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25.901 Installation. 25.951 General. 25.991 Fuel pumps. 25.1011 General. 25.1041 General. 25.1091 Air induction. 25.1121 General. Powerplant Controls and Accessories 25.1141 Powerplant controls: general. 25.1181 Designated fire zones; regions included. |
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(a) For the purpose of this part, the airplane powerplant installation includes each component that-- (1) Is necessary for propulsion; (2) Affects the control of the major propulsive units; or (3) Affects the safety of the major propulsive units between normal inspections or overhauls. (b) For each powerplant-- (1) The installation must comply with-- (i) The installation instructions provided under Sec. 33.5 of this chapter; and (ii) The applicable provisions of this subpart; (2) The components of the installation must be constructed, arranged, and installed so as to ensure their continued safe operation between normal inspections or overhauls; (3) The installation must be accessible for necessary inspections and maintenance; and (4) The major components of the installation must be electrically bonded to the other parts of the airplane. (c) For each powerplant and auxiliary power unit installation, it must be established that no single failure or malfunction or probable combination of failures will jeopardize the safe operation of the airplane except that the failure of structural elements need not be considered if the probability of such failure is extremely remote. (d) Each auxiliary power unit installation must meet the applicable provisions of this subpart. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR 5676, Apr. 8, 1970; Amdt. 25-40, 42 FR 15042, Mar. 17, 1977; Amdt. 25-46, 43 FR 50597, Oct. 30, 1978] Sec. 25.903 Engines. (a) Engine type certificate. (1) Each engine must have a type certificate and must meet the applicable requirements of part 34 of this chapter. (2) Each turbine engine must either-- (i) Comply with Sec. 33.77 of this chapter in effect on October 31, 1974, or as subsequently amended; or (ii) Be shown to have a foreign object ingestion service history in similar installation locations which has not resulted in any unsafe condition. (b) Engine isolation. The powerplants must be arranged and isolated from each other to allow operation, in at least one configuration, so that the failure or malfunction of any engine, or of any system that can affect the engine, will not-- (1) Prevent the continued safe operation of the remaining engines; or (2) Require immediate action by any crewmember for continued safe operation. (c) Control of engine rotation. There must be means for stopping the rotation of any engine individually in flight, except that, for turbine engine installations, the means for stopping the rotation of any engine need be provided only where continued rotation could jeopardize the safety of the airplane. Each component of the stopping and restarting system on the engine side of the firewall that might be exposed to fire must be at least fire resistant. If hydraulic propeller feathering systems are used for this purpose, the feathering lines must be at least fire resistant under the operating conditions that may be expected to exist during feathering. (d) Turbine engine installations. For turbine engine installations-- (1) Design precautions must be taken to minimize the hazards to the airplane in the event of an engine rotor failure or of a fire originating within the engine which burns through the engine case. (2) The powerplant systems associated with engine control devices, systems, and instrumentation, must be designed to give reasonable assurance that those engine operating limitations that adversely affect turbine rotor structural integrity will not be exceeded in service. (e) Restart capability. (1) Means to restart any engine in flight must be provided. (2) An altitude and airspeed envelope must be established for in-flight engine restarting, and each engine must have a restart capability within that envelope. (3) For turbine engine powered airplanes, if the minimum windmilling speed of the engines, following the inflight shutdown of all engines, is insufficient to provide the necessary electrical power for engine ignition, a power source independent of the engine-driven electrical power generating system must be provided to permit in-flight engine ignition for restarting. (f) Auxiliary Power Unit. Each auxiliary power unit must be approved or meet the requirements of the category for its intended use. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR 5676, Apr. 8, 1970; Amdt. 25-40, 42 FR 15042, Mar. 17, 1977; Amdt. 25-57, 49 FR 6848, Feb. 23, 1984; Amdt. 25-72, 55 FR 29784, July 20, 1990; Amdt. 25-73, 55 FR 32861, Aug. 10, 1990; 55 FR 35139, Aug. 28, 1990] Sec. 25.904 Automatic takeoff thrust control system (ATTCS). Each applicant seeking approval for installation of an engine power control system that automatically resets the power or thrust on the operating engine(s) when any engine fails during the takeoff must comply with the requirements of Appendix I of this part. [Amdt. 25-62, 52 FR 43156, Nov. 9, 1987] Sec. 25.905 Propellers. (a) Each propeller must have a type certificate. (b) Engine power and propeller shaft rotational speed may not exceed the limits for which the propeller is certificated. (c) Each component of the propeller blade pitch control system must meet the requirements of Sec. 35.42 of this chapter. (d) Design precautions must be taken to minimize the hazards to the airplane in the event a propeller blade fails or is released by a hub failure. The hazards which must be considered include damage to structure and vital systems due to impact of a failed or released blade and the unbalance created by such failure or release. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-54, 45 FR 60173, Sept. 11, 1980; Amdt. 25-57, 49 FR 6848, Feb. 23, 1984; Amdt 25-72, 55 FR 29784, July 20, 1990] Sec. 25.907 Propeller vibration. (a) The magnitude of the propeller blade vibration stresses under any normal condition of operation must be determined by actual measurement or by comparison with similar installations for which these measurements have been made. (b) The determined vibration stresses may not exceed values that have been shown to be safe for continuous operation. Sec. 25.925 Propeller clearance. Unless smaller clearances are substantiated, propeller clearances with the airplane at maximum weight, with the most adverse center of gravity, and with the propeller in the most adverse pitch position, may not be less than the following: (a) Ground clearance. There must be a clearance of at least seven inches (for each airplane with nose wheel landing gear) or nine inches (for each airplane with tail wheel landing gear) between each propeller and the ground with the landing gear statically deflected and in the level takeoff, or taxiing attitude, whichever is most critical. In addition, there must be positive clearance between the propeller and the ground when in the level takeoff attitude with the critical tire(s) completely deflated and the corresponding landing gear strut bottomed. (b) Water clearance. There must be a clearance of at least 18 inches between each propeller and the water, unless compliance with Sec. 25.239(a) can be shown with a lesser clearance. (c) Structural clearance. There must be-- (1) At least one inch radial clearance between the blade tips and the airplane structure, plus any additional radial clearance necessary to prevent harmful vibration; (2) At least one-half inch longitudinal clearance between the propeller blades or cuffs and stationary parts of the airplane; and (3) Positive clearance between other rotating parts of the propeller or spinner and stationary parts of the airplane. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-72, 55 FR 29784, July 20, 1990] Sec. 25.929 Propeller deicing. (a) For airplanes intended for use where icing may be expected, there must be a means to prevent or remove hazardous ice accumulation on propellers or on accessories where ice accumulation would jeopardize engine performance. (b) If combustible fluid is used for propeller deicing, Secs. 25.1181 through 25.1185 and 25.1189 apply. Sec. 25.933 Reversing systems. (a) For turbojet reversing systems-- (1) Each system intended for ground operation only must be designed so that during any reversal in flight the engine will produce no more than flight idle thrust. In addition, it must be shown by analysis or test, or both, that-- (i) Each operable reverser can be restored to the forward thrust position; and (ii) The airplane is capable of continued safe flight and landing under any possible position of the thrust reverser. (2) Each system intended for inflight use must be designed so that no unsafe condition will result during normal operation of the system, or from any failure (or reasonably likely combination of failures) of the reversing system, under any anticipated condition of operation of the airplane including ground operation. Failure of structural elements need not be considered if the probability of this kind of failure is extremely remote. (3) Each system must have means to prevent the engine from producing more than idle thrust when the reversing system malfunctions, except that it may produce any greater forward thrust that is shown to allow directional control to be maintained, with aerodynamic means alone, under the most critical reversing condition expected in operation. (b) For propeller reversing systems-- (1) Each system intended for ground operation only must be designed so that no single failure (or reasonably likely combination of failures) or malfunction of the system will result in unwanted reverse thrust under any expected operating condition. Failure of structural elements need not be considered if this kind of failure is extremely remote. (2) Compliance with this section may be shown by failure analysis or testing, or both, for propeller systems that allow propeller blades to move from the flight low-pitch position to a position that is substantially less than that at the normal flight low-pitch position. The analysis may include or be supported by the analysis made to show compliance with the requirements of Sec. 35.21 of this chapter for the propeller and associated installation components. [Doc. No. 24344, Amdt. 25-72, 55 FR 29784, July 20, 1990] Sec. 25.934 Turbojet engine thrust reverser system tests. Thrust reversers installed on turbojet engines must meet the requirements of Sec. 33.97 of this chapter. [Amdt. 25-23, 35 FR 5677, Apr. 8, 1970] Sec. 25.937 