ED Decision 2018/007/R
(a) General. The rotor drive system includes any part necessary to transmit power from the engines to the rotor hubs. This includes gearboxes, shafting, universal joints, couplings, rotor brake assemblies, clutches, supporting bearings for shafting, any attendant accessory pads or drives, lubricating systems for drive system gearboxes, oil coolers and any cooling fans that are a part of, attached to, or mounted on the rotor drive system.
(b) Design assessment. A design assessment must be performed to ensure that the rotor drive system functions safely over the full range of conditions for which certification is sought. The design assessment must include a detailed failure analysis to identify all failures that will prevent continued safe flight or safe landing, and must identify the means to minimise the likelihood of their occurrence.
(c) Arrangement. Rotor drive systems must be arranged as follows:
(1) Each rotor drive system of multi- engine rotorcraft must be arranged so that each rotor necessary for operation and control will continue to be driven by the remaining engines if any engine fails.
(2) For single-engine rotorcraft, each rotor drive system must be so arranged that each rotor necessary for control in autorotation will continue to be driven by the main rotors after disengagement of the engine from the main and auxiliary rotors.
(3) Each rotor drive system must incorporate a unit for each engine to automatically disengage that engine from the main and auxiliary rotors if that engine fails.
(4) If a torque limiting device is used in the rotor drive system, it must be located so as to allow continued control of the rotorcraft when the device is operating.
(5) If the rotors must be phased for intermeshing, each system must provide constant and positive phase relationship under any operating condition.
(6) If a rotor dephasing device is incorporated, there must be means to keep the rotors locked in proper phase before operation.
[Amdt No: 29/5]
AMC1 29.917 Rotor drive system design
ED Decision 2021/016/R
VIBRATION HEALTH MONITORING
This AMC provides further guidance and acceptable means of compliance to supplement Federal Aviation Administration (FAA) Advisory Circular (AC) 29-2C, § AC 29.917. As such, it should be used in conjunction with the FAA AC.
This AMC clarifies the scope of complying with CS 29.1465, where the applicant uses vibration health monitoring as a compensating provision to meet CS 29.917(b).
Where vibration health monitoring is used as a compensating provision to meet CS 29.917(b), the competent authority should approve the design and performance of the vibration health monitoring system by requesting compliance with CS 29.1465(a).
[Amdt No: 29/5]
[Amdt No: 29/10]
AMC2 29.917 Rotor drive system design
ED Decision 2021/016/R
LUBRICATION SYSTEMS
This AMC provides further guidance and acceptable means of compliance to supplement Federal Aviation Administration (FAA) Advisory Circular (AC) 29 2C, § AC 29.917(b). As such, it should be used in conjunction with the FAA AC.
This AMC addresses the applicant’s dedicated safety assessment of the rotor drive system’s lubrication system and details how to use this assessment to help the applicant comply with CS 29.927(c).
For lubrication systems: a dedicated safety assessment should be performed that addresses all the lubrication systems of rotor drive system gearboxes and, in particular, the following:
(a) Identification of any single failure, malfunction, or reasonably conceivable combinations of failures that may result in a loss of oil pressure, a loss of oil supply to the dynamic components or a loss of the oil scavenge function. This normally takes the form of a failure mode and effects analysis. Compensating provisions should be identified to minimise the likelihood of occurrence of these failures. The safety assessment should also consider potential assembly or maintenance errors that cannot be readily detected during specified functional checks.
(b) The safety assessment should consider any specific design features which are subject to variability in manufacture or wear/degradation in service and which could have an appreciable effect on the maximum period of operation following loss of lubrication. Any features that may have a significant influence on the behaviour of the residual oil or the auxiliary lubrication system should be taken into account when determining the configuration of test articles.
(c) Identification of the most severe failure mode that results in the shortest duration of time in which the gearbox should be able to operate following the indication to the flight crew of a normal-use lubrication system failure. This should be used for simulating lubrication failure during the loss-of-lubrication test described in CS 29.927(c).
(d) Auxiliary lubrication system: Where compliance with CS 29.927(c) is reliant upon the operation of an auxiliary lubrication system, sufficient independence between the normal-use and auxiliary lubrication systems should be substantiated. Common-cause failure analysis, including common-mode, particular-risk, and zonal safety analyses, should be performed. It should be established that no single failure or identified common-cause failure will prevent the operation of both the normal-use and the auxiliary lubrication systems, apart from any failures that are determined to be extremely remote lubrication failures. The effects of inadvertent operation of the auxiliary lubrication system should also be considered.
(e) Definitions
(1) Lubrication system failure: in the context of CS 29.917(b), references to a failure of the lubrication system should be interpreted as any failure that results in a loss of pressure and an associated low oil pressure warning, within the duration of one flight.
(2) Most severe failure mode: the failure mode of the normal use lubrication system that results in the shortest duration of time in which the gearbox is expected to operate following an indication to the flight crew.
(3) Normal-use lubrication system: the lubrication system relied upon during normal operation.
(4) Auxiliary lubrication system: any lubrication system that is independent of the normal use lubrication system.
(5) Independent: an auxiliary lubrication system should be able to function after a failure of the normal-use lubrication system. Failure modes which may result in the subsequent failure of both the auxiliary and the normal-use lubrication systems and which may prevent continued safe flight or safe landing should be shown to be extremely remote lubrication failures.
(6) Extremely remote lubrication failure: a lubrication failure where the likelihood of occurrence has been minimised, either by structural analysis in accordance with CS 29.571 or laboratory testing. Alternatively, in-service experience or other means can be used which indicate a level of reliability comparable with one failure per 10 million hours. Failure modes including failures of external pipes, fittings, coolers, or hoses, and any components that require periodic removal by maintainers, should not be considered as extremely remote lubrication failures.
