certifications will be accurately and rigorously observed during ground and flight operations in an aircraft. This paragraph is intended to promote this concept.
(2) Procedures. Primary engine limitations in the form of time, gas temperature, torque, and rotational speed and their corresponding allowable transient values are defined in the approved engine installation manual. The rotorcraft manufacturer must provide reliable, accurate means to assure that these limitations are not exceeded. These means may be in the form of automatic limiters or by crew monitoring of appropriately marked instruments. The FAA/AUTHORITY powerplant certification engineer and the rotorcraft manufacturer's staff should verify these aspects by:
Evaluating all applicable instrument, indicator, or warning devices, including transmitters, and limiting devices, if any, for system tolerances.
(Ii) Closely reviewing the component qualification reports of items in c(2)(1) above to verify that these devices are properly qualified and that any deviations are acceptable.
(ill) Assuring that maintenance data are provided for functional checks and calibration of instruments and devices which are used to monitor or protect critical turbine rotor limitations. Preflight checks for automatic limiter devices may be appropriate.
(iv) Verifying that instrument markings are clear and relatively simple, that corresponding flight manual instructions and descriptions are straightforward and complete, and that instruments are located and orientated to minimize the probability of reading error.
398A. $29.903 (Amendment 29-26) ENGINES.
a. Explanation. Amendment 29-26 adds § 29.903(a) that requires reciprocating engines used in rotorcraft to be certified in accordance with the rotorcraft engine testing requirements in § 33.49(d). This change is incorporated to ensure that certification requirements are not overlooked when reciprocating engines are installed in rotorcraft to be certified under Part 29 requirements. Section 29.903(b)(2) was revised to identify and clarify crew action; i.e., normal pilot action allowable with primary flight controls, in determining if adequate powerplant systems isolation is provided. This change eliminates any possible confusion that may exist regarding the acceptability of modifying optimum flight control manipulation to protect engine parameters.
Section 29.903(c)(3) was added and requires engine restart capability to be available throughout the flight envelope appropriate to the rotorcraft. This will avoid the concept that an in-flight engine restart envelope constitutes acceptable compliance with this rule.
618 Par 398
b. Procedures.
(1) Engine type certification. All engines installed in rotorcraft should have a type certificate. The specific certification requirements for installation of reciprocating engines in rotorcraft are found in Part 33. Engines certificated under other approved certification rules (CAR Part 13 and FAR § 21.29, for imported engines) are also eligible. If a component, system, or arrangement is certified under Part 33 or other requirement, the applicant is not relieved of the necessity to comply with the requirements of Part 29. If the component, system, or arrangement, supplied as a part of a certificated engine, meets the Part 33 and Part 29 requirements, subsequent changes to these components, systems, or arrangements could negate compliance with Part 29. For example, an engine may initially be equipped by the engine manufacturer with an oil tank filler cap that complies with the Category A requirements of § 29.1013(c)(2) but is subsequently changed to a simpler and less expensive cap that complies with § 33.71(c)(4). The airframe manufacturer should ensure that the requirements of § 29.1013(c)(2) are maintained.
(2) Category A: control of engine rotation. Section 29.903(c)(3) requires an engine restart capability which is appropriate to the rotorcraft. The minimum envelope for the restart capability should be equal to or better than the rotorcraft takeoff/landing maximum altitude and temperature limits. Compliance is usually shown by conducting actual in-flight restarts during flight tests and/or other tests in accordance with an approved test plan. Restarts should be conducted at various altitudes, ambient that the engine start can be accomplished without jeopardizing continued safe operation of the rotorcraft. Pilot workload for a preexisting one-engine-inoperative situation, the location of the restart system controls, and the availability of a second pilot should be considered. The emergency/malfunction instruction sections of the rotorcraft flight manual (RFM) should present a detailed definition of the approved restart envelope and detailed instructions for the restart. Eligible ambient atmospheric conditions, prestart requirements (to allow for waste fuel drainage), starter duty cycle (if different from the ground start duty cycle), and prestart situation analysis should be included. The prestart situation analysis should consider the following questions:
- Should a restart be attempted in view of the cause for initial shutdown?
Is inlet system ice ingestion a possibility?
Is reignition of fuel in the engine nacelle a possibility?
Is sufficient restart time available?
- Is power available?
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- Is altitude sufficient to maintain terrain clearance?
