Part 33 — CCAR-33 航空发动机适航标准 航空发动机（活塞式、涡喷、涡扇、涡桨）的适航审定标准。 FAR Part 33 原文 Part 33 Authority: Source: § 33.1 Applicability. (a) This part prescribes airworthiness standards for the issue of type certificates and changes to those certificates, for aircraft engines. (b) Each person who applies under part 21 for such a certificate or change must show compliance with the applicable requirements of this part and the applicable requirements of part 34 of this chapter. § 33.3 General. Each applicant must show that the aircraft engine concerned meets the applicable requirements of this part. § 33.4 Instructions for Continued Airworthiness. The applicant must prepare Instructions for Continued Airworthiness in accordance with appendix A to this part that are acceptable to the Administrator. The instructions may be incomplete at type certification if a program exists to ensure their completion prior to delivery of the first aircraft with the engine installed, or upon issuance of a standard certificate of airworthiness for the aircraft with the engine installed, whichever occurs later. § 33.5 Instruction manual for installing and operating the engine. Each applicant must prepare and make available to the Administrator prior to the issuance of the type certificate, and to the owner at the time of delivery of the engine, approved instructions for installing and operating the engine. The instructions must include at least the following: (a) Installation instructions. (1) The location of engine mounting attachments, the method of attaching the engine to the aircraft, and the maximum allowable load for the mounting attachments and related structure. (2) The location and description of engine connections to be attached to accessories, pipes, wires, cables, ducts, and cowling. (3) An outline drawing of the engine including overall dimensions. (4) A definition of the physical and functional interfaces with the aircraft and aircraft equipment, including the propeller when applicable. (5) Where an engine system relies on components that are not part of the engine type design, the interface conditions and reliability requirements for those components upon which engine type certification is based must be specified in the engine installation instructions directly or by reference to appropriate documentation. (6) A list of the instruments necessary for control of the engine, including the overall limits of accuracy and transient response required of such instruments for control of the operation of the engine, must also be stated so that the suitability of the instruments as installed may be assessed. (b) Operation instructions. (1) The operating limitations established by the Administrator. (2) The power or thrust ratings and procedures for correcting for nonstandard atmosphere. (3) The recommended procedures, under normal and extreme ambient conditions for— (i) Starting; (ii) Operating on the ground; and (iii) Operating during flight. (4) For rotorcraft engines having one or more OEI ratings, applicants must provide data on engine performance characteristics and variability to enable the aircraft manufacturer to establish aircraft power assurance procedures. (5) A description of the primary and all alternate modes, and any back-up system, together with any associated limitations, of the engine control system and its interface with the aircraft systems, including the propeller when applicable. (c) Safety analysis assumptions. The assumptions of the safety analysis as described in § 33.75(d) with respect to the reliability of safety devices, instrumentation, early warning devices, maintenance checks, and similar equipment or procedures that are outside the control of the engine manufacturer. § 33.7 Engine ratings and operating limitations. (a) Engine ratings and operating limitations are established by the Administrator and included in the engine certificate data sheet specified in § 21.41 of this chapter, including ratings and limitations based on the operating conditions and information specified in this section, as applicable, and any other information found necessary for safe operation of the engine. (b) For reciprocating engines, ratings and operating limitations are established relating to the following: (1) Horsepower or torque, r.p.m., manifold pressure, and time at critical pressure altitude and sea level pressure altitude for— (i) Rated maximum continuous power (relating to unsupercharged operation or to operation in each supercharger mode as applicable); and (ii) Rated takeoff power (relating to unsupercharged operation or to operation in each supercharger mode as applicable). (2) Fuel grade or specification. (3) Oil grade or specification. (4) Temperature of the— (i) Cylinder; (ii) Oil at the oil inlet; and (iii) Turbosupercharger turbine wheel inlet gas. (5) Pressure of— (i) Fuel at the fuel inlet; and (ii) Oil at the main oil gallery. (6) Accessory drive torque and overhang moment. (7) Component life. (8) Turbosupercharger turbine wheel r.p.m. (c) For turbine engines, ratings and operating limitations are established relating to the following: (1) Horsepower, torque, or thrust, r.p.m., gas temperature, and time for— (i) Rated maximum continuous power or thrust (augmented); (ii) Rated maximum continuous power or thrust (unaugmented); (iii) Rated takeoff power or thrust (augmented); (iv) Rated takeoff power or thrust (unaugmented); (v) Rated 30-minute OEI power; (vi) Rated 2 1/2 -minute OEI power; (vii) Rated continuous OEI power; and (viii) Rated 2-minute OEI Power; (ix) Rated 30-second OEI power; and (x) Auxiliary power unit (APU) mode of operation. (2) Fuel designation or specification. (3) Oil grade or specification. (4) Hydraulic fluid specification. (5) Temperature of— (i) Oil at a location specified by the applicant; (ii) Induction air at the inlet face of a supersonic engine, including steady state operation and transient over-temperature and time allowed; (iii) Hydraulic fluid of a supersonic engine; (iv) Fuel at a location specified by the applicant; and (v) External surfaces of the engine, if specified by the applicant. (6) Pressure of— (i) Fuel at the fuel inlet; (ii) Oil at a location specified by the applicant; (iii) Induction air at the inlet face of a supersonic engine, including steady state operation and transient overpressure and time allowed; and (iv) Hydraulic fluid. (7) Accessory drive torque and overhang moment. (8) Component life. (9) Fuel filtration. (10) Oil filtration. (11) Bleed air. (12) The number of start-stop stress cycles approved for each rotor disc and spacer. (13) Inlet air distortion at the engine inlet. (14) Transient rotor shaft overspeed r.p.m., and number of overspeed occurrences. (15) Transient gas overtemperature, and number of overtemperature occurrences. (16) Transient engine overtorque, and number of overtorque occurrences. (17) Maximum engine overtorque for turbopropeller and turboshaft engines incorporating free power turbines. (18) For engines to be used in supersonic aircraft, engine rotor windmilling rotational r.p.m. (d) In determining the engine performance and operating limitations, the overall limits of accuracy of the engine control system and of the necessary instrumentation as defined in § 33.5(a)(6) must be taken into account. § 33.8 Selection of engine power and thrust ratings. (a) Requested engine power and thrust ratings must be selected by the applicant. (b) Each selected rating must be for the lowest power or thrust that all engines of the same type may be expected to produce under the conditions used to determine that rating. § 33.11 Applicability. This subpart prescribes the general design and construction requirements for reciprocating and turbine aircraft engines. § 33.13 § 33.15 Materials. The suitability and durability of materials used in the engine must— (a) Be established on the basis of experience or tests; and (b) Conform to approved specifications (such as industry or military specifications) that ensure their having the strength and other properties assumed in the design data. § 33.17 Fire protection. (a) The design and construction of the engine and the materials used must minimize the probability of the occurrence and spread of fire during normal operation and failure conditions, and must minimize the effect of such a fire. In addition, the design and construction of turbine engines must minimize the probability of the occurrence of an internal fire that could result in structural failure or other hazardous effects. (b) Except as provided in paragraph (c) of this section, each external line, fitting, and other component, which contains or conveys flammable fluid during normal engine operation, must be fire resistant or fireproof, as determined by the Administrator. Components must be shielded or located to safeguard against the ignition of leaking flammable fluid. (c) A tank, which contains flammable fluids and any associated shut-off means and supports, which are part of and attached to the engine, must be fireproof either by construction or by protection unless damage by fire will not cause leakage or spillage of a hazardous quantity of flammable fluid. For a reciprocating engine having an integral oil sump of less than 23.7 liters capacity, the oil sump need not be fireproof or enclosed by a fireproof shield. (d) An engine component designed, constructed, and installed to act as a firewall must be: (1) Fireproof; (2) Constructed so that no hazardous quantity of air, fluid or flame can pass around or through the firewall; and, (3) Protected against corrosion; (e) In addition to the requirements of paragraphs (a) and (b) of this section, engine control system components that are located in a designated fire zone must be fire resistant or fireproof, as determined by the Administrator. (f) Unintentional accumulation of hazardous quantities of flammable fluid within the engine must be prevented by draining and venting. (g) Any components, modules, or equipment, which are susceptible to or are potential sources of static discharges or electrical fault currents must be designed and constructed to be properly grounded to the engine reference, to minimize the risk of ignition in external areas where flammable fluids or vapors could be present. § 33.19 Durability. (a) Engine design and construction must minimize the development of an unsafe condition of the engine between overhaul periods. The design of the compressor and turbine rotor cases must provide for the containment of damage from rotor blade failure. Energy levels and trajectories of fragments resulting from rotor blade failure that lie outside the compressor and turbine rotor cases must be defined. (b) Each component of the propeller blade pitch control system which is a part of the engine type design must meet the requirements of §§ 35.21, 35.23, 35.42 and 35.43 of this chapter. § 33.21 Engine cooling. Engine design and construction must provide the necessary cooling under conditions in which the airplane is expected to operate. § 33.23 Engine mounting attachments and structure. (a) The maximum allowable limit and ultimate loads for engine mounting attachments and related engine structure must be specified. (b) The engine mounting attachments and related engine structure must be able to withstand— (1) The specified limit loads without permanent deformation; and (2) The specified ultimate loads without failure, but may exhibit permanent deformation. § 33.25 Accessory attachments. The engine must operate properly with the accessory drive and mounting attachments loaded. Each engine accessory drive and mounting attachment must include provisions for sealing to prevent contamination of, or unacceptable leakage from, the engine interior. A drive and mounting attachment requiring lubrication for external drive splines, or coupling by engine oil, must include provisions for sealing to prevent unacceptable loss of oil and to prevent contamination from sources outside the chamber enclosing the drive connection. The design of the engine must allow for the examination, adjustment, or removal of each accessory required for engine operation. § 33.27 Turbine, compressor, fan, and turbosupercharger rotor overspeed. (a) For each fan, compressor, turbine, and turbosupercharger rotor, the applicant must establish by test, analysis, or a combination of both, that each rotor will not burst when operated in the engine for 5 minutes at whichever of the conditions defined in paragraph (b) of this section is the most critical with respect to the integrity of such a rotor. (1) Test rotors used to demonstrate compliance with this section that do not have the most adverse combination of material properties and dimensional tolerances must be tested at conditions which have been adjusted to ensure the minimum specification rotor possesses the required overspeed capability. This can be accomplished by increasing test speed, temperature, and/or loads. (2) When an engine test is being used to demonstrate compliance with the overspeed conditions listed in paragraph (b)(3) or (b)(4) of this section and the failure of a component or system is sudden and transient, it may not be possible to operate the engine for 5 minutes after the failure. Under these circumstances, the actual overspeed duration is acceptable if the required maximum overspeed is achieved. (b) When determining the maximum overspeed condition applicable to each rotor in order to comply with paragraphs (a) and (c) of this section, the applicant must evaluate the following rotor speeds taking into consideration the part's operating temperatures and temperature gradients throughout the engine's operating envelope: (1) 120 percent of the maximum permissible rotor speed associated with any of the engine ratings except one-engine-inoperative (OEI) ratings of less than 2 1/2 minutes. (2) 115 percent of the maximum permissible rotor speed associated with any OEI ratings of less than 2 1/2 minutes. (3) 105 percent of the highest rotor speed that would result from either: (i) The failure of the component or system which, in a representative installation of the engine, is the most critical with respect to overspeed when operating at any rating condition except OEI ratings of less than 2 1/2 minutes, or (ii) The failure of any component or system in a representative installation of the engine, in combination with any other failure of a component or system that would not normally be detected during a routine pre-flight check or during normal flight operation, that is the most critical with respect to overspeed, except as provided by paragraph (c) of this section, when operating at any rating condition except OEI ratings of less than 2 1/2 minutes. (4) 100 percent of the highest rotor speed that would result from the failure of the component or system which, in a representative installation of the engine, is the most critical with respect to overspeed when operating at any OEI rating of less than 2 1/2 minutes. (c) The highest overspeed that results from a complete loss of load on a turbine rotor, except as provided by paragraph (f) of this section, must be included in the overspeed conditions considered by paragraphs (b)(3)(i), (b)(3)(ii), and (b)(4) of this section, regardless of whether that overspeed results from a failure within the engine or external to the engine. The overspeed resulting from any other single failure must be considered when selecting the most limiting overspeed conditions applicable to each rotor. Overspeeds resulting from combinations of failures must also be considered unless the applicant can show that the probability of occurrence is not greater than extremely remote (probability range of 10 −7 to 10 −9 per engine flight hour). (d) In addition, the applicant must demonstrate that each fan, compressor, turbine, and turbosupercharger rotor complies with paragraphs (d)(1) and (d)(2) of this section for the maximum overspeed achieved when subjected to the conditions specified in paragraphs (b)(3) and (b)(4) of this section. The applicant must use the approach in paragraph (a) of this section which specifies the required test conditions. (1) Rotor Growth must not cause the engine to: (i) Catch fire, (ii) Release high-energy debris through the engine casing or result in a hazardous failure of the engine casing, (iii) Generate loads greater than those ultimate loads specified in § 33.23(a), or (iv) Lose the capability of being shut down. (2) Following an overspeed event and after continued operation, the rotor may not exhibit conditions such as cracking or distortion which preclude continued safe operation. (e) The design and functioning of engine control systems, instruments, and other methods not covered under § 33.28 must ensure that the engine operating limitations that affect turbine, compressor, fan, and turbosupercharger rotor structural integrity will not be exceeded in service. (f) Failure of a shaft section may be excluded from consideration in determining the highest overspeed that would result from a complete loss of load on a turbine rotor if the applicant: (1) Identifies the shaft as an engine life-limited-part and complies with § 33.70. (2) Uses material and design features that are well understood and that can be analyzed by well-established and validated stress analysis techniques. (3) Determines, based on an assessment of the environment surrounding the shaft section, that environmental influences are unlikely to cause a shaft failure. This assessment must include complexity of design, corrosion, wear, vibration, fire, contact with adjacent components or structure, overheating, and secondary effects from other failures or combination of failures. (4) Identifies and declares, in accordance with § 33.5, any assumptions regarding the engine installation in making the assessment described above in paragraph (f)(3) of this section. (5) Assesses, and considers as appropriate, experience with shaft sections of similar design. (6) Does not exclude the entire shaft. (g) If analysis is used to meet the overspeed requirements, then the analytical tool must be validated to prior overspeed test results of a similar rotor. The tool must be validated for each material. The rotor being certified must not exceed the boundaries of the rotors being used to validate the analytical tool in terms of geometric shape, operating stress, and temperature. Validation includes the ability to accurately predict rotor dimensional growth and the burst speed. The predictions must also show that the rotor being certified does not have lower burst and growth margins than rotors used to validate the tool. § 33.28 Engine control systems. (a) Applicability. These requirements are applicable to any system or device that is part of engine type design, that controls, limits, or monitors engine operation, and is necessary for the continued airworthiness of the engine. (b) Validation —(1) Functional aspects. The applicant must substantiate by tests, analysis, or a combination thereof, that the engine control system performs the intended functions in a manner which: (i) Enables selected values of relevant control parameters to be maintained and the engine kept within the approved operating limits over changing atmospheric conditions in the declared flight envelope; (ii) Complies with the operability requirements of §§ 33.51, 33.65 and 33.73, as appropriate, under all likely system inputs and allowable engine power or thrust demands, unless it can be demonstrated that failure of the control function results in a non-dispatchable condition in the intended application; (iii) Allows modulation of engine power or thrust with adequate sensitivity over the declared range of engine operating conditions; and (iv) Does not create unacceptable power or thrust oscillations. (2) Environmental limits. The applicant must demonstrate, when complying with §§ 33.53 or 33.91, that the engine control system functionality will not be adversely affected by declared environmental conditions, including electromagnetic interference (EMI), High Intensity Radiated Fields (HIRF), and lightning. The limits to which the system has been qualified must be documented in the engine installation instructions. (c) Control transitions. (1) The applicant must demonstrate that, when fault or failure results in a change from one control mode to another, from one channel to another, or from the primary system to the back-up system, the change occurs so that: (i) The engine does not exceed any of its operating limitations; (ii) The engine does not surge, stall, or experience unacceptable thrust or power changes or oscillations or other unacceptable characteristics; and (iii) There is a means to alert the flight crew if the crew is required to initiate, respond to, or be aware of the control mode change. The means to alert the crew must be described in the engine installation instructions, and the crew action must be described in the engine operating instructions; (2) The magnitude of any change in thrust or power and the associated transition time must be identified and described in the engine installation instructions and the engine operating instructions. (d) Engine control system failures. The applicant must design and construct the engine control system so that: (1) The rate for Loss of Thrust (or Power) Control (LOTC/LOPC) events, consistent with the safety objective associated with the intended application can be achieved; (2) In the full-up configuration, the system is single fault tolerant, as determined by the Administrator, for electrical or electronic failures with respect to LOTC/LOPC events; (3) Single failures of engine control system components do not result in a hazardous engine effect; and (4) Foreseeable failures or malfunctions leading to local events in the intended aircraft installation, such as fire, overheat, or failures leading to damage to engine control system components, do not result in a hazardous engine effect due to engine control system failures or malfunctions. (e) S ystem safety assessment. When complying with this section and § 33.75, the applicant must complete a System Safety Assessment for the engine control system. This assessment must identify faults or failures that result in a change in thrust or power, transmission of erroneous data, or an effect on engine operability producing a surge or stall together with the predicted frequency of occurrence of these faults or failures. (f) Protection systems. (1) The design and functioning of engine control devices and systems, together with engine instruments and operating and maintenance instructions, must provide reasonable assurance that those engine operating limitations that affect turbine, compressor, fan, and turbosupercharger rotor structural integrity will not be exceeded in service. (2) When electronic overspeed protection systems are provided, the design must include a means for testing, at least once per engine start/stop cycle, to establish the availability of the protection function. The means must be such that a complete test of the system can be achieved in the minimum number of cycles. If the test is not fully automatic, the requirement for a manual test must be contained in the engine instructions for operation. (3) When overspeed protection is provided through hydromechanical or mechanical means, the applicant must demonstrate by test or other acceptable means that the overspeed function remains available between inspection and maintenance periods. (g) Software. The applicant must design, implement, and verify all associated software to minimize the existence of errors by using a method, approved by the FAA, consistent with the criticality of the performed functions. (h) Aircraft-supplied data. Single failures leading to loss, interruption or corruption of aircraft-supplied data (other than thrust or power command signals from the aircraft), or data shared between engines must: (1) Not result in a hazardous engine effect for any engine; and (2) Be detected and accommodated. The accommodation strategy must not result in an unacceptable change in thrust or power or an unacceptable change in engine operating and starting characteristics. The applicant must evaluate and document in the engine installation instructions the effects of these failures on engine power or thrust, engine operability, and starting characteristics throughout the flight envelope. (i) Aircraft-supplied electrical power. (1) The applicant must design the engine control system so that the loss, malfunction, or interruption of electrical power supplied from the aircraft to the engine control system will not result in any of the following: (i) A hazardous engine effect, or (ii) The unacceptable transmission of erroneous data. (2) When an engine dedicated power source is required for compliance with paragraph (i)(1) of this section, its capacity should provide sufficient margin to account for engine operation below idle where the engine control system is designed and expected to recover engine operation automatically. (3) The applicant must identify and declare the need for, and the characteristics of, any electrical power supplied from the aircraft to the engine control system for starting and operating the engine, including transient and steady state voltage limits, in the engine instructions for installation. (4) Low voltage transients outside the power supply voltage limitations declared in paragraph (i)(3) of this section must meet the requirements of paragraph (i)(1) of this