(f) Determination of the Most Severe Failure Mode
(1) The objective of the loss-of-lubrication test is to demonstrate the operation of a rotor drive system gearbox following the most severe failure mode of the normal-use lubrication system. The determination of the most severe failure mode may not be immediately obvious, as leakage rates vary, and system performance following leaks from different areas varies as well. Thus, a careful analysis of the potential failure modes should be conducted, taking into account the effects of flight conditions if relevant.
(2) The starting point for the determination of the most severe failure mode should be an assessment of all the potential lubrication system failure modes. This should be accomplished as part of the CS 29.917(b) design assessment, and should include leaks from any connections between components that are assembled together, such as threaded connections, hydraulic inserts, gaskets, seals, and packing (O-rings). Failure modes, such as failures of external lines, failures of component retention hardware and wall-through cracks that have not been substantiated for CS 29.307, CS 29.571 and CS 29.923(m) should also be considered. The determination that a failure is an extremely remote lubrication failure, when used to eliminate a potential failure mode from being considered as a candidate most severe failure mode, should be substantiated. Where leakage rates or the effect of failure modes cannot be easily determined, then a laboratory test should be conducted. Once the most severe failure mode has been determined, this should form the basis of the conditions for the start of the test.
(g) Use of an auxiliary lubrication system
The use of an auxiliary lubrication system may be an acceptable means of providing extended operating time after a loss of lubrication. The auxiliary lubrication system should be designed to provide sufficient independence from the normal-use lubrication system. Since the auxiliary lubrication system is by definition integral to the same gearbox as the normal-use lubrication system, it may be impractical for it to be completely independent. Therefore, designs should be conceived such that shared components or interfaces between the normal-use and auxiliary lubrication systems are minimised and comply with the design assessment provisions of CS 29.917(b). A failure of any common feature shared by both the normal-use and auxiliary lubrication systems that could result in the failure of both systems, and would consequently reduce the maximum period of operation following loss of lubrication, should be shown to be an extremely remote lubrication failure. If compliance with CS 29.927(c) is reliant on the functioning of an auxiliary lubrication system, then:
(1) in the unlikely event of a combined failure of both the normal-use lubrication system and the auxiliary lubrication system, the RFM emergency procedures should instruct the flight crew to ‘LAND IMMEDIATELY’ unless testing representing this failure mode has been performed in order to substantiate that an increased duration is justified; and
(2) a means of verifying that the auxiliary lubrication system is functioning properly should be provided during normal operation of the rotorcraft on either a periodic, pre-flight or continual basis. Following a failure of the normal-use lube system and activation of an auxiliary lubrication system, the flight crew should be alerted in the event of any system malfunction.
(h) Independence of the auxiliary lubrication system.
(1) In order to ensure that the auxiliary lubrication system is sufficiently independent:
(i) a failure of any pressurised portion of the normal-use lubrication system should not result in a subsequent failure of the auxiliary lubrication system;
(ii) common failure modes shown to defeat both the normal-use and the auxiliary lubrication systems should be shown to be extremely remote lubrication failures, unless it is demonstrated by testing conducted to comply with CS 29.927(c) that the failure mode does not compromise the Maximum period of operation following loss of lubrication; and
(iii) control systems, logic and health-reporting systems should not be shared; consideration should be given to the design process to ensure appropriate segregation of the control and warning systems in the system architecture.
(2) Methods which should be used to demonstrate that failure modes of common areas are extremely remote include:
(i) field experience of the exact design with an exact application;
(ii) field experience with a similar design/application with supporting test data to allow a comparison;
(iii) demonstration by test of extremely low leakage rates;
(iv) redundancy of design;
(v) structural substantiation with a high safety margin for elements of the lubrication systems assessed against CS 29.571; and
(vi) assessment of the potential dormant failure modes of the auxiliary lubrication system, and in order to minimise the risk of dormant failures, determination of the health of the auxiliary lubrication system prior to each flight.
[Amdt No: 29/5]
[Amdt No: 29/10]
AMC3 29.917 Rotor drive system design
ED Decision 2021/016/R
CHIP DETECTION SYSTEM
This AMC provides further guidance and acceptable means of compliance to supplement Federal Aviation Administration (FAA) Advisory Circular (AC) 29 2C, § AC 29.917(b). As such, it should be used in conjunction with the FAA AC.
This AMC contains additional considerations for each chip detection system that the applicant uses as a compensating provision to meet CS 29.917(b). For each chip detection system that the applicant uses as a compensating provision for hazardous or catastrophic failures to meet CS 29.917(b), this section introduces AMC to substantiate the chip detection system that is specified in CS 29.1337(e) as an appropriate compensating provision.
(a) The applicant may identify a chip detection system that is installed on a rotor drive system transmission or gearbox as a compensating provision in the rotor drive system design assessment to comply with CS 29.1337(e). The chip detection system that is used as a compensating provision is intended to minimise the likelihood of occurrence of certain failures in transmissions and gearboxes, including hazardous and catastrophic failures.
(b) To be accepted as an appropriate compensating provision, the chip detection system should effectively indicate the presence of ferromagnetic particles that are released due to damage or excessive wear. That damage or excessive wear could lead to the failures whose likelihood of occurrence the chip detection system is intended to minimise. As a result, to demonstrate compliance with CS 29.917(b), the applicant should substantiate the effectiveness of the chip detection system for all the identified hazardous and catastrophic failure modes through full scale test evidence.
(c) The test(s) that are performed to demonstrate compliance with CS 29.917(b) should address all those areas of the rotor drive system that are associated with the failures for which the chip detection system is identified as a compensating provision. AMC1 29.1337 provides further guidance on the use of full-scale testing as a means to demonstrate the compliance of the chip detection system. It also defines performance objectives that the applicant should meet to demonstrate the general level of effectiveness of the system. However, the applicant should specifically assess the amount of ferromagnetic particles and use the value of 60 mg that is provided in AMC1 29.1337(e) only if supported by that assessment. This means that an amount of particles is justified to be released with sufficient margin before a hazardous or catastrophic failure occurs.
