Title 14--Aeronautics and Space
CHAPTER I--FEDERAL AVIATION ADMINISTRATION, DEPARTMENT OF TRANSPORTATION
SUBCHAPTER C--AIRCRAFT
PART 25--AIRWORTHINESS STANDARDS: TRANSPORT CATEGORY AIRPLANES
Special Federal Aviation Regulation No. 13
Subpart A--General
Sec. 25.1 Applicability.
Sec. 25.2 Special retroactive requirements.
Subpart B--Flight
General
Sec. 25.21 Proof of compliance.
Sec. 25.23 Load distribution limits.
Sec. 25.25 Weight limits.
Sec. 25.27 Center of gravity limits.
Sec. 25.29 Empty weight and corresponding center of gravity.
Sec. 25.31 Removable ballast.
Sec. 25.33 Propeller speed and pitch limits.
Performance
Sec. 25.101 General.
Sec. 25.103 Stalling speed.
Sec. 25.105 Takeoff.
Sec. 25.107 Takeoff speeds.
Sec. 25.109 Accelerate-stop distance.
Sec. 25.111 Takeoff path.
Sec. 25.113 Takeoff distance and takeoff run.
Sec. 25.115 Takeoff flight path.
Sec. 25.117 Climb: general.
Sec. 25.119 Landing climb: All-engine-operating.
Sec. 25.121 Climb: One-engine-inoperative.
Sec. 25.123 En route flight paths.
Sec. 25.125 Landing.
Controllability and Maneuverability
Sec. 25.143 General.
Sec. 25.145 Longitudinal control.
Sec. 25.147 Directional and lateral control.
Sec. 25.149 Minimum control speed.
Trim
Sec. 25.161 Trim.
Stability
Sec. 25.171 General.
Sec. 25.173 Static longitudinal stability.
Sec. 25.175 Demonstration of static longitudinal stability.
Sec. 25.177 Static lateral-directional stability.
Sec. 25.181 Dynamic stability.
Stalls
Sec. 25.201 Stall demonstration.
Sec. 25.203 Stall characteristics.
Sec. 25.205 [Removed. Amdt. 25-72, 55 FR 29775, July 20, 1990]
Sec. 25.207 Stall warning.
Ground and Water Handling Characteristics
Sec. 25.231 Longitudinal stability and control.
Sec. 25.233 Directional stability and control.
Sec. 25.235 Taxiing condition.
Sec. 25.237 Wind velocities.
Sec. 25.239 Spray characteristics, control, and stability on
water.
Miscellaneous Flight Requirements
Sec. 25.251 Vibration and buffeting.
Sec. 25.253 High-speed characteristics.
Sec. 25.255 Out-of-trim characteristics.
Subpart C--Structure
General
Sec. 25.301 Loads.
Sec. 25.303 Factor of safety.
Sec. 25.305 Strength and deformation.
Sec. 25.307 Proof of structure.
Flight Loads
Sec. 25.321 General.
Flight Maneuver and Gust Conditions
Sec. 25.331 General.
Sec. 25.333 Flight envelope.
Sec. 25.335 Design airspeeds.
Sec. 25.337 Limit maneuvering load factors.
Sec. 25.341 Gust loads.
Sec. 25.343 Design fuel and oil loads.
Sec. 25.345 High lift devices.
Sec. 25.349 Rolling conditions.
Sec. 25.351 Yawing conditions.
Supplementary Conditions
Sec. 25.361 Engine torque.
Sec. 25.363 Side load on engine mount.
Sec. 25.365 Pressurized cabin loads.
Sec. 25.367 Unsymmetrical loads due to engine failure.
Sec. 25.371 Gyroscopic loads.
Sec. 25.373 Speed control devices.
Control Surface and System Loads
Sec. 25.391 Control surface loads: general.
Sec. 25.393 Loads parallel to hinge line.
Sec. 25.395 Control system.
Sec. 25.397 Control system loads.
Sec. 25.399 Dual control system.
Sec. 25.405 Secondary control system.
Sec. 25.407 Trim tab effects.
Sec. 25.409 Tabs.
Sec. 25.415 Ground gust conditions.
Sec. 25.427 Unsymmetrical loads.
Sec. 25.445 Outboard fins.
Sec. 25.457 Wing flaps.
Sec. 25.459 Special devices.
Ground Loads
Sec. 25.471 General.
Sec. 25.473 Ground load conditions and assumptions.
Sec. 25.477 Landing gear arrangement.
Sec. 25.479 Level landing conditions.
Sec. 25.481 Tail-down landing conditions.
Sec. 25.483 One-wheel landing conditions.
Sec. 25.485 Side load conditions.
Sec. 25.487 Rebound landing condition.
Sec. 25.489 Ground handling conditions.
Sec. 25.491 Takeoff run.
Sec. 25.493 Braked roll conditions.
Sec. 25.495 Turning.
Sec. 25.497 Tail-wheel yawing.
Sec. 25.499 Nose-wheel yaw.
Sec. 25.503 Pivoting.
Sec. 25.507 Reversed braking.
Sec. 25.509 Towing loads.
Sec. 25.511 Ground load: unsymmetrical loads on multiple-wheel
units.
Water Loads
Sec. 25.521 General.
Sec. 25.523 Design weights and center of gravity positions.
Sec. 25.525 Application of loads.
Sec. 25.527 Hull and main float load factors.
Sec. 25.529 Hull and main float landing conditions.
Sec. 25.531 Hull and main float takeoff condition.
Sec. 25.533 Hull and main float bottom pressures.
Sec. 25.535 Auxiliary float loads.
Sec. 25.537 Seawing loads.
Emergency Landing Conditions
Sec. 25.561 General.
Sec. 25.562 Emergency landing dynamic conditions.
Sec. 25.563 Structural ditching provisions.
Fatigue Evaluation
Sec. 25.571 Damage--tolerance and fatigue evaluation of
structure.
Lightning Protection
Sec. 25.581 Lightning protection.
Subpart D--Design and Construction
General
Sec. 25.601 General.
Sec. 25.603 Materials.
Sec. 25.605 Fabrication methods.
Sec. 25.607 Fasteners.
Sec. 25.609 Protection of structure.
Sec. 25.611 Accessibility provisions.
Sec. 25.613 Material strength properties and design values.
Sec. 25.615 [Removed. Amdt. 25-72, 55 FR 29776, July 20, 1990]
Sec. 25.619 Special factors.
Sec. 25.621 Casting factors.
Sec. 25.623 Bearing factors.
Sec. 25.625 Fitting factors.
Sec. 25.629 Aeroelastic stability requirements.
Sec. 25.631 Bird strike damage.
Control Surfaces
Sec. 25.651 Proof of strength.
Sec. 25.655 Installation.
Sec. 25.657 Hinges.
Control Systems
Sec. 25.671 General.
Sec. 25.672 Stability augmentation and automatic and power-
operated systems.
Sec. 25.673 [Removed. Amdt. 25-72, 55 FR 29777, July 20, 1990]
Sec. 25.675 Stops.
Sec. 25.677 Trim systems.
Sec. 25.679 Control system gust locks.
Sec. 25.681 Limit load static tests.
Sec. 25.683 Operation tests.
Sec. 25.685 Control system details.
Sec. 25.689 Cable systems.
Sec. 25.693 Joints.
Sec. 25.697 Lift and drag devices, controls.
Sec. 25.699 Lift and drag device indicator.
Sec. 25.701 Flap and slat interconnection.
Sec. 25.703 Takeoff warning system.
Landing Gear
Sec. 25.721 General.
Sec. 25.723 Shock absorption tests.
Sec. 25.725 Limit drop tests.
Sec. 25.727 Reserve energy absorption drop tests.
Sec. 25.729 Retracting mechanism.
Sec. 25.731 Wheels.
Sec. 25.733 Tires.
Sec. 25.735 Brakes.
Sec. 25.737 Skis.
Floats and Hulls
Sec. 25.751 Main float buoyancy.
Sec. 25.753 Main float design.
Sec. 25.755 Hulls.
Personnel and Cargo Accommodations
Sec. 25.771 Pilot compartment.
Sec. 25.772 Pilot compartment doors.
Sec. 25.773 Pilot compartment view.
Sec. 25.775 Windshields and windows.
Sec. 25.777 Cockpit controls.
Sec. 25.779 Motion and effect of cockpit controls.
Sec. 25.781 Cockpit control knob shape.
Sec. 25.783 Doors.
Sec. 25.785 Seats, berths, safety belts, and harnesses.
Sec. 25.787 Stowage compartments.
Sec. 25.789 Retention of items of mass in passenger and crew
compartments and galleys.
Sec. 25.791 Passenger information signs and placards.
Sec. 25.793 Floor surfaces.
Emergency Provisions
Sec. 25.801 Ditching.
Sec. 25.803 Emergency evacuation.
Sec. 25.805 [Removed. 55 FR 29781, July 20, 1990]
Sec. 25.807 Emergency exits.
Sec. 25.809 Emergency exit arrangement.
Sec. 25.810 Emergency egress assist means and escape routes.
Sec. 25.811 Emergency exit marking.
Sec. 25.812 Emergency lighting.
Sec. 25.813 Emergency exit access.
Sec. 25.815 Width of aisle.
Sec. 25.817 Maximum number of seats abreast.
Sec. 25.819 Lower deck service compartments (including
galleys).
Ventilation and Heating
Sec. 25.831 Ventilation.
Sec. 25.832 Cabin ozone concentration.
Sec. 25.833 Combustion heating systems.
Pressurization
Sec. 25.841 Pressurized cabins.
Sec. 25.843 Tests for pressurized cabins.
Fire Protection
Sec. 25.851 Fire extinguishers.
Sec. 25.853 Compartment interiors.
Sec. 25.854 Lavatory fire protection.
Sec. 25.855 Cargo or baggage compartments.
Sec. 25.857 Cargo compartment classification.
Sec. 25.858 Cargo compartment fire detection systems.
Sec. 25.859 Combustion heater fire protection.
Sec. 25.863 Flammable fluid fire protection.
Sec. 25.865 Fire protection of flight controls, engine mounts,
and other flight structure.
Sec. 25.867 Fire protection: other components.
Sec. 25.869 Fire protection: systems.
Miscellaneous
Sec. 25.871 Leveling means.
Sec. 25.875 Reinforcement near propellers.
Subpart E--Powerplant
General
Sec. 25.901 Installation.
Sec. 25.903 Engines.
Sec. 25.904 Automatic takeoff thrust control system (ATTCS).
Sec. 25.905 Propellers.
Sec. 25.907 Propeller vibration.
Sec. 25.925 Propeller clearance.
Sec. 25.929 Propeller deicing.
Sec. 25.933 Reversing systems.
Sec. 25.934 Turbojet engine thrust reverser system tests.
Sec. 25.937 Turbopropeller-drag limiting systems.
Sec. 25.939 Turbine engine operating characteristics.
Sec. 25.941 Inlet, engine, and exhaust compatibility.
Sec. 25.943 Negative acceleration.
Sec. 25.945 Thrust or power augmentation system.
Fuel System
Sec. 25.951 General.
Sec. 25.952 Fuel system analysis and test.
Sec. 25.953 Fuel system independence.
Sec. 25.954 Fuel system lightning protection.
Sec. 25.955 Fuel flow.
Sec. 25.957 Flow between interconnected tanks.
Sec. 25.959 Unusable fuel supply.
Sec. 25.961 Fuel system hot weather operation.
Sec. 25.963 Fuel tanks: general.
Sec. 25.965 Fuel tank tests.
Sec. 25.967 Fuel tank installations.
Sec. 25.969 Fuel tank expansion space.
Sec. 25.971 Fuel tank sump.
Sec. 25.973 Fuel tank filler connection.
Sec. 25.975 Fuel tank vents and carburetor vapor vents.
Sec. 25.977 Fuel tank outlet.
Sec. 25.979 Pressure fueling system.
Sec. 25.981 Fuel tank temperature.
Fuel System Components
Sec. 25.991 Fuel pumps.
Sec. 25.993 Fuel system lines and fittings.
Sec. 25.994 Fuel system components.
Sec. 25.995 Fuel valves.
Sec. 25.997 Fuel strainer or filter.
Sec. 25.999 Fuel system drains.
Sec. 25.1001 Fuel jettisoning system.
Oil System
Sec. 25.1011 General.
Sec. 25.1013 Oil tanks.
Sec. 25.1015 Oil tank tests.
Sec. 25.1017 Oil lines and fittings.
Sec. 25.1019 Oil strainer or filter.
Sec. 25.1021 Oil system drains.
Sec. 25.1023 Oil radiators.
Sec. 25.1025 Oil valves.
Sec. 25.1027 Propeller feathering system.
Cooling
Sec. 25.1041 General.
Sec. 25.1043 Cooling tests.
Sec. 25.1045 Cooling test procedures.
Induction System
Sec. 25.1091 Air induction.
Sec. 25.1093 Induction system icing protection.
Sec. 25.1101 Carburetor air preheater design.
Sec. 25.1103 Induction system ducts and air duct systems.
Sec. 25.1105 Induction system screens.
Sec. 25.1107 Inter-coolers and after-coolers.
Exhaust System
Sec. 25.1121 General.
Sec. 25.1123 Exhaust piping.
Sec. 25.1125 Exhaust heat exchangers.
Sec. 25.1127 Exhaust driven turbo-superchargers.
Powerplant Controls and Accessories
Sec. 25.1141 Powerplant controls: general.
Sec. 25.1142 Auxiliary power unit controls.
Sec. 25.1143 Engine controls.
Sec. 25.1145 Ignition switches.
Sec. 25.1147 Mixture controls.
Sec. 25.1149 Propeller speed and pitch controls.
Sec. 25.1153 Propeller feathering controls.
Sec. 25.1155 Reverse thrust and propeller pitch settings below
the flight regime.
Sec. 25.1157 Carburetor air temperature controls.
Sec. 25.1159 Supercharger controls.
Sec. 25.1161 Fuel jettisoning system controls.
Sec. 25.1163 Powerplant accessories.
Sec. 25.1165 Engine ignition systems.
Sec. 25.1167 Accessory gearboxes.
Powerplant Fire Protection
Sec. 25.1181 Designated fire zones; regions included.
Sec. 25.1182 Nacelle areas behind firewalls, and engine pod
attaching structures containing flammable fluid lines.
Sec. 25.1183 Flammable fluid-carrying components.
Sec. 25.1185 Flammable fluids.
Sec. 25.1187 Drainage and ventilation of fire zones.
Sec. 25.1189 Shutoff means.
Sec. 25.1191 Firewalls.
Sec. 25.1192 Engine accessory section diaphragm.
Sec. 25.1193 Cowling and nacelle skin.
Sec. 25.1195 Fire extinguishing systems.
Sec. 25.1197 Fire extinguishing agents.
Sec. 25.1199 Extinguishing agent containers.
Sec. 25.1201 Fire extinguishing system materials.
Sec. 25.1203 Fire detector system.
Sec. 25.1207 Compliance.
Subpart F--Equipment
General
Sec. 25.1301 Function and installation.
Sec. 25.1303 Flight and navigation instruments.
Sec. 25.1305 Powerplant instruments.
Sec. 25.1307 Miscellaneous equipment.
Sec. 25.1309 Equipment, systems, and installations.
Instruments: Installation
Sec. 25.1321 Arrangement and visibility.
Sec. 25.1322 Warning, caution, and advisory lights.
Sec. 25.1323 Airspeed indicating system.
Sec. 25.1325 Static pressure systems.
Sec. 25.1326 Pitot heat indication systems.
Sec. 25.1327 Magnetic direction indicator.
Sec. 25.1329 Automatic pilot system.
Sec. 25.1331 Instruments using a power supply.
Sec. 25.1333 Instrument systems.
Sec. 25.1335 Flight director systems.
Sec. 25.1337 Powerplant instruments.
Electrical Systems and Equipment
Sec. 25.1351 General.
Sec. 25.1353 Electrical equipment and installations.
Sec. 25.1355 Distribution system.
Sec. 25.1357 Circuit protective devices.
Sec. 25.1359 [Removed. 55 FR 29785, July 20, 1990]
Sec. 25.1363 Electrical system tests.
Lights
Sec. 25.1381 Instrument lights.
Sec. 25.1383 Landing lights.
Sec. 25.1385 Position light system installation.
Sec. 25.1387 Position light system dihedral angles.
Sec. 25.1389 Position light distribution and intensities.
Sec. 25.1391 Minimum intensities in the horizontal plane of
forward and rear position lights.
Sec. 25.1393 Minimum intensities in any vertical plane of
forward and rear position lights.
Sec. 25.1395 Maximum intensities in overlapping beams of
forward and rear position lights.
Sec. 25.1397 Color specifications.
Sec. 25.1399 Riding light.
Sec. 25.1401 Anticollision light system.
Sec. 25.1403 Wing icing detection lights.
Safety Equipment
Sec. 25.1411 General.
Sec. 25.1413 [Removed. 55 FR 29785, July 20, 1990]
Sec. 25.1415 Ditching equipment.@
Sec. 25.1416 [Removed. 55 FR 29785, July 20, 1985]
Sec. 25.1419 Ice protection.
Sec. 25.1421 Megaphones.
Miscellaneous Equipment
Sec. 25.1423 Public address system.
Sec. 25.1431 Electronic equipment.
Sec. 25.1433 Vacuum systems.
Sec. 25.1435 Hydraulic systems.
Sec. 25.1438 Pressurization and pneumatic systems.
Sec. 25.1439 Protective breathing equipment.
Sec. 25.1441 Oxygen equipment and supply.
Sec. 25.1443 Minimum mass flow of supplemental oxygen.
Sec. 25.1445 Equipment standards for the oxygen distributing
system.
Sec. 25.1447 Equipment standards for oxygen dispensing units.
Sec. 25.1449 Means for determining use of oxygen.
Sec. 25.1450 Chemical oxygen generators.
Sec. 25.1451 [Removed. 55 FR 29786, July 20, 1990]
Sec. 25.1453 Protection of oxygen equipment from rupture.
Sec. 25.1455 Draining of fluids subject to freezing.
Sec. 25.1457 Cockpit voice recorders.
Sec. 25.1459 Flight recorders.
Sec. 25.1461 Equipment containing high energy rotors.
Subpart G--Operating Limitations and Information
Sec. 25.1501 General.
Operating Limitations
Sec. 25.1503 Airspeed limitations: general.
Sec. 25.1505 Maximum operating limit speed.
Sec. 25.1507 Maneuvering speed.
Sec. 25.1511 Flap extended speed.
Sec. 25.1513 Minimum control speed.
Sec. 25.1515 Landing gear speeds.
Sec. 25.1519 Weight, center of gravity, and weight
distribution.
Sec. 25.1521 Powerplant limitations.
Sec. 25.1522 Auxiliary power unit limitations.
Sec. 25.1523 Minimum flight crew.
Sec. 25.1525 Kinds of operation.
Sec. 25.1527 Maximum operating altitude.
Sec. 25.1529 Instructions for Continued Airworthiness.
Sec. 25.1531 Maneuvering flight load factors.
Sec. 25.1533 Additional operating limitations.
Markings and Placards
Sec. 25.1541 General.
Sec. 25.1543 Instrument markings: general.
Sec. 25.1545 Airspeed limitation information.
Sec. 25.1547 Magnetic direction indicator.
Sec. 25.1549 Powerplant and auxiliary power unit instruments.
Sec. 25.1551 Oil quantity indication.
Sec. 25.1553 Fuel quantity indicator.
Sec. 25.1555 Control markings.
Sec. 25.1557 Miscellaneous markings and placards.
Sec. 25.1561 Safety equipment.
Sec. 25.1563 Airspeed placard.
Airplane Flight Manual
Sec. 25.1581 General.
Sec. 25.1583 Operating limitations.
Sec. 25.1585 Operating procedures.
Sec. 25.1587 Performance information.
Appendix A
Appendix B
Appendix C to Part 25
Appendix D to Part 25
Appendix E to Part 25
Appendix F to Part 25
Appendix G to Part 25--Continuous Gust Design Criteria
Appendix H to Part 25--Instructions for Continued Airworthiness
Appendix I to Part 25--Installation of an Automatic Takeoff Thrust
Control System (ATTCS)
Appendix J to Part 25--Emergency Demonstration
Special Federal Aviation Regulation No. 13
1. Applicability. Contrary provisions of the Civil Air Regulations
regarding certification notwithstanding,1 this regulation shall provide the
basis for approval by the Administrator of modifications of individual
Douglas DC-3 and Lockheed L-18 airplanes subsequent to the effective date of
this regulation.
NOTE 1 It is not intended to waive compliance with such airworthiness
requirements as are included in the operating parts of the Civil Air
Regulations for specific types of operation.
2. General modifications. Except as modified in sections 3 and 4 of this
regulation, an applicant for approval of modifications to a DC-3 or L-18
airplane which result in changes in design or in changes to approved
limitations shall show that the modifications were accomplished in accordance
with the rules of either Part 4a or Part 4b in effect on September 1, 1953,
which are applicable to the modification being made: Provided, That an
applicant may elect to accomplish a modification in accordance with the rules
of Part 4b in effect on the date of application for the modification in lieu
of Part 4a or Part 4b as in effect on September 1, 1953: And provided
further, That each specific modification must be accomplished in accordance
with all of the provisions contained in the elected rules relating to the
particular modification.
3. Specific conditions for approval. An applicant for any approval of the
following specific changes shall comply with section 2 of this regulation as
modified by the applicable provisions of this section.
(a) Increase in take-off power limitation--1,200 to 1,350 horsepower. The
engine take-off power limitation for the airplane may be increased to more
than 1,200 horsepower but not to more than 1,350 horsepower per engine if the
increase in power does not adversely affect the flight characteristics of the
airplane.
(b) Increase in take-off power limitation to more than 1,350 horsepower.
The engine take-off power limitation for the airplane may be increased to
more than 1,350 horsepower per engine if compliance is shown with the flight
characteristics and ground handling requirements of Part 4b.
(c) Installation of engines of not more than 1,830 cubic inches
displacement and not having a certificated take-off rating of more than 1,350
horsepower. Engines of not more than 1,830 cubic inches displacement and not
having a certificated take-off rating of more than 1,350 horsepower which
necessitate a major modification of redesign of the engine installation may
be installed, if the engine fire prevention and fire protection are
equivalent to that on the prior engine installation.
(d) Installation of engines of more than 1,830 cubic inches displacement or
having certificated take-off rating of more than 1,350 horsepower. Engines of
more than 1,830 cubic inches displacement or having certificated take-off
rating of more than 1,350 horsepower may be installed if compliance is shown
with the engine installation requirements of Part 4b: Provided, That where
literal compliance with the engine installation requirements of Part 4b is
extremely difficult to accomplish and would not contribute materially to the
objective sought, and the Administrator finds that the experience with the
DC-3 or L-18 airplanes justifies it, he is authorized to accept such measures
of compliance as he finds will effectively accomplish the basic objective.
4. Establishment of new maximum certificated weights. An applicant for
approval of new maximum certificated weights shall apply for an amendment of
the airworthiness certificate of the airplane and shall show that the weights
sought have been established, and the appropriate manual material obtained,
as provided in this section.
Note: Transport category performance requirements result in the
establishment of maximum certificated weights for various altitudes.
(a) Weights-25,200 to 26,900 for the DC-3 and 18,500 to 19,500 for the L-
18. New maximum certificated weights of more than 25,200 but not more than
26,900 pounds for DC-3 and more than 18,500 but not more than 19,500 pounds
for L-18 airplanes may be established in accordance with the transport
category performance requirements of either Part 4a or Part 4b, if the
airplane at the new maximum weights can meet the structural requirements of
the elected part.
(b) Weights of more than 26,900 for the DC-3 and 19,500 for the L-18. New
maximum certificated weights of more than 26,900 pounds for DC-3 and 19,500
pounds for L-18 airplanes shall be established in accordance with the
structural performance, flight characteristics, and ground handling
requirements of Part 4b: Provided, That where literal compliance with the
structural requirements of Part 4b is extremely difficult to accomplish and
would not contribute materially to the objective sought, and the
Administrator finds that the experience with the DC-3 or L-18 airplanes
justifies it, he is authorized to accept such measures of compliance as he
finds will effectively accomplish the basic objective.
(c) Airplane flight manual-performance operating information. An approved
airplane flight manual shall be provided for each DC-3 and L-18 airplane
which has had new maximum certificated weights established under this
section. The airplane flight manual shall contain the applicable performance
information prescribed in that part of the regulations under which the new
certificated weights were established and such additional information as may
be necessary to enable the application of the take-off, en route, and landing
limitations prescribed for transport category airplanes in the operating
parts of the Civil Air Regulations.
(d) Performance operating limitations. Each airplane for which new maximum
certificated weights are established in accordance with paragraphs (a) or (b)
of this section shall be considered a transport category airplane for the
purpose of complying with the performance operating limitations applicable to
the operations in which it is utilized.
5. Reference. Unless otherwise provided, all references in this regulation
to Part 4a and Part 4b are those parts of the Civil Air Regulations in effect
on September 1, 1953.
This regulation supersedes Special Civil Air Regulation SR-398 and shall
remain effective until superseded or rescinded by the Board.
[19 FR 5039, Aug. 11, 1954. Redesignated at 29 FR 19099, Dec. 30, 1964]
Subpart A--General
Sec. 25.1 Applicability.
(a) This part prescribes airworthiness standards for the issue of type
certificates, and changes to those certificates, for transport category
airplanes.
(b) Each person who applies under Part 21 for such a certificate or change
must show compliance with the applicable requirements in this part.
Sec. 25.2 Special retroactive requirements.
The following special retroactive requirements are applicable to an
airplane for which the regulations referenced in the type certificate predate
the sections specified below--
(a) Irrespective of the date of application, each applicant for a
supplemental type certificate (or an amendment to a type certificate)
involving an increase in passenger seating capacity to a total greater than
that for which the airplane has been type certificated must show that the
airplane concerned 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), and 25.1359(c) in effect on October 24, 1967, and
(2) Sections 25.803(b) and 25.803(c)(1) in effect on April 23, 1969.
(b) Irrespective of the date of application, each applicant for a
supplemental type certificate (or an amendment to a type certificate) for an
airplane manufactured after October 16, 1987, must show that the airplane
meets the requirements of Sec. 25.807(c)(7) in effect on July 24, 1989.
(c) Compliance with subsequent revisions to the sections specified in
paragraph (a) or (b) above may be elected in accordance with Sec.
21.101(a)(2) of this chapter or may be required in accordance with Sec.
21.101(b) of this chapter.
[Doc. No. 24344, Amdt. 25-72, 55 FR 29773, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Subpart B--Flight
General
Sec. 25.21 Proof of compliance.
(a) Each requirement of this subpart must be met at each appropriate
combination of weight and center of gravity within the range of loading
conditions for which certification is requested. This must be shown--
(1) By tests upon an airplane of the type for which certification is
requested, or by calculations based on, and equal in accuracy to, the results
of testing; and
(2) By systematic investigation of each probable combination of weight and
center of gravity, if compliance cannot be reasonably inferred from
combinations investigated.
(b) [Reserved]
(c) The controllability, stability, trim, and stalling characteristics of
the airplane must be shown for each altitude up to the maximum expected in
operation.
(d) Parameters critical for the test being conducted, such as weight,
loading (center of gravity and inertia), airspeed, power, and wind, must be
maintained within acceptable tolerances of the critical values during flight
testing.
(e) If compliance with the flight characteristics requirements is dependent
upon a stability augmentation system or upon any other automatic or power-
operated system, compliance must be shown with Secs. 25.671 and 25.672.
(f) In meeting the requirements of Secs. 25.105(d), 25.125, 25.233, and
25.237, the wind velocity must be measured at a height of 10 meters above the
surface, or corrected for the difference between the height at which the wind
velocity is measured and the 10-meter height.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5671, Apr. 8, 1970; Amdt. 25-42, 43 FR 2320, Jan. 16, 1978; Amdt. 25-72, 55
FR 29774, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.23 Load distribution limits.
(a) Ranges of weights and centers of gravity within which the airplane may
be safely operated must be established. If a weight and center of gravity
combination is allowable only within certain load distribution limits (such
as spanwise) that could be inadvertently exceeded, these limits and the
corresponding weight and center of gravity combinations must be established.
(b) The load distribution limits may not exceed--
(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 is shown.
Sec. 25.25 Weight limits.
(a) Maximum weights. Maximum weights corresponding to the airplane
operating conditions (such as ramp, ground or water taxi, takeoff, en route,
and landing), environmental conditions (such as altitude and temperature),
and loading conditions (such as zero fuel weight, center of gravity position
and weight distribution) must be established so that they are not more than--
(1) The highest weight selected by the applicant for the particular
conditions; or
(2) The highest weight at which compliance with each applicable structural
loading and flight requirement is shown, except that for airplanes equipped
with standby power rocket engines the maximum weight must not be more than
the highest weight established in accordance with Appendix E of this part; or
(3) The highest weight at which compliance is shown with the certification
requirements of Part 36 of this chapter.
(b) Minimum weight. The minimum weight (the lowest weight at which
compliance with each applicable requirement of this part is shown) must be
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 structural loading condition of this part is shown); or
(3) The lowest weight at which compliance with each applicable flight
requirement is shown.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5671, Apr. 8, 1970; Amdt. 25-63, 53 FR 16365, May 6, 1988]
Sec. 25.27 Center of gravity limits.
The extreme forward and the extreme aft center of gravity limitations must
be established for each 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.
Sec. 25.29 Empty weight and corresponding center of gravity.
(a) The empty weight and corresponding center of gravity must be determined
by weighing the airplane with--
(1) Fixed ballast;
(2) Unusable fuel determined under Sec. 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, lavatory precharge water, and fluids intended for
injection in the engine.
(b) The condition of the airplane at the time of determining empty weight
must be one that is well defined and can be easily repeated.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-42, 43 FR
2320, Jan. 16, 1978; Amdt. 25-72, 55 FR 29774, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.31 Removable ballast.
Removable ballast may be used on showing compliance with the flight
requirements of this subpart.
Sec. 25.33 Propeller speed and pitch limits.
(a) The propeller speed and pitch must be limited to values that will
ensure-
(1) Safe operation under normal operating conditions; and
(2) Compliance with the performance requirements of Secs. 25.101 through
25.125.
(b) There must be a propeller speed limiting means at the governor. It must
limit the maximum possible governed engine speed to a value not exceeding the
maximum allowable r.p.m.
(c) The means used to limit the low pitch position of the propeller blades
must be set so that the engine does not exceed 103 percent of the maximum
allowable engine rpm or 99 percent of an approved maximum overspeed,
whichever is greater, with--
(1) The propeller blades at the low pitch limit and governor inoperative;
(2) The airplane stationary under standard atmospheric conditions with no
wind; and
(3) The engines operating at the takeoff manifold pressure limit for
reciprocating engine powered airplanes or the maximum takeoff torque limit
for turbopropeller engine-powered airplanes.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-57, 49 FR
6848, Feb. 23, 1984; Amdt. 25-72, 55 FR 29774, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Performance
Sec. 25.101 General.
(a) Unless otherwise prescribed, airplanes must meet the applicable
performance requirements of this subpart for ambient atmospheric conditions
and still air.
(b) The performance, as affected by engine power or thrust, must be based
on the following relative humidities;
(1) For turbine engine powered airplanes, a relative humidity of--
(i) 80 percent, at and below standard temperatures; and
(ii) 34 percent, at and above standard temperatures plus 50 deg. F.
Between these two temperatures, the relative humidity must vary linearly.
(2) For reciprocating engine powered airplanes, a relative humidity of 80
percent in a standard atmosphere. Engine power corrections for vapor pressure
must be made in accordance with the following table:
Specific
humidity
Vapor w (Lb.
pressure moisture
Altitude e (In. per lb. Density ratio
H (ft.) Hg.) dry air) <rho>/<sigma>=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
(c) The performance must correspond to the propulsive thrust available
under the particular ambient atmospheric conditions, the particular flight
condition, and t@ relative humidity specified in paragraph (b) of this
section. The available propulsive thrust must correspond to engine power or
thrust, not exceeding the approved power or thrust less--
(1) Installation losses; and
(2) The power or equivalent thrust absorbed by the accessories and services
appropriate to the particular ambient atmospheric conditions and the
particular flight condition.
(d) Unless otherwise prescribed, the applicant must select the takeoff, en
route, approach, and landing configurations for the airplane.
(e) The airplane configurations may vary with weight, altitude, and
temperature, to the extent they are compatible with the operating procedures
required by paragraph (f) of this section.
(f) Unless otherwise prescribed, in determining the accelerate-stop
distances, takeoff flight paths, takeoff distances, and landing distances,
changes in the airplane's configuration, speed, power, and thrust, must be
made in accordance with procedures established by the applicant for operation
in service.
(g) Procedures for the execution of balked landings and missed approaches
associated with the conditions prescribed in Secs. 25.119 and 25.121(d) must
be established.
(h) The procedures established under paragraphs (f) and (g) of this section
must--
(1) Be able to be consistently executed in service by crews of average
skill;
(2) Use methods or devices that are safe and reliable; and
(3) Include allowance for any time delays, in the execution of the
procedures, that may reasonably be expected in service.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 41 FR
55466, Dec. 20, 1976]
Sec. 25.103 Stalling speed.
(a) VS is the calibrated stalling speed, or the minimum steady flight
speed, in knots, at which the airplane is controllable, with--
(1) Zero thrust at the stalling speed, or, if the resultant thrust has no
appreciable effect on the stalling speed, with engines idling and throttles
closed;
(2) Propeller pitch controls (if applicable) in the position necessary for
compliance with paragraph (a)(1) of this section and the airplane in other
respects (such as flaps and landing gear) in the condition existing in the
test in which VS is being used;
(3) The weight used when VS is being used as a factor to determine
compliance with a required performance standard; and
(4) The most unfavorable center of gravity allowable.
(b) The stalling speed VS is the minimum speed obtained as follows:
(1) Trim the airplane for straight flight at any speed not less than 1.2 VS
or more than 1.4 VS At a speed sufficiently above the stall speed to ensure
steady conditions, apply the elevator control at a rate so that the airplane
speed reduction does not exceed one knot per second.
(2) Meet the flight characteristics provisions of Sec. 25.203.
Sec. 25.105 Takeoff.
(a) The takeoff speeds described in Sec. 25.107, the accelerate-stop
distance described in Sec. 25.109, the takeoff path described in Sec. 25.111,
and the takeoff distance and takeoff run described in Sec. 25.113, must be
determined--
(1) At each weight, altitude, and ambient temperature within the
operational limits selected by the applicant; and
(2) In the selected configuration for takeoff.
(b) No takeoff made to determine the data required by this section may
require exceptional piloting skill or alertness.
(c) The takeoff data must be based on--
(1) A smooth, dry, hard-surfaced runway, in the case of land planes and
amphibians;
(2) Smooth water, in the case of seaplanes and amphibians; and
(3) Smooth, dry snow, in the case of skiplanes.
(d) The takeoff data must include, within the established operational
limits of the airplane, the following operational correction factors:
(1) Not more than 50 percent of nominal wind components along the takeoff
path opposite to the direction of takeoff, and not less than 150 percent of
nominal wind components along the takeoff path in the direction of takeoff.
(2) Effective runway gradients.
Sec. 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 selected by the applicant, but may not be less than VmcG
determined under Sec. 25.149(e).
(2) V1, in terms of calibrated airspeed, is the takeoff decision speed
selected by the applicant; however, V1 may not be less than VEF plus the
speed gained with the critical engine inoperative during the time interval
between the instant at which the critical engine is failed, and the instant
at which the pilot recognizes and reacts to the engine failure, as indicated
by the pilot's application of the first retarding means during accelerate-
stop tests.
(b) V2MIN, in terms of calibrated airspeed, may not be less than--
(1) 1.2 VS for--
(i) Two-engine and three-engine turbopropeller and reciprocating engine
powered airplanes; and
(ii) Turbojet powered airplanes without provisions for obtaining a
significant reduction in the one-engine-inoperative power-on stalling speed;
(2) 1.15 VS for--
(i) Turbopropeller and reciprocating engine powered airplanes with more
than three engines; and
(ii) Turbojet powered airplanes with provisions for obtaining a significant
reduction in the one-engine-inoperative power-on stalling speed; and
(3) 1.10 times VMC established under Sec. 25.149.
(c) V2, in terms of calibrated airspeed, must be selected by the applicant
to provide at least the gradient of climb required by Sec. 25.121(b) but may
not be less than--
(1) V2MIN, and
(2) VR plus the speed increment attained (in accordance with Sec. 25.111
(c)(2)) before reaching a height of 35 feet above the takeoff surface.
(d) VMU is the calibrated airspeed at and above which the airplane can
safely lift off the ground, and continue the takeoff. VMU speeds must be
selected by the applicant throughout the range of thrust-to-weight ratios to
be certificated. These speeds may be established from free air data if these
data are verified by ground takeoff tests.
(e) VR, in terms of calibrated airspeed, must be selected in accordance
with the conditions of paragraphs (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 Sec. 25.111(c)(2)) that
allows reaching V2 before reaching a height of 35 feet above the takeoff
surface; or
(iv) A speed that, if the airplane is rotated at its maximum practicable
rate, will result in a VLOF of not less than 110 percent of VMU in the all-
engines-operating condition and not less than 105 percent of VMU determined
at the thrust-to-weight ratio corresponding to the one-engine-inoperative
condition.
(2) For any given set of conditions (such as weight, configuration, and
temperature), a single value of VR, obtained in accordance with this
paragraph, must be used to show compliance with both the one-engine-
inoperative and the all-engines-operating takeoff provisions.
(3) It must be shown that the one-engine-inoperative takeoff distance,
using a rotation speed of 5 knots less than VR established in accordance with
paragraphs (e)(1) and (2) of this section, does not exceed the corresponding
one-engine-inoperative takeoff distance using the established VR. The takeoff
distances must be determined in accordance with Sec. 25.113(a)(1).
(4) Reasonably expected variations in service from the established takeoff
procedures for the operation of the airplane (such as over-rotation of the
airplane and out-of-trim conditions) may not result in unsafe flight
characteristics or in marked increases in the scheduled takeoff distances
established in accordance with Sec. 25.113(a).
(f) VLOF is the calibrated airspeed at which the airplane first becomes
airborne.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 41 FR
55466, Dec. 20, 1976; Amdt. 25-42, 43 FR 2320, Jan. 16, 1978]
Sec. 25.109 Accelerate-stop distance.
(a) The accelerate-stop distance is the greater of the following distances:
(1) The sum of the distances necessary to--
(i) Accelerate the airplane from a standing start to VEF with all engines
operating;
(ii) Accelerate the airplane from VEF to V1 and continue the acceleration
for 2.0 seconds after V1 is reached, assuming the critical engine fails at
VEF; and
(iii) Come to a full stop from the point reached at the end of the
acceleration period prescribed in paragraph (a)(1)(ii) of this section,
assuming that the pilot does not apply any means of retarding the airplane
until that point is reached and that the critical engine is still
inoperative.
(2) The sum of the distances necessary to--
(i) Accelerate the airplane from a standing start to V1 and continue the
acceleration for 2.0 seconds after V1 is reached with all engines operating;
and
(ii) Come to a full stop from the point reached at the end of the
acceleration period prescribed in paragraph (a)(2)(i) of this section,
assuming that the pilot does not apply any means of retarding the airplane
until that point is reached and that all engines are still operating.
(b) Means other than wheel brakes may be used to determine the accelerate-
stop distance if that means--
(1) Is safe and reliable;
(2) Is used so that consistent results can be expected under normal
operating conditions; and
(3) Is such that exceptional skill is not required to control the airplane.
(c) The landing gear must remain extended throughout the accelerate-stop
distance.
(d) If the accelerate-stop distance includes a stopway with surface
characteristics substantially different from those of a smooth hard-surfaced
runway, the takeoff data must include operational correction factors for the
accelerate-stop distance. The correction factors must account for the
particular surface characteristics of the stopway and the variations in these
characteristics with seasonal weather conditions (such as temperature, rain,
snow, and ice) within the established operational limits.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-42, 43 FR
2321, Jan. 16, 1978]
Sec. 25.111 Takeoff path.
(a) The takeoff path extends from a standing start to a point in the
takeoff at which the airplane is 1,500 feet above the takeoff surface, or at
which the transition from the takeoff to the en route configuration is
completed and a speed is reached at which compliance with Sec. 25.121(c) is
shown, whichever point is higher. In addition--
(1) The takeoff path must be based on the procedures prescribed in Sec.
25.101(f);
(2) The airplane must be accelerated on the ground to VEF, at which point
the critical engine must be made inoperative and remain inoperative for the
rest of the takeoff; and
(3) After reaching VEF, the airplane must be accelerated to V2.
(b) During the acceleration to speed V2, the nose gear may be raised off
the ground at a speed not less than VR. However, landing gear retraction may
not be begun until the airplane is airborne.
(c) During the takeoff path determination in accordance with paragraphs (a)
and (b) of this section--
(1) The slope of the airborne part of the takeoff path must be positive at
each point;
(2) The airplane must reach V2 before it is 35 feet above the takeoff
surface and must continue at a speed as close as practical to, but not less
than V2, until it is 400 feet above the takeoff surface;
(3) At each point along the takeoff path, starting at the point at which
the airplane reaches 400 feet above the takeoff surface, the available
gradient 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; and
(4) Except for gear retraction and propeller feathering, the airplane
configuration may not be changed, and no change in power or thrust that
requires action by the pilot may be made, until the airplane is 400 feet
above the takeoff surface.
(d) The takeoff path must be determined by a continuous demonstrated
takeoff or by synthesis from segments. If the takeoff path is determined by
the segmental method--
(1) The segments must be clearly defined and must be related to the
distinct changes in the configuration, power or thrust, and speed;
(2) The weight of the airplane, the configuration, and the power or thrust
must be constant throughout each segment and must correspond to the most
critical condition prevailing in the segment;
(3) The flight path must be based on the airplane's performance without
ground effect; and
(4) The takeoff path data must be checked by continuous demonstrated
takeoffs up to the point at which the airplane is out of ground effect and
its speed is stabilized, to ensure that the path is conservative relative to
the continous path.
The airplane is considered to be out of the ground effect when it reaches a
height equal to its wing span.
(e) For airplanes equipped with standby power rocket engines, the takeoff
path may be determined in accordance with section II of Appendix E.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended 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 FR 29774, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.113 Takeoff distance and takeoff run.
(a) Takeoff distance is the greater of--
(1) The horizontal distance along the takeoff path from the start of the
takeoff to the point at which the airplane is 35 feet above the takeoff
surface, determined under Sec. 25.111; or
(2) 115 percent of the horizontal distance along the takeoff path, with all
engines operating, from the start of the takeoff to the point at which the
airplane is 35 feet above the takeoff surface, as determined by a procedure
consistent with Sec. 25.111.
(b) If the takeoff distance includes a clearway, the takeoff run is the
greater of--
(1) The horizontal distance along the takeoff path from the start of the
takeoff to a point equidistant between the point at which VLOF is reached and
the point at which the airplane is 35 feet above the takeoff surface, as
determined under Sec. 25.111; or
(2) 115 percent of the horizontal distance along the takeoff path, with all
engines operating, from the start of the takeoff to a point equidistant
between the point at which VLOF is reached and the point at which the
airplane is 35 feet above the takeoff surface, determined by a procedure
consistent with Sec. 25.111.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5671, Apr. 8, 1970]
Sec. 25.115 Takeoff flight path.
(a) The takeoff flight path begins 35 feet above the takeoff surface at the
end of the takeoff distance determined in accordance with Sec. 25.113(a).
(b) The net takeoff flight path data must be determined so that they
represent the actual takeoff flight paths (determined in accordance with Sec.
25.111 and with paragraph (a) of this section) reduced at each point by a
gradient of climb equal to--
(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 acceleration along that part of the takeoff flight
path at which the airplane is accelerated in level flight.
Sec. 25.117 Climb: general.
Compliance with the requirements of Secs. 25.119 and 25.121 must be shown
at each weight, altitude, and ambient temperature within the operational
limits established for the airplane and with the most unfavorable center of
gravity for each configuration.
Sec. 25.119 Landing climb: All-engine-operating.
In the landing configuration, the steady gradient of climb may not be less
than 3.2 percent, with--
(a) The engines at the power or thrust that is available eight seconds
after initiation of movement of the power or thrust controls from the minimum
flight idle to the takeoff position; and
(b) A climb speed of not more than 1.3 VS.
Sec. 25.121 Climb: One-engine-inoperative.
(a) Takeoff; landing gear extended. In the critical takeoff configuration
existing along the flight path (between the points at which the airplane
reaches VLOF and at which the landing gear is fully retracted) and in the
configuration used in Sec. 25.111 but without ground effect, the steady
gradient of climb must be positive for two-engine airplanes, and not less
than 0.3 percent for three-engine airplanes or 0.5 percent for four-engine
airplanes, at VLOF and with--
(1) The critical engine inoperative and the remaining engines at the power
or thrust available when retraction of the landing gear is begun in
accordance with Sec. 25.111 unless there is a more critical power operating
condition existing later along the flight path but before the point at which
the landing gear is fully retracted; and
(2) The weight equal to the weight existing when retraction of the landing
gear is begun, determined under Sec. 25.111.
(b) Takeoff; landing gear retracted. In the takeoff configuration existing
at the point of the flight path at which the landing gear is fully retracted,
and in the configuration used in Sec. 25.111 but without ground effect, the
steady gradient of climb may not be less than 2.4 percent for two-engine
airplanes, 2.7 percent for three-engine airplanes, and 3.0 percent for four-
engine airplanes, at V2 and with--
(1) The critical engine inoperative, the remaining engines at the takeoff
power or thrust available at the time the landing gear is fully retracted,
determined under Sec. 25.111, unless there is a more critical power operating
condition existing later along the flight path but before the point where the
airplane reaches a height of 400 feet above the takeoff surface; and
(2) The weight equal to the weight existing when the airplane's landing
gear is fully retracted, determined under Sec. 25.111.
(c) Final takeoff. In the en route configuration at the end of the takeoff
path determined in accordance with Sec. 25.111, the steady gradient of climb
may not be less than 1.2 percent for two-engine airplanes, 1.5 percent for
three-engine airplanes, and 1.7 percent for four-engine airplanes, at not
less than 1.25 VS and with--
(1) The critical engine inoperative and the remaining engines at the
available maximum continuous power or thrust; and
(2) The weight equal to the weight existing at the end of the takeoff path,
determined under Sec. 25.111.
(d) Approach. In the approach configuration corresponding to the normal
all-engines-operating procedure in which VS for this configuration does not
exceed 110 percent of the VS for the related landing configuration, the
steady gradient of climb may not be less than 2.1 percent for two-engine
airplanes, 2.4 percent for three-engine airplanes, and 2.7 percent for four-
engine airplanes, with--
(1) The critical engine inoperative, the remaining engines at the available
takeoff power or thrust;
(2) The maximum landing weight; and
(3) A climb speed established in connection with normal landing procedures,
but not exceeding 1.5 VS.
Sec. 25.123 En route flight paths.
(a) For the en route configuration, the flight paths prescribed in
paragraphs (b) and (c) of this section must be determined at each weight,
altitude, and ambient temperature, within the operating limits established
for the airplane. The variation of weight along the flight path, accounting
for the progressive consumption of fuel and oil by the operating engines, may
be included in the computation. The flight paths must be determined at any
selected speed, 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 adequate cooling in the hot-day condition.
(b) The one-engine-inoperative net flight path data must represent the
actual climb performance diminished by a gradient of climb of 1.1 percent for
two-engine airplanes, 1.4 percent for three-engine airplanes, and 1.6 percent
for four-engine airplanes.
(c) For three- or four-engine airplanes, the two-engine-inoperative net
flight path data must represent the actual climb performance diminished by a
gradient of climb of 0.3 percent for three-engine airplanes and 0.5 percent
for four-engine airplanes.
Sec. 25.125 Landing.
(a) The horizontal distance necessary to land and to come to a complete
stop (or to a speed of approximately 3 knots for water landings) from a point
50 feet above the landing surface must be determined (for standard
temperatures, at each weight, altitude, and wind within the operational
limits established by the applicant for the airplane) as follows:
(1) The airplane must be in the landing configuration.
(2) A stabilized approach, with a calibrated airspeed of not less than
1.3 VS, must be maintained down to the 50 foot height.
(3) Changes in configuration, power or thrust, and speed, must be made in
accordance with the established procedures for service operation.
(4) The landing must be made without excessive vertical acceleration,
tendency to bounce, nose over, ground loop, porpoise, or water loop.
(5) The landings may not require exceptional piloting skill or alertness.
(b) For landplanes and amphibians, the landing distance on land must be
determined on a level, smooth, dry, hard-surfaced runway. In addition--
(1) The pressures on the wheel braking systems may not exceed those
specified by the brake manufacturer;
(2) The brakes may not be used so as to cause excessive wear of brakes or
tires; and
(3) Means other than wheel brakes may be used if that means--
(i) Is safe and reliable;
(ii) Is used so that consistent results can be expected in service; and
(iii) Is such that exceptional skill is not required to control the
airplane.
(c) For seaplanes and amphibians, the landing distance on water must be
determined on smooth water.
(d) For skiplanes, the landing distance on snow must be determined on
smooth, dry, snow.
(e) The landing distance data must include correction factors for not more
than 50 percent of the nominal wind components along the landing path
opposite to the direction of landing, and not less than 150 percent of the
nominal wind components along the landing path in the direction of landing.
(f) If any device is used that depends on the operation of any engine, and
if the landing distance would be noticeably increased when a landing is made
with that engine inoperative, the landing distance must be determined with
that engine inoperative unless the use of compensating means will result in a
landing distance not more than that with each engine operating.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-72, 55 FR
29774, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Controllability and Maneuverability
Sec. 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 transition from one flight
condition to any other flight condition without exceptional piloting skill,
alertness, or strength, and without danger of exceeding the airplane limit-
load factor under any probable 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 engine when the airplane is in the en route, approach, or
landing configuration and is trimmed with the critical engine inoperative;
and
(3) Configuration changes, including deployment or retraction of
deceleration devices.
(c) If, during the testing required by paragraphs (a) and (b) of this
section, marginal conditions exist with regard to required pilot strength,
the "strength of pilots" limits may not exceed the limits prescribed in the
following table:
Values in pound of force
as applied to the control
wheel or rudder pedals Pitch Roll Yaw
For temporary application 75 60 150
For prolonged application 10 5 20
(d) In showing the temporary control force limitations of paragraph (c) of
this section, approved operating procedures or conventional operating
practices must be followed (including being as nearly trimmed as possible at
the next preceding steady flight condition, except that, in the case of
takeoff, the airplane must be trimmed in accordance with approved operating
procedures).
(e) For the purpose of complying with the prolonged control force
limitations of paragraph (c) of this section, the airplane must be as nearly
trimmed as possible.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-42, 43 FR
2321, Jan. 16, 1978]
Sec. 25.145 Longitudinal control.
(a) It must be possible at any speed between the trim speed prescribed in
Sec. 25.103(b)(1) and Vs, to pitch the nose downward so that the acceleration
to this selected trim speed is prompt with--
(1) The airplane trimmed at the trim speed prescribed in Sec. 25.103(b)(1).
(2) The landing gear extended;
(3) The wing flaps (i) retracted and (ii) extended; 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 pounds control force (representative of the maximum temporary
force that readily can be applied by one hand) may be required for the
following maneuvers:
(1) With power off, flaps retracted, and the airplane trimmed at 1.4 VS1,
extend the flaps as rapidly as possible while maintaining the airspeed at
approximately 40 percent above the stalling speed existing at each instant
throughout the maneuver.
(2) Repeat paragraph (b)(1) except initially extend the flaps and then
retract them as rapidly as possible.
(3) Repeat paragraph (b)(2) except with takeoff power.
(4) With power off, flaps retracted, and the airplane trimmed at 1.4 VS1,
apply takeoff power rapidly while maintaining the same airspeed.
(5) Repeat paragraph (b)(4) except with flaps extended.
(6) With power off, flaps extended, and the airplane trimmed at 1.4 VS1,
obtain and maintain airspeeds between 1.1 VS1, and either 1.7 VS1, or VFE,
whichever is lower.
(c) Is must be possible, without exceptional piloting skill, to prevent
loss of altitude when complete retraction of the high lift devices from any
position is begun during steady, straight, level flight at 1.1 VS1 for
propeller powered airplanes, or 1.2 VS1 for turbojet powered airplanes,
with--
(1) Simultaneous application of not more than takeoff power taking into
account the critical engine operating conditions;
(2) The landing gear extended; and
(3) The critical combinations of landing weights and altitudes.
If gated high-lift device control positions are provided, retraction must be
shown from any position from the maximum landing position to the first gated
position, between gated positions, and from the last gated position to the
full retraction position. In addition, the first gated control position from
the landing position must correspond with the high-lift devices configuration
used to establish the go-around procedure from the landing configuration.
Each gated control position must require a separate and distinct motion of
the control to pass through the gated position and must have features to
prevent inadvertent movement of the control through the gated position.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5671, Apr. 8, 1970; Amdt. 25-72, 55 FR 29774, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.147 Directional and lateral control.
(a) Directional control; general. It must be possible, with the wings
level, to yaw into the operative engine and to safely make a reasonably
sudden change in heading of up to 15 degrees in the direction of the critical
inoperative engine. This must be shown at 1.4Vs1 for heading changes up to 15
degrees (except that the heading change at which the rudder pedal force is
150 pounds need not be exceeded), and with--
(1) The critical engine inoperative and its propeller in the minimum drag
position;
(2) The power required for level flight at 1.4 VS1, but not more than
maximum continuous power;
(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 engines must meet the requirements of paragraph (a) of this
section except that--
(1) The two critical engines must be inoperative with their propellers (if
applicable) in the minimum drag 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 deg. banked
turns, with and against the inoperative engine, from steady flight at a speed
equal to 1.4 VS1, with--
(1) The critical engine inoperative and its propeller (if applicable) in
the minimum drag position;
(2) The remaining engines at maximum continuous 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 engines must be able to make 20 deg. banked turns, with and
against the inoperative engines, from steady flight at a speed equal to 1.4
VS1, with maximum continuous power, and with the airplane in the
configuration prescribed by paragraph (b) of this section.
(e) Lateral control; all engines operating. With the engines operating,
roll response must allow normal maneuvers (such as recovery from upsets
produced by gusts and the initiation of evasive maneuvers). There must be
enough excess lateral control in sideslips (up to sideslip angles that might
be required in normal operation), to allow a limited amount of maneuvering
and to correct for gusts. Lateral control must be enough at any speed up to
VFC/MFC to provide a peak roll rate necessary for safety, without excessive
control forces or travel.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-42, 43 FR
2321, Jan. 16, 1978; Amdt. 25-72, 55 FR 29774, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.149 Minimum control speed.
(a) In establishing the minimum control speeds required by this section,
the method used to simulate critical engine failure must represent the most
critical mode of powerplant failure with respect to controllability expected
in service.
(b) VMC is the calibrated airspeed at which, when the critical engine is
suddenly made inoperative, it is possible to maintain control of the airplane
with that engine still inoperative and maintain straight flight with an angle
of bank of not more than 5 degrees.
(c) VMC may not exceed 1.2 VS 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 configuration existing along
the flight path after the airplane becomes airborne, except with the landing
gear retracted;
(6) The airplane airborne and the ground effect negligible; and
(7) If applicable, the propeller of the inoperative engine--
(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 compliance with the climb requirements of Sec. 25.121.
(d) The rudder forces required to maintain control at VMC may not exceed
150 pounds nor may it be necessary to reduce power or thrust of the operative
engines. During recovery, the airplane may not assume any dangerous attitude
or require exceptional piloting skill, alertness, or strength to prevent a
heading change of more than 20 degrees.
(e) VMCG, the minimum control speed on the ground, is the calibrated
airspeed during the takeoff run at which, when the critical engine is
suddenly made inoperative, it is possible to maintain control of the airplane
using the rudder control alone (without the use of nosewheel steering), as
limited by 150 pounds of force, and the lateral control to the extent of
keeping the wings level to enable the takeoff to be safely continued using
normal piloting skill. In the determination of VMCG, assuming that the path
of the airplane accelerating with all engines operating is along the
centerline of the runway, its path from the point at which the critical
engine is made inoperative to the point at which recovery to a direction
parallel to the centerline is completed may not deviate more than 30 feet
laterally from the centerline at any point. VMCG must be established with--
(1) The airplane in each takeoff configuration or, at the option of the
applicant, in the most critical takeoff configuration;
(2) Maximum available takeoff power or thrust on 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 landing approach with all
engines operating, is the calibrated airspeed at which, when the critical
engine is suddenly made inoperative, it is possible to maintain control of
the airplane with that engine still inoperative and maintain straight flight
with an angle of bank of not more than 5 degrees. VMCL must be established
with--
(1) The airplane in the most critical configuration for approach with all
engines operating;
(2) The most unfavorable center of gravity;
(3) The airplane trimmed for approach with all engines operating;
(4) The maximum sea level landing weight (or any lesser weight necessary to
show VMCL); and
(5) Maximum available takeoff power or thrust on the operating engines.
(g) For airplanes with three or more engines, VMCL-2, the minimum control
speed during landing approach with one critical engine inoperative, is the
calibrated airspeed at which, when a second critical engine is suddenly made
inoperative, it is possible to maintain control of the airplane with both
engines still inoperative and maintain straight flight with an angle of bank
of not more than 5 degrees. VMCL-2 must be established with--
(1) The airplane in the most critical configuration for approach with the
critical engine inoperative;
(2) The most unfavorable center of gravity;
(3) The airplane trimmed for approach with the critical engine inoperative;
(4) The maximum sea level landing weight (or any lesser weight necessary to
show VMCL-2);
(5) The power or thrust on the operating engines required to maintain an
approach path angle of 3 degrees when one critical engine is inoperative; and
(6) The power or thrust on the operating engines rapidly changed,
immediately after the second critical engine is made inoperative, from the
power or thrust prescribed in paragraph (g)(5) of this section to--
(i) Minimum available power or thrust; and
(ii) Maximum available takeoff power or thrust.
(h) The rudder control forces required to maintain control at VMCL and
VMCL-2 may not exceed 150 pounds, nor may it be necessary to reduce the power
or thrust of the operating engines. In addition, the airplane may not assume
any dangerous attitudes or require exceptional piloting skill, alertness, or
strength to prevent a divergence in the approach flight path that would
jeopardize continued safe approach when--
(1) The critical engine is suddenly made inoperative; and
(2) For the determination of VMCL-2, the power or thrust on the operating
engines is changed in accordance with paragraph (g)(6) of this section.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended 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]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Trim
Sec. 25.161 Trim.
(a) General. Each airplane must meet the trim requirements of this section
after being trimmed, and without further pressure upon, or movement of,
either the primary controls or their corresponding trim controls by the pilot
or the automatic pilot.
(b) Lateral and directional trim. The airplane must maintain lateral and
directional trim with the most adverse lateral displacement of the center of
gravity within the relevant operating limitations, during normally expected
conditions of operation (including operation at any speed from 1.4 VS1 to
VMO/MMO).
(c) Longitudinal trim. The airplane must maintain longitudinal trim
during--
(1) A climb with maximum continuous power at a speed not more than 1.4 VS1,
with the landing gear retracted, and the flaps (i) retracted and (ii) in the
takeoff position;
(2) A glide with power off at a speed not more than 1.4 VS1, with the
landing gear extended, the wing flaps (i) retracted and (ii) extended, the
most unfavorable center of gravity position approved for landing with the
maximum landing weight, and with the most unfavorable center of gravity
position approved for landing regardless of weight; and
(3) Level flight at any speed from 1.4 VS1, to VMO/MMO, with the landing
gear and flaps retracted, and from 1.4 VS1 to VLE with the landing gear
extended.
(d) Longitudinal, directional, and lateral trim. The airplane must maintain
longitudinal, directional, and lateral trim (and for the lateral trim, the
angle of bank may not exceed five degrees) at 1.4 VS1 during climbing flight
with--
(1) The critical engine inoperative;
(2) The remaining engines at maximum continuous power; and
(3) The landing gear and flaps retracted.
(e) Airplanes with four or more engines. Each airplane with four or more
engines must maintain trim in rectilinear flight--
(1) At the climb speed, configuration, and power required by Sec. 25.123(a)
for the purpose of establishing the rate of climb;
(2) With the most unfavorable center of gravity position; and
(3) At the weight at which the two-engine-inoperative climb is equal to at
least 0.013 VS02 at an altitude of 5,000 feet.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5671, Apr. 8, 1970; Amdt. 25-38, 41 FR 55466, Dec. 20, 1976]
Stability
Sec. 25.171 General.
The airplane must be longitudinally, directionally, and laterally stable in
accordance with the provisions of Secs. 25.173 through 25.177. In addition,
suitable stability and control feel (static stability) is required in any
condition normally encountered in service, if flight tests show it is
necessary for safe operation.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-7, 30 FR
13117, Oct. 15, 1965]
Sec. 25.173 Static longitudinal stability.
Under the conditions specified in Sec. 25.175, the characteristics of the
elevator control forces (including friction) must be as follows:
(a) A pull must be required to obtain and maintain speeds below the
specified trim speed, and a push must be required to obtain and maintain
speeds above the specified trim speed. This must be shown at any speed that
can be obtained except speeds higher than the landing gear or wing flap
operating limit speeds or VFC/MFC, whichever is appropriate, or lower than
the minimum speed for steady unstalled flight.
(b) The airspeed must return to within 10 percent of the original trim
speed for the climb, approach, and landing conditions specified in Sec.
25.175 (a), (c), and (d), and must return to within 7.5 percent of the
original trim speed for the cruising condition specified in Sec. 25.175(b),
when the control force is slowly released from any speed within the range
specified in paragraph (a) of this section.
(c) The average gradient of the stable slope of the stick force versus
speed curve may not be less than 1 pound for each 6 knots.
(d) Within the free return speed range specified in paragraph (b) of this
section, it is permissible for the airplane, without control forces, to
stabilize on speeds above or below the desired trim speeds if exceptional
attention on the part of the pilot is not required to return to and maintain
the desired trim speed and altitude.
[Amdt. 25-7, 30 FR 13117, Oct. 15, 1965]
Sec. 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 percent 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 power or thrust selected by the applicant as an operating
limitation 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 than 1.4 VS1.
(b) Cruise. Static longitudinal stability must be shown in the cruise
condition as follows:
(1) With the landing gear retracted at high speed, the stick force curve
must have a stable slope at all speeds within a range which is the greater of
15 percent of the trim speed plus the resulting free return speed range, or
50 knots plus the resulting free return speed range, above and below the trim
speed (except that the speed range need not include speeds less than 1.4 VS1,
nor speeds greater than VFC/MFC, nor speeds that require a stick force of
more than 50 pounds), with--
(i) The wing flaps retracted;
(ii) The center of gravity in the most adverse position (see Sec. 25.27);
(iii) The most critical weight between the maximum 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 applicant as
an operating limitation (see Sec. 25.1521), except that the power need not
exceed that required at VMO/MMO; and
(v) The airplane trimmed for level flight with the power required in
paragraph (b)(1)(iv) of this section.
