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Subpart D--Design and Construction 25.601 General. 25.651 Proof of strength. 25.671 General. 25.721 General. 25.751 Main float buoyancy. Personnel and Cargo Accommodations 25.771 Pilot compartment. 25.801 Ditching. 25.831 Ventilation. 25.841 Pressurized cabins. 25.851 Fire extinguishers. 25.871 Leveling means. |
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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] 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] 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 abearing 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] 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] 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] |
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(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] |
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(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. 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] 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] 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] |
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(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] 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 [...??equation goes here] (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: [...??equation goes here] (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, [...??equation goes here] [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] 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] 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] 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: [...??equation goes here] (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] 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 |
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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. |
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(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] 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] 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 (2) Secondary.Controls Motion and effect (b) Powerplant and auxiliary controls: (1) Powerplant Motion and effect (2) Auxiliary Motion and effect [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-72, 55 FR 29778, July 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: - Flap Control Knob [ ...Illustration appears here...
] [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-72, 55 FR 29778, July 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] 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] 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 amember 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] 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] |
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(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] 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 Jof 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] 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. 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)III III IV 1 through 9 1 Additional exits are required for passenger seating configurations greater than 179 seats in accordance with the following table: Additional emergency exits Increase in passenger (each side seating of configuration fuselage) allowed Type A 110 (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. Apassenger 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] 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] 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 |