Turbopropeller-drag limiting systems. Turbopropeller power airplane propeller-drag limiting systems must be designed so that no single failure or malfunction of any of the systems during normal or emergency operation results in propeller drag in excess of that for which the airplane was designed under Sec. 25.367. Failure of structural elements of the drag limiting systems need not be considered if the probability of this kind of failure is extremely remote. Sec. 25.939 Turbine engine operating characteristics. (a) Turbine engine operating characteristics must be investigated in flight to determine that no adverse characteristics (such as stall, surge, or flameout) are present, to a hazardous degree, during normal and emergency operation within the range of operating limitations of the airplane and of the engine. (b) [Reserved] (c) The turbine engine air inlet system may not, as a result of air flow distortion during normal operation, cause vibration harmful to the engine. [Amdt. 25-11, 32 FR 6912, May 5, 1967, as amended by Amdt. 25-40, 42 FR 15043, Mar. 17, 1977] Sec. 25.941 Inlet, engine, and exhaust compatibility. For airplanes using variable inlet or exhaust system geometry, or both-- (a) The system comprised of the inlet, engine (including thrust augmentation systems, if incorporated), and exhaust must be shown to function properly under all operating conditions for which approval is sought, including all engine rotating speeds and power settings, and engine inlet and exhaust configurations; (b) The dynamic effects of the operation of these (including consideration of probable malfunctions) upon the aerodynamic control of the airplane may not result in any condition that would require exceptional skill, alertness, or strength on the part of the pilot to avoid exceeding an operational or structural limitation of the airplane; and (c) In showing compliance with paragraph (b) of this section, the pilot strength required may not exceed the limits set forth in Sec. 25.143(c), subject to the conditions set forth in paragraphs (d) and (e) of Sec. 25.143. [Amdt. 25-38, 41 FR 55467, Dec. 20, 1976] Sec. 25.943 Negative acceleration. No hazardous malfunction of an engine, an auxiliary power unit approved for use in flight, or any component or system associated with the powerplant or auxiliary power unit may occur when the airplane is operated at the negative accelerations within the flight envelopes prescribed in Sec. 25.333. This must be shown for the greatest duration expected for the acceleration. [Amdt. 25-40, 42 FR 15043, Mar. 17, 1977] Sec. 25.945 Thrust or power augmentation system. (a) General. Each fluid injection system must provide a flow of fluid at the rate and pressure established for proper engine functioning under each intended operating condition. If the fluid can freeze, fluid freezing may not damage the airplane or adversely affect airplane performance. (b) Fluid tanks. Each augmentation system fluid tank must meet the following requirements: (1) Each tank must be able to withstand without failure the vibration, inertia, fluid, and structural loads that is may be subject to in operation. (2) The tanks as mounted in the airplane must be able to withstand without failure or leakage an internal pressure 1.5 times the maximum operating pressure. (3) If a vent is provided, the venting must be effective under all normal flight conditions. (4) [Reserved] (c) Augmentation system drains must be designed and located in accordance with Sec. 25.1455 if-- (1) The augmentation system fluid is subject to freezing; and (2) The fluid may be drained in flight or during ground operation. (d) The augmentation liquid tank capacity available for the use of each engine must be large enough to allow operation of the airplane under the approved procedures for the use of liquid-augmented power. The computation of liquid consumption must be based on the maximum approved rate appropriate for the desired engine output and must include the effect of temperature on engine performance as well as any other factors that might vary the amount of liquid required. (e) This section does not apply to fuel injection systems. [Amdt. 25-40, 42 FR 15043, Mar. 17, 1977, as amended by Amdt. 25-72, 55 FR 29785, July 20, 1990] |
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(a) Each fuel system must be constructed and arranged to ensure a flow of fuel at a rate and pressure established for proper engine and auxiliary power unit functioning under each likely operating condition, including any maneuver for which certification is requested and during which the engine or auxiliary power unit is permitted to be in operation. (b) Each fuel system must be arranged so that any air which is introduced into the system will not result in-- (1) Power interruption for more than 20 seconds for reciprocating engines; or (2) Flameout for turbine engines. (c) Each fuel system for a turbine engine must be capable of sustained operation throughout its flow and pressure range with fuel initially saturated with water at 80 deg. F and having 0.75cc of free water per gallon added and cooled to the most critical condition for icing likely to be encountered in operation. (d) Each fuel system for a turbine engine powered airplane must meet the applicable fuel venting requirements of part 34 of this chapter. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR 5677, Apr. 8, 1970; Amdt. 25-36, 39 FR 35460, Oct. 1, 1974; Amdt. 25-38, 41 FR 55467, Dec. 20, 1976; Amdt. 25-73, 55 FR 32861, Aug. 10, 1990; 55 FR 35139, Aug. 28, 1990] Sec. 25.952 Fuel system analysis and test. (a) Proper fuel system functioning under all probable operating conditions must be shown by analysis and those tests found necessary by the Administrator. Tests, if required, must be made using the airplane fuel system or a test article that reproduces the operating characteristics of the portion of the fuel system to be tested. (b) The likely failure of any heat exchanger using fuel as one of its fluids may not result in a hazardous condition. [Amdt. 25-40, 42 FR 15043, Mar. 17, 1977] Sec. 25.953 Fuel system independence. Each fuel system must meet the requirements of Sec. 25.903(b) by-- (a) Allowing the supply of fuel to each engine through a system independent of each part of the system supplying fuel to any other engine; or (b) Any other acceptable method. Sec. 25.954 Fuel system lightning protection. The fuel system must be designed and arranged to prevent the ignition of fuel vapor within the system by-- (a) Direct lightning strikes to areas having a high probability of stroke attachment; (b) Swept lightning strokes to areas where swept strokes are highly probable; and (c) Corona and streamering at fuel vent outlets. [Amdt. 25-14, 32 FR 11629, Aug. 11, 1967] Sec. 25.955 Fuel flow. (a) Each fuel system must provide at least 100 percent of the fuel flow required under each intended operating condition and maneuver. Compliance must be shown as follows: (1) Fuel must be delivered to each engine at a pressure within the limits specified in the engine type certificate. (2) The quantity of fuel in the tank may not exceed the amount established as the unusable fuel supply for that tank under the requirements of Sec. 25.959 plus that necessary to show compliance with this section. (3) Each main pump must be used that is necessary for each operating condition and attitude for which compliance with this section is shown, and the appropriate emergency pump must be substituted for each main pump so used. (4) If there is a fuel flowmeter, it must be blocked and the fuel must flow through the meter or its bypass. (b) If an engine can be supplied with fuel from more than one tank, the fuel system must-- (1) For each reciprocating engine, supply the full fuel pressure to that engine in not more than 20 seconds after switching to any other fuel tank containing usable fuel when engine malfunctioning becomes apparent due to the depletion of the fuel supply in any tank from which the engine can be fed; and (2) For each turbine engine, in addition to having appropriate manual switching capability, be designed to prevent interruption of fuel flow to that engine, without attention by the flight crew, when any tank supplying fuel to that engine is depleted of usable fuel during normal operation, and any other tank, that normally supplies fuel to that engine alone, contains usable fuel. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-11, 32 FR 6912, May 5, 1967] Sec. 25.957 Flow between interconnected tanks. If fuel can be pumped from one tank to another in flight, the fuel tank vents and the fuel transfer system must be designed so that no structural damage to the tanks can occur because of overfilling. Sec. 25.959 Unusable fuel supply. The unusable fuel quantity for each fuel tank and its fuel system components must be established at not less than the quantity at which the first evidence of engine malfunction occurs under the most adverse fuel feed condition for all intended operations and flight maneuvers involving fuel feeding from that tank. Fuel system component failures need not be considered. [Amdt. 25-23, 35 FR 5677, Apr. 8, 1970, as amended by Amdt. 25-40, 42 FR 15043, Mar. 17, 1977] Sec. 25.961 Fuel system hot weather operation. (a) The fuel system must perform satisfactorily in hot weather operation. This must be shown by showing that the fuel system from the tank outlets to each engine is pressurized, under all intended operations, so as to prevent vapor formation, or must be shown by climbing from the altitude of the airport elected by the applicant to the maximum altitude established as an operating limitation under Sec. 25.1527. If a climb test is elected, there may be no evidence of vapor lock or other malfunctioning during the climb test conducted under the following conditions: (1) For reciprocating engine powered airplanes, the engines must operate at maximum continuous power, except that takeoff power must be used for the altitudes from 1,000 feet below the critical altitude through the critical altitude. The time interval during which takeoff power is used may not be less than the takeoff time limitation. (2) For turbine engine powered airplanes, the engines must operate at takeoff power for the time interval selected for showing the takeoff flight path, and at maximum continuous power for the rest of the climb. (3) The weight of the airplane must be the weight with full fuel tanks, minimum crew, and the ballast necessary to maintain the center of gravity within allowable limits. (4) The climb airspeed may not exceed-- (i) For reciprocating engine powered airplanes, the maximum airspeed established