source.

(c) The design of the engine ignition systemmust account for—

(1) The condition of an inoperative generator;(2) The condition of a completely depleted bat-

tery with the generator running at its normal oper-ating speed; and

(3) The condition of a completely depleted bat-tery with the generator operating at idling speed, ifthere is only one battery.

(d) Magneto ground wiring (for separate igni-tion circuits) that lies on the engine side of the firewall, must be installed, located, or protected, tominimize the probability of simultaneous failure oftwo or more wires as a result of mechanical dam-age, electrical faults, or other cause.

(e) No ground wire for any engine may berouted through a fire zone of another engine un-less each part of that wire within that zone is fire-proof.

(f) Each ignition system must be independentof any electrical circuit, not used for assisting,controlling, or analyzing the operation of thatsystem.

(g) There must be means to warn appropriateflight crewmembers if the malfunctioning of anypart of the electrical system is causing the contin-uous discharge of any battery necessary for en-gine ignition.

(h) Each engine ignition system of a turbinepowered airplane must be considered an essen-tial electrical load.


§25.1167 Accessory gearboxes.For airplanes equipped with an accessory gear-

box that is not certificated as part of an engine—(a) The engine with gearbox and connecting

transmissions and shafts attached must be sub-jected to the tests specified in §33.49 or §33.87 ofthis chapter, as applicable;

(b) The accessory gearbox must meet the re-quirements of §§33.25 and 33.53 or 33.91 of thischapter, as applicable; and

(c) Possible misalignments and torsional load-ings of the gearbox, transmission, and shaft sys-tem, expected to result under normal operatingconditions must be evaluated.


POWERPLANT FIRE PROTECTION

§25.1181 Designated fire zones; regions included.(a) Designated fire zones are—(1) The engine power section;(2) The engine accessory section;(3) Except for reciprocating engines, any com-

plete powerplant compartment in which no isola-tion is provided between the engine power sectionand the engine accessory section;

(4) Any auxiliary power unit compartment;(5) Any fuel-burning heater and other combus-

tion equipment installation described in §25.859;(6) The compressor and accessory sections of

turbine engines; and(7) Combustor, turbine, and tailpipe sections of

turbine engine installations that contain lines orcomponents carrying flammable fluids or gases.

(b) Each designated fire zone must meet therequirements of §§25.867, and 25.1185 through25.1203.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–11, 32 FR 6913, May 5, 1967;Amdt. 25–23, 35 FR 5677, April 8, 1970; Amdt. 25–72,55 FR 29785, July 20, 1990]

§25.1182 Nacelle areas behind firewalls, and engine pod attaching structures containing flammable fluid lines.(a) Each nacelle area immediately behind the

firewall, and each portion of any engine pod at-taching structure containing flammable fluid lines,must meet each requirement of §§25.1103(b),25.1165 (d) and (e), 25.1183, 25.1185(c),25.1187, 25.1189, and 25.1195 through 25.1203,including those concerning designated fire zones.However, engine pod attaching structures neednot contain fire detection or extinguishing means.

(b) For each area covered by paragraph (a) ofthis section that contains a retractable landinggear, compliance with that paragraph need onlybe shown with the landing gear retracted.


§25.1183 Flammable fluid-carrying components.(a) Except as provided in paragraph (b) of this

section, each line, fitting, and other componentcarrying flammable fluid in any area subject to en-gine fire conditions, and each component whichconveys or contains flammable fluid in a desig-nated fire zone must be fire resistant, except thatflammable fluid tanks and supports in a desig-nated fire zone must be fireproof or be enclosedby a fireproof shield unless damage by fire to anynon-fireproof part will not cause leakage or spill-age of flammable fluid. Components must be


112 ASA

shielded or located to safeguard against the igni-tion of leaking flammable fluid. An integral oilsump of less than 25-quart capacity on a recipro-cating engine need not be fireproof nor be en-closed by a fireproof shield.

(b) Paragraph (a) of this section does not applyto—

(1) Lines, fittings, and components which arealready approved as part of a type certificated en-gine; and

(2) Vent and drain lines, and their fittings,whose failure will not result in, or add to, a firehazard.

(c) All components, including ducts, within adesignated fire zone must be fireproof if, when ex-posed to or damaged by fire, they could —

(1) Result in fire spreading to other regions ofthe airplane; or

(2) Cause unintentional operation of, or inabilityto operate, essential services or equipment.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended 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; Amdt. 25–101, 65 FR 79710,Dec. 19, 2000]

§25.1185 Flammable fluids.(a) Except for the integral oil sumps specified in

§25.1183 (a), no tank or reservoir that is a part ofa system containing flammable fluids or gasesmay be in a designated fire zone unless the fluidcontained, the design of the system, the materialsused in the tank, the shut-off means, and all con-nections, lines, and control provide a degree ofsafety equal to that which would exist if the tank orreservoir were outside such a zone.

(b) There must be at least one-half inch of clearairspace between each tank or reservoir andeach firewall or shroud isolating a designated firezone.

(c) Absorbent materials close to flammablefluid system components that might leak must becovered or treated to prevent the absorption ofhazardous quantities of fluids.


§25.1187 Drainage and ventilation of fire zones.(a) There must be complete drainage of each

part of each designated fire zone to minimize thehazards resulting from failure or malfunctioning ofany component containing flammable fluids. Thedrainage means must be—

(1) Effective under conditions expected to pre-vail when drainage is needed; and

(2) Arranged so that no discharged fluid willcause an additional fire hazard.

(b) Each designated fire zone must be venti-lated to prevent the accumulation of flammablevapors.

(c) No ventilation opening may be where itwould allow the entry of flammable fluids, vapors,or flame from other zones.

(d) Each ventilation means must be arrangedso that no discharged vapors will cause an addi-tional fire hazard.

(e) Unless the extinguishing agent capacityand rate of discharge are based on maximum airflow through a zone, there must be means to al-low the crew to shut off sources of forced ventila-tion to any fire zone except the engine power sec-tion of the nacelle and the combustion heater ven-tilating air ducts.

§25.1189 Shutoff means.(a) Each engine installation and each fire zone

specified in §25.1181(a) (4) and (5) must have ameans to shut off or otherwise prevent hazardousquantities of fuel, oil, deicer, and other flammablefluids, from flowing into, within, or through anydesignated fire zone, except that shutoff meansare not required for—

(1) Lines, fittings, and components forming anintegral part of an engine; and

(2) Oil systems for turbine engine installationsin which all components of the system in a desig-nated fire zone, including oil tanks, are fireproofor located in areas not subject to engine fire con-ditions.

(b) The closing of any fuel shutoff valve for anyengine may not make fuel unavailable to the re-maining engines.

(c) Operation of any shutoff may not interferewith the later emergency operation of other equip-ment, such as the means for feathering the pro-peller.

(d) Each flammable fluid shutoff means andcontrol must be fireproof or must be located andprotected so that any fire in a fire zone will not af-fect its operation.

(e) No hazardous quantity of flammable fluidmay drain into any designated fire zone aftershutoff.

(f) There must be means to guard against inad-vertent operation of the shutoff means and tomake it possible for the crew to reopen the shutoffmeans in flight after it has been closed.

(g) Each tank-to-engine shutoff valve must belocated so that the operation of the valve will notbe affected by powerplant or engine mount struc-tural failure.

(h) Each shutoff valve must have a means torelieve excessive pressure accumulation unless ameans for pressure relief is otherwise provided inthe system.


ASA 113

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§25.1191 Firewalls.(a) Each engine, auxiliary power unit, fuel-

burning heater, other combustion equipment in-tended for operation in flight, and the combustion,turbine, and tailpipe sections of turbine engines,must be isolated from the rest of the airplane byfirewalls, shrouds, or equivalent means.

(b) Each firewall and shroud must be—(1) Fireproof;(2) Constructed so that no hazardous quantity

of air, fluid, or flame can pass from the compart-ment to other parts of the airplane;

(3) Constructed so that each opening is sealedwith close fitting fireproof grommets, bushings, orfirewall fittings; and

(4) Protected against corrosion.

§25.1192 Engine accessory section diaphragm.For reciprocating engines, the engine power

section and all portions of the exhaust systemmust be isolated from the engine accessory com-partment by a diaphragm that complies with thefirewall requirements of §25.1191.


§25.1193 Cowling and nacelle skin.(a) Each cowling must be constructed and sup-

ported so that it can resist any vibration, inertia,and air load to which it may be subjected in oper-ation.

(b) Cowling must meet the drainage and venti-lation requirements of §25.1187.

(c) On airplanes with a diaphragm isolating theengine power section from the engine accessorysection, each part of the accessory section cowl-ing subject to flame in case of fire in the enginepower section of the powerplant must—

(1) Be fireproof; and(2) Meet the requirements of §25.1191.(d) Each part of the cowling subject to high

temperatures due to its nearness to exhaust sys-tem parts or exhaust gas impingement must befireproof.

(e) Each airplane must—(1) Be designed and constructed so that no fire

originating in any fire zone can enter, eitherthrough openings or by burning through externalskin, any other zone or region where it would cre-ate additional hazards;

(2) Meet paragraph (e)(1) of this section withthe landing gear retracted (if applicable); and

(3) Have fireproof skin in areas subject to flameif a fire starts in the engine power or accessorysections.

§25.1195 Fire extinguishing systems.(a) Except for combustor, turbine, and tail pipe

sections of turbine engine installations that con-tain lines or components carrying flammable flu-ids or gases for which it is shown that a fire origi-nating in these sections can be controlled, theremust be a fire extinguisher system serving eachdesignated fire zone.

(b) The fire extinguishing system, the quantityof the extinguishing agent, the rate of discharge,and the discharge distribution must be adequateto extinguish fires. It must be shown by either ac-tual or simulated flights tests that under criticalairflow conditions in flight the discharge of the ex-tinguishing agent in each designated fire zonespecified in paragraph (a) of this section will pro-vide an agent concentration capable of extin-guishing fires in that zone and of minimizing theprobability of reignition. An individual “one-shot”system may be used for auxiliary power units, fuelburning heaters, and other combustion equip-ment. For each other designated fire zone, twodischarges must be provided each of which pro-duces adequate agent concentration.

(c) The fire extinguishing system for a nacellemust be able to simultaneously protect each zoneof the nacelle for which protection is provided.


§25.1197 Fire extinguishing agents.(a) Fire extinguishing agents must—(1) Be capable of extinguishing flames emanat-

ing from any burning of fluids or other combustiblematerials in the area protected by the fire extin-guishing system; and

(2) Have thermal stability over the temperaturerange likely to be experienced in the compartmentin which they are stored.

(b) If any toxic extinguishing agent is used, pro-visions must be made to prevent harmful concen-trations of fluid or fluid vapors (from leakage dur-ing normal operation of the airplane or as a resultof discharging the fire extinguisher on the groundor in flight) from entering any personnel compart-ment, even though a defect may exist in the extin-guishing system. This must be shown by test ex-cept for built-in carbon dioxide fuselage compart-ment fire extinguishing systems for which—

(1) Five pounds or less of carbon dioxide will bedischarged, under established fire control proce-dures, into any fuselage compartment; or

(2) There is protective breathing equipment foreach flight crewmember on flight deck duty.


114 ASA

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–38, 41 FR 55467, Dec. 20, 1976;Amdt. 25–40, 42 FR 15044, March 17, 1977]

§25.1199 Extinguishing agent containers.(a) Each extinguishing agent container must

have a pressure relief to prevent bursting of thecontainer by excessive internal pressures.

(b) The discharge end of each discharge linefrom a pressure relief connection must be locatedso that discharge of the fire extinguishing agentwould not damage the airplane. The line mustalso be located or protected to prevent cloggingcaused by ice or other foreign matter.

(c) There must be a means for each fire extin-guishing agent container to indicate that the con-tainer has discharged or that the charging pres-sure is below the established minimum necessaryfor proper functioning.

(d) The temperature of each container must bemaintained, under intended operating conditions,to prevent the pressure in the container from—

(1) Falling below that necessary to provide anadequate rate of discharge; or

(2) Rising high enough to cause premature dis-charge.

(e) If a pyrotechnic capsule is used to dis-charge the extinguishing agent, each containermust be installed so that temperature conditionswill not cause hazardous deterioration of the pyro-technic capsule.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5678, April 8, 1970;Amdt. 25–40, 42 FR 15044, March 17, 1977]

§25.1201 Fire extinguishing system materials.(a) No material in any fire extinguishing system

may react chemically with any extinguishing agentso as to create a hazard.

(b) Each system component in an engine com-partment must be fireproof.

§25.1203 Fire detector system.(a) There must be approved, quick acting fire or

overheat detectors in each designated fire zone,and in the combustion, turbine, and tailpipe sec-tions of turbine engine installations, in numbersand locations ensuring prompt detection of fire inthose zones.

(b) Each fire detector system must be con-structed and installed so that—

(1) It will withstand the vibration, inertia, andother loads to which it may be subjected in oper-ation;

(2) There is a means to warn the crew in theevent that the sensor or associated wiring within adesignated fire zone is severed at one point, un-

less the system continues to function as a satis-factory detection system after the severing; and

(3) There is a means to warn the crew in theevent of a short circuit in the sensor or associatedwiring within a designated fire zone, unless thesystem continues to function as a satisfactory de-tection system after the short circuit.

(c) No fire or overheat detector may be affectedby any oil, water, other fluids or fumes that mightbe present.

(d) There must be means to allow the crew tocheck, in flight, the functioning of each fire oroverheat detector electric circuit.

(e) Components of each fire or overheat detec-tor system in a fire zone must be at least fire-re-sistant.

(f) No fire or overheat detector system compo-nent for any fire zone may pass through anotherfire zone, unless—

(1) It is protected against the possibility of falsewarnings resulting from fires in zones throughwhich it passes; or

(2) Each zone involved is simultaneously pro-tected by the same detector and extinguishingsystem.

(g) Each fire detector system must be con-structed so that when it is in the configuration forinstallation it will not exceed the alarm activationtime approved for the detectors using the re-sponse time criteria specified in the appropriateTechnical Standard Order for the detector.

(h) EWIS for each fire or overheat detector sys-tem in a fire zone must meet the requirements of§25.1731.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5678, April 8, 1970;Amdt. 25–26, 36 FR 5493, March 24, 1971; Amdt. 25–123, 72 FR 63405, Nov. 8, 2007]

§25.1207 Compliance.Unless otherwise specified, compliance with

the requirements of §§25.1181 through 25.1203must be shown by a full scale fire test or by one ormore of the following methods:

(a) Tests of similar powerplant configurations;(b) Tests of components;(c) Service experience of aircraft with similar

powerplant configurations;(d) Analysis.



ASA 115

25

Subpart F—Equipment

GENERAL

§25.1301 Function and installation.(a) Each item of installed equipment must—(1) Be of a kind and design appropriate to its in-

tended function;(2) Be labeled as to its identification, function,

or operating limitations, or any applicable combi-nation of these factors;

(3) Be installed according to limitations speci-fied for that equipment; and

(4) Function properly when installed.(b) EWIS must meet the requirements of Sub-

part H of this part.


§25.1303 Flight and navigation instruments.(a) The following flight and navigation instru-

ments must be installed so that the instrument isvisible from each pilot station:

(1) A free air temperature indicator or an air-temperature indicator which provides indicationsthat are convertible to free-air temperature.

(2) A clock displaying hours, minutes, and sec-onds with a sweep-second pointer or digital pre-sentation.

(3) A direction indicator (nonstabilized mag-netic compass).

(b) The following flight and navigation instru-ments must be installed at each pilot station:

(1) An airspeed indicator. If airspeed limitationsvary with altitude, the indicator must have a maxi-mum allowable airspeed indicator showing thevariation of VMO with altitude.

(2) An altimeter (sensitive).(3) A rate-of-climb indicator (vertical speed).(4) A gyroscopic rate-of-turn indicator com-

bined with an integral slip-skid indicator (turn-and-bank indicator) except that only a slip-skid indica-tor is required on large airplanes with a third atti-tude instrument system usable through flight atti-tudes of 360° of pitch and roll and installed in ac-cordance with §121.305(k) of this title.

(5) A bank and pitch indicator (gyroscopicallystabilized).

(6) A direction indicator (gyroscopically stabi-lized, magnetic or nonmagnetic).

(c) The following flight and navigation instru-ments are required as prescribed in this para-graph:

(1) A speed warning device is required for tur-bine engine powered airplanes and for airplaneswith VMO/MMO greater than 0.8 VDF/MDF or 0.8VD/MD. The speed warning device must give ef-fective aural warning (differing distinctively from

aural warnings used for other purposes) to thepilots, whenever the speed exceeds VMO plus 6knots or MMO +0.01. The upper limit of the pro-duction tolerance for the warning device may notexceed the prescribed warning speed.

(2) A machmeter is required at each pilot sta-tion for airplanes with compressibility limitationsnot otherwise indicated to the pilot by the air-speed indicating system required under para-graph (b)(1) of this section.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5678, April 8, 1970;Amdt. 25–24, 35 FR 7108, May 6, 1970; Amdt. 25–38,41 FR 55467, Dec. 20, 1976; Amdt. 25–90, 62 FR13253, March 19, 1997]

§25.1305 Powerplant instruments.The following are required powerplant instru-

ments:(a) For all airplanes.(1) A fuel pressure warning means for each en-

gine, or a master warning means for all engineswith provision for isolating the individual warningmeans from the master warning means.

(2) A fuel quantity indicator for each fuel tank.(3) An oil quantity indicator for each oil tank.(4) An oil pressure indicator for each indepen-

dent pressure oil system of each engine.(5) An oil pressure warning means for each en-

gine, or a master warning means for all engineswith provision for isolating the individual warningmeans from the master warning means.

(6) An oil temperature indicator for each engine.(7) Fire-warning indicators.(8) An augmentation liquid quantity indicator

(appropriate for the manner in which the liquid isto be used in operation) for each tank.

(b) For reciprocating engine-powered air-planes. In addition to the powerplant instrumentsrequired by paragraph (a) of this section, the fol-lowing powerplant instruments are required:

(1) A carburetor air temperature indicator foreach engine.

(2) A cylinder head temperature indicator foreach air-cooled engine.

(3) A manifold pressure indicator for each en-gine.

(4) A fuel pressure indicator (to indicate thepressure at which the fuel is supplied) for eachengine.

(5) A fuel flowmeter, or fuel mixture indicator,for each engine without an automatic altitude mix-ture control.

(6) A tachometer for each engine.(7) A device that indicates, to the flight crew

(during flight), any change in the power output, foreach engine with—

(i) An automatic propeller feathering system,whose operation is initiated by a power outputmeasuring system; or


116 ASA

(ii) A total engine piston displacement of 2,000cubic inches or more.

(8) A means to indicate to the pilot when thepropeller is in reverse pitch, for each reversingpropeller.

(c) For turbine engine-powered airplanes. Inaddition to the powerplant instruments requiredby paragraph (a) of this section, the followingpowerplant instruments are required:

(1) A gas temperature indicator for each en-gine.

(2) A fuel flowmeter indicator for each engine.(3) A tachometer (to indicate the speed of the

rotors with established limiting speeds) for eachengine.

(4) A means to indicate, to the flight crew, theoperation of each engine starter that can be oper-ated continuously but that is neither designed forcontinuous operation nor designed to preventhazard if it failed.

(5) An indicator to indicate the functioning of thepowerplant ice protection system for each engine.

(6) An indicator for the fuel strainer or filter re-quired by §25.997 to indicate the occurrence ofcontamination of the strainer or filter before itreaches the capacity established in accordancewith §25.997(d).

(7) A warning means for the oil strainer or filterrequired by §25.1019, if it has no bypass, to warnthe pilot of the occurrence of contamination of thestrainer or filter screen before it reaches the capac-ity established in accordance with §25.1019(a)(2).

(8) An indicator to indicate the proper function-ing of any heater used to prevent ice clogging offuel system components.

(d) For turbojet engine powered airplanes. Inaddition to the powerplant instruments requiredby paragraphs (a) and (c) of this section, the fol-lowing powerplant instruments are required:

(1) An indicator to indicate thrust, or a parame-ter that is directly related to thrust, to the pilot. Theindication must be based on the direct measure-ment of thrust or of parameters that are directlyrelated to thrust. The indicator must indicate achange in thrust resulting from any engine mal-function, damage, or deterioration.

(2) A position indicating means to indicate tothe flight crew when the thrust reversing device isin the reverse thrust position, for each engine us-ing a thrust reversing device.

(3) An indicator to indicate rotor system unbal-ance.

(e) For turbopropeller-powered airplanes. Inaddition to the powerplant instruments requiredby paragraphs (a) and (c) of this section, the fol-lowing powerplant instruments are required:

(1) A torque indicator for each engine.(2) Position indicating means to indicate to the

flight crew when the propeller blade angle is be-low the flight low pitch position, for each propeller.

(f) For airplanes equipped with fluid systems(other than fuel) for thrust or power augmenta-tion, an approved means must be provided to in-dicate the proper functioning of that system to theflight crew.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5678, April 8, 1970;Amdt. 25–35, 39 FR 1831, Jan. 15, 1974; Amdt. 25–36,39 FR 35461, Oct. 1, 1974; Amdt. 25–38, 41 FR 55467,Dec. 20, 1976; Amdt. 25–54, 45 FR 60173, Sept. 11,1980; Amdt. 25–72, 55 FR 29785, July 20, 1990]

§25.1307 Miscellaneous equipment.The following is required miscellaneous equip-

ment:(a) [Reserved](b) Two or more independent sources of electri-

cal energy.(c) Electrical protective devices, as prescribed

in this part.(d) Two systems for two-way radio communica-

tions, with controls for each accessible from eachpilot station, designed and installed so that failureof one system will not preclude operation of theother system. The use of a common antenna sys-tem is acceptable if adequate reliability is shown.

(e) Two systems for radio navigation, with con-trols for each accessible from each pilot station,designed and installed so that failure of one sys-tem will not preclude operation of the other sys-tem. The use of a common antenna system is ac-ceptable if adequate reliability is shown.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5678, April 8, 1970;Amdt. 25–46, 43 FR 50598, Oct. 30, 1978; Amdt. 25–54,45 FR 60173, Sept. 11, 1980; Amdt. 25–72, 55 FR29785, July 20, 1990]

§25.1309 Equipment, systems, and installations.(a) The equipment, systems, and installations

whose functioning is required by this subchapter,must be designed to ensure that they performtheir intended functions under any foreseeableoperating condition.

(b) The airplane systems and associated com-ponents, considered separately and in relation toother systems, must be designed so that—

(1) The occurrence of any failure conditionwhich would prevent the continued safe flight andlanding of the airplane is extremely improbable,and

(2) The occurrence of any other failure condi-tions which would reduce the capability of the air-plane or the ability of the crew to cope with ad-verse operating conditions is improbable.

(c) Warning information must be provided toalert the crew to unsafe system operating condi-tions, and to enable them to take appropriate cor-


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rective action. Systems, controls, and associatedmonitoring and warning means must be designedto minimize crew errors which could create addi-tional hazards.

(d) Compliance with the requirements of para-graph (b) of this section must be shown by analy-sis, and where necessary, by appropriate ground,flight, or simulator tests. The analysis must con-sider—

(1) Possible modes of failure, including mal-functions and damage from external sources.

(2) The probability of multiple failures and un-detected failures.

(3) The resulting effects on the airplane and oc-cupants, considering the stage of flight and oper-ating conditions, and

(4) The crew warning cues, corrective action re-quired, and the capability of detecting faults.

(e) In showing compliance with paragraphs (a)and (b) of this section with regard to the electricalsystem and equipment design and installation,critical environmental conditions must be consid-ered. For electrical generation, distribution, andutilization equipment required by or used in com-plying with this chapter, except equipment coveredby Technical Standard Orders containing environ-mental test procedures, the ability to provide con-tinuous, safe service under foreseeable environ-mental conditions may be shown by environmentaltests, design analysis, or reference to previouscomparable service experience on other aircraft.

(f) EWIS must be assessed in accordance withthe requirements of §25.1709.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5679, April 8, 1970;Amdt. 25–38, 41 FR 55467, Dec. 20, 1976; Amdt. 25–41, 42 FR 36970, July 18, 1977; Amdt. 25–123, 72 FR63405, Nov. 8, 2007]

§25.1310 Power source capacity and distribution.(a) Each installation whose functioning is re-

quired for type certification or under operatingrules and that requires a power supply is an "es-sential load" on the power supply. The powersources and the system must be able to supplythe following power loads in probable operatingcombinations and for probable durations:

(1) Loads connected to the system with thesystem functioning normally.

(2) Essential loads, after failure of any oneprime mover, power converter, or energy storagedevice.

(3) Essential loads after failure of—(i) Any one engine on two-engine airplanes;

and(ii) Any two engines on airplanes with three or

more engines.

(4) Essential loads for which an alternatesource of power is required, after any failure ormalfunction in any one power supply system, dis-tribution system, or other utilization system.

(b) In determining compliance with paragraphs(a)(2) and (3) of this section, the power loads maybe assumed to be reduced under a monitoringprocedure consistent with safety in the kinds ofoperation authorized. Loads not required in con-trolled flight need not be considered for the two-engine-inoperative condition on airplanes withthree or more engines.


§25.1316 System lightning protection.(a) For functions whose failure would contribute

to or cause a condition that would prevent thecontinued safe flight and landing of the airplane,each electrical and electronic system that per-forms these functions must be designed and in-stalled to ensure that the operation and opera-tional capabilities of the systems to perform thesefunctions are not adversely affected when the air-plane is exposed to lightning.

(b) For functions whose failure would contrib-ute to or cause a condition that would reduce thecapability of the airplane or the ability of the flight-crew to cope with adverse operating conditions,each electrical and electronic system that per-forms these functions must be designed and in-stalled to ensure that these functions can be re-covered in a timely manner after the airplane isexposed to lightning.

(c) Compliance with the lightning protectioncriteria prescribed in paragraphs (a) and (b) ofthis section must be shown for exposure to a se-vere lightning environment. The applicant mustdesign for and verify that aircraft electrical/elec-tronic systems are protected against the effectsof lightning by:

(1) Determining the lightning strike zones forthe airplane;

(2) Establishing the external lightning environ-ment for the zones;

(3) Establishing the internal environment;(4) Identifying all the electrical and electronic

systems that are subject to the requirements ofthis section, and their locations on or within theairplane;

(5) Establishing the susceptibility of the sys-tems to the internal and external lightning envi-ronment;

(6) Designing protection; and(7) Verifying that the protection is adequate.

[Docket No. 25912, 59 FR 22116, April 28, 1994]


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§25.1317 High-Intensity Radiated Fields (HIRF) Protection.(a) Except as provided in paragraph (d) of this

section, each electrical and electronic system thatperforms a function whose failure would preventthe continued safe flight and landing of the air-plane must be designed and installed so that—

(1) The function is not adversely affected dur-ing and after the time the airplane is exposed toHIRF environment I, as described in appendix Lto this part;

(2) The system automatically recovers normaloperation of that function, in a timely manner, af-ter the airplane is exposed to HIRF environment I,as described in appendix L to this part, unless thesystem's recovery conflicts with other operationalor functional requirements of the system; and

(3) The system is not adversely affected duringand after the time the airplane is exposed to HIRFenvironment II, as described in appendix L to thispart.

(b) Each electrical and electronic system thatperforms a function whose failure would signifi-cantly reduce the capability of the airplane or theability of the flightcrew to respond to an adverseoperating condition must be designed and in-stalled so the system is not adversely affectedwhen the equipment providing these functions isexposed to equipment HIRF test level 1 or 2, asdescribed in appendix L to this part.

(c) Each electrical and electronic system thatperforms a function whose failure would reducethe capability of the airplane or the ability of theflightcrew to respond to an adverse operatingcondition must be designed and installed so thesystem is not adversely affected when the equip-ment providing the function is exposed to equip-ment HIRF test level 3, as described in appendixL to this part.

(d) Before December 1, 2012, an electrical orelectronic system that performs a function whosefailure would prevent the continued safe flight andlanding of an airplane may be designed and in-stalled without meeting the provisions of para-graph (a) provided—

(1) The system has previously been shown tocomply with special conditions for HIRF, pre-scribed under §21.16, issued before December 1,2007;

(2) The HIRF immunity characteristics of thesystem have not changed since compliance withthe special conditions was demonstrated; and

(3) The data used to demonstrate compliancewith the special conditions is provided.


INSTRUMENTS: INSTALLATION

§25.1321 Arrangement and visibility.(a) Each flight, navigation, and powerplant in-

strument for use by any pilot must be plainly visi-ble to him from his station with the minimum prac-ticable deviation from his normal position and lineof vision when he is looking forward along theflight path.

(b) The flight instruments required by §25.1303must be grouped on the instrument panel andcentered as nearly as practicable about the verti-cal plane of the pilot’s forward vision. In addition—

(1) The instrument that most effectively indi-cates attitude must be on the panel in the top cen-ter position;

(2) The instrument that most effectively indi-cates airspeed must be adjacent to and directly tothe left of the instrument in the top center position:

(3) The instrument that most effectively indi-cates altitude must be adjacent to and directly tothe right of the instrument in the top center posi-tion; and

(4) The instrument that most effectively indi-cates direction of flight must be adjacent to anddirectly below the instrument in the top center po-sition.

(c) Required powerplant instruments must beclosely grouped on the instrument panel. Inaddition—

(1) The location of identical powerplant instru-ments for the engines must prevent confusion asto which engine each instrument relates; and

(2) Powerplant instruments vital to the safe op-eration of the airplane must be plainly visible tothe appropriate crewmembers.

(d) Instrument panel vibration may not damageor impair the accuracy of any instrument.

(e) If a visual indicator is provided to indicatemalfunction of an instrument, it must be effectiveunder all probable cockpit lighting conditions.


§25.1322 Warning, caution, and advisory lights.If warning, caution or advisory lights are in-

stalled in the cockpit, they must, unless otherwiseapproved by the Administrator, be—

(a) Red, for warning lights (lights indicating ahazard which may require immediate correctiveaction);

(b) Amber, for caution lights (lights indicatingthe possible need for future corrective action);

(c) Green, for safe operation lights; and(d) Any other color, including white, for lights

not described in paragraphs (a) through (c) of thissection, provided the color differs sufficiently from


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the colors prescribed in paragraphs (a) through(c) of this section to avoid possible confusion.


§25.1323 Airspeed indicating system.For each airspeed indicating system, the fol-

lowing apply:(a) Each airspeed indicating instrument must

be approved and must be calibrated to indicatetrue airspeed (at sea level with a standard atmo-sphere) with a minimum practicable instrumentcalibration error when the corresponding pitot andstatic pressures are applied.

(b) Each system must be calibrated to deter-mine the system error (that is, the relation be-tween IAS and CAS) in flight and during the ac-celerated takeoff ground run. The ground run cali-bration must be determined—

(1) From 0.8 of the minimum value of V1 to themaximum value of V2, considering the approvedranges of altitude and weight; and

(2) With the flaps and power settings corre-sponding to the values determined in the estab-lishment of the takeoff path under §25.111 as-suming that the critical engine fails at the mini-mum value of V1.

(c) The airspeed error of the installation, ex-cluding the airspeed indicator instrument calibra-tion error, may not exceed three percent or fiveknots, whichever is greater, throughout the speedrange, from—

(1) VMO to 1.23 VSR1, with flaps retracted; and(2) 1.23 VSR0 to VFE with flaps in the landing

position.(d) From 1.23 VSR to the speed at which stall

warning begins, the IAS must change perceptiblywith CAS and in the same sense, and at speedsbelow stall warning speed the IAS must notchange in an incorrect sense.

(e) From VMO to VMO + 2/3 (VDF – VMO), theIAS must change perceptibly with CAS and in thesame sense, and at higher speeds up to VDF theIAS must not change in an incorrect sense.

(f) There must be no indication of airspeed thatwould cause undue difficulty to the pilot during thetakeoff between the initiation of rotation and theachievement of a steady climbing condition.

(g) The effects of airspeed indicating systemlag may not introduce significant takeoff indicatedairspeed bias, or significant errors in takeoff or ac-celerate-stop distances.

(h) Each system must be arranged, so far aspracticable, to prevent malfunction or serious er-ror due to the entry of moisture, dirt, or other sub-stances.

(i) Each system must have a heated pitot tubeor an equivalent means of preventing malfunctiondue to icing.

(j) Where duplicate airspeed indicators are re-quired, their respective pitot tubes must be farenough apart to avoid damage to both tubes in acollision with a bird.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–57, 49 FR 6849, Feb. 23, 1984;Amdt. 25–108, 67 FR 70828, Nov. 26, 2002; Amdt. 25–109, 67 FR 76656, Dec. 12, 2002]

§25.1325 Static pressure systems.(a) Each instrument with static air case con-

nections must be vented to the outside atmo-sphere through an appropriate piping system.

(b) Each static port must be designed and lo-cated in such manner that the static pressure sys-tem performance is least affected by airflow varia-tion, or by moisture or other foreign matter, andthat the correlation between air pressure in thestatic pressure system and true ambient atmo-spheric static pressure is not changed when theairplane is exposed to the continuous and inter-mittent maximum icing conditions defined in Ap-pendix C of this part.

(c) The design and installation of the staticpressure system must be such that—

(1) Positive drainage of moisture is provided;chafing of the tubing and excessive distortion orrestriction at bends in the tubing is avoided; andthe materials used are durable, suitable for thepurpose intended, and protected against corro-sion; and

(2) It is airtight except for the port into the atmo-sphere. A proof test must be conducted to dem-onstrate the integrity of the static pressure systemin the following manner:

(i) Unpressurized airplanes. Evacuate thestatic pressure system to a pressure differential ofapproximately 1 inch of mercury or to a readingon the altimeter, 1,000 feet above the airplane el-evation at the time of the test. Without additionalpumping for a period of 1 minute, the loss of indi-cated altitude must not exceed 100 feet on the al-timeter.

(ii) Pressurized airplanes. Evacuate the staticpressure system until a pressure differentialequivalent to the maximum cabin pressure differ-ential for which the airplane is type certificated isachieved. Without additional pumping for a periodof 1 minute, the loss of indicated altitude must notexceed 2 percent of the equivalent altitude of themaximum cabin differential pressure or 100 feet,whichever is greater.

(d) Each pressure altimeter must be approvedand must be calibrated to indicate pressure alti-tude in a standard atmosphere, with a minimumpracticable calibration error when the correspond-ing static pressures are applied.

(e) Each system must be designed and in-stalled so that the error in indicated pressure alti-


120 ASA

tude, at sea level, with a standard atmosphere,excluding instrument calibration error, does notresult in an error of more than ±30 feet per 100knots speed for the appropriate configuration inthe speed range between 1.23 VSR0 with flaps ex-tended and 1.7 VSR1 with flaps retracted. How-ever, the error need not be less than ±30 feet.

(f) If an altimeter system is fitted with a devicethat provides corrections to the altimeter indica-tion, the device must be designed and installed insuch manner that it can be bypassed when it mal-functions, unless an alternate altimeter system isprovided. Each correction device must be fittedwith a means for indicating the occurrence of rea-sonably probable malfunctions, including powerfailure, to the flight crew. The indicating meansmust be effective for any cockpit lighting conditionlikely to occur.

(g) Except as provided in paragraph (h) of thissection, if the static pressure system incorporatesboth a primary and an alternate static pressuresource, the means for selecting one or the othersource must be designed so that—

(1) When either source is selected, the other isblocked off; and

(2) Both sources cannot be blocked off simulta-neously.

(h) For unpressurized airplanes, paragraph(g)(1) of this section does not apply if it can bedemonstrated that the static pressure system cal-ibration, when either static pressure source is se-lected, is not changed by the other static pressuresource being open or blocked.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–5, 30 FR 8261, June 29, 1965;Amdt. 25–12, 32 FR 7587, May 24, 1967; Amdt. 25–41,42 FR 36970, July 18, 1977; Amdt. 25–108, 67 FR70828, Nov. 26, 2002]

§25.1326 Pitot heat indication systems.If a flight instrument pitot heating system is in-

stalled, an indication system must be provided toindicate to the flight crew when that pitot heatingsystem is not operating. The indication systemmust comply with the following requirements:

(a) The indication provided must incorporate anamber light that is in clear view of a flight crew-member.

(b) The indication provided must be designedto alert the flight crew if either of the following con-ditions exist:

(1) The pitot heating system is switched “off.”(2) The pitot heating system is switched “on”

and any pitot tube heating element is inoperative.


§25.1327 Magnetic direction indicator.(a) Each magnetic direction indicator must be

installed so that its accuracy is not excessively af-fected by the airplane’s vibration or magneticfields.

(b) The compensated installation may not havea deviation, in level flight, greater than 10 degreeson any heading.

§25.1329 Flight guidance system.(a) Quick disengagement controls for the auto-

pilot and autothrust functions must be provided foreach pilot. The autopilot quick disengagementcontrols must be located on both control wheels(or equivalent). The autothrust quick disengage-ment controls must be located on the thrust con-trol levers. Quick disengagement controls must bereadily accessible to each pilot while operatingthe control wheel (or equivalent) and thrust con-trol levers.

(b) The effects of a failure of the system to dis-engage the autopilot or autothrust functions whenmanually commanded by the pilot must be as-sessed in accordance with the requirements of§25.1309.

(c) Engagement or switching of the flight guid-ance system, a mode, or a sensor may not causea transient response of the airplane’s flight pathany greater than a minor transient, as defined inparagraph (n)(1) of this section.

(d) Under normal conditions, the disengage-ment of any automatic control function of a flightguidance system may not cause a transient re-sponse of the airplane’s flight path any greaterthan a minor transient.

(e) Under rare normal and non-normal condi-tions, disengagement of any automatic controlfunction of a flight guidance system may not resultin a transient any greater than a significant tran-sient, as defined in paragraph (n)(2) of this sec-tion.

(f) The function and direction of motion of eachcommand reference control, such as heading se-lect or vertical speed, must be plainly indicatedon, or adjacent to, each control if necessary toprevent inappropriate use or confusion.

(g) Under any condition of flight appropriate toits use, the flight guidance system may not pro-duce hazardous loads on the airplane, nor createhazardous deviations in the flight path. This ap-plies to both fault-free operation and in the eventof a malfunction, and assumes that the pilot be-gins corrective action within a reasonable periodof time.

(h) When the flight guidance system is in use, ameans must be provided to avoid excursions be-yond an acceptable margin from the speed rangeof the normal flight envelope. If the airplane expe-riences an excursion outside this range, a means


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must be provided to prevent the flight guidancesystem from providing guidance or control to anunsafe speed.

(i) The flight guidance system functions, con-trols, indications, and alerts must be designed tominimize flightcrew errors and confusion concern-ing the behavior and operation of the flight guid-ance system. Means must be provided to indicatethe current mode of operation, including anyarmed modes, transitions, and reversions. Selec-tor switch position is not an acceptable means ofindication. The controls and indications must begrouped and presented in a logical and consistentmanner. The indications must be visible to eachpilot under all expected lighting conditions.

(j) Following disengagement of the autopilot, awarning (visual and auditory) must be provided toeach pilot and be timely and distinct from all othercockpit warnings.

(k) Following disengagement of the autothrustfunction, a caution must be provided to each pilot.

(l) The autopilot may not create a potential haz-ard when the flightcrew applies an override forceto the flight controls.

(m) During autothrust operation, it must bepossible for the flightcrew to move the thrust le-vers without requiring excessive force. The auto-thrust may not create a potential hazard when theflightcrew applies an override force to the thrustlevers.

(n) For purposes of this section, a transient is adisturbance in the control or flight path of the air-plane that is not consistent with response to flight-crew inputs or environmental conditions.

(1) A minor transient would not significantly re-duce safety margins and would involve flightcrewactions that are well within their capabilities. A mi-nor transient may involve a slight increase inflightcrew workload or some physical discomfortto passengers or cabin crew.

(2) A significant transient may lead to a signifi-cant reduction in safety margins, an increase inflightcrew workload, discomfort to the flightcrew,or physical distress to the passengers or cabincrew, possibly including non-fatal injuries. Signifi-cant transients do not require, in order to remainwithin or recover to the normal flight envelope,any of the following:

(i) Exceptional piloting skill, alertness, orstrength.

(ii) Forces applied by the pilot which are greaterthan those specified in §25.143(c).

