Subpart E — Powerplant

GENERAL

ED Decision 2003/16/RM

(a)For the purpose of this Code, the powerplant installation includes each part of the rotorcraft (other than the main and auxiliary rotor structures) that:

(1)Is necessary for propulsion;

(2)Affects the control of the major propulsive units; or

(3)Affects the safety of the major propulsive units between normal inspections or overhauls.

(b)For each powerplant installation:

(1)The installation must comply with:

(i)The installation instructions provided under CS-E; and

(ii)The applicable provisions of this Subpart.

(2)Each component of the installation must be constructed, arranged, and installed to ensure its continued safe operation between normal inspections or overhauls for the range of temperature and altitude for which approval is requested.

(3)Accessibility must be provided to allow any inspection and maintenance necessary for continued airworthiness.

(4)Electrical interconnections must be provided to prevent differences of potential between major components of the installation and the rest of the rotorcraft.

(5)Axial and radial expansion of turbine engines may not affect the safety of the installation; and

(6)Design precautions must be taken to minimise the possibility of incorrect assembly of components and equipment essential to safe operation of the rotorcraft, except where operation with the incorrect assembly can be shown to be extremely improbable.

(c)For each powerplant and auxiliary power unit installation, it must be established that no single failure or malfunction or probable combination of failures will jeopardise the safe operation of the rotorcraft except that the failure of structural elements need not be considered if the probability of any such failure is extremely remote.

(d)Each auxiliary power unit installation must meet the applicable provisions of this Subpart.

ED Decision 2003/16/RM

(a)(Reserved)

(b)Category A; engine isolation. For each Category A rotorcraft, the powerplants must be arranged and isolated from each other to allow operation, in at least one configuration, so that the failure or malfunction of any engine, or the failure of any system that can affect any engine, will not –

(1)Prevent the continued safe operation of the remaining engines; or

(2)Require immediate action, other than normal pilot action with primary flight controls, by any crew member to maintain safe operation.

(c)Category A; control of engine rotation. For each Category A rotorcraft, there must be a means for stopping the rotation of any engine individually in flight, except that, for turbine engine installations, the means for stopping the engine need be provided only where necessary for safety. In addition –

(1)Each component of the engine stopping system that is located on the engine side of the firewall, and that might be exposed to fire, must be at least fire resistant; or

(2)Duplicate means must be available for stopping the engine and the controls must be where all are not likely to be damaged at the same time in case of fire.

(d)Turbine engine installation. For turbine engine installations,

(1)Design precautions must be taken to minimise the hazards to the rotorcraft in the event of an engine rotor failure; and,

(2)The powerplant systems associated with engine control devices, systems, and instrumentation must be designed to give reasonable assurance that those engine operating limitations that adversely affect engine rotor structural integrity will not be exceeded in service.

(e)Restart capability:

(1)A means to restart any engine in flight must be provided.

(2)Except for the in-flight shutdown of all engines, engine restart capability must be demonstrated throughout a flight envelope for the rotorcraft.

(3)Following the in-flight shutdown of all engines, in-flight engine restart capability must be provided.

ED Decision 2023/001/R

FRAGMENT CONTAINMENT

This AMC supplements FAA AC 29.903 with regard to the credit that can be taken from engine manufacturer data substantiating the capability of the engine to contain fragments.

(a)Blade containment

Single blade radial containment is a CS-E / CS-APU requirement. Full credit is given to engine certification for blade containment, and no specific certification activity is required at helicopter level for blade failure. This approach is supported by the in-service experience.

(b)Small debris containment at engine level

Some engine designs feature the capability to retain radially small debris, featuring, for instance, a reinforced casing or blade shedding capability.

The engine uncontained model features a small debris over a ±15° spread angle. Small fragments can be a collateral effect of either large or intermediate fragment release, but are released over larger spread angles, typically ±15°. Therefore, from a CS 29.903(d) point of view, no credit can be given to engine radial containment for small debris, which might however have other safety benefits.

