14 C.F.R. Subpart E—Powerplant
Title 14 - Aeronautics and Space
(a) For the purpose of this part, the airplane powerplant installation includes each component that— (1) Is necessary for propulsion; and (2) Affects the safety of the major propulsive units. (b) Each powerplant installation must be constructed and arranged to— (1) Ensure safe operation to the maximum altitude for which approval is requested. (2) Be accessible for necessary inspections and maintenance. (c) Engine cowls and nacelles must be easily removable or openable by the pilot to provide adequate access to and exposure of the engine compartment for preflight checks. (d) Each turbine engine installation must be constructed and arranged to— (1) Result in carcass vibration characteristics that do not exceed those established during the type certification of the engine. (2) Ensure that the capability of the installed engine to withstand the ingestion of rain, hail, ice, and birds into the engine inlet is not less than the capability established for the engine itself under §23.903(a)(2). (e) The installation must comply with— (1) The instructions provided under the engine type certificate and the propeller type certificate. (2) The applicable provisions of this subpart. (f) Each auxiliary power unit installation must meet the applicable portions of this part. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–7, 34 FR 13092, Aug. 13, 1969; Amdt. 23–18, 42 FR 15041, Mar. 17, 1977; Amdt. 23–29, 49 FR 6846, Feb. 23, 1984; Amdt. 23–34, 52 FR 1832, Jan. 15, 1987; Amdt. 23–34, 52 FR 34745, Sept. 14, 1987; Amdt. 23–43, 58 FR 18970, Apr. 9, 1993; Amdt. 23–51, 61 FR 5136, Feb. 9, 1996; Amdt. 23–53, 63 FR 14797, Mar. 26, 1998] (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 and its installation must comply with one of the following: (i) Sections 33.76, 33.77 and 33.78 of this chapter in effect on December 13, 2000, or as subsequently amended; or (ii) Sections 33.77 and 33.78 of this chapter in effect on April 30, 1998, or as subsequently amended before December 13, 2000; or (iii) Section 33.77 of this chapter in effect on October 31, 1974, or as subsequently amended before April 30, 1998, unless that engine's foreign object ingestion service history has resulted in an unsafe condition; or (iv) Be shown to have a foreign object ingestion service history in similar installation locations which has not resulted in any unsafe condition. Note: §33.77 of this chapter in effect on October 31, 1974, was published in 14 CFR parts 1 to 59, Revised as of January 1, 1975. See 39 FR 35467, October 1, 1974. (b) 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 inside 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 operating limitations that adversely affect turbine rotor structural integrity will not be exceeded in service. (c) 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 the failure or malfunction (including destruction by fire in the engine compartment) of any system that can affect an engine (other than a fuel tank if only one fuel tank is installed), will not: (1) Prevent the continued safe operation of the remaining engines; or (2) Require immediate action by any crewmember for continued safe operation of the remaining engines. (d) Starting and stopping (piston engine). (1) The design of the installation must be such that risk of fire or mechanical damage to the engine or airplane, as a result of starting the engine in any conditions in which starting is to be permitted, is reduced to a minimum. Any techniques and associated limitations for engine starting must be established and included in the Airplane Flight Manual, approved manual material, or applicable operating placards. Means must be provided for— (i) Restarting any engine of a multiengine airplane in flight, and (ii) Stopping any engine in flight, after engine failure, if continued engine rotation would cause a hazard to the airplane. (2) In addition, for commuter category airplanes, the following apply: (i) Each component of the stopping system on the engine side of the firewall that might be exposed to fire must be at least fire resistant. (ii) 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. (e) Starting and stopping (turbine engine). Turbine engine installations must comply with the following: (1) The design of the installation must be such that risk of fire or mechanical damage to the engine or the airplane, as a result of starting the engine in any conditions in which starting is to be permitted, is reduced to a minimum. Any techniques and associated limitations must be established and included in the Airplane Flight Manual, approved manual material, or applicable operating placards. (2) There must be means for stopping combustion within any engine and for stopping the rotation of any engine if continued rotation would cause a hazard to the airplane. Each component of the engine stopping system located in any fire zone must be fire resistant. If hydraulic propeller feathering systems are used for stopping the engine, the hydraulic feathering lines or hoses must be fire resistant. (3) It must be possible to restart an engine in flight. Any techniques and associated limitations must be established and included in the Airplane Flight Manual, approved manual material, or applicable operating placards. (4) It must be demonstrated in flight that when restarting engines following a false start, all fuel or vapor is discharged in such a way that it does not constitute a fire hazard. (f) Restart envelope. An altitude and airspeed envelope must be established for the airplane for in-flight engine restarting and each installed engine must have a restart capability within that envelope. (g) Restart capability. For turbine engine powered airplanes, if the minimum windmilling speed of the engines, following the in-flight 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. [Amdt. 23–14, 38 FR 31822, Nov. 19, 1973] Editorial Note: For If installed, an automatic power reserve (APR) system that automatically advances the power or thrust on the operating engine(s), when any engine fails during takeoff, must comply with appendix H of this part. [Doc. No. 26344, 58 FR 18970, Apr. 9, 1993] (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 featherable propeller must have a means to unfeather it in flight. (d) Each component of the propeller blade pitch control system must meet the requirements of §35.42 of this chapter. (e) All areas of the airplane forward of the pusher propeller that are likely to accumulate and shed ice into the propeller disc during any operating condition must be suitably protected to prevent ice formation, or it must be shown that any ice shed into the propeller disc will not create a hazardous condition. (f) Each pusher propeller must be marked so that the disc is conspicuous under normal daylight ground conditions. (g) If the engine exhaust gases are discharged into the pusher propeller disc, it must be shown by tests, or analysis supported by tests, that the propeller is capable of continuous safe operation. (h) All engine cowling, access doors, and other removable items must be designed to ensure that they will not separate from the airplane and contact the pusher propeller. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–26, 45 FR 60171, Sept. 11, 1980; Amdt. 23–29, 49 FR 6847, Feb. 23, 1984; Amdt. 23–43, 58 FR 18970, Apr. 9, 1993] (a) Each propeller other than a conventional fixed-pitch wooden propeller must be shown to have vibration stresses, in normal operating conditions, that do not exceed values that have been shown by the propeller manufacturer to be safe for continuous operation. This must be shown by— (1) Measurement of stresses through direct testing of the propeller; (2) Comparison with similar installations for which these measurements have been made; or (3) Any other acceptable test method or service experience that proves the safety of the installation. (b) Proof of safe vibration characteristics for any type of propeller, except for conventional, fixed-pitch, wood propellers must be shown where necessary. [Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23–51, 61 FR 5136, Feb. 9, 1996] (a) Each turbocharger must be approved under the engine type certificate or it must be shown that the turbocharger system, while in its normal engine installation and operating in the engine environment— (1) Can withstand, without defect, an endurance test of 150 hours that meets the applicable requirements of §33.49 of this subchapter; and (2) Will have no adverse effect upon the engine. (b) Control system malfunctions, vibrations, and abnormal speeds and temperatures expected in service may not damage the turbocharger compressor or turbine. (c) Each turbocharger case must be able to contain fragments of a compressor or turbine that fails at the highest speed that is obtainable with normal speed control devices inoperative. (d) Each intercooler installation, where provided, must comply with the following— (1) The mounting provisions of the intercooler must be designed to withstand the loads imposed on the system; (2) It must be shown that, under the installed vibration environment, the intercooler will not fail in a manner allowing portions of the intercooler to be ingested by the engine; and (3) Airflow through the intercooler must not discharge directly on any airplane component (e.g., windshield) unless such discharge is shown to cause no hazard to the airplane under all operating conditions. (e) Engine power, cooling characteristics, operating limits, and procedures affected by the turbocharger system installations must be evaluated. Turbocharger operating procedures and limitations must be included in the Airplane Flight Manual in accordance with §23.1581. [Amdt. 23–7, 34 FR 13092, Aug. 13, 1969, as amended by Amdt. 23–43, 58 FR 18970, Apr. 9, 1993] Unless smaller clearances are substantiated, propeller clearances, with the airplane at the most adverse combination of weight and 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, normal takeoff, or taxing attitude, whichever is most critical. In addition, for each airplane with conventional landing gear struts using fluid or mechanical means for absorbing landing shocks, there must be positive clearance between the propeller and the ground in the level takeoff attitude with the critical tire completely deflated and the corresponding landing gear strut bottomed. Positive clearance for airplanes using leaf spring struts is shown with a deflection corresponding to 1.5g. (b) Aft-mounted propellers. In addition to the clearances specified in paragraph (a) of this section, an airplane with an aft mounted propeller must be designed such that the propeller will not contact the runway surface when the airplane is in the maximum pitch attitude attainable during normal takeoffs and landings. (c) Water clearance. There must be a clearance of at least 18 inches between each propeller and the water, unless compliance with §23.239 can be shown with a lesser clearance. (d) 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. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–43, 58 FR 18971, Apr. 9, 1993; Amdt. 23–51, 61 FR 5136, Feb. 9, 1996; Amdt. 23–48, 61 FR 5148, Feb. 9, 1996] Propellers (except wooden propellers) and other components of complete engine installations must be protected against the accumulation of ice as necessary to enable satisfactory functioning without appreciable loss of thrust when operated in the icing conditions for which certification is requested. [Amdt. 23–14, 33 FR 31822, Nov. 19, 1973, as amended by Amdt. 23–51, 61 FR 5136, Feb. 9, 1996] (a) For turbojet and turbofan 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; or (ii) The airplane is capable of continued safe flight and landing under any possible position of the thrust reverser. (2) Each system intended for in-flight use must be designed so that no unsafe condition will result during normal operation of the system, or from any failure, or likely combination of failures, of the reversing system under any operating condition including ground operation. Failure of structural elements need not be considered if the probability of this type of failure is extremely remote. (3) Each system must have a means to prevent the engine from producing more than idle thrust when the reversing system malfunctions; except that it may produce any greater 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 must be designed so that no single failure, likely combination of failures or malfunction of the system will result in unwanted reverse thrust under any operating condition. Failure of structural elements need not be considered if the probability of this type of failure is extremely remote. (2) Compliance with paragraph (b)(1) of this section must be shown by failure analysis, or testing, or both, for propeller systems that allow the propeller blades to move from the flight low-pitch position to a position that is substantially less than the normal flight, low-pitch position. The analysis may include or be supported by the analysis made to show compliance with §35.21 for the type certification of the propeller and associated installation components. Credit will be given for pertinent analysis and testing completed by the engine and propeller manufacturers. [Doc. No. 26344, 58 FR 18971, Apr. 9, 1993, as amended by Amdt. 23–51, 61 FR 5136, Feb. 9, 1996] Thrust reverser systems of turbojet or turbofan engines must meet the requirements of §33.97 of this chapter or it must be demonstrated by tests that engine operation and vibratory levels are not affected. [Doc. No. 26344, 58 FR 18971, Apr. 9, 1993] (a) Turbopropeller-powered 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 the structural requirements of this part. Failure of structural elements of the drag limiting systems need not be considered if the probability of this kind of failure is extremely remote. (b) As used in this section, drag limiting systems include manual or automatic devices that, when actuated after engine power loss, can move the propeller blades toward the feather position to reduce windmilling drag to a safe level. [Amdt. 23–7, 34 FR 13093, Aug. 13, 1969, as amended by Amdt. 23–43, 58 FR 18971, Apr. 9, 1993] (a) Turbine engine powerplant 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) Turbocharged reciprocating engine operating characteristics must be investigated in flight to assure that no adverse characteristics, as a result of an inadvertent overboost, surge, flooding, or vapor lock, are present during normal or emergency operation of the engine(s) throughout the range of operating limitations of both airplane and engine. (c) For turbine engines, the air inlet system must not, as a result of airflow distortion during normal operation, cause vibration harmful to the engine. [Amdt. 23–7, 34 FR 13093 Aug. 13, 1969, as amended by Amdt. 23–14, 38 FR 31823, Nov. 19, 1973; Amdt. 23–18, 42 FR 15041, Mar. 17, 1977; Amdt. 23–42, 56 FR 354, Jan. 3, 1991] 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 §23.333. This must be shown for the greatest value and duration of the acceleration expected in service. [Amdt. 23–18, 42 FR 15041, Mar. 17, 1977, as amended by Amdt. 23–43, 58 FR 18971, Apr. 9, 1993]
Title 14: Aeronautics and Space
PART 23—AIRWORTHINESS STANDARDS: NORMAL, UTILITY, ACROBATIC, AND COMMUTER CATEGORY AIRPLANES
Subpart E—Powerplant
General
§ 23.901 Installation.
§ 23.903 Engines.
§ 23.904 Automatic power reserve system.
§ 23.905 Propellers.
§ 23.907 Propeller vibration.
§ 23.909 Turbocharger systems.
§ 23.925 Propeller clearance.
§ 23.929 Engine installation ice protection.
§ 23.933 Reversing systems.
§ 23.934 Turbojet and turbofan engine thrust reverser systems tests.
§ 23.937 Turbopropeller-drag limiting systems.
§ 23.939 Powerplant operating characteristics.
§ 23.943 Negative acceleration.
Fuel System
§ 23.951 General.
(a) Each fuel system must be constructed and arranged to ensure fuel flow 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—
(1) No fuel pump can draw fuel from more than one tank at a time; or
(2) There are means to prevent introducing air into the system.
(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 °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.
[Amdt. 23–15, 39 FR 35459, Oct. 1, 1974, as amended by Amdt. 23–40, 55 FR 32861, Aug. 10, 1990; Amdt. 23–43, 58 FR 18971, Apr. 9, 1993]
§ 23.953 Fuel system independence.
(a) Each fuel system for a multiengine airplane must be arranged so that, in at least one system configuration, the failure of any one component (other than a fuel tank) will not result in the loss of power of more than one engine or require immediate action by the pilot to prevent the loss of power of more than one engine.
(b) If a single fuel tank (or series of fuel tanks interconnected to function as a single fuel tank) is used on a multiengine airplane, the following must be provided:
(1) Independent tank outlets for each engine, each incorporating a shut-off valve at the tank. This shutoff valve may also serve as the fire wall shutoff valve required if the line between the valve and the engine compartment does not contain more than one quart of fuel (or any greater amount shown to be safe) that can escape into the engine compartment.
(2) At least two vents arranged to minimize the probability of both vents becoming obstructed simultaneously.
(3) Filler caps designed to minimize the probability of incorrect installation or inflight loss.
(4) A fuel system in which those parts of the system from each tank outlet to any engine are independent of each part of the system supplying fuel to any other engine.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–7, 34 FR 13093 Aug. 13, 1969; Amdt. 23–43, 58 FR 18971, Apr. 9, 1993]
§ 23.954 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 on areas where swept strokes are highly probable; and
(c) Corona or streamering at fuel vent outlets.
[Amdt. 23–7, 34 FR 13093, Aug. 13, 1969]
§ 23.955 Fuel flow.
(a) General. The ability of the fuel system to provide fuel at the rates specified in this section and at a pressure sufficient for proper engine operation must be shown in the attitude that is most critical with respect to fuel feed and quantity of unusable fuel. These conditions may be simulated in a suitable mockup. In addition—
(1) The quantity of fuel in the tank may not exceed the amount established as the unusable fuel supply for that tank under §23.959(a) plus that quantity necessary to show compliance with this section.
(2) If there is a fuel flowmeter, it must be blocked during the flow test and the fuel must flow through the meter or its bypass.
(3) If there is a flowmeter without a bypass, it must not have any probable failure mode that would restrict fuel flow below the level required for this fuel demonstration.
(4) The fuel flow must include that flow necessary for vapor return flow, jet pump drive flow, and for all other purposes for which fuel is used.
(b) Gravity systems. The fuel flow rate for gravity systems (main and reserve supply) must be 150 percent of the takeoff fuel consumption of the engine.
(c) Pump systems. The fuel flow rate for each pump system (main and reserve supply) for each reciprocating engine must be 125 percent of the fuel flow required by the engine at the maximum takeoff power approved under this part.
(1) This flow rate is required for each main pump and each emergency pump, and must be available when the pump is operating as it would during takeoff;
(2) For each hand-operated pump, this rate must occur at not more than 60 complete cycles (120 single strokes) per minute.
(3) The fuel pressure, with main and emergency pumps operating simultaneously, must not exceed the fuel inlet pressure limits of the engine unless it can be shown that no adverse effect occurs.
(d) Auxiliary fuel systems and fuel transfer systems. Paragraphs (b), (c), and (f) of this section apply to each auxiliary and transfer system, except that—
(1) The required fuel flow rate must be established upon the basis of maximum continuous power and engine rotational speed, instead of takeoff power and fuel consumption; and
(2) If there is a placard providing operating instructions, a lesser flow rate may be used for transferring fuel from any auxiliary tank into a larger main tank. This lesser flow rate must be adequate to maintain engine maximum continuous power but the flow rate must not overfill the main tank at lower engine powers.
(e) Multiple fuel tanks. For reciprocating engines that are supplied with fuel from more than one tank, if engine power loss becomes apparent due to fuel depletion from the tank selected, it must be possible after switching to any full tank, in level flight, to obtain 75 percent maximum continuous power on that engine in not more than—
(1) 10 seconds for naturally aspirated single-engine airplanes;
(2) 20 seconds for turbocharged single-engine airplanes, provided that 75 percent maximum continuous naturally aspirated power is regained within 10 seconds; or
(3) 20 seconds for multiengine airplanes.
(f) Turbine engine fuel systems. Each turbine engine fuel system must provide at least 100 percent of the fuel flow required by the engine under each intended operation condition and maneuver. The conditions may be simulated in a suitable mockup. This flow must—
(1) Be shown with the airplane in the most adverse fuel feed condition (with respect to altitudes, attitudes, and other conditions) that is expected in operation; and
(2) For multiengine airplanes, notwithstanding the lower flow rate allowed by paragraph (d) of this section, be automatically uninterrupted with respect to any engine until all the fuel scheduled for use by that engine has been consumed. In addition—
(i) For the purposes of this section, “fuel scheduled for use by that engine” means all fuel in any tank intended for use by a specific engine.
(ii) The fuel system design must clearly indicate the engine for which fuel in any tank is scheduled.
(iii) Compliance with this paragraph must require no pilot action after completion of the engine starting phase of operations.
(3) For single-engine airplanes, require no pilot action after completion of the engine starting phase of operations unless means are provided that unmistakenly alert the pilot to take any needed action at least five minutes prior to the needed action; such pilot action must not cause any change in engine operation; and such pilot action must not distract pilot attention from essential flight duties during any phase of operations for which the airplane is approved.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–7, 34 FR 13093, Aug. 13, 1969; Amdt. 23–43, 58 FR 18971, Apr. 9, 1993; Amdt. 23–51, 61 FR 5136, Feb. 9, 1996]
§ 23.957 Flow between interconnected tanks.
(a) It must be impossible, in a gravity feed system with interconnected tank outlets, for enough fuel to flow between the tanks to cause an overflow of fuel from any tank vent under the conditions in §23.959, except that full tanks must be used.
