SUBPART E - POWERPLANT
ROTOR DRIVE SYSTEM AC 29.917. § 29.917 (Amendment 29-12) DESIGN.
a. Section 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. Section 29.917(b) Arrangement:
(1) Explanation.
(i) 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.
(ii) Section 29.917(b)(2) pertains to single-engine rotorcraft and is similar to the requirement of paragraph AC 29.917b(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.
(iii) 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.
(ii) 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.
(iii) 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.
(v) 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.
AC 29.917A. § 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). Section 29.917(a) sets forth a definition of the rotor drive system and its associated components and § 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) Section 29.917(a) General. The method of compliance for this section is unchanged.
(2) Section 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 out the design assessment, the results of the certification ground and flight Page E - 34
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 § 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--
(i) 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;
(iii) 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. Reduce to the least possible amount by means that can be shown to be both technically feasible and economically justifiable.
(6) Health Monitoring. A Vibration Health Monitoring System (VHM) is used to acquire and process helicopter drive system vibration signals.
(i) The principal purpose of a VHM is to increase the likelihood of detection of incipient faults in the rotor drive system that could prevent continued safe flight and safe landing by providing timely warning of potential failures to the pilot and maintenance personnel.
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AC 29-2C, Chg 3 9/30/2008
(ii) VHM data can be used to improve the helicopter's monitoring practices to mitigate a major or a hazardous/severe failure. The VHM data can also be used to improve maintenance.
(iii) A VHM system can be used to monitor all or some critical components of the rotor drive systems. Critical components include the input driveshaft to the main gearbox from the engine, bearings, tail rotor drive shaft, hanger bearings, one way clutch, main rotor mast and tail rotor mast. The supplier should state the component coverage and the fault detection capability for all affected components. The health monitoring effectiveness should be validated by tests or analysis or both.
(iv) Typically, a VHM system consists of sensors (e.g., accelerometers and tachometer), signal acquisition, signal processing, data management, VHM alert generation and management, a pilot interface, and a maintenance interface.
(v) Signal Processing: The helicopter's rotors, its drive systems and engines are a mixture of complex and simple mechanical elements. Therefore, the sensors and signal processing and analysis techniques utilized should reflect the complexity of the mechanical elements and their vibratory modes.
(vi) AC 29 MG-15 provides airworthiness approval guidance for rotorcraft health usage monitoring systems. This guidance can be used for incorporating VHM.
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 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:
(i) Each item/component function(s).
(ii) Item/component failure modes and their causes.
(iii) The most critical operational phase/mode associated with the failure mode.
(iv) 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.
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(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 worstcase 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.
(1) The second stage of the design assessment is to summarize the hazardous and catastrophic failures and appropriately substantiate the compensating provisions that 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: All parts with catastrophic failure modes and critical characteristics are to be identified as Critical Parts and be subject to a Critical Parts Plan (see AC 29.602.). Where a high level of integrity is used as a compensating provision, parts with a hazardous failure mode which would prevent continued safe flight may be included in a Critical Parts Plan or subjected to other enhancements to the normal control procedures for parts.
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AC 29-2C, Chg 3 9/30/2008
(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.
(viii) Special assembly procedures or functional tests for the avoidance of assembly errors which could be safety critical.
(ix) Safety devices or use of vibration health monitoring systems are recommended in addition to those provisions identified in paragraphs e.(2)(vi) and
e.(2)(vii).
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AC 29.921. § 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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b. General. This rule requires (1) that any limitations on the use of the rotor brake must be established, and (2) that the control for the brake must be guarded to prevent inadvertent operation.
c. Limitations.
(1) The limitations on the use of the rotor brake should first be defined by the applicant and will normally consist of merely the maximum rotor speed eligible for application of the brake. In some installations, limitations associated with engine operation may be specified. For example, some “free power section” type turbine engines can be safely operated within certain low limits with the rotor brake engaged, while other engines cannot tolerate this condition. At least one manufacturer has included a maximum rotor speed for emergency rotor brake application. This is considered an enhancing safety consideration and is recommended.