(iii) Accelerations or attitudes in the airplanethat might result in further hazard to secured ornon-secured occupants.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–46, 43 FR 50598, Oct. 30, 1978;Amdt. 25–119, 71 FR 18191, April 11, 2006]

§25.1331 Instruments using a power supply.(a) For each instrument required by §25.1303(b)

that uses a power supply, the following apply:(1) Each instrument must have a visual means

integral with, the instrument, to indicate whenpower adequate to sustain proper instrument per-formance is not being supplied. The power mustbe measured at or near the point where it entersthe instruments. For electric instruments, thepower is considered to be adequate when thevoltage is within approved limits.

(2) Each instrument must, in the event of thefailure of one power source, be supplied by an-other power source. This may be accomplishedautomatically or by manual means.

(3) If an instrument presenting navigation datareceives information from sources external to thatinstrument and loss of that information would ren-der the presented data unreliable, the instrumentmust incorporate a visual means to warn thecrew, when such loss of information occurs, thatthe presented data should not be relied upon.

(b) As used in this section, “instrument” in-cludes devices that are physically contained inone unit, and devices that are composed of two ormore physically separate units or componentsconnected together (such as a remote indicatinggyroscopic direction indicator that includes amagnetic sensing element, a gyroscopic unit, anamplifier and an indicator connected together).


§25.1333 Instrument systems.For systems that operate the instruments re-

quired by §25.1303(b) which are located at eachpilot’s station—

(a) Means must be provided to connect the re-quired instruments at the first pilot’s station to op-erating systems which are independent of the op-erating systems at other flight crew stations, orother equipment;

(b) The equipment, systems, and installationsmust be designed so that one display of the infor-mation essential to the safety of flight which isprovided by the instruments, including attitude, di-rection, airspeed, and altitude will remain avail-able to the pilots, without additional crewmemberaction, after any single failure or combination offailures that is not shown to be extremely improb-able; and

(c) Additional instruments, systems, or equip-ment may not be connected to the operating sys-tems for the required instruments, unless provi-sions are made to ensure the continued normalfunctioning of the required instruments in theevent of any malfunction of the additional instru-


122 ASA

ments, systems, or equipment which is not shownto be extremely improbable.


§25.1337 Powerplant instruments.(a) Instruments and instrument lines.(1) Each powerplant and auxiliary power unit

instrument line must meet the requirements of§§25.993 and 25.1183.

(2) Each line carrying flammable fluids underpressure must—

(i) Have restricting orifices or other safety de-vices at the source of pressure to prevent the es-cape of excessive fluid if the line fails; and

(ii) Be installed and located so that the escapeof fluids would not create a hazard.

(3) Each powerplant and auxiliary power unitinstrument that utilizes flammable fluids must beinstalled and located so that the escape of fluidwould not create a hazard.

(b) Fuel quantity indicator. There must bemeans to indicate to the flight crewmembers, thequantity, in gallons or equivalent units, of usablefuel in each tank during flight. In addition—

(1) Each fuel quantity indicator must be cali-brated to read “zero” during level flight when thequantity of fuel remaining in the tank is equal tothe unusable fuel supply determined under§25.959;

(2) Tanks with interconnected outlets and air-spaces may be treated as one tank and need nothave separate indicators; and

(3) Each exposed sight gauge, used as a fuelquantity indicator, must be protected againstdamage.

(c) Fuel flowmeter system. If a fuel flowmetersystem is installed, each metering componentmust have a means for bypassing the fuel supplyif malfunction of that component severely restrictsfuel flow.

(d) Oil quantity indicator. There must be a stickgauge or equivalent means to indicate the quan-tity of oil in each tank. If an oil transfer or reserveoil supply system is installed, there must be ameans to indicate to the flight crew, in flight, thequantity of oil in each tank.

(e) Turbopropeller blade position indicator. Re-quired turbopropeller blade position indicatorsmust begin indicating before the blade movesmore than eight degrees below the flight low pitchstop. The source of indication must directly sensethe blade position.

(f) Fuel pressure indicator. There must bemeans to measure fuel pressure, in each systemsupplying reciprocating engines, at a point down-stream of any fuel pump except fuel injectionpumps. In addition—

(1) If necessary for the maintenance of properfuel delivery pressure, there must be a connectionto transmit the carburetor air intake static pres-sure to the proper pump relief valve connection;and

(2) If a connection is required under paragraph(f)(1) of this section, the gauge balance lines mustbe independently connected to the carburetor in-let pressure to avoid erroneous readings.


ELECTRICAL SYSTEMS AND EQUIPMENT

§25.1351 General.(a) Electrical system capacity. The required

generating capacity, and number and kinds ofpower sources must—

(1) Be determined by an electrical load analy-sis; and

(2) Meet the requirements of §25.1309.(b) Generating system. The generating system

includes electrical power sources, main powerbusses, transmission cables, and associated con-trol, regulation, and protective devices. It must bedesigned so that—

(1) Power sources function properly when inde-pendent and when connected in combination;

(2) No failure or malfunction of any powersource can create a hazard or impair the ability ofremaining sources to supply essential loads;

(3) The system voltage and frequency (as ap-plicable) at the terminals of all essential loadequipment can be maintained within the limits forwhich the equipment is designed, during anyprobable operating condition; and

(4) System transients due to switching, faultclearing, or other causes do not make essentialloads inoperative, and do not cause a smoke orfire hazard.

(5) There are means accessible, in flight, to ap-propriate crewmembers for the individual and col-lective disconnection of the electrical powersources from the system.

(6) There are means to indicate to appropriatecrewmembers the generating system quantitiesessential for the safe operation of the system,such as the voltage and current supplied by eachgenerator.

(c) External power. If provisions are made forconnecting external power to the airplane, andthat external power can be electrically connectedto equipment other than that used for enginestarting, means must be provided to ensure thatno external power supply having a reverse polar-ity, or a reverse phase sequence, can supplypower to the airplane’s electrical system.

(d) Operation without normal electrical power.It must be shown by analysis, tests, or both, that


ASA 123

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the airplane can be operated safely in VFR condi-tions, for a period of not less than five minutes,with the normal electrical power (electrical powersources excluding the battery) inoperative, withcritical type fuel (from the standpoint of flameoutand restart capability), and with the airplane ini-tially at the maximum certificated altitude. Parts ofthe electrical system may remain on if—

(1) A single malfunction, including a wire bun-dle or junction box fire, cannot result in loss ofboth the part turned off and the part turned on;and

(2) The parts turned on are electrically and me-chanically isolated from the parts turned off.


§25.1353 Electrical equipment and installations.(a) Electrical equipment and controls must be

installed so that operation of any one unit or sys-tem of units will not adversely affect the simulta-neous operation of any other electrical unit or sys-tem essential to safe operation. Any electrical in-terference likely to be present in the airplane mustnot result in hazardous effects on the airplane orits systems.

(b) Storage batteries must be designed and in-stalled as follows:

(1) Safe cell temperatures and pressures mustbe maintained during any probable charging ordischarging condition. No uncontrolled increasein cell temperature may result when the battery isrecharged (after previous complete discharge)—

(i) At maximum regulated voltage or power;(ii) During a flight of maximum duration; and(iii) Under the most adverse cooling condition

likely to occur in service.(2) Compliance with paragraph (b)(1) of this

section must be shown by test unless experiencewith similar batteries and installations has shownthat maintaining safe cell temperatures and pres-sures presents no problem.

(3) No explosive or toxic gases emitted by anybattery in normal operation, or as the result of anyprobable malfunction in the charging system orbattery installation, may accumulate in hazardousquantities within the airplane.

(4) No corrosive fluids or gases that may es-cape from the battery may damage surroundingairplane structures or adjacent essential equip-ment.

(5) Each nickel cadmium battery installationmust have provisions to prevent any hazardouseffect on structure or essential systems that maybe caused by the maximum amount of heat thebattery can generate during a short circuit of thebattery or of individual cells.

(6) Nickel cadmium battery installations musthave—

(i) A system to control the charging rate of thebattery automatically so as to prevent batteryoverheating;

(ii) A battery temperature sensing and over-temperature warning system with a means for dis-connecting the battery from its charging source inthe event of an over-temperature condition; or

(iii) A battery failure sensing and warning sys-tem with a means for disconnecting the batteryfrom its charging source in the event of batteryfailure.

(c) Electrical bonding must provide an ade-quate electrical return path under both normaland fault conditions, on airplanes havinggrounded electrical systems.


§25.1355 Distribution system.(a) The distribution system includes the distri-

bution busses, their associated feeders, and eachcontrol and protective device.

(b) [Reserved](c) If two independent sources of electrical

power for particular equipment or systems are re-quired by this chapter, in the event of the failure ofone power source for such equipment or system,another power source (including its separatefeeder) must be automatically provided or bemanually selectable to maintain equipment orsystem operation.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–23, 35 FR 5679, April 8, 1970;Amdt. 25–38, 41 FR 55468, Dec. 20, 1976]

§25.1357 Circuit protective devices.(a) Automatic protective devices must be used

to minimize distress to the electrical system andhazard to the airplane in the event of wiring faultsor serious malfunction of the system or connectedequipment.

(b) The protective and control devices in thegenerating system must be designed to de-ener-gize and disconnect faulty power sources andpower transmission equipment from their associ-ated busses with sufficient rapidity to provide pro-tection from hazardous over-voltage and othermalfunctioning.

(c) Each resettable circuit protective devicemust be designed so that, when an overload orcircuit fault exists, it will open the circuit irrespec-tive of the position of the operating control.

(d) If the ability to reset a circuit breaker or re-place a fuse is essential to safety in flight, that cir-cuit breaker or fuse must be located and identifiedso that it can be readily reset or replaced in flight.Where fuses are used, there must be spare fuses


124 ASA

for use in flight equal to at least 50% of the num-ber of fuses of each rating required for completecircuit protection.

(e) Each circuit for essential loads must haveindividual circuit protection. However, individualprotection for each circuit in an essential load sys-tem (such as each position light circuit in a sys-tem) is not required.

(f) For airplane systems for which the ability toremove or reset power during normal operationsis necessary, the system must be designed sothat circuit breakers are not the primary means toremove or reset system power unless specificallydesigned for use as a switch.

(g) Automatic reset circuit breakers may beused as integral protectors for electrical equip-ment (such as thermal cut-outs) if there is circuitprotection to protect the cable to the equipment.


§25.1360 Precautions against injury.(a) Shock. The electrical system must be de-

signed to minimize risk of electric shock to crew,passengers, and servicing personnel and tomaintenance personnel using normal precau-tions.

(b) Burns. The temperature of any part thatmay be handled by a crewmember during normaloperations must not cause dangerous inadvertentmovement by the crewmember or injury to thecrewmember.


§25.1362 Electrical supplies for emergency conditions.A suitable electrical supply must be provided to

those services required for emergency proce-dures after an emergency landing or ditching. Thecircuits for these services must be designed, pro-tected, and installed so that the risk of the ser-vices being rendered ineffective under theseemergency conditions is minimized.


§25.1363 Electrical system tests.(a) When laboratory tests of the electrical sys-

tem are conducted—(1) The tests must be performed on a mock-up

using the same generating equipment used in theairplane;

(2) The equipment must simulate the electricalcharacteristics of the distribution wiring and con-nected loads to the extent necessary for valid testresults; and

(3) Laboratory generator drives must simulatethe actual prime movers on the airplane with re-spect to their reaction to generator loading, in-cluding loading due to faults.

(b) For each flight condition that cannot be sim-ulated adequately in the laboratory or by groundtests on the airplane, flight tests must be made.

§25.1365 Electrical appliances, motors, and transformers.(a) Domestic appliances must be designed and

installed so that in the event of failures of the elec-trical supply or control system, the requirementsof §25.1309(b), (c), and (d) will be satisfied. Do-mestic appliances are items such as cooktops,ovens, coffee makers, water heaters, refrigera-tors, and toilet flush systems that are placed onthe airplane to provide service amenities to pas-sengers.

(b) Galleys and cooking appliances must be in-stalled in a way that minimizes risk of overheat orfire.

(c) Domestic appliances, particularly those ingalley areas, must be installed or protected so asto prevent damage or contamination of otherequipment or systems from fluids or vapors whichmay be present during normal operation or as aresult of spillage, if such damage or contamina-tion could create a hazardous condition.

(d) Unless compliance with §25.1309(b) is pro-vided by the circuit protective device required by§25.1357(a), electric motors and transformers, in-cluding those installed in domestic systems, musthave a suitable thermal protection device to pre-vent overheating under normal operation and fail-ure conditions, if overheating could create asmoke or fire hazard.


LIGHTS

§25.1381 Instrument lights.(a) The instrument lights must—(1) Provide sufficient illumination to make each

instrument, switch and other device necessary forsafe operation easily readable unless sufficient il-lumination is available from another source; and

(2) Be installed so that—(i) Their direct rays are shielded from the pilot’s

eyes; and(ii) No objectionable reflections are visible to

the pilot.(b) Unless undimmed instrument lights are sat-

isfactory under each expected flight condition,there must be a means to control the intensity ofillumination.



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§25.1383 Landing lights.(a) Each landing light must be approved, and

must be installed so that—(1) No objectionable glare is visible to the pilot;(2) The pilot is not adversely affected by hala-

tion; and(3) It provides enough light for night landing.(b) Except when one switch is used for the

lights of a multiple light installation at one location,there must be a separate switch for each light.

(c) There must be a means to indicate to thepilots when the landing lights are extended.

§25.1385 Position light system installation.(a) General. Each part of each position light

system must meet the applicable requirements ofthis section and each system as a whole mustmeet the requirements of §§25.1387 through25.1397.

(b) Forward position lights. Forward positionlights must consist of a red and a green lightspaced laterally as far apart as practicable and in-stalled forward on the airplane so that, with theairplane in the normal flying position, the red lightis on the left side and the green light is on the rightside. Each light must be approved.

(c) Rear position light. The rear position lightmust be a white light mounted as far aft as practi-cable on the tail or on each wing tip, and must beapproved.

(d) Light covers and color filters. Each lightcover or color filter must be at least flame resistantand may not change color or shape or lose any ap-preciable light transmission during normal use.


§25.1387 Position light system dihedral angles.(a) Except as provided in paragraph (e) of this

section, each forward and rear position light must,as installed, show unbroken light within the dihe-dral angles described in this section.

(b) Dihedral angle L (left) is formed by two in-tersecting vertical planes, the first parallel to thelongitudinal axis of the airplane, and the other at110 degrees to the left of the first, as viewed whenlooking forward along the longitudinal axis.

(c) Dihedral angle R (right) is formed by two in-tersecting vertical planes, the first parallel to thelongitudinal axis of the airplane, and the other at110 degrees to the right of the first, as viewedwhen looking forward along the longitudinal axis.

(d) Dihedral angle A (aft) is formed by two inter-secting vertical planes making angles of 70 de-grees to the right and to the left, respectively, to avertical plane passing through the longitudinal

axis, as viewed when looking aft along the longi-tudinal axis.

(e) If the rear position light, when mounted asfar aft as practicable in accordance with§25.1385(c), cannot show unbroken light withindihedral angle A (as defined in paragraph (d) ofthis section), a solid angle or angles of obstructedvisibility totaling not more than 0.04 steradians isallowable within that dihedral angle, if such solidangle is within a cone whose apex is at the rearposition light and whose elements make an angleof 30° with a vertical line passing through the rearposition light.


§25.1389 Position light distribution and intensities.(a) General. The intensities prescribed in this

section must be provided by new equipment withlight covers and color filters in place. Intensitiesmust be determined with the light source operat-ing at a steady value equal to the average lumi-nous output of the source at the normal operatingvoltage of the airplane. The light distribution andintensity of each position light must meet the re-quirements of paragraph (b) of this section.

(b) Forward and rear position lights. The lightdistribution and intensities of forward and rear po-sition lights must be expressed in terms of mini-mum intensities in the horizontal plane, minimumintensities in any vertical plane, and maximum in-tensities in overlapping beams, within dihedral an-gles L, R, and A, and must meet the following re-quirements:

(1) Intensities in the horizontal plane. Each in-tensity in the horizontal plane (the plane contain-ing the longitudinal axis of the airplane and per-pendicular to the plane of symmetry of the air-plane) must equal or exceed the values in§25.1391.

(2) Intensities in any vertical plane. Each inten-sity in any vertical plane (the plane perpendicularto the horizontal plane) must equal or exceed theappropriate value in §25.1393, where I is the min-imum intensity prescribed in §25.1391 for the cor-responding angles in the horizontal plane.

(3) Intensities in overlaps between adjacentsignals. No intensity in any overlap between adja-cent signals may exceed the values given in§25.1395, except that higher intensities in over-laps may be used with main beam intensities sub-stantially greater than the minima specified in§§25.1391 and 25.1393 if the overlap intensitiesin relation to the main beam intensities do not ad-versely affect signal clarity. When the peak inten-sity of the forward position lights is more than 100candles, the maximum overlap intensities be-tween them may exceed the values given in


126 ASA

§25.1395 if the overlap intensity in Area A is notmore than 10 percent of peak position light inten-sity and the overlap intensity in Area B is notgreater than 2.5 percent of peak position light in-tensity.

§25.1391 Minimum intensities in the horizontal plane of forward and rear position lights.

Each position light intensity must equal or exceedthe applicable values in the following table:

§25.1393 Minimum intensities in any vertical plane of forward and rear position lights.Each position light intensity must equal or ex-

ceed the applicable values in the following table:

§25.1395 Maximum intensities in overlapping beams of forward and rear position lights.No position light intensity may exceed the ap-

plicable values in the following table, except asprovided in §25.1389(b)(3).

Where—(a) Area A includes all directions in the adja-

cent dihedral angle that pass through the lightsource and intersect the common boundary planeat more than 10 degrees but less than 20 de-grees; and

(b) Area B includes all directions in the adja-cent dihedral angle that pass through the lightsource and intersect the common boundary planeat more than 20 degrees.

§25.1397 Color specifications.Each position light color must have the applica-

ble International Commission on Illuminationchromaticity coordinates as follows:

(a) Aviation red—

“y” is not greater than 0.335; and“z” is not greater than 0.002.

(b) Aviation green—

“x” is not greater than 0.440 – 0.320 y;“x” is not greater than y – 0.170; and“y” is not less than 0.390 – 0.170 x.

(c) Aviation white—

“x” is not less than 0.300 and not greater than0.540;

“y” is not less than “x – 0.040” or “y0 – 0.010”,whichever is the smaller; and

“y” is not greater than “x+0.020” nor “0.636–0.400 x”;

Where “y0” is the “y” coordinate of the Planck-ian radiator for the value of “x” considered.


§25.1399 Riding light.(a) Each riding (anchor) light required for a sea-

plane or amphibian must be installed so that itcan—

(1) Show a white light for at least 2 nauticalmiles at night under clear atmospheric conditions;and

(2) Show the maximum unbroken light practica-ble when the airplane is moored or drifting on thewater.

(b) Externally hung lights may be used.

§25.1401 Anticollision light system.(a) General. The airplane must have an anticol-

lision light system that—(1) Consists of one or more approved anticolli-

sion lights located so that their light will not impairthe crew’s vision or detract from the conspicuity ofthe position lights; and

Dihedral angle (light included)

Angle from right or left of longitudinal axis, measured from dead ahead

Intensity (candles)

L and R (forward red and green)

0° to 10°

10° to 20°20° to 110°

40

305

A (rear white) 110° to 180° 20

Angle above or below the horizontal plane Intensity, I

0° 1.00

0° to 5° 0.90

5° to 10° 0.80

10° to 15° 0.70

15° to 20° 0.50

20° to 30° 0.30

30° to 40° 0.10

40° to 90° 0.05

OverlapsMaximum Intensity

Area A (candles)

Area B (candles)

Green in dihedral angle L 10 1

Red in dihedral angle R 10 1

Green in dihedral angle A 5 1

Red in dihedral angle A 5 1

Rear white in dihedral angle L 5 1

Rear white in dihedral angle R 5 1


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(2) Meets the requirements of paragraphs (b)through (f) of this section.

(b) Field of coverage. The system must consistof enough lights to illuminate the vital areasaround the airplane considering the physical con-figuration and flight characteristics of the airplane.The field of coverage must extend in each direc-tion within at least 75 degrees above and 75 de-grees below the horizontal plane of the airplane,except that a solid angle or angles of obstructedvisibility totaling not more than 0.03 steradians isallowable within a solid angle equal to 0.15 stera-dians centered about the longitudinal axis in therearward direction.

(c) Flashing characteristics. The arrangementof the system, that is, the number of light sources,beam width, speed of rotation, and other charac-teristics, must give an effective flash frequency ofnot less than 40, nor more than 100 cycles perminute. The effective flash frequency is the fre-quency at which the airplane’s complete anticolli-sion light system is observed from a distance, andapplies to each sector of light including any over-laps that exist when the system consists of morethan one light source. In overlaps, flash frequen-cies may exceed 100, but not 180 cycles perminute.

(d) Color. Each anticollision light must be eitheraviation red or aviation white and must meet theapplicable requirements of §25.1397.

(e) Light intensity. The minimum light intensitiesin all vertical planes, measured with the red filter(if used) and expressed in terms of “effective” in-tensities, must meet the requirements of para-graph (f) of this section. The following relationmust be assumed:

where:

Ie = effective intensity (candles).I(t) = instantaneous intensity as a function of time.t2 – t1 = flash time interval (seconds).

Normally, the maximum value of effective intensityis obtained when t2 and t1 are chosen so that theeffective intensity is equal to the instantaneous in-tensity at t2 and t1.

(f) Minimum effective intensities for anticollisionlights. Each anticollision light effective intensitymust equal or exceed the applicable values in thefollowing table:


§25.1403 Wing icing detection lights.Unless operations at night in known or forecast

icing conditions are prohibited by an operatinglimitation, a means must be provided for illuminat-ing or otherwise determining the formation of iceon the parts of the wings that are critical from thestandpoint of ice accumulation. Any illuminationthat is used must be of a type that will not causeglare or reflection that would handicap crewmem-bers in the performance of their duties.


SAFETY EQUIPMENT

§25.1411 General.(a) Accessibility. Required safety equipment to

be used by the crew in an emergency must bereadily accessible.

(b) Stowage provisions. Stowage provisions forrequired emergency equipment must be fur-nished and must—

(1) Be arranged so that the equipment is di-rectly accessible and its location is obvious; and

(2) Protect the safety equipment from inadvert-ent damage.

(c) Emergency exit descent device. The stow-age provisions for the emergency exit descent de-vices required by §25.810(a) must be at each exitfor which they are intended.

(d) Liferafts.(1) The stowage provisions for the liferafts de-

scribed in §25.1415 must accommodate enoughrafts for the maximum number of occupants forwhich certification for ditching is requested.

(2) Liferafts must be stowed near exits throughwhich the rafts can be launched during an un-planned ditching.

(3) Rafts automatically or remotely releasedoutside the airplane must be attached to the air-plane by means of the static line prescribed in§25.1415.

Ie

l t( ) tdt1

t2∫

0.2 t2 t1–( )+---------------------------------=

Angle above or below the horizontal plane

Effective Intensity (candles)

0° to 5° 400

5° to 10° 240

10° to 20° 80

20° to 30° 40

30° to 75° 20


128 ASA

(4) The stowage provisions for each portableliferaft must allow rapid detachment and removalof the raft for use at other than the intended exits.

(e) Long-range signaling device. The stowageprovisions for the long-range signaling device re-quired by §25.1415 must be near an exit availableduring an unplanned ditching.

(f) Life preserver stowage provisions. The stow-age provisions for life preservers described in§25.1415 must accommodate one life preserverfor each occupant for which certification for ditch-ing is requested. Each life preserver must bewithin easy reach of each seated occupant.

(g) Life line stowage provisions. If certificationfor ditching under §25.801 is requested, theremust be provisions to store life lines. These provi-sions must—

(1) Allow one life line to be attached to eachside of the fuselage; and

(2) Be arranged to allow the life lines to be usedto enable the occupants to stay on the wing afterditching.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–32, 37 FR 3972, Feb. 24, 1972;Amdt. 25–46, 43 FR 50598, Oct. 30, 1978; Amdt. 25–53,45 FR 41593, June 19, 1980; Amdt. 25–70, 54 FR43925, Oct. 27, 1989; Amdt. 25–79, 58 FR 45229, Aug.26, 1993; Amdt. 25–116, 69 FR 62789, Oct. 27, 2004]

§25.1415 Ditching equipment.(a) Ditching equipment used in airplanes to be

certificated for ditching under §25.801, and re-quired by the operating rules of this chapter, mustmeet the requirements of this section.

(b) Each liferaft and each life preserver must beapproved. In addition—

(1) Unless excess rafts of enough capacity areprovided, the buoyancy and seating capacity be-yond the rated capacity of the rafts must accom-modate all occupants of the airplane in the event ofa loss of one raft of the largest rated capacity; and

(2) Each raft must have a trailing line, and musthave a static line designed to hold the raft nearthe airplane but to release it if the airplane be-comes totally submerged.

(c) Approved survival equipment must be at-tached to each liferaft.

(d) There must be an approved survival typeemergency locator transmitter for use in one liferaft.

(e) For airplanes not certificated for ditching un-der §25.801 and not having approved life preserv-ers, there must be an approved flotation meansfor each occupant. This means must be withineasy reach of each seated occupant and must bereadily removable from the airplane.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–29, 36 FR 18722, Sept. 21, 1971;Amdt 25–50, 45 FR 38348, June 9, 1980; Amdt. 25–72,55 FR 29785, July 20, 1990; as Amdt. 25–82, 59 FR32057, June 21, 1994]

§25.1419 Ice protection.If the applicant seeks certification for flight in ic-

ing conditions, the airplane must be able to safelyoperate in the continuous maximum and intermit-tent maximum icing conditions of appendix C. Toestablish this—

(a) An analysis must be performed to establishthat the ice protection for the various componentsof the airplane is adequate, taking into account thevarious airplane operational configurations; and

(b) To verify the ice protection analysis, tocheck for icing anomalies, and to demonstratethat the ice protection system and its componentsare effective, the airplane or its components mustbe flight tested in the various operational configu-rations, in measured natural atmospheric icingconditions and, as found necessary, by one ormore of the following means:

(1) Laboratory dry air or simulated icing tests,or a combination of both, of the components ormodels of the components.

(2) Flight dry air tests of the ice protection sys-tem as a whole, or of its individual components.

(3) Flight tests of the airplane or its compo-nents in measured simulated icing conditions.

(c) Caution information, such as an amber cau-tion light or equivalent, must be provided to alertthe flightcrew when the anti-ice or de-ice systemis not functioning normally.

(d) For turbine engine powered airplanes, theice protection provisions of this section are con-sidered to be applicable primarily to the airframe.For the powerplant installation, certain additionalprovisions of subpart E of this part may be foundapplicable.

(e) One of the following methods of icing detec-tion and activation of the airframe ice protectionsystem must be provided:

(1) A primary ice detection system that auto-matically activates or alerts the flightcrew to acti-vate the airframe ice protection system;

(2) A definition of visual cues for recognition ofthe first sign of ice accretion on a specified sur-face combined with an advisory ice detection sys-tem that alerts the flightcrew to activate the air-frame ice protection system; or

(3) Identification of conditions conducive to air-frame icing as defined by an appropriate static ortotal air temperature and visible moisture for useby the flightcrew to activate the airframe ice pro-tection system.

(f) Unless the applicant shows that the airframeice protection system need not be operated dur-ing specific phases of flight, the requirements of


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paragraph (e) of this section are applicable to allphases of flight.

(g) After the initial activation of the airframe iceprotection system—

(1) The ice protection system must be designedto operate continuously;

(2) The airplane must be equipped with a sys-tem that automatically cycles the ice protectionsystem; or

(3) An ice detection system must be providedto alert the flightcrew each time the ice protectionsystem must be cycled.

(h) Procedures for operation of the ice protec-tion system, including activation and deactivation,must be established and documented in the Air-plane Flight Manual.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–72, 55 FR 29785, July 20, 1990;Amdt. 25–121, 72 FR 44669, Aug. 8, 2007; Amdt. 25–129, 74 FR 38339, Aug. 3, 2009]

§25.1421 Megaphones.If a megaphone is installed, a restraining means

must be provided that is capable of restraining themegaphone when it is subjected to the ultimate in-ertia forces specified in §25.561(b)(3).


MISCELLANEOUS EQUIPMENT

§25.1423 Public address system.A public address system required by this chap-

ter must—(a) Be powerable when the aircraft is in flight or

stopped on the ground, after the shutdown or fail-ure of all engines and auxiliary power units, or thedisconnection or failure of all power sources de-pendent on their continued operation, for—

(1) A time duration of at least 10 minutes, in-cluding an aggregate time duration of at least 5minutes of announcements made by flight andcabin crewmembers, considering all other loadswhich may remain powered by the same sourcewhen all other power sources are inoperative; and

(2) An additional time duration in its standbystate appropriate or required for any other loadsthat are powered by the same source and that areessential to safety of flight or required duringemergency conditions.

(b) Be capable of operation within 10 secondsby a flight attendant at those stations in the pas-senger compartment from which they system isaccessible.

(c) Be intelligible at all passenger seats, lavato-ries, and flight attendant seats and work stations.

(d) Be designed so that no unused, unstowedmicrophone will render the system inoperative.

(e) Be capable of functioning independently ofany required crewmember interphone system.

(f) Be accessible for immediate use from eachof two flight crewmember stations in the pilot com-partment.

(g) For each required floor-level passengeremergency exit which has an adjacent flight at-tendant seat, have a microphone which is readilyaccessible to the seated flight attendant, exceptthat one microphone may serve more than oneexit, provided the proximity of the exits allows un-assisted verbal communication between seatedflight attendants.

[Docket No. 26003, 58 FR 45229, Aug. 26, 1993]

§25.1431 Electronic equipment.(a) In showing compliance with §25.1309 (a)

and (b) with respect to radio and electronic equip-ment and their installations, critical environmentalconditions must be considered.

(b) Radio and electronic equipment must besupplied with power under the requirements of§25.1355(c).

(c) Radio and electronic equipment, controls,and wiring must be installed so that operation ofany one unit or system of units will not adverselyaffect the simultaneous operation of any other ra-dio or electronic unit, or system of units, requiredby this chapter.

(d) Electronic equipment must be designedand installed such that it does not cause essentialloads to become inoperative as a result of electri-cal power supply transients or transients fromother causes.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Docket Nos. FAA–2001–9634, FAA–2001–9633, FAA–2001–9638, FAA–2001–9637, Amdt. 25–113; 69 FR 12529, March 16, 2004]

§25.1433 Vacuum systems.There must be means, in addition to the normal

pressure relief, to automatically relieve the pres-sure in the discharge lines from the vacuum airpump when the delivery temperature of the air be-comes unsafe.


§25.1435 Hydraulic systems.(a) Element design. Each element of the hy-

draulic system must be designed to:(1) Withstand the proof pressure without per-

manent deformation that would prevent it fromperforming its intended functions, and the ultimatepressure without rupture. The proof and ultimatepressures are defined in terms of the design oper-ating pressure (DOP) as follows:


130 ASA

(2) Withstand, without deformation that wouldprevent it from performing its intended function,the design operating pressure in combination withlimit structural loads that may be imposed;

(3) Withstand, without rupture, the design oper-ating pressure multiplied by a factor of 1.5 in com-bination with ultimate structural load that can rea-sonably occur simultaneously;

(4) Withstand the fatigue effects of all cyclicpressures, including transients, and associatedexternally induced loads, taking into account theconsequences of element failure; and

(5) Perform as intended under all environmen-tal conditions for which the airplane is certificated.

(b) System design. Each hydraulic systemmust:

(1) Have means located at a flightcrew stationto indicate appropriate system parameters, if

(i) It performs a function necessary for contin-ued safe flight and landing; or

(ii) In the event of hydraulic system malfunc-tion, corrective action by the crew to ensure con-tinued safe flight and landing is necessary;

(2) Have means to ensure that system pres-sures, including transient pressures and pres-sures from fluid volumetric changes in elementsthat are likely to remain closed long enough forsuch changes to occur, are within the design ca-pabilities of each element, such that they meetthe requirements defined in §25.1435(a)(1)through (a)(5);

(3) Have means to minimize the release ofharmful or hazardous concentrations of hydraulicfluid or vapors into the crew and passenger com-partments during flight;

(4) Meet the applicable requirements of§§25.863, 25.1183, 25.1185, and 25.1189 if aflammable hydraulic fluid is used; and

(5) Be designed to use any suitable hydraulicfluid specified by the airplane manufacturer, whichmust be identified by appropriate markings as re-quired by §25.1541.

(c) Tests. Tests must be conducted on the hy-draulic system(s), and/or subsystem(s) and ele-ments, except that analysis may be used in placeof or to supplement testing, where the analysis isshown to be reliable and appropriate. All internal

and external influences must be taken into ac-count to an extent necessary to evaluate their ef-fects, and to assure reliable system and elementfunctioning and integration. Failure or unaccept-able deficiency of an element or system must becorrected and be sufficiently retested, where nec-essary.

(1) The system(s), subsystem(s), or element(s)must be subjected to performance, fatigue, andendurance tests representative of airplane groundand flight operations.

(2) The complete system must be tested to de-termine proper functional performance and rela-tion to the other systems, including simulation ofrelevant failure conditions, and to support or vali-date element design.

(3) The complete hydraulic system(s) must befunctionally tested on the airplane in normal oper-ation over the range of motion of all associateduser systems. The test must be conducted at thesystem relief pressure or 1.25 times the DOP if asystem pressure relief device is not part of the sys-tem design. Clearances between hydraulic systemelements and other systems or structural elementsmust remain adequate and there must be no detri-mental effects.

[Docket No. 28617, 66 FR 27402, May 16, 2001]

Element Proof (xDOP)

Ultimate (xDOP)

1. Tubes and fittings 1.5 3.0

2. Pressure vessels containing gas: High pressure (e.g., accumulators) Low pressure (e.g., reservoirs)

3.01.5

4.03.0

3. Hoses 2.0 4.0

4. All other elements 1.5 2.0


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§25.1438 Pressurization and pneumatic systems.(a) Pressurization system elements must be

burst pressure tested to 2.0 times, and proof pres-sure tested to 1.5 times, the maximum normal op-erating pressure.

(b) Pneumatic system elements must be burstpressure tested to 3.0 times, and proof pressuretested to 1.5 times, the maximum normal operat-ing pressure.

(c) An analysis, or a combination of analysisand test, may be substituted for any test requiredby paragraph (a) or (b) of this section if the Ad-ministrator finds it equivalent to the required test.


§25.1439 Protective breathing equipment.(a) If there is a class A, B, or E cargo compart-

ment, protective breathing equipment must be in-stalled for the use of appropriate crewmembers.In addition, protective breathing equipment mustbe installed in each isolated separate compart-ment in the airplane, including upper and lowerlobe galleys, in which crewmember occupancy ispermitted during flight for the maximum numberof crewmembers expected to be in the area dur-ing any operation.

(b) For protective breathing equipment re-quired by paragraph (a) of this section or by anyoperating rule of this chapter, the following apply:

(1) The equipment must be designed to protectthe flight crew from smoke, carbon dioxide, andother harmful gases while on flight deck duty andwhile 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 com-municate with each other, while at their assignedduty stations.

(4) The part of the equipment protecting theeyes may not cause any appreciable adverse ef-fect on vision and must allow corrective glasses tobe worn.

(5) The equipment must supply protective oxy-gen of 15 minutes duration per crewmember at apressure altitude of 8,000 feet with a respiratoryminute volume of 30 liters per minute BTPD. If ademand oxygen system is used, a supply of 300liters of free oxygen at 70°F and 760 mm. Hg.pressure is considered to be of 15-minute dura-tion at the prescribed altitude and minute volume.If a continuous flow protective breathing system isused (including a mask with a standard re-breather bag) a flow rate of 60 liters per minute at8,000 feet (45 liters per minute at sea level) and asupply of 600 liters of free oxygen at 70°F and760 mm. Hg. pressure is considered to be of 15-minute duration at the prescribed altitude andminute volume. BTPD refers to body temperatureconditions (that is, 37°C, at ambient pressure,dry).

(6) The equipment must meet the requirementsof paragraphs (b) and (c) of §25.1441.


§25.1441 Oxygen equipment and supply.(a) If certification with supplemental oxygen

equipment is requested, the equipment must meetthe requirements of this section and §§25.1443through 25.1453.

(b) The oxygen system must be free from haz-ards in itself, in its method of operation, and in itseffect upon other components.

(c) There must be a means to allow the crew toreadily determine, during flight, the quantity of ox-ygen available in each source of supply.

(d) The oxygen flow rate and the oxygen equip-ment for airplanes for which certification for oper-ation above 40,000 feet is requested must be ap-proved.

§25.1443 Minimum mass flow of supplemental oxygen.(a) If continuous flow equipment is installed for

use by flight crewmembers, the minimum massflow of supplemental oxygen required for eachcrewmember may not be less than the flow re-quired to maintain, during inspiration, a mean tra-cheal oxygen partial pressure of 149 mm. Hg.when breathing 15 liters per minute, BTPS, andwith a maximum tidal volume of 700 cc. with aconstant time interval between respirations.

(b) If demand equipment is installed for use byflight crewmembers, the minimum mass flow ofsupplemental oxygen required for each crew-member may not be less than the flow required tomaintain, during inspiration, a mean tracheal oxy-gen partial pressure of 122 mm. Hg., up to and in-cluding a cabin pressure altitude of 35,000 feet,and 95 percent oxygen between cabin pressurealtitudes of 35,000 and 40,000 feet, when breath-


132 ASA

ing 20 liters per minute BTPS. In addition, theremust be means to allow the crew to use undilutedoxygen at their discretion.

(c) For passengers and cabin attendants, theminimum mass flow of supplemental oxygen re-quired for each person at various cabin pressurealtitudes may not be less than the flow required tomaintain, during inspiration and while using the ox-ygen equipment (including masks) provided, thefollowing mean tracheal oxygen partial pressures:

(1) At cabin pressure altitudes above 10,000feet up to and including 18,500 feet, a mean tra-cheal oxygen partial pressure of 100 mm. Hg.when breathing 15 liters per minute, BTPS, andwith a tidal volume of 700 cc. with a constant timeinterval between respirations.

(2) At cabin pressure altitudes above 18,500feet up to and including 40,000 feet, a mean tra-cheal oxygen partial pressure of 83.8 mm. Hg.when breathing 30 liters per minute, BTPS, andwith a tidal volume of 1,100 cc. with a constanttime interval between respirations.

(d) If first-aid oxygen equipment is installed, theminimum mass flow of oxygen to each user maynot be less than four liters per minute, STPD.However, there may be a means to decrease thisflow to not less than two liters per minute, STPD,at any cabin altitude. The quantity of oxygen re-quired is based upon an average flow rate of threeliters per minute per person for whom first-aid ox-ygen is required.

(e) If portable oxygen equipment is installed foruse by crewmembers, the minimum mass flow ofsupplemental oxygen is the same as specified inparagraph (a) or (b) of this section, whichever isapplicable.

§25.1445 Equipment standards for the oxygen distributing system.(a) When oxygen is supplied to both crew and

passengers, the distribution system must be de-signed for either—

(1) A source of supply for the flight crew on dutyand a separate source for the passengers andother crewmembers; or

(2) A common source of supply with means toseparately reserve the minimum supply requiredby the flight crew on duty.

(b) Portable walk-around oxygen units of thecontinuous flow, diluter-demand, and straight de-mand kinds may be used to meet the crew or pas-senger breathing requirements.

§25.1447 Equipment standards for oxygen dispensing units.If oxygen dispensing units are installed, the fol-

lowing apply:(a) There must be an individual dispensing unit

for each occupant for whom supplemental oxygenis to be supplied. Units must be designed to coverthe nose and mouth and must be equipped with asuitable means to retain the unit in position on theface. Flight crew masks for supplemental oxygenmust have provisions for the use of communica-tion equipment.

(b) If certification for operation up to and includ-ing 25,000 feet is requested, an oxygen supplyterminal and unit of oxygen dispensing equipmentfor the immediate use of oxygen by each crew-member must be within easy reach of that crew-member. For any other occupants, the supply ter-minals and dispensing equipment must be lo-cated to allow the use of oxygen as required bythe operating rules in this chapter.