(c)Rotor containment at engine or APU level

CS-APU has provisions to demonstrate rotor containment. For engines, while not required by CS-E, engine manufacturers might decide to design their engines featuring rotor containment systems, for all or specific rotating stages.

For engines, the containment capability is not required by CS-E and the corresponding data is not covered by the engine type certificate; the helicopter manufacturer should propose a mechanism to ensure that the data is valid, under their DOA or by validation through the engine type certificate whereas for an APU, CS-ETSO requirements are in place, and it can be expected that the data is covered by the ETSO issuance.

In-service experience has shown that such containment features successfully perform their intended purpose of retaining the biggest debris (large fragments). However, small debris can defeat the containment system, either by missing it or by exiting through damages caused by the large fragments. Rotor containment systems, as explained in paragraph f.(1) of AC 29.903C, still require some activity at helicopter level to ensure that the risks associated with uncontained engine or APU uncontained failure are adequately mitigated.

Note: For APUs, AMC 20.128A defines an acceptable model based upon debris exiting the containment system with a 1 % residual energy.

[Amdt No: 29/11]

ED Decision 2023/001/R

ENGINE RESTART CAPABILITY

This AMC replaces FAA AC 29-2C, § AC 29.903B and should be used when showing compliance with CS 29.903(e).

(a)Explanation

CS 29.903(e) requires that any engine must have a restart capability that has been demonstrated throughout a flight envelope to be certificated for the rotorcraft.

(b)Procedures

Compliance is usually shown by conducting actual in-flight restarts during flight tests or other tests in accordance with an approved test plan. However, CS 29.903(e)(2) does not require in-flight demonstration of restart capability for single-engine rotorcraft or for all-engine shutdown of multi-engine rotorcraft. In the past, engine restart capability for single-engine rotorcraft has been demonstrated on the ground taking into account altitude effects, warm engine characteristics, depleted battery, etc. However, latest-technology engines embody electronic engine controls (EEC or FADEC) that may have sophisticated starting or restarting laws. For these designs the engine restart capability demonstrated on ground may not provide the level of representativeness required and therefore applicants are encouraged to demonstrate the capability in flight. The minimum restart envelope for category A rotorcraft is discussed in AC 29.903A. 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 the engine in single-engine rotorcraft, or all engines in a multi-engine rotorcraft, is the primary requirement, and the means of providing this capability is left to the applicant. To minimise any potential altitude loss following the failure of one or more engines, engine restart should be available at the earliest opportunity. The engine certification should be checked to ensure that the flight manual instructions for in-flight restart are consistent with any specific engine restart requirements. If the procedure was only demonstrated on ground, this should be stated in the RFM.

[Amdt No: 29/11]

ED Decision 2003/16/RM

(a)Each engine must be installed to prevent the harmful vibration of any part of the engine or rotorcraft.

(b)The addition of the rotor and the rotor drive system to the engine may not subject the principal rotating parts of the engine to excessive vibration stresses. This must be shown by a vibration investigation.

ED Decision 2003/16/RM

For cooling fans that are a part of a powerplant installation the following apply:

(a)Category A. For cooling fans installed in Category A rotorcraft, it must be shown that a fan blade failure will not prevent continued safe flight either because of damage caused by the failed blade or loss of cooling air.

(b)Category B. For cooling fans installed in Category B rotorcraft, there must be means to protect the rotorcraft and allow a safe landing if a fan blade fails. It must be shown that :

(1)The fan blade would be contained in the case of a failure;

(2)Each fan is located so that a fan blade failure will not jeopardise safety; or

(3)Each fan blade can withstand an ultimate load of 1.5 times the centrifugal force expected in service, limited by either:

(i)The highest rotational speeds achievable under uncontrolled conditions; or

(ii)An overspeed limiting device.

(c)Fatigue evaluation. Unless a fatigue evaluation under CS 29.571 is conducted, it must be shown that cooling fan blades are not operating at resonant conditions within the operating limits of the rotorcraft.

