AC 21 .17-4 Appendix A SUBPART I—ELECTRIC ENGINE PL.3305 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 approval 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 engines having one or more engine inoperative ratings or other contingency 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 PL.3375(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.
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AC 21.17-4 Appendix A PL.3307 Engine ratings and operating limitations.
(a) Engine ratings and operating limitations are established by the Administrator and included in the type certificate data sheet specified in 14 CFR 21.41, 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 electric engines, ratings and operating limits are established relating to the following:
(1) Shaft power, torque, or thrust; rotational speed; temperature; and time for:
(i) Rated takeoff power o r thrust;
(ii) Rated maximum continuous power or thrust; and
(iii) Rated maximum temporary power or thrust a nd associated time limit.
(2) Duty cycle and the rating at that duty cycle. The duty cycle must be declared in the type certificate data sheet.
(3) Cooling fluid grade or specification.
(4) Power-supply requirements.
(5) Any other ratings or limitations that are necessary for the safe operation of the engine.
(c) In determining the engine performance and operating limitations, the overall limits of accuracy of the engine control system, of the engine electrical systems, and of the necessary instrumentation as defined in PL.3305(a)(6) must be taken into account.
PL.3308 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.
PL.3315 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.
PL.3317 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.
(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 A-29
AC 21.17-4 Appendix A protection unless damage by fire will not cause leakage or spillage of a hazardous quantity of flammable fluid. For an 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.
(h) High-voltage electrical wiring interconnection systems must be protected against arc faults that can lead to a fire that could result in hazardous engine effects as defined in PL.3375(g)(2).
Any non-protected electrical wiring interconnects must be analyzed to show that arc faults do not cause a hazardous engine effect.
PL.3319 Durability.
(a) The engine design and construction must minimize the development of an unsafe condition of the engine between maintenance intervals, overhaul periods, or mandatory actions described in the applicable Instructions for Continued Airworthiness.
(b) Each component of the propeller blade pitch control system which is a part of the engine type design must meet the requirements of sections PL.3521, PL.3523, PL.3542 and PL.3543.
PL.3321 Engine cooling.
(a) Engine design and construction must provide the necessary cooling under conditions in which the aircraft is expected to operate.
(b) If cooling is required to satisfy the safety analysis as described in PL.3375, the cooling system monitoring features and usage must be documented in accordance with the requirements in PL.3305.
PL.3323 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 A-30
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(2) The specified ultimate loads without failure, but may exhibit permanent deformation.
(c) Engine structure necessary to transmit loads from the engine-driven lift devices to the airframe must comply with all applicable design load conditions and safety factors defined in subpart C.
PL.3325 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.
PL.3326 Engine electrical systems.
(a) Applicability. Any system or device that provides, uses, conditions, or distributes electrical power, and is part of the engine type design, must provide for the continued airworthiness of the engine and must maintain electric engine ratings.
(b) Electrical systems. The electrical system must ensure the safe generation and transmission of power, and electrical load shedding, if required, and that the engine does not experience any unacceptable operating characteristics or exceed its operating limits. High voltage electrical wiring interconnection systems must be protected against arc faults that could result in hazardous engine effects, as defined in PL.3375(g)(2).
(c) Electrical-power distribution.
(1) The engine electrical-power distribution system must be designed to provide the safe transfer of electrical energy throughout the electric engine. The system must be designed to provide electrical power so that the loss, malfunction, or interruption of the electrical power source will not result in a hazardous engine effect, as defined in PL.3375(g)(2).
(2) The system must be designed and maintained to withstand normal and abnormal conditions during all ground and flight operations.
(3) The system must provide mechanical or automatic means of isolating a faulted electricalenergy generation or storage device from affecting the safe transmission of electric energy to the electric engine or detrimental engine effects in the intended aircraft application.
(d) Protection systems. The engine electrical system must be designed such that the loss, malfunction, interruption of the electrical power source, or power conditions that exceed design limits will not result in hazardous engine effects, as defined in PL.3375(g)(2).
(e) Electrical Power Characteristics. The applicant must identify, declare, and document in accordance with the requirements in P L.3305 the characteristics of any electrical power supplied from—
(1) The aircraft to the engine electrical system, for starting and operating the engine, including transient and steady-state voltage limits, and A-31
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(2) The engine to the aircraft via energy regeneration, and any other characteristics necessary for safe operation of the engine.
(f) Environmental limits. Environmental limits that cannot adequately be substantiated by endurance demonstration, validated analysis, or a combination thereof must be demonstrated by the system and component tests in PL.3391.