(c) If certification for operation above 25,000feet is requested, there must be oxygen dispensingequipment meeting the following requirements:

(1) There must be an oxygen dispensing unitconnected to oxygen supply terminals immediatelyavailable to each occupant, wherever seated, andat least two oxygen dispensing units connected tooxygen terminals in each lavatory. The total num-ber of dispensing units and outlets in the cabinmust exceed the number of seats by at least 10percent. The extra units must be as uniformly dis-tributed throughout the cabin as practicable. If cer-tification for operation above 30,000 feet is re-quested, the dispensing units providing the re-quired oxygen flow must be automaticallypresented to the occupants before the cabin pres-sure altitude exceeds 15,000 feet. The crew mustbe provided with a manual means of making thedispensing units immediately available in the eventof failure of the automatic system.

(2) Each flight crewmember on flight deck dutymust be provided with a quick-donning type oxy-gen dispensing unit connected to an oxygen sup-ply terminal. This dispensing unit must be immedi-ately available to the flight crewmember whenseated at his station, and installed so that it:

(i) Can be placed on the face from its ready po-sition, properly secured, sealed, and supplyingoxygen upon demand, with one hand, within fiveseconds and without disturbing eyeglasses orcausing delay in proceeding with emergency du-ties; and

(ii) Allows, while in place, the performance ofnormal communication functions.

(3) The oxygen dispensing equipment for theflight crewmembers must be:

(i) The diluter demand or pressure demand(pressure demand mask with a diluter demand


ASA 133

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pressure breathing regulator) type, or other ap-proved oxygen equipment shown to provide thesame degree of protection, for airplanes to be oper-ated above 25,000 feet.

(ii) The pressure demand (pressure demandmask with a diluter demand pressure breathingregulator) type with mask-mounted regulator, orother approved oxygen equipment shown to pro-vide the same degree of protection, for airplanesoperated at altitudes where decompressions thatare not extremely improbable may expose theflightcrew to cabin pressure altitudes in excess of34,000 feet.

(4) Portable oxygen equipment must be imme-diately available for each cabin attendant. Theportable oxygen equipment must have the oxygendispensing unit connected to the portable oxygensupply.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–41, 42 FR 36971, July 18, 1977;Amdt. 25–87, 61 FR 28696, June 5, 1996; Amdt. 25–116, 69 FR 62789, Oct. 27, 2004]

§25.1449 Means for determining use of oxygen.There must be a means to allow the crew to de-

termine whether oxygen is being delivered to thedispensing equipment.

§25.1450 Chemical oxygen generators.(a) For the purpose of this section, a chemical

oxygen generator is defined as a device whichproduces oxygen by chemical reaction.

(b) Each chemical oxygen generator must bedesigned and installed in accordance with the fol-lowing requirements:

(1) Surface temperature developed by the gen-erator during operation may not create a hazardto the airplane or to its occupants.

(2) Means must be provided to relieve any in-ternal pressure that may be hazardous.

(c) In addition to meeting the requirements inparagraph (b) of this section, each portable chem-ical oxygen generator that is capable of sustainedoperation by successive replacement of a genera-tor 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 el-

ement may be hot, unless the element construc-tion is such that the surface temperature cannotexceed 100 degrees F.


§25.1453 Protection of oxygen equipment from rupture.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.

§25.1455 Draining of fluids subject to freezing.If fluids subject to freezing may be drained

overboard in flight or during ground operation, thedrains must be designed and located to preventthe formation of hazardous quantities of ice on theairplane as a result of the drainage.


§25.1457 Cockpit voice recorders.(a) Each cockpit voice recorder required by the

operating rules of this chapter must be approvedand must be installed so that it will record the fol-lowing:

(1) Voice communications transmitted from orreceived in the airplane by radio.

(2) Voice communications of flight crewmem-bers on the flight deck.

(3) Voice communications of flight crewmem-bers on the flight deck, using the airplane’s inter-phone system.

(4) Voice or audio signals identifying navigationor approach aids introduced into a headset orspeaker.

(5) Voice communications of flight crewmem-bers using the passenger loudspeaker system, ifthere is such a system and if the fourth channel isavailable in accordance with the requirements ofparagraph (c)(4)(ii) of this section.

(6) If datalink communication equipment is in-stalled, all datalink communications, using an ap-proved data message set. Datalink messagesmust be recorded as the output signal from thecommunications unit that translates the signal intousable data.

(b) The recording requirements of paragraph(a)(2) of this section must be met by installing acockpit-mounted area microphone, located in thebest position for recording voice communicationsoriginating at the first and second pilot stationsand voice communications of other crewmemberson the flight deck when directed to those stations.The microphone must be so located and, if neces-sary, the preamplifiers and filters of the recordermust be so adjusted or supplemented, that the in-telligibility of the recorded communications is ashigh as practicable when recorded under flightcockpit noise conditions and played back. Re-


134 ASA

peated aural or visual playback of the record maybe used in evaluating intelligibility.

(c) Each cockpit voice recorder must be in-stalled so that the part of the communication oraudio signals specified in paragraph (a) of thissection obtained from each of the followingsources is recorded on a separate channel:

(1) For the first channel, from each boom,mask, or hand-held microphone, headset, orspeaker used at the first pilot station.

(2) For the second channel from each boom,mask, or hand-held microphone, headset, orspeaker used at the second pilot station.

(3) For the third channel — from the cockpit-mounted area microphone.

(4) For the fourth channel, from—(i) Each boom, mask, or hand-held micro-

phone, headset, or speaker used at the station forthe third and fourth crew members; or

(ii) If the stations specified in paragraph (c)(4)(i)of this section are not required or if the signal atsuch a station is picked up by another channel,each microphone on the flight deck that is usedwith the passenger loudspeaker system, if its sig-nals are not picked up by another channel.

(5) As far as is practicable all sounds receivedby the microphone listed in paragraphs (c)(1), (2),and (4) of this section must be recorded withoutinterruption irrespective of the position of the in-terphone-transmitter key switch. The design shallensure that sidetone for the flight crew is pro-duced only when the interphone, public addresssystem, or radio transmitters are in use.

(d) Each cockpit voice recorder must be in-stalled so that—

(1)(i) It receives its electrical power from thebus that provides the maximum reliability for oper-ation of the cockpit voice recorder without jeopar-dizing service to essential or emergency loads.

(ii) It remains powered for as long as possiblewithout jeopardizing emergency operation of theairplane.

(2) There is an automatic means to simulta-neously stop the recorder and prevent each era-sure feature from functioning, within 10 minutesafter crash impact;

(3) There is an aural or visual means for preflightchecking of the recorder for proper operation;

(4) Any single electrical failure external to the re-corder does not disable both the cockpit voice re-corder and the flight data recorder;

(5) It has an independent power source—(i) That provides 10 ±1 minutes of electrical

power to operate both the cockpit voice recorderand cockpit-mounted area microphone;

(ii) That is located as close as practicable to thecockpit voice recorder; and

(iii) To which the cockpit voice recorder andcockpit-mounted area microphone are switchedautomatically in the event that all other power to

the cockpit voice recorder is interrupted either bynormal shutdown or by any other loss of power tothe electrical power bus; and

(6) It is in a separate container from the flightdata recorder when both are required. If used tocomply with only the cockpit voice recorder re-quirements, a combination unit may be installed.

(e) The recorder container must be located andmounted to minimize the probability of rupture ofthe container as a result of crash impact and con-sequent heat damage to the recorder from fire.

(1) Except as provided in paragraph (e)(2) ofthis section, the recorder container must be lo-cated as far aft as practicable, but need not beoutside of the pressurized compartment, and maynot be located where aft-mounted engines maycrush the container during impact.

(2) If two separate combination digital flightdata recorder and cockpit voice recorder units areinstalled instead of one cockpit voice recorder andone digital flight data recorder, the combinationunit that is installed to comply with the cockpitvoice recorder requirements may be located nearthe cockpit.

(f) If the cockpit voice recorder has a bulk era-sure device, the installation must be designed tominimize the probability of inadvertent operationand actuation of the device during crash impact.

(g) Each recorder container must—(1) Be either bright orange or bright yellow;(2) Have reflective tape affixed to its external

surface to facilitate its location under water; and(3) Have an underwater locating device, when

required by the operating rules of this chapter, onor adjacent to the container which is secured insuch manner that they are not likely to be sepa-rated during crash impact.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–2, 30 FR 3932, March 26, 1965;Amdt. 25–16, 32 FR 13914, Oct. 6, 1967; Amdt. 25–41,42 FR 36971, July 18, 1977; Amdt. 25–65, 53 FR 26143,July 11, 1988; Amdt. 25–124, 73 FR 12563, March 7,2008; Amdt. 25–124, 74 FR 32800, July 9, 2009]


ASA 135

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§25.1459 Flight data recorders.(a) Each flight recorder required by the operat-

ing rules of this chapter must be installed sothat—

(1) It is supplied with airspeed, altitude, and di-rectional data obtained from sources that meetthe accuracy requirements of §§25.1323,25.1325, and 25.1327, as appropriate;

(2) The vertical acceleration sensor is rigidly at-tached, and located longitudinally either within theapproved center of gravity limits of the airplane, orat a distance forward or aft of these limits thatdoes not exceed 25 percent of the airplane’smean aerodynamic chord;

(3)(i) It receives its electrical power from thebus that provides the maximum reliability for oper-ation of the flight data recorder without jeopardiz-ing service to essential or emergency loads.

(ii) It remains powered for as long as possiblewithout jeopardizing emergency operation of theairplane.

(4) There is an aural or visual means for pre-flight checking of the recorder for proper recordingof data in the storage medium;

(5) Except for recorders powered solely by theengine-driven electrical generator system, thereis an automatic means to simultaneously stop arecorder that has a data erasure feature and pre-vent each erasure feature from functioning, within10 minutes after crash impact;

(6) There is a means to record data from whichthe time of each radio transmission either to orfrom ATC can be determined;

(7) Any single electrical failure external to therecorder does not disable both the cockpit voicerecorder and the flight data recorder; and

(8) It is in a separate container from the cockpitvoice recorder when both are required. If used tocomply with only the flight data recorder require-ments, a combination unit may be installed. If acombination unit is installed as a cockpit voice re-corder to comply with §25.1457(e)(2), a combina-tion unit must be used to comply with this flightdata recorder requirement.

(b) Each nonejectable record container mustbe located and mounted so as to minimize theprobability of container rupture resulting fromcrash impact and subsequent damage to therecord from fire. In meeting this requirement therecord container must be located as far aft aspracticable, but need not be aft of the pressurizedcompartment, and may not be where aft-mountedengines may crush the container upon impact.

(c) A correlation must be established betweenthe flight recorder readings of airspeed, altitude,and heading and the corresponding readings(taking into account correction factors) of the firstpilot’s instruments. The correlation must cover theairspeed range over which the airplane is to beoperated, the range of altitude to which the air-plane is limited, and 360 degrees of heading. Cor-relation may be established on the ground as ap-propriate.

(d) Each recorder container must—(1) Be either bright orange or bright yellow;(2) Have reflective tape affixed to its external

surface to facilitate its location under water; and(3) Have an underwater locating device, when

required by the operating rules of this chapter, onor adjacent to the container which is secured insuch a manner that they are not likely to be sepa-rated during crash impact.

(e) Any novel or unique design or operationalcharacteristics of the aircraft shall be evaluated todetermine if any dedicated parameters must berecorded on flight recorders in addition to or inplace of existing requirements.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–8, 31 FR 127, Jan. 6, 1966;Amdt. 25–25, 35 FR 13192, Aug. 19, 1970; Amdt. 25–37,40 FR 2577, Jan. 14, 1975; Amdt. 25–41, 42 FR 36971,July 18, 1977; Amdt. 25–65, 53 FR 26144, July 11, 1988;Amdt. 25–124, 73 FR 12563, March 7, 2008; Amdt. 25–124, 74 FR 32800, July 9, 2009]


136 ASA

§25.1461 Equipment containing high energy rotors.(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 bymalfunctions, vibration, abnormal speeds, andabnormal temperatures. In addition—

(1) Auxiliary rotor cases must be able to con-tain damage caused by the failure of high energyrotor blades; and

(2) Equipment control devices, systems, andinstrumentation must reasonably ensure that nooperating limitations affecting the integrity of highenergy rotors will be exceeded in service.

(c) It must be shown by test that equipmentcontaining high energy rotors can contain any fail-ure of a high energy rotor that occurs at the high-est speed obtainable with the normal speed con-trol devices inoperative.

(d) Equipment containing high energy rotorsmust be located where rotor failure will neither en-danger the occupants nor adversely affect contin-ued safe flight.



ASA 137

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Subpart G —Operating Limitations

and Information

§25.1501 General.(a) Each operating limitation specified in

§§25.1503 through 25.1533 and other limitationsand information necessary for safe operationmust be established.

(b) The operating limitations and other informa-tion necessary for safe operation must be madeavailable to the crewmembers as prescribed in§§25.1541 through 25.1587.


OPERATING LIMITATIONS

§25.1503 Airspeed limitations: general.When airspeed limitations are a function of

weight, weight distribution, altitude, or Mach num-ber, limitations corresponding to each criticalcombination of these factors must be established.

§25.1505 Maximum operating limit speed.The maximum operating limit speed (VMO/MMO

airspeed or Mach Number, whichever is critical ata particular altitude) is a speed that may not bedeliberately exceeded in any regime of flight(climb, cruise, or descent), unless a higher speedis authorized for flight test or pilot training opera-tions. VMO/MMO must be established so that it isnot greater than the design cruising speed VC andso that it is sufficiently below VD/MD or VDF/MDF,to make it highly improbable that the latter speedswill be inadvertently exceeded in operations. Thespeed margin between VMO/MMO and VD/MD orVDF/MDF may not be less than that determinedunder §25.335(b) or found necessary during theflight tests conducted under §25.253.


§25.1507 Maneuvering speed.The maneuvering speed must be established

so that it does not exceed the design maneuver-ing speed VA determined under §25.335(c).

§25.1511 Flap extended speed.The established flap extended speed VFE must

be established so that it does not exceed the de-sign flap speed VF chosen under §§25.335(e) and25.345, for the corresponding flap positions andengine powers.

§25.1513 Minimum control speed.The minimum control speed VMC determined

under §25.149 must be established as an operat-ing limitation.

§25.1515 Landing gear speeds.(a) The established landing gear operating

speed or speeds, VLO, may not exceed the speedat which it is safe both to extend and to retract thelanding gear, as determined under §25.729 or byflight characteristics. If the extension speed is notthe same as the retraction speed, the two speedsmust be designated as VLO(EXT) and VLO(RET), re-spectively.

(b) The established landing gear extendedspeed VLE may not exceed the speed at which itis safe to fly with the landing gear secured in thefully extended position, and that determined un-der §25.729.


§ 25.1516 Other speed limitations.Any other limitation associated with speed

must be established.

[Docket No. FAA–2000–8511, 66 FR 34024, June 26,2001]

§25.1517 Rough air speed, VRA.A rough air speed, VRA, for use as the recom-

mended turbulence penetration airspeed in§25.1585(a)(8), must be established, which—

(1) Is not greater than the design airspeed formaximum gust intensity, selected for VB; and

(2) Is not less than the minimum value of VBspecified in §25.335(d); and

(3) Is sufficiently less than VMO to ensure thatlikely speed variation during rough air encounterswill not cause the overspeed warning to operatetoo frequently. In the absence of a rational investi-gation substantiating the use of other values, VRAmust be less than VMO — 35 knots (TAS).

[Docket No. 27902, 61 FR 5222, Feb. 9, 1996]

§25.1519 Weight, center of gravity, and weight distribution.The airplane weight, center of gravity, and

weight distribution limitations determined under§§25.23 through 25.27 must be established asoperating limitations.

§25.1521 Powerplant limitations.(a) General. The powerplant limitations pre-

scribed in this section must be established so thatthey do not exceed the corresponding limits forwhich the engines or propellers are type certifi-cated and do not exceed the values on which


138 ASA

compliance with any other requirement of this partis based.

(b) Reciprocating engine installations. Operat-ing limitations relating to the following must be es-tablished for reciprocating engine installations:

(1) Horsepower or torque, r.p.m., manifoldpressure, and time at critical pressure altitude andsea level pressure altitude for—

(i) Maximum continuous power (relating to un-supercharged operation or to operation in eachsupercharger mode as applicable); and

(ii) Takeoff power (relating to unsuperchargedoperation or to operation in each superchargermode as applicable).

(2) Fuel grade or specification.(3) Cylinder head and oil temperatures.(4) Any other parameter for which a limitation

has been established as part of the engine typecertificate except that a limitation need not be es-tablished for a parameter that cannot be exceededduring normal operation due to the design of theinstallation or to another established limitation.

(c) Turbine engine installations. Operating limi-tations relating to the following must be estab-lished for turbine engine installations:

(1) Horsepower, torque or thrust, r.p.m., gastemperature, and time for—

(i) Maximum continuous power or thrust (relat-ing to augmented or unaugmented operation asapplicable).

(ii) Takeoff power or thrust (relating to aug-mented or unaugmented operation as applicable).

(2) Fuel designation or specification.(3) Any other parameter for which a limitation

has been established as part of the engine typecertificate except that a limitation need not be es-tablished for a parameter that cannot be exceededduring normal operation due to the design of theinstallation or to another established limitation.

(d) Ambient temperature. An ambient tempera-ture limitation (including limitations for winterizationinstallations, if applicable) must be established asthe maximum ambient atmospheric temperatureestablished in accordance with §25.1043(b).


§25.1522 Auxiliary power unit limitations.If an auxiliary power unit is installed in the air-

plane, limitations established for the auxiliarypower unit, including categories of operation,must be specified as operating limitations for theairplane.


§25.1523 Minimum flight crew.The minimum flight crew must be established

so that it is sufficient for safe operation, consider-ing—

(a) The workload on individual crewmembers;(b) The accessibility and ease of operation of

necessary controls by the appropriate crewmem-ber; and

(c) The kind of operation authorized under§25.1525. The criteria used in making the deter-minations required by this section are set forth inAppendix D.


§25.1525 Kinds of operation.The kinds of operation to which the airplane is

limited are established by the category in which itis eligible for certification and by the installedequipment.

§ 25.1527 Ambient air temperature and operating altitude.The extremes of the ambient air temperature

and operating altitude for which operation is al-lowed, as limited by flight, structural, powerplant,functional, or equipment characteristics, must beestablished.


§25.1529 Instructions for Continued Airworthiness.The applicant must prepare Instructions for

Continued Airworthiness in accordance with Ap-pendix H to this part that are acceptable to theAdministrator. The instructions may be incompleteat type certification if a program exists to ensuretheir completion prior to delivery of the first air-plane or issuance of a standard certificate of air-worthiness, whichever occurs later.


§25.1531 Maneuvering flight load factors.Load factor limitations, not exceeding the posi-

tive limit load factors determined from the maneu-vering diagram in §25.333(b), must be estab-lished.

§25.1533 Additional operating limitations.(a) Additional operating limitations must be es-

tablished as follows:(1) The maximum takeoff weights must be es-

tablished as the weights at which compliance is


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shown with the applicable provisions of this part(including the takeoff climb provisions of§25.121(a) through (c), for altitudes and ambienttemperatures).

(2) The maximum landing weights must be es-tablished as the weights at which compliance isshown with the applicable provisions of this part(including the landing and approach climb provi-sions of §§25.119 and 25.121(d) for altitudes andambient temperatures).

(3) The minimum takeoff distances must be es-tablished as the distances at which compliance isshown with the applicable provisions of this part(including the provisions of §§25.109 and 25.113,for weights, altitudes, temperatures, wind compo-nents, runway surface conditions (dry and wet),and runway gradients for smooth, hard-surfacedrunways. Additionally, at the option of the appli-cant, wet runway takeoff distances may be estab-lished for runway surfaces that have beengrooved or treated with a porous friction course,and may be approved for use on runways wheresuch surfaces have been designed, constructed,and maintained in a manner acceptable to the Ad-ministrator.

(b) The extremes for variable factors (such asaltitude, temperature, wind, and runway gradi-ents) are those at which compliance with the ap-plicable provisions of this part is shown.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–38, 41 FR 55468, Dec. 20, 1976;Amdt. 25–72, 55 FR 29786, July 20, 1990; Amdt. 25–92,63 FR 8321, Feb. 18, 1998]

§25.1535 ETOPS approval.Except as provided in §25.3, each applicant

seeking ETOPS type design approval must com-ply with the provisions of Appendix K of this part.

[Docket No. FAA–2002–6717, 72 FR 1873, Jan. 16,2007]

MARKINGS AND PLACARDS

§25.1541 General.(a) The airplane must contain—(1) The specified markings and placards; and(2) Any additional information, instrument

markings, and placards required for the safe oper-ation if there are unusual design, operating, orhandling characteristics.

(b) Each marking and placard prescribed inparagraph (a) of this section—

(1) Must be displayed in a conspicuous place;and

(2) May not be easily erased, disfigured, or ob-scured.

§25.1543 Instrument markings: general.For each instrument—(a) When markings are on the cover glass of

the instrument, there must be means to maintainthe correct alignment of the glass cover with theface of the dial; and

(b) Each instrument marking must be clearlyvisible to the appropriate crewmember.


§25.1545 Airspeed limitation information.The airspeed limitations required by §25.1583

(a) must be easily read and understood by theflight crew.

§25.1547 Magnetic direction indicator.(a) A placard meeting the requirements of this

section must be installed on, or near, the mag-netic direction indicator.

(b) The placard must show the calibration ofthe instrument in level flight with the engines op-erating.

(c) The placard must state whether the calibra-tion was made with radio receivers on or off.

(d) Each calibration reading must be in termsof magnetic heading in not more than 45 degreeincrements.

§25.1549 Powerplant and auxiliary power unit instruments.For each required powerplant and auxiliary

power unit instrument, as appropriate to the typeof instrument—

(a) Each maximum and, if applicable, minimumsafe operating limit must be marked with a red ra-dial or a red line;

(b) Each normal operating range must bemarked with a green arc or green line, not extend-ing beyond the maximum and minimum safe limits;

(c) Each takeoff and precautionary range mustbe marked with a yellow arc or a yellow line; and

(d) Each engine, auxiliary power unit, or pro-peller speed range that is restricted because ofexcessive vibration stresses must be marked withred arcs or red lines.


§25.1551 Oil quantity indication.Each oil quantity indicating means must be

marked to indicate the quantity of oil readily andaccurately.



140 ASA

§25.1553 Fuel quantity indicator.If the unusable fuel supply for any tank exceeds

one gallon, or five percent of the tank capacity,whichever is greater, a red arc must be marked onits indicator extending from the calibrated zeroreading to the lowest reading obtainable in levelflight.

§25.1555 Control markings.(a) Each cockpit control, other than primary

flight controls and controls whose function is obvi-ous, must be plainly marked as to its function andmethod of operation.

(b) Each aerodynamic control must be markedunder the requirements of §§25.677 and 25.699.

(c) For powerplant fuel controls—(1) Each fuel tank selector control must be

marked to indicate the position corresponding toeach tank and to each existing cross feed position;

(2) If safe operation requires the use of anytanks in a specific sequence, that sequence mustbe marked on, or adjacent to, the selector forthose tanks; and

(3) Each valve control for each engine must bemarked to indicate the position corresponding toeach engine controlled.

(d) For accessory, auxiliary, and emergencycontrols —

(1) Each emergency control (including eachfuel jettisoning and fluid shutoff must be coloredred; and

(2) Each visual indicator required by §25.729(e)must be marked so that the pilot can determine atany time when the wheels are locked in either ex-treme position, if retractable landing gear is used.

§25.1557 Miscellaneous markings and placards.(a) Baggage and cargo compartments and bal-

last location. Each baggage and cargo compart-ment, and each ballast location must have a plac-ard stating any limitations on contents, includingweight, that are necessary under the loading re-quirements. However, underseat compartmentsdesigned for the storage of carry-on articlesweighing not more than 20 pounds need not havea loading limitation placard.

(b) Powerplant fluid filler openings. The follow-ing apply:

(1) Fuel filler openings must be marked at ornear the filler cover with—

(i) The word “fuel”;(ii) For reciprocating engine powered airplanes,

the minimum fuel grade;(iii) For turbine engine powered airplanes, the

permissible fuel designations; and(iv) For pressure fueling systems, the maxi-

mum permissible fueling supply pressure and themaximum permissible defueling pressure.

(2) Oil filler openings must be marked at ornear the filler cover with the word “oil.”

(3) Augmentation fluid filler openings must bemarked at or near the filler cover to identify the re-quired fluid.

(c) Emergency exit placards. Each emergencyexit placard must meet the requirements of§25.811.

(d) Doors. Each door that must be used in or-der to reach any required emergency exit musthave a suitable placard stating that the door is tobe latched in the open position during takeoff andlanding.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–32, 37 FR 3972, Feb. 24, 1972;Amdt. 25–38, 41 FR 55468, Dec. 20, 1976; Amdt. 25–72, 55 FR 29786, July 20, 1990]

§25.1561 Safety equipment.(a) Each safety equipment control to be oper-

ated by the crew in emergency, such as controlsfor automatic liferaft releases, must be plainlymarked as to its method of operation.

(b) Each location, such as a locker or compart-ment, that carries any fire extinguishing, signal-ing, or other life saving equipment must bemarked accordingly.

(c) Stowage provisions for required emergencyequipment must be conspicuously marked toidentify the contents and facilitate the easy re-moval of the equipment.

(d) Each liferaft must have obviously markedoperating instructions.

(e) Approved survival equipment must bemarked for identification and method of operation.


§25.1563 Airspeed placard.A placard showing the maximum airspeeds for

flap extension for the takeoff, approach, and land-ing positions must be installed in clear view ofeach pilot.

AIRPLANE FLIGHT MANUAL

§25.1581 General.(a) Furnishing information. An Airplane Flight

Manual must be furnished with each airplane, andit must contain the following:

(1) Information required by §§25.1583 through25.1587.

(2) Other information that is necessary for safeoperation because of design, operating, or han-dling characteristics.

(3) Any limitation, procedure, or other informa-tion established as a condition of compliance withthe applicable noise standards of part 36 of thischapter.


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(b) Approved information. Each part of themanual listed in §§25.1583 through 25.1587, thatis appropriate to the airplane, must be furnished,verified, and approved, and must be segregated,identified, and clearly distinguished from each un-approved part of that manual.

(c) [Reserved](d) Each Airplane Flight Manual must include a

table of contents if the complexity of the manualindicates a need for it.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–42, 43 FR 2323, Jan. 16, 1978;Amdt. 25–72, 55 FR 29786, July 20, 1990]

§25.1583 Operating limitations.(a) Airspeed limitations. The following airspeed

limitations and any other airspeed limitations nec-essary for safe operation must be furnished:

(1) The maximum operating limit speedVMO/MMO and a statement that this speed limitmay not be deliberately exceeded in any regimeof flight (climb, cruise, or descent) unless a higherspeed is authorized for flight test or pilot training.

(2) If an airspeed limitation is based upon com-pressibility effects, a statement to this effect andinformation as to any symptoms, the probable be-havior of the airplane, and the recommended re-covery procedures.

(3) The maneuvering speed VA and a state-ment that full application of rudder and aileroncontrols, as well as maneuvers that involve anglesof attack near the stall, should be confined tospeeds below this value.

(4) The flap extended speed VFE and the perti-nent flap positions and engine powers.

(5) The landing gear operating speed orspeeds, and a statement explaining the speedsas defined in §25.1515(a).

(6) The landing gear extended speed VLE, ifgreater than VLO, and a statement that this is themaximum speed at which the airplane can besafely flown with the landing gear extended.

(b) Powerplant limitations. The following infor-mation must be furnished:

(1) Limitations required by §25.1521 and§25.1522.

(2) Explanation of the limitations, when appro-priate.

(3) Information necessary for marking the instru-ments required by §§25.1549 through 25.1553.

(c) Weight and loading distribution. The weightand center of gravity limitations established under§25.1519 must be furnished in the Airplane FlightManual. All of the following information, includingthe weight distribution limitations established un-der §25.1519, must be presented either in the Air-plane Flight Manual or in a separate weight andbalance control and loading document that is in-

corporated by reference in the Airplane FlightManual:

(1) The condition of the airplane and the itemsincluded in the empty weight as defined in accor-dance with §25.29.

(2) Loading instructions necessary to ensureloading of the airplane within the weight and cen-ter of gravity limits, and to maintain the loadingwithin these limits in flight.

(3) If certification for more than one center ofgravity range is requested, the appropriate limita-tions, with regard to weight and loading proce-dures, for each separate center of gravity range.

(d) Flight crew. The number and functions ofthe minimum flight crew determined under§25.1523 must be furnished.

(e) Kinds of operation. The kinds of operationapproved under §25.1525 must be furnished.

(f) Ambient air temperatures and operating alti-tudes. The extremes of the ambient air tempera-tures and operating altitudes established under§25.1527 must be furnished.

(g) [Reserved](h) Additional operating limitations. The operat-

ing limitations established under §25.1533 mustbe furnished.

(i) Maneuvering flight load factors. The positivemaneuvering limit load factors for which the struc-ture is proven, described in terms of accelera-tions, must be furnished.

[Docket No. 5066, 29 FR 1891, Dec. 24, 1964; asamended by Amdt. 25–38, 41 FR 55468, Dec, 20, 1976;Amdt. 25–42, 43 FR 2323, Jan. 16, 1978; Amdt. 25–46,43 FR 50598, Oct. 30, 1978; Amdt. 25–72, 55 FR 29787,July 20, 1990; Amdt. 25–105, 66 FR 34024, June 26,2001]

§25.1585 Operating procedures.(a) Operating procedures must be furnished

for—(1) Normal procedures peculiar to the particu-

lar type or model encountered in connection withroutine operations;

(2) Non-normal procedures for malfunctioncases and failure conditions involving the use ofspecial systems or the alternative use of regularsystems; and

(3) Emergency procedures for foreseeable butunusual situations in which immediate and pre-cise action by the crew may be expected to sub-stantially reduce the risk of catastrophe.

(b) Information or procedures not directly re-lated to airworthiness or not under the control ofthe crew, must not be included, nor must any pro-cedure that is accepted as basic airmanship.

(c) Information identifying each operating con-dition in which the fuel system independence pre-scribed in §25.953 is necessary for safety mustbe furnished, together with instructions for placing


142 ASA

the fuel system in a configuration used to showcompliance with that section.

(d) The buffet onset envelopes, determined un-der §25.251 must be furnished. The buffet onsetenvelopes presented may reflect the center ofgravity at which the airplane is normally loadedduring cruise if corrections for the effect of differ-ent center of gravity locations are furnished.

(e) Information must be furnished that indicatesthat when the fuel quantity indicator reads “zero”in level flight, any fuel remaining in the fuel tankcannot be used safely in flight.

(f) Information on the total quantity of usablefuel for each fuel tank must be furnished.


§25.1587 Performance information.(a) Each Airplane Flight Manual must contain

information to permit conversion of the indicatedtemperature to free air temperature if other than afree air temperature indicator is used to complywith the requirements of §25.1303(a)(1).

(b) Each Airplane Flight Manual must containthe performance information computed under theapplicable provisions of this part (including§§25.115, 25.123, and 25.125 for the weights, al-titudes, temperatures, wind components, and run-way gradients, as applicable) within the opera-tional limits of the airplane, and must contain thefollowing:

(1) In each case, the conditions of power, con-figuration, and speeds, and the procedures forhandling the airplane and any system having asignificant effect on the performance information.

(2) VSR determined in accordance with§25.103.

(3) The following performance information (de-termined by extrapolation and computed for therange of weights between the maximum landingweight and the maximum takeoff weight):

(i) Climb in the landing configuration.(ii) Climb in the approach configuration.(iii) Landing distance.(4) Procedures established under §25.101(f)

and (g) that are related to the limitations and infor-mation required by §25.1533 and by this para-graph (b) in the form of guidance material, includ-ing any relevant limitations or information.

(5) An explanation of significant or unusualflight or ground handling characteristics of the air-plane.

(6) Corrections to indicated values of airspeed,altitude, and outside air temperature.

(7) An explanation of operational landing run-way length factors included in the presentation ofthe landing distance, if appropriate.

[Docket No. FAA–2000–8511, 66 FR 34024, June 26,2001; Amdt. 25–108, 67 FR 70828, Nov. 26, 2002]


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Subpart H — Electrical Wiring Interconnection

Systems (EWIS)Source: FAA–2004–18379, 72 FR 63406, Nov. 8, 2007,unless otherwise noted.

§25.1701 Definition.(a) As used in this chapter, electrical wiring

interconnection system (EWIS) means any wire,wiring device, or combination of these, includingtermination devices, installed in any area of theairplane for the purpose of transmitting electricalenergy, including data and signals, between twoor more intended termination points. This in-cludes:

(1) Wires and cables.(2) Bus bars.(3) The termination point on electrical devices,

including those on relays, interrupters, switches,contactors, terminal blocks and circuit breakers,and other circuit protection devices.

(4) Connectors, including feed-through connec-tors.

(5) Connector accessories.(6) Electrical grounding and bonding devices

and their associated connections.(7) Electrical splices.(8) Materials used to provide additional protec-

tion for wires, including wire insulation, wire sleev-ing, and conduits that have electrical terminationfor the purpose of bonding.

(9) Shields or braids.(10) Clamps and other devices used to route

and support the wire bundle.(11) Cable tie devices.(12) Labels or other means of identification.(13) Pressure seals.(14) EWIS components inside shelves, panels,

racks, junction boxes, distribution panels, andback-planes of equipment racks, including, butnot limited to, circuit board back-planes, wire inte-gration units, and external wiring of equipment.

(b) Except for the equipment indicated in para-graph (a)(14) of this section, EWIS componentsinside the following equipment, and the externalconnectors that are part of that equipment, areexcluded from the definition in paragraph (a) ofthis section:

(1) Electrical equipment or avionics that arequalified to environmental conditions and testingprocedures when those conditions and proce-dures are—

(i) Appropriate for the intended function and op-erating environment, and

(ii) Acceptable to the FAA.(2) Portable electrical devices that are not part

of the type design of the airplane. This includes

personal entertainment devices and laptop com-puters.

(3) Fiber optics.

§25.1703 Function and installation: EWIS.(a) Each EWIS component installed in any area

of the aircraft must:(1) Be of a kind and design appropriate to its in-

tended function.(2) Be installed according to limitations speci-

fied for the EWIS components.(3) Perform the function for which it was in-

tended without degrading the airworthiness of theairplane.

(4) Be designed and installed in a way that willminimize mechanical strain.

(b) Selection of wires must take into accountknown characteristics of the wire in relation toeach installation and application to minimize therisk of wire damage, including any arc trackingphenomena.

(c) The design and installation of the mainpower cables (including generator cables) in thefuselage must allow for a reasonable degree ofdeformation and stretching without failure.

(d) EWIS components located in areas ofknown moisture accumulation must be protectedto minimize any hazardous effects due to mois-ture.

§25.1705 Systems and functions: EWIS.(a) EWIS associated with any system required

for type certification or by operating rules must beconsidered an integral part of that system andmust be considered in showing compliance withthe applicable requirements for that system.

(b) For systems to which the following rules ap-ply, the components of EWIS associated withthose systems must be considered an integralpart of that system or systems and must be con-sidered in showing compliance with the applicablerequirements for that system.

(1) §25.773(b)(2) Pilot compartment view.(2) §25.981 Fuel tank ignition prevention.(3) §25.1165 Engine ignition systems.(4) §25.1310 Power source capacity and distri-

bution.(5) §25.1316 System lightning protection.(6) §25.1331(a)(2) Instruments using a power

supply.(7) §25.1351 General.(8) §25.1355 Distribution system.(9) §25.1360 Precautions against injury.(10) §25.1362 Electrical supplies for emer-

gency conditions.(11) §25.1365 Electrical appliances, motors,

and transformers.(12) §25.1431(c) and (d) Electronic equipment.


144 ASA

§25.1707 System separation: EWIS.(a) Each EWIS must be designed and installed

with adequate physical separation from otherEWIS and airplane systems so that an EWIScomponent failure will not create a hazardouscondition. Unless otherwise stated, for the pur-poses of this section, adequate physical separa-tion must be achieved by separation distance orby a barrier that provides protection equivalent tothat separation distance.

(b) Each EWIS must be designed and installedso that any electrical interference likely to bepresent in the airplane will not result in hazardouseffects upon the airplane or its systems.

(c) Wires and cables carrying heavy current,and their associated EWIS components, must bedesigned and installed to ensure adequate physi-cal separation and electrical isolation so thatdamage to circuits associated with essential func-tions will be minimized under fault conditions.

(d) Each EWIS associated with independentairplane power sources or power sources con-nected in combination must be designed and in-stalled to ensure adequate physical separationand electrical isolation so that a fault in any oneairplane power source EWIS will not adversely af-fect any other independent power sources. In ad-dition:

(1) Airplane independent electrical powersources must not share a common ground termi-nating location.

(2) Airplane system static grounds must notshare a common ground terminating location withany of the airplane’s independent electrical powersources.

(e) Except to the extent necessary to provideelectrical connection to the fuel systems compo-nents, the EWIS must be designed and installedwith adequate physical separation from fuel linesand other fuel system components, so that:

(1) An EWIS component failure will not create ahazardous condition.

(2) Any fuel leakage onto EWIS componentswill not create a hazardous condition.

(f) Except to the extent necessary to provideelectrical connection to the hydraulic systemscomponents, EWIS must be designed and in-stalled with adequate physical separation fromhydraulic lines and other hydraulic system compo-nents, so that:


(2) Any hydraulic fluid leakage onto EWIS com-ponents will not create a hazardous condition.

(g) Except to the extent necessary to provideelectrical connection to the oxygen systems com-ponents, EWIS must be designed and installedwith adequate physical separation from oxygenlines and other oxygen system components, so

that an EWIS component failure will not create ahazardous condition.

(h) Except to the extent necessary to provideelectrical connection to the water/waste systemscomponents, EWIS must be designed and in-stalled with adequate physical separation fromwater/waste lines and other water/waste systemcomponents, so that:


(2) Any water/waste leakage onto EWIS com-ponents will not create a hazardous condition.

(i) EWIS must be designed and installed withadequate physical separation between the EWISand flight or other mechanical control systems ca-bles and associated system components, so that:

(1) Chafing, jamming, or other interference areprevented.


(3) Failure of any flight or other mechanicalcontrol systems cables or systems componentswill not damage the EWIS and create a hazard-ous condition.

(j) EWIS must be designed and installed withadequate physical separation between the EWIScomponents and heated equipment, hot air ducts,and lines, so that:


(2) Any hot air leakage or heat generated ontoEWIS components will not create a hazardouscondition.

(k) For systems for which redundancy is re-quired, by certification rules, by operating rules, oras a result of the assessment required by§25.1709, EWIS components associated withthose systems must be designed and installedwith adequate physical separation.

(l) Each EWIS must be designed and installedso there is adequate physical separation betweenit and other aircraft components and aircraft struc-ture, and so that the EWIS is protected fromsharp edges and corners, to minimize potentialfor abrasion/chafing, vibration damage, and othertypes of mechanical damage.

§25.1709 System safety: EWIS.Each EWIS must be designed and installed so

that:(a) Each catastrophic failure condition—(1) Is extremely improbable; and(2) Does not result from a single failure.(b) Each hazardous failure condition is ex-

tremely remote.


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§25.1711 Component identification: EWIS.(a) EWIS components must be labeled or

otherwise identified using a consistent methodthat facilitates identification of the EWIS compo-nent, its function, and its design limitations, if any.

(b) For systems for which redundancy is re-quired, by certification rules, by operating rules, oras a result of the assessment required by§25.1709, EWIS components associated withthose systems must be specifically identified withcomponent part number, function, and separationrequirement for bundles.

(1) The identification must be placed along thewire, cable, or wire bundle at appropriate intervalsand in areas of the airplane where it is readily vis-ible to maintenance, repair, or alteration person-nel.

(2) If an EWIS component cannot be markedphysically, then other means of identification mustbe provided.

(c) The identifying markings required by para-graphs (a) and (b) of this section must remain leg-ible throughout the expected service life of theEWIS component.

(d) The means used for identifying each EWIScomponent as required by this section must nothave an adverse effect on the performance of thatcomponent throughout its expected service life.