(b) 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 any airplane component can occur because of overfilling of any tank.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–43, 58 FR 18972, Apr. 9, 1993]
§ 23.959 Unusable fuel supply.
(a) The unusable fuel supply for each tank must be established as not less than that quantity at which the first evidence of malfunctioning occurs under the most adverse fuel feed condition occurring under each intended operation and flight maneuver involving that tank. Fuel system component failures need not be considered.
(b) The effect on the usable fuel quantity as a result of a failure of any pump shall be determined.
[Amdt. 23–7, 34 FR 13093, Aug. 13, 1969, as amended by Amdt. 23–18, 42 FR 15041, Mar. 17, 1977; Amdt. 23–51, 61 FR 5136, Feb. 9, 1996]
§ 23.961 Fuel system hot weather operation.
Each fuel system must be free from vapor lock when using fuel at its critical temperature, with respect to vapor formation, when operating the airplane in all critical operating and environmental conditions for which approval is requested. For turbine fuel, the initial temperature must be 110 °F, −0°, +5 °F or the maximum outside air temperature for which approval is requested, whichever is more critical.
[Doc. No. 26344, 58 FR 18972, Apr. 9, 1993; 58 FR 27060, May 6, 1993]
§ 23.963 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) Each flexible fuel tank liner must be shown to be suitable for the particular application.
(c) Each integral fuel tank must have adequate facilities for interior inspection and repair.
(d) The total usable capacity of the fuel tanks must be enough for at least one-half hour of operation at maximum continuous power.
(e) Each fuel quantity indicator must be adjusted, as specified in §23.1337(b), to account for the unusable fuel supply determined under §23.959(a).
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt 23–34, 52 FR 1832, Jan. 15, 1987; Amdt. 23–43, 58 FR 18972, Apr. 9, 1993; Amdt. 23–51, 61 FR 5136, Feb. 9, 1996]
§ 23.965 Fuel tank tests.
(a) Each fuel tank must be able to withstand the following pressures without failure or leakage:
(1) For each conventional metal tank and nonmetallic tank with walls not supported by the airplane structure, a pressure of 3.5 p.s.i., or that pressure developed during maximum ultimate acceleration with a full tank, whichever is greater.
(2) For each integral tank, the pressure developed during the maximum limit acceleration of the airplane with a full tank, with simultaneous application of the critical limit structural loads.
(3) For each nonmetallic tank with walls supported by the airplane structure and constructed in an acceptable manner using acceptable basic tank material, and with actual or simulated support conditions, a pressure of 2 p.s.i. for the first tank of a specific design. The supporting structure must be designed for the critical loads occurring in the flight or landing strength conditions combined with the fuel pressure loads resulting from the corresponding accelerations.
(b) Each fuel tank with large, unsupported, or unstiffened flat surfaces,whose failure or deformation could cause fuel leakage, must be able to withstand the following test without leakage, failure, or excessive deformation of the tank walls:
(1) Each complete tank assembly and its support 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 a total displacement of not less than
(3) The test frequency of vibration must be as follows:
(i) If no frequency of vibration resulting from any rpm within the normal operating range of engine or propeller speeds is critical, the test frequency of vibration is:
(A) The number of cycles per minute obtained by multiplying the maximum continuous propeller speed in rpm by 0.9 for propeller-driven airplanes, and
(B) For non-propeller driven airplanes the test frequency of vibration is 2,000 cycles per minute.
(ii) If only one frequency of vibration resulting from any rpm within the normal operating range of engine or propeller speeds is critical, that frequency of vibration must be the test frequency.
(iii) If more than one frequency of vibration resulting from any rpm within the normal operating range of engine or propeller speeds is critical, the most critical of these frequencies must be the test frequency.
(4) Under paragraph (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 a rate of 16 to 20 complete cycles per minute, through an angle of 15° on either side of the horizontal (30° total), about an axis parallel to the axis of the fuselage, for 25 hours.
(c) Each integral tank using methods of construction and sealing not previously proven to be adequate by test data or service experience must be able to withstand the vibration test specified in paragraphs (b)(1) through (4) of this section.
(d) Each tank with a nonmetallic liner must be subjected to the sloshing test outlined in paragraph (b)(5) of this section, with the fuel at room temperature. In addition, a specimen liner of the same basic construction as that to be used in the airplane must, when installed in a suitable test tank, withstand the sloshing test with fuel at a temperature of 110 °F.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–43, 58 FR 18972, Apr. 9, 1993; Amdt. 23–43, 61 FR 253, Jan. 4, 1996; Amdt. 23–51, 61 FR 5136, Feb. 9, 1996]
§ 23.967 Fuel tank installation.
(a) Each fuel tank must be supported so that tank loads are not concentrated. In addition—
(1) There must be pads, if necessary, to prevent chafing between each tank and its supports;
(2) Padding must be nonabsorbent or treated to prevent the absorption of fuel;
(3) If a flexible tank liner is used, it must be supported so that it is not required to withstand fluid loads;
(4) Interior surfaces adjacent to the liner must be smooth and free from projections that could cause wear, unless—
(i) Provisions are made for protection of the liner at those points; or
(ii) The construction of the liner itself provides such protection; and
(5) A positive pressure must be maintained within the vapor space of each bladder cell under any condition of operation, except for a particular condition for which it is shown that a zero or negative pressure will not cause the bladder cell to collapse; and
(6) Syphoning of fuel (other than minor spillage) or collapse of bladder fuel cells may not result from improper securing or loss of the fuel filler cap.
(b) Each tank compartment must be ventilated and drained to prevent the accumulation of flammable fluids or vapors. Each compartment adjacent to a tank that is an integral part of the airplane structure must also be ventilated and drained.
(c) No fuel tank may be on the engine side of the firewall. There must be at least one-half inch of clearance between the fuel tank and the firewall. No part of the engine nacelle skin that lies immediately behind a major air opening from the engine compartment may act as the wall of an integral tank.
(d) Each fuel tank must be isolated from personnel compartments by a fume-proof and fuel-proof enclosure that is vented and drained to the exterior of the airplane. The required enclosure must sustain any personnel compartment pressurization loads without permanent deformation or failure under the conditions of §§23.365 and 23.843 of this part. A bladder-type fuel cell, if used, must have a retaining shell at least equivalent to a metal fuel tank in structural integrity.
(e) Fuel tanks must be designed, located, and installed so as to retain fuel:
(1) When subjected to the inertia loads resulting from the ultimate static load factors prescribed in §23.561(b)(2) of this part; and
(2) Under conditions likely to occur when the airplane lands on a paved runway at a normal landing speed under each of the following conditions:
(i) The airplane in a normal landing attitude and its landing gear retracted.
(ii) The most critical landing gear leg collapsed and the other landing gear legs extended.
In showing compliance with paragraph (e)(2) of this section, the tearing away of an engine mount must be considered unless all the engines are installed above the wing or on the tail or fuselage of the airplane.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–7, 34 FR 13903, Aug. 13, 1969; Amdt. 23–14, 38 FR 31823, Nov. 19, 1973; Amdt. 23–18, 42 FR 15041, Mar. 17, 1977; Amdt. 23–26, 45 FR 60171, Sept. 11, 1980; Amdt. 23–36, 53 FR 30815, Aug. 15, 1988; Amdt. 23–43, 58 FR 18972, Apr. 9, 1993]
§ 23.969 Fuel tank expansion space.
Each fuel tank must have an expansion space of not less than two percent of the tank capacity, unless the tank vent discharges clear of the airplane (in which case no expansion space is required). It must be impossible to fill the expansion space inadvertently with the airplane in the normal ground attitude.
§ 23.971 Fuel tank sump.
(a) Each fuel tank must have a drainable sump with an effective capacity, in the normal ground and flight attitudes, of 0.25 percent of the tank capacity, or
(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 normal ground attitude.
(c) Each reciprocating engine fuel system must have a sediment bowl or chamber that is accessible for drainage; has a capacity of 1 ounce for every 20 gallons of fuel tank capacity; and each fuel tank outlet is located so that, in the normal flight attitude, water will drain from all parts of the tank except the sump to the sediment bowl or chamber.
(d) Each sump, sediment bowl, and sediment chamber drain required by paragraphs (a), (b), and (c) of this section must comply with the drain provisions of §23.999(b)(1) and (b)(2).
[Doc. No. 26344, 58 FR 18972, Apr. 9, 1993; 58 FR 27060, May 6, 1993]
§ 23.973 Fuel tank filler connection.
(a) Each fuel tank filler connection must be marked as prescribed in §23.1557(c).
(b) Spilled fuel must be prevented from entering the fuel tank compartment or any part of the airplane other than the tank itself.
(c) Each filler cap must provide a fuel-tight seal for the main filler opening. However, there may be small openings in the fuel tank cap for venting purposes or for the purpose of allowing passage of a fuel gauge through the cap provided such openings comply with the requirements of §23.975(a).
(d) Each fuel filling point, except pressure fueling connection points, must have a provision for electrically bonding the airplane to ground fueling equipment.
(e) For airplanes with engines requiring gasoline as the only permissible fuel, the inside diameter of the fuel filler opening must be no larger than 2.36 inches.
(f) For airplanes with turbine engines, the inside diameter of the fuel filler opening must be no smaller than 2.95 inches.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23–18, 42 FR 15041, Mar. 17, 1977; Amdt. 23–43, 58 FR 18972, Apr. 9, 1993; Amdt. 23–51, 61 FR 5136, Feb. 9, 1996]
§ 23.975 Fuel tank vents and carburetor vapor vents.