(2) With the landing gear retracted at low speed, the stick force curve
must have a stable slope at all speeds within a range which is the greater of
15 percent of the trim speed plus the resulting free return speed range, or
50 knots plus the resulting free return speed range, above and below the trim
speed (except that the speed range need not include speeds less than 1.4 VS1,
nor speeds greater than the minimum speed of the applicable speed range
prescribed in paragraph (b)(1), nor speeds that require a stick force of more
than 50 pounds), with--
(i) Wing flaps, center of gravity position, and weight as specified in
paragraph (b)(1) of this section;
(ii) Power required for level flight at a speed equal to VMO + 1.4 VS1/2;
and
(iii) The airplane trimmed for level flight with the power required in
paragraph (b)(2)(ii) of this section.
(3) With the landing gear extended, the stick force curve must have a
stable slope at all speeds within a range which is the greater of 15 percent
of the trim speed plus the resulting free return speed range, or 50 knots
plus the resulting free return speed range, above and below the trim speed
(except that the speed range need not include speeds less than 1.4 VS1, nor
speeds greater than VLE, nor speeds that require a stick force of more than
50 pounds), with--
(i) Wing flap, center of gravity position, and weight as specified in
paragraph (b)(1) of this section;
(ii) 75 percent of maximum continuous power for reciprocating engines or,
for turbine engines, the maximum cruising power selected by the applicant as
an operating limitation, except that the power need not exceed that required
for level flight at VLE; and
(iii) The aircraft trimmed for level flight with the power required in
paragraph (b)(3)(ii) of this section.
(c) Approach. The stick force curve must have a stable slope at speeds
between 1.1 VS1 and 1.8 VS1, with--
(1) Wing flaps in the approach position;
(2) Landing gear retracted;
(3) Maximum landing weight; and
(4) The airplane trimmed at 1.4 VS1 with enough power to maintain level
flight at this speed.
(d) Landing. The stick force curve must have a stable slope, and the stick
force may not exceed 80 pounds, at speeds between 1.1 VS0 and 1.3 VS0 with--
(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 1.4 VS0 with power or thrust off.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-7, 30 FR
13117, Oct. 15, 1965]
Sec. 25.177 Static lateral-directional stability.
(a) [Reserved]
(b) [Reserved]
(c) In straight, steady sideslips, the aileron and rudder control movements
and forces must be substantially proportional to the angle of sideslip in a
stable sense; and the factor of proportionality must lie between limits found
necessary for safe operation throughout the range of sideslip angles
appropriate to the operation of the airplane. At greater angles, up to the
angle at which full rudder is used or a rudder force of 180 pounds is
obtained, the rudder pedal forces may not reverse; and increased rudder
deflection must be needed for increased angles of sideslip. Compliance with
this paragraph must be demonstrated for all landing gear and flap positions
and symmetrical power conditions at speeds from 1.2 VS1 to VFE, VLE, or VFC/
MFC, as appropriate.
(d) The rudder gradi@ts must meet the requirements of paragraph (c) at
speeds between VMO/MMO and VFC/MFC except that the dihedral effect (aileron
deflection opposite the corresponding rudder input) may be negative provided
the divergence is gradual, easily recognized, and easily controlled by the
pilot.
[Doc. No. 24344, Amdt. 25-72, 55 FR 29774, July 20, 1990; 55 FR 37607,
Sept. 12, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.181 Dynamic stability.
(a) Any short period oscillation, not including combined lateral-
directional oscillations, occurring between 1.2 VS and maximum allowable
speed appropriate to the configuration of the airplane must be heavily
damped with the primary controls--
(1) Free; and
(2) In a fixed position.
(b) Any combined lateral-directional oscillations ("Dutch roll") occurring
between 1.2 VS and maximum allowable speed appropriate to the configuration
of the airplane must be positively damped with controls free, and must be
controllable with normal use of the primary controls without requiring
exceptional pilot skill.
[Amdt. 25-42, 43 FR 2322, Jan. 16, 1978, as amended by Amdt. 25-72, 55 FR
29775, July 20, 1990; 55 FR 37607, Sept. 12, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Stalls
Sec. 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 1.6 VS1 (where VS1
corresponds to the stalling speed with flaps in the approach position, the
landing gear retracted, and maximum landing weight).
(b) In either condition required by paragraph (a) of this section, it must
be possible to meet the applicable requirements of Sec. 25.203 with--
(1) Flaps and landing gear in any likely combination of positions;
(2) Representative weights within the range for which certification is
requested; and
(3) The most adverse center of gravity for recovery.
(c) The following procedure must be used to show compliance with Sec.
25.203:
(1) With the airplane trimmed for straight flight at the speed prescribed
in Sec. 25.103(b)(1), reduce the speed with the elevator control until it is
steady at slightly above stalling speed. Apply elevator control so that the
speed reduction does not exceed one knot per second until (i) the airplane is
stalled, or (ii) the control reaches the stop.
(2) As soon as the airplane is stalled, recover by normal recovery
techniques.
(d) Occurrence of stall is defined as follows:
(1) The airplane may be considered stalled when, at an angle of attack
measurably greater than that for maximum lift, the inherent flight
characteristics give a clear and distinctive indication to the pilot that the
airplane is stalled. Typical indications of a stall, occurring either
individually or in combination, are--
(i) A nose-down pitch that cannot be readily arrested;
(ii) A roll that cannot be readily arrested; or
(iii) If clear enough, a loss of control effectiveness, an abrupt change in
control force or motion, or a distinctive shaking of the pilot's controls.
(2) For any configuration in which the airplane demonstrates an
unmistakable inherent aerodynamic warning of a magnitude and severity that is
a strong and effective deterrent to further speed reduction, the airplane may
be considered stalled when it reaches the speed at which the effective
deterrent is clearly manifested.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 41 FR
55466, Dec. 20, 1976; Amdt. 25-42, 43 FR 2322, Jan. 16, 1978]
Sec. 25.203 Stall characteristics.
(a) It must be possible to produce and to correct roll and yaw by
unreversed use of the aileron and rudder controls, up to the time the
airplane is stalled. No abnormal nose-up pitching may occur. The longitudinal
control force must be positive up to and throughout the stall. In addition,
it must be possible to promptly prevent stalling and to recover from a stall
by normal use of the controls.
(b) For level wing stalls, the roll occurring between the stall and the
completion of the recovery may not exceed approximately 20 degrees.
(c) For turning flight stalls, the action of the airplane after the stall
may not be so violent or extreme as to make it difficult, with normal
piloting skill, to effect a prompt recovery and to regain control of the
airplane.
Sec. 25.205 [Removed. Amdt. 25-72, 55 FR 29775, July 20, 1990]
EDITORIAL NOTE: For the convenience of the user, the removed text is
set out below.
Sec. 25.205 Stalls: Critical engine inoperative.
(a) It must be possible to safely recover from a stall with the critical
engine inoperative--
(1) Without applying power to the inoperative engine;
(2) With flaps and landing gear retracted; and
(3) With the remaining engines at up to 75 percent of maximum continuous
power, or up to the power at which the wings can be held level with the use
of maximum control travel, whichever is less.
(b) The operating engines may be throttled back during stall recovery from
stalls with the critical engine inoperative.
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.207 Stall warning.
(a) Stall warning with sufficient margin to prevent inadvertent stalling
with the flaps and landing gear in any normal position must be clear and
distinctive to the pilot in straight and turning flight.
(b) The warning may be furnished either through the inherent aerodynamic
qualities of the airplane or by a device that will give clearly
distinguishable indications under expected conditions of flight. However, a
visual stall warning device that requires the attention of the crew within
the cockpit is not acceptable by itself. If a warning device is used, it must
provide a warning in each of the airplane configuations prescribed in
paragraph (a) of this section at the speed prescribed in paragraph (c) of
this section.
(c) The stall warning must begin at a speed exceeding the stalling speed
(i.e., the speed at which the airplane stalls or the minimum speed
demonstrated, whichever is applicable under the provisions of Sec. 25.201(d))
by seven percent or at any lesser margin if the stall warning has enough
clarity, duration, distinctiveness, or similar properties.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-7, 30 FR
13118, Oct. 15, 1965; Amdt. 25-42, 43 FR 2322, Jan. 16, 1978]
Ground and Water Handling Characteristics
Sec. 25.231 Longitudinal stability and control.
(a) Landplanes may have no uncontrollable tendency to nose over in any
reasonably expected operating condition or when rebound occurs during landing
or takeoff. In addition--
(1) Wheel brakes must operate smoothly and may not cause any undue tendency
to nose over; and
(2) If a tail-wheel landing gear is used, it must be possible, during the
takeoff ground run on concrete, to maintain any altitude up to thrust line
level, at 80 percent of VS1.
(b) For seaplanes and amphibians, the most adverse water conditions safe
for takeoff, taxiing, and landing, must be established.
Sec. 25.233 Directional stability and control.
(a) There may be no uncontrollable ground-looping tendency in 90 deg. cross
winds, up to a wind velocity of 20 knots or 0.2 VS0, whichever is greater,
except that the wind velocity need not exceed 25 knots. At any speed at which
the airplane may be expected to be operated on the ground. This may be shown
while establishing the 90 deg. cross component of wind velocity required by
Sec. 25.237.
(b) Landplanes must be satisfactorily controllable, without exceptional
piloting skill or alertness, in power-off landings at normal landing speed,
without using brakes or engine power to maintain a straight path. This may be
shown during power-off landings made in conjunction with other tests.
(c) The airplane must have adequate directional control during taxiing.
This may be shown during taxiing prior to takeoffs made in conjunction with
other tests.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5671, Apr. 8, 1970; Amdt. 25-42, 43 FR 2322, Jan. 16, 1978]
Sec. 25.235 Taxiing condition.
The shock absorbing mechanism may not damage the structure of the airplane
when the airplane is taxied on the roughest ground that may reasonably be
expected in normal operation.
Sec. 25.237 Wind velocities.
(a) For landplanes and amphibians, a 90-degree cross component of wind
velocity, demonstrated to be safe for takeoff and landing, must be
established for dry runways and must be at least 20 knots or 0.2 VS0,
whichever is greater, except that it need not exceed 25 knots.
(b) For seaplanes and amphibians, the following applies:
(1) A 90-degree cross component of wind velocity, up to which takeoff and
landing is safe under all water conditions that may reasonably be expected in
normal operation, must be established and must be at least 20 knots or 0.2
Vs0, whichever is greater, except that it need not exceed 25 knots.
(2) A wind velocity, for which taxiing is safe in any direction under all
water conditions that may reasonably be expected in normal operation, must be
established and must be at least 20 knots or 0.2 VS0, whichever is greater,
except that it need not exceed 25 knots.
[Amdt. 25-42, 43 FR 2322, Jan. 16, 1978]
Sec. 25.239 Spray characteristics, control, and stability on water.
(a) For seaplanes and amphibians, during takeoff, taxiing, and landing, and
in the conditions set forth in paragraph (b) of this section, there may be
no--
(1) Spray characteristics that would impair the pilot's view, cause damage,
or result in the taking in 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 designed to withstand the resulting water loads.
(b) Compliance with the requirements of paragraph (a) of this section must
be shown--
(1) In water conditions, from smooth to the most adverse condition
established in accordance with Sec. 25.231;
(2) In wind and cross-wind velocities, water currents, and associated waves
and swells that may reasonably be expected in operation on water;
(3) At speeds that may reasonably be expected in operation on water;
(4) With sudden failure of the critical engine at any time while on water;
and
(5) At each weight and center of gravity position, relevant to each
operating condition, within the range of loading conditions for which
certification is requested.
(c) In the water conditions of paragraph (b) of this section, and in the
corresponding wind conditions, the seaplane or amphibian must be able to
drift for five minutes with engines inoperative, aided, if necessary, by a
sea anchor.
Miscellaneous Flight Requirements
Sec. 25.251 Vibration and buffeting.
(a) The airplane must be demonstrated in flight to be free from any
vibration and buffeting that would prevent continued safe flight in any
likely operating condition.
(b) Each part of the airplane must be demonstrated in flight to be free
from excessive vibration under any appropriate speed and power conditions up
to VDF/MDF. The maximum speeds shown must be used in establishing the
operating limitations of the airplane in accordance with Sec. 25.1505.
(c) Except as provided in paragraph (d) of this section, there may be no
buffeting condition, in normal flight, including configuration changes during
cruise, severe enough to interfere with the control of the airplane, to cause
excessive fatigue to the crew, or to cause structural damage. Stall warning
buffeting within these limits is allowable.
(d) There may be no perceptible buffeting condition in the cruise
configuration in straight flight at any speed up to VMO/MMO, except that
stall warning buffeting is allowable.
(e) For an airplane with MD greater than .6 or with a maximum operating
altitude greater than 25,000 feet, the positive maneuvering load factors at
which the onset of perceptible buffeting occurs must be determined with the
airplane in the cruise configuration for the ranges of airspeed or Mach
number, weight, and altitude for which the airplane is to be certificated.
The envelopes of load factor, speed, altitude, and weight must provide a
sufficient range of speeds and load factors for normal operations. Probable
inadvertent excursions beyond the boundaries of the buffet onset envelopes
may not result in unsafe conditions.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5671, Apr. 8, 1970; Amdt. 25-72, 55 FR 29775, July 20, 1990; Amdt. 25-77, 57
FR 28949, June 29, 1992]
*****************************************************************************
57 FR 28946, No. 125, June 29, 1992
SUMMARY: This amendment revises the airworthiness standards of the Federal
Aviation Regulations (FAR) for type certification of transport category
airplanes concerning vibration, buffet, flutter and divergence. It clarifies
the requirement to consider flutter and divergence when treating certain
damage and failure conditions required by other sections of the FAR and
adjusts the safety margins related to aeroelastic stabiity to make them more
appropriate for the conditions to which they apply. These changes are made to
provide consistency with other sections of the FAR and to take into account
advances in technology and the evolution of the design of transport
airplanes.
EFFECTIVE DATE: July 29, 1992.
*****************************************************************************
Sec. 25.253 High-speed characteristics.
(a) Speed increase and recovery characteristics. The following speed
increase and recovery characteristics must be met:
(1) Operating conditions and characteristics likely to cause inadvertent
speed increases (including upsets in pitch and roll) must be simulated with
the airplane trimmed at any likely cruise speed up to VMO/MMO. These
conditions and characteristics include gust upsets, inadvertent control
movements, low stick force gradient in relation to control friction,
passenger movement, leveling off from climb, and descent from Mach to
airspeed limit altitudes.
(2) Allowing for pilot reaction time after effective inherent or artificial
speed warning occurs, it must be shown that the airplane can be recovered to
a normal attitude and its speed reduced to VMO/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 for recovery.
(3) With the airplane trimmed at any speed up to VMO /MMO, there must be no
reversal of the response to control input about any axis at any speed up to
VDF/MDF. Any tendency to pitch, roll, or yaw must be mild and readily
controllable, using normal piloting techniques. When the airplane is trimmed
at VMO/MMO, the slope of the elevator control force versus speed curve need
not be stable at speeds greater than VFC/MFC, but there must be a push force
at all speeds up to VDF/MDF and there must be no sudden or excessive
reduction of elevator control force as VDF/MDF is reached.
(b) Maximum speed for stability characteristics, VFC/MFC. VFC/MFC is the
maximum speed at which the requirements of Secs. 25.147(e), 25.175(b)(1),
25.177, and 25.181 must be met with flaps and landing gear retracted. It may
not be less than a speed midway between VMO/MMO and VDF/MDF, except that, for
altitudes where Mach number is the limiting factor, MFC need not exceed the
Mach number at which effective speed warning occurs.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5671, Apr. 8, 1970; Amdt. 25-54, 45 FR 60172, Sept. 11, 1980; Amdt. 25-72, 55
FR 29775, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.255 Out-of-trim characteristics.
(a) From an initial condition with the airplane trimmed at cruise speeds up
to VMO/MMO, the airplane must have satisfactory maneuvering stability and
controllability with the degree of out-of-trim in both the airplane nose-up
and nose-down directions, which results from the greater of--
(1) A three-second movement of the longitudinal trim sy@em at its normal
rate for the particular flight condition with no aerodynamic load (or an
equivalent degree of trim for airplanes that do not have a power-operated
trim system), except as limited by stops in the trim system, including those
required by Sec. 25.655(b) for adjustable stabilizers; or
(2) The maximum mistrim that can be sustained by the autopilot while
maintaining level flight in the high speed cruising condition.
(b) In the out-of-trim condition specified in paragraph (a) of this
section, when the normal acceleration is varied from +1 g to the positive and
negative values specified in paragraph (c) of this section--
(1) The stick force vs. g curve must have a positive slope at any speed up
to and including VFC/MFC; and
(2) At speeds between VFC/MFC and VDF/MDF the direction of the primary
longitudinal control force may not reverse.
(c) Except as provided in paragraphs (d) and (e) of this section,
compliance with the provisions of paragraph (a) of this section must be
demonstrated 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 acceptable 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 compliance and marginal conditions exist during flight test
with regard to reversal of primary longitudinal control force, flight tests
must be accomplished from the normal acceleration at which a marginal
condition is found to exist to the applicable limit specified in paragraph
(b)(1) of this section.
(e) During flight tests required by paragraph (a) of this section, the
limit maneuvering load factors prescribed in Secs. 25.333(b) and 25.337, and
the maneuvering load factors associated with probable inadvertent excursions
beyond the boundaries of the buffet onset envelopes determined under Sec.
25.251(e), need not be exceeded. In addition, the entry speeds for flight
test demonstrations at normal acceleration values less than 1 g must be
limited to the extent necessary to accomplish a recovery without exceeding
VDF/MDF.
(f) In the out-of-trim condition specified in paragraph (a) of this
section, it must be possible from an overspeed condition at VDF/MDF to
produce at least 1.5 g for recovery by applying not more than 125 pounds of
longitudinal control force using either the primary longitudinal control
alone or the primary longitudinal control and the longitudinal trim system.
If the longitudinal trim is used to assist in producing the required load
factor, it must be shown at VDF/MDF that the longitudinal trim can be
actuated in the airplane nose-up direction with the primary surface loaded to
correspond to the least of the following airplane nose-up control forces:
(1) The maximum control forces expected in service as specified in Secs.
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 a strong deterrent to further application of primary
longitudinal control force.
[Amdt. No. 25-42, 43 FR 2322, Jan. 16, 1978]
Subpart C--Structure
General
Sec. 25.301 Loads.
(a) Strength requirements are specified in terms of limit loads (the
maximum loads to be expected in service) and ultimate loads (limit loads
multiplied by prescribed factors of safety). Unless otherwise provided,
prescribed loads are limit loads.
(b) Unless otherwise provided, the specified air, ground, and water loads
must be placed in equilibrium with inertia forces, considering each item of
mass in the airplane. These loads must be distributed to conservatively
approximate or closely represent actual conditions. Methods used to determine
load intensities and distribution must be validated by flight load
measurement unless the methods used for determining those loading conditions
are shown to be reliable.
(c) If deflections under load would significantly change the distribution
of external or internal loads, this redistribution must be taken into
account.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5672, Apr. 8, 1970]
Sec. 25.303 Factor of safety.
Unless otherwise specified, a factor of safety of 1.5 must be applied to
the prescribed limit load which are considered external loads on the
structure. When a loading condition is prescribed in terms of ultimate loads,
a factor of safety need not be applied unless otherwise specified.
[Amdt. 25-23, 35 FR 5672, Apr. 8, 1970]
Sec. 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 may not
interfere with safe operation.
(b) The structure must be able to support ultimate loads without failure
for at least 3 seconds. However, when proof of strength is shown by dynamic
tests simulating actual load conditions, the 3-second limit does not apply.
Static tests conducted to ultimate load must include the ultimate deflections
and ultimate deformation induced by the loading. When analytical methods are
used to show compliance with the ultimate load strength requirements, it must
be shown that--
(1) The effects of deformation are not significant;
(2) The deformations involved are fully accounted for in the analysis; or
(3) The methods and assumptions used are sufficient to cover the effects of
these deformations.
(c) Where structural flexibility is such that any rate of load application
likely to occur in the operating conditions might produce transient stresses
appreciably higher than those corresponding to static loads, the effects of
this rate of application must be considered.
(d) The dynamic response of the airplane to vertical and lateral continuous
turbulence must be taken into account. The continuous gust design criteria of
Appendix G of this part must be used to establish the dynamic response unless
more rational criteria are shown.
(e) The airplane must be designed to withstand any vibration and buffeting
that might occur in any likely operating condition up to VD/MD, including
stall and probable inadvertent excursions beyond the boundaries of the buffet
onset envelope. This must be shown by analysis, flight tests, or other tests
found necessary by the Administrator.
(f) Unless shown to be extremely improbable, the airplane must be designed
to withstand any forced structural vibration resulting from any failure,
malfunction or adverse condition in the flight control system. These must be
considered limit loads and must be investigated at airspeeds up to VC/MC.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5672, Apr. 8, 1970; Amdt. 25-54, 45 FR 60172, Sept. 11, 1980; Amdt. 25-77, 57
FR 28949, June 29, 1992]
*****************************************************************************
57 FR 28946, No. 125, June 29, 1992
SUMMARY: This amendment revises the airworthiness standards of the Federal
Aviation Regulations (FAR) for type certification of transport category
airplanes concerning vibration, buffet, flutter and divergence. It clarifies
the requirement to consider flutter and divergence when treating certain
damage and failure conditions required by other sections of the FAR and
adjusts the safety margins related to aeroelastic stabiity to make them more
appropriate for the conditions to which they apply. These changes are made to
provide consistency with other sections of the FAR and to take into account
advances in technology and the evolution of the design of transport
airplanes.
EFFECTIVE DATE: July 29, 1992.
*****************************************************************************
Sec. 25.307 Proof of structure.
(a) Compliance with the strength and deformation requirements of this
subpart must be shown for each critical loading condition. Structural
analysis may be used only if the structure conforms to that for which
experience has shown this method to be reliable. The Administrator may
require ultimate load tests in cases where limit load tests may be
inadequate.
(b) [Reserved]
(c) [Reserved]
(d) When static or dynamic tests are used to show compliance with the
requirements of Sec. 25.305(b) for flight structures, appropriate material
correction factors must be applied to the test results, unless the structure,
or part thereof, being tested has features such that a number of elements
contribute to the total strength of the structure and the failure of one
element results in the redistribution of the load through alternate load
paths.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5672, Apr. 8, 1970; Amdt. 25-54, 45 FR 60172, Sept. 11, 1980; Amdt. 25-72, 55
FR 29775, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Flight Loads
Sec. 25.321 General.
(a) Flight load factors represent the ratio of the aerodynamic force
component (acting normal to the assumed longitudinal axis of the airplane) to
the weight of the airplane. A positive load factor is one in which the
aerodynamic force acts upward with respect to the airplane.
(b) Considering compressibility effects at each speed, compliance with the
flight load requirements of this subpart must be shown--
(1) At each critical altitude within the range of altitudes selected by the
applicant;
(2) At each weight from the design minimum weight to the design maximum
weight appropriate to each particular flight load condition; and
(3) For each required altitude and weight, for any practicable distribution
of disposable load within the operating limitations recorded in the Airplane
Flight Manual.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5672, Apr. 8, 1970]
Flight Maneuver and Gust Conditions
Sec. 25.331 General.
(a) Procedure. The analysis of symmetrical flight must include at least the
conditions specified in paragraphs (b) through (d) of this section. The
following procedure must be used:
(1) Enough points on the maneuvering and gust envelopes must be
investigated to ensure that the maximum load for each part of the airplane
structure is obtained. A conservative combined envelope may be used.
(2) The significant forces acting on the airplane must be placed in
equilibrium in a rational or conservative manner. The linear inertia forces
must be considered in equilibrium with thrust and all aerodynamic loads,
while the angular (pitching) inertia forces must be considered in equilibrium
with thrust and all aerodynamic moments, including moments due to loads on
components such as tail surfaces and nacelles. Critical thrust values in the
range from zero to maximum continuous thrust must be considered.
(3) Where sudden displacement of a control is specified, the assumed rate
of control surface displacement may not be less than the rate that could be
applied by the pilot through the control system.
(4) In determining elevator angles and chordwise load distribution (in the
maneuvering conditions of paragraphs (b) and (c) of this section) in turns
and pullups, the effect of corresponding pitching velocities must be taken
into account. The in-trim and out-of-trim flight conditions specified in Sec.
25.255 must be considered.
(b) Maneuvering balanced conditions. Assuming the airplane to be in
equilibrium with zero pitching acceleration, the maneuvering conditions A
through I on the maneuvering envelope in Sec. 25.333(b) must be investigated.
(c) Maneuvering pitching conditions. The following conditions involving
pitching acceleration must be investigated:
(1) Maximum elevator displacement at VA. The airplane is assumed to be
flying in steady level flight (point A1, Sec. 25.333(b)) and, except as
limited by pilot effort in accordance with Sec. 25.397(b), the pitching
control is suddenly moved to obtain extreme positive pitching acceleration
(nose up). The dynamic response or, at the option of the applicant, the
transient rigid body response of the airplane, must be taken into account in
determining the tail load. Airplane loads which occur subsequent to the
normal acceleration at the center of gravity exceeding the maximum positive
limit maneuvering load factor, n, need not be considered.
(2) Specified control displacement. A checked maneuver, based on a rational
pitching control motion vs. time profile, must be established in which the
design limit load factor specified in Sec. 25.337 will not be exceeded.
Unless lesser values cannot be exceeded, the airplane response must result in
pitching accelerations not less than the following:
(i) A positive pitching acceleration (nose up) is assumed to be reached
concurrently with the airplane load factor of 1.0 (Points A1 to D1, Sec.
25.333(b)). The positive acceleration must be equal to at least
39n
---- (n-1.5), (Radians/sec./2/ )
v
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 (nose down) is assumed to be reached
concurrently with the positive maneuvering load factor (Points A2 to D2,
Sec. 25.333(b)). This negative pitching acceleration must be equal to at
least
-26n
------ (n-1.5), (Radians/sec./2/ )
v
where--
n is the positive load factor at the speed under consideration; and V is the
airplane equivalent speed in knots.
(d) Gust conditions. The gust conditions B' through J' Sec. 25.333(c), must
be investigated. The following provisions apply:
(1) The air load increment due to a specified gust must be added to the
initial balancing tail load corresponding to steady level flight.
(2) The alleviating effect of wing down-wash and of the airplane's motion
in response to the gust may be included in computing the tail gust load
increment.
(3) Instead of a rational investigation of the airplane response, the gust
alleviation factor Kg may be applied to the specified gust intensity for the
horizontal tail.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5672, Apr. 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]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.333 Flight envelope.
(a) General. The strength requirements must be met at each combination of
airspeed and load factor on and within the boundaries of the representative
maneuvering and gust envelopes (V-n diagrams) of paragraphs (b) and (c) of
this section. These envelopes must also be used in determining the airplane
structural operating limitations as specified in Sec. 25.1501.
(b) Maneuvering envelope.
[ ...Illustration appears here... ]
(c) Gust envelope.
[ ...Illustration appears here... ]
Sec. 25.335 Design airspeeds.
The selected design airspeeds are equivalent airspeeds (EAS). Estimated
values of VS0 and VS1 must be conservative.
(a) Design cruising speed, VC. For VC, the following apply:
(1) The minimum value of VC must be sufficiently greater than VB to provide
for inadvertent speed increases likely to occur as a result of severe
atmospheric turbulence.
(2) In the absence of a rational investigation substantiating the use of
other values, VC may not be less than VB+43 knots. However, it need not
exceed the maximum speed in level flight at maximum continuous power for the
corresponding altitude.
(3) At altitudes where VD is limited by Mach number, VC may be limited to a
selected Mach number.
(b) Design dive speed, VD. VD must be selected so that VC/MC is not greater
than 0.8 VD/MD, or so that the minimum speed margin between VC/MC and VD/MD
is the greater of the following values:
(1) From an initial condition of stabilized flight at VC/MC, the airplane
is upset, flown for 20 seconds along a flight path 7.5 deg. below the initial
path, and then pulled up at a load factor of 1.5 g (0.5 g acceleration
increment). The speed increase occurring in this maneuver may be calculated
if reliable or conservative aerodynamic data is used. Power as specified in
Sec. 25.175(b)(1)(iv) is assumed until the pullup is initiated, at which time
power reduction and the use of pilot controlled drag devices may be assumed;
(2) The minimum speed margin must be enough to provide for atmospheric
variations (such as horizontal gusts, and penetration of jet streams and cold
fronts) and for instrument errors and airframe production variations. These
factors may be considered on a probability basis. However, the margin at
altitude where MC is limited by compressibility effects may not be less than
0.05 M.
(c) Design maneuvering speed VA. For VA, the following apply:
(1) VA may not be less than VS1 <radical>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 CN max
curve intersects the positive maneuver load factor line, whichever is less.
(d) Design speed for maximum gust intensity, VB. For VB, the following
apply:
(1) VB may not be less than the speed determined by the intersection of the
line representing the maximum position lift CN max and the line representing
the rough air gust velocity on the gust V-n diagram, or (<radical>ng) VS1,
whichever is less, where--
(i) ng is the positive airplane gust load factor due to gust, at speed VC
(in accordance with Sec. 25.341), and at the particular weight under
consideration; and
(ii) VS1 is the stalling speed with the flaps retracted at the particular
weight under consideration.
(2) 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 Sec. 25.697(a)) must be sufficiently greater than the operating speed
recommended for the corresponding stage of flight (including balked landings)
to allow for probable variations in control of airspeed and for transition
from one flap position to another.
(2) If an automatic flap positioning or load limiting device is used, the
speeds and corresponding flap positions programmed or allowed by the device
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 selected design speed for each drag
device must be sufficiently greater than the speed recommended for the
operation of the device to allow for probable variations in speed control.
For drag devices intended for use in high speed descents, VDD may not be less
than VD. When an automatic drag device positioning or load limiting means is
used, the speeds and corresponding drag device positions programmed or
allowed by the automatic means must be used for design.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5672, Apr. 8, 1970]
Sec. 25.337 Limit maneuvering load factors.
(a) Except where limited by maximum (static) lift coefficients, the
airplane is assumed to be subjected to symmetrical maneuvers resulting in the
limit maneuvering load factors prescribed in this section. Pitching
velocities appropriate to the corresponding pull-up and steady turn maneuvers
must be taken into account.
(b) The positive limit maneuvering load factor "n" for any speed up to Vn
may not be less than 2.1+24,000/ (W +10,000) except that "n" may not be less
than 2.5 and need not be greater than 3.8--where "W" is the design maximum
takeoff weight.
(c) The negative limit maneuvering load factor--
(1) May not be less than -1.0 at speeds up to VC; and
(2) Must vary linearly with speed from the value at VC to zero at VD.
(d) Maneuvering load factors lower than those specified in this section may
be used if the airplane has design features that make it impossible to exceed
these values in flight.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5672, Apr. 8, 1970]
Sec. 25.341 Gust loads.
(a) The airplane is assumed to be subjected to symmetrical vertical gusts
in level flight. The resulting limit load factors must correspond to the
conditions determined as follows:
(1) Positive (up) and negative (down) rough air gusts of 66 fps at VB must
be considered at altitudes between sea level and 20,000 feet. The gust
velocity may be reduced linearly from 66 fps at 20,000 feet to 38 fps at
50,000 feet.
(2) Positive and negative gusts of 50 fps at VC must be considered at
altitudes between sea level and 20,000 feet. The gust velocity may be reduced
linearly from 50 fps at 20,000 feet to 25 fps at 50,000 feet.
(3) Positive and negative gusts of 25 fps at VD must be considered at
altitudes between sea level and 20,000 feet. The gust velocity may be reduced
linearly from 25 fps at 20,000 feet to 12.5 fps at 50,000 feet.
(b) The following assumptions must be made:
(1) The shape of the gust is
Ude 2<pi>s
U = --- (1-cos ------ )
2 25C
where--
s=distance penetrated into gust (ft);
C=mean geometric chord of wing (ft); and
Ude=derived gust velocity referred to in paragraph (a) (fps).
(2) Gust load factors vary linearly between the specified conditions B'
through G', as shown on the gust envelope in Sec. 25.333(c).
(c) In the absence of a more rational analysis, the gust load factors must
be computed as follows:
KgUdeVa
n=1 + ---------
498 (W/S)
where--
0.88<mu>g
Kg = ----------- = gust alleviation factor;
5.3+<mu>g
2(W/S)
<mu>g = ------ = airplane mass ratio:
rCag
Ude=derived gust velocities referred to in paragraph (a) (fps);
r=density of air (slugs cu. ft.);
W/S=wing loading (psf);
C=mean geometric chord (ft);
g=acceleration due to gravity (ft/sec**2);
V=airplane equivalent speed (knots); and
a=slope of the airplane normal force coefficient curve CNA per radian if the
gust loads are applied to the wings and horizontal method. The wing lift
curve slope CAL per radian may be used when the gust load is applied to
the wings only and the horizontal tail gust loads are treated as a
separate condition.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-72, 55
FR 29775, July 20, 1990; 55 FR 37607, Sept. 12, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.343 Design fuel and oil loads.
(a) The disposable load combinations must include each fuel and oil load in
the range from zero fuel and oil to the selected maximum fuel and oil load. A
structural reserve fuel condition, not exceeding 45 minutes of fuel under the
operating conditions in Sec. 25.1001(e) and (f), as applicable, may be
selected.
(b) If a structural reserve fuel condition is selected, it must be used as
the minimum fuel weight condition for showing compliance with the flight load
requirements as prescribed in this subpart. In addition--
(1) The structure must be designed for a condition of zero fuel and oil in
the wing at limit loads corresponding to--
(i) A maneuvering load factor of +2.25; and
(ii) Gust intensities equal to 85 percent of the values prescribed in Sec.
25.341; and
(2) Fatigue evaluation of the structure must account for any increase in
operating stresses resulting from the design condition of paragraph (b)(1) of
this section; and
(3) The flutter, deformation, and vibration requirements must also be met
with zero fuel.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-18, 33 FR
12226, Aug. 30, 1968; Amdt. 25-72, 55 FR 29775, July 20, 1990; 55 FR 37607,
Sept. 12, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.345 High lift devices.
(a) If flaps are to be used during takeoff, approach, or landing, at the
design flap speeds established for these stages of flight under Sec.
25.335(e) and with the flaps in the corresponding positions, the airplane is
assumed to be subjected to symmetrical maneuvers and gusts within the range
determined by--
(1) Maneuvering to a positive limit load factor of 2.0; and
(2) Positive and negative 25 fps derived gusts acting normal to the flight
path in level flight.
(b) The airplane must be designed for the conditions prescribed in
paragraph (a) of this section, except that the airplane load factor need not
exceed 1.0, taking into account, as separate conditions, the effects of--
(1) Propeller slipstream corresponding to maximum continuous power at the
design flap speeds VF, and with takeoff power at not less than 1.4 times the
stalling speed for the particular flap position and associated maximum
weight; and
(2) A head-on gust of 25 feet per second velocity (EAS).
(c) If flaps or similar high lift devices are to be used in en route
conditions, and with flaps in the appropriate position at speeds up to the
flap design speed chosen for these conditions, the airplane is assumed to be
subjected to symmetrical maneuvers and gusts within the range determined by--
(1) Maneuvering to a positive limit load factor as prescribed in Sec.
25.337(b); and
(2) Positive and negative derived gusts as prescribed in Sec. 25.341 acting
normal to the flight path in level flight.
(d) The airplane must be designed for landing at the maximum takeoff weight
with a maneuvering load factor of 1.5g and the flaps and similar high lift
devices in the landing configuration.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by 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 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.349 Rolling conditions.
The airplane must be designed for rolling loads resulting from the
conditions specified in paragraphs (a) and (b) of this section. Unbalanced
aerodynamic moments about the center of gravity must be reacted in a rational
or conservative manner, considering the principal masses furnishing the
reacting inertia forces.
(a) Maneuvering. The following conditions, speeds, and aileron deflections
(except as the deflections may be limited by pilot effort) must be considered
in combination with an airplane load factor of zero and of two-thirds of the
positive maneuvering factor used in design. In determining the required
aileron deflections, the torsional flexibility of the wing must be considered
in accordance with Sec. 25.301(b):
(1) Conditions corresponding to steady rolling velocities must be
investigated. In addition, conditions corresponding to maximum angular
acceleration must be investigated for airplanes with engines or other weight
concentrations outboard of the fuselage. For the angular acceleration
conditions, zero rolling velocity may be assumed in the absence of a rational
time history investigation of the maneuver.
(2) At VA, a sudden deflection of the aileron to the stop is assumed.
(3) At VC, the aileron deflection must be that required to produce a rate
of roll not less than that obtained in paragraph (a)(2) of this section.
(4) At VD, the aileron deflection must be that required to produce a rate
of roll not less than one-third of that in paragraph (a)(2) of this section.
(b) Unsymmetrical gusts. The condition of unsymmetrical gusts must be
considered by modifying the symmetrical flight conditions B' or C' (in Sec.
25.333(c)) whichever produces the critical load. It is assumed that 100
percent of the wing air load acts on one side of the airplane and 80 percent
acts on the other side.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5672, Apr. 8, 1970]
Sec. 25.351 Yawing conditions.
The airplane must be designed for loads resulting from the conditions
specified in paragraphs (a) and (b) of this section. Unbalanced aerodynamic
moments about the center of gravity must be reacted in a rational or
conservative manner considering the principal masses furnishing the reacting
inertia forces:
(a) Maneuvering. At speeds from VMC to VD, the following maneuvers must be
considered. In computing the tail loads, the yawing velocity may be assumed
to be zero:
(1) With the airplane in unaccelerated flight at zero yaw, it is assumed
that the rudder control is suddenly displaced to the maximum deflection, as
limited by the control surface stops, or by a 300-pound rudder pedal force,
whichever is less.
(2) With the rudder deflected as specified in paragraph (a)(1) of this
section, it is assumed that the airplane yaws to the resulting sideslip
angle.
(3) With the airplane yawed to the static sideslip angle corresponding to
the rudder deflection specified in paragraph (a)(1) of this section, it is
assumed that the rudder is returned to neutral.
(b) Lateral gusts. The airplane is assumed to encounter derived gusts
normal to the plane of symmetry while in unaccelerated flight. The derived
gusts and airplane speeds corresponding to conditions B' through J' (in Sec.
25.333(c)) (as determined by Secs. 25.341 and 25.345(a)(2) or Sec.
25.345(c)(2)) must be investigated. The shape of the gust must be as
specified in Sec. 25.341. In the absence of a rational investigation of the
airplane's response to a gust, the gust loading on the vertical tail surfaces
must be computed as follows:
KgtUdeVatSt
Lt = -----------
498
where--
Lt=vertical tail load (lbs.);
0.88<mu>gt
Kgt = ------------ = gust alleviation factor;
5.3+<mu>gt
2W K
<mu>gt = -------- ( ----- )**2 =lateral mass ratio;
pCtgatSt lt
Ude=derived gust velocity (fps);
p=air density (slugs/cu. ft.);
W=airplane weight (lbs.);
St=area of vertical tail (ft.**2);
Ct=mean geometric chord of vertical surface (ft.);
at=lift curve slope of vertical tail (per radian);
K=radius of gyration in yaw (ft).;
lt=distance from airplane c.g., to lift center of vertical surface (ft.);
g=acceleration due to gravity (ft./sec.**2); and
V=airplane equivalent speed (knots).
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5672, Apr. 8, 1970; Amdt. 25-46, 43 FR 50595, Oct. 30, 1978; Amdt. 25-72, 55
FR 29775, July 20, 1990; 55 FR 37608, Sept. 12, 1990; 55 FR 41415, Oct. 11,
1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Supplementary Conditions
Sec. 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 takeoff power and propeller
speed acting simultaneously with 75 percent of the limit loads from flight
condition A of Sec. 25.333(b);
(2) A limit torque corresponding to the maximum continuous power and
propeller speed, acting simultaneously with the limit loads from flight
condition A of Sec. 25.333(b); and
(3) For turbopropeller installations, in addition to the conditions
specified in paragraphs (a)(1) and (2) of this section, a limit engine torque
corresponding to takeoff power and propeller speed, multiplied by a factor
accounting for propeller control system malfunction, including quick
feathering, acting simultaneously with 1g level flight loads. In the absence
of a rational analysis, a factor of 1.6 must be used.
(b) For turbine engine installations, the engine mounts and supporting
structure must be designed to withstand each of the following:
(1) A limit engine torque load imposed by sudden engine stoppage due to
malfunction or structural failure (such as compressor jamming).
(2) A limit engine torque load imposed by the maximum acceleration of the
engine.
(c) The limit engine torque to be considered under paragraph (a) of this
section must be obtained by multiplying mean torque for the specified power
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.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5672, Apr. 8, 1970; Amdt. 25-46, 43 FR 50595, Oct. 30, 1978; Amdt. 25-72, 55
FR 29776, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.363 Side load on engine mount.
(a) Each engine mount and its supporting structure must be designed for a
limit load factor in a lateral direction, for the side load on the engine
mount, at least equal to the maximum load factor obtained in the yawing
conditions but not less than--
(1) 1.33; or
(2) One-third of the limit load factor for flight condition A as prescribed
in Sec. 25.333(b).
(b) The side load prescribed in paragraph (a) of this section may be
assumed to be independent of other flight conditions.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5672, Apr. 8, 1970]
Sec. 25.365 Pressurized cabin 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 with pressure differential loads from zero up to the maximum
relief valve setting.
(b) The external pressure distribution in flight, and stress concentrations
and fatigue effects must be accounted for.
(c) If landings may be made with the compartment pressurized, landing loads
must be combined with pressure differential loads from zero up to the maximum
allowed during landing.
(d) The airplane structure must be strong enough to withstand the pressure
differential loads corresponding to the maximum relief valve setting
multiplied by a factor of 1.33, omitting other loads.
(e) Any structure, component or part, inside or outside a pressurized
compartment, the failure of which could interfere with continued safe flight
and landing, must be designed to withstand the effects of a sudden release of
pressure through an opening in any compartment at any operating altitude
resulting from each of the following conditions:
(1) The penetration of the compartment by a portion of an engine following
an engine disintegration;
(2) Any opening in any pressurized compartment up to the size Ho in square
feet; however, small compartments may be combined with an adjacent
pressurized compartment and both considered as a single compartment for
openings that cannot reasonably be expected to be confined to the small
compartment. The size Ho must be computed by the following formula:
Ho=PAs
where,
Ho=Maximum opening in square feet, need not exceed 20 square feet.
As
P = ---- +.024
6240
As=Maximum cross-sectional area of the pressurized shell normal to the
longitudinal axis, in square feet; and
(3) The maximum opening caused by airplane or equipment failures not shown
to be extremely improbable.
(f) In complying with paragraph (e) of this section, the fail-safe features
of the design may be considered in determining the probability of failure or
penetration and probable size of openings, provided that possible improper
operation of closure devices and inadvertent door openings are also
considered. Furthermore, the resulting differential pressure loads must be
combined in a rational and conservative manner with 1-g level flight loads
and any loads arising from emergency depressurization conditions. These loads
may be considered as ultimate conditions; however, any deformations
associated with these conditions must not interfere with continued safe
flight and landing. The pressure relief provided by intercompartment venting
may also be considered.
(g) Bulkheads, floors, and partitions in pressurized compartments for
occupants must be designed to withstand the conditions specified in paragraph
(e) of this section. In addition, reasonable design precautions must be taken
to minimize the probability of parts becoming detached and injuring occupants
while in their seats.
EDITORIAL NOTE: At 55 FR 29776, July 20, 1990, Sec. 25.365 was amended by
removing the words "for occupants" from the introductory sentence, but could
not be executed because of an intervening amendment.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-54, 45 FR
60172, Sept. 11, 1980; Amdt. 25-71, 55 FR 13477, Apr. 10, 1990; Amdt. 25-72,
55 FR 29776, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.367 Unsymmetrical loads due to engine failure.
(a) The airplane must be designed for the unsymmetrical loads resulting
from the failure of the critical engine. Turbopropeller airplanes must be
designed for the following conditions in combination with a single
malfunction of the propeller drag limiting system, considering the probable
pilot corrective action on the flight controls:
(1) At speeds between VMC and VD, the loads resulting from power failure
because of fuel flow interruption are considered to be limit loads.
(2) At speeds between VMC and VC, the loads resulting from the
disconnection of the engine compressor from the turbine or from loss of the
turbine blades are considered to be ultimate loads.
(3) The time history of the thrust decay and drag build-up occurring as a
result of the prescribed engine failures must be substantiated by test or
other data applicable to the particular engine-propeller combination.
(4) The timing and magnitude of the probable pilot corrective action must
be conservatively estimated, considering the characteristics of the
particular engine-propeller-airplane combination.
(b) Pilot corrective action may be assumed to be initiated at the time
maximum yawing velocity is reached, but not earlier than two seconds after
the engine failure. The magnitude of the corrective action may be based on
the control forces specified in Sec. 25.397(b) except that lower forces may
be assumed where it is shown by anaylsis or test that these forces can
control the yaw and roll resulting from the prescribed engine failure
conditions.
Sec. 25.371 Gyroscopic loads.
The structure supporting the engines must be designed for gyroscopic loads
associated with the conditions specified in Secs. 25.331, 25.349, and 25.351,
with the engines at maximum continuous r.p.m.
Sec. 25.373 Speed control devices.
If speed control devices (such as spoilers and drag flaps) are installed
for use in en route conditions--
(a) The airplane must be designed for the symmetrical maneuvers and gusts
prescribed in Secs. 25.333, 25.337, and 25.341, and the yawing maneuvers and
lateral gusts in Sec. 25.351, at each setting and the maximum speed
associated with that setting; and
(b) If the device has automatic operating or load limiting features, the
airplane must be designed for the maneuver and gust conditions prescribed in
paragraph (a) of this section, at the speeds and corresponding device
positions that the mechanism allows.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-72, 55 FR
29776, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Control Surface and System Loads
Sec. 25.391 Control surface loads: general.
The control surfaces must be designed for the limit loads resulting from
the flight conditions in Secs. 25.331, 25.349, and 25.351 and the ground gust
conditions in Sec. 25.415, considering the requirements for--
(a) Loads parallel to hinge line, in Sec. 25.393;
(b) Pilot effort effects, in Sec. 25.397;
(c) Trim tab effects, in Sec. 25.407;
(d) Unsymmetrical loads, in Sec. 25.427; and
(e) Outboard fins, in Sec. 25.445.
Sec. 25.393 Loads parallel to hinge line.
(a) Control surfaces and supporting hinge brackets must be designed for
inertia loads acting parallel to the hinge line.
(b) In the absence of more rational data, the inertia 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.
Sec. 25.395 Control system.
(a) Longitudinal, lateral, directional, and drag control system and their
supporting structures must be designed for loads corresponding to 125 percent
of the computed hinge moments of the movable control surface in the
conditions prescribed in Sec. 25.391.
(b) The system limit loads, except the loads resulting from ground gusts,
need not exceed the loads that can be produced by the pilot (or pilots) and
by automatic or power devices operating the controls.
(c) The loads must not be less than those resulting from application of the
minimum forces prescribed in Sec. 25.397(c).
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5672, Apr. 8, 1970; Amdt. 25-72, 55 FR 29776, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 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 control grips or
pads (in a manner simulating flight conditions) and to be reacted at the
attachment of the control system to the control surface horn.
(b) Pilot effort effects. In the control surface flight loading condition,
the air loads on movable surfaces and the corresponding deflections need not
exceed those that would result in flight from the application of any pilot
force within the ranges specified in paragraph (c) of this section. Two-
thirds of the maximum values specified for the aileron and elevator may be
used if control surface hinge moments are based on reliable data. In applying
this criterion, the effects of servo mechanisms, tabs, and automatic pilot
systems, must be considered.
(c) Limit pilot forces and torques. The limit pilot forces and torques are
as follows:
Minimum
Maximum forces forces or
Control or torques torques
Aileron:
Stick 100 lbs 40 lbs.
Wheel /1/ 80 D in.-lbs /2/ 40 D in.-lbs.
Elevator:
Stick 250 lbs 100 lbs.
Wheel (symmetrical) 300 lbs 100 lbs.
Wheel (unsymmetrical) /3/ 100 lbs.
Rudder 300 lbs 130 lbs.
/1/ The critical parts of the aileron control system must
be designed for a single tangential force with a limit
value equal to 1.25 times the couple force determined from
these criteria.
/2/ D= wheel diameter (inches).
/3/ The unsymmetrical forces must be applied at one of the
normal handgrip points on the periphery of the control
wheel.
[Doc. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 41 FR
55466, Dec. 20, 1976; Amdt. 25-72, 55 FR 29776, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.399 Dual control system.
(a) Each dual control system must be designed for the pilots operating in
opposition, using individual pilot forces not less than--
(1) 0.75 times those obtained under Sec. 25.395; or
(2) The minimum forces specified in Sec. 25.397(c).
(b) The control system must be designed for pilot forces applied in the
same direction, using individual pilot forces not less than 0.75 times those
obtained under Sec. 25.395.
Sec. 25.405 Secondary control system.
Secondary controls, such as wheel brake, spoiler, and tab controls, must be
designed for the maximum forces that a pilot is likely to apply to those
controls. The following values may be used:
Pilot Control Force Limits (Secondary Controls)
Control Limit pilot forces
Miscellaneous: 1+R
*Crank, wheel, or (------) x 50 lbs., but
lever. 3
not less than 50 lbs. nor more than 150 lbs. (R=radius).
(Applicable to any angle within 20 deg. of plane of
control).
Twist 133 in.-lbs.
Push-pull To be chosen by applicant.
*Limited to flap, tab, stabilizer, spoiler, and landing gear operation
controls.
Sec. 25.407 Trim tab effects.
The effects of trim tabs on the control surface design conditions must be
accounted for only where the surface loads are limited by maximum pilot
effort. In these cases, the tabs are considered to be deflected in the
direction that would assist the pilot, and the deflections are--
(a) For elevator trim tabs, those required to trim the airplane at any
point within the positive portion of the pertinent flight envelope in Sec.
25.333(b), except as limited by the stops; and
(b) For aileron and rudder trim tabs, those required to trim the airplane
in the critical unsymmetrical power and loading conditions, with appropriate
allowance for rigging tolerances.
Sec. 25.409 Tabs.
(a) Trim tabs. Trim tabs must be designed to withstand loads arising from
all likely combinations of tab setting, primary control position, and
airplane speed (obtainable without exceeding the flight load conditions
prescribed for the airplane as a whole), when the effect of the tab is
opposed by pilot effort forces up to those specified in Sec. 25.397(b).
(b) Balancing tabs. Balancing tabs must be designed for deflections
consistent with the primary control surface loading conditions.
(c) Servo tabs. Servo tabs must be designed for deflections consistent with
the primary control surface loading conditions obtainable within the pilot
maneuvering effort, considering possible opposition from the trim tabs.
Sec. 25.415 Ground gust conditions.
(a) The control system must be designed as follows for control surface
loads due to ground gusts and taxiing downwind:
(1) The control system between the stops nearest the surfaces and the
cockpit controls must be designed for loads corresponding to the limit hinge
moments H of paragraph (a)(2) of this section. These loads need not exceed--
(i) The loads corresponding to the maximum pilot loads in Sec. 25.397(c)
for each pilot alone; or
(ii) 0.75 times these maximum loads for each pilot when the pilot forces
are applied in the same direction.
(2) The control system stops nearest the surfaces, the control system
locks, and the parts of the systems (if any) between these stops and locks
and the control surface horns, must be designed for limit hinge moments H
obtained from the formula, H=KcSsq, where--
H=limit hinge moment (ft. lbs.);
c=mean chord of the control surface aft of the hinge line (ft.);
Ss=area of the control surface aft of the hinge line (sq. ft.);
q=dynamic pressure (p.s.f.) based on a design speed not less than 14.6(W/S)
1/2 +14.6 (f.p.s.), except that the design speed need not exceed 88
f.p.s. (W/S is wing loading based on maximum airplane weight and wing
area); and
K=limit hinge moment factor for ground gusts derived in paragraph (b) of this
section.
(b) The limit hinge moment factor K for ground gusts must be derived as
follows:
Surface K Position of controls
(a) Aileron 0.75 Control column locked or lashed in mid-position.
(b) ......do /1/ 1 +/-0.50 Ailerons at full throw.
(c) Elevator /1/ 1 +/-0.75 (c) Elevator full down.
(d) ......do /1/ 1 +/-0.75 (d) Elevator full up.
(e) Rudder 0.75 (e) Rudder in neutral.
(f) ......do 0.75 (f) Rudder at full throw.
/1/ A positive value of K indicates a moment tending to depress the surface,
while a negative value of K indicates a moment tending to raise the surface.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-72, 55 FR
29776, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.427 Unsymmetrical loads.
(a) Horizontal tail surfaces and their supporting structure must be
designed for unsymmetrical loads arising from yawing and slipstream effects,
in combination with the prescribed flight conditions.
(b) In the absence of more rational data, the following apply:
(1) For airplanes that are conventional in regard to location of
propellers, wings, tail surfaces, and fuselage shape--
(i) 100 percent of the maximum loading from the symmetrical flight
conditions may be assumed to act on the surface on one side of the plane of
symmetry; and
(ii) 80 percent of this loading may be assumed to act on the other side.
(2) For empennage arrangements where the horizontal tail surfaces have
appreciable dihedral or are supported by the vertical tail surfaces, the
surfaces and supporting structure must be designed for the combined vertical
and horizontal surface loads resulting from each prescribed flight load
condition considered separately.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5673, Apr. 8, 1970]
Sec. 25.445 Outboard fins.
(a) If outboard fins are on the horizontal tail surface, the tail surfaces
must be designed for the maximum horizontal surface load in combination with
the corresponding loads induced on the vertical surfaces by endplate effects.
These induced effects need not be combined with other vertical surface loads.
(b) To provide for unsymmetrical loading when outboard fins extend above
and below the horizontal surface, the critical vertical surface loading (load
per unit area) determined under Sec. 25.391 must also be applied as follows:
(1) 100 percent to the area of the vertical surfaces above (or below) the
horizontal surface.
(2) 80 percent to the area below (or above) the horizontal surface.
Sec. 25.457 Wing flaps.
Wing flaps, their operating mechanisms, and their supporting structures
must be designed for critical loads occurring in the conditions prescribed in
Sec. 25.345, accounting for the loads occurring during transition from one
flap position and airspeed to another.
Sec. 25.459 Special devices.
The loading for special devices using aerodynamic surfaces (such as slots,
slats, and spoilers) must be determined from test data.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-72, 55 FR
29776, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Ground Loads
Sec. 25.471 General.
(a) Loads and equilibrium. For limit ground loads--
(1) Limit ground loads obtained under this subpart are considered to be
external forces applied to the airplane structure; and
(2) In each specified ground load condition, the external loads must be
placed in equilibrium with the linear and angular inertia loads in a rational
or conservative manner.
(b) Critical centers of gravity. The critical centers of gravity within the
range for which certification is requested must be selected so that the
maximum design loads are obtained in each landing gear element. Fore and aft,
vertical, and lateral airplane centers of gravity must be considered. Lateral
displacements of the c.g. from the airplane centerline which would result in
main gear loads not greater than 103 percent of the critical design load for
symmetrical loading conditions may be selected without considering the
effects of these lateral c.g. displacements on the loading of the main gear
elements, or on the airplane structure provided--
(1) The lateral displacement of the c.g. results from random passenger or
cargo disposition within the fuselage or from random unsymmetrical fuel
loading or fuel usage; and
(2) Appropriate loading instructions for random disposable loads are
included under the provisions of Sec. 25.1583(c)(1) to ensure that the
lateral displacement of the center of gravity is maintained within these
limits.
(c) Landing gear dimension data. Figure 1 of Appendix A contains the basic
landing gear dimension data.
[Amdt. 25-23, 35 FR 5673, Apr. 8, 1970]
Sec. 25.473 Ground load conditions and assumptions.
(a) For the landing conditions specified in Secs. 25.479 through 25.485,
the following apply:
(1) The selected limit vertical inertia load factors at the center of
gravity of the airplane may not be less than the values that would be
obtained--
(i) In the attitude and subject to the drag loads associated with the
particular landing condition;
(ii) With a limit descent velocity of 10 f.p.s. at the design landing
weight (the maximum weight for landing conditions at the maximum descent
velocity); and
(iii) With a limit descent velocity of 6 f.p.s. at the design takeoff
weight (the maximum weight for landing conditions at a reduced descent
velocity).
(2) Airplane lift, not exceeding the airplane weight, may be assumed to
exist throughout the landing impact and to act through the center of gravity
of the airplane.
(b) The prescribed descent velocities may be modified if it is shown that
the airplane has design features that make it impossible to develop these
velocities.
(c) The minimum limit inertia load factors corresponding to the required
limit descent velocities must be determined in accordance with Sec.
25.723(a).
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5673, Apr. 8, 1970]
Sec. 25.477 Landing gear arrangement.
Sections 25.479 through 25.485 apply to airplanes with conventional
arrangements of main and nose gears, or main and tail gears, when normal
operating techniques are used.
Sec. 25.479 Level landing conditions.
(a) In the level attitude, the airplane is assumed to contact the ground at
forward velocity components, ranging from VL1 to 1.25 VL2 parallel to the
ground, and to be subjected to the load factors prescribed in Sec.
25.473(a)(1) with--
(1) VL1 equal to VS0 (TAS) at the appropriate landing weight and in
standard sea level conditions; and
(2) VL2 equal to VS0 (TAS) at the appropriate landing weight and altitudes
in a hot day temperature of 41 degrees F. above standard.
(b) The effects of increased contact speeds must be investigated if
approval of downwind landings exceeding 10 knots is desired.
(c) Assuming that the following combinations of vertical and drag
components act at the axle centerline, the following apply:
(1) For the condition of maximum wheel spin-up load, drag components
simulating the forces required to accelerate the wheel rolling assembly up to
the specified ground speed must be combined with the vertical ground
reactions existing at the instant of peak drag loads. The coefficient of
friction between the tires and the ground may be established by considering
the effects of skidding velocity and tire pressure. However, this coefficient
of friction need not be more than 0.8. This condition must be applied to the
landing gear, directly affected attaching structure, and large mass items
such as external fuel tanks and nacelles.
(2) For the condition of maximum wheel vertical load, an aft acting drag
component of not less than 25 percent of the maximum vertical ground reaction
must be combined with the maximum ground reaction of Sec. 25.473.
(3) For the condition of maximum springback load, forward-acting horizontal
loads resulting from a rapid reduction of the spin-up drag loads must be
combined with the vertical ground reactions at the instant of the peak
forward load. This condition must be applied to the landing gear, directly
affected attaching structure, and large mass items such as external fuel
tanks and nacelles.
(d) For the level landing attitude for airplanes with tail wheels, the
conditions specified in paragraphs (a) through (c) of this section must be
investigated with the airplane horizontal reference line horizontal in
accordance with figure 2 of Appendix A.
(e) For the level landing attitude for airplanes with nose wheels, shown in
figure 2 of Appendix A, the conditions specified in paragraphs (a) through
(c) of this section must be investigated, assuming the following attitudes:
(1) An attitude in which the main wheels are assumed to contact the ground
with the nose wheel just clear of the ground.
(2) If reasonably attainable at the specified descent and forward
velocities, an attitude in which the nose and main wheels are assumed to
contact the ground simultaneously. For this attitude--
(i) The nose and main gear may be separately investigated under the
conditions in paragraph (c) (1) and (3) of this section; and
(ii) The pitching moment is assumed, under the condition in paragraph
(c)(2) of this section, to be resisted by the nose gear.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5673, Apr. 8, 1970]
Sec. 25.481 Tail-down landing conditions.
(a) In the tail-down attitude, the airplane is assumed to contact the
ground at forward velocity components, ranging from VL1 to VL2, parallel to
the ground, and is subjected to the load factors prescribed in Sec.
25.473(a)(1) with--
(1) VL1 equal to VS0 (TAS) at the appropriate landing weight and in
standard sea level conditions; and
(2) VL2 equal to VS0 (TAS) at the appropriate landing weight and altitudes
in a hot day temperature of 41 degrees F. above standard.
The combination of vertical and drag components specified in Sec. 25.479(c)
(1) and (3) is considered to be acting at the main wheel axle centerline.
(b) For the tail-down landing condition for airplanes with tail wheels, the
main and tail wheels are assumed to contact the ground simultaneously, in
accordance with figure 3 of Appendix A. Ground reaction conditions on the
tail wheel are 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 airplanes with nose wheels, the
airplane is assumed to be at an attitude corresponding to either the stalling
angle or the maximum angle allowing clearance with the ground by each part of
the airplane other than the main wheels, in accordance with figure 3 of
Appendix A, whichever is less.
Sec. 25.483 One-wheel landing conditions.
For the one-wheel landing condition, the airplane is assumed to be in the
level attitude and to contact the ground on one side of the main landing
gear, in accordance with Figure 4 of Appendix A. In this attitude--
(a) The ground reactions must be the same as those obtained on that side
under Sec. 25.479(c)(2); and
(b) Each unbalanced external load must be reacted by airplane inertia in a
rational or conservative manner.
Sec. 25.485 Side load conditions.
(a) For the side load condition, the airplane is assumed to be in the level
attitude with only the main wheels contacting the ground, in accordance with
figure 5 of Appendix A.
(b) Side loads of 0.8 of the vertical reaction (on one side) acting inward
and 0.6 of the vertical reaction (on the other side) acting outward must be
combined with one-half of the maximum vertical ground reactions obtained in
the level landing conditions. These loads are assumed to be applied at the
ground contact point and to be resisted by the inertia of the airplane. The
drag loads may be assumed to be zero.
Sec. 25.487 Rebound landing condition.
(a) The landing gear and its supporting structure must be investigated for
the loads occurring during rebound of the airplane from the landing surface.
(b) With the landing gear fully extended and not in contact with the
ground, a load factor of 20.0 must act on the unsprung weights of the landing
gear. This load factor must act in the direction of motion of the unsprung
weights as they reach their limiting positions in extending with relation to
the sprung parts of the landing gear.
Sec. 25.489 Ground handling conditions.
Unless otherwise prescribed, the landing gear and airplane structure must
be investigated for the conditions in Secs. 25.491 through 25.509 with the
airplane at the design ramp weight (the maximum weight for ground handling
conditions). No wing lift may be considered. The shock absorbers and tires
may be assumed to be in their static position.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5673, Apr. 8, 1970]
Sec. 25.491 Takeoff run.
The landing gear and the airplane structure are assumed to be subjected to
loads not less than those obtained under conditions described in Sec. 25.235.
Sec. 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 main wheels, in accordance with figure 6 of Appendix A.
The limit vertical load factor is 1.2 at the design landing weight and 1.0 at
the design ramp weight. A drag reaction equal to the vertical reaction
multiplied by a coefficient of friction of 0.8, must be combined with the
vertical ground reaction and applied at the ground contact point.