Note: the applicant should not consider that demonstrating the effectiveness of a chip detection system to comply with CS 29.917(b) and CS 29.1337(e) is an alternative to providing a robust and reliable design, or a means to relieve the applicant of demonstrating compliance with other necessary compensating provisions.
[Amdt No: 29/10]
ED Decision 2003/16/RM
If there is a means to control the rotation of the rotor drive system independently of the engine, any limitations on the use of that means must be specified, and the control for that means must be guarded to prevent inadvertent operation.
CS 29.923 Rotor drive system and control mechanism tests
ED Decision 2003/16/RM
(a) Endurance tests, general. Each rotor drive system and rotor control mechanism must be tested, as prescribed in sub-paragraphs (b) to (n) and (p), for at least 200 hours plus the time required to meet the requirements of sub-paragraphs (b)(2), (b)(3) and (k). These tests must be conducted as follows:
(1) Ten-hour test cycles must be used, except that the test cycle must be extended to include the OEI test of sub-paragraphs (b)(2) and (k), if OEI ratings are requested.
(2) The tests must be conducted on the rotorcraft.
(3) The test torque and rotational speed must be:
(i) Determined by the powerplant limitations; and
(ii) Absorbed by the rotors to be approved for the rotorcraft.
(b) Endurance tests, take-off run. The take- off run must be conducted as follows:
(1) Except as prescribed in sub- paragraphs (b)(2) and (b)(3), the take-off torque run must consist of 1 hour of alternate runs of 5 minutes at take-off torque and the maximum speed for use with take-off torque, and 5 minutes at as low an engine idle speed as practicable. The engine must be declutched from the rotor drive system, and the rotor brake, if furnished and so intended, must be applied during the first minute of the idle run. During the remaining 4 minutes of the idle run, the clutch must be engaged so that the engine drives the rotors at the minimum practical rpm. The engine and the rotor drive system must be accelerated at the maximum rate. When declutching the engine, it must be decelerated rapidly enough to allow the operation of the overrunning clutch.
(2) For helicopters for which the use of a 2½-minute OEI rating is requested, the take- off run must be conducted as prescribed in subparagraph (b)(1), except for the third and sixth runs for which the take-off torque and the maximum speed for use with take-off torque are prescribed in that paragraph. For these runs, the following apply:
(i) Each run must consist of at least one period of 2½ minutes with take- off torque and the maximum speed for use with take-off torque on all engines.
(ii) Each run must consist of at least one period, for each engine in sequence, during which that engine simulates a power failure and the remaining engines are run at the 2½- minutes OEI torque and the maximum speed for use with 2½-minute OEI torque for 2½ minutes.
(3) For multi-engine, turbine-powered rotorcraft for which the use of 30-second/2-minute OEI power is requested, the take-off run must be conducted as prescribed in sub- paragraph (b)(1) except for the following:
(i) Immediately following any one 5-minute power-on run required by sub-paragraph (b)(1), simulate a failure, for each power source in turn, and apply the maximum torque and the maximum speed for use with the 30-second OEI power to the remaining affected drive system power inputs for not less than 30 seconds. Each application of 30-second OEI power must be followed by two applications of the maximum torque and the maximum speed for use with the 2 minute OEI power for not less than 2 minutes each; the second application must follow a period at stabilised continuous or 30-minute OEI power (whichever is requested by the applicant.) At least one run sequence must be conducted from a simulated ‘flight idle’ condition. When conducted on a bench test, the test sequence must be conducted following stabilisation at take-off power.
(ii) For the purpose of this paragraph, an affected power input includes all parts of the rotor drive system which can be adversely affected by the application of higher or asymmetric torque and speed prescribed by the test.
(iii) This test may be conducted on a representative bench test facility when engine limitations either preclude repeated use of this power or would result in premature engine removals during the test. The loads, the vibration frequency, and the methods of application to the affected rotor drive system components must be representative of rotorcraft conditions. Test components must be those used to show compliance with the remainder of this paragraph.
(c) Endurance tests, maximum continuous run. Three hours of continuous operation at maximum continuous torque and the maximum speed for use with maximum continuous torque must be conducted as follows:
(1) The main rotor controls must be operated at a minimum of 15 times each hour through the main rotor pitch positions of maximum vertical thrust, maximum forward thrust component, maximum aft thrust component, maximum left thrust component, and maximum right thrust component, except that the control movements need not produce loads or blade flapping motion exceeding the maximum loads of motions encountered in flight.
(2) The directional controls must be operated at a minimum of 15 times each hour through the control extremes of maximum right turning torque, neutral torque as required by the power applied to the main rotor, and maximum left turning torque.
(3) Each maximum control position must be held for at least 10 seconds, and the rate of change of control position must be at least as rapid as that for normal operation.
(d) Endurance tests: 90% of maximum continuous run. One hour of continuous operation at 90% of maximum continuous torque and the maximum speed for use with 90% of maximum continuous torque must be conducted.
(e) Endurance tests; 80% of maximum continuous run. One hour of continuous operation at 80% of maximum continuous torque and the minimum speed for use with 80% of maximum continuous torque must be conducted.
(f) Endurance tests; 60% of maximum continuous run. Two hours or, for helicopters for which the use of either 30-minute OEI power or continuous OEI power is requested, 1 hour of continuous operation at 60% of maximum continuous torque and the minimum speed for use with 60% of maximum continuous torque must be conducted.