(e) Identification for EWIS modifications to thetype design must be consistent with the identifica-tion scheme of the original type design.

§25.1713 Fire protection: EWIS.(a) All EWIS components must meet the appli-

cable fire and smoke protection requirements of§25.831(c) of this part.

(b) EWIS components that are located in des-ignated fire zones and are used during emer-gency procedures must be fire resistant.

(c) Insulation on electrical wire and electricalcable, and materials used to provide additionalprotection for the wire and cable, installed in anyarea of the airplane, must be self-extinguishingwhen tested in accordance with the applicableportions of Appendix F, part I, of 14 CFR part 25.

§25.1715 Electrical bonding and protection against static electricity: EWIS.(a) EWIS components used for electrical bond-

ing and protection against static electricity mustmeet the requirements of §25.899.

(b) On airplanes having grounded electricalsystems, electrical bonding provided by EWIScomponents must provide an electrical returnpath capable of carrying both normal and faultcurrents without creating a shock hazard or dam-

age to the EWIS components, other airplane sys-tem components, or airplane structure.

§25.1717 Circuit protective devices: EWIS.Electrical wires and cables must be designed

and installed so they are compatible with the cir-cuit protection devices required by §25.1357, sothat a fire or smoke hazard cannot be created un-der temporary or continuous fault conditions.

§25.1719 Accessibility provisions: EWIS.Access must be provided to allow inspection

and replacement of any EWIS component as nec-essary for continued airworthiness.

§25.1721 Protection of EWIS.(a) No cargo or baggage compartment may

contain any EWIS whose damage or failure mayaffect safe operation, unless the EWIS is pro-tected so that:

(1) It cannot be damaged by movement ofcargo or baggage in the compartment.

(2) Its breakage or failure will not create a firehazard.

(b) EWIS must be designed and installed tominimize damage and risk of damage to EWIS bymovement of people in the airplane during allphases of flight, maintenance, and servicing.

(c) EWIS must be designed and installed tominimize damage and risk of damage to EWIS byitems carried onto the aircraft by passengers orcabin crew.

§25.1723 Flammable fluid fire protection: EWIS.EWIS components located in each area where

flammable fluid or vapors might escape by leak-age of a fluid system must be considered a poten-tial ignition source and must meet the require-ments of §25.863.

§25.1725 Powerplants: EWIS.(a) EWIS associated with any powerplant must

be designed and installed so that the failure of anEWIS component will not prevent the continuedsafe operation of the remaining powerplants or re-quire immediate action by any crewmember forcontinued safe operation, in accordance with therequirements of §25.903(b).

(b) Design precautions must be taken to mini-mize hazards to the airplane due to EWIS dam-age in the event of a powerplant rotor failure or afire originating within the powerplant that burnsthrough the powerplant case, in accordance withthe requirements of §25.903(d)(1).


146 ASA

§25.1727 Flammable fluid shutoff means: EWIS.EWIS associated with each flammable fluid

shutoff means and control must be fireproof ormust be located and protected so that any fire in afire zone will not affect operation of the flammablefluid shutoff means, in accordance with the re-quirements of §25.1189.

§25.1729 Instructions for Continued Airworthiness: EWIS.The applicant must prepare Instructions for

Continued Airworthiness applicable to EWIS inaccordance with Appendix H sections H25.4 andH25.5 to this part that are approved by the FAA.

§25.1731 Powerplant and APU fire detector system: EWIS.(a) EWIS that are part of each fire or overheat

detector system in a fire zone must be fire-resis-tant.

(b) No EWIS component of any fire or overheatdetector system for any fire zone may passthrough another fire zone, unless:

(1) It is protected against the possibility of falsewarnings resulting from fires in zones throughwhich it passes; or

(2) Each zone involved is simultaneously pro-tected by the same detector and extinguishingsystem.

(c) EWIS that are part of each fire or overheatdetector system in a fire zone must meet the re-quirements of §25.1203.

§25.1733 Fire detector systems, general: EWIS.EWIS associated with any installed fire protec-

tion system, including those required by §§25.854and 25.858, must be considered an integral partof the system in showing compliance with the ap-plicable requirements for that system.

Part 25: Airworthiness Standards: Transport Category Appendix A to Part 25

ASA 147

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APPENDIX A TO PART 25

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148 ASA

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ASA 149

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THE AIRPLANE INERTIA LOADS REQUIREDTO BALANCE THE EXTERNAL FORCES

SINGLE WHEEL LOADFROM 2 WHEEL LEVELLANDING CONDITION.

NOSE OR TAIL WHEEL TYPE

FIGURE 4

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0.80VM 0.60VM

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* NOSE GEAR GROUND REACTION = 0

NOSE OR TAIL WHEEL TYPE AIRPLANE IN LEVEL ALTITUDE

= ONE-HALF THE MAXIMUM VERTICAL GROUND REACTION CONTAINED AT EACH MAIN GEAR IN THE LEVEL LANDING CONDITIONS.

FIGURE 5—Lateral drift landing.

Appendix A to Part 25 Federal Aviation Regulations

150 ASA

I

1.2W

(AT

DE

SIG

N L

AN

DIN

G W

EIG

HT

)1.

0W (

AT D

ES

IGN

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)

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ASA 151

25

FIGURE 8—Pivoting, nose or tail wheel type.

CENTER OF ROTATION

T

W

VM VN VM

VN and VM are static ground reactions. For tail wheel type the airplane is inthe three point attitude. Pivoting is assumed to take place about one mainlanding gear unit.

Appendix B to Part 25 Federal Aviation Regulations

152 ASA

APPENDIX B TO PART 25

CL CL

ß = ßk

Unflared Bottom

FIGURE 1 — Pictorial definition of angles, dimensions, and directions on a seaplane.

Flared Bottom

ß

ßk

Flare

y

Y

Z

X

Forebody Afterbody

Z

Part 25: Airworthiness Standards: Transport Category Appendix B to Part 25

ASA 153

25

FIGURE 2 — Hull station weighing factor.

1.5

1.0

0.375

1.0

CG

Afterbody Length LaForebody Length Lf

K1 (Vertical Loads)

2.0

1.00.5

1.0

Afterbody Length LaForebody Length Lf

0.75

Lf /2

K2 (Bottom Pressures)

Appendix B to Part 25 Federal Aviation Regulations

154 ASA

FIGURE 3 — Transverse pressure distributions.

UNFLARED SYMMETRICAL

Local Pressure Distributed Pressure

FLARED UNSYMMETRICAL

Pk

0.75 P kP P

FLARE

Pch

Pk

0.75 P k

P

P/2

Part 25: Airworthiness Standards: Transport Category Appendix C to Part 25

ASA 155

25

APPENDIX C TO PART 25PART I — ATMOSPHERIC ICING CONDITIONS

(a) Continuous maximum icing. The maximumcontinuous intensity of atmospheric icing condi-tions (continuous maximum icing) is defined bythe variables of the cloud liquid water content, themean effective diameter of the cloud droplets, theambient air temperature, and the interrelationshipof these three variables as shown in figure 1 ofthis appendix. The limiting icing envelope in termsof altitude and temperature is given in figure 2 ofthis appendix. The inter-relationship of cloud liq-uid water content with drop diameter and altitudeis determined from figures 1 and 2. The cloud liq-uid water content for continuous maximum icingconditions of a horizontal extent, other than 17.4nautical miles, is determined by the value of liquidwater content of figure 1, multiplied by the appro-priate factor from figure 3 of this appendix.

(b) Intermittent maximum icing. The intermit-tent maximum intensity of atmospheric icing con-ditions (intermittent maximum icing) is defined bythe variables of the cloud liquid water content, themean effective diameter of the cloud droplets, theambient air temperature, and the interrelationshipof these three variables as shown in figure 4 ofthis appendix. The limiting icing envelope in termsof altitude and temperature is given in figure 5 ofthis appendix. The inter-relationship of cloud liq-uid water content with drop diameter and altitudeis determined from figures 4 and 5. The cloud liq-uid water content for intermittent maximum icingconditions of a horizontal extent, other than 2.6nautical miles, is determined by the value of cloudliquid water content of figure 4 multiplied by theappropriate factor in figure 6 of this appendix.

(c) Takeoff maximum icing. The maximum in-tensity of atmospheric icing conditions for takeoff(takeoff maximum icing) is defined by the cloudliquid water content of 0.35 g/m3, the mean effec-tive diameter of the cloud droplets of 20 microns,and the ambient air temperature at ground level ofminus 9 degrees Celsius (-9°C). The takeoff max-imum icing conditions extend from ground level toa height of 1,500 feet above the level of the take-off surface.

PART II — AIRFRAME ICE ACCRETIONS FOR SHOWING COMPLIANCE WITH SUBPART B.(a) Ice accretions—General. The most critical

ice accretion in terms of airplane performanceand handling qualities for each flight phase mustbe used to show compliance with the applicableairplane performance and handling requirementsin icing conditions of subpart B of this part. Appli-cants must demonstrate that the full range of at-mospheric icing conditions specified in part I ofthis appendix have been considered, including

the mean effective drop diameter, liquid watercontent, and temperature appropriate to the flightconditions (for example, configuration, speed, an-gle-of-attack, and altitude). The ice accretions foreach flight phase are defined as follows:

(1) Takeoff ice is the most critical ice accretionon unprotected surfaces and any ice accretion onthe protected surfaces appropriate to normal iceprotection system operation, occurring betweenliftoff and 400 feet above the takeoff surface, as-suming accretion starts at liftoff in the takeoffmaximum icing conditions of part I, paragraph (c)of this appendix.

(2) Final takeoff ice is the most critical ice ac-cretion on unprotected surfaces, and any ice ac-cretion on the protected surfaces appropriate tonormal ice protection system operation, between400 feet and either 1,500 feet above the takeoffsurface, or the height at which the transition fromthe takeoff to the en route configuration is com-pleted and VFTO is reached, whichever is higher.Ice accretion is assumed to start at liftoff in thetakeoff maximum icing conditions of part I, para-graph (c) of this appendix.

(3) En route ice is the critical ice accretion onthe unprotected surfaces, and any ice accretionon the protected surfaces appropriate to normalice protection system operation, during the enroute phase.

(4) Holding ice is the critical ice accretion onthe unprotected surfaces, and any ice accretionon the protected surfaces appropriate to normalice protection system operation, during the hold-ing flight phase.

(5) Approach ice is the critical ice accretion onthe unprotected surfaces, and any ice accretionon the protected surfaces appropriate to normalice protection system operation following exit fromthe holding flight phase and transition to the mostcritical approach configuration.

(6) Landing ice is the critical ice accretion onthe unprotected surfaces, and any ice accretionon the protected surfaces appropriate to normalice protection system operation following exit fromthe approach flight phase and transition to the fi-nal landing configuration.

(b) In order to reduce the number of ice accre-tions to be considered when demonstrating com-pliance with the requirements of §25.21(g), any ofthe ice accretions defined in paragraph (a) of thissection may be used for any other flight phase if itis shown to be more critical than the specific iceaccretion defined for that flight phase. Configura-tion differences and their effects on ice accretionsmust be taken into account.

(c) The ice accretion that has the most adverseeffect on handling qualities may be used for air-plane performance tests provided any differencein performance is conservatively taken into ac-count.

Appendix C to Part 25 Federal Aviation Regulations

156 ASA

(d) For both unprotected and protected parts,the ice accretion for the takeoff phase may be de-termined by calculation, assuming the takeoffmaximum icing conditions defined in appendix C,and assuming that:

(1) Airfoils, control surfaces and, if applicable,propellers are free from frost, snow, or ice at thestart of the takeoff;

(2) The ice accretion starts at liftoff;(3) The critical ratio of thrust/power-to-weight;(4) Failure of the critical engine occurs at VEF;

and(5) Crew activation of the ice protection system

is in accordance with a normal operating proce-dure provided in the Airplane Flight Manual, ex-cept that after beginning the takeoff roll, it must beassumed that the crew takes no action to activatethe ice protection system until the airplane is atleast 400 feet above the takeoff surface.

(e) The ice accretion before the ice protectionsystem has been activated and is performing itsintended function is the critical ice accretionformed on the unprotected and normally pro-tected surfaces before activation and effective op-eration of the ice protection system in continuousmaximum atmospheric icing conditions. This iceaccretion only applies in showing compliance to§§25.143(j) and 25.207(h), and 25.207(i).

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–121, 72 FR 44669, Aug. 8, 2007;Amdt. 25–129, 74 FR 38340, Aug. 3, 2009]


ASA 157

25

0

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158 ASA

FIGURE 2

AM

BIE

NT

TE

MP

ER

ATU

RE

– °

F

PRESSURE ALTITUDE – 1000 FT.

CONTINUOUS MAXIMUM (STRATIFORM CLOUDS)ATMOSPHERIC ICING CONDITIONS

AMBIENT TEMPERATURE VS PRESSURE ALTITUDE

SOURCE OF DATA

NACA TN NO. 2569

0 4 8 12 16 20 24

32

26

20

14

8

2

-4

-10

-16

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ASA 159

25

FIG

UR

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Liquid Water Content Factor, F-Dimensionless

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273

8

300

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160 ASA

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LIQUID WATER CONTENT – GRAMS PER CUBIC METER

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ASA 161

25

PRESSURE ALTITUDE – 1000 FEET

0 4 8 12 16 20 24 28 30

NOTE:DASHED LINES INDICATEPOSSIBLE EXTENT OFLIMITS.

AM

BIE

NT

TE

MP

ER

ATU

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– °

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20

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INTERMITTENT MAXIMUM (CUMULIFORM CLOUDS)ATMOSPHERIC ICING CONDITIONS

AMBIENT TEMPERATURE VS PRESSURE ALTITUDE

SOURCE OF DATA

NACA TN NO. 2569

FIGURE 5


162 ASA

Liquid Water Content Factor, F-Dimensionless

CLO

UD

HO

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ON

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EX

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NT

— N

AU

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0.3

0.4

0.5

0.6

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1.0

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2.0

3.0

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0.9

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1.35

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No.

273

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FIG

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Part 25: Airworthiness Standards: Transport Category Appendix D to Part 25

ASA 163

25

APPENDIX D TO PART 25Criteria for determining minimum flight crew. The

following are considered by the Agency in determin-ing the minimum flight crew under §25.1523:

(a) Basic workload functions. The following ba-sic workload functions are considered:

(1) Flight path control.(2) Collision avoidance.(3) Navigation.(4) Communications.(5) Operation and monitoring of aircraft en-

gines and systems.(6) Command decisions.(b) Workload factors. The following workload

factors are considered significant when analyzingand demonstrating workload for minimum flightcrew determination:

(1) The accessibility, ease, and simplicity of op-eration of all necessary flight, power, and equip-ment controls, including emergency fuel shutoffvalves, electrical controls, electronic controls, pres-surization system controls, and engine controls.

(2) The accessibility and conspicuity of all nec-essary instruments and failure warning devicessuch as fire warning, electrical system malfunc-tion, and other failure or caution indicators. Theextent to which such instruments or devices directthe proper corrective action is also considered.

(3) The number, urgency, and complexity of op-erating procedures with particular considerationgiven to the specific fuel management scheduleimposed by center of gravity, structural or otherconsiderations of an airworthiness nature, and tothe ability of each engine to operate at all timesfrom a single tank or source which is automati-cally replenished if fuel is also stored in othertanks.

(4) The degree and duration of concentratedmental and physical effort involved in normal op-eration and in diagnosing and coping with mal-functions and emergencies.

(5) The extent of required monitoring of thefuel, hydraulic, pressurization, electrical, elec-tronic, deicing, and other systems while en route.

(6) The actions requiring a crewmember to beunavailable at his assigned duty station, includ-ing: observation of systems, emergency operationof any control, and emergencies in any compart-ment.

(7) The degree of automation provided in theaircraft systems to afford (after failures or mal-functions) automatic crossover or isolation of diffi-culties to minimize the need for flight crew actionto guard against loss of hydraulic or electric powerto flight controls or to other essential systems.

(8) The communications and navigation work-load.

(9) The possibility of increased workload asso-ciated with any emergency that may lead to otheremergencies.

(10) Incapacitation of a flight crewmemberwhenever the applicable operating rule requires aminimum flight crew of at least two pilots.

(c) Kind of operation authorized. The determi-nation of the kind of operation authorized requiresconsideration of the operating rules under whichthe airplane will be operated. Unless an applicantdesires approval for a more limited kind of opera-tion. It is assumed that each airplane certificatedunder this Part will operate under IFR conditions.


Appendix E to Part 25 Federal Aviation Regulations

164 ASA

APPENDIX E TO PART 25I — LIMITED WEIGHT CREDIT FOR AIRPLANES

EQUIPPED WITH STANDBY POWER

(a) Each applicant for an increase in the maxi-mum certificated takeoff and landing weights ofan airplane equipped with a type-certificatedstandby power rocket engine may obtain an in-crease as specified in paragraph (b) if—

(1) The installation of the rocket engine hasbeen approved and it has been established byflight test that the rocket engine and its controlscan be operated safely and reliably at the in-crease in maximum weight; and

(2) The Airplane Flight Manual, or the placard,markings or manuals required in place thereof,set forth in addition to any other operating limita-tions the Administrator may require, the increasedweight approved under this regulation and a pro-hibition against the operation of the airplane atthe approved increased weight when—

(i) The installed standby power rocket engineshave been stored or installed in excess of thetime limit established by the manufacturer of therocket engine (usually stenciled on the enginecasing); or

(ii) The rocket engine fuel has been expendedor discharged.

(b) The currently approved maximum takeoffand landing weights at which an airplane is certif-icated without a standby power rocket engine in-stallation may be increased by an amount thatdoes not exceed any of the following:

(1) An amount equal in pounds to 0.014 IN,where I is the maximum usable impulse in pounds-seconds available from each standby power rocketengine and N is the number of rocket engines in-stalled.

(2) An amount equal to 5 percent of the maxi-mum certificated weight approved in accordancewith the applicable airworthiness regulations with-out standby power rocket engines installed.

(3) An amount equal to the weight of the rocketengine installation.

(4) An amount that, together with the currentlyapproved maximum weight, would equal the max-imum structural weight established for the air-plane without standby rocket engines installed.

II — PERFORMANCE CREDIT FOR TRANSPORT CATEGORY AIRPLANES EQUIPPED

WITH STANDBY POWER

The Administrator may grant performancecredit for the use of standby power on transportcategory airplanes. However, the performancecredit applies only to the maximum certificatedtakeoff and landing weights, the takeoff distance,and the takeoff paths, and may not exceed thatfound by the Administrator to result in an overall

level of safety in the takeoff, approach, and land-ing regimes of flight equivalent to that prescribedin the regulations under which the airplane wasoriginally certificated without standby power. Forthe purposes of this Appendix, “standby power” ispower or thrust, or both, obtained from rocket en-gines for a relatively short period and actuatedonly in cases of emergency. The following provi-sions apply:

(1) Takeoff; general. The takeoff data pre-scribed in paragraphs (2) and (3) of this Appendixmust be determined at all weights and altitudes,and at ambient temperatures if applicable, atwhich performance credit is to be applied.

(2) Takeoff path.(a) The one-engine-inoperative takeoff path

with standby power in use must be determined inaccordance with the performance requirements ofthe applicable airworthiness regulations.

(b) The one-engine-inoperative takeoff path(excluding that part where the airplane is on orjust above the takeoff surface) determined in ac-cordance with paragraph (a) of this section mustlie above the one-engine-inoperative takeoff pathwithout standby power at the maximum takeoffweight at which all of the applicable airworthinessrequirements are met. For the purpose of thiscomparison, the flight path is considered to ex-tend to at least a height of 400 feet above thetakeoff surface.

(c) The takeoff path with all engines operating,but without the use of standby power, must reflecta conservatively greater overall level of perfor-mance than the one-engine-inoperative takeoffpath established in accordance with paragraph(a) of this section. The margin must be estab-lished by the Administrator to insure safe day-to-day operations, but in no case may it be less than15 percent. The all-engines-operating takeoff pathmust be determined by a procedure consistentwith that established in complying with paragraph(a) of this section.

(d) For reciprocating-engine-powered air-planes, the takeoff path to be scheduled in theAirplane Flight Manual must represent the one-engine-operative takeoff path determined in ac-cordance with paragraph (a) of this section andmodified to reflect the procedure (see paragraph(6)) established by the applicant for flap retractionand attainment of the en route speed. The sched-uled takeoff path must have a positive slope at allpoints of the airborne portion and at no point mustit lie above the takeoff path specified in paragraph(a) of this section.

(3) Takeoff distance. The takeoff distance mustbe the horizontal distance along the one-engine-inoperative take off path determined in accor-dance with paragraph (2)(a) from the start of thetakeoff to the point where the airplane attains aheight of 50 feet above the takeoff surface for re-

Part 25: Airworthiness Standards: Transport Category Appendix E to Part 25

ASA 165

25

ciprocating-engine-powered airplanes and aheight of 35 feet above the takeoff surface for tur-bine-powered airplanes.

(4) Maximum certificated takeoff weights. Themaximum certificated takeoff weights must be de-termined at all altitudes, and at ambient tempera-tures, if applicable, at which performance credit isto be applied and may not exceed the weights es-tablished in compliance with paragraphs (a) and(b) of this section.

(a) The conditions of paragraphs (2) (b)through (d) must be met at the maximum certifi-cated takeoff weight.

(b) Without the use of standby power, the air-plane must meet all of the en route requirementsof the applicable airworthiness regulations underwhich the airplane was originally certificated. Inaddition, turbine-powered airplanes without theuse of standby power must meet the final takeoffclimb requirements prescribed in the applicableairworthiness regulations.

(5) Maximum certificated landing weights.(a) The maximum certificated landing weights

(one-engine-inoperative approach and all-engine-operating landing climb) must be determined at allaltitudes, and at ambient temperatures if applica-ble, at which performance credit is to be appliedand must not exceed that established in compli-ance with paragraph (b) of this section.

(b) The flight path, with the engines operatingat the power or thrust, or both, appropriate to theairplane configuration and with standby power inuse, must lie above the flight path without standbypower in use at the maximum weight at which allof the applicable airworthiness requirements aremet. In addition, the flight paths must comply withsubparagraphs (i) and (ii) of this paragraph.

(i) The flight paths must be established withoutchanging the appropriate airplane configuration.

(ii) The flight paths must be carried out for aminimum height of 400 feet above the point wherestandby power is actuated.

(6) Airplane configuration, speed, and powerand thrust; general. Any change in the airplane’sconfiguration, speed, and power or thrust, or both,must be made in accordance with the proceduresestablished by the applicant for the operation ofthe airplane in service and must comply withparagraphs (a) through (c) of this section. In addi-tion, procedures must be established for the exe-cution of balked landings and missed ap-proaches.

(a) The Administrator must find that the proce-dure can be consistently executed in service bycrews of average skill.

(b) The procedure may not involve methods orthe use of devices which have not been proven tobe safe and reliable.

(c) Allowances must be made for such time de-lays in the execution of the procedures as may bereasonably expected to occur during service.

(7) Installation and operation; standby power.The standby power unit and its installation mustcomply with paragraphs (a) and (b) of this section.

(a) The standby power unit and its installationmust not adversely affect the safety of the airplane.

(b) The operation of the standby power unit andits control must have proven to be safe and reli-able.


Appendix F to Part 25 Federal Aviation Regulations

166 ASA

APPENDIX F TO PART 25

PART I — TEST CRITERIA AND PROCEDURES FOR SHOWING COMPLIANCE

WITH §25.853, OR §25.855.(a) Material test criteria—(1) Interior compartments occupied by crew or

passengers.(i) Interior ceiling panels, interior wall panels,

partitions, galley structure, large cabinet walls,structural flooring, and materials used in the con-struction of stowage compartments (other than un-derseat stowage compartments and compart-ments for stowing small items such as magazinesand maps) must be self-extinguishing when testedvertically in accordance with the applicable por-tions of part I of this appendix. The average burnlength may not exceed 6 inches and the averageflame time after removal of the flame source maynot exceed 15 seconds. Drippings from the testspecimen may not continue to flame for more thanan average of 3 seconds after falling.

(ii) Floor covering, textiles (including draperiesand upholstery), seat cushions, padding, decora-tive and nondecorative coated fabrics, leather,trays and galley furnishings, electrical conduit, airducting, joint and edge covering, liners of Class Band E cargo or baggage compartments, floorpanels of Class B, C, D, or E cargo or baggagecompartments, cargo covers and transparencies,molded and thermoformed parts, air ductingjoints, and trim strips (decorative and chafing),that are constructed of materials not covered insubparagraph (iv) below, must be self-extinguish-ing when tested vertically in accordance with theapplicable portions of part I of this appendix orother approved equivalent means. The averageburn length may not exceed 8 inches, and the av-erage flame time after removal of the flamesource may not exceed 15 seconds. Drippingsfrom the test specimen may not continue to flamefor more than an average of 5 seconds after fall-ing.

(iii) Motion picture film must be safety filmmeeting the Standard Specifications for SafetyPhotographic Film PHI.25 (available from theAmerican National Standards Institute, 1430Broadway, New York, NY 10018). If the film travelsthrough ducts, the ducts must meet the require-ments of subparagraph (ii) of this paragraph.

(iv) Clear plastic windows and signs, parts con-structed in whole or in part of elastomeric materi-als, edge lighted instrument assemblies consist-ing of two or more instruments in a commonhousing, seat belts, shoulder harnesses, andcargo and baggage tiedown equipment, includingcontainers, bins, pallets, etc., used in passengeror crew compartments, may not have an averageburn rate greater than 2.5 inches per minute when

tested horizontally in accordance with the applica-ble portions of this appendix.

(v) Except for small parts (such as knobs, han-dles, rollers, fasteners, clips, grommets, rubstrips, pulleys, and small electrical parts) thatwould not contribute significantly to the propaga-tion of a fire and for electrical wire and cable insu-lation, materials in items not specified in para-graphs (a)(1)(i), (ii), (iii), or (iv) of part I of this ap-pendix may not have a burn rate greater than 4.0inches per minute when tested horizontally in ac-cordance with the applicable portions of this ap-pendix.

(2) Cargo and baggage compartments not oc-cupied by crew or passengers.

(i) [Reserved](ii) A cargo or baggage compartment defined in

§25.857 as Class B or E must have a liner con-structed of materials that meet the requirementsof paragraph (a)(1)(ii) of part I of this appendixand separated from the airplane structure (exceptfor attachments). In addition, such liners must besubjected to the 45 degree angle test. The flamemay not penetrate (pass through) the materialduring application of the flame or subsequent toits removal. The average flame time after removalof the flame source may not exceed 15 seconds,and the average glow time may not exceed 10seconds.

(iii) A cargo or baggage compartment definedin §25.857 as Class B, C, D, or E must have floorpanels constructed of materials which meet therequirements of paragraph (a)(1)(ii) of part I ofthis appendix and which are separated from theairplane structure (except for attachments). Suchpanels must be subjected to the 45 degree angletest. The flame may not penetrate (pass through)the material during application of the flame orsubsequent to its removal. The average flametime after removal of the flame source may not ex-ceed 15 seconds, and the average glow time maynot exceed 10 seconds.

(iv) Insulation blankets and covers used to pro-tect cargo must be constructed of materials thatmeet the requirements of paragraph (a)(1)(ii) ofpart I of this appendix. Tiedown equipment (in-cluding containers, bins, and pallets) used in eachcargo and baggage compartment must be con-structed of materials that meet the requirementsof paragraph (a)(1)(v) of part I of this appendix.

(3) Electrical system components. Insulationon electrical wire or cable installed in any area ofthe fuselage must be self-extinguishing whensubjected to the 60 degree test specified in part Iof this appendix. The average burn length may notexceed 3 inches, and the average flame time afterremoval of the flame source may not exceed 30seconds. Drippings from the test specimen maynot continue to flame for more than an average of3 seconds after falling.

Part 25: Airworthiness Standards: Transport Category Appendix F to Part 25

ASA 167

25

(b) Test Procedures—(1) Conditioning. Specimens must be condi-

tioned to 70 ± 5 F, and at 50 percent ±5 percentrelative humidity until moisture equilibrium isreached or for 24 hours. Each specimen must re-main in the conditioning environment until it issubjected to the flame.

(2) Specimen configuration. Except for smallparts and electrical wire and cable insulation, ma-terials must be tested either as section cut from afabricated part as installed in the airplane or as aspecimen simulating a cut section, such as aspecimen cut from a flat sheet of the material or amodel of the fabricated part. The specimen maybe cut from any location in a fabricated part; how-ever, fabricated units, such as sandwich panels,may not be separated for test. Except as notedbelow, the specimen thickness must be no thickerthan the minimum thickness to be qualified for usein the airplane. Test specimens of thick foamparts, such as seat cushions, must be 1⁄2-inch inthickness. Test specimens of materials that mustmeet the requirements of paragraph (a)(1)(v) ofpart I of this appendix must be no more than 1⁄8-inch in thickness. Electrical wire and cable speci-mens must be the same size as used in the air-plane. In the case of fabrics, both the warp and filldirection of the weave must be tested to deter-mine the most critical flammability condition.Specimens must be mounted in a metal frame sothat the two long edges and the upper edge areheld securely during the vertical test prescribed insubparagraph (4) of this paragraph and the twolong edges and the edge away from the flame areheld securely during the horizontal test prescribedin subparagraph (5) of this paragraph. The ex-posed area of the specimen must be at least 2inches wide and 12 inches long, unless the actualsize used in the airplane is smaller. The edge towhich the burner flame is applied must not consistof the finished or protected edge of the specimenbut must be representative of the actual cross-section of the material or part as installed in theairplane. The specimen must be mounted in ametal frame so that all four edges are held se-curely and the exposed area of the specimen is atleast 8 inches by 8 inches during the 45° test pre-scribed in subparagraph (6) of this paragraph.

(3) Apparatus. Except as provided in subpara-graph (7) of this paragraph, tests must be con-ducted in a draft-free cabinet in accordance withFederal Test Method Standard 191 Model 5903(revised Method 5902) for the vertical test, orMethod 5906 for horizontal test (available fromthe General Services Administration, BusinessService Center, Region 3, Seventh & D StreetsSW., Washington, DC 20407). Specimens whichare too large for the cabinet must be tested in sim-ilar draft-free conditions.

(4) Vertical test. A minimum of three specimensmust be tested and results averaged. For fabrics,the direction of weave corresponding to the mostcritical flammability conditions must be parallel tothe longest dimension. Each specimen must besupported vertically. The specimen must be ex-posed to a Bunsen or Tirrill burner with a nominal3⁄8-inch I.D. tube adjusted to give a flame of 11⁄2inches in height. The minimum flame temperaturemeasured by a calibrated thermocouple pyrome-ter in the center of the flame must be 1550°F Thelower edge of the specimen must be 3⁄4-inchabove the top edge of the burner. The flame mustbe applied to the center line of the lower edge ofthe specimen. For materials covered by para-graph (a)(1)(i) of part I of this appendix, the flamemust be applied for 60 seconds and then re-moved. For materials covered by paragraph(a)(1)(ii) of part I of this appendix, the flame mustbe applied for 12 seconds and then removed.Flame time, burn length, and flaming time of drip-pings, if any, may be recorded. The burn lengthdetermined in accordance with subparagraph (7)of this paragraph must be measured to the near-est tenth of an inch.

(5) Horizontal test. A minimum of three speci-mens must be tested and the results averaged.Each specimen must be supported horizontally.The exposed surface, when installed in the air-craft, must be face down for the test. The speci-men must be exposed to a Bunsen or Tirrill burnerwith a nominal 3⁄8-inch I.D. tube adjusted to give aflame of 11⁄2 inches in height. The minimum flametemperature measured by a calibrated thermo-couple pyrometer in the center of the flame mustbe 1550°F. The specimen must be positioned sothat the edge being tested is centered 3⁄4-inchabove the top of the burner. The flame must beapplied for 15 seconds and then removed. A min-imum of 10 inches of specimen must be used fortiming purposes, approximately 11⁄2 inches mustburn before the burning front reaches the timingzone, and the average burn rate must be re-corded.

(6) Forty-five degree test. A minimum of threespecimens must be tested and the results aver-aged. The specimens must be supported at anangle of 45° to a horizontal surface. The exposedsurface when installed in the aircraft must be facedown for the test. The specimens must be ex-posed to a Bunsen or Tirrill burner with a nominal3⁄8-inch I.D. tube adjusted to give a flame of 11⁄2inches in height. The minimum flame temperaturemeasured by a calibrated thermocouple pyrome-ter in the center of the flame must be 1550°F. Suit-able precautions must be taken to avoid drafts.The flame must be applied for 30 seconds withone-third contacting the material at the center ofthe specimen and then removed. Flame time,


168 ASA

glow time, and whether the flame penetrates(passes through) the specimen must be recorded.

(7) Sixty degree test. A minimum of three spec-imens of each wire specification (make and size)must be tested. The specimen of wire or cable (in-cluding insulation) must be placed at an angle of60° with the horizontal in the cabinet specified insubparagraph (3) of this paragraph with the cabi-net door open during the test, or must be placedwithin a chamber approximately 2 feet high by 1foot by 1 foot, open at the top and at one verticalside (front), and which allows sufficient flow of airfor complete combustion, but which is free fromdrafts. The specimen must be parallel to and ap-proximately 6 inches from the front of the cham-ber. The lower end of the specimen must be heldrigidly clamped. The upper end of the specimenmust pass over a pulley or rod and must have anappropriate weight attached to it so that the spec-imen is held tautly throughout the flammabilitytest. The test specimen span between lowerclamp and upper pulley or rod must be 24 inchesand must be marked 8 inches from the lower endto indicate the central point for flame application.A flame from a Bunsen or Tirrill burner must beapplied for 30 seconds at the test mark. Theburner must be mounted underneath the testmark on the specimen, perpendicular to the spec-imen and at an angle of 30° to the vertical planeof the specimen. The burner must have a nominalbore of 3⁄8-inch and be adjusted to provide a 3-inch high flame with an inner cone approximatelyone-third of the flame height. The minimum tem-perature of the hottest portion of the flame, asmeasured with a calibrated thermocouple pyrom-eter, may not be less than 1750°F. The burnermust be positioned so that the hottest portion ofthe flame is applied to the test mark on the wire.Flame time, burn length, and flaming time of drip-pings, if any, must be recorded. The burn lengthdetermined in accordance with paragraph (8) ofthis paragraph must be measured to the nearesttenth of an inch. Breaking of the wire specimensis not considered a failure.

(8) Burn length. Burn length is the distancefrom the original edge to the farthest evidence ofdamage to the test specimen due to flame im-pingement, including areas of partial or completeconsumption, charring, or embrittlement, but notincluding areas sooted, stained, warped, or dis-colored, nor areas where material has shrunk ormelted away from the heat source.

PART II — FLAMMABILITY OF SEAT CUSHIONS

(a) Criteria for Acceptance. Each seat cushionmust meet the following criteria:

(1) At least three sets of seat bottom and seatback cushion specimens must be tested.

(2) If the cushion is constructed with a fireblocking material, the fire blocking material mustcompletely enclose the cushion foam core mate-rial.

(3) Each specimen tested must be fabricatedusing the principal components (i.e., foam core,flotation material, fire blocking material, if used,and dress covering) and assembly processes(representative seams and closures) intended foruse in the production articles. If a different mate-rial combination is used for the back cushion thanfor the bottom cushion, both material combina-tions must be tested as complete specimen sets,each set consisting of a back cushion specimenand a bottom cushion specimen. If a cushion, in-cluding outer dress covering, is demonstrated tomeet the requirements of this appendix using theoil burner test, the dress covering of that cushionmay be replaced with a similar dress coveringprovided the burn length of the replacement cov-ering, as determined by the test specified in§25.853(c), does not exceed the correspondingburn length of the dress covering used on thecushion subjected to the oil burner test.

(4) For at least two-thirds of the total number ofspecimen sets tested, the burn length from theburner must not reach the side of the cushion op-posite the burner. The burn length must not ex-ceed 17 inches. Burn length is the perpendiculardistance from the inside edge of the seat frameclosest to the burner to the farthest evidence ofdamage to the test specimen due to flame im-pingement, including areas of partial or completeconsumption, charring, or embrittlement, but notincluding areas sooted, stained, warped, or dis-colored, or areas where material has shrunk ormelted away from the heat source.

(5) The average percentage weight loss mustnot exceed 10 percent. Also, at least two-thirds ofthe total number of specimen sets tested must notexceed 10 percent weight loss. All droppings fall-ing from the cushions and mounting stand are tobe discarded before the after-test weight is deter-mined. The percentage weight loss for a speci-men set is the weight of the specimen set beforetesting less the weight of the specimen set aftertesting expressed as the percentage of the weightbefore testing.

(b) Test Conditions. Vertical air velocity shouldaverage 25 fpm ±10 fpm at the top of the backseat cushion. Horizontal air velocity should be be-low 10 fpm just above the bottom seat cushion.Air velocities should be measured with the ventila-tion hood operating and the burner motor off.

(c) Test Specimens.(1) For each test, one set of cushion specimens

representing a seat bottom and seat back cushionmust be used.

(2) The seat bottom cushion specimen must be18 ± 1⁄8 inches (457 ± 3 mm) wide by 20 ± 1⁄8


ASA 169

25

inches (508 ± 3 mm) deep by 4 ± 1⁄8 inches(102 ± 3 mm) thick, exclusive of fabric closuresand seam overlap.

(3) The seat back cushion specimen must be18 ± 1⁄8 inches (432 ± 3 mm) wide by 25 ± 1⁄8inches (635 ± 3 mm) high by 2 ± 1⁄8 inches (51 ± 3mm) thick, exclusive of fabric closures and seamoverlap.

(4) The specimens must be conditioned at70 ±5°F (21 ± 2°C) 55% ± 10% relative humidityfor at least 24 hours before testing.

(d) Test Apparatus. The arrangement of the testapparatus is shown in Figures 1 through 5 andmust include the components described in thissection. Minor details of the apparatus may vary,depending on the model burner used.

(1) Specimen Mounting Stand. The mountingstand for the test specimens consists of steel an-gles, as shown in Figure 1. The length of themounting stand legs is 12 ± 1⁄8 inches (305 ± 3mm). The mounting stand must be used formounting the test specimen seat bottom and seatback, as shown in Figure 2. The mounting standshould also include a suitable drip pan lined withaluminum foil, dull side up.

(2) Test Burner. The burner to be used in test-ing must—

(i) Be a modified gun type;(ii) Have an 80-degree spray angle nozzle

nominally rated for 2.25 gallons/hour at 100 psi;(iii) Have a 12-inch (305 mm) burner cone in-

stalled at the end of the draft tube, with an open-ing 6 inches (152 mm) high and 11 inches (280mm) wide, as shown in Figure 3; and

(iv) Have a burner fuel pressure regulator thatis adjusted to deliver a nominal 2.0 gallon/hour of# 2 Grade kerosene or equivalent required for thetest.

Burner models which have been used success-fully in testing are the Lennox Model OB-32, Car-lin Model 200 CRD, and Park Model DPL 3400.FAA published reports pertinent to this type ofburner are: (1) Powerplant Engineering ReportNo. 3A, Standard Fire Test Apparatus and Proce-dure for Flexible Hose Assemblies, dated March1978; and (2) Report No. DOT/FAA/RD/76/213,Reevaluation of Burner Characteristics for FireResistance Tests, dated January 1977.

(3) Calorimeter.(i) The calorimeter to be used in testing must

be a (0-15.0 BTU/ft2 -sec. 0-17.0 W/cm2) calorim-eter, accurate ±3%, mounted in a 6-inch by 12-inch (152 by 305 mm) by 3⁄4-inch (19 mm) thickcalcium silicate insulating board which is attachedto a steel angle bracket for placement in the teststand during burner calibration, as shown in Fig-ure 4.

(ii) Because crumbling of the insulating boardwith service can result in misalignment of the cal-

orimeter, the calorimeter must be monitored andthe mounting shimmed, as necessary, to ensurethat the calorimeter face is flush with the exposedplane of the insulating board in a plane parallel tothe exit of the test burner cone.