ROTOR DRIVE SYSTEM

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]

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]

ED Decision 2023/001/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)for the unlikely event of a combined failure of both the normal-use lubrication system and the auxiliary lubrication system, the applicant should perform additional loss of lubrication tests simulating this condition. The aim is to substantiate additional RFM emergency procedures for this combined failure to ensure the capability of the drive system to sustain a minimum duration of safe operation. These procedures should instruct the flight crew to ‘LAND IMMEDIATELY’ unless the additional tests performed representing this failure mode demonstrate 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 lubrication 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]

[Amdt No: 29/11]

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.

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 2023/001/R

(a)Introduction

This AMC supplements FAA AC 29-2C, § AC 29.923 and should be used in conjunction with that AC when demonstrating compliance with CS 29.923.

(b)30-minute power rating

(1)Explanation

The option to establish a 30-minute power rating for turbine engines for rotorcraft has been introduced in CS-E Amendment 5 (published on 14 December 2018) with the creation of CS-E 40(b)(4). Means to demonstrate compliance with this requirement are provided in the associated AMC E 40(b)(3) and (b)(4) 30-Second OEI, 2-Minute OEI and 30-minute Power Ratings.

In particular, AMC E 40(b)(3) and (b)(4) mentions that ‘The 30-Minute Power rating may be set at any level between the Maximum Continuous up to and including the take-off rating, and may be used for multiple periods of up to 30 minutes each, at any time between the take-off and landing phases in any flight.’ In addition, CS-E 740(c)(2)(i) specifies additional running time for the endurance test for engines for rotorcraft for which approval with this rating is sought.

In comparison, the endurance test programme specified in CS 29.923 for rotorcraft rotor drive systems and control mechanisms:

—addresses the take-off power rating, which is ‘limited in use to a continuous period of not more than 5 minutes’ according to CS-Definitions, through the test runs specified in CS 29.923(b), and

—currently does not address the 30-minute power rating.

(2)Procedures

For applications including a 30-minute power rating, the applicant should consider that the approval of such rating should be supported by additional tests to be agreed with Agency, with the aim of determining that the rotor drive mechanism is safe considering the use of this specific power rating. In this context, the applicant may consider running additional test phases and/or extending the running time and/or increasing the minimum torque and speed conditions defined in CS 29.923 to include testing of this power rating.

[Amdt No: 29/11]

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]

ED Decision 2023/001/R

(a)Introduction

This AMC supplements FAA AC 29-2C, § AC 29.927 and should be used in conjunction with that AC when demonstrating compliance with CS 29.927.

(b)Variable rotor speed (NR)

(1)Explanation

The variable rotor speed (NR) function allows running at different NR levels to achieve, for instance, lower noise levels and better rotorcraft performance.

In addition to the endurance test prescribed in CS 29.923, additional tests may be necessary to demonstrate that rotor drive systems of rotorcraft with a variable NR are safe.

(2)Procedure

In order to substantiate an acceptable vibration and dynamic behaviour of rotor drive systems when using the available range of rotor speeds within the variable NR function, the applicant should consider performing specific test investigations, as prescribed in CS 29.927(a). The need for representative test runs at the different torque and rotor speed combinations, covering steady states and transient conditions to be encountered in operation, should be evaluated by and agreed with the Agency.

[Amdt No: 29/11]

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]

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.

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.

FUEL SYSTEMS

ED Decision 2003/16/RM

(a)Each fuel system must be constructed and arranged to ensure a flow of fuel at a rate and pressure established for proper engine and auxiliary power unit functioning under any likely operating conditions, including the manoeuvres for which certification is requested and during which the engine or auxiliary power unit is permitted to be in operation.

(b)Each fuel system must be arranged so that:

(1)No engine or fuel pump can draw fuel from more than one tank at a time; or

(2)There are means to prevent introducing air into the system.