(g) Endurance tests: engine malfunctioning run. It must be determined whether malfunctioning of components, such as the engine fuel or ignition systems, or whether unequal engine power can cause dynamic conditions detrimental to the drive system. If so, a suitable number of hours of operation must be accomplished under those conditions, 1 hour of which must be included in each cycle, and the remaining hours of which must be accomplished at the end of the 20 cycles. If no detrimental condition results, an additional hour of operation in compliance with sub-paragraph (b) must be conducted in accordance with the run schedule of sub-paragraph (b)(1) without consideration of sub-paragraph (b)(2).
(h) Endurance tests; overspeed run. One hour of continuous operation must be conducted at maximum continuous torque and the maximum power-on overspeed expected in service, assuming that speed and torque limiting devices, if any, function properly.
(i) Endurance tests: rotor control positions. When the rotor controls are not being cycled during the endurance tests, the rotor must be operated, using the procedures prescribed in subparagraph (c), to produce each of the maximum thrust positions for the following percentages of test time (except that the control positions need not produce loads or blade flapping motion exceeding the maximum loads or motions encountered in flight):
(1) For full vertical thrust, 20%.
(2) For the forward thrust component, 50%
(3) For the right thrust component, 10%.
(4) For the left thrust component, 10%.
(5) For the aft thrust component, 10%.
(j) Endurance tests, clutch and brake engagements. A total of at least 400 clutch and brake engagements, including the engagements of sub-paragraph (b), must be made during the take-off torque runs and, if necessary, at each change of torque and speed throughout the test. In each clutch engagement, the shaft on the driven side of the clutch must be accelerated from rest. The clutch engagements must be accomplished at the speed and by the method prescribed by the applicant. During deceleration after each clutch engagement, the engines must be stopped rapidly enough to allow the engines to be automatically disengaged from the rotors and rotor drives. If a rotor brake is installed for stopping the rotor, the clutch, during brake engagements, must be disengaged above 40% of maximum continuous rotor speed and the rotors allowed to decelerate to 40% of maximum continuous rotor speed, at which time the rotor brake must be applied. If the clutch design does not allow stopping the rotors with the engine running, or if no clutch is provided, the engine must be stopped before each application of the rotor brake, and then immediately be started after the rotors stop.
(k) Endurance tests, OEI power run.
(1) For rotorcraft for which the use of 30-minute OEI power is requested, a run at 30-minute OEI torque and the maximum speed for use with 30-minute OEI torque must be conducted as follows. For each engine, in sequence, that engine must be inoperative and the remaining engines must be run for a 30-minute period.
(2) For rotorcraft for which the use of continuous OEI power is requested, a run at continuous OEI torque and the maximum speed for use with continuous OEI torque must be conducted as follows. For each engine, in sequence, that engine must be inoperative and the remaining engines must be run for 1 hour.
(3) The number of periods prescribed in sub-paragraph (k)(1) or (k)(2) may not be less than the number of engines, nor may it be less than two.
(l) Reserved.
(m) Any components that are affected by manoeuvring and gust loads must be investigated for the same flight conditions as are the main rotors, and their service lives must be determined by fatigue tests or by other acceptable methods. In addition, a level of safety equal to that of the main rotors must be provided for:
(1) Each component in the rotor drive system whose failure would cause an uncontrolled landing;
(2) Each component essential to the phasing of rotors on multi-rotor rotorcraft, or that furnishes a driving link for the essential control of rotors in autorotation; and
(3) Each component common to two or more engines on multi-engine rotorcraft.
(n) Special tests. Each rotor drive system designed to operate at two or more gear ratios must be subjected to special testing for durations necessary to substantiate the safety of the rotor drive system.
(o) Each part tested as prescribed in this paragraph must be in a serviceable condition at the end of the tests. No intervening disassembly which might affect test results may be conducted.
(p) Endurance tests; operating lubricants. To be approved for use in rotor drive and control systems, lubricants must meet the specifications of lubricants used during the tests prescribed by this paragraph. Additional or alternate lubricants may be qualified by equivalent testing or by comparative analysis of lubricant specifications and rotor drive and control system characteristics. In addition:
(1) At least three 10-hour cycles required by this paragraph must be conducted with transmission and gearbox lubricant temperatures, at the location prescribed for measurement, not lower than the maximum operating temperature for which approval is requested;
(2) For pressure lubricated systems, at least three 10-hour cycles required by this paragraph must be conducted with the lubricant pressure, at the location prescribed for measurement, not higher than the minimum operating pressure for which approval is requested; and
(3) The test conditions of sub-paragraphs (p)(1) and (p)(2) must be applied simultaneously and must be extended to include operation at any one-engine-inoperative rating for which approval is requested.
ED Decision 2018/007/R
(a) Any additional dynamic, endurance, and operational tests, and vibratory investigations necessary to determine that the rotor drive mechanism is safe, must be performed.
(b) If turbine engine torque output to the transmission can exceed the highest engine or transmission torque limit, and that output is not directly controlled by the pilot under normal operating conditions (such as where the primary engine power control is accomplished through the flight control), the following test must be made:
(1) Under conditions associated with all engines operating, make 200 applications, for 10 seconds each, of torque that is at least equal to the lesser of:
(i) The maximum torque used in meeting CS 29.923 plus 10%; or
(ii) The maximum torque attainable under probable operating conditions, assuming that torque limiting devices, if any, function properly.
(2) For multi-engine rotorcraft under conditions associated with each engine, in turn, becoming inoperative, apply to the remaining transmission torque inputs the maximum torque attainable under probable operating conditions, assuming that torque limiting devices, if any, function properly. Each transmission input must be tested at this maximum torque for at least 15 minutes.
(c) Lubrication system failure. For rotor drive system gearboxes required for continued safe flight or safe landing which have a pressurised normal-use lubrication system, the following apply:
(1) Category A. Confidence shall be established that the rotor drive system has an in-flight operational endurance capability of at least 30 minutes following a failure of any one pressurised normal-use lubrication system.