(2) Control guard mechanisms to prevent inadvertent operation may be conventional. A cockpit evaluation should be conducted by flight test personnel to affirm the function of the guard and the brake, and that markings, if any, are adequate and that both latched and unlatched positions of the control do not interfere with other cockpit functions.
d. General qualification aspects should include:
(1) The 400 applications required by § 29.923(j) conducted as a part of the § 29.923 endurance test.
(2) Torsional vibration measurements of the loads in the brake components and the rotor drive system during a critical brake engagement procedure, with appropriate consideration in the fatigue evaluation for these components. Brake engagements should be conducted with and without collective control displacement as authorized by the flight manual or a training manual.
(3) Brake component temperature measurements during a critical brake application in conjunction with an evaluation of the general brake compartment for compliance with §§ 29.863 and 29.1183.
(4) Placards, decals, and flight manual limitations and instructions appropriate to operate the rotor brake safely.
(5) An evaluation for hazardous failure modes as required by § 29.901(c). If the brake hydraulic system is integral with the rotorcraft hydraulic system, failure modes of pressure regulators and control valves, including valve leakage, will be of interest.
Mechanical cams, calipers, and levers may be prone to seize or fail to release the brake due, in part, to corrosion and lack of lubrication to be expected when brake components encounter high temperature cycling Page E - 38
NOTE: Most rotor brakes include nonmetallic pucks or liners, usually included in nonrotating brake components, which are subject to wear in proportion to the number of applications. Replacement of these pucks during the § 29.923 endurance test has been found acceptable provided the reason for replacement is simply wear and not because of any change in brake loading, disk temperature, or vibratory characteristics which can be expected in service. Verify that the maintenance manual includes a routine check for excessive puck or liner wear.
e. Other comments. Rotor brakes may be added to the basic design as a postcertification program without necessarily reconducting the complete § 29.923 endurance test provided:
(1) Steady and vibratory stresses in brake components, the rotor drive system, and in the rotor system itself are determined and shown to be acceptable.
NOTE: Moments, stresses, etc., from brake operation apply loads to the drive system in the reverse direction to normal powered flight. Advise the airframe engineer to require evaluation of chordwise bending loads in the hub and blade components of the main rotor system.
(2) The 400 brake engagements of § 29.923(j) should be accomplished with a complete rotor and rotor drive system, followed by disassembly sufficient to verify that all components subject to loads from the brake remain serviceable. Since this test may be so short as not to cause appreciable wear patterns to appear, special pretest coatings such as black oxide or Du-Lite may be needed on gear teeth and bearing races to distinguish and evaluate the contact patterns. Information on maximum deceleration rates should be supplied to the manufacturer of the engines to be used in the rotorcraft for evaluation of the acceptability of backloading or motoring of turbines, fuel control components, torque meters, etc.
AC 29.923. § 29.923 (Amendment 29-17) ROTOR DRIVE SYSTEM AND CONTROL MECHANISM TESTS.
a. Explanation.
(1) This rule is intended to require demonstration that the rotor drive system, as defined in § 29.917(a), is capable of normal operation within the limitations proposed, without hazard of failure from excessive wear or deterioration due to mechanical loads.
The basic test is not designed and should not be expected to demonstrate safety from oscillatory stresses normally investigated under §§ 29.571 and 29.907, although any data generated by these tests, which are in fact applicable to showing compliance with §§ 29.571 and 29.907, may be used. Some variations in the endurance test plan to generate data applicable to the vibration substantiative effort or other qualification aspects may be acceptable if the basic requirements of the endurance test are preserved.