(4) Thermocouples. The seven thermocouplesto be used for testing must be 1⁄16- to 1⁄8-inchmetal sheathed, ceramic packed, type K,grounded thermocouples with a nominal 22 to 30American wire gage (AWG)-size conductor. Theseven thermocouples must be attached to a steelangle bracket to form a thermocouple rake forplacement in the test stand during burner calibra-tion, as shown in Figure 5.

(5) Apparatus Arrangement. The test burnermust be mounted on a suitable stand to positionthe exit of the burner cone a distance of 4 ± 1⁄8inches (102 ± 3 mm) from one side of the speci-men mounting stand. The burner stand shouldhave the capability of allowing the burner to beswung away from the specimen mounting standduring warmup periods.

(6) Data Recording. A recording potentiometeror other suitable calibrated instrument with an ap-propriate range must be used to measure andrecord the outputs of the calorimeter and the ther-mocouples.

(7) Weight Scale. Weighing Device — A devicemust be used that with proper procedures maydetermine the before and after test weights ofeach set of seat cushion specimens within 0.02pound (9 grams). A continuous weighing systemis preferred.

(8) Timing Device. A stopwatch or other device(calibrated to ±1 second) must be used to mea-sure the time of application of the burner flameand self-extinguishing time or test duration.

(e) Preparation of Apparatus. Before calibra-tion, all equipment must be turned on and theburner fuel must be adjusted as specified in para-graph (d)(2).

(f) Calibration. To ensure the proper thermaloutput of the burner, the following test must bemade:

(1) Place the calorimeter on the test stand asshown in Figure 4 at a distance of 4 ± 1⁄8 inches(102 ±3 mm) from the exit of the burner cone.

(2) Turn on the burner, allow it to run for 2 min-utes for warmup, and adjust the burner air intakedamper to produce a reading of 10.5 ± 0.5 BTU/ft2-sec. (11.9 ± 0.6 w/cm2) on the calorimeter to en-sure steady state conditions have been achieved.Turn off the burner.

(3) Replace the calorimeter with the thermo-couple rake (Figure 5).

(4) Turn on the burner and ensure that the ther-mocouples are reading 1900 ±100°F (1038±38°C) to ensure steady state conditions havebeen achieved.


170 ASA

(5) If the calorimeter and thermocouples do notread within range, repeat steps in paragraphs 1through 4 and adjust the burner air intake damperuntil the proper readings are obtained. The ther-mocouple rake and the calorimeter should beused frequently to maintain and record calibratedtest parameters. Until the specific apparatus hasdemonstrated consistency, each test should becalibrated. After consistency has been confirmed,several tests may be conducted with the pre-testcalibration before and a calibration check after theseries.

(g) Test Procedure. The flammability of eachset of specimens must be tested as follows:

(1) Record the weight of each set of seat bot-tom and seat back cushion specimens to betested to the nearest 0.02 pound (9 grams).

(2) Mount the seat bottom and seat back cush-ion test specimens on the test stand as shown inFigure 2, securing the seat back cushion speci-men to the test stand at the top.

(3) Swing the burner into position and ensurethat the distance from the exit of the burner coneto the side of the seat bottom cushion specimen is4 ± 1⁄8 inches (102 ± 3 mm).

(4) Swing the burner away from the test posi-tion. Turn on the burner and allow it to run for 2minutes to provide adequate warmup of theburner cone and flame stabilization.

(5) To begin the test, swing the burner into thetest position and simultaneously start the timingdevice.

(6) Expose the seat bottom cushion specimento the burner flame for 2 minutes and then turn offthe burner. Immediately swing the burner awayfrom the test position. Terminate test 7 minutes af-ter initiating cushion exposure to the flame by useof a gaseous extinguishing agent (i.e., Halon orCO2).

(7) Determine the weight of the remains of theseat cushion specimen set left on the mountingstand to the nearest 0.02 pound (9 grams) exclud-ing all droppings.

(h) Test Report. With respect to all specimensets tested for a particular seat cushion for whichtesting of compliance is performed, the followinginformation must be recorded:

(1) An identification and description of thespecimens being tested.

(2) The number of specimen sets tested.(3) The initial weight and residual weight of

each set, the calculated percentage weight loss ofeach set, and the calculated average percentageweight loss for the total number of sets tested.

(4) The burn length for each set tested.


ASA 171

25

SIDE VIEW

STEEL FLAT STOCK11⁄2" x 1⁄8"

(38 x 3 mm)

FRONT VIEW

STEEL ANGLE1" x 1" x 1⁄8"

(25 x 25 x 3 mm)

33 ±

1⁄ 8"

(838

±3 m

m)

12 ±

1⁄ 8"

(304

±3 m

m)

TOP VIEW

221 ⁄ 8

± 1

⁄ 8"(5

61±3

mm

)

181⁄8 ± 1⁄8"(450±3 mm)

NOTE:

ALL JOINTS WELDEDFLAT STOCK BUTT WELDEDALL MEASUREMENTS INSIDE

Figure 1


172 ASA

Figure 2

SEATBACK

CUSHION

22 ± 1⁄8"(559 ± 3 mm)

20 ± 1⁄8"(509 ± 3 mm)

SIDE VIEW

10"(254 mm)

25 ± 1⁄8"(635 ± 3 mm)

2"(51mm)

2"(51 mm)

4 ± 1⁄8"(102 ± 3 mm)

SEAT BOTTOMCUSHION

TOP VIEW

SEAT CUSHIONMOUNTING

CL

18 ± 1⁄8"(457 ± 3 mm)

BURNER CONE

10"(254 mm)


ASA 173

25

Figure 3

NOTE:ONE HALF (1⁄2) OF TUBEEXTENSION SHOWN. SECONDHALF MATES AT SPOTWELDOVERLAPS.

MATERIAL: 0.050 STAINLESS STEEL

BOLT HOLES

121⁄4"

TO DRAFT TUBE.1⁄2 SECTION OFCONNECTING

FLANGE

DRAFT TUBE EXTENSIONFOR FAA HOSE TESTBURNER

1⁄2"

27⁄64" SPACES

33⁄4"

71⁄2"

41⁄4"

11"

6"

CONNECTING FLANGE

BOLTS

1⁄2"

1"

27⁄64" SPACES

131⁄2"

19⁄32" SPACES4"5"

151⁄2"

10° BENDSON BROKEN LINES

OV

ER

LAP

FO

R S

PO

TW

ELD

OV

ER

LAP

F

OR

SP

OT

WE

LD

9 8 7 6 5 4 3 2 9 8 7 6 5 4 3 2

9 8 7 6 5 4 3 2 1 2 3 4 5 6 7 8 9

13"


174 ASA

Figure 4

(152 x 305 x 19 mm)6" x 12" x 3⁄4"

MARINITE BLOCK

(254 ± 3 mm)10 ± 1⁄8"

SIDE VIEW

(25 mm)1" DIAMETER HOLE FORCALORIMETER MOUNTING

6 ± 1⁄8"(152 ± 3 mm)

12 ± 1⁄8"(305 ± 3 mm)

21.5 ± 1⁄8"(546 ± 3 mm)

3 ± 1⁄8"(76 ± 3 mm)

BURNER CONE4 ± 1⁄8"(102 ± 3 mm)

(25 mm) 1" DIAMETER

(19 mm)3⁄4"

12 ± 1⁄8"(305 ± 3 mm)

WATER-COOLEDCALORIMETER

RACK FITS INSIDESEAT FRAME

STEEL ANGLE1" x 1" x 1⁄8"(25 x 25 x 3 mm)

TOP VIEWCALORIMETER BRACKET


ASA 175

25

Figure 5

21.5 ± 1⁄8"(305 ± 3 mm)

SIDE VIEW7 THERMOCOUPLES

10 ± 1⁄8"(254 ± 3 mm)

3 ± 1⁄8" (76 ± 3 mm)

1"(25 mm)

1" (25 mm)

BURNER CONE

RACKT FITS INSIDE SEAT FRAME

TOP VIEWTHERMOCOUPLE RAKE BRACKET

1"(25 mm)

12 ± 1⁄8"(305 ± 3 mm)

4 ± 1⁄8" (102 ± 3 mm)


(25 x 25 x 3mm)


176 ASA

PART III — TEST METHOD TO DETERMINE FLAME PENETRATION RESISTANCE OF CARGO

COMPARTMENT LINERS.(a) Criteria for Acceptance.(1) At least three specimens of cargo compart-

ment sidewall or ceiling liner panels must betested.

(2) Each specimen tested must simulate thecargo compartment sidewall or ceiling liner panel,including any design features, such as joints,lamp assemblies, etc., the failure of which wouldaffect the capability of the liner to safely contain afire.

(3) There must be no flame penetration of anyspecimen within 5 minutes after application of theflame source, and the peak temperature mea-sured at 4 inches above the upper surface of thehorizontal test sample must not exceed 400°F.

(b) Summary of Method. This method providesa laboratory test procedure for measuring the ca-pability of cargo compartment lining materials toresist flame penetration with a 2 gallon per hour(GPH) #2 Grade kerosene or equivalent burnerfire source. Ceiling and sidewall liner panels maybe tested individually provided a baffle is used tosimulate the missing panel. Any specimen thatpasses the test as a ceiling liner panel may beused as a sidewall liner panel.

(c) Test Specimens.(1) The specimen to be tested must measure

16 ± 1⁄8 inches (406 ± 3 mm) by 24+1⁄8 inches(610 ± 3 mm).

(2) The specimens must be conditioned at 70°F± 5°F (21°C ± 2°C) and 55% ± 5% humidity for atleast 24 hours before testing.

(d) Test Apparatus. The arrangement of the testapparatus, which is shown in Figure 3 of Part IIand Figures 1 through 3 of this part of Appendix F,must include the components described in thissection. Minor details of the apparatus may vary,depending on the model of the burner used.

(1) Specimen Mounting Stand. The mountingstand for the test specimens consists of steel an-gles as shown in Figure 1.

(2) Test Burner. The burner to be used in test-ing must—

(i) Be a modified gun type.(ii) Use a suitable nozzle and maintain fuel

pressure to yield a 2 GPH fuel flow. For example:an 80 degree nozzle nominally rated at 2.25 GPHand operated at 85 pounds per square inch (PSI)gage to deliver 2.03 GPH.

(iii) Have a 12 inch (305 mm) burner extensioninstalled at the end of the draft tube with an open-ing 6 inches (152 mm) high and 11 inches (280mm) wide as shown in Figure 3 of Part II of thisappendix.

(iv) Have a burner fuel pressure regulator thatis adjusted to deliver a nominal 2.0 GPH of #2Grade kerosene or equivalent.

Burner models which have been used success-fully in testing are the Lenox Model OB-32, CarlinModel 200 CRD and Park Model DPL. The basicburner is described in FAA Powerplant Engineer-ing Report No. 3A, Standard Fire Test Apparatusand Procedure for Flexible Hose Assemblies,dated March 1978; however, the test settingsspecified in this appendix differ in some instancesfrom those specified in the report.

(3) Calorimeter.(i) The calorimeter to be used in testing must be

a total heat flux Foil Type Gardon Gage of an ap-propriate range (approximately 0 to 15.0 Britishthermal unit (BTU) per ft.2 sec., 0-17.0 watts/cm2).The calorimeter must be mounted in a 6 inch by 12inch (152 by 305 mm) by 3⁄4 inch (19 mm) thick in-sulating block which is attached to a steel anglebracket for placement in the test stand duringburner calibration as shown in Figure 2 of this partof this appendix.

(ii) The insulating block must be monitored fordeterioration and the mounting shimmed as nec-essary to ensure that the calorimeter face is par-allel to the exit plane of the test burner cone.

(4) Thermocouples. The seven thermocouplesto be used for testing must be 1⁄16 inch ceramicsheathed, type K, grounded thermocouples with anominal 30 American wire gage (AWG) size con-ductor. The seven thermocouples must be at-tached to a steel angle bracket to form a thermo-couple rake for placement in the test stand duringburner calibration as shown in Figure 3 of this partof this appendix.

(5) Apparatus Arrangement. The test burnermust be mounted on a suitable stand to positionthe exit of the burner cone a distance of 8 inchesfrom the ceiling liner panel and 2 inches from thesidewall liner panel. The burner stand shouldhave the capability of allowing the burner to beswung away from the test specimen during warm-up periods.

(6) Instrumentation. A recording potentiometer orother suitable instrument with an appropriate rangemust be used to measure and record the outputs ofthe calorimeter and the thermocouples.

(7) Timing Device. A stopwatch or other devicemust be used to measure the time of flame applica-tion and the time of flame penetration, if it occurs.

(e) Preparation of Apparatus. Before calibra-tion, all equipment must be turned on and allowedto stabilize, and the burner fuel flow must be ad-justed as specified in paragraph (d)(2).

(f) Calibration. To ensure the proper thermaloutput of the burner the following test must bemade:

(1) Remove the burner extension from the endof the draft tube. Turn on the blower portion of the


ASA 177

25

burner without turning the fuel or igniters on. Mea-sure the air velocity using a hot wire anemometerin the center of the draft tube across the face ofthe opening. Adjust the damper such that the airvelocity is in the range of 1550 to 1800 ft./min. Iftabs are being used at the exit of the draft tube,they must be removed prior to this measurement.Reinstall the draft tube extension cone.

(2) Place the calorimeter on the test stand asshown in Figure 2 at a distance of 8 inches (203mm) from the exit of the burner cone to simulatethe position of the horizontal test specimen.

(3) Turn on the burner, allow it to run for 2 min-utes for warm-up, and adjust the damper to pro-duce a calorimeter reading of 8.0 ±0.5 BTU perft.2 sec. (9.1 ± 0.6 Watts/cm2).

(4) Replace the calorimeter with the thermo-couple rake (see Figure 3).

(5) Turn on the burner and ensure that each ofthe seven thermocouples reads 1700°F ±100°F(927°C ±38°C.) to ensure steady state conditionshave been achieved. If the temperature is out ofthis range, repeat steps 2 through 5 until properreadings are obtained.

(6) Turn off the burner and remove the thermo-couple rake.

(7) Repeat (1) to ensure that the burner is inthe correct range.

(g) Test Procedure.(1) Mount a thermocouple of the same type as

that used for calibration at a distance of 4 inches(102 mm) above the horizontal (ceiling) test spec-imen. The thermocouple should be centered overthe burner cone.

(2) Mount the test specimen on the test standshown in Figure 1 in either the horizontal or verti-cal position. Mount the insulating material in theother position.

(3) Position the burner so that flames will notimpinge on the specimen, turn the burner on, andallow it to run for 2 minutes. Rotate the burner toapply the flame to the specimen and simulta-neously start the timing device.

(4) Expose the test specimen to the flame for 5minutes and then turn off the burner. The test maybe terminated earlier if flame penetration is ob-served.

(5) When testing ceiling liner panels, record thepeak temperature measured 4 inches above thesample.

(6) Record the time at which flame penetrationoccurs if applicable.

(h) Test Report. The test report must includethe following:

(1) A complete description of the materialstested including type, manufacturer, thickness,and other appropriate data.

(2) Observations of the behavior of the testspecimens during flame exposure such as delam-ination, resin ignition, smoke, etc., including thetime of such occurrence.

(3) The time at which flame penetration occurs,if applicable, for each of the three specimenstested.

(4) Panel orientation (ceiling or sidewall).[SEE FIGURES BEGINNING ON THE NEXT PAGE]


178 ASA

HORIZONTAL SPEC. POSITION

TOP VIEW

VERTICAL SPEC. POSITION

HORIZONTAL AND VERTICAL SPECIMENS ARECLAMPED IN PLACE ON ALL EDGES BETWEENANGLES AS SHOWN IN VIEW A - A

TEST STAND FRAME

VERTICAL SPEC.

HORIZONTAL SPEC.

SUPPORT ANGLE

VIEW A - A (Typical)

24"

A

A

4"

48"

16"

BURNER CONE

BURNER ASSEMBLY

BURNER SHIELD

2"

8"

FRONT VIEW

CL

1" x 3" x 1⁄8"

STEEL “U” CHANNEL

TEST STAND IS CONSTRUCTED WITH 1" x 1" x 1⁄8" STEEL ANGLES, ALL JOINTS WELDED

SUPPORT ANGLES ARE 1" x 1" x 1⁄8" CUT TO FIT

FIGURE 1. Test apparatus for horizontal and vertical mounting.

SUPPORT BRACE

SIDE VIEW

CL

16"


ASA 179

25


180 ASA

BURNER CONE

CL

SIDE VIEW

FIGURE 3. Thermocouple rake bracket

8"


1"

TOP VIEW

NOTE: BRACKET IS CLAMPED TO TESTSTAND WITH THERMOCOUPLESOFF CENTER OF BURNER CONEBY ONE INCH.

CL

CL

1"1"

24"

SEVEN THERMOCOUPLES

12 1⁄2"


ASA 181

25

PART IV — TEST METHOD TO DETERMINE THE HEAT RELEASE RATE FROM CABIN MATERIALS

EXPOSED TO RADIANT HEAT.(a) Summary of Method. Three or more speci-

mens representing the completed aircraft compo-nent are tested. Each test specimen is injectedinto an environmental chamber through which aconstant flow of air passes. The specimen’s expo-sure is determined by a radiant heat source ad-justed to produce, on the specimen, the desiredtotal heat flux of 3.5 W/cm2. The specimen istested with the exposed surface vertical. Combus-tion is initiated by piloted ignition. The combustionproducts leaving the chamber are monitored in or-der to calculate the release rate of heat.

(b) Apparatus. The Ohio State University(OSU) rate of heat release apparatus, as de-scribed below, is used. This is a modified versionof the rate of heat release apparatus standardizedby the American Society of Testing and Materials(ASTM), ASTM E-906.

(1) This apparatus is shown in Figures 1A and1B of this part IV. All exterior surfaces of the appa-ratus, except the holding chamber, must be insu-lated with 1 inch (25 mm) thick, low density, hightemperature, fiberglass board insulation. A gas-keted door, through which the sample injectionrod slides, must be used to form an airtight clo-sure on the specimen hold chamber.

(2) Thermopile. The temperature difference be-tween the air entering the environmental chamberand that leaving must be monitored by a thermo-pile having five hot, and five cold, 24-gaugeChromel-Alumel junctions. The hot junctions mustbe spaced across the top of the exhaust stack,.38 inches (10 mm) below the top of the chimney.The thermocouples must have a .050 ±.010 inch(1.3 ±.3 mm) diameter, ball-type, welded tip. Onethermocouple must be located in the geometriccenter, with the other four located 1.18 inch (30mm) from the center along the diagonal towardeach of the corners (Figure 5 of this part IV). Thecold junctions must be located in the pan belowthe lower air distribution plate (see paragraph(b)(4) of this part IV). Thermopile hot junctionsmust be cleared of soot deposits as needed tomaintain the calibrated sensitivity.

(3) Radiation Source. A radiant heat source in-corporating four Type LL silicon carbide elements,20 inches (508 mm) long by .63 inch (16 mm)O.D., must be used, as shown in Figures 2A and2B of this part IV. The heat source must have anominal resistance of 1.4 ohms and be capable ofgenerating a flux up to 100 kW/m2. The siliconecarbide elements must be mounted in the stain-less steel panel box by inserting them through .63inch (16 mm) holes in .03 inch (1 mm) thick ce-ramic fiber or calcium-silicate millboard. Locationsof the holes in the pads and stainless steel cover

plates are shown in Figure 2B of this part IV. Thetruncated diamond-shaped mask of .042 ±.002inch (1.07 ±.05 mm) stainless steel must beadded to provide uniform heat flux density overthe area occupied by the vertical sample.

(4) Air Distribution System. The air entering theenvironmental chamber must be distributed by a.25 inch (6.3 mm) thick aluminum plate havingeight No. 4 drill-holes, located 2 inches (51 mm)from sides on 4 inch (102 mm) centers, mountedat the base of the environmental chamber. A sec-ond plate of 18 gauge stainless steel having 120,evenly spaced, No. 28 drill holes must bemounted 6 inches (152 mm) above the aluminumplate. A well-regulated air supply is required. Theair-supply manifold at the base of the pyramidalsection must have 48, evenly spaced, No. 26 drillholes located .38 inch (10 mm) from the inneredge of the manifold, resulting in an airflow split ofapproximately three to one within the apparatus.

(5) Exhaust Stack. An exhaust stack, 5.25 x2.75 inches (133 x 70 mm) in cross section, and10 inches (254 mm) long, fabricated from 28gauge stainless steel must be mounted on theoutlet of the pyramidal section. A 1.0 x 3.0 inch(25 x 76 mm) baffle plate of 0.18 ±.002 inch (.50±.05 mm) stainless steel must be centered insidethe stack, perpendicular to the air flow, 3 inches(76 mm) above the base of the stack.

(6) Specimen Holders. (i) The specimen mustbe tested in a vertical orientation. The specimenholder (Figure 3 of this part IV) must incorporate aframe that touches the specimen (which iswrapped with aluminum foil as required by para-graph (d)(3) of this Part) along only the .25 inch (6mm) perimeter. A “V” shaped spring is used tohold the assembly together. A detachable .50 x.50 x 5.91 inch (12 x 12 x 150 mm) drip pan andtwo .020 inch (.5 mm) stainless steel wires (asshown in Figure 3 of this part IV) must be used fortesting materials prone to melting and dripping.The positioning of the spring and frame may bechanged to accommodate different specimenthicknesses by inserting the retaining rod in differ-ent holes on the specimen holder.

(ii) Since the radiation shield described inASTM E-906 is not used, a guide pin must beadded to the injection mechanism. This fits into aslotted metal plate on the injection mechanismoutside of the holding chamber. It can be used toprovide accurate positioning of the specimen faceafter injection. The front surface of the specimenmust be 3.9 inches (100 mm) from the closed ra-diation doors after injection.

(iii) The specimen holder clips onto themounted bracket (Figure 3 of this part IV). Themounting bracket must be attached to the injec-tion rod by three screws that pass through a wide-area washer welded onto a 1⁄2-inch (13 mm) nut.The end of the injection rod must be threaded to


182 ASA

screw into the nut, and a .020 inch (5.1 mm) thickwide area washer must be held between two 1⁄2-inch (13 mm) nuts that are adjusted to tightlycover the hole in the radiation doors throughwhich the injection rod or calibration calorimeterpass.

(7) Calorimeter. A total-flux type calorimetermust be mounted in the center of a 1⁄2-inch Kao-wool “M” board inserted in the sample holder tomeasure the total heat flux. The calorimeter musthave a view angle of 180 degrees and be cali-brated for incident flux. The calorimeter calibrationmust be acceptable to the Administrator.

(8) Pilot-Flame Positions. Pilot ignition of thespecimen must be accomplished by simulta-neously exposing the specimen to a lower pilotburner and an upper pilot burner, as described inparagraph (b)(8)(i) and (b)(8)(ii) or (b)(8)(iii) of thispart IV, respectively. Since intermittent pilot flameextinguishment for more than 3 seconds would in-validate the test results, a spark ignitor may be in-stalled to ensure that the lower pilot burner re-mains lighted.

(i) Lower Pilot Burner. The pilot-flame tubingmust be .25 inch (6.3 mm) O.D., .03 inch (0.8 mm)wall, stainless steel tubing. A mixture of 120cm3/min. of methane and 850 cm3/min. of airmust be fed to the lower pilot flame burner. Thenormal position of the end of the pilot burner tub-ing is .40 inch (10 mm) from and perpendicular tothe exposed vertical surface of the specimen. Thecenterline at the outlet of the burner tubing mustintersect the vertical centerline of the sample at apoint .20 inch (5 mm) above the lower exposededge of the specimen.

(ii) Standard Three-Hole Upper Burner. Thepilot burner must be a straight length of .25 inch(6.3 mm) O.D., .03 inch (0.8 mm) wall, stainlesssteel tubing that is 14 inches (360 mm) long. Oneend of the tubing must be closed, and three No.40 drill holes must be drilled into the tubing, 2.38inch (60 mm) apart, for gas ports, all radiating inthe same direction. The first hole must be .19 inch(5 mm) from the closed end of the tubing. Thetube must be positioned .75 inch (19 mm) aboveand .75 inch (19 mm) behind the exposed upperedge of the specimen. The middle hole must be inthe vertical plane perpendicular to the exposedsurface of the specimen which passes through itsvertical centerline and must be pointed toward theradiation source. The gas supplied to the burnermust be methane and must be adjusted to pro-duce flame lengths of 1 inch (25 mm).

(iii) Optional Fourteen-Hole Upper Pilot Burner.This burner may be used in lieu of the standardthree-hole burner described in paragraph (b)(8)(ii)of this part IV. The pilot burner must be a straightlength of .25 inch (6.3 mm) O.D., .03 inch (0.8mm) wall, stainless steel tubing that is 15.75inches (400 mm) long. One end of the tubing must

be closed, and 14 No. 59 drill holes must bedrilled into the tubing, .50 inch (13 mm) apart, forgas ports, all radiating in the same direction. Thefirst hole must be .50 inch (13 mm) from theclosed end of the tubing. The tube must be posi-tioned above the specimen holder so that theholes are placed above the specimen as shown inFigure 1B of this part IV. The fuel supplied to theburner must be methane mixed with air in a ratioof approximately 50/50 by volume. The total gasflow must be adjusted to produce flame lengths of1 inch (25 mm). When the gas/air ratio and theflow rate are properly adjusted, approximately .25inch (6 mm) of the flame length appears yellow incolor.

(c) Calibration of Equipment.(1) Heat Release Rate. A calibration burner, as

shown in Figure 4, must be placed over the end ofthe lower pilot flame tubing using a gas tight con-nection. The flow of gas to the pilot flame must beat least 99 percent methane and must be accu-rately metered. Prior to usage, the wet test metermust be properly leveled and filled with distilledwater to the tip of the internal pointer while no gasis flowing. Ambient temperature and pressure ofthe water are based on the internal wet test metertemperature. A baseline flow rate of approxi-mately 1 liter/min. must be set and increased tohigher preset flows of 4, 6, 8, 6 and 4 liters/min.Immediately prior to recording methane flowrates, a flow rate of 8 liters/min. must be used for2 minutes to precondition the chamber. This is notrecorded as part of calibration. The rate must bedetermined by using a stopwatch to time a com-plete revolution of the wet test meter for both thebaseline and higher flow, with the flow returned tobaseline before changing to the next higher flow.The thermopile baseline voltage must be mea-sured. The gas flow to the burner must be in-creased to the higher preset flow and allowed toburn for 2.0 minutes, and the thermopile voltagemust be measured. The sequence must be re-peated until all five values have been determined.The average of the five values must be used asthe calibration factor. The procedure must be re-peated if the percent relative standard deviation isgreater than 5 percent. Calculations are shown inparagraph (f) of this part IV.

(2) Flux Uniformity. Uniformity of flux over thespecimen must be checked periodically and aftereach heating element change to determine if it iswithin acceptable limits of plus or minus 5 percent.

(3) As noted in paragraph (b)(2) of this part IV,thermopile hot junctions must be cleared of sootdeposits as needed to maintain the calibratedsensitivity.

(d) Preparation of Test Specimens.(1) The test specimens must be representative

of the aircraft component in regard to materialsand construction methods. The standard size for


ASA 183

25

the test specimens is 5.91 ±.03 x 5.91 ±.03inches (149 ±1 x 149 ±1 mm). The thickness ofthe specimen must be the same as that of the air-craft component it represents up to a maximumthickness of 1.75 inches (45 mm). Test specimensrepresenting thicker components must be 1.75inches (45 mm).

(2) Conditioning. Specimens must be condi-tioned as described in Part 1 of this appendix.

(3) Mounting. Each test specimen must bewrapped tightly on all sides of the specimen, ex-cept for the one surface that is exposed with a sin-gle layer of .001 inch (.025 mm) aluminum foil.

(e) Procedure.(1) The power supply to the radiant panel must

be set to produce a radiant flux of 3.5 ±.05 W/cm2,as measured at the point the center of the speci-men surface will occupy when positioned for thetest. The radiant flux must be measured after theair flow through the equipment is adjusted to thedesired rate.

(2) After the pilot flames are lighted, their posi-tion must be checked as described in paragraph(b)(8) of this part IV.

(3) Air flow through the apparatus must be con-trolled by a circular plate orifice located in a 1.5inch (38.1 mm) I.D. pipe with two pressure mea-suring points, located 1.5 inches (38 mm) up-stream and .75 inches (19 mm) downstream ofthe orifice plate. The pipe must be connected to amanometer set at a pressure differential of 7.87inches (200 mm) of Hg. (See Figure 1B of thispart IV.) The total air flow to the equipment is ap-proximately .04 m3/seconds. The stop on the ver-tical specimen holder rod must be adjusted sothat the exposed surface of the specimen is posi-tioned 3.9 inches (100 mm) from the entrancewhen injected into the environmental chamber.

(4) The specimen must be placed in the holdchamber with the radiation doors closed. The air-tight outer door must be secured, and the record-ing devices must be started. The specimen mustbe retained in the hold chamber for 60 seconds,plus or minus 10 seconds, before injection. Thethermopile “zero” value must be determined dur-ing the last 20 seconds of the hold period. Thesample must not be injected before completion ofthe “Zero” value determination.

(5) When the specimen is to be injected, the ra-diation doors must be opened. After the specimenis injected into the environmental chamber, the ra-diation doors must be closed behind the speci-men.

(6) [Reserved](7) Injection of the specimen and closure of the

inner door marks time zero. A record of the ther-mopile output with at least one data point per sec-ond must be made during the time the specimenis in the environmental chamber.

(8) The test duration time is five minutes. Thelower pilot burner and the upper pilot burner mustremain lighted for the entire duration of the test,except that there may be intermittent flame extin-guishment for periods that do not exceed 3 sec-onds. Furthermore, if the optional three-hole up-per burner is used, at least two flamelets must re-main lighted for the entire duration of the test,except that there may be intermittent flame extin-guishment of all three flamelets for periods that donot exceed 3 seconds.

(9) A minimum of three specimens must betested.

(f) Calculations.(1) The calibration factor is calculated as fol-

lows:

F0 = flow of methane at baseline (1pm)F1 = higher preset flow of methane (1pm)V0 = thermopile voltage at baseline (mv)V1 = thermopile voltage at higher flow (mv)Ta = Ambient temperature (K)P = Ambient pressure (mm Hg)Pv = Water vapor pressure (mm Hg)

(2) Heat release rates may be calculated fromthe reading of the thermopile output voltage atany instant of time as:

HRR = Heat release rate (kw/m2)Vb = baseline voltage (mv)Vm = measured thermopile voltage (mv)Kh = calibration factor (kw/mv)

(3) The integral of the heat release rate is thetotal heat release as a function of time and is cal-culated by multiplying the rate by the data sam-pling frequency in minutes and summing the timefrom zero to two minutes.

(g) Criteria. The total positive heat release overthe first two minutes of exposure for each of thethree or more samples tested must be averaged,and the peak heat release rate for each of thesamples must be averaged. The average totalheat release must not exceed 65 kilowatt-minutes

Kh

F1 F0–( )V 1 V 0–( )

------------------------210.8 22–( )kcal

mole---------------------------------------- 273

T a---------

P PV–

760---------------- mole CH4STP

22.41----------------------------------- WATT min

.01433 kcal---------------------------- kw

1000w----------------××××××=

HRRV m V b–( )Kn

.02323m2

---------------------------------=


184 ASA

per square meter, and the average peak heat re-lease rate must not exceed 65 kilowatts persquare meter.

(h) Report. The test report must include the fol-lowing for each specimen tested:

(1) Description of the specimen.(2) Radiant heat flux to the specimen, ex-

pressed in W/cm2.(3) Data giving release rates of heat (in kW/m2)

as a function of time, either graphically or tabu-lated at intervals no greater than 10 seconds. Thecalibration factor (kn) must be recorded.

(4) If melting, sagging, delaminating, or otherbehavior that affects the exposed surface area orthe mode of burning occurs, these behaviors mustbe reported, together with the time at which suchbehaviors were observed.

(5) The peak heat release and the 2-minute in-tegrated heat release rate must be reported.


ASA 185

25

5.2 ± 0.1 in(132 ± 2 mm)

Thermocouplelevel

Thermocouple placementin chimney1.2 in (30 mm) fromcenter on diagonal

2.8 ± 0.1 in(71 ± 2 mm)

7.0 ± 0.1 in(178 ± 3 mm)

0.39 ± 0.02 in(10 ± 0.5 mm)

Chimney0.018 ± 0.002 in(0.38 ± 0.05 mm) thick

Baffle1.00 ± 0.03 in x 3.00 ± 0.03 in(25 ± 1 mm x 76 ± 1 mm)0.018 ± 0.002 in(0.38 ± 0.05 mm) thick

Outer cone0.031 ± 0.002 in(0.79 ± 0.05 mm) thick

Inner cone0.018 ± 0.002 in(0.38 ± 0.05 mm) thick

12.20 ± 0.25 in(310 ± 6 mm)

Air manifoldwith 48 no. 26 holes

13.0 ± 0.25 in(330 ± 6mm)

0.75 ± 0.02 in(19.0 ± 0.5 mm)

Mask0.042 ± 0.002 inthick 7.4 ± 0.1 in

(188 ± 3mm)

2.00 ± 0.05 in(52 ± 1 mm)

1.5 in (38mm)nominalinsidediameter

31.0 ± 0.5 in(787 ± 12 mm)

Second stageair distributionplate with120 no. 28 holes

16.00 ± 0.25 in(406 ± 6 mm)

Air distribution platewith 8 no. 4 holes

6.00 ± 0.25 in(152 ± 6 mm)

Air distribution lower chamber0.049 ± 0.002 in(1.24 ± 0.05 mm) thick

Thermopile Air flow

0.25 ± 0.01 in(6 ± 0.3 mm)

4.00 ± 0.25 in(102 ± 6 mm)

Figure 1A Rate of Heat Release Apparatus


186 ASA

0.39 ± 0.02 in(10.0 ± 0.5 mm)

A

Baffle

NOTESeal and secure with bolts or clamps

A

Baffle Air manifold with48 no. 26 holes

Section A-A

+ +

Specimen top

Air manifold with48 no. 26 holes

16.00 ± 0.25 in(406 ± 6 mm)

Window frame

Specimen face

3.1 in(80 mm)

0.79 in(20 mm)

Hinged flap

0.25 ± 0.01 in(6 ± 0.3 mm)

13.25 ± 0.25 in(337 ± 0.6 mm)

Lowerpilotburner

3.9 in(100 mm)

0.39 in(10 mm)

0.25 om (6.4 mm)nominal outsidediameter

Hinge (typ)

+

26 in (711 mm)minimum

11 in (279mm)minimum

1.50 ± 0.03 in(38 ± 1 mm)

0.75 ± 0.02 in(19 ± 1 mm)

End view

Lowerthermopileconnection

1.5 in (38 mm)nominal insidediameter

0.750 ± 0.002 in(19.1 ± 0.1 mm) orifice24-gauge stainless steel

+

8.00 ± 0.25 in(203 ± 6 mm)

No. 4 holes

Air distribution platewith 8 no. 4 holes

Figure 1B Rate of Heat Release Apparatus


ASA 187

25

Top

Mask

50

78

150115

156

(Unless denoted otherwise all dimensions are in millimeters.)Figure 2A. “Globar” Radiant Panel


188 ASA

Top

25

115

2620

80

Reflector, adjust slope,top and bottom, foruniform heat flux onsample

Mask 19Gauge

1⁄4" - 20 Machine Screw,75 lg

39

39

78

20

80

(Unless denoted otherwise all dimensions are in millimeters.)Figure 2B. “Globar” Radiant Panel


ASA 189

25

SPRING

ANGLE IRONRETAINER FRAME

3/8 3/8

2 Radius Flange, 1 Long

.020StainlessSteelWire

156

6578 156

15

45

Spo

t Wel

d Fr

ame

35

6MM

3

5

SAMPLE HOLDER15

MOUNTING BRACKET

(Unless denoted otherwise,all dimensions are in millimeters.)

Figure 3.

165

165

24 2a Steel

10-32

MachineScrew

1⁄2 - 13 Nut

2 Radius, 5 Flange

Washer 50 mm OD13 mm ID

Drill &Tap for10-32 Machine Screw

Weld1⁄2" Nutto Washer

Weld

Washer

5

7

50 mm OD13 mm ID


190 ASA

125 mm

25

No. 32 Drill Hole

Leak-Free Seal on 6.35Pilot Tubing

110

(Unless denoted otherwise, all dimensions are in millimeters.)

Figure 4.

25

25

25

25

9.5 O.D.

9.5 Tubing


ASA 191

25

30MM

10 MM

Figure 5. Thermocouple Position


192 ASA

PART V — TEST METHOD TO DETERMINE THE SMOKE EMISSION

CHARACTERISTICS OF CABIN MATERIALS

(a) Summary of Method. The specimens mustbe constructed, conditioned, and tested in theflaming mode in accordance with American Soci-ety of Testing and Materials (ASTM) StandardTest Method ASTM F814-83.

(b) Acceptance Criteria. The specific opticalsmoke density (DS), which is obtained by averag-ing the reading obtained after 4 minutes with eachof the three specimens, shall not exceed 200.

PART VI — TEST METHOD TO DETERMINE THE FLAMMABILITY AND FLAME PROPAGATION

CHARACTERISTICS OF THERMAL/ACOUSTIC INSULATION MATERIALS

Use this test method to evaluate the flammabil-ity and flame propagation characteristics of ther-mal/acoustic insulation when exposed to both aradiant heat source and a flame.

(a) Definitions.“Flame propagation” means the furthest dis-

tance of the propagation of visible flame towardsthe far end of the test specimen, measured fromthe midpoint of the ignition source flame. Measurethis distance after initially applying the ignitionsource and before all flame on the test specimenis extinguished. The measurement is not a deter-mination of burn length made after the test.

“Radiant heat source” means an electric or airpropane panel.

“Thermal/acoustic insulation” means a mate-rial or system of materials used to provide thermaland/or acoustic protection. Examples include fi-berglass or other batting material encapsulatedby a film covering and foams.

“Zero point” means the point of application ofthe pilot burner to the test specimen.

(b) Test apparatus.

(1) Radiant panel test chamber. Conduct testsin a radiant panel test chamber (see figure 1above). Place the test chamber under an exhausthood to facilitate clearing the chamber of smokeafter each test. The radiant panel test chambermust be an enclosure 55 inches (1397 mm) longby 19.5 (495 mm) deep by 28 (710 mm) to 30inches (maximum) (762 mm) above the test spec-imen. Insulate the sides, ends, and top with a fi-brous ceramic insulation, such as Kaowool MTMboard. On the front side, provide a 52 by 12-inch(1321 by 305 mm) draft-free, high-temperature,glass window for viewing the sample during test-ing. Place a door below the window to provide ac-cess to the movable specimen platform holder.The bottom of the test chamber must be a slidingsteel platform that has provision for securing thetest specimen holder in a fixed and level position.The chamber must have an internal chimney withexterior dimensions of 5.1 inches (129 mm) wide,by 16.2 inches (411 mm) deep by 13 inches (330mm) high at the opposite end of the chamber fromthe radiant energy source. The interior dimen-sions must be 4.5 inches (114 mm) wide by 15.6inches (395 mm) deep. The chimney must extendto the top of the chamber (see figure 2).


ASA 193

25

(2) Radiant heat source. Mount the radiant heatenergy source in a cast iron frame or equivalent.An electric panel must have six, 3-inch wide emit-ter strips. The emitter strips must be perpendicu-lar to the length of the panel. The panel must havea radiation surface of 12-7/8 by 18-1/2 inches(327 by 470 mm). The panel must be capable ofoperating at temperatures up to 1300°F (704°C).An air propane panel must be made of a porousrefractory material and have a radiation surface of12 by 18 inches (305 by 457 mm). The panel mustbe capable of operating at temperatures up to1,500°F (816°C). See figures 3a and 3b.

(i) Electric radiant panel. The radiant panelmust be 3-phase and operate at 208 volts. A sin-gle-phase, 240 volt panel is also acceptable. Usea solid-state power controller and microproces-sor-based controller to set the electric panel oper-ating parameters.

(ii) Gas radiant panel. Use propane (liquid pe-troleum gas—2.1 UN 1075) for the radiant panelfuel. The panel fuel system must consist of a ven-turi-type aspirator for mixing gas and air at ap-proximately atmospheric pressure. Provide suit-able instrumentation for monitoring and control-ling the flow of fuel and air to the panel. Include anair flow gauge, an air flow regulator, and a gaspressure gauge.