Although restart capability from an all-engines-out flight condition is not required, special instructions for restarting from this situation should also be included commensurate with the system capability to accomplish the starts.
398B. § 29.903 (Amendment 29-31) ENGINES.
a. Explanation. Amendment 29-31 clarified the requirements for control of engine rotation and in-flight restart of engines. Section 29.903(c)(1) was changed by adding the word "or" at the end of the paragraph, which provided an option on how to protect the engine stopping system from fire.
b. Procedures. All of the policy material pertaining to this section remains in effect. Additionally, the new § 29.903(e) requires that any engine should have a restart capability that has been demonstrated throughout a flight envelope to be certificated for the rotorcraft. The minimum restart envelope for Category A rotorcraft is discussed in Paragraph 398A of this AC. The restart capability can consider windmilling of the engine as part of this restart capability; however, most rotorcraft airspeeds and the locations of the engines do not support engine windmilling up to start speeds. Only electrical power requirements were considered for restarting; however, other factors that may affect this capability are permitted to be considered. Engine restart capability following an in-flight shutdown of all engines is the primary requirement, and the means of providing this capability is left to the applicant.
398C. § 29,903 (Amendment 29-36) TURBINE ENGINE INSTALLATION.
a. Explanation. Amendment 29-36 revises § 29.903(d) to require that design precautions should be taken to minimize hazards to the rotorcraft in the event of an engine failure.
b. Procedures. Appendix 3 to this AC provides information relative to methods of compliance for this section.
399. § 29.907 ENGINE VIBRATION.
a. Explanation. This very generalized requirement is authority to require substantiation of the effects of vibration on any part of the engine or the rotorcraft. In normal certification practice, the vibration effects of concern to the powerplant engineer are the vibratory loads or stresses in the engine and in the rotor drive system. Vibration effects on the rotor drive system are of concern if the corresponding loads or stresses result in fatigue damage. This aspect, however, is adequately addressed in § 29.571.
Vibration effects on the engine are usually categorized as "installation vibration" and 620 Par 398A
"torsional vibration." Methods of evaluation and limitations of these vibrations are established by the engine manufacturer.
b. Procedures. Review Order 8110.9, Handbook on Vibration Substantiation and Fatigue Evaluation of Helicopter and other Power Transmission Systems. Note that the mechanical coupling of the engines to the rotor drive system creates, for torsional vibration considerations, one, rather complicated, drive system which responds to any forced or resonant frequency. Antinodes or nodes and frequencies may exist in the engine shaft which are absent when the engine is operated on a test stand; therefore, the vibration investigation conducted under Part 33 is not conclusive with respect to torsionals. As noted in Order 8110.9, the engine manufacturers' assistance is necessary to find compliance. Section 29.571 was amended by Amendment 29-13 to include "rotor drive systems between the engines and the rotor hubs" as part of the flight structure. This rule supplements § 29.907 and requires coordination with the structures certification engineer to avoid duplication of effort by the rotorcraft manufacturer. Advisory Circular 20-95, Fatigue Evaluation of Rotorcraft Structure, which provides acceptable methods of compliance with § 29.571, may also be used to find compliance with § 29.907. In addition to basic drive system components such as main and auxiliary rotor drive shafts, the vibratory evaluation should include couplings, gear teeth, gear cases and splines, and should consider, where appropriate, low cycle fatigue associated with ground-air-ground cycles.
400. § 29.908 (Amendment 29-13) COOLING FANS.
a. 29.908(a):
(1) Explanation. This paragraph applies to Category A rotorcraft and is intended to require that powerplant area cooling fans be designed and installed to enable continued safe operation of the rotorcraft after failure of a cooling fan blade.
The phrase "except that the loss of cooling need not be considered" at the end of this paragraph is intended to make clear that for the purposes of this section, the FAA/AUTHORITY is concerned only with the fragmentation effect of a fan blade failure (reference Preamble Item 3-64 of Amendment 29-12).