section. The engine control system must be capable of resuming normal operation when aircraft-supplied power returns to within the declared limits. (j) Air pressure signal. The applicant must consider the effects of blockage or leakage of the signal lines on the engine control system as part of the System Safety Assessment of paragraph (e) of this section and must adopt the appropriate design precautions. (k) Automatic availability and control of engine power for 30-second OEI rating. Rotorcraft engines having a 30-second OEI rating must incorporate a means, or a provision for a means, for automatic availability and automatic control of the 30-second OEI power within its operating limitations. (l) Engine shut down means. Means must be provided for shutting down the engine rapidly. (m) Programmable logic devices. The development of programmable logic devices using digital logic or other complex design technologies must provide a level of assurance for the encoded logic commensurate with the hazard associated with the failure or malfunction of the systems in which the devices are located. The applicant must provide evidence that the development of these devices has been done by using a method, approved by the FAA, that is consistent with the criticality of the performed function. § 33.29 Instrument connection. (a) Unless it is constructed to prevent its connection to an incorrect instrument, each connection provided for powerplant instruments required by aircraft airworthiness regulations or necessary to insure operation of the engine in compliance with any engine limitation must be marked to identify it with its corresponding instrument. (b) A connection must be provided on each turbojet engine for an indicator system to indicate rotor system unbalance. (c) Each rotorcraft turbine engine having a 30-second OEI rating and a 2-minute OEI rating must have a means or a provision for a means to: (1) Alert the pilot when the engine is at the 30-second OEI and the 2-minute OEI power levels, when the event begins, and when the time interval expires; (2) Automatically record each usage and duration of power at the 30-second OEI and 2-minute OEI levels; (3) Alert maintenance personnel in a positive manner that the engine has been operated at either or both of the 30-second and 2-minute OEI power levels, and permit retrieval of the recorded data; and (4) Enable routine verification of the proper operation of the above means. (d) The means, or the provision for a means, of paragraphs (c)(2) and (c)(3) of this section must not be capable of being reset in flight. (e) The applicant must make provision for the installation of instrumentation necessary to ensure operation in compliance with engine operating limitations. Where, in presenting the safety analysis, or complying with any other requirement, dependence is placed on instrumentation that is not otherwise mandatory in the assumed aircraft installation, then the applicant must specify this instrumentation in the engine installation instructions and declare it mandatory in the engine approval documentation. (f) As part of the System Safety Assessment of § 33.28(e), the applicant must assess the possibility and subsequent effect of incorrect fit of instruments, sensors, or connectors. Where necessary, the applicant must take design precautions to prevent incorrect configuration of the system. (g) The sensors, together with associated wiring and signal conditioning, must be segregated, electrically and physically, to the extent necessary to ensure that the probability of a fault propagating from instrumentation and monitoring functions to control functions, or vice versa, is consistent with the failure effect of the fault. (h) The applicant must provide instrumentation enabling the flight crew to monitor the functioning of the turbine cooling system unless appropriate inspections are published in the relevant manuals and evidence shows that: (1) Other existing instrumentation provides adequate warning of failure or impending failure; (2) Failure of the cooling system would not lead to hazardous engine effects before detection; or (3) The probability of failure of the cooling system is extremely remote. § 33.31 Applicability. This subpart prescribes additional design and construction requirements for reciprocating aircraft engines. § 33.33 Vibration. The engine must be designed and constructed to function throughout its normal operating range of crankshaft rotational speeds and engine powers without inducing excessive stress in any of the engine parts because of vibration and without imparting excessive vibration forces to the aircraft structure. § 33.34 Turbocharger rotors. Each turbocharger case must be designed and constructed to 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. § 33.35 Fuel and induction system. (a) The fuel system of the engine must be designed and constructed to supply an appropriate mixture of fuel to the cylinders throughout the complete operating range of the engine under all flight and atmospheric conditions. (b) The intake passages of the engine through which air or fuel in combination with air passes for combustion purposes must be designed and constructed to minimize the danger of ice accretion in those passages. The engine must be designed and constructed to permit the use of a means for ice prevention. (c) The type and degree of fuel filtering necessary for protection of the engine fuel system against foreign particles in the fuel must be specified. The applicant must show that foreign particles passing through the prescribed filtering means will not critically impair engine fuel system functioning. (d) Each passage in the induction system that conducts a mixture of fuel and air must be self-draining, to prevent a liquid lock in the cylinders, in all attitudes that the applicant establishes as those the engine can have when the aircraft in which it is installed is in the static ground attitude. (e) If provided as part of the engine, the applicant must show for each fluid injection (other than fuel) system and its controls that the flow of the injected fluid is adequately controlled. § 33.37 Ignition system. Each spark ignition engine must have a dual ignition system with at least two spark plugs for each cylinder and two separate electric circuits with separate sources of electrical energy, or have an ignition system of equivalent in-flight reliability. § 33.39 Lubrication system. (a) The lubrication system of the engine must be designed and constructed so that it will function properly in all flight attitudes and atmospheric conditions in which the airplane is expected to operate. In wet sump engines, this requirement must be met when only one-half of the maximum lubricant supply is in the engine. (b) The lubrication system of the engine must be designed and constructed to allow installing a means of cooling the lubricant. (c) The crankcase must be vented to the atmosphere to preclude leakage of oil from excessive pressure in the crankcase. § 33.41 Applicability. This subpart prescribes the block tests and inspections for reciprocating aircraft engines. § 33.42 General. Before each endurance test required by this subpart, the adjustment setting and functioning characteristic of each component having an adjustment setting and a functioning characteristic that can be established independent of installation on the engine must be established and recorded. § 33.43 Vibration test. (a) Each engine must undergo a vibration survey to establish the torsional and bending vibration characteristics of the crankshaft and the propeller shaft or other output shaft, over the range of crankshaft speed and engine power, under steady state and transient conditions, from idling speed to either 110 percent of the desired maximum continuous speed rating or 103 percent of the maximum desired takeoff speed rating, whichever is higher. The survey must be conducted using, for airplane engines, the same configuration of the propeller type which is used for the endurance test, and using, for other engines, the same configuration of the loading device type which is used for the endurance test. (b) The torsional and bending vibration stresses of the crankshaft and the propeller shaft or other output shaft may not exceed the endurance limit stress of the material from which the shaft is made. If the maximum stress in the shaft cannot be shown to be below the endurance limit by measurement, the vibration frequency and amplitude must be measured. The peak amplitude must be shown to produce a stress below the endurance limit; if not, the engine must be run at the condition producing the peak amplitude until, for steel shafts, 10 million stress reversals have been sustained without fatigue failure and, for other shafts, until it is shown that fatigue will not occur within the endurance limit stress of the material. (c) Each accessory drive and mounting attachment must be loaded, with the loads imposed by each accessory used only for an aircraft service being the limit load specified by the applicant for the drive or attachment point. (d) The vibration survey described in paragraph (a) of this section must be repeated with that cylinder not firing which has the most adverse vibration effect, in order to establish the conditions under which the engine can be operated safely in that abnormal state. However, for this vibration survey, the engine speed range need only extend from idle to the maximum desired takeoff speed, and compliance with paragraph (b) of this section need not be shown. § 33.45 Calibration tests. (a) Each engine must be subjected to the calibration tests necessary to establish its power characteristics and the conditions for the endurance test specified in § 33.49. The results of the power characteristics calibration tests form the basis for establishing the characteristics of the engine over its entire operating range of crankshaft rotational speeds, manifold pressures, fuel/air mixture settings, and altitudes. Power ratings are based upon standard atmospheric conditions with only those accessories installed which are essential for engine functioning. (b) A power check at sea level conditions must be accomplished on the endurance test engine after the endurance test. Any change in power characteristics which occurs during the endurance test must be determined. Measurements taken during the final portion of the endurance test may be used in showing compliance with the requirements of this paragraph. § 33.47 Detonation test. Each engine must be tested to establish that the engine can function without detonation throughout its range of intended conditions of operation. § 33.49 Endurance test. (a) General. Each engine must be subjected to an endurance test that includes a total of 150 hours of operation (except as provided in paragraph (e)(1)(iii) of this section) and, depending upon the type and contemplated use of the engine, consists of one of the series of runs specified in paragraphs (b) through (e) of this section, as applicable. The runs must be made in the order found appropriate by the Administrator for the particular engine being tested. During the endurance test the engine power and the crankshaft rotational speed must be kept within ±3 percent of the rated values. During the runs at rated takeoff power and for at least 35 hours at rated maximum continuous power, one cylinder must be operated at not less than the limiting temperature, the other cylinders must be operated at a temperature not lower than 50 degrees F. below the limiting temperature, and the oil inlet temperature must be maintained within ±10 degrees F. of the limiting temperature. An engine that is equipped with a propeller shaft must be fitted for the endurance test with a propeller that thrust-loads the engine to the maximum thrust which the engine is designed to resist at each applicable operating condition specified in this section. Each accessory drive and mounting attachment must be loaded. During operation at rated takeoff power and rated maximum continuous power, the load imposed by each accessory used only for an aircraft service must be the limit load specified by the applicant for the engine drive or attachment point. (b) Unsupercharged engines and engines incorporating a gear-driven single-speed supercharger. For engines not incorporating a supercharger and for engines incorporating a gear-driven single-speed supercharger the applicant must conduct the following runs: (1) A 30-hour run consisting of alternate periods of 5 minutes at rated takeoff power with takeoff speed, and 5 minutes at maximum best economy cruising power or maximum recommended cruising power. (2) A 20-hour run consisting of alternate periods of 1 1/2 hours at rated maximum continuous power with maximum continuous speed, and 1/2 hour at 75 percent rated maximum continuous power and 91 percent maximum continuous speed. (3) A 20-hour run consisting of alternate periods of 1 1/2 hours at rated maximum continuous power with maximum continuous speed, and 1/2 hour at 70 percent rated maximum continuous power and 89 percent maximum continuous speed. (4) A 20-hour run consisting of alternate periods of 1 1/2 hours at rated maximum continuous power with maximum continuous speed, and 1/2 hour at 65 percent rated maximum continuous power and 87 percent maximum continuous speed. (5) A 20-hour run consisting of alternate periods of 1 1/2 hours at rated maximum continuous power with maximum continuous speed, and 1/2 hour at 60 percent rated maximum continuous power and 84.5 percent maximum continuous speed. (6) A 20-hour run consisting of alternate periods of 1 1/2 hours at rated maximum continuous power with maximum continuous speed, and 1/2 hour at 50 percent rated maximum continuous power and 79.5 percent maximum continuous speed. (7) A 20-hour run consisting of alternate periods of 2 1/2 hours at rated maximum continuous power with maximum continuous speed, and 2 1/2 hours at maximum best economy cruising power or at maximum recommended cruising power. (c) Engines incorporating a gear-driven two-speed supercharger. For engines incorporating a gear-driven two-speed supercharger the applicant must conduct the following runs: (1) A 30-hour run consisting of alternate periods in the lower gear ratio of 5 minutes at rated takeoff power with takeoff speed, and 5 minutes at maximum best economy cruising power or at maximum recommended cruising power. If a takeoff power rating is desired in the higher gear ratio, 15 hours of the 30-hour run must be made in the higher gear ratio in alternate periods of 5 minutes at the observed horsepower obtainable with the takeoff critical altitude manifold pressure and takeoff speed, and 5 minutes at 70 percent high ratio rated maximum continuous power and 89 percent high ratio maximum continuous speed. (2) A 15-hour run consisting of alternate periods in the lower gear ratio of 1 hour at rated maximum continuous power with maximum continuous speed, and 1/2 hour at 75 percent rated maximum continuous power and 91 percent maximum continuous speed. (3) A 15-hour run consisting of alternate periods in the lower gear ratio of 1 hour at rated maximum continuous power with maximum continuous speed, and 1/2 hour at 70 percent rated maximum continuous power and 89 percent maximum continuous speed. (4) A 30-hour run in the higher gear ratio at rated maximum continuous power with maximum continuous speed. (5) A 5-hour run consisting of alternate periods of 5 minutes in each of the supercharger gear ratios. The first 5 minutes of the test must be made at maximum continuous speed in the higher gear ratio and the observed horsepower obtainable with 90 percent of maximum continuous manifold pressure in the higher gear ratio under sea level conditions. The condition for operation for the alternate 5 minutes in the lower gear ratio must be that obtained by shifting to the lower gear ratio at constant speed. (6) A 10-hour run consisting of alternate periods in the lower gear ratio of 1 hour at rated maximum continuous power with maximum continuous speed, and 1 hour at 65 percent rated maximum continuous power and 87 percent maximum continuous speed. (7) A 10-hour run consisting of alternate periods in the lower gear ratio of 1 hour at rated maximum continuous power with maximum continuous speed, and 1 hour at 60 percent rated maximum continuous power and 84.5 percent maximum continuous speed. (8) A 10-hour run consisting of alternate periods in the lower gear ratio of 1 hour at rated maximum continuous power with maximum continuous speed, and 1 hour at 50 percent rated maximum continuous power and 79.5 percent maximum continuous speed. (9) A 20-hour run consisting of alternate periods in the lower gear ratio of 2 hours at rated maximum continuous power with maximum continuous speed, and 2 hours at maximum best economy cruising power and speed or at maximum recommended cruising power. (10) A 5-hour run in the lower gear ratio at maximum best economy cruising power and speed or at maximum recommended cruising power and speed. Where simulated altitude test equipment is not available when operating in the higher gear ratio, the runs may be made at the observed horsepower obtained with the critical altitude manifold pressure or specified percentages thereof, and the fuel-air mixtures may be adjusted to be rich enough to suppress detonation. (d) Helicopter engines. To be eligible for use on a helicopter each engine must either comply with paragraphs (a) through (j) of § 29.923 of this chapter, or must undergo the following series of runs: (1) A 35-hour run consisting of alternate periods of 30 minutes each at rated takeoff power with takeoff speed, and at rated maximum continuous power with maximum continuous speed. (2) A 25-hour run consisting of alternate periods of 2 1/2 hours each at rated maximum continuous power with maximum continuous speed, and at 70 percent rated maximum continuous power with maximum continuous speed. (3) A 25-hour run consisting of alternate periods of 2 1/2 hours each at rated maximum continuous power with maximum continuous speed, and at 70 percent rated maximum continuous power with 80 to 90 percent maximum continuous speed. (4) A 25-hour run consisting of alternate periods of 2 1/2 hours each at 30 percent rated maximum continuous power with takeoff speed, and at 30 percent rated maximum continuous power with 80 to 90 percent maximum continuous speed. (5) A 25-hour run consisting of alternate periods of 2 1/2 hours each at 80 percent rated maximum continuous power with takeoff speed, and at either rated maximum continuous power with 110 percent maximum continuous speed or at rated takeoff power with 103 percent takeoff speed, whichever results in the greater speed. (6) A 15-hour run at 105 percent rated maximum continuous power with 105 percent maximum continuous speed or at full throttle and corresponding speed at standard sea level carburetor entrance pressure, if 105 percent of the rated maximum continuous power is not exceeded. (e) Turbosupercharged engines. For engines incorporating a turbosupercharger the following apply except that altitude testing may be simulated provided the applicant shows that the engine and supercharger are being subjected to mechanical loads and operating temperatures no less severe than if run at actual altitude conditions: (1) For engines used in airplanes the applicant must conduct the runs specified in paragraph (b) of this section, except— (i) The entire run specified in paragraph (b)(1) of this section must be made at sea level altitude pressure; (ii) The portions of the runs specified in paragraphs (b)(2) through (7) of this section at rated maximum continuous power must be made at critical altitude pressure, and the portions of the runs at other power must be made at 8,000 feet altitude pressure; and (iii) The turbosupercharger used during the 150-hour endurance test must be run on the bench for an additional 50 hours at the limiting turbine wheel inlet gas temperature and rotational speed for rated maximum continuous power operation unless the limiting temperature and speed are maintained during 50 hours of the rated maximum continuous power operation. (2) For engines used in helicopters the applicant must conduct the runs specified in paragraph (d) of this section, except— (i) The entire run specified in paragraph (d)(1) of this section must be made at critical altitude pressure; (ii) The portions of the runs specified in paragraphs (d)(2) and (3) of this section at rated maximum continuous power must be made at critical altitude pressure and the portions of the runs at other power must be made at 8,000 feet altitude pressure; (iii) The entire run specified in paragraph (d)(4) of this section must be made at 8,000 feet altitude pressure; (iv) The portion of the runs specified in paragraph (d)(5) of this section at 80 percent of rated maximum continuous power must be made at 8,000 feet altitude pressure and the portions of the runs at other power must be made at critical altitude pressure; (v) The entire run specified in paragraph (d)(6) of this section must be made at critical altitude pressure; and (vi) The turbosupercharger used during the endurance test must be run on the bench for 50 hours at the limiting turbine wheel inlet gas temperature and rotational speed for rated maximum continuous power operation unless the limiting temperature and speed are maintained during 50 hours of the rated maximum continuous power operation. § 33.51 Operation test. The operation test must include the testing found necessary by the Administrator to demonstrate backfire characteristics, starting, idling, acceleration, overspeeding, functioning of propeller and ignition, and any other operational characteristic of the engine. If the engine incorporates a multispeed supercharger drive, the design and construction must allow the supercharger to be shifted from operation at the lower speed ratio to the higher and the power appropriate to the manifold pressure and speed settings for rated maximum continuous power at the higher supercharger speed ratio must be obtainable within five seconds. § 33.53 Engine system and component tests. (a) For those systems and components that cannot be adequately substantiated in accordance with endurance testing of § 33.49, the applicant must conduct additional tests to demonstrate that systems or components are able to perform the intended functions in all declared environmental and operating conditions. (b) Temperature limits must be established for each component that requires temperature controlling provisions in the aircraft installation to assure satisfactory functioning, reliability, and durability. § 33.55 Teardown inspection. After completing the endurance test— (a) Each engine must be completely disassembled; (b) Each component having an adjustment setting and a functioning characteristic that can be established independent of installation on the engine must retain each setting and functioning characteristic within the limits that were established and recorded at the beginning of the test; and (c) Each engine component must conform to the type design and be eligible for incorporation into an engine for continued operation, in accordance with information submitted in compliance with § 33.4. § 33.57 General conduct of block tests. (a) The applicant may, in conducting the block tests, use separate engines of identical design and construction in the vibration, calibration, detonation, endurance, and operation tests, except that, if a separate engine is used for the endurance test it must be subjected to a calibration check before starting the endurance test. (b) The applicant may service and make minor repairs to the engine during the block tests in accordance with the service and maintenance instructions submitted in compliance with § 33.4. If the frequency of the service is excessive, or the number of stops due to engine malfunction is excessive, or a major repair, or replacement of a part is found necessary during the block tests or as the result of findings from the teardown inspection, the engine or its parts may be subjected to any additional test the Administrator finds necessary. (c) Each applicant must furnish all testing facilities, including equipment and competent personnel, to conduct the block tests. § 33.61 Applicability. This subpart prescribes additional design and construction requirements for turbine aircraft engines. § 33.62 Stress analysis. A stress analysis must be performed on each turbine engine showing the design safety margin of each turbine engine rotor, spacer, and rotor shaft. § 33.63 Vibration. Each engine must be designed and constructed to function throughout its declared flight envelope and operating range of rotational speeds and power/thrust, without inducing excessive stress in any engine part because of vibration and without imparting excessive vibration forces to the aircraft structure. § 33.64 Pressurized engine static parts. (a) Strength. The applicant must establish by test, validated analysis, or a combination of both, that all static parts subject to significant gas or liquid pressure loads for a stabilized period of one minute will not: (1) Exhibit permanent distortion beyond serviceable limits or exhibit leakage that could create a hazardous condition when subjected to the greater of the following pressures: (i) 1.1 times the maximum working pressure; (ii) 1.33 times the normal working pressure; or (iii) 35 kPa (5 p.s.i.) above the normal working pressure. (2) Exhibit fracture or burst when subjected to the greater of the following pressures: (i) 1.15 times the maximum possible pressure; (ii) 1.5 times the maximum working pressure; or (iii) 35 kPa (5 p.s.i.) above the maximum possible pressure. (b) Compliance with this section must take into account: (1) The operating temperature of the part; (2) Any other significant static loads in addition to pressure loads; (3) Minimum properties representative of both the material and the processes used in the construction of the part; and (4) Any adverse geometry conditions allowed by the type design. § 33.65 Surge and stall characteristics. When the engine is operated in accordance with operating instructions required by § 33.5(b), starting, a change of power or thrust, power or thrust augmentation, limiting inlet air distortion, or inlet air temperature may not cause surge or stall to the extent that flameout, structural failure, overtemperature, or failure of the engine to recover power or thrust will occur at any point in the operating envelope. § 33.66 Bleed air system. The engine must supply bleed air without adverse effect on the engine, excluding reduced thrust or power output, at all conditions up to the discharge flow conditions established as a limitation under § 33.7(c)(11). If bleed air used for engine anti-icing can be controlled, provision must be made for a means to indicate the functioning of the engine ice protection system. § 33.67 Fuel system. (a) With fuel supplied to the engine at the flow and pressure specified by the applicant, the engine must function properly under each operating condition required by this part. Each fuel control adjusting means that may not be manipulated while the fuel control device is mounted on the engine must be secured by a locking device and sealed, or otherwise be inaccessible. All other fuel control adjusting means must be accessible and marked to indicate the function of the adjustment unless the function is obvious. (b) There must be a fuel strainer or filter between the engine fuel inlet opening and the inlet of either the fuel metering device or the engine-driven positive displacement pump whichever is nearer the engine fuel inlet. In addition, the following provisions apply to each strainer or filter required by this paragraph (b): (1) It must be accessible for draining and cleaning and must incorporate a screen or element that is easily removable. (2) It must 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. (3) It must 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, unless adequate strength margins under all loading conditions are provided in the lines and connections. (4) It must have the type and degree of fuel filtering specified as necessary for protection of the engine fuel system against foreign particles in the fuel. The applicant must show: (i) That foreign particles passing through the specified filtering means do not impair the engine fuel system functioning; and (ii) That the fuel system is capable of sustained operation throughout its flow and pressure range with the fuel initially saturated with water at 80 °F (27 °C) and having 0.025 fluid ounces per gallon (0.20 milliliters per liter) of free water added and cooled to the most critical condition for icing likely to be encountered in operation. However, this requirement may be met by demonstrating the effectiveness of specified approved fuel anti-icing additives, or that the fuel system incorporates a fuel heater which maintains the fuel temperature at the fuel strainer or fuel inlet above 32 °F (0 °C) under the most critical conditions. (5) The applicant must demonstrate that the filtering means has the capacity (with respect to engine operating limitations) to ensure that the engine will continue to operate within approved limits, with fuel contaminated to the maximum degree of particle size and density likely to be encountered in service. Operation under these conditions must be demonstrated for a period acceptable to the Administrator, beginning when indication of impending filter blockage is first given by either: (i) Existing engine instrumentation; or (ii) Additional means incorporated into the engine fuel system. (6) Any strainer or filter bypass must be designed and constructed 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. (c) If provided as part of the engine, the applicant must show for each fluid injection (other than fuel) system and its controls that the flow of the injected fluid is adequately controlled. § 33.68 Induction system icing. Each engine, with all icing protection systems operating, must: (a) Operate throughout its flight power range, including the minimum descent idle rotor speeds achievable in flight, in the icing conditions defined for turbojet, turbofan, and turboprop engines in Appendices C and O of part 25 of this chapter, and Appendix D of this part, and for turboshaft engines in Appendix C of part 29 of this chapter, without the accumulation of ice on the engine components that: (1) Adversely affects engine operation or that causes an unacceptable permanent loss of power or thrust or unacceptable increase in engine operating temperature; or (2) Results in unacceptable temporary power loss or engine damage; or (3) Causes a stall, surge, or flameout or loss of engine controllability. The applicant must account for in-flight ram effects in any critical point analysis or test demonstration of these flight conditions. (b) Operate throughout its flight power range, including minimum descent idle rotor speeds achievable in flight, in the icing conditions defined for turbojet, turbofan, and turboprop engines in Appendices C and O of part 25 of this chapter, and for turboshaft engines in Appendix C of part 29 of this chapter. In addition: (1) It must be shown through Critical Point Analysis (CPA) that the complete ice envelope has been analyzed, and that the most critical points must be demonstrated by engine test, analysis, or a combination of the two to operate acceptably. Extended flight in critical flight conditions such as hold, descent, approach, climb, and cruise, must be addressed, for the ice conditions defined in these appendices. (2) It must be shown by engine test, analysis, or a combination of the two that the engine can operate acceptably for the following durations: (i) At engine powers that can sustain level flight: A duration that achieves repetitive, stabilized operation for turbojet, turbofan, and turboprop engines in the icing conditions defined in Appendices C and O of part 25 of this chapter, and for turboshaft engines in the icing conditions defined in Appendix C of part 29 of this chapter. (ii) At engine power below that which can sustain level flight: (A) Demonstration in altitude flight simulation test facility: A duration of 10 minutes consistent with a simulated flight descent of 10,000 ft (3 km) in altitude while operating in Continuous Maximum icing conditions defined in Appendix C of part 25 of this chapter for turbojet, turbofan, and turboprop engines, and for turboshaft engines in the icing conditions defined in Appendix C of part 29 of this chapter, plus 40 percent liquid water content margin, at the critical level of airspeed and air temperature; or (B) Demonstration in ground test facility: A duration of 3 cycles of alternating icing exposure corresponding to the liquid water content levels and standard cloud lengths starting in Intermittent Maximum and then in Continuous Maximum icing conditions defined in Appendix C of part 25 of this chapter for turbojet, turbofan, and turboprop engines, and for turboshaft engines in the icing conditions defined in Appendix C of part 29 of this chapter, at the critical level of air temperature. (c) In addition to complying with paragraph (b) of this section, the following conditions shown in Table 1 of this section unless replaced by similar CPA test conditions that are more critical or produce an equivalent level of severity, must be demonstrated by an engine test: (d) Operate at ground idle speed for a minimum of 30 minutes at each of the following icing conditions shown in Table 2 of this section with the available air bleed for icing protection at its critical condition, without adverse effect, followed by acceleration to takeoff power or thrust. During the idle operation, the engine may be run up periodically to a moderate power or thrust setting in a manner acceptable to the Administrator. Analysis may be used to show ambient temperatures below the tested temperature are less critical. The applicant must document any demonstrated run ups and minimum ambient temperature capability in the engine operating manual as mandatory in icing conditions. The applicant must demonstrate, with consideration of expected airport elevations, the following: (e) Demonstrate by test, analysis, or combination of the two, acceptable operation for turbojet, turbofan, and turboprop engines in mixed phase and ice crystal icing conditions throughout Appendix D of this part, icing envelope throughout its flight power range, including minimum descent idling speeds. § 33.69 Ignitions system. Each engine must be equipped with an ignition system for starting the engine on the ground and in flight. An electric ignition system must have at least two igniters and two separate secondary electric circuits, except that only one igniter is required for fuel burning augmentation systems. § 33.70 Engine life-limited parts. By a procedure approved by the FAA, operating limitations must be established which specify the maximum allowable number of flight cycles for each engine life-limited part. Engine life-limited parts are rotor and major static structural parts whose primary failure is likely to result in a hazardous engine effect. Typically, engine life-limited parts include, but are not limited to disks, spacers, hubs, shafts, high-pressure casings, and non-redundant mount components. For the purposes of this section, a hazardous engine effect is any of the conditions listed in § 33.75 of this part. The applicant will establish the integrity of each engine life-limited part by: (a) An engineering plan that contains the steps required to ensure each engine life-limited part is withdrawn from service at an approved life before hazardous engine effects can occur. These steps include validated analysis, test, or service experience which ensures that the combination of loads, material properties, environmental influences and operating conditions, including the effects of other engine parts influencing these parameters, are sufficiently well known and predictable so that the operating limitations can be established and maintained for each engine life-limited part. Applicants must perform appropriate damage tolerance assessments to address the potential for failure from material, manufacturing, and service induced anomalies within the approved life of the part. Applicants must publish a list of the life-limited engine parts and the approved life for each part in the Airworthiness Limitations Section of the Instructions for Continued Airworthiness as required by § 33.4 of this part. (b) A manufacturing plan that identifies the specific manufacturing constraints necessary to consistently produce each engine life-limited part with the attributes required by the engineering plan. (c) A service management plan that defines in-service processes for maintenance and the limitations to repair for each engine life-limited part that will maintain attributes consistent with those required by the engineering plan. These processes and limitations will become part of the Instructions for Continued Airworthiness. § 33.71 Lubrication system. (a) General. Each lubrication system must function properly in the flight attitudes and atmospheric conditions in which an aircraft is expected to operate. (b) Oil strainer or filter. There must be an oil strainer or filter through which all of the engine oil flows. In addition: (1) Each strainer or filter required by this paragraph 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 element completely blocked. (2) The type and degree of filtering necessary for protection of the engine oil system against foreign particles in the oil must be specified. The applicant must demonstrate that foreign particles passing through the specified filtering means do not impair engine oil system functioning. (3) Each strainer or filter required by this paragraph must have the capacity (with respect to operating limitations established for the engine) to ensure that engine oil system functioning is not impaired with the oil contaminated to a degree (with respect to particle size and density) that is greater than that established for the engine in paragraph (b)(2) of this section. (4) For each strainer or filter required by this paragraph, except the strainer or filter at the oil tank outlet, there must be means to indicate contamination before it reaches the capacity established in accordance with paragraph (b)(3) of this section. (5) Any filter bypass must be designed and constructed so that the release of collected contaminants is minimized by appropriate location of the bypass to ensure that the collected contaminants are not in the bypass flow path. (6) Each strainer or filter required by this paragraph that has no bypass, except the strainer or filter at an oil tank outlet or for a scavenge pump, must have provisions for connection with a warning means to warn the pilot of the occurance of contamination of the screen before it reaches the capacity established in accordance with paragraph (b)(3) of this section. (7) Each strainer or filter required by this paragraph must be accessible for draining and cleaning. (c) Oil tanks. (1) Each oil tank must have an expansion space of not less than 10 percent of the tank capacity. (2) It must be impossible to inadvertently fill the oil tank expansion space. (3) Each recessed oil tank filler connection that can retain any appreciable quantity of oil must have provision for fitting a drain. (4) Each oil tank cap must provide an oil-tight seal. For an applicant seeking eligibility for an engine to be installed on an airplane approved for ETOPS, the oil tank must be designed to prevent a hazardous loss of oil due to an incorrectly installed oil tank cap. (5) Each oil tank filler must be marked with the word “oil.” (6) Each oil tank must be vented from the top part of the expansion space, with the vent so arranged that condensed water vapor that might freeze and obstruct the line cannot accumulate at any point. (7) There must be means to prevent entrance into the oil tank or into any oil tank outlet, of any object that might obstruct the flow of oil through the system. (8) There must be a shutoff valve at the outlet of each oil tank, unless the external portion of the oil system (including oil tank supports) is fireproof. (9) Each unpressurized oil tank may not leak when subjected to a maximum operating temperature and an internal pressure of 5 p.s.i., and each pressurized oil tank must meet the requirements of § 33.64. (10) Leaked or spilled oil may not accumulate between the tank and the remainder of the engine. (11) Each oil tank must have an oil quantity indicator or provisions for one. (12) If the propeller feathering system depends on engine oil— (i) There must be means to trap an amount of oil in the tank if the supply becomes depleted due to failure of any part of the lubricating system other than the tank itself; (ii) The amount of trapped oil must be enough to accomplish the feathering opeation and must be available only to the feathering pump; and (iii) Provision must be made to prevent sludge or other foreign matter from affecting the safe operation of the propeller feathering system. (d) Oil drains. A drain (or drains) must be provided to allow safe drainage of the oil system. Each drain must— (1) Be accessible; and (2) Have manual or automatic means for positive locking in the closed position. (e) Oil radiators. Each oil radiator must withstand, without failure, any vibration, inertia, and oil pressure load to which it is subjected during the block tests. § 33.72 Hydraulic actuating systems. Each hydraulic actuating system must function properly under all conditions in which the engine is expected to operate. Each filter or screen must be accessible for servicing and each tank must meet the design criteria of § 33.71. § 33.73 Power or thrust response. The design and construction of the engine must enable an increase— (a) From minimum to rated takeoff power or thrust with the maximum bleed air and power extraction to be permitted in an aircraft, without overtemperature, surge, stall, or other detrimental factors occurring to the engine whenever the power control lever is moved from the minimum to the maximum position in not more than 1 second, except that the Administrator may allow additional time increments for different regimes of control operation requiring control scheduling; and (b) From the fixed minimum flight idle power lever position when provided, or if not provided, from not more than 15 percent of the rated takeoff power or thrust available to 95 percent rated takeoff power or thrust in not over 5 seconds. The 5-second power or thrust response must occur from a stabilized static condition using only the bleed air and accessories loads necessary to run the engine. This takeoff rating is specified by the applicant and need not include thrust augmentation. § 33.74 Continued rotation. If any of the engine main rotating systems continue to rotate after the engine is shutdown for any reason while in flight, and if means to prevent that continued rotation are not provided, then any continued rotation during the maximum period of flight, and in the flight conditions expected to occur with that engine inoperative, may not result in any condition described in § 33.75(g)(2)(i) through (vi) of this part. § 33.75 Safety analysis. (a) (1) The applicant must analyze the engine, including the control system, to assess the likely consequences of all failures that can reasonably be expected to occur. This analysis will take into account, if applicable: (i) Aircraft-level devices and procedures assumed to be associated with a typical installation. Such assumptions must be stated in the analysis. (ii) Consequential secondary failures and latent failures. (iii) Multiple failures referred to in paragraph (d) of this section or that result in the hazardous engine effects defined in paragraph (g)(2) of this section. (2) The applicant must summarize those failures that could result in major engine effects or hazardous engine effects, as defined in paragraph (g) of this section, and estimate the probability of occurrence of those effects. Any engine part the failure of which could reasonably result in a hazardous engine effect must be clearly identified in this summary. (3) The applicant must show that hazardous engine effects are predicted to occur at a rate not in excess of that defined as extremely remote (probability range of 10 −7 to 10 −9 per engine flight hour). Since the estimated probability for individual failures may be insufficiently precise to enable the applicant to assess the total rate for hazardous engine effects, compliance may be shown by demonstrating that the probability of a hazardous engine effect arising from an individual failure can be predicted to be not greater than 10 −8 per engine flight hour. In dealing with probabilities of this low order of magnitude, absolute proof is not possible, and compliance may be shown by reliance on engineering judgment and previous experience combined with sound design and test philosophies. (4) The applicant must show that major engine effects are predicted to occur at a rate not in excess of that defined as remote (probability range of 10 −5 to 10 −7 per engine flight hour). (b) The FAA may require that any assumption as to the effects of failures and likely combination of failures be verified by test. (c) The primary failure of certain single elements cannot be sensibly estimated in numerical terms. If the failure of such elements is likely to result in hazardous engine effects, then compliance may be shown by reliance on the prescribed integrity requirements of §§ 33.15, 33.27, and 33.70 as applicable. These instances must be stated in the safety analysis. (d) If reliance is placed on a safety system to prevent a failure from progressing to hazardous engine effects, the possibility of a safety system failure in combination with a basic engine failure must be included in the analysis. Such a safety system may include safety devices, instrumentation, early warning devices, maintenance checks, and other similar equipment or procedures. If items of a safety system are outside the control of the engine manufacturer, the assumptions of the safety analysis with respect to the reliability of these parts must be clearly stated in the analysis and identified in the installation instructions under § 33.5 of this part. (e) If the safety analysis depends on one or more of the following items, those items must be identified in the analysis and appropriately substantiated. (1) Maintenance actions being carried out at stated intervals. This includes the verification of the serviceability of items that could fail in a latent manner. When necessary to prevent hazardous engine effects, these maintenance actions and intervals must be published in the instructions for continued airworthiness required under § 33.4 of this part. Additionally, if errors in maintenance of the engine, including the control system, could lead to hazardous engine effects, the appropriate procedures must be included in the relevant engine manuals. (2) Verification of the satisfactory functioning of safety or other devices at pre-flight or other stated periods. The details of this satisfactory functioning must be published in the appropriate manual. (3) The provisions of specific instrumentation not otherwise required. (4) Flight crew actions to be specified in the operating instructions established under § 33.5. (f) If applicable, the safety analysis must also include, but not be limited to, investigation of the following: (1) Indicating equipment; (2) Manual and automatic controls; (3) Compressor bleed systems; (4) Refrigerant injection systems; (5) Gas temperature control systems; (6) Engine speed, power, or thrust governors and fuel control systems; (7) Engine overspeed, overtemperature, or topping limiters; (8) Propeller control systems; and (9) Engine or propeller thrust reversal systems. (g) Unless otherwise approved by the FAA and stated in the safety analysis, for compliance with part 33, the following failure definitions apply to the engine: (1) An engine failure in which the only consequence is partial or complete loss of thrust or power (and associated engine services) from the engine will be regarded as a minor engine effect. (2) The following effects will be regarded as hazardous engine effects: (i) Non-containment of high-energy debris; (ii) Concentration of toxic products in the engine bleed air intended for the cabin sufficient to incapacitate crew or passengers; (iii) Significant thrust in the opposite direction to that commanded by the pilot; (iv) Uncontrolled fire; (v) Failure of the engine mount system leading to inadvertent engine separation; (vi) Release of the propeller by the engine, if applicable; and (vii) Complete inability to shut the engine down. (3) An effect whose severity falls between those effects covered in paragraphs (g)(1) and (g)(2) of this section will be regarded as a major engine effect. § 33.76 Bird ingestion. (a) General. Compliance with paragraphs (b) through (e) of this section shall be in accordance with the following: (1) Except as specified in paragraphs (d) and (e) of this section, all ingestion tests must be conducted with the engine stabilized at no less than 100 percent takeoff power or thrust, for test day ambient conditions prior to the ingestion. In addition, the demonstration of compliance must account for engine operation at sea level takeoff conditions on the hottest day that a minimum engine can achieve maximum rated takeoff thrust or power. (2) The engine inlet throat area as used in this section to determine the bird quantity and weights will be established by the applicant and identified as a limitation in the installation instructions required under § 33.5. (3) The impact to the front of the engine from the large single bird, the single largest medium bird which can enter the inlet, and the large flocking bird must be evaluated. Applicants must show that the associated components when struck under the conditions prescribed in paragraphs (b), (c) or (d) of this section, as applicable, will not affect the engine to the extent that the engine cannot comply with the requirements of paragraphs (b)(3), (c)(6) and (d)(4) of this section. (4) For an engine that incorporates an inlet protection device, compliance with this section shall be established with the device functioning. The engine approval will be endorsed to show that compliance with the requirements has been established with the device functioning. (5) Objects that are accepted by the Administrator may be substituted for birds when conducting the bird ingestion tests required by paragraphs (b) through (e) of this section. (6) If compliance with the requirements of this section is not established, the