Source: https://www.ecfr.gov/cgi-bin/retrieveECFR?gp=&mc=true&n=sp14.1.25.d&r=SUBPART&ty=HTML
Timestamp: 2020-07-11 21:58:37
Document Index: 658223566

Matched Legal Cases: ['art 25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', 'art 25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25', '§25']

Title 14 → Chapter I → Subchapter C → Part 25 → Subpart D
§25.601 General.
§25.603 Materials.
§25.605 Fabrication methods.
§25.607 Fasteners.
§25.609 Protection of structure.
§25.611 Accessibility provisions.
§25.613 Material strength properties and material design values.
§25.619 Special factors.
§25.621 Casting factors.
§25.623 Bearing factors.
§25.625 Fitting factors.
§25.631 Bird strike damage.
§25.651 Proof of strength.
§25.655 Installation.
§25.657 Hinges.
§25.675 Stops.
§25.677 Trim systems.
§25.679 Control system gust locks.
§25.681 Limit load static tests.
§25.683 Operation tests.
§25.685 Control system details.
§25.689 Cable systems.
§25.693 Joints.
§25.697 Lift and drag devices, controls.
§25.699 Lift and drag device indicator.
§25.701 Flap and slat interconnection.
§25.703 Takeoff warning system.
§25.721 General.
§25.723 Shock absorption tests.
§§25.725-25.727 [Reserved]
§25.729 Retracting mechanism.
§25.731 Wheels.
§25.733 Tires.
§25.735 Brakes and braking systems.
§25.737 Skis.
§25.751 Main float buoyancy.
§25.753 Main float design.
§25.755 Hulls.
§25.771 Pilot compartment.
§25.772 Pilot compartment doors.
§25.773 Pilot compartment view.
§25.775 Windshields and windows.
§25.777 Cockpit controls.
§25.779 Motion and effect of cockpit controls.
§25.781 Cockpit control knob shape.
§25.783 Fuselage doors.
§25.785 Seats, berths, safety belts, and harnesses.
§25.787 Stowage compartments.
§25.789 Retention of items of mass in passenger and crew compartments and galleys.
§25.791 Passenger information signs and placards.
§25.793 Floor surfaces.
§25.795 Security considerations.
§25.801 Ditching.
§25.803 Emergency evacuation.
§25.807 Emergency exits.
§25.810 Emergency egress assist means and escape routes.
§25.811 Emergency exit marking.
§25.813 Emergency exit access.
§25.815 Width of aisle.
§25.817 Maximum number of seats abreast.
§25.819 Lower deck service compartments (including galleys).
§25.820 Lavatory doors.
§25.831 Ventilation.
§25.832 Cabin ozone concentration.
§25.833 Combustion heating systems.
§25.843 Tests for pressurized cabins.
§25.871 Leveling means.
§25.875 Reinforcement near propellers.
§25.899 Electrical bonding and protection against static electricity.
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.
[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]
(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.
(b) Have provisions for ventilation and drainage where necessary for protection.
(a)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 §25.1529.
(b) EWIS must meet the accessibility requirements of §25.1719.
[Amdt. 25-23, 35 FR 5674, Apr. 8, 1970, as amended by Amdt. 25-123, 72 FR 63404, Nov. 8, 2007]
The factor of safety prescribed in §25.303 must be multiplied by the highest pertinent special factor of safety prescribed in §§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.
(a) General. For castings used in structural applications, 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.
(c) Critical castings. Each casting whose failure could preclude continued safe flight and landing of the airplane or could result in serious injury to occupants is a critical casting. Each critical casting must have a factor associated with it for showing compliance with strength and deformation requirements of §25.305, and must comply with the following criteria associated with that factor:
(1) A casting factor of 1.0 or greater may be used, provided that—
(i) It is demonstrated, in the form of process qualification, proof of product, and process monitoring that, for each casting design and part number, the castings produced by each foundry and process combination have coefficients of variation of the material properties that are equivalent to those of wrought alloy products of similar composition. Process monitoring must include testing of coupons cut from the prolongations of each casting (or each set of castings, if produced from a single pour into a single mold in a runner system) and, on a sampling basis, coupons cut from critical areas of production castings. The acceptance criteria for the process monitoring inspections and tests must be established and included in the process specifications to ensure the properties of the production castings are controlled to within levels used in design.
