14 C.F.R. Subpart B—Design and Construction; General
Title 14 - Aeronautics and Space
This subpart prescribes the general design and construction requirements for reciprocating and turbine aircraft engines.
Title 14: Aeronautics and Space
PART 33—AIRWORTHINESS STANDARDS: AIRCRAFT ENGINES
Subpart B—Design and Construction; General
§ 33.11 Applicability.
§ 33.13 [Reserved]
§ 33.14 Start-stop cyclic stress (low-cycle fatigue).
By a procedure approved by the FAA, operating limitations must be established which specify the maximum allowable number of start-stop stress cycles for each rotor structural part (such as discs, spacers, hubs, and shafts of the compressors and turbines), the failure of which could produce a hazard to the aircraft. A start-stop stress cycle consists of a flight cycle profile or an equivalent representation of engine usage. It includes starting the engine, accelerating to maximum rated power or thrust, decelerating, and stopping. For each cycle, the rotor structural parts must reach stabilized temperature during engine operation at a maximum rate power or thrust and after engine shutdown, unless it is shown that the parts undergo the same stress range without temperature stabilization.
[Amdt. 33–10, 49 FR 6850, Feb. 23, 1984]
§ 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.
Secs. 313(a), 601, and 603, 72 Stat. 759, 775, 49 U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 U.S.C. 1655(c))
[Amdt. 33–8, 42 FR 15047, Mar. 17, 1977, as amended by Amdt. 33–10, 49 FR 6850, Feb. 23, 1984]
§ 33.17 Fire prevention.
(a) The design and construction of the engine and the materials used must minimize the probability of the occurrence and spread of 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, overheating, or other hazardous conditions.
(b) Except as provided in paragraphs (c), (d), and (e) of this section, each external line, fitting, and other component, which contains or conveys flammable fluid must be fire resistant. Components must be shielded or located to safeguard against the ignition of leaking flammable fluid.
(c) Flammable fluid tanks and supports which are part of and attached to the engine must be fireproof or be enclosed by a fireproof shield unless damage by fire to any non-fireproof part will not cause leakage or spillage of flammable fluid. For a reciprocating engine having an integral oil sump of less than 25-quart capacity, the oil sump need not be fireproof nor be enclosed by fireproof shield.
(d) For turbine engines type certificated for use in supersonic aircraft, each external component which conveys or contains flammable fluid must be fireproof.
(e) Unwanted accumulation of flammable fluid and vapor must be prevented by draining and venting.
(Secs. 313(a), 601, and 603, 72 Stat. 759, 775, 49 U.S.C. 1354(a), 1421, and 1423; sec. 6(c), 49 U.S.C. 1655(c))
[Amdt. 33–6, 39 FR 35464, Oct. 1, 1974, as amended by Amdt. 33–8, 42 FR 15047, Mar. 17, 1977; Amdt. 33–10, 49 FR 6850, Feb. 23, 1984]
§ 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 §35.42 of this chapter.
[Doc. No. 3025, 29 FR 7453, June 10, 1964, as amended by Amdt. 33–9, 45 FR 60181, Sept. 11, 1980; Amdt. 33–10, 49 FR 6851, Feb. 23, 1984]
§ 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.
[Amdt. 33–10, 49 FR 6851, Feb. 23, 1984]
§ 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.
[Amdt. 33–10, 49 FR 6851, Feb. 23, 1984]
§ 33.27 Turbine, compressor, fan, and turbosupercharger rotors.
(a) Turbine, compressor, fan, and turbosupercharger rotors must have sufficient strength to withstand the test conditions specified in paragraph (c) of this section.
(b) The design and functioning of engine control devices, systems, and instruments must give reasonable assurance that those engine operating limitations that affect turbine, compressor, fan, and turbosupercharger rotor structural integrity will not be exceeded in service.
(c) The most critically stressed rotor component (except blades) of each turbine, compressor, and fan, including integral drum rotors and centrifugal compressors in an engine or turbosupercharger, as determined by analysis or other acceptable means, must be tested for a period of 5 minutes—
(1) At its maximum operating temperature, except as provided in paragraph (c)(2)(iv) of this section; and
(2) At the highest speed of the following, as applicable:
(i) 120 percent of its maximum permissible r.p.m. if tested on a rig and equipped with blades or blade weights.
(ii) 115 percent of its maximum permissible r.p.m. if tested on an engine.
(iii) 115 percent of its maximum permissible r.p.m. if tested on turbosupercharger driven by a hot gas supply from a special burner rig.
(iv) 120 percent of the r.p.m. at which, while cold spinning, it is subject to operating stresses that are equivalent to those induced at the maximum operating temperature and maximum permissible r.p.m.
(v) 105 percent of the highest speed that would result from failure of the most critical component or system in a representative installation of the engine.
(vi) The highest speed that would result from the failure of any component or system in a representative installation of the engine, in combination with any failure of a component or system that would not normally be detected during a routine preflight check or during normal flight operation.
Following the test, each rotor must be within approved dimensional limits for an overspeed condition and may not be cracked.
[Amdt. 33–10, 49 FR 6851, Feb. 23, 1984]
§ 33.28 Electrical and electronic engine control systems.
Each control system which relies on electrical and electronic means for normal operation must:
(a) Have the control system description, the percent of available power or trust controlled in both normal operation and failure conditions, and the range of control of other controlled functions, specified in the instruction manual required by §33.5 for the engine;
(b) Be designed and constructed so that any failure of aircraft-supplied power or data will not result in an unacceptable change in power or thrust, or prevent continued safe operation of the engine;
(c) Be designed and constructed so that no single failure or malfunction, or probable combination of failures of electrical or electronic components of the control system, results in an unsafe condition;
(d) Have environmental limits, including transients caused by lightning strikes, specified in the instruction manual; and
(e) Have all associated software designed and implemented to prevent errors that would result in an unacceptable loss of power or thrust, or other unsafe condition, and have the method used to design and implement the software approved by the Administrator.
[Doc. No. 24466, 58 FR 29095, May 18, 1993]
§ 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 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) Determine, in a positive manner, that the engine has been operated at each rating; and
(3) Automatically record each usage and duration of power at each rating.
[Amdt. 33–5, 39 FR 1831, Jan. 15, 1974, as amended by Amdt. 33–6, 39 FR 35465, Oct. 1, 1974; Amdt. 33–18, 61 FR 31328, June 19, 1996]
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