for climbing from takeoff to the maximum operating altitude with the airplane in the following configuration: (A) Landing gear retracted. (B) Wing flaps in the most favorable position. (C) Cowl flaps (or other means of controlling the engine cooling supply) in the position that provides adequate cooling in the hot-day condition. (D) Engine operating within the maximum continuous power limitations. (E) Maximum takeoff weight; and (ii) For turbine engine powered airplanes, the maximum airspeed established for climbing from takeoff to the maximum operating altitude. (5) The fuel temperature must be at least 110 deg. F. (b) The test prescribed in paragraph (a) of this section may be performed in flight or on the ground under closely simulated flight conditions. If a flight test is performed in weather cold enough to interfere with the proper conduct of the test, the fuel tank surfaces, fuel lines, and other fuel system parts subject to cold air must be insulated to simulate, insofar as practicable, flight in hot weather. [Amdt. 25-11, 32 FR 6912, May 5, 1967, as amended by Amdt. 25-57, 49 FR 6848, Feb. 23, 1984] Sec. 25.963 Fuel tanks: general. (a) Each fuel tank must be able to withstand, without failure, the vibration, inertia, fluid, and structural loads that it may be subjected to in operation. (b) Flexible fuel tank liners must be approved or must be shown to be suitable for the particular application. (c) Integral fuel tanks must have facilities for interior inspection and repair. (d) Fuel tanks within the fuselage contour must be able to resist rupture and to retain fuel, under the inertia forces prescribed for the emergency landing conditions in Sec. 25.561. In addition, these tanks must be in a protected position so that exposure of the tanks to scraping action with the ground is unlikely. (e) Fuel tank access covers must comply with the following criteria in order to avoid loss of hazardous quantities of fuel: (1) All covers located in an area where experience or analysis indicates a strike is likely must be shown by analysis or tests to minimize penetration and deformation by tire fragments, low energy engine debris, or other likely debris. (2) All covers must be fire resistant as defined in part 1 of this chapter. (f) For pressurized fuel tanks, a means with fail-safe features must be provided to prevent the buildup of an excessive pressure difference between the inside and the outside of the tank. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-40, 42 FR 15043, Mar. 17, 1977; Amdt. 25-69, 54 FR 40352, Sept. 29, 1989] Sec. 25.965 Fuel tank tests. (a) It must be shown by tests that the fuel tanks, as mounted in the airplane, can withstand, without failure or leakage, the more critical of the pressures resulting from the conditions specified in paragraphs (a)(1) and (2) of this section. In addition, it must be shown by either analysis or tests, that tank surfaces subjected to more critical pressures resulting from the condition of paragraphs (a)(3) and (4) of this section, are able to withstand the following pressures: (1) An internal pressure of 3.5 psi. (2) 125 percent of the maximum air pressure developed in the tank from ram effect. (3) Fluid pressures developed during maximum limit accelerations, and deflections, of the airplane with a full tank. (4) Fluid pressures developed during the most adverse combination of airplane roll and fuel load. (b) Each metallic tank with large unsupported or unstiffened flat surfaces, whose failure or deformation could cause fuel leakage, must be able to withstand the following test, or its equivalent, without leakage or excessive deformation of the tank walls: (1) Each complete tank assembly and its supports must be vibration tested while mounted to simulate the actual installation. (2) Except as specified in paragraph (b)(4) of this section, the tank assembly must be vibrated for 25 hours at an amplitude of not less than 1/32 of an inch (unless another amplitude is substantiated) while 2/3 filled with water or other suitable test fluid. (3) The test frequency of vibration must be as follows: (i) If no frequency of vibration resulting from any r.p.m. within the normal operating range of engine speeds is critical, the test frequency of vibration must be 2,000 cycles per minute. (ii) If only one frequency of vibration resulting from any r.p.m. within the normal operating range of engine speeds is critical, that frequency of vibration must be the test frequency. (iii) If more than one frequency of vibration resulting from any r.p.m. within the normal operating range of engine speeds is critical, the most critical of these frequencies must be the test frequency. (4) Under paragraphs (b)(3) (ii) and (iii) of this section, the time of test must be adjusted to accomplish the same number of vibration cycles that would be accomplished in 25 hours at the frequency specified in paragraph (b)(3)(i) of this section. (5) During the test, the tank assembly must be rocked at the rate of 16 to 20 complete cycles per minute, through an angle of 15 deg. on both sides of the horizontal (30 deg. total), about the most critical axis, for 25 hours. If motion about more than one axis is likely to be critical, the tank must be rocked about each critical axis for 12 1/2 hours. (c) Except where satisfactory operating experience with a similar tank in a similar installation is shown, nonmetallic tanks must withstand the test specified in paragraph (b)(5) of this section, with fuel at a temperature of 110 deg. F. During this test, a representative specimen of the tank must be installed in a supporting structure simulating the installation in the airplane. (d) For pressurized fuel tanks, it must be shown by analysis or tests that the fuel tanks can withstand the maximum pressure likely to occur on the ground or in flight. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-11, 32 FR 6913, May 5, 1967; Amdt. 25-40, 42 FR 15043, Mar. 17, 1977] Sec. 25.967 Fuel tank installations. (a) Each fuel tank must be supported so that tank loads (resulting from the weight of the fuel in the tanks) are not concentrated on unsupported tank surfaces. In addition-- (1) There must be pads, if necessary, to prevent chafing between the tank and its supports; (2) Padding must be nonabsorbent or treated to prevent the absorption of fluids; (3) If a flexible tank liner is used, it must be supported so that it is not required to withstand fluid loads; and (4) Each interior surface of the tank compartment must be smooth and free of projections that could cause wear of the liner unless-- (i) Provisions are made for protection of the liner at these points; or (ii) The construction of the liner itself provides that protection. (b) Spaces adjacent to tank surfaces must be ventilated to avoid fume accumulation due to minor leakage. If the tank is in a sealed compartment, ventilation may be limited to drain holes large enough to prevent excessive pressure resulting from altitude changes. (c) The location of each tank must meet the requirements of Sec. 25.1185(a). (d) No engine nacelle skin immediately behind a major air outlet from the engine compartment may act as the wall of an integral tank. (e) Each fuel tank must be isolated from personnel compartments by a fumeproof and fuelproof enclosure. Sec. 25.969 Fuel tank expansion space. Each fuel tank must have an expansion space of not less than 2 percent of the tank capacity. It must be impossible to fill the expansion space inadvertently with the airplane in the normal ground attitude. For pressure fueling systems, compliance with this section may be shown with the means provided to comply with Sec. 25.979(b). [Amdt. 25-11, 32 FR 6913, May 5, 1967] Sec. 25.971 Fuel tank sump. (a) Each fuel tank must have a sump with an effective capacity, in the normal ground attitude, of not less than the greater of 0.10 percent of the tank capacity or one-sixteenth of a gallon unless operating limitations are established to ensure that the accumulation of water in service will not exceed the sump capacity. (b) Each fuel tank must allow drainage of any hazardous quantity of water from any part of the tank to its sump with the airplane in the ground attitude. (c) Each fuel tank sump must have an accessible drain that-- (1) Allows complete drainage of the sump on the ground; (2) Discharges clear of each part of the airplane; and (3) Has manual or automatic means for positive locking in the closed position. Sec. 25.973 Fuel tank filler connection. Each fuel tank filler connection must prevent the entrance of fuel into any part of the airplane other than the tank itself. In addition-- (a) [Reserved] (b) Each recessed filler connection that can retain any appreciable quantity of fuel must have a drain that discharges clear of each part of the airplane; (c) Each filler cap must provide a fuel-tight seal; and (d) Each fuel filling point, except pressure fueling connection points, must have a provision for electrically bonding the airplane to ground fueling equipment. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-40, 42 FR 15043, Mar. 17, 1977; Amdt. 25-72, 55 FR 29785, July 20, 1990] Sec. 25.975 Fuel tank vents and carburetor vapor vents. (a) Fuel tank vents. Each fuel tank must be vented from the top part of the expansion space so that venting is effective under any normal flight condition. In addition-- (1) Each vent must be arranged to avoid stoppage by dirt or ice formation; (2) The vent arrangement must prevent siphoning of fuel during normal operation; (3) The venting capacity and vent pressure levels must maintain acceptable differences of pressure between the interior and exterior of the tank, during-- (i) Normal flight operation; (ii) Maximum rate of ascent and descent; and (iii) Refueling and defueling (where applicable); (4) Airspaces of tanks with interconnected outlets must be interconnected; (5) There may be no point in any vent line where moisture can accumulate with the airplane in the ground attitude or the level flight attitude, unless drainage is provided; and (6) No vent or drainage provision may end at any point-- (i) Where the discharge of fuel from the vent outlet would constitute a fire hazard; or (ii) From which fumes could enter personnel compartments. (b) Carburetor vapor vents. Each carburetor with vapor elimination connections must have a vent line to lead vapors back to one of the fuel tanks. In addition-- (1) Each vent system must have means to avoid stoppage by ice; and (2) If there is more than one fuel tank, and it is necessary to use the tanks in a definite sequence, each vapor vent return line must lead back to the fuel tank used for takeoff and landing. Sec. 