(iii) One casting undergoes a static test and is shown to meet the strength and deformation requirements of §25.305(a) and (b).
(2) A casting factor of 1.25 or greater may be used, provided that—
(i) Each casting receives:
(A) The strength requirements of §25.305(b) at an ultimate load corresponding to a casting factor of 1.25; and
(B) The deformation requirements of §25.305(a) at a load of 1.15 times the limit load.
(3) A casting factor of 1.50 or greater may be used, provided that—
(A) The strength requirements of §25.305(b) at an ultimate load corresponding to a casting factor of 1.50; and
(d) Non-critical castings. For each casting other than critical castings, as specified in paragraph (c) of this section, the following apply:
(iii) Three sample castings undergo static tests and are shown to meet the strength and deformation requirements of §25.305(a) and (b).
(5) The number of castings per production batch to be inspected by non-visual methods in accordance with paragraphs (d)(2) and (3) of this section may be reduced when an approved quality control procedure is established.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-139, 79 FR 59429, Oct. 2, 2014]
(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.
(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—
(b) No fitting factor need be used—
(d) For each seat, berth, safety belt, and harness, the fitting factor specified in §25.785(f)(3) applies.
(2) For the conditions described in §25.629(d) below, for all approved altitudes, any airspeed up to the greater airspeed defined by;
(i) The VD/MD envelope determined by §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.
(8) Any damage or failure condition, required or selected for investigation by §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 §25.571(e), and
(ii) A damage tolerance investigation in accordance with §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 §§25.631, 25.671, 25.672, and 25.1309.
(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.
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 §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.
(b) Compliance with the special factors requirements of §§25.619 through 25.625 and 25.657 for control surface hinges must be shown by analysis or individual load tests.
(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.
(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.
(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.
[Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 41 FR 55466, Dec. 20, 1976]
(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.
(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; Amdt. 25-115, 69 FR 40527, July 2, 2004]
(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.
(a) Compliance with the limit load requirements of this Part must be shown by tests in which—
(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.
(a) 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—
(b) It must be shown by analysis and, where necessary, by tests, that in the presence of deflections of the airplane structure due to the separate application of pitch, roll, and yaw limit maneuver loads, the control system, when loaded to obtain these limit loads and operated within its operational range of deflections, can be exercised about all control axes and remain free from—
(2) Excessive friction;
(3) Disconnection; and
(4) Any form of permanent damage.
(c) It must be shown that under vibration loads in the normal flight and ground operating conditions, no hazard can result from interference or contact with adjacent elements.
[Amdt. 25-139, 79 FR 59430, Oct. 2, 2014]
(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.
(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.
(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.
(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).
(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]
A takeoff warning system must be installed and must meet the following requirements:
(a) The system must provide to the pilots an aural warning that is automatically activated during the initial portion of the takeoff roll if the airplane is in a configuration, including any of the following, that would not allow a safe takeoff:
(1) The wing flaps or leading edge devices are not within the approved range of takeoff positions.
(2) Wing spoilers (except lateral control spoilers meeting the requirements of §25.671), speed brakes, or longitudinal trim devices are in a position that would not allow a safe takeoff.
(b) The warning required by paragraph (a) of this section must continue until—
(1) The configuration is changed to allow a safe takeoff;
(2) Action is taken by the pilot to terminate the takeoff roll;
(3) The airplane is rotated for takeoff; or
(4) The warning is manually deactivated by the pilot.
(c) The means used to activate the system must function properly throughout the ranges of takeoff weights, altitudes, and temperatures for which certification is requested.
(a) The landing gear system must be designed so that when it fails due to overloads during takeoff and landing, the failure mode is not likely to cause spillage of enough fuel to constitute a fire hazard. The overloads must be assumed to act in the upward and aft directions in combination with side loads acting inboard and outboard. In the absence of a more rational analysis, the side loads must be assumed to be up to 20 percent of the vertical load or 20 percent of the drag load, whichever is greater.
(b) The airplane must be designed to avoid any rupture leading to the spillage of enough fuel to constitute a fire hazard as a result of a wheels-up landing on a paved runway, under the following minor crash landing conditions:
(1) Impact at 5 feet-per-second vertical velocity, with the airplane under control, at Maximum Design Landing Weight—
(i) With the landing gear fully retracted; and
(ii) With any one or more landing gear legs not extended.