(g) Electrical-system failures. The engine electrical system must—
(1) Have a maximum rate of loss of power control (LOPC) that is suitable for the intended aircraft application;
(2) When in the full-up configuration, be single fault tolerant, as determined by the Administrator, for electrical, electrically detectable, and electronic failures involving LOPC events;
(3) Not have any single failure that results in hazardous engine effects as defined in PL.3375(g)(2); and
(4) Ensure any electrical system failures or malfunctions that lead to local events in the intended aircraft application do not result in hazardous engine effects as defined in PL.3375(g)(2).
(h) System safety assessment. The applicant must perform a system safety assessment. This assessment must identify faults or failures that affect normal operation, together with the predicted frequency of occurrence of these faults or failures. The intended aircraft application must be taken into account to assure the assessment of the engine system safety is valid. The rates of hazardous and major faults must be declared and documented in accordance with the requirements in PL.3305.
PL.3327 Overspeed.
(a) A rotor overspeed must not result in a burst, rotor growth, or damage that results in a hazardous engine effect, as defined in PL.3375(g)(2). Compliance with this paragraph must be shown by test, validated analysis, or a combination of both. Applicable assumed rotor speeds must be declared and justified.
(b) Rotors must possess sufficient strength with a margin to burst above approved operating conditions and above failure conditions leading to rotor overspeed. The margin to burst must be shown by test, validated analysis, or a combination thereof.
(c) The engine must not exceed the rotor-speed operational limitations that could affect rotor structural integrity.
PL.3328 Engine control systems.
(a) Applicability. These requirements apply to any system or device that is part of the engine type design that controls, limits, monitors, or protects engine operation and is necessary for the continued airworthiness of the engine.
(b) Engine control. The engine control system must ensure that the engine does not experience any unacceptable operating characteristics or exceed its operating limits, including in failure conditions where the 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, if applicable.
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(c) Design assurance. The software and complex electronic hardware, including programmable logic devices, must be—
(1) Designed and developed using a structured and systematic approach that provides a level of assurance for the logic commensurate with the hazard associated with the failure or malfunction of the systems in which the devices are located; and
(2) Substantiated by a verification methodology acceptable to the Administrator.
(d) Validation. All functional aspects of the control system must be substantiated by test, analysis, or a combination thereof, to show that the engine control system performs the intended functions throughout the approved flight envelope.
(e) Environmental limits. Environmental limits that cannot be adequately substantiated by endurance demonstration, validated analysis, or a combination thereof must be demonstrated by the system and component tests in PL.3391.
(f) Engine control system failures. The engine control system must—
(1) Have a maximum rate of LOPC that is suitable for the intended aircraft application. The estimated LOPC rate must be documented in accordance with the requirements in PL.3305;
(2) When in the full-up configuration, be single fault tolerant, as determined by the Administrator, for electrical, electrically detectable, and electronic failures involving LOPC events;
(3) Not have any single failure that results in hazardous engine effects as defined in PL.3375(g)(2); and
(4) Ensure failures or malfunctions that lead to local events in the aircraft do not result in hazardous engine effects as defined in PL.3375(g)(2) due to engine control system failures or malfunctions.
(g) System safety assessment. The applicant must perform a system safety assessment. This assessment must identify faults or failures that affect normal operation, together with the predicted frequency of occurrence of these faults or failures. The intended aircraft application must be taken into account to assure the assessment of the engine control system safety is valid.
The rates of hazardous and major faults must be documented in accordance with the requirements in PL.3305.
(h) Protection systems. The engine control devices and systems’ design and function, together with engine instruments, operating instructions, and maintenance instructions, must ensure that engine operating limits that can lead to a hazard will not be exceeded in-service.
(i) Aircraft-supplied data. Any single failure leading to loss, interruption, or corruption of aircraft-supplied data (other than power command signals from the aircraft), or aircraft-supplied data shared between engine systems within a single engine or between fully independent engine systems, must—
(1) Not result in a hazardous engine effect, as defined in PL.3375(g)(2), for any engine installed on the aircraft; and
(2) Be able to be detected and accommodated by the control system.
(j) Engine control system electrical power.
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AC 21.17-4 Appendix A
(1) The engine control system must be designed such that the loss, malfunction, or interruption of the control system electrical power source will not result in a hazardous engine effect, unacceptable transmission of erroneous data, or continued engine operation in the absence of the control function. Hazardous engine effects are defined in PL.3375(g)(2). The engine control system must be capable of resuming normal operation when aircraft-supplied power returns to within the declared limits.
