lubricant pressures and temperatures; time at ratings; and number of excursions at ratings.
Production and Overhauled Engines. Although discussion of production engines would normally be considered beyond the scope of type certification, the following information may be useful:
(1) Each engine newly produced under a type certificate (TC) only, or under a production certificate, must be tested to assure that engine output is not less than the TC ratings and within each rating's respective associated operating limitations.
(See References 3 and 4 above).
(2) It should be recognized that engine manufacturers and performance margin. overhaulers will typically deliver engines with some degree of tween the engine suplier This margin is usually based on a business agreement between the engine supplier and the installer or user.
(i) The important point is that there is no regulatory basis for engines to have more than minimum type certificated (specification) power/thrust available at delivery, whether in new or overhauled condition.
(ii) A further point is that in-service engines must always be capable of delivering the type certificate rating minimum performance, again, within the respective rating's operating limitations. Otherwise, the engine must be removed from further service and minimum performance restored before return to service.
(3) There are special cases wherein some in-service engine applications have been approved for subsequent operations to lesser performance levels than established during initial type certification. These cases are known as having deteriorated power operation approvals. Such cases have been approved only following extensive reevaluation of the engine and the aircraft to the deteriorated performance level, against the original type certification basis. This approval is limited to the specific initial operator's route structure.
19. Section 33.8, Selection of Engine Power and Thrust Ratings.
Section 33.8 Selection of engine power and thrust ratings.
(a) Requested engine power and thrust ratings must be selected by 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.
*** Chap 3 34 Par 18
Guidance. The INTENT of this section is self-evident.
Incorporations: None.
References: None.
basis, the power or thrust ratings which new engines, aed eratings restored to new condition, are required to produce. objective of this requirement is to establish, on a uniform are used in establishing aircraft performance for which consistent minimum performance is desired. Ratings selected by the applicant are established by FAA block tests. Ratings are typically based on specific atmospheric conditions, for the static case, with no customer bleed, no power extraction, and 100 percent recovery of inlet and exhaust losses. The applicant's experience on the calibrated output of a number of engines, of both the applicable new model and of engines of closely similar design characteristics, should be reviewed to evaluate the range of output expected in all other engines of that type. The selected ratings should be consistent with the output of the lowest output engine anticipated.
Chap 3 Par 19 35 (and 36)
SECTION 2. SUBPART B--DESIGN AND CONSTRUCTION; GENERAL
20. Section 33.14, Start-Stop Cyclic Stress (Low-Cycle Fatique).
Section 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 startstop stress cycle consists of a flight cycle profile or an equivalent representation of engine usage. It includes 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 maximum rated power or thrust and after engine shutdown, unless it is shown that the parts undergo the same stress range without temperature stabilization.
*** Guidance. The INTENT of this section is to establish cyclic life limits using FAA-approved methodology for rotating structural components whose failure could produce a hazard to the aircraft.
Incorporations: AC 33.3, Turbine and Compressor Rotors Type Certification Substantiation Procedures, dated September 9, 1968.
References: None.
The guidance embodied in the "Incorporations" above, with respect to determination of service life limits, is still considered valid.
a. Some applicants have pre-approved procedures on file with the FAA for establishing initial (and increases of) low cycle fatigue (ICF) lives. Such procedures should be utilized where they exist.
b. The FAA requires the applicant to define the ICF verification test cycle, based on the most severe mission cycle expected in service. The determination of the most severe mission cycle expected in service is based on the concurrence of the FAA engine certification project manager with the applicant's proposed definition.
(1) It is determined logically, based on the type of engine, types of airframe applications, and anticipated service Chap 3 Par 20 37
usage. In rare cases, specific airframe applications during the time of the engine certification program are not known, and, therefore, more speculative, but conservative, approaches may need to be used. In all cases, the FAA's installation office may be consulted to assist in determining the mission cycie.
The test cycle basis should, in part, incorporate some excursions to the redline limits of gas path temperature, rotational speeds, torques, pressures, etc., associated with the maximum ratings used in the mission cycle. Consider such excursions that could reasonably be expected to occur within the extremes of ambient temperature and takeoff altitude conditions encountered during the engine's service life.
(3) The supporting data required for ICF qualification includes mechanical and thermal stress analyses and tests, and material analyses and tests. Since there are various acceptable methods of determining maximum temperature gradients, the FAA does not advocate that any particular methodology be applied for thermal analyses. Test engine rotor structural parts should attain a stabilized temperature for each applied test cycle.
Increases in service life limits, similar to the determination of initial life limits, will need to be based representative severe operating cycle, identified from service experience. Programs to increase service life limits should be based on further cyclic testing and analysis of in-service parts that have attained a reasonable percentage of their life limits.
d. FAA findings of compliance with existing FAR Section 33.14 are based, in part, on the applicant's definition of the flight cycle profile, or equivalent representation of engine usage.
Implementation of this aspect of the rule, in the case of OEI rated engines, requires addressing the use of OEI ratings at a realistic frequency representative of actual (expected) inservice use.
(1) To date, the service experience with OEI rated engines is that use of OEI ratings, following a true case of one engine having failed, is an extremely rare event. Thus, it is considered totally unrealistic to mandate the use of OEI ratings as a constituent of each flight cycle.
(2) However, the applicant is still required, in all cases, to account for low cycie fatique effects, based on the anticipated usage of OEI ratings during the life of the engine.
This could be accomplished, for example, by adding a reasoned finite number of cycles to the expended life of the appropriate components for each OEI power excursion.
Chap 3 38 Par 20
21. Section 33.15, Materials.
Section 33.15 Materials.
The suitability and durability of materials used in the engine must--
(a) Be established on the basis of experience or
(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.
*** Guidance. The INTENT of this section is self-evident.
Incorporations: AC 33-4, 'Design Considerations Concerning the Use of Titanium in Aircraft Turbine Engines, dated July 28, 1983.
References: None.
The applicant should select approved specification materials in the design and construction of the engine to minimize the development of an unsafe condition of the engine. Although the regulations do not specifically address corrosion and deterioration, these factors need to be considered in the type certification process, particularly in the design stage. Examples are proposed use of non-corrosion resistant alloys in corrosionsusceptible areas and proposed use of galvanic materials in contact.
22. Section 33.17, Fire Prevention.
Section 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) or this section, each external line, fitting, anc ›ther component, which contains or conveys flammable fluic 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 Chap 3 Par 21 39
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