(a) Each fuel tank must be vented from the top part of the expansion space. In addition—
(1) Each vent outlet must be located and constructed in a manner that minimizes the possibility of its being obstructed by ice or other foreign matter;
(2) Each vent must be constructed to prevent siphoning of fuel during normal operation;
(3) The venting capacity must allow the rapid relief of excessive differences of pressure between the interior and exterior of the tank;
(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 either the ground or level flight attitudes, unless drainage is provided. Any drain valve installed must be accessible for drainage;
(6) No vent may terminate at a point where the discharge of fuel from the vent outlet will constitute a fire hazard or from which fumes may enter personnel compartments; and
(7) Vents must be arranged to prevent the loss of fuel, except fuel discharged because of thermal expansion, when the airplane is parked in any direction on a ramp having a one-percent slope.
(b) Each carburetor with vapor elimination connections and each fuel injection engine employing vapor return provisions must have a separate vent line to lead vapors back to the top of one of the fuel tanks. If there is more than one tank and it is necessary to use these tanks in a definite sequence for any reason, the vapor vent line must lead back to the fuel tank to be used first, unless the relative capacities of the tanks are such that return to another tank is preferable.
(c) For acrobatic category airplanes, excessive loss of fuel during acrobatic maneuvers, including short periods of inverted flight, must be prevented. It must be impossible for fuel to siphon from the vent when normal flight has been resumed after any acrobatic maneuver for which certification is requested.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23–18, 42 FR 15041, Mar. 17, 1977; Amdt. 23–29, 49 FR 6847, Feb. 23, 1984; Amdt. 23–43, 58 FR 18973, Apr. 9, 1993; Amdt. 23–51, 61 FR 5136, Feb. 9, 1996]
§ 23.977 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) The clear area of each fuel tank outlet strainer must be at least five times the area of the outlet line.
(c) The diameter of each strainer must be at least that of the fuel tank outlet.
(d) Each strainer must be accessible for inspection and cleaning.
[Amdt. 23–17, 41 FR 55465, Dec. 20, 1976, as amended by Amdt. 23–43, 58 FR 18973, Apr. 9, 1993]
§ 23.979 Pressure fueling systems.
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) For commuter category airplanes, indicate at each fueling station, a 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) All parts of the fuel system up to the tank which are subjected to fueling pressures must have a proof pressure of 1.33 times, and an ultimate pressure of at least 2.0 times, the surge pressure likely to occur during fueling.
[Amdt. 23–14, 38 FR 31823, Nov. 19, 1973, as amended by Amdt. 23–51, 61 FR 5137, Feb. 9, 1996]
Fuel System Components
§ 23.991 Fuel pumps.
(a) Main pumps. For main pumps, the following apply:
(1) For reciprocating engine installations having fuel pumps to supply fuel to the engine, at least one pump for each engine must be directly driven by the engine and must meet §23.955. This pump is a main pump.
(2) For turbine engine installations, 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. In addition—
(i) There must be at least one main pump for each turbine engine;
(ii) The power supply for the main pump for each engine must be independent of the power supply for each main pump for any other engine; and
(iii) For each main pump, provision must be made to allow the bypass of each positive displacement fuel pump other than a fuel injection pump approved as part of the engine.
(b) Emergency pumps. There must be an emergency pump immediately available to supply fuel to the engine if any main pump (other than a fuel injection pump approved as part of an engine) fails. The power supply for each emergency pump must be independent of the power supply for each corresponding main pump.
(c) Warning means. If both the main pump and emergency pump operate continuously, there must be a means to indicate to the appropriate flight crewmembers a malfunction of either pump.
(d) Operation of any fuel pump may not affect engine operation so as to create a hazard, regardless of the engine power or thrust setting or the functional status of any other fuel pump.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–7, 34 FR 13093, Aug. 13, 1969; Amdt. 23–26, 45 FR 60171, Sept. 11, 1980; Amdt. 23–43, 58 FR 18973, Apr. 9, 1993]
§ 23.993 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) Each flexible hose 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 shutdown.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–43, 58 FR 18973, Apr. 9, 1993]
§ 23.994 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. 23–29, 49 FR 6847, Feb. 23, 1984]
§ 23.995 Fuel valves and controls.
(a) There must be a means to allow appropriate flight crew members to rapidly shut off, in flight, the fuel to each engine individually.
(b) No shutoff valve may be on the engine side of any firewall. In addition, there must be means to—
(1) Guard against inadvertent operation of each shutoff valve; and
(2) Allow appropriate flight crew members to reopen each valve rapidly after it has been closed.
(c) Each valve and fuel system control must be supported so that loads resulting from its operation or from accelerated flight conditions are not transmitted to the lines connected to the valve.
(d) Each valve and fuel system control must be installed so that gravity and vibration will not affect the selected position.
(e) Each fuel valve handle and its connections to the valve mechanism must have design features that minimize the possibility of incorrect installation.
(f) Each check valve must be constructed, or otherwise incorporate provisions, to preclude incorrect assembly or connection of the valve.
(g) Fuel tank selector valves must—
(1) Require a separate and distinct action to place the selector in the “OFF” position; and
(2) Have the tank selector positions located in such a manner that it is impossible for the selector to pass through the “OFF” position when changing from one tank to another.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–14, 38 FR 31823, Nov. 19, 1973; Amdt. 23–17, 41 FR 55465, Dec. 20, 1976; Amdt. 23–18, 42 FR 15041, Mar. 17, 1977; Amdt. 23–29, 49 FR 6847, Feb. 23, 1984]
§ 23.997 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 during its type certification.
(e) In addition, for commuter category airplanes, unless means are provided in the fuel system to prevent the accumulation of ice on the filter, a means must be provided to automatically maintain the fuel flow if ice clogging of the filter occurs.
[Amdt. 23–15, 39 FR 35459, Oct. 1, 1974, as amended by Amdt. 23–29, 49 FR 6847, Feb. 23, 1984; Amdt. 23–34, 52 FR 1832, Jan. 15, 1987; Amdt. 23–43, 58 FR 18973, Apr. 9, 1993]
§ 23.999 Fuel system drains.
(a) There must be at least one drain to allow safe drainage of the entire fuel system with the airplane in its normal ground attitude.
(b) Each drain required by paragraph (a) of this section and §23.971 must—
(1) Discharge clear of all parts of the airplane;
(2) Have a drain valve—
(i) That has manual or automatic means for positive locking in the closed position;
(ii) That is readily accessible;
(iii) That can be easily opened and closed;
(iv) That allows the fuel to be caught for examination;
(v) That can be observed for proper closing; and
(vi) That is either located or protected to prevent fuel spillage in the event of a landing with landing gear retracted.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–17, 41 FR 55465, Dec. 20, 1976; Amdt. 23–43, 58 FR 18973, Apr. 9, 1993]
§ 23.1001 Fuel jettisoning system.
(a) If the design landing weight is less than that permitted under the requirements of §23.473(b), the airplane must have a fuel jettisoning system installed that is able to jettison enough fuel to bring the maximum weight down to the design landing weight. The average rate of fuel jettisoning must be at least 1 percent of the maximum weight per minute, except that the time required to jettison the fuel need not be less than 10 minutes.
(b) Fuel jettisoning must be demonstrated at maximum weight with flaps and landing gear up and in—
(1) A power-off glide at 1.4 V
(2) A climb, at the speed at which the one-engine-inoperative enroute climb data have been established in accordance with §23.69(b), with the critical engine inoperative and the remaining engines at maximum continuous power; and
(3) Level flight at 1.4 V
(c) During the flight tests prescribed in paragraph (b) 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.
(d) For reciprocating engine powered airplanes, the jettisoning system must be designed so that it is not possible to jettison 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 all the fuel.
(e) For turbine engine powered airplanes, the jettisoning system must be designed so that it is not possible to jettison 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.
(f) The fuel jettisoning valve must be designed to allow flight crewmembers to close the valve during any part of the jettisoning operation.
(g) 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.
(h) 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.
[Amdt. 23–7, 34 FR 13094, Aug. 13, 1969, as amended by Amdt. 23–43, 58 FR 18973, Apr. 9, 1993; Amdt. 23–51, 61 FR 5137, Feb. 9, 1996]
Oil System
§ 23.1011 General.
(a) For oil systems and components that have been approved under the engine airworthiness requirements and where those requirements are equal to or more severe than the corresponding requirements of subpart E of this part, that approval need not be duplicated. Where the requirements of subpart E of this part are more severe, substantiation must be shown to the requirements of subpart E of this part.
(b) 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.
(c) The usable oil tank capacity may not be less than the product of the endurance of the airplane under critical operating conditions and the maximum oil consumption of the engine under the same conditions, plus a suitable margin to ensure adequate circulation and cooling.
(d) For an oil system without an oil transfer system, only the usable oil tank capacity may be considered. The amount of oil in the engine oil lines, the oil radiator, and the feathering reserve, may not be considered.
(e) If an oil transfer system is used, and the transfer pump can pump some of the oil in the transfer lines into the main engine oil tanks, the amount of oil in these lines that can be pumped by the transfer pump may be included in the oil capacity.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–43, 58 FR 18973, Apr. 9, 1993]
§ 23.1013 Oil tanks.
(a) Installation. Each oil tank must be installed to—
(1) Meet the requirements of §23.967 (a) and (b); and
(2) Withstand any vibration, inertia, and fluid loads expected in operation.
(b) Expansion space. Oil tank expansion space must be provided so that—
(1) Each oil tank used with a reciprocating engine has 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 has an expansion space of not less than 10 percent of the tank capacity; and
(2) It is impossible to fill the expansion space inadvertently with the airplane in the normal ground attitude.
(c) Filler connection. Each oil tank filler connection must be marked as specified in §23.1557(c). Each recessed oil tank filler connection of an oil tank used with a turbine engine, that can retain any appreciable quantity of oil, must have provisions for fitting a drain.