(2) Procedures. If a fan shroud is provided, the applicant may demonstrate that the shroud configuration and strength are adequate to contain a failed fan blade and any other fan blades, guide vanes, etc., which can be expected to fail sequentially to the initial blade failure. The demonstration can be facilitated by making a saw slot at the root of a blade sufficiently deep to weaken the blade retention strength and create a failure while the fan is rotating at the maximum speed established for the test. If the fan is driven by the rotor drive system, the test speed should be equal to or above the maximum transient speed to be expected with the rotor system. If the fan is driven by other means; i.e., bleed air turbine, hydraulic motor, engine N, turbine, etc., the rotational speed for the blade failure demonstration should be based on a critical analysis of speed regimes to be expected. Containment is not required if the fan is Par 399 621
located so that blade failure (and any sequential fan component failure) will not jeopardize safety. This may be shown by test or analysis. Segment shielding would likely be involved.
b. § 29,908(b):
(1) Explanation. This paragraph applies to Category B rotorcraft and is intended to provide safety to the rotorcraft in the event of an assumed cooling fan blade failure or to prescribe a test to show that the cooling fan blade retention means is sufficient that blade failure is not a consideration.
(2) Procedures.
compliance with this section. If $ 29.908(b)(1) is selected, follow the procedures The applicant may select § 29.908(b)(1), (b)(2), or (b)(3) to show outlined above for Category A rotorcraft.
Section 29.909(b)(2) may be selected; however, without containment, damage to any component or structure in the plane of the fan rotor or any other trajectory to be expected should not cause the loss of any function essential to a controlled landing.
(ili) If § 29.908(b)(3) is selected, a spin test at 122.5 percent of the maximum speed associated with either engine terminal speed or an overspeed limiting device would be acceptable to show compliance. No failure should occur, and distortion should not result in fan element contact with housings or other adjacent components. (Note: 150 percent of the centrifugal force is achieved at 122.5 percent of the rotational speed.)
400A. § 29.908 (Amendment 29-26) COOLING FANS.
a. Explanation. Amendment 29-26 requires that cooling fans be designed and installed to enable continued safe flight and adequate cooling of the rotorcraft following a fan blade failure. Compliance with the previous requirements could have resulted in hazards to the rotorcraft with the loss of cooling air to critical powerplant components.
A new section was also added to the rule for cooling fans, which are not part of the powerplant installation and therefore not subject to the fatigue evaluation under § 29.571. It should be determined that no cooling fan blade resonant conditions exist within the operating limits of the rotorcraft unless a fatigue evaluation is conducted.
b. Procedures. Neither mechanical damage nor loss of cooling air should prevent "continued safe flight." The definition of "continued safe flight" is contained in Appendix 1 of AC 20-136 and is quoted as follows:
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Continued safe flight and landing. This phrase means that the aircraft is capable of safely aborting or continuing a takeoff; continuing controlled flight and landing, possibly using emergency procedures but without requiring exceptional pilot skill or strength. Some aircraft damage may occur as a result of the failure condition or upon landing. For airplanes, the safe landing must be accomplished at a suitable airport. For rotorcraft, this means maintaining the ability of the rotorcraft to cope with adverse operating conditions and to land safely at a suitable site.
The FAA/AUTHORITY has determined that for Category A rotorcraft the phrase, "continued safe flight" means that the rotorcraft retains the capability to return and land safely at the point of departure or continue and land safely at the original intended destination or a suitable alternate site.
(1) This section is intended to ensure that a cooling fan blade failure will not jeopardize safety of the rotorcraft. Three ways to show compliance with this section are as follows:
A demonstration should be conducted to show that at the maximum fan speed to be expected, a failed blade will be contained within a housing or shroud which is included in the proposed type design and is designated as the containment shield;
(H) It should be shown that the installed cooling fan is located such that a blade failure will not jeopardize the safety of the rotorcraft or its ability to continue safe flight (Category A) or land safely (Category B); or,
(iil) It should be shown that the cooling fan blades can withstand an ultimate load 1.5 times the maximum centrifugal force that may be expected in service.
The maximum centrifugal forces will occur at the maximum cooling fan rotational speeds. The maximum fan rotational speeds may be related to an overspeed limiting device or to the maximum transient speed to be expected from analysis or test of the engine, system, or component which drives the fan. The maximum rotational speed will be as follows:
(A) For fans driven directly by the engine:
(1) The terminal engine rotational speed that will occur under uncontrolled conditions; such as output shaft disconnect; or
(2) The maximum engine rotational speed that would be controlled by a reliable, approved engine overspeed limiting device.
(B) For fans driven by the rotor drive system, the maximum rotor drive system rotational speed to be expected in service including transients. (Note:
Capability to withstand the ultimate load of 1.5 times the centrifugal force means that no Par 400A 623
failure would occur and distortion should not result in fan element contact with housings or other adjacent components during the 122.5 percent spin test which equates to 150 percent centrifugal force.)