(2) Sliding on the ground, with—
(i) The landing gear fully retracted and with up to a 20° yaw angle; and
(ii) Any one or more landing gear legs not extended and with 0° yaw angle.
(c) For configurations where the engine nacelle is likely to come into contact with the ground, the engine pylon or engine mounting must be designed so that when it fails due to overloads (assuming the overloads to act predominantly in the upward direction and separately, predominantly in the aft direction), the failure mode is not likely to cause the spillage of enough fuel to constitute a fire hazard.
(a) The analytical representation of the landing gear dynamic characteristics that is used in determining the landing loads must be validated by energy absorption tests. A range of tests must be conducted to ensure that the analytical representation is valid for the design conditions specified in §25.473.
(c) In lieu of the tests prescribed in this section, changes in previously approved design weights and minor changes in design may be substantiated by analyses based on previous tests conducted on the same basic landing gear system that has similar energy absorption characteristics.
[Doc. No. 1999-5835, 66 FR 27394, May 16, 2001]
(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 1.23VSR (with the wing-flaps in take-off position at design take-off weight), occurring during retraction and extension at any airspeed up to 1.5 VSR1 (with the wing-flaps in the approach position at design landing weight), and
(iii) Any load factor up to those specified in §25.345(a) for the wing-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.
(c) Emergency operation. There must be an emergency means for extending the landing gear in the event of—
(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.
(7) A flightcrew alert must be provided whenever the landing gear position is not consistent with the landing gear selector lever position.
(f) Protection of equipment on landing gear and in wheel wells. Equipment that is essential to the safe operation of the airplane and that is located on the landing gear and in wheel wells must be protected from the damaging effects of—
(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; Amdt. 25-136, 77 FR 1617, Jan. 11, 2012]
(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—
(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.
(d) Overpressure burst prevention. Means must be provided in each wheel to prevent wheel failure and tire burst that may result from excessive pressurization of the wheel and tire assembly.
(e) Braked wheels. Each braked wheel must meet the applicable requirements of §25.735.
(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—
(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—
(f) Kinetic energy capacity—(1) Design landing stop. The design landing stop is an operational landing stop at maximum landing weight. The design landing stop brake kinetic energy absorption requirement of each wheel, brake, and tire assembly must be determined. It must be substantiated by dynamometer testing that the wheel, brake and tire assembly is capable of absorbing not less than this level of kinetic energy throughout the defined wear range of the brake. The energy absorption rate derived from the airplane manufacturer's braking requirements must be achieved. The mean deceleration must not be less than 10 fps2.
(2) Maximum kinetic energy accelerate-stop. The maximum kinetic energy accelerate-stop is a rejected takeoff for the most critical combination of airplane takeoff weight and speed. The accelerate-stop brake kinetic energy absorption requirement of each wheel, brake, and tire assembly must be determined. It must be substantiated by dynamometer testing that the wheel, brake, and tire assembly is capable of absorbing not less than this level of kinetic energy throughout the defined wear range of the brake. The energy absorption rate derived from the airplane manufacturer's braking requirements must be achieved. The mean deceleration must not be less than 6 fps2.
(g) Brake condition after high kinetic energy dynamometer stop(s). Following the high kinetic energy stop demonstration(s) required by paragraph (f) of this section, with the parking brake promptly and fully applied for at least 3 minutes, it must be demonstrated that for at least 5 minutes from application of the parking brake, no condition occurs (or has occurred during the stop), including fire associated with the tire or wheel and brake assembly, that could prejudice the safe and complete evacuation of the airplane.
[Doc. No. FAA-1999-6063, 67 FR 20420, Apr. 24, 2002, as amended by Amdt. 25-108, 67 FR 70827, Nov. 26, 2002; 68 FR 1955, Jan. 15, 2003]
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.
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.
Each main float must be approved and must meet the requirements of §25.521.
(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.
(a) Each pilot compartment and its equipment must allow the minimum flight crew (established under §25.1523) to perform their duties without unreasonable concentration or fatigue.