For each rotor drive system gearbox necessary for continued safe flight or safe landing, a test shall be conducted simulating the effect of the most severe failure mode of the normal-use lubrication system as determined by the failure analysis of CS 29.917(b). The duration of the test shall be dependent upon the number of tests and the component condition after the test. The test shall be conducted such that it begins upon the indication to the flight crew that a lubrication failure has occurred, and its loading is consistent with 1 minute at maximum continuous power, followed by the minimum power needed for continued flight at the rotorcraft maximum gross weight. The test shall end with a 45-second out of ground effect (OGE) hover to simulate a landing phase. Test results must substantiate the maximum period of operation following loss of lubrication by means of an extended test duration, multiple test specimens, or another approach prescribed by the applicant and accepted by EASA, and must support the procedures published in the rotorcraft flight manual (RFM). Flight durations longer than 30 minutes may be demonstrated by means of a correspondingly longer test with appropriate margin and substantiation.
(2) Category B. Confidence shall be established that the rotor drive system has an in-flight operational endurance capability to complete an autorotation descent and landing following a failure of any one pressurised normal-use lubrication system.
For each rotor drive system gearbox necessary for safe autorotation descent or safe landing, a test of at least 16 minutes and 15 seconds following the most severe failure mode of the normal-use lubrication system as determined by the failure analysis of CS 29.917(b) shall be conducted. The test shall be conducted such that it begins upon the indication to the flight crew that a lubrication failure has occurred and its loading is consistent with 1 minute at maximum continuous power, after which the input torque should be reduced to simulate autorotation for 15 minutes. The test shall be completed by the application of an input torque to simulate a minimum power landing for approximately 15 seconds.
(d) Overspeed test. The rotor drive system must be subjected to 50 overspeed runs, each 30 ± 3 seconds in duration, at not less than either the higher of the rotational speed to be expected from an engine control device failure or 105% of the maximum rotational speed, including transients, to be expected in service. If speed and torque limiting devices are installed, are independent of the normal engine control, and are shown to be reliable, their rotational speed limits need not be exceeded. These runs must be conducted as follows:
(1) Overspeed runs must be alternated with stabilising runs of from 1 to 5 minutes duration each at 60 to 80% of maximum continuous speed.
(2) Acceleration and deceleration must be accomplished in a period not longer than 10 seconds (except where maximum engine acceleration rate will require more than 10 seconds), and the time for changing speeds may not be deducted from the specified time for the overspeed runs.
(3) Overspeed runs must be made with the rotors in the flattest pitch for smooth operation.
(e) The tests prescribed in sub-paragraphs (b) and (d) must be conducted on the rotorcraft and the torque must be absorbed by the rotors to be installed, except that other ground or flight test facilities with other appropriate methods of torque absorption may be used if the conditions of support and vibration closely simulate the conditions that would exist during a test on the rotorcraft.
(f) Each test prescribed by this paragraph must be conducted without intervening disassembly and, except for the lubrication system failure test required by sub-paragraph (c) , each part tested must be in a serviceable condition at the conclusion of the test.
[Amdt No: 29/5]
AMC1 29.927(c) Additional tests
ED Decision 2021/016/R
This AMC replaces item a. (Section 29.927(c)) of FAA AC 29.927 (Amendment 29-26).
(a) Explanation
(1) AMC 29.927 revises the rotor drive systems loss of lubrication test provisions for Category A rotorcraft, as defined in CS 29.927(c). This changes the related requirement to show a capability through testing of at least 36 minutes’ duration. Additionally, minimum periods and load conditions are now defined directly in the provision. The failure condition to be simulated is the most severe loss of lubrication failure mode of the normal-use lubrication system, which is defined in AMC2 29.917(b). In addition, the term ‘unless such failures are extremely remote’ has been removed from the requirement. Assessment of the lubrication system reliability is now addressed under 29.917(b).
(2) CS 29.927(c) is intended to apply to pressurised lubrication systems, as the likelihood of loss of lubrication is significantly greater for gearboxes that use pressurised lubrication and external cooling. This is due to the increased complexity of the lubrication system, the external components that circulate oil outside the gearbox, and the resultant rapid leakages that may occur with a pressurised system. A pressurised lubrication system is more commonly used in the rotorcraft’s main gearbox, but one may also be used in other rotor drive system gearboxes. The need for dedicated loss of lubrication testing for gearboxes using non-pressurised (splash) lubrication systems is determined by the design assessment carried out in accordance with 29.917(b).
(3) This provision is applicable to any pressurised lubrication gearbox that is necessary for continued safe flight or safe landing. Accordingly, this provision is not applicable to gearboxes that are not essential for continued safe flight or safe landing and which have a lubrication system which is independent of other essential gearboxes.
(4) The lubricating system has two primary functions. The first is to provide lubricating oil to contacting or rubbing surfaces to reduce the heat energy generated by friction. The second is to dissipate the heat energy generated by the friction of meshing gears and bearings, thus maintaining surface and component temperatures. Accordingly, a loss of lubrication leads to increased friction between components and increased component surface temperatures. With increased component surface temperatures, surface hardness may be lost, resulting in the inability of the component to carry or transmit loads appropriately. Thermal expansion in gearbox components may eventually lead to the mechanical failure of bearings, journals, gears, shafts, and clutches that are subjected to high loads and rotational speeds. A loss of lubrication may result from either internal or external failures.
(5) The intent of the rule change for Category A rotorcraft is to provide confidence in the continued flight capability of the rotorcraft, which should be of at least 30 minutes’ duration after the loss of lubricant pressure in any single rotorcraft drive system gearbox, with the aim of optimising the eventual landing opportunities. In order to enable the crew to determine the safest action in the event of a loss of gearbox oil, the emergency procedures of the rotorcraft flight manual (RFM) should include instructions that define the maximum time period within which the rotorcraft should land. This AMC provides guidance for the completion of the loss of lubrication test and for how to demonstrate confidence in the margin of safety associated with the maximum period of operation following loss of lubrication, and associated period defined in the RFM emergency procedures. This margin of safety is intended to substantiate a period of operation that has been evaluated as likely to be safer than making a forced landing over hostile terrain.
