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 a 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.
*** Guidance. The INTENT of this section is to assure that design, materials, and construction techniques are used to minimize the occurrence and spread of fire.
Incorporations: AC 33-4, Design Considerations Concerning the Use of Titanium in Aircraft Turbine Engines, dated July 28, 1983.
References:
SAE AS1055B, 'Fire Testing of Flexible Hose, Tube Assemblies, Coils, Fittings, and Similar System Components, March 1, 1978.
2. SAE AIR 1377A, Fire Test Equipment for Flexible Hose and Tube Assemblies, January 1980.
3. FAA Powerplant Engineering Report No. 3A, ¿'¿Standard Fire rest Apparatus and Procedure, Revised March 1978.
4. FAA Report No. FAA-RD-79-51, Titanium Combustion in Turbine Engines," dated July 1979; and JAR-E, Section 2, Paragraph ACJE530 (C) , "Titanium Fires."
5. SAE Report No. 690436, ¿¿Ignition of Aircraft Fluids on High Temperature Engine Surfaces, by W.T. Westfield of the FAA.
6. FAA Report No. FAA-RD-75-155, 'Ignition and Propagation Rates for Flames in a Fuel Mist, October 1975, by C.E.
Polymeropoulos.
7. AC 20-135, 'Powerplant Installation and Propulsion System Component Fire Protection Test Methods, Standards and Criteria,i dated February 6, 1990.
proof are very broad and are not quantified in terms of flame The FAR Part 1 definitions of "fire resistant" and "fire temperature and time of immersion. For purposes of engine certification, fire resistant should be associated with a test flame temperature of 2000 °F, for at least a 5-minute duration;
Chap 3 40 Par 22
fireproof should be associated with a test flame temperature of 2000 °F, for at least a 15-minute duration.
b. Fire test conditions to be applied should be established consistent with the intended use of the engine component or subsystem being evaluated. Examples of this are: sheet stock designed to be firewalls or fireshrouds; formed parts designed as oil tanks; castings designed as fuel pump bodies; seamless tubing designed as fuel and oil supply lines; etc.
The test flame should be applied at the component or subsystem location that is determined by analysis and/or inspection to be most critical to fire susceptibility.
Additionally, the operating characteristics of the component or subsystem should be established consistent with those which would normally be expected to exist during actual engine fire conditions.
Fire susceptibility means not only the probable location of a fire, but also the specific location of the component or subsystem within the fire zone that may be least likely to survive the effects of a fire. Such a determination should consider, at least, the following potential factors: materials characteristics; parts geometry; local torching effects; joint seals; vibration characteristics; internal fluid levels, pressures, and flows; wicking of surface coatings; engine design requirement for installed engine, through nacelle exhaust ejector driven component cooling airflows; characteristics of probable leakage in forms of sprays, drops, streams; etc.
e. The fire test burner to be used should provide the desired 2000 °F flame temperature of sufficient size, and at an energy level (BTU/hr) appropriate to the geometry of the engine component or subsystem being evaluated. Further general guidance on burner configuration can be found in References 1, 2, and 3 above.
f. In all cases, fire test acceptance criteria should show no support of combustion by the constituent material being tested; no burn-through; no leakage of fluid-carrying or fluid-holding parts sufficient to increase the fire hazard; and fire-protecting parts should be capable of satisfactorily performing the functions for which they were designed.
g. Compliance with the requirements of Section 33.17(a) requires that 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. (See also Reference 4 above).
Chap 3 Par 22 41
h. Available references relating to hot surface ignition (HSI) criteria for fuels are consistent in that the HSI temperature for JP-type fuels is in the range of 425-500 °F (minimum) for the non-ventilated case. HSI temperature increases dramatically with increasing airflow. It is also influenced by the characteristics of the secondary cooling air and engine surface geometry, ventilation rates of the engine nacelle, direction and changes of direction of the cooling air, and changes in engine power while leakage is occurring. Further information on this topic is available in References 5 and 6 above.
23. Section 33.19, Durability.
Section 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 Section 35.42 of this chapter.
*** Guidance. The INTENT of this section is self-evident Incorporations:
1 • AC 33-4, 'Design Considerations Concerning the Use of Titanium in Aircraft Turbine Engines, dated July 28, 1983.
2. AC 33-5, 'Turbine Engine Rotor Blade Containment/Durability, dated June 18, 1990.
References: None.
a. This section addresses the generic safety requirement of engines between "overhaul" periods, in addition to specific design requirements for rotor blade containment and for propeller blade pitch control.
(1) The generic requirement is fulfilled by compliance not only to each of the appropriate sections of FAR Part 33, but also to the "grandfather" Sections of 21.33, "Inspection and tests," and 21.16, "Special conditions."