25.977 Fuel tank outlet. (a) There must be a fuel strainer for the fuel tank outlet or for the booster pump. This strainer must-- (1) For reciprocating engine powered airplanes, have 8 to 16 meshes per inch; and (2) For turbine engine powered airplanes, prevent the passage of any object that could restrict fuel flow or damage any fuel system component. (b) [Reserved] (c) The clear area of each fuel tank outlet strainer must be at least five times the area of the outlet line. (d) The diameter of each strainer must be at least that of the fuel tank outlet. (e) Each finger strainer must be accessible for inspection and cleaning. [Amdt. 25-11, 32 FR 6913, May 5, 1967, as amended by Amdt. 25-36, 39 FR 35460, Oct. 1, 1974] Sec. 25.979 Pressure fueling system. For pressure fueling systems, the following apply: (a) Each pressure fueling system fuel manifold connection must have means to prevent the escape of hazardous quantities of fuel from the system if the fuel entry valve fails. (b) An automatic shutoff means must be provided to prevent the quantity of fuel in each tank from exceeding the maximum quantity approved for that tank. This means must-- (1) Allow checking for proper shutoff operation before each fueling of the tank; and (2) Provide indication at each fueling station of failure of the shutoff means to stop the fuel flow at the maximum quantity approved for that tank. (c) A means must be provided to prevent damage to the fuel system in the event of failure of the automatic shutoff means prescribed in paragraph (b) of this section. (d) The airplane pressure fueling system (not including fuel tanks and fuel tank vents) must withstand an ultimate load that is 2.0 times the load arising from the maximum pressures, including surge, that is likely to occur during fueling. The maximum surge pressure must be established with any combination of tank valves being either intentionally or inadvertently closed. (e) The airplane defueling system (not including fuel tanks and fuel tank vents) must withstand an ultimate load that is 2.0 times the load arising from the maximum permissible defueling pressure (positive or negative) at the airplane fueling connection. [Amdt. 25-11, 32 FR 6913, May 5, 1967, as amended by Amdt. 25-38, 41 FR 55467, Dec. 20, 1976; Amdt. 25-72, 55 FR 29785, July 20, 1990] Sec. 25.981 Fuel tank temperature. (a) The highest temperature allowing a safe margin below the lowest expected auto ignition temperature of the fuel in the fuel tanks must be determined. (b) No temperature at any place inside any fuel tank where fuel ignition is possible may exceed the temperature determined under paragraph (a) of this section. This must be shown under all probable operating, failure, and malfunction conditions of any component whose operation, failure, or malfunction could increase the temperature inside the tank. [Amdt. 25-11, 32 FR 6913, May 5, 1967] |
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(a) Main pumps. Each fuel pump required for proper engine operation, or required to meet the fuel system requirements of this subpart (other than those in paragraph (b) of this section, is a main pump. For each main pump, provision must be made to allow the bypass of each positive displacement fuel pump other than a fuel injection pump (a pump that supplies the proper flow and pressure for fuel injection when the injection is not accomplished in a carburetor) approved as part of the engine. (b) Emergency pumps. There must be emergency pumps or another main pump to feed each engine immediately after failure of any main pump (other than a fuel injection pump approved as part of the engine). Sec. 25.993 Fuel system lines and fittings. (a) Each fuel line must be installed and supported to prevent excessive vibration and to withstand loads due to fuel pressure and accelerated flight conditions. (b) Each fuel line connected to components of the airplane between which relative motion could exist must have provisions for flexibility. (c) Each flexible connection in fuel lines that may be under pressure and subjected to axial loading must use flexible hose assemblies. (d) Flexible hose must be approved or must be shown to be suitable for the particular application. (e) No flexible hose that might be adversely affected by exposure to high temperatures may be used where excessive temperatures will exist during operation or after engine shut-down. (f) Each fuel line within the fuselage must be designed and installed to allow a reasonable degree of deformation and stretching without leakage. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-15, 32 FR 13266, Sept. 20, 1967] Sec. 25.994 Fuel system components. Fuel system components in an engine nacelle or in the fuselage must be protected from damage which could result in spillage of enough fuel to constitute a fire hazard as a result of a wheels-up landing on a paved runway. [Amdt. 25-57, 49 FR 6848, Feb. 23, 1984] Sec. 25.995 Fuel valves. In addition to the requirements of Sec. 25.1189 for shutoff means, each fuel valve must-- (a) [Reserved] (b) Be supported so that no loads resulting from their operation or from accelerated flight conditions are transmitted to the lines attached to the valve. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-40, 42 FR 15043, Mar. 17, 1977] Sec. 25.997 Fuel strainer or filter. There must be a fuel strainer or filter between the fuel tank outlet and the inlet of either the fuel metering device or an engine driven positive displacement pump, whichever is nearer the fuel tank outlet. This fuel strainer or filter must-- (a) Be accessible for draining and cleaning and must incorporate a screen or element which is easily removable; (b) Have a sediment trap and drain except that it need not have a drain if the strainer or filter is easily removable for drain purposes; (c) Be mounted so that its weight is not supported by the connecting lines or by the inlet or outlet connections of the strainer or filter itself, unless adequate strength margins under all loading conditions are provided in the lines and connections; and (d) Have the capacity (with respect to operating limitations established for the engine) to ensure that engine fuel system functioning is not impaired, with the fuel contaminated to a degree (with respect to particle size and density) that is greater than that established for the engine in Part 33 of this chapter. [Amdt. No. 25-36, 39 FR 35460, Oct. 1, 1974, as amended by Amdt. 25-57, 49 FR 6848, Feb. 23, 1984] Sec. 25.999 Fuel system drains. (a) Drainage of the fuel system must be accomplished by the use of fuel strainer and fuel tank sump drains. (b) Each drain required by paragraph (a) of this section must-- (1) Discharge clear of all parts of the airplane; (2) Have manual or automatic means for positive locking in the closed position; and (3) Have a drain valve-- (i) That is readily accessible and which can be easily opened and closed; and (ii) That is either located or protected to prevent fuel spillage in the event of a landing with landing gear retracted. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 41 FR 55467, Dec. 20, 1976] Sec. 25.1001 Fuel jettisoning system. (a) A fuel jettisoning system must be installed on each airplane unless it is shown that the airplane meets the climb requirements of Secs. 25.119 and 25.121(d) at maximum takeoff weight, less the actual or computed weight of fuel necessary for a 15-minute flight comprised of a takeoff, go-around, and landing at the airport of departure with the airplane configuration, speed, power, and thrust the same as that used in meeting the applicable takeoff, approach, and landing climb performance requirements of this part. (b) If a fuel jettisoning system is required it must be capable of jettisoning enough fuel within 15 minutes, starting with the weight given in paragraph (a) of this section, to enable the airplane to meet the climb requirements of Secs. 25.119 and 25.121(d), assuming that the fuel is jettisoned under the conditions, except weight, found least favorable during the flight tests prescribed in paragraph (c) of this section. (c) Fuel jettisoning must be demonstrated beginning at maximum takeoff weight with flaps and landing gear up and in-- (1) A power-off glide at 1.4 Vs1; (2) A climb at the one-engine inoperative best rate-of-climb speed, with the critical engine inoperative and the remaining engines at maximum continuous power; and (3) Level flight at 1.4 Vs1; if the results of the tests in the conditions specified in paragraphs (c) (1) and (2) of this section show that this condition could be critical. (d) During the flight tests prescribed in paragraph (c) of this section, it must be shown that-- (1) The fuel jettisoning system and its operation are free from fire hazard; (2) The fuel discharges clear of any part of the airplane; (3) Fuel or fumes do not enter any parts of the airplane; and (4) The jettisoning operation does not adversely affect the controllability of the airplane. (e) For reciprocating engine powered airplanes, means must be provided to prevent jettisoning the fuel in the tanks used for takeoff and landing below the level allowing 45 minutes flight at 75 percent maximum continuous power. However, if there is an auxiliary control independent of the main jettisoning control, the system may be designed to jettison the remaining fuel by means of the auxiliary jettisoning control. (f) For turbine engine powered airplanes, means must be provided to prevent jettisoning the fuel in the tanks used for takeoff and landing below the level allowing climb from sea level to 10,000 feet and thereafter allowing 45 minutes cruise at a speed for maximum range. However, if there is an auxiliary control independent of the main jettisoning control, the system may be designed to jettison the remaining fuel by means of the auxiliary jettisoning control. (g) The fuel jettisoning valve must be designed to allow flight personnel to close the valve during any part of the jettisoning operation. (h) Unless it is shown that using any means (including flaps, slots, and slats) for changing the airflow across or around the wings does not adversely affect fuel jettisoning, there must be a placard, adjacent to the jettisoning control, to warn flight crewmembers against jettisoning fuel while the means that change the airflow are being used. (i) The fuel jettisoning system must be designed so that any reasonably probable single malfunction in the system will not result in a hazardous condition due to unsymmetrical jettisoning of, or inability to jettison, fuel. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25-18, 33 FR 12226, Aug. 30, 1968; Amdt.