(iii) Radiant panel placement. Mount the panelin the chamber at 30° to the horizontal specimenplane, and 7-1/2 inches above the zero point ofthe specimen.

(3) Specimen holding system.(i) The sliding platform serves as the housing

for test specimen placement. Brackets may be at-tached (via wing nuts) to the top lip of the platformin order to accommodate various thicknesses oftest specimens. Place the test specimens on asheet of Kaowool MTM board or 1260 StandardBoard (manufactured by Thermal Ceramics andavailable in Europe), or equivalent, either restingon the bottom lip of the sliding platform or on thebase of the brackets. It may be necessary to usemultiple sheets of material based on the thicknessof the test specimen (to meet the sample heightrequirement). Typically, these non-combustiblesheets of material are available in 1/4 inch (6 mm)thicknesses. See figure 4. A sliding platform thatis deeper than the 2-inch (50.8mm) platformshown in figure 4 is also acceptable as long as thesample height requirement is met.


194 ASA

(ii) Attach a 1/2 inch (13 mm) piece of KaowoolMTM board or other high temperature materialmeasuring 41-1/2 by 8-1/4 inches (1054 by 210mm) to the back of the platform. This boardserves as a heat retainer and protects the testspecimen from excessive preheating. The heightof this board must not impede the sliding platformmovement (in and out of the test chamber). If theplatform has been fabricated such that the backside of the platform is high enough to prevent ex-cess preheating of the specimen when the slidingplatform is out, a retainer board is not necessary.

(iii) Place the test specimen horizontally on thenon-combustible board(s). Place a steel retain-ing/securing frame fabricated of mild steel, havinga thickness of 1/8 inch (3.2 mm) and overall di-mensions of 23 by 13-1/8 inches (584 by 333 mm)with a specimen opening of 19 by 10-3/4 inches(483 by 273 mm) over the test specimen. Thefront, back, and right portions of the top flange ofthe frame must rest on the top of the sliding plat-form, and the bottom flanges must pinch all 4sides of the test specimen. The right bottomflange must be flush with the sliding platform. Seefigure 5.


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(4) Pilot Burner. The pilot burner used to ignitethe specimen must be a BernzomaticTM commer-cial propane venturi torch with an axially symmet-ric burner tip and a propane supply tube with anorifice diameter of 0.006 inches (0.15 mm). Thelength of the burner tube must be 2-7/8 inches (71mm). The propane flow must be adjusted via gaspressure through an in-line regulator to produce ablue inner cone length of 3/4 inch (19 mm). A 3/4inch (19 mm) guide (such as a thin strip of metal)may be soldered to the top of the burner to aid insetting the flame height. The overall flame lengthmust be approximately 5 inches long (127 mm).Provide a way to move the burner out of the igni-

tion position so that the flame is horizontal and atleast 2 inches (50 mm) above the specimenplane. See figure 6.


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(5) Thermocouples. Install a 24 American WireGauge (AWG) Type K (Chromel-Alumel) thermo-couple in the test chamber for temperature moni-toring. Insert it into the chamber through a smallhole drilled through the back of the chamber.Place the thermocouple so that it extends 11inches (279 mm) out from the back of the cham-ber wall, 11-1/2 inches (292 mm) from the rightside of the chamber wall, and is 2 inches (51 mm)below the radiant panel. The use of other thermo-couples is optional.

(6) Calorimeter. The calorimeter must be aone-inch cylindrical water-cooled, total heat fluxdensity, foil type Gardon Gage that has a range of0 to 5 BTU/ft2-second (0 to 5.7 Watts/cm2).

(7) Calorimeter calibration specification andprocedure.

(i) Calorimeter specification.(A) Foil diameter must be 0.25 ±0.005 inches

(6.35 ±0.13 mm).(B) Foil thickness must be 0.0005 ±0.0001

inches (0.013 ±0.0025 mm).(C) Foil material must be thermocouple grade

Constantan.(D) Temperature measurement must be a Cop-

per Constantan thermocouple.(E) The copper center wire diameter must be

0.0005 inches (0.013 mm).(F) The entire face of the calorimeter must be

lightly coated with “Black Velvet” paint having anemissivity of 96 or greater.

(ii) Calorimeter calibration.(A) The calibration method must be by compar-

ison to a like standardized transducer.(B) The standardized transducer must meet the

specifications given in paragraph VI(b)(6) of thisappendix.

(C) Calibrate the standard transducer against aprimary standard traceable to the National Insti-tute of Standards and Technology (NIST).

(D) The method of transfer must be a heatedgraphite plate.

(E) The graphite plate must be electricallyheated, have a clear surface area on each side ofthe plate of at least 2 by 2 inches (51 by 51 mm),and be 1/8 inch ±1/16 inch thick (3.2 ±1.6 mm).

(F) Center the 2 transducers on opposite sidesof the plates at equal distances from the plate.

(G) The distance of the calorimeter to the platemust be no less than 0.0625 inches (1.6 mm), norgreater than 0.375 inches (9.5 mm).

(H) The range used in calibration must be atleast 0-3.5 BTUs/ft2 second (0-3.9 Watts/cm2)and no greater than 0-5.7 BTUs/ft2 second (0-6.4Watts/cm2).

(I) The recording device used must record the 2transducers simultaneously or at least within 1/10of each other.

(8) Calorimeter fixture. With the sliding platformpulled out of the chamber, install the calorimeterholding frame and place a sheet of non-combusti-ble material in the bottom of the sliding platformadjacent to the holding frame. This will preventheat losses during calibration. The frame must be13-1/8 inches (333 mm) deep (front to back) by 8inches (203 mm) wide and must rest on the top ofthe sliding platform. It must be fabricated of 1/8inch (3.2 mm) flat stock steel and have an openingthat accommodates a 1/2 inch (12.7 mm) thickpiece of refractory board, which is level with thetop of the sliding platform. The board must havethree 1-inch (25.4 mm) diameter holes drilledthrough the board for calorimeter insertion. Thedistance to the radiant panel surface from the cen-terline of the first hole (“zero” position) must be 7-1/2 ±1/8 inches (191 ±3 mm). The distance be-tween the centerline of the first hole to the center-line of the second hole must be 2 inches (51 mm).It must also be the same distance from the center-line of the second hole to the centerline of the thirdhole. See figure 7. A calorimeter holding framethat differs in construction is acceptable as long asthe height from the centerline of the first hole to theradiant panel and the distance between holes isthe same as described in this paragraph.


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(9) Instrumentation. Provide a calibrated re-cording device with an appropriate range or acomputerized data acquisition system to measureand record the outputs of the calorimeter and thethermocouple. The data acquisition system mustbe capable of recording the calorimeter output ev-ery second during calibration.

(10) Timing device. Provide a stopwatch orother device, accurate to ±1 second/hour, to mea-sure the time of application of the pilot burnerflame.

(c) Test specimens.(1) Specimen preparation. Prepare and test a

minimum of three test specimens. If an orientedfilm cover material is used, prepare and test boththe warp and fill directions.

(2) Construction. Test specimens must includeall materials used in construction of the insulation(including batting, film, scrim, tape etc.). Cut apiece of core material such as foam or fiberglass,and cut a piece of film cover material (if used)large enough to cover the core material. Heatsealing is the preferred method of preparing fiber-glass samples, since they can be made withoutcompressing the fiberglass (“box sample”). Covermaterials that are not heat sealable may be sta-pled, sewn, or taped as long as the cover materialis over-cut enough to be drawn down the sideswithout compressing the core material. The fas-tening means should be as continuous as possi-ble along the length of the seams. The specimenthickness must be of the same thickness as in-stalled in the airplane.

(3) Specimen Dimensions. To facilitate properplacement of specimens in the sliding platformhousing, cut non-rigid core materials, such as fi-berglass, 12-1/2 inches (318mm) wide by 23inches (584mm) long. Cut rigid materials, such asfoam, 11-1/2 ±1/4 inches (292 mm ±6mm) wideby 23 inches (584mm) long in order to fit properlyin the sliding platform housing and provide a flat,exposed surface equal to the opening in the hous-ing.

(d) Specimen conditioning. Condition the testspecimens at 70 ±5°F (21 ±2°C) and 55% ±10%relative humidity, for a minimum of 24 hours priorto testing.

(e) Apparatus Calibration.(1) With the sliding platform out of the chamber,

install the calorimeter holding frame. Push theplatform back into the chamber and insert the cal-orimeter into the first hole (“zero” position). Seefigure 7. Close the bottom door located below thesliding platform. The distance from the centerlineof the calorimeter to the radiant panel surface atthis point must be 7-1/2 inches ±1/8 (191 mm ±3).Prior to igniting the radiant panel, ensure that thecalorimeter face is clean and that there is waterrunning through the calorimeter.

(2) Ignite the panel. Adjust the fuel/air mixtureto achieve 1.5 BTUs/ft2-second ±5% (1.7Watts/cm2 ±5%) at the “zero” position. If using anelectric panel, set the power controller to achievethe proper heat flux. Allow the unit to reach steadystate (this may take up to 1 hour). The pilot burnermust be off and in the down position during thistime.


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(3) After steady-state conditions have beenreached, move the calorimeter 2 inches (51 mm)from the “zero” position (first hole) to position 1and record the heat flux. Move the calorimeter toposition 2 and record the heat flux. Allow enoughtime at each position for the calorimeter to stabi-lize. Table 1 depicts typical calibration values atthe three positions.

Table 1Calibration Table

(4) Open the bottom door, remove the calorim-eter and holder fixture. Use caution as the fixtureis very hot.

(f) Test Procedure.(1) Ignite the pilot burner. Ensure that it is at

least 2 inches (51 mm) above the top of the plat-form. The burner must not contact the specimenuntil the test begins.

(2) Place the test specimen in the sliding plat-form holder. Ensure that the test sample surfaceis level with the top of the platform. At “zero” point,the specimen surface must be 7-1/2 inches ±1/8inch (191 mm ±3) below the radiant panel.

(3) Place the retaining/securing frame over thetest specimen. It may be necessary (due to com-pression) to adjust the sample (up or down) in or-der to maintain the distance from the sample tothe radiant panel (7-1/2 inches ±1/8 inch (191 mm±3) at “zero” position). With film/fiberglass assem-blies, it is critical to make a slit in the film cover topurge any air inside. This allows the operator tomaintain the proper test specimen position (levelwith the top of the platform) and to allow ventila-tion of gases during testing. A longitudinal slit, ap-proximately 2 inches (51mm) in length, must becentered 3 inches ±1/2 inch (76mm ±13mm) fromthe left flange of the securing frame. A utility knifeis acceptable for slitting the film cover.

(4) Immediately push the sliding platform intothe chamber and close the bottom door.

(5) Bring the pilot burner flame into contact withthe center of the specimen at the “zero” point andsimultaneously start the timer. The pilot burnermust be at a 27° angle with the sample and be ap-proximately 1/2 inch (12 mm) above the sample.See figure 7. A stop, as shown in figure 8, allowsthe operator to position the burner correctly eachtime.

(6) Leave the burner in position for 15 secondsand then remove to a position at least 2 inches(51 mm) above the specimen.

(g) Report.(1) Identify and describe the test specimen.(2) Report any shrinkage or melting of the test

specimen.(3) Report the flame propagation distance. If

this distance is less than 2 inches, report this as apass (no measurement required).

(4) Report the after-flame time.(h) Requirements.(1) There must be no flame propagation be-

yond 2 inches (51 mm) to the left of the centerlineof the pilot flame application.

(2) The flame time after removal of the pilotburner may not exceed 3 seconds on any speci-men.

PART VII —TEST METHOD TO DETERMINE THE BURNTHROUGH RESISTANCE OF

THERMAL/ACOUSTIC INSULATION MATERIALS

Use the following test method to evaluate theburnthrough resistance characteristics of aircraftthermal/acoustic insulation materials when ex-posed to a high intensity open flame.

(a) Definitions.Burnthrough time means the time, in sec-

onds, for the burner flame to penetrate the testspecimen, and/or the time required for the heatflux to reach 2.0 Btu/ft2sec (2.27 W/cm2) on theinboard side, at a distance of 12 inches (30.5 cm)from the front surface of the insulation blanket test

Position BTUs / ft 2sec Watts / cm2

“Zero” Position 1.5 1.7

Position 1 1.51–1.50–1.49 1.71–1.70–1.69

Position 2 1.43–1.44 1.62–1.63


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frame, whichever is sooner. The burnthrough timeis measured at the inboard side of each of the in-sulation blanket specimens.

Insulation blanket specimen means one oftwo specimens positioned in either side of the testrig, at an angle of 30° with respect to vertical.

Specimen set means two insulation blanketspecimens. Both specimens must represent thesame production insulation blanket constructionand materials, proportioned to correspond to thespecimen size.

(b) Apparatus.(1) The arrangement of the test apparatus is

shown in figures 1 and 2 and must include the ca-pability of swinging the burner away from the testspecimen during warm-up.SEE FIGURE 1 AT THE END OF PART VII OF THIS APPENDIX

(2) Test burner. The test burner must be a mod-ified gun-type such as the Park Model DPL 3400.Flame characteristics are highly dependent onactual burner setup. Parameters such as fuelpressure, nozzle depth, stator position, and intakeairflow must be properly adjusted to achieve thecorrect flame output.SEE FIGURE 2 AT THE END OF PART VII OF THIS APPENDIX

(i) Nozzle. A nozzle must maintain the fuelpressure to yield a nominal 6.0 gal/hr (0.378L/min) fuel flow. A Monarch-manufactured 80° PL(hollow cone) nozzle nominally rated at 6.0 gal/hrat 100 lb/in2 (0.71 MPa) delivers a proper spraypattern.

(ii) Fuel Rail. The fuel rail must be adjusted toposition the fuel nozzle at a depth of 0.3125 inch(8 mm) from the end plane of the exit stator, whichmust be mounted in the end of the draft tube.

(iii) Internal Stator. The internal stator, locatedin the middle of the draft tube, must be positionedat a depth of 3.75 inches (95 mm) from the tip ofthe fuel nozzle. The stator must also be posi-tioned such that the integral igniters are located atan angle midway between the 10 and 11 o’clockposition, when viewed looking into the draft tube.Minor deviations to the igniter angle are accept-able if the temperature and heat flux requirementsconform to the requirements of paragraph VII(e)of this appendix.

(iv) Blower Fan. The cylindrical blower fan usedto pump air through the burner must measure5.25 inches (133 mm) in diameter by 3.5 inches(89 mm) in width.

(v) Burner cone. Install a 12 +0.125-inch (305±3 mm) burner extension cone at the end of thedraft tube. The cone must have an opening 6±0.125-inch (152 ±3 mm) high and 11 ±0.125-inch (280 ±3 mm) wide (see figure 3).

(vi) Fuel. Use JP-8, Jet A, or their internationalequivalent, at a flow rate of 6.0 ±0.2 gal/hr (0.378±0.0126 L/min). If this fuel is unavailable, ASTMK2 fuel (Number 2 grade kerosene) or ASTM D2

fuel (Number 2 grade fuel oil or Number 2 dieselfuel) are acceptable if the nominal fuel flow rate,temperature, and heat flux measurements con-form to the requirements of paragraph VII(e) ofthis appendix.

(vii) Fuel pressure regulator. Provide a fuelpressure regulator, adjusted to deliver a nominal6.0 gal/hr (0.378 L/min) flow rate. An operatingfuel pressure of 100 lb/in2 (0.71 MPa) for a nomi-nally rated 6.0 gal/hr 80° spray angle nozzle (suchas a PL type) delivers 6.0 ±0.2 gal/hr (0.378±0.0126 L/min).SEE FIGURE 3 AT THE END OF PART VII OF THIS APPENDIX

(3) Calibration rig and equipment.(i) Construct individual calibration rigs to incor-

porate a calorimeter and thermocouple rake forthe measurement of heat flux and temperature.Position the calibration rigs to allow movement ofthe burner from the test rig position to either theheat flux or temperature position with minimal dif-ficulty.

(ii) Calorimeter. The calorimeter must be a totalheat flux, foil type Gardon Gage of an appropriaterange such as 0-20 Btu/ft 2-sec (0-22.7 W/cm2),accurate to ±3% of the indicated reading. Theheat flux calibration method must be in accor-dance with paragraph VI(b)(7) of this appendix.

(iii) Calorimeter mounting. Mount the calorime-ter in a 6- by 12- ±0.125 inch (152- by 305- ±3mm) by 0.75 ±0.125 inch (19 mm ±3 mm) thick in-sulating block which is attached to the heat fluxcalibration rig during calibration (figure 4). Monitorthe insulating block for deterioration and replace itwhen necessary. Adjust the mounting as neces-sary to ensure that the calorimeter face is parallelto the exit plane of the test burner cone.SEE FIGURES 4 AND 5 AT THE END OF PART VII OF THISAPPENDIX

(iv) Thermocouples. Provide seven 1/8-inch(3.2 mm) ceramic packed, metal sheathed, type K(Chromel-alumel), grounded junction thermocou-ples with a nominal 24 American Wire Gauge(AWG) size conductor for calibration. Attach thethermocouples to a steel angle bracket to form athermocouple rake for placement in the calibra-tion rig during burner calibration (figure 5).

(v) Air velocity meter. Use a vane-type air ve-locity meter to calibrate the velocity of air enteringthe burner. An Omega Engineering Model HH30Ais satisfactory. Use a suitable adapter to attachthe measuring device to the inlet side of theburner to prevent air from entering the burnerother than through the measuring device, whichwould produce erroneously low readings. Use aflexible duct, measuring 4 inches wide (102 mm)by 20 feet long (6.1 meters), to supply fresh air tothe burner intake to prevent damage to the air ve-locity meter from ingested soot. An optional airboxpermanently mounted to the burner intake area


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can effectively house the air velocity meter andprovide a mounting port for the flexible intakeduct.

(4) Test specimen mounting frame. Make themounting frame for the test specimens of 1/8-inch(3.2 mm) thick steel as shown in figure 1, exceptfor the center vertical former, which should be 1/4-inch (6.4 mm) thick to minimize warpage. Thespecimen mounting frame stringers (horizontal)should be bolted to the test frame formers (verti-cal) such that the expansion of the stringers willnot cause the entire structure to warp. Use themounting frame for mounting the two insulationblanket test specimens as shown in figure 2.

(5) Backface calorimeters. Mount two total heatflux Gardon type calorimeters behind the insula-tion test specimens on the back side (cold) areaof the test specimen mounting frame as shown infigure 6. Position the calorimeters along the sameplane as the burner cone centerline, at a distanceof 4 inches (102 mm) from the vertical centerlineof the test frame.SEE FIGURE 6 AT THE END OF PART VII OF THIS APPENDIX

(i) The calorimeters must be a total heat flux,foil type Gardon Gage of an appropriate rangesuch as 0-5 Btu/ft2-sec (0-5.7 W/cm2), accurate to±3% of the indicated reading. The heat flux cali-bration method must comply with paragraphVI(b)(7) of this appendix.

(6) Instrumentation. Provide a recording poten-tiometer or other suitable calibrated instrumentwith an appropriate range to measure and recordthe outputs of the calorimeter and the thermocou-ples.

(7) Timing device. Provide a stopwatch or otherdevice, accurate to ±1%, to measure the time ofapplication of the burner flame and burnthroughtime.

(8) Test chamber. Perform tests in a suitablechamber to reduce or eliminate the possibility oftest fluctuation due to air movement. The cham-ber must have a minimum floor area of 10 by 10feet (305 by 305 cm).

(i) Ventilation hood. Provide the test chamberwith an exhaust system capable of removing theproducts of combustion expelled during tests.

(c) Test Specimens.(1) Specimen preparation. Prepare a minimum

of three specimen sets of the same constructionand configuration for testing.

(2) Insulation blanket test specimen.(i) For batt-type materials such as fiberglass,

the constructed, finished blanket specimen as-semblies must be 32 inches wide by 36 incheslong (81.3 by 91.4 cm), exclusive of heat sealedfilm edges.

(ii) For rigid and other non-conforming types ofinsulation materials, the finished test specimens

must fit into the test rig in such a manner as toreplicate the actual in-service installation.

(3) Construction. Make each of the specimenstested using the principal components (i.e., insu-lation, fire barrier material if used, and moisturebarrier film) and assembly processes (representa-tive seams and closures).

(i) Fire barrier material. If the insulation blanketis constructed with a fire barrier material, placethe fire barrier material in a manner reflective ofthe installed arrangement For example, if the ma-terial will be placed on the outboard side of the in-sulation material, inside the moisture film, place itthe same way in the test specimen.

(ii) Insulation material. Blankets that utilizemore than one variety of insulation (composition,density, etc.) must have specimen sets con-structed that reflect the insulation combinationused. If, however, several blanket types use simi-lar insulation combinations, it is not necessary totest each combination if it is possible to bracketthe various combinations.

(iii) Moisture barrier film. If a production blanketconstruction utilizes more than one type of mois-ture barrier film, perform separate tests on eachcombination. For example, if a polyimide film isused in conjunction with an insulation in order toenhance the burnthrough capabilities, also testthe same insulation when used with a polyvinylfluoride film.

(iv) Installation on test frame. Attach the blan-ket test specimens to the test frame using 12 steelspring type clamps as shown in figure 7. Use theclamps to hold the blankets in place in both of theouter vertical formers, as well as the center verti-cal former (4 clamps per former). The clamp sur-faces should measure 1 inch by 2 inches (25 by51 mm). Place the top and bottom clamps 6inches (15.2 cm) from the top and bottom of thetest frame, respectively. Place the middle clamps8 inches (20.3 cm) from the top and bottomclamps.SEE FIGURE 7 AT THE END OF PART VII OF THIS APPENDIX

(Note: For blanket materials that cannot be in-stalled in accordance with figure 7 above, theblankets must be installed in a manner approvedby the FAA.)

(v) Conditioning. Condition the specimens at70° ±5°F (21° ±2°C) and 55% ±10% relative hu-midity for a minimum of 24 hours prior to testing.

(d) Preparation of apparatus.(1) Level and center the frame assembly to en-

sure alignment of the calorimeter and/or thermo-couple rake with the burner cone.

(2) Turn on the ventilation hood for the testchamber. Do not turn on the burner blower. Mea-sure the airflow of the test chamber using a vaneanemometer or equivalent measuring device. Thevertical air velocity just behind the top of the upper


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insulation blanket test specimen must be 100 ±50ft/min (0.51 ±0.25 m/s). The horizontal air velocityat this point must be less than 50 ft/min (0.25m/s).

(3) If a calibrated flow meter is not available,measure the fuel flow rate using a graduated cyl-inder of appropriate size. Turn on the burner mo-tor/fuel pump, after insuring that the igniter sys-tem is turned off. Collect the fuel via a plastic orrubber tube into the graduated cylinder for a 2-minute period. Determine the flow rate in gallonsper hour. The fuel flow rate must be 6.0 ±0.2 gal-lons per hour (0.378 ±0.0126 L/min).

(e) Calibration.(1) Position the burner in front of the calorime-

ter so that it is centered and the vertical plane ofthe burner cone exit is 4 ±0.125 inches (102 ±3mm) from the calorimeter face. Ensure that thehorizontal centerline of the burner cone is offset 1inch below the horizontal centerline of the calo-rimeter (figure 8). Without disturbing the calorime-ter position, rotate the burner in front of the ther-mocouple rake, such that the middle thermocou-ple (number 4 of 7) is centered on the burnercone.SEE FIGURE 8 AT THE END OF PART VII OF THIS APPENDIX

Ensure that the horizontal centerline of theburner cone is also offset 1 inch below the hori-zontal centerline of the thermocouple tips. Re-check measurements by rotating the burner toeach position to ensure proper alignment be-tween the cone and the calorimeter and thermo-couple rake. (Note: The test burner mounting sys-tem must incorporate “detents” that ensure propercentering of the burner cone with respect to boththe calorimeter and the thermocouple rakes, sothat rapid positioning of the burner can beachieved during the calibration procedure.)

(2) Position the air velocity meter in the adapteror airbox, making certain that no gaps exist whereair could leak around the air velocity measuringdevice. Turn on the blower/motor while ensuringthat the fuel solenoid and igniters are off. Adjustthe air intake velocity to a level of 2150 ft/min,(10.92 m/s) then turn off the blower/motor. (Note:The Omega HH30 air velocity meter measures2.625 inches in diameter. To calculate the intakeairflow, multiply the cross-sectional area (0.03758ft2) by the air velocity (2150 ft/min) to obtain 80.80ft3/min. An air velocity meter other than the HH30unit can be used, provided the calculated airflowof 80.80 ft3/min (2.29 m3/min) is equivalent.)

(3) Rotate the burner from the test position tothe warm-up position. Prior to lighting the burner,ensure that the calorimeter face is clean of sootdeposits, and there is water running through thecalorimeter. Examine and clean the burner coneof any evidence of buildup of products of combus-tion, soot, etc. Soot buildup inside the burner

cone may affect the flame characteristics andcause calibration difficulties. Since the burnercone may distort with time, dimensions should bechecked periodically.

(4) While the burner is still rotated to the warm-up position, turn on the blower/motor, igniters andfuel flow, and light the burner. Allow it to warm upfor a period of 2 minutes. Move the burner into thecalibration position and allow 1 minute for calorim-eter stabilization, then record the heat flux onceevery second for a period of 30 seconds. Turn offburner, rotate out of position, and allow to cool.Calculate the average heat flux over this 30-sec-ond duration. The average heat flux should be16.0 ±0.8 Btu/ft2 sec (18.2 ±0.9 W/cm2).

(5) Position the burner in front of the thermo-couple rake. After checking for proper alignment,rotate the burner to the warm-up position, turn onthe blower/motor, igniters and fuel flow, and lightthe burner. Allow it to warm up for a period of 2minutes. Move the burner into the calibration posi-tion and allow 1 minute for thermocouple stabili-zation, then record the temperature of each of the7 thermocouples once every second for a periodof 30 seconds. Turn off burner, rotate out of posi-tion, and allow to cool. Calculate the average tem-perature of each thermocouple over this 30-sec-ond period and record. The average temperatureof each of the 7 thermocouples should be 1900°F±100°F (1038 ±56°C).

(6) If either the heat flux or the temperaturesare not within the specified range, adjust theburner intake air velocity and repeat the proce-dures of paragraphs (4) and (5) above to obtainthe proper values. Ensure that the inlet air velocityis within the range of 2150 ft/min ±50 ft/min (10.92±0.25 m/s).

(7) Calibrate prior to each test until consistencyhas been demonstrated. After consistency hasbeen confirmed, several tests may be conductedwith calibration conducted before and after a se-ries of tests.

(f) Test procedure.(1) Secure the two insulation blanket test spec-

imens to the test frame. The insulation blanketsshould be attached to the test rig center verticalformer using four spring clamps positioned asshown in figure 7 (according to the criteria ofparagraph (c)(3)(iv) of this part of this appendix).

(2) Ensure that the vertical plane of the burnercone is at a distance of 4 ±0.125 inch (102 ±3mm) from the outer surface of the horizontalstringers of the test specimen frame, and that theburner and test frame are both situated at a 30°angle with respect to vertical.

(3) When ready to begin the test, direct theburner away from the test position to the warm-upposition so that the flame will not impinge on thespecimens prematurely. Turn on and light theburner and allow it to stabilize for 2 minutes.


202 ASA

(4) To begin the test, rotate the burner into thetest position and simultaneously start the timingdevice.

(5) Expose the test specimens to the burnerflame for 4 minutes and then turn off the burner.Immediately rotate the burner out of the test posi-tion.

(6) Determine (where applicable) the burn-through time, or the point at which the heat fluxexceeds 2.0 Btu/ft2-sec (2.27 W/cm2).

(g) Report.(1) Identify and describe the specimen being

tested.(2) Report the number of insulation blanket

specimens tested.(3) Report the burnthrough time (if any), and

the maximum heat flux on the back face of the in-sulation blanket test specimen, and the time atwhich the maximum occurred.

(h) Requirements.(1) Each of the two insulation blanket test spec-

imens must not allow fire or flame penetration inless than 4 minutes.

(2) Each of the two insulation blanket test spec-imens must not allow more than 2.0 Btu/ft2-sec(2.27 W/cm2) on the cold side of the insulationspecimens at a point 12 inches (30.5 cm) from theface of the test rig.

SEE PART VII FIGURES BEGINNING ON THE NEXT PAGE.


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[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; as amended by Amdt. 25–32, 37 FR 3972, Feb. 24, 1972; 37 FR5284, March 14, 1972; Amdt. 25–55, 47 FR 13315, March 29, 1982; Amdt. 25–59, 49 FR 43193, Oct. 26, 1984;Amdt. 25–60, 51 FR 18243, May 16, 1986; Amdt. 25–61, 51 FR 26214, July 21, 1986; 51 FR 28322, Aug. 7, 1986;Amdt. 25–66, 53 FR 32573, Aug. 25, 1988; 53 FR 37542, 37671, Sept. 27, 1988; Amdt. 25–72, 55 FR 29787, July 20,1990; Amdt. 25–83, 60 FR 6623, Feb. 2, 1995; Amdt. 25–83, 60 FR 11194, March 1, 1995; Amdt. 25–94, 63 FR8848, Feb. 23, 1998; Docket No. FAA–2000–7909, Amdt. 25–111, 68 FR 45059, July 31, 2003; Amdt. 25–128, 74 FR25645, May 29, 2009]

Part 25: Airworthiness Standards: Transport Category Appendix G to Part 25

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APPENDIX G TO PART 25CONTINUOUS GUST DESIGN CRITERIA

The continuous gust design criteria in this ap-pendix must be used in establishing the dynamicresponse of the airplane to vertical and lateralcontinuous turbulence unless a more rational cri-teria is used. The following gust load require-ments apply to mission analysis and design enve-lope analysis:

(a) The limit gust loads utilizing the continuousturbulence concept must be determined in accor-dance with the provisions of either paragraph (b)or paragraphs (c) and (d) of this appendix.

(b) Design envelope analysis. The limit loadsmust be determined in accordance with the fol-lowing:

(1) All critical altitudes, weights, and weight dis-tributions, as specified in §25.321(b), and all criti-cal speeds within the ranges indicated in para-graph (b)(3) of this appendix must be considered.

(2) Values of A8 (ratio of root-mean-square in-cremental load root-mean-square gust velocity)must be determined by dynamic analysis. Thepower spectral density of the atmospheric turbu-lence must be as given by the equation—

where:φ = power-spectral density (ft./sec.)2/rad./ft.σ = root-mean-square gust velocity, ft./sec.Ω = reduced frequency, radians per foot.L = 2,500 ft.

(3) The limit loads must be obtained by multi-plying the A values determined by the dynamicanalysis by the following values of the gust veloc-ity Uσ:

(i) At speed VC: Uσ = 85 fps true gust velocity inthe interval 0 to 30,000 ft. altitude and is linearlydecreased to 30 fps true gust velocity at 80,000 ft.altitude. Where the Administrator finds that a de-sign is comparable to a similar design with exten-sive satisfactory service experience, it will be ac-ceptable to select Uσ at VC, less than 85 fps, butnot less than 75 fps, with linear decrease fromthat value at 20,000 feet to 30 fps at 80,000 feet.The following factors will be taken into accountwhen assessing comparability to a similar design:

(1) The transfer function of the new designshould exhibit no unusual characteristics as com-pared to the similar design which will significantlyaffect response to turbulence; e.g., coalescenceof modal response in the frequency regime whichcan result in a significant increase of loads.

(2) The typical mission of the new airplane issubstantially equivalent to that of the similar de-sign.

(3) The similar design should demonstrate theadequacy of the Uσ selected.

(ii) At speed VB: Uσ is equal to 1.32 times thevalues obtained under paragraph (b)(3)(i) of thisappendix.

(iii) At speed VD: Uσ is equal to 1⁄2 the values ob-tained under paragraph (b)(3)(i) of this appendix.

(iv) At speeds between VB and VC and betweenVC and VD: Uσ is equal to a value obtained by lin-ear interpolation.

(4) When a stability augmentation system is in-cluded in the analysis, the effect of system nonlin-earities on loads at the limit load level must be re-alistically or conservatively accounted for.

(c) Mission analysis. Limit loads must be deter-mined in accordance with the following:

(1) The expected utilization of the airplane mustbe represented by one or more night profiles inwhich the load distribution and the variation withtime of speed, altitude, gross weight, and centerof gravity position are defined. These profilesmust be divided into mission segments or blocks,for analysis, and average or effective values of thepertinent parameters defined for each segment.

(2) For each of the mission segments definedunder paragraph (c)(1) of this appendix, values ofA and NO must be determined by analysis. A isdefined as the ratio of root-mean-square incre-mental load to root-mean-square gust velocityand NO is the radius of gyration of the load powerspectral density function about zero frequency.

φ Ω( ) σ2L/π 1 + 8/3 (1.339 LΩ )2

1 + 1.339LΩ( )2[ ] 11 6⁄-------------------------------------------------------=

Appendix G to Part 25 Federal Aviation Regulations

212 ASA

The power spectral density of the atmosphericturbulence must be given by the equation set forthin paragraph (b)(2) of this appendix.

(3) For each of the load and stress quantitiesselected, the frequency of exceedance must bedetermined as a function of load level by meansof the equation—

where:

t = selected time interval.y = net value of the load or stress.Yone – g = value of the load or stress in one-g level

flight.N(y) = average number of exceedances of the

indicated value of the load or stress in unit time.

Σ = symbol denoting summation over all mission segments.

No, A = parameters determined by dynamic analysis as defined in paragraph (c)(2) of this appendix.

P1, P2, b1, b2 = parameters defining the probability distributions of root-mean-square gust velocity, to be read from Figures 1 and 2 of this appendix.

The limit gust loads must be read form the fre-quency of exceedance curves at a frequency ofexceedance of 2 x 10-5 exceedances per hour.Both positive and negative load directions mustbe considered in determining the limit loads.

(4) If a stability augmentation system is utilizedto reduce the gust loads, consideration must begiven to the fraction of flight time that the systemmay be inoperative. The flight profiles of para-graph (c)(1) of this appendix must include flightwith the system inoperative for this fraction of theflight time. When a stability augmentation systemis included in the analysis, the effect of systemnonlinearities on loads at the limit load level mustbe conservatively accounted for.

(d) Supplementary design envelope analysis.In addition to the limit loads defined by paragraph(c) of this appendix, limit loads must also be de-termined in accordance with paragraph (b) of thisappendix, except that—

(1) In paragraph (b)(3)(i) of this appendix, thevalue of Uσ = 85 fps true gust velocity is replacedby Uσ = 60 fps true gust velocity on the interval 0to 30,000 ft. altitude, and is linearly decreased to25 fps true gust velocity at 80,000 ft. altitude; and

(2) In paragraph (b) of this appendix, the refer-ence to paragraphs (b)(3)(i) through (b)(3)(iii) ofthis appendix is to be understood as referring tothe paragraph as modified by paragraph (d)(1).

N y( ) ΣtN0 P1

y yone g––

b1A-----------------------------–

P2

y yone g––

b2A-----------------------------

exp+exp=

Part 25: Airworthiness Standards: Transport Category Appendix G to Part 25

ASA 213

25

80 70 50 40 30 20 10

060

ALTITUDE, 1000 ft

.000

001

.000

01.0

001

.001

.1.0

11

10

P1

P2

P 1 A

ND

P2

FIG

UR

E 1

P 1

AN

D P

2 V

ALU

ES

Appendix G to Part 25 Federal Aviation Regulations

214 ASA

ALTITUDE, 1000 ft

FIG

UR

E 2

b 1

AN

D b

2 V

ALU

ES

b 1 A

ND

b2

80

70

60

50

40

30

20

10 0

01

23

45

67

89

10

11

12

b 1

b 2

Part 25: Airworthiness Standards: Transport Category Appendix H to Part 25

ASA 215

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APPENDIX H TO PART 25INSTRUCTIONS FOR CONTINUED

AIRWORTHINESS

H25.1 GENERAL.(a) This appendix specifies requirements for

preparation of Instructions for Continued Airwor-thiness as required by §§25.1529, 25.1729, andapplicable provisions of parts 21 and 26 of thischapter.

(b) The Instructions for Continued Airworthi-ness for each airplane must include the Instruc-tions for Continued Airworthiness for each engineand propeller (hereinafter designated “products”),for each appliance required by this chapter, andany required information relating to the interfaceof those appliances and products with the air-plane. If Instructions for Continued Airworthinessare not supplied by the manufacturer of an appli-ance or product installed in the airplane, the In-structions for Continued Airworthiness for the air-plane must include the information essential tothe continued airworthiness of the airplane.

(c) The applicant must submit to the FAA a pro-gram to show how changes to the Instructions forContinued Airworthiness made by the applicant orby the manufacturers or products and appliancesinstalled in the airplane will be distributed.

H25.2 FORMAT.(a) The Instructions for Continued Airworthiness

must be in the form of a manual or manuals as ap-propriate for the quantity of data to be provided.

(b) The format of the manual or manuals mustprovide for a practical arrangement.

H25.3 CONTENT.The contents of the manual or manuals must

be prepared in the English language. The Instruc-tions for Continued Airworthiness must containthe following manuals or sections, as appropriate,and information:

(a) Airplane maintenance manual or section.(1) Introduction information that includes an ex-

planation of the airplane’s features and data to theextent necessary for maintenance or preventivemaintenance.

(2) A description of the airplane and its sys-tems and installations including its engines, pro-pellers, and appliances.

(3) Basic control and operation information de-scribing how the airplane components and systemsare controlled and how they operate, including anyspecial procedures and limitations that apply.

(4) Servicing information that covers details re-garding servicing points, capacities of tanks, res-ervoirs, types of fluids to be used, pressures appli-cable to the various systems, location of access

panels for inspection and servicing, locations of lu-brication points, lubricants to be used, equipmentrequired for servicing, tow instructions and limita-tions, mooring, jacking, and leveling information.

(b) Maintenance instructions.(1) Scheduling information for each part of the

airplane and its engines, auxiliary power units,propellers, accessories, instruments, and equip-ment that provides the recommended periods atwhich they should be cleaned, inspected, ad-justed, tested, and lubricated, and the degree ofinspection, the applicable wear tolerances, andwork recommended at these periods. However,the applicant may refer to an accessory, instru-ment, or equipment manufacturer as the source ofthis information if the applicant shows that theitem has an exceptionally high degree of com-plexity requiring specialized maintenance tech-niques, test equipment, or expertise. The recom-mended overhaul periods and necessary crossreferences to the Airworthiness Limitations sec-tion of the manual must also be included. In addi-tion, the applicant must include an inspection pro-gram that includes the frequency and extent of theinspections necessary to provide for the contin-ued airworthiness of the airplane.

(2) Troubleshooting information describingprobable malfunctions, how to recognize thosemalfunctions, and the remedial action for thosemalfunctions.

(3) Information describing the order andmethod of removing and replacing products andparts with any necessary precautions to be taken.

(4) Other general procedural instructions in-cluding procedures for system testing duringground running, symmetry checks, weighing anddetermining the center of gravity, lifting and shor-ing, and storage limitations.

(c) Diagrams of structural access plates and in-formation needed to gain access for inspectionswhen access plates are not provided.

(d) Details for the application of special inspec-tion techniques including radiographic and ultra-sonic testing where such processes are specified.

(e) Information needed to apply protectivetreatments to the structure after inspection.

(f) All data relative to structural fasteners suchas identification, discard recommendations, andtorque values.

(g) A list of special tools needed.

Appendix H to Part 25 Federal Aviation Regulations

216 ASA

H25.4 AIRWORTHINESS LIMITATIONS SECTION.

(a) The Instructions for Continued Airworthi-ness must contain a section titled AirworthinessLimitations that is segregated and clearly distin-guishable from the rest of the document. This sec-tion must set forth—

(1) Each mandatory replacement time, struc-tural inspection interval, and related structural in-spection procedures approved under §25.571.

(2) Each mandatory replacement time, inspec-tion interval, related inspection procedure, and allcritical design configuration control limitations ap-proved under §25.981 for the fuel tank system.

(3) Any mandatory replacement time of EWIScomponents as defined in §25.1701.