(b) For an airplane with a nose wheel the limit vertical load factor is 1.2
at the design landing weight, and 1.0 at the design ramp weight. A drag
reaction equal to the vertical reaction, multiplied by a coefficient of
friction of 0.8, must be combined with the vertical reaction and applied at
the ground contact point of each wheel with brakes. The following two
attitudes, in accordance with figure 6 of Appendix A, must be considered:
(1) The level attitude with the wheels contacting the ground and the loads
distributed between the main and nose gear. Zero pitching acceleration is
assumed.
(2) The level attitude with only the main gear contacting the ground and
with the pitching moment resisted by angular acceleration.
(c) A drag reaction lower than that prescribed in paragraphs (a) and (b) of
this section may be used if it is substantiated that an effective drag force
of 0.8 times the vertical reaction cannot be attained under any likely
loading condition.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5673, Apr. 8, 1970]
Sec. 25.495 Turning.
In the static position, in accordance with figure 7 of Appendix A, the
airplane is assumed to execute a steady turn by nose gear steering, or by
application of sufficient differential power, so that the limit load factors
applied at the center of gravity are 1.0 vertically and 0.5 laterally. The
side ground reaction of each wheel must be 0.5 of the vertical reaction.
Sec. 25.497 Tail-wheel yawing.
(a) A vertical ground reaction equal to the static load on the tail wheel,
in combination with a side component of equal magnitude, is assumed.
(b) If there is a swivel, the tail wheel is assumed to be swiveled 90 deg.
to the airplane longitudinal axis with the resultant load passing through the
axle.
(c) If there is a lock, steering device, or shimmy damper the tail wheel is
also assumed to be in the trailing position with the side load acting at the
ground contact point.
Sec. 25.499 Nose-wheel yaw.
(a) A vertical load factor of 1.0 at the airplane center of gravity, and a
side component at the nose wheel ground contact equal to 0.8 of the vertical
ground reaction at that point are assumed.
(b) With the airplane assumed to be in static equilibrium with the loads
resulting from the use of brakes on one side of the main landing gear, the
nose gear, its attaching structure, and the fuselage structure forward of the
center of gravity must be designed for the following loads:
(1) A vertical load factor at the center of gravity of 1.0.
(2) A forward acting load at the airplane center of gravity of 0.8 times
the vertical load on one main gear.
(3) Side and vertical loads at the ground contact point on the nose gear
that are required for static equilibrium.
(4) A side load factor at the airplane center of gravity of zero.
(c) If the loads prescribed in paragraph (b) of this section result in a
nose gear side load higher than 0.8 times the vertical nose gear load, the
design nose gear side load @y be limited to 0.8 times the vertical load,
with unbalanced yawing moments assumed to be resisted by airplane inertia
forces.
(d) For other than the nose gear, its attaching structure, and the forward
fuselage structure, the loading conditions are those prescribed in paragraph
(b) of this s@tion, except that--
(1) A lower drag reaction may be used if an effective drag force of 0.8
times the vertical reaction cannot be reached under any likely loading
condition; and
(2) The forward acting load at the center of gravity need not exceed the
maximum drag reaction on one main gear, determined in accordance with Sec.
25.493(b).
(e) With the airplane at design ramp weight, and the nose gear in any
steerable position, the combined application of full normal steering torque
and a vertical force equal to the maximum static reaction on the nose gear
must be considered in designing the nose gear, its attaching structure, and
the forward fuselage structure.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5673, Apr. 8, 1970; Amdt. 25-46, 43 FR 50595, Oct. 30, 1978]
Sec. 25.503 Pivoting.
(a) The airplane is assumed to pivot about one side of the main gear with
the brakes on that side locked. The limit vertical load factor must be 1.0
and the coefficient of friction 0.8.
(b) The airplane is assumed to be in static equilibrium, with the loads
being applied at the ground contact points, in accordance with figure 8 of
Appendix A.
Sec. 25.507 Reversed braking.
(a) The airplane must be in a three point static ground attitude.
Horizontal reactions parallel to the ground and directed forward must be
applied at the ground contact point of each wheel with brakes. The limit
loads must be equal to 0.55 times the vertical load at each wheel or to the
load developed by 1.2 times the nominal maximum static brake torque,
whichever is less.
(b) For airplanes with nose wheels, the pitching moment must be balanced by
rotational inertia.
(c) For airplanes with tail wheels, the resultant of the ground reactions
must pass through the center of gravity of the airplane.
Sec. 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 fittings and
must act parallel to the ground. In addition--
(1) A vertical load factor equal to 1.0 must be considered acting at the
center of gravity;
(2) The shock struts and tires must be in their static positions; and
(3) With WT as the design ramp weight, the towing load, FTOW, is--
(i) 0.3 WT for WT less than 30,000 pounds;
(ii) (6WT+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 specified for
the auxiliary gear apply. For towing points located outboard of the main
gear, the drag and side tow load components specified for the main gear
apply. Where the specified angle of swivel cannot be reached, the maximum
obtainable angle 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 at the main gear must be reacted
by a side force at the static ground line of the wheel to which the load is
applied.
(2) The towing loads at the auxiliary gear and the drag components of the
towing loads at the main gear must be reacted as follows:
(i) A reaction with a maximum value equal to the vertical reaction must be
applied at the axle of the wheel to which the load is applied. Enough
airplane inertia to achieve equilibrium must be applied.
(ii) The loads must be reacted by airplane inertia.
(d) The prescribed towing loads are as follows:
Load
Tow point Position Magnitude No. Direction
Main gear 0.75 FTOW per 1 Forward,
main gear 2 parallel to
unit 3 drag axis.
4 Forward, at
30 deg. to
drag axis.
Aft,
parallel to
drag axis.
Aft, at 30
deg. to drag
axis.
Auxiliary Swiveled 1.0 FTOW 5 Forward.
gear forward 6 Aft.
Swiveled aft ......do 7 Forward.
8 Aft.
Swiveled 45 0.5 FTOW 9 Forward, in
deg. from 10 plane of
forward wheel.
Aft, in
plane of
wheel.
Swiveled 45 ......do 11 Forward, in
deg. from 12 plane of
aft wheel.
Aft, in
plane of
wheel.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5673, Apr. 8, 1970]
Sec. 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 ground loads prescribed in this subpart under paragraphs (b)
through (f) of this section. In addition--
(1) A tandem strut gear arrangement is a multiple-wheel unit; and
(2) In determining the total load on a gear unit with respect to the
provisions of paragraphs (b) through (f) of this section, the transverse
shift in the load centroid, due to unsymmetrical load distribution on the
wheels, may be neglected.
(b) Distribution of limit loads to wheels; tires inflated. The distribution
of the limit loads among the wheels of the landing gear must be established
for each landing, taxiing, and ground handling condition, taking into account
the effects of the following factors:
(1) The number of wheels and their physical arrangements. For truck type
landing gear units, the effects of any seesaw motion of the truck during the
landing impact must be considered in determining the maximum design loads for
the fore and aft wheel pairs.
(2) Any differentials in tire diameters resulting from a combination of
manufacturing tolerances, tire growth, and tire wear. A maximum tire-diameter
differential equal to 2/3 of the most unfavorable combination of diameter
variations that is obtained when taking into account manufacturing
tolerances, tire growth, and tire wear, may be assumed.
(3) Any unequal tire inflation pressure, assuming the maximum variation to
be +/-5 percent of the nominal tire inflation pressure.
(4) A runway crown of zero and a runway crown having a convex upward shape
that may be approximated by a slope of 1 1/2 percent with the horizontal.
Runway crown effects must be considered with the nose gear unit on either
slope of the crown.
(5) The airplane attitude.
(6) Any structural deflections.
(c) Deflated tires. The effect of deflated tires on the structure must be
considered with respect to the loading conditions specified in paragraphs (d)
through (f) of this section, taking into account the physical arrangement of
the gear components. In addition--
(1) The deflation of any one tire for each multiple wheel landing gear
unit, and the deflation of any two critical tires for each landing gear unit
using four or more wheels per unit, must be considered; and
(2) The ground reactions must be applied to the wheels with inflated tires
except that, for multiple-wheel gear units with more than one shock strut, a
rational distribution of the ground reactions between the deflated and
inflated tires, accounting for the differences in shock strut extensions
resulting from a deflated tire, may be used.
(d) Landing conditions. For one and for two deflated tires, the applied
load to each gear unit is assumed to be 60 percent and 50 percent,
respectively, of the limit load applied to each gear for each of the
prescribed landing conditions. However, for the drift landing condition of
Sec. 25.485, 100 percent of the vertical load must be applied.
(e) Taxiing and ground handling conditions. For one and for two deflated
tires--
(1) The applied side or drag load factor, or both factors, at the center of
gravity must be the most critical value up to 50 percent and 40 percent,
respectively, of the limit side or drag load factors, or both factors,
corresponding to the most severe condition resulting from consideration of
the prescribed taxiing and ground handling conditions;
(2) For the braked roll conditions of Sec. 25.493 (a) and (b)(2), the drag
loads on each inflated tire may not be less than those at each tire for the
symmetrical load distribution with no deflated tires;
(3) The vertical load factor at the center of gravity must be 60 percent
and 50 percent, respectively, of the factor with no deflated tires, except
that it may not be less than 1g; and
(4) Pivoting need not be considered.
(f) Towing conditions. For one and for two deflated tires, the towing load,
FTOW, must be 60 percent and 50 percent, respectively, of the load
prescribed.
Water Loads
Sec. 25.521 General.
(a) Seaplanes must be designed for the water loads developed during takeoff
and landing, with the seaplane in any attitude likely to occur in normal
operation, and at the appropriate forward and sinking velocities under the
most severe sea conditions likely to be encountered.
(b) Unless a more rational analysis of the water loads is made, or the
standards in ANC-3 are used, Secs. 25.523 through 25.537 apply.
(c) The requirements of this section and Secs. 25.523 through 25.537 apply
also to amphibians.
Sec. 25.523 Design weights and center of gravity positions.
(a) Design weights. The water load requirements must be met at each
operating weight up to the design landing weight except that, for the takeoff
condition prescribed in Sec. 25.531, the design water takeoff weight (the
maximum weight for water taxi and takeoff run) must be used.
(b) Center of gravity positions. The critical centers of gravity within the
limits for which certification is requested must be considered to reach
maximum design loads for each part of the seaplane structure.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5673, Apr. 8, 1970]
Sec. 25.525 Application of loads.
(a) Unless otherwise prescribed, the seaplane as a whole is assumed to be
subjected to the loads corresponding to the load factors specified in Sec.
25.527.
(b) In applying the loads resulting from the load factors prescribed in
Sec. 25.527, the loads may be distributed over the hull or main float bottom
(in order to avoid excessive local shear loads and bending moments at the
location of water load application) using pressures not less than those
prescribed in Sec. 25.533(b).
(c) For twin float seaplanes, each float must be treated as an equivalent
hull on a fictitious seaplane with a weight equal to one-half the weight of
the twin float seaplane.
(d) Except in the takeoff condition of Sec. 25.531, the aerodynamic lift on
the seaplane during the impact is assumed to be 2/3 of the weight of the
seaplane.
Sec. 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
C1VS02
nw = ----------------
(Tan2/3 b)W1/3
(2) For the bow and stern landing cases
C1VS02 K1
nw = ---------------- x ----------
(Tan2/3 b)W1/3 (1+rx2)2/3
(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 be less than that necessary to obtain the minimum value
of step load factor of 2.33).
(3) VS0=seaplane stalling speed in knots with flaps extended in the
appropriate landing position and with no slipstream effect.
(4) <beta>=angle of dead rise at the longitudinal station at which the load
factor is being determined in accordance with figure 1 of Appendix B.
(5) W=seaplane design landing weight in pounds.
(6) K1=empirical hull station weighing factor, in accordance with figure 2
of Appendix B.
(7) rx=ratio of distance, measured parallel to hull reference axis, from
the center of gravity of the seaplane to the hull longitudinal station at
which the load factor is being computed to the radius of gyration in pitch of
the seaplane, the hull reference axis being a straight line, in the plane of
symmetry, tangential to the keel at the main step.
(c) For a twin float seaplane, because of the effect of flexibility of the
attachment of the floats to the seaplane, the factor K1 may be reduced at the
bow and stern to 0.8 of the value shown in figure 2 of Appendix B. This
reduction applies only to the design of the carrythrough and seaplane
structure.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5673, Apr. 8, 1970]
Sec. 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 those computed
under Sec. 25.527. In addition--
(1) For symmetrical step landings, the resultant water load must be applied
at the keel, through the center of gravity, and must be directed
perpendicularly to the keel line;
(2) For symmetrical bow landings, the resultant water load must be applied
at the keel, one-fifth of the longitudinal distance from the bow to the step,
and must be directed perpendicularly to the keel line; and
(3) For symmetrical stern landings, the resultant water load must be
applied at the keel, at a point 85 percent of the longitudinal distance from
the step to the stern post, and must be directed perpendicularly to the keel
line.
(b) Unsymmetrical landing for hull and single float seaplanes.
Unsymmetrical step, bow, and stern landing conditions must be investigated.
In addition--
(1) The loading for each condition consists of an upward component and a
side component equal, respectively, to 0.75 and 0.25 tan <beta> times the
resultant load in the corresponding symmetrical landing condition; and
(2) The point of application and direction of the upward component of the
load is the same as that in the symmetrical condition, and the point of
application of the side component is at the same longitudinal station as the
upward component but is directed inward perpendicularly to the plane of
symmetry at a point midway between the keel and chine lines.
(c) Unsymmetrical landing; twin float seaplanes. The unsymmetrical loading
consists of an upward load at the step of each float of 0.75 and a side load
of 0.25 tan <beta> at one float times the step landing load reached under
Sec. 25.527. The side load is directed inboard, perpendicularly to the plane
of symmetry midway between the keel and chine lines of the float, at the same
longitudinal station as the upward load.
Sec. 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:
CTOVS12
n = ----------------
(tan2/3 b)W1/3
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;
<beta>=angle of dead rise at the main step (degrees); and
W=design water takeoff weight in pounds.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5673, Apr. 8, 1970]
Sec. 25.533 Hull and main float bottom pressures.
(a) General. The hull and main float structure, including frames and
bulkheads, stringers, and bottom plating, must be designed under this
section.
(b) Local pressures. For the design of the bottom plating and stringers and
their attachments to the supporting structure, the following pressure
distributions must be applied:
(1) For an unflared bottom, the pressure at the chine is 0.75 times the
pressure at the keel, and the pressures between the keel and chine vary
linearly, in accordance with figure 3 of Appendix B. The pressure at the keel
(psi) is computed as follows:
K2VS12
Pk = C2 x ------
tan bk
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 design water takeoff weight with
flaps extended in the appropriate takeoff position; and
<beta>k=angle of dead rise at keel, in accordance with figure 1 of Appendix
B.
(2) For a flared bottom, the pressure at the beginning of the flare is the
same as that for an unflared bottom, and the pressure between the chine and
the beginning of the flare varies linearly, in accordance with figure 3 of
Appendix B. The pressure distribution is the same as that prescribed in
paragraph (b)(1) of this section for an unflared bottom except that the
pressure at the chine is computed as follows:
K2VS12
Pch = C3 x ------
tan b
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
<beta>= angle of dead rise at appropriate station.
The area over which these pressures are applied must simulate pressures
occurring during high localized impacts on the hull or float, but need not
extend over an area that would induce critical stresses in the frames or in
the overall structure.
(c) Distributed pressures. For the design of the frames, keel, and chine
structure, the following pressure distributions apply:
(1) Symmetrical pressures are computed as follows:
K2VS02
P = C4 x ------
tan b
where--
P =pressure (p.s.i.);
C4 =0.078 C1 (with C1 computed under Sec. 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 <beta>=angle of dead rise at
appropriate station.
(2) The unsymmetrical pressure distribution consists of the pressures
prescribed in paragraph (c)(1) of this section on one side of the hull or
main float centerline and one-half of that pressure on the other side of the
hull or main float centerline, in accordance with figure 3 of Appendix B.
These pressures are uniform and must be applied simultaneously over the
entire hull or main float bottom. The loads obtained must be carried into the
sidewall structure of the hull proper, but need not be transmitted in a fore
and aft direction as shear and bending loads.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5673, Apr. 8, 1970]
Sec. 25.535 Auxiliary float loads.
(a) General. Auxiliary floats and their attachments and supporting
structures must be designed for the conditions prescribed in this section. In
the cases specified in paragraphs (b) through (e) of this section, the
prescribed water loads may be distributed over the float bottom to avoid
excessive local loads, using bottom pressures not less than those prescribed
in paragraph (g) of this section.
(b) Step loading. The resultant water load must be applied in the plane of
symmetry of the float at a point three-fourths of the distance from the bow
to 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
times the weight of the displaced water when the float is completely
submerged:
C5 VSO**2 W 2/3
L= -----------------------------------
tan 2/3 <beta>S (1 + ry**2) **2/3
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;
<beta>S=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 must be applied in the plane of
symmetry of the float at a point one-fourth of the distance from the bow to
the step and must be perpendicular to the tangent to the keel line at that
point. The magnitude of the resultant load is that specified in paragraph (b)
of this section.
(d) Unsymmetrical step loading. The resultant water load consists of a
component equal to 0.75 times the load specified in paragraph (a) of this
section and a side component equal to 3.25 tan <beta> times the load
specified in paragraph (b) of this section. The side load must be applied
perpendicularly to the plane of symmetry of the float at a point midway
between the keel and the chine.
(e) Unsymmetrical bow loading. The resultant water load consists of a
component equal to 0.75 times the load specified in paragraph (b) of this
section and a side component equal to 0.25 tan <beta> times the load
specified in paragraph (c) of this section. The side load must be applied
perpendicularly to the plane of symmetry at a point midway between the keel
and the chine.
(f) Immersed float condition. The resultant load must be applied at the
centroid of the cross section of the float at a point one-third of the
distance from the bow to the step. The limit load components are as follows:
vertical = <rho>gVSO
aft = Cx2 <rho>V**2/3 (KVSO)**2.
side = Cy2 <rho>V**2/3 (KVSO)**2.
where--
<rho>=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 bottom pressures must be established
under Sec. 25.533, except that the value of K2 in the formulae may be taken
as 1.0. The angle of dead rise to be used in determining the float bottom
pressures is set forth in paragraph (b) of this section.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5673, Apr. 8, 1970]
Sec. 25.537 Seawing loads.
Seawing design loads must be based on applicable test data.
Emergency Landing Conditions
Sec. 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
to protect each occupant under those conditions.
(b) The structure must be designed to give each occupant every reasonable
chance of escaping serious injury in a minor crash landing when--
(1) Proper use is made of seats, belts, and all other safety design
provisions;
(2) The wheels are retracted (where applicable); and
(3) The occupant experiences the following ultimate inertia forces acting
separately relative to the 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) The supporting structure must be designed to restrain, under all loads
up to those specified in paragraph (b)(3) of this section, each item of mass
that could injure an occupant if it came loose in a minor crash landing.
(d) Seats and items of mass (and their supporting structure) must not
deform under any loads up to those specified in paragraph (b)(3) of this
section in any manner that would impede subsequent rapid evacuation of
occupants.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5673, Apr. 8, 1970; Amdt. 25-64, 53 FR 17646, May 17, 1988]
Sec. 25.562 Emergency landing dynamic conditions.
(a) The seat and restraint system in the airplane must be designed as
prescribed in this section to protect each occupant during an emergency
landing condition when--
(1) Proper use is made of seats, safety belts, and shoulder harnesses
provided for in the design; and
(2) The occupant is exposed to loads resulting from the conditions
prescribed in this section.
(b) Each seat type design approved for crew or passenger occupancy during
takeoff and landing must successfully complete dynamic tests or be
demonstrated by rational analysis based on dynamic tests of a similar type
seat, in accordance with each of the following emergency landing conditions.
The tests must be conducted with an occupant simulated by a 170-pound
anthropomorphic test dummy, as defined by 49 CFR Part 572, Subpart B, or its
equivalent, sitting in the normal upright position.
(1) A change in downward vertical velocity (Dv) of not less than 35 feet
per second, with the airplane's longitudinal axis canted downward 30 degrees
with respect to the horizontal plane and with the wings level. Peak floor
deceleration must occur in not more than 0.08 seconds after impact and must
reach a minimum of 14g.
(2) A change in forward longitudinal velocity (Dv) of not less than 44 feet
per second, with the airplane's longitudinal axis horizontal and yawed 10
degrees either right or left, whichever would cause the greatest likelihood
of the upper torso restraint system (where installed) moving off the
occupant's shoulder, and with the wings level. Peak floor deceleration must
occur in not more than 0.09 seconds after impact and must reach a minimum of
16g. Where floor rails or floor fittings are used to attach the seating
devices to the test fixture, the rails or fittings must be misaligned with
respect to the adjacent set of rails or fittings by at least 10 degrees
vertically (i.e., out of Parallel) with one rolled 10 degrees.
(c) The following performance measures must not be exceeded during the
dynamic tests conducted in accordance with paragraph (b) of this section:
(1) Where upper torso straps are used for crewmembers, tension loads in
individual straps must not exceed 1,750 pounds. If dual straps are used for
restraining the upper torso, the total strap tension loads must not exceed
2,000 pounds.
(2) The maximum compressive load measured between the pelvis and the lumbar
column of the anthropomorphic dummy must not exceed 1,500 pounds.
(3) The upper torso restraint straps (where installed) must remain on the
occupant's shoulder during the impact.
(4) The lap safety belt must remain on the occupant's pelvis during the
impact.
(5) Each occupant must be protected from serious head injury under the
conditions prescribed in paragraph (b) of this section. Where head contact
with seats or other structure can occur, protection must be provided so that
the head impact does not exceed a Head Injury Criterion (HIC) of 1,000 units.
The level of HIC is defined by the equation:
1 t22.5
HCI = { [ (t2-t1) [-------- I a(t)dt ] }
(t2-t1) t1 max
Where:
t1 is the initial integration time,
t2 is the final integration time, and
a(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 contact with seats or other
structure, protection must be provided to prevent axially compressive loads
exceeding 2,250 pounds in each femur.
(7) The seat must remain attached at all points of attachment, although the
structure may have yielded.
(8) Seats must not yield under the tests specified in paragraphs (b)(1) and
(b)(2) of this section to the extent they would impede rapid evacuation of
the airplane occupants.
[Amdt. 25-64, 53 FR 17646, May 17, 1988]
Sec. 25.563 Structural ditching provisions.
Structural strength considerations of ditching provisions must be in
accordance with Sec. 25.801(e).
Fatigue Evaluation
Sec. 25.571 Damage--tolerance and fatigue evaluation of structure.
(a) General. An evaluation of the strength, detail design, and fabrication
must show that catastrophic failure due to fatigue, corrosion, or accidental
damage, will be avoided throughout the operational life of the airplane. This
evaluation must be conducted in accordance with the provisions of paragraphs
(b) and (e) of this section, except as specified in paragraph (c) of this
section, for each part of the structure which could contribute to a
catastrophic failure (such as wing, empennage, control surfaces and their
systems, the fuselage, engine mounting, landing gear, and their related
primary attachments). Advisory Circular AC No. 25.571-1 contains guidance
information relating to the requirements of this section (copies of the
advisory circular may be obtained from the U.S. Department of Transportation,
Publications Section M443.1, Washington, D.C. 20590). For turbojet powered
airplanes, those parts which could contribute to a catastrophic failure must
also be evaluated under paragraph (d) of this section. In addition, the
following apply:
(1) Each evaluation required by this section must include--
(i) The typical loading spectra, temperatures, and humidities expected in
service;
(ii) The identification of principal structural elements and detail design
points, the failure of which could cause catastrophic failure of the
airplane; and
(iii) An analysis, supported by test evidence, of the principal structural
elements and detail design points identified in paragraph (a)(1)(ii) of this
section.
(2) The service history of airplanes of similar structural design, taking
due account of differences in operating conditions and procedures, may be
used in the evaluations required by this section.
(3) Based on the evaluations required by this section, inspections or other
procedures must be established as necessary to prevent catastrophic failure,
and must be included in the Airworthiness Limitations section of the
Instruction for Continued Airworthiness required by Sec. 25.1529.
(b) Damage-tolerance evaluation. The evaluation must include a
determination of the probable locations and modes of damage due to fatigue,
corrosion, or accidental damage. The determination must be by analysis
supported by test evidence and (if available) service experience. Damage at
multiple sites due to prior fatigue exposure must be included where the
design is such that this type of damage can be expected to occur. The
evaluation must incorporate repeated load and static analyses supported by
test evidence. The extent of damage for residual strength evaluation at any
time within the operational life must be consistent with the initial
detectability and subsequent growth under repeated loads. The residual
strength evaluation must show that the remaining structure is able to
withstand loads (considered as static ultimate loads) corresponding to the
following conditions:
(1) The limit symmetrical maneuvering conditions specified in Sec. 25.337
at VC and in Sec. 25.345.
(2) The limit gust condition specified in Secs. 25.305(d), 25.341, and
25.351(b) at the specified speeds up to Vc, and in Sec. 25.345.
(3) The limit rolling conditions specified in Sec. 25.349 and the limit
unsymmetrical conditions specified in Secs. 25.367 and 25.427, at speeds up
to VC.
(4) The limit yaw maneuvering conditions specified in Sec. 25.351(a) at the
specified speeds up to VC.
(5) For pressurized cabins, the following conditions:
(i) The normal operating differential pressure combined with the expected
external aerodynamic pressures applied simultaneously with the flight loading
conditions specified in paragraphs (b) (1) through (4) of this section, if
they have a significant effect.
(ii) The expected external aerodynamic pressures in 1 g flight combined
with a cabin differential pressure equal to 1.1 times the normal operating
differential pressure without any other load.
(6) For landing gear and directly-affected airframe structure, the limit
ground loading conditions specified in Secs. 25.473, 25.491, and 25.493.
If significant changes in structural stiffness or geometry, or both, follow
from a structural failure, or partial failure, the effect on damage tolerance
must be further investigated.
(c) Fatigue (safe-life) evaluation. Compliance with the damage-tolerance
requirements of paragraph (b) of this section is not required if the
applicant establishes that their application for particular structure is
impractical. This structure must be shown by analysis, supported by test
evidence, to be able to withstand the repeated loads of variable magnitude
expected during its service life without detectable cracks. Appropriate safe-
life scatter factors must be applied.
(d) Sonic fatigue strength. It must be shown by analysis, supported by test
evidence, or by the service history of airplanes of similar structural design
and sonic excitation environment, that--
(1) Sonic fatigue cracks are not probable in any part of the flight
structure subject to sonic excitation; or
(2) Catastrophic failure caused by sonic cracks is not probable assuming
that the loads prescribed in paragraph (b) of this section are applied to all
areas affected by those cracks.
(e) Damage-tolerance (discrete source) evaluation. The airplane must be
capable of successfully completing a flight during which likely structural
damage occurs as a result of--
(1) Impact with a 4-pound bird at Vc at sea level to 8,000 feet;
(2) Uncontained fan blade impact;
(3) Uncontained engine failure; or
(4) Uncontained high energy rotating machinery failure.
The damaged structure must be able to withstand the static loads (considered
as ultimate loads) which are reasonably expected to occur on the flight.
Dynamic effects on these static loads need not be considered. Corrective
action to be taken by the pilot following the incident, such as limiting
maneuvers, avoiding turbulence, and reducing speed, must be considered. If
significant changes in structural stiffness or geometry, or both, follow from
a structural failure or partial failure, the effect on damage tolerance must
be further investigated.
[Amdt. 25-45, 43 FR 46242, Oct. 5, 1978, as amended by Amdt. 25-54, 45 FR
60173, Sept. 11, 1980; Amdt. 25-72, 55 FR 29776, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Lightning Protection
Sec. 25.581 Lightning protection.
(a) The airplane must be protected against catastrophic 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 airframe; 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 shown by--
(1) Designing the components to minimize the effect of a strike; or
(2) Incorporating acceptable means of diverting the resulting electrical
current so as not to endanger the airplane.
[Amdt. 25-23, 35 FR 5674, Apr. 8, 1970]
Subpart D--Design and Construction
General
Sec. 25.601 General.
The airplane may not have design features or details that experience has
shown to be hazardous or unreliable. The suitability of each questionable
design detail and part must be established by tests.
Sec. 25.603 Materials.
The suitability and durability of materials used for parts, the failure of
which could adversely affect safety, must--
(a) Be established on the basis of experience or tests;
(b) Conform to approved specifications (such as industry or military
specifications, or Technical Standard Orders) that ensure their having the
strength and other properties assumed in the design data; and
(c) Take into account the effects of environmental conditions, such as
temperature and humidity, expected in service.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 41 FR
55466, Dec. 20 1976; Amdt. 25-46, 43 FR 50595, Oct. 30, 1978]
Sec. 25.605 Fabrication methods.
(a) The methods of fabrication used must produce a consistently sound
structure. If a fabrication process (such as gluing, spot welding, or heat
treating) requires close control to reach this objective, the process must be
performed under an approved process specification.
(b) Each new aircraft fabrication method must be substantiated by a test
program.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-46, 43 FR
50595, Oct. 30, 1978]
Sec. 25.607 Fasteners.
(a) Each removable bolt, screw, nut, pin, or other removable fastener must
incorporate two separate locking devices if--
(1) Its loss could preclude continued flight and landing within the design
limitations of the airplane using normal pilot skill and strength; or
(2) Its loss could result in reduction in pitch, yaw, or roll control
capability or response below that required by Subpart B of this chapter.
(b) The fasteners specified in paragraph (a) of this section and their
locking devices may not be adversely affected by the environmental conditions
associated with the particular installation.
(c) No self-locking nut may be used on any bolt subject to rotation in
operation unless a nonfriction locking device is used in addition to the
self-locking device.
[Amdt. 25-23, 35 FR 5674, Apr. 8, 1970]
Sec. 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.
Sec. 25.611 Accessibility provisions.
Means must be provided to allow inspection (including inspection of
principal structural elements and control systems), replacement of parts
normally requiring replacement, adjustment, and lubrication as necessary for
continued airworthiness. The inspection means for each item must be
practicable for the inspection interval for the item. Nondestructive
inspection aids may be used to inspect structural elements where it is
impracticable to provide means for direct visual inspection if it is shown
that the inspection is effective and the inspection procedures are specified
in the maintenance manual required by Sec. 25.1529.
[Amdt. 25-23, 35 FR 5674, Apr. 8, 1970]
Sec. 25.613 Material strength properties and design values.
(a) Material strength properties must be based on enough tests of material
meeting approved specifications to establish design values on a statistical
basis.
(b) Design values must be chosen to minimize the probability of structural
failures due to material variability. Except as provided in paragraph (e) of
this section, compliance with this paragraph must be shown by selecting
design values which assure material strength with the following probability:
(1) Where applied loads are eventually distributed through a single member
within an assembly, the failure of which would result in loss of structural
integrity of the component, 99 percent probability with 95 percent
confidence.
(2) For redundant structure, in which the failure of individual elements
would result in applied loads being safely distributed to other load carrying
members, 90 percent probability with 95 percent confidence.
(c) The effects of temperature on allowable stresses used for design in an
essential component or structure must be considered where thermal effects are
significant under normal operating conditions.
(d) The strength, detail design, and fabrication of the structure must
minimize the probability of disastrous fatigue failure, particularly at
points of stress concentration.
(e) Greater design values may be used if a "premium selection" of the
material is made in which a specimen of each individual item is tested before
use to determine that the actual strength properties of that particular item
will equal or exceed those used in design.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-46, 43 FR
50595, Oct. 30, 1978; Amdt. 25-72, 55 FR 29776, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.615 [Removed. Amdt. 25-72, 55 FR 29776, July 20, 1990]
EDITORIAL NOTE: For the convenience of the user, the removed text is
set out below.
Sec. 25.615 Design properties.
(a) Design properties outlined in MIL-HDBK-5 may be used subject to the
following conditions:
(1) Where applied loads are eventually distributed through a single member
within an assembly, the failure of which would result in the loss of the
structural integrity of the component involved, the guaranteed minimum design
mechanical properties ("A" values) when listed in MIL-HDBK-5 must be met.
(2) Redundant structures, in which the failure of individual elements would
result in applied loads being safely distributed to other load-carrying
members, may be designed on the basis of the "90 percent probability ("B"
values)" when listed in MIL-HDBK-5.
(b) Design values greater than the guaranteed minimums required by
paragraph (a) of this section may be used where only guaranteed minimum
values are normally allowed if a "premium selection" of the material is made
in which a specimen of each individual item is tested before use to determine
that the actual strength properties of that particular item will equal or
exceed those used in design.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5674. Apr. 8, 1970]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.619 Special factors.
The factor of safety prescribed in Sec. 25.303 must be multiplied by the
highest pertinent special factor of safety prescribed in Secs. 25.621 through
25.625 for each part of the structure whose strength is--
(a) Uncertain;
(b) Likely to deteriorate in service before normal replacement; or
(c) Subject to appreciable variability because of uncertainties in
manufacturing processes or inspection methods.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5674, Apr. 8, 1970]
Sec. 25.621 Casting factors.
(a) General. The factors, tests, and inspections specified in paragraphs
(b) through (d) of this section must be applied in addition to those
necessary to establish foundry quality control. The inspections must meet
approved specifications. Paragraphs (c) and (d) of this section apply to any
structural castings except castings that are pressure tested as parts of
hydraulic or other fluid systems and do not support structural loads.
(b) Bearing stresses and surfaces. The casting factors specified in
paragraphs (c) and (d) of this section--
(1) Need not exceed 1.25 with respect to bearing stresses regardless of the
method of inspection used; and
(2) Need not be used with respect to the bearing surfaces of a part whose
bearing factor is larger than the applicable casting factor.
(c) Critical castings. For each casting whose failure would preclude
continued safe flight and landing of the airplane or result in serious injury
to occupants, 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 penetrant inspection methods or approved equivalent
nondestructive inspection methods.
(2) For each critical casting with a casting factor less than 1.50, three
sample castings must be static tested and shown to meet--
(i) The strength requirements of Sec. 25.305 at an ultimate load
corresponding to a casting factor of 1.25; and
(ii) The deformation requirements of Sec. 25.305 at a load of 1.15 times
the limit load.
(3) Examples of these castings are structural attachment fittings, parts of
flight control systems, control surface hinges and balance weight
attachments, seat, berth, safety belt, and fuel and oil tank supports and
attachments, and cabin pressure valves.
(d) Noncritical castings. For each casting other than those specified in
paragraph (c) of this section, the following apply:
(1) Except as provided in paragraphs (d) (2) and (3) of this section, the
casting factors and corresponding inspections must meet the following table:
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.
(2) The percentage of castings inspected by nonvisual methods may be
reduced below that specified in paragraph (d)(1) of this section when an
approved quality control procedure is established.
(3) For castings procured to a specification that guarantees the mechanical
properties of the material in the casting and provides for demonstration of
these properties by test of coupons cut from 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 factors of "1.25 through 1.50" and tested under paragraph
(c)(2) of this section.
Sec. 25.623 Bearing factors.
(a) Except as provided in paragraph (b) of this section, each part that has
clearance (free fit), and that is subject to pounding or vibration, must have
a bearing factor large enough to provide for the effects of normal relative
motion.
(b) No bearing factor need be used for a part for which any larger special
factor is prescribed.
Sec. 25.625 Fitting factors.
For each fitting (a part or terminal used to join one structural member to
another), the following apply:
(a) For each fitting whose strength is not proven by limit and ultimate
load tests in which actual stress conditions are simulated in the fitting and
surrounding structures, a fitting factor of at least 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 as continuous joints in metal plating, welded joints, and
scarf joints in wood); or
(2) With respect to any bearing surface for which a larger special factor
is used.
(c) For each integral fitting, the part must be treated as a fitting up to
the point at which the section properties become typical of the member.
(d) For each seat, berth, safety belt, and harness, the fitting factor
specified in Sec. 25.785(f)(3) applies.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5674, Apr. 8, 1970; Amdt. 25-72, 55 FR 29776, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.629 Aeroelastic stability requirements.
(a) General. The aeroelastic stability evaluations required under this
section include flutter, divergence, control reversal and any undue loss of
stability and control as a result of structural deformation. The aeroelastic
evaluation must include whirl modes associated with any propeller or rotating
device that contributes significant dynamic forces. Compliance with this
section must be shown by analyses, wind tunnel tests, ground vibration tests,
flight tests, or other means found necessary by the Administrator.
(b) Aeroelastic stability envelopes. The airplane must be designed to be
free from aeroelastic instability for all configurations and design
conditions within the aeroelastic stability envelopes as follows:
(1) For normal conditions without failures, malfunctions, or adverse
conditions, all combinations of altitudes and speeds encompassed by the VD/MD
versus altitude envelope enlarged at all points by an increase of 15 percent
in equivalent airspeed at both constant Mach number and constant altitude. In
addition, a proper margin of stability must exist at all speeds up to VD/MD
and, there must be no large and rapid reduction in stability as VD/MD is
approached. The enlarged envelope may be limited to Mach 1.0 when MD is less
than 1.0 at all design altitudes, and
(2) For the conditions described in Sec. 25.629(d) below, for all approved
altitudes, any airspeed up to the greater airspeed defined by;
(i) The VD/MD envelope determined by Sec. 25.335(b); or,
(ii) An altitude-airspeed envelope defined by a 15 percent increase in
equivalent airspeed above VC at constant altitude, from sea level to the
altitude of the intersection of 1.15 VC with the extension of the constant
cruise Mach number line, MC, then a linear variation in equivalent airspeed
to MC+.05 at the altitude of the lowest VC/MC intersection; then, at higher
altitudes, up to the maximum flight altitude, the boundary defined by a .05
Mach increase in MC at constant altitude.
(c) Balance weights. If concentrated balance weights are used, their
effectiveness and strength, including supporting structure, must be
substantiated.
(d) Failures, malfunctions, and adverse conditions. The failures,
malfunctions, and adverse conditions which must be considered in showing
compliance with this section are:
(1) Any critical fuel loading conditions, not shown to be extremely
improbable, which may result from mismanagement of fuel.
(2) Any single failure in any flutter damper system.
(3) For airplanes not approved for operation in icing conditions, the
maximum likely ice accumulation expected as a result of an inadvertent
encounter.
(4) Failure of any single element of the structure supporting any engine,
independently mounted propeller shaft, large auxiliary power unit, or large
externally mounted aerodynamic body (such as an external fuel tank).
(5) For airplanes with engines that have propellers or large rotating
devices capable of significant dynamic forces, any single failure of the
engine structure that would reduce the rigidity of the rotational axis.
(6) The absence of aerodynamic or gyroscopic forces resulting from the most
adverse combination of feathered propellers or other rotating devices capable
of significant dynamic forces. In addition, the effect of a single feathered
propeller or rotating device must be coupled with the failures of paragraphs
(d)(4) and (d)(5) of this section.
(7) Any single propeller or rotating device capable of significant dynamic
forces rotating at the highest likely overspeed.
(8) Any damage or failure condition, required or selected for investigation
by Sec. 25.571. The single structural failures described in paragraphs (d)(4)
and (d)(5) of this section need not be considered in showing compliance with
this section if;
(i) The structural element could not fail due to discrete source damage
resulting from the conditions described in Sec. 25.571(e), and
(ii) A damage tolerance investigation in accordance with Sec. 25.571(b)
shows that the maximum extent of damage assumed for the purpose of residual
strength evaluation does not involve complete failure of the structural
element.
(9) Any damage, failure, or malfunction considered under Secs. 25.631,
25.671, 25.672, and 25.1309.
(10) Any other combination of failures, malfunctions, or adverse conditions
not shown to be extremely improbable.
(e) Flight flutter testing. Full scale flight flutter tests at speeds up to
VDF/MDF must be conducted for new type designs and for modifications to a
type design unless the modifications have been shown to have an insignificant
effect on the aeroelastic stability. These tests must demonstrate that the
airplane has a proper margin of damping at all speeds up to VDF/MDF, and that
there is no large and rapid reduction in damping as VDF/MDF, is approached.
If a failure, malfunction, or adverse condition is simulated during flight
test in showing compliance with paragraph (d) of this section, the maximum
speed investigated need not exceed VFC/MFC if it is shown, by correlation of
the flight test data with other test data or analyses, that the airplane is
free from any aeroelastic instability at all speeds within the altitude-
airspeed envelope described in paragraph (b)(2) of this section.
[57 FR 28949, June 29, 1992]
*****************************************************************************
57 FR 28946, No. 125, June 29, 1992
SUMMARY: This amendment revises the airworthiness standards of the Federal
Aviation Regulations (FAR) for type certification of transport category
airplanes concerning vibration, buffet, flutter and divergence. It clarifies
the requirement to consider flutter and divergence when treating certain
damage and failure conditions required by other sections of the FAR and
adjusts the safety margins related to aeroelastic stabiity to make them more
appropriate for the conditions to which they apply. These changes are made to
provide consistency with other sections of the FAR and to take into account
advances in technology and the evolution of the design of transport
airplanes.
EFFECTIVE DATE: July 29, 1992.
*****************************************************************************
Sec. 25.631 Bird strike damage.
The empennage structure must be designed to assure capability of continued
safe flight and landing of the airplane after impact with an 8-pound bird
when the velocity of the airplane (relative to the bird along the airplane's
flight path) is equal to VC at sea level, selected under Sec. 25.335(a).
Compliance with this section by provision of redundant structure and
protected location of control system elements or protective devices such as
splitter plates or energy absorbing material is acceptable. Where compliance
is shown by analysis, tests, or both, use of data on airplanes having similar
structural design is acceptable.
[Amdt. 25-23, 35 FR 5674, Apr. 8, 1970]
Control Surfaces
Sec. 25.651 Proof of strength.
(a) Limit load tests of control surfaces are required. These tests must
include the horn or fitting to which the control system is attached.
(b) Compliance with the special factors requirements of Secs. 25.619
through 25.625 and 25.657 for control surface hinges must be shown by
analysis or individual load tests.
Sec. 25.655 Installation.
(a) Movable tail surfaces must be installed so that there is no
interference between any surfaces when one is held in its extreme position
and the others are operated through their full angular movement.
(b) If an adjustable stabilizer is used, it must have stops that will limit
its range of travel to the maximum for which the airplane is shown to meet
the trim requirements of Sec. 25.161.
Sec. 25.657 Hinges.
(a) For control surface hinges, including ball, roller, and self-lubricated
bearing hinges, the approved rating of the bearing may not be exceeded. For
nonstandard bearing hinge configurations, the rating must be established on
the basis of experience or tests and, in the absence of a rational
investigation, a factor of safety of not less than 6.67 must be used with
respect to the ultimate bearing strength of the softest material used as a
bearing.
(b) Hinges must have enough strength and rigidity for loads parallel to the
hinge line.
[Amdt. 25-23, 35 FR 5674, Apr. 8, 1970]
Control Systems
Sec. 25.671 General.
(a) Each control and control system must operate with the ease, smoothness,
and positiveness appropriate to its function.
(b) Each element of each flight control system must be designed, or
distinctively and permanently marked, to minimize the probability of
incorrect assembly that could result in the malfunctioning of the system.
(c) The airplane must be shown by analysis, tests, or both, to be capable
of continued safe flight and landing after any of the following failures or
jamming in the flight control system and surfaces (including trim, lift,
drag, and feel systems), within the normal flight envelope, without requiring
exceptional piloting skill or strength. Probable malfunctions must have only
minor effects on control system operation and must be capable of being
readily counteracted by the pilot.
(1) Any single failure, excluding jamming (for example, disconnection or
failure of mechanical elements, or structural failure of hydraulic
components, such as actuators, control spool housing, and valves).
(2) Any combination of failures not shown to be extremely improbable,
excluding jamming (for example, dual electrical or hydraulic system failures,
or any single failure in combination with any probable hydraulic or
electrical failure).
(3) Any jam in a control position normally encountered during takeoff,
climb, cruise, normal turns, descent, and landing unless the jam is shown to
be extremely improbable, or can be alleviated. A runaway of a flight control
to an adverse position and jam must be accounted for if such runaway and
subsequent jamming is not extremely improbable.
(d) The airplane must be designed so that it is controllable if all engines
fail. Compliance with this requirement may be shown by analysis where that
method has been shown to be reliable.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5674, Apr. 8, 1970]
Sec. 25.672 Stability augmentation and automatic and power-operated systems.
If the functioning of stability augmentation or other automatic or power-
operated systems is necessary to show compliance with the flight
characteristics requirements of this part, such systems must comply with Sec.
25.671 and the following:
(a) A warning which is clearly distinguishable to the pilot under expected
flight conditions without requiring his attention must be provided for any
failure in the stability augmentation system or in any other automatic or
power-operated system which could result in an unsafe condition if the pilot
were not aware of the failure. Warning systems must not activate the control
systems.
(b) The design of the stability augmentation system or of any other
automatic or power-operated system must permit initial counteraction of
failures of the type specified in Sec. 25.671(c) without requiring
exceptional pilot skill or strength, by either the deactivation of the
system, or a failed portion thereof, or by overriding the failure by movement
of the flight controls in the normal sense.
(c) It must be shown that after any single failure of the stability
augmentation system or any other automatic or power-operated system--
(1) The airplane is safely controllable when the failure or malfunction
occurs at any speed or altitude within the approved operating limitations
that is critical for the type of failure being considered;
(2) The controllability and maneuverability requirements of this part are
met within a practical operational flight envelope (for example, speed,
altitude, normal acceleration, and airplane configurations) which is
described in the Airplane Flight Manual; and
(3) The trim, stability, and stall characteristics are not impaired below a
level needed to permit continued safe flight and landing.
[Amdt. 25-23, 35 FR 5675 Apr. 8, 1970]
Sec. 25.673 [Removed. Amdt. 25-72, 55 FR 29777, July 20, 1990]
EDITORIAL NOTE: For the convenience of the user, the removed text is set
out below.
Sec. 25.673 Two-control airplanes.
Two-control airplanes must be able to continue safely in flight and landing
if any one connecting element in the directional-lateral flight control
system fails.
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.675 Stops.
(a) Each control system must have stops that positively limit the range of
motion of each movable aerodynamic surface controlled by the system.
(b) Each stop must be located so that wear, slackness, or take-up
adjustments will not adversely affect the control characteristics of the
airplane because of a change in the range of surface travel.
(c) Each stop must be able to withstand any loads corresponding to the
design conditions for the control system.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 41 FR
55466, Dec. 20, 1976]
Sec. 25.677 Trim systems.
(a) Trim controls must be designed to prevent inadvertent or abrupt
operation and to operate in the plane, and with the sense of motion, of the
airplane.
(b) There must be means adjacent to the trim control to indicate the
direction of the control movement relative to the airplane motion. In
addition, there must be clearly visible means to indicate the position of the
trim device with respect to the range of adjustment.
(c) Trim control systems must be designed to prevent creeping in flight.
Trim tab controls must be irreversible unless the tab is appropriately
balanced 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 irreversible unit to the airplane structure must
consist of a rigid connection.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5675, Apr. 8, 1970]
Sec. 25.679 Control system gust locks.
(a) There must be a device to prevent damage to the control surfaces
(including tabs), and to the control system, from gusts striking the airplane
while it is on the ground or water. If the device, when engaged, prevents
normal operation of the control surfaces by the pilot, it must--
(1) Automatically disengage when the pilot operates the primary flight
controls in a normal manner; or
(2) Limit the operation of the airplane so that the pilot receives
unmistakable warning at the start of takeoff.
(b) The device must have means to preclude the possibility of it becoming
inadvertently engaged in flight.
Sec. 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 attaching the system to the
main structure is included.
(b) Compliance must be shown (by analyses or individual load tests) with
the special factor requirements for control system joints subject to angular
motion.
Sec. 25.683 Operation tests.
It must be shown by operation tests that when portions of the control
system subject to pilot effort loads are loaded to 80 percent of the limit
load specified for the system and the powered portions of the control system
are loaded to the maximum load expected in normal operation, the system is
free from--
(a) Jamming;
(b) Excessive friction; and
(c) Excessive deflection.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5675, Apr. 8, 1970]
Sec. 25.685 Control system details.
(a) Each detail of each control system must be designed and installed to
prevent jamming, chafing, and interference from cargo, passengers, loose
objects, or the freezing of moisture.
(b) There must be means in the cockpit to prevent the entry of foreign
objects into places where they would jam the system.
(c) There must be means to prevent the slapping of cables or tubes against
other parts.
(d) Sections 25.689 and 25.693 apply to cable systems and joints.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 41 FR
55466, Dec. 20, 1976]
Sec. 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 rudder systems; and
(2) Each cable system must be designed so that there will be no hazardous
change in cable tension throughout the range of travel under operating
conditions and temperature variations.
(b) Each kind and size of pulley must correspond to the cable with which it
is used. Pulleys and sprockets must have closely fitted guards to prevent the
cables and chains from being displaced or fouled. Each pulley must lie in the
plane passing through the cable so that the cable does not rub against the
pulley flange.
(c) Fairleads must be installed so that they do not cause a change in cable
direction of more than three degrees.
(d) Clevis pins subject to load or motion and retained only by cotter pins
may not be used in the control system.
(e) Turnbuckles must be attached to parts having angular motion in a manner
that will positively prevent binding throughout the range of travel.
(f) There must be provisions for visual inspection of fairleads, pulleys,
terminals, and turnbuckles.
Sec. 25.693 Joints.
Control system joints (in push-pull systems) that are subject to angular
motion, except those in ball and roller bearing systems, must have a special
factor of safety of not less than 3.33 with respect to the ultimate bearing
strength of the softest material used as a bearing. This factor may be
reduced to 2.0 for joints in cable control systems. For ball or roller
bearings, the approved ratings may not be exceeded.
[Doc. No. 24344, Amdt. 25-72, 55 FR 29777, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 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
established under Sec. 25.101(d). Lift and drag devices must maintain the
selected positions, except for movement produced by an automatic positioning
or load limiting device, without further attention by the pilots.
(b) Each lift and drag device control must be designed and located to make
inadvertent operation improbable. Lift and drag devices intended for ground
operation only must have means to prevent the inadvertant operation of their
controls in flight if that operation could be hazardous.
(c) The rate of motion of the surfaces in response to the operation of the
control and the characteristics of the automatic positioning or load limiting
device must give satisfactory flight and performance characteristics under
steady or changing conditions of airspeed, engine power, and airplane
attitude.
(d) The lift device control must be designed to retract the surfaces from
the fully extended position, during steady flight at maximum continuous
engine power at any speed below VF +9.0 (knots).
[Amdt. 25-23, 35 FR 5675, Apr. 8, 1970, as amended by Amdt. 25-46, 43 FR
50595, Oct. 30, 1978; Amdt. 25-57, 49 FR 6848, Feb. 23, 1984]
Sec. 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 having a separate control in the cockpit to adjust its
position. In addition, an indication of unsymmetrical operation or other
malfunction in the lift or drag device systems must be provided when such
indication is necessary to enable the pilots to prevent or counteract an
unsafe flight or ground condition, considering the effects on flight
characteristics and performance.
(b) There must be means to indicate to the pilots the takeoff, en route,
approach, and landing lift device positions.
(c) If any extension of the lift and drag devices beyond the landing
position is possible, the controls must be clearly marked to identify this
range of extension.
[Amdt. 25-23, 35 FR 5675, Apr. 8, 1970]
Sec. 25.701 Flap and slat interconnection.
(a) Unless the airplane has safe flight characteristics with the flaps or
slats retracted on one side and extended on the other, the motion of flaps or
slats on opposite sides of the plane of symmetry must be synchronized by a
mechanical interconnection or approved equivalent means.
(b) If a wing flap or slat interconnection or equivalent means is used, it
must be designed to account for the applicable unsymmetrical loads, including
those resulting from flight with the engines on one side of the plane of
symmetry inoperative and the remaining engines at takeoff power.
(c) For airplanes with flaps or slats that are not subjected to slipstream
conditions, the structure must be designed for the loads imposed when the
wing flaps or slats on one side are carrying the most severe load occurring
in the prescribed symmetrical conditions and those on the other side are
carrying not more than 80 percent of that load.
(d) The interconnection must be designed for the loads resulting when
interconnected flap or slat surfaces on one side of the plane of symmetry are
jammed and immovable while the surfaces on the other side are free to move
and the full power of the surface actuating system is applied.
[Doc. No. 24344, Amdt. 25-72, 55 FR 29777, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 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 aural warning that is
automatically activated during the initial portion of the takeoff roll if the
airplane is in a configuration, including any of the following, that would
not allow a safe takeoff:
(1) The wing flaps or leading edge devices are not within the approved
range of takeoff positions.
(2) Wing spoilers (except lateral control spoilers meeting the requirements
of Sec. 25.671), speed brakes, or longitudinal trim devices are in a position
that would not allow a safe takeoff.
(b) The warning required by paragraph (a) of this section must continue
until--
(1) The configuration is changed to allow a safe takeoff;
(2) Action is taken by the pilot to terminate the takeoff roll;
(3) The airplane is rotated for takeoff; or
(4) The warning is manually deactivated by the pilot.
(c) The means used to activate the system must function properly throughout
the ranges of takeoff weights, altitudes, and temperatures for which
certification is requested.
[Amdt. 25-42, 43 FR 2323, Jan. 16, 1978]
Landing Gear
Sec. 25.721 General.
(a) The main landing gear system must be designed so that if it fails due
to overloads during takeoff and landing (assuming the overloads to act in the
upward and aft directions), the failure mode is not likely to cause--
(1) For airplanes that have passenger seating configuration, excluding
pilots seats, of nine seats or less, the spillage of enough fuel from any
fuel system in the fuselage to constitute a fire hazard; and
(2) For airplanes that have a passenger seating configuration, excluding
pilots seats, of 10 seats or more, the spillage of enough fuel from any part
of the fuel system to constitute a fire hazard.
(b) Each airplane that has a passenger seating configuration excluding
pilots seats, of 10 seats or more must be designed so that with the airplane
under control it can be landed on a paved runway with any one or more landing
gear legs not extended without sustaining a structural component failure that
is likely to cause the spillage of enough fuel to constitute a fire hazard.
(c) Compliance with the provisions of this section may be shown by analysis
or tests, or both.
[Amdt. 25-32, 37 FR 3969, Feb. 24, 1972]
Sec. 25.723 Shock absorption tests.
(a) It must be shown that the limit load factors selected for design in
accordance with Sec. 25.473 for takeoff and landing weights, respectively,
will not be exceeded. This must be shown by energy absorption tests except
that analyses based on earlier tests conducted on the same basic landing gear
system which has similar energy absorption characteristics may be used for
increases in previously approved takeoff and landing weights.
(b) The landing gear may not fail in a test, demonstrating its reserve
energy absorption capacity, simulating a descent velocity of 12 f.p.s. at
design landing weight, assuming airplane lift not greater than the airplane
weight acting during the landing impact.
[Amdt. 25-23, 35 FR 5675, Apr. 8, 1970, as amended by Amdt. 25-46, 43 FR
50595, Oct. 30, 1978; Amdt. 25-72, 55 FR 29777, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.725 Limit drop tests.
(a) If compliance with Sec. 25.723(a) is shown by free drop tests, these
tests must be made on the complete airplane, or on units consisting of a
wheel, tire, and shock absorber, in their proper positions, from free drop
heights not less than--
(1) 18.7 inches for the design landing weight conditions; and
(2) 6.7 inches for the design takeoff weight conditions.
(b) If airplane lift is simulated by air cylinders or by other mechanical
means, the weight used for the drop must be equal to W. If the effect of
airplane lift is represented in free drop tests by an equivalent reduced
mass, the landing gear must be dropped with an effective mass equal to
h+(1-L)d
We = W x ------------
h+d
where--
We= the effective weight to be used in the drop test (lbs.);
h =specified free drop height (inches);
d =deflection under impact of the tire (at the approved inflation pressure)
plus the vertical component of the axle travel relative to the drop mass
(inches);
W=WM for main gear units (lbs.), equal to the static weight on that unit with
the airplane in the level attitude (with the nose wheel clear in the case
of nose wheel type airplanes);
W=WT for tail gear units (lbs.), equal to the static weight on the tail unit
with the airplane in the tail-down attitude;
W=WN for nose wheel units (lbs.), equal to the vertical component of the
static reaction that would exist at the nose wheel, assuming that the
mass of the airplane acts at the center of gravity and exerts a force of
1.0 g downward and 0.25 g forward; and
L = ratio of the assumed airplane lift to the airplane weight, but not more
than 1.0.
(c) The drop test attitude of the landing gear unit and the application of
appropriate drag loads during the test must simulate the airplane landing
conditions in a manner consistent with the development of a rational or
conservative limit load factor value.
(d) The value of d used in the computation of We in paragraph (b) of this
section may not exceed the value actually obtained in the drop test.
(e) The limit inertia load factor n must be determined from the free drop
test in paragraph (b) of this section according to the following formula:
We
n = nj ---- + L
W
where--
nj =the load factor developed in the drop test (that is, the acceleration dv/
dt in g's recorded in the drop test) plus 1.0; and
We, W, and L are the same as in the drop test computation.
(f) The value of n determined in paragraph (e) of this section may not be
more than the limit inertia load factor used in the landing conditions in
Sec. 25.473.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5675, Apr. 8, 1970]
Sec. 25.727 Reserve energy absorption drop tests.
(a) If compliance with the reserve energy absorption condition specified in
Sec. 25.723(b) is shown by free drop tests, the drop height may not be less
than 27 inches.
(b) If airplane lift is simulated by air cylinders or by other mechanical
means, the weight used for the drop must be equal to W. If the effect of
airplane lift is represented in free drop tests by an equivalent reduced
mass, the landing gear must be dropped with an effective mass,
Wh
We = ------
h+d
where the symbols and other details are the same as in Sec. 25.725(b).
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5675, Apr. 8, 1970]
Sec. 25.729 Retracting mechanism.
(a) General. For airplanes with retractable landing gear, the following
apply:
(1) The landing gear retracting mechanism, wheel well doors, and supporting
structure, must be designed for--
(i) The loads occurring in the flight conditions when the gear is in the
retracted position,
(ii) The combination of friction loads, inertia loads, brake torque loads,
air loads, and gyroscopic loads resulting from the wheels rotating at a
peripheral speed equal to 1.3 Vs (with the flaps in takeoff position at
design takeoff weight), occurring during retraction and extension at any
airspeed up to 1.6 Vs1 (with the flaps in the approach position at design
landing weight), and
(iii) Any load factor up to those specified in Sec. 25.345(a) for the flaps
extended condition.
(2) Unless there are other means to decelerate the airplane in flight at
this speed, the landing gear, the retracting mechanism, and the airplane
structure (including wheel well doors) must be designed to withstand the
flight loads occurring with the landing gear in the extended position at any
speed up to 0.67 VC.
(3) Landing gear doors, their operating mechanism, and their supporting
structures must be designed for the yawing maneuvers prescribed for the
airplane in addition to the conditions of airspeed and load factor prescribed
in paragraphs (a) (1) and (2) of this section.
(b) Landing gear lock. There must be positive means to keep the landing
gear extended, in flight and on the ground.
(c) Emergency operation. There must be an emergency means for extending the
landing gear in the event of--
(1) Any reasonably probable failure in the normal retraction system; or
(2) The failure of any single source of hydraulic, electric, or equivalent
energy supply.
(d) Operation test. The proper functioning of the retracting mechanism must
be shown by operation tests.
(e) Position indicator and warning device. If a retractable landing gear is
used, there must be a landing gear position indicator (as well as necessary
switches to actuate the indicator) or other means to inform the pilot that
the gear is secured in the extended (or retracted) position. This means must
be designed as follows:
(1) If switches are used, they must be located and coupled to the landing
gear mechanical systems in a manner that prevents an erroneous indication of
"down and locked" if the landing gear is not in a fully extended position, or
of "up and locked" if the landing gear is not in the fully retracted
position. The switches may be located where they are operated by the actual
landing gear locking latch or device.
(2) The flightcrew must be given an aural warning that functions
continuously, or is periodically repeated, if a landing is attempted when the
landing gear is not locked down.
(3) The warning must be given in sufficient time to allow the landing gear
to be locked down or a go-around to be made.
(4) There must not be a manual shut-off means readily available to the
flightcrew for the warning required by paragraph (e)(2) of this section such
that it could be operated instinctively, inadvertently, or by habitual
reflexive action.
(5) The system used to generate the aural warning must be designed to
eliminate false or inappropriate alerts.
(6) Failures of systems used to inhibit the landing gear aural warning,
that would prevent the warning system from operating, must be improbable.
(f) Protection of equipment in wheel wells. Equipment that is essential to
safe operation of the airplane and that is located in wheel wells must be
protected from the damaging effects of--
(1) A bursting tire, unless it is shown that a tire cannot burst from
overheat; and
(2) A loose tire tread, unless it is shown that a loose tire tread cannot
cause damage.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5676, Apr. 8, 1970; Amdt. 25-42, 43 FR 2323, Jan. 16, 1978; Amdt. 25-72, 55
FR 29777, July 20, 1990; Amdt. 25-75, 56 FR 63762, Dec. 5, 1991]
*****************************************************************************
56 FR 63760, No. 234, Dec. 5, 1991
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the airworthiness standards for landing gear aural warning systems in
transport category airplanes to reflect current design practices. They
require that if a landing is attempted when the landing gear is not locked
down, the flightcrew must be given an aural warning in sufficient time to
allow the landing gear to be locked down or a go-around to be made. These
amendments state the intent of the current regulations in more objective
terms to eliminate nuisance warnings and to simplify the certification
process.
EFFECTIVE DATE: January 6, 1992.
*****************************************************************************
Sec. 25.731 Wheels.
(a) Each main and nose wheel must be approved.
(b) The maximum static load rating of each wheel may not be less than the
corresponding static ground reaction with--
(1) Design maximum weight; and
(2) Critical center of gravity.
(c) The maximum limit load rating of each wheel must equal or exceed the
maximum radial limit load determined under the applicable ground load
requirements of this part.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-72, 55
FR 29777, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.733 Tires.
(a) When a landing gear axle is fitted with a single wheel and tire
assembly, the wheel must be fitted with a suitable tire of proper fit with a
speed rating approved by the Administrator that is not exceeded under
critical conditions and with a load rating approved by the Administrator that
is not exceeded under--
(1) The loads on the main wheel tire, corresponding to the most critical
combination of airplane weight (up to maximum weight) and center of gravity
position, and
(2) The loads corresponding to the ground reactions in paragraph (b) of
this section, on the nose wheel tire, except as provided in paragraphs (b)(2)
and (b)(3) of this section.
(b) The applicable ground reactions for nose wheel tires are as follows:
(1) The static ground reaction for the tire corresponding to the most
critical combination of airplane weight (up to maximum ramp weight) and
center of gravity position with a force of 1.0g acting downward at the center
of gravity. This load may not exceed the load rating of the tire.
(2) The ground reaction of the tire corresponding to the most critical
combination of airplane weight (up to maximum landing weight) and center of
gravity position combined with forces of 1.0g downward and 0.31g forward
acting at the center of gravity. The reactions in this case must be
distributed to the nose and main wheels by the principles of statics with a
drag reaction equal to 0.31 times the vertical load at each wheel with brakes
capable of producing this ground reaction. This nose tire load may not exceed
1.5 times the load rating of the tire.