Chap 3 42 Par 22
Collectively, these speak to not only the basis for type certification, but more specifically to: (1) materials and products specifications; (2) conformance of parts to the drawings in the type design; (3) specification of manufacturing processes;
and (4) construction and assembly specifications in the engine type design (drawings). Each of these items should be traceable in the type certification design documentation from the engine top assembly drawing.
b. This section also addresses the specific design requirements for rotor blade containment, whereas FAR Section
33.94 addresses the demonstration requirements; both are addressed in Incorporation 2. Essentially, the design (analysis) is borne out by testing to substantiate that the potential energy of the containment structure is greater than the kinetic energy of the released rotor blade debris.
24. Section 33.21, Engine Cooling.
Section 33.21 Engine cooling.
Engine design and construction must provide the necessary cooling under conditions in which the airplane is expected to operate.
*** Guidance. The INTENT of this section is self-evident.
Incorporations: None.
References: None.
Interpretations of this regulation have historically been limited to operating fluid, external engine surface, and component heat rejection considerations, particularly of reciprocating engines. However, inasmuch as the rule applies to ¿Design and Construction-General, it should be interpreted to apply to both reciprocating and turbine engines, and to all engine cooling considerations.
Evaluation of the engine cooling systems design should not only address operating fluid and external component cooling, but also internal flow path cooling of turbochargers, rotors, spacers, nozzles, combustors, cases, etc. Particular attention should be given to ensure that continued safe operation and integrity of the critical structural components, while operating in the intended environment, is maintained throughout the design life and/or time to inspection, as appropriate.
Chap 3 Par 23 43
25. Section 33.23, Engine Mounting Attachments and Structure.
Section 33.23 Engine mounting attachments and structure.
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.
*** Guidance. The INTENT of this section is to specify the allowable limit and ultimate loads, as defined, for the engine mounting attachments and related engine structure.
Incorporations: None.
References: None.
Engine attachment (mount) limit loads are the maximum to be expected in service and include, in combination, static mount loads, operating engine maximum reactive torque loads, gyroscopic loads, and maximum ground and flight loads.
Ultimate load capability of the engine attachment mount structure is substantiated by analysis and by limited rig and component testing, which should also serve to validate the applicant's analytical methodology.
b. Engine mounting attachments and structure design should include consideration of deterioration, and corrosion effects expected during the design life of the engine. Suitable methods of detection of pending faults and failures at, or between, inspection intervals to be approved, should also be given consideration.
26. Section 33.25, Accessory Attachments.
Section 33.25 Accessory attachments.
The engine must operate properly with the accessory drive and mounting attachments loaded.
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 Chap 3 44 Par 25
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.
*** Guidance. The INTENT of this section is that design and construction of engine accessory drives and their mount pads permit proper engine operation when loaded; prevent contamination and excessive loss of oil; and permit inspection, adjustment, and removal of accessories.
Incorporations: None.
References: None.
If engine accessory drives are designed to protect the engine from potential unsafe conditions, such as those due to accessory seizure overtorque, by designed failure of a drive shaft shear section; then, it should be shown that the shear section will not fail under all normal operating load conditions, or otherwise result in damage or hazard to the engine or its subsequent aircraft installation.
27. Section 33.27, Turbine, Compressor, Fan and Turbosupercharger Rotors.
Section 33.27 Turbine, compressor, fan and turbosupercharger rotors.
(a) Turbine, compressor, fan, and turbosupercharger 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, 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.
Chap 3 Par 26 45
(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 a 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.
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.
*** Guidance. The INTENT of this section is to assure engine rotor structural integrity, by design and functioning of engine control and instrumentation systems, to inhibit exceedances of operating limitations; and, to demonstrate that each engine rotor has sufficient strength to withstand a maximum overspeed condition without cracking.
Incorporations: None.
References: None.
If the engine bill-of-material includes instrumentation (including sensors), then such must be designed and must function so as to assure safe operation, (structural integrity of the inservice engine). Note that the rule requires the most critically stressed rotor component (including discs, spacers, cooling lates) of each turbine, compressor, and tan to be tested at a naximum overspeed condition for 5 minutes. Determination of the maximum overspeed condition needs to consider the 6 cases defined in the rule -- including shaft failures which decouple a turbine rotor from its load. Rotor shafts are not considered prime reliable in that it is assumed that shaft failures will occur, both as a single failure and in combination with another undetected or dormant failure.