(b) Procedures
(1) CS 29.927(c) prescribes a test that is intended to demonstrate that no hazardous failure or malfunction will occur within a defined period, and in a specified reduced-power condition, in the event of a significant failure of the rotor drive lubrication system. The failure of the lubrication system should not impair the ability of the crew to continue the safe operation of Category A rotorcraft for the defined period after an indication of the failure has been provided to the flight crew. For Category B rotorcraft, safe operation under autorotative conditions should be possible for a period of at least 15 minutes. For both Category A and B rotorcraft, some damage to the rotor drive system components is acceptable after completion of the lubrication system testing. However, the condition of the components will influence the maximum period of operation following loss of lubrication.
(2) Since this is a test of the capability of the gearbox to operate with residual oil or oil supplied from an auxiliary lubrication system, the method for draining the oil and the operating conditions are also defined in the provision. The entry condition for the test should also be representative, and is defined in this AMC. For Category B rotorcraft, it is necessary to simulate an autorotation for a period of 15 minutes, followed by a minimum-power landing.
(c) Definitions
For the purposes of this test and the assessment of continued operation after a loss of lubrication, the following definitions apply:
(1) Maximum period of operation following loss of lubrication: The maximum period of time following a loss of oil pressure warning, within which the rotorcraft should land. The period stated in the associated RFM emergency procedures should not exceed the maximum period of operation following loss of lubrication.
(2) Residual oil: the oil present in the gearbox after experiencing the most severe failure mode, beginning at the time the pilot receives an indication of the failure. (Note: the amount of residual oil may decrease with time, and test conditions should take into account the possible effects of flight conditions where relevant. Also, when the lubrication system incorporates an auxiliary lubrication system, this will supplement the residual oil in the event of a failure of the normal-use lubrication system).
(d) Certification test configuration
Each gearbox lubricated by a pressurised system that is necessary for continued safe flight or safe landing should be tested. Deviations from the gearbox configuration being certified may be allowed where necessary for the installation of test instrumentation or equipment to facilitate simulation of the most severe failure mode. If any specific design features are identified in the safety assessment that may have a significant influence on the behaviour of the residual oil or the auxiliary lubrication system, they should be taken into account when determining the configuration of the test articles.
(e) Loss of lubrication test
(1) Category A rotorcraft
(i) Test entry condition: the test starting condition should be 100 % of the torque associated with all engines operative (AEO) maximum continuous power (MCP) and at the nominal speed for use with MCP. In addition, the torque necessary for the anti-torque function should be simulated for straight and level flight at the same flight conditions. The oil temperature should be stabilised at the maximum oil temperature limit for normal operation.
(ii) Draining of oil: once the oil temperature has stabilised at the maximum declared oil temperature limit for normal operation, the oil should be drained simulating the most severe failure mode of the normal-use lubrication system. The most severe failure mode should be determined by the failure analysis of CS 29.917(b). The location and rate of oil drainage should be representative of the mode being simulated and the drainage should continue throughout the test.
(iii) Depleted-oil run: upon illumination of the ‘low oil pressure’ warning or other indication, as required by CS 29.1305, continue to operate at AEO MCP and the nominal speed for use in this condition for 1 minute. Then, reduce the torque values to be greater than or equal to those necessary to sustain flight at the maximum gross weight and the most efficient flight conditions under standard atmospheric conditions (Vy). This condition should be maintained during the time determined necessary by the applicant to justify the maximum period of operation following loss of lubrication taking into account the applicable reduction factors. When determining the torque values to sustain flight at the maximum gross weight and the most efficient flight conditions (Vy), it should be assumed that the condition starts at 100 % maximum take-off weight (MTOW), and, thereafter, consideration for the fuel burn during the test is allowed.
(iv) Simulated landing: to complete the test, power should be applied to the gearbox for at least 45 seconds to simulate an out of ground effect (OGE) hover.
(v) Test conditions: for (i) to (iv) above, the input and output shaft torques should be reacted appropriately and the corresponding input and output shaft loads should be applied. As the efficiency of the gearbox may change during the test, the input loads may need to be adjusted in order to maintain the correct output shaft torque during the test. The vertical load of the main gearbox should be applied at the mast, and should be equal to the maximum gross weight of the rotorcraft at 1 g.
(vi) This test may be conducted on a representative bench test rig. The test should be performed with all the accessory loads represented by a load associated with normal cruise conditions. The test should not be performed with an ambient temperature in the test cell lower than ISA conditions. No additional ventilation that could reduce the gearbox temperature should be used which could result in temperatures which are lower than those which are likely to be experienced on the helicopter operating at ISA conditions.
(vii) A successful demonstration may involve limited damage to the rotor drive system; however, the gearbox should continue to transmit the necessary torque to the output shafts throughout the duration of the test. The loss of drive to accessories that are necessary for continued safe flight or safe landing should constitute a test failure.
(2) Category B rotorcraft
(i) The provisions for Category A apply, except that the rotor drive system need only perform a depleted-oil run for 15 minutes operating at a torque and speed to simulate autorotative conditions.
(ii) A successful demonstration may involve limited damage to the rotor drive system provided that it is established that the autorotative capabilities of the rotorcraft would not be significantly impaired. If compliance with Category A provisions is demonstrated, Category B provisions will be considered to have been met.
(3) The test parameters described in (e)(1) above have been chosen to represent an occurrence of loss of oil in flight, namely a reaction/transition period for the crew to be able to reduce power, followed by an extended period at reduced power for continued flight at Vy. When determining the torque necessary for the reduced-power segment of this test, an international standard atmosphere (ISA) sea level condition is considered to be acceptable.