(3) The ground reaction of the tire corresponding to the most critical
combination of airplane weight (up to maximum ramp weight) and center of
gravity position combined with forces of 1.0g downward and 0.20g forward
acting at the center of gravity. The reactions in this case must be
distributed to the nose and main wheels by the principles of statics with a
drag reaction equal to 0.20 times the vertical load at each wheel with brakes
capable of producing this ground reaction. This nose tire load may not exceed
1.5 times the load rating of the tire.
(c) When a landing gear axle is fitted with more than one wheel and tire
assembly, such as dual or dual-tandem, each wheel must be fitted with a
suitable tire of proper fit with a speed rating approved by the Administrator
that is not exceeded under critical conditions, and with a load rating
approved by the Administrator that is not exceeded by--
(1) The loads on each main wheel tire, corresponding to the most critical
combination of airplane weight (up to maximum weight) and center of gravity
position, when multiplied by a factor of 1.07; and
(2) Loads specified in paragraphs (a)(2), (b)(1), (b)(2), and (b)(3) of
this section on each nose wheel tire.
(d) Each tire installed on a retractable landing gear system must, at the
maximum size of the tire type expected in service, have a clearance to
surrounding structure and systems that is adequate to prevent unintended
contact between the tire and any part of the structure or systems.
(e) For an airplane with a maximum certificated takeoff weight of more than
75,000 pounds, tires mounted on braked wheels must be inflated with dry
nitrogen or other gases shown to be inert so that the gas mixture in the tire
does not contain oxygen in excess of 5 percent by volume, unless it can be
shown that the tire liner material will not produce a volatile gas when
heated or that means are provided to prevent tire temperatures from reaching
unsafe levels.
[Amdt. 25-48, 44 FR 68752, Nov. 29, 1979, as amended by Amdt. 25-72, 55 FR
29777, July 20, 1990; Amdt. 25-78, 58 FR 11781, Feb. 26, 1993]
*****************************************************************************
58 FR 11778, No. 37, Feb. 26, 1993
SUMMARY: This amendment to the Federal Aviation Regulations (FAR) requires
that an inert gas, such as nitrogen, be used in lieu of air, for inflation of
tires on certain transport category airplanes. This action is prompted by at
least three cases in which the oxygen in air-filled tires combined with
volatile gases given off by a severely overheated tire and exploded upon
reaching autoignition temperature. The use of an inert gas for tire inflation
will eliminate the possibility of a tire explosion.
EFFECTIVE DATE: March 29, 1993.
*****************************************************************************
Sec. 25.735 Brakes.
(a) Each brake must be approved.
(b) The brake system and associated systems must be designed and
constructed so that if any electrical, pneumatic, hydraulic, or mechanical
connecting or transmitting element (excluding the operating pedal or handle)
fails, or if any single source of hydraulic or other brake operating energy
supply is lost, it is possible to bring the airplane to rest under conditions
specified in Sec. 25.125, with a mean deceleration during the landing roll of
at least 50 percent of that obtained in determining the landing distance as
prescribed in that section. Subcomponents within the brake assembly, such as
brake drum, shoes, and actuators (or their equivalents), shall be considered
as connecting or transmitting elements, unless it is shown that leakage of
hydraulic fluid resulting from failure of the sealing elements in these
subcomponents within the brake assembly would not reduce the braking
effectiveness below that specified in this paragraph.
(c) Brake controls may not require excessive control force in their
operation.
(d) The airplane must have a parking control that, when set by the pilot,
will without further attention, prevent the airplane from rolling on a paved,
level runway with takeoff power on the critical engine.
(e) If antiskid devices are installed, the devices and associated systems
must be designed so that no single probable malfunction will result in a
hazardous loss of braking ability or directional control of the airplane.
(f) The brake kinetic energy capacity rating of each main wheel-brake
assembly may not be less than the kinetic energy absorption requirements
determined under either of the following methods:
(1) The brake kinetic energy absorption requirements must be based on a
rational analysis of the sequence of events expected during operational
landings at maximum landing weight. This analysis must include conservative
values of airplane speed at which the brakes are applied, braking coefficient
of friction between tires and runway, aerodynamic drag, propeller drag or
power-plant forward thrust, and (if more critical) the most adverse single
engine or propeller malfunction.
(2) Instead of a rational analysis, the kinetic energy absorption
requirements for each main wheel brake assembly may be derived from the
following formula, which assumes an equal distribution of braking between
main wheels:
WV 2
KE = 0.0443 ----
N
where--
KE=Kinetic energy per wheel (ft.-lb.);
W=Design landing weight (lb.);
V=Airplane speed in knots. V must be not less than VS 0, the poweroff
stalling speed of the airplane at sea level, at the design landing
weight, and in the landing configuration; and
N=Number of main wheels with brakes.
The formula must be modified in cases of unequal braking distribution.
(g) The minimum stalling speed rating of each main wheel-brake assembly
(that is, the initial speed used in the dynamometer tests) may not be more
than the V used in the determination of kinetic energy in accordance with
paragraph (f) of this section, assuming that the test procedures for wheel-
brake assemblies involve a specified rate of deceleration, and, therefore,
for the same amount of kinetic energy, the rate of energy absorption (the
power absorbing ability of the brake) varies inversely with the initial
speed.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5676, Apr. 8, 1970; Amdt. 25-48, 44 FR 68742, Nov. 29, 1979; Amdt. 25-72, 55
FR 29777, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.737 Skis.
Each ski must be approved. The maximum limit load rating of each ski must
equal or exceed the maximum limit load determined under the applicable ground
load requirements of this part.
Floats and Hulls
Sec. 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 the seaplane or amphibian in fresh water; and
(b) Not less than five watertight compartments approximately equal in
volume.
Sec. 25.753 Main float design.
Each main float must be approved and must meet the requirements of Sec.
25.521.
Sec. 25.755 Hulls.
(a) Each hull must have enough watertight compartments so that, with any
two adjacent compartments flooded, the buoyancy of the hull and auxiliary
floats (and wheel tires, if used) provides a margin of positive stability
great enough to minimize the probability of capsizing in rough, fresh water.
(b) Bulkheads with watertight doors may be used for communication between
compartments.
Personnel and Cargo Accommodations
Sec. 25.771 Pilot compartment.
(a) Each pilot compartment and its equipment must allow the minimum flight
crew (established under Sec. 25.1523) to perform their duties without
unreasonable concentration or fatigue.
(b) The primary controls listed in Sec. 25.779(a), excluding cables and
control rods, must be located with respect to the propellers so that no
member of the minimum flight crew (established under Sec. 25.1523), or part
of the controls, lies in the region between the plane of rotation of any
inboard propeller and the surface generated by a line passing through the
center of the propeller hub making an angle of five degrees forward or aft of
the plane of rotation of the propeller.
(c) If provision is made for a second pilot, the airplane must be
controllable with equal safety from either pilot seat.
(d) The pilot compartment must be constructed so that, when flying in rain
or snow, it will not leak in a manner that will distract the crew or harm the
structure.
(e) Vibration and noise characteristics of cockpit equipment may not
interfere with safe operation of the airplane.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-4, 30 FR
6113, Apr. 30, 1965]
Sec. 25.772 Pilot compartment doors.
For an airplane that has a maximum passenger seating configuration of more
than 20 seats and that has a lockable door installed between the pilot
compartment and the passenger compartment:
(a) The emergency exit configuration must be designed so that neither
crewmembers nor passengers need use that door in order to reach the emergency
exits provided for them; and
(b) Means must be provided to enable flight crewmembers to directly enter
the passenger compartment from the pilot compartment if the cockpit door
becomes jammed.
[Doc. No. 24344, Admt. 25-72, 55 FR 29777, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.773 Pilot compartment view.
(a) Nonprecipitation conditions. For nonprecipitation conditions, the
following apply:
(1) Each pilot compartment must be arranged to give the pilots a
sufficiently extensive, clear, and undistorted view, to enable them to safely
perform any maneuvers within the operating limitations of the airplane,
including taxiing takeoff, approach, and landing.
(2) Each pilot compartment must be free of glare and reflection that could
interfere with the normal duties of the minimum flight crew (established
under Sec. 25.1523). This must be shown in day and night flight tests under
nonprecipitation conditions.
(b) Precipitation conditions. For precipitation conditions, the following
apply:
(1) The airplane must have a means to maintain a clear portion of the
windshield, during precipitation conditions, sufficient for both pilots to
have a sufficiently extensive view along the flight path in normal flight
attitudes of the airplane. This means must be designed to function, without
continuous attention on the part of the crew, in--
(i) Heavy rain at speeds up to 1.6 Vs1 with lift and drag devices
retracted; and
(ii) The icing conditions specified in Sec. 25.1419 if certification with
ice protection provisions is requested.
(2) The first pilot must have--
(i) A window that is openable under the conditions prescribed in paragraph
(b)(1) of this section when the cabin is not pressurized, provides the view
specified in that paragraph, and gives sufficient protection from the
elements against impairment of the pilot's vision; or
(ii) An alternate means to maintain a clear view under the conditions
specified in paragraph (b)(1) of this section, considering the probable
damage due to a severe hail encounter.
(c) Internal windshield and window fogging. The airplane must have a means
to prevent fogging of the internal portions of the windshield and window
panels over an area which would provide the visibility specified in paragraph
(a) of this section under all internal and external ambient conditions,
including precipitation conditions, in which the airplane is intended to be
operated.
(d) Fixed markers or other guides must be installed at each pilot station
to enable the pilots to position themselves in their seats for an optimum
combination of outside visibility and instrument scan. If lighted markers or
guides are used they must comply with the requirements specified in Sec.
25.1381.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5676, Apr. 8, 1970; Amdt. 25-46, 43 FR 50595, Oct. 30, 1978; Amdt. 25-72, 55
FR 29778, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.775 Windshields and windows.
(a) Internal panes must be made of nonsplintering material.
(b) Windshield panes directly in front of the pilots in the normal conduct
of their duties, and the supporting structures for these panes, must
withstand, without penetration, the impact of a four-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, selected under Sec.
25.335(a).
(c) Unless it can be shown by analysis or tests that the probability of
occurrence of a critical windshield fragmentation condition is of a low
order, the airplane must have a means to minimize the danger to the pilots
from flying windshield fragments due to bird impact. This must be shown for
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 longitudinal axis of the
airplane; and
(3) Has any part of the pane located where its fragmentation will
constitute a hazard to the pilots.
(d) The design of windshields and windows in pressurized airplanes must be
based on factors peculiar to high altitude operation, including the effects
of continuous and cyclic pressurization loadings, the inherent
characteristics of the material used, and the effects of temperatures and
temperature differentials. The windshield and window panels must be capable
of withstanding the maximum cabin pressure differential loads combined with
critical aerodynamic pressure and temperature effects after any single
failure in the installation or associated systems. It may be assumed that,
after a single failure that is obvious to the flight crew (established under
Sec. 25.1523), the cabin pressure differential is reduced from the maximum,
in accordance with appropriate operating limitations, to allow continued safe
flight of the airplane with a cabin pressure altitude of not more than 15,000
feet.
(e) The windshield panels in front of the pilots must be arranged so that,
assuming the loss of vision through any one panel, one or more panels remain
available for use by a pilot seated at a pilot station to permit continued
safe flight and landing.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5676, Apr. 8, 1970; Amdt. 25-38, 41 FR 55466, Dec. 20, 1976]
Sec. 25.777 Cockpit controls.
(a) Each cockpit control must be located to provide convenient operation
and to prevent confusion and inadvertent operation.
(b) The direction of movement of cockpit controls must meet the
requirements of Sec. 25.779. Wherever practicable, the sense of motion
involved in the operation of other controls must correspond to the sense of
the effect of the operation upon the airplane or upon the part operated.
Controls of a variable nature using a rotary motion must move clockwise from
the off position, through an increasing range, to the full on position.
(c) The controls must be located and arranged, with respect to the pilots'
seats, so that there is full and unrestricted movement of each control
without interference from the cockpit structure or the clothing of the
minimum flight crew (established under Sec. 25.1523) when any member of this
flight crew, from 5'2'' to 6'3'' in height, is seated with the seat belt and
shoulder harness (if provided) fastened.
(d) Identical powerplant controls for each engine must be located to
prevent confusion as to the engines they control.
(e) Wing flap controls and other auxiliary lift device controls must be
located on top of the pedestal, aft of the throttles, centrally or to the
right of the pedestal centerline, and not less than 10 inches aft of the
landing gear control.
(f) The landing gear control must be located forward of the throttles and
must be operable by each pilot when seated with seat belt and shoulder
harness (if provided) fastened.
(g) Control knobs must be shaped in accordance with Sec. 25.781. In
addition, the knobs must be of the same color, and this color must contrast
with the color of control knobs for other purposes and the surrounding
cockpit.
(h) If a flight engineer is required as part of the minimum flight crew
(established under Sec. 25.1523), the airplane must have a flight engineer
station located and arranged so that the flight crewmembers can perform their
functions efficiently and without interfering with each other.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-46, 43 FR
50596, Oct. 30, 1978]
Sec. 25.779 Motion and effect of cockpit controls.
Cockpit controls must be designed so that they operate in accordance with
the following movement and actuation:
(a) Aerodynamic controls:
(1) Primary.
Controls Motion and effect
Aileron Right (clockwise) for right wing down.
Elevator Rearward for nose up.
Rudder Right pedal forward for nose right.
(2) Secondary.
Controls Motion and effect
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.
(b) Powerplant and auxiliary controls:
(1) Powerplant.
Controls Motion and effect
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.
(2) Auxiliary.
Motion and
Controls effect
Landing gear Down to extend.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-72, 55
FR 29778, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.781 Cockpit control knob shape.
Cockpit control knobs must conform to the general shapes (but not
necessarily the exact sizes or specific proportions) in the following figure:
[ ...Illustration appears here... ]
Flap Control Knob
[ ...Illustration appears here... ]
Landing Gear Control Knob
[ ...Illustration appears here... ]
Mixture Control Knob
[ ...Illustration appears here... ]
Supercharger Control Knob
[ ...Illustration appears here... ]
Power or Thrust Knob
[ ...Illustration appears here... ]
Propeller Control Knob
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-72, 55 FR
29778, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.783 Doors.
(a) Each cabin must have at least one easily accessible external door.
(b) There must be a means to lock and safeguard each external door against
opening in flight (either inadvertently by persons or as a result of
mechanical failure or failure of a single structural element either during or
after closure). Each external door must be openable from both the inside and
the outside, even though persons may be crowded against the door on the
inside of the airplane. Inward opening doors may be used if there are means
to prevent occupants from crowding against the door to an extent that would
interfere with the opening of the door. The means of opening must be simple
and obvious and must be arranged and marked so that it can be readily located
and operated, even in darkness. Auxiliary locking devices may be used.
(c) Each external door must be reasonably free from jamming as a result of
fuselage deformation in a minor crash.
(d) Each external door must be located where persons using them will not be
endangered by the propellers when appropriate operating procedures are used.
(e) There must be a provision for direct visual inspection of the locking
mechanism to determine if external doors, for which the initial opening
movement is not inward (including passenger, crew, service, and cargo doors),
are fully closed and locked. The provision must be discernible under
operational lighting conditions by appropriate crewmembers using a flashlight
or equivalent lighting source. In addition, there must be a visual warning
means to signal the appropriate flight crewmembers if any external door is
not fully closed and locked. The means must be designed such that any failure
or combination of failures that would result in an erroneous closed and
locked indication is improbable for doors for which the initial opening
movement is not inward.
(f) External doors must have provisions to prevent the initiation of
pressurization of the airplane to an unsafe level if the door is not fully
closed and locked. In addition, it must be shown by safety analysis that
inadvertent opening is extemely improbable.
(g) Cargo and service doors not suitable for use as emergency exits need
only meet paragraphs (e) and (f) of this section and be safeguarded against
opening in flight as a result of mechanical failure or failure of a single
structural element.
(h) Each passenger entry door in the side of the fuselage must qualify as a
Type A, Type I, or Type II passenger emergency exit and must meet the
requirements of Secs. 25.807 through 25.813 that apply to that type of
passenger emergency exit.
(i) If an integral stair is installed in a passenger entry door that is
qualified as a passenger emergency exit, the stair must be designed so that
under the following conditions the effectiveness of passenger emergency
egress will not be impaired:
(1) The door, integral stair, and operating mechanism have been subjected
to the inertia forces specified in Sec. 25.561(b)(3), acting separately
relative to the surrounding structure.
(2) The airplane is in the normal ground attitude and in each of the
attitudes corresponding to collapse of one or more legs of the landing gear.
(j) All lavatory doors must be designed to preclude anyone from becoming
trapped inside the lavatory, and if a locking mechanism is installed, it be
capable of being unlocked from the outside without the aid of special tools.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-15, 32 FR
13262, Sept. 20, 1967; Amdt. 25-23, 35 FR 5676, Apr. 8, 1970; Amdt. 25-54, 45
FR 60173, Sept. 11, 1980; Amdt. 25-72, 55 FR 29780, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.785 Seats, berths, safety belts, and harnesses.
(a) A seat (or berth for a nonambulant person) must be provided for each
occupant who has reached his or her second birthday.
(b) Each seat, berth, safety belt, harness, and adjacent part of the
airplane at each station designated as occupiable during takeoff and landing
must be designed so that a person making proper use of these facilities will
not suffer serious injury in an emergency landing as a result of the inertia
forces specified in Secs. 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 plane containing the airplane centerline must be protected from
head injury by a safety belt and an energy absorbing rest that will support
the arms, shoulders, head, and spine, or by a safety belt and shoulder
harness that will prevent the head from contacting any injurious object. Each
occupant of any other seat must be protected from head injury by a safety
belt and, as appropriate to the type, location, and angle of facing of each
seat, by one or more of the following:
(1) A shoulder harness that will prevent the head from contacting any
injurious object.
(2) The elimination of any injurious object within striking radius of the
head.
(3) An energy absorbing rest that will support the arms, shoulders, head,
and spine.
(e) Each berth must be designed so that the forward part has a padded end
board, canvas diaphragm, or equivalent means, that can withstand the static
load reaction of the occupant when subjected to the forward inertia force
specified in Sec. 25.561. Berths must be free from corners and protuberances
likely to cause injury to a person occupying the berth during emergency
conditions.
(f) Each seat or berth, and its supporting structure, and each safety belt
or harness and its anchorage must be designed for an occupant weight of 170
pounds, considering the maximum load factors, inertia forces, and reactions
among the occupant, seat, safety belt, and harness for each relevant flight
and ground load condition (including the emergency landing conditions
prescribed in Sec. 25.561). In addition--
(1) The structural analysis and testing of the seats, berths, and their
supporting structures may be determined by assuming that the critical load in
the forward, sideward, downward, upward, and rearward directions (as
determined from the prescribed flight, ground, and emergency landing
conditions) acts separately or using selected combinations of loads if the
required strength in each specified direction is substantiated. The forward
load factor need not be applied to safety belts for berths.
(2) Each pilot seat must be designed for the reactions resulting from the
application of the pilot forces prescribed in Sec. 25.395.
(3) The inertia forces specified in Sec. 25.561 must be multiplied by a
factor of 1.33 (instead of the fitting factor prescribed in Sec. 25.625) in
determining the strength of the attachment of each seat to the structure and
each belt or harness to the seat or structure.
(g) Each seat at a flight deck station must have a restraint system
consisting of a combined safety belt and shoulder harness with a single-point
release that permits the flight deck occupant, when seated with the restraint
system fastened, to perform all of the occupant's necessary flight deck
functions. There must be a means to secure each combined restraint system
when not in use to prevent interference with the operation of the airplane
and with rapid egress in an emergency.
(h) Each seat located in the passenger compartment and designated for use
during takeoff and landing by a flight attendant required by the operating
rules of this chapter must be:
(1) Near a required floor level emergency exit, except that another
location is acceptable if the emergency egress of passengers would be
enhanced with that location. A flight attendant seat must be located adjacent
to each Type A emergency exit. Other flight attendant seats must be evenly
distributed among the required floor level emergency exits to the extent
feasible.
(2) To the extent possible, without compromising proximity to a required
floor level emergency exit, located to provide a direct view of the cabin
area for which the flight attendant is responsible.
(3) Positioned so that the seat will not interfere with the use of a
passageway or exit when the seat is not in use.
(4) Located to minimize the probability that occupants would suffer injury
by being struck by items dislodged from service areas, stowage compartments,
or service equipment.
(5) Either forward or rearward facing with an energy absorbing rest that is
designed to support the arms, shoulders, head, and spine.
(6) Equipped with a restraint system consisting of a combined safety belt
and shoulder harness unit with a single point release. There must be means to
secure each restraint system when not in use to prevent interference with
rapid egress in an emergency.
(i) Each safety belt must be equipped with a metal to metal latching
device.
(j) If the seat backs do not provide a firm handhold, there must be a
handgrip or rail along each aisle to enable persons to steady themselves
while using the aisles in moderately rough air.
(k) Each projecting object that would injure persons seated or moving about
the airplane in normal flight must be padded.
(l) Each forward observer's seat required by the operating rules must be
shown to be suitable for use in conducting the necessary enroute inspection.
[Doc. No. 24344, Amdt. 25-72, 55 FR 29780, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 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 compartment must be
designed for its placarded maximum weight of contents and for the critical
load distribution at the appropriate maximum load factors corresponding to
the specified flight and ground load conditions, and to the emergency landing
conditions of Sec. 25.561(b), except that the forces specified in the
emergency landing conditions need not be applied to compartments located
below, or forward, of all occupants in the airplane. If the airplane has a
passenger seating configuration, excluding pilots seats, of 10 seats or more,
each stowage compartment in the passenger cabin, except for underseat and
overhead compartments for passenger convenience, must be completely enclosed.
(b) There must be a means to prevent the contents in the compartments from
becoming a hazard by shifting, under the loads specified in paragraph (a) of
this section. For stowage compartments in the passenger and crew cabin, if
the means used is a latched door, the design must take into consideration the
wear and deterioration expected in service.
(c) If cargo compartment lamps are installed, each lamp must be installed
so as to prevent contact between lamp bulb and cargo.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended 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]
Sec. 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 design) in a passenger or crew compartment or galley from
becoming a hazard by shifting under the appropriate maximum load factors
corresponding to the specified flight and ground load conditions, and to the
emergency landing conditions of Sec. 25.561(b).
(b) Each interphone restraint system must be designed so that when
subjected to the load factors specified in Sec. 25.561(b)(3), the interphone
will remain in its stowed position.
[Amdt. 25-32, 37 FR 3969, Feb. 24, 1972, as amended by Amdt. 25-46, 43 FR
50596, Oct. 30, 1978]
Sec. 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 to each person seated in the cabin. If smoking is to
be allowed, and if the crew compartment is separated from the passenger
compartment, there must be at least one sign notifying when smoking is
prohibited. Signs which notify when smoking is prohibited must be operable by
a member of the flightcrew and, when illuminated, must be legible under all
probable conditions of cabin illumination to each person seated in the cabin.
(b) Signs that notify when seat belts should be fastened and that are
installed to comply with the operating rules of this chapter must be operable
by a member of the flightcrew and, when illuminated, must be legible under
all probable conditions of cabin illumination to each person seated in the
cabin.
(c) A placard must be located on or adjacent to the door of each receptacle
used for the disposal of flammable waste materials to indicate that use of
the receptacle for disposal of cigarettes, etc., is prohibited.
(d) Lavatories must have "No Smoking" or "No Smoking in Lavatory" placards
conspicuously located on or adjacent to each side of the entry door.
(e) Symbols that clearly express the intent of the sign or placard may be
used in lieu of letters.
[Doc. No. 24344, Amdt. 25-72, 55 FR 29780, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.793 Floor surfaces.
The floor surface of all areas which are likely to become wet in service
must have slip resistant properties.
[Amdt. 25-51, 45 FR 7755, Feb. 4, 1980]
Emergency Provisions
Sec. 25.801 Ditching.
(a) If certification with ditching provisions is requested, the airplane
must meet the requirements of this section and Secs. 25.807(e), 25.1411, and
25.1415(a).
(b) Each practicable design measure, compatible with the general
characteristics of the airplane, must be taken to minimize the probability
that in an emergency landing on water, the behavior of the airplane would
cause immediate injury to the occupants or would make it impossible for them
to escape.
(c) The probable behavior of the airplane in a water landing must be
investigated by model tests or by comparison with airplanes of similar
configuration for which the ditching characteristics are known. Scoops,
flaps, projections, and any other factor likely to affect the hydrodynamic
characteristics of the airplane, must be considered.
(d) It must be shown that, under reasonably probable water conditions, the
flotation time and trim of the airplane will allow the occupants to leave the
airplane and enter the liferafts required by Sec. 25.1415. If compliance with
this provision is shown by buoyancy and trim computations, appropriate
allowances must be made for probable structural damage and leakage. If the
airplane has fuel tanks (with fuel jettisoning provisions) that can
reasonably be expected to withstand a ditching without leakage, the
jettisonable volume of fuel may be considered as buoyancy volume.
(e) Unless the effects of the collapse of external doors and windows are
accounted for in the investigation of the probable behavior of the airplane
in a water landing (as prescribed in paragraphs (c) and (d) of this section),
the external doors and windows must be designed to withstand the probable
maximum local pressures.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-72, 55 FR
29781, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.803 Emergency evacuation.
(a) Each crew and passenger area must have emergency means to allow rapid
evacuation in crash landings, with the landing gear extended as well as with
the landing gear retracted, considering 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 that the maximum seating capacity, including the number of
crewmembers required by the operating rules for which certification is
requested, can be evacuated from the airplane to the ground under simulated
emergency conditions within 90 seconds. Compliance with this requirement must
be shown by actual demonstration using the test criteria outlined in appendix
J of this part unless the Administrator finds that a combination of analysis
and testing will provide data equivalent to that which would be obtained by
actual demonstration.
(d) [Reserved]
(e) [Reserved]
[Doc. No. 5066, 29 FR 18291 Dec. 24, 1964, as amended by Amdt. 25-15, 32 FR
13262, Sept. 20, 1967; Amdt. 25-20, 34 FR 5544, Mar. 22, 1969; Amdt. 25-32,
37 FR 3969, Feb. 24, 1972; Amdt. 25-46, 43 FR 50596, Oct. 30, 1978; Amdt.
25-72, 55 FR 29781, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.805 [Removed. 55 FR 29781, July 20, 1990]
EDITORIAL NOTE: For the convenience of the user, the removed text is
set out below.
Sec. 25.805 Flight crew emergency exits.
Except for airplanes with a passenger capacity of 20 or less in which the
proximity of passenger emergency exits to the flight crew area offers a
convenient and readily accessible means of evacuation for the flight crew,
the following apply:
(a) There must be either one exit on each side of the airplane or a top
hatch, in the flight crew area.
(b) Each exit must be of sufficient size and must be located so as to allow
rapid evacuation of the crew. An exit size and shape of other than at least
19 by 20 inches unobstructed rectangular opening may be used only if exit
utility is satisfactorily shown, by a typical flight crewmember, to the
Administrator.
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 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 inches wide by 48 inches high, with corner radii not greater
than one-third the width of the exit.
(2) Type II. This type is a rectangular opening of not less than 20 inches
wide by 44 inches high, with corner radii not greater than one-third the
width of the exit. Type II exits must be floor level exits unless located
over the wing, in which case they may not have a step-up inside the airplane
of more than 10 inches nor a step-down outside the airplane of more than 17
inches.
(3) Type III. This type is a rectangular opening of not less than 20 inches
wide by 36 inches high, with corner radii not greater than one-third the
width of the exit, and with a step-up inside the airplane of not more than 20
inches. If the exit is located over the wing, the step-down outside the
airplane may not exceed 27 inches.
(4) Type IV. This type is a rectangular opening of not less than 19 inches
wide by 26 inches high, with corner radii not greater than one-third the
width of the exit, located over the wing, with a step-up inside the airplane
of not more than 29 inches and a step-down outside the airplane of not more
than 36 inches.
(5) Ventral. This type is an exit from the passenger compartment through
the pressure shell and the bottom fuselage skin. The dimensions and physical
configuration of this type of exit must allow at least the same rate of
egress as a Type I exit with the airplane in the normal ground attitude, with
landing gear extended.
(6) Tail cone. This type is an aft exit from the passenger compartment
through the pressure shell and through an openable cone of the fuselage aft
of the pressure shell. The means of opening the tailcone must be simple and
obvious and must employ a single operation.
(7) Type A. This type is a floor level exit with a rectangular opening of
not less than 42 inches wide by 72 inches high with corner radii not greater
than one-sixth of the width of the exit.
(b) Step down distance. Step down distance, as used in this section, means
the actual distance between the bottom of the required opening and a usable
foot hold, extending out from the fuselage, that is large enough to be
effective without searching by sight or feel.
(c) Over-sized exits. Openings larger than those specified in this section,
whether or not of rectangular shape, may be used if the specified rectangular
opening can be inscribed within the opening and the base of the inscribed
rectangular opening meets the specified step-up and step-down heights.
(d) Passenger emergency exits. Except as provided in paragraphs (d) (3)
through (7) of this section, the minimum number and type of passenger
emergency exits is as follows:
(1) For passenger seating configurations of 1 through 299 seats:
Emergency exits for
each side of the
fuselage
Passenger
seating
configuration
(crewmember
seats not Type Type Type Type
included) I II III IV
1 through 9 1
10 through 19 1
20 through 39 1 1
40 through 79 1 1
80 through 109 1 2
110 through 139 2 1
140 through 179 2 2
Additional exits are required for passenger seating configurations greater
than 179 seats in accordance with the following table:
Additional
emergency Increase in
exits passenger
(each side seating
of configuration
fuselage) allowed
Type A 110
Type I 45
Type II 40
Type III 35
(2) For passenger seating configurations greater than 299 seats, each
emergency exit in the side of the fuselage must be either a Type A or Type I.
A passenger seating configuration of 110 seats is allowed for each pair of
Type A exits and a passenger seating configuration of 45 seats is allowed for
each pair of Type I exits.
(3) If a passenger ventral or tail cone exit is installed and that exit
provides at least the same rate of egress as a Type III exit with the
airplane in the most adverse exit opening condition that would result from
the collapse of one or more legs of the landing gear, an increase in the
passenger seating configuration beyond the limits specified in paragraph (d)
(1) or (2) of this section may be allowed as follows:
(i) For a ventral exit, 12 additional passenger seats.
(ii) For a tail cone exit incorporating a floor level opening of not less
than 20 inches wide by 60 inches high, with corner radii not greater than
one-third the width of the exit, in the pressure shell and incorporating an
approved assist means in accordance with Sec. 25.809(h), 25 additional
passenger seats.
(iii) For a tail cone exit incorporating an opening in the pressure shell
which is at least equivalent to a Type III emergency exit with respect to
dimensions, step-up and step-down distance, and with the top of the opening
not less than 56 inches from the passenger compartment floor, 15 additional
passenger seats.
(4) For airplanes on which the vertical location of the wing does not allow
the installation of overwing exits, an exit of at least the dimensions of a
Type III exit must be installed instead of each Type IV exit required by
subparagraph (1) of this paragraph.
(5) An alternate emergency exit configuration may be approved in lieu of
that specified in paragraph (d) (1) or (2) of this section provided the
overall evacuation capability is shown to be equal to or greater than that of
the specified emergency exit configuration.
(6) The following must also meet the applicable emergency exit requirements
of Secs. 25.809 through 25.813:
(i) Each emergency exit in the passenger compartment in excess of the
minimum number of required emergency exits.
(ii) Any other floor level door or exit that is accessible from the
passenger compartment and is as large or larger than a Type II exit, but less
than 46 inches wide.
(iii) Any other passenger ventral or tail cone exit.
(7) For an airplane that is required to have more than one passenger
emergency exit for each side of the fuselage, no passenger emergency exit
shall be more than 60 feet from any adjacent passenger emergency exit on the
same side of the same deck of the fuselage, as measured parallel to the
airplane's longitudinal axis between the nearest exit edges.
(e) Ditching emergency exits for passengers. Ditching emergency exits must
be provided in accordance with the following requirements whether or not
certification with ditching provisions is requested:
(1) For airplanes that have a passenger seating configuration of nine seats
or less, excluding pilots seats, one exit above the waterline in each side of
the airplane, meeting at least the dimensions of a Type IV exit.
(2) For airplanes that have a passenger seating configuration of 10 seats
or more, excluding pilots seats, one exit above the waterline in a side of
the airplane, meeting at least the dimensions of a Type III exit for each
unit (or part of a unit) of 35 passenger seats, but no less than two such
exits in the passenger cabin, with one on each side of the airplane. The
passenger seat/exit ratio may be increased through the use of larger exits,
or other means, provided it is shown that the evacuation capability during
ditching has been improved accordingly.
(3) If it is impractical to locate side exits above the waterline, the side
exits must be replaced by an equal number of readily accessible overhead
hatches of not less than the dimensions of a Type III exit, except that for
airplanes with a passenger configuration of 35 seats or less, excluding
pilots seats, the two required Type III side exits need be replaced by only
one overhead hatch.
(f) Flightcrew emergency exits. For airplanes in which the proximity of
passenger emergency exits to the flightcrew area does not offer a convenient
and readily accessible means of evacuation of the flightcrew, and for all
airplanes having a passenger seating capacity greater than 20, flightcrew
exits shall be located in the flightcrew area. Such exits shall be of
sufficient size and so located as to permit rapid evacuation by the crew. One
exit shall be provided on each side of the airplane; or, alternatively, a top
hatch shall be provided. Each exit must encompass an unobstructed rectangular
opening of at least 19 by 20 inches unless satisfactory exit utility can be
demonstrated by a typical crewmember.
[Doc. No. 24344, Admt. 25-72, 55 FR 29781, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.809 Emergency exit arrangement.
(a) Each emergency exit, including a flight crew emergency exit, must be a
movable door or hatch in the external walls of the fuselage, allowing
unobstructed opening to the outside.
(b) Each emergency exit must be openable from the inside and the outside
except that sliding window emergency exits in the flight crew area need not
be openable from the outside if other approved exits are convenient and
readily accessible to the flight crew area. Each emergency exit must be
capable of being opened, when there is no fuselage deformation--
(1) With the airplane in the normal ground attitude and in each of the
attitudes corresponding to collapse of one or more legs of the landing gear;
and
(2) Within 10 seconds measured from the time when the opening means is
actuated to the time when the exit is fully opened.
(c) The means of opening emergency exits must be simple and obvious and may
not require exceptional effort. Internal exit-opening means involving
sequence operations (such as operation of two handles or latches or the
release of safety catches) may be used for flight crew emergency exits if it
can be reasonably established that these means are simple and obvious to
crewmembers trained in their use.
(d) If a single power-boost or single power-operated system is the primary
system for operating more than one exit in an emergency, each exit must be
capable of meeting the requirements of paragraph (b) of this section in the
event of failure of the primary system. Manual operation of the exit (after
failure of the primary system) is acceptable.
(e) Each emergency exit must be shown by tests, or by a combination of
analysis and tests, to meet the requirements of paragraphs (b) and (c) of
this section.
(f) There must be a means to lock each emergency exit and to safeguard
against its opening in flight, either inadvertently by persons or as a result
of mechanical failure. In addition, there must be a means for direct visual
inspection of the locking mechanism by crewmembers to determine that each
emergency exit, for which the initial opening movement is outward, is fully
locked.
(g) There must be provisions to minimize the probability of jamming of the
emergency exits resulting from fuselage deformation in a minor crash landing.
(h) When required by the operating rules for any large passenger-carrying
turbojet-powered airplane, each ventral exit and tailcone exit must be--
(1) Designed and constructed so that it cannot be opened during flight; and
(2) Marked with a placard readable from a distance of 30 inches and
installed at a conspicuous location near the means of opening the exit,
stating that the exit has been designed and constructed so that it cannot be
opened during flight.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended 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 FR 50597, Oct. 30, 1978; Amdt.
25-47, 44 FR 61325, Oct. 25, 1979; Amdt. 25-72, 55 FR 29782, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.810 Emergency egress assist means and escape routes.
(a) Each nonoverwing landplane emergency exit more than 6 feet from the
ground with the airplane on the ground and the landing gear extended and each
nonoverwing Type A exit must have an approved means to assist the occupants
in descending to the ground.
(1) The assisting means for each passenger emergency exit must be a self-
supporting slide or equivalent; and, in the case of a Type A exit, it must be
capable of carrying simultaneously two parallel lines of evacuees. In
addition, the assisting means must be designed to meet the following
requirements:
(i) It must be automatically deployed and deployment must begin during the
interval between the time the exit opening means is actuated from inside the
airplane and the time the exit is fully opened. However, each passenger
emergency exit which is also a passenger entrance door or a service door must
be provided with means to prevent deployment of the assisting means when it
is opened from either the inside or the outside under nonemergency conditions
for normal use.
(ii) It must be automatically erected within 10 seconds after deployment is
begun.
(iii) It must be of such length after full deployment that the lower end is
self-supporting on the ground and provides safe evacuation of occupants to
the ground after collapse of one or more legs of the landing gear.
(iv) It must have the capability, in 25-knot winds directed from the most
critical angle, to deploy and, with the assistance of only one person, to
remain usable after full deployment to evacuate occupants safely to the
ground.
(v) For each system installation (mockup or airplane installed), five
consecutive deployment and inflation tests must be conducted (per exit)
without failure, and at least three tests of each such five-test series must
be conducted using a single representative sample of the device. The sample
devices must be deployed and inflated by the system's primary means after
being subjected to the inertia forces specified in Sec. 25.561(b). If any
part of the system fails or does not function properly during the required
tests, the cause of the failure or malfunction must be corrected by positive
means and after that, the full series of five consecutive deployment and
inflation tests must be conducted without failure.
(2) The assisting means for flightcrew emergency exits may be a rope or any
other means demonstrated to be suitable for the purpose. If the assisting
means is a rope, or an approved device equivalent to a rope, it must be--
(i) Attached to the fuselage structure at or above the top of the emergency
exit opening, or, for a device at a pilot's emergency exit window, at another
approved location if the stowed device, or its attachment, would reduce the
pilot's view in flight;
(ii) Able (with its attachment) to withstand a 400-pound static load.
(b) Assist means from the cabin to the wing are required for each Type A
exit located above the wing and having a stepdown unless the exit without an
assist means can be shown to have a rate of passenger egress at least equal
to that of the same type of nonoverwing exit. If an assist means is required,
it must be automatically deployed and automatically erected, concurrent with
the opening of the exit and self-supporting within 10 seconds.
(c) An escape route must be established from each overwing emergency exit,
and (except for flap surfaces suitable as slides) covered with a slip
resistant surface. Except where a means for channeling the flow of evacuees
is provided--
(1) The escape route must be at least 42 inches wide at Type A passenger
emergency exits and must be at least 2 feet wide at all other passenger
emergency exits, and
(2) The escape route surface must have a reflectance of at least 80
percent, and must be defined by markings with a surface-to-marking contrast
ratio of at least 5:1.
(d) If the place on the airplane structure at which the escape route
required in paragraph (c) of this section terminates, is more than 6 feet
from the ground with the airplane on the ground and the landing gear
extended, means to reach the ground must be provided to assist evacuees who
have used the escape route. If the escape route is over a flap, the height of
the terminal edge must be measured with the flap in the takeoff or landing
position, whichever is higher from the ground. The assisting means must be
usable and self-supporting with one or more landing gear legs collapsed and
under a 25-knot wind directed from the most critical angle. The assisting
means provided for each escape route leading from a Type A emergency exit
must be capable of carrying simultaneously two parallel lines of evacuees.
For other than Type A exits, the assist means must be capable of carrying
simultaneously as many parallel lines of evacuees as there are required
escape routes.
[Doc. No. 24344, Amdt. 25-72, 55 FR 29782, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.811 Emergency exit marking.
(a) Each passenger emergency exit, its means of access, and its means of
opening must be conspicuously marked.
(b) The identity and location of each passenger emergency exit must be
recognizable from a distance equal to the width of the cabin.
(c) Means must be provided to assist the occupants in locating the exits in
conditions of dense smoke.
(d) The location of each passenger emergency exit must be indicated by a
sign visible to occupants approaching along the main passenger aisle (or
aisles). There must be--
(1) A passenger emergency exit locator sign above the aisle (or aisles)
near each passenger emergency exit, or at another overhead location if it is
more practical because of low headroom, except that one sign may serve more
than one exit if each exit can be seen readily from the sign;
(2) A passenger emergency exit marking sign next to each passenger
emergency exit, except that one sign may serve two such exits if they both
can be seen readily from the sign; and
(3) A sign on each bulkhead or divider that prevents fore and aft vision
along the passenger cabin to indicate emergency exits beyond and obscured by
the bulkhead or divider, except that if this is not possible the sign may be
placed at another appropriate location.
(e) The location of the operating handle and instructions for opening exits
from the inside of the airplane must be shown in the following manner:
Each Type III passenger emergency exit operating handle must be self-
illuminated with an initial brightness of at least 160 microlamberts. If the
operating handle is covered, self-illuminated cover removal instructions
having an initial brightness of at least 160 microlamberts must also be
provided.
(1) Each passenger emergency exit must have, on or near the exit, a marking
that is readable from a distance of 30 inches.
(2) Each passenger emergency exit operating handle and the cover removal
instructions, if the operating handle is covered, must--
(i) Be self-illuminated with an initial brightness of at least 160
microlamberts; or
(ii) Be conspicuously located and well illuminated by the emergency
lighting even in conditions of occupant crowding at the exit.
(3) [Reserved]
(4) Each Type A, Type I, and Type II passenger emergency exit with a
locking mechanism released by rotary motion of the handle must be marked--
(i) With a red arrow, with a shaft at least three-fourths of an inch wide
and a head twice the width of the shaft, extending along at least 70 degrees
of arc at a radius approximately equal to three-fourths of the handle length.
(ii) So that the centerline of the exit handle is within +/-1 inch of the
projected point of the arrow when the handle has reached full travel and has
released the locking mechanism, and
(iii) With the word "open" in red letters 1 inch high, placed horizontally
near the head of the arrow.
(f) Each emergency exit that is required to be openable from the outside,
and its means of opening, must be marked on the outside of the airplane. In
addition, the following apply:
(1) The outside marking for each passenger emergency exit in the side of
the fuselage must include a 2-inch colored band outlining the exit.
(2) Each outside marking including the band, must have color contrast to be
readily distinguishable from the surrounding fuselage surface. The contrast
must be such that if the reflectance of the darker color is 15 percent or
less, the reflectance of the lighter color must be at least 45 percent.
"Reflectance" is the ratio of the luminous flux reflected by a body to the
luminous flux it receives. When the reflectance of the darker color is
greater than 15 percent, at least a 30-percent difference between its
reflectance and the reflectance of the lighter color must be provided.
(3) In the case of exists other than those in the side of the fuselage,
such as ventral or tail cone exists, the external means of opening, including
instructions if applicable, must be conspicuously marked in red, or bright
chrome yellow if the background color is such that red is inconspicuous. When
the opening means is located on only one side of the fuselage, a conspicuous
marking to that effect must be provided on the other side.
(g) Each sign required by paragraph (d) of this section may use the word
"exit" in its legend in place of the term "emergency exit".
[Amdt. 25-15, 32 FR 13264, Sept. 20, 1967, as amended by 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]
*****************************************************************************
58 FR 45224, No. 164, Aug. 26, 1993
SUMMARY: These amendments to the airworthiness standards for transport
category airplanes and the operating rules for air carrier operators of such
airplanes modify the procedures for conducting an emergency evacuation
demonstration. These include a requirement that the flightcrew take no active
role in the demonstration, and a change to the age/sex distribution
requirement for demonstration participants. In addition, the airworthiness
standards are amended to standardize the illumination requirements for the
handles of the various types of passenger emergency exits, and to add a
requirement to prevent the inadvertent disabling of the public address system
because of an unstowed microphone. These amendments are intended to enhance
the provisions for egress of occupants of transport category airplanes under
emergency conditions.
EFFECTIVE DATE: September 27, 1993.
*****************************************************************************
Sec. 25.812 Emergency lighting.
(a) An emergency lighting system, independent of the main lighting system,
must be installed. However, the sources of general cabin illumination may be
common to both the emergency and the main lighting systems if the power
supply to the emergency lighting system is independent of the power supply to
the main lighting system. The emergency lighting system must include:
(1) Illuminated emergency exit marking and locating signs, sources of
general cabin illumination, interior lighting in emergency exit areas, and
floor proximity escape path marking.
(2) Exterior emergency lighting.
(b) Emergency exit signs--
(1) For airplanes that have a passenger seating configuration, excluding
pilot seats, of 10 seats or more must meet the following requirements:
(i) Each passenger emergency exit locator sign required by Sec.
25.811(d)(1) and each passenger emergency exit marking sign required by Sec.
25.811(d)(2) must have red letters at least 1 1/2 inches high on an
illuminated white background, and must have an area of at least 21 square
inches excluding the letters. The lighted background-to-letter contrast must
be at least 10 : 1. The letter height to stroke-width ratio may not be more
than 7 : 1 nor less than 6 : 1. These signs must be internally electrically
illuminated with a background brightness of at least 25 foot-lamberts and a
high-to-low background contrast no greater than 3 : 1.
(ii) Each passenger emergency exit sign required by Sec. 25.811(d)(3) must
have red letters at least 1 1/2 inches high on a white background having an
area of at least 21 square inches excluding the letters. These signs must be
internally electrically illuminated or self-illuminated by other than
electrical means and must have an initial brightness 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 seating configuration, excluding
pilot seats, of nine seats or less, that are required by Sec. 25.811(d) (1),
(2), and (3) must have red letters at least 1 inch high on a white background
at least 2 inches high. These signs may be internally electrically
illuminated, or self-illuminated by other than electrical means, with an
initial brightness of at least 160 microlamberts. The colors may be reversed
in the case of a sign that is self-illuminated by other than electrical
means.
(c) General illumination in the passenger cabin must be provided so that
when measured along the centerline of main passenger aisle(s), and cross
aisle(s) between main aisles, at seat arm-rest height and at 40-inch
intervals, the average illumination is not less than 0.05 foot-candle and the
illumination at each 40-inch interval is not less than 0.01 foot-candle. A
main passenger aisle(s) is considered to extend along the fuselage from the
most forward passenger emergency exit or cabin occupant seat, whichever is
farther forward, to the most rearward passenger emergency exit or cabin
occupant seat, whichever is farther aft.
(d) The floor of the passageway leading to each floor-level passenger
emergency exit, between the main aisles and the exit openings, must be
provided with illumination that is not less than 0.02 foot-candle measured
along a line that is within 6 inches of and parallel to the floor and is
centered on the passenger evacuation path.
(e) Floor proximity emergency escape path marking must provide emergency
evacuation guidance for passengers when all sources of illumination more than
4 feet above the cabin aisle floor are totally obscured. In the dark of the
night, the floor proximity emergency escape path marking must enable each
passenger to--
(1) After leaving the passenger seat, visually identify the emergency
escape path along the cabin aisle floor to the first exits or pair of exits
forward and aft of the seat; and
(2) Readily identify each exit from the emergency escape path by reference
only to markings and visual features not more than 4 feet above the cabin
floor.
(f) Except for subsystems provided in accordance with paragraph (h) of this
section that serve no more than one assist means, are independent of the
airplane's main emergency lighting system, and are automatically activated
when the assist means is erected, the emergency lighting system must be
designed as follows.
(1) The lights must be operable manually from the flight crew station and
from a point in the passenger compartment that is readily accessible to a
normal flight attendant seat.
(2) There must be a flight crew warning light which illuminates when power
is on in the airplane and the emergency lighting control device is not armed.
(3) The cockpit control device must have an "on," "off," and "armed"
position so that when armed in the cockpit or turned on at either the cockpit
or flight attendant station the lights will either light or remain lighted
upon interruption (except an interruption caused by a transverse vertical
separation of the fuselage during crash landing) of the airplane's normal
electric power. There must be a means to safeguard against inadvertent
operation of the control device from the "armed" or "on" positions.
(g) Exterior emergency lighting must be provided as follows:
(1) At each overwing emergency exit the illumination must be--
(i) Not less than 0.03 foot-candle (measured normal to the direction of the
incident light) on a 2-square-foot area where an evacuee is likely to make
his first step outside the cabin;
(ii) Not less than 0.05 foot-candle (measured normal to the direction of
the incident light) for a minimum width of 42 inches for a Type A overwing
emergency exit and of 2 feet for all other overwing emergency exits along the
30 percent of the slip-resistant portion of the escape route required in Sec.
25.803(e) that is farthest from the exit; and
(iii) Not less than 0.03 foot-candle on the ground surface with the landing
gear extended (measured normal to the direction of the incident light) where
an evacuee using the established escape route would normally make first
contact with the ground.
(2) At each non-overwing emergency exit not required by Sec. 25.809(f) to
have descent assist means the illumination must be not less than 0.03 foot-
candle (measured normal to the direction of the incident light) on the ground
surface with the landing gear extended where an evacuee is likely to make his
first contact with the ground outside the cabin.
(h) The means required in Sec. 25.809 (f)(1) and (h) to assist the
occupants in descending to the ground must be illuminated so that the erected
assist means is visible from the airplane.
(1) If the assist means is illuminated by exterior emergency lighting, it
must provide illumination of not less than 0.03 foot-candle (measured normal
to the direction of the incident light) at the ground end of the erected
assist means where an evacuee using the established escape route would
normally make first contact with the ground, with the airplane in each of the
attitudes corresponding to the collapse of one or more legs of the landing
gear.
(2) If the emergency lighting subsystem illuminating the assist means
serves no other assist means, is independent of the airplane's main emergency
lighting system, and is automatically activated when the assist means is
erected, the lighting provisions--
(i) May not be adversely affected by stowage; and
(ii) Must provide illumination of not less than 0.03 foot-candle (measured
normal to the direction of incident light) at the ground and of the erected
assist means where an evacuee would normally make first contact with the
ground, with the airplane in each of the attitudes corresponding to the
collapse of one or more legs of the landing gear.
(i) The energy supply to each emergency lighting unit must provide the
required level of illumination for at least 10 minutes at the critical
ambient conditions after emergency landing.
(j) If storage batteries are used as the energy supply for the emergency
lighting system, they may be recharged from the airplane's main electric
power system: Provided, That, the charging circuit is designed to preclude
inadvertent battery discharge into charging circuit faults.
(k) Components of the emergency lighting system, including batteries,
wiring relays, lamps, and switches must be capable of normal operation after
having been subjected to the inertia forces listed in Sec. 25.561(b).
(l) The emergency lighting system must be designed so that after any single
transverse vertical separation of the fuselage during crash landing--
(1) Not more than 25 percent of all electrically illuminated emergency
lights required by this section are rendered inoperative, in addition to the
lights that are directly damaged by the separation;
(2) Each electrically illuminated exit sign required under Sec.
25.811(d)(2) remains operative exclusive of those that are directly damaged
by the separation; and
(3) At least one required exterior emergency light for each side of the
airplane remains operative exclusive of those that are directly damaged by
the separation.
[Amdt. 25-15, 32 FR 13265, Sept. 20, 1967, as amended by 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]
Sec. 25.813 Emergency exit access.
Each required emergency exit must be accessible to the passengers and
located where it will afford an effective means of evacuation. Emergency exit
distribution must be as uniform as practical, taking passenger distribution
into account; however, the size and location of exits on both 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 ventral emergency
exit, the floor level exit must be in the rearward part of the passenger
compartment, unless another location affords a more effective means of
passenger evacuation. Where more than one floor level exit per side is
prescribed, at least one floor level exit per side must be located near each
end of the cabin, except that this provision does not apply to combination
cargo/passenger configurations. In addition--
(a) There must be a passageway leading from the nearest main aisle to each
Type I, Type II, or Type A emergency exit and between individual passenger
areas. Each passageway leading to a Type A exit must be unobstructed and at
least 36 inches wide. Passageways between individual passenger areas and
those leading to Type I and Type II emergency exits must be unobstructed and
at least 20 inches wide. Unless there are two or more main aisles, each Type
A exit must be located so that there is passenger flow along the main aisle
to that exit from both the forward and aft directions. If two or more main
aisles are provided, there must be unobstructed cross-aisles at least 20
inches wide between main aisles. There must be--
(1) A cross-aisle which leads directly to each passageway between the
nearest main aisle and a Type A exit; and
(2) A cross-aisle which leads to the immediate vicinity of each passageway
between the nearest main aisle and a Type 1, Type II, or Type III exit;
except that when two Type III exits are located within three passenger rows
of each other, a single cross-aisle may be used if it leads to the vicinity
between the passageways from the nearest main aisle to each exit.
(b) Adequate space to allow crewmember(s) to assist in the evacuation of
passengers must be provided as follows:
(1) The assist space must not reduce the unobstructed width of the
passageway below that required for the exit.
(2) For each Type A exit, assist space must be provided at each side of the
exit regardless of whether the exit is covered by Sec. 25.810(a).
(3) For any other type exit that is covered by Sec. 25.810(a), space must
at least be provided at one side of the passageway.
(c) The following must be provided for each Type III or Type IV exit--(1)
There must be access from the nearest aisle to each exit. In addition, for
each Type III exit in an airplane that has a passenger seating configuration
of 60 or more--
(i) Except as provided in paragraph (c)(1)(ii), the access must be provided
by an unobstructed passageway that is at least 10 inches in width for
interior arrangements in which the adjacent seat rows on the exit side of the
aisle contain no more than two seats, or 20 inches in width for interior
arrangements in which those rows contain three seats. The width of the
passageway must be measured with adjacent seats adjusted to their most
adverse position. The centerline of the required passageway width must not be
displaced more than 5 inches horizontally from that of the exit.
(ii) In lieu of one 10- or 20-inch passageway, there may be two
passageways, between seat rows only, that must be at least 6 inches in width
and lead to an unobstructed space adjacent to each exit. (Adjacent exits must
not share a common passageway.) The width of the passageways must be measured
with adjacent seats adjusted to their most adverse position. The unobstructed
space adjacent to the exit must extend vertically from the floor to the
ceiling (or bottom of sidewall stowage bins), inboard from the exit for a
distance not less than the width of the narrowest passenger seat installed on
the airplane, and from the forward edge of the forward passageway to the aft
edge of the aft passageway. The exit opening must be totally within the fore
and aft bounds of the unobstructed space.
(2) In addition to the access--
(i) For airplanes that have a passenger seating configuration of 20 or
more, the projected opening of the exit provided must not be obstructed and
there must be no interference in opening the exit by seats, berths, or other
protrusions (including any seatback in the most adverse position) for a
distance from that exit not less than the width of the narrowest passenger
seat installed on the airplane.
(ii) For airplanes that have a passenger seating configuration of 19 or
fewer, there may be minor obstructions in this region, if there are
compensating factors to maintain the effectiveness of the exit.
(3) For each Type III exit, regardless of the passenger capacity of the
airplane in which it is installed, there must be placards that--
(i) Are readable by all persons seated adjacent to and facing a passageway
to the exit;
(ii) Accurately state or illustrate the proper method of opening the exit,
including the use of handholds; and
(iii) If the exit is a removable hatch, state the weight of the hatch and
indicate an appropriate location to place the hatch after removal.
(d) If it is necessary to pass through a passageway between passenger
compartments to reach any required emergency exit from any seat in the
passenger cabin, the passageway must be unobstructed. However, curtains may
be used if they allow free entry through the passageway.
(e) No door may be installed in any partition between passenger
compartments.
(f) If it is necessary to pass through a doorway separating the passenger
cabin from other areas to reach any required emergency exit from any
passenger seat, the door must have a means to latch it in open position. The
latching means must be able to withstand the loads imposed upon it when the
door is subjected to the ultimate inertia forces, relative to the surrounding
structure, listed in Sec. 25.561(b).
[Amdt. 25-1, 30 FR 3204, Mar. 9, 1965, as amended by Amdt. 25-15, 32 FR
13265, Sept. 20, 1967; Amdt. 25-32, 37 FR 3971, Feb. 24, 1972; Amdt. 25-46,
43 FR 50597, Oct. 30, 1978; Amdt. 25-72, 55 FR 29783, July 20, 1990; Amdt.
25-76, 57 FR 19244, May 4, 1992; 57 FR 29120, June 30, 1992]
*****************************************************************************
57 FR 19220, No. 86, May 4, 1992
Corrected. 57 FR 29120, No. 126, June 30, 1992
SUMMARY: This amendment revises the Federal Aviation Regulations (FAR) to
require improved access to the Type III emergency exits (typically smaller
over-wing exits) in transport category airplanes with 60 or more passenger
seats. These changes are the result of tests that were conducted at the
FAA's Civil Aeromedical Institute (CAMI), and are intended to improve the
ability of occupants to evacuate an airplane under emergency conditions.
They affect air carriers and commercial operators of transport category
airplanes as well as the manufacturers of such airplanes.
EFFECTIVE DATE: June 3, 1992.
*****************************************************************************
Sec. 25.815 Width of aisle.
The passenger aisle width at any point between seats must equal or exceed
the values in the following table:
Minimum
passenger
aisle width
(inches)
25
Less in.
than and
Passenger 25 in. more
seating from from
capacity floor floor
10 or less /1/ 12 15
11 through 19 12 20
20 or more 15 20
/1/ A narrower width not less than 9
inches may be approved when
substantiated by tests found necessary
by the Administrator.
[Amdt. 25-15, 32 FR 13265, Sept. 20, 1967, as amended by Amdt. 25-38, 41 FR
55466, Dec. 20, 1976]
Sec. 25.817 Maximum number of seats abreast.
On airplanes having only one passenger aisle, no more than three seats
abreast may be placed on each side of the aisle in any one row.
[Amdt. 25-15, 32 FR 13265, Sept. 20, 1967]
Sec. 25.819 Lower deck service compartments (including galleys).
For airplanes with a service compartment located below the main deck, which
may be occupied during taxi or flight but not during takeoff or landing, the
following apply:
(a) There must be at least two emergency evacuation routes, one at each end
of each lower deck service compartment or two having sufficient separation
within each compartment, which could be used by each occupant of the lower
deck service compartment to rapidly evacuate to the main deck under normal
and emergency lighting conditions. The routes must provide for the evacuation
of incapacitated persons, with assistance. The use of the evacuation routes
may not be dependent on any powered device. The routes must be designed to
minimize the possibility of blockage which might result from fire, mechanical
or structural failure, or persons standing on top of or against the escape
routes. In the event the airplane's main power system or compartment main
lighting system should fail, emergency illumination for each lower deck
service compartment must be automatically provided.
(b) There must be a means for two-way voice communication between the
flight deck and each lower deck service compartment.
(c) There must be an aural emergency alarm system, audible during normal
and emergency conditions, to enable crewmembers on the flight deck and at
each required floor level emergency exit to alert occupants of each lower
deck service compartment of an emergency situation.
(d) There must be a means, readily detectable by occupants of each lower
deck service compartment, that indicates when seat belts should be fastened.
(e) If a public address system is installed in the airplane, speakers must
be provided in each lower deck service compartment.
(f) For each occupant permitted in a lower deck service compartment, there
must be a forward or aft facing seat which meets the requirements of Sec.
25.785(c) and must be able to withstand maximum flight loads when occupied.
(g) For each powered lift system installed between a lower deck service
compartment and the main deck for the carriage of persons or equipment, or
both, the system must meet the following requirements:
(1) Each lift control switch outside the lift, except emergency stop
buttons, must be designed to prevent the activation of the life if the lift
door, or the hatch required by paragraph (g)(3) of this section, or both are
open.
(2) An emergency stop button, that when activated will immediately stop the
lift, must be installed within the lift and at each entrance to the lift.
(3) There must be a hatch capable of being used for evacuating persons from
the lift that is openable from inside and outside the lift without tools,
with the lift in any position.
[Amdt. 25-53, 45 FR 41593, June 19, 1980; 45 FR 43154, June 26, 1980]
Ventilation and Heating
Sec. 25.831 Ventilation.
(a) Each passenger and crew compartment must be ventilated, and each crew
compartment must have enough fresh air (but not less than 10 cu. ft. per
minute per crewmember) to enable crewmembers to perform their duties without
undue discomfort or fatigue.
(b) Crew and passenger compartment air must be free from harmful or
hazardous concentrations of gases or vapors. In meeting this requirement, the
following apply:
(1) Carbon monoxide concentrations in excess of 1 part in 20,000 parts of
air are considered hazardous. For test purposes, any acceptable carbon
monoxide detection method may be used.
(2) Carbon dioxide in excess of three percent by volume (sea level
equivalent) is considered hazardous in the case of crewmembers. Higher
concentrations of carbon dioxide may be allowed in crew compartments if
appropriate protective breathing equipment is available.
(c) There must be provisions made to ensure that the conditions prescribed
in paragraph (b) of this section are met after reasonably probable failures
or malfunctioning of the ventilating, heating, pressurization, or other
systems and equipment.
(d) If accumulation of hazardous quantities of smoke in the cockpit area is
reasonably probable, smoke evacuation must be readily accomplished, starting
with full pressurization and without depressurizing beyond safe limits.
(e) Except as provided in paragraph (f) of this section, means must be
provided to enable the occupants of the following compartments and areas to
control the temperature and quantity of ventilating air supplied to their
compartment or area independently of the temperature and quantity of air
supplied to other compartments and areas:
(1) The flight crew compartment.
(2) Crewmember compartments and areas other than the flight crew
compartment unless the crewmember compartment or area is ventilated by air
interchange with other compartments or areas under all operating conditions.
(f) Means to enable the flight crew to control the temperature and quantity
of ventilating air supplied to the flight crew compartment independently of
the temperature and quantity of ventilating air supplied to other
compartments are not required if all of the following conditions are met:
(1) The total volume of the flight crew and passenger compartments is 800
cubic feet or less.
(2) The air inlets and passages for air to flow between flight crew and
passenger compartments are arranged to provide compartment temperatures
within 5 degrees F. of each other and adequate ventilation to occupants in
both compartments.
(3) The temperature and ventilation controls are accessible to the flight
crew.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-41, 42 FR
36970, July 18, 1977]
Sec. 25.832 Cabin ozone concentration.
(a) The airplane cabin ozone concentration during flight must be shown now
to exceed--
(1) 0.25 parts per million by volume, sea level equivalent, at any time
above flight level 320; and
(2) 0.01 parts per million by volume, sea level equivalent, time-weighted
average during any 3-hour interval above flight level 270.
(b) For the purpose of this section, "sea level equivalent" refers to
conditions of 25 deg. C and 760 millimeters of mercury pressure.
(c) Compliance with this section must be shown by analysis or tests based
on airplane operational procedures and performance limitations, that
demonstrate that either--
(1) The airplane cannot be operated at an altitude which would result in
cabin ozone concentrations exceeding the limits prescribed by paragraph (a)
of this section; or
(2) The airplane ventilation system, including any ozone control equipment,
will maintain cabin ozone concentrations at or below the limits prescribed by
paragraph (a) of this section.
[Amdt. 25-50, 45 FR 3883, Jan. 1, 1980, as amended by Amdt. 25-56, 47 FR
58489, Dec. 30, 1982]
Sec. 25.833 Combustion heating systems.