25-57, 49 FR 6848, Feb. 23, 1984] |
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(a) Each engine must have an independent oil system that can supply it with an appropriate quantity of oil at a temperature not above that safe for continuous operation. (b) The usable oil capacity may not be less than the product of the endurance of the airplane under critical operating conditions and the approved maximum allowable oil consumption of the engine under the same conditions, plus a suitable margin to ensure system circulation. Instead of a rational analysis of airplane range for the purpose of computing oil requirements for reciprocating engine powered airplanes, the following fuel/ oil ratios may be used: (1) For airplanes without a reserve oil or oil transfer system, a fuel/oilratio of 30:1 by volume. (2) For airplanes with either a reserve oil or oil transfer system, a fuel/ oil ratio of 40:1 by volume. (c) Fuel/oil ratios higher than those prescribed in paragraphs (b) (1) and (2) of this section may be used if substantiated by data on actual engine oil consumption. Sec. 25.1013 Oil tanks. (a) Installation. Each oil tank installation must meet the requirements of Sec. 25.967. (b) Expansion space. Oil tank expansion space must be provided as follows: (1) Each oil tank used with a reciprocating engine must have an expansion space of not less than the greater of 10 percent of the tank capacity or 0.5 gallon, and each oil tank used with a turbine engine must have an expansion space of not less than 10 percent of the tank capacity. (2) Each reserve oil tank not directly connected to any engine may have an expansion space of not less than two percent of the tank capacity. (3) It must be impossible to fill the expansion space inadvertently with the airplane in the normal ground attitude. (c) Filler connection. Each recessed oil tank filler connection that can retain any appreciable quantity of oil must have a drain that discharges clear of each part of the airplane. In addition, each oil tank filler cap must provide an oil-tight seal. (d) Vent. Oil tanks must be vented as follows: (1) Each oil tank must be vented from the top part of the expansion space so that venting is effective under any normal flight condition. (2) Oil tank vents must be arranged so that condensed water vapor that might freeze and obstruct the line cannot accumulate at any point. (e) Outlet. There must be means to prevent entrance into the tank itself, or into the tank outlet, of any object that might obstruct the flow of oil through the system. No oil tank outlet may be enclosed by any screen or guard that would reduce the flow of oil below a safe value at any operating temperature. There must be a shutoff valve at the outlet of each oil tank used with a turbine engine, unless the external portion of the oil system (including the oil tank supports) is fireproof. (f) Flexible oil tank liners. Each flexible oil tank liner must be approved or must be shown to be suitable for the particular application. [Doc. No. 5066, 29 FR 18291, Dec. 24, as amended by Amdt. 25-19, 33 FR 15410, Oct. 17, 1968; Amdt. 25-23, 35 FR 5677, Apr. 8, 1970; Amdt. 25-36, 39 FR 35460, Oct. 1, 1974; Amdt. 25-57, 49 FR 6848, Feb. 23, 1984; Amdt. 25-72, 55 FR 29785, July 20, 1990] Sec. 25.1015 Oil tank tests. Each oil tank must be designed and installed so that-- (a) It can withstand, without failure, each vibration, inertia, and fluid load that it may be subjected to in operation; and (b) It meets the provisions of Sec. 25.965, except-- (1) The test pressure-- (i) For pressurized tanks used with a turbine engine, may not be less than 5p.s.i. plus the maximum operating pressure of the tank instead of the pressure specified in Sec. 25.965(a); and (ii) For all other tanks may not be less than 5 p.s.i. instead of the pressure specified in Sec. 25.965(a); and (2) The test fluid must be oil at 250 deg. F. instead of the fluid specified in Sec. 25.965(c). [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-36, 39 FR 35461, Oct. 1, 1974] Sec. 25.1017 Oil lines and fittings. (a) Each oil line must meet the requirements of Sec. 25.993 and each oil line and fitting in any designated fire zone must meet the requirements of Sec. 25.1183. (b) Breather lines must be arranged so that-- (1) Condensed water vapor that might freeze and obstruct the line cannot accumulate at any point; (2) The breather discharge does not constitute a fire hazard if foaming occurs or causes emitted oil to strike the pilot's windshield; and (3) The breather does not discharge into the engine air induction system. Sec. 25.1019 Oil strainer or filter. (a) Each turbine engine installation must incorporate an oil strainer or filter through which all of the engine oil flows and which meets the following requirements: (1) Each oil strainer or filter that has a bypass must be constructed and installed so that oil will flow at the normal rate through the rest of the system with the strainer or filter completely blocked. (2) The oil strainer or filter must have the capacity (with respect to operating limitations established for the engine) to ensure that engine oil system functioning is not impaired when the oil is contaminated to a degree (with respect to particle size and density) that is greater than that established for the engine under Part 33 of this chapter. (3) The oil strainer or filter, unless it is installed at an oil tank outlet, must incorporate an indicator that will indicate contamination before it reaches the capacity established in accordance with paragraph (a)(2) of this section. (4) The bypass of a strainer or filter must be constructed and installed so that the release of collected contaminants is minimized by appropriate location of the bypass to ensure that collected contaminants are not in the bypass flow path. (5) An oil strainer or filter that has no bypass, except one that is installed at an oil tank outlet, must have a means to connect it to the warning system required in Sec. 25.1305(c)(7). (b) Each oil strainer or filter in a powerplant installation using reciprocating engines must be constructed and installed so that oil will flow at the normal rate through the rest of the system with the strainer or filter element completely blocked. [Amdt. 25-36, 39 FR 35461, Oct. 1, 1974, as amended by Amdt. 25-57, 49 FR 6848, Feb. 23, 1984] Sec. 25.1021 Oil system drains. A drain (or drains) must be provided to allow safe drainage of the oil system. Each drain must-- (a) Be accessible; and (b) Have manual or automatic means for positive locking in the closed position. [Amdt. 25-57, 49 FR 6848, Feb. 23, 1984] Sec. 25.1023 Oil radiators. (a) Each oil radiator must be able to withstand, without failure, any vibration, inertia, and oil pressure load to which it would be subjected in operation. (b) Each oil radiator air duct must be located so that, in case of fire, flames coming from normal openings of the engine nacelle cannot impinge directly upon the radiator. Sec. 25.1025 Oil valves. (a) Each oil shutoff must meet the requirements of Sec. 25.1189. (b) The closing of oil shutoff means may not prevent propeller feathering. (c) Each oil valve must have positive stops or suitable index provisions in the "on" and "off" positions and must be supported so that no loads resulting from its operation or from accelerated flight conditions are transmitted to the lines attached to the valve. Sec. 25.1027 Propeller feathering system. (a) If the propeller feathering system depends on engine oil, there must be means to trap an amount of oil in the tank if the supply becomes depleted due to failure of any part of the lubricating system other than the tank itself. (b) The amount of trapped oil must be enough to accomplish the feathering operation and must be available only to the feathering pump. (c) The ability of the system to accomplish feathering with the trapped oil must be shown. This may be done on the ground using an auxiliary source of oil for lubricating the engine during operation. (d) Provision must be made to prevent sludge or other foreign matter from affecting the safe operation of the propeller feathering system. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25-38, 41 FR 55467, Dec. 20, 1976] |
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The powerplant and auxiliary power unit cooling provisions must be able to maintain the temperatures of powerplant components, engine fluids, and auxiliary power unit components and fluids within the temperature limits established for these components and fluids, under ground, water, and flight operating conditions, and after normal engine or auxiliary power unit shutdown, or both. [Amdt. 25-38, 41 FR 55467, Dec. 20, 1976] Sec. 25.1043 Cooling tests. (a) General. Compliance with Sec. 25.1041 must be shown by tests, under critical ground, water, and flight operating conditions. For these tests, the following apply: (1) If the tests are conducted under conditions deviating from the maximum ambient atmospheric temperature, the recorded powerplant temperatures must be corrected under paragraphs (c) and (d) of this section. (2) No corrected temperatures determined under paragraph (a)(1) of this section may exceed established limits. (3) For reciprocating engines, the fuel used during the cooling tests must be the minimum grade approved for the engines, and the mixture settings must be those normally used in the flight stages for which the cooling tests are conducted. The test procedures must be as prescribed in Sec. 25.1045. (b) Maximum ambient atmospheric temperature. A maximum ambient atmospheric temperature corresponding to sea level conditions of at least 100 degrees F must be established. The assumed temperature lapse rate is 3.6 degrees F per thousand feet of altitude above sea level until a temperature of -69.7 degrees F is reached, above which altitude the temperature is considered constant at -69.7 degrees F. However, for winterization installations, the applicant may select a maximum ambient atmospheric temperature corresponding to sea level conditions of less than 100 degrees F. (c) Correction factor (except cylinder barrels). Unless a more rational correction applies, temperatures of engine fluids and powerplant components (except cylinder barrels) for which temperature limits are established, must be corrected by adding to them the difference between the maximum ambient atmospheric temperature and the temperature of the ambient air at the time of the first occurrence of the maximum component or fluid temperature recorded during the cooling test. (d) Correction factor for cylinder barrel temperatures. Unless a more rational correction applies, cylinder barrel temperatures must be corrected by adding to them 0.7 times the difference between the maximum ambient atmospheric temperature and the temperature of the ambient air at the time of the first occurrence of the maximum cylinder barrel temperature recorded during the cooling test. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-42, 43 FR 2323, Jan. 16, 1978] Sec. 25.1045 Cooling test procedures. (a) Compliance with Sec. 25.1041 must be shown for the takeoff, climb, en route, and landing stages of flight that correspond to the applicable performance requirements. The cooling tests must be conducted with the airplane in the configuration, and operating under the conditions, that are critical relative to cooling during each stage of flight. For the cooling tests, a temperature is "stabilized" when its rate of change is less than two degrees F. per minute. (b) Temperatures must be stabilized under the conditions from which entry is made into each stage of flight being investigated, unless the entry condition normally is not one during which component and the engine fluid temperatures would stabilize (in which case, operation through the full entry condition must be conducted before entry into the stage of flight being investigated in order to allow temperatures to reach their natural levels at the time of entry). The takeoff cooling test must be preceded by a period during which the powerplant component and engine fluid temperatures are stabilized with the engines at ground idle. (c) Cooling tests for each stage of flight must be continued until-- (1) The component and engine fluid temperatures stabilize; (2) The stage of flight is completed; or (3) An operating limitation is reached. (d) For reciprocating engine powered airplanes, it may be assumed, for cooling test purposes, that the takeoff stage of flight is complete when the airplane reaches an altitude of 1,500 feet above the takeoff surface or reaches a point in the takeoff where the transition from the takeoff to the en route configuration is completed and a speed is reached at which compliance with Sec. 25.121(c) is shown, whichever point is at a higher altitude. The airplane must be in the following configuration: (1) Landing gear retracted. (2) Wing flaps in the most favorable position. (3) Cowl flaps (or other means of controlling the engine cooling supply) in the position that provides adequate cooling in the hot-day condition. (4) Critical engine inoperative and its propeller stopped. (5) Remaining engines at the maximum continuous power available for the altitude. (e) For hull seaplanes and amphibians, cooling must be shown during taxiing downwind for 10 minutes, at five knots above step speed. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as
amended by Amdt. 25-57, 49 FR 6848, Feb. 23, 1984] |
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(a) The air induction system for each engine and auxiliary power unit must supply-- (1) The air required by that engine and auxiliary power unit under each operating condition for which certification is requested; and (2) The air for proper fuel metering and mixture distribution with the induction system valves in any position. (b) Each reciprocating engine must have an alternate air source that prevents the entry of rain, ice, or any other foreign matter. (c) Air intakes may not open within the cowling, unless-- (1) That part of the cowling is isolated from the engine accessory section by means of a fireproof diaphragm; or (2) For reciprocating engines, there are means to prevent the emergence of backfire flames. (d) For turbine engine powered airplanes and airplanes incorporating auxiliary power units-- (1) There must be means to prevent hazardous quantities of fuel leakage or overflow from drains, vents, or other components of flammable fluid systems from entering the engine or auxiliary power unit intake system; and (2) The airplane must be designed to prevent water or slush on the runway, taxiway, or other airport operating surfaces from being directed into the engine or auxiliary power unit air inlet ducts in hazardous quantities, and the air inlet ducts must be located or protected so as to minimize the ingestion of foreign matter during takeoff, landing, and taxiing. (e) If the engine induction system contains parts or components that could be damaged by foreign objects entering the air inlet, it must be shown by tests or, if appropriate, by analysis that the induction system design can withstand the foreign object ingestion test conditions of Sec. 33.77 of this chapter without failure of parts or components that could create a hazard. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 41 FR 55467, Dec. 20, 1976; Amdt. 25-40, 42 FR 15043, Mar. 17, 1977; Amdt. 25-57, 49 FR 6849, Feb. 23, 1984] Sec. 25.1093 Induction system icing protection. (a) Reciprocating engines. Each reciprocating engine air induction system must have means to prevent and eliminate icing. Unless this is done by other means, it must be shown that, in air free of visible moisture at a temperature of 30 deg. F., each airplane with altitude engines using-- (1) Conventional venturi carburetors have a preheater that can provide a heat rise of 120 deg. F. with the engine at 60 percent of maximum continuous power; or (2) Carburetors tending to reduce the probability of ice formation has a preheater that can provide a heat rise of 100 deg. F. with the engine at 60 percent of maximum continuous power. (b) Turbine engines. (1) Each turbine engine must operate throughout the flight power range of the engine (including idling), without the accumulation of ice on the engine, inlet system components, or airframe components that would adversely affect engine operation or cause a serious loss of power or thrust-- (i) Under the icing conditions specified in appendix C, and (ii) In falling and blowing snow within the limitations established for the airplane for such operation. (2) Each turbine engine must idle for 30 minutes on the ground, with the air bleed available for engine icing protection at its critical condition, without adverse effect, in an atmosphere that is at a temperature between 15 deg. and 30 deg. F (between -9 deg. and -1 deg. C) and has a liquid water content not less than 0.3 grams per cubic meter in the form of drops having a mean effective diameter not less than 20 microns, followed by momentary operation at takeoff power or thrust. During the 30 minutes of idle operation, the engine may be run up periodically to a moderate power or thrust setting in a manner acceptable to the Administrator. (c) Supercharged reciprocating engines. For each engine having a supercharger to pressurize the air before it enters the carburetor, the heat rise in the air caused by that supercharging at any altitude may be utilized in determining compliance with paragraph (a) of this section if the heat rise utilized is that which will be available, automatically, for the applicable altitude and operating condition because of supercharging. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 41 FR 55467, Dec. 20, 1976; Amdt. 25-40, 42 FR 15043, Mar. 17, 1977; Amdt. 25-57, 49 FR 6849, Feb. 23, 1984; Amdt. 25-72, 55 FR 29785, July 20, 1990] Sec. 25.1101 Carburetor air preheater design. Each carburetor air preheater must be designed and constructed to-- (a) Ensure ventilation of the preheater when the engine is operated in cold air; (b) Allow inspection of the exhaust manifold parts that it surrounds; and (c) Allow inspection of critical parts of the preheater itself. Sec. 25.1103 Induction system ducts and air duct systems. (a) Each induction system duct upstream of the first stage of the engine supercharger and of the auxiliary power unit compressor must have a drain to prevent the hazardous accumulation of fuel and moisture in the ground attitude. No drain may discharge where it might cause a fire hazard. (b) Each induction system duct must be-- (1) Strong enough to prevent induction system failures resulting from normal backfire conditions; and (2) Fire-resistant if it is in any fire zone for which a fire-extinguishing system is required, except that ducts for auxiliary power units must be fireproof within the auxiliary power unit fire zone. (c) Each duct connected to components between which relative motion could exist must have means for flexibility. (d) For turbine engine and auxiliary power unit bleed air duct systems, no hazard may result if a duct failure occurs at any point between the air duct source and the airplane unit served by the air. (e) Each auxiliary power unit induction system duct must be fireproof for a sufficient distance upstream of the auxiliary power unit compartment to prevent hot gas reverse flow from burning through auxiliary power unit ducts and entering any other compartment or area of the airplane in which a hazard would be created resulting from the entry of hot gases. The materials used to form the remainder of the induction system duct and plenum chamber of the auxiliary power unit must be capable of resisting the maximum heat conditions likely to occur. (f) Each auxiliary power unit induction system duct must be constructed of materials that will not absorb or trap hazardous quantities of flammable fluids that could be ignited in the event of a surge or reverse flow condition. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-46, 43 FR 50597, Oct. 30, 1978] Sec. 25.1105 Induction system screens. If induction system screens are used-- (a) Each screen must be upstream of the carburetor; (b) No screen may be in any part of the induction system that is the only passage through which air can reach the engine, unless it can be deiced by heated air; (c) No screen may be deiced by alcohol alone; and (d) It must be impossible for fuel to strike any screen. Sec. 25.1107 Inter-coolers and after-coolers. Each inter-cooler and after-cooler must be able to
withstand any vibration, inertia, and air pressure load to
which it would be subjected in operation. |