(2) The applicant must identify, and document in accordance with the requirements in P L.3305, the characteristics of any electrical power supplied from the aircraft to the engine control system, including transient and steady-state voltage limits, and any other characteristics necessary for safe operation of the engine.
PL.3329 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) In addition, as part of the system safety assessment of PL.3328(g) and PL.3326(h), the applicant must assess the possibility and subsequent effect of incorrect fit of instruments, sensors, or connectors. Where practicable, the applicant must take design precautions to prevent incorrect configuration of the system.
(c) 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.
(d) The applicant must provide instrumentation enabling the flight crew to monitor the functioning of the engine cooling system unless 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, as defined in PL.3375(g)(2), before detection; or
(3) The probability of failure of the cooling system is extremely remote.
(e) 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.
PL.3362 Stress analysis.
(a) A mechanical and thermal stress analysis, as well as an analysis of the stress caused by electromagnetic forces, must show a sufficient design margin to prevent unacceptable operating characteristics and hazardous engine effects as defined in PL.3375(g)(2).
(b) Maximum stresses in the engine must be determined by test, validated analysis, or a combination thereof, and must be shown not to exceed minimum material properties.
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AC 21 .17-4 Appendix A PL.3363 Vibration.
The engine must be designed and constructed to function throughout its operating range of rotational speeds and engine output power/thrust, including defined exceedances, without inducing excessive stress in any of the engine parts because of vibration and without imparting excessive vibration forces to the aircraft structure.
PL.3364 Liquid and Gas Systems.
(a) Each system used for lubrication or cooling of engine components must be designed and constructed to function properly in all flight attitudes and atmospheric conditions in which the engine is expected to operate.
(b) If a system used for lubrication or cooling of engine components is not self-contained, the interfaces to that system must be defined and documented in accordance with the requirements in PL.3305.
(c) The applicant must establish by test, validated analysis, or a combination of both, that all static parts subject to significant pressure loads will not:
(1) Exhibit permanent distortion beyond serviceable limits or exhibit leakage that could create a hazardous condition when subjected to normal and maximum working pressure with margin.
(2) Exhibit fracture or burst when subjected to the greater of maximum possible pressures with margin.
(d) Compliance with paragraph (c) of 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 physical geometry conditions allowed by the type design, such as minimum material and minimum radii.
(e) Approved coolants and lubricants must be documented in accordance with the requirements in PL.3305.
PL.3370 Critical and life-limited parts.
(a) The applicant must show, by a safety analysis or means acceptable to the Administrator, whether rotating or moving components, bearings, shafts, static parts, and non-redundant mount components should be classified, designed, manufactured, and managed throughout their service life as critical or life-limited parts.
(1) Critical part means a part that must meet prescribed integrity specifications to avoid its primary failure, which is likely to result in a hazardous engine effect as defined in PL.3375(g)(2).
(2) Life-limited parts may include but are not limited to a rotor or major structural static part, the failure of which can result in a hazardous engine effect, as defined in PL.3375(g)(2), due to lowcycle fatigue (LCF) mechanism. A life limit is an operational limitation that specifies the maximum allowable number of flight cycles that a part can endure before the applicant must remove it from the engine.
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(3) For critical and life-limited parts of engines powering engine-driven lift devices, the related engine structure necessary to transmit loads from the device to the airframe must comply with loading spectra defined in PL.2210.
(4) For critical and life-limited parts of engines powering engine-driven lift devices, the related non-metallic engine structure necessary to transmit loads from the device to the airframe must comply with all applicable fatigue and damage tolerance requirements defined in PL.2240.
(b) In establishing the integrity of each critical part or life-limited part, the applicant must provide to the Administrator the following three plans for approval:
(1) 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. In addition, the engineering plan must contain the steps to ensure each engine critical part is withdrawn from service 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 and critical 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 PL.1529.
(2) A manufacturing plan that identifies the specific manufacturing constraints necessary to consistently produce each engine life-limited and critical part with the attributes required by the engineering plan.
(3) A service management plan that defines in-service processes for maintenance and the limitations to repair for each engine life-limited and critical 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.
PL.3371 Lubrication system.
(a) The lubrication system must be designed and constructed to function properly between scheduled maintenance intervals in all flight attitudes and atmospheric conditions in which the engine is expected to operate.
(b) The lubrication system must be designed to prevent contamination of the engine bearings and lubrication system components.
(c) The applicant must demonstrate by test, validated analysis, or a combination thereof, the unique lubrication attributes and functional capability of paragraphs (a) and (b) of this section.
PL.3373 Power or thrust response.