(b) The primary controls listed in §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 §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]
For an airplane that has a lockable door installed between the pilot compartment and the passenger compartment:
(a) For airplanes with a maximum passenger seating configuration of more than 20 seats, the emergency exit configuration must be designed so that neither crewmembers nor passengers require use of the flightdeck 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.
(c) There must be an emergency means to enable a flight attendant to enter the pilot compartment in the event that the flightcrew becomes incapacitated.
[Doc. No. 24344, 55 FR 29777, July 20, 1990, as amended by Amdt. 25-106, 67 FR 2127, Jan. 15, 2002]
(a) Nonprecipitation conditions. For nonprecipitation conditions, the following apply:
(1) Each pilot compartment must be arranged to give the pilots a sufficiently extensive, clear, and undistorted view, to enable them to safely perform any maneuvers within the operating limitations of the airplane, including taxiing takeoff, approach, and landing.
(2) Each pilot compartment must be free of glare and reflection that could interfere with the normal duties of the minimum flight crew (established under §25.1523). This must be shown in day and night flight tests under nonprecipitation conditions.
(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 1.5 VSR1 with lift and drag devices retracted; and
(ii) The icing conditions specified in Appendix C of this part and the following icing conditions specified in Appendix O of this part, if certification for flight in icing conditions is sought:
(A) For airplanes certificated in accordance with §25.1420(a)(1), the icing conditions that the airplane is certified to safely exit following detection.
(B) For airplanes certificated in accordance with §25.1420(a)(2), the icing conditions that the airplane is certified to safely operate in and the icing conditions that the airplane is certified to safely exit following detection.
(C) For airplanes certificated in accordance with §25.1420(a)(3) and for airplanes not subject to §25.1420, all icing conditions.
(2) No single failure of the systems used to provide the view required by paragraph (b)(1) of this section must cause the loss of that view by both pilots in the specified precipitation conditions.
(i) Any system failure or combination of failures which is not extremely improbable, in accordance with §25.1309, under the precipitation conditions specified in paragraph (b)(1) of this section.
(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 §25.1381.
(e) Vision systems with transparent displays. A vision system with a transparent display surface located in the pilot's outside field of view, such as a head up-display, head mounted display, or other equivalent display, must meet the following requirements in nonprecipitation and precipitation conditions:
(1) While the vision system display is in operation, it must compensate for interference with the pilot's outside field of view such that the combination of what is visible in the display and what remains visible through and around it, enables the pilot to perform the maneuvers and normal duties of paragraph (a) of this section.
(2) The pilot's view of the external scene may not be distorted by the transparent display surface or by the vision system imagery. When the vision system displays imagery or any symbology that is referenced to the imagery and outside scene topography, including attitude symbology, flight path vector, and flight path angle reference cue, that imagery and symbology must be aligned with, and scaled to, the external scene.
(3) The vision system must provide a means to allow the pilot using the display to immediately deactivate and reactivate the vision system imagery, on demand, without removing the pilot's hands from the primary flight controls or thrust controls.
(4) When the vision system is not in operation it may not restrict the pilot from performing the maneuvers specified in paragraph (a)(1) of this section or the pilot compartment from meeting the provisions of paragraph (a)(2) of this section.
[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; Amdt. 25-108, 67 FR 70827, Nov. 26, 2002; Amdt. 25-121, 72 FR 44669, Aug. 8, 2007; Amdt. 25-136, 77 FR 1618, Jan. 11, 2012; Amdt. 25-140, 79 FR 65525, Nov. 4, 2014; Docket FAA-2013-0485, Amdt. 25-144, 81 FR 90169, Dec. 13, 2016]
(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 §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 §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 §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 §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]
Cockpit controls must be designed so that they operate in accordance with the following movement and actuation:
(a) Aerodynamic controls:
(1) Primary.
Aileron Right (clockwise) for right wing down.
Elevator Rearward for nose up.
Rudder Right pedal forward for nose right.
(2) Secondary.
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.
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.
Landing gear Down to extend.
(a) General. This section applies to fuselage doors, which includes all doors, hatches, openable windows, access panels, covers, etc., on the exterior of the fuselage that do not require the use of tools to open or close. This also applies to each door or hatch through a pressure bulkhead, including any bulkhead that is specifically designed to function as a secondary bulkhead under the prescribed failure conditions of part 25. These doors must meet the requirements of this section, taking into account both pressurized and unpressurized flight, and must be designed as follows:
(5) Each removable bolt, screw, nut, pin, or other removable fastener must meet the locking requirements of §25.607.