engine type certification documentation will show that the engine shall be limited to aircraft installations in which it is shown that a bird cannot strike the engine, or be ingested into the engine, or adversely restrict airflow into the engine. (b) Large single bird. Compliance with the large bird ingestion requirements shall be in accordance with the following: (1) The large bird ingestion test shall be conducted using one bird of a weight determined from Table 1 aimed at the most critical exposed location on the first stage rotor blades and ingested at a bird speed of 200-knots for engines to be installed on airplanes, or the maximum airspeed for normal rotorcraft flight operations for engines to be installed on rotorcraft. (2) Power lever movement is not permitted within 15 seconds following ingestion of the large bird. (3) Ingestion of a single large bird tested under the conditions prescribed in this section may not result in any condition described in § 33.75(g)(2) of this part. (4) Compliance with the large bird ingestion requirements of this paragraph may be shown by demonstrating that the requirements of § 33.94(a) constitute a more severe demonstration of blade containment and rotor unbalance than the requirements of this paragraph. (c) Small and medium flocking bird. Compliance with the small and medium bird ingestion requirements shall be in accordance with the following: (1) Analysis or component test, or both, acceptable to the Administrator, shall be conducted to determine the critical ingestion parameters affecting power loss and damage. Critical ingestion parameters shall include, but are not limited to, the effects of bird speed, critical target location, and first stage rotor speed. The critical bird ingestion speed should reflect the most critical condition within the range of airspeeds used for normal flight operations up to 1,500 feet above ground level, but not less than V 1 minimum for airplanes. (2) Medium bird engine tests shall be conducted so as to simulate a flock encounter, and will use the bird weights and quantities specified in Table 2. When only one bird is specified, that bird will be aimed at the engine core primary flow path; the other critical locations on the engine face area must be addressed, as necessary, by appropriate tests or analysis, or both. When two or more birds are specified in Table 2, the largest of those birds must be aimed at the engine core primary flow path, and a second bird must be aimed at the most critical exposed location on the first stage rotor blades. Any remaining birds must be evenly distributed over the engine face area. (3) In addition, except for rotorcraft engines, it must also be substantiated by appropriate tests or analysis or both, that when the full fan assembly is subjected to the ingestion of the quantity and weights of bird from Table 3, aimed at the fan assembly's most critical location outboard of the primary core flowpath, and in accordance with the applicable test conditions of this paragraph, that the engine can comply with the acceptance criteria of this paragraph. (4) A small bird ingestion test is not required if the prescribed number of medium birds pass into the engine rotor blades during the medium bird test. (5) Small bird ingestion tests shall be conducted so as to simulate a flock encounter using one 85 gram (0.187 lb.) bird for each 0.032 square-meter (49.6 square-inches) of inlet area, or fraction thereof, up to a maximum of 16 birds. The birds will be aimed so as to account for any critical exposed locations on the first stage rotor blades, with any remaining birds evenly distributed over the engine face area. (6) Ingestion of small and medium birds tested under the conditions prescribed in this paragraph may not cause any of the following: (i) More than a sustained 25-percent power or thrust loss; (ii) The engine to be shut down during the required run-on demonstration prescribed in paragraphs (c)(7) or (c)(8) of this section; (iii) The conditions defined in paragraph (b)(3) of this section. (iv) Unacceptable deterioration of engine handling characteristics. (7) Except for rotorcraft engines, the following test schedule shall be used: (i) Ingestion so as to simulate a flock encounter, with approximately 1 second elapsed time from the moment of the first bird ingestion to the last. (ii) Followed by 2 minutes without power lever movement after the ingestion. (iii) Followed by 3 minutes at 75-percent of the test condition. (iv) Followed by 6 minutes at 60-percent of the test condition. (v) Followed by 6 minutes at 40-percent of the test condition. (vi) Followed by 1 minute at approach idle. (vii) Followed by 2 minutes at 75-percent of the test condition. (viii) Followed by stabilizing at idle and engine shut down. (ix) The durations specified are times at the defined conditions with the power being changed between each condition in less than 10 seconds. (8) For rotorcraft engines, the following test schedule shall be used: (i) Ingestion so as to simulate a flock encounter within approximately 1 second elapsed time between the first ingestion and the last. (ii) Followed by 3 minutes at 75-percent of the test condition. (iii) Followed by 90 seconds at descent flight idle. (iv) Followed by 30 seconds at 75-percent of the test condition. (v) Followed by stabilizing at idle and engine shut down. (vi) The durations specified are times at the defined conditions with the power being changed between each condition in less than 10 seconds. (9) Engines intended for use in multi-engine rotorcraft are not required to comply with the medium bird ingestion portion of this section, providing that the appropriate type certificate documentation is so endorsed. (10) If any engine operating limit(s) is exceeded during the initial 2 minutes without power lever movement, as provided by paragraph (c)(7)(ii) of this section, then it shall be established that the limit exceedence will not result in an unsafe condition. (d) Large flocking bird. An engine test will be performed as follows: (1) Large flocking bird engine tests will be performed using the bird mass and weights in Table 4, and ingested at a bird speed of 200 knots. (2) Prior to the ingestion, the engine must be stabilized at no less than the mechanical rotor speed of the first exposed stage or stages that, on a standard day, would produce 90 percent of the sea level static maximum rated takeoff power or thrust. (3) The bird must be targeted on the first exposed rotating stage or stages at a blade airfoil height of not less than 50 percent measured at the leading edge. (4) Ingestion of a large flocking bird under the conditions prescribed in this paragraph must not cause any of the following: (i) A sustained reduction of power or thrust to less than 50 percent of maximum rated takeoff power or thrust during the run-on segment specified under paragraph (d)(5)(i) of this section. (ii) Engine shutdown during the required run-on demonstration specified in paragraph (d)(5) of this section. (iii) The conditions specified in paragraph (b)(3) of this section. (5) The following test schedule must be used: (i) Ingestion followed by 1 minute without power lever movement. (ii) Followed by 13 minutes at not less than 50 percent of maximum rated takeoff power or thrust. (iii) Followed by 2 minutes between 30 and 35 percent of maximum rated takeoff power or thrust. (iv) Followed by 1 minute with power or thrust increased from that set in paragraph (d)(5)(iii) of this section, by between 5 and 10 percent of maximum rated takeoff power or thrust. (v) Followed by 2 minutes with power or thrust reduced from that set in paragraph (d)(5)(iv) of this section, by between 5 and 10 percent of maximum rated takeoff power or thrust. (vi) Followed by a minimum of 1 minute at ground idle then engine shutdown. The durations specified are times at the defined conditions. Power lever movement between each condition will be 10 seconds or less, except that power lever movements allowed within paragraph (d)(5)(ii) of this section are not limited, and for setting power under paragraph (d)(5)(iii) of this section will be 30 seconds or less. (6) Compliance with the large flocking bird ingestion requirements of this paragraph (d) may also be demonstrated by: (i) Incorporating the requirements of paragraph (d)(4) and (d)(5) of this section, into the large single bird test demonstration specified in paragraph (b)(1) of this section; or (ii) Use of an engine subassembly test at the ingestion conditions specified in paragraph (b)(1) of this section if: (A) All components critical to complying with the requirements of paragraph (d) of this section are included in the subassembly test; (B) The components of paragraph (d)(6)(ii)(A) of this section are installed in a representative engine for a run-on demonstration in accordance with paragraphs (d)(4) and (d)(5) of this section; except that section (d)(5)(i) is deleted and section (d)(5)(ii) must be 14 minutes in duration after the engine is started and stabilized; and (C) The dynamic effects that would have been experienced during a full engine ingestion test can be shown to be negligible with respect to meeting the requirements of paragraphs (d)(4) and (d)(5) of this section. (7) Applicants must show that an unsafe condition will not result if any engine operating limit is exceeded during the run-on period. (e) Core flocking bird test. Except as provided in paragraph (e)(4) of this section, for turbofan engines, an engine test must be performed in accordance with either paragraph (e)(1) or (2) of this section. The test specified in paragraph (e)(2) must be conducted if testing or validated analysis shows that no bird material will be ingested into the engine core during the test under the conditions specified in paragraph (e)(1). (1) Climb flocking bird test. (i) Test requirements are as follows: (A) Before ingestion, the engine must be stabilized at the mechanical rotor speed of the first exposed stage or stages that produce the lowest expected power or thrust required during climb through 3,000 feet above mean sea level (MSL) at standard day conditions. (B) The climb flocking bird test shall be conducted using one bird of the highest weight specified in table 2 to this section for the engine inlet area. (C) Ingestion must be at 261-knots true airspeed. (D) The bird must be aimed at the first exposed rotating stage or stages, at the blade airfoil height, as measured at the leading edge that will result in maximum bird material ingestion into the engine core. (ii) Ingestion of a flocking bird into the engine core under the conditions prescribed in paragraph (e)(1)(i) of this section must not cause any of the following: (A) Sustained power or thrust reduction to less than 50 percent maximum rated takeoff power or thrust during the run-on segment specified under paragraph (e)(1)(iii)(B) of this section, that cannot be restored only by movement of the power lever. (B) Sustained power or thrust reduction to less than flight idle power or thrust during the run-on segment specified under paragraph (e)(1)(iii)(B) of this section. (C) Engine shutdown during the required run-on demonstration specified in paragraph (e)(1)(iii) of this section. (D) Any condition specified in § 33.75(g)(2). (iii) The following test schedule must be used (power lever movement between conditions must occur within 10 seconds or less, unless otherwise noted): Note 1 to paragraph (e)(1)(iii) introductory text. Durations specified are times at the defined conditions in paragraphs (e)(1)(iii)(A) through (I) of this section. (A) Ingestion. (B) Followed by 1 minute without power lever movement. (C) Followed by power lever movement to increase power or thrust to not less than 50 percent maximum rated takeoff power or thrust, if the initial bird ingestion resulted in a reduction in power or thrust below that level. (D) Followed by 13 minutes at not less than 50 percent maximum rated takeoff power or thrust. Power lever movement in this condition is unlimited. (E) Followed by 2 minutes at 30-35 percent maximum rated takeoff power or thrust. (F) Followed by 1 minute with power or thrust increased from that set in paragraph (e)(1)(iii)(E) of this section, by 5-10 percent maximum rated takeoff power or thrust. (G) Followed by 2 minutes with power or thrust reduced from that set in paragraph (e)(1)(iii)(F) of this section, by 5-10 percent maximum rated takeoff power or thrust. (H) Followed by 1 minute minimum at ground idle. (I) Followed by engine shutdown. (2) Approach flocking bird test. (i) Test requirements are as follows: (A) Before ingestion, the engine must be stabilized at the mechanical rotor speed of the first exposed stage or stages that produce approach idle thrust when descending through 3,000 feet MSL at standard day conditions. (B) The approach flocking bird test shall be conducted using one bird of the highest weight specified in table 2 to this section for the engine inlet area. (C) Ingestion must be at 209-knots true airspeed. (D) The bird must be aimed at the first exposed rotating stage or stages, at the blade airfoil height measured at the leading edge that will result in maximum bird material ingestion into the engine core. (ii) Ingestion of a flocking bird into the engine core under the conditions prescribed in paragraph (e)(2)(i) of this section may not cause any of the following: (A) Power or thrust reduction to less than flight idle power or thrust during the run-on segment specified under paragraph (e)(2)(iii)(B) of this section. (B) Engine shutdown during the required run-on demonstration specified in paragraph (e)(2)(iii) of this section. (C) Any condition specified in § 33.75(g)(2). (iii) The following test schedule must be used (power lever movement between conditions must occur within 10 seconds or less, unless otherwise noted): Note 2 to paragraph (e)(2)(iii) introductory text. Durations specified are times at the defined conditions in paragraphs (e)(2)(iii)(A) through (H) of this section. (A) Ingestion. (B) Followed by 1 minute without power lever movement. (C) Followed by 2 minutes at 30-35 percent maximum rated takeoff power or thrust. Power lever movement in this condition is unlimited. (D) Followed by 1 minute with power or thrust increased from that set in paragraph (e)(2)(iii)(C) of this section, by 5-10 percent maximum rated takeoff power or thrust. (E) Followed by 2 minutes with power or thrust reduced from that set in paragraph (e)(2)(iii)(D) of this section, by 5-10 percent maximum rated takeoff power or thrust. (F) Followed by 1 minute minimum at ground idle. (G) Followed by engine shutdown. (H) Power lever movement between each condition must be 10 seconds or less, except that any power lever movements are allowed within the time period of paragraph (e)(2)(iii)(C) of this section. (3) Results of exceeding engine-operating limits. Applicants must show that an unsafe condition will not result if any engine-operating limit is exceeded during the run-on period. (4) Combining tests. The climb flocking bird test of paragraph (e)(1) of this section may be combined with the medium flocking bird test of paragraph (c) of this section, if the climb first stage rotor speed calculated in paragraph (e)(1) of this section is within 3 percent of the first stage rotor speed required by paragraph (c)(1) of this section. As used in this paragraph (e)(4), “combined” means that, instead of separately conducting the tests specified in paragraphs (c) and (e)(1) of this section, the test conducted under paragraph (c) of this section satisfies the requirements of paragraph (e) of this section if the bird aimed at the core of the engine meets the bird ingestion speed criteria of paragraph (e)(1)(i)(C) of this section. § 33.77 Foreign object ingestion—ice. (a) Compliance with the requirements of this section must be demonstrated by engine ice ingestion test or by validated analysis showing equivalence of other means for demonstrating soft body damage tolerance. (b) [Reserved] (c) Ingestion of ice under the conditions of this section may not— (1) Cause an immediate or ultimate unacceptable sustained power or thrust loss; or (2) Require the engine to be shutdown. (d) For an engine that incorporates a protection device, compliance with this section need not be demonstrated with respect to ice formed forward of the protection device if it is shown that— (1) Such ice is of a size that will not pass through the protective device; (2) The protective device will withstand the impact of the ice; and (3) The ice stopped by the protective device will not obstruct the flow of induction air into the engine with a resultant sustained reduction in power or thrust greater than those values defined by paragraph (c) of this section. (e) Compliance with the requirements of this section must be demonstrated by engine ice ingestion test under the following ingestion conditions or by validated analysis showing equivalence of other means for demonstrating soft body damage tolerance. (1) The minimum ice quantity and dimensions will be established by the engine size as defined in Table 1 of this section. (2) The ingested ice dimensions are determined by linear interpolation between table values, and are based on the actual engine's inlet hilite area. (3) The ingestion velocity will simulate ice from the inlet being sucked into the engine. (4) Engine operation will be at the maximum cruise power or thrust unless lower power is more critical. § 33.78 Rain and hail ingestion. (a) All engines. (1) The ingestion of large hailstones (0.8 to 0.9 specific gravity) at the maximum true air speed, up to 15,000 feet (4,500 meters), associated with a representative aircraft operating in rough air, with the engine at maximum continuous power, may not cause unacceptable mechanical damage or unacceptable power or thrust loss after the ingestion, or require the engine to be shut down. One-half the number of hailstones shall be aimed randomly over the inlet face area and the other half aimed at the critical inlet face area. The hailstones shall be ingested in a rapid sequence to simulate a hailstone encounter and the number and size of the hailstones shall be determined as follows: (i) One 1-inch (25 millimeters) diameter hailstone for engines with inlet areas of not more than 100 square inches (0.0645 square meters). (ii) One 1-inch (25 millimeters) diameter and one 2-inch (50 millimeters) diameter hailstone for each 150 square inches (0.0968 square meters) of inlet area, or fraction thereof, for engines with inlet areas of more than 100 square inches (0.0645 square meters). (2) In addition to complying with paragraph (a)(1) of this section and except as provided in paragraph (b) of this section, it must be shown that each engine is capable of acceptable operation throughout its specified operating envelope when subjected to sudden encounters with the certification standard concentrations of rain and hail, as defined in appendix B to this part. Acceptable engine operation precludes flameout, run down, continued or non-recoverable surge or stall, or loss of acceleration and deceleration capability, during any three minute continuous period in rain and during any 30 second continuous period in hail. It must also be shown after the ingestion that there is no unacceptable mechanical damage, unacceptable power or thrust loss, or other adverse engine anomalies. (b) Engines for rotorcraft. As an alternative to the requirements specified in paragraph (a)(2) of this section, for rotorcraft turbine engines only, it must be shown that each engine is capable of acceptable operation during and after the ingestion of rain with an overall ratio of water droplet flow to airflow, by weight, with a uniform distribution at the inlet plane, of at least four percent. Acceptable engine operation precludes flameout, run down, continued or non-recoverable surge or stall, or loss of acceleration and deceleration capability. It must also be shown after the ingestion that there is no unacceptable mechanical damage, unacceptable power loss, or other adverse engine anomalies. The rain ingestion must occur under the following static ground level conditions: (1) A normal stabilization period at take-off power without rain ingestion, followed immediately by the suddenly commencing ingestion of rain for three minutes at takeoff power, then (2) Continuation of the rain ingestion during subsequent rapid deceleration to minimum idle, then (3) Continuation of the rain ingestion during three minutes at minimum idle power to be certified for flight operation, then (4) Continuation of the rain ingestion during subsequent rapid acceleration to takeoff power. (c) Engines for supersonic airplanes. In addition to complying with paragraphs (a)(1) and (a)(2) of this section, a separate test for supersonic airplane engines only, shall be conducted with three hailstones ingested at supersonic cruise velocity. These hailstones shall be aimed at the engine's critical face area, and their ingestion must not cause unacceptable mechanical damage or unacceptable power or thrust loss after the ingestion or require the engine to be shut down. The size of these hailstones shall be determined from the linear variation in diameter from 1-inch (25 millimeters) at 35,000 feet (10,500 meters) to 1/4 -inch (6 millimeters) at 60,000 feet (18,000 meters) using the diameter corresponding to the lowest expected supersonic cruise altitude. Alternatively, three larger hailstones may be ingested at subsonic velocities such that the kinetic energy of these larger hailstones is equivalent to the applicable supersonic ingestion conditions. (d) For an engine that incorporates or requires the use of a protection device, demonstration of the rain and hail ingestion capabilities of the engine, as required in paragraphs (a), (b), and (c) of this section, may be waived wholly or in part by the Administrator if the applicant shows that: (1) The subject rain and hail constituents are of a size that will not pass through the protection device; (2) The protection device will withstand the impact of the subject rain and hail constituents; and (3) The subject of rain and hail constituents, stopped by the protection device, will not obstruct the flow of induction air into the engine, resulting in damage, power or thrust loss, or other adverse engine anomalies in excess of what would be accepted in paragraphs (a), (b), and (c) of this section. § 33.79 Fuel burning thrust augmentor. Each fuel burning thrust augmentor, including the nozzle, must— (a) Provide cutoff of the fuel burning thrust augmentor; (b) Permit on-off cycling; (c) Be controllable within the intended range of operation; (d) Upon a failure or malfunction of augmentor combustion, not cause the engine to lose thrust other than that provided by the augmentor; and (e) Have controls that function compatibly with the other engine controls and automatically shut off augmentor fuel flow if the engine rotor speed drops below the minimum rotational speed at which the augmentor is intended to function. § 33.81 Applicability. This subpart prescribes the block tests and inspections for turbine engines. § 33.82 General. Before each endurance test required by this subpart, the adjustment setting and functioning characteristic of each component having an adjustment setting and a functioning characteristic that can be established independent of installation on the engine must be established and recorded. § 33.83 Vibration test. (a) Each engine must undergo vibration surveys to establish that the vibration characteristics of those components that may be subject to mechanically or aerodynamically induced vibratory excitations are acceptable throughout the declared flight envelope. The engine surveys shall be based upon an appropriate combination of experience, analysis, and component test and shall address, as a minimum, blades, vanes, rotor discs, spacers, and rotor shafts. (b) The surveys shall cover the ranges of power or thrust, and both the physical and corrected rotational speeds for each rotor system, corresponding to operations throughout the range of ambient conditions in the declared flight envelope, from the minimum rotational speed up to 103 percent of the maximum physical and corrected rotational speed permitted for rating periods of two minutes or longer, and up to 100 percent of all other permitted physical and corrected rotational speeds, including those that are overspeeds. If there is any indication of a stress peak arising at the highest of those required physical or corrected rotational speeds, the surveys shall be extended sufficiently to reveal the maximum stress values present, except that the extension need not cover more than a further 2 percentage points increase beyond those speeds. (c) Evaluations shall be made of the following: (1) The effects on vibration characteristics of operating with scheduled changes (including tolerances) to variable vane angles, compressor bleeds, accessory loading, the most adverse inlet air flow distortion pattern declared by the manufacturer, and the most adverse conditions in the exhaust duct(s); and (2) The aerodynamic and aeromechanical factors which might induce or influence flutter in those systems susceptible to that form of vibration. (d) Except as provided by paragraph (e) of this section, the vibration stresses associated with the vibration characteristics determined under this section, when combined with the appropriate steady