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(2) The construction of this rule is such that a series of runs, each at least (but not limited to) 10 hours in length must be repeated 20 times, for a total of (at least) 200 hours of test, not including time required to adjust power or to stabilize operating conditions for those conditions that require stabilization. Extension of the total test beyond 200 hours (or extension of test runs beyond 10 hours) will occur if qualification for the 2½-minute one-engine-inoperative (OEI) optional rating is proposed by the applicant. The 30- minute OEI rating qualification test will extend the test beyond 200 hours for rotorcraft equipped with three or more engines. Also, compliance with § 29.923(g) may result in extended endurance tests if dynamic or malfunction conditions exist which adversely affect the endurance tolerance of the rotor drive system.
Section 29.923(a) should be interpreted as requiring test runs or cycles to be repeated in essentially the same sequence, although more than 10 hours may be needed to complete a run or cycle. This section also requires the test to be conducted “on the rotorcraft.” This means a rotorcraft in conformity to the design for which approval is requested. However, many nonconformity features, such as doors, some cowling and instrumentation, fuel tanks (alternate external fuel supply may be utilized), interior features, fire detectors, extinguishers, inlet ducts, exhaust baffles, etc., may be acceptable provided each item is technically considered and found to be unimportant to the test results. Any significant deviations from the conformed rotorcraft configuration should be coordinated with the cognizant FAA/AUTHORITY engineering staff and if found acceptable, documented as such. The restraint (tie- down) arrangement used during the test will necessarily be arranged to react rotor thrust loads in lateral as well as vertical directions. However, the restraint should permit normal deflections due to rotor thrust in the engine and drive system support arrangement.
(3) Safety cables may be installed normal to the tail boom at the tail rotor gearbox location; however, restraint may be provided to keep airframe deflections from exceeding those expected in normal and accelerated flight.
(4) The test torque requirements of § 29.923(a)(3)(i) mean the torque values for which approval is requested but not to exceed the values approved for the respective limits for the engine to be used. However, an applicant should be allowed to qualify the rotor drive system for torque values higher than those for which approval is requested if the engines actually used are capable of the torque and can be shown by an output shaft torsional investigation to be equivalent or conservative with respect to torsional vibration to the engines proposed for the initial certification configuration. Variations in rotational speed from the certification values should not be allowed except where careful evaluation of vibration aspects, bearing loads, centrifugal stiffening effects, and torque variations are conducted.
(5) The rotor configuration required by § 29.923(a)(3)(ii) is intended to assure that lift, torque, and vibration loads to be expected in service are introduced into the endurance test, although the presence of the vibration aspects does not normally satisfy the vibration evaluations required by §§ 29.571 and 29.907. In fact, vibration modes may be changed and amplified by the tie-down restraints and the increased thrust to be Page E - 40
expected from in-ground effects on the rotor system. These effects, although unquantified, are intended as a normal part of endurance testing. Preproduction rotor blades have been successfully used in endurance tests but only after specific investigations of blade properties such as stiffness, inertia and inertia distribution, thrust and blade bending, and torsional frequency response have been carefully compared to assure validity of the test. The endurance test includes testing of the rotor control mechanism. Conformity of the rotors may be very significant to this aspect of the test.
b. Procedures.
(1) Section 29.923(b)(1) prescribes the takeoff portion of the endurance test.
This test involves a series of 5-minute repetitive runs at the torque and at the engine/rotor rotational speed selected by the applicant for the takeoff limit for the rotorcraft. These values of torque (manifold pressure, for reciprocating engines) and RPM should correspond to the red radials on the corresponding powerplant instruments, except on installations where uncompensated engine governor “droop” results in a higher rotational speed for lower powers. The requirement in this section for declutching the engine may be difficult to achieve if engine deceleration and rotor system deceleration rates are similar. In some cases, the engine fuel control deceleration schedule may be adjusted to achieve clutch disengagement, otherwise, an engine shaft brake mechanism may be needed.