(d) Vent. Oil tanks must be vented as follows:
(1) Each oil tank must be vented to the engine from the top part of the expansion space so that the vent connection is not covered by oil 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.
(3) For acrobatic category airplanes, there must be means to prevent hazardous loss of oil during acrobatic maneuvers, including short periods of inverted flight.
(e) Outlet. 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. No oil tank outlet diameter may be less than the diameter of the engine oil pump inlet. Each oil tank used with a turbine engine must have 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. 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 oil tank supports) is fireproof.
(f) Flexible liners. Each flexible oil tank liner must be of an acceptable kind.
(g) Each oil tank filler cap of an oil tank that is used with an engine must provide an oiltight seal.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–15, 39 FR 35459 Oct. 1, 1974; Amdt. 23–43, 58 FR 18973, Apr. 9, 1993; Amdt. 23–51, 61 FR 5137, Feb. 9, 1996]
§ 23.1015 Oil tank tests.
Each oil tank must be tested under §23.965, except that—
(a) The applied pressure must be five p.s.i. for the tank construction instead of the pressures specified in §23.965(a);
(b) For a tank with a nonmetallic liner the test fluid must be oil rather than fuel as specified in §23.965(d), and the slosh test on a specimen liner must be conducted with the oil at 250 °F.; and
(c) For pressurized tanks used with a turbine engine, the test pressure may not be less than 5 p.s.i. plus the maximum operating pressure of the tank.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–15, 39 FR 35460, Oct. 1, 1974]
§ 23.1017 Oil lines and fittings.
(a) Oil lines. Oil lines must meet §23.993 and must accommodate a flow of oil at a rate and pressure adequate for proper engine functioning under any normal operating condition.
(b) Breather lines. Breather lines must be arranged so that—
(1) Condensed water vapor or oil that might freeze and obstruct the line cannot accumulate at any point;
(2) The breather discharge will not constitute a fire hazard if foaming occurs, or cause emitted oil to strike the pilot's windshield;
(3) The breather does not discharge into the engine air induction system; and
(4) For acrobatic category airplanes, there is no excessive loss of oil from the breather during acrobatic maneuvers, including short periods of inverted flight.
(5) The breather outlet is protected against blockage by ice or foreign matter.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–7, 34 FR 13094, Aug. 13, 1969; Amdt. 23–14, 38 FR 31823, Nov. 19, 1973]
§ 23.1019 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 for its type certification.
(3) The oil strainer or filter, unless it is installed at an oil tank outlet, must incorporate a means to 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 §23.1305(c)(9).
(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. 23–15, 39 FR 35460, Oct. 1, 1974, as amended by Amdt. 23–29, 49 FR 6847, Feb. 23, 1984; Amdt. 23–43, 58 FR 18973, Apr. 9, 1993]
§ 23.1021 Oil system drains.
A drain (or drains) must be provided to allow safe drainage of the oil system. Each drain must—
(a) Be accessible;
(b) Have drain valves, or other closures, employing manual or automatic shut-off means for positive locking in the closed position; and
(c) Be located or protected to prevent inadvertent operation.
[Amdt. 23–29, 49 FR 6847, Feb. 23, 1984, as amended by Amdt. 23–43, 58 FR 18973, Apr. 9, 1993]
§ 23.1023 Oil radiators.
Each oil radiator and its supporting structures must be able to withstand the vibration, inertia, and oil pressure loads to which it would be subjected in operation.
§ 23.1027 Propeller feathering system.
(a) If the propeller feathering system uses engine oil and that oil supply can become depleted due to failure of any part of the oil system, a means must be incorporated to reserve enough oil to operate the feathering system.
(b) The amount of reserved oil must be enough to accomplish feathering and must be available only to the feathering pump.
(c) The ability of the system to accomplish feathering with the reserved oil must be shown.
(d) Provision must be made to prevent sludge or other foreign matter from affecting the safe operation of the propeller feathering system.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–14, 38 FR 31823, Nov. 19, 1973; Amdt. 23–43, 58 FR 18973, Apr. 9, 1993]
Cooling
§ 23.1041 General.
The powerplant and auxiliary power unit cooling provisions must maintain the temperatures of powerplant components and engine fluids, and auxiliary power unit components and fluids within the limits established for those components and fluids under the most adverse ground, water, and flight operations to the maximum altitude and maximum ambient atmospheric temperature conditions for which approval is requested, and after normal engine and auxiliary power unit shutdown.
[Doc. No. 26344, 58 FR 18973, Apr. 9, 1993, as amended by Amdt. 23–51, 61 FR 5137, Feb. 9, 1996]
§ 23.1043 Cooling tests.
(a) General. Compliance with §23.1041 must be shown on the basis of tests, for which the following apply:
(1) If the tests are conducted under ambient atmospheric temperature conditions deviating from the maximum for which approval is requested, the recorded powerplant temperatures must be corrected under paragraphs (c) and (d) of this section, unless a more rational correction method is applicable.
(2) No corrected temperature determined under paragraph (a)(1) of this section may exceed established limits.
(3) The fuel used during the cooling tests must be of the minimum grade approved for the engine.
(4) For turbocharged engines, each turbocharger must be operated through that part of the climb profile for which operation with the turbocharger is requested.
(5) For a reciprocating engine, the mixture settings must be the leanest recommended for climb.
(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). 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 for the relevant altitude for which approval has been requested and the temperature of the ambient air at the time of the first occurrence of the maximum fluid or component temperature recorded during the cooling test.
(d) Correction factor for cylinder barrel temperatures. Cylinder barrel temperatures must be corrected by adding to them 0.7 times the difference between the maximum ambient atmospheric temperature for the relevant altitude for which approval has been requested 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. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–7, 34 FR 13094, Aug. 13, 1969; Amdt. 23–21, 43 FR 2319, Jan. 16, 1978; Amdt. 23–51, 61 FR 5137, Feb. 9, 1996]
§ 23.1045 Cooling test procedures for turbine engine powered airplanes.
(a) Compliance with §23.1041 must be shown for all phases of operation. The airplane must be flown in the configurations, at the speeds, and following the procedures recommended in the Airplane Flight Manual for the relevant stage of flight, that correspond to the applicable performance requirements that are critical to cooling.
(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 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.
[Amdt. 23–7, 34 FR 13094, Aug. 13, 1969, as amended by Amdt. 23–51, 61 FR 5137, Feb. 9, 1996]
§ 23.1047 Cooling test procedures for reciprocating engine powered airplanes.
Compliance with §23.1041 must be shown for the climb (or, for multiengine airplanes with negative one-engine-inoperative rates of climb, the descent) stage of flight. The airplane must be flown in the configurations, at the speeds and following the procedures recommended in the Airplane Flight Manual, that correspond to the applicable performance requirements that are critical to cooling.
[Amdt. 23–51, 61 FR 5137, Feb. 9, 1996]
Liquid Cooling
§ 23.1061 Installation.
(a) General. Each liquid-cooled engine must have an independent cooling system (including coolant tank) installed so that—
(1) Each coolant tank is supported so that tank loads are distributed over a large part of the tank surface;
(2) There are pads or other isolation means between the tank and its supports to prevent chafing.
(3) Pads or any other isolation means that is used must be nonabsorbent or must be treated to prevent absorption of flammable fluids; and
(4) No air or vapor can be trapped in any part of the system, except the coolant tank expansion space, during filling or during operation.
(b) Coolant tank. The tank capacity must be at least one gallon, plus 10 percent of the cooling system capacity. In addition—
(1) Each coolant tank must be able to withstand the vibration, inertia, and fluid loads to which it may be subjected in operation;
(2) Each coolant tank must have an expansion space of at least 10 percent of the total cooling system capacity; and
(3) It must be impossible to fill the expansion space inadvertently with the airplane in the normal ground attitude.
(c) Filler connection. Each coolant tank filler connection must be marked as specified in §23.1557(c). In addition—
(1) Spilled coolant must be prevented from entering the coolant tank compartment or any part of the airplane other than the tank itself; and
(2) Each recessed coolant filler connection must have a drain that discharges clear of the entire airplane.
(d) Lines and fittings. Each coolant system line and fitting must meet the requirements of §23.993, except that the inside diameter of the engine coolant inlet and outlet lines may not be less than the diameter of the corresponding engine inlet and outlet connections.
(e) Radiators. Each coolant radiator must be able to withstand any vibration, inertia, and coolant pressure load to which it may normally be subjected. In addition—
(1) Each radiator must be supported to allow expansion due to operating temperatures and prevent the transmittal of harmful vibration to the radiator; and
(2) If flammable coolant is used, the air intake duct to the coolant radiator must be located so that (in case of fire) flames from the nacelle cannot strike the radiator.
(f) Drains. There must be an accessible drain that—
(1) Drains the entire cooling system (including the coolant tank, radiator, and the engine) when the airplane is in the normal ground altitude;
(2) Discharges clear of the entire airplane; and
(3) Has means to positively lock it closed.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–43, 58 FR 18973, Apr. 9, 1993]
§ 23.1063 Coolant tank tests.
Each coolant tank must be tested under §23.965, except that—
(a) The test required by §23.965(a)(1) must be replaced with a similar test using the sum of the pressure developed during the maximum ultimate acceleration with a full tank or a pressure of 3.5 pounds per square inch, whichever is greater, plus the maximum working pressure of the system; and
(b) For a tank with a nonmetallic liner the test fluid must be coolant rather than fuel as specified in §23.965(d), and the slosh test on a specimen liner must be conducted with the coolant at operating temperature.