(c)Each fuel system for a turbine engine must be capable of sustained operation throughout its flow and pressure range with fuel initially saturated with water at 27°C (80°F) and having 0.20 cm3 of free water per litre (0.75 cc per US-gallon) added and cooled to the most critical condition for icing likely to be encountered in operation.

ED Decision 2003/16/RM

Unless other means acceptable to the Agency are employed to minimise the hazard of fuel fires to occupants following an otherwise survivable impact (crash landing), the fuel systems must incorporate the design features of this paragraph. These systems must be shown to be capable of sustaining the static and dynamic deceleration loads of this paragraph, considered as ultimate loads acting alone, measured at the system component’s centre of gravity without structural damage to the system components, fuel tanks, or their attachments that would leak fuel to an ignition source.

(a)Drop test requirements. Each tank, or the most critical tank, must be drop-tested as follows:

(1)The drop height must be at least 15.2m (50 ft).

(2)The drop impact surface must be non deforming.

(3)The tanks must be filled with water to 80% of the normal, full capacity.

(4)The tank must be enclosed in a surrounding structure representative of the installation unless it can be established that the surrounding structure is free of projections or other design features likely to contribute to rupture of the tank.

(5)The tank must drop freely and impact in a horizontal position ± 10°.

(6)After the drop test, there must be no leakage.

(b)Fuel tank load factors. Except for fuel tanks located so that tank rupture with fuel release to either significant ignition sources, such as engines, heaters, and auxiliary power units, or occupants is extremely remote, each fuel tank must be designed and installed to retain its contents under the following ultimate inertial load factors, acting alone.

(1)For fuel tanks in the cabin –

(i)Upward – 4 g.

(ii)Forward – 16 g.

(iii)Sideward – 8 g.

(iv)Downward – 20 g.

(2)For fuel tanks located above or behind the crew or passenger compartment that, if loosened, could injure an occupant in an emergency landing –

(i)Upward – 1.5 g.

(ii)Forward – 8 g.

(iii)Sideward – 2 g.

(iv)Downward – 4 g.

(3)For fuel tanks in other areas –

(i)Upward –1.5 g.

(ii)Forward – 4 g.

(iii)Sideward – 2 g.

(iv)Downward – 4 g.

(c)Fuel line self-sealing breakaway couplings. Self-sealing breakaway couplings must be installed unless hazardous relative motion of fuel system components to each other or to local rotorcraft structure is demonstrated to be extremely improbable or unless other means are provided. The couplings or equivalent devices must be installed at all fuel tank- to-fuel line connections, tank-to-tank interconnects, and at other points in the fuel system where local structural deformation could lead to release of fuel.

(1)The design and construction of self- sealing breakaway couplings must incorporate the following design features:

(i)The load necessary to separate a breakaway coupling must be between 25 and 50% of the minimum ultimate failure load (ultimate strength) of the weakest component in the fluid-carrying line. The separation load must in no case be less than 1334 N (300 pounds), regardless of the size of the fluid line.

(ii)A breakaway coupling must separate whenever its ultimate load (as defined in sub-paragraph (c)(1)(i)) is applied in the failure modes most likely to occur.

(iii)All breakaway coupling must incorporate design provisions to visually ascertain that the coupling is locked together (leak-free) and is open during normal installation and service.

(iv)All breakaway couplings must incorporate design provisions to prevent uncoupling or unintended closing due to operational shocks, vibrations, or accelerations.

(v)No breakaway coupling design may allow the release of fuel once the coupling has performed its intended function.

(2)All individual breakaway couplings, coupling fuel feed systems, or equivalent means must be designed, tested, installed, and maintained so inadvertent fuel shutoff in flight is improbable in accordance with CS 29.955(a) and must comply with the fatigue evaluation requirements of CS 29.571 without leaking.

(3)Alternate, equivalent means to the use of breakaway couplings must not create a survivable impact-induced load on the fuel line to which it is installed greater than 25 to 50% of the ultimate load (strength) of the weakest component in the line and must comply with the fatigue requirements of CS 29.571 without leaking.