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For powerplant and auxiliary power unit installations the following apply: (a) Each exhaust system must ensure safe disposal of exhaust gases without fire hazard or carbon monoxide contamination in any personnel compartment. For test purposes, any acceptable carbon monoxide detection method may be used to show the absence of carbon monoxide. (b) Each exhaust system part with a surface hot enough to ignite flammable fluids or vapors must be located or shielded so that leakage from any system carrying flammable fluids or vapors will not result in a fire caused by impingement of the fluids or vapors on any part of the exhaust system including shields for the exhaust system. (c) Each component that hot exhaust gases could strike, or that could be subjected to high temperatures from exhaust system parts, must be fireproof. All exhaust system components must be separated by fireproof shields from adjacent parts of the airplane that are outside the engine and auxiliary power unit compartments. (d) No exhaust gases may discharge so as to cause a fire hazard with respect to any flammable fluid vent or drain. (e) No exhaust gases may discharge where they will cause a glare seriously affecting pilot vision at night. (f) Each exhaust system component must be ventilated to prevent points of excessively high temperature. (g) Each exhaust shroud must be ventilated or insulated to avoid, during normal operation, a temperature high enough to ignite any flammable fluids or vapors external to the shroud. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-40, 42 FR 15043, Mar. 17, 1977] Sec. 25.1123 Exhaust piping. For powerplant and auxiliary power unit installations, the following apply: (a) Exhaust piping must be heat and corrosion resistant, and must have provisions to prevent failure due to expansion by operating temperatures. (b) Piping must be supported to withstand any vibration and inertia loads to which it would be subjected in operation; and (c) Piping connected to components between which relative motion could exist must have means for flexibility. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-40, 42 FR 15044, Mar. 17, 1977] Sec. 25.1125 Exhaust heat exchangers. For reciprocating engine powered airplanes, the following apply: (a) Each exhaust heat exchanger must be constructed and installed to withstand each vibration, inertia, and other load to which it would be subjected in operation. In addition-- (1) Each exchanger must be suitable for continued operation at high temperatures and resistant to corrosion from exhaust gases; (2) There must be means for the inspection of the critical parts of each exchanger; (3) Each exchanger must have cooling provisions wherever it is subject to contact with exhaust gases; and (4) No exhaust heat exchanger or muff may have any stagnant areas or liquid traps that would increase the probability of ignition of flammable fluids or vapors that might be present in case of the failure or malfunction of components carrying flammable fluids. (b) If an exhaust heat exchanger is used for heating ventilating air- (1) There must be a secondary heat exchanger between the primary exhaust gas heat exchanger and the ventilating air system; or (2) Other means must be used to preclude the harmful contamination of the ventilating air. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-38, 41 FR 55467, Dec. 20, 1976] Sec. 25.1127 Exhaust driven turbo-superchargers. (a) Each exhaust driven turbo-supercharger must be approved or shown to be suitable for the particular application. It must be installed and supported to ensure safe operation between normal inspections and overhauls. In addition, there must be provisions for expansion and flexibility between exhaust conduits and the turbine. (b) There must be provisions for lubricating the turbine and for cooling turbine parts where temperatures are critical. (c) If the normal turbo-supercharger control system
malfunctions, the turbine speed may not exceed its maximum
allowable value. Except for the waste gate operating
components, the components provided for meeting this
requirement must be independent of the normal
turbo-supercharger controls. |
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Each powerplant control must be located, arranged, and designed under Secs. 25.777 through 25.781 and marked under Sec. 25.1555. In addition, it must meet the following requirements: (a) Each control must be located so that it cannot be inadvertently operated by persons entering, leaving, or moving normally in, the cockpit. (b) Each flexible control must be approved or must be shown to be suitable for the particular application. (c) Each control must have sufficient strength and rigidity to withstand operating loads without failure and without excessive deflection. (d) Each control must be able to maintain any set position without constant attention by flight crewmembers and without creep due to control loads or vibration. (e) The portion of each powerplant control located in a designated fire zone that is required to be operated in the event of fire must be at least fire resistant. (f) Powerplant valve controls located in the cockpit must have-- (1) For manual valves, positive stops or in the case of fuel valves suitable index provisions, in the open and closed position; and (2) For power-assisted valves, a means to indicate to the flight crew when the valve-- (i) Is in the fully open or fully closed position; or (ii) Is moving between the fully open and fully closed position. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-40, 42 FR 15044, Mar. 17, 1977; Amdt. 25-72, 55 FR 29785, July 20, 1990] Sec. 25.1142 Auxiliary power unit controls. Means must be provided on the flight deck for starting, stopping, and emergency shutdown of each installed auxiliary power unit. [Amdt. 25-46, 43 FR 50598, Oct. 30, 1978] Sec. 25.1143 Engine controls. (a) There must be a separate power or thrust control for each engine. (b) Power and thrust controls must be arranged to allow-- (1) Separate control of each engine; and (2) Simultaneous control of all engines. (c) Each power and thrust control must provide a positive and immediately responsive means of controlling its engine. (d) For each fluid injection (other than fuel) system and its controls not provided and approved as part of the engine, the applicant must show that the flow of the injection fluid is adequately controlled. (e) If a power or thrust control incorporates a fuel shutoff feature, the control must have a means to prevent the inadvertent movement of the control into the shutoff position. The means must-- (1) Have a positive lock or stop at the idle position; and (2) Require a separate and distinct operation to place the control in the shutoff position. [Amdt. 25-23, 35 FR 5677, Apr. 8, 1970, as amended by Amdt. 25-38, 41 FR 55467, Dec. 20, 1976; Amdt. 25-57, 49 FR 6849, Feb. 23, 1984] Sec. 25.1145 Ignition switches. (a) Ignition switches must control each engine ignition circuit on each engine. (b) There must be means to quickly shut off all ignition by the grouping of switches or by a master ignition control. (c) Each group of ignition switches, except ignition switches for turbine engines for which continuous ignition is not required, and each master ignition control must have a means to prevent its inadvertent operation. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-40, 42 FR 15044 Mar. 17, 1977] Sec. 25.1147 Mixture controls. (a) If there are mixture controls, each engine must have a separate control. The controls must be grouped and arranged to allow-- (1) Separate control of each engine; and (2) Simultaneous control of all engines. (b) Each intermediate position of the mixture controls that corresponds to anormal operating setting must be identifiable by feel and sight. (c) The mixture controls must be accessible to both pilots. However, if there is a separate flight engineer station with a control panel, the controls need be accessible only to the flight engineer. Sec. 25.1149 Propeller speed and pitch controls. (a) There must be a separate propeller speed and pitch control for each propeller. (b) The controls must be grouped and arranged to allow-- (1) Separate control of each propeller; and (2) Simultaneous control of all propellers. (c) The controls must allow synchronization of all propellers. (d) The propeller speed and pitch controls must be to the right of, and at least one inch below, the pilot's throttle controls. Sec. 25.1153 Propeller feathering controls. (a) There must be a separate propeller feathering control for each propeller. The control must have means to prevent its inadvertent operation. (b) If feathering is accomplished by movement of the propeller pitch or speed control lever, there must be means to prevent the inadvertent movement of this lever to the feathering position during normal operation. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-11, 32 FR 6913, May 5, 1967] Sec. 25.1155 Reverse thrust and propeller pitch settings below the flight regime. Each control for reverse thrust and for propeller pitch settings below the flight regime must have means to prevent its inadvertent operation. The means must have a positive lock or stop at the flight idle position and must require a separate and distinct operation by the crew to displace the control from the flight regime (forward thrust regime for turbojet powered airplanes). [Amdt. 25-11, 32 FR 6913, May 5, 1967] Sec. 25.1157 Carburetor air temperature controls. There must be a separate carburetor air temperature control for each engine. Sec. 25.1159 Supercharger controls. Each supercharger control must be accessible to the pilots or, if there is aseparate flight engineer station with a control panel, to the flight engineer. Sec. 25.1161 Fuel jettisoning system controls. Each fuel jettisoning system control must have guards to prevent inadvertent operation. No control may be near any fire extinguisher control or other control used to combat fire. Sec. 25.1163 Powerplant accessories. (a) Each engine mounted accessory must-- (1) Be approved for mounting on the engine involved; (2) Use the provisions on the engine for mounting; and (3) Be sealed to prevent contamination of the engine oil system and the accessory system. (b) Electrical equipment subject to arcing or sparking must be installed to minimize the probability of contact with any flammable fluids or vapors that might be present in a free state. (c) If continued rotation of an engine-driven cabin supercharger or of any remote accessory driven by the engine is hazardous if malfunctioning occurs, there must be means to prevent rotation without interfering with the continued operation of the engine. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-57, 49 FR 6849, Feb. 23, 1984] Sec. 25.1165 Engine ignition systems. (a) Each battery ignition system must be supplemented by a generator that is automatically available as an alternate source of electrical energy to allow continued engine operation if any battery becomes depleted. (b) The capacity of batteries and generators must be large enough to meet the simultaneous demands of the engine ignition system and the greatest demands of any electrical system components that draw electrical energy from the same source. (c) The design of the engine ignition system must account for-- (1) The condition of an inoperative generator; (2) The condition of a completely depleted battery with the generator running at its normal operating speed; and (3) The condition of a completely depleted battery with the generator operating at idling speed, if there is only one battery. (d) Magneto ground wiring (for separate ignition circuits) that lies on the engine side of the fire wall, must be installed, located, or protected, to minimize the probability of simultaneous failure of two or more wires as a result of mechanical damage, electrical faults, or other cause. (e) No ground wire for any engine may be routed through a fire zone of another engine unless each part of that wire within that zone is fireproof. (f) Each ignition system must be independent of any electrical circuit, not used for assisting, controlling, or analyzing the operation of that system. (g) There must be means to warn appropriate flight crewmembers if the malfunctioning of any part of the electrical system is causing the continuous discharge of any battery necessary for engine ignition. (h) Each engine ignition system of a turbine powered airplane must be considered an essential electrical load. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR 5677, Apr. 8, 1970; Amdt. 25-72, 55 FR 29785, July 20, 1990] Sec. 25.1167 Accessory gearboxes. For airplanes equipped with an accessory gearbox that is not certificated as part of an engine-- (a) The engine with gearbox and connecting transmissions and shafts attached must be subjected to the tests specified in Sec. 33.49 or Sec. 33.87 of this chapter, as applicable; (b) The accessory gearbox must meet the requirements of Secs. 33.25 and 33.53 or 33.91 of this chapter, as applicable; & (c) Possible misalignments and torsional loadings of the gearbox, transmission, and shaft system, expected to result under normal operating conditions must be evaluated. [Amdt. 25-38, 41 FR 55467, Dec. 20, 1976] |