stresses, must be less than the endurance limits of the materials concerned, after making due allowances for operating conditions for the permitted variations in properties of the materials. The suitability of these stress margins must be justified for each part evaluated. If it is determined that certain operating conditions, or ranges, need to be limited, operating and installation limitations shall be established. (e) The effects on vibration characteristics of excitation forces caused by fault conditions (such as, but not limited to, out-of balance, local blockage or enlargement of stator vane passages, fuel nozzle blockage, incorrectly schedule compressor variables, etc.) shall be evaluated by test or analysis, or by reference to previous experience and shall be shown not to create a hazardous condition. (f) Compliance with this section shall be substantiated for each specific installation configuration that can affect the vibration characteristics of the engine. If these vibration effects cannot be fully investigated during engine certification, the methods by which they can be evaluated and methods by which compliance can be shown shall be substantiated and defined in the installation instructions required by § 33.5. § 33.84 Engine overtorque test. (a) If approval of a maximum engine overtorque is sought for an engine incorporating a free power turbine, compliance with this section must be demonstrated by testing. (1) The test may be run as part of the endurance test requirement of § 33.87. Alternatively, tests may be performed on a complete engine or equivalent testing on individual groups of components. (2) Upon conclusion of tests conducted to show compliance with this section, each engine part or individual groups of components must meet the requirements of § 33.93(a)(1) and (a)(2). (b) The test conditions must be as follows: (1) A total of 15 minutes run at the maximum engine overtorque to be approved. This may be done in separate runs, each being of at least 2 1/2 minutes duration. (2) A power turbine rotational speed equal to the highest speed at which the maximum overtorque can occur in service. The test speed may not be more than the limit speed of take-off or OEI ratings longer than 2 minutes. (3) For engines incorporating a reduction gearbox, a gearbox oil temperature equal to the maximum temperature when the maximum engine overtorque could occur in service; and for all other engines, an oil temperature within the normal operating range. (4) A turbine entry gas temperature equal to the maximum steady state temperature approved for use during periods longer than 20 seconds when operating at conditions not associated with 30-second or 2 minutes OEI ratings. The requirement to run the test at the maximum approved steady state temperature may be waived by the FAA if the applicant can demonstrate that other testing provides substantiation of the temperature effects when considered in combination with the other parameters identified in paragraphs (b)(1), (b)(2) and (b)(3) of this section. § 33.85 Calibration tests. (a) Each engine must be subjected to those calibration tests necessary to establish its power characteristics and the conditions for the endurance test specified § 33.87. The results of the power characteristics calibration tests form the basis for establishing the characteristics of the engine over its entire operating range of speeds, pressures, temperatures, and altitudes. Power ratings are based upon standard atmospheric conditions with no airbleed for aircraft services and with only those accessories installed which are essential for engine functioning. (b) A power check at sea level conditions must be accomplished on the endurance test engine after the endurance test and any change in power characteristics which occurs during the endurance test must be determined. Measurements taken during the final portion of the endurance test may be used in showing compliance with the requirements of this paragraph. (c) In showing compliance with this section, each condition must stabilize before measurements are taken, except as permitted by paragraph (d) of this section. (d) In the case of engines having 30-second OEI, and 2-minute OEI ratings, measurements taken during the applicable endurance test prescribed in § 33.87(f) (1) through (8) may be used in showing compliance with the requirements of this section for these OEI ratings. § 33.87 Endurance test. (a) General. Each engine must be subjected to an endurance test that includes a total of at least 150 hours of operation and, depending upon the type and contemplated use of the engine, consists of one of the series of runs specified in paragraphs (b) through (g) of this section, as applicable. For engines tested under paragraphs (b), (c), (d), (e) or (g) of this section, the prescribed 6-hour test sequence must be conducted 25 times to complete the required 150 hours of operation. Engines for which the 30-second OEI and 2-minute OEI ratings are desired must be further tested under paragraph (f) of this section. The following test requirements apply: (1) The runs must be made in the order found appropriate by the FAA for the particular engine being tested. (2) Any automatic engine control that is part of the engine must control the engine during the endurance test except for operations where automatic control is normally overridden by manual control or where manual control is otherwise specified for a particular test run. (3) Except as provided in paragraph (a)(5) of this section, power or thrust, gas temperature, rotor shaft rotational speed, and, if limited, temperature of external surfaces of the engine must be at least 100 percent of the value associated with the particular engine operation being tested. More than one test may be run if all parameters cannot be held at the 100 percent level simultaneously. (4) The runs must be made using fuel, lubricants and hydraulic fluid which conform to the specifications specified in complying with § 33.7(c). (5) Maximum air bleed for engine and aircraft services must be used during at least one-fifth of the runs, except for the test required under paragraph (f) of this section, provided the validity of the test is not compromised. However, for these runs, the power or thrust or the rotor shaft rotational speed may be less than 100 percent of the value associated with the particular operation being tested if the FAA finds that the validity of the endurance test is not compromised. (6) Each accessory drive and mounting attachment must be loaded in accordance with paragraphs (a)(6)(i) and (ii) of this section, except as permitted by paragraph (a)(6)(iii) of this section for the test required under paragraph (f) of this section. (i) The load imposed by each accessory used only for aircraft service must be the limit load specified by the applicant for the engine drive and attachment point during rated maximum continuous power or thrust and higher output. (ii) The endurance test of any accessory drive and mounting attachment under load may be accomplished on a separate rig if the validity of the test is confirmed by an approved analysis. (iii) The applicant is not required to load the accessory drives and mounting attachments when running the tests under paragraphs (f)(1) through (f)(8) of this section if the applicant can substantiate that there is no significant effect on the durability of any accessory drive or engine component. However, the applicant must add the equivalent engine output power extraction from the power turbine rotor assembly to the engine shaft output. (7) During the runs at any rated power or thrust the gas temperature and the oil inlet temperature must be maintained at the limiting temperature except where the test periods are not longer than 5 minutes and do not allow stabilization. At least one run must be made with fuel, oil, and hydraulic fluid at the minimum pressure limit and at least one run must be made with fuel, oil, and hydraulic fluid at the maximum pressure limit with fluid temperature reduced as necessary to allow maximum pressure to be attained. (8) If the number of occurrences of either transient rotor shaft overspeed, transient gas overtemperature or transient engine overtorque is limited, that number of the accelerations required by paragraphs (b) through (g) of this section must be made at the limiting overspeed, overtemperature or overtorque. If the number of occurrences is not limited, half the required accelerations must be made at the limiting overspeed, overtemperature or overtorque. (9) For each engine type certificated for use on supersonic aircraft the following additional test requirements apply: (i) To change the thrust setting, the power control lever must be moved from the initial position to the final position in not more than one second except for movements into the fuel burning thrust augmentor augmentation position if additional time to confirm ignition is necessary. (ii) During the runs at any rated augmented thrust the hydraulic fluid temperature must be maintained at the limiting temperature except where the test periods are not long enough to allow stabilization. (iii) During the simulated supersonic runs the fuel temperature and induction air temperature may not be less than the limiting temperature. (iv) The endurance test must be conducted with the fuel burning thrust augmentor installed, with the primary and secondary exhaust nozzles installed, and with the variable area exhaust nozzles operated during each run according to the methods specified in complying with § 33.5(b). (v) During the runs at thrust settings for maximum continuous thrust and percentages thereof, the engine must be operated with the inlet air distortion at the limit for those thrust settings. (b) Engines other than certain rotorcraft engines. For each engine except a rotorcraft engine for which a rating is desired under paragraph (c), (d), or (e) of this section, the applicant must conduct the following runs: (1) Takeoff and idling. One hour of alternate five-minute periods at rated takeoff power or thrust and at idling power or thrust. The developed powers or thrusts at takeoff and idling conditions and their corresponding rotor speed and gas temperature conditions must be as established by the power control in accordance with the schedule established by the applicant. The applicant may, during any one period, manually control the rotor speed, power, or thrust while taking data to check performance. For engines with augmented takeoff power ratings that involve increases in turbine inlet temperature, rotor speed, or shaft power, this period of running at takeoff must be at the augmented rating. For engines with augmented takeoff power ratings that do not materially increase operating severity, the amount of running conducted at the augmented rating is determined by the FAA. In changing the power setting after each period, the power-control lever must be moved in the manner prescribed in paragraph (b)(5) of this section. (2) Rated maximum continuous and takeoff power or thrust. Thirty minutes at— (i) Rated maximum continuous power or thrust during fifteen of the twenty-five 6-hour endurance test cycles; and (ii) Rated takeoff power or thrust during ten of the twenty-five 6-hour endurance test cycles. (3) Rated maximum continuous power or thrust. One hour and 30 minutes at rated maximum continuous power or thrust. (4) Incremental cruise power or thrust. Two hours and 30 minutes at the successive power lever positions corresponding to at least 15 approximately equal speed and time increments between maximum continuous engine rotational speed and ground or minimum idle rotational speed. For engines operating at constant speed, the thrust and power may be varied in place of speed. If there is significant peak vibration anywhere between ground idle and maximum continuous conditions, the number of increments chosen may be changed to increase the amount of running made while subject to the peak vibrations up to not more than 50 percent of the total time spent in incremental running. (5) Acceleration and deceleration runs. 30 minutes of accelerations and decelerations, consisting of six cycles from idling power or thrust to rated takeoff power or thrust and maintained at the takeoff power lever position for 30 seconds and at the idling power lever position for approximately four and one-half minutes. In complying with this paragraph, the power-control lever must be moved from one extreme position to the other in not more than one second, except that, if different regimes of control operations are incorporated necessitating scheduling of the power-control lever motion in going from one extreme position to the other, a longer period of time is acceptable, but not more than two seconds. (6) Starts. One hundred starts must be made, of which 25 starts must be preceded by at least a two-hour engine shutdown. There must be at least 10 false engine starts, pausing for the applicant's specified minimum fuel drainage time, before attempting a normal start. There must be at least 10 normal restarts with not longer than 15 minutes since engine shutdown. The remaining starts may be made after completing the 150 hours of endurance testing. (c) Rotorcraft engines for which a 30-minute OEI power rating is desired. For each rotorcraft engine for which a 30-minute OEI power rating is desired, the applicant must conduct the following series of tests: (1) Takeoff and idling. One hour of alternate 5-minute periods at rated takeoff power and at idling power. The developed powers at takeoff and idling conditions and their corresponding rotor speed and gas temperature conditions must be as established by the power control in accordance with the schedule established by the applicant. During any one period, the rotor speed and power may be controlled manually while taking data to check performance. For engines with augmented takeoff power ratings that involve increases in turbine inlet temperature, rotor speed, or shaft power, this period of running at rated takeoff power must be at the augmented power rating. In changing the power setting after each period, the power control lever must be moved in the manner prescribed in paragraph (c)(6) of this section. (2) Rated maximum continuous and takeoff power. Thirty minutes at— (i) Rated maximum continuous power during fifteen of the twenty-five 6-hour endurance test cycles; and (ii) Rated takeoff power during ten of the twenty-five 6-hour endurance test cycles. (3) Rated maximum continuous power. One hour at rated maximum continuous power. (4) Rated 30-minute OEI power. Thirty minutes at rated 30-minute OEI power. (5) Incremental cruise power. Two hours and 30 minutes at the successive power lever positions corresponding with not less than 15 approximately equal speed and time increments between maximum continuous engine rotational speed and ground or minimum idle rotational speed. For engines operating at constant speed, power may be varied in place of speed. If there are significant peak vibrations anywhere between ground idle and maximum continuous conditions, the number of increments chosen must be changed to increase the amount of running conducted while subject to peak vibrations up to not more than 50 percent of the total time spent in incremental running. (6) Acceleration and deceleration runs. Thirty minutes of accelerations and decelerations, consisting of six cycles from idling power to rated takeoff power and maintained at the takeoff power lever position for 30 seconds and at the idling power lever position for approximately 4 1/2 minutes. In complying with this paragraph, the power control lever must be moved from one extreme position to the other in not more than one second. If, however, different regimes of control operations are incorporated that necessitate scheduling of the power control lever motion from one extreme position to the other, then a longer period of time is acceptable, but not more than two seconds. (7) Starts. One hundred starts, of which 25 starts must be preceded by at least a two-hour engine shutdown. There must be at least 10 false engine starts, pausing for the applicant's specified minimum fuel drainage time, before attempting a normal start. There must be at least 10 normal restarts not more than 15 minutes after engine shutdown. The remaining starts may be made after completing the 150 hours of endurance testing. (d) Rotorcraft engines for which a continuous OEI rating is desired. For each rotorcraft engine for which a continuous OEI power rating is desired, the applicant must conduct the following series of tests: (1) Takeoff and idling. One hour of alternate 5-minute periods at rated takeoff power and at idling power. The developed powers at takeoff and idling conditions and their corresponding rotor speed and gas temperature conditions must be as established by the power control in accordance with the schedule established by the applicant. During any one period the rotor speed and power may be controlled manually while taking data to check performance. For engines with augmented takeoff power ratings that involve increases in turbine inlet temperature, rotor speed, or shaft power, this period of running at rated takeoff power must be at the augmented power rating. In changing the power setting after each period, the power control lever must be moved in the manner prescribed in paragraph (d)(6) of this section. (2) Rated maximum continuous and takeoff power. Thirty minutes at— (i) Rated maximum continuous power during fifteen of the twenty-five 6-hour endurance test cycles; and (ii) Rated takeoff power during ten of the twenty-five 6-hour endurance test cycles. (3) Rated continuous OEI power. One hour at rated continuous OEI power. (4) Rated maximum continuous power. One hour at rated maximum continuous power. (5) Incremental cruise power. Two hours at the successive power lever positions corresponding with not less than 12 approximately equal speed and time increments between maximum continuous engine rotational speed and ground or minimum idle rotational speed. For engines operating at constant speed, power may be varied in place of speed. If there are significant peak vibrations anywhere between ground idle and maximum continuous conditions, the number of increments chosen must be changed to increase the amount of running conducted while being subjected to the peak vibrations up to not more than 50 percent of the total time spent in incremental running. (6) Acceleration and deceleration runs. Thirty minutes of accelerations and decelerations, consisting of six cycles from idling power to rated takeoff power and maintained at the takeoff power lever position for 30 seconds and at the idling power lever position for approximately 4 1/2 minutes. In complying with this paragraph, the power control lever must be moved from one extreme position to the other in not more than 1 second, except that if different regimes of control operations are incorporated necessitating scheduling of the power control lever motion in going from one extreme position to the other, a longer period of time is acceptable, but not more than 2 seconds. (7) Starts. One hundred starts, of which 25 starts must be preceded by at least a 2-hour engine shutdown. There must be at least 10 false engine starts, pausing for the applicant's specified minimum fuel drainage time, before attempting a normal start. There must be at least 10 normal restarts with not longer than 15 minutes since engine shutdown. The remaining starts may be made after completing the 150 hours of endurance testing. (e) Rotorcraft engines for which a 2 1/2 -minute OEI power rating is desired. For each rotorcraft engine for which a 2 1/2 -minute OEI power rating is desired, the applicant must conduct the following series of tests: (1) Takeoff, 2 1/2 -minute OEI, and idling. One hour of alternate 5-minute periods at rated takeoff power and at idling power except that, during the third and sixth takeoff power periods, only 2 1/2 minutes need be conducted at rated takeoff power, and the remaining 2 1/2 minutes must be conducted at rated 2 1/2 -minute OEI power. The developed powers at takeoff, 2 1/2 -minute OEI, and idling conditions and their corresponding rotor speed and gas temperature conditions must be as established by the power control in accordance with the schedule established by the applicant. The applicant may, during any one period, control manually the rotor speed and power while taking data to check performance. For engines with augmented takeoff power ratings that involve increases in turbine inlet temperature, rotor speed, or shaft power, this period of running at rated takeoff power must be at the augmented rating. In changing the power setting after or during each period, the power control lever must be moved in the manner prescribed in paragraph (b)(5), (c)(6), or (d)(6) of this section, as applicable. (2) The tests required in paragraphs (b)(2) through (b)(6), or (c)(2) through (c)(7), or (d)(2) through (d)(7) of this section, as applicable, except that in one of the 6-hour test sequences, the last 5 minutes of the 30 minutes at takeoff power test period of paragraph (b)(2) of this section, or of the 30 minutes at 30-minute OEI power test period of paragraph (c)(4) of this section, or of the l hour at continuous OEI power test period of paragraph (d)(3) of this section, must be run at 2 1/2 -minute OEI power. (f) Rotorcraft Engines for which 30-second OEI and 2-minute OEI ratings are desired. For each rotorcraft engine for which 30-second OEI and 2-minute OEI power ratings are desired, and following completion of the tests under paragraphs (b), (c), (d), or (e) of this section, the applicant may disassemble the tested engine to the extent necessary to show compliance with the requirements of § 33.93(a). The tested engine must then be reassembled using the same parts used during the test runs of paragraphs (b), (c), (d), or (e) of this section, except those parts described as consumables in the Instructions for Continued Airworthiness. Additionally, the tests required in paragraphs (f)(1) through (f)(8) of this section must be run continuously. If a stop occurs during these tests, the interrupted sequence must be repeated unless the applicant shows that the severity of the test would not be reduced if it were continued. The applicant must conduct the following test sequence four times, for a total time of not less than 120 minutes: (1) Takeoff power. Three minutes at rated takeoff power. (2) 30-second OEI power. Thirty seconds at rated 30-second OEI power. (3) 2-minute OEI power. Two minutes at rated 2-minute OEI power. (4) 30-minute OEI power, continuous OEI power, or maximum continuous power. Five minutes at whichever is the greatest of rated 30-minute OEI power, rated continuous OEI power, or rated maximum continuous power, except that, during the first test sequence, this period shall be 65 minutes. However, where the greatest rated power is 30-minute OEI power, that sixty-five minute period shall consist of 30 minutes at 30-minute OEI power followed by 35 minutes at whichever is the greater of continuous OEI power or maximum continuous power. (5) 50 percent takeoff power. One minute at 50 percent takeoff power. (6) 30-second OEI power. Thirty seconds at rated 30-second OEI power. (7) 2-minute OEI power. Two minutes at rated 2-minute OEI power. (8) Idle. One minute at flight idle. (g) Supersonic aircraft engines. For each engine type certificated for use on supersonic aircraft the applicant must conduct the following: (1) Subsonic test under sea level ambient atmospheric conditions. Thirty runs of one hour each must be made, consisting of— (i) Two periods of 5 minutes at rated takeoff augmented thrust each followed by 5 minutes at idle thrust; (ii) One period of 5 minutes at rated takeoff thrust followed by 5 minutes at not more than 15 percent of rated takeoff thrust; (iii) One period of 10 minutes at rated takeoff augmented thrust followed by 2 minutes at idle thrust, except that if rated maximum continuous augmented thrust is lower than rated takeoff augmented thrust, 5 of the 10-minute periods must be at rated maximum continuous augmented thrust; and (iv) Six periods of 1 minute at rated takeoff augmented thrust each followed by 2 minutes, including acceleration and deceleration time, at idle thrust. (2) Simulated supersonic test. Each run of the simulated supersonic test must be preceded by changing the inlet air temperature and pressure from that attained at subsonic condition to the temperature and pressure attained at supersonic velocity, and must be followed by a return to the temperature attained at subsonic condition. Thirty runs of 4 hours each must be made, consisting of— (i) One period of 30 minutes at the thrust obtained with the power control lever set at the position for rated maximum continuous augmented thrust followed by 10 minutes at the thrust obtained with the power control lever set at the position for 90 percent of rated maximum continuous augmented thrust. The end of this period in the first five runs must be made with the induction air temperature at the limiting condition of transient overtemperature, but need not be repeated during the periods specified in paragraphs (g)(2)(ii) through (iv) of this section; (ii) One period repeating the run specified in paragraph (g)(2)(i) of this section, except that it must be followed by 10 minutes at the thrust obtained