Induction System
§ 23.1091 Air induction system.
(a) The air induction system for each engine and auxiliary power unit and their accessories must supply the air required by that engine and auxiliary power unit and their accessories under the operating conditions for which certification is requested.
(b) Each reciprocating engine installation must have at least two separate air intake sources and must meet the following:
(1) Primary air intakes may open within the cowling if that part of the cowling is isolated from the engine accessory section by a fire-resistant diaphragm or if there are means to prevent the emergence of backfire flames.
(2) Each alternate air intake must be located in a sheltered position and may not open within the cowling if the emergence of backfire flames will result in a hazard.
(3) The supplying of air to the engine through the alternate air intake system may not result in a loss of excessive power in addition to the power loss due to the rise in air temperature.
(4) Each automatic alternate air door must have an override means accessible to the flight crew.
(5) Each automatic alternate air door must have a means to indicate to the flight crew when it is not closed.
(c) For turbine engine powered airplanes—
(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 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 intake ducts in hazardous quantities. The air intake ducts must be located or protected so as to minimize the hazard of ingestion of foreign matter during takeoff, landing, and taxiing.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–7, 34 FR 13095, Aug. 13, 1969; Amdt. 23–43, 58 FR 18973, Apr. 9, 1993; 58 FR 27060, May 6, 1993; Amdt. 23–51, 61 FR 5137, Feb. 9, 1996]
§ 23.1093 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 °F.—
(1) Each airplane with sea level engines using conventional venturi carburetors has a preheater that can provide a heat rise of 90 °F. with the engines at 75 percent of maximum continuous power;
(2) Each airplane with altitude engines using conventional venturi carburetors has a preheater that can provide a heat rise of 120 °F. with the engines at 75 percent of maximum continuous power;
(3) Each airplane with altitude engines using fuel metering device tending to prevent icing has a preheater that, with the engines at 60 percent of maximum continuous power, can provide a heat rise of—
(i) 100 °F.; or
(ii) 40 °F., if a fluid deicing system meeting the requirements of §§23.1095 through 23.1099 is installed;
(4) Each airplane with sea level engine(s) using fuel metering device tending to prevent icing has a sheltered alternate source of air with a preheat of not less than 60 °F with the engines at 75 percent of maximum continuous power;
(5) Each airplane with sea level or altitude engine(s) using fuel injection systems having metering components on which impact ice may accumulate has a preheater capable of providing a heat rise of 75 °F when the engine is operating at 75 percent of its maximum continuous power; and
(6) Each airplane with sea level or altitude engine(s) using fuel injection systems not having fuel metering components projecting into the airstream on which ice may form, and introducing fuel into the air induction system downstream of any components or other obstruction on which ice produced by fuel evaporation may form, has a sheltered alternate source of air with a preheat of not less than 60 °F with the engines at 75 percent of its maximum continuous power.
(b) Turbine engines. (1) Each turbine engine and its air inlet system must operate throughout the flight power range of the engine (including idling), without the accumulation of ice on engine or inlet system 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 of part 25 of this chapter; and
(ii) In snow, both falling and blowing, 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° and 30 °F (between −9° and −1 °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) Reciprocating engines with Superchargers. For airplanes with reciprocating engines having superchargers to pressurize the air before it enters the fuel metering device, 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 altitudes and operating condition because of supercharging.
[Amdt. 23-7, 34 FR 13095, Aug. 13, 1969, as amended by Amdt. 23–15, 39 FR 35460, Oct. 1, 1974; Amdt. 23–17, 41 FR 55465, Dec. 20, 1976; Amdt. 23–18, 42 FR 15041, Mar. 17, 1977; Amdt. 23–29, 49 FR 6847, Feb. 23, 1984; Amdt. 23–43, 58 FR 18973, Apr. 9, 1993; Amdt. 23–51, 61 FR 5137, Feb. 9, 1996]
§ 23.1095 Carburetor deicing fluid flow rate.
(a) If a carburetor deicing fluid system is used, it must be able to simultaneously supply each engine with a rate of fluid flow, expressed in pounds per hour, of not less than 2.5 times the square root of the maximum continuous power of the engine.
(b) The fluid must be introduced into the air induction system—
(1) Close to, and upstream of, the carburetor; and
(2) So that it is equally distributed over the entire cross section of the induction system air passages.
§ 23.1097 Carburetor deicing fluid system capacity.
(a) The capacity of each carburetor deicing fluid system—
(1) May not be less than the greater of—
(i) That required to provide fluid at the rate specified in §23.1095 for a time equal to three percent of the maximum endurance of the airplane; or
(ii) 20 minutes at that flow rate; and
(2) Need not exceed that required for two hours of operation.
(b) If the available preheat exceeds 50 °F. but is less than 100 °F., the capacity of the system may be decreased in proportion to the heat rise available in excess of 50 °F.
§ 23.1099 Carburetor deicing fluid system detail design.
Each carburetor deicing fluid system must meet the applicable requirements for the design of a fuel system, except as specified in §§23.1095 and 23.1097.
§ 23.1101 Induction air preheater design.
Each exhaust-heated, induction air preheater must be designed and constructed to—
(a) Ensure ventilation of the preheater when the induction air preheater is not being used during engine operation;
(b) Allow inspection of the exhaust manifold parts that it surrounds; and
(c) Allow inspection of critical parts of the preheater itself.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–43, 58 FR 18974, Apr. 9, 1993]
§ 23.1103 Induction system ducts.
(a) Each induction system duct must have a drain to prevent the accumulation of fuel or moisture in the normal ground and flight attitudes. No drain may discharge where it will cause a fire hazard.
(b) Each duct connected to components between which relative motion could exist must have means for flexibility.
(c) Each flexible induction system duct must be capable of withstanding the effects of temperature extremes, fuel, oil, water, and solvents to which it is expected to be exposed in service and maintenance without hazardous deterioration or delamination.
(d) For reciprocating engine installations, each induction system duct must be—
(1) Strong enough to prevent induction system failures resulting from normal backfire conditions; and
(2) Fire resistant in any compartment for which a fire extinguishing system is required.
(e) Each inlet system duct for an auxiliary power unit must be—
(1) Fireproof within the auxiliary power unit compartment;
(2) Fireproof for a sufficient distance upstream of the auxiliary power unit compartment to prevent hot gas reverse flow from burning through the duct and entering any other compartment of the airplane in which a hazard would be created by the entry of the hot gases;
(3) Constructed of materials suitable to the environmental conditions expected in service, except in those areas requiring fireproof or fire resistant materials; and
(4) Constructed of materials that will not absorb or trap hazardous quantities of flammable fluids that could be ignited by a surge or reverse-flow condition.
(f) Induction system ducts that supply air to a cabin pressurization system must be suitably constructed of material that will not produce hazardous quantities of toxic gases or isolated to prevent hazardous quantities of toxic gases from entering the cabin during a powerplant fire.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–7, 34 FR 13095, Aug. 13, 1969; Amdt. 23–43, 58 FR 18974, Apr. 9, 1993]
§ 23.1105 Induction system screens.
If induction system screens are used—
(a) Each screen must be upstream of the carburetor or fuel injection system.
(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—
(1) The available preheat is at least 100 °F.; and
(2) The screen 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.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1996, as amended by Amdt. 23–51, 61 FR 5137, Feb. 9, 1996]
§ 23.1107 Induction system filters.
If an air filter is used to protect the engine against foreign material particles in the induction air supply—
(a) Each air filter must be capable of withstanding the effects of temperature extremes, rain, fuel, oil, and solvents to which it is expected to be exposed in service and maintenance; and
(b) Each air filter shall have a design feature to prevent material separated from the filter media from interfering with proper fuel metering operation.
[Doc. No. 26344, 58 FR 18974, Apr. 9, 1993, as amended by Amdt. 23–51, 61 FR 5137, Feb. 9, 1996]
§ 23.1109 Turbocharger bleed air system.
The following applies to turbocharged bleed air systems used for cabin pressurization:
(a) The cabin air system may not be subject to hazardous contamination following any probable failure of the turbocharger or its lubrication system.
(b) The turbocharger supply air must be taken from a source where it cannot be contaminated by harmful or hazardous gases or vapors following any probable failure or malfunction of the engine exhaust, hydraulic, fuel, or oil system.
[Amdt. 23–42, 56 FR 354, Jan. 3, 1991]
§ 23.1111 Turbine engine bleed air system.
For turbine engine bleed air systems, the following apply:
(a) No hazard may result if duct rupture or failure occurs anywhere between the engine port and the airplane unit served by the bleed air.
(b) The effect on airplane and engine performance of using maximum bleed air must be established.
(c) Hazardous contamination of cabin air systems may not result from failures of the engine lubricating system.
[Amdt. 23–7, 34 FR 13095, Aug. 13, 1969, as amended by Amdt. 23–17, 41 FR 55465, Dec. 20, 1976]
Exhaust System
§ 23.1121 General.
For powerplant and auxiliary power unit installations, the following apply—
(a) Each exhaust system must ensure safe disposal of exhaust gases without fire hazard or carbon monoxide contamination in any personnel compartment.
(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 exhaust system must be separated by fireproof shields from adjacent flammable parts of the airplane that are outside of the engine and auxiliary power unit compartments.
(d) No exhaust gases may discharge dangerously near any fuel or oil system drain.
(e) No exhaust gases may be discharged 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) If significant traps exist, each turbine engine and auxiliary power unit exhaust system must have drains discharging clear of the airplane, in any normal ground and flight attitude, to prevent fuel accumulation after the failure of an attempted engine or auxiliary power unit start.
(h) Each exhaust heat exchanger must incorporate means to prevent blockage of the exhaust port after any internal heat exchanger failure.