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(a) Designated fire zones are-- (1) The engine power section; (2) The engine accessory section; (3) Except for reciprocating engines, any complete powerplant compartment in which no isolation is provided between the engine power section and the engine accessory section; (4) Any auxiliary power unit compartment; (5) Any fuel-burning heater and other combustion equipment installation described in Sec. 25.859; (6) The compressor and accessory sections of turbine engines; and (7) Combustor, turbine, and tailpipe sections of turbine engine installations that contain lines or components carrying flammable fluids or gases. (b) Each designated fire zone must meet the requirements of Secs. 25.867, and 25.1185 through 25.1203. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-11, 32 FR 6913, May 5, 1967; Amdt. 25-23, 35 FR 5677, Apr. 8, 1970; Amdt. 25-72, 55 FR 29785, July 20, 1990] Sec. 25.1182 Nacelle areas behind firewalls, and engine pod attaching structures containing flammable fluid lines. (a) Each nacelle area immediately behind the firewall, and each portion of any engine pod attaching structure containing flammable fluid lines, must meet each requirement of Secs. 25.1103(b), 25.1165 (d) and (e), 25.1183, 25.1185(c), 25.1187, 25.1189, and 25.1195 through 25.1203, including those concerning designated fire zones. However, engine pod attaching structures need not contain fire detection or extinguishing means. (b) For each area covered by paragraph (a) of this section that contains a retractable landing gear, compliance with that paragraph need only be shown with the landing gear retracted. [Amdt. 25-11, 32 FR 6913, May 5, 1967] Sec. 25.1183 Flammable fluid-carrying components. (a) Except as provided in paragraph (b) of this section, each line, fitting, and other component carrying flammable fluid in any area subject to engine fire conditions, and each component which conveys or contains flammable fluid in a designated fire zone must be fire resistant, except that flammable fluid tanks and supports in a designated fire zone must be fireproof or be enclosed by a fireproof shield unless damage by fire to any non-fireproof part will not cause leakage or spillage of flammable fluid. Components must be shielded or located to safeguard against the ignition of leaking flammable fluid. An integral oil sump of less than 25-quart capacity on a reciprocating engine need not be fireproof nor be enclosed by a fireproof shield. (b) Paragraph (a) of this section does not apply to-- (1) Lines, fittings, and components which are already approved as part of a type certificated engine; and (2) Vent and drain lines, and their fittings, whose failure will not result in, or add to, a fire hazard. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-11, 32 FR 6913, May 5, 1967; Amdt. 25-36, 39 FR 35461, Oct. 1, 1974; Amdt. 25-57, 49 FR 6849, Feb. 23, 1984] Sec. 25.1185 Flammable fluids. (a) Except for the integral oil sumps specified in Sec. 25.1013 (a), no tank or reservoir that is a part of a system containing flammable fluids or gases may be in a designated fire zone unless the fluid contained, the design of the system, the materials used in the tank, the shut-off means, and all connections, lines, and control provide a degree of safety equal to that which would exist if the tank or reservoir were outside such a zone. (b) There must be at least one-half inch of clear airspace between each tank or reservoir and each firewall or shroud isolating a designated fire zone. (c) Absorbent materials close to flammable fluid system components that might leak must be covered or treated to prevent the absorption of hazardous quantities of fluids. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964 as amended by Amdt. 25-19, 33 FR 15410, Oct. 17, 1968] Sec. 25.1187 Drainage and ventilation of fire zones. (a) There must be complete drainage of each part of each designated fire zone to minimize the hazards resulting from failure or malfunctioning of any component containing flammable fluids. The drainage means must be-- (1) Effective under conditions expected to prevail when drainage is needed; and (2) Arranged so that no discharged fluid will cause an additional fire hazard. (b) Each designated fire zone must be ventilated to prevent the accumulation of flammable vapors. (c) No ventilation opening may be where it would allow the entry of flammable fluids, vapors, or flame from other zones. (d) Each ventilation means must be arranged so that no discharged vapors will cause an additional fire hazard. (e) Unless the extinguishing agent capacity and rate of discharge are based on maximum air flow through a zone, there must be means to allow the crew to shut off sources of forced ventilation to any fire zone except the engine power section of the nacelle and the combustion heater ventilating air ducts. Sec. 25.1189 Shutoff means. (a) Each engine installation and each fire zone specified in Sec. 25.1181(a) (4) and (5) must have a means to shut off or otherwise prevent hazardous quantities of fuel, oil, deicer, and other flammable fluids, from flowing into, within, or through any designated fire zone, except that shutoff means are not required for-- (1) Lines, fittings, and components forming an integral part of an engine; and (2) Oil systems for turbine engine installations in which all components of the system in a designated fire zone, including oil tanks, are fireproof or located in areas not subject to engine fire conditions. (b) The closing of any fuel shutoff valve for any engine may not make fuel unavailable to the remaining engines. (c) Operation of any shutoff may not interfere with the later emergency operation of other equipment, such as the means for feathering the propeller. (d) Each flammable fluid shutoff means and control must be fireproof or must be located and protected so that any fire in a fire zone will not affect its operation. (e) No hazardous quantity of flammable fluid may drain into any designated fire zone after shutoff. (f) There must be means to guard against inadvertent operation of the shutoff means and to make it possible for the crew to reopen the shutoff means in flight after it has been closed. (g) Each tank-to-engine shutoff valve must be located so that the operation of the valve will not be affected by powerplant or engine mount structural failure. (h) Each shutoff valve must have a means to relieve excessive pressure accumulation unless a means for pressure relief is otherwise provided in the system. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-23, 35 FR 5677, Apr. 8, 1970; Amdt. 25-57, 49 FR 6849, Feb. 23, 1984] Sec. 25.1191 Firewalls. (a) Each engine, auxiliary power unit, fuel-burning heater, other combustion equipment intended for operation in flight, and the combustion, turbine, and tailpipe sections of turbine engines, must be isolated from the rest of the airplane by firewalls, shrouds, or equivalent means. (b) Each firewall and shroud must be-- (1) Fireproof; (2) Constructed so that no hazardous quantity of air, fluid, or flame can pass from the compartment to other parts of the airplane; (3) Constructed so that each opening is sealed with close fitting fireproof grommets, bushings, or firewall fittings; and (4) Protected against corrosion. Sec. 25.1192 Engine accessory section diaphragm. For reciprocating engines, the engine power section and all portions of the exhaust system must be isolated from the engine accessory compartment by a diaphragm that complies with the firewall requirements of Sec. 25.1191. [Amdt. 25-23, 35 FR 5678, Apr. 8, 1970] Sec. 25.1193 Cowling and nacelle skin. (a) Each cowling must be constructed and supported so that it can resist any vibration, inertia, and air load to which it may be subjected in operation. (b) Cowling must meet the drainage and ventilation requirements of Sec. 25.1187. (c) On airplanes with a diaphragm isolating the engine power section from the engine accessory section, each part of the accessory section cowling subject to flame in case of fire in the engine power section of the powerplant must-- (1) Be fireproof; and (2) Meet the requirements of Sec. 25.1191. (d) Each part of the cowling subject to high temperatures due to its nearness to exhaust system parts or exhaust gas impingement must be fireproof. (e) Each airplane must-- (1) Be designed and constructed so that no fire originating in any fire zone can enter, either through openings or by burning through external skin, any other zone or region where it would create additional hazards; (2) Meet paragraph (e)(1) of this section with the landing gear retracted (if applicable); and (3) Have fireproof skin in areas subject to flame if a fire starts in the engine power or accessory sections. Sec. 25.1195 Fire extinguishing systems. (a) Except for combustor, turbine, and tail pipe sections of turbine engine installations that contain lines or components carrying flammable fluids or gases for which it is shown that a fire originating in these sections can be controlled, there must be a fire extinguisher system serving each designated fire zone. (b) The fire extinguishing system, the quantity of the extinguishing agent, the rate of discharge, and the discharge distribution must be adequate to extinguish fires. It must be shown by either actual or simulated flights tests that under critical airflow conditions in flight the discharge of the extinguishing agent in each designated fire zone specified in paragraph (a) of this section will provide an agent concentration capable of extinguishing fires in that zone and of minimizing the probability of reignition. An individual "one-shot" system may be used for auxiliary power units, fuel burning heaters, and other combustion equipment. For each other designated fire zone, two discharges must be provided each of which produces adequate agent concentration. (c) The fire extinguishing system for a nacelle must be able to simultaneously protect each zone of the nacelle for which protection is provided. [Doc. No. 5066, 29 FR 18291, Dec. 24, 1964, as amended by Amdt. 25-46, 43 FR 50598, Oct. 30, 1978] Sec. 25.1197 Fire extinguishing agents. (a) Fire extinguishing agents must-- (1) Be capable of extinguishing flames emanating from any burning of fluids or other combustible materials in the area protected by the fire extinguishing system; and (2) Have thermal stability over the temperature range likely to be experienced in the compartment in which they are stored. (b) If any toxic extinguishing agent is used, provisions must be made to prevent harmful concentrations of fluid or |