Combustion heaters must be approved.
[Doc. No. 24344, Amdt. 25-72, 55 FR 29783, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Pressurization
Sec. 25.841 Pressurized cabins.
(a) Pressurized cabins and compartments to be occupied must be equipped to
provide a cabin pressure altitude of not more than 8,000 feet at the maximum
operating altitude of the airplane under normal operating conditions. If
certification for operation over 25,000 feet is requested, the airplane must
be able to maintain a cabin pressure altitude of not more than 15,000 feet in
the event of any reasonably probable failure or malfunction in the
pressurization system.
(b) Pressurized cabins must have at least the following valves, controls,
and indicators for controlling cabin pressure:
(1) Two pressure relief valves to automatically limit the positive pressure
differential to a predetermined value at the maximum rate of flow delivered
by the pressure source. The combined capacity of the relief valves must be
large enough so that the failure of any one valve would not cause an
appreciable rise in the pressure differential. The pressure differential is
positive when the internal pressure is greater than the external.
(2) Two reverse pressure differential relief valves (or their equivalents)
to automatically prevent a negative pressure differential that would damage
the structure. One valve is enough, however, if it is of a design that
reasonably precludes its malfunctioning.
(3) A means by which the pressure differential can be rapidly equalized.
(4) An automatic or manual regulator for controlling the intake or exhaust
airflow, or both, for maintaining the required internal pressures and airflow
rates.
(5) Instruments at the pilot or flight engineer station to show the
pressure differential, the cabin pressure altitude, and the rate of change of
the cabin pressure altitude.
(6) Warning indication at the pilot or flight engineer station to indicate
when the safe or preset pressure differential and cabin pressure altitude
limits are exceeded. Appropriate warning markings on the cabin pressure
differential indicator meet the warning requirement for pressure differential
limits and an aural or visual signal (in addition to cabin altitude
indicating means) meets the warning requirement for cabin pressure altitude
limits if it warns the flight crew when the cabin pressure altitude exceeds
10,000 feet.
(7) A warning placard at the pilot or flight engineer station if the
structure is not designed for pressure differentials up to the maximum relief
valve setting in combination with landing loads.
(8) The pressure sensors necessary to meet the requirements of paragraphs
(b)(5) and (b)(6) of this section and Sec. 25.1447(c), must be located and
the sensing system designed so that, in the event of loss of cabin pressure
in any passenger or crew compartment (including upper and lower lobe
galleys), the warning and automatic presentation devices, required by those
provisions, will be actuated without any delay that would significantly
increase the hazards resulting from decompression.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 41 FR
55466, Dec. 20, 1976]
Sec. 25.843 Tests for pressurized cabins.
(a) Strength test. The complete pressurized cabin, including doors,
windows, and valves, must be tested as a pressure vessel for the pressure
differential specified in Sec. 25.365(d).
(b) Functional tests. The following functional tests must be performed:
(1) Tests of the functioning and capacity of the positive and negative
pressure differential valves, and of the emergency release valve, to
stimulate the effects of closed regulator valves.
(2) Tests of the pressurization system to show proper functioning under
each possible condition of pressure, temperature, and moisture, up to the
maximum altitude for which certification is requested.
(3) Flight tests, to show the performance of the pressure supply, pressure
and flow regulators, indicators, and warning signals, in steady and stepped
climbs and descents at rates corresponding to the maximum attainable within
the operating limitations of the airplane, up to the maximum altitude for
which certification is requested.
(4) Tests of each door and emergency exit, to show that they operate
properly after being subjected to the flight tests prescribed in paragraph
(b)(3) of this section.
Fire Protection
Sec. 25.851 Fire extinguishers.
(a) Hand fire extinguishers. (1) The following minimum number of hand fire
extinguishers must be conveniently located and evenly distributed in
passenger compartments:
Passenger No. of
capacity extinguishers
7 through 30 1
31 through 60 2
61 through 200 3
201 through 300 4
301 through 400 5
401 through 500 6
501 through 600 7
601 through 700 8
(2) At least one hand fire extinguisher must be conveniently located in the
pilot compartment.
(3) At least one readily accessible hand fire extinguisher must be
available for use in each Class A or Class B cargo or baggage compartment and
in each Class E cargo or baggage compartment that is accessible to
crewmembers in flight.
(4) At least one hand fire extinguisher must be located in, or readily
accessible for use in, each galley located above or below the passenger
compartment.
(5) Each hand fire extinguisher must be approved.
(6) At least one of the required fire extinguishers located in the
passenger compartment of an airplane with a passenger capacity of at least 31
and not more than 60, and at least two of the fire extinguishers located in
the passenger compartment of an airplane with a passenger capacity of 61 or
more must contain Halon 1211 (bromochlorodifluoromethane CBrC1F2), or
equivalent, as the extinguishing agent. The type of extinguishing agent used
in any other extinguisher required by this section must be appropriate for
the kinds of fires likely to occur where used.
(7) The quantity of extinguishing agent used in each extinguisher required
by this section must be appropriate for the kinds of fires likely to occur
where used.
(8) Each extinguisher intended for use in a personnel compartment must be
designed to minimize the hazard of toxic gas concentration.
(b) Built-in fire extinguishers. If a built-in fire extinguisher is
provided--
(1) Each built-in fire extinguishing system must be installed so that--
(i) No extinguishing agent likely to enter personnel compartments will be
hazardous to the occupants; and
(ii) No discharge of the extinguisher can cause structural damage.
(2) The capacity of each required built-in fire extinguishing system must
be adequate for any fire likely to occur in the compartment where used,
considering the volume of the compartment and the ventilation rate.
[56 FR 15456, April 16, 1991]
*****************************************************************************
56 FR 15450, No. 73, Apr. 16, 1991
SUMMARY: This amendment provides improved cabin fire protection for
transport category airplanes by requiring: (1) Each lavatory in an airplane
with a passenger seating capacity of 20 or more to be equipped with a smoke
detector system that provides warning to the cockpit or to the passenger
cabin crew; (2) each lavatory trash receptacle in an airplane with a seating
capacity of 20 or more to be equipped with a fire extinguisher that
discharges automatically upon the occurrence of a fire within the
receptacle; (3) the number of hand fire extinguishers in the cabins of
airplanes with passenger seating capacities greater than 200 to be
increased; (4) a specified number of the hand fire extinguishers in the
cabin to contain Halon 1211 or equivalent as the extinguishing agent; and
(5) one hand fire extinguisher in each galley that is located above or below
the passenger compartment. In addition, one hand fire extinguisher would be
required for certain all-cargo airplanes. These safety protections against
possible inflight fires are currently required for operation of airplanes
used in air carrier or commercial service. This amendment adopts these
requirements as design standards for transport category airplanes.
EFFECTIVE DATE: May 16, 1991.
*****************************************************************************
Sec. 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 the applicable test criteria prescribed in part I of
appendix F of this part or other approved equivalent methods.
(b) In addition to meeting the requirements of paragraph (a), seat
cushions, except those on flight crewmember seats, must meet the test
requirements of part II of appendix F of this part, or equivalent.
(c) For airplanes with passenger capacities of 20 or more, interior ceiling
and wall panels (other than lighting lenses), partitions, and the outer
surfaces of galleys, large cabinets and stowage compartments (other than
underseat stowage compartments and compartments for stowing small items, such
as magazines and maps) must also meet the test requirements of parts IV and V
of appendix F of this part, or other approved equivalent method, in addition
to the flammability requirements prescribed in paragraph (a) of this section.
(d) Smoking is not to be allowed in lavatories. If smoking is to be allowed
in any compartment occupied by the crew or passengers, an adequate number of
self-contained, removable ashtrays must be provided for all seated occupants,
and
(e) Regardless of whether smoking is allowed in any other part of the
airplane, lavatories must have self-contained removable ashtrays located
conspicuously on or near the entry side of each lavatory door, except that
one ashtray may serve more than one lavatory door if the ashtray can be seen
readily from the cabin side of each lavatory served.
(f) Each receptacle used for the disposal of flammable waste material must
be fully enclosed, constructed of at least fire resistant materials, and must
contain fires likely to occur in it under normal use. The ability of the
receptacle to contain those fires under all probable conditions of wear,
misalignment, and ventilation expected in service must be demonstrated by
test.
[Doc. No. 24344, Amdt. 25-72, 55 FR 29783, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 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 provides a warning light in the cockpit, or provides a
warning light or audible warning in the passenger cabin that would be readily
detected by a flight attendant; and
(b) Each lavatory must be equipped with a built-in fire extinguisher for
each disposal receptacle for towels, paper, or waste, located within the
lavatory. The extinguisher must be designed to discharge automatically into
each disposal receptacle upon occurrence of a fire in that receptacle.
[56 FR 15456, April 16, 1991]
*****************************************************************************
56 FR 15450, No. 73, Apr. 16, 1991
SUMMARY: This amendment provides improved cabin fire protection for
transport category airplanes by requiring: (1) Each lavatory in an airplane
with a passenger seating capacity of 20 or more to be equipped with a smoke
detector system that provides warning to the cockpit or to the passenger
cabin crew; (2) each lavatory trash receptacle in an airplane with a seating
capacity of 20 or more to be equipped with a fire extinguisher that
discharges automatically upon the occurrence of a fire within the
receptacle; (3) the number of hand fire extinguishers in the cabins of
airplanes with passenger seating capacities greater than 200 to be
increased; (4) a specified number of the hand fire extinguishers in the
cabin to contain Halon 1211 or equivalent as the extinguishing agent; and
(5) one hand fire extinguisher in each galley that is located above or below
the passenger compartment. In addition, one hand fire extinguisher would be
required for certain all-cargo airplanes. These safety protections against
possible inflight fires are currently required for operation of airplanes
used in air carrier or commercial service. This amendment adopts these
requirements as design standards for transport category airplanes.
EFFECTIVE DATE: May 16, 1991.
*****************************************************************************
Sec. 25.855 Cargo or baggage compartments.
For each cargo and baggage compartment not occupied by crew or passengers,
the following apply:
(a) The compartment must meet one of the class requirements of Sec. 25.857.
(b) Class B through Class E cargo or baggage compartments, as defined in
Sec. 25.857, must have a liner, and the liner must be separate from (but may
be attached to) the airplane structure.
(c) Ceiling and sidewall liner panels of Class C and D compartments must
meet the test requirements of part III of appendix F of this part or other
approved equivalent methods.
(d) All other materials used in the construction of the cargo or baggage
compartment must meet the applicable test criteria prescribed in part I of
appendix F of this part or other approved equivalent methods.
(e) No compartment may contain any controls, wiring, lines, equipment, or
accessories whose damage or failure would affect safe operation, unless those
items are protected so that--
(1) They cannot be damaged by the movement of cargo in the compartment, and
(2) Their breakage or failure will not create a fire hazard.
(f) There must be means to prevent cargo or baggage from interfering with
the functioning of the fire protective features of the compartment.
(g) Sources of heat within the compartment must be shielded and insulated
to prevent igniting the cargo or baggage.
(h) Flight tests must be conducted to show compliance with the provisions
of Sec. 25.857 concerning--
(1) Compartment accessibility,
(2) The entries of hazardous quantities of smoke or extinguishing agent
into compartments occupied by the crew or passengers, and
(3) The dissipation of the extinguishing agent in Class C compartments.
(i) During the above tests, it must be shown that no inadvertent operation
of smoke or fire detectors in any compartment would occur as a result of fire
contained in any other compartment, either during or after extinguishment,
unless the extinguishing system floods each such compartment simultaneously.
[Doc. No. 24344, Amdt. 25-72, 55 FR 29784, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.857 Cargo compartment classification.
(a) Class A; A Class A cargo or baggage compartment is one in which--
(1) The presence of a fire would be easily discovered by a crewmember while
at his station; and
(2) Each part of the compartment is easily accessible in flight.
(b) Class B. A Class B cargo or baggage compartment is one in which--
(1) There is sufficient access in flight to enable a crewmember to
effectively reach any part of the compartment with the contents of a hand
fire extinguisher;
(2) When the access provisions are being used, no hazardous quantity of
smoke, flames, or extinguishing agent, will enter any compartment occupied by
the crew or passengers;
(3) There is a separate approved smoke detector or fire detector system to
give warning at the pilot or flight engineer station.
(c) Class C. A Class C cargo or baggage compartment is one not meeting the
requirements for either a Class A or B compartment but in which--
(1) There is a separate approved smoke detector or fire detector system to
give warning at the pilot or flight engineer station;
(2) There is an approved built-in fire-extinguishing system controllable
from the pilot or flight engineer stations;
(3) There are means to exclude hazardous quantities of smoke, flames, or
extinguishing agent, from any compartment occupied by the crew or passengers;
(4) There are means to control ventilation and drafts within the
compartment so that the extinguishing agent used can control any fire that
may start within the compartment.
(d) Class D. A Class D cargo or baggage compartment is one in which--
(1) A fire occurring in it will be completely confined without endangering
the safety of the airplane or the occupants;
(2) There are means to exclude hazardous quantities of smoke, flames, or
other noxious gases, from any compartment occupied by the crew or passengers;
(3) Ventilation and drafts are controlled within each compartment so that
any fire likely to occur in the compartment will not progress beyond safe
limits; and
(4) [Reserved]
(5) Consideration is given to the effect of heat within the compartment on
adjacent critical parts of the airplane. For compartments of 500 cu. ft. or
less, an airflow of 1500 cu. ft. per hour is acceptable.
(6) The compartment volume does not exceed 1,000 cubic feet.
(e) Class E. A Class E cargo compartment is one on airplanes used only for
the carriage of cargo and in which--
(1) [Reserved]
(2) There is a separate approved smoke or fire detector system to give
warning at the pilot or flight engineer station;
(3) There are means to shut off the ventilating airflow to, or within, the
compartment, and the controls for these means are accessible to the flight
crew in the crew compartment;
(4) There are means to exclude hazardous quantities of smoke, flames, or
noxious gases, from the flight crew compartment; and
(5) The required crew emergency exits are accessible under any cargo
loading condition.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-32, 37 FR
3972, Feb. 24, 1972; Amdt. 25-60, 51 FR 18243, May 16, 1986]
Sec. 25.858 Cargo compartment fire detection systems.
If certification with cargo compartment fire detection provisions is
requested, the following must be met for each cargo compartment with those
provisions:
(a) The detection system must provide a visual indication to the flight
crew within one minute after the start of a fire.
(b) The system must be capable of detecting a fire at a temperature
significantly below that at which the structural integrity of the airplane is
substantially decreased.
(c) There must be means to allow the crew to check in flight, the
functioning of each fire detector circuit.
(d) The effectiveness of the detection system must be shown for all
approved operating configurations and conditions.
[Amdt. 25-54, 45 FR 60173, Sept. 11, 1980]
Sec. 25.859 Combustion heater fire protection.
(a) Combustion heater fire zones. The following combustion heater fire
zones must be protected from fire in accordance with the applicable
provisions of Secs. 25.1181 through 25.1191 and Secs. 25.1195 through
25.1203;
(1) The region surrounding the heater, if this region contains any
flammable fluid system components (excluding the heater fuel system), that
could--
(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 that surrounds the combustion
chamber. However, no fire extinguishment is required in cabin ventilating air
passages.
(b) Ventilating air ducts. Each ventilating air duct passing through any
fire zone must be fireproof. In addition--
(1) Unless isolation is provided by fireproof valves or by equally
effective means, the ventilating air duct downstream of each heater must be
fireproof for a distance great enough to ensure that any fire originating in
the heater can be contained in the duct; and
(2) Each part of any ventilating duct passing through any region having a
flammable fluid system must be constructed or isolated from that system so
that the malfunctioning of any component of that system cannot introduce
flammable fluids or vapors into the ventilating airstream.
(c) Combustion air ducts. Each combustion air duct must be fireproof for a
distance great enough to prevent damage from backfiring or reverse flame
propagation. In addition--
(1) No combustion air duct may have a common opening with the ventilating
airstream unless flames from backfires or reverse burning cannot enter the
ventilating airstream under any operating condition, including reverse flow
or malfunctioning of the heater or its associated components; and
(2) No combustion air duct may restrict the prompt relief of any backfire
that, if so restricted, could cause heater failure.
(d) Heater controls; general. Provision must be made to prevent the
hazardous accumulation of water or ice on or in any heater control component,
control system tubing, or safety control.
(e) Heater safety controls. For each combustion heater there must be the
following safety control means:
(1) Means independent of the components provided for the normal continuous
control of air temperature, airflow, and fuel flow must be provided, for each
heater, to automatically shut off the ignition and fuel supply to that heater
at a point remote from that heater when any of the following occurs:
(i) The heat exchanger temperature exceeds safe limits.
(ii) The ventilating air temperature exceeds safe limits.
(iii) The combustion airflow becomes inadequate for safe operation.
(iv) The ventilating airflow becomes inadequate for safe operation.
(2) The means of complying with paragraph (e)(1) of this section for any
individual heater must--
(i) Be independent of components serving any other heater whose heat output
is essential for safe operation; and
(ii) Keep the heater off until restarted by the crew.
(3) There must be means to warn the crew when any heater whose heat output
is essential for safe operation has been shut off by the automatic means
prescribed in paragraph (e)(1) of this section.
(f) Air intakes. Each combustion and ventilating air intake must be located
so that no flammable fluids or vapors can enter the heater system under 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
Secs. 25.1121 and 25.1123. In addition, there must be provisions in the
design of the heater exhaust system to safely expel the products of
combustion to prevent the occurrence of--
(1) Fuel leakage from the exhaust to surrounding compartments;
(2) Exhaust gas impingement on surrounding equipment or structure;
(3) Ignition of flammable fluids by the exhaust, if the exhaust is in a
compartment containing flammable fluid lines; and
(4) Restriction by the exhaust of the prompt relief of backfires that, if
so restricted, could cause heater failure.
(h) Heater fuel systems. Each heater fuel system must meet each powerplant
fuel system requirement affecting safe heater operation. Each heater fuel
system component within the ventilating airstream must be protected by
shrouds so that no leakage from those components can enter the ventilating
airstream.
(i) Drains. There must be means to safely drain fuel that might accumulate
within the combustion chamber or the heat exchanger. In addition--
(1) Each part of any drain that operates at high temperatures must be
protected in the same manner as heater exhausts; and
(2) Each drain must be protected from hazardous ice accumulation under any
operating condition.
[Doc. No. 5066, 29 FR 18291, Dec. 24 1964, as amended by Amdt. 25-11, 32 FR
6912, May 5, 1967; Amdt. 25-23, 35 FR 5676, Apr. 8, 1970]
Sec. 25.863 Flammable fluid fire protection.
(a) In each area where flammable fluids or vapors might escape by leakage
of a fluid system, there must be means to minimize the probability of
ignition of the fluids and vapors, and the resultant hazards if ignition does
occur.
(b) Compliance with paragraph (a) of this section must be shown by analysis
or tests, and the following factors must be considered:
(1) Possible sources and paths of fluid leakage, and means of detecting
leakage.
(2) Flammability characteristics of fluids, including effects of any
combustible or absorbing materials.
(3) Possible ignition sources, including electrical faults, overheating of
equipment, and malfunctioning of protective devices.
(4) Means available for controlling or extinguishing a fire, such as
stopping flow of fluids, shutting down equipment, fireproof containment, or
use of extinguishing agents.
(5) Ability of airplane components that are critical to safety of flight to
withstand fire and heat.
(c) If action by the flight crew is required to prevent or counteract a
fluid fire (e.g., equipment shutdown or actuation of a fire extinguisher)
quick acting means must be provided to alert the crew.
(d) Each area where flammable fluids or vapors might escape by leakage of a
fluid system must be identified and defined.
[Amdt. 25-23, 35 FR 5676, Apr. 8, 1970, as amended by Amdt. 25-46, 43 FR
50597, Oct. 30, 1978]
Sec. 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 fire zones or in adjacent areas which would be
subjected to the effects of fire in the fire zone must be constructed of
fireproof material or shielded so that they are capable of withstanding the
effects of fire.
[Amdt. 25-23, 35 FR 5676, Apr. 8, 1970]
Sec. 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 apply to tail surfaces to the
rear of the nacelles that could not be readily affected by heat, flames, or
sparks coming from a designated fire zone or engine compartment of any
nacelle.
[Amdt. 25-23, 35 FR 5676, Apr. 8, 1970]
Sec. 25.869 Fire protection: systems.
(a) Electrical system components:
(1) Components of the electrical system must meet the applicable fire and
smoke protection requirements of Secs. 25.831(c) and 25.863.
(2) Electrical cables, terminals, and equipment in designated fire zones,
that are used during emergency procedures, must be at least fire resistant.
(3) Main power cables (including generator cables) in the fuselage must be
designed to allow a reasonable degree of deformation and stretching without
failure and must be--
(i) Isolated from flammable fluid lines; or
(ii) Shrouded by means of electrically insulated, flexible conduit, or
equivalent, which is in addition to the normal cable insulation.
(4) Insulation on electrical wire and electrical cable installed in any
area of the fuselage must be self-extinguishing when tested in accordance
with the applicable portions of part I, appendix F of this part.
(b) Each vacuum air system line and fitting on the discharge side of the
pump that might contain flammable vapors or fluids must meet the requirements
of Sec. 25.1183 if the line or fitting is in a designated fire zone. Other
vacuum air systems components in designated fire zones must be at least fire
resistant.
(c) Oxygen equipment and lines must--
(1) Not be located in any designated fire zone,
(2) Be protected from heat that may be generated in, or escape from, any
designated fire zone, and
(3) Be installed so that escaping oxygen cannot cause ignition of grease,
fluid, or vapor accumulations that are present in normal operation or as a
result of failure or malfunction of any system.
[Doc. No. 24344, Amdt. 25-72, 55 FR 29784, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Miscellaneous
Sec. 25.871 Leveling means.
There must be means for determining when the airplane is in a level
position on the ground.
[Amdt. 25-23, 35 FR 5676, Apr. 8, 1970]
Sec. 25.875 Reinforcement near propellers.
(a) Each part of the airplane near the propeller tips must be strong and
stiff enough to withstand the effects of the induced vibration and of ice
thrown from the propeller.
(b) No window may be near the propeller tips unless it can withstand the
most severe ice impact likely to occur.
General
Sec. 25.901 Installation.
(a) For the purpose of this part, the airplane powerplant installation
includes each component that--
(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 Sec. 33.5 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 ensure their continued safe operation between normal
inspections or overhauls;
(3) The installation must be accessible for necessary inspections and
maintenance; and
(4) The major components of the installation must be electrically bonded to
the other parts of the airplane.
(c) For each powerplant and auxiliary power unit installation, it must be
established that no single failure or malfunction or probable combination of
failures will jeopardize the safe operation of the airplane except that the
failure of structural elements need not be considered if the probability of
such failure is extremely remote.
(d) Each auxiliary power unit installation must meet the applicable
provisions of this subpart.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5676, Apr. 8, 1970; Amdt. 25-40, 42 FR 15042, Mar. 17, 1977; Amdt. 25-46, 43
FR 50597, Oct. 30, 1978]
Sec. 25.903 Engines.
(a) Engine type certificate.
(1) Each engine must have a type certificate and must meet the applicable
requirements of part 34 of this chapter.
(2) Each turbine engine must either--
(i) Comply with Sec. 33.77 of this chapter in effect on October 31, 1974,
or as subsequently amended; or
(ii) Be shown to have a foreign object ingestion service history in similar
installation locations which has not resulted in any unsafe condition.
(b) Engine isolation. The powerplants must be arranged and isolated from
each other to allow operation, in at least one configuration, so that the
failure or malfunction of any engine, or of any system that can affect the
engine, will not--
(1) Prevent the continued safe operation of the remaining engines; or
(2) Require immediate action by any crewmember for continued safe
operation.
(c) Control of engine rotation. There must be means for stopping the
rotation of any engine individually in flight, except that, for turbine
engine installations, the means for stopping the rotation of any engine need
be provided only where continued rotation could jeopardize the safety of the
airplane. Each component of the stopping and restarting system on the engine
side of the firewall that might be exposed to fire must be at least fire-
resistant. If hydraulic propeller feathering systems are used for this
purpose, the feathering lines must be at least fire resistant under the
operating conditions that may be expected to exist during feathering.
(d) Turbine engine installations. For turbine engine installations--
(1) Design precautions must be taken to minimize the hazards to the
airplane in the event of an engine rotor failure or of a fire originating
within the engine which burns through the engine case.
(2) The powerplant systems associated with engine control devices, systems,
and instrumentation, must be designed to give reasonable assurance that those
engine operating limitations that adversely affect turbine rotor structural
integrity 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, and each engine must have a restart capability within that
envelope.
(3) For turbine engine powered airplanes, if the minimum windmilling speed
of the engines, following the inflight shutdown of all engines, is
insufficient to provide the necessary electrical power for engine ignition, a
power source independent of the engine-driven electrical power generating
system must be provided to permit in-flight engine ignition for restarting.
(f) Auxiliary Power Unit. Each auxiliary power unit must be approved or
meet the requirements of the category for its intended use.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5676, Apr. 8, 1970; Amdt. 25-40, 42 FR 15042, Mar. 17, 1977; Amdt. 25-57, 49
FR 6848, Feb. 23, 1984; Amdt. 25-72, 55 FR 29784, July 20, 1990; Amdt. 25-73,
55 FR 32861, Aug. 10, 1990; 55 FR 35139, Aug. 28, 1990]
EFFECTIVE DATE NOTE: At Amdt. 25-72, 55 FR 29784, July 20, 1990, Sec.
25.903 was amended by adding paragraph (f) effective Aug. 20, 1990.
*****************************************************************************
55 FR 32856, No. 155, Aug. 10, 1990
SUMMARY: This final rule codifies as new part 34 all of the applicable
aircraft engine fuel venting and exhaust emission requirements of Special
Federal Aviation Regulation (SFAR) 27-5, and the test procedures specified
under the regulations implementing the Clean Air Act. This rule consolidates
all of the requirements and test procedures into this part, and inserts into
other affected parts the requirements to comply with new part 34. New part 34
does not alter any of the requirements specified under SFAR 27-5 or the
regulations implementing the Clean Air Act.
EFFECTIVE DATE: September 10, 1990.
*****************************************************************************
Sec. 25.904 Automatic takeoff thrust control system (ATTCS).
Each applicant seeking approval for installation of an engine power control
system that automatically resets the power or thrust on the operating
engine(s) when any engine fails during the takeoff must comply with the
requirements of Appendix I of this part.
[Amdt. 25-62, 52 FR 43156, Nov. 9, 1987]
Sec. 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 propeller is certificated.
(c) Each component of the propeller blade pitch control system must meet
the requirements of Sec. 35.42 of this chapter.
(d) Design precautions must be taken to minimize the hazards to the
airplane in the event a propeller blade fails or is released by a hub
failure. The hazards which must be considered include damage to structure and
vital systems due to impact of a failed or released blade and the unbalance
created by such failure or release.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended 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]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.907 Propeller vibration.
(a) The magnitude of the propeller blade vibration stresses under any
normal condition of operation must be determined by actual measurement or by
comparison with similar installations for which these measurements have been
made.
(b) The determined vibration stresses may not exceed values that have been
shown to be safe for continuous operation.
Sec. 25.925 Propeller clearance.
Unless smaller clearances are substantiated, propeller clearances with the
airplane at maximum weight, with the most adverse center of gravity, and with
the propeller in the most adverse pitch position, may not be less than the
following:
(a) Ground clearance. There must be a clearance of at least seven inches
(for each airplane with nose wheel landing gear) or nine inches (for each
airplane with tail wheel landing gear) between each propeller and the ground
with the landing gear statically deflected and in the level takeoff, or
taxiing attitude, whichever is most critical. In addition, there must be
positive clearance between the propeller and the ground when in the level
takeoff attitude with the critical tire(s) completely deflated and the
corresponding landing gear strut bottomed.
(b) Water clearance. There must be a clearance of at least 18 inches
between each propeller and the water, unless compliance with Sec. 25.239(a)
can be 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 any additional radial clearance necessary to prevent
harmful vibration;
(2) At least one-half inch longitudinal clearance between the propeller
blades or cuffs and stationary parts of the airplane; and
(3) Positive clearance between other rotating parts of the propeller or
spinner and stationary parts of the airplane.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-72, 55
FR 29784, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.929 Propeller deicing.
(a) For airplanes intended for use where icing may be expected, there must
be a means to prevent or remove hazardous ice accumulation on propellers or
on accessories where ice accumulation would jeopardize engine performance.
(b) If combustible fluid is used for propeller deicing, Secs. 25.1181
through 25.1185 and 25.1189 apply.
Sec. 25.933 Reversing systems.
(a) For turbojet reversing systems--
(1) Each system intended for ground operation only must be designed so that
during any reversal in flight the engine will produce no more than flight
idle thrust. In addition, it must be shown by analysis or test, or both,
that--
(i) Each operable reverser can be restored to the forward thrust position;
and
(ii) The airplane is capable of continued safe flight and landing under any
possible position of the thrust reverser.
(2) Each system intended for inflight use must be designed so that no
unsafe condition will result during normal operation of the system, or from
any failure (or reasonably likely combination of failures) of the reversing
system, under any anticipated condition of operation of the airplane
including ground operation. Failure of structural elements need not be
considered if the probability of this kind of failure is extremely remote.
(3) Each system must have means to prevent the engine from producing more
than idle thrust when the reversing system malfunctions, except that it may
produce any greater forward thrust that is shown to allow directional control
to be maintained, with aerodynamic means alone, under the most critical
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 (or reasonably likely combination of failures) or
malfunction of the system will result in unwanted reverse thrust under any
expected operating condition. Failure of structural elements need not be
considered if this kind of failure is extremely remote.
(2) Compliance with this section may be shown by failure analysis or
testing, or both, for propeller systems that allow propeller blades to move
from the flight low-pitch position to a position that is substantially less
than that at the normal flight low-pitch position. The analysis may include
or be supported by the analysis made to show compliance with the requirements
of Sec. 35.21 of this chapter for the propeller and associated installation
components.
[Doc. No. 24344, Amdt. 25-72, 55 FR 29784, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.934 Turbojet engine thrust reverser system tests.
Thrust reversers installed on turbojet engines must meet the requirements
of Sec. 33.97 of this chapter.
[Amdt. 25-23, 35 FR 5677, Apr. 8, 1970]
Sec. 25.937 Turbopropeller-drag limiting systems.
Turbopropeller power airplane propeller-drag limiting systems must be
designed so that no single failure or malfunction of any of the systems
during normal or emergency operation results in propeller drag in excess of
that for which the airplane was designed under Sec. 25.367. Failure of
structural elements of the drag limiting systems need not be considered if
the probability of this kind of failure is extremely remote.
Sec. 25.939 Turbine engine operating characteristics.
(a) Turbine engine operating characteristics must be investigated in flight
to determine that no adverse characteristics (such as stall, surge, or
flameout) are present, to a hazardous degree, during normal and emergency
operation within the range of operating limitations of the airplane and of
the engine.
(b) [Reserved]
(c) The turbine engine air inlet system may not, as a result of air flow
distortion during normal operation, cause vibration harmful to the engine.
[Amdt. 25-11, 32 FR 6912, May 5, 1967, as amended by Amdt. 25-40, 42 FR
15043, Mar. 17, 1977]
Sec. 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 incorporated), and exhaust must be shown to function
properly under all operating conditions for which approval is sought,
including all engine rotating speeds and power settings, and engine inlet and
exhaust configurations;
(b) The dynamic effects of the operation of these (including consideration
of probable malfunctions) upon the aerodynamic control of the airplane may
not result in any condition that would require exceptional skill, alertness,
or strength on the part of the pilot to avoid exceeding an operational or
structural limitation of the airplane; and
(c) In showing compliance with paragraph (b) of this section, the pilot
strength required may not exceed the limits set forth in Sec. 25.143(c),
subject to the conditions set forth in paragraphs (d) and (e) of Sec. 25.143.
[Amdt. 25-38, 41 FR 55467, Dec. 20, 1976]
Sec. 25.943 Negative acceleration.
No hazardous malfunction of an engine, an auxiliary power unit approved for
use in flight, or any component or system associated with the powerplant or
auxiliary power unit may occur when the airplane is operated at the negative
accelerations within the flight envelopes prescribed in Sec. 25.333. This
must be shown for the greatest duration expected for the acceleration.
[Amdt. 25-40, 42 FR 15043, Mar. 17, 1977]
Sec. 25.945 Thrust or power augmentation system.
(a) General. Each fluid injection system must provide a flow of fluid at
the rate and pressure established for proper engine functioning under each
intended operating condition. If the fluid can freeze, fluid freezing may not
damage the airplane or adversely affect airplane performance.
(b) Fluid tanks. Each augmentation system fluid tank must meet the
following requirements:
(1) Each tank must be able to withstand without failure the vibration,
inertia, fluid, and structural loads that is may be subject to in operation.
(2) The tanks as mounted in the airplane must be able to withstand without
failure or leakage an internal pressure 1.5 times the maximum operating
pressure.
(3) If a vent is provided, the venting must be effective under all normal
flight conditions.
(4) [Reserved]
(c) Augmentation system drains must be designed and located in accordance
with Sec. 25.1455 if--
(1) The augmentation system fluid is subject to freezing; and
(2) The fluid may be drained in flight or during ground operation.
(d) The augmentation liquid tank capacity available for the use of each
engine must be large enough to allow operation of the airplane under the
approved procedures for the use of liquid-augmented power. The computation of
liquid consumption must be based on the maximum approved rate appropriate for
the desired engine output and must include the effect of temperature on
engine performance as well as any other factors that might vary the amount of
liquid required.
(e) This section does not apply to fuel injection systems.
[Amdt. 25-40, 42 FR 15043, Mar. 17, 1977, as amended by Amdt. 25-72, 55
FR 29785, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Fuel System
Sec. 25.951 General.
(a) Each fuel system must be constructed and arranged to ensure a flow of
fuel at a rate and pressure established for proper engine and auxiliary power
unit functioning under each likely operating condition, including any
maneuver for which certification is requested and during which the engine or
auxiliary power unit is permitted to be in operation.
(b) Each fuel system must be arranged so that any air which is introduced
into the system will not result in--
(1) Power interruption for more than 20 seconds for reciprocating engines;
or
(2) Flameout for turbine engines.
(c) Each fuel system for a turbine engine must be capable of sustained
operation throughout its flow and pressure range with fuel initially
saturated with water at 80 deg. F and having 0.75cc of free water per gallon
added and cooled to the most critical condition for icing likely to be
encountered in operation.
(d) Each fuel system for a turbine engine powered airplane must meet the
applicable fuel venting requirements of part 34 of this chapter.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5677, Apr. 8, 1970; Amdt. 25-36, 39 FR 35460, Oct. 1, 1974; Amdt. 25-38, 41
FR 55467, Dec. 20, 1976; Amdt. 25-73, 55 FR 32861, Aug. 10, 1990; 55 FR
35139, Aug. 28, 1990]
*****************************************************************************
55 FR 32856, No. 155, Aug. 10, 1990
SUMMARY: This final rule codifies as new part 34 all of the applicable
aircraft engine fuel venting and exhaust emission requirements of Special
Federal Aviation Regulation (SFAR) 27-5, and the test procedures specified
under the regulations implementing the Clean Air Act. This rule consolidates
all of the requirements and test procedures into this part, and inserts into
other affected parts the requirements to comply with new part 34. New part 34
does not alter any of the requirements specified under SFAR 27-5 or the
regulations implementing the Clean Air Act.
EFFECTIVE DATE: September 10, 1990.
*****************************************************************************
Sec. 25.952 Fuel system analysis and test.
(a) Proper fuel system functioning under all probable operating conditions
must be shown by analysis and those tests found necessary by the
Administrator. Tests, if required, must be made using the airplane fuel
system or a test article that reproduces the operating characteristics of the
portion of the fuel system to be tested.
(b) The likely failure of any heat exchanger using fuel as one of its
fluids may not result in a hazardous condition.
[Amdt. 25-40, 42 FR 15043, Mar. 17, 1977]
Sec. 25.953 Fuel system independence.
Each fuel system must meet the requirements of Sec. 25.903(b) by--
(a) Allowing the supply of fuel to each engine through a system independent
of each part of the system supplying fuel to any other engine; or
(b) Any other acceptable method.
Sec. 25.954 Fuel system lightning protection.
The fuel system must be designed and arranged to prevent the ignition of
fuel vapor within the system by--
(a) Direct lightning strikes to areas having a high probability of stroke
attachment;
(b) Swept lightning strokes to areas where swept strokes are highly
probable; and
(c) Corona and streamering at fuel vent outlets.
[Amdt. 25-14, 32 FR 11629, Aug. 11, 1967]
Sec. 25.955 Fuel flow.
(a) Each fuel system must provide at least 100 percent of the fuel flow
required under each intended operating condition and maneuver. Compliance
must be shown as follows:
(1) Fuel must be delivered to each engine at a pressure within the limits
specified in the engine type certificate.
(2) The quantity of fuel in the tank may not exceed the amount established
as the unusable fuel supply for that tank under the requirements of Sec.
25.959 plus that necessary to show compliance with this section.
(3) Each main pump must be used that is necessary for each operating
condition and attitude for which compliance with this section is shown, and
the appropriate emergency pump must be substituted for each main pump so
used.
(4) If there is a fuel flowmeter, it must be blocked and the fuel must flow
through the meter or its bypass.
(b) If an engine can be supplied with fuel from more than one tank, the
fuel system must--
(1) For each reciprocating engine, supply the full fuel pressure to that
engine in not more than 20 seconds after switching to any other fuel tank
containing usable fuel when engine malfunctioning becomes apparent due to the
depletion of the fuel supply in any tank from which the engine can be fed;
and
(2) For each turbine engine, in addition to having appropriate manual
switching capability, be designed to prevent interruption of fuel flow to
that engine, without attention by the flight crew, when any tank supplying
fuel to that engine is depleted of usable fuel during normal operation, and
any other tank, that normally supplies fuel to that engine alone, contains
usable fuel.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-11, 32 FR
6912, May 5, 1967]
Sec. 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 structural
damage to the tanks can occur because of overfilling.
Sec. 25.959 Unusable fuel supply.
The unusable fuel quantity for each fuel tank and its fuel system
components must be established at not less than the quantity at which the
first evidence of engine malfunction occurs under the most adverse fuel feed
condition for all intended operations and flight maneuvers involving fuel
feeding from that tank. Fuel system component failures need not be
considered.
[Amdt. 25-23, 35 FR 5677, Apr. 8, 1970, as amended by Amdt. 25-40, 42 FR
15043, Mar. 17, 1977]
Sec. 25.961 Fuel system hot weather operation.
(a) The fuel system must perform satisfactorily in hot weather operation.
This must be shown by showing that the fuel system from the tank outlets to
each engine is pressurized, under all intended operations, so as to prevent
vapor formation, or must be shown by climbing from the altitude of the
airport elected by the applicant to the maximum altitude established as an
operating limitation under Sec. 25.1527. If a climb test is elected, there
may be no evidence of vapor lock or other malfunctioning during the climb
test conducted under the following conditions:
(1) For reciprocating engine powered airplanes, the engines must operate at
maximum continuous power, except that takeoff power must be used for the
altitudes from 1,000 feet below the critical altitude through the critical
altitude. The time interval during which takeoff power is used may not be
less than the takeoff time limitation.
(2) For turbine engine powered airplanes, the engines must operate at
takeoff power for the time interval selected for showing the takeoff flight
path, and at maximum continuous power for the rest of the climb.
(3) The weight of the airplane must be the weight with full fuel tanks,
minimum crew, and the ballast necessary to maintain the center of gravity
within allowable limits.
(4) The climb airspeed may not exceed--
(i) For reciprocating engine powered airplanes, the maximum airspeed
established for climbing from takeoff to the maximum operating altitude with
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 provides adequate cooling in the hot-day condition.
(D) Engine operating within the maximum continuous power limitations.
(E) Maximum takeoff weight; and
(ii) For turbine engine powered airplanes, the maximum airspeed established
for climbing from takeoff to the maximum operating altitude.
(5) The fuel temperature must be at least 110 deg. F.
(b) The test prescribed in paragraph (a) of this section may be performed
in flight or on the ground under closely simulated flight conditions. If a
flight test is performed in weather cold enough to interfere with the proper
conduct of the test, the fuel tank surfaces, fuel lines, and other fuel
system parts subject to cold air must be insulated to simulate, insofar as
practicable, flight in hot weather.
[Amdt. 25-11, 32 FR 6912, May 5, 1967, as amended by Amdt. 25-57, 49 FR 6848,
Feb. 23, 1984]
Sec. 25.963 Fuel tanks: general.
(a) Each fuel tank must be able to withstand, without failure, the
vibration, inertia, fluid, and structural loads that it may be subjected to
in operation.
(b) Flexible fuel tank liners must be approved or must be shown to be
suitable for the particular application.
(c) Integral fuel tanks must have facilities for interior inspection and
repair.
(d) Fuel tanks within the fuselage contour must be able to resist rupture
and to retain fuel, under the inertia forces prescribed for the emergency
landing conditions in Sec. 25.561. In addition, these tanks must be in a
protected position so that exposure of the tanks to scraping action with the
ground is unlikely.
(e) Fuel tank access covers must comply with the following criteria in
order to avoid loss of hazardous quantities of fuel:
(1) All covers located in an area where experience or analysis indicates a
strike is likely must be shown by analysis or tests to minimize penetration
and deformation by tire fragments, low energy engine debris, or other likely
debris.
(2) All covers must be fire resistant as defined in part 1 of this chapter.
(f) For pressurized fuel tanks, a means with fail-safe features must be
provided to prevent the buildup of an excessive pressure difference between
the inside and the outside of the tank.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-40, 42 FR
15043, Mar. 17, 1977; Amdt. 25-69, 54 FR 40352, Sept. 29, 1989]
Sec. 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 the
pressures resulting from the conditions specified in paragraphs (a)(1) and
(2) of this section. In addition, it must be shown by either analysis or
tests, that tank surfaces subjected to more critical pressures resulting from
the condition of paragraphs (a)(3) and (4) of this section, are able to
withstand the 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 airplane with a full tank.
(4) Fluid pressures developed during the most adverse combination of
airplane roll and fuel load.
(b) Each metallic tank with large unsupported or unstiffened flat surfaces,
whose failure or deformation could cause fuel leakage, must be able to
withstand the following test, or its equivalent, without leakage or excessive
deformation of the tank walls:
(1) Each complete tank assembly and its supports must be vibration tested
while mounted to simulate the actual installation.
(2) Except as specified in paragraph (b)(4) of this section, the tank
assembly must be vibrated for 25 hours at an amplitude of not less than 1/32
of an inch (unless another amplitude is substantiated) while 2/3 filled with
water or other suitable test fluid.
(3) The test frequency of vibration must be as follows:
(i) If no frequency of vibration resulting from any r.p.m. within the
normal operating range of engine speeds is critical, the test frequency of
vibration must be 2,000 cycles per minute.
(ii) If only one frequency of vibration resulting from any r.p.m. within
the normal operating range of engine speeds is critical, that frequency of
vibration must be the test frequency.
(iii) If more than one frequency of vibration resulting from any r.p.m.
within the normal operating range of engine speeds is critical, the most
critical of these frequencies must be the test frequency.
(4) Under paragraphs (b)(3) (ii) and (iii) of this section, the time of
test must be adjusted to accomplish the same number of vibration cycles that
would be accomplished in 25 hours at the frequency specified in paragraph
(b)(3)(i) of this section.
(5) During the test, the tank assembly must be rocked at the rate of 16 to
20 complete cycles per minute, through an angle of 15 deg. on both sides of
the horizontal (30 deg. total), about the most critical axis, for 25 hours.
If motion about more than one axis is likely to be critical, the tank must be
rocked about each critical axis for 12 1/2 hours.
(c) Except where satisfactory operating experience with a similar tank in a
similar installation is shown, nonmetallic tanks must withstand the test
specified in paragraph (b)(5) of this section, with fuel at a temperature of
110 deg. F. During this test, a representative specimen of the tank must be
installed in a supporting structure simulating the installation in the
airplane.
(d) For pressurized fuel tanks, it must be shown by analysis or tests that
the fuel tanks can withstand the maximum pressure likely to occur on the
ground or in flight.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-11, 32 FR
6913, May 5, 1967; Amdt. 25-40, 42 FR 15043, Mar. 17, 1977]
Sec. 25.967 Fuel tank installations.
(a) Each fuel tank must be supported so that tank loads (resulting from the
weight of the fuel in the tanks) are not concentrated on unsupported tank
surfaces. In addition--
(1) There must be pads, if necessary, to prevent chafing between the tank
and its supports;
(2) Padding must be nonabsorbent or treated to prevent the absorption of
fluids;
(3) If a flexible tank liner is used, it must be supported so that it is
not required to withstand fluid loads; and
(4) Each interior surface of the tank compartment must be smooth and free
of projections that could cause wear of the liner unless--
(i) Provisions are made for protection of the liner at these points; or
(ii) The construction of the liner itself provides that protection.
(b) Spaces adjacent to tank surfaces must be ventilated to avoid fume
accumulation due to minor leakage. If the tank is in a sealed compartment,
ventilation may be limited to drain holes large enough to prevent excessive
pressure resulting from altitude changes.
(c) The location of each tank must meet the requirements of Sec.
25.1185(a).
(d) No engine nacelle skin immediately behind a major air outlet from the
engine compartment may act as the wall of an integral tank.
(e) Each fuel tank must be isolated from personnel compartments by a
fumeproof and fuelproof enclosure.
Sec. 25.969 Fuel tank expansion space.
Each fuel tank must have an expansion space of not less than 2 percent of
the tank capacity. It must be impossible to fill the expansion space
inadvertently with the airplane in the normal ground attitude. For pressure
fueling systems, compliance with this section may be shown with the means
provided to comply with Sec. 25.979(b).
[Amdt. 25-11, 32 FR 6913, May 5, 1967]
Sec. 25.971 Fuel tank sump.
(a) Each fuel tank must have a sump with an effective capacity, in the
normal ground attitude, of not less than the greater of 0.10 percent of the
tank capacity or one-sixteenth of a gallon unless operating limitations are
established to ensure that the accumulation of water in service will not
exceed the sump capacity.
(b) Each fuel tank must allow drainage of any hazardous quantity of water
from any part of the tank to its sump with the airplane in the ground
attitude.
(c) Each fuel tank sump must have an accessible drain that--
(1) Allows complete drainage of the sump on the ground;
(2) Discharges clear of each part of the airplane; and
(3) Has manual or automatic means for positive locking in the closed
position.
Sec. 25.973 Fuel tank filler connection.
Each fuel tank filler connection must prevent the entrance of fuel into any
part of the airplane other than the tank itself. In addition--
(a) [Reserved]
(b) Each recessed filler connection that can retain any appreciable
quantity of fuel must have a drain that discharges clear of each part of the
airplane;
(c) Each filler cap must provide a fuel-tight seal; and
(d) Each fuel filling point, except pressure fueling connection points,
must have a provision for electrically bonding the airplane to ground fueling
equipment.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-40, 42 FR
15043, Mar. 17, 1977; Amdt. 25-72, 55 FR 29785, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 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 space so that venting is effective under any normal flight
condition. In addition--
(1) Each vent must be arranged to avoid stoppage by dirt or ice formation;
(2) The vent arrangement must prevent siphoning of fuel during normal
operation;
(3) The venting capacity and vent pressure levels must maintain acceptable
differences of pressure 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 outlets must be interconnected;
(5) There may be no point in any vent line where moisture can accumulate
with the airplane in the ground attitude or the level flight attitude, unless
drainage is provided; and
(6) No vent or drainage provision may end at any point--
(i) Where the discharge of fuel from the vent outlet would constitute a
fire hazard; or
(ii) From which fumes could enter personnel compartments.
(b) Carburetor vapor vents. Each carburetor with vapor elimination
connections must have a vent line to lead vapors back to one of the fuel
tanks. In addition--
(1) Each vent system must have means to avoid stoppage by ice; and
(2) If there is more than one fuel tank, and it is necessary to use the
tanks in a definite sequence, each vapor vent return line must lead back to
the fuel tank used for takeoff and landing.
Sec. 25.977 Fuel tank outlet.
(a) There must be a fuel strainer for the fuel tank outlet or for the
booster pump. This strainer must--
(1) For reciprocating engine powered airplanes, have 8 to 16 meshes per
inch; and
(2) For turbine engine powered airplanes, prevent the passage of any object
that could restrict fuel 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 the outlet line.
(d) The diameter of each strainer must be at least that of the fuel tank
outlet.
(e) Each finger strainer must be accessible for inspection and cleaning.
[Amdt. 25-11, 32 FR 6913, May 5, 1967, as amended by Amdt. 25-36, 39 FR
35460, Oct. 1, 1974]
Sec. 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 escape of hazardous quantities of fuel from the system if the
fuel entry valve fails.
(b) An automatic shutoff means must be provided to prevent the quantity of
fuel in each tank from exceeding the maximum quantity approved for that tank.
This means must--
(1) Allow checking for proper shutoff operation before each fueling of the
tank; and
(2) Provide indication at each fueling station of failure of the shutoff
means to stop the fuel flow at the maximum quantity approved for that tank.
(c) A means must be provided to prevent damage to the fuel system in the
event of failure of the automatic shutoff means prescribed in paragraph (b)
of this section.
(d) The airplane pressure fueling system (not including fuel tanks and fuel
tank vents) must withstand an ultimate load that is 2.0 times the load
arising from the maximum pressures, including surge, that is likely to occur
during fueling. The maximum surge pressure must be established with any
combination of tank valves being either intentionally or inadvertently
closed.
(e) The airplane defueling system (not including fuel tanks and fuel tank
vents) must withstand an ultimate load that is 2.0 times the load arising
from the maximum permissible defueling pressure (positive or negative) at the
airplane fueling connection.
[Amdt. 25-11, 32 FR 6913, May 5, 1967, as amended by Amdt. 25-38, 41 FR
55467, Dec. 20, 1976; Amdt. 25-72, 55 FR 29785, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.981 Fuel tank temperature.
(a) The highest temperature allowing a safe margin below the lowest
expected auto ignition temperature of the fuel in the fuel tanks must be
determined.
(b) No temperature at any place inside any fuel tank where fuel ignition is
possible may exceed the temperature determined under paragraph (a) of this
section. This must be shown under all probable operating, failure, and
malfunction conditions of any component whose operation, failure, or
malfunction could increase the temperature inside the tank.
[Amdt. 25-11, 32 FR 6913, May 5, 1967]
Fuel System Components
Sec. 25.991 Fuel pumps.
(a) Main pumps. Each fuel pump required for proper engine operation, or
required to meet the fuel system requirements of this subpart (other than
those in paragraph (b) of this section, is a main pump. For each main pump,
provision must be made to allow the bypass of each positive displacement fuel
pump other than a fuel injection pump (a pump that supplies the proper flow
and pressure for fuel injection when the injection is not accomplished in a
carburetor) approved as part of the engine.
(b) Emergency pumps. There must be emergency pumps or another main pump to
feed each engine immediately after failure of any main pump (other than a
fuel injection pump approved as part of the engine).
Sec. 25.993 Fuel system lines and fittings.
(a) Each fuel line must be installed and supported to prevent excessive
vibration and to withstand loads due to fuel pressure and accelerated flight
conditions.
(b) Each fuel line connected to components of the airplane between which
relative motion could exist must have provisions for flexibility.
(c) Each flexible connection in fuel lines that may be under pressure and
subjected to axial loading must use flexible hose assemblies.
(d) Flexible hose must be approved or must be shown to be suitable for the
particular application.
(e) No flexible hose that might be adversely affected by exposure to high
temperatures may be used where excessive temperatures will exist during
operation or after engine shut-down.
(f) Each fuel line within the fuselage must be designed and installed to
allow a reasonable degree of deformation and stretching without leakage.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-15, 32 FR
13266, Sept. 20, 1967]
Sec. 25.994 Fuel system components.
Fuel system components in an engine nacelle or in the fuselage must be
protected from damage which could result in spillage of enough fuel to
constitute a fire hazard as a result of a wheels-up landing on a paved
runway.
[Amdt. 25-57, 49 FR 6848, Feb. 23, 1984]
Sec. 25.995 Fuel valves.
In addition to the requirements of Sec. 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 conditions are transmitted to the lines attached to the
valve.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-40, 42 FR
15043, Mar. 17, 1977]
Sec. 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 fuel metering device or an engine driven positive
displacement pump, whichever is nearer the fuel tank outlet. This fuel
strainer or filter must--
(a) Be accessible for draining and cleaning and must incorporate a screen
or element which is easily removable;
(b) Have a sediment trap and drain except that it need not have a drain if
the strainer or filter is easily removable for drain purposes;
(c) Be mounted so that its weight is not supported by the connecting lines
or by the inlet or outlet connections of the strainer or filter itself,
unless adequate strength margins under all loading conditions are provided in
the lines and connections; and
(d) Have the capacity (with respect to operating limitations established
for the engine) to ensure that engine fuel system functioning is not
impaired, with the fuel contaminated to a degree (with respect to particle
size and density) that is greater than that established for the engine in
Part 33 of this chapter.
[Amdt. No. 25-36, 39 FR 35460, Oct. 1, 1974, as amended by Amdt. 25-57, 49 FR
6848, Feb. 23, 1984]
Sec. 25.999 Fuel system drains.
(a) Drainage of the fuel system must be accomplished by the use of fuel
strainer and fuel tank sump drains.
(b) Each drain required by paragraph (a) of this section must--
(1) Discharge clear of all parts of the airplane;
(2) Have manual or automatic means for positive 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 landing gear retracted.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 41 FR
55467, Dec. 20, 1976]
Sec. 25.1001 Fuel jettisoning system.
(a) A fuel jettisoning system must be installed on each airplane unless it
is shown that the airplane meets the climb requirements of Secs. 25.119 and
25.121(d) at maximum takeoff weight, less the actual or computed weight of
fuel necessary for a 15-minute flight comprised of a takeoff, go-around, and
landing at the airport of departure with the airplane configuration, speed,
power, and thrust the same as that used in meeting the applicable takeoff,
approach, and landing climb performance requirements of this part.
(b) If a fuel jettisoning system is required it must be capable of
jettisoning enough fuel within 15 minutes, starting with the weight given in
paragraph (a) of this section, to enable the airplane to meet the climb
requirements of Secs. 25.119 and 25.121(d), assuming that the fuel is
jettisoned under the conditions, except weight, found least favorable during
the flight tests prescribed in paragraph (c) of this section.
(c) Fuel jettisoning must be demonstrated beginning at maximum takeoff
weight with flaps and landing gear up and in--
(1) A power-off glide at 1.4 Vs1;
(2) A climb at the one-engine inoperative best rate-of-climb speed, with
the critical engine inoperative and the remaining engines at maximum
continuous power; and
(3) Level flight at 1.4 Vs1; if the results of the tests in the conditions
specified in paragraphs (c) (1) and (2) of this section show that this
condition could be critical.
(d) During the flight tests prescribed in paragraph (c) of this section, it
must be shown that--
(1) The fuel jettisoning system and its operation are free from fire
hazard;
(2) The fuel discharges clear of any part of the airplane;
(3) Fuel or fumes do not enter any parts of the airplane; and
(4) The jettisoning operation does not adversely affect the controllability
of the airplane.
(e) For reciprocating engine powered airplanes, means must be provided to
prevent jettisoning the fuel in the tanks used for takeoff and landing below
the level allowing 45 minutes flight at 75 percent maximum continuous power.
However, if there is an auxiliary control independent of the main jettisoning
control, the system may be designed to jettison the remaining fuel by means
of the auxiliary jettisoning control.
(f) For turbine engine powered airplanes, means must be provided to prevent
jettisoning the fuel in the tanks used for takeoff and landing below the
level allowing climb from sea level to 10,000 feet and thereafter allowing 45
minutes cruise at a speed for maximum range. However, if there is an
auxiliary control independent of the main jettisoning control, the system may
be designed to jettison the remaining fuel by means of the auxiliary
jettisoning control.
(g) The fuel jettisoning valve must be designed to allow flight personnel
to close the valve during any part of the jettisoning operation.
(h) Unless it is shown that using any means (including flaps, slots, and
slats) for changing the airflow across or around the wings does not adversely
affect fuel jettisoning, there must be a placard, adjacent to the jettisoning
control, to warn flight crewmembers against jettisoning fuel while the means
that change the airflow are being used.
(i) The fuel jettisoning system must be designed so that any reasonably
probable single malfunction in the system will not result in a hazardous
condition due to unsymmetrical jettisoning of, or inability to jettison,
fuel.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-18, 33 FR
12226, Aug. 30, 1968; Amdt. 25-57, 49 FR 6848, Feb. 23, 1984]
Oil System
Sec. 25.1011 General.
(a) Each engine must have an independent oil system that can supply it with
an appropriate quantity of oil at a temperature not above that safe for
continuous operation.
(b) The usable oil capacity may not be less than the product of the
endurance of the airplane under critical operating conditions and the
approved maximum allowable oil consumption of the engine under the same
conditions, plus a suitable margin to ensure system circulation. Instead of a
rational analysis of airplane range for the purpose of computing oil
requirements for reciprocating engine powered airplanes, the following fuel/
oil ratios may be used:
(1) For airplanes without a reserve oil or oil transfer system, a fuel/oil
ratio of 30:1 by volume.
(2) For airplanes with either a reserve oil or oil transfer system, a fuel/
oil ratio of 40:1 by volume.
(c) Fuel/oil ratios higher than those prescribed in paragraphs (b) (1) and
(2) of this section may be used if substantiated by data on actual engine oil
consumption.
Sec. 25.1013 Oil tanks.
(a) Installation. Each oil tank installation must meet the requirements of
Sec. 25.967.
(b) Expansion space. Oil tank expansion space must be provided as follows:
(1) Each oil tank used with a reciprocating engine must have an expansion
space of not less than the greater of 10 percent of the tank capacity or 0.5
gallon, and each oil tank used with a turbine engine must have an expansion
space of not less than 10 percent of the tank capacity.
(2) Each reserve oil tank not directly connected to any engine may have an
expansion space of not less than two percent of the tank capacity.
(3) It must be impossible to fill the expansion space inadvertently with
the airplane in the normal ground attitude.
(c) Filler connection. Each recessed oil tank filler connection that can
retain any appreciable quantity of oil must have a drain that discharges
clear of each part of the airplane. In addition, each oil 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 effective under any normal flight condition.
(2) Oil tank vents must be arranged so that condensed water vapor that
might freeze and obstruct the line cannot accumulate at any point.
(e) Outlet. There must be means to prevent entrance into the tank itself,
or into the tank outlet, of any object that might obstruct the flow of oil
through the system. No oil tank outlet may be enclosed by any screen or guard
that would reduce the flow of oil below a safe value at any operating
temperature. There must be a shutoff valve at the outlet of each oil tank
used with a turbine engine, unless the external portion of the oil system
(including the oil tank supports) is fireproof.
(f) Flexible oil tank liners. Each flexible oil tank liner must be approved
or must be shown to be suitable for the particular application.
[Doc. No. 5066, 29 FR 18291, Dec. 24, as amended by Amdt. 25-19, 33 FR 15410,
Oct. 17, 1968; Amdt. 25-23, 35 FR 5677, Apr. 8, 1970; Amdt. 25-36, 39 FR
35460, Oct. 1, 1974; Amdt. 25-57, 49 FR 6848, Feb. 23, 1984; Amdt. 25-72, 55
FR 29785, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.1015 Oil tank tests.
Each oil tank must be designed and installed so that--
(a) It can withstand, without failure, each vibration, inertia, and fluid
load that it may be subjected to in operation; and
(b) It meets the provisions of Sec. 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 maximum operating pressure of the tank instead of the
pressure specified in Sec. 25.965(a); and
(ii) For all other tanks may not be less than 5 p.s.i. instead of the
pressure specified in Sec. 25.965(a); and
(2) The test fluid must be oil at 250 deg. F. instead of the fluid
specified in Sec. 25.965(c).
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-36, 39 FR
35461, Oct. 1, 1974]
Sec. 25.1017 Oil lines and fittings.
(a) Each oil line must meet the requirements of Sec. 25.993 and each oil
line and fitting in any designated fire zone must meet the requirements of
Sec. 25.1183.
(b) Breather lines must be arranged so that--
(1) Condensed water vapor that might freeze and obstruct the line cannot
accumulate at any point;
(2) The breather discharge does not constitute a fire hazard if foaming
occurs or causes emitted oil to strike the pilot's windshield; and
(3) The breather does not discharge into the engine air induction system.
Sec. 25.1019 Oil strainer or filter.
(a) Each turbine engine installation must incorporate an oil strainer or
filter through which all of the engine oil flows and which meets the
following requirements:
(1) Each oil strainer or filter that has a bypass must be constructed and
installed so that oil will flow at the normal rate through the rest of the
system with the strainer or filter completely blocked.
(2) The oil strainer or filter must have the capacity (with respect to
operating limitations established for the engine) to ensure that engine oil
system functioning is not impaired when the oil is contaminated to a degree
(with respect to particle size and density) that is greater than that
established for the engine under Part 33 of this chapter.
(3) The oil strainer or filter, unless it is installed at an oil tank
outlet, must incorporate an indicator that will indicate contamination before
it reaches the capacity established in accordance with paragraph (a)(2) of
this section.
(4) The bypass of a strainer or filter must be constructed and installed so
that the release of collected contaminants is minimized by appropriate
location of the bypass to ensure that collected contaminants 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
warning system required in Sec. 25.1305(c)(7).
(b) Each oil strainer or filter in a powerplant installation using
reciprocating engines must be constructed and installed so that oil will flow
at the normal rate through the rest of the system with the strainer or filter
element completely blocked.
[Amdt. 25-36, 39 FR 35461, Oct. 1, 1974, as amended by Amdt. 25-57, 49 FR
6848, Feb. 23, 1984]
Sec. 25.1021 Oil system drains.
A drain (or drains) must be provided to allow safe drainage of the oil
system. Each drain must--
(a) Be accessible; and
(b) Have manual or automatic means for positive locking in the closed
position.
[Amdt. 25-57, 49 FR 6848, Feb. 23, 1984]
Sec. 25.1023 Oil radiators.
(a) Each oil radiator must be able to withstand, without failure, any
vibration, inertia, and oil pressure load to which it would be subjected in
operation.
(b) Each oil radiator air duct must be located so that, in case of fire,
flames coming from normal openings of the engine nacelle cannot impinge
directly upon the radiator.
Sec. 25.1025 Oil valves.
(a) Each oil shutoff must meet the requirements of Sec. 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" positions and must be supported so that no loads resulting
from its operation or from accelerated flight conditions are transmitted to
the lines attached to the valve.
Sec. 25.1027 Propeller feathering system.
(a) If the propeller feathering system depends on engine oil, there must be
means to trap an amount of oil in the tank if the supply becomes depleted due
to failure of any part of the lubricating system other than the tank itself.
(b) The amount of trapped oil must be enough to accomplish the feathering
operation and must be available only to the feathering pump.
(c) The ability of the system to accomplish feathering with the trapped oil
must be shown. This may be done on the ground using an auxiliary source of
oil for lubricating the engine during operation.