with the power control lever set at the position for 80 percent of rated maximum continuous augmented thrust; (iii) One period repeating the run specified in paragraph (g)(2)(i) of this section, except that it must be followed by 10 minutes at the thrust obtained with the power control lever set at the position for 60 percent of rated maximum continuous augmented thrust and then 10 minutes at not more than 15 percent of rated takeoff thrust; (iv) One period repeating the runs specified in paragraphs (g)(2)(i) and (ii) of this section; and (v) One period of 30 minutes with 25 of the runs made at the thrust obtained with the power control lever set at the position for rated maximum continuous augmented thrust, each followed by idle thrust and with the remaining 5 runs at the thrust obtained with the power control lever set at the position for rated maximum continuous augmented thrust for 25 minutes each, followed by subsonic operation at not more than 15 percent or rated takeoff thrust and accelerated to rated takeoff thrust for 5 minutes using hot fuel. (3) Starts. One hundred starts must be made, of which 25 starts must be preceded by an engine shutdown of at least 2 hours. There must be at least 10 false engine starts, pausing for the applicant's specified minimum fuel drainage time before attempting a normal start. At least 10 starts must be normal restarts, each made no later than 15 minutes after engine shutdown. The starts may be made at any time, including the period of endurance testing. § 33.88 Engine overtemperature test. (a) Each engine must run for 5 minutes at maximum permissible rpm with the gas temperature at least 75 °F (42 °C) higher than the maximum rating's steady-state operating limit, excluding maximum values of rpm and gas temperature associated with the 30-second OEI and 2-minute OEI ratings. Following this run, the turbine assembly must be within serviceable limits. (b) In addition to the test requirements in paragraph (a) of this section, each engine for which 30-second OEI and 2-minute OEI ratings are desired, that incorporates a means for automatic temperature control within its operating limitations in accordance with § 33.28(k), must run for a period of 4 minutes at the maximum power-on rpm with the gas temperature at least 35 °F (19 °C) higher than the maximum operating limit at 30-second OEI rating. Following this run, the turbine assembly may exhibit distress beyond the limits for an overtemperature condition provided the engine is shown by analysis or test, as found necessary by the FAA, to maintain the integrity of the turbine assembly. (c) A separate test vehicle may be used for each test condition. § 33.89 Operation test. (a) The operation test must include testing found necessary by the Administrator to demonstrate— (1) Starting, idling, acceleration, overspeeding, ignition, functioning of the propeller (if the engine is designated to operate with a propeller); (2) Compliance with the engine response requirements of § 33.73; and (3) The minimum power or thrust response time to 95 percent rated takeoff power or thrust, from power lever positions representative of minimum idle and of minimum flight idle, starting from stabilized idle operation, under the following engine load conditions: (i) No bleed air and power extraction for aircraft use. (ii) Maximum allowable bleed air and power extraction for aircraft use. (iii) An intermediate value for bleed air and power extraction representative of that which might be used as a maximum for aircraft during approach to a landing. (4) If testing facilities are not available, the determination of power extraction required in paragraph (a)(3)(ii) and (iii) of this section may be accomplished through appropriate analytical means. (b) The operation test must include all testing found necessary by the Administrator to demonstrate that the engine has safe operating characteristics throughout its specified operating envelope. § 33.90 Initial maintenance inspection test. Each applicant, except an applicant for an engine being type certificated through amendment of an existing type certificate or through supplemental type certification procedures, must complete one of the following tests on an engine that substantially conforms to the type design to establish when the initial maintenance inspection is required: (a) An approved engine test that simulates the conditions in which the engine is expected to operate in service, including typical start-stop cycles. (b) An approved engine test conducted in accordance with § 33.201 (c) through (f). § 33.91 Engine system and component tests. (a) For those systems or components that cannot be adequately substantiated in accordance with endurance testing of § 33.87, the applicant must conduct additional tests to demonstrate that the systems or components are able to perform the intended functions in all declared environmental and operating conditions. (b) Temperature limits must be established for those components that require temperature controlling provisions in the aircraft installation to assure satisfactory functioning, reliability, and durability. (c) Each unpressurized hydraulic fluid tank may not fail or leak when subjected to a maximum operating temperature and an internal pressure of 5 p.s.i., and each pressurized hydraulic fluid tank must meet the requirements of § 33.64. (d) For an engine type certificated for use in supersonic aircraft, the systems, safety devices, and external components that may fail because of operation at maximum and minimum operating temperatures must be identified and tested at maximum and minimum operating temperatures and while temperature and other operating conditions are cycled between maximum and minimum operating values. § 33.92 Rotor locking tests. If continued rotation is prevented by a means to lock the rotor(s), the engine must be subjected to a test that includes 25 operations of this means under the following conditions: (a) The engine must be shut down from rated maximum continuous thrust or power; and (b) The means for stopping and locking the rotor(s) must be operated as specified in the engine operating instructions while being subjected to the maximum torque that could result from continued flight in this condition; and (c) Following rotor locking, the rotor(s) must be held stationary under these conditions for five minutes for each of the 25 operations. § 33.93 Teardown inspection. (a) After completing the endurance testing of § 33.87 (b), (c), (d), (e), or (g) of this part, each engine must be completely disassembled, and (1) Each component having an adjustment setting and a functioning characteristic that can be established independent of installation on the engine must retain each setting and functioning characteristic within the limits that were established and recorded at the beginning of the test; and (2) Each engine part must conform to the type design and be eligible for incorporation into an engine for continued operation, in accordance with information submitted in compliance with § 33.4. (b) After completing the endurance testing of § 33.87(f), each engine must be completely disassembled, and (1) Each component having an adjustment setting and a functioning characteristic that can be established independent of installation on the engine must retain each setting and functioning characteristic within the limits that were established and recorded at the beginning of the test; and (2) Each engine may exhibit deterioration in excess of that permitted in paragraph (a)(2) of this section, including some engine parts or components that may be unsuitable for further use. The applicant must show by inspection, analysis, test, or by any combination thereof as found necessary by the FAA, that structural integrity of the engine is maintained; or (c) In lieu of compliance with paragraph (b) of this section, each engine for which the 30-second OEI and 2-minute OEI ratings are desired, may be subjected to the endurance testing of §§ 33.87 (b), (c), (d), or (e) of this part, and followed by the testing of § 33.87(f) without intervening disassembly and inspection. However, the engine must comply with paragraph (a) of this section after completing the endurance testing of § 33.87(f). § 33.94 Blade containment and rotor unbalance tests. (a) Except as provided in paragraph (b) of this section, it must be demonstrated by engine tests that the engine is capable of containing damage without catching fire and without failure of its mounting attachments when operated for at least 15 seconds, unless the resulting engine damage induces a self shutdown, after each of the following events: (1) Failure of the most critical compressor or fan blade while operating at maximum permissible r.p.m. The blade failure must occur at the outermost retention groove or, for integrally-bladed rotor discs, at least 80 percent of the blade must fail. (2) Failure of the most critical turbine blade while operating at maximum permissible r.p.m. The blade failure must occur at the outermost retention groove or, for integrally-bladed rotor discs, at least 80 percent of the blade must fail. The most critical turbine blade must be determined by considering turbine blade weight and the strength of the adjacent turbine case at case temperatures and pressures associated with operation at maximum permissible r.p.m. (b) Analysis based on rig testing, component testing, or service experience may be substitute for one of the engine tests prescribed in paragraphs (a)(1) and (a)(2) of this section if— (1) That test, of the two prescribed, produces the least rotor unbalance; and (2) The analysis is shown to be equivalent to the test. § 33.95 Engine-propeller systems tests. If the engine is designed to operate with a propeller, the following tests must be made with a representative propeller installed by either including the tests in the endurance run or otherwise performing them in a manner acceptable to the Administrator: (a) Feathering operation: 25 cycles. (b) Negative torque and thrust system operation: 25 cycles from rated maximum continuous power. (c) Automatic decoupler operation: 25 cycles from rated maximum continuous power (if repeated decoupling and recoupling in service is the intended function of the device). (d) Reverse thrust operation: 175 cycles from the flight-idle position to full reverse and 25 cycles at rated maximum continuous power from full forward to full reverse thrust. At the end of each cycle the propeller must be operated in reverse pitch for a period of 30 seconds at the maximum rotational speed and power specified by the applicant for reverse pitch operation. § 33.96 Engine tests in auxiliary power unit (APU) mode. If the engine is designed with a propeller brake which will allow the propeller to be brought to a stop while the gas generator portion of the engine remains in operation, and remain stopped during operation of the engine as an auxiliary power unit (“APU mode”), in addition to the requirements of § 33.87, the applicant must conduct the following tests: (a) Ground locking: A total of 45 hours with the propeller brake engaged in a manner which clearly demonstrates its ability to function without adverse effects on the complete engine while the engine is operating in the APU mode under the maximum conditions of engine speed, torque, temperature, air bleed, and power extraction as specified by the applicant. (b) Dynamic braking: A total of 400 application-release cycles of brake engagements must be made in a manner which clearly demonstrates its ability to function without adverse effects on the complete engine under the maximum conditions of engine acceleration/deceleration rate, speed, torque, and temperature as specified by the applicant. The propeller must be stopped prior to brake release. (c) One hundred engine starts and stops with the propeller brake engaged. (d) The tests required by paragraphs (a), (b), and (c) of this section must be performed on the same engine, but this engine need not be the same engine used for the tests required by § 33.87. (e) The tests required by paragraphs (a), (b), and (c) of this section must be followed by engine disassembly to the extent necessary to show compliance with the requirements of § 33.93(a) and § 33.93(b). § 33.97 Thrust reversers. (a) If the engine incorporates a reverser, the endurance, calibration, operation, and vibration tests prescribed in this subpart must be run with the reverser installed. In complying with this section, the power control lever must be moved from one extreme position to the other in not more than one second except, if regimes of control operations are incorporated necessitating scheduling of the power-control lever motion in going from one extreme position to the other, a longer period of time is acceptable but not more than three seconds. In addition, the test prescribed in paragraph (b) of this section must be made. This test may be scheduled as part of the endurance run. (b) 175 reversals must be made from flight-idle forward thrust to maximum reverse thrust and 25 reversals must be made from rated takeoff thrust to maximum reverse thrust. After each reversal the reverser must be operated at full reverse thrust for a period of one minute, except that, in the case of a reverser intended for use only as a braking means on the ground, the reverser need only be operated at full reverse thrust for 30 seconds. § 33.99 General conduct of block tests. (a) Each applicant may, in making a block test, use separate engines of identical design and construction in the vibration, calibration, endurance, and operation tests, except that, if a separate engine is used for the endurance test it must be subjected to a calibration check before starting the endurance test. (b) Each applicant may service and make minor repairs to the engine during the block tests in accordance with the service and maintenance instructions submitted in compliance with § 33.4. If the frequency of the service is excessive, or the number of stops due to engine malfunction is excessive, or a major repair, or replacement of a part is found necessary during the block tests or as the result of findings from the teardown inspection, the engine or its parts must be subjected to any additional tests the Administrator finds necessary. (c) Each applicant must furnish all testing facilities, including equipment and competent personnel, to conduct the block tests. § 33.201 Design and test requirements for Early ETOPS eligibility. An applicant seeking type design approval for an engine to be installed on a two-engine airplane approved for ETOPS without the service experience specified in part 25, appendix K, K25.2.1 of this chapter, must comply with the following: (a) The engine must be designed using a design quality process acceptable to the FAA, that ensures the design features of the engine minimize the occurrence of failures, malfunctions, defects, and maintenance errors that could result in an IFSD, loss of thrust control, or other power loss. (b) The design features of the engine must address problems shown to result in an IFSD, loss of thrust control, or other power loss in the applicant's other relevant type designs approved within the past 10 years, to the extent that adequate service data is available within that 10-year period. An applicant without adequate service data must show experience with and knowledge of problem mitigating design practices equivalent to that gained from actual service experience in a manner acceptable to the FAA. (c) Except as specified in paragraph (f) of this section, the applicant must conduct a simulated ETOPS mission cyclic endurance test in accordance with an approved test plan on an engine that substantially conforms to the type design. The test must: (1) Include a minimum of 3,000 representative service start-stop mission cycles and three simulated diversion cycles at maximum continuous thrust or power for the maximum diversion time for which ETOPS eligibility is sought. Each start-stop mission cycle must include the use of take-off, climb, cruise, descent, approach, and landing thrust or power and the use of thrust reverse (if applicable). The diversions must be evenly distributed over the duration of the test. The last diversion must be conducted within 100 cycles of the completion of the test. (2) Be performed with the high speed and low speed main engine rotors independently unbalanced to obtain a minimum of 90 percent of the recommended field service maintenance vibration levels. For engines with three main engine rotors, the intermediate speed rotor must be independently unbalanced to obtain a minimum of 90 percent of the recommended production acceptance vibration level. The required peak vibration levels must be verified during a slow acceleration and deceleration run of the test engine covering the main engine rotor operating speed ranges. (3) Include a minimum of three million vibration cycles for each 60 rpm incremental step of the typical high-speed rotor start-stop mission cycle. The test may be conducted using any rotor speed step increment from 60 to 200 rpm provided the test encompasses the typical service start-stop cycle speed range. For incremental steps greater than 60 rpm, the minimum number of vibration cycles must be linearly increased up to ten million cycles for a 200 rpm incremental step. (4) Include a minimum of 300,000 vibration cycles for each 60 rpm incremental step of the high-speed rotor approved operational speed range between minimum flight idle and cruise power not covered by paragraph (c)(3) of this section. The test may be conducted using any rotor speed step increment from 60 to 200 rpm provided the test encompasses the applicable speed range. For incremental steps greater than 60 rpm the minimum number of vibration cycles must be linearly increased up to 1 million for a 200 rpm incremental step. (5) Include vibration surveys at periodic intervals throughout the test. The equivalent value of the peak vibration level observed during the surveys must meet the minimum vibration requirement of § 33.201(c)(2). (d) Prior to the test required by paragraph (c) of this section, the engine must be subjected to a calibration test to document power and thrust characteristics. (e) At the conclusion of the testing required by paragraph (c) of this section, the engine must: (1) Be subjected to a calibration test at sea-level conditions. Any change in power or thrust characteristics must be within approved limits. (2) Be visually inspected in accordance with the on-wing inspection recommendations and limits contained in the Instructions for Continued Airworthiness submitted in compliance with § 33.4. (3) Be completely disassembled and inspected— (i) In accordance with the applicable inspection recommendations and limits contained in the Instructions for Continued Airworthiness submitted in compliance with § 33.4; (ii) With consideration of the causes of IFSD, loss of thrust control, or other power loss identified by paragraph (b) of this section; and (iii) In a manner to identify wear or distress conditions that could result in an IFSD, loss of thrust control, or other power loss not specifically identified by paragraph (b) of this section or addressed within the Instructions for Continued Airworthiness. (4) Not show wear or distress to the extent that could result in an IFSD, loss of thrust control, or other power loss within a period of operation before the component, assembly, or system would likely have been inspected or functionally tested for integrity while in service. Such wear or distress must have corrective action implemented through a design change, a change to maintenance instructions, or operational procedures before ETOPS eligibility is granted. The type and frequency of wear and distress that occurs during the engine test must be consistent with the type and frequency of wear and distress that would be expected to occur on ETOPS eligible engines. (f) An alternative mission cycle endurance test that provides an equivalent demonstration of the unbalance and vibration specified in paragraph (c) of this section may be used when approved by the FAA. (g) For an applicant using the simulated ETOPS mission cyclic endurance test to comply with § 33.90, the test may be interrupted so that the engine may be inspected by an on-wing or other method, using criteria acceptable to the FAA, after completion of the test cycles required to comply with § 33.90(a). Following the inspection, the ETOPS test must be resumed to complete the requirements of this section. CCAR-33 原文 CCAR-33 来源 : CAAC官网 信息公开 - 民航规章 中国民用航空总局关于修订《航空发动机适航标准》的决定 中国民用航空总局令 第109号 《中国民用航空总局关于修订《航空发动机适航标准》的决定 》已经2002年3月20日中国民用航空总局局务会议通过，现予公布，自2002年4月19日起施行。 局长 刘剑锋 二〇〇二年三月二十日 中国民用航空总局依据《中华人民共和国民用航空法》第三十四条，决定对《航空发动机适航标准》（CCAR-33）作如下修订： 一、规章名称“航空发动机适航标准”修改为“航空发动机适航规定”。 ** 二、原规章中A分部、B分部、C分部、D分部、E分部、F分部分别改为A章、B章、C章、D章、E章、F章。 ** 三、原规章中关于条的序号的表述“§……”改为“第……条”。 四、增加第33.28条，内容如下: 依靠电气和电子装置进行正常工作的每一控制系统必须满足下列要求： （a）在第33.5条所要求的发动机安装和使用说明手册中应对控制系统进行说明、并应规定在正常工作和失效状态所控制的可用功率或推力的百分比、以及其他被控制的功能的控制范围； （b） 控制系统的设计和构造应能保证由飞机提供的电源或数据的任何失效不应导致功率或推力发生不可接受的变化，或妨碍发动机继续安全运转； （c） 控制系统的设计和构造应能保证不会由于控制系统电气或电子部件的单个失效或故障，或可能发生的组合失效，而导致不安全状态的发生； （d） 在该安装和使用说明手册中应规定环境限制，包括雷击引起的瞬变状态；并且 （e） 所有相关软件的设计和执行应具有防止导致不可接受的功率或推力损失或其他不安全状态的防错功能，并且，软件的设计和实施方法须经中国民用航空总局批准。 五、增加第33.74条，内容如下 ： 第33.74条 持续转动 由于飞行中的任何原因使发动机停车，如果发动机的任何主转动系统仍持续转动并且没有提供阻止持续转动的装置，那么在最长的飞行周期内和在预期该发动机不工作的飞行条件下，任何持续的转动不得导致第33.75条(a)至(c)所描述的任何情况。 六、增加第33.76条，内容如下： （a）概述 为符合本条（b）、（c），应遵照下列规定： （1）吸鸟试验应在吸鸟前的试验天气环境条件下，发动机稳定在不小于100%的起飞功率或推力的状态下进行。另外，符合性的验证必须考虑在海平面最热天气的起飞条件下最差的发动机能够达到最大额定起飞功率或推力的运转情况。 （2）应由申请人来确定在本条中用来决定鸟的数量和重量的发动机进气道喉道面积，并且将其确认为第33.5条所要求的安装说明中的一个限制。 （3）必须对可能进入进气道的单只大鸟和单只最大的中鸟对发动机前部的撞击进行评估。必须证明，当按本条（b）或（c）的规定的条件（如适用）撞击相关部件时，不会影响发动机，使之达到不符合本条（b）（3）和（c）（6）要求的程度。 （4） 对于采用进气道防护装置的发动机，本条的符合性验证应在该防护 装置起作用的情况下进行。发动机的批准文件上应注明对这些要求的符合性验证是在防护装置起作用的情况下进行的。 （5）按本条（b）和（c）的要求进行吸鸟试验时，可用中国民用航空总局可接受的物体代替鸟。 （6）如果本条中各项要求的符合性未被验证，在发动机的型号审定文件中应说明该发动机应仅限于安装在不可能发生鸟撞击发动机，或者发动机不会吸入鸟，或者鸟不会对进入发动机的气流产生不利限制的航空器上。 （b） 大鸟为符合大鸟吸入的要求，应遵照下列规定： （1） 大鸟吸入试验应使用表1规定重量的1只鸟。该鸟应投向第一级旋转叶片最关键的暴露位置。对于安装在固定翼飞机上的发动机，吸入鸟的速度应为370公里/小时（200节）；对于安装在旋翼航空器上的发动机，吸入鸟的速度应为旋翼航空器正常飞行时的最大的空速。 （2）在大鸟吸入后的15秒内不允许移动功率杆。 （3）在本条规定的条件下进行单只大鸟的吸鸟试验时，不得导致发动机出现下列情况之一： （i）着火； （ii）危险的碎片穿透发动机机匣飞出； （iii）产生的载荷大于33.23(a)中规定的极限载荷； （iv）失去停车能力。 （4）对本款中大鸟吸入要求的符合性验证也可以通过验证第33.94条(a)中在叶片包容性和转子不平衡性方面的各项要求比本条的各项要求更为严格来证明。 表1 大鸟的重量要求 （c） 中鸟和小鸟 为符合中鸟和小鸟吸入的要求，应遵照下列规定： （1） 应采用中国民用航空总局可接受的分析方法或部件试验或两者的组合，来确定影响功率损失和造成损坏的关键吸鸟参数。关键吸鸟参数应包括，但不限于，鸟速、关键目标位置和第一级转子转速的影响。吸鸟临界速度应反映从地面到地面上460米（1500英尺）的正常飞行高度所使用的空速范围内的最严酷条件，但不应小于飞机的V1最小速度。 （2） 应进行吸中鸟的发动机试验以便模拟遭遇鸟群，表2中规定了使用鸟的数量和重量。当规定只用1只鸟时，这只鸟应投在发动机核心机流通道上；必要时，应通过合适的试验或分析或两者的组合来确定发动机前迎风表面上的其他关键位置。在表2中规定使用2只或2只以上的鸟时，其中最大的1只鸟应投向发动机核心机流通道上，而次重的1只鸟应投向第一级转子叶片的最关键的暴露位置上，其余的鸟必须均匀地分布在整个发动机的前表面上。 （3）此外，除旋翼航空器发动机外，也必须通过适当的试验或分析或两者的组合来证明，当根据本款适用的试验条件，用表3规定数量和重量的鸟，投向核心机主流道外侧风扇组件的最关键位置，而使整个风扇组件经受吸鸟试验时，发动机应能符合本款的验收准则。 （4） 在中鸟试验期间，如果规定数量的中鸟通过了发动机转子叶片，则不再要求作小鸟吸入试验。 （5） 应进行小鸟吸入试验以便模拟遭遇鸟群。试验时鸟的数量应按在每0.032平方米(49.6平方英寸) 进气道面积或其余数部分使用1只85克(0.187磅) 的鸟计算，但最多不超过16只鸟。在对准这些鸟的打击位置时应考虑到第一级转子叶片上的任何关键打击位置，而其余的鸟应均匀地分布在整个发动机前表面上。 （6） 在按本款中规定条件下进行试验时，吸入小鸟和中鸟不得引起下列的任何情况： (i) 持续的功率或推力损失超过25%； (ii) 在本条( c)(7) 或(c)(8) 规定的要求连续验证期间发动机停车； (iii) 出现本条(b)(3) 定义的各种情况； (iv) 不可接受的发动机操纵特性的降低。 （7） 除旋翼航空器发动机外，应采用下列试验程序： （i） 为模拟遭遇鸟群，从吸入第1只鸟的时刻到吸入最后1只鸟经过的时间应为大约1秒钟； (ii) 吸鸟之后2分钟内，不能移动功率杆； (iii) 随后3分钟，在试验状态的75%； (iv) 随后6分钟，在试验状态的60%； (v) 随后6分钟，在试验状态的40%； (vi) 随后1分钟，在进场慢车位置； (vii) 随后2分钟，在试验状态的75%； (viii) 随后稳定在慢车位置并使发动机停车。规定的持续时间是指，当功率杆在每个状态之间移动的时间不超过10秒时所定义的状态的工作时间。 （8） 对于旋翼航空器发动机，使用下列试验程序 (i) 为模拟遭遇鸟群，从吸入第1只鸟的时刻到吸入最后1只鸟经过的时间应为大约1秒钟； (ii) 随后3分钟，在试验状态的75%； (iii) 随后90秒钟，在下降的飞行慢车位置； (iv) 随后30秒钟，在试验状态的75%； (v) 随后稳定在慢车位置并使发动机停车。规定的持续时间是指，当功率杆在每个状态之间移动的时间不超过10秒时所定义的状态的工作时间。 （9）如果相应的型号审定文件中注明不要求预期在多发旋翼航空器上使用的发动机遵守本条的中鸟吸入部分，则这类发动机可以不遵守本条的中鸟吸入部分的要求。 （10）如果发生按本条(c)(7)(ii)的规定，在不移动功率杆的情况下，在最初的2分钟期间，出现发动机超过任何工作限制的情况，则应确认该超限情况不会导致出现不安全状态。 表2 中鸟群的数量和重量要求 表3 附加的完整性评估 七、增加第33.78条，内容如下 ： 第33.78条 吸雨和吸雹 （a）所有发动机 （1）当航空器在最大高度达4,500米（15,000英尺）的颠簸气流中飞行的典型飞行条件下，发动机在最大连续功率状态下以最大真实空速吸入大冰雹（比重在0.8-0.9）之后，不得引起不可接受的机械损坏或不可接受的功率或推力损失或者要求发动机停车。此时，一半数量的冰雹应随机投向整个进气道正前方的区域，而另一半则应投向进气道正前方的关键区域。应快速连续地吸入冰雹来模拟遭遇冰雹的情况，并且冰雹的数量和尺寸应按以下列方式确定： （i） 对于进气道面积不大于0.064平方米（100平方英寸）的发动机，为1颗25毫米（1英寸）直径的冰雹； （ii）对于进气道面积大于0.064平方米（100平方英寸）的发动机，每0.0968平方米（150平方英寸）的进气道面积或其余数，为1颗25毫米（1英寸）直径和1颗50毫米（2英寸）直径的冰雹。 （2） 除了遵照本条（a）（1）的规定外，但本条（b）的规定除外，每型发动机必须证明当其突然遭遇浓度达到本规定附录B中定义的审定标准的雨和冰雹时，在其整个规定的工作包线范围内仍有可接受的工作能力。发动机可接受的工作能力是指在任何连续3分钟的降雨周期内，和任何连续30秒的降冰雹周期内，发动机不熄火、不降转、不发生持续或不可恢复的喘振或失速、或不失去加速和减速的能力。还必须证明吸入之后没有不可接受的机械损坏，不可接受的功率或推力损失或其他不利的发动机异常情况。 （b） 旋翼航空器发动机 作为对本条（a）（2）规定要求的另一种验证方法仅适用于旋翼航空器涡轮发动机。当吸入的雨在进气道平面上均匀分布、水滴流量与空气流量的总重量比至少为4%时，必须证明每型发动机在吸雨期间和之后，具有满意的工作能力，即发动机不熄火、不降转、不发生持续或不可恢复的喘振或失速、或不失去加速和减速的能力。还必须证明吸雨之后没有不可接受的机械损坏，不可接受的功率损失或其他不利的发动机异常情况。吸雨必须在下列地面静止条件下进行： （1） 在无吸雨条件下在起飞功率状态稳定一正常的时间周期，随后立即 在起飞功率状态突然开始吸雨3分钟，然后 （2） 在快速减速到最小慢车期间持续吸雨，然后 （3） 在审定的最小空中慢车功率状态运转3分钟期间持续吸雨，然后 （4） 在快速加速到起飞功率期间持续吸雨。 （c） 超音速飞机发动机 除了符合本条（a）（1）和（a） （2）的规定外，应仅对超音速飞机发动机进行单独的试验。试验时发动机应以超音速巡航速度吸入不同的3颗冰雹。这些冰雹应投向发动机正面的关键区域，并且吸雹后不能造成不可接受的机械损坏、或不可接受的功率或推力损失或要求发动机停车。试验冰雹的尺寸应根据在10,500米（35,000英尺）时冰雹直径为25毫米（1英寸），到18,000米（60,000英尺）时冰雹直径为6毫米(1/4英寸)的线性关系来确定。所使用的冰雹直径应与所预期的最低超音速巡航高度相对应。另一种替代方法是，在亚音速下吸入三颗较大的冰雹，但这三颗冰雹的动能应与超音速时吸入的冰雹的动能等效。 （d）对于已安装或要求使用防护装置的发动机，如果申请人能证明符合下列条件，则中国民用航空总局可以全部或部分地免除本条（a）、（b）和（c）中关于发动机吸雨和吸雹能力的验证要求： （1）所遭遇的雨和冰雹构成物的尺寸大到不能通过该防护装置。 （2）该防护装置能够承受所遭遇的雨和冰雹构成物的打击。并且 （3） 防护装置阻挡的雨和冰雹构成物，不会阻碍进入发动机的空气流量，至使所造成的损坏、功率或推力损失、或其他对发动机不利的情况超过本条（a）、（b）和（c）中可接受的水平。 