(i) For the purpose of compliance with §23.603, the failure of any part of the exhaust system will be considered to adversely affect safety.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–7, 34 FR 13095, Aug. 13, 1969; Amdt. 23–18, 42 FR 15042, Mar. 17, 1977; Amdt. 23–43, 58 FR 18974, Apr. 9, 1993; Amdt. 23–51, 61 FR 5137, Feb. 9, 1996]
§ 23.1123 Exhaust system.
(a) Each exhaust system must be fireproof and corrosion-resistant, and must have means to prevent failure due to expansion by operating temperatures.
(b) Each exhaust system must be supported to withstand the vibration and inertia loads to which it may be subjected in operation.
(c) Parts of the system connected to components between which relative motion could exist must have means for flexibility.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–43, 58 FR 18974, Apr. 9, 1993]
§ 23.1125 Exhaust heat exchangers.
For reciprocating engine powered airplanes the following apply:
(a) Each exhaust heat exchanger must be constructed and installed to withstand the vibration, inertia, and other loads that it may be subjected to in normal 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 inspection of critical parts of each exchanger; and
(3) Each exchanger must have cooling provisions wherever it is subject to contact with exhaust gases.
(b) Each heat exchanger used for heating ventilating air must be constructed so that exhaust gases may not enter the ventilating air.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–17, 41 FR 55465, Dec. 20, 1976]
Powerplant Controls and Accessories
§ 23.1141 Powerplant controls: General.
(a) Powerplant controls must be located and arranged under §23.777 and marked under §23.1555(a).
(b) Each flexible control must be shown to be suitable for the particular application.
(c) Each control must be able to maintain any necessary position without—
(1) Constant attention by flight crew members; or
(2) Tendency to creep due to control loads or vibration.
(d) Each control must be able to withstand operating loads without failure or excessive deflection.
(e) For turbine engine powered airplanes, no single failure or malfunction, or probable combination thereof, in any powerplant control system may cause the failure of any powerplant function necessary for safety.
(f) The portion of each powerplant control located in the engine compartment that is required to be operated in the event of fire must be at least fire resistant.
(g) 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. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–7, 34 FR 13095, Aug. 13, 1969; Amdt. 23–14, 38 FR 31823, Nov. 19, 1973; Amdt. 23–18, 42 FR 15042, Mar. 17, 1977; Amdt. 23–51, 61 FR 5137, Feb. 9, 1996]
§ 23.1142 Auxiliary power unit controls.
Means must be provided on the flight deck for the starting, stopping, monitoring, and emergency shutdown of each installed auxiliary power unit.
[Doc. No. 26344, 58 FR 18974, Apr. 9, 1993]
§ 23.1143 Engine controls.
(a) There must be a separate power or thrust control for each engine and a separate control for each supercharger that requires a control.
(b) Power, thrust, and supercharger controls must be arranged to allow—
(1) Separate control of each engine and each supercharger; and
(2) Simultaneous control of all engines and all superchargers.
(c) Each power, thrust, or supercharger control must give a positive and immediate responsive means of controlling its engine or supercharger.
(d) The power, thrust, or supercharger controls for each engine or supercharger must be independent of those for every other engine or supercharger.
(e) 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.
(f) If a power, thrust, or a fuel control (other than a mixture control) incorporates a fuel shutoff feature, the control must have a means to prevent the inadvertent movement of the control into the off 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.
(g) For reciprocating single-engine airplanes, each power or thrust control must be designed so that if the control separates at the engine fuel metering device, the airplane is capable of continued safe flight and landing.
[Amdt. 23–7, 34 FR 13095, Aug. 13, 1969, as amended by Amdt. 23–17, 41 FR 55465, Dec. 20, 1976; Amdt. 23–29, 49 FR 6847, Feb. 23, 1984; Amdt. 23–43, 58 FR 18974, Apr. 9, 1993; Amdt. 23–51, 61 FR 5137, Feb. 9, 1996]
§ 23.1145 Ignition switches.
(a) Ignition switches must control and shut off each ignition circuit on each engine.
(b) There must be means to quickly shut off all ignition on multiengine airplanes 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. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23–18, 42 FR 15042, Mar. 17, 1977; Amdt. 23–43, 58 FR 18974, Apr. 9, 1993]
§ 23.1147 Mixture controls.
(a) If there are mixture controls, each engine must have a separate control, and each mixture control must have guards or must be shaped or arranged to prevent confusion by feel with other controls.
(1) The controls must be grouped and arranged to allow—
(i) Separate control of each engine; and
(ii) Simultaneous control of all engines.
(2) The controls must require a separate and distinct operation to move the control toward lean or shut-off position.
(b) For reciprocating single-engine airplanes, each manual engine mixture control must be designed so that, if the control separates at the engine fuel metering device, the airplane is capable of continued safe flight and landing.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–7, 34 FR 13096, Aug. 13, 1969; Amdt. 23–33, 51 FR 26657, July 24, 1986; Amdt. 23–43, 58 FR 18974, Apr. 9, 1993]
§ 23.1149 Propeller speed and pitch controls.
(a) If there are propeller speed or pitch controls, they must be grouped and arranged to allow—
(1) Separate control of each propeller; and
(2) Simultaneous control of all propellers.
(b) The controls must allow ready synchronization of all propellers on multiengine airplanes.
§ 23.1153 Propeller feathering controls.
If there are propeller feathering controls installed, it must be possible to feather each propeller separately. Each control must have a means to prevent inadvertent operation.
[Doc. No. 27804, 61 FR 5138, Feb. 9, 1996]
§ 23.1155 Turbine engine reverse thrust and propeller pitch settings below the flight regime.
For turbine engine installations, 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. 23–7, 34 FR 13096, Aug. 13, 1969]
§ 23.1157 Carburetor air temperature controls.
There must be a separate carburetor air temperature control for each engine.
§ 23.1163 Powerplant accessories.
(a) Each engine mounted accessory must—
(1) Be approved for mounting on the engine involved and use the provisions on the engines for mounting; or
(2) Have torque limiting means on all accessory drives in order to prevent the torque limits established for those drives from being exceeded; and
(3) In addition to paragraphs (a)(1) or (a)(2) of this section, 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) Each generator rated at or more than 6 kilowatts must be designed and installed to minimize the probability of a fire hazard in the event it malfunctions.
(d) If the continued rotation of any accessory remotely driven by the engine is hazardous when malfunctioning occurs, a means to prevent rotation without interfering with the continued operation of the engine must be provided.
(e) Each accessory driven by a gearbox that is not approved as part of the powerplant driving the gearbox must—
(1) Have torque limiting means to prevent the torque limits established for the affected drive from being exceeded;
(2) Use the provisions on the gearbox for mounting; and
(3) Be sealed to prevent contamination of the gearbox oil system and the accessory system.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–14, 38 FR 31823, Nov. 19, 1973; Amdt. 23–29, 49 FR 6847, Feb. 23, 1984; Amdt. 23–34, 52 FR 1832, Jan. 15, 1987; Amdt. 23–42, 56 FR 354, Jan. 3, 1991]
§ 23.1165 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 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) There must be means to warn appropriate crewmembers if malfunctioning of any part of the electrical system is causing the continuous discharge of any battery used for engine ignition.
(e) Each turbine engine ignition system must be independent of any electrical circuit that is not used for assisting, controlling, or analyzing the operation of that system.
(f) In addition, for commuter category airplanes, each turbopropeller ignition system must be an essential electrical load.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–17, 41 FR 55465 Dec. 20, 1976; Amdt. 23–34, 52 FR 1833, Jan. 15, 1987]
Powerplant Fire Protection
§ 23.1181 Designated fire zones; regions included.
Designated fire zones are—
(a) For reciprocating engines—
(1) The power section;
(2) The accessory section;
(3) Any complete powerplant compartment in which there is no isolation between the power section and the accessory section.
(b) For turbine engines—
(1) The compressor and accessory sections;
(2) The combustor, turbine and tailpipe sections that contain lines or components carrying flammable fluids or gases.
(3) Any complete powerplant compartment in which there is no isolation between compressor, accessory, combustor, turbine, and tailpipe sections.
(c) Any auxiliary power unit compartment; and
(d) Any fuel-burning heater, and other combustion equipment installation described in §23.859.
[Doc. No. 26344, 58 FR 18975, Apr. 9, 1993, as amended by Amdt. 23–51, 61 FR 5138, Feb. 9, 1996]
§ 23.1182 Nacelle areas behind firewalls.
Components, lines, and fittings, except those subject to the provisions of §23.1351(e), located behind the engine-compartment firewall must be constructed of such materials and located at such distances from the firewall that they will not suffer damage sufficient to endanger the airplane if a portion of the engine side of the firewall is subjected to a flame temperature of not less than 2000 °F for 15 minutes.
[Amdt. 23–14, 38 FR 31816, Nov. 19, 1973]
§ 23.1183 Lines, fittings, and components.
(a) Except as provided in paragraph (b) of this section, each component, line, and fitting carrying flammable fluids, gas, or air in any area subject to engine fire conditions must be at least fire resistant, except that flammable fluid tanks and supports which are part of and attached to the engine 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 so as to safeguard against the ignition of leaking flammable fluid. Flexible hose assemblies (hose and end fittings) must be shown to be suitable for the particular application. 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. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–5, 32 FR 6912, May 5, 1967; Amdt. 23–15, 39 FR 35460, Oct. 1, 1974; Amdt. 23–29, 49 FR 6847, Feb. 23, 1984; Amdt. 23–51, 61 FR 5138, Feb. 9, 1996]
§ 23.1189 Shutoff means.