(2) The torque and speed requirements for the optional 2½-minute one-engine-inoperative (OEI) tests should be interpreted as described above for the takeoff runs. If the test is conducted during warm ambient conditions, excessive engine gas temperatures may be required to achieve the torque and speed conditions required by this part of the test. Minor adjustments in the run schedule may be allowed to take advantage of cooler nighttime ambient temperatures. Addition of water/alcohol systems to increase engine hot-day power may be appropriate in some instances. Liquid nitrogen spray into engine inlets has also been used to depress inlet temperatures sufficiently to obtain test conditions.
(3) In § 29.923(c), (d), (e), and (f), the torque requirements should be interpreted as above; i.e., the run should be made with maximum continuous torque or percentage thereof, as specified by the subparagraph, and the rotational speed should be maximum continuous for paragraph (d) and the lowest permissible “power-on” speed for paragraphs (e) and (f). Rotor control cycling must be accomplished during the “maximum continuous” portion of the endurance run. The controls of concern are the flight controls; i.e., cyclic and directional controls for rotorcraft with tail rotor and single main rotor. The collective control is normally used to set power and is not involved in control cycling. During control cycling the controls may be cycled from stop to stop, or a limited travel may be accepted if the travel produces the maximum fore and aft, left and right, and yaw thrust components of the rotors as measured in flight. One method of determining the required control displacement is to measure main rotor mast bending in level forward flight at maximum continuous power for the forward control displacement limit, and in level rearward flight at maximum continuous power (or the power Page E - 41
associated with the maximum rearward flight speed to be expected) for the aft control displacement limit. Using the same mast bending instrumentation, with the rotorcraft in the ground tie-down situation, and with collective control set for maximum continuous power, displace the cyclic fore and aft to obtain the same mast bending as measured in flight. Similar measurements and control displacements may be used for sideward thrust components. Yaw control displacement should consider maneuver requirements in conjunction with sideward flight. Critical gross weight and center of gravity should be used to establish test conditions. These same procedures may be used to establish limited control positions required to comply with § 29.923(i) except that typical flight conditions to be used would be stabilized level flight at maximum continuous power, climb at maximum continuous power, and hovering, including stabilized sideward and rearward flight. Note that for § 29.923(i)(1) vertical thrust is required. Depending on the mast angle and center of gravity, this condition may not necessarily involve zero mast bending loads. Vertical thrust may be used during the takeoff run, including the runs at 2½-minute power and the overspeed run of § 29.923(h). One- engine-inoperative runs (§ 29.923(k)) should be conducted with the cyclic set for maximum forward thrust. For these runs and any run that does not specify the position for the yaw control, that control should be set to react main rotor torque.
(4) Section § 29.923(g) provides for introducing special tests into the endurance tests to demonstrate that the transmission and drive system can tolerate certain engine malfunctions to be expected. This was originally directed at demonstrating safety in the event of spark plug or magneto failures of reciprocating engines. Turbine engines normally do not exhibit failure modes suitable for substantiation by endurance testing; however, severe or abusive operating conditions which must be expected to occur in service should be defined and included in this test.
Conditions or phenomena to be considered should include but not be limited to moderate engine surge, abusive clutch engagements, torque mismatching, anticipated control mishandling, and so forth. Alternatively, repeating the takeoff run of § 29.923(b) may be appropriate. It is not intended that the special testing for 2½ -minute power be repeated if a rerun of the takeoff power run is required by § 29.923(g).
(5) Section 29.923(h) requires overspeed testing at the torque which will produce maximum continuous power and at the maximum rotational speed to be expected. Normally this would be the maximum transient, power -on rotor speed available with speed controls operating. Special control adjustments for test purposes may be needed to achieve the required test conditions.
(6) Section 29.923(i) requires stabilized flight control displacement according to a prescribed schedule. The control displacement should be the same as derived to show compliance with § 29.923(c)(2).
(7) Section 29.923(j) requires 400 clutch and brake engagements. These tests are prescribed to establish a level of reliability of clutch and brake components installed as a part of the rotor drive system of rotorcraft. The clutch tests apply to all clutches installed to comply with § 29.917(b)(3), and each such clutch must be tested. A rotor Page E - 42