(4) Should the applicant wish to establish a positive safety margin for a Category A rotorcraft for a maximum period of operation following loss of lubrication longer than 30 minutes, it will be necessary to extend the test duration representing flight at Vy, described in (e)(1)(iii) above.
(f) Determination of the maximum period of operation following loss of lubrication
In order to enable the flight crew to determine the safest action in the event of a loss of gearbox oil, the RFM emergency procedures should include instructions defining the maximum period of time, for each gearbox subject to 29.927(c), within which the rotorcraft should land. This period starts at the low pressure warning. Specific instructions can be prescribed by the applicant as an alternative to, or in addition to, defining the maximum period of operation following loss of lubrication, in order to maintain a continued safe flight and safe landing capability. The flight time allowance listed in the RFM should be based on the OEM's determination of what is appropriate, using guidance from the available test data, but it should be no greater than what is substantiated per the acceptable means of compliance (AMC) prescribed below. Accordingly, it is necessary to demonstrate reasonable confidence in the ability of the gearbox to continue operation enabling safe flight and safe landing after experiencing a loss of oil or a lubrication failure. (f)(1) to (f)(4) below describe acceptable means of compliance (AMC) to demonstrate this level of confidence, for a specified period at given operating conditions. This AMC explains how the test duration, the number of tests, the condition of the gearbox components upon completion of the tests, and the behaviour of the gearbox during these tests may be combined to establish a positive safety margin when determining the maximum period of operation following loss of lubrication.
(1) Certification test duration
The duration of the loss of lubrication certification test, as defined in (e) above, should be used as the starting point for the determination of the maximum period of operation following loss of lubrication and should be reduced as described in the following paragraphs as appropriate. The start of the test is considered to be the time at which the lubrication failure is indicated to the pilot.
(2) Reduction factor
In order to substantiate the maximum period of operation following loss of lubrication, a suitable reduction factor should be applied to correlate the test duration with the maximum period of operation following loss of lubrication. Suitable reduction factors should be used as follows:
(i) 0.6 where the certification test has no supporting data to provide understanding of the gearbox behaviour and confidence in the repeatability of the certification test data.
(ii) 0.8 where the certification test is corroborated by one representative full-scale test (certification or development test). The corroborating test results should show consistency of the temperature history, and demonstrate good correlation with the certification test.
(iii) 0.9 where the certification test is corroborated by two or more representative full-scale tests (certification or development tests) or by one representative full scale and one or more modular tests, historical data, or simulation results. The corroborating data should show consistency of the temperature history, and demonstrate good correlation with the certification test. In addition the behaviour of the limiting design characteristics is established and supported by repeatable test data.
Note: Specific testing, simulation or representative development test data from other programmes are examples of data that can be used to support the application of this Kr factor.
(iv) When two or more tests are submitted to show compliance with this provision, the test of shortest duration will be considered to be the certification test and should be used as the basis for demonstrating the maximum period of operation following loss of lubrication. If excessive variation is experienced between tests, it should be investigated and explained.
(v) The intent of using data from multiple tests is that the parts replaced between tests are those that potentially limit the performance of the gearbox when operating under residual oil or oil supplied from an auxiliary lubrication system. Where particular design characteristics are known to be critical to residual oil performance, parts should be selected at the most severe end of the tolerance range of the dimensions/specifications impacting these characteristics. Additionally, the objective of multiple tests is to evaluate the consistency between tests (using different gearbox components). When using multiple (full scale or modular) test results to corroborate the certification test duration and, thus, support the determination of the maximum period of operation following loss of lubrication, the criteria for the reconciliation between the corroborating test data and an official certification test should include:
a. the test conditions, i.e. loads, entry point and test profile, should be duplicated on the development test as for the official test, and any deviations should be substantiated;
b. the representativeness of parts should be demonstrated and documented;
c. the test equipment and instrumentation should be qualified and calibrated;
d. the correlation between development and official test should be demonstrated by absolute temperatures and temperature rates of change; and
e. in addition for modular tests, the lubrication conditions should be conservatively simulated to avoid that the isolated module benefits from secondary lubrication from the boundaries of the module, which may not be representative of the module conditions in a full test.
(vi) When determining the appropriate reduction factor, consideration should be given to any factors that may reflect the health or stability of gearbox components during the test(s). These factors are addressed below and include: temperature history, maximum temperatures achieved with respect to physical limitations of the material, simulation results, and the time difference between the demonstrated duration up to a test failure and the duration of the certification test.
a. Temperature rate of change during test. Gearboxes operating after loss of lubrication sometimes exhibit portions of the test where the thermal response is either stable (approaching to zero rate of change) or meta-stable (with a ‘small’ rate of change). It is considered that confidence in the behaviour of the gearbox may be greater for a maximum absolute temperature measured under these conditions in the context of the certification test or an official test. Portions of the test that exhibit a larger temperature rate of change should be investigated and substantiated.
b. Maximum temperature reached during test. Similarly to the rate of temperature change, general experience from ‘total loss of lubrication’ tests performed has shown that successful tests do not exceed certain values of temperature measured at critical locations of the gearbox. The applicant should record temperature measurements from critical points of the gearbox or at related locations in order to compare with previous experience. This data should be used to validate analysis models and to support the application of a high Kr value when determining the maximum period of operation following loss of lubrication.
c. Models/simulations. Numerical simulation of loss of lubrication conditions is not considered sufficient to demonstrate confidence in absolute temperature values achieved during the certification test, when applied to the prediction of the maximum period of operation following loss of lubrication. However, it may be possible to apply numerical simulation (0-3 dimensional) to extrapolate test results to other boundary or entry conditions.
d. Extended operation. The applicant is encouraged to perform tests in order to evaluate the time difference between the point at which the certification test was concluded and the likely time of gearbox failure (if the certification test had continued). Of equal importance is the identification of the gearbox design features which are most likely to initiate gearbox failure in the event of extended operation after loss of lubrication.