(6) Certain doors, as specified by §25.807(h), must also meet the applicable requirements of §§25.809 through 25.812 for emergency exits.
(b) Opening by persons. There must be a means to safeguard each door against opening during flight due to inadvertent action by persons. In addition, design precautions must be taken to minimize the possibility for a person to open a door intentionally during flight. If these precautions include the use of auxiliary devices, those devices and their controlling systems must be designed so that—
(1) No single failure will prevent more than one exit from being opened; and
(c) Pressurization prevention means. There must be a provision to prevent pressurization of the airplane to an unsafe level if any door subject to pressurization is not fully closed, latched, and locked.
(2) Doors that meet the conditions described in paragraph (h) of this section are not required to have a dedicated pressurization prevention means if, from every possible position of the door, it will remain open to the extent that it prevents pressurization or safely close and latch as pressurization takes place. This must also be shown with any single failure and malfunction, except that—
(i) With failures or malfunctions in the latching mechanism, it need not latch after closing; and
(d) Latching and locking. The latching and locking mechanisms must be designed as follows:
(3) Each door subject to pressurization, and for which the initial opening movement is not inward, must—
(i) Have an individual lock for each latch;
(6) It must not be possible to unlatch the latches with the locks in the locked position. Locks must be designed to withstand the limit loads resulting from—
(i) The maximum operator effort when the latches are operated manually;
(ii) The powered latch actuators, if installed; and
(iii) The relative motion between the latch and the structural counterpart.
(e) Warning, caution, and advisory indications. Doors must be provided with the following indications:
(3) There must be a visual means on the flight deck to signal the pilots if any door is not fully closed, latched, and locked. The means must be designed such that any failure or combination of failures that would result in an erroneous closed, latched, and locked indication is improbable for—
(i) Each door that is subject to pressurization and for which the initial opening movement is not inward; or
(ii) Each door that could be a hazard if unlatched.
(f) Visual inspection provision. Each door for which unlatching of the door could be a hazard must have a provision for direct visual inspection to determine, without ambiguity, if the door is fully closed, latched, and locked. The provision must be permanent and discernible under operational lighting conditions, or by means of a flashlight or equivalent light source.
(g) Certain maintenance doors, removable emergency exits, and access panels. Some doors not normally opened except for maintenance purposes or emergency evacuation and some access panels need not comply with certain paragraphs of this section as follows:
(h) Doors that are not a hazard. For the purposes of this section, a door is considered not to be a hazard in the unlatched condition during flight, provided it can be shown to meet all of the following conditions:
(1) Doors in pressurized compartments would remain in the fully closed position if not restrained by the latches when subject to a pressure greater than 1⁄2 psi. Opening by persons, either inadvertently or intentionally, need not be considered in making this determination.
(5) The airplane would meet the structural design requirements with the door open. This assessment must include the aeroelastic stability requirements of §25.629, as well as the strength requirements of subpart C of this part.
(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 §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 §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 §25.395.
(3) The inertia forces specified in §25.561 must be multiplied by a factor of 1.33 (instead of the fitting factor prescribed in §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.
(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 or B 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.
(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.
(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.
[Amdt. 25-72, 55 FR 29780, July 20, 1990, as amended by Amdt. 25-88, 61 FR 57956, Nov. 8, 1996]
(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 those emergency landing conditions of §25.561(b)(3) for which the breaking loose of the contents of such compartments in the specified direction could—
(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 §25.561(b).
(b) Each interphone restraint system must be designed so that when subjected to the load factors specified in §25.561(b)(3), the interphone will remain in its stowed position.
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]
(where H0 is defined under §25.365(e)(2) of this part and D need not exceed 5.05 feet (1.54 meters)). The sphere is applied everywhere within the fuselage—limited by the forward bulkhead and the aft bulkhead of the passenger cabin and cargo compartment beyond which only one-half the sphere is applied.
(e) Exceptions. Airplanes used solely to transport cargo only need to meet the requirements of paragraphs (b)(1), (b)(3), and (c)(2) of this section.