(b) If the Instructions for Continued Airworthi-ness consist of multiple documents, the sectionrequired by this paragraph must be included in theprincipal manual. This section must contain a leg-ible statement in a prominent location that reads:“The Airworthiness Limitations section is FAA-ap-proved and specifies maintenance required under§§43.16 and 91.403 of the Federal Aviation Regu-lations, unless an alternative program has beenFAA approved.”

H25.5 ELECTRICAL WIRING INTERCONNECTION SYSTEM (EWIS)

INSTRUCTIONS FOR CONTINUED AIRWORTHINESS.

(a) The applicant must prepare Instructions forContinued Airworthiness (ICA) applicable toEWIS as defined by §25.1701 that are approvedby the FAA and include the following:

(1) Maintenance and inspection requirementsfor the EWIS developed with the use of an en-hanced zonal analysis procedure that includes:

(i) Identification of each zone of the airplane.(ii) Identification of each zone that contains

EWIS.(iii) Identification of each zone containing EWIS

that also contains combustible materials.(iv) Identification of each zone in which EWIS is

in close proximity to both primary and back-up hy-draulic, mechanical, or electrical flight controlsand lines.

(v) Identification of—(A) Tasks, and the intervals for performing

those tasks, that will reduce the likelihood of igni-tion sources and accumulation of combustiblematerial, and

(B) Procedures, and the intervals for perform-ing those procedures, that will effectively cleanthe EWIS components of combustible material ifthere is not an effective task to reduce the likeli-hood of combustible material accumulation.

(vi) Instructions for protections and caution in-formation that will minimize contamination and ac-cidental damage to EWIS, as applicable, duringperformance of maintenance, alteration, or re-pairs.

(2) Acceptable EWIS maintenance practices ina standard format.

(3) Wire separation requirements as deter-mined under §25.1707.

(4) Information explaining the EWIS identifica-tion method and requirements for identifying anychanges to EWIS under §25.1711.

(5) Electrical load data and instructions for up-dating that data.

(b) The EWIS ICA developed in accordancewith the requirements of H25.5(a)(1) must be inthe form of a document appropriate for the infor-mation to be provided, and they must be easilyrecognizable as EWIS ICA. This document musteither contain the required EWIS ICA or specifi-cally reference other portions of the ICA that con-tain this information.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–54, 45 FR 60177, Sept. 11, 1980;Amdt. 25–68, 54 FR 34329, Aug. 18, 1989; Amdt. 25–102, 66 FR 23130, May 7, 2001; Amdt. 25–123, 72 FR63408, Nov. 8, 2007]

Part 25: Airworthiness Standards: Transport Category Appendix I to Part 25

ASA 217

25

APPENDIX I TO PART 25INSTALLATION OF AN AUTOMATIC TAKEOFF

THRUST CONTROL SYSTEM (ATTCS)

I25.1 GENERAL.(a) This appendix specifies additional require-

ments for installation of an engine power controlsystem that automatically resets thrust or poweron operating engine(s) in the event of any one en-gine failure during takeoff.

(b) With the ATTCS and associated systemsfunctioning normally as designed, all applicablerequirements of Part 25, except as provided inthis appendix, must be met without requiring anyaction by the crew to increase thrust or power.

I25.2 DEFINITIONS.(a) Automatic Takeoff Thrust Control System

(ATTCS). An ATTCS is defined as the entire auto-matic system used on takeoff, including all de-vices, both mechanical and electrical, that senseengine failure, transmit signals, actuate fuel con-trols or power levers or increase engine power byother means on operating engines to achievescheduled thrust or power increases, and furnishcockpit information on system operation.

(b) Critical Time Interval. When conducting anATTCS takeoff, the critical time interval is be-tween V1 minus 1 second and a point on the min-imum performance, all-engine flight path where,assuming a simultaneous occurrence of an en-gine and ATTCS failure, the resulting minimumflight path thereafter intersects the Part 25 re-quired actual flight path at no less than 400 feetabove the takeoff surface. This time interval isshown in the following illustration:

Appendix I to Part 25 Federal Aviation Regulations

218 ASA

1 sec

V1

Critical time interval

Heightaboverunwaysurface(ft.)

All E

ngin

e Fl

ight

Pat

h

Actua

l flt.

path

(25.

115b

)

one-

engi

ne in

oper

ative

w/

ATTC

ope

ratin

g

400'

Engine andATTC failure

Flight path with ATTCSand engine failure

Part 25: Airworthiness Standards: Transport Category Appendix I to Part 25

ASA 219

25

I25.3 PERFORMANCE AND SYSTEM RELIABILITY REQUIREMENTS.

The applicant must comply with the perfor-mance and ATTCS reliability requirements as fol-lows:

(a) An ATTCS failure or a combination of fail-ures in the ATTCS during the critical time interval:

(1) Shall not prevent the insertion of the maxi-mum approved takeoff thrust or power, or must beshown to be an improbable event.

(2) Shall not result in a significant loss or reduc-tion in thrust or power, or must be shown to be anextremely improbable event.

(b) The concurrent existence of an ATTCS fail-ure and an engine failure during the critical time in-terval must be shown to be extremely improbable.

(c) All applicable performance requirements ofPart 25 must be met with an engine failure occur-ring at the most critical point during takeoff withthe ATTCS system functioning.

I25.4 THRUST SETTING.The initial takeoff thrust or power setting on

each engine at the beginning of the takeoff rollmay not be less than any of the following:

(a) Ninety (90) percent of the thrust or powerset by the ATTCS (the maximum takeoff thrust orpower approved for the airplane under existingambient conditions);

(b) That required to permit normal operation ofall safety-related systems and equipment depen-dent upon engine thrust or power lever position; or

(c) That shown to be free of hazardous engineresponse characteristics when thrust or power isadvanced from the initial takeoff thrust or power tothe maximum approved takeoff thrust or power.

I25.5 POWERPLANT CONTROLS.(a) In addition to the requirements of §25.1141,

no single failure or malfunction, or probable combi-nation thereof, of the ATTCS, including associatedsystems, may cause the failure of any powerplantfunction necessary for safety.

(b) The ATTCS must be designed to:(1) Apply thrust or power on the operating en-

gine(s), following any one engine failure duringtakeoff, to achieve the maximum approved takeoffthrust or power without exceeding engine operat-ing limits;

(2) Permit manual decrease or increase inthrust or power up to the maximum takeoff thrustor power approved for the airplane under existingconditions through the use of the power lever. Forairplanes equipped with limiters that automaticallyprevent engine operating limits from being ex-ceeded under existing ambient conditions, othermeans may be used to increase the thrust orpower in the event of an ATTCS failure provided

the means is located on or forward of the powerlevers; is easily identified and operated under alloperating conditions by a single action of eitherpilot with the hand that is normally used to actuatethe power levers; and meets the requirements of§25.777 (a), (b), and (c);

(3) Provide a means to verify to the flightcrewbefore takeoff that the ATTCS is in a condition tooperate; and

(4) Provide a means for the flightcrew to deacti-vate the automatic function. This means must bedesigned to prevent inadvertent deactivation.

I25.6 POWERPLANT INSTRUMENTS.In addition to the requirements of §25.1305:(a) A means must be provided to indicate when

the ATTCS is in the armed or ready condition; and(b) If the inherent flight characteristics of the

airplane do not provide adequate warning that anengine has failed, a warning system that is inde-pendent of the ATTCS must be provided to givethe pilot a clear warning of any engine failure dur-ing takeoff.


Appendix J to Part 25 Federal Aviation Regulations

220 ASA

APPENDIX J TO PART 25EMERGENCY DEMONSTRATION

The following test criteria and procedures mustbe used for showing compliance with §25.803:

(a) The emergency evacuation must be con-ducted with exterior ambient light levels of nogreater than 0.3 foot-candles prior to the activa-tion of the airplane emergency lighting system.The source(s) of the initial exterior ambient lightlevel may remain active or illuminated during theactual demonstration. There must, however, be noincrease in the exterior ambient light level exceptfor that due to activation of the airplane emer-gency lighting system.

(b) The airplane must be in a normal attitudewith landing gear extended.

(c) Unless the airplane is equipped with an off-wing descent means, stands or ramps may beused for descent from the wing to the ground.Safety equipment such as mats or inverted liferafts may be placed on the floor or ground to pro-tect participants. No other equipment that is notpart of the emergency evacuation equipment ofthe airplane may be used to aid the participants inreaching the ground.

(d) Except as provided in paragraph (a) of thisAppendix, only the airplane’s emergency lightingsystem may provide illumination.

(e) All emergency equipment required for theplanned operation of the airplane must be in-stalled.

(f) Each internal door or curtain must be in thetakeoff configuration.

(g) Each crewmember must be seated in thenormally assigned seat for takeoff and must re-main in the seat until receiving the signal for com-mencement of the demonstration. Each crew-member must be a person having knowledge ofthe operation of exits and emergency equipmentand, if compliance with §121.291 is also beingdemonstrated, each flight attendant must be amember of a regularly scheduled line crew.

(h) A representative passenger load of personsin normal health must be used as follows:

(1) At least 40 percent of the passenger loadmust be female.

(2) At least 35 percent of the passenger loadmust be over 50 years of age.

(3) At least 15 percent of the passenger loadmust be female and over 50 years of age.

(4) Three life-size dolls, not included as part ofthe total passenger load, must be carried by pas-sengers to simulate live infants 2 years old oryounger.

(5) Crewmembers, mechanics, and trainingpersonnel, who maintain or operate the airplanein the normal course of their duties, may not beused as passengers.

(i) No passenger may be assigned a specificseat except as the Administrator may require. Ex-cept as required by subparagraph (g) of this para-graph, no employee of the applicant may beseated next to an emergency exit.

(j) Seat belts and shoulder harnesses (as re-quired) must be fastened.

(k) Before the start of the demonstration, ap-proximately one-half of the total average amountof carry-on baggage, blankets, pillows, and othersimilar articles must be distributed at several loca-tions in aisles and emergency exit access ways tocreate minor obstructions.

(l) No prior indication may be given to anycrewmember or passenger of the particular exitsto be used in the demonstration.

(m) The applicant may not practice, rehearse,or describe the demonstration for the participantsnor may any participant have taken part in thistype of demonstration within the preceding 6months.

(n) Prior to entering the demonstration aircraft,the passengers may also be advised to follow di-rections of crewmembers but may not be in-structed on the procedures to be followed in thedemonstration, except with respect to safety pro-cedures in place for the demonstration or whichhave to do with the demonstration site. Prior tothe start of the demonstration, the pre-takeoffpassenger briefing required by §121.571 may begiven. Flight attendants may assign demonstra-tion subjects to assist persons from the bottom ofa slide, consistent with their approved trainingprogram.

(o) The airplane must be configured to preventdisclosure of the active emergency exits to dem-onstration participants in the airplane until thestart of the demonstration.

(p) Exits used in the demonstration must con-sist of one exit from each exit pair. The demon-stration may be conducted with the escape slides,if provided, inflated and the exits open at the be-ginning of the demonstration. In this case, all exitsmust be configured such that the active exits arenot disclosed to the occupants. If this method isused, the exit preparation time for each exit uti-lized must be accounted for, and exits that are notto be used in the demonstration must not be indi-cated before the demonstration has started. Theexits to be used must be representative of all ofthe emergency exits on the airplane and must bedesignated by the applicant, subject to approvalby the Administrator. At least one floor level exitmust be used.

Part 25: Airworthiness Standards: Transport Category Appendix J to Part 25

ASA 221

25

(q) Except as provided in paragraph (c) of thissection, all evacuees must leave the airplane by ameans provided as part of the airplane’s equip-ment.

(r) The applicant’s approved procedures mustbe fully utilized, except the flightcrew must take noactive role in assisting others inside the cabin dur-ing the demonstration.

(s) The evacuation time period is completedwhen the last occupant has evacuated the air-plane and is on the ground. Provided that the ac-ceptance rate of the stand or ramp is no greaterthan the acceptance rate of the means availableon the airplane for descent from the wing duringan actual crash situation, evacuees using standsor ramps allowed by paragraph (c) of this Appen-dix are considered to be on the ground when theyare on the stand or ramp.

[Docket No. 5066, 29 FR 18291, Dec. 24, 1964; asamended by Amdt. 25–72, 55 FR 29788, July 20, 1990;Amdt. 25–79, Aug. 26, 1993; Amdt. 25–117, 69 FR67499, Nov. 17, 2004]

Appendix K to Part 25 Federal Aviation Regulations

222 ASA

APPENDIX K TO PART 25EXTENDED OPERATIONS (ETOPS)

Source: Docket No. FAA–2002–6717, 72 FR 1873, Jan.16, 2007, unless otherwise noted.

This appendix specifies airworthiness require-ments for the approval of an airplane-engine com-bination for extended operations (ETOPS). Fortwo-engine airplanes, the applicant must complywith sections K25.1 and K25.2 of this appendix.For airplanes with more than two engines, the ap-plicant must comply with sections K25.1 andK25.3 of this appendix.

K25.1 DESIGN REQUIREMENTS.K25.1.1 Part 25 compliance.

The airplane-engine combination must complywith the requirements of part 25 considering themaximum flight time and the longest diversiontime for which the applicant seeks approval.K25.1.2 Human factors.

An applicant must consider crew workload, op-erational implications, and the crew’s and passen-gers’ physiological needs during continued opera-tion with failure effects for the longest diversiontime for which it seeks approval.

K25.1.3 Airplane systems.(a) Operation in icing conditions.(1) The airplane must be certificated for opera-

tion in icing conditions in accordance with§25.1419.

(2) The airplane must be able to safely conductan ETOPS diversion with the most critical ice ac-cretion resulting from:

(i) Icing conditions encountered at an altitudethat the airplane would have to fly following an en-gine failure or cabin decompression.

(ii) A 15-minute hold in the continuous maxi-mum icing conditions specified in Appendix C ofthis part with a liquid water content factor of 1.0.

(iii) Ice accumulated during approach and land-ing in the icing conditions specified in Appendix Cof this part.

(b) Electrical power supply. The airplane mustbe equipped with at least three independentsources of electrical power.

(c) Time limited systems. The applicant mustdefine the system time capability of each ETOPSsignificant system that is time-limited.K25.1.4 Propulsion systems.

(a) Fuel system design. Fuel necessary tocomplete an ETOPS flight (including a diversionfor the longest time for which the applicant seeksapproval) must be available to the operating en-gines at the pressure and fuel-flow required by§25.955 under any airplane failure condition notshown to be extremely improbable. Types of fail-ures that must be considered include, but are notlimited to: crossfeed valve failures, automatic fuel

management system failures, and normal electri-cal power generation failures.

(1) If the engine has been certified for limitedoperation with negative engine-fuel-pump-inletpressures, the following requirements apply:

(i) Airplane demonstration-testing must coverworst case cruise and diversion conditions involv-ing:

(A) Fuel grade and temperature.(B) Thrust or power variations.(C) Turbulence and negative G.(D) Fuel system components degraded within

their approved maintenance limits.(ii) Unusable-fuel quantity in the suction feed

configuration must be determined in accordancewith §25.959.

(2) For two-engine airplanes to be certificatedfor ETOPS beyond 180 minutes, one fuel boostpump in each main tank and at least one cross-feed valve, or other means for transferring fuel,must be powered by an independent electricalpower source other than the three power sourcesrequired to comply with section K25.1.3(b) of thisappendix. This requirement does not apply if thenormal fuel boost pressure, crossfeed valve actu-ation, or fuel transfer capability is not provided byelectrical power.

(3) An alert must be displayed to the flightcrewwhen the quantity of fuel available to the enginesfalls below the level required to fly to the destina-tion. The alert must be given when there isenough fuel remaining to safely complete a diver-sion. This alert must account for abnormal fuelmanagement or transfer between tanks, and pos-sible loss of fuel. This paragraph does not apply toairplanes with a required flight engineer.

(b) APU design. If an APU is needed to complywith this appendix, the applicant must demon-strate that:

(1) The reliability of the APU is adequate tomeet those requirements; and

(2) If it is necessary that the APU be able tostart in flight, it is able to start at any altitude up tothe maximum operating altitude of the airplane, or45,000 feet, whichever is lower, and run for the re-mainder of any flight.

(c) Engine oil tank design. The engine oil tankfiller cap must comply with §33.71(c)(4) of thischapter.K25.1.5 Engine-condition monitoring.

Procedures for engine-condition monitoringmust be specified and validated in accordancewith Part 33, Appendix A, paragraph A33.3(c) ofthis chapter.K25.1.6 Configuration, maintenance, and pro-cedures.

The applicant must list any configuration, oper-ating and maintenance requirements, hardwarelife limits, MMEL constraints, and ETOPS ap-proval in a CMP document.

Part 25: Airworthiness Standards: Transport Category Appendix K to Part 25

ASA 223

25

K25.1.7 Airplane flight manual.The airplane flight manual must contain the fol-

lowing information applicable to the ETOPS typedesign approval:

(a) Special limitations, including any limitationassociated with operation of the airplane up to themaximum diversion time being approved.

(b) Required markings or placards.(c) The airborne equipment required for ex-

tended operations and flightcrew operating proce-dures for this equipment.

(d) The system time capability for the following:(1) The most limiting fire suppression system

for Class C cargo or baggage compartments.(2) The most limiting ETOPS significant system

other than fire suppression systems for Class Ccargo or baggage compartments.

(e) This statement: “The type-design reliabilityand performance of this airplane-engine combi-nation has been evaluated under 14 CFR 25.1535and found suitable for (identify maximum ap-proved diversion time) extended operations(ETOPS) when the configuration, maintenance,and procedures standard contained in (identifythe CMP document) are met. The actual maxi-mum approved diversion time for this airplanemay be less based on its most limiting systemtime capability. This finding does not constituteoperational approval to conduct ETOPS.’’

K25.2. TWO-ENGINE AIRPLANES.An applicant for ETOPS type design approval

of a two-engine airplane must use one of themethods described in section K25.2.1, K25.2.2,or K25.2.3 of this appendix.K25.2.1 Service experience method.

An applicant for ETOPS type design approvalusing the service experience method must com-ply with sections K25.2.1(a) and K25.2.1(b) of thisappendix before conducting the assessmentsspecified in sections K25.2.1(c) and K25.2.1(d) ofthis appendix, and the flight test specified in sec-tion K25.2.1(e) of this appendix.

(a) Service experience. The world fleet for theairplane-engine combination must accumulate aminimum of 250,000 engine-hours. The FAA mayreduce this number of hours if the applicant iden-tifies compensating factors that are acceptable tothe FAA. The compensating factors may includeexperience on another airplane, but experienceon the candidate airplane must make up a signifi-cant portion of the total service experience.

(b) In-flight shutdown (IFSD) rates. The demon-strated 12-month rolling average IFSD rate for theworld fleet of the airplane-engine combinationmust be commensurate with the level of ETOPSapproval being sought.

(1) For type design approval up to and includ-ing 120 minutes: An IFSD rate of 0.05 or less per

1,000 world-fleet engine-hours, unless otherwiseapproved by the FAA. Unless the IFSD rate is0.02 or less per 1,000 world-fleet engine-hours,the applicant must provide a list of corrective ac-tions in the CMP document specified in sectionK25.1.6 of this appendix, that, when taken, wouldresult in an IFSD rate of 0.02 or less per 1,000fleet engine-hours.

(2) For type design approval up to and includ-ing 180 minutes: An IFSD rate of 0.02 or less per1,000 world-fleet engine-hours, unless otherwiseapproved by the FAA. If the airplane-engine com-bination does not meet this rate by compliancewith an existing 120-minute CMP document, thennew or additional CMP requirements that the ap-plicant has demonstrated would achieve this IFSDrate must be added to the CMP document.

(3) For type design approval beyond 180 min-utes: An IFSD rate of 0.01 or less per 1,000 fleetengine-hours unless otherwise approved by theFAA. If the airplane-engine combination does notmeet this rate by compliance with an existing 120-minute or 180-minute CMP document, then newor additional CMP requirements that the applicanthas demonstrated would achieve this IFSD ratemust be added to the CMP document.

(c) Propulsion system assessment.(1) The applicant must conduct a propulsion

system assessment based on the following datacollected from the world-fleet of the airplane-en-gine combination:

(i) A list of all IFSDs, unplanned ground engineshutdowns, and occurrences (both ground and in-flight) when an engine was not shut down, but en-gine control or the desired thrust or power levelwas not achieved, including engine flameouts.Planned IFSDs performed during flight trainingneed not be included. For each item, the applicantmust provide—

(A) Each airplane and engine make, model,and serial number;

(B) Engine configuration, and major alterationhistory;

(C) Engine position;(D) Circumstances leading up to the engine

shutdown or occurrence;(E) Phase of flight or ground operation;(F) Weather and other environmental condi-

tions; and(G) Cause of engine shutdown or occurrence.(ii) A history of unscheduled engine removal

rates since introduction into service (using 6- and12-month rolling averages), with a summary ofthe major causes for the removals.

(iii) A list of all propulsion system events(whether or not caused by maintenance or flight-crew error), including dispatch delays, cancella-tions, aborted takeoffs, turnbacks, diversions, andflights that continue to destination after the event.


224 ASA

(iv) The total number of engine hours and cy-cles, the number of hours for the engine with thehighest number of hours, the number of cycles forthe engine with the highest number of cycles, andthe distribution of hours and cycles.

(v) The mean time between failures (MTBF) ofpropulsion system components that affect reliabil-ity.

(vi) A history of the IFSD rates since introduc-tion into service using a 12-month rolling average.

(2) The cause or potential cause of each itemlisted in K25.2.1(c)(1)(i) must have a correctiveaction or actions that are shown to be effective inpreventing future occurrences. Each correctiveaction must be identified in the CMP documentspecified in section K25.1.6. A corrective action isnot required:

(i) For an item where the manufacturer is un-able to determine a cause or potential cause.

(ii) For an event where it is technically unfeasi-ble to develop a corrective action.

(iii) If the world-fleet IFSD rate—(A) Is at or below 0.02 per 1,000 world-fleet en-

gine-hours for approval up to and including 180-minute ETOPS; or

(B) Is at or below 0.01 per 1,000 world-fleet en-gine-hours for approval greater than 180-minuteETOPS.

(d) Airplane systems assessment. The appli-cant must conduct an airplane systems assess-ment. The applicant must show that the airplanesystems comply with §25.1309(b) using availablein-service reliability data for ETOPS significantsystems on the candidate airplane-engine combi-nation. Each cause or potential cause of a rele-vant design, manufacturing, operational, andmaintenance problem occurring in service musthave a corrective action or actions that are shownto be effective in preventing future occurrences.Each corrective action must be identified in theCMP document specified in section K25.1.6 ofthis appendix. A corrective action is not required ifthe problem would not significantly impact thesafety or reliability of the airplane system in-volved. A relevant problem is a problem with anETOPS group 1 significant system that has orcould result in, an IFSD or diversion. The appli-cant must include in this assessment relevantproblems with similar or identical equipment in-stalled on other types of airplanes to the extentsuch information is reasonably available.

(e) Airplane flight test. The applicant must con-duct a flight test to validate the flightcrew’s abilityto safely conduct an ETOPS diversion with an in-operative engine and worst-case ETOPS Signifi-cant System failures and malfunctions that couldoccur in service. The flight test must validate theairplane’s flying qualities and performance withthe demonstrated failures and malfunctions.

K25.2.2 Early ETOPS method.An applicant for ETOPS type design approval

using the Early ETOPS method must comply withthe following requirements:

(a) Assessment of relevant experience with air-planes previously certificated under part 25. Theapplicant must identify specific corrective actionstaken on the candidate airplane to prevent rele-vant design, manufacturing, operational, andmaintenance problems experienced on airplanespreviously certificated under part 25 manufac-tured by the applicant. Specific corrective actionsare not required if the nature of a problem is suchthat the problem would not significantly impact thesafety or reliability of the airplane system in-volved. A relevant problem is a problem with anETOPS group 1 significant system that has orcould result in an IFSD or diversion. The applicantmust include in this assessment relevant prob-lems of supplier-provided ETOPS group 1 signifi-cant systems and similar or identical equipmentused on airplanes built by other manufacturers tothe extent such information is reasonably avail-able.

(b) Propulsion system design.(1) The engine used in the applicant’s airplane

design must be approved as eligible for EarlyETOPS in accordance with §33.201 of this chap-ter.

(2) The applicant must design the propulsionsystem to preclude failures or malfunctions thatcould result in an IFSD. The applicant must showcompliance with this requirement by analysis,test, in-service experience on other airplanes, orother means acceptable to the FAA. If analysis isused, the applicant must show that the propulsionsystem design will minimize failures and malfunc-tions with the objective of achieving the followingIFSD rates:

(i) An IFSD rate of 0.02 or less per 1,000 world-fleet engine-hours for type design approval up toand including 180 minutes.

(ii) An IFSD rate of 0.01 or less per 1,000world-fleet engine-hours for type design approvalbeyond 180 minutes.

(c) Maintenance and operational procedures.The applicant must validate all maintenance andoperational procedures for ETOPS significantsystems. The applicant must identify, track, andresolve any problems found during the validationin accordance with the problem tracking and reso-lution system specified in section K25.2.2(h) ofthis appendix.

(d) Propulsion system validation test.(1) The installed engine configuration for which

approval is being sought must comply with§33.201(c) of this chapter. The test engine mustbe configured with a complete airplane nacellepackage, including engine-mounted equipment,except for any configuration differences neces-


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sary to accommodate test stand interfaces withthe engine nacelle package. At the conclusion ofthe test, the propulsion system must be—

(i) Visually inspected according to the appli-cant’s on-wing inspection recommendations andlimits; and

(ii) Completely disassembled and the propul-sion system hardware inspected to determinewhether it meets the service limits specified in theInstructions for Continued Airworthiness submit-ted in compliance with §25.1529.

(2) The applicant must identify, track, and re-solve each cause or potential cause of IFSD, lossof thrust control, or other power loss encounteredduring this inspection in accordance with theproblem tracking and resolution system specifiedin section K25.2.2 (h) of this appendix.

(e) New technology testing. Technology new tothe applicant, including substantially new manu-facturing techniques, must be tested to substanti-ate its suitability for the airplane design.

(f) APU validation test. If an APU is needed tocomply with this appendix, one APU of the type tobe certified with the airplane must be tested for3,000 equivalent airplane operational cycles. Fol-lowing completion of the test, the APU must bedisassembled and inspected. The applicant mustidentify, track, and resolve each cause or potentialcause of an inability to start or operate the APU inflight as intended in accordance with the problemtracking and resolution system specified in sec-tion K25.2.2(h) of this appendix.

(g) Airplane demonstration. For each airplane-engine combination to be approved for ETOPS,the applicant must flight test at least one airplaneto demonstrate that the airplane, and its compo-nents and equipment are capable of functioningproperly during ETOPS flights and diversions ofthe longest duration for which the applicant seeksapproval. This flight testing may be performed inconjunction with, but may not substitute for theflight testing required by §21.35(b)(2) of this chap-ter.

(1) The airplane demonstration flight test pro-gram must include:

(i) Flights simulating actual ETOPS, includingflight at normal cruise altitude, step climbs, and, ifapplicable, APU operation.

(ii) Maximum duration flights with maximum du-ration diversions.

(iii) Maximum duration engine-inoperative di-versions distributed among the engines installedon the airplanes used for the airplane demonstra-tion flight test program. At least two one-engine-inoperative diversions must be conducted at max-imum continuous thrust or power using the sameengine.

(iv) Flights under non-normal conditions todemonstrate the flightcrew’s ability to safely con-duct an ETOPS diversion with worst-case ETOPS

significant system failures or malfunctions thatcould occur in service.

(v) Diversions to airports that represent air-ports of the types used for ETOPS diversions.

(vi) Repeated exposure to humid and inclementweather on the ground followed by a long-durationflight at normal cruise altitude.

(2) The airplane demonstration flight test pro-gram must validate the adequacy of the airplane’sflying qualities and performance, and the flight-crew’s ability to safely conduct an ETOPS diver-sion under the conditions specified in sectionK25.2.2(g)(1) of this appendix.

(3) During the airplane demonstration flight testprogram, each test airplane must be operatedand maintained using the applicant’s recom-mended operating and maintenance procedures.

(4) At the completion of the airplane demon-stration flight test program, each ETOPS signifi-cant system must undergo an on-wing inspectionor test in accordance with the tasks defined in theproposed Instructions for Continued Airworthi-ness to establish its condition for continued safeoperation. Each engine must also undergo a gaspath inspection. These inspections must be con-ducted in a manner to identify abnormal condi-tions that could result in an IFSD or diversion. Theapplicant must identify, track and resolve any ab-normal conditions in accordance with the problemtracking and resolution system specified in sec-tion K25.2.2(h) of this appendix.

(h) Problem tracking and resolution system.(1) The applicant must establish and maintain a

problem tracking and resolution system. The sys-tem must:

(i) Contain a process for prompt reporting tothe responsible FAA aircraft certification office ofeach occurrence reportable under §21.4(a)(6) en-countered during the phases of airplane and en-gine development used to assess Early ETOPSeligibility.

(ii) Contain a process for notifying the responsi-ble FAA aircraft certification office of each pro-posed corrective action that the applicant deter-mines necessary for each problem identified fromthe occurrences reported under section K25.2.2.(h)(1)(i) of this appendix. The timing of the notifi-cation must permit appropriate FAA review beforetaking the proposed corrective action.

(2) If the applicant is seeking ETOPS type de-sign approval of a change to an airplane-enginecombination previously approved for ETOPS, theproblem tracking and resolution system need onlyaddress those problems specified in the followingtable, provided the applicant obtains prior authori-zation from the FAA:


226 ASA

(i) Acceptance criteria. The type and frequencyof failures and malfunctions on ETOPS significantsystems that occur during the airplane flight testprogram and the airplane demonstration flight testprogram specified in section K25.2.2(g) of this ap-pendix must be consistent with the type and fre-quency of failures and malfunctions that would beexpected to occur on currently certificated air-planes approved for ETOPS.K25.2.3. Combined service experience andEarly ETOPS method.

An applicant for ETOPS type design approvalusing the combined service experience and EarlyETOPS method must comply with the followingrequirements.

(a) A service experience requirement of notless than 15,000 engine-hours for the world fleetof the candidate airplane-engine combination.

(b) The Early ETOPS requirements of K25.2.2,except for the airplane demonstration specified insection K25.2.2(g) of this appendix; and

(c) The flight test requirement of sectionK25.2.1(e) of this appendix.

K25.3. AIRPLANES WITH MORE THAN TWO ENGINES.

An applicant for ETOPS type design approvalof an airplane with more than two engines mustuse one of the methods described in sectionK25.3.1, K25.3.2, or K25.3.3 of this appendix.K25.3.1 Service experience method.

An applicant for ETOPS type design approvalusing the service experience method must com-ply with section K25.3.1(a) of this appendix beforeconducting the airplane systems assessmentspecified in K25.3.1(b), and the flight test speci-fied in section K25.3.1(c) of this appendix.

(a) Service experience. The world fleet for theairplane-engine combination must accumulate aminimum of 250,000 engine-hours. The FAA mayreduce this number of hours if the applicant iden-tifies compensating factors that are acceptable tothe FAA. The compensating factors may includeexperience on another airplane, but experienceon the candidate airplane must make up a signifi-

cant portion of the total required service experi-ence.

(b) Airplane systems assessment. The appli-cant must conduct an airplane systems assess-ment. The applicant must show that the airplanesystems comply with the §25.1309(b) using avail-able in-service reliability data for ETOPS signifi-cant systems on the candidate airplane-enginecombination. Each cause or potential cause of arelevant design, manufacturing, operational ormaintenance problem occurring in service musthave a corrective action or actions that are shownto be effective in preventing future occurrences.Each corrective action must be identified in theCMP document specified in section K25.1.6 ofthis appendix. A corrective action is not required ifthe problem would not significantly impact thesafety or reliability of the airplane system in-volved. A relevant problem is a problem with anETOPS group 1 significant system that has orcould result in an IFSD or diversion. The applicantmust include in this assessment relevant prob-lems with similar or identical equipment installedon other types of airplanes to the extent such in-formation is reasonably available.

(c) Airplane flight test. The applicant must con-duct a flight test to validate the flightcrew’s abilityto safely conduct an ETOPS diversion with an in-operative engine and worst-case ETOPS signifi-cant system failures and malfunctions that couldoccur in service. The flight test must validate theairplane’s flying qualities and performance withthe demonstrated failures and malfunctions.K25.3.2 Early ETOPS method.

An applicant for ETOPS type design approvalusing the Early ETOPS method must comply withthe following requirements:

(a) Maintenance and operational procedures.The applicant must validate all maintenance andoperational procedures for ETOPS significantsystems. The applicant must identify, track and re-solve any problems found during the validation inaccordance with the problem tracking and resolu-tion system specified in section K25.3.2(e) of thisappendix.

(b) New technology testing. Technology new tothe applicant, including substantially new manu-facturing techniques, must be tested to substanti-ate its suitability for the airplane design.

(c) APU validation test. If an APU is needed tocomply with this appendix, one APU of the type tobe certified with the airplane must be tested for3,000 equivalent airplane operational cycles. Fol-lowing completion of the test, the APU must bedisassembled and inspected. The applicant mustidentify, track, and resolve each cause or potentialcause of an inability to start or operate the APU inflight as intended in accordance with the problemtracking and resolution system specified in sec-tion K25.3.2(e) of this appendix.

If the change does not require a new airplane type certificate and…

Then the Problem Tracking and Resolution System must address…

(i) Requires a new engine type certificate.

All problems applicable to the new engine installation, and for the remainder of the airplane, problems in changed systems only.

(ii) Does not require a new engine type certificate.

Problems in changed systems only.


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(d) Airplane demonstration. For each airplane-engine combination to be approved for ETOPS,the applicant must flight test at least one airplaneto demonstrate that the airplane, and its compo-nents and equipment are capable of functioningproperly during ETOPS flights and diversions ofthe longest duration for which the applicant seeksapproval. This flight testing may be performed inconjunction with, but may not substitute for theflight testing required by §21.35(b)(2).

(1) The airplane demonstration flight test pro-gram must include:

(i) Flights simulating actual ETOPS includingflight at normal cruise altitude, step climbs, and, ifapplicable, APU operation.

(ii) Maximum duration flights with maximum du-ration diversions.

(iii) Maximum duration engine-inoperative di-versions distributed among the engines installedon the airplanes used for the airplane demonstra-tion flight test program. At least two one engine-inoperative diversions must be conducted at max-imum continuous thrust or power using the sameengine.

(iv) Flights under non-normal conditions to vali-date the flightcrew’s ability to safely conduct anETOPS diversion with worst-case ETOPS signifi-cant system failures or malfunctions that could oc-cur in service.

(v) Diversions to airports that represent air-ports of the types used for ETOPS diversions.

(vi) Repeated exposure to humid and inclementweather on the ground followed by a long durationflight at normal cruise altitude.

(2) The airplane demonstration flight test pro-gram must validate the adequacy of the airplane’sflying qualities and performance, and the flight-crew’s ability to safely conduct an ETOPS diver-sion under the conditions specified in sectionK25.3.2(d)(1) of this appendix.

(3) During the airplane demonstration flight testprogram, each test airplane must be operatedand maintained using the applicant’s recom-mended operating and maintenance procedures.

(4) At the completion of the airplane demon-stration, each ETOPS significant system must un-dergo an on-wing inspection or test in accordancewith the tasks defined in the proposed Instruc-tions for Continued Airworthiness to establish itscondition for continued safe operation. Each en-gine must also undergo a gas path inspection.These inspections must be conducted in a man-ner to identify abnormal conditions that could re-sult in an IFSD or diversion. The applicant mustidentify, track and resolve any abnormal condi-tions in accordance with the problem tracking andresolution system specified in section K25.3.2(e)of this appendix.

(e) Problem tracking and resolution system.

(1) The applicant must establish and maintain aproblem tracking and resolution system. The sys-tem must:

(i) Contain a process for prompt reporting tothe responsible FAA aircraft certification office ofeach occurrence reportable under §21.4(a)(6) en-countered during the phases of airplane and en-gine development used to assess Early ETOPSeligibility.

(ii) Contain a process for notifying the responsi-ble FAA aircraft certification office of each pro-posed corrective action that the applicant deter-mines necessary for each problem identified fromthe occurrences reported under sectionK25.3.2(h)(1)(i) of this appendix. The timing of thenotification must permit appropriate FAA reviewbefore taking the proposed corrective action.

(2) If the applicant is seeking ETOPS type de-sign approval of a change to an airplane-enginecombination previously approved for ETOPS, theproblem tracking and resolution system need onlyaddress those problems specified in the followingtable, provided the applicant obtains prior authori-zation from the FAA:

(f) Acceptance criteria. The type and frequencyof failures and malfunctions on ETOPS significantsystems that occur during the airplane flight testprogram and the airplane demonstration flight testprogram specified in section K25.3.2(d) of this ap-pendix must be consistent with the type and fre-quency of failures and malfunctions that would beexpected to occur on currently certificated air-planes approved for ETOPS.K25.3.3 Combined service experience andEarly ETOPS method.

An applicant for ETOPS type design approvalusing the Early ETOPS method must comply withthe following requirements:

(a) A service experience requirement of lessthan 15,000 engine-hours for the world fleet of thecandidate airplane-engine combination;

(b) The Early ETOPS requirements of sectionK25.3.2 of this appendix, except for the airplanedemonstration specified in section K25.3.2(d) ofthis appendix; and

(c) The flight test requirement of sectionK25.3.1(c) of this appendix.

If the change does not require a new airplane type certificate and…

Then the Problem Tracking and Resolution System must address…

(i) Requires a new engine type certificate.

All problems applicable to the new engine installation, and for the remainder of the airplane, problems in changed systems only.

(ii) Does not require a new engine type certificate.

Problems in changed systems only.

Appendix L to Part 25 Federal Aviation Regulations

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APPENDIX L TO PART 25HIRF ENVIRONMENTS AND EQUIPMENT HIRF

TEST LEVELS

This appendix specifies the HIRF environmentsand equipment HIRF test levels for electrical andelectronic systems under §25.1317. The fieldstrength values for the HIRF environments andequipment HIRF test levels are expressed in root-mean-square units measured during the peak ofthe modulation cycle.

(a) HIRF environment I is specified in the fol-lowing table:

In this table, the higher field strength applies atthe frequency band edges.

(b) HIRF environment II is specified in the fol-lowing table:

In this table, the higher field strength applies atthe frequency band edges.

Table I.— HIRF Environment I

FrequencyField strength (volts / meter)

Peak Average

10 kHz–2 MHz 50 50

2 MHz–30 MHz 100 100

30 MHz–100 MHz 50 50

100 MHz–400 MHz 100 100

400 MHz–700 MHz 700 50

700 MHz–1 GHz 700 100

1 GHz–2 GHz 2,000 200

2 GHz–6 GHz 3,000 200

6 GHz–8 GHz 1,000 200

8 GHz–12 GHz 3,000 300

12 GHz–18 GHz 2,000 200

18 GHz–40 GHz 600 200

Table II.— HIRF Environment II

FrequencyField strength (volts / meter)

Peak Average

10 kHz–500 kHz 20 20

500 kHz–2 MHz 30 30

2 MHz–30 MHz 100 100

30 MHz–100 MHz 10 10

100 MHz–200 MHz 30 10

200 MHz–400 MHz 10 10

400 MHz–1 GHz 700 40

1 GHz–2 GHz 1,300 160

2 GHz–4 GHz 3,000 120

4 GHz–6 GHz 3,000 160

6 GHz–8 GHz 400 170

8 GHz–12 GHz 1,230 230

12 GHz–18 GHz 730 190

18 GHz–40 GHz 600 150

Part 25: Airworthiness Standards: Transport Category Appendix L to Part 25

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(c) Equipment HIRF Test Level 1.(1) From 10 kilohertz (kHz) to 400 megahertz

(MHz), use conducted susceptibility tests withcontinuous wave (CW) and 1 kHz square wavemodulation with 90 percent depth or greater. Theconducted susceptibility current must start at aminimum of 0.6 milliamperes (mA) at 10 kHz, in-creasing 20 decibels (dB) per frequency decadeto a minimum of 30 mA at 500 kHz.

(2) From 500 kHz to 40 MHz, the conductedsusceptibility current must be at least 30 mA.

(3) From 40 MHz to 400 MHz, use conductedsusceptibility tests, starting at a minimum of 30mA at 40 MHz, decreasing 20 dB per frequencydecade to a minimum of 3 mA at 400 MHz.