(d) Provision must be made to prevent sludge or other foreign matter from
affecting the safe operation of the propeller feathering system.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 41 FR
55467, Dec. 20, 1976]
Cooling
Sec. 25.1041 General.
The powerplant and auxiliary power unit cooling provisions must be able to
maintain the temperatures of powerplant components, engine fluids, and
auxiliary power unit components and fluids within the temperature limits
established for these components and fluids, under ground, water, and flight
operating conditions, and after normal engine or auxiliary power unit
shutdown, or both.
[Amdt. 25-38, 41 FR 55467, Dec. 20, 1976]
Sec. 25.1043 Cooling tests.
(a) General. Compliance with Sec. 25.1041 must be shown by tests, under
critical ground, water, and flight operating conditions. For these tests, the
following apply:
(1) If the tests are conducted under conditions deviating from the maximum
ambient atmospheric temperature, the recorded powerplant temperatures must be
corrected under paragraphs (c) and (d) of this section.
(2) No corrected temperatures determined under paragraph (a)(1) of this
section may exceed established limits.
(3) For reciprocating engines, the fuel used during the cooling tests must
be the minimum grade approved for the engines, and the mixture settings must
be those normally used in the flight stages for which the cooling tests are
conducted. The test procedures must be as prescribed in Sec. 25.1045.
(b) Maximum ambient atmospheric temperature. A maximum ambient atmospheric
temperature corresponding to sea level conditions of at least 100 degrees F
must be established. The assumed temperature lapse rate is 3.6 degrees F per
thousand feet of altitude above sea level until a temperature of -69.7
degrees F is reached, above which altitude the temperature is considered
constant at -69.7 degrees F. However, for winterization installations, the
applicant may select a maximum ambient atmospheric temperature corresponding
to sea level conditions of less than 100 degrees F.
(c) Correction factor (except cylinder barrels). Unless a more rational
correction applies, temperatures of engine fluids and powerplant components
(except cylinder barrels) for which temperature limits are established, must
be corrected by adding to them the difference between the maximum ambient
atmospheric temperature and the temperature of the ambient air at the time of
the first occurrence of the maximum component or fluid temperature recorded
during the cooling test.
(d) Correction factor for cylinder barrel temperatures. Unless a more
rational correction applies, cylinder barrel temperatures must be corrected
by adding to them 0.7 times the difference between the maximum ambient
atmospheric temperature and the temperature of the ambient air at the time of
the first occurrence of the maximum cylinder barrel temperature recorded
during the cooling test.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-42, 43 FR
2323, Jan. 16, 1978]
Sec. 25.1045 Cooling test procedures.
(a) Compliance with Sec. 25.1041 must be shown for the takeoff, climb, en
route, and landing stages of flight that correspond to the applicable
performance requirements. The cooling tests must be conducted with the
airplane in the configuration, and operating under the conditions, that are
critical 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 F. per minute.
(b) Temperatures must be stabilized under the conditions from which entry
is made into each stage of flight being investigated, unless the entry
condition normally is not one during which component and the engine fluid
temperatures would stabilize (in which case, operation through the full entry
condition must be conducted before entry into the stage of flight being
investigated in order to allow temperatures to reach their natural levels at
the time of entry). The takeoff cooling test must be preceded by a period
during which the powerplant component and engine fluid temperatures are
stabilized with the engines at ground idle.
(c) Cooling tests for each stage of flight must be continued until--
(1) The component and engine fluid temperatures stabilize;
(2) The stage of flight is completed; or
(3) An operating limitation is reached.
(d) For reciprocating engine powered airplanes, it may be assumed, for
cooling test purposes, that the takeoff stage of flight is complete when the
airplane reaches an altitude of 1,500 feet above the takeoff surface or
reaches a point in the takeoff where the transition from the takeoff to the
en route configuration is completed and a speed is reached at which
compliance with Sec. 25.121(c) is shown, whichever point is at a higher
altitude. The airplane must be in the following 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 provides adequate cooling in the hot-day condition.
(4) Critical engine inoperative and its propeller stopped.
(5) Remaining engines at the maximum continuous power available for the
altitude.
(e) For hull seaplanes and amphibians, cooling must be shown during taxiing
downwind for 10 minutes, at five knots above step speed.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-57, 49 FR
6848, Feb. 23, 1984]
Induction System
Sec. 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 for which certification is requested; and
(2) The air for proper fuel metering and mixture distribution with the
induction system valves in any position.
(b) Each reciprocating engine must have an alternate 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 the engine accessory section
by means of a fireproof diaphragm; or
(2) For reciprocating engines, there are means to prevent the emergence of
backfire flames.
(d) For turbine engine powered airplanes and airplanes incorporating
auxiliary power units--
(1) There must be means to prevent hazardous quantities of fuel leakage or
overflow from drains, vents, or other components of flammable fluid systems
from entering the engine or auxiliary power unit intake system; and
(2) The airplane must be designed to prevent water or slush on the runway,
taxiway, or other airport operating surfaces from being directed into the
engine or auxiliary power unit air inlet ducts in hazardous quantities, and
the air inlet ducts must be located or protected so as to minimize the
ingestion of foreign matter during takeoff, landing, and taxiing.
(e) If the engine induction system contains parts or components that could
be damaged by foreign objects entering the air inlet, it must be shown by
tests or, if appropriate, by analysis that the induction system design can
withstand the foreign object ingestion test conditions of Sec. 33.77 of this
chapter without failure of parts or components that could create a hazard.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 41 FR
55467, Dec. 20, 1976; Amdt. 25-40, 42 FR 15043, Mar. 17, 1977; Amdt. 25-57,
49 FR 6849, Feb. 23, 1984]
Sec. 25.1093 Induction system icing protection.
(a) Reciprocating engines. Each reciprocating engine air induction system
must have means to prevent and eliminate icing. Unless this is done by other
means, it must be shown that, in air free of visible moisture at a
temperature of 30 deg. F., each airplane with altitude engines using--
(1) Conventional venturi carburetors have a preheater that can provide a
heat rise of 120 deg. F. with the engine at 60 percent of maximum continuous
power; or
(2) Carburetors tending to reduce the probability of ice formation has a
preheater that can provide a heat rise of 100 deg. F. with the engine at 60
percent of maximum continuous power.
(b) Turbine engines. (1) Each turbine engine must operate throughout the
flight power range of the engine (including idling), without the accumulation
of ice on the engine, inlet system components, or airframe components that
would adversely affect engine operation or cause a serious loss of power or
thrust--
(i) Under the icing conditions specified in appendix C, and
(ii) In falling and blowing snow within the limitations established for the
airplane for such operation.
(2) Each turbine engine must idle for 30 minutes on the ground, with the
air bleed available for engine icing protection at its critical condition,
without adverse effect, in an atmosphere that is at a temperature between 15
deg. and 30 deg. F (between -9 deg. and -1 deg. C) and has a liquid water
content not less than 0.3 grams per cubic meter in the form of drops having a
mean effective diameter not less than 20 microns, followed by momentary
operation at takeoff power or thrust. During the 30 minutes of idle
operation, the engine may be run up periodically to a moderate power or
thrust setting in a manner acceptable to the Administrator.
(c) Supercharged reciprocating engines. For each engine having a
supercharger to pressurize the air before it enters the carburetor, the heat
rise in the air caused by that supercharging at any altitude may be utilized
in determining compliance with paragraph (a) of this section if the heat rise
utilized is that which will be available, automatically, for the applicable
altitude and operating condition because of supercharging.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 41 FR
55467, Dec. 20, 1976; Amdt. 25-40, 42 FR 15043, Mar. 17, 1977; Amdt. 25-57,
49 FR 6849, Feb. 23, 1984; Amdt. 25-72, 55 FR 29785, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.1101 Carburetor air preheater design.
Each carburetor air preheater must be designed 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 preheater itself.
Sec. 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 the auxiliary power unit compressor must have a drain to
prevent the hazardous accumulation of fuel and moisture in the ground
attitude. No drain may discharge where it might cause a fire hazard.
(b) Each induction system duct must be--
(1) Strong enough to prevent induction system failures resulting from
normal backfire conditions; and
(2) Fire-resistant if it is in any fire zone for which a fire-extinguishing
system is required, except that ducts for auxiliary power units must be
fireproof within the auxiliary power unit fire zone.
(c) Each duct connected to components between which relative motion could
exist must have means for flexibility.
(d) For turbine engine and auxiliary power unit bleed air duct systems, no
hazard may result if a duct failure occurs at any point between the air duct
source and the airplane unit served by the air.
(e) Each auxiliary power unit induction system duct must be fireproof for a
sufficient distance upstream of the auxiliary power unit compartment to
prevent hot gas reverse flow from burning through auxiliary power unit ducts
and entering any other compartment or area of the airplane in which a hazard
would be created resulting from the entry of hot gases. The materials used to
form the remainder of the induction system duct and plenum chamber of the
auxiliary power unit must be capable of resisting the maximum heat conditions
likely to occur.
(f) Each auxiliary power unit induction system duct must be constructed of
materials that will not absorb or trap hazardous quantities of flammable
fluids that could be ignited in the event of a surge or reverse flow
condition.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-46, 43 FR
50597, Oct. 30, 1978]
Sec. 25.1105 Induction system screens.
If induction system screens are used--
(a) Each screen must be upstream of the carburetor;
(b) No screen may be in any part of the induction system that is the only
passage through which air can reach the engine, unless it can be deiced by
heated air;
(c) No screen may be deiced by alcohol alone; and
(d) It must be impossible for fuel to strike any screen.
Sec. 25.1107 Inter-coolers and after-coolers.
Each inter-cooler and after-cooler must be able to withstand any vibration,
inertia, and air pressure load to which it would be subjected in operation.
Exhaust System
Sec. 25.1121 General.
For powerplant and auxiliary power unit installations the following apply:
(a) Each exhaust system must ensure safe disposal of exhaust gases without
fire hazard or carbon monoxide contamination in any personnel compartment.
For test purposes, any acceptable carbon monoxide detection method may be
used to show the absence of carbon monoxide.
(b) Each exhaust system part with a surface hot enough to ignite flammable
fluids or vapors must be located or shielded so that leakage from any system
carrying flammable fluids or vapors will not result in a fire caused by
impingement of the fluids or vapors on any part of the exhaust system
including shields for the exhaust system.
(c) Each component that hot exhaust gases could strike, or that could be
subjected to high temperatures from exhaust system parts, must be fireproof.
All exhaust system components must be separated by fireproof shields from
adjacent parts of the airplane that are outside the engine and auxiliary
power unit compartments.
(d) No exhaust gases may discharge so as to cause a fire hazard with
respect to any flammable fluid vent or drain.
(e) No exhaust gases may discharge where they will cause a glare seriously
affecting pilot vision at night.
(f) Each exhaust system component must be ventilated to prevent points of
excessively high temperature.
(g) Each exhaust shroud must be ventilated or insulated to avoid, during
normal operation, a temperature high enough to ignite any flammable fluids or
vapors external to the shroud.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-40, 42 FR
15043, Mar. 17, 1977]
Sec. 25.1123 Exhaust piping.
For powerplant and auxiliary power unit installations, the following apply:
(a) Exhaust piping must be heat and corrosion resistant, and must have
provisions to prevent failure due to expansion by operating temperatures.
(b) Piping must be supported to withstand any vibration and inertia loads
to which it would be subjected in operation; and
(c) Piping connected to components between which relative motion could
exist must have means for flexibility.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-40, 42 FR
15044, Mar. 17, 1977]
Sec. 25.1125 Exhaust heat exchangers.
For reciprocating engine powered airplanes, the following apply:
(a) Each exhaust heat exchanger must be constructed and installed to
withstand each vibration, inertia, and other load to which it would be
subjected in operation. In addition--
(1) Each exchanger must be suitable for continued operation at high
temperatures and resistant to corrosion from exhaust gases;
(2) There must be means for the inspection of the critical parts of each
exchanger;
(3) Each exchanger must have cooling provisions wherever it is subject to
contact with exhaust gases; and
(4) No exhaust heat exchanger or muff may have any stagnant areas or liquid
traps that would increase the probability of ignition of flammable fluids or
vapors that might be present in case of the failure or malfunction of
components carrying flammable fluids.
(b) If an exhaust heat exchanger is used for heating ventilating air--
(1) There must be a secondary heat exchanger between the primary exhaust
gas heat exchanger and the ventilating air system; or
(2) Other means must be used to preclude the harmful contamination of the
ventilating air.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 41 FR
55467, Dec. 20, 1976]
Sec. 25.1127 Exhaust driven turbo-superchargers.
(a) Each exhaust driven turbo-supercharger must be approved or shown to be
suitable for the particular application. It must be installed and supported
to ensure safe operation between normal inspections and overhauls. In
addition, there must be provisions for expansion and flexibility between
exhaust conduits and the turbine.
(b) There must be provisions for lubricating the turbine and for cooling
turbine parts where temperatures are critical.
(c) If the normal turbo-supercharger control system malfunctions, the
turbine speed may not exceed its maximum allowable value. Except for the
waste gate operating components, the components provided for meeting this
requirement must be independent of the normal turbo-supercharger controls.
Powerplant Controls and Accessories
Sec. 25.1141 Powerplant controls: general.
Each powerplant control must be located, arranged, and designed under Secs.
25.777 through 25.781 and marked under Sec. 25.1555. In addition, it must
meet the following requirements:
(a) Each control must be located so that it cannot be inadvertently
operated by persons entering, leaving, or moving normally in, the cockpit.
(b) Each flexible control must be approved or must be shown to be suitable
for the particular application.
(c) Each control must have sufficient strength and rigidity to withstand
operating loads without failure and without excessive deflection.
(d) Each control must be able to maintain any set position without constant
attention by flight crewmembers and without creep due to control loads or
vibration.
(e) The portion of each powerplant control located in a designated fire
zone that is required to be operated in the event of fire must be at least
fire resistant.
(f) Powerplant valve controls located in the cockpit must have--
(1) For manual valves, positive stops or in the case of fuel valves
suitable index provisions, in the open and closed position; and
(2) For power-assisted valves, a means to indicate 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.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-40, 42 FR
15044, Mar. 17, 1977; Amdt. 25-72, 55 FR 29785, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.1142 Auxiliary power unit controls.
Means must be provided on the flight deck for starting, stopping, and
emergency shutdown of each installed auxiliary power unit.
[Amdt. 25-46, 43 FR 50598, Oct. 30, 1978]
Sec. 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 of controlling its engine.
(d) For each fluid injection (other than fuel) system and its controls not
provided and approved as part of the engine, the applicant must show that the
flow of the injection fluid is adequately controlled.
(e) If a power or thrust control incorporates a fuel shutoff feature, the
control must have a means to prevent the inadvertent movement of the control
into the shutoff position. The means must--
(1) Have a positive lock or stop at the idle position; and
(2) Require a separate and distinct operation to place the control in the
shutoff position.
[Amdt. 25-23, 35 FR 5677, Apr. 8, 1970, as amended by Amdt. 25-38, 41 FR
55467, Dec. 20, 1976; Amdt. 25-57, 49 FR 6849, Feb. 23, 1984]
Sec. 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 master ignition control.
(c) Each group of ignition switches, except ignition switches for turbine
engines for which continuous ignition is not required, and each master
ignition control must have a means to prevent its inadvertent operation.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-40, 42 FR
15044 Mar. 17, 1977]
Sec. 25.1147 Mixture controls.
(a) If there are mixture controls, each engine must have a separate
control. The controls must be 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 operating setting must be identifiable by feel and sight.
(c) The mixture controls must be accessible to both pilots. However, if
there is a separate flight engineer station with a control panel, the
controls need be accessible only to the flight engineer.
Sec. 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, the pilot's throttle controls.
Sec. 25.1153 Propeller feathering controls.
(a) There must be a separate propeller feathering control for each
propeller. The control must have means to prevent its inadvertent operation.
(b) If feathering is accomplished by movement of the propeller pitch or
speed control lever, there must be means to prevent the inadvertent movement
of this lever to the feathering position during normal operation.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-11, 32 FR
6913, May 5, 1967]
Sec. 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 have means to prevent its inadvertent operation. The means
must have a positive lock or stop at the flight idle position and must
require a separate and distinct operation by the crew to displace the control
from the flight regime (forward thrust regime for turbojet powered
airplanes).
[Amdt. 25-11, 32 FR 6913, May 5, 1967]
Sec. 25.1157 Carburetor air temperature controls.
There must be a separate carburetor air temperature control for each
engine.
Sec. 25.1159 Supercharger controls.
Each supercharger control must be accessible to the pilots or, if there is
a separate flight engineer station with a control panel, to the flight
engineer.
Sec. 25.1161 Fuel jettisoning system controls.
Each fuel jettisoning system control must have guards to prevent
inadvertent operation. No control may be near any fire extinguisher control
or other control used to combat fire.
Sec. 25.1163 Powerplant accessories.
(a) Each engine mounted accessory must--
(1) Be approved for mounting on the engine involved;
(2) Use the provisions on the engine for mounting; 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 probability of contact with any flammable fluids or vapors that
might be present in a free state.
(c) If continued rotation of an engine-driven cabin supercharger or of any
remote accessory driven by the engine is hazardous if malfunctioning occurs,
there must be means to prevent rotation without interfering with the
continued operation of the engine.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-57, 49 FR
6849, Feb. 23, 1984]
Sec. 25.1165 Engine ignition systems.
(a) Each battery ignition system must be supplemented by a generator that
is automatically available as an alternate source of electrical energy to
allow continued engine operation if any battery becomes depleted.
(b) The capacity of batteries and generators must be large enough to meet
the simultaneous demands of the engine ignition system and the greatest
demands of any electrical system components that draw electrical energy from
the same source.
(c) The design of the engine ignition system must account for--
(1) The condition of an inoperative generator;
(2) The condition of a completely depleted battery with the generator
running at its normal operating speed; and
(3) The condition of a completely depleted battery with the generator
operating at idling speed, if there is only one battery.
(d) Magneto ground wiring (for separate ignition circuits) that lies on the
engine side of the fire wall, must be installed, located, or protected, to
minimize the probability of simultaneous failure of two or more wires as a
result of mechanical damage, electrical faults, or other cause.
(e) No ground wire for any engine may be routed through a fire zone of
another engine unless each part of that wire within that zone is fireproof.
(f) Each ignition system must be independent of any electrical circuit, not
used for assisting, controlling, or analyzing the operation of that system.
(g) There must be means to warn appropriate flight crewmembers if the
malfunctioning of any part of the electrical system is causing the continuous
discharge of any battery necessary for engine ignition.
(h) Each engine ignition system of a turbine powered airplane must be
considered an essential electrical load.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5677, Apr. 8, 1970; Amdt. 25-72, 55 FR 29785, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.1167 Accessory gearboxes.
For airplanes equipped with an accessory gearbox that is not certificated
as part of an engine--
(a) The engine with gearbox and connecting transmissions and shafts
attached must be subjected to the tests specified in Sec. 33.49 or Sec. 33.87
of this chapter, as applicable;
(b) The accessory gearbox must meet the requirements of Secs. 33.25 and
33.53 or 33.91 of this chapter, as applicable; and
(c) Possible misalignments and torsional loadings of the gearbox,
transmission, and shaft system, expected to result under normal operating
conditions must be evaluated.
[Amdt. 25-38, 41 FR 55467, Dec. 20, 1976]
Powerplant Fire Protection
Sec. 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 complete powerplant compartment
in which no isolation is provided between the engine power section and the
engine accessory section;
(4) Any auxiliary power unit compartment;
(5) Any fuel-burning heater and other combustion equipment installation
described in Sec. 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 or components carrying flammable fluids or
gases.
(b) Each designated fire zone must meet the requirements of Secs. 25.867,
and 25.1185 through 25.1203.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-11, 32 FR
6913, May 5, 1967; Amdt. 25-23, 35 FR 5677, Apr. 8, 1970; Amdt. 25-72, 55 FR
29785, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 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 attaching structure containing flammable fluid lines, must
meet each requirement of Secs. 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
need not contain fire detection or extinguishing means.
(b) For each area covered by paragraph (a) of this section that contains a
retractable landing gear, compliance with that paragraph need only be shown
with the landing gear retracted.
[Amdt. 25-11, 32 FR 6913, May 5, 1967]
Sec. 25.1183 Flammable fluid-carrying components.
(a) Except as provided in paragraph (b) of this section, each line,
fitting, and other component carrying flammable fluid in any area subject to
engine fire conditions, and each component which conveys or contains
flammable fluid in a designated fire zone must be fire resistant, except that
flammable fluid tanks and supports in a designated fire zone must be
fireproof or be enclosed by a fireproof shield unless damage by fire to any
non-fireproof part will not cause leakage or spillage of flammable fluid.
Components must be shielded or located to safeguard against the ignition of
leaking flammable fluid. An integral oil sump of less than 25-quart capacity
on a reciprocating engine need not be fireproof nor be enclosed by a
fireproof shield.
(b) Paragraph (a) of this section does not apply to--
(1) Lines, fittings, and components which are already approved as part of a
type certificated engine; and
(2) Vent and drain lines, and their fittings, whose failure will not result
in, or add to, a fire hazard.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-11, 32 FR
6913, May 5, 1967; Amdt. 25-36, 39 FR 35461, Oct. 1, 1974; Amdt. 25-57, 49 FR
6849, Feb. 23, 1984]
Sec. 25.1185 Flammable fluids.
(a) Except for the integral oil sumps specified in Sec. 25.1013 (a), no
tank or reservoir that is a part of a system containing flammable fluids or
gases may be in a designated fire zone unless the fluid contained, the design
of the system, the materials used in the tank, the shut-off means, and all
connections, lines, and control provide a degree of safety equal to that
which would exist if the tank or reservoir were outside such a zone.
(b) There must be at least one-half inch of clear airspace between each
tank or reservoir and each firewall or shroud isolating a designated fire
zone.
(c) Absorbent materials close to flammable fluid system components that
might leak must be covered or treated to prevent the absorption of hazardous
quantities of fluids.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964 as amended by Amdt. 25-19, 33 FR
15410, Oct. 17, 1968]
Sec. 25.1187 Drainage and ventilation of fire zones.
(a) There must be complete drainage of each part of each designated fire
zone to minimize the hazards resulting from failure or malfunctioning of any
component containing flammable fluids. The drainage means must be--
(1) Effective under conditions expected to prevail when drainage is needed;
and
(2) Arranged so that no discharged fluid will cause an additional fire
hazard.
(b) Each designated fire zone must be ventilated to prevent the
accumulation of flammable vapors.
(c) No ventilation opening may be where it would allow the entry of
flammable fluids, vapors, or flame from other zones.
(d) Each ventilation means must be arranged so that no discharged vapors
will cause an additional fire hazard.
(e) Unless the extinguishing agent capacity and rate of discharge are based
on maximum air flow through a zone, there must be means to allow the crew to
shut off sources of forced ventilation to any fire zone except the engine
power section of the nacelle and the combustion heater ventilating air ducts.
Sec. 25.1189 Shutoff means.
(a) Each engine installation and each fire zone specified in Sec.
25.1181(a) (4) and (5) must have a means to shut off or otherwise prevent
hazardous quantities of fuel, oil, deicer, and other flammable fluids, from
flowing into, within, or through any designated fire zone, except that
shutoff means are not required for--
(1) Lines, fittings, and components forming an integral part of an engine;
and
(2) Oil systems for turbine engine installations in which all components of
the system in a designated fire zone, including oil tanks, are fireproof or
located in areas not subject to engine fire conditions.
(b) The closing of any fuel shutoff valve for any engine may not make fuel
unavailable to the remaining engines.
(c) Operation of any shutoff may not interfere with the later emergency
operation of other equipment, such as the means for feathering the propeller.
(d) Each flammable fluid shutoff means and control must be fireproof or
must be located and protected so that any fire in a fire zone will not affect
its operation.
(e) No hazardous quantity of flammable fluid may drain into any designated
fire zone after shutoff.
(f) There must be means to guard against inadvertent operation of the
shutoff means and to make it possible for the crew to reopen the shutoff
means in flight after it has been closed.
(g) Each tank-to-engine shutoff valve must be located so that the operation
of the valve will not be affected by powerplant or engine mount structural
failure.
(h) Each shutoff valve must have a means to relieve excessive pressure
accumulation unless a means for pressure relief is otherwise provided in the
system.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5677, Apr. 8, 1970; Amdt. 25-57, 49 FR 6849, Feb. 23, 1984]
Sec. 25.1191 Firewalls.
(a) Each engine, auxiliary power unit, fuel-burning heater, other
combustion equipment intended for operation in flight, and the combustion,
turbine, and tailpipe sections of turbine engines, must be isolated from the
rest of the airplane by firewalls, 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 compartment to other parts of the airplane;
(3) Constructed so that each opening is sealed with close fitting fireproof
grommets, bushings, or firewall fittings; and
(4) Protected against corrosion.
Sec. 25.1192 Engine accessory section diaphragm.
For reciprocating engines, the engine power section and all portions of the
exhaust system must be isolated from the engine accessory compartment by a
diaphragm that complies with the firewall requirements of Sec. 25.1191.
[Amdt. 25-23, 35 FR 5678, Apr. 8, 1970]
Sec. 25.1193 Cowling and nacelle skin.
(a) Each cowling must be constructed and supported so that it can resist
any vibration, inertia, and air load to which it may be subjected in
operation.
(b) Cowling must meet the drainage and ventilation requirements of Sec.
25.1187.
(c) On airplanes with a diaphragm isolating the engine power section from
the engine accessory section, each part of the accessory section cowling
subject to flame in case of fire in the engine power section of the
powerplant must--
(1) Be fireproof; and
(2) Meet the requirements of Sec. 25.1191.
(d) Each part of the cowling subject to high temperatures due to its
nearness to exhaust system parts or exhaust gas impingement must be
fireproof.
(e) Each airplane must--
(1) Be designed and constructed so that no fire originating in any fire
zone can enter, either through openings or by burning through external skin,
any other zone or region where it would create additional hazards;
(2) Meet paragraph (e)(1) of this section with the landing gear retracted
(if applicable); and
(3) Have fireproof skin in areas subject to flame if a fire starts in the
engine power or accessory sections.
Sec. 25.1195 Fire extinguishing systems.
(a) Except for combustor, turbine, and tail pipe sections of turbine engine
installations that contain lines or components carrying flammable fluids or
gases for which it is shown that a fire originating in these sections can be
controlled, there must be a fire extinguisher system serving each designated
fire zone.
(b) The fire extinguishing system, the quantity of the extinguishing agent,
the rate of discharge, and the discharge distribution must be adequate to
extinguish fires. It must be shown by either actual or simulated flights
tests that under critical airflow conditions in flight the discharge of the
extinguishing agent in each designated fire zone specified in paragraph (a)
of this section will provide an agent concentration capable of extinguishing
fires in that zone and of minimizing the probability of reignition. An
individual "one-shot" system may be used for auxiliary power units, fuel
burning heaters, and other combustion equipment. For each other designated
fire zone, two discharges must be provided each of which produces adequate
agent concentration.
(c) The fire extinguishing system for a nacelle must be able to
simultaneously protect each zone of the nacelle for which protection is
provided.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-46, 43 FR
50598, Oct. 30, 1978]
Sec. 25.1197 Fire extinguishing agents.
(a) Fire extinguishing agents must--
(1) Be capable of extinguishing flames emanating from any burning of fluids
or other combustible materials in the area protected by the fire
extinguishing system; and
(2) Have thermal stability over the temperature range likely to be
experienced in the compartment in which they are stored.
(b) If any toxic extinguishing agent is used, provisions must be made to
prevent harmful concentrations of fluid or fluid vapors (from leakage during
normal operation of the airplane or as a result of discharging the fire
extinguisher on the ground or in flight) from entering any personnel
compartment, even though a defect may exist in the extinguishing system. This
must be shown by test except for built-in carbon dioxide fuselage compartment
fire extinguishing systems for which--
(1) Five pounds or less of carbon dioxide will be discharged, under
established fire control procedures, into any fuselage compartment; or
(2) There is protective breathing equipment for each flight crewmember on
flight deck duty.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 41 FR
55467, Dec. 20, 1976; Amdt. 25-40, 42 FR 15044, Mar. 17, 1977]
Sec. 25.1199 Extinguishing agent containers.
(a) Each extinguishing agent container must have a pressure relief to
prevent bursting of the container by excessive internal pressures.
(b) The discharge end of each discharge line from a pressure relief
connection must be located so that discharge of the fire extinguishing agent
would not damage the airplane. The line must also be located or protected to
prevent clogging caused by ice or other foreign matter.
(c) There must be a means for each fire extinguishing agent container to
indicate that the container has discharged or that the charging pressure is
below the established minimum necessary for proper functioning.
(d) The temperature of each container must be maintained, under intended
operating conditions, to prevent the pressure in the container from--
(1) Falling below that necessary to provide an adequate rate of discharge;
or
(2) Rising high enough to cause premature discharge.
(e) If a pyrotechnic capsule is used to discharge the extinguishing agent,
each container must be installed so that temperature conditions will not
cause hazardous deterioration of the pyrotechnic capsule.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5678, Apr. 8, 1970; Amdt. 25-40, 42 FR 15044, Mar. 17, 1977]
Sec. 25.1201 Fire extinguishing system materials.
(a) No material in any fire extinguishing system may react chemically with
any extinguishing agent so as to create a hazard.
(b) Each system component in an engine compartment must be fireproof.
Sec. 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 sections
of turbine engine installations, in numbers and locations ensuring prompt
detection of fire in those zones.
(b) Each fire detector system must be constructed and installed so that--
(1) It will withstand the vibration, inertia, and other loads to which it
may be subjected in operation;
(2) There is a means to warn the crew in the event that the sensor or
associated wiring within a designated fire zone is severed at one point,
unless the system continues to function as a satisfactory detection system
after the severing; and
(3) There is a means to warn the crew in the event of a short circuit in
the sensor or associated wiring within a designated fire zone, unless the
system continues to function as a satisfactory detection system after the
short circuit.
(c) No fire or overheat detector may be affected by any oil, water, other
fluids or fumes that might be present.
(d) There must be means to allow the crew to check, in flight, the
functioning of each fire or overheat detector electric circuit.
(e) Wiring and other components of each fire or overheat detector system in
a fire zone must be at least fire-resistant.
(f) No fire or overheat detector system component for any fire zone may
pass through another fire zone, unless--
(1) It is protected against the possibility of false warnings resulting
from fires in zones through which it passes; or
(2) Each zone involved is simultaneously protected by the same detector and
extinguishing system.
(g) Each fire detector system must be constructed so that when it is in the
configuration for installation it will not exceed the alarm activation time
approved for the detectors using the response time criteria specified in the
appropriate Technical Standard Order for the detector.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5678, Apr. 8, 1970; Amdt. 25-26, 36 FR 5493, Mar. 24, 1971]
Sec. 25.1207 Compliance.
Unless otherwise specified, compliance with the requirements of Secs.
25.1181 through 25.1203 must be shown by a full scale fire test or by one or
more 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.
[Amdt. 25-46, 43 FR 50598, Oct. 30, 1978]
Subpart F--Equipment
General
Sec. 25.1301 Function and installation.
Each item of installed equipment must--
(a) Be of a kind and design appropriate to its intended function;
(b) Be labeled as to its identification, function, or operating
limitations, or any applicable combination of these factors;
(c) Be installed according to limitations specified for that equipment; and
(d) Function properly when installed.
Sec. 25.1303 Flight and navigation instruments.
(a) The following flight and navigation instruments must be installed so
that the instrument is visible from each pilot station:
(1) A free air temperature indicator or an air-temperature indicator which
provides indications that are convertible to free-air temperature.
(2) A clock displaying hours, minutes, and seconds with a sweep-second
pointer or digital presentation.
(3) A direction indicator (nonstabilized magnetic compass).
(b) The following flight and navigation instruments must be installed at
each pilot station:
(1) An airspeed indicator. If airspeed limitations vary with altitude, the
indicator must have a maximum allowable airspeed indicator showing the
variation of VMO with altitude.
(2) An altimeter (sensitive).
(3) A rate-of-climb indicator (vertical speed).
(4) A gyroscopic rate-of-turn indicator combined with an integral slip-skid
indicator (turn-and-bank indicator) except that only a slip-skid indicator is
required on large airplanes with a third attitude instrument system useable
through flight attitudes of 360 deg. of pitch and roll and installed in
accordance with Sec. 121.305(j) of this title.
(5) A bank and pitch indicator (gyroscopically stabilized).
(6) A direction indicator (gyroscopically stabilized, magnetic or
nonmagnetic).
(c) The following flight and navigation instruments are required as
prescribed in this paragraph:
(1) A speed warning device is required for turbine engine powered airplanes
and for airplanes with VMO/MMO greater than 0.8 VDF/MDF or 0.8 V D/MD. The
speed warning device must give effective aural warning (differing
distinctively from aural warnings used for other purposes) to the pilots,
whenever the speed exceeds VMO plus 6 knots or MMO +0.01. The upper limit of
the production tolerance for the warning device may not exceed the prescribed
warning speed.
(2) A machmeter is required at each pilot station for airplanes with
compressibility limitations not otherwise indicated to the pilot by the
airspeed indicating system required under paragraph (b)(1) of this section.
[Amdt. 25-23, 35 FR 5678, Apr. 8, 1970, as amended by Amdt. 25-24, 35 FR
7108, May 6, 1970; Amdt. 25-38, 41 FR 55467, Dec. 20, 1976]
Sec. 25.1305 Powerplant instruments.
The following are required powerplant instruments:
(a) For all airplanes. (1) A fuel pressure warning means for each engine,
or a master warning means for all engines with provision for isolating the
individual warning means 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 independent pressure oil system of
each engine.
(5) An oil pressure warning means for each engine, or a master warning
means for all engines with provision for isolating the individual warning
means 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 is to be used in operation) for each tank.
(b) For reciprocating engine-powered airplanes. In addition to the
powerplant instruments required by paragraph (a) of this section, the
following powerplant instruments are required:
(1) A carburetor air temperature indicator for each engine.
(2) A cylinder head temperature indicator for each air-cooled engine.
(3) A manana creia
(2) mu
(2) mu
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(2) mu
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(2) mu
(2) mu
(2) mu
(2) mu
(2) mu
(2) mu
(2) mu
(2) mu
(2) mu
(2) mu
(2) mu
(2) mu
(2) mu
(2) mu
(2) mu
(2) mu
(2) mu
(2) mu
(2) mu
(2) mu
(2) mu
(2) mu
(2) mu
(2) mu
(2) mu
(2) mu
(2) mu
(2) mu
(2) mu
(2) mu
(2) mu
(2) mu
uments r
(2) mu
y paragraphs (a) and (c)
(2) mu
the following
(2) mu
lant i
(2) mu
nts ar
(2) mu
:
(1) An
(2) mu
(2) mu
(2) mu
thrust
(2) mu
(2) mu
eter
(2) mu
irectly
related to th
(2) mu
(2) mu
ilot
(2) mu
ndication must be
(2) mu
on the direct
(2) mu
ure
(2) mu
hrust
(2) mu
ara
(2) mu
that are directly re
(2) mu
o thrust.
The in
(2) mu
(2) mu
(2) mu
a c
(2) mu
thrust resulting from any engine
malfunction, damage
(2) mu
eter
(2) mu
ion.
(2) A positio
(2) mu
ating m
(2) mu
(2) mu
to the fli
(2) mu
ew w
(2) mu
th
(2) mu
(2) mu
device is in the reverse thrust position
(2) mu
engine
us
(2) mu
hrust
(2) mu
g device
(2) mu
(2) mu
(2) mu
rotor system
(2) mu
alance.
(e) For turbopropeller-powered ai
(2) mu
(2) mu
ition to the p
(2) mu
(2) mu
ument
(2) mu
d by p
(2) mu
a)
(2) mu
(2) mu
ection, the following
powerplant
(2) mu
nts ar
(2) mu
(2) mu
torqu, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.1307 Miscellaneous equipment.
The following is required miscellaneous equipment:
(a) [Reserved]
(b) Two or more independent sources of electrical energy.
(c) Electrical protective devices, as prescribed in this part.
(d) Two systems for two-way radio communications, with controls for each
accessible from each pilot station, designed and installed so that failure of
one system will not preclude operation of the other system. The use of a
common antenna system is acceptable if adequate reliability is shown.
(e) Two systems for radio navigation, with controls for each accessible
from each pilot station, designed and installed so that failure of one system
will not preclude operation of the other system. The use of a common antenna
system is acceptable if adequate reliability is shown.
[Amdt. 25-23, 35 FR 5678, Apr. 8, 1970, as amended by Amdt. 25-46, 43 FR
50598, Oct. 30, 1978; Amdt. 25-54, 45 FR 60173, Sept. 11, 1980; Amdt. 25-72,
55 FR 29785, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 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 perform their
intended functions under any foreseeable operating condition.
(b) The airplane systems and associated components, considered separately
and in relation to other systems, must be designed so that--
(1) The occurrence of any failure condition which would prevent the
(2) mu
ued safe flight and landing of the airplane is extremely improbable,
and
(2) The
(2) mu
ce of any other failure conditions which would reduce the
capability of the airplane or the ability of the crew to cope with adverse
operating conditions is improbable.
(c) Warning information must be provided to alert the crew to unsafe system
operating conditions, and to enable them to take appropriate corrective
action. Systems, controls, and associated monitoring and
(2) mu
s must
be
(2) mu
to minimize crew errors which could create additional hazards.
(d) Compli
(2) mu
the requirements of paragraph (b) of this section must
be shown by analysis, and where necessary, by appropriate ground, flight, or
simulator tests. The analysis must consider--
(1) Possible modes of failure, including malfunctions and damage from
external sources.
(2) The probability of multiple failures and undetected failures.
(3) The resulting effects on the airplane and occupants, considering the
stage of flight and operating conditions, and
(4) The crew warning cues, corrective action required, and the capability
of detecting faults.
(e) Each installation
(2) mu
se functioning i
(2) mu
d by this subchapter, and
that requires a power supply, is an "essential load" on the power supply. The
power sources and the system must be able to supply
(2) mu
g power loads
in probable operating combinations and for probable durations:
(1) Loads connected to the system with the system functioning normally.
(2) Essential loads, after failure of any one prime mover, power conve
(2) mu
,
or energy storage device.
(3) Essential loads after failure of--
(i) Any one engine on two-engine airplanes; and
(ii) Any two engines on three-or-more-engine airplanes.
(4) Essential loads for which an alternate source of power is required by
this chapter, after any failure or malfunction in any one power supply
system, distribution system, or other utilization system.
(f) In determining compliance with paragraphs (e) (2) and (3) of this
sect
(2) mu
he power loads may be assumed to be reduced under a monitoring
procedure consistent with safety in the kinds of operation authorized. Loads
not required in controlled flight need not be considered for the two-engine-
inoperative condition on airplanes with three or more engines.
(g) In showing compli
(2) mu
paragraphs (a) and (b) of this section with
regard to the electrical system and equipment design and installation,
critical environmental conditions must be considered. For electrical
generation, distribution, and utilization equipment required by or used in
complying with this chapter, except equipment covered by Technical Standard
Orders containing environmental test procedures, the ability to provide
(2) mu
uous, safe service under foreseeable environmental conditions may be
shown by environmental tests, design analysis, or reference to previous
comparable service experience on other aircraft.
[Amdt. 25-23, 35 FR 5679, Apr. 8, 1970, as amended by Amdt. 25-38, 41 FR
55467, Dec. 20, 1976; Amdt. 25-41, 42 FR 36970, July 18, 1977]
Inst
(2) mu
: Installation
Sec. 25.1321 Arrangement and visibility.
(a) Each flight, navigation, and powerplant
(2) mu
ument for use by any pilot
must be plainly visible to him from his station with the minimum practicable
deviation from his normal position and line of vision when he is looking
forward along the flight path.
(b) Th
(2) mu
(2) mu
uments
(2) mu
ec. 25.1303 must be grouped on the
(2) mu
ument panel and centered as nearly as practicable about the vertical
plane of the pilot's forward vision. In addition--
(1) The
(2) mu
nt that most effectively
(2) mu
s attitude must be on the
panel in the top center positio
;
(2) The
(2) mu
nt that most effectively in
(2) mu
s airspeed must be
adjacent to and di
(2) mu
to the left of the instrument in the top center
position:
(3) The
(2) mu
nt that most effectively i
(2) mu
altitude must be
adjacent to and directly to the right of the instrument in the top center
position; and
(4) The inst
(2) mu
that most effectively in
icates direction of flight must
be adjacent to and di
(2) mu
below the inst
ument in the top center position.
(c) Required
(2) mu
(2) mu
nts must be closely grouped on the
(2) mu
ument panel. In addition--
(1) The location of identical
(2) mu
struments for the engines must
prevent confusion as to which engine each i
(2) mu
nt relates; and
(2) Powerplant
(2) mu
nts vital to the safe operation of the airplane must
be plainly visible to the appropriate crewmembers.
(d) Ins
(2) mu
t panel vibration may not damage or impair the accuracy of any
(2) mu
ument.
(e) If a visual
(2) mu
is provided t
(2) mu
cate malfunction of an
(2) mu
ument, it must be effective under all probable cockpit lighting
conditions.
[Amdt. 25-23, 35 FR 5679, Apr. 8, 1970, as amended by Amdt. 25-41, 42 FR
36970, July 18, 1977]
Sec. 25.1322 Warning, caution, and advisory lights.
If warning, caution or advisory lights are installed in the cockpit, they
must, unless otherwise approved by the Administrator, be--
(a) Red, for warning lights (lights
(2) mu
g a hazard which may require
immediate corrective action);
(b) Amber, for caution lights (lights
ndicating the 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) o
(2) mu
on, provided the color differs
sufficiently from the colors prescribed in paragraphs (a) through (c) o
this
sect
on to avoid possible confusion.
[Amdt. 25-38, 41 FR 55467, Dec. 20, 1976]
Sec. 25.1323 Airspeed indicating system.
For each airspeed indicating system,
(2) mu
g apply:
(a) Each airspeed indicating instrument must be approved and must be
calibrated to
(2) mu
rue airspeed (at sea level with a standard
atmosphere) with a minimum practicable inst
ument calibration error when the
corresponding pitot and static pressures are applied.
(b) Each system must be calibrated to determine the system error (that is,
the relation between IAS and CAS) in flight and during the accelerated
takeoff ground run. The ground run calibration must be determined--
(1) From 0.8 of the minimum value of V1 to the maximum value of V2,
considering the approved ranges of altitude and weight; and
(2) With the flaps and power settings corresponding to the values
determined in the establish
(2) mu
he takeoff path under Sec. 25.111
assuming that the critical engine fails at the minimum value of V1.
(c) The airspeed error of the installation, excluding the airspeed
indicator
(2) mu
ument calibration error, may not exceed three percent or five
knots, whichever is greater, throughout the speed range, from--
(1) VMO to 1.3 VS1, with flaps retracted; and
(2) 1.3 VS0 to VFE with flaps in the landing position.
(d) Each system must be arranged, so far as practicable, to prevent
malfunction or serious error du
(2) mu
entry of moisture, dirt, or other
substances.
(e) Each system must have a heated pitot tube or an equivalent means of
preventing malfunction due to icing.
(f) Where duplicate airspeed indicators are required, their respective
pitot tubes must be far enough apart to avoid damage to both tubes in a
collision with a bird.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-57, 49 FR
6849, Feb. 23, 1984]
Sec. 25.1325 Static pressure systems.
(a) Each i
(2) mu
nt with static air case connections must be vented to the
outside atmosphere through an appropriate piping system.
(b) Each static port must be designed and located in such manner that the
static pressure system performance is least affected by airflow variation, or
by moisture or other foreign matter, and that the correlation between air
pressure in the static pressure system and true ambient atmospheric static
pressure is not changed when the airplane is exposed to the continuous and
intermittent maximum icing conditions defined in Appendix C of this part.
(c) The design and installation of the static pressure system must be such
that--
(1) Positive drainage of moisture is provided; chafing of the tubing and
excessive distortion or restriction at bends in the tubing is avoided; and
the materials used are durable, suitable for the purpose intended, and
protected against corrosion; and
(2) It is airtight except for the port into the atmosphere. A proof test
must be conducted to demonstrate the integrity of the static pressure system
in the following manner:
(i) Unpressurized air
(2) mu
Evacuat
(2) mu
tatic pressure system to a
pressure differential of approximately 1 inch of mercury or to a reading on
the altim
(2) mu
1,000 feet above the airplane elevation at the time of the
test. Without additional pumping for a period of 1 minute, the loss of
indicated altitude must not exceed 100 feet on the altimeter.
(ii) Pressurized airplanes. Evacuat
the static pressure system until a
pressure differential equivalent to the maximum cabin pressure differential
for which the airplane is type certificated is achieved. Without additional
pumping for a period of 1 minute, the loss of indicated altitude must not
exceed 2 percent of the equivalent altitude of the maximum cabin differential
pressure or 100 feet, whichever is greater.
(d) Each pressure altim
ter must be approved and must be calibr
(2) mu
o
(2) mu
pressure altitude in a standard atmosphere, with a minimum
practicable calibration error when the corresponding static pressures are
applied.
(e) Each system must be designed and installed so that the error in
indicated pressure altitude, at sea level, with a standard atmosphere,
excluding
(2) mu
nt calibration error, does not result in an error of more
than +/-30 feet per 100 knots speed for the appropriate configuration in the
speed range between 1.3 VS0 with flaps extended and 1.8 VS1 with flaps
retracted. However, the error need not be less than +/-30 feet.
(f) If an altimeter system is fitted with a
(2) mu
hat provides
corrections to the altimeter indication, the device must be designed and
installed in such manner that it can be bypassed when it malfunctions, unless
an alternate altimeter system is provided. Each correction device must be
fitted with a means for indicating the occurrence of reasonably probable
malfunctions, including power failure, to the fli
ht crew. The indicating
means must be effective for any cockpit lighting condition likely to occur.
(g) Except as provided in paragraph (h)
(2) mu
if the static
pressure system incorporates both a primary and an alternate static pressure
source, the means for selecting one or the other source must be designed so
that--
(1) When either source is selected
(2) mu
ther is blocked off; and
(2) Both sources cannot be blocked off simultaneously.
(h) For unpressurized airplanes, paragraph (g)(1) of this section does not
apply if it can be demonstr
(2) mu
hat the static pressure system calibration,
when either static pressure source is selected
is not
(2) mu
ed by the other
static pressure source being open or blocked.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-5, 30 FR
8261, June 29, 1965; Amdt. 25-12, 32 FR 7587, May 24, 1967; Amdt. 25-41, 42
FR 36970, July 18, 1977]
Sec. 25.1326 Pitot heat
(2) mu
on systems.
If a flight i
(2) mu
nt pitot heating system is installed, an
(2) mu
on
system must be provided to
(2) mu
to the fli
h
(2) mu
(2) mu
hat pitot
heating system is not operating
(2) mu
ndication system must comply with the
following requirements:
(a) The indication provided must incorporate an amber light that is in
clear view of a flight crewmember.
(b) The
(2) mu
on provided must be designed to alert the flight crew if
either of the following conditions 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.
[Amdt. 25-43, 43 FR 10339, Mar. 13, 1978]
Sec. 25.1327 Magnetic directio
(2) mu
ator.
(a) Each magnetic directio
(2) mu
must be installed so that its
accuracy is not excessively affected by the airplane's vibration or magnetic
fields.
(b) The compensated installation
may not have a deviation, in level flight,
greater than 10 degrees on any heading.
Sec. 25.1329 Automatic
(2) mu
system.
(a) Each automatic
(2) mu
system must be approved and must be designed so
that the
(2) mu
atic
(2) mu
(2) mu
quickly and positively disengaged by the
pilots to prevent it from interfering with their control of the airplane.
(b) Unless there is automatic synchronization, each system must have a
means to readily
(2) mu
to the pilot the alignment of the actuati
(2) mu
evice
in relatio
(2) mu
he control system it operates.
(c) Each manually operated control for the system must be readily
accessible to the pilots.
(d) Quick release (emergency) controls must be on both control wheels, on
the side of each wheel opposite the throttles.
(e) Attitude controls must operate in the plane and sense of motion
specified in Secs. 25.777(b) and 25.779(a) for cockpit controls. The
direction of mot
(2) mu
e plainly
(2) mu
d on, or adjacent to, each
control.
(f) The system must be designed and adjusted so that, within the range of
adjustment available to the human pilot, it cannot produce hazardous loads on
the airplane, or create hazardous deviations in the flight path, under any
condition of flight appropriate to its use, either during normal operation or
in the event of a malfunction, assuming that corrective action begins within
a reasonable period of time.
(g) If the automatic pilot integrates signals from auxiliary controls or
furnishes signals for operation of other equipment, there must be positive
interlocks and sequencing of engagement to prevent improper operation.
Protection against adverse interaction of integrated components, resulting
from a malfunction, is also required.
(h) If the auto
(2) mu
pilot system
(2) mu
coupled to airborne navigation
equipment, means must be provided t
(2) mu
to
(2) mu
ht crew the c
(2) mu
(2) mu
mode of operation. Selector switch positio
is not
acceptable as a means of
indication.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-46, 43 FR
50598, Oct. 30, 1978]
Sec. 25.1331 Instruments using a power supply.
(a) For each
(2) mu
nt
(2) mu
ec. 25.1303(b) that uses a power
supply
(2) mu
wing apply:
(1) Each i
(2) mu
nt must have a visual
means integral with, the
(2) mu
nt,
t
(2) mu
cate when power adequate to sustain proper
(2) mu
ument performance is
not being supplied. The power must be measured at or near the point where it
enters the ins
(2) mu
ts. For electric
(2) mu
uments,
(2) mu
is considered to
be adequate
(2) mu
he voltage is within approved limits.
(2) Each ins
(2) mu
t must, in the event of the failure of one power source,
be supplied by another power source. This may be accomplished automatically
or by manual means.
(3) If an i
(2) mu
nt presenting navigation data receives information from
sources external to that instrument and loss of that information would render
the presented data unreliable, the ins
(2) mu
must incorporate a visual
means
to warn the crew, when such loss of information occurs, that the presented
data should not be relied upon.
(b) As used in this section, "
(2) mu
nt" includes devices that are
physically contained in one unit, and devices that are composed of two or
more physically separate units or components connected together (such as a
remot
(2) mu
cating gyroscopic directio
(2) mu
hat includes a magnetic
sensing element, a gyroscopic unit, an amplifier and a
(2) mu
ator connected
together).
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-41, 42 FR
36970, July 18, 1977]
Sec. 25.1333 Ins
(2) mu
t systems.
For systems that operate the ins
ument
(2) mu
Sec. 25.1303(b) which
are located at each pilot's station--
(a) Means must be provided to connect the required inst
(2) mu
at the first
pilot's station to operating systems which are independent of the operating
systems at other flight crew stations, or other equipment;
(b) Th
equipment, systems, and installations must be designed so that one
display of the information essential to the safety of flight which is
provided by the inst
uments, including attitude, directio
, airspeed, and
altitude will remain available to the pilots, without additional crewmember
action, after any single failure or combination of failures that is not shown
to be extremely improbable; and
(c) Additional i
(2) mu
nts, systems, or equipment may not be connected to
the operating systems for the required i
(2) mu
nts, unless provisions are
made to ensure the continued normal functioni
(2) mu
required
(2) mu
nts
in the event of any malfunction of the additional
(2) mu
uments, systems, or
equipment which is not shown to be extremely improbable.
[Amdt. 25-23, 35 FR 5679, Apr. 8, 1970, as amended by Amdt. 25-41, 42 FR
36970, July 18, 1977]
Sec. 25.1335 Flight director systems.
If a flight director system is installed, means must be provided to
(2) mu
to the fli
h
crew its current mode of operation. Selector switch
position is not acceptable as a means of
(2) mu
on.
[Amdt. 25-41, 42 FR 36970, July 18, 1977]
Sec. 25.1337 Powerplant
(2) mu
nts.
(a) I
(2) mu
nts and inst
(2) mu
lines.
(1) Each
(2) mu
and auxiliary power unit i
(2) mu
nt line must meet the
requirements of Secs. 25.993 and 25.1183.
(2) Each line carrying flammable fluids under pressure must--
(i) Have restricting orifices or other safety devices at the source of
pressure to prevent the escape of excessive fluid if the line fails; and
(ii) Be installed and located so that the escape of fluids would not create
a hazard.
(3) Each powerplant and auxiliary power unit
(2) mu
nt that utilizes
flammable fluids must be installed and located so that the escape of fluid
would not create a hazard.
(b) Fuel quantity indicator. There must be
(2) mu
(2) mu
to the flight
crewmembers, the quantity, in gallons or equivalent units, of usable fuel in
each tank during flight. In addition--
(1) Each fuel quantity
(2) mu
must be calibr
(2) mu
o read "zero" during
level flight when the quantity of fuel remaining in the tank is equal to the
unusable fuel supply determined under Sec. 25.959;
(2) Tanks with interconnected outlets and airspaces may be treated as one
tank and need not have separate indicators; and
(3) Each exposed sight gauge, used as a fuel quantity indicator, must be
protected against damage.
(c) Fuel flowmeter system. If a fuel flowmeter system is installed, each
meter
ng component must have a means for bypassing the fuel supply if
(2) mu
lfunction of that component severely restricts fuel flow.
(d) Oil quantity in
(2) mu
. There must be a stick gauge or equivalent means
to
(2) mu
e quantity of oil in each tank. If an oil transfer or reserve
oil supply system is installed, there must be a
(2) mu
(2) mu
to the
flight crew, in flight, the quantity of oil in each tank.
(e) Turb
(2) mu
eller blade positio
indicator. Required turbopropeller blade
position indicators must begin
(2) mu
g before the blade moves more than
eight degrees below the flight low pitch stop. The source of indication must
di
(2) mu
sense the blade position.
(f) Fuel
(2) mu
ndicator. There must be means to measure fuel pressure,
in each system supplying reciprocating engines, at a point downstream of any
fuel pump except fuel injection pumps.
(2) mu
ition--
(1) If necessary for the maintenance of proper fuel delivery pressure,
there must be a connection to transmit the carburetor air intake static
pressure 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 must be independently connected to the carburetor inlet
pressure to avoid erroneous readings.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-40, 42 FR
15044, Mar. 17, 1977]
Electrical Systems and Equipment
Sec. 25.1351 General.
(a) Electrical system capacity. The required generating capacity, and
number and kinds of power sources must--
(1) Be determined by an electrical load analysis; and
(2) Meet the requirements of Sec. 25.1309.
(b) Generating system. The generating system includes electrical power
sources, main power busses, transmission cables, and associated control,
regulation, and protective devices. It must be designed so that--
(1) Power sources function properly when independent and when connected in
combination;
(2) No failure or malfunction of any power source can create a hazard or
impair the ability of remaining sources to supply essential loads;
(3) The system voltage and frequency (as applicable) at the terminals of
all essential load equipment can be maintained within the limits for which
the equipment is designed, during any probable operating condition; and
(4) System transients due to switching, fault clearing, or other causes do
not make essential loads inoperative, and do not cause a smoke or fire
hazard.
(5) There are means accessible, in flight, to appropriate crewmembers for
the individual and collective disconnection of the electrical power sources
from the system.
(6) There are means to i
(2) mu
to appropriate crewmembers the generating
system quantities essential for the safe opera
(2) mu
ystem, such as the
voltage and c
(2) mu
t supplied by each generator.
(c) External power. If provisions are made for connecting external power to
the airplane, and that external power can be electrically connected to
equipment other than that used for engine starting, means must be provided to
ensure that no external power supply having a reverse polarity
(2) mu
(2) mu
phase sequence, can supply power to the airplane's electrical system.
(d) Operation without normal electrical power. It must be shown by
analysis, tests, or both, that the airplane can be operated safely i
VFR
conditions, for a period of not less than five minutes, with the normal
electrical power (electrical power sources excluding the battery)
inoperative, with critical type fuel (from the standpoint of flameout and
restart capability), and with the airplane initially at the maximum
certificated altitude. Parts of the electrical system may remain on if--
(1) A single malfunction, including a wire bundle or junction box fire,
cannot result in loss of both the part turned off and the part turned on; and
(2) The parts turned on are electrically and mechanically isolated from the
parts turned off.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-41, 42 FR
36970, July 18, 1977; Amdt. 25-72, 55 FR 29785, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.1353 Electrical equipment and installations.
(a) Electrical equipment, controls, and wiring must be installed so that
operation of any one unit or system of units will not adversely affect the
simultaneous operation of any other electrical unit or system essential to
the safe opera
ion.
(b) Cables must be grouped, routed, and spaced so that damage to essential
circuits will be minimized if there are faults in heavy c
(2) mu
t-carrying
cables.
(c) Storage batteries must be designed and installed as follows:
(1) Safe cell temperatures and pressures must be maintained during any
probable charging or discharging condition. No uncontrolled increase in cell
temperature may r
(2) mu
(2) mu
he battery is recharged (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 (c)(1) of this section must be shown by test
unless experience with similar batteries and installations has shown that
maintaining safe cell temperatures and pressures presents no problem.
(3) No explosive or toxic gases emitted by any battery in normal operation,
or as the result of any probable malfunction in the charging system or
battery installation, may accumulate in hazardous quantities within the
airplane.
(4) No corrosive fluids or gases that may escape from the battery may
damage surrounding airplane structures or adjacent essential equipment.
(5) Each nickel cadmium battery installation capable of being used to start
an engine or auxiliary power unit must have provisions to prevent any
hazardous effect on structure or essential systems that may be caused by the
maximum amount of heat the battery can generate during a short circuit of the
battery or of its individual cells.
(6) Nickel cadmium battery installations capable of being used to start an
engine or auxiliary power unit must have--
(i) A system to control the charging rate of the battery auto
atically so
as to prevent battery overheating;
(ii) A battery temperature sensing and over-temperature warning system with
a mea
(2) mu
or disconnecting the battery from its charging source in the eve
(2) mu
of an over-temperature condition
(2) mu
(2) mu
battery failure sensing and
arning system with a mea
(2) mu
or
disconnecting the battery from its charging source in the event of battery
failure.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-41, 42 FR
36970, July 18, 1977; Amdt. 25-42, 43 FR 2323, Jan. 16, 1978]
Sec. 25.1355 Distribution system.
(a) The distribution system includes the distribution busses, their
associated feeders, and each control and protective
(2) mu
(b) [Reserved]
(c) If two independent sources of electrical power for particular equipment
or systems are required by this chapter, in the event of the failure of one
power source for such equipment or system, another power source (including
its separate feeder) must be autor system,u
ally provided or be manually selectable
to maintain equipment or system operation.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR
5679, Apr. 8, 1970; Amdt. 25-38, 41 FR 55468, Dec. 20, 1976]
Sec. 25.1357 Circuit protective devices.
(a) Automatic protective
evices must be used to minimize distress to the
electrical system and hr system,u
to the airplane in the event of wiring faults or
serious malfunction of the system or connected equipment.
(b) Th
protective and control devices in the generating system must be
r system,u
to de-energize and disconnect faulty power sources and power
transmission equipment from their associated busses with sufficient rapidity
to provide protection from hazardous over-voltage and other malfunctioning.
(c) Each resettable circuit protective
evice must be designed so that,
when an overload or circuit fault exists, it will open the circuit
irrespective of the position of the operating control.
(d) If the ability to reset a circuit breaker or replace a fuse is
essential to safety in flight, that circuit breaker or fuse must be locar system,u
and identified so that it can be readily reset or replaced in flight.
(e) Each circuit for essential loads must have individual circuit
protection. However, individual protection for each circuit in an essential
load system (such as each position light circuit in a system) is not
required.
(f) If fuses are used, there must be spare fuses for use in flight equal to
at least 50 percent of the number of fuses of each rating required for
complete circuit protection.
(g) Autor system,u
reset circuit breakers may be used as integral protectors for
electrical equipment (such as thermal cut-outs) if there is circuit
protection to protect the cable to the equipment.
Sec. 25.1359 [Removed. 55 FR 29785, July 20, 1990]
EDITORIAL NOTE: For the convenience of the user, the removed text is
set out below.
Sec. 25.1359 Electrical system fire and smoke protection.
(a) Components of the electrical system must meet the applicable fire and
smoke protection requirements of Secs. 25.831(c), 25.863, and 25.867.
(b) Electrical cables, terminals, and equipment in designated fire zones,
that are used during emergency procedures, must be at least fire-resistant.
(c) Main power cables (including generator cables) in the fuselage must be
designed to allow a reasonable degree of deformation and stretching without
failure and must--
(1) Be isolated from flammable fluid lines; or
(2) Be shrouded by means of electrically insulated flexible conduit, or
equivalent, which is in adr system,u
o the normal cable insulation.
(d) Insulation on electrical wire and electrical cable installed in any
area of the fuselage must be self-extinguishing when tested at an angle of 60
deg. in accordance with the applicable portions of Appendix F of this part,
or other approved equivalent methods. The average burn length may not exceed
3 inches and the average flame time after removal of the flame source may not
exceed 30 seconds. Drippings from the test specimen may not continue to flame
for more than an average of 3 seconds after falling.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-17, 33 FR
9066, June 20, 1968; Amdt. 25-32, 37 FR 3972, Feb. 24, 1972; Amdt. 25-57, 49
FR 6849, Feb. 23, 1984]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.1363 Electrical system tests.
(a) When laboratory tests of the electrical system are conducted--
(1) The tests must be performed on a mock-up using the same generating
equipment used in the airplane;
(2) The equipment must simulate the electrical characteristics of the
distribution wiring and connected loads to the extent necessary for valid
test results; and
(3) Laboratory generator drives must simulate the actual prime movers on
the airplane with respect to their reaction to generator loading, including
loading due to faults.
(b) For each flight condition that cannot be simulated adequately i
the
laboratory or by ground tests on the airplane, flight tests must be made.
Lights
Sec. 25.1381 Inst
ument lights.
(a) Ther system,u
ument lights must--
(1) Provide sufficient illumination to make each i
strument, switch and
other device necessary for safe operation easily readable unless sufficient
illumination is available frr system,u
other source; and
(2) Be installed so that-r system,u
Their direct rays are shielded from the pilot's eyes; and
(ii) No objectionable reflections are visible to thr system,u
ot.
(b) Unless undimmedr system,u
ument lights are satisfactory under each expected
flight condition, there must be a means to control the intensity of
illumination.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-72, 55 FR
29785, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.1383 Landing lights.
(a) Each landing light must be approved, and must be installed so that--
(1) No objectionable glare is visiblr system,u
pilot;
(2) The r system,u
is not adversely affected by halation; 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 r system,u
r system,u
to the pilots when the landing lights
are extended.
Sec. 25.1385 Position light system installation.
(a) General. Each part of each position light system must meet the
applicable requirements of this section and each system as a whole must meet
the requirements of Secs. 25.1387 through 25.1397.
(b) Forward positio
lights. Forward position lights must consist of a red
and a green light spaced laterally as far apart as practicable and installed
forward on the airplane so that, with the airplane in the normal flyir system,u
osition, the red light is on the left side and the green light is on the
right side. Each light must be approved.