(d)Frangible or deformable structural attachments. Unless hazardous relative motion of fuel tanks and fuel system components to local rotorcraft structure is demonstrated to be extremely improbable in an otherwise survivable impact, frangible or locally deformable attachments of fuel tanks and fuel system components to local rotorcraft structure must be used. The attachment of fuel tanks and fuel system components to local rotorcraft structure. whether frangible or locally deformable, must be designed such that its separation or relative local deformation will occur without rupture or local tearout of the fuel tank or fuel system component that will cause fuel leakage. The ultimate strength of frangible or deformable attachments must be as follows:

(1)The load required to separate a frangible attachment from its support structure, or deform a locally deformable attachment relative to its support structure, must be between 25 and 50% of the minimum ultimate load (ultimate strength) of the weakest component in the attached system. In no case may the load be less than 1334 N (300 pounds).

(2)A frangible or locally deformable attachment must separate or locally deform as intended whenever its ultimate load (as defined in sub-paragraph (d)(1)) is applied in the modes most likely to occur.

(3)All frangible or locally deformable attachments must comply with the fatigue requirements of CS 29.571.

(e)Separation of fuel and ignition sources. To provide maximum crash resistance, fuel must be located as far as practicable from all occupiable areas and from all potential ignition sources.

(f)Other basic mechanical design criteria. Fuel tanks, fuel lines, electrical wires and electrical devices must be designed, constructed, and installed, as far as practicable, to be crash resistant.

(g)Rigid or semi-rigid fuel tanks. Rigid or semi-rigid fuel tank or bladder walls must be impact and tear resistant.

ED Decision 2003/16/RM

(a)For Category A rotorcraft:

(1)The fuel system must meet the requirements of CS 29.903(b); and

(2)Unless other provisions are made to meet sub-paragraph (a)(1) , the fuel system must allow fuel to be supplied to each engine through a system independent of those parts of each system supplying fuel to other engines.

(b)Each fuel system for a multi-engine Category B rotorcraft must meet the requirements of sub-paragraph (a)(2). However, separate fuel tanks need not be provided for each engine.

ED Decision 2003/16/RM

The fuel system must be designed and arranged to prevent the ignition of fuel vapour within the system by:

(a)Direct lightning strikes to areas having a high probability of stroke attachment;

(b)Swept lightning strokes to areas where swept strokes are highly probable; and

(c)Corona and streamering at fuel vent outlets.

ED Decision 2012/022/R

(a)General. The fuel system for each engine must provide the engine with at least 100% of the fuel required under all operating and manoeuvring conditions to be approved for the rotorcraft, including, as applicable, the fuel required to operate the engines under the test conditions required by CS 29.927. Unless equivalent methods are used, compliance must be shown by test during which the following provisions are met, except that combinations of conditions which are shown to be improbable need not be considered.

(1)The fuel pressure, corrected for accelerations (load factors), must be within the limits specified by the engine type certificate data sheet.

(2)The fuel level in the tank may not exceed that established as the unusable fuel supply for that tank under CS 29.959, plus that necessary to conduct the test.

(3)The fuel head between the tank and the engine must be critical with respect to rotorcraft flight attitudes.

(4)The fuel flow transmitter, if installed, and the critical fuel pump (for pump-fed systems) must be installed to produce (by actual or simulated failure) the critical restriction to fuel flow to be expected from component failure.

(5)Critical values of engine rotational speed, electrical power, or other sources of fuel pump motive power must be applied.

(6)Critical values of fuel properties which adversely affect fuel flow are applied during demonstrations of fuel flow capability.

(7)The fuel filter required by CS 29.997 is blocked to the degree necessary to simulate the accumulation of fuel contamination required to activate the indicator required by CS 29.1305(a)(18).

(b)Fuel transfer system. If normal operation of the fuel system requires fuel to be transferred to another tank, the transfer must occur automatically via a system which has been shown to maintain the fuel level in the receiving tank within acceptable limits during flight or surface operation of the rotorcraft.