(2) Fatigue. If the cooling fan is not included in the fatigue evaluation under § 29.571, it should be shown that the cooling fan blades are not operating at resonant conditions within the normal operating limits of the rotorcraft.
401.-420. RESERVED.
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SECTION 25. ROTOR DRIVE SYSTEM
421. § 29.917 (Amendment 29-12) DESIGN
a. § 29.917 (a) General:
(1) Explanation. This paragraph sets forth a definition of the rotor drive system and its associated components. The intent of this paragraph is to clarify and/or establish the identification of components to be considered in other rules which are applicable to the rotor drive system.
(2) Procedures. Coordinate with other certification personnel to ensure that other rules pertaining to rotor drive systems are properly addressed.
b. § 29.917(b) Arrangement:
(1) Explanation Section 29.917(b)(1) pertains to multiengine rotorcraft and requires the drive system arrangement to be such that the rotors will continue to be driven by the remaining engines in order to ensure that lift and control to be expected from the rotors are available if an engine fails.
Section 29.917(b)(2) pertains to single-engine rotorcraft and is similar to the requirement of paragraph 421b(1)(i) except that it requires each rotor necessary for operation and control to be driven by the main rotor(s) after disengagement of the engine from the main and auxiliary rotors.
Section 29.917(b)(3) is intended to require a design which allows the rotor system to be protected from the torsional drag of an inoperative engine.
(iv) Section 29.917(b)(4) pertains to optional torque limiting means (shear sections or clutches) and prohibits these devices from being located in the cross-shafting system between rotors.
(v) Section 29.917(b)(5) is intended to ensure that the design prevents rotors from contacting each other if intermeshing is possible.
(vi) Section 29.917(b)(6) is intended to ensure that locking devices are installed to keep rotors in proper phase if dephasing is possible.
(2) Procedures.
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(i) Section 29.917(b)(1) is normally complied with by cross-shafting between rotors, usually via one or more transmissions or gear boxes, to optimize the mechanical simplicity and weight aspects. Individual engine input arrangements are required.
(її) Section 29.917(b)(2) may be complied with by cross-shafting between rotors. Usually this involves driving the antitorque rotor via a drive shaft from the main transmission.
(ili) Section 29.917(b)(3) may be complied with by installing "free-wheel" or "one-way" clutches in the engine output shaft or transmission input quill. Note that the output section of "free power turbine" engines is not an acceptable method of compliance.
(iv) Section 29.917(b)(4). Any torque limiting devices in the rotor system should be located in the engine output or transmission input quill to ensure that any disconnect from overtorque does not preclude continued normal function and relation of the rotors.
Section 29.917(b)(5). Phase control of intermeshing rotors should utilize positive mechanical drive components. Deflections in both shafting (torsional) and rotors (blade chordwise bending) should be considered in establishing compliance.
(vi) Section 29.917(b)(6). Reconnection of dephased rotors should employ positive mechanical locking pins with secure locking methods.
421A. § 29.917 (Amendment 29-40) DESIGN.
a. Explanation. Amendment 29-40 introduces a new § 29.917(b). The previous § 29.917(b) has been redesignated as § 29.917(c). FAR 29.917(a) sets forth a definition of the rotor drive system and its associated components and FAR 29.917(b) requires a design assessment to be performed. The intent of this paragraph (b) is to identify the critical components and to establish and/or clarify their design integrity to show that the basic airworthiness requirements, which are applicable to the rotor drive system, will be met.
b. Procedures.
(1) § 29.917(a) General. The method of compliance for this section is unchanged.
(2) § 29.917(b) Design Assessment. A design assessment of the rotor drive system should be carried out in order to substantiate that the system is of a safe design and that compensating provisions are made available to prevent failures classified as hazardous and catastrophic in the sense specified in Paragraph (c) below. In carrying 646 Par 421
out the design assessment the results of the certification ground and flight testing (including any failures or degradation) should be taken into consideration. Previous service experience with similar designs should also be taken into account (see also FAR 29.601(a)).
c. Definitions. For the purposes of this assessment, failure conditions may be classified according to the severity of their effects as follows:
(1) Minor. Failure conditions which would not significantly reduce rotorcraft safety, and which involve crew actions that are well within their capabilities. Minor failure conditions may include, for example, a slight reduction in safety margins or functional capabilities, a slight increase in crew workload, such as routine flight plan changes, or some inconvenience to occupants.