(4) From 100 MHz to 400 MHz, use radiatedsusceptibility tests at a minimum of 20 volts permeter (V/m) peak with CW and 1 kHz squarewave modulation with 90 percent depth or greater.

(5) From 400 MHz to 8 gigahertz (GHz), use ra-diated susceptibility tests at a minimum of 150V/m peak with pulse modulation of 4 percent dutycycle with a 1 kHz pulse repetition frequency. Thissignal must be switched on and off at a rate of 1Hz with a duty cycle of 50 percent.

(d) Equipment HIRF Test Level 2. EquipmentHIRF test level 2 is HIRF environment II in table IIof this appendix reduced by acceptable aircrafttransfer function and attenuation curves. Testingmust cover the frequency band of 10 kHz to 8GHz.

(e) Equipment HIRF Test Level 3.(1) From 10 kHz to 400 MHz, use conducted

susceptibility tests, starting at a minimum of 0.15mA at 10 kHz, increasing 20 dB per frequency de-cade to a minimum of 7.5 mA at 500 kHz.

(2) From 500 kHz to 40 MHz, use conductedsusceptibility tests at a minimum of 7.5 mA.

(3) From 40 MHz to 400 MHz, use conductedsusceptibility tests, starting at a minimum of 7.5mA at 40 MHz, decreasing 20 dB per frequencydecade to a minimum of 0.75 mA at 400 MHz.

(4) From 100 MHz to 8 GHz, use radiated sus-ceptibility tests at a minimum of 5 V/m.


Appendix M to Part 25 Federal Aviation Regulations

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APPENDIX M TO PART 25FUEL TANK SYSTEM FLAMMABILITY

REDUCTION MEANS

Source: Docket No. FAA–2005–22997, 73 FR 42494,July 21, 2008, unless otherwise noted.

M25.1 FUEL TANK FLAMMABILITY EXPOSURE REQUIREMENTS.

(a) The Fleet Average Flammability Exposureof each fuel tank, as determined in accordancewith Appendix N of this part, may not exceed 3percent of the Flammability Exposure EvaluationTime (FEET), as defined in Appendix N of thispart. As a portion of this 3 percent, if flammabilityreduction means (FRM) are used, each of the fol-lowing time periods may not exceed 1.8 percentof the FEET:

(1) When any FRM is operational but the fueltank is not inert and the tank is flammable; and

(2) When any FRM is inoperative and the tankis flammable.

(b) The Fleet Average Flammability Exposure,as defined in Appendix N of this part, of each fueltank may not exceed 3 percent of the portion ofthe FEET occurring during either ground or take-off/climb phases of flight during warm days. Theanalysis must consider the following conditions.

(1) The analysis must use the subset of thoseflights that begin with a sea level ground ambienttemperature of 80°F (standard day plus 21°F at-mosphere) or above, from the flammability expo-sure analysis done for overall performance.

(2) For the ground and takeoff/climb phases offlight, the average flammability exposure must becalculated by dividing the time during the specificflight phase the fuel tank is flammable by the totaltime of the specific flight phase.

(3) Compliance with this paragraph may beshown using only those flights for which the air-plane is dispatched with the flammability reduc-tion means operational.

M25.2 SHOWING COMPLIANCE.(a) The applicant must provide data from analy-

sis, ground testing, and flight testing, or any com-bination of these, that:

(1) Validate the parameters used in the analy-sis required by paragraph M25.1 of this appendix;

(2) Substantiate that the FRM is effective at lim-iting flammability exposure in all compartments ofeach tank for which the FRM is used to showcompliance with paragraph M25.1 of this appen-dix; and

(3) Describe the circumstances under whichthe FRM would not be operated during eachphase of flight.

(b) The applicant must validate that the FRMmeets the requirements of paragraph M25.1 ofthis appendix with any airplane or engine configu-ration affecting the performance of the FRM forwhich approval is sought.

M25.3 RELIABILITY INDICATIONS AND MAINTENANCE ACCESS.

(a) Reliability indications must be provided toidentify failures of the FRM that would otherwisebe latent and whose identification is necessary toensure the fuel tank with an FRM meets the fleetaverage flammability exposure requirementslisted in paragraph M25.1 of this appendix, includ-ing when the FRM is inoperative.

(b) Sufficient accessibility to FRM reliability in-dications must be provided for maintenance per-sonnel or the flightcrew.

(c) The access doors and panels to the fueltanks with FRMs (including any tanks that com-municate with a tank via a vent system), and toany other confined spaces or enclosed areas thatcould contain hazardous atmosphere under nor-mal conditions or failure conditions, must be per-manently stenciled, marked, or placarded to warnmaintenance personnel of the possible presenceof a potentially hazardous atmosphere.

Part 25: Airworthiness Standards: Transport Category Appendix M to Part 25

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M25.4 AIRWORTHINESS LIMITATIONS AND PROCEDURES.

(a) If FRM is used to comply with paragraphM25.1 of this appendix, Airworthiness Limitationsmust be identified for all maintenance or inspec-tion tasks required to identify failures of compo-nents within the FRM that are needed to meetparagraph M25.1 of this appendix.

(b) Maintenance procedures must be devel-oped to identify any hazards to be consideredduring maintenance of the FRM. These proce-dures must be included in the instructions for con-tinued airworthiness (ICA).

M25.5 RELIABILITY REPORTING.The effects of airplane component failures on

FRM reliability must be assessed on an on-goingbasis. The applicant/holder must do the following:

(a) Demonstrate effective means to ensure col-lection of FRM reliability data. The means mustprovide data affecting FRM reliability, such ascomponent failures.

(b) Unless alternative reporting procedures areapproved by the FAA Oversight Office, as definedin part 26 of this subchapter, provide a report tothe FAA every six months for the first five years af-ter service introduction. After that period, contin-ued reporting every six months may be replacedwith other reliability tracking methods found ac-ceptable to the FAA or eliminated if it is estab-lished that the reliability of the FRM meets, andwill continue to meet, the exposure requirementsof paragraph M25.1 of this appendix.

(c) Develop service instructions or revise theapplicable airplane manual, according to a sched-ule approved by the FAA Oversight Office, as de-fined in part 26 of this subchapter, to correct anyfailures of the FRM that occur in service thatcould increase any fuel tank’s Fleet AverageFlammability Exposure to more than that requiredby paragraph M25.1 of this appendix.

Appendix N to Part 25 Federal Aviation Regulations

232 ASA

APPENDIX N TO PART 25FUEL TANK FLAMMABILITY EXPOSURE AND

RELIABILITY ANALYSIS

Source: Docket No. FAA–2005–22997, 73 FR 42495,July 21, 2008, unless otherwise noted.

N25.1 GENERAL.(a) This appendix specifies the requirements

for conducting fuel tank fleet average flammabilityexposure analyses required to meet §25.981(b)and Appendix M of this part. For fuel tanks in-stalled in aluminum wings, a qualitative assess-ment is sufficient if it substantiates that the tank isa conventional unheated wing tank.

(b) This appendix defines parameters affectingfuel tank flammability that must be used in per-forming the analysis. These include parametersthat affect all airplanes within the fleet, such as astatistical distribution of ambient temperature, fuelflash point, flight lengths, and airplane descentrate. Demonstration of compliance also requiresapplication of factors specific to the airplanemodel being evaluated. Factors that need to be in-cluded are maximum range, cruise mach number,typical altitude where the airplane begins initialcruise phase of flight, fuel temperature duringboth ground and flight times, and the performanceof a flammability reduction means (FRM) if in-stalled.

(c) The following definitions, input variables,and data tables must be used in the program todetermine fleet average flammability exposure fora specific airplane model.

N25.2 DEFINITIONS.(a) Bulk Average Fuel Temperature means

the average fuel temperature within the fuel tankor different sections of the tank if the tank is sub-divided by baffles or compartments.

(b) Flammability Exposure Evaluation Time(FEET). The time from the start of preparing theairplane for flight, through the flight and landing,until all payload is unloaded, and all passengersand crew have disembarked. In the Monte Carloprogram, the flight time is randomly selected fromthe Flight Length Distribution (Table 2), the pre-flight times are provided as a function of the flighttime, and the post-flight time is a constant 30 min-utes.

(c) Flammable. With respect to a fluid or gas,flammable means susceptible to igniting readily orto exploding (14 CFR Part 1, Definitions). A non-flammable ullage is one where the fuel-air vaporis too lean or too rich to burn or is inert as definedbelow. For the purposes of this appendix, a fueltank that is not inert is considered flammablewhen the bulk average fuel temperature within thetank is within the flammable range for the fuel type

being used. For any fuel tank that is subdividedinto sections by baffles or compartments, the tankis considered flammable when the bulk averagefuel temperature within any section of the tank,that is not inert, is within the flammable range forthe fuel type being used.

(d) Flash Point. The flash point of a flammablefluid means the lowest temperature at which theapplication of a flame to a heated sample causesthe vapor to ignite momentarily, or “flash.” Table 1of this appendix provides the flash point for thestandard fuel to be used in the analysis.

(e) Fleet average flammability exposure isthe percentage of the flammability exposure eval-uation time (FEET) each fuel tank ullage is flam-mable for a fleet of an airplane type operatingover the range of flight lengths in a world-widerange of environmental conditions and fuel prop-erties as defined in this appendix.

(f) Gaussian Distribution is another name forthe normal distribution, a symmetrical frequencydistribution having a precise mathematical for-mula relating the mean and standard deviation ofthe samples. Gaussian distributions yield bell-shaped frequency curves having a preponder-ance of values around the mean with progres-sively fewer observations as the curve extendsoutward.

(g) Hazardous atmosphere. An atmospherethat may expose maintenance personnel, pas-sengers or flight crew to the risk of death, inca-pacitation, impairment of ability to self-rescue(that is, escape unaided from a confined space),injury, or acute illness.

(h) Inert. For the purpose of this appendix, thetank is considered inert when the bulk averageoxygen concentration within each compartment ofthe tank is 12 percent or less from sea level up to10,000 feet altitude, then linearly increasing from12 percent at 10,000 feet to 14.5 percent at40,000 feet altitude, and extrapolated linearlyabove that altitude.

(i) Inerting. A process where a noncombusti-ble gas is introduced into the ullage of a fuel tankso that the ullage becomes non-flammable.

(j) Monte Carlo Analysis. The analyticalmethod that is specified in this appendix as thecompliance means for assessing the fleet aver-age flammability exposure time for a fuel tank.

(k) Oxygen evolution occurs when oxygendissolved in the fuel is released into the ullage asthe pressure and temperature in the fuel tank arereduced.

(l) Standard deviation is a statistical measureof the dispersion or variation in a distribution,equal to the square root of the arithmetic mean ofthe squares of the deviations from the arithmeticmeans.

(m) Transport Effects. For purposes of thisappendix, transport effects are the change in fuel

Part 25: Airworthiness Standards: Transport Category Appendix N to Part 25

ASA 233

25

vapor concentration in a fuel tank caused by lowfuel conditions and fuel condensation and vapor-ization.

(n) Ullage. The volume within the fuel tank notoccupied by liquid fuel.

N25.3 FUEL TANK FLAMMABILITY EXPOSURE ANALYSIS.

(a) A flammability exposure analysis must beconducted for the fuel tank under evaluation todetermine fleet average flammability exposure forthe airplane and fuel types under evaluation. Forfuel tanks that are subdivided by baffles or com-partments, an analysis must be performed eitherfor each section of the tank, or for the section ofthe tank having the highest flammability expo-sure. Consideration of transport effects is not al-lowed in the analysis. The analysis must be donein accordance with the methods and proceduresset forth in the Fuel Tank Flammability Assess-ment Method User’s Manual, dated May 2008,document number DOT/FAA/AR-05/8 (incorpo-rated by reference, see §25.5). The parametersspecified in sections N25.3(b) and (c) of this ap-pendix must be used in the fuel tank flammabilityexposure “Monte Carlo” analysis.

(b) The following parameters are defined in theMonte Carlo analysis and provided in paragraphN25.4 of this appendix:

(1) Cruise Ambient Temperature, as defined inthis appendix.

(2) Ground Ambient Temperature, as defined inthis appendix.

(3) Fuel Flash Point, as defined in this appen-dix.

(4) Flight Length Distribution, as defined in Ta-ble 2 of this appendix.

(5) Airplane Climb and Descent Profiles, as de-fined in the Fuel Tank Flammability AssessmentMethod User’s Manual, dated May 2008, docu-ment number DOT/FAA/AR-05/8 (incorporated byreference in §25.5).

(c) Parameters that are specific to the particu-lar airplane model under evaluation that must beprovided as inputs to the Monte Carlo analysisare:

(1) Airplane cruise altitude.(2) Fuel tank quantities. If fuel quantity affects

fuel tank flammability, inputs to the Monte Carloanalysis must be provided that represent the ac-tual fuel quantity within the fuel tank or compart-ment of the fuel tank throughout each of theflights being evaluated. Input values for this datamust be obtained from ground and flight test dataor the approved FAA fuel management proce-dures.

(3) Airplane cruise mach number.(4) Airplane maximum range.

(5) Fuel tank thermal characteristics. If fueltemperature affects fuel tank flammability, inputsto the Monte Carlo analysis must be provided thatrepresent the actual bulk average fuel tempera-ture within the fuel tank at each point in timethroughout each of the flights being evaluated.For fuel tanks that are subdivided by baffles orcompartments, bulk average fuel temperature in-puts must be provided for each section of thetank. Input values for these data must be obtainedfrom ground and flight test data or a thermalmodel of the tank that has been validated byground and flight test data.

(6) Maximum airplane operating temperaturelimit, as defined by any limitations in the airplaneflight manual.

(7) Airplane Utilization. The applicant must pro-vide data supporting the number of flights per dayand the number of hours per flight for the specificairplane model under evaluation. If there is no ex-isting airplane fleet data to support the airplanebeing evaluated, the applicant must provide sub-stantiation that the number of flights per day andthe number of hours per flight for that airplanemodel is consistent with the existing fleet datathey propose to use.

(d) Fuel Tank FRM Model. If FRM is used, anFAA approved Monte Carlo program must beused to show compliance with the flammability re-quirements of §25.981 and Appendix M of thispart. The program must determine the time peri-ods during each flight phase when the fuel tank orcompartment with the FRM would be flammable.The following factors must be considered in es-tablishing these time periods:

(1) Any time periods throughout the flammabil-ity exposure evaluation time and under the fullrange of expected operating conditions, when theFRM is operating properly but fails to maintain anon-flammable fuel tank because of the effects ofthe fuel tank vent system or other causes,

(2) If dispatch with the system inoperative un-der the Master Minimum Equipment List (MMEL)is requested, the time period assumed in the reli-ability analysis (60 flight hours must be used for a10-day MMEL dispatch limit unless an alternativeperiod has been approved by the Administrator),

(3) Frequency and duration of time periods ofFRM inoperability, substantiated by test or analy-sis acceptable to the FAA, caused by latent orknown failures, including airplane system shut-downs and failures that could cause the FRM toshut down or become inoperative.

(4) Effects of failures of the FRM that could in-crease the flammability exposure of the fuel tank.

(5) If an FRM is used that is affected by oxygenconcentrations in the fuel tank, the time periodswhen oxygen evolution from the fuel results in thefuel tank or compartment exceeding the inertlevel. The applicant must include any times when

Appendix N to Part 25 Federal Aviation Regulations

234 ASA

oxygen evolution from the fuel in the tank or com-partment under evaluation would result in a flam-mable fuel tank. The oxygen evolution rate thatmust be used is defined in the Fuel Tank Flamma-bility Assessment Method User’s Manual, datedMay 2008, document number DOT/FAA/AR-05/8(incorporated by reference in §25.5).

(6) If an inerting system FRM is used, the ef-fects of any air that may enter the fuel tank follow-ing the last flight of the day due to changes in am-bient temperature, as defined in Table 4, during a12-hour overnight period.

(e) The applicant must submit to the FAA Over-sight Office for approval the fuel tank flammabilityanalysis, including the airplane-specific parame-ters identified under paragraph N25.3(c) of thisappendix and any deviations from the parametersidentified in paragraph N25.3(b) of this appendixthat affect flammability exposure, substantiatingdata, and any airworthiness limitations and otherconditions assumed in the analysis.

N25.4 VARIABLES AND DATA TABLES.The following data must be used when con-

ducting a flammability exposure analysis to deter-mine the fleet average flammability exposure.Variables used to calculate fleet flammability ex-posure must include atmospheric ambient tem-peratures, flight length, flammability exposureevaluation time, fuel flash point, thermal charac-teristics of the fuel tank, overnight temperaturedrop, and oxygen evolution from the fuel into theullage.

(a) Atmospheric Ambient Temperatures andFuel Properties.

(1) In order to predict flammability exposureduring a given flight, the variation of ground ambi-ent temperatures, cruise ambient temperatures,and a method to compute the transition fromground to cruise and back again must be used.The variation of the ground and cruise ambienttemperatures and the flash point of the fuel is de-

fined by a Gaussian curve, given by the 50 per-cent value and a 1-standard deviation value.

(2) Ambient Temperature: Under the program,the ground and cruise ambient temperatures arelinked by a set of assumptions on the atmo-sphere. The temperature varies with altitude fol-lowing the International Standard Atmosphere(ISA) rate of change from the ground ambienttemperature until the cruise temperature for theflight is reached. Above this altitude, the ambienttemperature is fixed at the cruise ambient temper-ature. This results in a variation in the upper atmo-spheric temperature. For cold days, an inversionis applied up to 10,000 feet, and then the ISA rateof change is used.

(3) Fuel properties:(i) For Jet A fuel, the variation of flash point of

the fuel is defined by a Gaussian curve, given bythe 50 percent value and a 1-standard deviation,as shown in Table 1 of this appendix.

(ii) The flammability envelope of the fuel thatmust be used for the flammability exposure analy-sis is a function of the flash point of the fuel se-lected by the Monte Carlo for a given flight. Theflammability envelope for the fuel is defined by theupper flammability limit (UFL) and lower flamma-bility limit (LFL) as follows:

(A) LFL at sea level = flash point temperature ofthe fuel at sea level minus 10°F. LFL decreasesfrom sea level value with increasing altitude at arate of 1°F per 808 feet.

(B) UFL at sea level = flash point temperatureof the fuel at sea level plus 63.5°F. UFL decreasesfrom the sea level value with increasing altitude ata rate of 1°F per 512 feet.

(4) For each flight analyzed, a separate randomnumber must be generated for each of the threeparameters (ground ambient temperature, cruiseambient temperature, and fuel flash point) usingthe Gaussian distribution defined in Table 1 of thisappendix.

TABLE 1—GAUSSIAN DISTRIBUTION FOR GROUND AMBIENT TEMPERATURE, CRUISE AMBIENT TEMPERATURE, AND FUEL FLASH POINT

(b) The Flight Length Distribution defined inTable 2 must be used in the Monte Carlo analysis.

ParameterTemperature in deg F

Ground ambient temperature

Cruise ambient temperature Fuel flash point (FP)

Mean Temp 59.95 -70 120

Neg 1 std dev 20.14 8 8

Pos 1 std dev 17.28 8 8

Part 25: Airworthiness Standards: Transport Category Appendix N to Part 25

ASA 235

25

TABLE 2—FLIGHT LENGTH DISTRIBUTION

Flight length (NM) Airplane maximum range—nautical miles (NM)

From To 1000 2000 3000 4000 5000 6000 7000 8000 9000 10000

Distribution of flight lengths (percentage of total)

0 200 11.7 7.5 6.2 5.5 4.7 4.0 3.4 3.0 2.6 2.3

200 400 27.3 19.9 17.0 15.2 13.2 11.4 9.7 8.5 7.5 6.7

400 600 46.3 40.0 35.7 32.6 28.5 24.9 21.2 18.7 16.4 14.8

600 800 10.3 11.6 11.0 10.2 9.1 8.0 6.9 6.1 5.4 4.8

800 1000 4.4 8.5 8.6 8.2 7.4 6.6 5.7 5.0 4.5 4.0

1000 1200 0.0 4.8 5.3 5.3 4.8 4.3 3.8 3.3 3.0 2.7

1200 1400 0.0 3.6 4.4 4.5 4.2 3.8 3.3 3.0 2.7 2.4

1400 1600 0.0 2.2 3.3 3.5 3.3 3.1 2.7 2.4 2.2 2.0

1600 1800 0.0 1.2 2.3 2.6 2.5 2.4 2.1 1.9 1.7 1.6

1800 2000 0.0 0.7 2.2 2.6 2.6 2.5 2.2 2.0 1.8 1.7

2000 2200 0.0 0.0 1.6 2.1 2.2 2.1 1.9 1.7 1.6 1.4

2200 2400 0.0 0.0 1.1 1.6 1.7 1.7 1.6 1.4 1.3 1.2

2400 2600 0.0 0.0 0.7 1.2 1.4 1.4 1.3 1.2 1.1 1.0

2600 2800 0.0 0.0 0.4 0.9 1.0 1.1 1.0 0.9 0.9 0.8

2800 3000 0.0 0.0 0.2 0.6 0.7 0.8 0.7 0.7 0.6 0.6

3000 3200 0.0 0.0 0.0 0.6 0.8 0.8 0.8 0.8 0.7 0.7

3200 3400 0.0 0.0 0.0 0.7 1.1 1.2 1.2 1.1 1.1 1.0

3400 3600 0.0 0.0 0.0 0.7 1.3 1.6 1.6 1.5 1.5 1.4

3600 3800 0.0 0.0 0.0 0.9 2.2 2.7 2.8 2.7 2.6 2.5

3800 4000 0.0 0.0 0.0 0.5 2.0 2.6 2.8 2.8 2.7 2.6

4000 4200 0.0 0.0 0.0 0.0 2.1 3.0 3.2 3.3 3.2 3.1

4200 4400 0.0 0.0 0.0 0.0 1.4 2.2 2.5 2.6 2.6 2.5

4400 4600 0.0 0.0 0.0 0.0 1.0 2.0 2.3 2.5 2.5 2.4

4600 4800 0.0 0.0 0.0 0.0 0.6 1.5 1.8 2.0 2.0 2.0

4800 5000 0.0 0.0 0.0 0.0 0.2 1.0 1.4 1.5 1.6 1.5

5000 5200 0.0 0.0 0.0 0.0 0.0 0.8 1.1 1.3 1.3 1.3

5200 5400 0.0 0.0 0.0 0.0 0.0 0.8 1.2 1.5 1.6 1.6

5400 5600 0.0 0.0 0.0 0.0 0.0 0.9 1.7 2.1 2.2 2.3

5600 5800 0.0 0.0 0.0 0.0 0.0 0.6 1.6 2.2 2.4 2.5

5800 6000 0.0 0.0 0.0 0.0 0.0 0.2 1.8 2.4 2.8 2.9

6000 6200 0.0 0.0 0.0 0.0 0.0 0.0 1.7 2.6 3.1 3.3

6200 6400 0.0 0.0 0.0 0.0 0.0 0.0 1.4 2.4 2.9 3.1

6400 6600 0.0 0.0 0.0 0.0 0.0 0.0 0.9 1.8 2.2 2.5

6600 6800 0.0 0.0 0.0 0.0 0.0 0.0 0.5 1.2 1.6 1.9

6800 7000 0.0 0.0 0.0 0.0 0.0 0.0 0.2 0.8 1.1 1.3

7000 7200 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.4 0.7 0.8

7200 7400 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.3 0.5 0.7

7400 7600 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.2 0.5 0.6

7600 7800 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.1 0.5 0.7

7800 8000 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.1 0.6 0.8

8000 8200 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.5 0.8

8200 8400 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.5 1.0

8400 8600 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.6 1.3

8600 8800 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.4 1.1

8800 9000 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.2 0.8

9000 9200 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.5

9200 9400 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.2

9400 9600 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.1

9600 9800 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.1

9800 10000 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.0 0.1


236 ASA

(c) Overnight Temperature Drop. For air-planes on which FRM is installed, the overnighttemperature drop for this appendix is defined us-ing:

(1) A temperature at the beginning of the over-night period that equals the landing temperatureof the previous flight that is a random value basedon a Gaussian distribution; and

(2) An overnight temperature drop that is a ran-dom value based on a Gaussian distribution.

(3) For any flight that will end with an overnightground period (one flight per day out of an aver-age number of flights per day, depending on utili-zation of the particular airplane model being eval-uated), the landing outside air temperature (OAT)is to be chosen as a random value from the follow-ing Gaussian curve:

TABLE 3—LANDING OUTSIDE AIR TEMPERATURE

(4) The outside ambient air temperature (OAT)overnight temperature drop is to be chosen as arandom value from the following Gaussian curve:

TABLE 4—OUTSIDE AIR TEMPERATURE (OAT) DROP

(d) Number of Simulated Flights Required in Analysis. In order for the Monte Carlo analysis tobe valid for showing compliance with the fleet average and warm day flammability exposure require-ments, the applicant must run the analysis for a minimum number of flights to ensure that the fleet av-erage and warm day flammability exposure for the fuel tank under evaluation meets the applicableflammability limits defined in Table 5 of this appendix.

TABLE 5—FLAMMABILITY EXPOSURE LIMIT

Parameter Landing outside air temperature °F

Mean Temperature 58.68

negative 1 std dev 20.55

positive 1 std dev 13.21

Parameter OAT drop temperature °F

Mean Temp 12.0

1 std dev 6.0

Minimum number of flights in Monte Carlo analysis

Maximum acceptable Monte Carlo average fuel tank flammability exposure (percent) to meet 3

percent requirements

Maximum acceptable Monte Carlo average fuel tank flammability exposure (percent) to meet 7 percent Part 26 requirements

10,000 2.91 6.79

100,000 2.98 6.96

1,000,000 3.00 7.00

Send additional terms to [email protected]

ASA • Part 25 Index 237

AAccelerate-stop distance §25.109 ............................. 17Accessory gearboxes §25.1167 .............................. 111After-cooler §25.1107.............................................. 108Air duct systems §25.1103 ...................................... 107Air induction §25.1091............................................. 106Aircraft structure

control surface and system loads §25.391–25.459 ............................... 43–45

emergency landing conditions §25.561–25.563 ....................... 54–56

fatigue evaluation §25.571.................................... 56flight maneuver and gust conditions

§25.331–25.351 ........................................... 36–41general load factors, limits

§25.301–25.307 ................................................. 35ground loads §25.471–25.519 ........................ 45–51lightning protection §25.581.................................. 57water loads §25.521–25.537........................... 51–54

Airplane flight manual §25.1581–25.1587....... 139–141Airspeed limitations

flap extended speed §25.1511............................ 136general §25.1503 ................................................ 136landing gear §25.1515 ........................................ 136maneuvering speed §25.1507 ............................ 136maximum operating §25.1505 ............................ 136minimum control speed §25.1513....................... 136rough air §25.1517.............................................. 136

Airworthiness standardsflight

compliance, proof of §25.21 ............................. 13controllability, maneuverability

§25.143–25.149......................................... 23–27ground and water handling

§25.231–25.239........................................ 31–32load distribution limits §25.23 ........................... 13stability §25.171–25.181............................. 27–29stalls §25.201–25.207................................. 29–30trim §25.161 ...................................................... 27

high-speed characteristics §25.253 ...................... 33out-of-trim characteristics §25.255........................ 33transport category airplanes

applicability, special requirements §25.1, 25.2...................................................... 12

vibration and buffeting §25.251............................. 32Automatic Takeoff Thrust Control

System (ATTCS) Appendix I............................... 216Auxiliary power unit

controls §25.1142 ............................................... 109

BBatteries §25.1353 .................................................. 123

CCarburetor air, preheater design

§25.1101................................................................ 107Carburetor air temperature controls §25.1157 ........ 110Center of gravity limits §25.27................................... 14Climb, general §25.117 ............................................. 21Climb, one engine inoperative §25.121..................... 21Cockpit voice recorders §25.1457........................... 133Continued airworthiness

instructions §25.1529, Appendix H ............. 137, 214Continuous gust design criteria

Appendix G........................................................... 210Controllability and maneuverability,

general §25.143 .................................................... 23

Coolinggeneral §25.1041................................................ 105test procedures §25.1045 ................................... 106tests §25.1043 .................................................... 105

Cowling §25.1193 ................................................... 113

DDesign airspeeds §25.335 ........................................ 37Design and construction

accessibility §25.611 ............................................ 58aeroelastic stability §25.629 ................................. 60bearing factors §25.623 ........................................ 59bird strike damage §25.631 .................................. 61casting factors §25.621 ........................................ 59control surfaces

hinges §25.657 ................................................. 61installation §25.655........................................... 61proof of strength §25.651 ................................. 61

control systemscables §25.689 ................................................. 63details §25.685 ................................................. 63flaps or slats §25.701 ....................................... 64gust locks §25.679............................................ 62joints §25.693 ................................................... 63lift/drag devices §25.697, 25.699................ 63, 64limit load tests §25.681..................................... 63operation tests §25.683 .................................... 63stability augmentation §25.672......................... 62stops §25.675 ................................................... 62takeoff warning §25.703 ................................... 64trim systems §25.677 ....................................... 62

control systems §25.671–25.703.................... 61–64emergency provisions §25.801–25.819.......... 76–85fabrication methods for §25.605 ........................... 58fasteners §25.607 ................................................. 58fire protection

cargo or baggage compartments §25.855–25.858 ....................................... 90–91

compartment interiors §25.853......................... 89components §25.867 ........................................ 93extinguishers §25.851 ...................................... 88flammable fluids §25.863.................................. 92flight controls §25.865 ...................................... 92heaters §25.859................................................ 91lavatories §25.854 ............................................ 89

fire protection §25.851–25.869 ....................... 88–93fitting factors §25.625 ........................................... 60floats and hulls §25.751–25.755........................... 68general standards §25.601–25.631 ................ 58–61landing gear

specifications §25.721–25.737.................. 64–68material strength §25.613 ..................................... 58materials for §25.603 ............................................ 58personnel and cargo accommodations

§25.771–25.793........................................... 68–75pressurization §25.841, 25.843 ...................... 87, 88protection of structure §25.609 ............................. 58static electricity

protection against §25.899 ............................... 93thermal/acoustic insulation materials §25.856...... 90ventilation, heating §25.831–25.833............... 86–87

Design fuel and oil loads §25.343 ............................. 39Designated fire zones §25.1181 ............................. 111

238 ASA • Part 25 Index

EEarly ETOPS method Appendix K................... 223, 225Electrical appliances, motors and

transformers §25.1365........................................ 124Electrical supply

for emergency services §25.1362....................... 124Electrical systems and equipment

circuit protection §25.1357.................................. 123distribution system §25.1355 .............................. 123general, airworthiness §25.1351......................... 122installation §25.1353 ........................................... 123lights

anticollision §25.1401 ..................................... 126instrument §25.1381 ....................................... 124landing §25.1383 ............................................ 125position §25.1385–25.1397 .................... 125–126riding §25.1399 ............................................... 126

tests §25.1363 .................................................... 124wing icing detection lights §25.1403 ................... 127

Electrical Wiring Interconnection System (EWIS) Appendix H............................................. 215

Electrical Wiring Interconnection Systems (EWIS) §25.1701–1733 .............................. 142–145

Electronic equipment §25.1431............................... 129Empty weight §25.29................................................. 14En route flight paths §25.123 .................................... 22Engine

accessory sectiondiaphragm §25.1192 ...................................... 113

controls §25.1143 ............................................... 109ignition systems §25.1165 .................................. 110torque §25.361...................................................... 41

Equipmentairworthiness standards

§25.1301–25.1461 ................................... 115–135function and installation §25.1301 ...................... 115instruments, flight and navigation §25.1303 ....... 115instruments, powerplant §25.1305...................... 115miscellaneous §25.1307 ..................................... 116systems, installation §25.1309............................ 116

ETOPS design requirements Appendix K ............... 221ETOPS type design §25.3......................................... 12ETOPS Appendix K................................................. 221EWIS (electrical wiring interconnection

system) Appendix H............................................ 215EWIS (electrical wiring interconnection

systems) §25.1701–1733............................ 142–145Exhaust driven turbo-supercharger §25.1127 ......... 109Exhaust heat exchanger §25.1125.......................... 108Exhaust piping §25.1123......................................... 108Exhaust system, general §25.1121 ......................... 108Extended operations (ETOPS) Appendix K ............ 221Extinguishing agent containers §25.1199 ............... 114

FFire

detector system §25.1203 .................................. 114extinguishers §25.851 .......................................... 88extinguishing agents §25.1197 ........................... 113extinguishing system materials §25.1201........... 114extinguishing systems §25.1195 ........................ 113protection, compliance §25.1207........................ 114zones

drainage of §25.1187...................................... 112ventilation of §25.1187 ................................... 112

Fire walls §25.1191 ................................................. 113Flammability Exposure Evaluation Time (FEET)

§25.981, Appendices M and N...............101, 229, 231Flammable

fluid-carrying components §25.1183 .................. 111fluids §25.1185 ................................................... 112

Flight, controllability,maneuverability §25.143–25.149 ................... 23–27

Flight data recorders §25.1459 ............................... 134Flight guidance system §25.1329 ........................... 120Flight instruments §25.1303.................................... 115Flight load factors §25.321........................................ 35Flight maneuvering envelope §25.333 ...................... 37Flight manual, airplane

operating limitations §25.1583............................ 140operating procedures §25.1585.......................... 140

Fluid, draining of §25.1455...................................... 133Fuel

jettisoning system controls §25.1161 ................. 110jettisoning system §25.1001 ............................... 103

Fuel pumps §25.991 ............................................... 102Fuel system

airworthiness standards §25.951–25.981..... 97–102analysis, tests §25.952 ......................................... 97components §25.991–25.1001 ................... 102–103fuel flow §25.955, 25.957 ..................................... 98hot weather operation §25.961 ............................. 98lightning protection §25.954 ................................. 98pressure fueling §25.979 .................................... 101tanks §25.963–25.977 .................................. 99–101

Fuel tank explosion prevention §25.981 ................. 101Fuel Tank Flammability Assessment Method User’s

Manual §25.5, 25.981, Appendix N .......12, 102, 232Fuel tank flammability exposure Appendix N .......... 231Fuel tank system flammability Appendix M ............. 229

GGust and turbulence loads §25.341 .......................... 39

HHigh energy rotors, in equipment §25.1461 ............ 135High lift devices §25.345 ........................................... 40High-intensity radiated fields (HIRF)

protection from §25.1317.................................... 118HIRF environments and equipment

test levels Appendix L......................................... 227Hulls §25.755 ............................................................ 68Hydraulic systems §25.1435 ................................... 129

Send additional terms to [email protected]

ASA • Part 25 Index 239

IIcing conditions

maximum intensities, tables Appendix C ............ 154Ignition switches §25.1145 ...................................... 110Induction system

ducts §25.1103 ................................................... 107icing protection §25.1093.................................... 107screens §25.1105 ............................................... 108

In-flight shutdown (IFSD) rates Appendix K ............ 222Injury, precautions §25.1360 ................................... 124Installation of

airspeed indication system §25.1323.................. 119flight guidance system §25.1329 ........................ 120instrument systems §25.1333 ............................. 121instruments using a power supply §25.1331....... 121magnetic direction indicator §25.1327 ................ 120pitot heat indication system §25.1326................. 120powerplant instruments §25.1337....................... 122static pressure system §25.1325 ........................ 119warning, caution, advisory lights §25.1322 ......... 118

Instrument markings §25.1543................................ 138Instruments

installation, arrangement and visibility§25.1321 ........................................................... 118

Inter-coolers §25.1107 ............................................ 108

LLanding §25.125........................................................ 22Landing climb §25.119 .............................................. 21Lightning protection, for systems §25.1316............. 117Load factors

limit maneuvering §25.337.................................... 38Loads

gyroscopic §25.371............................................... 43unsymmetrical §25.367, 25.427...................... 42, 45

MMarkings and placards

airspeed information §25.1545 ........................... 138cockpit control §25.1555 ..................................... 139magnetic direction indicator §25.1547 ................ 138powerplant, APU instruments §25.1549 ............. 138safety equipment §25.1561................................. 139specifications §25.1541–25.1563 ............... 138–139

Minimum flight crewcriteria for determining Appendix D..................... 162

Mixture controls §25.1147 ....................................... 110

NNacelle areas behind firewalls §25.1182................. 111Nacelle skin §25.1193 ............................................. 113Navigation instruments §25.1303............................ 115

OOil

filter §25.1019 ..................................................... 105fittings §25.1017 ................................................. 104lines §25.1017 .................................................... 104radiators §25.1023.............................................. 105strainer §25.1019................................................ 105system drains §25.1021 ..................................... 105system, general §25.1011 .................................. 104tank tests §25.1015 ............................................ 104tanks §25.1013 ................................................... 104valves §25.1025 ................................................. 105

Operating limitationsairspeed §25.1503–25.1517 ............................... 136auxiliary power unit §25.1522 ............................. 137for airworthiness, transport category

airplanes §25.1501–1533...................... 136–137maneuvering flight, load factors §25.1531.......... 137minimum flight crew

§25.1523, Appendix D ............................. 137, 162powerplant §25.1521 .......................................... 136weight, center of gravity and

weight distribution §25.1519 ......................... 136Oxygen

distributing system, standards §25.1445 ............ 131distributing units, standards §25.1447 ................ 132equipment protection from rupture §25.1453 ..... 133generators, chemical §25.1450 .......................... 132means for determining use §25.1449 ................. 132supplemental, flow of §25.1443 .......................... 131

Oxygen equipment, supply §25.1441...................... 131

PPerformance, general §25.101.................................. 14Position lights

color specifications §25.1397 ............................. 126dihedral angle required §25.1387 ....................... 125distribution, intensity §25.1389 ........................... 125installation §25.1385........................................... 125minimum, maximum intensities required (tables)

§25.1391, 25.1393, 25.1395............................ 126Power source capacity §25.1310 ............................ 117Powerplant

accessories §25.1163......................................... 110airworthiness standards, transport category

airplanes §25.901–25.945......................... 94–97augmentation systems §25.945............................ 97automatic takeoff thrust control system

(ATTCS) §25.904 ............................................ 95controls, general §25.1141 ................................. 109engines §25.903 ................................................... 94inlet, engine, exhaust §25.941.............................. 96installation §25.901............................................... 94negative acceleration §25.943.............................. 97propeller clearance §25.925 ................................. 95propeller deicing §25.929 ..................................... 95propeller vibration and fatigue §25.907 ................ 95propeller-drag limiting systems §25.937 ............... 96propellers §25.905 ................................................ 95thrust reversers, systems §25.933, 25.934 .......... 96turbine engine operating characteristics

§25.939............................................................... 96

240 ASA • Part 25 Index

Pressurization, pneumatic systems §25.1438......... 130Pressurized compartment loads §25.365.................. 41Propeller

feathering controls §25.1153 .............................. 110feathering system §25.1027................................ 105pitch controls §25.1149....................................... 110pitch settings below flight regime §25.1155 ........ 110reinforcement §25.875 .......................................... 93speed controls §25.1149..................................... 110speed, pitch limits §25.33 ..................................... 14

Protective breathing equipment §25.1439............... 130Public address system §25.1423 ............................ 129

RReverse thrust below the flight regime §25.1155 .... 110Rolling conditions §25.349 ........................................ 40

SSafety equipment

ditching §25.1415................................................ 128general requirements §25.1411 .......................... 127ice protection §25.1419....................................... 128megaphones §25.1421 ....................................... 128

Shutoff means §25.1189 ......................................... 112Side load on engine §25.363..................................... 41Signs §25.791 ........................................................... 75Speed control devices §25.373 ................................. 43Stall speed §25.103................................................... 15Stall warning §25.207................................................ 30Static electricity,

protection against §25.899.................................... 93Supercharger controls §25.1159 ............................. 110Symmetric maneuvering conditions §25.331 ............ 36

TTakeoff

distance and run §25.113 ..................................... 20flight path §25.115 ................................................ 21path §25.111......................................................... 19speeds §25.107 .................................................... 16

Takeoff §25.105 ........................................................ 15Test criteria, showing compliance

Appendix F ............................................................ 165Trim §25.161 ............................................................. 27

VVacuum systems §25.1433..................................... 129

WWeight limits §25.25.................................................. 13

Far Part25

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