八、增加附件B，内容如下： 附件B 合格审定标准大气降雨和冰雹的浓度 为了按照第33.78条（a）（2）的要求进行合格审定，图B1、表B1、表B2、表B3、表B4规定了雨和冰雹的大气浓度和尺寸分布。只要申请人能表明所使用的替代方法没有降低试验的严格程度，在通常通过喷洒液态水模拟降雨以及投掷冰块制造的冰雹模拟降冰雹的情况下，允许使用不同于本规定附录B规定的这些水滴和冰雹的形状、尺寸和尺寸分布，或者允许使用尺寸和形状单一的水滴或冰雹。 图B1雨和冰雹的征兆图表,利用表B1和B2可获得合格审定浓度 表B1 合格审定标准的大气雨浓度 注：在其他高度上雨的水含量的值可以由线性内插的方法确定。 表B2 合格审定标准的大气冰雹浓度 注：在其他高度上的冰雹水含量值可以用线性内插法确定。低于2,230米(7,300英尺)和大于8,840米(29,000英尺)的冰雹征兆可根据线性外插数据获得。 表B3 合格审定标准的大气雨滴尺寸分布 注：雨滴的平均直径为2.66毫米 表B4 合格审定标准的大气冰雹尺寸分布 注：冰雹的平均直径为16毫米 九、§33.1改为第33.1条后，条款修改为： 第33.1条 适用范围 (a) 本规定规定颁发和更改航空发动机型号合格证用的适航标准。 (b) 按照中国民用航空规章《民用航空产品和零部件合格审定规定》（CCAR-21）的规定申请航空发动机型号合格证或申请对该合格证进行更改的法人，必须表明符合本规定中适用的要求，并且必须表明符合中国民用航空规章《涡轮发动机飞机燃油排泄和排气排出物规定》（CCAR-34）。 十、§33.7改为第33.7条后，条款修改为： 第33.7条 发动机额定值和使用限制 (a) 发动机额定值和使用限制由中国民用航空总局认定，并包含在中国民用航空规章《民用航空产品和零部件合格审定规定》（CCAR-21）规定的发动机型号合格证数据单中，其中包括按本条规定的各种适用的使用条件和资料确定的额定值和限制以及为发动机安全使用所必需的任何其他资料。 (b) 对于活塞式发动机，额定值和使用限制的确定与下列因素有关： (1) 下列功率状态值在临界压力高度与海平面压力高度下的功率或扭矩、转速(转/分)、进气压力和时间： (i) 额定最大连续功率(与非增压使用状态或与适用的增压器各种使用状态有关)。 (ii) 额定起飞功率(与非增压使用状态或与适用的增压器各种使用状态有关)。 (2) 燃油牌号或规格； (3) 滑油品级或规格； (4) 下列各项温度: (i) 气缸温度； (ii) 滑油进口温度； (iii) 涡轮增压器的涡轮进气温度。 (5) 下列各项压力： (i) 燃油进口压力； (ii) 主滑油腔的滑油压力。 (6) 附件传动扭矩和悬臂力矩； (7) 部件寿命； (8) 涡轮增压器的涡轮转速(转/分)。 (c) 对于涡轮发动机，额定值和使用限制的确定与下列因素有关： (1) 下列状态的功率、扭矩或推力、转速(转/分)、燃气温度和时间： (i) 额定最大连续功率或推力(加力的)； (ii) 额定最大连续功率或推力(不加力的)； (iii) 额定起飞功率或推力(加力的)； (iv) 额定起飞功率或推力(不加力的)； (v) 额定30分钟一台发动机不工作（OEI）功率； (vi) 额定2 1/2分钟一台发动机不工作（OEI）功率； (vii) 额定连续一台发动机不工作（OEI）功率； (viii) 额定2分钟一台发动机不工作（OEI）功率； (ix) 额定30秒钟一台发动机不工作（OEI）功率； (x) 辅助动力装置(APU)的工作方式。 (2) 燃油牌号或规格； (3) 滑油品级或规格； (4) 液压油规格； (5) 下列各项温度： (i) 在申请人规定部位上的滑油温度； (ii) 超音速发动机进口截面上的进气温度，包括稳态工作时的温度和瞬时超温温度及其允许超温的时间； (iii) 超音速发动机的液压油温度； (iv) 在申请人规定部位上的燃油温度； (v) 申请人如有规定的发动机的外表面温度。 (6) 下列各项压力： (i) 燃油进口压力； (ii) 在申请人规定部位上的滑油压力； (iii) 超音速发动机进口截面上的进气压力，包括稳态工作时的压力和瞬时超压压力及其允许超压的时间； (iv) 液压油压力。 (7) 附件传动的扭矩和悬臂力矩； (8) 部件寿命； (9) 燃油过滤； (10) 滑油过滤； (11) 引气； (12) 每一转子盘和隔圈被批准的起动—停车应力循环次数； (13) 发动机进气畸变； (14) 转子轴的瞬时超转转速(转/分)和超转出现的次数； (15) 燃气的瞬时超温温度和超温出现的次数； (16) 超音速航空器发动机的转子风车转速(转/分)。 十一、§33.29改为第33.29条后，条款修改为： (a) 除非在结构上能防止错接仪表，否则，按航空器适航标准要求的动力装置仪表所设置的每个连接件或者为保证发动机工作符合任何发动机使用限制所必需的每个连接件，都必须作标记，以标明与相应的仪表一致。 (b) 每台涡轮发动机必须为指示转子系统不平衡的显示系统提供接头。 (c) 具有30秒钟一台发动机不工作（OEI）功率额定值和2分钟一台发动机不工作（OEI）功率额定值的旋翼航空器涡轮发动机应采取以下措施： (1) 当发动机处于30秒钟一台发动机不工作（OEI）和2分钟一台发动机不工作（OEI）功率状态及状态开始和该时间间隔结束时，应提示飞行员； (2) 使用可靠的方式确认发动机是在每一额定功率状态运转；并且 (3) 自动记录每次使用的次数和在每一额定功率状态的持续时间。 十二、§33.63改为第33.63条后，条款修改为： 每型发动机的设计和构造必须使发动机在其声明的整个飞行包线和整个转速和功率或推力的工作范围内正常工作，而不应导致因振动而使发动机的任何零部件应力过大，并且也不应导致将过大的振动力传给航空器结构。 十三、§33.67改为第33.67条后，条款修改为： (a) 在按申请人规定的流量和压力对发动机供给燃油的情况下，该发动机必须在本规章规定的各种工作状态下都能正常地工作。不可再调整的每个燃油控制调节装置装于发动机上时必须用锁紧装置固定并且必须是铅封的，否则应是不可达的。所有其他的燃油控制调节装置必须是可达的，并且作标记以指明调节功能，除非该功能是显而易见的。 (b) 在发动机燃油进口与燃油计量装置进口，或与发动机传动的正排量泵进口(两种进口中取距发动机燃油进口较近者)之间，必须设置燃油滤或滤网。此外下列规定适用于本款(b)要求的每个燃油滤或滤网： (1) 必须是便于放泄和清洗，并必须采用易于拆卸的网件或滤芯； (2) 除非滤网或油滤易于拆卸进行放油，而不需设置放油装置，否则必须具有沉淀槽和放油嘴； (3) 除非导管或接头在所有载荷情况下均具有足够的强度裕量，否则，油滤或滤网的重量不能由相连的导管或其入口或出口的接头支承。 (4) 必须规定为防止燃油中外来颗粒进入发动机燃油系统所必需的燃油滤的类型和过滤度。申请人必须表明符合下列要求： (i) 通过规定过滤装置的外来颗粒不会损害发动机燃油系统的功能； (ii) 在27℃(80°F)的含水的初始饱和燃油中每升加进0.2毫升游离水(每加仑含0.025液英两)，并冷却到工作中可能遇到的最危险的结冰条件下，燃油系统在其整个流量和压力范围内能持续工作。然而，这一要求可以通过验证特定的经批准的燃油防冰添加剂的有效性来满足；或者燃油系统带有燃油加热器，它能在最危险结冰条件下将燃油滤或燃油进口处的燃油温度保持在0℃(32°F)以上。 (5) 申请人必须验证在燃油被污染到工作中可能遇到的最大程度的颗粒尺寸和密度时，过滤装置具有保证发动机在其批准的极限内继续运转的能力(与发动机使用限制相对应)。必须验证发动机在这些条件下，按中国民用航空总局可接受的一段时间内工作，这段时间由下列装置开始指示过滤器临近阻塞时算起： (i) 现有的发动机仪表； (ii) 装在发动机燃油系统的附加装置。 (6) 任何滤网或油滤旁路装置的设计与构造，必须通过其适当设置使积聚的污物逸出最少，以确保积聚的污物不致进入旁通油路。 (c) 对于每个流体喷射(除燃油)系统和其控制装置，如果作为发动机的一部分，申请人必须表明喷射流体量是充分可控的。 (d) 具有30秒钟一台发动机不工作（OEI）功率额定值的发动机，必须具有30秒钟一台发动机不工作（OEI）功率的自动可用性和自动控制装置。 十四、§33.77改为第33.77条后，条款修改为： [(a) 备用] [(b) 备用] (c) 在本条（e）的条件下吸冰时不得出现以下情况： (1) 引起持续的功率或推力损失；或 (2) 要求发动机停车。 (d) 对于采用防护装置的发动机，如果能证明符合下列各项要求，则无需验证在本条（e）规定的条件下外来物吸入是否符合本条规定： （1）该外来物的尺寸大到使它不能通过该防护装置； （2）该防护装置将能经受该外来物的撞击； （3）被防护装置阻挡的该外来物或若干外来物不会阻碍空气流入发动机，从而造成数值超过本条（c）所要求的功率或推力减少。 (e) 在下列吸入条件下，必须通过发动机试验证明符合本条（c）的要求： （1）冰的数量应是由于滞后2分钟开启防冰系统而在典型的进气道整流罩和发动机正面积聚的最多数量的冰；或者使用质量和厚度与该发动机的尺寸可比拟的一块冰。 （2）吸冰速度应能模拟被吸入发动机进气道的冰块的速度。 （3）发动机应工作在最大巡航功率或推力状态。 （4）吸冰试验应能模拟在-4?C（25?F）时遇到的最大连续结冰条件。 十五、§33.83改为第33.83条后，条款修改为： (a) 每型发动机必须进行振动测试，以确定可能受机械或空气动力导致激振的部件的振动特性在整个声明的飞行包线范围内是可接受的。发动机测试应该以经验、分析和部件试验适当的结合为基础，并且应至少涉及转子叶片、静子叶片、转子盘、隔圈和转子轴。 (b) 测试应覆盖对应于声明的整个飞行包线环境条件范围内的功率或推力、每个转子系统的物理和换算转速，从最小转速直到允许工作2分钟或更长的额定时间的最大物理转速和换算转速的103％，并直到所有其他允许工作的物理或换算转速的100％，包括超转转速。 如果测试结果表明应力峰值出现在这些要求的物理或换算转速的最大转速处，则应将测试范围充分扩大到足以找到存在的最大应力值，但该转速范围的扩大不必包括比那些转速再增加2％以上的转速。 (c) 应该对下列情况进行评估： (1) 在改变可调静子叶片角度（包括其调节容差）、压气机引气、附件加载、发动机制造商声明的最恶劣的进气道进气流场畸变以及在（各）排气管内最恶劣条件等情况下对振动特性的影响；而且 (2) 在对颤振敏感的系统中，可能导致或影响颤振的气动力学和航空力学因素。 (d) 除本条（e）规定的以外，为在各种工作条件下允许材料的性能变化留出适当的容差后，与本条确定的振动特性有关的振动应力与适当的稳态应力相加后之和，必须小于有关材料的持久极限。对于每一个被评估的零件，必须证明这些应力裕度的适用性是合理的。如果确定某些工作状态或范围需要加以限制，则应该制定使用和安装限制。 (e) 应该通过试验或分析，或参考以往的经验，评估失效情况（例如，但不限于，失去平衡，静子叶片通道局部堵塞或扩大，燃油喷嘴堵塞，不正确的压气机调节变量等等）所引起的激振力对振动特性的影响，并且证明不会产生有害的情况。 (f) 应对可能影响发动机振动特性的每一具体安装构型进行对本条的符合性验证。如果在发动机型号合格审查期间不能完全地查明这些振动影响，应该对评估的方法和证明符合性的方法加以验证，并应在第33.5条要求的安装说明中定义这些方法。 十六、§33.85改为第33.85条后，条款修改为： 第33.85条 校准试验 (a) 每型发动机必须进行为确定第33.87条规定的有关持久试验的发动机功率特性和条件所必需的校准试验。功率特性校准试验的结果是确定在整个转速、压力、温度和高度工作范围内发动机特性的依据。功率额定值以标准大气条件为基准，无供航空器使用的引气，并且只装有发动机正常工作所必需的那些附件。 (b) 进行持久试验的发动机在持久试验后必须进行在海平面条件下的功率检查，必须确定在持久试验期间出现的任何功率特性变化。在持久试验最后阶段取得的测量值可以用于证明对本款要求的符合性。 (c) 在证明对本条的符合性时，除本条（d）允许的情况外，在进行测量前，发动机在每一状态必须是稳定的。 (d) 在发动机有30秒钟一台发动机不工作（OEI）和2分钟一台发动机不工作（OEI）功率额定值的情况下，可以使用第33.87条(f)(1)至(8)规定的适用的持久试验所取得的测量结果，以证明符合本条对这些一台发动机不工作（OEI）额定值的要求。 十七、§33.87改为第33.87条后，条款修改为： 第33.87条 持久试验 (a) 概述 每型发动机必须进行持久试验，它包括总时数至少为150小时的试验，并且，根据发动机型号和预期使用情况，持久试验（凡适用时）应由本条(b)至(g)中规定的系列运转中的某一个运转组成。对于按本条(b)、(c)、(d)、(e)或(g)进行试验的发动机，必须进行25次规定的6小时试验程序，以完成要求的总时数为150小时的试验。对要求有30秒钟一台发动机不工作（OEI）和2分钟一台发动机不工作（OEI）功率额定值的发动机必须按本条(f)进一步试验。试验按下列要求进行： (1) 对于待试的特定发动机，各项运转须按中国民用航空总局认为合适的顺序进行； (2) 除了一般须由手动控制超控自动控制的那些发动机工作状态，或者必须另外规定进行手动控制的某些特定试验运转情况以外，在持久试验期间，发动机必须在属于发动机组成部分的发动机自动控制装置的控制之下。 (3) 除了本条(a)(5)的规定，发动机功率或推力、燃气温度、转子轴的转速，以及如果有限制时，包括发动机外表面的温度，必须至少是被试的特定发动机相应规定值的100%。如果所有参数值不能同时保持在100%的水平，则可以进行若干次试验； (4) 在进行发动机运转时必须使用符合第33.7条(c)规定规格的燃油、润滑油和液压油； (5) 在至少1/5的运转期间，必须使用供发动机和航空器使用的最大引气量。但是，若中国民用航空总局发现在进行这样的运转时，持久试验的有效性没有受到影响，则功率、推力或转子轴转速可以比被试的特定工作状态的相应规定值的100%低； (6) 必须对每一附件传动装置和安装构件加载。在仅供航空器使用的每一个附件上所加的载荷，必须是在额定的最大连续功率或推力和更高的功率输出时，由申请人为发动机传动装置和安装点所规定的极限载荷。在对任何附件传动装置和安装构件加载条件下的持久试验也可以在单独的试验台上进行，但试验的有效性必须使用经过批准的分析方法来证实。 (7) 除了试验时间不超过5分钟和不允许稳定的场合外，在以任何额定功率或推力运转期间，燃气温度和滑油进口温度必须保持在限制温度。至少有一次运转必须在燃油、滑油和液压油的最小限制压力下进行；并且至少有一次运转必须在燃油、滑油和液压油最大限制压力下进行，同时，必要时可以降低油液温度以便允许获得最大压力； (8) 如果转子轴瞬时超转或燃气瞬时超温的出现次数有限制，则本条(b)至(g)所规定的加速次数必须在限制超转或超温的情况下进行。如果出现上述超转或超温的次数没有限制，则所规定的加速次数中有一半必须在限制超转或超温的情况下进行； (9) 下列附加试验要求适用于装在超音速航空器上的每型发动机的型号合格审定： (i) 为了改变推力调定值，功率控制杆必须在不超过1秒的时间内从初始位置推到最终位置，但如果为确保点火必须增加时间，以便将功率控制杆推到用燃油产生加力推力的加力位置的情况除外。 (ii) 在以任何额定加力推力的运转期间，除了试验时间不足以使温度稳定的场合外，液压油温度必须保持在限制温度下。 (iii) 在模拟超音速运转期间，燃油温度和进气温度不得低于限制温度； (iv) 持久试验必须在装有燃料加力装置和主尾喷管、副尾喷管并在使用可调面积喷管的情况下进行。在每次运转期间，按第33.5条(b)规定的方法实施。 (v) 在以最大连续推力和其相应百分比的推力调定值进行运转期间，发动机必须在上述推力调定值的极限进气畸变条件下工作。 (b) 除某些旋翼航空器发动机以外的发动机 除了本条(c)、(d)或(e)中要求额定值的旋翼机发动机外，对于每型发动机，申请人必须进行下列运转： (1) 起飞和慢车 1小时试验，由5分钟额定起飞功率或推力及5分钟慢车功率或推力交替组成。在起飞和慢车状态及其相应的转子转速和燃气温度条件下发出的功率或推力必须通过用功率控制杆按制造者确定的程序加以调定。在任一个运转周期内，申请人可以在录取检查性能数据时，手动控制转子转速、功率或推力。对于具有加大起飞功率额定值，包括提高涡轮前温度、转子转速或轴功率的发动机，在以起飞功率运转的该周期必须在加大功率额定值的情况下进行。对于实质上不会增加工作苛刻程度的具有加大起飞功率额定值的发动机，以加大功率额定值进行运转的次数由中国民用航空总局决定。在每次5分钟周期后更改功率调定值时，必须按本条(b)(5)规定的方式移动功率控制杆。 (2) 额定最大连续和起飞功率和推力 在下列情况下各运转30分钟： (i) 在25次6小时持久试验循环中的15次期间，应在额定最大连续功率或推力下进行运转。 (ii) 在25次6小时持久试验循环中的10次期间，应在额定起飞功率或推力下进行运转。 (3) 额定最大连续功率和推力 应以额定最大连续功率和推力进行1小时30分钟运转。 (4) 递增的巡航功率和推力 在最大连续发动机转速和地面或最小慢车转速之间应至少分成15个大致相同的转速和时间增量，依次在与这15个转速和时间增量相对应的功率控制杆位置连续进行2小时30分钟的试验。对于以恒定转速工作的发动机，可以用改变推力和功率来代替改变转速。如果在地面慢车和最大连续之间任何状态有显著的峰值振动，则可以变更所选择的增量个数，以便使承受峰值振动影响的运转时数增加到不超过递增运转总时数的50%。 (5) 加速和减速运转 30分钟加速和减速运转应由6个循环组成，而每个循环应由慢车功率或推力到额定起飞功率或推力所组成，并且须在起飞功率控制杆位置保持30秒，在慢车功率控制杆位置保持约4 1/2分钟。为符合本款规定，功率控制杆必须在不超过1秒内从一个极端位置推到另一极端位置；但是，如果采用了必须按时间程序把功率控制杆从一个极端位置移动到另一极端位置的不同的调节工作方式，允许使用较长时间的情况除外。但移动功率杆的时间最长不能超过2秒。 (6) 起动 必须进行100次起动试验，其中的25次必须在发动机停车至少2小时后进行。其中必须至少有10次发动机假起动。每次假起动后准备正常起动前，按申请人规定的最短排油时间暂停起动。其中至少有10次正常再起动必须在发动机停车后 15分钟内进行。其余的起动可以在150小时的持久试验完成后进行。 (c) 要求30分钟一台发动机不工作（OEI）功率额定值的旋翼航空器发动机 对于要求30分钟一台发动机不工作（OEI）功率额定值的每型旋翼航空器发动机，申请人必须进行下列一系列试验： (1) 起飞和慢车 1小时试验，由5分钟额定起飞功率及5分钟慢车功率交替组成。在起飞和慢车状态及其相应的转子转速和燃气温度条件下发出的功率必须通过功率控制杆按制造者规定的程序加以确定。在任何一个运转周期内，可以在录取检查性能的数据时，手动控制转子转速和功率和推力。具有加大起飞功率额定值包括增加涡轮进气温度、转子转速或轴功率的发动机，在以额定起飞功率运转期间，必须以加大额定值进行。在每次5分钟试验后变更功率调定值时，必须按本条(c)(5)规定的方式移动功率控制杆。 (2) 额定30分钟一台发动机不工作（OEI）功率 以额定30分钟一台发动机不工作（OEI）功率进行30分钟试验； (3) 额定最大连续功率 以额定最大连续功率和推力运转2小时； (4) 递增的巡航功率 在最大连续发动机转速和地面或最小慢车转速之间应至少分成12个大致相同的转速和时间增量，依次在与这12个转速和时间增量相对应的功率控制杆位置连续进行2小时的试验。对于以恒定转速工作的发动机，可以用改变功率来代替改变转速。如果在地面慢车和最大连续功率之间任何状态有显著的峰值振动，则可以变更所选择的增量个数，以便使承受峰值振动影响的运转时数增加到不超过递增运转总时数的50%。 (5) 加速和减速运转 30分钟加速和减速运转应由6个循环组成，而每个循环应由慢车功率到额定起飞功率所组成，并且须在起飞功率控制杆位置保持30秒，在慢车功率控制杆位置保持约4 1/2分钟。为符合本款规定，功率控制杆必须在不超过1秒内从一个极端位置推到另一极端位置；但是，如果采用了必须按时间程序把功率控制杆从一个极端位置移动到另一极端位置的不同的调节工作方式，允许使用较长时间的情况除外。但移动功率杆的时间最长不能超过2秒。 (6)起动 必须进行100次起动试验，其中的25次必须在发动机停车至少2小时后进行。其中必须至少有10次发动机假起动。每次假起动后准备正常起动前，按申请人规定的最短排油时间暂停起动。其中至少有10次正常再起动必须在发动机停车后 15分钟内进行。其余的起动可以在150小时的持久试验完成后进行。 (d) 要求连续一台发动机不工作（OEI）功率额定值的旋翼航空器发动机 对于要求连续一台发动机不工作（OEI）功率额定值的每型旋翼航空器发动机，申请人必须进行下述一系列试验： (1) 起飞和慢车 1小时试验，由5分钟额定起飞功率及5分钟慢车功率交替组成。在起飞和慢车状态及其相应的转子转速和燃气温度条件下发出的功率和推力必须通过功率控制杆按制造者规定的程序加以确定。在任何一个运转周期内，可以在录取检查性能的数据时，手动控制转子转速和功率。具有加大起飞功率额定值包括增加涡轮进气温度、转子转速或轴功率的发动机，在以额定起飞功率运转期间，必须以加大额定值进行。在每次5分钟试验后变更功率调定值时，必须按本条(c)(5)规定的方式移动功率控制杆。 (2) 额定最大连续功率和起飞功率 在下列情况下各运转30分钟： (i) 在25次6小时持久试验循环中的15次期间，以额定最大连续功率进行运转，以及 (ii) 在25次6小时持久试验循环中的10次期间，以额定起飞功率进行运转。 (3) 额定连续一台发动机不工作（OEI）功率 以额定连续一台发动机不工作（OEI）功率运转1小时。 (4) 额定最大连续功率 以额定最大连续功率运转1小时。 (5) 递增的巡航功率 在最大连续发动机转速和地面或最小慢车转速之间应至少分成12个大致相同的转速和时间增量，依次在与这12个转速和时间增量相对应的功率控制杆位置连续进行2小时的试验。对于以恒定转速工作的发动机，可以用改变功率来代替改变转速。如果在地面慢车和最大连续功率之间任何状态有显著的峰值振动，则可以变更所选择的增量个数，以便使承受峰值振动影响的运转时数增加到不超过递增运转总时数的50%。 (6) 加速和减速运转 30分钟加速和减速运转应由6个循环组成，而每个循环应由慢车功率到额定起飞功率所组成，并且须在起飞功率控制杆位置保持30秒，在慢车功率控制杆位置保持约4 1/2分钟。为符合本款规定，功率控制杆必须在不超过1秒内从一个极端位置推到另一极端位置；但是，如果采用了必须按时间程序把功率控制杆从一个极端位置移动到另一极端位置的不同的调节工作方式，允许使用较长时间的情况除外。移动功率杆的时间最长不能超过2秒。 (7) 起动 必须进行100次起动试验，其中的25次必须在发动机停车至少2小时后进行。其中必须至少有10次发动机假起动。每次假起动后准备正常起动前，按申请人规定的最短排油时间暂停起动。其中至少有10次正常再起动必须在发动机停车后 15分钟内进行。其余的起动可以在150小时的持久试验完成后进行。 (e) 要求2 1/2分钟一台发动机不工作（OEI）功率额定值的旋翼航空器发动机 对于要求2 1/2分钟一台发动机不工作（OEI）功率额定值的旋翼航空器发动机，申请人必须进行以下一系列试验： (1) 起飞， 2 1/2分钟一台发动机不工作（OEI）功率和慢车。 1小时试验，由5分钟额定起飞功率及5分钟慢车功率交替组成。但是，在第3次和第6次起飞功率期间，仅需以额定起飞功率试验2 1/2分钟，余下的2 1/2分钟必须以额定的2 1/2分钟OEI功率进行试验的情况除外。在发动机起飞、2 1/2分钟OEI和慢车状态及其相应的转子转速和燃气温度状态下发出的功率，必须通过使用功率控制杆按制造者确定的程序加以调定。在任一个运转期间，申请人在录取检查性能用的数据时，可以手动控制转子转速和功率。具有加大起飞功率额定值，包括增加涡轮前温度、转子转速或轴功率的发动机，在以额定起飞功率运转期间，必须以加大额定值进行。在每次5分钟试验后或试验期间变更功率调定值时，必须按本条(d)(6)规定的方式移动功率控制杆。 (2) 除了25次在每6小时试验程序中的1次外，以及除了在本条(b)(2)规定的30分钟起飞功率试验周期内的最后5分钟，或本条(c)(2)规定的以30分钟OEI功率进行30分钟试验周期内的最后5分钟，或本条(d)(3)规定的1小时连续OEI功率试验周期内的最后5分钟外，按本条(b)(2)至(b)(6)，或(c)(2)至(c)(6)，或(d)(2)至(d)(7)所要求的试验，在适用时，必须在2 1/2分钟OEI功率状态运转。 （f） 要求30秒钟一台发动机不工作（OEI）和2分钟一台发动机不工作（OEI）功率额定值的旋翼航空器发动机 对于要求30秒钟一台发动机不工作（OEI）和2分钟一台发动机不工作（OEI）功率额定值的旋翼航空器发动机，在完成了本条(b)、(c)、(d)或(e)规定的试验后，申请人可以分解试验后的发动机至能证明符合第33.93条(a)的要求所需要的程度。此试验发动机必须用按本条(b)、(c)、(d)或(e)试验用的相同零部件重新装配，但持续适航性说明文件规定的消耗件除外。然后，申请人必须进行下列试验程序4次，总时数不低于120分钟： (1) 起飞功率 以额定起飞功率进行3分钟运转。 (2) 30秒钟一台发动机不工作（OEI）功率 以额定30秒钟一台发动机不工作（OEI）功率进行30秒钟运转。 (3) 2分钟一台发动机不工作（OEI）功率 以额定2分钟一台发动机不工作（OEI）功率进行2分钟运转。 (4) 30分钟一台发动机不工作（OEI）功率、连续一台发动机不工作（OEI）功率或最大连续功率 以额定30分钟一台发动机不工作（OEI）功率、额定连续一台发动机不工作（OEI）功率或额定最大连续功率（取大者）进行5分钟运转。但是，第一次试验程序期间，该时间周期应该为65分钟的情况除外。 (5) 50％起飞功率 以50％起飞功率进行1分钟运转。 (6) 30秒钟一台发动机不工作（OEI）功率 以额定30秒钟一台发动机不工作（OEI）功率进行30秒钟运转。 (7) 2分钟一台发动机不工作（OEI）功率 以额定2分钟一台发动机不工作（OEI）功率进行2分钟运转。 (8) 慢车 以慢车功率进行1分钟运转。 (g) 超音速航空器发动机 对于用于超音速航空器的每型发动机的型号合格审定，申请人必须进行下列试验： (1) 在海平面环境大气条件下的亚音速试验 必须进行每阶段1小时共30阶段的运转，每阶段运转由下列各项组成： (i) 2次5分钟的额定起飞加力推力，每次接着5分钟的慢车推力； (ii) 1次5分钟的额定起飞推力，接着5分钟的不超过15%额定起飞推力； (iii) 1次10分钟的额定起飞加力推力，接着2分钟的慢车推力。但是，如果额定最大连续加力推力低于额定起飞加力推力，则10分钟周期中的5分钟为额定最大连续加力推力的情况除外； (iv) 6次1分钟的额定起飞加力推力，每次接着2分钟的慢车推力，包括加速和减速的时间在内。 (2) 模拟超音速试验 必须在模拟超音速试验的每次运转前，把亚音速状态所达到的进气温度和压力变换到超音速所达到的温度和压力，随后必须再返回到亚音速状态所达到的温度。必须进行每阶段4小时共计30阶段的运转，每次运转由下列各项组成： (i) 一个以功率控制杆在额定最大连续加力推力位置上所获得的推力进行30分钟运转周期，接着以功率控制杆在90%额定最大连续加力推力位置上所获得的推力进行10分钟运转 。在前5个阶段该运转周期的末尾，空气进气温度必须在瞬时超温的极限条件下进行，但在本条(g)(2)(ii)至(iv)中规定的试验期间不必重复该运转； (ii) 重复进行一次本条(g)(2)(i)规定的运转周期。但是，必须接着以功率控制杆在80%额定最大连续加力推力位置上所获得的推力进行10分钟运转的情况除外； (iii) 重复进行一次本条(g)(2)(i)规定的运转周期。但是，必须接着以功率操纵杆在60%额定最大连续加力推力位置上所获得的推力进行10分钟运转，然后以不超过15%的额定起飞推力运转10分钟的情况除外； (iv) 重复进行本条(g)(2)(i)和(ii)规定的运转各一次； (v) 进行一次30分钟的运转周期，30个阶段中的25个运转阶段以功率控制杆在额定最大连续加力推力位置上所获得的推力进行，并且每阶段运转后接着在慢车推力状态下工作；其余的5个运转阶段以功率控制杆在额定最大连续加力推力位置上所获得的推力试验25分钟，每阶段接着用热燃油以不大于15%的额定起飞推力进行亚音速工作，并加速到额定起飞推力工作5分钟。 (3) 起动 必须进行100次起动试验，其中的25次必须在发动机停车至少2小时后进行。其中必须至少有10次发动机假起动。每次假起动后准备正常起动前，按申请人规定的最短排油时间暂停起动。其中至少有10次正常再起动必须在发动机停车后 15分钟内进行。起动可以在包括持久试验期间的任何时候进行。 十八、§33.88改为第33.88条后，条款修改为： (a) 每型发动机必须在比最大额定功率下的稳态工作限制温度高至少42℃(75oF)的燃气温度下，以最大允许转速运转5分钟。但不包括对应30秒钟一台发动机不工作（OEI） 和2分钟一台发动机不工作（OEI）的转速和燃气温度的最大值。在此运转后，涡轮部件必须在可使用的限制范围内。 (b) 每型要求30秒钟一台发动机不工作（OEI）和2分钟一台发动机不工作（OEI）功率额定值的发动机，在不安装温度限制装置的情况下，必须在超过30秒钟一台发动机不工作（OEI）功率额定值工作限制温度至少42oC（75oF）时，以接通最大功率转速运转5分钟。在此运转后，只要通过中国民用航空总局认为必要的分析或试验证明发动机能保持涡轮部件的完整性，则在涡轮部件上可以有超出该超温条件限制范围的损伤。 (c) 要求30秒钟一台发动机不工作（OEI）和2分钟一台发动机不工作（OEI）功率额定值的每型发动机，在安装温度限制装置的情况下，必须在超过30秒钟一台发动机不工作（OEI）功率额定值工作限制温度至少20oC（35oF）时，以接通最大功率转速运转4分钟。在此运转后，只要通过中国民用航空总局认为必要的分析或试验表明发动机能保持涡轮部件的完整性，则在涡轮部件上可以有超出该超温条件限制范围的损伤。 (d) 对每一试验条件，可以使用单独的试验设备。 十九、§33.91改为第33.91条后，条款修改为： (a) 对于不能按本规定第33.87条规定的持久试验予以充分验证的那些系统，必须进行另外的试验，以确定在所有正常预期的飞行和大气条件下，这些部件能可靠地工作。 (b) 必须确定在航空器安装中要求温度控制措施的那些部件的温度限制，以确保其良好的功能、可靠性和耐久性。 (c) 每个不增压的液压油油箱在受到最大工作温度和34.5千帕（5磅/英寸2）的内部压力时，不得出现失效或泄漏。每个增压的液压油油箱在受到最大工作温度和不低于34.5千帕（5磅/英寸2）的内部压力加上油箱的最大工作压力时，不得出现失效或泄漏。 (d) 对于超音速航空器的发动机型号合格审定，必须确定由于在最高和最低工作温度时可能会发生失效的发动机系统、安全装置及外部附件。并且必须在最高和最低工作温度以及当温度和其他使用条件在最高和最低使用值之间循环时进行试验。 二十、§33.92改为第33.92条后，条款修改为： 如果采用锁定转子装置以阻止发动机持续转动，则发动机必须在以下条件下进行包括该装置工作25次的试验： (a) 发动机必须从额定最大连续推力或功率状态停车；并且 (b) 必须在承受在该状态下持续飞行可能引起的最大扭矩的情况下，按发动机使用说明的规定操作停止和锁定转子的装置，并且 (c) 在25次工作中，每一次转子锁定后，转子必须在这些状态下保持静止5分钟。 二十一、§33.93改为第33.93条后，条款修改为： 第33.93条 分解检查 (a) 在完成本章第33.87条(b)、(c)、(d)、(e)或(g)的持久试验后，每台发动机必须完全分解，并满足下列要求： (1) 不论是否安装在发动机上即可确定其调整位置和功能特性的每个部件，必须使其每个调整位置和功能特性保持在试验开始时确定和记录的限制范围内。 (2) 按第33.4条提交的资料，每个发动机零部件必须符合型号设计并且应仍然可以安装在发动机上继续使用。 (b) 在完成本章第33.87条(f)的持久试验后，每台发动机必须完全分解，并满足下列要求： (1) 不论是否安装在发动机上即可确定其调整位置和功能特性的每个部件，必须使其每个调整位置和功能特性保持在试验开始时确定和记录的限制范围内；并且 (2) 每型发动机可以有超出本条(a)(2)允许的损伤，包括某些不适合于进一步使用的发动机零件或部件。当中国民用航空总局认为必要时，申请人必须通过分析和、或试验，证明发动机以及包括安装节、机匣、轴承座、轴和转子的结构完整性得到了保持；或者 (c) 代替本条(b)的符合性，可以在要求30秒钟一台发动机不工作（OEI）和2分钟一台发动机不工作（OEI）功率额定值的发动机上进行本规定第33.87条(b)、(c)、(d)或(e)规定的持久试验，接着进行第33.87条(f)规定的试验，但中间不进行分解和检查。在完成第33.87条(f)的持久试验后，发动机必须满足本条(a)的要求。 1988年2月9日发布施行的《航空发动机适航标准》（CCAR-33）根据本决定作相应修订，重新发布。 本决定自2002年4月19日起施行。 航空发动机适航规定 A章 总 则 (a) 本规定规定颁发和更改航空发动机型号合格证用的适航标准。 (b) 按照中国民用航空规章《民用航空产品和零部件合格审定规定》（CCAR-21）的规定申请航空发动机型号合格证或申请对该合格证进行更改的法人，必须表明符合本规定中适用的要求，并且必须表明符合中国民用航空规章《涡轮发动机飞机燃油排泄和排气排出物规定》（CCAR-34）。 [2002年4月19日第一次修订] 第33.3条 概述 每一个申请人必须表明该型航空发动机符合本规定中适用的要求。 第33.4条 持续适航文件 申请人必须根据本规定附件A编制中国民用航空总局可接受的持续适航文件。如果有计划保证在交付第一架装有该发动机的航空器之前或者在为装有该发动机的航空器颁发适航证之前完成这些文件，则这些文件在型号合格审定时可以是不完备的。 第33.5条 发动机安装和使用说明手册 每一个申请人必须备有在型号合格证颁发之前可供中国民用航空总局应用，在发动机交付时可供用户应用的经批准的发动机安装和使用说明手册。该说明手册必须至少包括下列内容： (a) 安装说明 (1) 发动机安装构件的位置，将发动机装接到航空器上的方法及安装构件和相关结构的最大允许载荷； (2) 发动机与附件、管件、导线和电缆、钢索、导管及整流罩连接的位置和说明； (3) 包括总体尺寸的发动机轮廓图。 (b) 使用说明 (1) 中国民用航空总局认定的使用限制； (2) 功率或推力的额定值及在非标准大气条件下的修正程序； (3) 在一般和极端环境条件下，对下列情况的荐用程序: (i) 起动； (ii) 地面运转； (iii)飞行中的运转； 第33.7条 发动机额定值和使用限制 (a) 发动机额定值和使用限制由中国民用航空总局认定，并包含在中国民用航空规章《民用航空产品和零部件合格审定规定》（CCAR-21）规定的发动机型号合格证数据单中，其中包括按本条规定的各种适用的使用条件和资料确定的额定值和限制以及为发动机安全使用所必需的任何其他资料。 (b) 对于活塞式发动机，额定值和使用限制的确定与下列因素有关： (1) 下列功率状态值在临界压力高度与海平面压力高度下的功率或扭矩、转速(转/分)、进气压力和时间： (i) 额定最大连续功率(与非增压使用状态或与适用的增压器各种使用状态有关)。 (ii) 额定起飞功率(与非增压使用状态或与适用的增压器各种使用状态有关)。 (2) 燃油牌号或规格； (3) 滑油品级或规格； (4) 下列各项温度: (i) 气缸温度； (ii) 滑油进口温度； (iii) 涡轮增压器的涡轮进气温度。 (5) 下列各项压力： (i) 燃油进口压力； (ii) 主滑油腔的滑油压力。 (6) 附件传动扭矩和悬臂力矩； (7) 部件寿命； (8) 涡轮增压器的涡轮转速(转/分)。 (c) 对于涡轮发动机，额定值和使用限制的确定与下列因素有关： (1) 下列状态的功率、扭矩或推力、转速(转/分)、燃气温度和时间： (i) 额定最大连续功率或推力(加力的)； (ii) 额定最大连续功率或推力(不加力的)； (iii) 额定起飞功率或推力(加力的)； (iv) 额定起飞功率或推力(不加力的)； (v) 额定30分钟一台发动机不工作（OEI）功率； (vi) 额定2 1/2分钟一台发动机不工作（OEI）功率； (vii) 额定连续一台发动机不工作（OEI）功率； (viii) 额定2分钟一台发动机不工作（OEI）功率； (ix) 额定30秒钟一台发动机不工作（OEI）功率； (x) 辅助动力装置(APU)的工作方式。 (2) 燃油牌号或规格； (3) 滑油品级或规格； (4) 液压油规格； (5) 下列各项温度： (i) 在申请人规定部位上的滑油温度； (ii) 超音速发动机进口截面上的进气温度，包括稳态工作时的温度和瞬时超温温度及其允许超温的时间； (iii) 超音速发动机的液压油温度； (iv) 在申请人规定部位上的燃油温度； (v) 申请人如有规定的发动机的外表面温度。 (6) 下列各项压力： (i) 燃油进口压力； (ii) 在申请人规定部位上的滑油压力； (iii) 超音速发动机进口截面上的进气压力，包括稳态工作时的压力和瞬时超压压力及其允许超压的时间； (iv) 液压油压力。 (7) 附件传动的扭矩和悬臂力矩； (8) 部件寿命； (9) 燃油过滤； (10) 滑油过滤； (11) 引气； (12) 每一转子盘和隔圈被批准的起动—停车应力循环次数； (13) 发动机进气畸变； (14) 转子轴的瞬时超转转速(转/分)和超转出现的次数； (15) 燃气的瞬时超温温度和超温出现的次数； (16) 超音速航空器发动机的转子风车转速(转/分)。 [2002年4月19日第一次修订] 第33.8条 发动机功率和推力额定值的选定 (a) 必须由申请人选定所申请的发动机功率和推力额定值。 (b) 选定的每种额定值必须是所有同型号发动机在用来确定此额定值的条件下预计能产生的最低功率或推力。 B章 设计与构造 第33.11条 适用范围 本章规定航空活塞式和涡轮发动机的一般设计与构造要求。 [第33.13条 备用] 第33.14条 起动一停车循环应力(低循环疲劳) 根据中国民用航空总局批准的程序，必须确定使用限制。该使用限制规定发生失效后可能危及航空器安全的每一转子结构件(压气机和涡轮的盘、隔圈、轮毂、轴)起动—停车应力循环的最大允许次数。起动—停车应力循环由飞行循环剖面图或由发动机当量使用图表组成，它包括发动机起动、加速到最大额定功率或推力、减速和停车。对于每次循环，除非表明转子结构件在温度没有稳定的情况下经受了相同的应力范围，否则，在发动机以最大额定功率或推力运转期间及发动机停车后，转子结构件必须达到稳定的温度。 第33.15条 材料 发动机所用材料的适用性和耐久性必须满足下列要求： (a) 建立在经验或试验的基础上； (b) 符合经批准的规范(如工业或军用规范)，保证这些材料具有设计资料中采用的强度和其他性能。 第33.17条 防火 (a) 发动机的设计和构造及所用的材料必须使着火和火焰蔓延的可能性减至最小。此外，涡轮发动机的设计和构造必须使出现导致结构失效、过热或其他危险状态的内部着火的可能性减至最小。 (b) 除本条(c)、(d)和(e)的规定外，存留或输送易燃液体的每一外部管路、接头和其他部件，均必须是耐火的。上述部件必须防护或设置以防止点燃泄漏的易燃液体。 (c) 属于发动机部分并与发动机相连的易燃液体箱和支架必须是防火的或用防火罩防护，任一非防火的零部件被火烧坏后不会引起易燃液体泄漏或溅出则除外，活塞式发动机上容量小于23.7升(25夸脱)的整体湿油池，既不必是防火的，也不需用防火罩防护。 (d) 对于超音速航空器的涡轮发动机的型号合格审定，要求每一个输送或存留易燃液体的外部部件必须是防火的。 (e) 必须用排放和通风的方法防止易燃液体和蒸汽的有害积聚。 第33.19条 耐久性 (a) 发动机的设计与构造必须使得发动机在翻修周期之间不安全状态的发展减至最小。压气机和涡轮转子机匣的设计必须对因转子叶片失效而引起的破坏具有包容性。必须确定由于转子叶片失效，穿透压气机和涡轮转子机匣后的转子叶片碎片的能量水平和轨迹。 (b) 属于发动机型号设计部分的螺旋桨桨距调节系统的每一个部件必须满足中国民用航空规章第35.42条的要求。 第33.21条 发动机冷却 发动机的设计与构造必须在飞机预定工作条件下提供必要的冷却。 第33.23条 发动机的安装构件和结构 (a) 必须规定发动机安装构件和相关的发动机结构的最大允许的限制载荷和极限载荷。 (b) 该发动机安装构件和相关的发动机结构必须能承受下列载荷： (1) 规定的限制载荷并且没有永久变形； (2) 规定的极限载荷并且没有破坏，但可以出现永久变形。 第33.25条 附件连接装置 发动机在附件传动装置和安装构件受载的情况下，必须能正常地运转。每一个发动机附件传动装置和安装构件必须具有密封措施以防止发动机内部的污染或来自发动机内部的不可接受的泄漏。要求用发动机滑油润滑外部传动花键或联轴节的传动装置和安装构件，必须采用密封措施以防止不可接受的滑油流失和防止来自封闭传动连接件腔室外的污染。发动机的设计必须能对发动机运转所需的每个附件进行检查、调整或更换。 第33.27条 涡轮、压气机、风扇和涡轮增压器转子 (a) 涡轮、压气机、风扇和涡轮增压器转子必须具有足够的强度以便能承受本条(c)款规定的试验条件。 (b) 发动机控制装置、系统和仪表的设计和功能必须给予合理的保证，使影响涡轮、压气机、风扇和涡轮增压器转子结构完整性的发动机使用限制在使用中不会超出。 (c) 根据分析或其他可接受的方法确定的每个涡轮、压气机和风扇中经受最关键应力的转子部件(除叶片外)，其中包括发动机或涡轮增压器中的整体鼓筒转子和离心式压气机，必须在下列条件下试验5分钟： (1) 除了本条(c)(2)(iv)的规定外，以其最大工作温度进行； (2) 以下列适用的最高转速进行： (i) 如果在试验台上试验并且转子部件装有叶片或叶片配重块，则以其最大允许转速的120%进行； (ii) 如果试验在发动机上进行，则以其最大允许转速的115%进行； (iii) 如果试验在涡轮增压器上进行，由一特制燃烧室试验台提供炽热燃气驱动，则以其最大允许转速115%进行； (iv) 以120%的某个转速进行，冷转时，转子部件承受的工作应力相当于最高工作温度和最大允许转速导致的应力； (v) 以105%的最高转速进行。此最高转速是发动机典型安装方式中导致最关键的部件或系统失效时的转速； (vi) 在发动机典型安装方式中，任一部件或系统失效并和飞行前例行检查中或正常飞行使用期间一般不予以检测的部件或系统发生的任一故障相组合时，所导致的最高转速。 试验后，在某种超转情况下的每个转子必须在批准的尺寸限制内，并且不得有裂纹。 第33.28条 发动机电气和电子控制系统 依靠电气和电子装置进行正常工作的每一控制系统必须满足下列要求： （a）在第33.5条所要求的发动机安装和使用说明手册中应对控制系统进行说明、并应规定在正常工作和失效状态所控制的可用功率或推力的百分比、以及其他被控制的功能的控制范围； （b）控制系统的设计和构造应能保证由飞机提供的电源或数据的任何失效不应导致功率或推力发生不可接受的变化，或妨碍发动机继续安全运转； （c） 控制系统的设计和构造成应能保证不会由于控制系统电气或电子部件的单个失效或故障，或可能发生的组合失效，而导致不安全状态的发生； （d）在该安装和使用说明手册中应规定环境限制，包括雷击引起的瞬变状态；并且 （e） 所有相关软件的设计和执行应具有防止导致不可接受的功率或推力损失或其他不安全状态的防错功能，并且，软件的设计和实施方法须经中国民用航空总局批准。 [2002年4月19日第一次修订] 第33.29条 仪表连接 (a) 除非在结构上能防止错接仪表，否则，按航空器适航标准要求的动力装置仪表所设置的每个连接件或者为保证发动机工作符合任何发动机使用限制所必需的每个连接件，都必须作标记，以标明与相应的仪表一致。 (b) 每台涡轮发动机必须为指示转子系统不平衡的显示系统提供接头。 (c) 具有30秒钟一台发动机不工作（OEI）功率额定值和2分钟一台发动机不工作（OEI）功率额定值的旋翼航空器涡轮发动机应采取以下措施： (1) 当发动机处于30秒钟一台发动机不工作（OEI）和2分钟一台发动机不工作（OEI）功率状态及状态开始和该时间间隔结束时，应提示飞行员； (2) 使用可靠的方式确认发动机是在每一额定功率状态运转；并且 (3) 自动记录每次使用的次数和在每一额定功率状态的持续时间。 [2002年4月19日第一次修订] C章 设计与构造: 活塞式航空发动机 本章规定活塞式航空发动机附加的设计与构造要求。 第33.33条 振动 发动机的设计与构造必须能使发动机在其曲轴转速和发动机功率的整个正常工作范围内运转，不会由于振动而引起发动机任何零部件的过大应力，并且也不会将过大的振动力传给航空器结构。 第33.35条 燃油和进气系统 (a) 发动机燃油系统的设计与构造必须能在所有飞行和大气条件下的整个发动机工作范围内向气缸提供适当的燃油混合物。 (b) 用于燃烧的空气或油气的混合物所通过的发动机进气通道的设计与构造，必须使冰在这些通道里积聚的危险减至最小。发动机的设计与构造必须允许采用防冰的措施。 (c) 必须规定为防止燃油中外来颗粒进入发动机燃油系统所必需的燃油滤的类型和过滤度。申请人必须表明通过规定的过滤装置的外来颗粒将不会严重地损害发动机燃油系统的功能。 (d) 当装该发动机的航空器在地面静止状态时，在申请人所确定的发动机可能有的所有姿态下，进气系统中，引导油气混合物的每一条通道，都必须是自身可以放泄的，以防止气缸内的液锁。 (e) 对于每个流体喷射(除了燃油)系统和其控制装置，如果作为发动机的一部分，申请人必须表明喷射流体的流量是充分可控的。 第33.37条 点火系统 火花点火发动机必须装有双点火系统，每个气缸至少有两只火花塞，并具有电源分开的两条独立电路；或者装有在飞行中可靠性相当的点火系统。 第33.39条 润滑系统 (a) 发动机的润滑系统的设计与构造，必须使该系统在飞机预期使用中的所有飞行姿态和大气条件下能正常地工作。装有湿油池的发动机，当发动机里的滑油只有最大滑油量的一半时，必须仍能满足这一要求。 (b) 发动机润滑系统的设计与构造必须能安装滑油冷却装置。 (c) 曲轴机匣应与大气相通，以消除曲轴机匣中压力过高时的滑油泄漏。 D章 台架试验: 活塞式航空发动机 本章规定活塞式航空发动机的台架试验和检验。 第33.42条 概述 在本章规定的每项持久试验前，不经装机即可确定其调整位置和功能特性的每个部件，必须确定和记录其调整位置和功能特性。 第33.43条 振动试验 (a) 每型发动机必须进行振动测试，以确定曲轴和螺旋桨轴或其他输出轴在整个曲轴转速和发动机功率范围之内，在稳定状态和瞬时状态下，从慢车转速到所要求的最大连续转速额定值的110%或到所要求的最大起飞转速额定值的103%(两者中取较大者)时的扭转和弯曲振动特性。对于飞机用的发动机，该项测试必须采用与持久试验所用的螺旋桨型号相同的结构形式，对于其他发动机，则采用与持久试验所用的负载装置型号相同的结构型式。 (b) 曲轴和螺旋桨轴或者其他输出轴的扭转和弯曲振动应力，不得超过制轴材料的持久极限应力。如果不能通过测量表明轴的最大应力低于持久极限，则必须测量振动频率和振幅。必须表明峰值振幅所产生的应力低于持久极限；否则，发动机必须在产生峰值振幅的状态下运转，对于钢轴，直到承受住一千万次应力交变而不发生疲劳损坏为止；对于其他材料的轴，直到表明在材料的持久极限应力范围之内不发生疲劳为止。 (c) 必须对每一附件传动装置和安装构件加载，该载荷由仅供航空器使用的每一附件装置所施加，并且是申请人为该传动装置或安装点规定的限制载荷。 (d) 本条(a)规定的振动测试必须在最不利振动效应的那只气缸不点火的情况下重复进行，以确定发动机在此非正常状态下安全使用的条件。但对此振动测试，发动机转速范围仅需从慢车到要求的最大起飞转速，并不必表明符合本条(b)。 第33.45条 校准试验 (a) 每型发动机必须进行为确定第33.49条规定的有关持久试验的发动机功率特性和条件所必需的校准试验。功率特性校准试验的结果构成确定整个使用范围内曲轴转速、进气压力、燃油/空气混合比调定值和高度的发动机特性。功率额定值以标准大气条件下只装有为发动机功能所必需的那些附件时为基准。 (b) 进行持久试验的发动机在持久试验后必须进行海平面状态时的功率检查。必须确定在持久试验期间出现的任何功率特性变化。在持久试验最后阶段取得的测量值可以用于表明符合本款的要求。 第33.47条 爆震试验 每型发动机必须试验，以确定在其预定的整个工作状态范围内，发动机能工作而不会发生爆震。 第33.49条 持久试验 (a) 概述 每型发动机必须进行持久试验，它包括总时数为150小时的试车(除本条(e)(1)(iii)中规定的外)，并根据发动机型号和预期工作情况由本条(b)至(e)中规定的一个适用的试验系列组成。对于待试的特定发动机必须按中国民用航空总局认为合适的程序进行试验。在持久试验期间，该发动机功率和曲轴转速必须保持在额定值的±3%的范围内。在以额定起飞功率和至少35小时额定最大连续功率运转期间，一只气缸必须在不低于限制温度下工作，其余气缸必须在不低于限制温度28℃(50oF)范围内工作，并且滑油进口温度必须保持在限制温度±5.5℃(10oF)范围内。装有螺旋桨轴的发动机必须装螺旋桨做持久试验，并且在本条规定的各种适用运转条件下，该螺旋桨要对发动机加载到其设计能承受的最大拉力载荷。必须对每个附件传动装置和安装构件加载。在以额定起飞功率和额定最大连续功率运转期间，由仅供飞机使用的每种附件所施加的载荷，必须是申请人为发动机传动装置或安装点规定的限制载荷。 (b) 非增压的发动机和采用齿轮传动单速增压器的发动机 对于不采用增压器的发动机和采用齿轮传动单速增压器的发动机，申请人必须作下列试验： (1) 30小时试验，由5分钟起飞转速下额定起飞功率和5分钟最大最经济巡航功率或荐用的最大巡航功率交替组成； (2) 20小时试验，由1 1/2小时最大连续转速下额定最大连续功率的和1/2小时75%的额定最大连续功率及91%最大连续转速交替组成； (3) 20小时试验，由1 1/2小时最大连续转速下额定最大连续功率和1/2小时70%额定最大连续功率及89%最大连续转速交替组成； (4) 20小时试验，由1 1/2小时最大连续转速下额定最大连续功率和1/2小时65%额定最大连续功率及87%最大连续转速交替组成； (5) 20小时试验，由1 1/2小时最大连续转速下额定最大连续功率和1/2小时60%额定最大连续功率及84.5%最大连续转速交替组成； (6) 20小时试验，由1 1/2小时最大连续转速下额定最大连续功率和1/2小时50%额定最大连续功率及79.5%最大连续转速交替组成； (7) 20小时试验，由2 1/2小时最大连续转速下额定最大连续功率和2 1/2小时最大最经济巡航功率或荐用的最大巡航功率交替组成。 (c) 采用齿轮传动双速增压器的发动机 对于采用齿轮传动双速增压器的发动机，申请人必须进行下列试验： (1) 30小时试验，由低传动比的5分钟额定起飞转速下额定起飞功率和5分钟最大最经济巡航功率或最大荐用巡航功率交替组成。如果在高传动比中要求起飞功率额定值，则30小时试验中的15小时试验必须在高传动比下进行，并由5分钟的在起飞临界高度进气压力和起飞转速下获得的功率测量值及5分钟70%高传动比额定最大连续功率和89%高传动比最大连续转速交替组成； (2) 15小时试验，由低传动比的1小时最大连续转速下额定最大连续功率和1/2小时75%额定最大连续功率及91%最大连续转速交替组成； (3) 15小时试验，由低传动比的1小时最大连续转速下额定最大连续功率和1/2小时70%额定最大连续功率及89%最大连续转速交替组成； (4) 30小时试验，以高传动比的最大连续转速下额定最大连续功率进行； (5) 5小时试验，由增压器的每个传动比各5分钟交替组成。该试验的第一个5分钟必须以高传动比下的最大连续转速及在海平面条件下以高传动比的90%的最大连续进气压力获得的测量功率进行。在低传动比下的5分钟交替试验状态必须是在恒定转速下转换到低传动比时所获得的状态； (6) 10小时试验，由低传动比1小时最大连续转速下额定最大连续功率和1小时65%额定最大连续功率及87%最大连续转速交替组成； (7) 10小时试验，由低传动比1小时最大连续转速下额定最大连续功率和1小时60%额定最大连续功率及84.5%最大连续转速交替组成； (8) 10小时试验，由低传动比1小时最大连续转速下额定最大连续功率和1小时50%额定最大连续功率及79.5%最大连续转速交替组成； (9) 20小时试验，由低传动比2小时最大连续转速下额定最大连续功率和2小时最大最经济巡航功率和转速或荐用的最大巡航功率和转速交替组成； (10) 5小时试验，在低传动比下以最大最经济巡航功率和转速或荐用的最大巡航功率和转速 进行；以高传动比运转时，在没有模拟高空试验装置的地方，这些试验可以用在临界高度进气压力或由此规定的百分数压力下获得的测量功率进行，并可将燃油/空气混合比调整到足以抑制爆震的富油混合气。 (d) 直升机发动机 为了适合于在直升机上的使用，每型发动机必须符合中国民用航空规章第29.923条(a)至(j)，或者必须进行以下一系列试验： (1) 35小时试验，由各30分钟的起飞转速下额定起飞功率和最大连续转速下额定最大连续功率交替组成； (2) 25小时试验，由各2 1/2小时的最大连续转速下额定最大连续功率和最大连续转速下70%额定最大连续功率交替组成； (3) 25小时试验，由各2 1/2小时的最大连续转速下额定最大连续功率和80%至90%最大连续转速下70%额定最大连续功率交替组成； (4) 25小时试验，由各2 1/2小时的起飞转速下30%额定最大连续功率和80%至90%最大连续转速下30%额定最大连续功率交替组成； (5) 25小时试验，由各2 1/2小时的起飞转速下80%额定最大连续功率和110%最大连续转速下额定最大连续功率或103%起飞转速下额定起飞功率(两者中取转速较大者)交替组成； (6) 15小时试验，以105%最大连续转速下105%额定最大连续功率进行，或者，如果不能超过105%额定最大连续功率时，则以全油门及在标准海平面汽化器出口压力下的相应转速进行。 (e) 涡轮增压的发动机 对于装有涡轮增压器的发动机，如果申请人表明在模拟高空试验中，发动机和增压器承受的机械载荷和工作温度不低于在实际高空条件下运转时的机械载荷和工作温度，则除了高空试验可以模拟外，按下列规定进行： (1) 对用于飞机的发动机，申请人必须实施本条(b)规定的试验，但下列情况除外： (i) 本条(b)(1)规定的整个试验必须在海平面高度压力下进行； (ii) 本条(b)(2)到(7)中所规定的以额定最大连续功率运转的部分必须在临界高度压力下进行；而以其他功率进行试验部分必须在2,440米(8,000英尺)高度压力下进行； (iii) 在150小时持久试验期间使用的涡轮增压器必须以额定最大连续功率运转时的涡轮进口燃气限制温度和转速增加50小时台架试验，除非在50小时额定最大连续功率运转中保持该限制温度和转速。 (2) 对用于直升机的发动机，申请人必须实施本条(d)款规定的试验，但下列情况除外： (i) 本条(d)(1)中规定的整个试验必须在临界高度压力下进行； (ii) 本条(d)(2)和(3)中规定的以额定最大连续功率进行试验的部分，必须在临界高度压力下进行；而以其他功率进行试验的部分，必须在2,440米(8,000英尺)高度压力下进行； (iii) 本条(d)(4)中规定的整个试验，必须在2,440米(8,000英尺)高度压力下进行； (iv) 本条(d)(5)规定的以80%额定最大连续功率进行试验的部分，必须在2,440米(8,000英尺)高度压力下进行，而以其他功率进行试验的部分，必须在临界高度压力下进行； (v) 本条(d)(6)规定的整个试验，必须在临界高度压力下进行； (vi) 在持久试验期间使用的涡轮增压器，必须以额定最大连续功率运转时的涡轮进口燃气限制温度和转速进行50小时台架试验，除非在50小时额定最大连续功率运转中保持该限制温度和转速。 第33.51条 工作试验 工作试验必须包括中国民用航空总局认为必要的试验，以验证发动机的回火特性、起动、慢车、加速、超转、螺旋桨功能和点火及任何其他工作特性。如果发动机装有多速增压器传动装置，则设计与构造必须允许增压器的运转从低速比转向高速比，并且在增压器高转速比下与额定最大连续功率所具有的进气压力和转速调定值相对应的功率，必须在5秒内达到。 第33.53条 发动机部件试验 (a) 对于不能按第33.49条用持久试验方法进行充分验证的每型发动机，申请人必须进行附加的试验，以确定那些部件在所有正常预期飞行和大气条件下都能可靠地工作。 (b) 必须确定在航空器安装中要求温度控制措施的每一部件的温度限制，以保证其良好的功能、可靠性和耐久性。 第33.55条 分解检查 在完成持久试验后，满足下列要求： (a) 每台发动机必须完全分解。 (b) 不经装机即可确定其调整位置和功能特性的每一部件的调整位置和功能特性必须保持在试验开始时已确定并记录的限制范围内。 (c) 按照第33.4条提交的资料，发动机每个部件必须符合型号设计要求，并且适宜于装在发动机上继续工作。 第33.57条 台架试验的一般实施 (a) 在台架试验时，申请人可用同一设计和结构的几台发动机分别进行振动、校准、爆震、持久和工作试验。如果用一台发动机单独进行持久试验，则该发动机在开始持久试验之前，必须经过校准检查。 (b) 申请人根据符合本规定第33.4条要求提交的维修和维护说明书，可以对在台架试验期间的发动机进行维护和小修。如果这类维护频次过高，或由于发动机故障停车次数过多，或在台架试车期间或分解检查的结果认为有必要大修或更换零件的话，则发动机或其零部件可能进行中国民用航空总局认为必要的任何附加试验。 (c) 每个申请人必须提供所有试验条件，包括设备和胜任的人员，以实施台架试验。 E章 设计与构造： 第33.61条 适用范围 本章规定航空涡轮发动机附加的设计与构造要求。 第33.62条 应力分析 必须对每型涡轮发动机进行应力分析，表明每个涡轮发动机转子、隔圈和转子轴的设计安全裕度。 第33.63条 振动 每型发动机的设计和构造必须使发动机在其声明的整个飞行包线和整个转速和功率或推力的工作范围内正常工作，而不应导致因振动而使发动机的任何零部件应力过大，并且也不应导致将过大的振动力传给航空器结构。 [2002年4月19日第一次修订] 第33.65条 喘振和失速特性 发动机按第33.5条(b)规定的使用说明运转时，即在发动机工作包线内的任何一点上，起动、功率或推力的变化、功率的增大或推力的加力，极限的进气畸变或进气温度，不得引起喘振或失速达到出现熄火、结构失效、超温或发动机功率或推力不能恢复的程度。 第33.66条 引气系统 在第33.7条(c)(11)中规定的极限引气状态的所有条件下，发动机必须提供引气而不会对发动机产生除推力或功率输出降低外的不利影响。如果能控制发动机防冰的引气，则必须设置指示发动机防冰系统功能的装置。 第33.67条 燃油系统 (a) 在按申请人规定的流量和压力对发动机供给燃油的情况下，该发动机必须在本规定规定的各种工作状态下都能正常地工作。不可再调整的每个燃油控制调节装置装于发动机上时必须用锁紧装置固定并且必须是铅封的，否则应是不可达的。所有其他的燃油控制调节装置必须是可达的，并且作标记以指明调节功能，除非该功能是显而易见的。 (b) 在发动机燃油进口与燃油计量装置进口，或与发动机传动的正排量泵进口(两种进口中取距发动机燃油进口较近者)之间，必须设置燃油滤或滤网。此外下列规定适用于本款(b)要求的每个燃油滤或滤网： (1) 必须是便于放泄和清洗，并必须采用易于拆卸的网件或滤芯； (2) 除非滤网或油滤易于拆卸进行放油，而不需设置放油装置，否则必须具有沉淀槽和放油嘴； (3) 除非导管或接头在所有载荷情况下均具有足够的强度裕量，否则，油滤或滤网的重量不能由相连的导管或其入口或出口的接头支承。 (4) 必须规定为防止燃油中外来颗粒进入发动机燃油系统所必需的燃油滤的类型和过滤度。申请人必须表明符合下列要求： (i) 通过规定过滤装置的外来颗粒不会损害发动机燃油系统的功能； (ii) 在27℃(80°F)的含水的初始饱和燃油中每升加进0.2毫升游离水(每加仑含0.025液英两)，并冷却到工作中可能遇到的最危险的结冰条件下，燃油系统在其整个流量和压力范围内能持续工作。然而，这一要求可以通过验证特定的经批准的燃油防冰添加剂的有效性来满足；或者燃油系统带有燃油加热器，它能在最危险结冰条件下将燃油滤或燃油进口处的燃油温度保持在0℃(32°F)以上。 (5) 申请人必须验证在燃油被污染到工作中可能遇到的最大程度的颗粒尺寸和密度时，过滤装置具有保证发动机在其批准的极限内继续运转的能力(与发动机使用限制相对应)。必须验证发动机在这些条件下，按中国民用航空总局可接受的一段时间内工作，这段时间由下列装置开始指示过滤器临近阻塞时算起： (i) 现有的发动机仪表； (ii) 装在发动机燃油系统的附加装置。 (6) 任何滤网或油滤旁路装置的设计与构造，必须通过其适当设置使积聚的污物逸出最少，以确保积聚的污物不致进入旁通油路。 (c) 对于每个流体喷射(除燃油)系统和其控制装置，如果作为发动机的一部分，申请人必须表明喷射流体量是充分可控的。 (d) 具有30秒钟一台发动机不工作（OEI）功率额定值的发动机，必须具有30秒钟一台发动机不工作（OEI）功率的自动可用性和自动控制装置。 [2002年4月19日第一次修订] 第33.68条 进气系统的结冰 在所有防冰系统工作时，每型发动机必须满足下列要求： (a) 在中国民用航空规章第25部附件C中规定的连续最大或间断最大结冰状态下，发动机在其整个飞行功率范围(包括慢车)内的工作中，在发动机部件上不应出现影响发动机工作或引起功率或推力严重损失的结冰情况。 (b) 在临界状态进行引气防冰时，地面慢车30分钟，不出现不利影响，此时大气的温度在－9°～－１℃之间(15°～30°F之间)，每立方米含液态水不少于0.3克并且以平均有效直径不小于20微米的水珠形式存在，接着发动机以起飞功率或推力进行短暂的运转。在30分钟慢车运转期间，该发动机可以以中国民用航空总局接受的方式周期性地加速运转到中等功率或推力调定值。 第33.69条 点火系统 每型发动机必须安装有地面和飞行中起动发动机的点火系统。除了燃油加力燃烧系统只要求一个点火器外，电点火系统必须至少有二个点火器和二条独立的次级电路。 第33.71条 润滑系统 (a) 概述 每一润滑系统在航空器预期使用的飞行姿态和大气条件下，必须能正常地工作。 (b) 滑油滤网或滑油滤 必须有一个供发动机所有滑油通过的滤网或油滤，此外还应满足下列要求： (1) 本款要求的具有旁路的滑油滤网或滑油滤，其构造和安装必须使得在该滤网或油滤元件完全堵塞的情况下，滑油仍能以正常的流量流经系统的其余部分； (2) 必须规定为防止滑油中外来颗粒进入发动机滑油系统所必需的滑油滤类型和过滤度。申请人必须表明通过规定的过滤装置的外来颗粒将不会损害发动机滑油系统的功能； (3) 当滑油污染程度大于本条(b)(2)的规定时(就颗粒的尺寸和密度而言)，本款要求的每个滤网或油滤必须具有保证发动机滑油系统功能不受损害的容量(就确定的发动机使用限制而言)； (4) 除了滑油箱出口的滤网或油滤，对于本款要求的每个滤网或油滤，必须具有在污染达到本条(b)(3)规定的容量之前能予以指示的装置； (5) 任何油滤旁路装置的设计与构造，必须通过其适当设置使积聚的污物逸出最少，以确保积聚的污物不致进入旁通油路； (6) 除了滑油箱出口或回油泵的滤网或油滤外，本款规定的没有旁路的每个滤网或油滤，必须具有一报警器连接装置，以便在滤网的污染达到本条(b)(3)确定的容量之前警告驾驶员； (7) 本款要求的每个滤网或油滤必须便于放泄和清洗。 (c) 滑油箱 (1) 每个滑油箱必须具有不小于油箱容量10%的膨胀空间； (2) 必须避免因疏忽而注满滑油箱膨胀空间的可能性； (3) 每个能存留一定数量滑油的凹型滑油箱加油接头，必须具有安装放油的装置； (4) 每个滑油箱盖必须有滑油密封件； (5) 每个滑油箱加油口应标上“滑油”字样； (6) 每个滑油箱必须在膨胀空间的顶部通气，通气口的布置应使可能冻结并阻塞管道的冷凝水蒸汽不能在任何部位积聚； (7) 必须有防止任何可能防碍滑油在系统中流通的物体进入滑油箱或任何滑油箱出口的装置； (8) 除非滑油系统的外部(包括滑油箱支架)是防火的，否则，在每个滑油箱出口必须有一个切断阀; (9) 每个不增压的滑油箱在受到最大工作温度和34.5千帕(0.35公斤/厘米2；5磅/英寸2)的内部压力时不得发生泄漏；而每个增压的滑油箱在受到最大工作温度和不低于34.5千帕(0.35公斤/厘米2；5磅/英寸2)的内部压力加上该油箱的最大工作压力时不得发生泄漏； (10) 漏出或溢出的滑油不得在油箱和发动机其他零部件之间积聚； (11) 每个滑油箱必须有滑油量指示器或相应的装置； (12) 如果螺旋桨顺桨系统使用发动机滑油，则应满足下列要求： (i) 如果不是油箱本身的失效而是由于润滑系统任一部分的失效使滑油供给量枯竭，则滑油箱必须具有一种能截留一定量滑油的装置； (ii) 被截留的滑油量必须足以完成顺桨工作，并且必须仅供顺桨泵使用； (iii) 必须设有用以防止油泥或其他外来物影响螺旋桨顺桨系统的安全工作的装置。 (d) 滑油放油装置 必须配备一个(或多个)放油嘴，以使滑油系统能安全放泄，每个放油装置必须满足下列要求： (1) 是可达的； (2) 有手动或自动装置确保锁定在关闭位置。 (e) 滑油散热器 每个滑油散热器必须能承受在台架试验中产生的任何振动、惯性和滑油压力载荷而不出现失效。 第33.72条 液压作动系统 在发动机所有预期的工作状态下，每个液压作动系统必须能正常工作。每个油滤或滤网必须便于维修并且每个油箱必须符合本规定第33.71的设计准则。 第33.73条 功率或推力响应 发动机的设计与构造必须满足下列要求： (a) 当功率控制杆在不超过1秒内从最小位置推到最大位置时，在航空器所允许的最大引气和功率提取状态下，从最小功率或推力增大到额定起飞功率或推力，不会出现发动机超温、喘振、失速或其他的有害因素，除非工作方式要求不同的控制程序，则中国民用航空总局可以允许增加额外的时间。 (b) 在不超过5秒时间内，保证从固定最小飞行慢车功率控制杆位置的功率或推力(如无该位置，从不超过15%的额定起飞功率或推力位置)增加至95%额定起飞功率或推力。该5秒种的功率或推力响应必须在仅使用发动机运转所必需的引气和附件载荷的稳定静态下产生。该起飞额定值由申请人规定并且不需包括加力推力值。 第33.74条 持续转动 由于飞行中的任何原因使发动机停车，如果发动机的任何主转动系统仍持续转动并且没有提供阻止持续转动的装置，那么在最长的飞行周期内和在预期该发动机不工作的飞行条件下，任何持续的转动不得导致第33.75条(a)至(c)所描述的任何情况。 [2002年4月19日第一次修订] 第33.75条 安全分析 必须用分析的方法表明，任何可能的发动机故障或单一或多重失效，或任何可能的不正常操纵，不会引起发动机出现下列情况之一： (a) 着火。 (b) 破裂(危险碎片穿透发动机机匣飞出)。 (c) 产生的载荷大于第33.23条(a)中规定的极限载荷。 (d) 失去停车能力。 第33.76条 吸鸟 （a）概述 为符合本条（b）、（c）的要求，应遵照下列规定： （1）吸鸟试验应在吸鸟前的试验天气环境条件下，发动机稳定在不小于100%的起飞功率或推力的状态下进行。另外，符合性的验证必须考虑在海平面最热天气的起飞条件下最差的发动机能够达到最大额定起飞功率或推力的运转情况。 （2）应由申请人来确定在本条中用来决定鸟的数量和重量的发动机进气道喉道面积，并且将其确认为第33.5条所要求的安装说明中的一个限制。 （3）必须对可能进入进气道的单只大鸟和单只最大的中鸟对发动机前部的撞击进行评估。必须证明，当按本条（b）或（c）的规定的条件（如适用）撞击相关部件时，不会影响发动机，使之达到不符合本条（b）（3）和（c）（6）要求的程度。 （4） 对于采用进气道防护装置的发动机，本条的符合性验证应在该防护 装置起作用的情况下进行。发动机的批准文件上应注明对这些要求的符合性验证是在防护装置起作用的情况下进行的。 （5）按本条（b）和（c）的要求进行吸鸟试验时，可用中国民用航空总局可接受的物体代替鸟。 （6）如果本条中各项要求的符合性未被验证，在发动机的型号审定文件中应说明该发动机应仅限于安装在不可能发生鸟撞击发动机,或者发动机不会吸入鸟,或者鸟不会对进入发动机的气流产生不利限制的航空器上。 （b） 大鸟 为符合大鸟吸入的要求，应遵照下列规定： （1） 大鸟的吸入试验应使用表1规定重量的1只鸟。该鸟应投向第一级旋转叶片最关键的暴露位置。对于安装在固定翼飞机上的发动机，吸入鸟的速度应为370公里/小时（200节）；对于安装在旋翼航空器上的发动机，吸入鸟的速度应为旋翼航空器正常飞行时的最大的空速。 （2）在大鸟吸入后的15秒内不允许移动功率杆。 （3）在本条规定的条件下进行单只大鸟的吸鸟试验时，不得导致发动机出现下列情况之一： （i）着火； （ii）危险的碎片穿透发动机机匣飞出； （iii）产生的载荷大于第33.23条(a)中规定的极限载荷； （iv）失去停车能力。 （4）对本款中大鸟吸入要求的符合性验证也可以通过验证第33.94条(a)中在叶片包容性和转子不平衡性方面的各项要求比本条的各项要求更为严格来证明。 表1 大鸟的重量要求 （c） 中鸟和小鸟 为符合中鸟和小鸟吸入的要求，应遵照下列规定： （1） 应采用中国民用航空总局可接受的分析方法或部件试验或是两者的组合，来确定影响功率损失和造成损坏的关键吸鸟参数。关键吸鸟参数应包括，但不限于，鸟速、关键目标位置和第一级转子转速的影响。吸鸟临界速度应反映从地面到地面上460米（1500英尺）的正常飞行高度所使用的空速范围内的最严酷条件，但不应小于飞机的V1最小速度。 （2） 应进行吸中鸟的发动机试验以便模拟遭遇鸟群，表2中规定了使用鸟的数量和重量。当规定只用1只鸟时，这只鸟应投在发动机核心机流通道上；必要时，应通过合适的试验或分析或两者的组合来确定发动机前迎风表面上的其他关键位置。在表2中规定使用2只或2只以上的鸟时，其中最大的1只鸟应投向发动机核心机流通道上，而次重的1只鸟应投向第一级转子叶片的最关键的暴露位置上，其余的鸟必须均匀地分布在整个发动机的前表面上。 （3）此外，除旋翼航空器发动机外，也必须通过适当的试验或分析或两者的组合来证明，当根据本款适用的试验条件，用表3规定数量和重量的鸟，投向核心机主流道外侧风扇组件的最关键位置，而使整个风扇组件经受吸鸟试验时，发动机应能符合本款的验收准则。 （4） 在中鸟试验期间，如果规定数量的中鸟通过了发动机转子叶片，则不再要求作小鸟吸入试验。 （5） 应进行小鸟吸入试验以便模拟遭遇鸟群。试验时鸟的数量应按在每0.032平方米(49.6平方英寸) 进气道面积或其余数部分使用1只85克(0.187磅) 的鸟计算，但最多不超过16只鸟。在对准这些鸟的打击位置时应考虑到第一级转子叶片上的任何关键打击位置，而其余的鸟应均匀地分布在整个发动机前表面上。 （6） 在按本款中规定条件下进行试验时，吸入小鸟和中鸟不得引起下列的任何情况： (i) 持续的功率或推力损失超过25%； (ii) 在本条( c)(7) 或(c)(8) 规定的要求连续验证期间发动机停车； (iii) 出现本条(b)(3) 定义的各种情况； (iv) 不可接受的发动机操纵特性的降低。 （7） 除旋翼航空器发动机外，应采用下列试验程序： （i） 为模拟遭遇鸟群，从吸入第1只鸟的时刻到吸入最后1只鸟经过的时间应为大约1秒钟； (ii) 吸鸟之后2分钟内，不能移动功率杆； (iii) 随后3分钟，在试验状态的75%； (iv) 随后6分钟，在试验状态的60%； (v) 随后6分钟，在试验状态的40%； (vi) 随后1分钟，在进场慢车位置； (vii) 随后2分钟，在试验状态的75%； (viii) 随后稳定在慢车位置并使发动机停车。规定的持续时间是指，当功率杆在每个状态之间移动的时间不超过10秒时所定义的状态的工作时间。 （8）对于旋翼航空器发动机，使用下列试验程序 (i) 为模拟遭遇鸟群，从吸入第1只鸟的时刻到吸入最后1只鸟经过的时间应为大约1秒钟； (ii) 随后3分钟，在试验状态的75%； (iii) 随后90秒钟，在下降的飞行慢车位置； (iv) 随后30秒钟，在试验状态的75%； (v) 随后稳定在慢车位置并使发动机停车。规定的持续时间是指，当功率杆在每个状态之间移动的时间不超过10秒时所定义的状态的工作时间。 （9）如果相应的型号审定文件中注明不要求预期在多发旋翼航空器上使用的发动机遵守本条的中鸟吸入部分，则这类发动机可以不遵守本条的中鸟吸入部分的要求。 （10）如果发生按本条(c)(7)(ii)的规定，在不移动功率杆的情况下，在最初的2分钟期间，出现发动机超过任何工作限制的情况，则应确认该超限情况不会导致出现不安全状态。 表2 中鸟群的数量和重量要求 表3 附加的完整性评估 [2002年4月19日第一次修订] 第33.77条 外物吸入—冰 [(a) 备用] [(b) 备用］ (c) 在本条（e）的条件下吸冰时不得出现以下情况： (1) 引起持续的功率或推力损失；或 (2) 要求发动机停车。 (d) 对于采用防护装置的发动机，如果能证明符合下列各项要求，则无需验证在本条（e）规定的条件下外来物吸入是否符合本条规定： （1）该外来物的尺寸大到使它不能通过该防护装置； （2）该防护装置将能经受该外来物的撞击； （3）被防护装置阻挡的该外来物或若干外来物不会阻碍空气流入发动机，从而造成数值超过本条（c）所要求的功率或推力减少。 (e) 在下列吸入条件下，必须通过发动机试验证明符合本条（c）款的要求： （1）冰的数量应是由于滞后2分钟开启防冰系统而在典型的进气道整流罩和发动机正面积聚的最多数量的冰；或者使用质量和厚度与该发动机的尺寸可比拟的一块冰。 （2）吸冰速度应能模拟被吸入发动机进气道的冰块的速度。 （3）发动机应工作在最大巡航功率或推力状态。 （4）吸冰试验应能模拟在-4?C（25?F）时遇到的最大连续结冰条件。 [2002年4月19日第一次修订] 第33.78条 吸雨和吸雹 （a）所有发动机 （1）当航空器在最大高度达4,500米（15,000英尺）的颠簸气流中飞行的典型飞行条件下，发动机在最大连续功率状态下以最大真实空速吸入大冰雹（比重在0.8—0.9）之后，不得引起不可接受的机械损坏或不可接受的功率或推力损失或者要求发动机停车。此时，一半数量的冰雹应随机投向整个进气道正前方的区域，而另一半则应投向进气道正前方的关键区域。应快速连续地吸入冰雹来模拟遭遇冰雹的情况，并且冰雹的数量和尺寸应按以下列方式确定： （i） 对于进气道面积不大于0.064平方米（100平方英寸）的发动机，为1颗25毫米（1英寸）直径的冰雹； （ii）对于进气道面积大于0.064平方米（100平方英寸）的发动机，每0.0968平方米（150平方英寸）的进气道面积或其余数，为1颗25毫米（1英寸）直径和1颗50毫米（2英寸）直径的冰雹。 （2） 除了遵照本条（a）（1）的规定外，但本条（b）的规定除外，每型发动机必须证明当其突然遭遇浓度达到本规定附录B中定义的审定标准的雨和冰雹时，在其整个规定的工作包线范围内仍有可接受的工作能力。发动机可接受的工作能力是指在任何连续3分钟的降雨周期内，和任何连续30秒的降冰雹周期内，发动机不熄火、不降转、不发生持续或不可恢复的喘振或失速、或不失去加速和减速的能力。还必须证明吸入之后没有不可接受的机械损坏，不可接受的功率或推力损失或其他不利的发动机异常情况。 （b） 旋翼航空器发动机 作为对本条（a）（2）规定要求的另一种验证方法仅适用于旋翼航空器涡轮发动机。当吸入的雨在进气道平面上均匀分布、水滴流量与空气流量的总重量比至少为4%时，必须证明每型发动机在吸雨期间和之后，具有满意的工作能力，即发动机不熄火、不降转、不发生持续或不可恢复的喘振或失速、或不失去加速和减速的能力。还必须证明吸雨之后没有不可接受的机械损坏，不可接受的功率损失或其他不利的发动机异常情况。吸雨必须在下列地面静止条件下进行： （1） 在无吸雨条件下在起飞功率状态稳定一正常的时间周期，随后立即 在起飞功率状态突然开始吸雨3分钟，然后 （5） 在快速减速到最小慢车期间持续吸雨，然后 （6） 在审定的最小空中慢车功率状态运转3分钟期间持续吸雨，然后 （7） 在快速加速到起飞功率期间持续吸雨。 （c） 超音速飞机发动机 除了符合本条（a）（1）和（a）（2）款的规定外，应仅对超音速飞机发动机进行单独的试验。试验时发动机应以超音速巡航速度吸入不同的3颗冰雹。这些冰雹应投向发动机正面的关键区域，并且吸雹后不能造成不可接受的机械损坏、或不可接受的功率或推力损失或要求发动机停车。试验冰雹的尺寸应根据在10,500米（35,000英尺）时冰雹直径为25毫米（1英寸），到18,000米（60,000英尺）时冰雹直径为6毫米(1/4英寸)的线性关系来确定。所使用的冰雹直径应与所预期的最低超音速巡航高度相对应。另一种替代方法是，在亚音速下吸入三颗较大的冰雹，但这三颗冰雹的动能应与超音速时吸入的冰雹的动能等效。 （d）对于已安装或要求使用防护装置的发动机，如果申请人能证明符合下列条件，则中国民用航空总局可以全部或部分地免除本条（a）、（b）和（c）中关于发动机吸雨和吸雹能力的验证要求： （1）所遭遇的雨和冰雹构成物的尺寸大到不能通过该防护装置。 （2）该防护装置能够承受所遭遇的雨和冰雹构成物的打击。并且 （3） 防护装置阻挡的雨和冰雹构成物，不会阻碍进入发动机的空气流量，至使所造成的损坏、功率或推力损失、或其他对发动机不利的情况超过本条（a）、（b）和（c）中可接受的水平。 [2002年4月19日第一次修订] 第33.79条 燃烧燃料加力装置 包括喷口的每个燃烧燃料加力装置，必须满足下列规定： (a) 设有燃烧燃料加力装置的切断装置； (b) 允许开—关交替进行； (c) 在预期的工作范围内可控制； (d) 除了加力装置提供的推力外，加力装置的失效或故障不能引起发动机推力损失； (e) 如果发动机转子转速下降到加力装置预期工作的最低转速以下时，应设有与发动机其他控制机构协调工作并自动切断提供加力装置燃料的控制机构。 F章 台架试验：航空涡轮发动机 本章规定涡轮发动机的台架试验和检验。 第33.82条 概述 在本章规定的每项持久试验前，必须确定和记录不经装机即可确定其调节器调整位置和功能特性的每个部件的调节器调整位置和功能特性。 第33.83条 振动试验 (a) 每型发动机必须进行振动测试，以确定可能受机械或空气动力导致激振的部件的振动特性在整个声明的飞行包线范围内是可接受的。发动机测试应该以经验、分析和部件试验适当的结合为基础，并且应至少涉及转子叶片、静子叶片、转子盘、隔圈和转子轴。 (b) 测试应覆盖对应于声明的整个飞行包线环境条件范围内的功率或推力、每个转子系统的物理和换算转速，从最小转速直到允许工作2分钟或更长的额定时间的最大物理转速和换算转速的103％，并直到所有其他允许工作的物理或换算转速的100％，包括超转转速。 如果测试结果表明应力峰值出现在这些要求的物理或换算转速的最大转速处，则应将测试范围充分扩大到足以找到存在的最大应力值，但该转速范围的扩大不必包括比那些转速再增加2％以上的转速。 (c) 应该对下列情况进行评估： (1) 在改变可调静子叶片角度（包括其调节容差）、压气机引气、附件加载、发动机制造商声明的最恶劣的进气道进气流场畸变以及在（各）排气管内最恶劣条件等情况下对振动特性的影响；而且 (2) 在对颤振敏感的系统中，可能导致或影响颤振的气动力学和航空力学因素。 (d) 除本条（e）规定的以外，为在各种工作条件下允许材料的性能变化留出适当的容差后，与本条确定的振动特性有关的振动应力与适当的稳态应力相加后之和，必须小于有关材料的持久极限。对于每一个被评估的零件，必须证明这些应力裕度的适用性是合理的。如果确定某些工作状态或范围需要加以限制，则应该制定使用和安装限制。 (e) 应该通过试验或分析，或参考以往的经验，评估失效情况（例如，但不限于，失去平衡，静子叶片通道局部堵塞或扩大，燃油喷嘴堵塞，不正确的压气机调节变量等等）所引起的激振力对振动特性的影响，并且证明不会产生有害的情况。 (f) 应对可能影响发动机振动特性的每一具体安装构型进行对本条的符合性验证。如果在发动机型号合格审查期间不能完全地查明这些振动影响，应该对评估的方法和证明符合性的方法加以验证，并应在第33.5条要求的安装说明中定义这些方法。 [2002年4月19日第一次修订] 第33.85条 校准试验 (a) 每型发动机必须进行为确定第33.87条规定的有关持久试验的发动机功率特性和条件所必需的校准试验。功率特性校准试验的结果是确定在整个转速、压力、温度和高度工作范围内发动机特性的依据。功率额定值以标准大气条件为基准，无供航空器使用的引气，并且只装有发动机正常工作所必需的那些附件。 (b) 进行持久试验的发动机在持久试验后必须进行在海平面条件下的功率检查，必须确定在持久试验期间出现的任何功率特性变化。在持久试验最后阶段取得的测量值可以用于证明对本款要求的符合性。 (c) 在证明对本条的符合性时，除本条（d）允许的情况外，在进行测量前，发动机在每一状态必须是稳定的。 (d) 在发动机有30秒钟一台发动机不工作（OEI）和2分钟一台发动机不工作（OEI）功率额定值的情况下，可以使用第33.87(f)(1)至(8)规定的适用的持久试验所取得的测量结果，以证明符合本条对这些一台发动机不工作（OEI）额定值的要求。 [2002年4月19日第一次修订] 第33.87条 持久试验 (a) 概述 每型发动机必须进行持久试验，它包括总时数至少为150小时的试验，并且，根据发动机型号和预期使用情况，持久试验（凡适用时）应由本条(b)至(g)中规定的系列运转中的某一个运转组成。对于按本条(b)、(c)、(d)、(e)或(g)进行试验的发动机，必须进行25次规定的6小时试验程序，以完成要求的总时数为150小时的试验。对要求有30秒钟一台发动机不工作（OEI）和2分钟一台发动机不工作（OEI）功率额定值的发动机必须按本条(f)进一步试验。试验按下列要求进行： (1) 对于待试的特定发动机，各项运转须按中国民用航空总局认为合适的顺序进行； (2) 除了一般须由手动控制超控自动控制的那些发动机工作状态，或者必须另外规定进行手动控制的某些特定试验运转情况以外，在持久试验期间，发动机必须在属于发动机组成部分的发动机自动控制装置的控制之下。 (3) 除了本条(a)(5)的规定，发动机功率或推力、燃气温度、转子轴的转速，以及如果有限制时，包括发动机外表面的温度，必须至少是被试的特定发动机相应规定值的100%。如果所有参数值不能同时保持在100%的水平，则可以进行若干次试验； (4) 在进行发动机运转时必须使用符合第33.7条(c)规定规格的燃油、润滑油和液压油； (5) 在至少1/5的运转期间，必须使用供发动机和航空器使用的最大引气量。但是，若中国民用航空总局发现在进行这样的运转时，持久试验的有效性没有受到影响，则功率、推力或转子轴转速可以比被试的特定工作状态的相应规定值的100%低； (6) 必须对每一附件传动装置和安装构件加载。在仅供航空器使用的每一个附件上所加的载荷，必须是在额定的最大连续功率或推力和更高的功率输出时，由申请人为发动机传动装置和安装点所规定的极限载荷。在对任何附件传动装置和安装构件加载条件下的持久试验也可以在单独的试验台上进行，但试验的有效性必须使用经过批准的分析方法来证实。 (7) 除了试验时间不超过5分钟和不允许稳定的场合外，在以任何额定功率或推力运转期间，燃气温度和滑油进口温度必须保持在限制温度。至少有一次运转必须在燃油、滑油和液压油的最小限制压力下进行；并且至少有一次运转必须在燃油、滑油和液压油最大限制压力下进行，同时，必要时可以降低油液温度以便允许获得最大压力； (8) 如果转子轴瞬时超转或燃气瞬时超温的出现次数有限制，则本条(b)至(g)所规定的加速次数必须在限制超转或超温的情况下进行。如果出现上述超转或超温的次数没有限制，则所规定的加速次数中有一半必须在限制超转或超温的情况下进行； (9) 下列附加试验要求适用于装在超音速航空器上的每型发动机的型号合格审定： (i) 为了改变推力调定值，功率控制杆必须在不超过1秒的时间内从初始位置推到最终位置，但如果为确保点火必须增加时间，以便将功率控制杆推到用燃油产生加力推力的加力位置的情况除。 (ii) 在以任何额定加力推力的运转期间，除了试验时间不足以使温度稳定的场合外，液压油温度必须保持在限制温度下。 (iii) 在模拟超音速运转期间，燃油温度和进气温度不得低于限制温度； (iv) 持久试验必须在装有燃料加力装置和主尾喷管、副尾喷管并在使用可调面积喷管的情况下进行。在每次运转期间，按第33.5(b)规定的方法实施。 (v) 在以最大连续推力和其相应百分比的推力调定值进行运转期间，发动机必须在上述推力调定值的极限进气畸变条件下工作。 (b) 除某些旋翼航空器发动机以外的发动机 除了本条(c)、(d)或(e)款中要求额定值的旋翼机发动机外，对于每型发动机，申请人必须进行下列运转： (1) 起飞和慢车 1小时试验，由5分钟额定起飞功率或推力及5分钟慢车功率或推力交替组成。在起飞和慢车状态及其相应的转子转速和燃气温度条件下发出的功率或推力必须通过用功率控制杆按制造者确定的程序加以调定。在任一个运转周期内，申请人可以在录取检查性能数据时，手动控制转子转速、功率或推力。对于具有加大起飞功率额定值，包括提高涡轮前温度、转子转速或轴功率的发动机，在以起飞功率运转的该周期必须在加大功率额定值的情况下进行。对于实质上不会增加工作苛刻程度的具有加大起飞功率额定值的发动机，以加大功率额定值进行运转的次数由中国民用航空总局决定。在每次5分钟周期后更改功率调定值时，必须按本条(b)(5)规定的方式移动功率控制杆。 (2) 额定最大连续和起飞功率或推力 在下列情况下各运转30分钟： (i) 在25次6小时持久试验循环中的15次期间，应在额定最大连续功率或推力下进行运转。 (ii) 在25次6小时持久试验循环中的10次期间，应在额定起飞功率或推力下进行运转。 (3) 额定最大连续功率或推力 应以额定最大连续功率或推力进行1小时30分钟运转。 (4) 递增的巡航功率或推力 在最大连续发动机转速和地面或最小慢车转速之间应至少分成15个大致相同的转速和时间增量，依次在与这15个转速和时间增量相对应的功率控制杆位置连续进行2小时30分钟的试验。对于以恒定转速工作的发动机，可以用改变推力和功率来代替改变转速。如果在地面慢车和最大连续之间任何状态有显著的峰值振动，则可以变更所选择的增量个数，以便使承受峰值振动影响的运转时数增加到不超过递增运转总时数的50%。 (5) 加速和减速运转 30分钟加速和减速运转应由6个循环组成，而每个循环应由慢车功率或推力到额定起飞功率或推力所组成，并且须在起飞功率控制杆位置保持30秒，在慢车功率控制杆位置保持约4 1/2分钟。为符合本款规定，功率控制杆必须在不超过1秒内从一个极端位置推到另一极端位置；但是，如果采用了必须按时间程序把功率控制杆从一个极端位置移动到另一极端位置的不同的调节工作方式，允许使用较长时间的情况除外。但移动功率杆的时间最长不能超过2秒。 (6) 起动 必须进行100次起动试验，其中的25次必须在发动机停车至少2小时后进行。其中必须至少有10次发动机假起动。每次假起动后准备正常起动前，按申请人规定的最短排油时间暂停起动。其中至少有10次正常再起动必须在发动机停车后 15分钟内进行。其余的起动可以在150小时的持久试验完成后进行。 （c）要求30分钟一台发动机不工作（OEI）功率额定值的旋翼航空器发动机 对于要求30分钟一台发动机不工作（OEI）功率额定值的每型旋翼航空器发动机，申请人必须进行下列一系列试验： (1) 起飞和慢车 1小时试验，由5分钟额定起飞功率及5分钟慢车功率交替组成。在起飞和慢车状态及其相应的转子转速和燃气温度条件下发出的功率必须通过功率控制杆按制造者规定的程序加以确定。在任何一个运转周期内，可以在录取检查性能的数据时，手动控制转子转速和功率和推力。具有加大起飞功率额定值包括增加涡轮进气温度、转子转速或轴功率的发动机，在以额定起飞功率运转期间，必须以加大额定值进行。在每次5分钟试验后变更功率调定值时，必须按本条(c)(5)规定的方式移动功率控制杆。 (2) 额定30分钟一台发动机不工作（OEI）功率 以额定30分钟一台发动机不工作（OEI）功率进行30分钟试验； (3) 额定最大连续功率 以额定最大连续功率和推力运转2小时； (4) 递增的巡航功率 在最大连续发动机转速和地面或最小慢车转速之间应至少分成12个大致相同的转速和时间增量，依次在与这12个转速和时间增量相对应的功率控制杆位置连续进行2小时的试验。对于以恒定转速工作的发动机，可以用改变功率来代替改变转速。如果在地面慢车和最大连续功率之间任何状态有显著的峰值振动，则可以变更所选择的增量个数，以便使承受峰值振动影响的运转时数增加到不超过递增运转总时数的50%。 (5) 加速和减速运转 30分钟加速和减速运转应由6个循环组成，而每个循环应由慢车功率到额定起飞功率所组成，并且须在起飞功率控制杆位置保持30秒，在慢车功率控制杆位置保持约4 1/2分钟。为符合本款规定，功率控制杆必须在不超过1秒内从一个极端位置推到另一极端位置；但是，如果采用了必须按时间程序把功率控制杆从一个极端位置移动到另一极端位置的不同的调节工作方式，允许使用较长时间的情况除外。但移动功率杆的时间最长不能超过2秒。 (6)起动 必须进行100次起动试验，其中的25次必须在发动机停车至少2小时后进行。其中必须至少有10次发动机假起动。每次假起动后准备正常起动前，按申请人规定的最短排油时间暂停起动。其中至少有10次正常再起动必须在发动机停车后 15分钟内进行。其余的起动可以在150小时的持久试验完成后进行。 （d）要求连续一台发动机不工作（OEI）功率额定值的旋翼航空器发动机 对于要求连续一台发动机不工作（OEI）功率额定值的每型旋翼航空器发动机，申请人必须进行下述一系列试验： (1) 起飞和慢车 1小时试验，由5分钟额定起飞功率及5分钟慢车功率交替组成。在起飞和慢车状态及其相应的转子转速和燃气温度条件下发出的功率和推力必须通过功率控制杆按制造者规定的程序加以确定。在任何一个运转周期内，可以在录取检查性能的数据时，手动控制转子转速和功率。具有加大起飞功率额定值包括增加涡轮进气温度、转子转速或轴功率的发动机，在以额定起飞功率运转期间，必须以加大额定值进行。在每次5分钟试验后变更功率调定值时，必须按本条(c)(5)规定的方式移动功率控制杆。 (2) 额定最大连续功率和起飞功率 在下列情况下各运转30分钟： (i) 在25次6小时持久试验循环中的15次期间，以额定最大连续功率进行运转，以及 (ii) 在25次6小时持久试验循环中的10次期间，以额定起飞功率进行运转。 (3) 额定连续一台发动机不工作（OEI）功率 以额定连续一台发动机不工作（OEI）功率运转1小时。 (4) 额定最大连续功率 以额定最大连续功率运转1小时。 (5) 递增的巡航功率 在最大连续发动机转速和地面或最小慢车转速之间应至少分成12个大致相同的转速和时间增量，依次在与这12个转速和时间增量相对应的功率控制杆位置连续进行2小时的试验。对于以恒定转速工作的发动机，可以用改变功率来代替改变转速。如果在地面慢车和最大连续功率之间任何状态有显著的峰值振动，则可以变更所选择的增量个数，以便使承受峰值振动影响的运转时数增加到不超过递增运转总时数的50%。 (6) 加速和减速运转 30分钟加速和减速运转应由6个循环组成，而每个循环应由慢车功率到额定起飞功率所组成，并且须在起飞功率控制杆位置保持30秒，在慢车功率控制杆位置保持约4 1/2分钟。为符合本款规定，功率控制杆必须在不超过1秒内从一个极端位置推到另一极端位置；但是，如果采用了必须按时间程序把功率控制杆从一个极端位置移动到另一极端位置的不同的调节工作方式，允许使用较长时间的情况除外。移动功率杆的时间最长不能超过2秒。 (7) 起动 必须进行100次起动试验，其中的25次必须在发动机停车至少2小时后进行。其中必须至少有10次发动机假起动。每次假起动后准备正常起动前，按申请人规定的最短排油时间暂停起动。其中至少有10次正常再起动必须在发动机停车后 15分钟内进行。其余的起动可以在150小时的持久试验完成后进行。 (e) 要求2 1/2分钟一台发动机不工作（OEI）功率额定值的旋翼航空器发动机 对于要求2 1/2分钟一台发动机不工作（OEI）功率额定值的旋翼航空器发动机，申请人必须进行以下一系列试验： (1) 起飞， 2 1/2分钟一台发动机不工作（OEI）功率和慢车。 1小时试验，由5分钟额定起飞功率及5分钟慢车功率交替组成。但是，在第3次和第6次起飞功率期间，仅需以额定起飞功率试验2 1/2分钟，余下的2 1/2分钟必须以额定的2 1/2分钟OEI功率进行试验的情况除外。在发动机起飞、2 1/2分钟OEI和慢车状态及其相应的转子转速和燃气温度状态下发出的功率，必须通过使用功率控制杆按制造者确定的程序加以调定。在任一个运转期间，申请人在录取检查性能用的数据时，可以手动控制转子转速和功率。具有加大起飞功率额定值，包括增加涡轮前温度、转子转速或轴功率的发动机，在以额定起飞功率运转期间，必须以加大额定值进行。在每次5分钟试验后或试验期间变更功率调定值时，必须按本条(d)(6)规定的方式移动功率控制杆。 (2) 除了25次在每6小时试验程序中的1次外，以及除了在本条(b)(2)规定的30分钟起飞功率试验周期内的最后5分钟，或本条(c)(2)规定的以30分钟OEI功率进行30分钟试验周期内的最后5分钟，或本条(d)(3)规定的1小时连续OEI功率试验周期内的最后5分钟外，按本条(b)(2)至(b)(6)，或(c)(2)至(c)(6)，或(d)(2)至(d)(7)所要求的试验，在适用时，必须在2 1/2分钟OEI功率状态运转。 （f） 要求30秒钟一台发动机不工作（OEI）和2分钟一台发动机不工作（OEI）功率额定值的旋翼航空器发动机 对于要求30秒钟一台发动机不工作（OEI）和2分钟一台发动机不工作（OEI）功率额定值的旋翼航空器发动机，在完成了本条(b)、(c)、(d)或(e)规定的试验后，申请人可以分解试验后的发动机至能证明符合第33.93(a)的要求所需要的程度。此试验发动机必须用按本条(b)、(c)、(d)或(e)试验用的相同零部件重新装配，但持续适航性说明文件规定的消耗件除外。然后，申请人必须进行下列试验程序4次，总时数不低于120分钟： (1) 起飞功率 以额定起飞功率进行3分钟运转。 (2) 30秒钟一台发动机不工作（OEI）功率 以额定30秒钟一台发动机不工作（OEI）功率进行30秒钟运转。 (3) 2分钟一台发动机不工作（OEI）功率 以额定2分钟一台发动机不工作（OEI）功率进行2分钟运转。 (4) 30分钟一台发动机不工作（OEI）功率、连续一台发动机不工作（OEI）功率或最大连续功率 以额定30分钟一台发动机不工作（OEI）功率、额定连续一台发动机不工作（OEI）功率或额定最大连续功率（取大者）进行5分钟运转。但是，第一次试验程序期间，该时间周期应该为65分钟的情况除外。 (5) 50％起飞功率 以50％起飞功率进行1分钟运转。 (6) 30秒钟一台发动机不工作（OEI）功率 以额定30秒钟一台发动机不工作（OEI）功率进行30秒钟运转。 (7) 2分钟一台发动机不工作（OEI）功率 以额定2分钟一台发动机不工作（OEI）功率进行2分钟运转。 (8) 慢车 以慢车功率进行1分钟运转。 (g) 超音速航空器发动机 对于用于超音速航空器的每型发动机的型号合格审定，申请人必须进行下列试验： (1) 在海平面环境大气条件下的亚音速试验 必须进行每阶段1小时共30阶段的运转，每阶段运转由下列各项组成： (i) 2次5分钟的额定起飞加力推力，每次接着5分钟的慢车推力； (ii) 1次5分钟的额定起飞推力，接着5分钟的不超过15%额定起飞推力； (iii) 1次10分钟的额定起飞加力推力，接着2分钟的慢车推力。但是，如果额定最大连续加力推力低于额定起飞加力推力，则10分钟周期中的5分钟为额定最大连续加力推力的情况除外； (iv) 6次1分钟的额定起飞加力推力，每次接着2分钟的慢车推力，包括加速和减速的时间在内。 (2) 模拟超音速试验 必须在模拟超音速试验的每次运转前，把亚音速状态所达到的进气温度和压力变换到超音速所达到的温度和压力，随后必须再返回到亚音速状态所达到的温度。必须进行每阶段4小时共计30阶段的运转，每次运转由下列各项组成： (i) 一个以功率控制杆在额定最大连续加力推力位置上所获得的推力进行30分钟运转周期，接着以功率控制杆在90%额定最大连续加力推力位置上所获得的推力进行10分钟运转 。在前5个阶段该运转周期的末尾，空气进气温度必须在瞬时超温的极限条件下进行，但在本条(g)(2)(ii)至(iv)中规定的试验期间不必重复该运转； (ii) 重复进行一次本条(g)(2)(i)规定的运转周期。但是，必须接着以功率控制杆在80%额定最大连续加力推力位置上所获得的推力进行10分钟运转的情况除外； (iii) 重复进行一次本条(g)(2)(i)规定的运转周期。但是，必须接着以功率操纵杆在60%额定最大连续加力推力位置上所获得的推力进行10分钟运转，然后以不超过15%的额定起飞推力运转10分钟的情况除外； (iv) 重复进行本条(g)(2)(i)和(ii)规定的运转各一次； (v) 进行一次30分钟的运转周期，30个阶段中的25个运转阶段以功率控制杆在额定最大连续加力推力位置上所获得的推力进行，并且每阶段运转后接着在慢车推力状态下工作；其余的5个运转阶段以功率控制杆在额定最大连续加力推力位置上所获得的推力试验25分钟，每阶段接着用热燃油以不大于15%的额定起飞推力进行亚音速工作，并加速到额定起飞推力工作5分钟。 (3) 起动 必须进行100次起动试验，其中的25次必须在发动机停车至少2小时后进行。其中必须至少有10次发动机假起动。每次假起动后准备正常起动前，按申请人规定的最短排油时间暂停起动。其中至少有10次正常再起动必须在发动机停车后 15分钟内进行。起动可以在包括持久试验期间的任何时候进行。 [2002年4月19日第一次修订] 第33.88条 发动机超温试验 (a) 每型发动机必须在比最大额定功率下的稳态工作限制温度高至少42℃(75oF)的燃气温度下，以最大允许转速运转5分钟。但不包括对应30秒钟一台发动机不工作（OEI） 和2分钟一台发动机不工作（OEI）的转速和燃气温度的最大值。在此运转后，涡轮部件必须在可使用的限制范围内。 (b) 每型要求30秒钟一台发动机不工作（OEI）和2分钟一台发动机不工作（OEI）功率额定值的发动机，在不安装温度限制装置的情况下，必须在超过30秒钟一台发动机不工作（OEI）功率额定值工作限制温度至少42oC（75oF）时，以接通最大功率转速运转5分钟。在此运转后，只要通过中国民用航空总局认为必要的分析或试验证明发动机能保持涡轮部件的完整性，则在涡轮部件上可以有超出该超温条件限制范围的损伤。 (c) 要求30秒钟一台发动机不工作（OEI）和2分钟一台发动机不工作（OEI）功率额定值的每型发动机，在安装温度限制装置的情况下，必须在超过30秒钟一台发动机不工作（OEI）功率额定值工作限制温度至少20oC（35oF）时，以接通最大功率转速运转4分钟。在此运转后，只要通过中国民用航空总局认为必要的分析或试验表明发动机能保持涡轮部件的完整性，则在涡轮部件上可以有超出该超温条件限制范围的损伤。 (d) 对每一试验条件，可以使用单独的试验设备。 [2002年4月19日第一次修订] 第33.89条 工作试验 (a) 工作试验必须包括中国民用航空总局认为必要的试验，以验证下列各项： (1) 起动、慢车、加速、超转、点火、螺旋桨功能(如果规定发动机装螺旋桨工作)； (2) 符合第33.73条发动机的响应要求； (3) 在下列发动机载荷条件下，从功率操纵杆代表的最小慢车和最小飞行慢车的位置由稳定的慢车工作状态开始到95%的额定起飞功率或推力状态的功率或推力最小响应时间： (i) 没有供航空器使用的引气和功率提取； (ii) 供航空器使用的最大允许引气和功率提取值； (iii) 代表航空器进场着陆期间使用的最大的引气和功率提取的某中间值。 (4) 如果没有合适的试验设备，则确定本条(a)(3)(ii)和(iii)规定的功率提取可以通过适当的分析方法进行。 (b) 工作试验必须包括中国民用航空总局认为必要的所有试验，以验证发动机在其规定的整个使用包线内所具有的安全工作特性。 第33.90条 初次维修检查 除了正在进行现有型号合格证更改或补充的型号合格审定的发动机外，每型发动机必须承受批准的运转试验，来模拟使用中所预期的发动机工作状态，包括典型的起动—停车循环，以确定要求初次维修检查的时限。运转试验必须在基本符合最终型号设计的发动机上进行。 第33.91条 发动机部件试验 (a) 对于不能按本规定第33.87条规定的持久试验予以充分验证的那些系统，必须进行另外的试验，以确定在所有正常预期的飞行和大气条件下，这些部件能可靠地工作。 (b) 必须确定在航空器安装中要求温度控制措施的那些部件的温度限制，以确保其良好的功能、可靠性和耐久性。 (c) 每个不增压的液压油油箱在受到最大工作温度和34.5千帕（5磅/英寸2）的内部压力时，不得出现失效或泄漏。每个增压的液压油油箱在受到最大工作温度和不低于34.5千帕（5磅/英寸2）的内部压力加上油箱的最大工作压力时，不得出现失效或泄漏。 (d) 对于超音速航空器的发动机型号合格审定，必须确定由于在最高和最低工作温度时可能会发生失效的发动机系统、安全装置及外部附件。并且必须在最高和最低工作温度以及当温度和其他使用条件在最高和最低使用值之间循环时进行试验。 [2002年4月19日第一次修订] 第33.92条 转子锁定试验 如果采用锁定转子装置以阻止发动机持续转动，则发动机必须在以下条件下进行包括该装置工作25次的试验： (a) 发动机必须从额定最大连续推力或功率状态停车；并且 (b) 必须在承受在该状态下持续飞行可能引起的最大扭矩的情况下，按发动机使用说明的规定操作停止和锁定转子的装置，并且 (c) 在25次工作中，每一次转子锁定后，转子必须在这些状态下保持静止5分钟。 [2002年4月19日第一次修订] 第33.93条 分解检查 (a) 在完成本章第33.87条(b)、(c)、(d)、(e)或(g)的持久试验后，每台发动机必须完全分解，并满足下列要求： (1) 不论是否安装在发动机上即可确定其调整位置和功能特性的每个部件，必须使其每个调整位置和功能特性保持在试验开始时确定和记录的限制范围内。 (2) 按第33.4条提交的资料，每个发动机零部件必须符合型号设计并且应仍然可以安装在发动机上继续使用。 (b) 在完成本章第33.87条(f)的持久试验后，每台发动机必须完全分解，并满足下列要求： (1) 不论是否安装在发动机上即可确定其调整位置和功能特性的每个部件，必须使其每个调整位置和功能特性保持在试验开始时确定和记录的限制范围内；并且 (2) 每型发动机可以有超出本条(a)(2)允许的损伤，包括某些不适合于进一步使用的发动机零件或部件。当中国民用航空总局认为必要时，申请人必须通过分析和、或试验，证明发动机以及包括安装节、机匣、轴承座、轴和转子的结构完整性得到了保持；或者 (c) 代替本条(b)的符合性，可以在要求30秒钟一台发动机不工作（OEI）和2分钟一台发动机不工作（OEI）功率额定值的发动机上进行本规定第33.87(b)、(c)、(d)或(e)规定的持久试验，接着进行第33.87(f)规定的试验，但中间不进行分解和检查。在完成第33.87(f)的持久试验后，发动机必须满足本条(a)的要求。 [2002年4月19日第一次修订] 第33.94条 叶片包容性和转子不平衡试验 (a) 除了本条(b)款规定外，除非在下列每一事故后发动机损坏的结果导致了自动停车，否则必须通过发动机试验验证：发动机能包容损坏件至少运转15秒不着火，并且其安装节也不失效。 (1) 在以最大允许转速运转期间，最危险的压气机或风扇的一个叶片失效。 该叶片失效必须出现在盘上最外层的固定榫槽处；或对于整体叶盘转子，叶片必须至少缺损80%。 (2) 在以最大允许转速运转期间，最危险的涡轮叶片失效。 该叶片失效必须出现在盘上最外部的固定榫槽处；或对于整体叶盘转子，该叶片必须至少缺损80%。必须根据涡轮叶片的重量和其邻近的涡轮机匣在与最大允许转速运转相关的温度和压力下的强度确定该最危险的涡轮叶片。 (b) 基于根据试验台试验、部件试验或使用经验的分析如果符合下列条件，可以代替本条(a)(1)和(a)(2)规定的发动机试验之一： (1) 某一试验（上述规定的两个试验之一）产生的转子不平衡量为最小； (2) 证明分析是等同于上述某一试验。 第33.95条 发动机—螺旋桨系统试验 如果设计的发动机是带螺旋桨工作的，则必须在装有一个有代表性的螺旋桨的情况下，进行下列试验，该试验可以包括在持久试验中；或者按中国民用航空总局接受的其他方法进行下列试验： (a) 顺桨试验 25次循环； (b) 负扭矩和推力系统试验 以额定最大连续功率作25次循环； (c) 自动退耦装置试验 以额定最大连续功率作25次循环(如果重复退耦和重新耦合是这种装置在使用中的预期功能)； (d) 负拉力 从飞行慢车位置到全负拉力175次循环；和以额定最大连续功率从全正拉力到全负拉力的25次循环。在每个循环结束时，螺旋桨必须在申请人对反桨距运转所规定的最大转速和功率下，用反桨距运转30秒。 第33.96条 以辅助动力装置(APU)方式工作的发动机试验 如果发动机设计成带螺旋桨制动器，而该制动器在发动机燃气发生器仍然工作期间，允许螺旋桨制动，并在发动机作为辅助动力装置（APU）方式工作期间保持制动，那么除了第33.87条的要求外，申请人必须做下列试验： (a) 地面锁定：螺旋桨制动器以某种方式耦合共45小时。这种方式在申请人规定的发动机转速、扭矩、温度、引气和功率提取的最大状态下，发动机处于APU方式工作时，能清楚地验证它的功能对全台发动机无有害的影响。 (b) 动态制动：制动器必须以某种方式进行共400个使用－放松耦合的循环。这种方式在申请人规定的发动机最大状态的加速/减速、转速、扭矩和温度时，能清楚地验证制动器的功能对全台发动机无有害的影响。制动器放松之前，螺旋桨必须制动。 (c) 螺旋桨制动器耦合时，进行100次发动机起动和停车。 (d) 本条(a)、(b)和(c)规定的试验必须在同一台发动机上进行，但这台发动机不必是第33.87规定试验中使用的同一台发动机。 (e) 必须在完成本条(a)、(b)和(c)规定的试验后，将发动机分解到为表 明符合第33.93(a)和第33.93(b)所必需的程度。 第33.97条 反推力装置 (a) 如果发动机装有反推力装置，则本章规定的持久、校准、工作和振动试验必须在安装了反推力装置的情况下进行。根据本条规定，功率操纵杆必须在不超过1秒的时间内从一个极端位置移到另一个极端位置，除非操纵方式需要功率操纵杆从一个极端位置移到另一个极端位置，有计划地进行，才允许有稍长的时间，但不能超过3秒。另外，本条(b)规定的试验也必须进行。这一试验可以作为持久试验的一部分。 (b) 必须进行从飞行慢车的向前推力到最大反推力的试验175次，以及必须从额定起飞推力到最大反推力进行25次反推力试验。在每次反推力后，反推力装置必须在全反推力下工作1分钟，除非反推力装置仅预备用作为地面制动装置，则该反推力装置只需在全反推力下工作30秒。 第33.99条 台架试验的一般实施 (a) 在作台架试验时，每个申请人可用同一设计和结构的几台发动机分别进行振动、校准、持久和工作试验。如果用一台发动机单独进行持久试验，则该发动机在持久试验开始之间，必须进行校准检查。 (b) 申请人根据符合第33.4条的要求提交维修和维护说明书，可以对在台架试验期间的发动机进行维护和小修。如果这类维护频次过高；或由于发动机故障，停车次数过多，或在台架试验期间或分解检查的结果认为有必要大修或更换零件的话，则发动机或其零部件可能要进行中国民用航空总局认为必要的任何附加试验。 (c) 每个申请人必须提供所有试验条件，包括设备和胜任的人员，以实施台架试验。 附件A 持续适航文件 (a) 本附录规定第33.4条所需要的持续适航文件的编制要求。 (b) 每一发动机持续适航文件必须包含所有发动机零部件的各种持续适航文件。如果发动机部件制造者未提供发动机零部件的持续适航文件，则发动机的持续适航文件必须包含对于发动机持续适航性必不可少的资料。 (c) 申请人必须向中国民用航空总局提交一份文件，说明如何分发由申请人或发动机零部件制造者对持续适航文件的更改资料。 第A33.2条 格式 (a) 必须根据所提供资料的数量，将持续适航文件编成一本或多本手册。 (b) 手册的编排格式必须实用。 第A33.3条 内容 手册的内容必须用中文编写。持续适航文件必须含有下列手册或条款(视适用而定)以及下列资料： (a) 发动机维护手册或条款 (1) 概述性资料，包括在维护或预防性维护所必需的对发动机特点和数据的说明； (2) 发动机及其部件、系统和安装的详细说明； (3) 安装说明，包括拆包、启封、验收、起吊和安装附件的正确程度及任何必要的检查； (4) 说明发动机部件、系统和装置如何使用的基本控制和使用资料，及说明发动机及其零部件起动、运转、试验和停车方法的资料，包括采用的特殊程序和限制； (5) 关于下列细节内容的维护资料：维护点、油箱和流体容器的容量、所用流体的类型、各系统所采用的压力、润滑点位置、所用的润滑剂和维护所需的设备； (6) 发动机每一零部件的定期维护资料，它给出发动机每一零部件的清洗、检查、调整、试验和润滑的荐用周期，并提供检查的程度、适用的磨损允差和在这些周期内推荐的工作内容。但是如果申请人表明某项附件、仪表或设备非常复杂，需要专业化的维护技术、测试设备或专家才能处理，则申请人可以指明向该件的制造厂商索取上述资料。荐用的翻修周期和与本文件适航性限制条款必要的互相参照也必须列入。此外，申请人必须提交一份包含发动机持续适航性所需检查频数和范围的检查大纲； (7) 说明可能发生的故障、如何判别这些故障以及对这些故障采取补救措施的检查排故资料； (8) 说明拆卸发动机及其零部件和更换零部件的顺序和方法及应采取的必要防护措施的资料。还必须包括正确的有关地面保管、装箱和运输的说明； (9) 维护所必需的工具和设备清单及其使用方法的说明。 (b) 发动机翻修手册或条款 (1) 分解资料包括翻修分解顺序和方法； (2) 清洗与检查说明包括翻修期间使用的材料和仪器、采用的方法和防护措施。还必须包括翻修检查的方法； (3) 有关翻修的所有公差与配合的明细表； (4) 磨损的或其他低于标准零部件详细的修理方法及其确定何时必须更换的必要资料； (5) 翻修时装配的顺序和方法； (6) 翻修后的试验说明； (7) 储存处理包括任何储存限制的说明； (8) 翻修需要的工具清单。 第A33.4条 适航限制条款 持续适航文件必须包含题为适航性限制的条款，该条应单独编排并与文件的其他部分明显地区分开来。该条必须规定强制性的更换时间、检查时间间隔和型号合格审定要求的有关程序。如持续适航文件由多本文件组成，则本节要求的条款必须编在主要手册中。必须在该条显著位置清晰说明：“本适航限制条款业经中国民用航空总局批准，规定了中国民用航空规章有关维护和营运的条款所要求的维护，如果中国民用航空总局已另行批准使用替代的大纲则除外”。 附件B 合格审定标准大气降雨和冰雹的浓度 为了按照第33.78条（a）（2）的要求进行合格审定，图B1、表B1、表B2、表B3、表B4规定了雨和冰雹的大气浓度和尺寸分布。只要申请人能表明所使用的替代方法没有降低试验的严格程度，在通常通过喷洒液态水模拟降雨以及投掷冰块制造的冰雹模拟降冰雹的情况下，允许使用不同于本规定附录B规定的这些水滴和冰雹的形状、尺寸和尺寸分布，或者允许使用尺寸和形状单一的水滴或冰雹。 图B1 雨和冰雹的征兆图表，利用表B1和B2可获得合格审定浓度 表B1 合格审定标准的大气雨浓度 注：在其他高度上雨的水含量的值可以由线性内插的方法确定。 表B2 合格审定标准的大气冰雹浓度 注：在其他高度上的冰雹水含量值可以用线性内插法确定。低于2,230米(7,300英尺)和大于8,840米(29,000英尺)的冰雹征兆可根据线性外插数据获得。 表B3 合格审定标准的大气雨滴尺寸分布 注：雨滴的平均直径为2.66毫米 表B4 合格审定标准的大气冰雹尺寸分布 注：冰雹的平均直径为16毫米 [2002年4月19日第一次修订] 关于《中国民用航空总局关于修订〈航空发动机适航标准〉的决定》的说明 ** ** 随着航空科学技术的发展，各种新技术不断应用于航空发动机，人们对安全标准的认识也在逐步提高。《航空发动机适航标准》（CCAR-33）主要是参考美国联邦航空条例FAR-33第11修正案制定的。目前FAR-33已修订到第20修正案。从第12到20修正案增加了多发旋翼航空器发动机一台发动机不工作（OEI）功率额定值、发动机电气和电子控制系统等新概念、新技术，更新了发动机的部分验证标准。 为保持我国适航标准与国际标准同步，防止国外不符合现行国际标准的发动机进入我国造成民用航空飞行隐患，配合国际民航组织安全审计和我国新支线飞机项目的开展，中国民用航空总局依据《中华人民共和国民用航空法》第三十四条，决定修订《航空发动机适航标准》，修订后的名称改为《航空发动机适航规定》。 二 修订的主要情况 在格式上，本次修订将原规章中A分部、B分部、C分部、D分部、E分部、F分部分别改为A章、B章、C章、D章、E章、F章；原规章中关于条的序号的表述“§……”改为“第……条”。 在内容上，本次修订主要参考了FAR-33第12至20修正案，修订内容主要包括： 1．引进了多发旋翼航空器发动机一台发动机不工作（OEI）功率额定值的概念； 2．新增加了对发动机电气和电子控制系统、持续转动、转子锁定试验、涡轮发动机飞机燃油排泄和排气排出物规定； 3．更新了对仪表连接、振动和振动试验、吸鸟、外物冰吸入、吸雨和吸雹、校准试验、持久试验、超温试验、发动机部件试验、分解检查等要求的验证，内容共涉及18个条款。 在文字处理上，本此修订尽量与1988年2月9日发布的《航空发动机适航标准》（CCAR-33）的文字保持一致。对于内容不做修订的原规章文字，虽不妥贴但含义仍然正确的，原则上不作改动；对于本此修订涉及的个别文字表达过于生涩的条款，在不影响原意的情况下进行了文字调整。 三 修订内容的说明 2．修订后的第33.7条（C）中增加了额定30分钟一台发动机不工作（OEI）功率、额定2 1/2分钟一台发动机不工作（OEI）功率、额定连续一台发动机不工作（OEI）功率、额定2分钟一台发动机不工作（OEI）功率、额定30秒钟一台发动机不工作（OEI）功率和辅助动力装置(APU)工作方式的要求。 3．增加第33.28条，提出了对电气和电子控制系统的具体要求。 4．修订后的第33.29条中增加了对于具有30秒钟一台发动机不工作（OEI）和2分钟一台发动机不工作（OEI）功率额定值的旋翼航空器涡轮发动机，在该状态工作时需提醒飞行员注意的要求。 5．修订后的第33.63条增加了发动机对运转的环境条件“飞行包线”的要求，在设计上建立发动机正常工作的边界。 6．修订后的第33.67条增加了对具有30秒钟一台发动机不工作（OEI）功率额定值的发动机的燃油系统的自动可用性和自动控制的要求。 7．增加第33.74条，提出了发动机空中停车对其转动系统的要求。 8．修订后的第33.77条规定了吸冰的要求，将原规章中吸鸟的要求列为第33.76条,将原规章中吸冰雹和吸雨的要求列为第33.78条。 9．修订后的第33.83条主要有下列变化： (b)对振动测试的物理转速和换算转速的范围作出新的规定。 (c)要求对列举的各因素对发动机振动的影响进行评估，并强调对颤振的评估。 (d)澄清振动应力与稳态应力的组合小于材料的持久极限。 10．修订后的第33.85条增加了（c）、（d）： （c）强调进行（a）、（b）规定的校准试验时每一状态必须稳定。 （d）增加了对具有30秒钟一台发动机不工作（OEI）和2分钟一台发动机不工作（OEI）功率额定值的旋翼航空器发动机的校准试验要求。 11．修订后的第33.87条增加了对具有连续一台发动机不工作（OEI）功率额定值、30秒钟一台发动机不工作（OEI）功率额定值和2分钟一台发动机不工作（OEI）功率额定值的旋翼航空器发动机的持久试验要求。 （a）增加了“对要求有30秒钟一台发动机不工作（OEI）和2分钟一台发动机不工作（OEI）功率额定值的发动机必须按本条(f)进一步试验”的要求。 （d）为增加的内容。增加了对于要求连续一台发动机不工作（OEI）功率额定值的旋翼航空器发动机的持久试验要求。 （e）编辑更改，其内容与原规章的§33.87（d）相同。 （f）为增加的内容。增加了对要求有30秒钟一台发动机不工作（OEI）和2分钟一台发动机不工作（OEI）功率额定值的旋翼航空器发动机的持久试验要求。 （g）编辑更改，其内容与原规章的§33.87（e）相同。 12．修订后的第33.88条增加了对要求有30秒钟一台发动机不工作（OEI）和2分钟一台发动机不工作（OEI）功率额定值的旋翼航空器发动机的超温试验要求，并且针对是否安装限制温度装置提出不同的试验要求。 （a）编辑更改，其内容与原规章的§33.88相同。 （b）增加了对不安装限制温度装置、要求有30秒钟一台发动机不工作（OEI）和2分钟一台发动机不工作（OEI）功率额定值的旋翼航空器发动机的超温试验要求。 （c）增加了对安装限制温度装置、要求有30秒钟一台发动机工作（OEI）和2分钟一台发动机不工作（OEI）功率额定值的旋翼航空器发动机的超温试验要求。 13．修订后的第33.91条增加（c）、（d），增加的内容为FAR-33第6修正案中增加的内容：航空器和航空发动机合格审定程序和型号合格审定标准。该修订案在美国于1974年10月31日生效。 14．修订后的第33.92条增加了转子锁定试验要求。 15．修订后的第33.93条增加了对要求有30秒钟一台发动机不工作（OEI）和2分钟一台发动机不工作（OEI）功率额定值的旋翼航空器发动机的分解检查的要求。 16．增加附件B，规定了符合第33.78条(a)(2)要求的雨和冰雹的大气浓度和尺寸分布。 四 修订参考资料 修正案编号 标题 生效日期 Amdt33-12 旋翼航空器规章评审大纲，第三号修正案1988.10.03 Amdt33-13 运行和飞行的一般规则修订（不适用）1990.08.18 Amdt33-14 涡轮发动机飞机燃油排泄和排气排出物的要求1990.09.10 Amdt33-15 适航标准：航空发动机电气和电子控制系统1993.08.16 Amdt33-16 权限援引的修订 （不适用）1995.12.28 Amdt33-17 适航标准：持续转动和转子锁定试验，振动试验1996.07.05 Amdt33-18 适航标准：旋翼航空器发动机一台发动机不工作（OEI）额定值的定义和型号审查标准1996.08.19 Amdt33-19 适航标准：吸雨和吸雹1998.04.30 Amdt33-20 适航标准：吸鸟和外物吸入—冰2000.12.13 五 CCAR-33本次修订涉及的条款 附件： 中国民用航空总局关于修订《航空发动机适航标准》的决定 民航局客户端 立即下载