(2) Major. Failure conditions which would reduce the capability of the rotorcraft or the ability of the crew to cope with adverse operating conditions to the extent that there would be, for example, a significant reduction in safety margins or functional capabilities, a significant increase in crew workload or in conditions impairing crew efficiency, or discomfort to occupants, possibly including injuries.
(3) Hazardous. Failure conditions which would reduce the capability of the rotorcraft or the ability of the crew to cope with adverse operating conditions to the extent that there would be - A large reduction in safety margins or functional capabilities;
(ii) Physical distress or higher workload such that the flight crew cannot be relied upon to perform their tasks accurately or completely;
(•1) Serious or fatal injury to a relatively small number of the occupants;
(iv) Loss of ability to continue safe flight to a suitable landing site.
(4) Catastrophic. Failure conditions which would prevent a safe landing.
(5) Minimize. Means to reduce to the least possible amount by means that can be shown to be both technically feasible and economically justifiable.
(6) Health Monitoring. Equipment, techniques, and/or procedures by which selected incipient failure or degradation can be determined.
d. Failure Analysis.
(1) The first stage of the design assessment should be the Failure Analysis, by which all the hazardous and catastrophic failure modes are identified. The failure Par 421A 647
analysis may consist of a structured, inductive bottom-up analysis, which is used to evaluate the effects of failures on the system and on the aircraft for each possible item or component failure. When properly formatted it will aid in identifying latent failures and the possible causes of each failure mode. The failure analysis should take into consideration all reasonably conceivable failure modes in accordance with the following:
Each item/component function(s).
(ii) Item/component failure modes and their causes.
(ill) The most critical operational phase/mode associated with the failure mode.
The effects of the failure mode on the item/component under analysis, the secondary effects on the rotor drive system and on the rotors, on other systems and on the rotorcraft. Combined effects of failures should be analyzed where a primary failure is likely to result in a secondary failure.
(v) The safety device or health monitoring means by which occurring or incipient failure modes are detected, or their effects mitigated. The analysis should consider the safety system failure.
(vi) The compensating provision(s) made available to circumvent or mitigate the effect of the failure mode (see also Paragraph (1) below).
(vii) The failure condition severity classification according to the definitions given in Paragraph (c) above.
(2) When deemed necessary for particular system failures of interest, the above analysis may be supplemented by a structured, deductive top-down analysis, which is used to determine which failure modes contribute to the system failure of interest.
(3) Dormant failure modes should be analyzed in conjunction with at least one other failure mode for the specific component or an interfacing component. This latter failure mode should be selected to represent a failure combination with potential worst case consequences.
(4) When significant doubt exists as to the effects of a failure, these effects may be required to be verified by tests.
e. Evaluation of Hazardous and Catastrophic Failures.
648 Par 421A
(1) The second stage of the design assessment is to summarize the hazardous and catastrophic failures and appropriately substantiate the compensating provisions which are made available to minimize the likelihood of their occurrence. Those failure conditions that are more severe should have a lower likelihood of occurrence associated with them than those that are less severe. The applicant should obtain early concurrence of the cognizant certificating authority with the compensating provisions for each hazardous or catastrophic failure.
(2) Compensating provisions may be selected from one or more of those listed below, but not necessarily limited to this list.
(i> Design features; i.e., safety factors, part-derating criteria, redundancies, etc.
(ii) A high level of integrity.
(iii) Fatigue tolerance evaluation.
(iv) Flight limitations.
(v) Emergency procedures.
(vi) An inspection or check that would detect the failure mode or evidence of conditions that could cause the failure mode.
(vii) A preventive maintenance action to minimize the likelihood of occurrence of the failure mode, including replacement actions and verification of serviceability of items which may be subject to a dormant failure mode.
(vill) Special assembly procedures or functional tests for the avoidance of assembly errors which could be safety critical.
(ix) Safety devices or health monitoring means beyond those identified in Paragraphs (vi) and (vii) above.
422. § 29.921 ROTOR BRAKE.
a. Background. Rotor brake safety requirements are intended not only to prevent adverse effects on aircraft performance due to brake drag but also to minimize the possibility of fires. These fires, caused by friction from a dragging rotor brake, have occurred both in flight and during ground operation with extremely hazardous consequences.
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