(f) Material Incorporated by Reference. You must use National Institute of Justice (NIJ) Standard 0101.04, Ballistic Resistance of Personal Body Armor, June 2001, Revision A, to establish ballistic resistance as required by paragraph (a)(3) of this section.
Type A 110
Type B 75
Type C 55
Type I 45
Type II 40
Type IV 9
(a) Each non over-wing Type A, Type B or Type C exit, and any other non over-wing landplane emergency exit more than 6 feet from the ground with the airplane on the ground and the landing gear extended, 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 Type A or Type B exits, 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) Except for assisting means installed at Type C exits, it must be automatically erected within 6 seconds after deployment is begun. Assisting means installed at Type C exits must be automatically erected within 10 seconds from the time the opening means of the exit is actuated.
(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 §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 or Type B 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 non over-wing exit. If an assist means is required, it must be automatically deployed and automatically erected concurrent with the opening of the exit. In the case of assist means installed at Type C exits, it must be self-supporting within 10 seconds from the time the opening means of the exits is actuated. For all other exit types, it must be self-supporting 6 seconds after deployment is begun.
(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 from each Type A or Type B passenger emergency exit, or any common escape route from two Type III passenger emergency exits, must be at least 42 inches wide; that from any other passenger emergency exit must be at least 24 inches wide; 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) Means must be provided to assist evacuees to reach the ground for all Type C exits located over the wing and, 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, for all other exit types.
(1) If the escape route is over the 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.
(2) 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.
(3) The assisting means provided for each escape route leading from a Type A or B emergency exit must be capable of carrying simultaneously two parallel lines of evacuees; and, the assisting means leading from any other exit type must be capable of carrying as many parallel lines of evacuees as there are required escape routes.
(4) The assisting means provided for each escape route leading from a Type C exit must be automatically erected within 10 seconds from the time the opening means of the exit is actuated, and that provided for the escape route leading from any other exit type must be automatically erected within 10 seconds after actuation of the erection system.
(e) If an integral stair is installed in a passenger entry door that is qualified as a passenger emergency exit, the stair must be designed so that, under the following conditions, the effectiveness of passenger emergency egress will not be impaired:
(1) The door, integral stair, and operating mechanism have been subjected to the inertia forces specified in §25.561(b)(3), acting separately relative to the surrounding structure.
(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.
[Amdt. 25-72, 55 FR 29782, July 20, 1990, as amended by Amdt. 25-88, 61 FR 57958, Nov. 8, 1996; 62 FR 1817, Jan. 13, 1997; Amdt. 25-114, 69 FR 24502, May 3, 2004]
(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—
(2) Each Type A, Type B, Type C or Type I passenger emergency exit operating handle must—
(4) Each Type A, Type B, Type C, Type I, or Type II passenger emergency exit with a locking mechanism released by rotary motion of the handle must be marked—
(iii) With the word “open” in red letters 1 inch high, placed horizontally near the head of the arrow.
(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”.
(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 §25.811(d)(1) and each passenger emergency exit marking sign required by §25.811(d)(2) must have red letters at least 11⁄2 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 §25.811(d)(3) must have red letters at least 11⁄2 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 §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.
(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 two feet for all other overwing emergency exits along the 30 percent of the slip-resistant portion of the escape route required in §25.810(c) that is farthest from the exit; and
(iii) Not less than 0.03 foot-candle on 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 §25.810(a) 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 first contact with the ground outside the cabin.
(h) The means required in §§25.810(a)(1) and (d) 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 §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 §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; Amdt. 25-88, 61 FR 57958, Nov. 8, 1996; Amdt. 25-116, 69 FR 62788, Oct. 27, 2004; Amdt. 25-128, 74 FR 25645, May 29, 2009]
(5) For any tailcone exit that qualifies for 25 additional passenger seats under the provisions of §25.807(g)(9)(ii), an assist space must be provided, if an assist means is required by §25.810(a).