Note: if, at the completion of the certification test landing simulation phase, the gearbox continues to transmit the necessary torque, it is acceptable to consider that the classification of component condition is Class 3 and can thus be considered a valid certification test result. Further component degradation resulting from continued running of the same test will not invalidate this result with respect to compliance with this requirement. Should an extended test be completed with a successful second landing simulation, the total duration can be considered applicable to the certification test result.
(3) Fixed time penalty.
Based on the condition of components necessary for continued safe flight or landing at the end of the certification test a fixed time penalty should be applied in accordance with the definitions below. This fixed time penalty should be 2 minutes for CLASS 1 (‘Good’ condition), 5 minutes for CLASS 2 (‘Fair’ condition), and 10 minutes for CLASS 3 (‘Imminent failure’ condition) with the CLASS defined based upon the following criteria.
CLASS 0 — Intact/serviceable
Parts in new condition. It is impractical to expect components to be in this condition after the test, but this classification is stated for reference only.
CLASS 1 — Good
— Parts are still well oil-wetted with little or no discolouration (light yellow to light/local blue).
— Local moderate scuffing of gear teeth and/or local moderate scorings on bearing-active surfaces is present.
— Hardened surfaces (gear teeth and bearing-active surfaces) may show slight/local reduction in hardness (maximum 2 points on the Rockwell C Hardness (HRC) scale).
— Normally, operation in these conditions should not significantly alter the vibration and noise signatures of the gearbox during test.
— Gearbox still transmits the required torque and rotates smoothly.
CLASS 2 — Fair
— Parts are almost completely dry, little residual oil in localised areas.
— Dark blue to brown discolouration is present, showing signs of uniform wear.
— Coatings such as silver plating are still visible but may be worn out locally or discoloured.
— Heavy localised scuffing on gear teeth as well wear on active surfaces of gear teeth are visible.
— Surface hardness may have been reduced more significantly (up to a maximum of 4 points on the HRC scale).
— Normally, operation in these conditions could cause moderate changes to the vibration and noise signatures of the gearbox during test.
— Gearbox still transmits the required torque.
CLASS 3 — Imminent failure
— Parts show evidence of plastic deformation or melting in local areas due to high temperatures.
— Macroscopic wear of some of the rolling elements of bearings and gear teeth, with appreciable alteration of dimensions and associated increases in clearances and play.
— Bearing cages are worn or with incipient breakage.
— Normally, operation in these conditions causes significant and audible changes to the vibration and noise signatures of the gearbox during test.
— The gearbox still transmits the required torque and is still capable of rotating immediately after test (after it has cooled down, it may be more difficult to rotate).
CLASS 4 — Failed
In this case, there is a complete and gross plastic deformation of parts, and bearing balls and rollers are melted. Parts in this conditions mean that the test specimen has failed, hence, this classification is also provided for reference only.
(4) Calculation of the maximum period of operation following loss of lubrication
Application of the factors described in (2) and (3) above can be represented by the following formula:
Td = ( Kr x Tc ) – Tp
where:
— Td is the Maximum Period of Operation Following Loss of Lubrication, for which confidence has been established and which is to be used as the basis for the period stated in the RFM emergency procedures. This period should not exceed Td;
— Kr is the confidence/reliability reduction factor defined in (2) above;
— Tc is the duration of the certification test (from low-pressure indication to end of test); and
— Tp is a fixed-time penalty to account for condition at the end of the test, as defined in (3) above.
(5) Secondary indication
Another possible means to increase confidence in the ability of the gearbox to continue to operate safely after suffering a loss of lubrication is to provide a secondary indication, which may indicate when the most critical mode of degradation has progressed to a level where gearbox functional failure may be imminent. If such a design feature is selected, the following considerations are necessary:
(i) evidence should be available, preferably from multiple tests, to provide confidence that the failure mode being monitored is always the most critical failure mode after a loss of lubrication, and that the rate of degradation up to the point of failure is understood;
(ii) if the oil pressure is normal, inhibition of the warning to the flight crew may be considered in order to reduce the likelihood of a false warning resulting in an instruction to ‘land immediately’; and
(iii) the availability/reliability of the warning should be justified; it should be possible to test the correct functioning of the sensor or warning during pre-flight/start-up checks or during routine maintenance.
(iv) noise and/or vibration detected by the crew should not be considered to be reliable secondary indications on their own.
[Amdt No: 29/5]
[Amdt No: 29/10]
CS 29.931 Shafting critical speed
ED Decision 2003/16/RM
(a) The critical speeds of any shafting must be determined by demonstration except that analytical methods may be used if reliable methods of analysis are available for the particular design.
(b) If any critical speed lies within, or close to, the operating ranges for idling, power-on, and autorotative conditions, the stresses occurring at that speed must be within safe limits. This must be shown by tests.
(c) If analytical methods are used and show that no critical speed lies within the permissible operating ranges, the margins between the calculated critical speeds and the limits of the allowable operating ranges must be adequate to allow for possible variations between the computed and actual values.
ED Decision 2003/16/RM
Each universal joint, slip joint, and other shafting joints whose lubrication is necessary for operation must have provision for lubrication.
CS 29.939 Turbine engine operating characteristics
ED Decision 2003/16/RM
(a) Turbine engine operating characteristics must be investigated in flight to determine that no adverse characteristics (such as stall, surge, or flameout) are present, to a hazardous degree, during normal and emergency operation within the range of operating limitations of the rotorcraft and of the engine.
(b) The turbine engine air inlet system may not, as a result of airflow distortion during normal operation, cause vibration harmful to the engine.
(c) For governor-controlled engines, it must be shown that there exists no hazardous torsional instability of the drive system associated with critical combinations of power, rotational speed, and control displacement.