(c) Rear position light. The rear position light must be a white light
mounted as far aft as practicable on the tail or on each wing tip, and must
be approved.
(d) Light covers and color system,u
s. Each light cover or color filter must
be
at least flame resistant and may notr system,u
e color or shape or lose any
appreciable light transmission during normal use.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 41 FR
55468, Dec. 20, 1976]
Sec. 25.1387 Position light system dihedral angles.
(a) Except as provided in paragraph (e) or system,u
on, each forward and
rear positio
light must, as installed, show r system,u
roken light within the
dihedral angles described in this section.
(b) Dihedral angle L (left) is formed by two intersecting vertical planes,
the first parallel to the longitudinal axis of the airplane, and the other at
110 degrees to the left of the first, as viewed when looking forward along
the longitudinal axis.
(c) Dihedral angle R (right) is formed by two intersecting vertical
lanes,
the first parallel to the longitudinal axis of the airplane, and the other at
110 degrees to the right of the first, as viewed when looking forward along
the longitudinal axis.
(d) Dihedral angle A (aft) is formed by two intersecting vertical planes
making angles of 70 degrees to the right and to the left, respectively, to a
vertical plane passing through the longitudinal axis, as viewed when looking
aft along the longitudinal axis.
(e) If the rear position light, when mounted as far aft as practicable in
ar system,u
nce with Sec. 25.1385(c), cannot show r system,u
roken light within dihedral
angle A (as defined in paragraph (d) or system,u
on), a solid angle or
angles of obstructed visibility totaling not more than 0.04 steradians is
allowable within that dihedral angle, if such solid angle is within a cone
whose apex is at the rear position light and whose elements make an angle of
30 deg. with a vertical line passing through the rear position light.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-30, 36 FR
21278, Nov. 5, 1971]
Sec. 25.1389 Position light distribution and intensities.
(a) General. The intensities prescribed in this section must be provided by
new equipment with light covers and color filters in place. Intensities must
be@determined with the light source operating at a steady value equal to the
average luminous or system,u
ut of the source at the normal operating voltage of the
airplane. The light distribution and intensity of each position light must
meet the requirements of paragraph (b) of this section.
(b) Forward and rear position lights. The light distribution and
intensities of forward and rear position lights must be expressed in terms of
minimum intensities in the horizontal plane, minimum intensities in any
vertical plane, and maximum intensities in overlapping beams, within dihedral
angles L, R, and A, and must meet r system,u
g requirements:
(1) Intensities in the horizontal plane. Each intensity in the horizontal
plane (the plane containing the longitudinal axis of the airplane and
perpendicular to the plane of symmetry of the airplane) must equal or exceed
the values in Sec. 25.1391.
(2) Intensities in any vertical plane. Each i
tensity in any vertical
plane (the plane perpendicular to the horizontal plane) must equal or exceed
the appropriate value in Sec. 25.1393, where I is the minimum intensr system,u
prescribed in Sec. 25.1391 for the corresponding angles in the horizon@l
plane.
(3) Intensrties in overlaps between adjacent signals. No intensity in any
overlap between adjacent signals may exceed the values given in Sec. 25.1395,
except that higher intensities in overlaps may be used with main beam
intensities substantially greater than the minima specified in Secs. 25.1391
and 25.1393 if the overlap intensities in relation to the main beam
intensities do not adversely affect signal clarity. When the peak intensity
of the forward position lights is more than 1r system,u
andles, the maximum overlap
intensities between them may exceed the values given in Sec. 25.1395 if the
overlap intensity in Area A is not more than 10 percent of peak positio
light intensrty and the overlap intensrty in Area B is not greater than 2.5
percent of peak position light intensity.
Sec. 25.1391 Minimum intensities in the horizontal plane of forward and rear
position lights.
Each position light intensity must equal or exceed the applicable values in
r system,u
g table:
Angle from right or left
Dihedral angle (light of longitudinal axis,
included) measured from dead ahead Intensity (candles)
L and R (forward red and 0 deg. to 10 deg. 40
green) 10 deg. to 20 deg. 30
20 deg. to 110 deg. 5
A (rear white) 110 deg. to 180 deg. 20
Sec. 25.1393 Minimum intensities in any vertical plane of forward and rear
position lights.
Each position light intensity must equal or exceed the applicable values in
the following table:
Angle above or
below the Intensity,
horizontal plane l
0 deg. 1.00
0 deg. to 5 deg. 0.90
5 deg. to 10 deg. 0.80
10 deg. to 15 deg. 0.70
15 deg. to 20 deg. 0.50
20 deg. to 30 deg. 0.30
30 deg. to 40 deg. 0.10
40 deg. to 90 deg. 0.05
Sec. 25.1395 Maximum intensrties in overlapping beams of forward and rear
position lights.
No position light intensrty may exceed the applicable values in the
following table, except as provided in Sec. 25.1389(b)(3).
Maximum intensity
Area A Area B
Overlaps (candles) (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
Where--
(a) Area A includes all directio
s in the adjacent dihedral angle that pass
through the light source and intersect the common boundary plane at more than
10 degrees but less than 20 degrees; and
(b) Area B includes all directio
s in the adjacent dihedral angle that pass
through the light source and intersect the common boundary plane at more than
20 degrees.
Sec. 25.1397 Color specifications.
Each position light color must have the applicable International Commission
on Illumination chror system,u
ity 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 than 0.540;
"y" is not
less than "x --0.040" or "y0--0.010", whichevr system,u
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 Planckian radiator for the value of
"x" considered.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-27, 36 FR
12972, July 10, 1971]
Sec. 25.1399 Riding light.
(a) Each riding (anchor) light required for a seaplane or amphibian must be
installed so that it can--
(1) Show a white light for at least 2 nautical miles at night under clear
atmospheric conditions; and
(2) Show the maximum r system,u
roken light practicable when the airplane is moored
or drifting on the water.
(b) Externally hung lights may be used.
Sec. 25.1401 Anticollision light system.
(a) General. The airplane must have an anticollision light system that--
(1) Consists of one or more approved anticollision lights located so that
their light will not impair the crew's vision or detract from the conspicuity
of the position lights; and
(2) Meets the requirements of paragraphs (b) through (f) of this section.
(b) Field of coverage. The system must consist of enough lights r system,u
lluminate the vital areas around the airplane considering the physical
configuration and flight characteristics of the airplane. The field of
coverage must extend in each direction within at least 75 degrees above and
75 degrees below the horizontal plane of the airplane, except that a solid
angle or angles of obstructed visibility totaling not more than 0.03
steradians is allowable within a solid angle equal to 0.15 steradians
centered about the longitudinal axr system,u
he rearward directio
.
(c) Flashing characteristics. The arrangement of the system,
hat is, the
number of light sources, beam width, speed of rotation, and other
characteristics, must give an effective flash frequency of not less than 40,
nor more than 1r system,u
ycles per minute. The effective flash frequency is the
frequency at which the airplane's complete anticollision light system is
observed from a distance, and applies to each sector of light including any
overlaps that existr system,u
he system consists of more than one light source.
In overlaps, flash frequencies may exceed 100, but not 180 cycles per minute.
(d) Color. Each anticollision light must be either aviation red or aviation
white and must meet the applicable requirements of Sec. 25.1397.
(e) Light intensity. The minimum light i
tensities in all vertical planes,
measured with the red filter (if used) and expressed in terms of "effective"
intensities, must meet the requirements of paragraph (fr system,u
ection. The
following relatio
must be assumed:
t2
INTEGRAL I(t)dt
t1
Ie =
--------------
0.2+(t2-t1)
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 intensity is obtained when t2 and t1
are chosen so that the effective intensrty is equal to the instantaneous
intensity at t2 and t1.
(f) Minimum effective intensities for anticollision lights. Each
anticollision light effective intensity must equal or exceed the applicable
values in the following table.
Angle above or Effective
below the intensity
horizontal plane (candles)
0 deg. to 5 deg. 400
5 deg. to 10 deg. 240
10 deg. to 20 deg. 80
20 deg. to 30 deg. 40
30 deg. to 75 deg. 20
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-27, 36 FR
12972, July 10, 1971; Amdt. 25-41, 42 FR 36970, July 18, 1977]
Sec. 25.1403 Wing icing detection lights.
Unless operations at night in known or forecast icing conditions are
prohibited by an operating limitation, a means must be provided for
illuminating or otherwise determining the formation of ice on the parts of
the wingr system,u
e critical from the standpoint of ice accumulation. Any
illur system,u
that is used must be of a type that will not cause glare or
reflection that would handicap crewmembers in the performance of their
duties.
[Amdt. 25-38, 41 FR 55468, Dec. 20, 1976]
Safety Equipment
Sec. 25.1411 General.
(a) Accessibility. Required safety equipment to be used by the crew in an
emergency must be readily accessible.
(b) Stowage provisions. Stowage provisions for required emergency equipment
must be furnished and must--
(1) Be arranged so that the equipment is directly accessible and its
location is obvious; and
(2) Protect the safety equipment from inadvertent damage.
(c) Emergency exit descent r system,u
The stowage provisions for the emergency
exit descent device r system,u
ec. 25.809(f) must be at the exits for which
they are intended.
(d) Liferafts. (1) The stowage provisions for the liferafts described in
Sec. 25.1415 must accommodate enough rafts for the maximum number of
occupants for which certification for ditching is requested.
(2) Liferafts must be stowed near exits through which the rafts can be
launched during an unplanned ditching.
(3) Rafts auto
atically or remotely released outside the airplane must be
attached to the airplane by means of the static line prescribed in Sec.
25.1415.
(4) The stowage provisions for each portable liferaft must allow rapid
detachment and removal of the raft for use at other than the intended exits.
(e) Long-range signalr system,u
vice
The stowage provisions for the long-range
signaling device r system,u
ec. 25.1415 must be near an exit available
during an unplanned ditching.
(f) Life preserver stowage provisions. The stowage provisions for life
preservers described in Sec. 25.1415 must accommodate one life preserver for
each occupant for which certification for ditching is requested. Each life
preserver must be within easy reach of each seated occupant.
(g) Life line stowage provisions. If certification for ditching under Sec.
25.801 is requested, there must be provisions to store life lines. These
provisions must--
(1) Allow one life line to be attached to each side of the fuselage; and
(2) Be arranged to allow the life lines to be used to enable the occupants
to stay on the wing after ditching.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended 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 FR 43925, Oct. 27, 1989; Amdt.
25-79, 58 FR 45229, Aug. 26, 1993]
*****************************************************************************
58 FR 45224, No. 164, Aug. 26, 1993
SUMMARY: These amendments to the airworthiness standards for transport
category airplanes and the operating rules for air carrier operators of such
airplanes modify the procedures for conducting an emergency evacuation
demonstration. These include a requirement that the flightcrew take no active
role in the demonstration, and ar system,u
e to the age/sex distribution
requirement for demonstration participants. In addition, the airworthiness
standards are amended to standardize the illumination requirements for the
handles of the various types of passenger emergency exits, and to add a
requirement to prevent the inadvertent disablir system,u
public address system
because of an unstowed microphone. These amendments are intended to enhance
the provisions for egress of occupants of transport category airplanes under
emergency conditions.
EFFECTIVE DATE: September 27, 1993.
*****************************************************************************
Sec. 25.1413 [Removed. 55 FR 29785, July 20, 1990]
EDITORIAL NOTE: For the convenience of the user, the removed text is
set out below.
Sec. 25.1413 Safety belts.
(a) If there are means to r system,u
o the passengers when safety belts
should be fastened, they must be installed to be operated from either pilot
seat.
(b) The rated strength of safety belts may not be less than that required
to withstand the ultimate load factors specified in Sec. 25.561, considering
the dimensional characteristics of the belt installation for the specific
seat or berth arrangement.
(c) Each belt and shoulder harness must be attached so that no part of the
anchorage can fail at a load lower than thar system,u
ich would result from the
application of ultimate load factors equal to those specified in Sec. 25.561,
multipliedr system,u
factor of 1.33. This factor must be used instead of the
fitting factor prescribed in Sec. 25.625. The forward load factor need not be
appliedrto safety belts for berths.
(d) Each safety belt must be equipped with a metal to metal latching
device.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-44, 43 FR
46233, Oct. 5, 1978; Amdt. 25-51, 45 FR 7755, Feb. 4, 1980]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.1415 Ditching equipment.
(a) Ditching equipment used in airplanes to be certificated for ditching
under Sec. 25.801, and required by the operating rules of this chapter, must
meet the requirements of this section.
(b) Each liferaft and each life preserver must be approved.r system,u
ition--
(1) Unless excess rafts of enough capacity are provided, the buoyancy and
seating capacity beyond the rated capacity of the rafts must accommodate all
occupants of the airplane in the eve
t of a loss of one raft of the largest
rated capacity; and
(2) Each raft must have a trailing line, and must have a static line
designed to hold the raft near the airplane r system,u
o release it if the airplane
becomes totally submerger system,u
(c) Approved survival equipment must be attached to each liferaft.
(d) There must be a survival type emergency locator transmitter that meets
the applicable requirements of TSO-C91 for use in one liferaftr system,u
r airplanes not certificated for ditching under Sec. 25.801 and not
having approved life preservers, there must be an approved flotation means
for each occupant. This means must be within easy reach of each seated
occupant and must be readily removable frrm the airplane.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended 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]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.1416 [Removed. 55 FR 29785, July 20, 1985]
EDITORIAL NOTE: For the convenience of the user, the removed text is
set out below.
Sec. 25.1416 Pneumatic de-icer boot system.
If certification with ice protection provisions is desired and a pnuer system,u
de-icer boot system is installed--
(a) The system must meet the requirements specified in Sec. 25.1419,
(b) Th
system and its components must be designed to perform their
intended function under any normal system operating temperature or pressure,
and
(c) Means to i
r system,u
tor system,u
hr system,u
that the pneur system,u
de-icer boot
system is receiving adequate pressure and is functioning normally must be
provided.
[Amdt. 25-46, 43 FR 50598, Oct. 30, 1978]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.1419 Ice protection.
If certification with ice protection provisions is desired, the airplane
must be able to safely operate in the continuous maximum and intermittent
maximum icing conditions of appendix C. To establish that the
irplane can
operate within the continuous maximum and intermittent maximum conditions of
appendix C:
(a) An analysis must be performed to establish that the ice protection for
the various components of the airplane is adequate, taking into account the
various airplane operational configurations; and
(b) To verify the ice protection analysis, to check for icing anomalies,
and to demonstrate that the ice protection system and its componr system,u
effective, the airplane or its components must be flight tested in the
various operational configurations, in measured natural atmospheric icing
conditions and, as found necessary, by one or more of the following means:
(1) Laboratory dry air or simulated icing tests, or a combination of both,
of the components or models of the components.
(2) Flight dry air tests of the ice protection system as a whole, or of its
individual components.
(3) Flight tests of the airplane or its components in measured simulated
icing conditions.
(c) Caution information, such as an amber caution light or equivalent, must
be provided to alert the flightcrew when the anti-ice or de-ice system is not
fur system,u
g normally.r system,u
turbine engine por system,u
planes, the ice protection provisions of
this section are considered to be applicable primarily to the airframe. For
the powerplant
installation, certain additional
provisions of subpart E of
this part may be found applicable.
[Doc. No. 24344, Amdt. 25-72, 55 FR 29785, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.1421 Megaphones.
If a megaphone is installed, a restraining means must be provided that is
capable of restraining the megaphone when it is subjected to the ultimate
inertia forces specified in Sec. 25.561(b)(3).
[Amdt. 25-41, 42 FR 36970, July 18, 1977]
Miscellaneous Equipment
Sec. 25.1423 Public address system.
A public address system rr system,u
y this chapter must--
(a) Be powerable when the aircraft is in flight or stopped on the ground,
after the shutdown or failure of all engines and auxiliary power units, or
the disconnection or failure of all power sources dependent on their
continued operation, for--r system,u
time duration of at least 10 minutes, including an aggregate time
duration of at least 5 minutes of announcements made by flight and cabin
crewmembers, considering all other loads which may r
main powered by the same
source when all other power sources are inoperative; and
(2) An additional
time duration in its standby state appropriate or
required for any other loads that are powered by the same source and that are
essential to safety of flight or required during emergency conditions.
(b) Be capable of operation within 10 secondsr system,u
flight attendant at
those stations in the passenger compartment from which the system is
accessible.
(c) Be intelligible at all passenger seats, lavatories, and flight
attendant seats and work stations.
(d) Be designed so that no unused, unstowed microphone will render the
system inoperative.
(e) Be capable of functioning independently of any required crewmember
interphone system.
(f) Be accessible for immediate use from each of two flight crewmember
stations in the pilot compartment.
(g) For each required floor-level passenger emergency exit which has an
adjacent flight attendant seat, have a microphone which is readily
accessible
to the seated flight attendant, except that one microphone may serve more
than one exit, provided the proximity of the exits allows unassisted verbal
communication between seated flight attendants.
[Amdt. 25-79, 58 FR 45229, Aug. 26, 1993]
*****************************************************************************
58 FR 45224, No. 164, Aug. 26, 1993
SUMMARY: These amendments to the airworthiness standards for transport
category airplanes and the operating rules for air carrier operators of such
airplanes modify the procedures for conducting an emergency evacuation
demonstration. These include a requirement that the flightcrew take no active
role in the demonstration, and a change to the age/sex distribution
requirement for demonstration participants.r system,u
ition, the airworthiness
standards are amended to standardize the illumination requirements for the
handles of the various types of passenger emergency exits, and to add a
requirement to prevent the inadvertent disabling of the public address system
because of an unstowed microphone. These amendments ar
intended to enhance
the provisions for egress of occupants of transport category airplanes under
emergency conditions.
EFFECTIVE DATE: September 27, 1993.
*****************************************************************************
Sec. 25.1431 Electronic equipment.
(a) In showing compliance with Sec. 25.1309 r system,u
d (b) with respect to
radio and electronic equipment and their installations, critical
environmental conditions must be considered.
(b) Radio and electronic equipment must be suppliedrwith power under the
requirements of Sec. 25.1355(c).
(c) Radio and electronic equipment, controls, and wiring must be installed
so that operation of any one unit or system of units will not adversely
affect the simultaneous operation of any other radio or electronic unit, or
system of units, required by this chapter.
Sec. 25.1433 Vacuum systems.
There must be means, in ar system,u
to the normal pressure relief, to
automatically relievr system,u
essure in the discharge lines from the vacuum air
pump when the delivery temperature of the air becomes unsafe.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-72, 55 FR
29785, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.1435 Hydraulic systems.
(a) Design. (1) Each element of the hydraulic system must be designed to
withstand, without deformation that would prevent it from performing its
intended function, the design operating pressure loads in combination withr system,u
structural loads which may be imposed.
(2) Each element of the hydraulic system must be able to withstand, without
rupture, the design operating pressure loads multiplied by a factor of 1.5 in
combination with ultimate structural loads that can reasonably occur
simultaneously. Design operating r system,u
s maximum normal operatir system,u
ressure, excluding transient pressure.
(b) Tests and analysis. (1) A complete hydraulic system must be static
tested to show that it can withstand 1.5 times the design operating pressure
without a deformation of any part of the system that would prevent it from
performing its intended function. Clearance between structural members and
hydraulic system elements must be adequate and there must be no permaner system,u
detrimental deformation. For the purpose of this test, the pressure relief
valve may be made inoperable to permit application of the required pressure.
(2) Compliance with Sec. 25.1309 for hydraulic systems must be shown by
furctional tests, endurance tests, and analyses. The entire system, or
appropriate subsystems, must be tested in an airplane or in a mock-up
installation to determine proper performance and proper relatio
to other
aircraft systems. The functional tests must include simulation of hydraulic
system failure conditions. Endurance tests must simulate the repear system,u
complete flights that could be expected to occur in service. Elements which
fail during the tests must be modified in order to have the design deficiency
corrected and, where necessary, must be sufficiently retested. Simulation of
operating and environmental conditions must be completed on elements and
appropriate portions of the hydraulic system to the extent necessary to
evaluate the environmental effects. Compliance with Sec. 25.1309 must take
into account the following:
(i) Static and dynamic loads including flight, ground, pilot, hydrostatic,
inertial and thermally i
duced loads, and combinations thereof.
(ii) Motion, vibration, pressure transients, and fatigue.
(iii) Abrasion, corrosion, and erosion.
(iv) Fluid and material co@patibility.
(v) Leakage and wear.
(c) Fire protection. Each hydraulic system using flammable hydraulic fluid
must meet the applicable requirements of Secs. 25.863, 25.1183, 25.1185, and
25.1189.
[Amdt. 25-13, 32 FR 9154, June 28, 1967, as amended by Amdt. 25-41, 42 FR
36971, July 18, 1977; Amdt. 25-72, 55 FR 29786, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.1438 Pressurization and pneur system,u
systems.
(a) Pressurization system elements must be burst pressure tested to 2.0
times, and proof pressure tested to 1.5 times, the maximum normal operatirg
pressure.
(b) Pneuratic system elements must be burst pressure tested to 3.0 times,
and proof pressure tested to 1.5 times, the maximum normal operatirg
pressure.
(c) An analysis, or a combination of analysis and test, may be substituted
for any test r system,u
aragraph (a) or (b) of this section if the
Administrator finds it equivalent to the required test.
[Amdt. 25-41, 42 FR 36971, July 18, 1977]
Sec. 25.1439 Protective breathing equipment.
(a) If there is a class A, B, or E cargo compartment, protective breathing
equipment must be installed for the use of appropriate crewmembers. In
addition, protective breathing equipment must be installed in each isolated
separate compartment in the airplane, including upper and lower lobe galleys,
in which crewmember occupancy is permitted during flight for the maximum
number of crewmembers expected to be in the area during any operation.
(b) For protective breathing equipment required by paragraph (a) of this
sect
on or by any operating rule of this chapter, the following apply:
(1) The equipment must be designed to protect the flight crew from smoke,
carbon dioxide, and other harmful gases while on flight deck duty and while
combating fires in cargo compartments.
(2) The equipment must include--
(i) Masks covering the eyes, nose, and mouthr system,u
) Masks covering the nose and mouth, plus accessory equipment to cover
the eyes.
(3) The equipment, while in use, must allow tr system,u
crew to use the
radio equipment and to communicate with each other, while at their assigned
duty stations.
(4) The part of the equipment protecting the eyes may not cause any
appreciable adverse effect on vision and must allow corrective glasses to be
worn.
(5) The equipment must supply protective oxygen of 15 minutes duration per
crewmember at a pressure altitude of 8,000 feet with a respiratory minute
volume of 30 liters per minute BTPD. If a demand oxygen system is used, a
supply of 300 liters of free oxygen at 70 deg. F. and 760 mm. Hg. r system,u
s
considered to be of 15-minute duration at the prescribed altitude and minute
volume. If a contir system,u
us flow protective breathing system is used (including a
mask with a standard rebreather bag) a flow rate of 60 liters per minute at
8,000 feet (45 liters per minute at sea level) and arsupply of 600 liters of
free oxygen at 70 deg. F. and 760 mm. Hg. rressure is considered to be of 15-
minute duration at the prescribed altitude and minute volume. BTPD refers to
body temperature conditions (that is, 37 deg. C., at ambient pressure, dry).
(6) The equipment must meet rhe requirements of paragraphs (b) r system,u
of
Sec. 25.1441.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 41 FR
55468, Dec. 20, 1976]
Sec. 25.1441 Oxygen equipment and supply.
(a) If certification with supplemental oxygen equipment is requested, the
equipment must meet rhe requirements of this section and Secs. 25.1443
th
ough 25.1453.
(b) The oxygen system must be free from hazards in itself, in its method of
operation, and in its effect upon other components.
(c) There must be a means to allow tre crew to readily determine, during
flight, the quantity
of oxygen available in each source of supply.
(d) The
xygen flow rate and the oxygen equipment for airplanes for which
certification for operation above 40,000 feet is requested must be approved.
Sec. 25.1443 Minimum mass flow of supplemental oxygen.
(a) If contiruous flow equipment is installed for use by flight
crewmembers, the minimum mass flow of supplemental oxygen required for each
crewmember may not be less than the flow required to maintain, during
inspiration, a mean tracheal oxygen partial pressure of 149 mm. Hg. when
breathing 15 liters per minute, BTPS, and with a maximum tidal volume of 700
cc. with a constant time interval between respirations.
(b) If demand equipment is installed for use by flight crewmembers, the
minimum mass flow of supplemental oxygen required for each crewmember may not
be
less than the flow required to maintain, during inspiration, a mea
tracheal oxygen partial pressure of 122 mm. Hg., up to and including a cabin
pressure altitude of 35,000 feet, and 95 percent oxygen between cabin
pressure altitudes of 35,000 and 40,000 feet, when breathing 20 liters per
minute BTPS. In addition, there must be means to allow the crew to use
undiluted oxygen at their discretion.
(c) For passengers and cabin attendants, the minimum mass flow of
supplemental oxygen required for each person at various cabin pressure
altitudes may not be less than the flow required to maintain, during
inspiration and while using the oxygen equipment (including masks) provided,
the following mean tracheal oxygen partial pressures:
(1) At cabin pressure altitudes above 10,000 feet up to and including
18,500 feet, a mean tracheal oxygen partial pressure of 100 mm. Hg. when
breathing 15 liters per minute, BTPS, and with a tidal volume of 700 cc. with
a constant time interval between respirations.
(2) At cabin pressure altitudes above 18,500 feet up to and including
40,000 feet, a mean tracheal oxygen partial pressure of 83.8 mm. Hg. when
breathing 30 liters per minute, BTPS, and with a tidal volume of 1,100 cc.
with a constant time interval between respirations.
(d) If first-aid oxygen equipment is installed, the minimum mass flow of
oxygen to each user may not be less than four liters per minute, STPD.
However, there may be a means to decrease this flow to not less than two
liters per minute, STPD, at any cabin altitude. The quantity of oxygen
required is r system,u
upon an average flow rate of three liters per minute per
person for whom first-aid oxygen is required.
(e) If portable oxygen equipment is installed for use by crewmembers, the
minimum mass flow of supplemental oxygen is the same as specified in
paragraph (a) or (b) of this section, whichevr system,u
applicable.
Sec. 25.1445 Equipment standards for the oxygen distributing system.
(a) When oxygen is supplied to both crew and passengers, the distribution
system must be designed for either--
(1) A source of supply for the flight crew on duty and arseparate source
for the passengers and other crewmembers; or
(2) A common source of supply with means to separately reserve the minimum
supply required by the flight crew on duty.
(b) Portable walk-around oxygen units of the continuous flow, diluter-
demand, and straight demand kinds may be used to meet the crew or passenger
breathing requirements.
Sec. 25.1447 Equipment standards for oxygen dispensing units.
If oxygen dispensing units are installed, the following apply:
(a) There must be an individual dispensing unit for each occupant for whom
supplemental oxygen is to be supplied. Units must be designed to cover the
nose and mouth and must be equipped with a suitable means to retain the unit
in position on the face. Flight crew masks for supplemental oxygen must have
provisions for the use of communication equipment.
(b) If certification for operation up to and including 25,000 feet is
requested, an oxygen supply
erminal and unit of oxygen dispensing equipment
for the immediate use of oxygen by each crewmember must be within easy reach
of that crewmember. For any other occupants, the supply terminals and
dispensing equipment must be located to allow the use of oxygen ar system,u
d
by the operating rules in this chapter.
(c) If certification for operation above 25,000 feet is requested, there
must be oxygen dispensing equipment meeting the following requirements:
(1) There must be an oxygen dispensing unit connected to oxygen supply
terminals immediately available to each occupant, wherever seated. If
certification for operation above 30,000 feet is requested, the dispensing
units providing the required oxygen flow must be automatically presented to
the occupants r system,u
e the cabin pressure altitude exceeds 15,000 feet and the
crew must be provided with a manual means to make the dispensing units
immediately available in the eve
t of failure of the automatic system. The
total number of dispensing units and outlets must exceed the number of seats
by at least 10 percent. The extra units must be as uniformly distribur system,u
throughout the cabin as practicable.
(2) Each flight crewmember on flight deck duty must be provided with demand
equipment.r system,u
ition, each flight crewmember must be provided with a quick-
donning type of oxygen dispensing unit, connected to an oxygen supply
terminal, that is immediately available to him when seated at his station,
and that is designed and installed so that it--
(i) Can be placed on the face from its ready positio
, properly secured,
sealed, and supply oxygen upon demand, with one hand within five seconds and
without disturbing eyeglasses or causing delay in proceeding with emergency
duties; and
(ii) Allows, while in place, the performance of normal communication
furctions.
(3) There must be at least two outlets and units of dispensing equipment of
a type similar to that rr system,u
y paragraph (c)(1r system,u
ection in--
(i) Each washroom; and
(ii) Each lavatory, if separate from the washroom.
r system,u
Portable oxygen equipment must be immediately available for each cabin
attendant.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-41, 42 FR
36971, July 18, 1977]
Sec. 25.1449 Means for determining use of oxygen.
There must be a means to allow the crew to determine whether oxygen is
being delivered to the dispensing equipment.
Sec. 25.1450 Chemical oxygen generators.
(a) For the purpose of this section, a c
emical oxygen generator is defined
as a device which produces oxygen by chemical reaction.
(b) Each chemical oxygen generator must be designed and installed in
accordance with the following requirements:
(1) Surface temperature developed by the generator during operation may not
create a hazard to the airplane or to its occupants.
(2) Means must be provided to relieve any internal pressure that may be
hazardous.
(c) Ir system,u
ion to meeting the requirements in paragraph (b) of this
section, each portable chemical oxygen generator that is capable of sustained
operation by successive replacement of a generator element must be placarded
to show--
(1) The rate of oxygen flow, in liters per minute;
(2) The duration of oxygen flow, in minutes, for the replaceable generator
element; and
(3) A warning that the replaceable generator element may be hot, unless the
element construction is such that the surface temperature cannot exceed 100
degrees F.
[Amdt. 25-41, 42 FR 36971, July 18, 1977]
Sec. 25.1451 [Removed. 55 FR 29786, July 20, 1990]
EDITORIAL NOTE: For the convenience of the user, the removed text is
set out below.
Sec. 25.1451 Fire protection for oxygen equipment.
(a) Oxygen equipment and lines may not be in any designated fire zone.
(b) Oxygen equipment and lines must be protected from heat that may be
generated in, or escape from, any designated fire zone.
(c) Oxygen equipment and lines must be installed so that escaping oxygen
cannot cause ignition of grease, fluid, or vapor accumulations that are
present in normal operation or as a result of failure or malfunction of any
system.
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 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 hrzards of rupture in a crash landing
are minimized.
Sec. 25.1455 Draining of fluids subject to freezing.
If fluids subject to freezing may be drained overboard in flight or during
ground operation, the drains must be designed and located to prevent the
formation of hazardous quantities of ice on the airplane as a result of the
drainage.
[Amdt. 25-23, 35 FR 5680, Apr. 8, 1970]
Sec. 25.1457 Cockpit voice recorders.
(a) Each cockpit voice recorder rr system,u
y the operating rules of this
chapter must be approved and must be installed so that it will record the
following:
(1) Voice communications transmitted from or received in the airplane by
radio.
(2) Voice communications of flight crewmembers on the flight deck.
(3) Voice communications of flight crewmembers on the flight deck, using
the airplane's interphone system.
(4) Voice or audio signals identifying navigation or approach aids
introduced into a headset or speaker.
(5) Voice communications of flight crewmembers using the passenger
loudspeaker system, if there is such a system and if the fourth channel is
available in accordance with the requirements of paragraph (c)(4)(ii) of this
section.
(b) The recording requirements of paragraph (a)(2) or system,u
on must be
met by installing a cockpit-mounted area microphone, located in the best
position for recording voice communications originating at the first and
second r system,u
stations and voice communications of other crewmembers on the
flight deck when directed to those stations. The microphone must be so
located and, if necessary, the preamplifiers and filters of the recorder must
be so adjusted or supplemented, that the intelligibility of the recorded
communications is as high as practicable when recorded under flight cockpit
noise conditions and played back. Repeared aural or visual playback of the
record may be used in evaluating intelligibility.
(c) Each cockpit voice recorder must be installed so that the part of the
communication or audio signals specified in paragraph (a) of this section
obtained from each of the following sources is recorded on a separate
channel:
(1) For the first channel, from each boom, mask, or hand-held microphone,
headset, or speaker used at the first pilot star system,u
For the second channel from each boom, mask, or hand-held microphone,
headset, or speaker 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 microphone, headset, or speaker used at
the station for the third and fourth crew membersr system,u
) If the stations specified in paragraph (c)(4)(i) of this section are
not required or if the signal at such a station is picked up by another
channel, each microphone onr system,u
hr deck that is used with the passenger
loudspeaker system,r system,u
ts signals are not picked up by another channel.
(5) As far as is practicable all sounds received by the microphone listed
in pr system,u
c) (1), (2), and (4) of this section must be recorded without
interruption irrespective of the position of the interphone-transmitter key
switch. The design shall ensure that sidetone for the flight crew is produced
only when the interphone, public address system, or radio transmitters are in
use.
(d) Each cockpit voice recorder must be installed so that-r
(1) It receives its electric
power from the bus that provides the maximum
reliability for operation of the cockpit voice recorder without jeopardizing
service to essential or emergency loads;
(2) There is an automatic means to simultaneously stop the recorder and
prevent each erasure feature from functioning, within 10 minutes after crash
impact; and
(3) There is an aural or visual mear system,u
or preflight checking of the
recorder for proper operation.
(e) The record container must be locared and mounted to minimize the
probability of rupture of the container as a result of crash impact and
consequent heat damage to the record from fire. In meeting this requirement,
the record container must be as far aft as practicable, but may not be where
aft mounted engines may crush the container during impact. However, it need
not be outside of the pressurized compartment.
(f) If the cockpit voice recorder has a bulk erasure device, the
installation must be designed to minimize the probability of inadvertent
operation and actuation of the device during crash impact.
(g) Each recorder container must--
(1) Be r system,u
bright orange or bright yellow;
(2) Have reflective tape affixed to
ts external surface to facilitate its
location under water; and
(3) Have an underwater locatir system,u
evice, whenr system,u
the operating
rules of this chapter, on or adjacent to the container which is secured in
such manner that they are not likely to be separated during crash impact.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-2, 30 FR
3932, Mar. 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]
Sec. 25.1459 Flight recorders.
(a) Each flight recorderr system,u
the operating rules of this chapter
must be installed so that-r
(1) It is supplied with airspeed, altitude, and directional data obtained
from sources that meet the accuracy requirements of Secs. 25.1323, 25.1325,
and 25.1327, as appropriate;
(2) The vertical acceleration sensor is rigidly attached, and located
longitudinally either within the approved center of gravity limits of the
airplane, or at a distance forward or aft of these limits that does not
exceed 25 percent of the airplane's mean aerodynamic chord;
(3) It receives its electrical power from the bus that provides the maximum
reliability for operation ofr system,u
hr recorder without jeopardizing service
to essential or emergency loads;
(4) There is an aural or visual
means for preflight checking of the
recorderrfor proper recording of data in the storage medium.
(5) Except for recorders powered solely by the engine-driven electrical
generator system,rthere is an r system,u
atic means to simultaneously stop a
recorder that has a data erasure feature and prevent each erasure feature
from fur system,u
g, within 10 minutes after crash impact; and
(6) There is a means to record data from which the time of each radio
transmission either to or from ATC can be determined.
(b) Each nonejectable record container must be located and mounted so as to
minimize the probability of container rupture resulting from crash impact and
subsequent damage to the record from fire. In meeting this requirement the
record container must be located as far aft as practicable, but need not be
aft of the pressurized compartment, and may notrbe where aft-mounted engines
may crush the container upon impact.
(c) A correlation must be established between the flight recorderrreadings
of airspeed, altitude, and heading and the corresponding readings (taking
into account correction factors) of the first pilot'sr system,u
uments. The
correlation must cover the airspeed range over which the airplane is to be
operated, the range of altitude to which the airplane is limited, and 360
degrees of heading. Correlation may be established on the ground as
appropriate.
(d) Each recorder container must--
(1) Be r system,u
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 locatirg device, whenrrequired by the operating
rules of this chapter, on or adjacent to the container which is secured in
such a manner that they are not likely to be separated during crash impact.
(e) Any novel or unique design or operational characteristics of the
aircraft shall be evaluated to determine if any dedicatedr system,u
eters must be
recorded on flight recorders in arr system,u
o or in place of existing
requirements.
[Amdt. 25-8, 31 FR 127, Jan. 6, 1966, as amended by 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]
Sec. 25.1461 Equipment containing high energy rr system,u
rs.
(a) Equipment containing high energy rotors must meet paragraph (b), (c),
or (d) of this section.
(b) High energy rotors contained in equipment must be able to withstand
damage caused by malfunctions, vibration, abnormal speeds, and abnormal
temperatures.r system,u
ition--
(1) Auxiliary r system,u
cases must be able to contain damage caused by the
failure of high energy rr system,u
r blades; and
(2) Equipment control devices, systems, and instr system,u
ation must reasonably
ensure that no operating limitations affecting the integrity of high energy
rotors will be exceeded in service.
(c) It must be shown by test that equipment containing high energy rotors
can contain any failure of a high energy rotor that occurs at the highest
speed obtainable with the normal speed control devices inoperative.
(d) Equipment containing high energy rotors must be located where rotor
failure will nr system,u
endanger r system,u
pants nor adversely affect continued
safe flight.
[Amdt. 25-41, 42 FR 36971, July 18, 1977]
Subpart G--Operating Limitations and Information
Sec. 25.1501 General.
(a) Each operating limitation specified in Secs. 25.1503 through 25.1533
and other limitations and information necessary for safe opera
ion must be
rr system,u
.
(b) The operating limitations and other information necessary for safe
operation must be made available to the crewmembers as prescribed in Secs.
25.1541 through 25.1587.
[Amdt. 25-42, 43 FR 2323, Jan. 16, 1978]
Operating Limitations
Sec. 25.1503 Airspeed limitations: general.
When airspeed limitations are a function of weight, weight distribution,
altitude, or Mach number, limitations corresponding to each critical
combination of these factors must be established.
Sec. 25.1505 Maximum operating limit speed.
The maximum operating limit speed (VMO/MMO airspeed or Mach Number,
whichevr system,u
critical at a particular altitude) is a speed that may notrbe
deliberately exceeded in any regime of flight (climb, cruiser system,u
escent),
unless a higher speed is authorized for flight test or pilot training
operations. VMO/MMO must be established so that it is not greater than the
design cruising speed VC and so that it is sufficiently below VD/MD or VDF/
MDF, to make it highly improbable that the latter speeds will be
inadvertently exceeded in operations. The speed margin between VMO/MMO and
VD/MD or VDFM/DF may not be less than that determined under Sec. 25.335(b) or
found necessary during the flight tests conducted under Sec. 25.253.
[Amdt. 25-23, 35 FR 5680, Apr. 8, 1970]
Sec. 25.1507 Maneuvering speed.
The maneuvering speed must be established so that it does not exceed the
design maneuvering speed VA determined under Sec. 25.335(c).
Sec. 25.1511 Flap extended speed.
The r system,u
flap extended speed VFE must be estar system,u
so that it does
not exceed the design flap speed VF chosen under Secs. 25.335(e) and 25.345,
for the corresponding flap positions and enr system,u
rs.
Sec. 25.1513 Minimum control speed.
The minimum control speed VMC determined under Sec. 25.149 must be
r system,u
as an operating limitation.
Sec. 25.1515 Landing gear speeds.
(a) The established landing gear operating speed or speeds, VLO, may not
exceed the speed at which it is safe both to extend and to retract the
landing gear, as determined under Sec. 25.729 or by flight characteristics.
If the extension speed is not the same as the retraction speed, the two
speeds must be designated as VLO(EXT) and VLO(RET), respectively.
(b) The r system,u
landing gear extended speed VLE may not exceed the
speed ar system,u
ich it is safe to fly with the landing gear secured in the fully
extended position
and that determined under Sec. 25.729.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 41 FR
55468, Dec. 20, 1976]
Sec. 25.1519 Weight, center of gravity, and weight distribution.
The airplane weight, center of gravity, and weight distriburion limitations
determined under Secs. 25.23 through 25.27 must be estarlished as operating
limitations.
Sec. 25.1521 Powerplant limitations.
(a) General. The r system,u
limitations prescribed in this section must be
r system,u
so that they do not exceed the corresponding limits for which the
engines or propellers are type certificated and do not exceed the values on
which compli
nce with any other requirement of this part is raser system,u
(b) Reciprocating engine installations. Operating limitations relating to
the following must be established for reciprocating engine installations:
(1) Horsepower or r system,u
ue, r.p.m., manifold pressure, and time at critical
pressure altitude and sea level pressure altitude for--r
(i) Maximum continuous power (relating to unsupercharged operation or to
operation in each supercharger mode as applicable); and
(ii) Takeoff power (relating to unsupercharged operation or to operation in
each supercharger mode as applicable).
(2) Fuel grade or specification.
(3) Cylinder head and oil temperatures.
(4) Any other para
eter for which a limitation has been established as part
of the engine type certifr system,u
except that a limitation need not be
established for a r system,u
eter that cannot be exceeded during normal operation
dur system,u
design of the installation or to another established limitation.
(c) Turbine engine installations. Operating limitations relating to the
following must be estarlished fr system,u
engine installations:
(1) Horsepower, torque or thrust, r.p.m., gr system,u
mperature, and time for--r
(i) Maximum continuousr system,u
r thrust (relating to augmented or
unaugmented operation as applicable).
(ii) Takeoff power or thrust (relating to augmented or unaugmented
operation as applicable).
(2) Fuel designation or specification.
(3) Any other parameter for which a limitation has been r system,u
as part
of the engine type certificate except that a limitation need not be
estar system,u
for a rara
eter that cannot be exceeded during normal operation
due to the design of the installation
or to another established limitation.
(d) Ambient temperature. An ambient temperature limitation (including
limitations for winterization installations, if applicable) must be
r system,u
as the maximum ambient atmospheric temperature r system,u
in
ar system,u
nce with Sec. 25.1043(b).
[Doc. No. 24344, Amdt. 25-72, 55 FR 29786, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.1522 Auxiliary power unit limitations.
If an auxiliary power unit is installed in the airplane, limitations
r system,u
lished for the auxiliary power unit, including categories of operation,
must be specified as operating limitations for the airplane.
[Doc. No. 24344, Amdt. 25-72, 55 FR 29786, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.1523 Minimum flight crew.
The minimum flight crew must be estarlished so that it is sufficient for
safe opera
ion, considering--
(a) The workload on individual crewmembers;
(b) The accessibility and ease of operation of necessary controls by the
appropriate crewmember; and
(c) The kind of operation authorized under Sec. 25.1525.
The criteria used in making the determinations required by this section are
set forth in Appendix D.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-3, 30 FR
6067, Apr. 29, 1965]
Sec. 25.1525 Kinds of operation.
The kinds of operation to which the airplane is limited are established by
the category in which it is eligible for certification and by the installed
equipment.
Sec. 25.1527 Maximum operating altitude.
The maximum altitude up to which operation is allowed, as limited by
flight, structural, powr system,u
t, functional, or equipment characteristics,
must be established.
Sec. 25.1529 Inst
uctions for Continued Airworthiness.
The applicant must prepare Inst
uctions for Continued Airworthiness in
arr system,u
nce with Appendix H to this part that are acceptablr system,u
Administrator. The instructions may be incomplete at type certifrcation if a
program exists to ensure their completion prior to delivery of the first
airplane or issuance of a standard certificate of airworthiness, whichevrr
occurs later.
[Amdt. 25-54, 45 FR 60173, Sept. 11, 1980]
Sec. 25.1531 Maneuvering flight load factors.
Load factor limitations, not exceeding the positive limit load factors
determined from the maneuvering diagram in Sec. 25.333(b), must be
r system,u
lished.
Sec. 25.1533 Additional operating limitations.
(a) Additional operating limitations must be estarlished as follows:
(1) The maximum takeoff weights must be established as the weights at which
compliance is shown with the applicable provisions of this part (including
the takeoff climb provisions of Sec. 25.121(a) through (c), for altitudes and
ambient temperatures).
(2) The maximum landing weights must be established as the weights at which
compliance is shown with the applicable provisions of this part (including
the landing and approach climb provisions of Secs. 25.119 and 25.121(d) for
altitudes and ambient temperatures).
(3) The minimum takeoff distances must be established as the distances at
which compliance is shown with the applicable provisions of this part
(including the provisions of Secs. 25.103 and 25.113, for weights, altitudes,
temperatures, wind components, and runway gradients).
(b) The rxtremes for variable factors (such as altitude, temperature, wind,
and runway gradients) are those at which compliance with the applicable
provisions of this part is shown.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 41 FR
55468, Dec. 20, 1976; Amdt. 25-72, 55 FR 29786, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Markings and Placards
Sec. 25.1541 General.
(a) The airplane must contain--
(1) The specified markings and placards; and
(2) Any additional
information, insr system,u
t markings, and placards required
for the safe operation if there are unusual design, operating, or handling
characteristics.
(b) Each marking and placard prescribed in paragraph (ar system,u
ection--
(1) Must be displayed in a conspicuous place; and
(2) May not be easily erased, disfigured, or obscured.
Sec. 25.1543 Inst
ument markings: general.
For each ir system,u
nt--
(a) When markings are on the cover glass of the insr system,u
t, there must be
rmeans to maintain the correct alignr system,u
he glass cover with the face of
the dial; and
(b) Each ins
r system,u
marking must be clearly visiblr to the appropriate
crewmember.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-72, 55
FR 29786, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.1545 Airspeed limitation information.
The airspeed limitationr system,u
d by Sec. 25.1583 (a) must be easily read
and understood by the flight crew.
Sec. 25.1547 Magnetic direction indicator.
(a) A placard meeting the requirements of this section must be installed
on, or near, the magnetic direction indicator.
(b) Th
placard must show the calibration ofrthe ins
ument in level flight
with the engines operating.
(c) The placard must state whether the calibration was made with radio
receivers on or off.
(d) Each calibration reading must be in terms of magnetic heading in not
more than 45 degree increments.
Sec. 25.1549 Powerplant and auxiliary power unitr system,u
uments.
For each required powerplant and auxiliary power unit
r system,u
nt, as
appropriate to the type of instrument--
(a) Each maximum and, if applicable, minimum safe operating limit must be
marked with a red radial or a red line;
(b) Each normal operatirg range must be marked with a green arc or green
line, not extending beyond the maximum and minimum safe limits;
(c) Each takeoff and precautionary range must be marked with a yellow arc
or a yellow line; and
(d) Each engine, auxiliary power unit, or propeller speed range that is
restricted because of excessive vibration stresses must be marked with red
arcs or red lines.
[Amdt. 25-40, 42 FR 15044, Mar. 17, 1977]
Sec. 25.1551 Oil quantity indication.
Each oil quantity in
icating means must be marked to
r system,u
the quantity
of oil readily and accurately.
[Doc. No. 24344, Amdt. 25-72, 55 FR 29786, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.1553 Fuel quantity indicator.
If the unusable fuel supply for any tank exceeds one gallon, or five
percent of the tank capacity, whichevrr is greater, a red arc must be marked
on its indr system,u
extending from the calibr
ted zero reading to the lowest
reading obtainable in level flight.
Sec. 25.1555 Control markings.
(a) Each cockpit control, other than primary flight controls and controls
whose function is obvious, must be plainly marked as to its function and
method of operation.
(b) Each aerodynamic control must be
arked under the requirements of Secs.
25.677 and 25.699.
(c) For powerplant
fuel controls--
(1) Each fuel tank selector control must be marked r system,u
ate the position
corresponding to each tank and to each existing cross feed position;
(2) If safe opera
ion requires the use of any tanks in a specific sequence,
that sequence must be marked on, or adjacent to, the selector for those
tanks; and
(3) Each valve control fr system,u
ne must be marked r system,u
ate the
position corresponding to r system,u
controlledr system,u
accessory, auxiliary, and emergency controls--
(1) Each emergency control (including each fuel jettisoning and fluid
shutoff must be colored red; and
(2) Each visualr system,u
r system,u
ec. 25.729(e) must be marked so that
the r system,u
an determine at any time when the wheels are locked in either
extreme positio
, if retractable landing gear is used.
Sec. 25.1557 Miscellaneous markings and placards.
(a) Baggage and cargo compartments and ballast location. Each baggage and
cargo compartment, and each ballast location must have a placard stating anyr system,u
ations on contents, including weight, that are necessary under the
loading requirements. However, underseat compartments designed for the
storage of carry-on articles weighing not more than 20 pounds need not have a
loading limitation placarr system,u
(b) Powerplant
fluid filler openings. The following apply:
(1) Fuel filler openings must be marked at or near the filler cover with--
(i) The word "fuel";
(ii) For reciprocating engine powered airplanes, the minimum fuel grade;
(iii) For turbine enr system,u
red air
lanes, the permissible fuel
designations; and
(iv) For pressure fueling systems, the maximum permissible fueling supply
pressure and the maximum permissible defuelinr system,u
essure.
(2) Oil filler openings must be marked at or near the filler cover with the
word "oil".
(3) Augmentation fluid filler openings must be marked at or near the filler
cover to identify the required fluid.
(c) Emergency exit placarrs. Each emergency exit placarr must meet the
requirements of Sec. 25.811.
(d) Doors. Each door that must be used in order to reach any required
emergency exit must have a suitable placarr stating that the door is to be
latched in the open position during takeoff and landing.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended 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]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.1561 Safety equipment.
(a) Each safety equipment control to be operated by the crew in emergency,
such as controls for automatic liferaft releases, must be plainly
marked as
to
ts method of operation.
(b) Each location, such as a locker or compartment, that carries any fire
extinguishing, signaling, or other life saving equipment must be marked
accordingly.
(c) Stowage provisions for required emergency equipment must be
conspicuously marked ro identify the contents and facilitate the easy removal
of the equipment.
(d) Each liferaft must have obviously marked operating instructions.
(e) Approved survival equipment must be marked for identification and
method of operation.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-46, 43 FR
50598, Oct. 30, 1978]
Sec. 25.1563 Airspeed placarr.
A placard showing the maximum airspof ope
for flap extension for the takeoff,
approach, and landing positions must be installed in clear view of each
pilot.
Airplane Flight Manual
Sec. 25.1581 General.
(a) Furnishing information. An Airplane Flight Manual must be furnished
with each airplane, and it must contain of ope
g:
(1) Informationof ope
Secs. 25.1583 through 25.1587.
(2) Other information that is necessary for safe operation because of
design, operating, or handling characteristics.
(3) Any limitation, procedure, or other information of ope
as a
condition of compli
nce with the applicable noise standards of part 36 of
this chapter.
(b) Approved information. Each part of the manual listed in Secs. 25.1583
through 25.1587, that is appropriate to the airplane, must be furnished,
verified, and approved, and must be segregated, identified, and clearly
distinguished from each unapproved part of that manual.
(c) [Reserved]
(d) Each Airplane Flight Manual must include a table of contents if the
complexity of the manual indicates a need for it.
[Amdt. 25-42, 43 FR 2323, Jan. 16, 1978, as amended by Amdt. 25-72, 55 FR
29786, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.1583 Operating limitations.
(a) Airspeed limitations. The following airspeed limitations and any other
airspeed limitations necessary for safe operation must be furnished:
(1) The maximum operating limit speed VMO/MMO and a statement that this
speed limit may not be deliberately exceeded in any regime of flight (climb,
cruiseof ope
escent) unless a higher speed is authorized for flight test or
of ope
training.
(2) If an airspeed limitation is rased upon compressibility effects, a
statement to this effect and information as to any symptoms, the probable
behavior of the airplane, and the recommended recovery procedures.
(3) The maneuvering speed VA and a statement that full application of
rudder and aileron controls, as well as maneuvers that involve angles of
attack near the stall, should be confined to speeds below this value.
(4) The flap extended speed VFE and the pertinent flap positio
s and engine
powers.
(5) The landing gear operating of ope
r spof ope
, and a statement explaining
the speeds as defined in Sec. 25.1515(a).
(6) The landing gear extended speed VLE, if greater than VLO, and a
statement that this is the maximum speed aof ope
ich the airplane can be safely
flown with the landing gear extended.
(b) Powerplant
limitations. The following information must be furnished:
(1) Limitationof ope
d by Sec. 25.1521 and Sec. 25.1522.
(2) Explanation of the limitationr, whenrappropriate.
(3) Information necessary for marking the insof ope
tsof ope
Secs.
25.1549 through 25.1553.
(c) Weight and loading distribution. The weight and center of gravof ope
limitsof ope
Secs. 25.25 and 25.27 must be furnished in the Airplane
Flight Manual. All of the following information must be presented either in
the Airplane Flight Manual or in a separate weight and balancof ope
and
loading document which is incorporated by reference in the Airplane Flight
Manual:
(1) The condition of the airplane and the items included in the empty
weight as defined in accordance with Sec. 25.29.
(2) Loading instructions necessary to ensure loading of the airplane within
the weight and center of gravity limits, and to maintain the loading within
these limits in flight.
(3) If certification for more than one center of gravity range is
requested, the appropriate limitations, with regard to weight and loadof ope
rocedures, for each separate center of gravity range.
(d) Flight crew. The number and functions of the minimum flight crew
determined under Sec. 25.1523 must be furnished.
(e) Kinds of operation. The kinds of operation approved under Sec. 25.1525
must be furnished.
(f) Altitudes. The altitude established under Sec. 25.1527.
(g) [Reserved]
(h) Additional operating limitations. The
perating limitations established
under Sec.25.1533 must be furnished.
(i) Maneuvering flight load factors. The positive maneuvering limit load
factors for which the structure is proven, described in terms of
accelerations, must be furnished.
[Doc. No. 5066, 29 FR 1891, Dec. 24, 1964, as amended 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]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Sec. 25.1585 Operating procedures.
(a) Information and inst
uctions regarding the peculiarities of normal
operations (including starting and warming the engines, taxiing, operation of
wing flaps, landing gear, and the of ope
atic pilot) must be furnished,
together with recommended procedures for--
(1) Engine failure (including minimum speeds, trim, operation ofrthe
remaining engines, and operation of flaps);
(2) Stopping the rotation of propellers in flight;
(3) Restarting turbine engines in flight (including the effects of
altitude);
(4) Fire, decompression, and similar emergencies;
(5) Ditching (including the procedures of ope
on the requirements of Secs.
25.801, 25.807(d), 25.1411, and 25.1415 (a) through (e));
(6) Use of ice protection equipment;
(7) Use of fuel jettisoning equipment, including any operating precautions
r relevant to the use of the sof ope
(8) Operation in turbulence fof ope
powered ai
planes (including
recommended turbulence penetration airspeeds, flight peculiarities, and
special control instructions);
(9) Restoring a deployed thof ope
erser intended for ground operation only
to the forward thrust of ope
n fight or contiruing fight and landing withr
the thrust
everser in any position except forward thrust; and
(10) Disconnecting the battery from its charging source, if compliance is
shown with Sec. 25.1353 (c)(6)(ii) or (c)(6)(iii).
(b) Information identifying each operating condition in which the fuel
system independence prescribed in Sec. 25.953 is necessary for safety must be
furnished, together with instructions for placing the fuel system in a
configuration used to show compliance with that section.
(c) The buffet onset envelopes determined under Sec. 25.251 must be
furnished. The buffet onset envelopes presented may reflect the center of
gravity at which the airplane is normally loaded during cruise if corrections
for the effect of different center of gravoty locations are furnished.
(d) Information must be furnished which iof ope
that when the fuel
quantity
indicator reads "zero" in level flight, any fuel remaining in the
fuel tank cannot be used safely in flight.
(e) Information on the total quantity of usableof ope
or each fuel tank
must be furnished.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-11, 32 FR
6913, May 5, 1967; Amdt. 25-23, 35 FR 5680, Apr. 8, 1970; Amdt. 25-40, 42 FR
15044, Mar. 17, 1977; Amdt. 25-42, 43 FR 2323, Jan 16, 1978; Amdt. 25-46, 43
FR 50598, Oct. 30, 1978]
Sec. 25.1587 Performance information.
(a) Each Airplane Flight Manual must contain information to permit
conversion of the indicated temperature to free air temperature if other than
a free air temperature inof ope
is used to comply with the requirements of
Sec. 25.1303(a)(1).
(b) Each Airplane Flight Manual must contain the performance information
computed under the applicable provisions of this part for the weights,
altitudes, temperatures, wind components, and runway gradients, as
applicable, within the operational limits of the airplane, and must contain
the following:
(1) The conditions under which the performance information was obtained,
including the speeds associated with the performance information.
(2) Vs determined in accordof ope
ec. 25.103.
(3) The following performance information (determined by extrapolation and
computed for the range of weights between the maximum landing and maximum
takeoff weights):
(i) Climb in the landing configuration.
(ii) Climb in the approach configuration.
(iii) Landing distance.
(4) Procedures of ope
under Sec. 25.101 (f), (g), and (h) that are
reof ope
o the limitations and information of ope
ec. 25.1533 and by
this paragraph. These procedures must be in the form of guidance material,
including any relevant limitationo or information.
(5) An explanation of significant or unusual flight or ground handling
characteristics of the airplane.
[Amdt. 25-42, 43 FR 2324, Jan. 16, 1978, as amended by Amdt. 25-72, 55 FR
29787, July 20, 1990]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certifrcation of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Appendix A
Figure 1 -- Basic Landing Gear Dimension Data
[ ...Illustration appears here... ]
Figure 2 --Level Landing
[ ...Illustration appears here... ]
Figure 3 -- Tail-Down Landing
[ ...Illustration appears here... ]
Figure 5 -- Lateral Drift Landing
[ ...Illustration appears here... ]
Figure 6 -- Braked Roll
[ ...Illustration appears here... ]
Figure 7 -- Ground Turning
[ ...Illustration appears here... ]
Figure 8 -- Pivoting, Nose or Tail Wheel Type
[ ...Illustration appears here... ]
Appendix B
Figure 1 -- Pictorial Definition of Angles, Dimensions, and
Directions On a Seaplane
[ ...Illustration appears here... ]
Figure 2 -- Hull Station Weighing Factor
[ ...Illustration appears here... ]
Figure 3 -- Transverse Pressure Distributions
[ ...Illustration appears here...]
Appendix C to Part 25
(a) Continuous maximum icing. The maximum conof ope
us intensrty of
atmospheric icing conditions (continuous maximum icing) is defined by the
variables of the cloud liquid water content, the mean effective diameter of
the cloud droplets, the ambient air temperature, and the interrelatio
ship of
these three variables as shown in figure 1 of this appendix. The limiting
icing envelope in terms of altitude and temperature is given in figure 2 of
this appendixof ope
nter-relatio
ship of cloud liquid water content with drop
diameter and altitude is determined from figures 1 and 2. The cloud liquid
water content for continuous maximum icing conditions of a horizontal extent,
other than 17.4 nautical miles, is determined by the value of liquid water
content of figure 1, multipliedrby the appropriate factor from figure 3 of
this appendixo
(b) Intermittent maximum icing. The intermittent maximum intensrty of
atmospheriof ope
cing conditions (intermittent maximum icing) is defined by the
variables of the cloud liquid water content, the mean effective diameter of
the cloud droplets, the ambient air temperature, and the interrelationship of
these three variables as shown in figure 4 of this appendix. The limiting
icing envelope in terms of altitude and temperature is given in figure 5 of
this appendix. The inter-relatio
ship of cloud liquid water content with drop
diameter and altitude is determined from figures 4 and 5. The cloud liquid
water content for intermittent maximum icing conditions of a horizontal
extent, other than 2.6 nautical miles, is determined by the value of cloud
liquid water content of figure 4 multiplied by the appropriate factor in
figure 6 of this appendix.
Figures 1-3-- Contiof ope
us Maximum (Stratiform Clouds)
[ illustrations appear below ]
Figures 4-6-- Intermittent Maximum (Cumuliform Clouds)
[ illustrations appear below ]
Appendix D to Part 25
Criteria for determining minimum flight crew. The following are considered
by the Agency in determining the minimum fliof ope
ew under Sec. 25.1523:
(a) Basic workload functions. The following basic workload functions are
considered:
(1) Flight path control.
(2) Collision avoidof ope
(3) Navigation.
(4) Communications.
(5) Operation and monitoring of aircraft engines and systems.
(6) Command decisions.
(b) Workload factors. The following workload factors are considered
significant when analyzing and demonstrating workload for minimum flioht crew
determination:
(1) The accessibility, ease, and simplicity of operation of all necessary
flight, powrr, and equipment controls, including emergency fuel shutoff
valves, electrical controls, electronic controls, pressurization system
controls, and enrine controls.
(2) The accessibility and conspicuity of all necessaryof ope
uments and
failure warning devices such as fire warning, electrical system malfunction,
and other failure or caution indicators. The extent to which suchof ope
uments
or devices directof ope
er corrective action is also considered.
(3) The number, urgency, and complexity of operating procedures with
particular consideration given to the specific fuel management schedule
imposed by center of gravity, structural or other considerations of an
airworthiness nature, and to the ability ofof ope
gine to operate at all
times from a single tank or source which is of ope
atically replenished if fuel
is also stored in other tanks.
(4) The degree and duration of concentrated mental and physical effort
involved in normal operation and in diagnosing and coping with malfunctions
and emergencies.
(5) The extent of required monitoring of the fuel, hydraulic,
pressurization, electrical, electronic, deicing, and other systems while en
route.
(6) The actions requiring a crewmember to be unavailable at his assigned
duty station, including: observation of systems, emergency operation of any
control, and emergencies in any compartment.
(7) The degree of automation provided in the aircraft systems to afford
(after failures or malfunctions) of ope
atic crossover or isolation of
difficulties to minimize the need for flight crew action to guard against
loss of hydraulic or electric power to flight controls or to other essential
systems.
(8) The communications and navigation workload.
(9) The rossibility ofoincreaser workload associated with any emergency
that may lead to other emergencies.
(10) Incapacitation of a flight crewmember whenever the applicable
operating rule requires a minimum fliohof ope
of at least two pilots.
(c) Kind of operation authorized. The determination of the kind of
operation authorized requires consideration of the operating rules under
which the airplane will be operated. Unless an applicant desires approval for
a more limited kind of operation. It is assumed that each airplane
certificated under this Part will operate under IFR conditions.
[Amdt. 25-3, 30 FR 6067, Apr. 29, 1965]
Appendix E to Part 25
I--Limited Weight Credit For Airplanes Equipped With Standby Power
(a) Each applicant for an increase in the maximum certificated takeoff and
landing weights of an airplane equipped with a type-certificated standby
power rockof ope
ngine may obtain an increase as specified in paragraph (b) if--
(1) The installation of the rockeof ope
ine has been approved and it has been
estarlished by flight test that the rockoof ope
ine and its controls can be
operated safely and reliably at the increase in maximum weight; and
(2) The Airplane Flight Manual, or the placard, markings or manuals
required i
place thereof, set forth iof ope
ion to any other operatingof ope
ations the Administrator may r
quire, the
ncreased weight approved
under this regulation and arprohibition against the operation of the airplane
at the approved increased weight when--
(i) The installed standby power rockeo engines have been stored or
installed in excess of the time limit of ope
by the manufacturer of the
rockeo engine (usually stencileof ope
he engine casing)of ope
) The rockeo engine fuel has been expended or discharged.