b. A rotor burst test, in lieu of the presently required 5- minute overspeed test, could be found as an acceptable equivalent means of compliance. In the case of output shaft (low) turbines, Chap 3 46 Par 27
the rotor burst test would be performed, in conjunction with an abrupt load decouple test, to demonstrate the existence of adequate margin between the peak speed and the burst speed. The actual burst speed must be corrected to account for the effects of Spect non actual material propertise, very poet fication timum material properties. Acceptance criteria must show at least a 5.0 percent speed margin between the abrupt decouple peak speed and correlates fe lovin taegity ovided sna hits qui reference, including significant differences:
Chap 3 Par 27 47
TABLE 27-1. ROTOR INTEGRITY REGULATIONS/GUIDANCE
REGULATION & CORRESPONDING SIGNIFICANT GUIDANCE
EFFECTIVITY DATE GUIDANCE & DIFFERENCES
ISSUE DATE FAR 33-13, 08/18/89 AC 33-3, • POST-TEST ACCEPTANCE FAR 33-12, 10/03/88 09/09/68 CRITERIA OF THE FAR TAKES FAR 33-11, 04/24/86 PRECEDENCE OVER THAT OF FAR 33-10, 03/26/84 THE AC.
FAR 33-9, 10/14/80 FAR 33-8, 05/02/77 FAR 33-7, 02/01/77 FAR 33-6, 10/31/74 FAR 33-5, 03/01/74 AC 33-3, • FAR DOES NOT FAR 33-4, 04/23/71 IDENTIFY POST-TEST ACCEPTANCE CRITERIA.
FAR 33-3, 04/03/67 AC 20-26, • POST-TEST ACCEPTANCE FAR 33-2, 07/06/66 CH.2 CRITERIA OF THE AC 07/08/65 REVISED BY CH.2.
(PARAGRAPH 4C).
FAR 33 & AC 20-26 • SUPERSEDES FAA 33-1, 02/01/65 CH. 1, 09/18/64; MEMORANDUM OF 02/09/59.
AC 20-26, • OVERSPEED TEST 07/22/64 CONDITION ANALYSIS BASED ON FAILURE (S) .
CAR 13-6, 04/22/64 FAA POLICY • ESTABLISHES CAR 13-5, 02/12/63 MEMORANDUM, REQUIREMENTS FOR CAR 13-4, 05/03/62 02/09/59 SUBSTANTIATION OF TURBINE CAR 13-3, 10/01/59 & COMPRESSOR ROTOR INTEGRITY. NOTE THAT THE
OVERSPEED TEST CONDITION
TO BE RUN IS BASED ON
ANALYSIS OF CONTROL
SYSTEM FAILURE(S) AND ON
A FACTOR OF TAKEOFF R.P.M., WHICHEVER YIELDS THE GREATEST STRESS.
• MUST DEMONSTRATE
OVERSTRESS MARGIN, BUT
DOES NOT OFFER ACCEPTANCE CRITERIA.
CAR 13-2, 05/17/58 NONE N/A CAR 13-1, 08/12/57 CAR 13, 06/15/56 Chap 3 48 Par 27
28. Section 33.29, Instrument Connection.
Section 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 of the engine in compliance with any engine limitation airworthiness regulations or necessary to insure operation no identity we 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.
*** Guidance. The INTENT of this section is to prevent misconnections of engine-required instrumentation, and to provide a drawing location for rotor unbalance sensing.
Incorporations: None.
References: None.
An acceptable method of complying with FAR Section 33.29 (b) is to Haterically, turbojet has ben interpreted to sure inelude ton Chap 3 Par 28 49 (and 50)
SECTION 3. SUBPART C--DESIGN AND CONSTRUCTION; RECIPROCATING
AIRCRAFT ENGINES
29. Section 33.33, Vibration.
Section 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.
*** Guidance. The INTENT of this section is self-evident.
Incorporations: None.
References: None.
30. Section 33.35, Fuel and Induction System.
Section 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.
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.
must show for each fluid injection (other than fuel) fluid is adequately controlled.
Chap 3 Par 29 51
Guidance. The INTENT of this section is self-evident.
Incorporations: None.
References: None.
31. Section 33.37, Ignition System.
Section 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.
*** Guidance. The INTENT of this section is self-evident.
Incorporations: None.
References: None.
The engine installation instructions should define the characteristics and interface requirements of all aircraft sources of electrical power required by the engine ignition system.
32. Section 33.39, Lubrication System.
Section 33.39 Lubrication system.
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. The lubrication system of the engine must be cutted to allow instaling a meane of cooling the lubricant.
(c) The crankcase must be vented to the atmosphere to preclude leakage of oil from excessive pressure in the crankcase.
*** Guidance. The INTENT of this section is self-evident.
Incorporations: None.
References: None.
Chap 3 52 Par 30