(c)Multiple fuel tanks. If an engine can be supplied with fuel from more than one tank, the fuel system, in addition to having appropriate manual switching capability, must be designed to prevent interruption of fuel flow to the engine, without attention by the flight crew, when any tank supplying fuel to that engine is depleted of usable fuel during normal operation and any other tank that normally supplies fuel to that engine alone contains usable fuel.

[Amdt 29/3]

ED Decision 2003/16/RM

(a)Where tank outlets are interconnected and allow fuel to flow between them due to gravity or flight accelerations, it must be impossible for fuel to flow between tanks in quantities great enough to cause overflow from the tank vent in any sustained flight condition.

(b)If fuel can be pumped from one tank to another in flight:

(1)The design of the vents and the fuel transfer system must prevent structural damage to tanks from overfilling; and

(2)There must be means to warn the crew before overflow through the vents occurs.

ED Decision 2003/16/RM

The unusable fuel supply for each tank must be established as not less than the quantity at which the first evidence of malfunction occurs under the most adverse fuel feed condition occurring under any intended operations and flight manoeuvres involving that tank.

ED Decision 2023/001/R

This AMC supplements FAA AC 29.959.

This AMC provides clarification on the acceptability of analyses and ground testing which could be used as means of compliance if supported by actual flight test data.

FAA AC 29-2C, § AC 29.959 provides some guidance by focusing on a flight/test demonstration as being directly in line with the rule intent to validate ‘… any intended operations and flight manoeuvres …’, but also provides for acceptability of analyses and ground testing.

In order to accept a demonstration by laboratory test with partial flight or ground test, the applicant should demonstrate the ability of the proposed substantiation method (bench testing, complemented by analysis and /or ground test) to cover the effects offered normally by the flight-testing environment.

In case the full flight-testing environment cannot be accurately simulated, it is necessary to either:

—revert to compliance demonstration based on flight test; or

—apply some conservatism factors on the unusable fuel quantity value resulting from the laboratory testing to determine the final unusable fuel value.

Any (steady or transitory) engine abnormal operation/malfunction has to be taken as an indication that the fuel in the tank is becoming unusable.

[Amdt No: 29/11]

ED Decision 2003/16/RM

Each suction lift fuel system and other fuel systems conducive to vapour formation must be shown to operate satisfactorily (within certification limits) when using fuel at the most critical temperature for vapour formation under critical operating conditions including, if applicable, the engine operating conditions defined by CS 29.927(b)(1) and (b)(2).

ED Decision 2003/16/RM

(a)Each fuel tank must be able to withstand, without failure, the vibration, inertia, fluid, and structural loads to which it may be subjected in operation.

(b)Each flexible fuel tank bladder or liner must be approved or shown to be suitable for the particular application and must be puncture resistant. Puncture resistance must be shown by meeting the ETSO-C80, paragraph 16.0, requirements using a minimum puncture force of 1646 N (370 pounds).

(c)Each integral fuel tank must have facilities for inspection and repair of its interior.

(d)The maximum exposed surface temperature of all components in the fuel tank must be less by a safe margin than the lowest expected auto-ignition temperature of the fuel or fuel vapour in the tank. Compliance with this requirement must be shown under all operating conditions and under all normal or malfunction conditions of all components inside the tank.

(e)Each fuel tank installed in personnel compartments must be isolated by fume-proof and fuel-proof enclosures that are drained and vented to the exterior of the rotorcraft. The design and construction of the enclosures must provide necessary protection for the tank, must be crash resistant during a survivable impact in accordance with CS 29.952, and must be adequate to withstand loads and abrasions to be expected in personnel compartments.

ED Decision 2003/16/RM

(a)Each fuel tank must be able to withstand the applicable pressure tests in this paragraph without failure or leakage. If practicable, test pressures may be applied in a manner simulating the pressure distribution in service.