(a) For each multiengine airplane the following apply:
(1) Each engine installation must have means to shut off or otherwise prevent hazardous quantities of fuel, oil, deicing fluid, and other flammable liquids from flowing into, within, or through any engine compartment, except in lines, fittings, and components forming an integral part of an engine.
(2) The closing of the fuel shutoff valve for any engine may not make any fuel unavailable to the remaining engines that would be available to those engines with that valve open.
(3) Operation of any shutoff means may not interfere with the later emergency operation of other equipment such as propeller feathering devices.
(4) Each shutoff must be outside of the engine compartment unless an equal degree of safety is provided with the shutoff inside the compartment.
(5) Not more than one quart of flammable fluid may escape into the engine compartment after engine shutoff. For those installations where the flammable fluid that escapes after shutdown cannot be limited to one quart, it must be demonstrated that this greater amount can be safely contained or drained overboard.
(6) There must be means to guard against inadvertent operation of each shutoff means, and to make it possible for the crew to reopen the shutoff means in flight after it has been closed.
(b) Turbine engine installations need not have an engine oil system shutoff if—
(1) The oil tank is integral with, or mounted on, the engine; and
(2) All oil system components external to the engine are fireproof or located in areas not subject to engine fire conditions.
(c) Power operated valves must have means to indicate to the flight crew when the valve has reached the selected position and must be designed so that the valve will not move from the selected position under vibration conditions likely to exist at the valve location.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–7, 34 FR 13096, Aug. 13, 1969; Amdt. 23–14, 38 FR 31823, Nov. 19, 1973; Amdt. 23–29, 49 FR 6847, Feb. 23, 1984; Amdt. 23–43, 58 FR 18975, Apr. 9, 1993]
§ 23.1191 Firewalls.
(a) Each engine, auxiliary power unit, fuel burning heater, and other combustion equipment, must be isolated from the rest of the airplane by firewalls, shrouds, or equivalent means.
(b) Each firewall or shroud must be constructed so that no hazardous quantity of liquid, gas, or flame can pass from the compartment created by the firewall or shroud to other parts of the airplane.
(c) Each opening in the firewall or shroud must be sealed with close fitting, fireproof grommets, bushings, or firewall fittings.
(d) [Reserved]
(e) Each firewall and shroud must be fireproof and protected against corrosion.
(f) Compliance with the criteria for fireproof materials or components must be shown as follows:
(1) The flame to which the materials or components are subjected must be 2,000 ±150 °F.
(2) Sheet materials approximately 10 inches square must be subjected to the flame from a suitable burner.
(3) The flame must be large enough to maintain the required test temperature over an area approximately five inches square.
(g) Firewall materials and fittings must resist flame penetration for at least 15 minutes.
(h) The following materials may be used in firewalls or shrouds without being tested as required by this section:
(1) Stainless steel sheet, 0.015 inch thick.
(2) Mild steel sheet (coated with aluminum or otherwise protected against corrosion) 0.018 inch thick.
(3) Terne plate, 0.018 inch thick.
(4) Monel metal, 0.018 inch thick.
(5) Steel or copper base alloy firewall fittings.
(6) Titanium sheet, 0.016 inch thick.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964, as amended by Amdt. 23–43, 58 FR 18975, Apr. 9, 1993; 58 FR 27060, May 6, 1993; Amdt. 23–51, 61 FR 5138, Feb. 9, 1996]
§ 23.1192 Engine accessory compartment diaphragm.
For aircooled radial engines, the engine power section and all portions of the exhaust sytem must be isolated from the engine accessory compartment by a diaphragm that meets the firewall requirements of §23.1191.
[Amdt. 23–14, 38 FR 31823, Nov. 19, 1973]
§ 23.1193 Cowling and nacelle.
(a) Each cowling must be constructed and supported so that it can resist any vibration, inertia, and air loads to which it may be subjected in operation.
(b) There must be means for rapid and complete drainage of each part of the cowling in the normal ground and flight attitudes. Drain operation may be shown by test, analysis, or both, to ensure that under normal aerodynamic pressure distribution expected in service each drain will operate as designed. No drain may discharge where it will cause a fire hazard.
(c) Cowling must be at least fire resistant.
(d) Each part behind an opening in the engine compartment cowling must be at least fire resistant for a distance of at least 24 inches aft of the opening.
(e) Each part of the cowling subjected to high temperatures due to its nearness to exhaust sytem ports or exhaust gas impingement, must be fire proof.
(f) Each nacelle of a multiengine airplane with supercharged engines must be designed and constructed so that with the landing gear retracted, a fire in the engine compartment will not burn through a cowling or nacelle and enter a nacelle area other than the engine compartment.
(g) In addition, for commuter category airplanes, the airplane must be designed so that no fire originating in any engine compartment can enter, either through openings or by burn-through, any other region where it would create additional hazards.
[Doc. No. 4080, 29 FR 17955, Dec. 18, 1964; 30 FR 258, Jan. 9, 1965, as amended by Amdt. 23–18, 42 FR 15042, Mar. 17, 1977; Amdt. 23–34, 52 FR 1833, Jan. 15, 1987; 58 FR 18975, Apr. 9, 1993]
§ 23.1195 Fire extinguishing systems.
(a) For commuter category airplanes, fire extinguishing systems must be installed and compliance shown with the following:
(1) Except for combustor, turbine, and tailpipe sections of turbine-engine installations that contain lines or components carrying flammable fluids or gases for which a fire originating in these sections is shown to be controllable, a fire extinguisher system must serve each engine compartment;
(2) The fire extinguishing system, the quantity of the extinguishing agent, the rate of discharge, and the discharge distribution must be adequate to extinguish fires. An individual “one shot” system may be used.
(3) The fire extinguishing system for a nacelle must be able to simultaneously protect each compartment of the nacelle for which protection is provided.
(b) If an auxiliary power unit is installed in any airplane certificated to this part, that auxiliary power unit compartment must be served by a fire extinguishing system meeting the requirements of paragraph (a)(2) of this section.
[Amdt. 23–34, 52 FR 1833, Jan. 15, 1987, as amended by Amdt. 23–43, 58 FR 18975, Apr. 9, 1993]
§ 23.1197 Fire extinguishing agents.
For commuter category airplanes, the following applies:
(a) Fire extinguishing agents must—
(1) Be capable of extinguishing flames emanating from any burning of fluids or other combustible materials in the area protected by the fire extinguishing system; and
(2) Have thermal stability over the temperature range likely to be experienced in the compartment in which they are stored.
(b) If any toxic extinguishing agent is used, provisions must be made to prevent harmful concentrations of fluid or fluid vapors (from leakage during normal operation of the airplane or as a result of discharging the fire extinguisher on the ground or in flight) from entering any personnel compartment, even though a defect may exist in the extinguishing system. This must be shown by test except for built-in carbon dioxide fuselage compartment fire extinguishing systems for which—
(1) Five pounds or less of carbon dioxide will be discharged, under established fire control procedures, into any fuselage compartment; or
(2) Protective breathing equipment is available for each flight crewmember on flight deck duty.
[Amdt. 23–34, 52 FR 1833, Jan. 15, 1987]
§ 23.1199 Extinguishing agent containers.
For commuter category airplanes, the following applies:
(a) Each extinguishing agent container must have a pressure relief to prevent bursting of the container by excessive internal pressures.
(b) The discharge end of each discharge line from a pressure relief connection must be located so that discharge of the fire extinguishing agent would not damage the airplane. The line must also be located or protected to prevent clogging caused by ice or other foreign matter.
(c) A means must be provided for each fire extinguishing agent container to indicate that the container has discharged or that the charging pressure is below the established minimum necessary for proper functioning.
(d) The temperature of each container must be maintained, under intended operating conditions, to prevent the pressure in the container from—
(1) Falling below that necessary to provide an adequate rate of discharge; or
(2) Rising high enough to cause premature discharge.
(e) If a pyrotechnic capsule is used to discharge the extinguishing agent, each container must be installed so that temperature conditions will not cause hazardous deterioration of the pyrotechnic capsule.
[Amdt. 23–34, 52 FR 1833, Jan. 15, 1987; 52 FR 34745, Sept. 14, 1987]
§ 23.1201 Fire extinguishing systems materials.
For commuter category airplanes, the following apply:
(a) No material in any fire extinguishing system may react chemically with any extinguishing agent so as to create a hazard.
(b) Each system component in an engine compartment must be fireproof.
[Amdt. 23–34, 52 FR 1833, Jan. 15, 1987; 52 FR 7262, Mar. 9, 1987]
§ 23.1203 Fire detector system.
(a) There must be means that ensure the prompt detection of a fire in—
(1) An engine compartment of—
(i) Multiengine turbine powered airplanes;
(ii) Multiengine reciprocating engine powered airplanes incorporating turbochargers;
(iii) Airplanes with engine(s) located where they are not readily visible from the cockpit; and
(iv) All commuter category airplanes.
(2) The auxiliary power unit compartment of any airplane incorporating an auxiliary power unit.
(b) Each fire detector must be constructed and installed to withstand the vibration, inertia, and other loads to which it may be subjected in operation.
(c) No fire detector may be affected by any oil, water, other fluids, or fumes that might be present.
(d) There must be means to allow the crew to check, in flight, the functioning of each fire detector electric circuit.
(e) Wiring and other components of each fire detector system in a designated fire zone must be at least fire resistant.
[Amdt. 23–18, 42 FR 15042, Mar. 17, 1977, as amended by Amdt. 23–34, 52 FR 1833, Jan. 15, 1987; Amdt. 23–43, 58 FR 18975, Apr. 9, 1993; Amdt. 23–51, 61 FR 5138, Feb. 9, 1996]
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