[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; Amdt. 25-76, 57 FR 29120, June 30, 1992; Amdt. 25-88, 61 FR 57958, Nov. 8, 1996; Amdt. 25-116, 69 FR 62788, Oct. 27, 2004; Amdt. 25-128, 74 FR 25645, May 29, 2009]
Minimum passenger aisle width (inches)
Less than 25 in. from floor
25 in. and more from floor
10 or less 112 15
11 through 19 12 20
20 or more 15 20
[Amdt. 25-15, 32 FR 13265, Sept. 20, 1967, as amended by Amdt. 25-38, 41 FR 55466, Dec. 20, 1976]
[Amdt. 25-15, 32 FR 13265, Sept. 20, 1967]
[Amdt. 25-53, 45 FR 41593, June 19, 1980; 45 FR 43154, June 26, 1980; Amdt. 25-110; 68 FR 36883, June 19, 2003]
All lavatory doors must be designed to preclude anyone from becoming trapped inside the lavatory. If a locking mechanism is installed, it must be capable of being unlocked from the outside without the aid of special tools.
[Doc. No. 2003-14193, 69 FR 24502, May 3, 2004]
(a) Under normal operating conditions and in the event of any probable failure conditions of any system which would adversely affect the ventilating air, the ventilation system must be designed to provide a sufficient amount of uncontaminated air to enable the crewmembers to perform their duties without undue discomfort or fatigue and to provide reasonable passenger comfort. For normal operating conditions, the ventilation system must be designed to provide each occupant with an airflow containing at least 0.55 pounds of fresh air per minute.
(b) Crew and passenger compartment air must be free from harmful or hazardous concentrations of gases or vapors. In meeting this requirement, the following apply:
(1) Carbon monoxide concentrations in excess of 1 part in 20,000 parts of air are considered hazardous. For test purposes, any acceptable carbon monoxide detection method may be used.
(2) Carbon dioxide concentration during flight must be shown not to exceed 0.5 percent by volume (sea level equivalent) in compartments normally occupied by passengers or crewmembers.
(c) There must be provisions made to ensure that the conditions prescribed in paragraph (b) of this section are met after reasonably probable failures or malfunctioning of the ventilating, heating, pressurization, or other systems and equipment.
(d) If accumulation of hazardous quantities of smoke in the cockpit area is reasonably probable, smoke evacuation must be readily accomplished, starting with full pressurization and without depressurizing beyond safe limits.
(e) Except as provided in paragraph (f) of this section, means must be provided to enable the occupants of the following compartments and areas to control the temperature and quantity of ventilating air supplied to their compartment or area independently of the temperature and quantity of air supplied to other compartments and areas:
(1) The flight crew compartment.
(2) Crewmember compartments and areas other than the flight crew compartment unless the crewmember compartment or area is ventilated by air interchange with other compartments or areas under all operating conditions.
(f) Means to enable the flight crew to control the temperature and quantity of ventilating air supplied to the flight crew compartment independently of the temperature and quantity of ventilating air supplied to other compartments are not required if all of the following conditions are met:
(1) The total volume of the flight crew and passenger compartments is 800 cubic feet or less.
(2) The air inlets and passages for air to flow between flight crew and passenger compartments are arranged to provide compartment temperatures within 5 degrees F. of each other and adequate ventilation to occupants in both compartments.
(3) The temperature and ventilation controls are accessible to the flight crew.
(g) The exposure time at any given temperature must not exceed the values shown in the following graph after any improbable failure condition.
(a) The airplane cabin ozone concentration during flight must be shown not to exceed—
(1) 0.25 parts per million by volume, sea level equivalent, at any time above flight level 320; and
(2) 0.1 parts per million by volume, sea level equivalent, time-weighted average during any 3-hour interval above flight level 270.
(b) For the purpose of this section, “sea level equivalent” refers to conditions of 25 °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; Amdt. 25-94, 63 FR 8848, Feb. 23, 1998]
Combustion heaters must be approved.
[Amdt. 25-72, 55 FR 29783, July 20, 1990]
(a) Strength test. The complete pressurized cabin, including doors, windows, and valves, must be tested as a pressure vessel for the pressure differential specified in §25.365(d).
(b) Functional tests. The following functional tests must be performed:
(1) Tests of the functioning and capacity of the positive and negative pressure differential valves, and of the emergency release valve, to 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.
There must be means for determining when the airplane is in a level position on the ground.
(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.
(a) Electrical bonding and protection against static electricity must be designed to minimize accumulation of electrostatic charge that would cause—
(1) Human injury from electrical shock,
(b) Compliance with paragraph (a) of this section may be shown by—