(a) The design and construction of the engine, including its control system, must enable an increase—
(1) From the minimum power o r thrust setting to the highest rated power or thrust without detrimental engine effects;
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(2) From the minimum obtainable power o r thrust while in flight, and while on the ground, to the highest rated power or thrust within a time interval determined to be appropriate for the intended aircraft application; and
(3) From the minimum torque to the highest rated torque without detrimental engine effects in the intended aircraft application.
(b) The transient response time for each respective power condition in paragraph (a)(1), (a)(2), and (a)(3) of this section must be documented in accordance with the requirements in P L.3305.
PL.3374 Continued rotation.
If the design allows any of the engine main rotating systems to continue to rotate after the engine is shut down while in-flight, this continued rotation must not result in any hazardous engine effects, as defined in PL.3375(g)(2).
PL.3375 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 the intended aircraft application. 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.
(b) The applicant must comply with paragraphs (a)(1) and (a)(2) of this section using the failure definitions in paragraph (g) of this section.
(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 such as PL.3315, PL.3327, PL.3370, 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 documented in accordance with the requirements in PL.3305.
(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.
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(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 ICA required under PL.1529. 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 PL.3305.
(f) The applicant must comply with paragraphs (d) and (e) of this section using the failure definitions in paragraph (g) of this section.
(g) Unless otherwise approved by the Administrator, the following definitions apply to the engine effects when showing compliance with these airworthiness criteria:
(1) A minor engine effect means the engine, despite its failure, performs its intended functions in accordance with paragraphs (g)(1)(i) through (g)(1)(iii) of this section and complies with the operability requirements of PL.3373 and PL.3389, as appropriate, with the exception of a slight reduction in functional capabilities.
(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 approved flight envelope;
(ii) Allows modulation of engine power or thrust with adequate sensitivity over the declared range of engine operating conditions; and
(iii) Does not create unacceptable power or thrust oscillation.
(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;
(vii) Complete inability to shut the engine down;
(viii) Electrocution of the crew, passengers, operators, maintainers, or others; and
(ix) Blockage of cooling systems that could cause the engine effects described in paragraphs
(g)(2)(i) through (g)(2)(viii) of this section.
(3) Any other engine effect is a major engine effect.
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(h) The applicant must take the intended aircraft application into account when performing the safety analysis.
(i) The results of the safety analysis, and the assumptions about the aircraft application used in the safety analysis, must be documented in accordance with the requirements in PL.3305.
PL.3377 Ingestion.
(a) Rain, ice, and hail ingestion must not result in an abnormal operation such as shutdown, power or thrust loss, erratic operation, or power oscillations throughout the engine operating range.
(b) Ingestion from other likely sources (birds, hailstones, foreign objects—ice slabs) must not result in unacceptable power or thrust loss or hazardous engine effects as defined in PL.3375(g)(2).
(c) If the design of the engine relies on features, attachments, or systems that the installer may supply, for the prevention of unacceptable power or thrust loss or hazardous engine effects as defined in PL.3375(g)(2) following potential ingestion, then the features, attachments, or systems must be documented in accordance with the requirements in P L.3305.
PL.3383 Vibration demonstration.
Each engine design must undergo a vibration survey to establish that the vibration characteristics of those components subject to induced vibration are acceptable throughout the approved flight envelope and engine operating range for the specific installation configuration. The possible sources of the induced vibration that the survey must assess are mechanical, aerodynamic, acoustical, internally induced electromagnetic, installation induced effects that can affect the engine vibration characteristics, and likely environmental effects. This survey must be shown by test, validated analysis, or a combination thereof.
PL.3384 Overtorque.
When approval is sought for a transient maximum engine overtorque, the applicant must demonstrate by test, validated analysis, or a combination thereof, that the engine can continue operation after operating at the maximum engine overtorque condition without maintenance action. Upon conclusion of overtorque tests conducted to show compliance with this subpart, or any other tests that are conducted in combination with the overtorque test, each engine part or individual groups of components must meet the requirements of PL.3393.
PL.3385 Calibration assurance.
Each engine must be subjected to calibration tests to establish its power characteristics and the conditions both before and after the endurance and durability demonstrations specified in PL.3387 and PL.3390.
PL.3387 Endurance demonstration.
(a) The applicant must subject the engine to an endurance demonstration, acceptable to the Administrator, to demonstrate the engine’s limit capabilities.
(b) The endurance demonstration must include increases and decreases of the engine’s power settings, energy regeneration, and dwellings at the power settings and energy regeneration for sufficient durations that produce the extreme physical conditions the engine experiences at rated A-39