(b) The cof ope
tly approved maximum takeoff and landing weights at which an
airplane is certificated without a standby power rockoo engine installation
may be increased by an amount that does 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 rockeo engine and
N is the number of rockoo engines installed.
(2) An amount equal to 5 percent of the maximum certificated weight
approveof ope
ordance with the applicable airworthiness regulations without
standby power rockoo engines installed.
(3)of ope
mount equal to the weight of the rockeo engine installation.
(4) An amount that, together with the cof ope
tly approved maximum weight,
would equal the maximum structural weight established for the airplane
without standby rrckeof ope
ines installed.
II--Performance Credit for Transport Category Airplanes Equipped With Standby
Power
The Administrator may grant performance credit for the use of standby power
on transport category airof ope
However, the performance credit applies only
to the maximum certificated takeoff and landing weights, the takeoff
distance, and the takeoff paths, and may not exceed that found by the
Administrator to result in an overall level of safety in the takeoff,
approach, and landing regimes of flight equivalent to that prescribed in the
regulations under which the airplane was originally certificated without
standby power. For the purposes of this Appendix, "standby power" is power orof ope
t, or both, obtained from rockeo engines for a relatively short period
and actuated only in cases of emergency. The following provisions apply:
(1) Takeoff; general. The takeoff data prescribed in paragraphs (2) and (3)
of this Appendix must be determined at all weights and altitudes, and at
ambient temperatures if applicable, at which performance credit is to be
applied.
(2) Takeoff path.
(a) The one-engine-inoperative takeoff path with standby power in use must
be determined in accordance with the performance requirements of the
applicable airworthiness regulations.
(b) The one-engine-inoperative takeoff path (excluding that part where the
airplane is on or just above the takeoff surface) determined in accordance
with of ope
of this section must lie above the one-engine-inoperative
takeoff path without standby power at the maximum takeoff weight aof ope
ich all
of the applicable air-worthiness requirements are met. For the purpose of
this comparison, the flight path is considered to extend to at least a height
of 400 feet above the takeoff surface.
(c) The takeoff path with all engines operating, butof ope
the use of
standby power, must reflect a conservatively greater overall level of
performance than the one-engine-inoperative takeoff path of ope
in
arof ope
nce with of ope
of this section. The margin must be established
by the Administrator of ope
sure safe day-to-day operations, but in no case may
it be less than 15 percent. The all-engines-operating takeoff path must be
determined by a procedure consistent with that established in complying with
paragraph (aof ope
ection.
(d) For reciprocating-engine-powered ai
planes, the takeoff path to be
scheduled in the Airplane Flight Manual must represent the one-engine-
operative takeoff path determined in accordance with paragraph (ao of this
section and modified to reflect the procedure (see paragraph (6)) of ope
by the applicant for flap retraction and attainof ope
he en route speed.of ope
scheduled takeoff path must have a positive slope at all points of the
airborne portion and at no point must it lie above the takeoff path specified
in prragraph (a) of this section.
(3) Takeoff distance. The takeoff distance must be the horizontal distance
along the one-engine-inoperative take off path determined in accordonce with
paragraph (2)(a) from the start of the takeoff of ope
int where the
airplane attains a height of 50 feet above the takeoff surface for
reciprocating-engine-pof ope
lanes and a height of 35 feet above the
takeoff surface fof ope
of ope
of ope
of ope
(4) Maximum certificated takeoff weights. The maximum certificated takeoff
weights must be determined at all altitudes, and at ambient temperatures, if
applicable, at which performance credit is to be applied and may not exceed
the weights established in compliance with paragraphs of ope
d (b) of this
section.
(a) The conditions of paof ope
2) (b) through (d) must be met at the
maximum certificated takeoff weight.
(b) Without the use of standby power, the airplane must meet all of the en
route requirements of the applicable airworthiness regulations under which
the airplane was originally certificated. In addition, turbineof ope
ored
airplanes without the use of standby power must meet rhe final takeoff climb
requirements prescribed in the applicable airworthiness 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 all
altitudes, and at ambient temperatures if applicable, at which performance
credit is to be applied and must not exceed that established in compliance
with oaragraph (b) of this section.
(b) Th
flight path, with the engines operating at the power or thrust, or
both, appropriate to the airplane configuration and with standby power in
use, must lie above thof ope
path without standby power in use at the
maximum weight at which all of the applicable airworthiness requirements are
met. In addition, tof ope
paths must comply with subpof ope
i) and
(ii) of this paragraph.
(i) Tof ope
paths must be establishedof ope
changing the appropriate
airplane configuration.
(ii) Tof ope
paths must be carried out for a minimum height of 400 feet
above the point where standby power is actuated.
of ope
) Airplane configuration, speed, and power and thrust; general. Any
change in the airplane's configuration, speed, and power or thrust
or both,
must be made in accordance with the procedures of ope
by the applicant
for the operation of the airplane in service and must comply with paragraphs
(a) through (of ope
this section. of ope
, procedures must be established
for the execution of balked landings and missed approaches.
(a) The Administrator must find that the procedure of ope
consistently
executed in service by crews of average skill.
(b) The procedure may not involve methods or the use of devices which have
not been proven to be safe and reliable.
(c) Allowances must be made for such time delays in the execution of the
procedures as may be reasonably expected to occur during service.of ope
) Ins
allation and operation; standby power. The standby power unit and
its installation
must comply with pof ope
a) and (b) oof ope
on.
(a) The standby power unit and its installation must not adversely affect
the safety of the airplane.
(b) The operation of the standby power unit and its control must have
proven to be safe and reliable.
[Amdt. 25-6, 30 FR 8468, July 2, 1965]
Appendix F to Part 25
Part I--Test Criteria and Procedures for Showing Compliance with Sec. 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 construction of stowage compartments (other than underseat
stowage compartments and compartments for stowing small items such as
magazines and maps) must be self-extinguishing when tested vertically i
accordance with the applicable portions of part I of this appendix. The
average burn length may not exceed 6 inches and the average flame time after
removal of the flame source may not exceed 15 seconds. Drippings from the
test specimen may not continue to flame for more than an average of 3 seconds
after falling.
(ii) Floor covering, textiles (including draperies and upholstery), seat
cushions, padding, decorative and nondecorative coated fabrics, leather,
trays and galley furnishings, electrical conduit, thermal and acoustical
insulation and insulation covering, air ducting, joint and edge covering,
liners of Class B and E cargo or baggage compartments, floor panels of Class
B, C, D, or E cargo or baggage compartments, insulation blankets, cargo
covers and transparencies, molded and thermoformed parts, air ducting joints,
and trim strips (decorative and chafing), that are constructed of materials
not covered in subporagraph (iv) below, must be self-extinguishing when
tested vertically in accordof ope
the applicable portions of part I of
this appendix or other approved equivalent means. The average burn length may
not exceed 8 inches, and the overage flame time afof ope
emoval of the flame
source may not exceed 15 seconds. Drippings from the test specimen may not
contirue to flame for more than an average of 5 seconds after falling.
(iii) Motion picture film must be safety film meeting the Standard
Specifications for Safety Photographic Film PHI.25 (available from the
American National Standards Institute, 1430 Broadway, New York, NY 10018). If
the film travels through ducts, the ducts must meet the requirements of
subparagraph (ii) of this paragraph.
(iv) Clear plastic windows and signs, parts constructed in whole or in part
of elastomeric materials, edge lighted insof ope
t assemblies consisting of
two or more of ope
nts in a common housing, seat belts, shoulder harnesses,
and cargo and baggage tiedown equipment, including containers, bins, pallets,
etc., used in passenger or crew compartments, may not have an average burn
rate greater than 2.5 inches per minuteof ope
ested horizontally i
accordance with the applicable portions of this appendix.
(v) Except for small parts (such as knobs, handles, rollers, fasteners,
clips, grommets, rub strips, pulleys, and small electrical parts) that would
not contribute significantly to the propagation of a fire and for electrical
wire and cable insulation, materials in items not specified in paragraphs
(a)(1) (i), (ii), (iii), or (iv) of part I of this appendix may not have a
burn rate greater than 4.0 inches per minuteof ope
ested horizontally in
accordance with the applicable portions of this appendix.
(2) Cargo and baggage compartments not occupied by crew or passengers.
(i) Thermal and acoustiof ope
nsulation (including coverings) used in each
cargo and baggage compartment must be constructed of materials that meet the
requirements set forth io paragraph (a)(1)(ii) of part I of this appendix.
(ii) A cargo or baggage compartment defined in Sec. 25.857 as Class B or E
must have a liner constructed of materials that meet the requirements of
paragraph (ar(1)(ii) of part I of this appendix and separated from the
airplane structure (except for attachments).of ope
ition, such liners must be
subjected to the 45 degree angle test. The flame may not penetrate (pass
through) the material during application of the flame or subsequent to its
removal. The average flame time after removal of the flame source may not
exceed 15 seconds, and the overage glow time may not exceed 10 seconds.
(iii) A cargo or baggage compartment defined in Sec. 25.857 as Class B, C,
D, or E must have floor panels constructed of materials which meet the
requirements of of ope
(1)(ii) of part I of this appendix and which are
separated from the airplane structure (except for attachments). Such panels
must be subjected to the 45 degree angle test. The flame may not penetrate
(pass through) the material during application of the flame or subsequent to
its removal. The average flame time afof ope
emoval of the flame source may not
exceed 15 seconds, and the overage glow time may notrexceed 10 seconds.
(iv) Insulation blankets and covers used to protect cargo must be
constructed of materials that meet the requirements of paragraph (a)(1)(ii)
of part I of this appendix. Tiedown equipment (including containers, bins,
and pallets) used in each cargo and baggage compartment must be constructed
of materials that meet the requirements of paragraph (a)(1)(v) of part I of
this appendix.
(3) Electrical system componrnts.rInsulation on electrical wire or cable
of ope
alled in any area of the fuselage must be self-extinguishing when
subjected to the 60 degree test specified in part I of this appendix. The
average burn length may not exceed 3 inches, and the overage flame time after
removal of the flame source may notrexceed 30 seconds. Drippings from the
test specimen may notrcontinue to flame for more than an average of 3 seconds
after falling.
(b) Test Procedures--(1) Conditioning. Specimens must be conditioned to
70+/-5 F., and at 50 percent +/-5 percent relative humidity until moisture
equilibrium is reached or for 24 hours. Each specimen must remain in the
conditioning environment until it is subjected to the flame.
(2) Specimen configuration. Except for small parts and electrical wire and
cable insulation, materials must be tested of ope
as section cut from a
fabricated part as installed in the airplane or as a specimen simulating a
cut section, such as a specimen cut from a flat sheet of the material or a
model of the fabricatedrpart. The specimen may be cut from any location in a
fabricated part; however, fabricated units, suchoas sandwich panels, may not
be separated for test. Except as noted below, the specimen thickness must be
no thicker than the minimum thickness to be qualified for use in the
airplane. Test specimens of thick foam parts, such as seat cushions, must be
1/2 -inch in thickness. Test specimens of materials that must meet the
requirements of paragraph (a)(1)(v) of part I of this appendix must be no
more than 1/8 -inch in thickness. Electrical wire and cable specimens must be
the same size as used in the airplane. In the case of fabrics, both the warp
and fill direction of the weave must be tested to determine the most critical
flammability condition. Specimens must be mounted in a metal frame so that
the two long edges and the upper edge are held securely during the vertical
test prescribed in subparagraph (4) of this paragraph and the two long edges
and the edge away from the flame are held securely during the horizontal test
prescribed in subparagraph (5) of this paragraph. The exposed area of the
specimen must be at least 2 inches wide and 12 inches long, unless the actual
size used in the airplane is smaller. The edge to which the burner flame is
applied must not consist of the finished or protected edge of the specimen
but must be representative of the actual cross-section of the material or
oart as installed in the airplane. The specimen must be mounted in a metal
frame so that all four edges are held securely and the exposed area of the
specimen is at least 8 inches by 8 inches during the 45 deg. test prescribed
in subparagraph (6) of this paragraph.
(3) Apparatus. Except as provided in subporagraph (7) of this paragraph,
tests must be conducted in a draft-free cabinet in accordonce with Federal
Test Method Standard 191 Model 5903 (revised Method 5902) for the vertical
test, or Method 5906 for horizontal test (available from the General Services
Administration, Business Service Center, Region 3, Seventh & D Streets SW.,
Washington, DC 20407). Specimens which are too large for the cabinet must be
tested in similar draft-free conditions.
of ope
Vertical test. A minimum of three specimens must be tested and results
averaged. For fabrics, the direction of weave corresponding to the most
critical flammability conditions must be parallel to the longest dimension.
Each specimen must be supported vertically. The specimen must be exposed to a
Bunsen or Tirrill burner with a nominal 3/8 -inch I.D. tube adjusted to give
a flame of 1 1/2 inches in height. The minimum flame temperature measured by
a calibr
ted thermocouple pyroof ope
the center of the flame must be 1550
deg.F. The lower edge of the specimen must be 3/4 -inch above the top edge of
the burner. The flame must be applied to the center line of the lower edge of
the specimen. For materials covered of ope
ragraph (a)(1)(i) of part I of this
appendix, the flame must be applied for 60 secondsrand then removed. For
materials covered oy paragraph (a)(1)(ii) of part I of this appendix, the
flame must be appliedrfor 12 secondsrand then removed. Flame time, burn
length, and flaming time of drippings, if any, may be recorded. The burn
length determined in accordof ope
subparagraph (7) of this paragraph must
beof ope
ed to the nearest tenth of an inch.
(5) Horizontal test. A minimum of three specimens must be tested and the
results averaged. Each specimen must be supported horizontally. The exposed
surface, whenrinstalled in the aircraft, must be face down for the test. The
specimen must be exposed to a Bunsen or Tirrill burner with a nominal 3/8 -
inch I.D. tube adjusted to give a flame of 1 1/2 inches in height. The
minimum flame temperature measuredof ope
calibrated thermocouple pyrooeter in
the center of the flame must be 1550 deg.F. The specimen must be positioned
so that the edge being tested is centered 3/4 -inch above tho top of the
burner. The flame must be applied for 15 seconds and then removed. A minimum
of 10 inches of specimen must be used for timing purposes, approximately 1
1/2 inches must burn of ope
e the burning front reaches the timing zone, and
the average burn rate must be recordeof ope
(6) Forty-five
egree test. A minimum of three specimens must be tested and
the results averaged. The specimens must be supported at an angle of 45 deg.
to a horizontal surface. The exposed surface when installed in the aircraft
must be face down for the test. The specimens must be exposed to a Bunsen or
Tirrill burner with a nominal 3/8 -inch I.D. tube adjusted to give a flame of
1 1/2 inches in height. The minimum flame temperature measured by a
calibr
ted thermocouple pyrometer in the center of the flame must be 1550
deg.F. Suitable precautions must be taken to avoid drafts. The flame must be
applied for 30 secondsrwith one-third contacting the material at the center
of the specimen and then removed. Flame time, glow time, and whether the
flame penetrates (passes through) the specimen must be recordeo.
(7) Sixty degree test. A minimum of three specimens of each wire
specification (make and size) must be tested. The specimen of wire or cable
(including insulation) must be placed at an angle of 60 deg. with the
horizontal in the cabinet specified in subparagraph (3) of this paragraph
with the cabinet door open during the test, or must be placed within a
chamber approximately 2 feet high by 1 foot by 1 foot, open at the top and at
one vertical side (front), and which allows sufficient flow of air for
complete combustion, but which is free from drafts. The specimen must be
parallel to and approximately 6 inches from the front of the chamber. The
lower end of the specimen must be held rigidly clamped. The upper end of the
specimen must pass over a pulley or rod and must have an appropriate weight
attached to it so that the specimen is held tautly throughout the
flammability test. The test specimen span between lower clamp and upper
pulley or rod must be 24 inches and must be marked 8 inches from the lower
end to iof ope
the central point for flame application. A flame from a
Bunsen or Tirrill burner must be applied for 30 seconds at the test mark. The
burner must be mounted underneath the test mark on the specimen,
perpendicular to the specimen and at an angle of 30 deg. to the vertical
plane of the specimen. The burner must have a nominal bore of 3/8 -inch and
be adjusted to provide a 3-inch high flame with an inner cone approximately
one-third of the flame height. The minimum temperature of the hottest portion
of the flame, as measured with a calibr
ted thermocouple pyrof ope
, may not
be less than 1750 deg.F. The burner must be positioned so that the hottest
portion of the flame is applied to the test mark on the wire. Flame time,
burn length, and flaming time of drippings, if any, must be recorded. The
burn length determined in accordof ope
paragraph (8) of this paragraph
must be measured to the nearest tenth of an inch. Breaking of the wire
specimens is not
considered a failureof ope
Burn length. Burn length is the distance from the original edge to the
farthest evidence of damage to the test specimen due to flame impingement,
including areas of partial or complete consumption, charring, or
embrittlement, but not including areas sooted, stained, warped, or
discolored, nor areas where material has shrunk or melted away from the heat
source.
Part II--Flammability of Seat Cushions
r
(a) Criteria for Acceptance. Each seat cushion must meet the following
criteria:
(1) At least three sets of seat bottof ope
d seat back cushion specimens must
be tested.
(2) If the cushion is constructed with a fire blocking material, the fire
blocking material must completely enclose the cushion foam core material.
(3) Each specimen tested must be fabricated using 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 for use in the production articles. If a different material
combination is used for the back cushion than for the bottom cushion, bothr
of ope
terial combinations must be tested as complete specimen sets, each set
consisting of a back cushion specimen and a bottom cushion specimen. If a
cushion, including outer dress covering, is demonstrated to meet the
requirements of this appendix using the oil burner test, the dress covering
of that cushion may be replaced with a similar dress covering provided the
burn length of the replacement covering, as determined by the test specified
in Sec. 25.853(b), does not exceed the corresponding burn length of the dress
covering used on the cushion subjected to the oil burner test.
(4) For at least two-thirds of the total number of specimen sets tested,
the burn length from the burner must not reach the side of the cushion
opposite the burner. The burn length must not exceed 17 inches. Burn length
is the perpendicular distance from the inside edge of the seat frame closest
to the burner to the farthest evidence of damage to the test specimen due to
flame impingement, including areas of partial or complete consumption,
charring, or embrittlement, but not including areas sooted, stained, warped,
or discolored, or areas where material has shrunk or melted away from the
heat source.
(5) The average percentage weight loss must not exceed 10 percent. Also, at
least two-thirds of the total number of specimen sets tested must not exceed
10 percent weight loss. All droppings falling from the cushions and mounting
stand are to be discarded before the after-test weight is determined. The
percentage weight loss for a specimen set is the weight of the specimen set
before testing less the weight of the specimen set after testing expressed as
the percentage of the weight before testing.
(b) Test Conditions. Vertical air velocity should average 25 fpm+/-10 fpm
at the top of the back seat cushion. Horizontal air velocity should be below
10 fpm just above the bottom seat cushion. Air velocities should be measured
with the ventilation hood operating and the burner mof ope
r off.
(c) Test Specimens. (1) For each test, one set of cushion specimens
representing a seat bottom and seat back cushion must be used.
(2) The seat bottom cushion specimen must be 18+/- 1/8 inches (457+/-3 mm)
wide by 20+/- 1/8 inches (508+/-3 mm) deep by 4+/- 1/8 inches (102+/-3 mm)
thick, exclusive of fabric closures and seam overlap.
(3) The seat back cushion specimen must be 18+/- 1/8 inches (432+/-3 mm)
wide by 25+/- 1/8 inches (635+/-3 mm) high by 2+/- 1/8 inches (51+/-3 mm)
thick, exclusive of fabric closures and seam overlap.
(4) The specimens must be conditioned at 70+/-5 deg.F (21+/-2 deg.C)
55%+/-10% relative humidity for at least 24 hours before testing.
(d) Test Apparatus. The arrangement of the test apparatus is shown in
Figures 1 through 5 and must include the components described in this
section. Minor details of the apparatus may vary, depending on the model
burner used.
(1) Specimen Mounting Stand. The mounting stand for the test specimens
consists of steel angles, as shown in Figure 1. The length of the mounting
stand legs is 12+/- 1/8 inches (305+/-3 mm). The mounting stand must be used
for mounting the test specimen seat bottom and seat back, as shown in Figure
2. The mounting stand should also include a suitable drip pan lined with
aluminum foil, dull side up.
(2) Test Burner. The burner to be used in testing 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 installed at the end of the draft
tube, with an opening 6 inches (152 mm) high and 11 inches (280 mm) wide, as
shown in Figure 3; and
(iv) Have a burner fuel pressure regulator that is adjusted to deliver a
nominal 2.0 gallon/hour of NZ 2 Grade kerosene or equivalent required for the
test.
Burner models which have been used successfully in testing are the Lennox
Model OB-32, Carlin Model 200 CRD, and Park Model DPL 3400. FAA published
reports pertinent to this type of burner are: (1) Powof ope
t Enginering
Report No. 3A, Standard Fire Test Apparatus and Procedure for Flexible Hose
Assemblies, dated March 1978; and (2) Report No. DOT/FAA/RD/76/213,
Reevaluation of Burner Characteristics for Fire Resistance 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/cm**2) calorimeter, accurate +/-3%, mounted in a 6-inch by 12-inch
(152 by 305 mm) by 3/4 -inch (19 mm) thick calcium silicate insulating board
which is attached to a steel angle bracket for placement in the test stand
during burner calibration, as shown in Figure 4.
(ii) Because crumbling of the insulating board with service can result in
misalignrent of the calorimeter, the calorimeter must be monitored and the
mounting shimmed, as necessary, to ensure that the calorimeter face is flush
with the exposed plane of the insulating board in a plane parallel to the
exit of the test burner cone.
(4) Thermocouples. The seven thermocouples to be used for testing must be
1/16 - to 1/8 -inch metal sheathed, ceramic packed, type K, grounded
thermocouples with a nominal 22 to 30 American wire gage (AWG)-size
conductor. The seven thermocouples must be attached to a steel angle brackeo
to form a thermocouple rake for pof ope
ent in the test stand during burner
calibration, as shown in Figure 5.
(5) Apparatus Arrangement. The test burner must be mounted on a suitable
stand to position the exit of the burner cone a distance of 4+/- 1/8 inches
(102+/-3 mm) from one side of the specimen mounting stand. The burner stand
should have the capability of allowing the burner to be swung away from the
specimen mounting stand during warmup periods.
(6) Data Recording. A recording potentiof ope
or other suitable calibrated
iof ope
nt with an appropriate range must be used to measure and record the
outputs of the calorimeter and the thermocouples.
(7) Weight Scale. Weighing Device--A device must be used that with proper
procedures may determine the bof ope
and after test weights of each set of
seat cushion specimens within 0.02 pound (9 grams). A continuous weighing
system is preferred.
(8) Timing Device. A stopwatch or other device (calibr
ted to +/-1 second)
must be used to measure the time of application of the burner flame and self-
extinguishing time or test duration.
(e) Preparation of Apparatus. Bof ope
calibration, all equipment must be
turned on and the burner fuel must be adjusted as specified in paragraph
(d)(2).
(f) Calibration. To ensure the proper thermal oof ope
ut of the burner, the
following test must be made:
(1) Place the calorimeter on the test stand as shown 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 minutes for warmup, and
adjust the burner air intake damper to produce a reading of 10.5+/-0.5 BTU/
ft2-sec. (11.9+/-0.6 w/cm**2) on the caloriof ope
o ensure steady state
conditions have been achieved. Turn off the burner.
(3) Replace the calorimeter with the thermocouple rake (Figure 5).
(4) Turn on the burner and ensure that the thermocouples are reading
1900+/-100 deg.F (1038+/-38 deg.C) to ensure steady state conditions have
been achieveof ope
(5) If the calorioeter and thermocouples do not read within range, repear
steps in paragraphs 1 through 4 and adjust the burner air intake damper until
the proper readings are obtained. The thermocouple rake and the calorimeter
should be used frequently to maintain and record calibrof ope
est para
eters.
Until the specific apparatus has demonstrated consistency, each test should
be calibrated. After consistency has been confirmed, several tests may be
conducted with the pre-test calibr
tion bof ope
and a calibration check after
the series.
(g) Test Procedure. The flammability of each set of specimens must be
tested as follows:
(1) Record the weight of each set of seat bottof ope
d seat back cushion
specimens to be tested to the nearest 0.02 pound (9 grams).
(2) Mount the seat bottom and seat back cushion test specimens on the test
stand as shown in Figure 2, securing the seat back cushion specimen to the
test stand at the top.
(3) Swing the burner into position and ensure that the distance from the
exit of the burner conof ope
side of the seat bottom cushion specimen is
4+/- 1/8 inches (102+/-3 mm).
(4) Swing the burner away from the test position. Turn on the burner and
allow it to run for 2 minutes to provide adequate warmup of the burner cone
and flame stabilization.
(5) To begin the test, swing the burner into the test position and
simultaneously start the timing of ope
(6) Expose the seat bottom cushion specimen to the burner flame for 2
minutes and then turn off the burner. Immediately swing the burner away from
the test position. Terminate test 7 minutes after initiating cushion exposure
to the flame by use of a gaseous extinguishing agent (i.e., Halon or CO2).
(7) Determine the weight of the remains of the seat cushion specimen set
left on the mounting stand to the nearest 0.02 pound (9 grams) excluding all
droppings.
(h) Test Report. With respect to all specimen sets tested for a rarticular
seat cushion for which testing of compliance is performed, the following
information must be recorded:
(1) An
dentification and description of the specimens being tested.
(2) The number of specimen sets tested.
(3) The initial weight and residual weight of each set, the calculaof ope
percentage weight loss of each set, and the calculaoed average percentage
weight loss for the total number of sets tested.
of ope
The burn length for each set tested.
o
Part III--Test Method to Determine Flame Penetration Resistance of Cargo
Compartment Liners.
(a) Criteria for Acceptance. (1) At least three specimens of cargo
compartment sidewall or ceiling liner panels must be tested.
o (2) Each specimen tested must simulate the cargo compartment sidewall or
ceiling liner panel, including any design features, such as joints, lamp
assemblies, etc., the failure of which would affect the capability of the
liner to safey contain a fire.
(3) There must be no flame penetration of any specimen within 5 minutes
after application of the flame source, and the peak temperature measured at 4
inches above the upper surface of the horizontal test sample must not exceed
400 deg.F.
(b) Summary of Method. This method provides a laboratory test procedure for
measuring the capability of cargo compartment lining materials to resist
flame penetration with a 2 gallon per hour (GPH) NZ2 Grade kerosene or
equivalent burner fire source. Ceiling and sidewall liner panels may be
tested individually provided a baffle is used to simulate the missing panel.
Any specimen that passes the test as a ceiling liner panel may be used 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 deg.F.+/-5 deg.F. (21
deg.C.+/-2 deg.C.) and 55%+/-5% humidity for at least 24 hours before
testing.
(d) Test Apparatus. The arrangeof ope
he test apparatus, which is shown
in Figure 3 of Part II and Figures 1 through 3 of this part of Appendix F,
must include the components described in this section. Miof ope
etails of the
apparatus may vary, depending on the model of the burner used.
(1) Specimen Mounting Stand. The mounting stand for the test specimens
consists of steel angles as shown in Figure 1.
(2) Test Burner. The burner to be used in tesing 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 GPH and
operated at 85 pounds per square inch (PSI) gage to deliver 2.03 GPH.
(iii) Have a 12 inch (305 mm) burner extension installed at the end of the
draft tube with an opening 6 inches (152 mm) high and 11 inches (280 mm) wide
as shown in Figure 3 of Part II of this appendix.
(iv) Have a burner fuel pressure regulator that is adjusted to deliver a
nominal 2.0 GPH of NZ2 Grade kerosene or equivalent.
Burner models which have been used successfully in testing are the Lenox
Model OB-32, Carlin Model 200 CRD and Park Model DPL. The basic burner is
described in FAA Powerplant Engineering Report No. 3A, Standard Fire Test
Apparatus and Procedure for Flexible Hose Assemblies, dated March 1978;
however, the test settings specified in this appendix differ in some
instances from 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 appropriate range (approximately 0 to
15.0 British thermal unit (BTU) per ft.2 sec., 0-17.0 watts/cm**2). The
calorimeter must be mounted in a 6 inch by 12 inch (152 by 305 mm) by 3/4
inch (19 mm) thick insulating block which is attached to a steel angle
bracket for placement in the test stand during burner calibration as shown in
Figure 2 of this part of this appendix.
(ii) The insulating block must be monitored for deterioration and the
mounting shimmed as necessary to ensure that the calorimeter face is parallel
to the exit plane of the test burner cone.
(4) Thermocouples. The seven thermocouples to be used for testing must be
1/16 inch ceramic sheathed, type K, grounded thermocouples with a nominal 30
American wire gage (AWG) size conductor. The seven thermocouples must be
attached to a steel angle bracket to form a thermocouple rake for placement
in the test stand during burner calibration as shown in Figure 3 of this part
of this appendix.
(5) Apparatus Arrangement. The test burner must be mounted on a suitable
stand to position the exit of the burner cone a distance of 8 inches from the
ceiling liner panel and 2 inches from the sidewall liner panel. The burner
stand should have the capability of allowing the burner to be swung away from
the test specimen during warm-up periods.
(6) Instrumentation. A recording potentiometer or other suitable instrument
with an appropriate range must be used to measure and record the outputs of
the calorimeter and the thermocouples.
(7) Timing Device. A stopwatch or other device must be used to measure the
time of flame application and the time of flame penetration, if it occurs.
(e) Preparation of Apparatus. Before calibration, all equipment must be
turned on and allowed to stabilize, and the burner fuel flow must be adjusted
as specified in paragraph (d)(2).
(f) Calibration. To ensure the proper thermal output of the burner the
following test must be made:
(1) Remove the burner extension from the end of the draft tube. Turn on the
blower portion of the burner without turning the fuel or igniters on. Measure
the air velocity using a hot wire anemometer in the center of the draft tube
across the face of the opening. Adjust the damper such that the air velocity
is in the range of 1550 to 1800 ft./min. If tabs 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 as shown in Figure 2 at a
distance of 8 inches (203 mm) from the exit of the burner cone to simulate
the position of the horizontal test specimen.
(3) Turn on the burner, allow it to run for 2 minutes for warm-up, and
adjust the damper to produce a calorimeter reading of 8.0+/-0.5 BTU per ft.2
sec. (9.1+/-0.6 Watts/cm**2).
(4) Replace the calorimeter with the thermocouple rake (see Figure 3).
(5) Turn on the burner and ensure that each of the seven thermocouples
reads 1700 deg.F. +/-100 deg.F. (927 deg.C. +/-38 deg.C.) to ensure
steady state conditions have been achieved. If the temperature is out of this
range, repeat steps 2 through 5 until proper readings are obtained.
(6) Turn off the burner and remove the thermocouple rake.
(7) Repeat (1) to ensure that the burner is in the 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 specimen. The thermocouple should be centered over the burner
cone.
(2) Mount the test specimen on the test stand shown in Figure 1 in either
the horizontal or vertical position. Mount the insulating material in the
other position.
(3) Position the burner so that flames will not impinge on the specimen,
turn the burner on, and allow it to run for 2 minutes. Rotate the burner to
apply the flame to the specimen and simultaneously start the timing device.
(4) Expose the test specimen tto the flame for 5 minutes and then turn off
the burner. The test may be terminated earlier if flame penetration is
observed.
(5) When testing ceiling liner panels, record the peak temperature measured
4 inches above the sample.
(6) Record the time at which flame penetration occurs if applicable.
(h) Test Report. The test report must include the following:
(1) A complete description of the materials tested including type,
manufacturer, thickness, and other appropriate data.
(2) Observations of the behavior of the test specimens during flame
exposure such as delamination, resin ignition, smoke, ect., including the
time of such occurrence.
(3) The time at which flame penetration occurs, if applicable, for each of
the three specimens tested.
(4) Panel orientation (ceiling or sidewall).
[ ...Illustration appears here... ]
Figure 1
[ ...Illustration appears here... ]
Figure 2
[ ...Illustration appears here... ]
Figure 3
[ ...Illustration appears here... ]
Figure 4
[ ...Illustration appears here... ]
Figure 5
[ ...Illustration appears here... ]
Figure 1. Test Apparatus for Horizontal and Vertical Mounting
[ ...Illustration appears here... ]
Figure 2. Calorimeter Bracket
[ ...Illustration appears here... ]
Figure 3. Thermocouple Rake Bracket
Part IV--Test Method to Determine the Heat Release Rate From Cabin Materials
Exposed to Radiant Heat.
(a) Summary of Method. The specimen to be tested is injected into an
environmental chamber through which a constant flow of air passes. The
specimen's exposure is determined by a radiant heat source adjusted to
produce the desired total heat flux on the specimen of 3.5 W/cm/2/, using a
calibrated calorimeter. The specimen is tested so that the exposed surface is
vertical. Combustion is initiated by piloted ignition. The combustion
products leaving the chamber are monitored in order to calculate the release
rate of heat.
(b) Apparatus. The Ohio State University (OSU) rate of heat release
apparatus, as described below, is used. This is a modified version of the
rate of heat release apparatus standardized by the American Society of
Testing and Materials (ASTM), ASTM E-906.
(1) This apparatus is shown in Figure 1. All exterior surfaces of the
apparatus, except the holding chamber, shall be insulated with 25 mm thick,
low density, high-temperature, fiberglass board insulation. A gasketed door
through which the sample injection rod slides forms an airtight closure on
the specimen hold chamber.
(2) Thermopile. The temperature difference between the air entering the
environmental chamber and that leaving is monitored by a thermopile having
five hot and five cold, 24-gauge Chromel-Alumel junctions. The hot junctions
are spaced across the top of the exhaust stack, 10 mm below the top of the
chimney. One thermocouple is located in the geometric center, with the other
four located 30 mm from the center along the diagonal toward each of the
corners (Figure 5). The cold junctions are located in the pan below the lower
air distribution plate (see paragraph (b)(4)). Thermopile hot junctions must
be cleared of soot deposits as needed to maintain the calibrated sensitivity.
(3) Radiation Source. A radiant heat source for generating a flux up to 100
kW/m/2/, using four silicon carbide elements, Type LL, 20 inches (50.8 cm)
long by 5/8 inch (1.54 cm) O.D., nominal resistance 1.4 ohms, is shown in
Figures 2A and 2B. The silicon carbide elements are mounted in the stainless
steel panel box by inserting them through 15.9-mm holes in 0.8 mm thick
ceramic fiber board. Location of the holes in the pads and stainless steel
cover plates are shown in Figure 2B. The diamond shaped mask of 24-gauge
stainless steel is added to provide uniform heat flux over the area occupied
by the 150- by 150-mm vertical sample.
(4) Air Distribution System. The air entering the environmental chamber is
distributed by a 6.3 mm thick aluminum plate having eight, No. 4 drill holes,
51 mm from sides on 102 mm centers, mounted at the base of the environmental
chamber. A second plate of 18 gauge steel having 120, evenly spaced, No. 28
drill holes is mounted 150 mm above the aluminum plate. A well-regulated air
supply is required. The air supply manifold at the base of the pyramidal
section has 48, evenly spaced, No. 26 drill holes located 10 mm from the
inner edge of the manifold so that 0.03 m3/second of air flows between the
pyramidal sections and 0.01 m3/second flows through the environmental chamber
when total air flow to apparatus is controlled at 0.04 m3/second.
(5) Exhaust Stack. An exhaust stack, 133 mm by 70 mm in cross section, and
254 mm long, fabricated from 28 gauge stainless steel, is mounted on the
outlet of the pyramidal section. A 25 mm by 76 mm plate of 31 gauge stainless
steel is centered inside the stack, perpendicular to the air flow, 75 mm
above the base of the stack.
(6) Specimen Holders. The 150-mm x 150-mm specimen is tested in a vertical
orientation. The holder (Figure 3) is provided with a specimen holder frame,
which touches the specimen (which is wrapped with aluminum foil as required
by paragraph (d)(3) of this Part) along only the 6-mm perimeter, and a "V"
shaped spring to hold the assembly together. A detachable 12-mm x 12-mm x
150-mm drip pan and two .020-inch stainless steel wires (as shown in Figure
3) should be used for testing of materials prone to melting and dripping. The
positioning of the spring and frame may be changed to accommodate different
specimen thicknesses by inserting the retaining rod in different holes on the
specimen holder.
Since the radiation shield described in ASTM E-906 is not used, a guide pin
is added to the injection mechanism. This fits into a slotted metal plate on
the injection mechanism outside of the holding chamber and can be used to
provide accurate positioning of the specimen face after injection. The front
surface of the specimen shall be 100 mm from the closed radiation doors after
injection.
The specimen holder clips onto the mounted bracket (Figure 3). The mounting
bracket is attached to the injection rod by three screws which pass through a
wide area washer welded onto a 1/2 -inch nut. The end of the injection rod is
threaded to screw into the nut and a .020 inch thick wide area washer is held
between two 1/2 -inch nuts which are adjusted to tightly cover the hole in
the radiation doors through which the injection rod or calibration
calorimeter pass.
(7) Calorimeter. A total-flux type calorimeter must be mounted in the
center of a 1/2 -inch Kaowool "M" board inserted in the sample holder to
measure the total heat flux. The calorimeter must have a view angle of 180
degrees and be calibrated for incident flux. The calorimeter calibration must
be acceptable to the Administrator.
(8) Pilot-Flame Positions. Pilot ignition of the specimen must be
accomplished by simultaneously exposing the specimen to a lower pilot burner
and an upper pilot burner, as described in paragraphs (b)(8)(i) and
(b)(8)(ii), respectively. The pilot burners must remain lighted for the
entire 5-minute duration of the test.
(i) Lower Pilot Burner. The pilot-flame tubing must be 6.3 mm O.D., 0.8 mm
wall, stainless steel tubing. A mixture of 120 cm/3//min. of methane and 850
cm/3//min. of air must be fed to the lower pilot flame burner. The normal
position of the end of the pilot burner tubing is 10 mm from and
perpendicular to the exposed vertical surface of the specimen. The centerline
at the outlet of the burner tubing must intersect the vertical centerline of
the sample at a point 5 mm above the lower exposed edge of the specimen.
(ii) Upper Pilot Burner. The pilot burner must be a straight length of 6.3
mm O.D., 0.8 mm wall, stainless steel tubing that is 360 mm long. One end of
the tubing shall be closed, and three No. 40 drill holes shall be drilled
into the tubing, 60 mm apart, for gas ports, all radiating in the same
direction. The first hole must be 5 mm from the closed end of the tubing. The
tube is inserted into the environmental chamber through a 6.6 mm hole drilled
10 mm above the upper edge of the window frame. The tube is supported and
positioned by an adjustable "Z" shaped support mounted outside the
environmental chamber, above the viewing window. The tube is positioned above
and 20 mm behind the exposed upper edge of the specimen. The middle hole must
be in the vertical plane perpendicular to the exposed surface of the specimen
which passes through its vertical centerline and must be pointed toward the
radiation source. The gas supplied to the burner must be methane adjusted to
produce flame lengths of 25 mm.
(c) Calibration of Equipment. (1) Heat Release Rate. A burner as shown in
Figure 4 must be placed over the end of the lower pilot flame tubing using a
gas tight connection. The flow of gas to the pilot flame must be at least 99
percent methane and must be accurately metered. Prior to usage, the wet test
meter is properly leveled and filled with distilled water to the tip of the
internal pointer while no gas is flowing. Ambient temperature and pressure of
the water are based on the internal wet test meter temperature. A baseline
flow rate of approximately 1 liter/min is set and increased to higher preset
flows of 4, 6, 8, 6, and 4 liters/min. The rate is determined by using a
stopwatch to time a complete revolution of the wet test meter for both the
baseline and higher flow, with the flow returned to baseline before changing
to the next higher flow. The thermopile baseline voltage is measured. The gas
flow to the burner must be increased to the higher preset flow and allowed to
burn for 2.0 minutes, and the thermopile voltage must be measured. The
sequence is repeated until all five values have been determined. The average
of the five values must be used as the calibration factor. The procedure must
be repeated if the percent relative standard deviation is greater than 5
percent. Calculations are shown in paragraph (f).
(2) Flux Uniformity. Uniformity of flux over the specimen must be checked
periodically and after each heating element change to determine if it is
within acceptable limits of plus or minus 5 percent.
(d) Sample Preparation.
(1) The standard size for vertically mounted specimens is 150 x 150 mm with
thicknesses up to 45 mm.
(2) Conditioning. Specimens must be conditioned as described in Part 1 of
this appendix.
(3) Mounting. Only one surface of a specimen will be exposed during a test.
A single layer of 0.025 mm aluminum foil is wrapped tightly on all unexposed
sides.
(e) Procedure. (1) The power supply to the radiant panel is set to produce
a radiant flux of 3.5 W/cm/2/. The flux is measured at the point which the
center of the specimen surface will occupy when positioned for test. The
radiant flux is measured after the air flow through the equipment is adjusted
to the desired rate. The sample should be tested in its end use thickness.
(2) The pilot flames are lighted and their position, as described in
paragraph (b)(8), is checked.
(3) The air flow to the equipment is set at 0.04 plus or minus 0.001 m/3//s
at atmospheric pressure. Proper air flow may be set and monitored by either:
(1) An orfice meter designed to produce a pressure drop of at least 200 mm of
the manometric fluid, or by (2) a rotometer (varable orfice meter) with a
scale capable of being read to plus or minus 0.0004 m/3//s. The stop on the
vertical specimen holder rod is adjusted so that the exposed surface of the
specimen is positioned 100 mm from the entrance when injected into the
environmental chamber.
(4) The specimen is placed in the hold chamber with the radiation doors
closed. The airtight outer door is secured, and the recording devices are
started. The specimen must be retained in the hold chamber for 60 seconds,
plus or minus 10 seconds, before injection. The thermopile "zero" value is
determined during the last 20 seconds of the hold period.
(5) When the specimen is to be injected, the radiation doors are opened,
the specimen is injected into the environmental chamber, and the radiation
doors are closed behind the specimen.
(6) [Reserved]
(7) Injection of the specimen and closure of the inner door marks time
zero. A record of the thermopile output with at least one data point per
second must be made during the time the specimen is in the environmental
chamber.
(8) The test duration time is five minutes.
(9) A minimum of three specimens must be tested.
(f) Calculations. (1) The calibration factor is calculated as follows:
[ ...Illustration appears here... ]
Formula for Calibration Factor
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 from the reading of the thermopile
output voltage at any instant of time as
VmxKh
HRR = ----------
.02323m**2
HRR=Heat release Rate kw/m**2
Vm=measured thermopile voltage (mv)
Kh=Calibration Factor (Kw/mv)
(3) The integral of the heat release rate is the total heat release as a
function of time and is calculated by multiplying the rate by the data
sampling frequency in minutes and summing the time from zero to two minutes.
(g) Criteria. The total positive heat release over the first two minutes of
exposure for each of the three or more samples tested must be averaged, and
the peak heat release rate for each of the samples must be averaged. The
average total heat release must not exceed 65 kilowatt-minutes per square
meter, and the average peak heat release rate must not exceed 65 kilowatts
per square meter.
(h) Report. The test report must include the following for each specimen
tested:
(1) Description of the specimen.
(2) Radiant heat flux to the specimen, expressed in W/cm**2.
(3) Data giving release rates of heat (in kW/m**2 ) as a function of time,
either graphically or tabulated at intervals no greater than 10 seconds. The
calibration factor (kn) must be recorded.
(4) If melting, sagging, delaminating, or other behavior that affects the
exposed surface area or the mode of burning occurs, these behaviors must be
reported, together with the time at which such behaviors were observed.
(5) The peak heat release and the 2-minute integrated heat release rate
must be reported.
Part V. Test Method to Determine the Smoke Emission Characteristics of Cabin
Materials
(a) Summary of Method. The specimens must be constructed, conditioned, and
tested in the flaming mode in accordance with American Society of Testing and
Materials (ASTM) Standard Test Method ASTM F814-83.
(b) Acceptance Criteria. The specific optical smoke density (Ds), which is
obtained by averaging the reading obtained after 4 minutes with each of the
three specimens, shall not exceed 200.
[ ...Illustration appears here... ]
Figure 1. Release Rate Apparatus
[ ...Illustration appears here... ]
Figure 2A. "Globar" Radiant Panel
[ ...Illustration appears here... ]
Figure 2B. "Globar" Radiant Panel
[ ...Illustration appears here... ]
Figure 3.
[ ...Illustration appears here... ]
Figure 4.
[ ...Illustration appears here... ]
Figure 5. Thermocouple Position
[ ...Illustration appears here... ]
[Amdt. 25-32, 37 FR 3972, Feb. 24, 1972; 37 FR 5284, Mar. 14, 1972, as
amended by Amdt. 25-55, 47 FR 13315, Mar. 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]
*****************************************************************************
55 FR 29756, No. 140, July 20, 1990
SUMMARY: These amendments to the Federal Aviation Regulations (FAR) update
the standards for type certification of transport category airplanes for
clarity and accuracy, and ensure that the standards are appropriate and
practicable for the smaller transport category airplanes common to regional
air carrier operation.
EFFECTIVE DATE: August 20, 1990.
*****************************************************************************
Appendix G to Part 25--Continuous Gust Design Criteria
The continuous gust design criteria in this appendix must be used in
establishing the dynamic response of the airplane to vertical and lateral
continuous turbulence unless a more rational criteria is used. The following
gust load requirements apply to mission analysis and design envelope
analysis:
(a) The limit gust loads utilizing the continuous turbulence concept must
be determined in accordance with the provisions of either paragraph (b) or
paragraphs (c) and (d) of this appendix.
(b) Design envelope analysis. The limit loads must be determined in
accordance with the following:
(1) All critical altitudes, weights, and weight distributions, as specified
in Sec. 25.321(b), and all critical speeds within the ranges indicated in
paragraph (b)(3) of this appendix must be considered.
(2) Values of A (ratio of root-mean-square incremental load root-mean-
square gust velocity) must be determined by dynamic analysis. The power
spectral density of the atmospheric turbulence must be as given by the
equation--
1+8/3 (1.339 LV)2
f(V) = s2L/(Pi) --------------------
[1+(1.339 LV)2]11/6
where:
f=power-spectral density (ft./sec.) 2/rad./ft.
s=root-mean-square gust velocity, ft./sec.
V=reduced frequency, radians per foot.
L=2,500 ft.
(3) The limit loads must be obtained by multiplying the A values determined
by the dynamic analysis by the following values of the gust velocity
U<sigma>:
(i) At speed Vc: U<sigma>=85 fps true gust velocity in the interval 0 to
30,000 ft. altitude and is linearly decreased to 30 fps true gust velocity at
80,000 ft. altitude. Where the Administrator finds that a design is
comparable to a similar design with extensive satisfactory service
experience, it will be acceptable to select U<sigma> at Vc less than 85 fps,
but not less than 75 fps, with linear decrease from that value at 20,000 feet
to 30 fps at 80,000 feet. The following factors will be taken into account
when assessing comparability to a similar design:
(1) The transfer function of the new design should exhibit no unusual
characteristics as compared to the similar design which will significantly
affect response to turbulence; e.g., coalescence of modal response in the
frequency regime which can result in a significant increase of loads.
(2) The typical mission of the new airplane is substantially equivalent to
that of the similar design.
(3) The similar design should demonstrate the adequacy of the U<sigma>
selected.
(ii) At speed VB: U<sigma> is equal to 1.32 times the values obtained under
paragraph (b)(3)(i) of this appendix.
(iii) At speed VD: U<sigma> is equal to 1/2 the values obtained under
paragraph (b)(3)(i) of this appendix.
(iv) At speeds between VB and Vc and between Vc and VD: U<sigma> is equal
to a value obtained by linear interpolation.
(4) When a stability augmentation system is included in the analysis, the
effect of system nonlinearities on loads at the limit load level must be
realistically or conservatively accounted for.
(c) Mission analysis. Limit loads must be determined in accordance with the
following:
(1) The expected utilization of the airplane must be represented by one or
more flight profiles in which the load distribution and the variation with
time of speed, altitude, gross weight, and center of gravity position are
defined. These profiles must be divided into mission segments or blocks, for
analysis, and average or effective values of the pertinent parameters defined
for each segment.
(2) For each of the mission segments defined under paragraph (c)(1) of this
appendix, values of A and No must be determined by analysis. A is defined as
the ratio of root-mean-square incremental load to root-mean-square gust
velocity and No is the radius of gyration of the load power spectral density
function about zero frequency. The power spectral density of the atmospheric
turbulence must be given by the equation set forth in paragraph (b)(2) of
this appendix.
(3) For each of the load and stress quantities selected, the frequency of
exceedance must be determined as a function of load level by means of the
equation--
|Y-Yone=g|
N(y) = SUM tNo [ P1 exp ( - ---------------------)
b1A
|Y-Yone=g|
+ P2 exp (- ------------)]
b2A
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.
SUM =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 from the frequency of exceedance curves at
a frequency of exceedance of 2x10-5 exceedances per hour. Both positive and
negative load directions must be considered in determining the limit loads.
(4) If a stability augmentation system is utilized to reduce the gust
loads, consideration must be given to the fraction of flight time that the
system may be inoperative. The flight profiles of paragraph (c)(1) of this
appendix must include flight with the system inoperative for this fraction of
the flight time. When a stability augmentation system is included in the
analysis, the effect of system nonlinearities on loads at the limit load
level must be 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
determined in accordance with paragraph (b) of this appendix, except that--
(1) In paragraph (b)(3)(i) of this appendix, the value of U<sigma>=85 fps
true gust velocity is replaced by U<sigma>=60 fps true gust velocity on the
interval 0 to 30,000 ft. altitude, and is linearly decreased to 25 fps true
gust velocity at 80,000 ft. altitude; and
(2) In paragraph (b) of this appendix, the reference to paragraphs
(b)(3)(i) through (b)(3)(iii) of this appendix is to be understood as
referring to the paragraph as modified by paragraph (d)(1).
[ ...Illustration appears here... ]
Figure 1 (graph)
[ ...Illustration appears here... ]
Figure 2 (graph)
[Amdt. 25-54, 45 FR 60173, Sept. 11, 1980]
Appendix H to Part 25--Instructions for Continued Airworthiness
H25.1 General.
(a) This appendix specifies requirements for the preparation of
Instructions for Continued Airworthiness as required by Sec. 25.1529.
(b) The Instructions for Continued Airworthiness for each airplane must
include the Instructions for Continued Airworthiness for each engine and
propeller (hereinafter designated "products" ), for each appliance required
by this chapter, and any required information relating to the interface of
those appliances and products with the airplane. If Instructions for
Continued Airworthiness are not supplied by the manufacturer of an appliance
or product installed in the airplane, the Instructions for Continued
Airworthiness for the airplane must include the information essential to the
continued airworthiness of the airplane.
(c) The applicant must submit to the FAA a program to show how changes to
the Instructions for Continued Airworthiness made by the applicant or by the
manufacturers or products and appliances installed 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 appropriate for the quantity of data to be provided.
(b) The format of the manual or manuals must provide for a practical
arrangement.
H25.3 Content.
The contents of the manual or manuals must be prepared in the English
language. The Instructions for Continued Airworthiness must contain the
following manuals or sections, as appropriate, and information:
(a) Airplane maintenance manual or section. (1) Introduction information
that includes an explanation of the airplane's features and data to the
extent necessary for maintenance or preventive maintenance.
(2) A description of the airplane and its systems and installations
including its engines, propellers, and appliances.
(3) Basic control and operation information describing how the airplane
components and systems are controlled and how they operate, including any
special procedures and limitations that apply.
(4) Servicing information that covers details regarding servicing points,
capacities of tanks, reservoirs, types of fluids to be used, pressures
applicable to the various systems, location of access panels for inspection
and servicing, locations of lubrication points, lubricants to be used,
equipment required for servicing, tow instructions and limitations, 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 equipment that provides the recommended periods at which
they should be cleaned, inspected, adjusted, tested, and lubricated, and the
degree of inspection, the applicable wear tolerances, and work recommended at
these periods. However, the applicant may refer to an accessory, instrument,
or equipment manufacturer as the source of this information if the applicant
shows that the item has an exceptionally high degree of complexity requiring
specialized maintenance techniques, test equipment, or expertise. The
recommended overhaul periods and necessary cross references to the
Airworthiness Limitations section of the manual must also be included. In
addition, the applicant must include an inspection program that includes the
frequency and extent of the inspections necessary to provide for the
continued airworthiness of the airplane.
(2) Troubleshooting information describing probable malfunctions, how to
recognize those malfunctions, and the remedial action for those malfunctions.
(3) Information describing the order and method of removing and replacing
products and parts with any necessary precautions to be taken.
(4) Other general procedural instructions including procedures for system
testing during ground running, symmetry checks, weighing and determining the
center of gravity, lifting and shoring, and storage limitations.
(c) Diagrams of structural access plates and information needed to gain
access for inspections when access plates are not provided.
(d) Details for the application of special inspection techniques including
radiographic and ultrasonic testing where such processes are specified.
(e) Information needed to apply protective treatments to the structure
after inspection.
(f) All data relative to structural fasteners such as identification,
discard recommendations, and torque values.
(g) A list of special tools needed.
H25.4 Airworthiness Limitations section.
The Instructions for Continued Airworthiness must contain a section titled
Airworthiness Limitations that is segregated and clearly distinguishable from
the rest of the document. This section must set forth each mandatory
replacement time, structural inspection interval, and related structural
inspection procedure approved under Sec. 25.571. If the Instructions for
Continued Airworthiness consist of multiple documents, the section required
by this paragraph must be included in the principal manual. This section must
contain a legible statement in a prominent location that reads: "The
Airworthiness Limitations section is FAA approved and specifies maintenance
required under Secs. 43.16 and 91.403 of the Federal Aviation Regulations
unless an alternative program has been FAA approved."
[Amdt. 25-54, 45 FR 60177, Sept. 11, 1980, as amended by Amdt. 25-68, 54 FR
34329, Aug. 18, 1989]
Effective Date Note: At 54 FR 34329, Aug. 18, 1989, Sec. H25.4 in Appendix
H, Part 25 was amended by changing the cross reference "Sec. 91.163" to "Sec.
91.403", effective August 18, 1990.
Appendix I to Part 25--Installation of an Automatic Takeoff Thrust Control
System (ATTCS)
I25.1 General.
(a) This appendix specifies additional requirements for installation of an
engine power control system that automatically resets thrust or power on
operating engine(s) in the event of any one engine failure during takeoff.
(b) With the ATTCS and associated systems functioning normally as designed,
all applicable requirements of Part 25, except as provided in this appendix,
must be met without requiring any action 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 automatic system used on takeoff, including all devices, both
mechanical and electrical, that sense engine failure, transmit signals,
actuate fuel controls or power levers or increase engine power by other means
on operating engines to achieve scheduled thrust or power increases, and
furnish cockpit information on system operation.
(b) Critical Time Interval. When conducting an ATTCS takeoff, the critical
time interval is between V1 minus 1 second and a point on the minimum
performance, all-engine flight path where, assuming a simultaneous occurrence
of an engine and ATTCS failure, the resulting minimum flight path thereafter
intersects the Part 25 required actual flight path at no less than 400 feet
above the takeoff surface. This time interval is shown in the following
illustration:
[ ...Illustration appears here... ]
I25.3 Performance and System Reliability Requirements.
The applicant must comply with the performance and ATTCS reliability
requirements as follows:
(a) An ATTCS failure or a combination of failures in the ATTCS during the
critical time interval:
(1) Shall not prevent the insertion of the maximum approved takeoff thrust
or power, or must be shown to be an improbable event.
(2) Shall not result in a significant loss or reduction in thrust or power,
or must be shown to be an extremely improbable event.
(b) The concurrent existence of an ATTCS failure and an engine failure
during the critical time interval must be shown to be extremely improbable.
(c) All applicable performance requirements of Part 25 must be met with an
engine failure occurring at the most critical point during takeoff with the
ATTCS system functioning.
I25.4 Thrust Setting.
The initial takeoff thrust or power setting on each engine at the
beginning of the takeoff roll may not be less than any of the following:
(a) Ninety (90) percent of the thrust or power set by the ATTCS (the
maximum takeoff thrust or power approved for the airplane under existing
ambient conditions);
(b) That required to permit normal operation of all safety-related systems
and equipment dependent upon engine thrust or power lever position; or
(c) That shown to be free of hazardous engine response characteristics when
thrust or power is advanced from the initial takeoff thrust or power to the
maximum approved takeoff thrust or power.
I25.5 Powerplant Controls.
(a) In addition to the requirements of Sec. 25.1141, no single failure or
malfunction, or probable combination thereof, of the ATTCS, including
associated systems, may cause the failure of any powerplant function
necessary for safety.
(b) The ATTCS must be designed to:
(1) Apply thrust or power on the operating engine(s), following any one
engine failure during takeoff, to achieve the maximum approved takeoff thrust
or power without exceeding engine operating limits;
(2) Permit manual decrease or increase in thrust or power up to the maximum
takeoff thrust or power approved for the airplane under existing conditions
through the use of the power lever. For airplanes equipped with limiters that
automatically prevent engine operating limits from being exceeded under
existing ambient conditions, other means may be used to increase the thrust
or power in the event of an ATTCS failure provided the means is located on or
forward of the power levers; is easily identified and operated under all
operating conditions by a single action of either pilot with the hand that is
normally used to actuate the power levers; and meets the requirements of Sec.
25.777 (a), (b), and (c);
(3) Provide a means to verify to the flightcrew before takeoff that the
ATTCS is in a condition to operate; and
(4) Provide a means for the flightcrew to deactivate the automatic
function. This means must be designed to prevent inadvertent deactivation.
I25.6 Powerplant Instruments.
In addition to the requirements of Sec. 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 an engine has failed, a warning system that is
independent of the ATTCS must be provided to give the pilot a clear warning
of any engine failure during takeoff.
[Amdt. 25-62, 52 FR 43156, Nov. 9, 1987]
Appendix J to Part 25--Emergency Demonstration
The following test criteria and procedures must be used for showing
compliance with Sec. 25.803:
(a) The emergency evacuation must be conducted either during the dark of
the night or during daylight with the dark of night simulated. If the
demonstration is conducted indoors during daylight hours, it must be
conducted with each window covered and each door closed to minimize the
daylight effect. Illumination on the floor or ground may be used, but it must
be kept low and shielded against shining into the airplane's windows or
doors.
(b) The airplane must be in a normal attitude with landing gear extended.
(c) Unless the airplane is equipped with an off-wing descent means, stands
or ramps may be used for descent from the wing to the ground. Safety
equipment such as mats or inverted life rafts may be placed on the floor or
ground to protect participants. No other equipment that is not part of the
emergency evacuation equipment of the airplane may be used to aid the
participants in reaching the ground.
(d) Except as provided in paragraph (a) of this Appendix, only the
airplane's emergency lighting system may provide illumination.
(e) All emergency equipment required for the planned operation of the
airplane must be installed.
(f) Each external door and exit, and each internal door or curtain, must be
in the takeoff configuration.
(g) Each crewmember must be seated in the normally assigned seat for
takeoff and must remain in the seat until receiving the signal for
commencement of the demonstration. Each crewmember must be a person having
knowledge of the operation of exits and emergency equipment and, if
compliance with Sec. 121.291 is also being demonstrated, each flight
attendant must be a member of a regularly scheduled line crew.
(h) A representative passenger load of persons in normal health must be
used as follows:
(1) At least 40 percent of the passenger load must be female.
(2) At least 35 percent of the passenger load must be over 50 years of age.
(3) At least 15 percent of the passenger load must be female and over 50
years of age.
(4) Three life-size dolls, not included as part of the total passenger
load, must be carried by passengers to simulate live infants 2 years old or
younger.
(5) Crewmembers, mechanics, and training personnel, who maintain or operate
the airplane in the normal course of their duties, may not be used as
passengers.
(i) No passenger may be assigned a specific seat except as the
Administrator may require. Except as required by subparagraph (g) of this
paragraph, no employee of the applicant may be seated next to an emergency
exit.
(j) Seat belts and shoulder harnesses (as required) must be fastened.
(k) Before the start of the demonstration, approximately one-half of the
total average amount of carry-on baggage, blankets, pillows, and other
similar articles must be distributed at several locations in aisles and
emergency exit access ways to create minor obstructions.
(l) No prior indication may be given to any crewmember or passenger of the
particular exits to be used in the demonstration.
(m) The applicant may not practice, rehearse, or describe the demonstration
for the participants nor may any participant have taken part in this type of
demonstration within the preceding 6 months.
(n) The pretakeoff passenger briefing required by Sec. 121.571 may be
given. The passengers may also be advised to follow directions of crewmembers
but not be instructed on the procedures to be followed in the demonstration.
(o) If safety equipment as allowed by paragraph (c) of this appendix is
provided, either all passenger and cockpit windows must be blacked out or all
of the emergency exits must have safety equipment in order to prevent
disclosure of the available emergency exits.
(p) Not more than 50 percent of the emergency exits in the sides of the
fuselage of an airplane that meets all of the requirements applicable to the
required emergency exits for that airplane may be used for the demonstration.
Exits that are not to be used in the demonstration must have the exit handle
deactivated or must be indicated by red lights, red tape, or other acceptable
means placed outside the exits to indicate fire or other reason why they are
unusable. The exits to be used must be representative of all of the emergency
exits on the airplane and must be designated by the applicant, subject to
approval by the Administrator. At least one floor level exit must be used.
(q) Except as provided in paragraph (c) of this section, all evacuees must
leave the airplane by a means provided as part of the airplane's equipment.
(r) The applicant's approved procedures must be fully utilized, except the
flightcrew must take no active role in assisting others inside the cabin
during the demonstration.
(s) The evacuation time period is completed when the last occupant has
evacuated the airplane and is on the ground. Provided that the acceptance
rate of the stand or ramp is no greater than the acceptance rate of the means
available on the airplane for descent from the wing during an actual crash
situation, evacuees using stands or ramps allowed by paragraph (c) of this
Appendix are considered to be on the ground when they are on the stand or
ramp.
[Doc. No. 24344, Amdt. 25-72, 55 FR 29788, July 20, 1990, as amended by Amdt.
25-79, 58 FR 45229, Aug. 26, 1993]
*****************************************************************************
58 FR 45224, No. 164, Aug. 26, 1993
SUMMARY: These amendments to the airworthiness standards for transport
category airplanes and the operating rules for air carrier operators of such
airplanes modify the procedures for conducting an emergency evacuation
demonstration. These include a requirement that the flightcrew take no active
role in the demonstration, and a change to the age/sex distribution
requirement for demonstration participants. In addition, the airworthiness
standards are amended to standardize the illumination requirements for the
handles of the various types of passenger emergency exits, and to add a
requirement to prevent the inadvertent disabling of the public address system
because of an unstowed microphone. These amendments are intended to enhance
the provisions for egress of occupants of transport category airplanes under
emergency conditions.
EFFECTIVE DATE: September 27, 1993.
*****************************************************************************