(b)Each conventional metal tank, each non- metallic tank with walls that are not supported by the rotorcraft structure, and each integral tank must be subjected to a pressure of 24 kPa (3.5 psi) unless the pressure developed during maximum limit acceleration or emergency deceleration with a full tank exceeds this value, in which case a hydrostatic head, or equivalent test, must be applied to duplicate the acceleration loads as far as possible. However, the pressure need not exceed 24 kPa (3.5 psi) on surfaces not exposed to the acceleration loading.

(c)Each non-metallic tank with walls supported by the rotorcraft structure must be subjected to the following tests:

(1)A pressure test of at least 14 kPa (2.0 psi). This test may be conducted on the tank alone in conjunction with the test specified in subparagraph (c)(2).

(2)A pressure test, with the tank mounted in the rotorcraft structure, equal to the load developed by the reaction of the contents, with the tank full, during maximum limit acceleration or emergency deceleration. However, the pressure need not exceed 14 kPa (2.0 psi) on surfaces not exposed to the acceleration loading.

(d)Each tank with large unsupported or unstiffened flat areas, or with other features whose failure or deformation could cause leakage, must be subjected to the following test or its equivalent:

(1)Each complete tank assembly and its supports must be vibration tested while mounted to simulate the actual installation.

(2)The tank assembly must be vibrated for 25 hours while two-thirds full of any suitable fluid. The amplitude of vibration may not be less than 0.8 mm (one thirty-second of an inch), unless otherwise substantiated.

(3)The test frequency of vibration must be as follows:

(i)If no frequency of vibration resulting from any rpm within the normal operating range of engine or rotor system speeds is critical, the test frequency of vibration, in number of cycles per minute, must, unless a frequency based on a more rational analysis is used, be the number obtained by averaging the maximum and minimum power-on engine speeds (rpm) for reciprocating engine powered rotorcraft or 2000 cpm for turbine engine powered rotorcraft.

(ii)If only one frequency of vibration resulting from any rpm within the normal operating range of engine or rotor system speeds is critical, that frequency of vibration must be the test frequency.

(iii)If more than one frequency of vibration resulting from any rpm within the normal operating range of engine or rotor system speeds is critical, the most critical of these frequencies must be the test frequency.

(4)Under sub-paragraph (d)(3)(ii) and (iii), the time of test must be adjusted to accomplish the same number of vibration cycles as would be accomplished in 25 hours at the frequency specified in sub-paragraph (d)(3)(i).

(5)During the test the tank assembly must be rocked at the rate of 16 to 20 complete cycles per minute through an angle of 15° on both sides of the horizontal (30° total), about the most critical axis, for 25 hours. If motion about more than one axis is likely to be critical, the tank must be rocked about each critical axis for 12½ hours.

ED Decision 2023/001/R

This AMC supplements FAA AC 29.965.

(a)Tests to be performed

CS 29.965 (a), (b) and (c) deal with the fuel tank pressure testing as follows:

Sub-paragraph (a) prescribes general testing conditions.

Sub-paragraph (b) prescribes testing conditions for conventional metal tanks, integral tanks and for non-metallic tanks with walls that are not supported by the rotorcraft structure.

Sub-paragraph (c) prescribes pressure testing for non-metallic tanks with walls supported by the rotorcraft structure.

CS 29.965(d) deals with fuel tank vibration & slosh testing with large unsupported or unstiffened flat areas. A clear definition of ‘large unsupported or unstiffened flat area’ is provided in FAA AC 29-2C, § AC 29.965.

The intent of the tests required in sub-paragraphs (a), (b) or (c) does not cover the intent of the test required in sub-paragraph (d) and vice versa.

Therefore pressure tests, as prescribed under (a), (b) or (c), and the vibration and slosh test, as prescribed under (d), should be performed.

(b)Use of MIL-T-6396

AC 29.965 (c)(6) recognises the use of MIL-T-6396 to support the demonstration of compliance with CS 29.965. However, few clarifications are required to appropriately make use of this standard.

Combined tests

To be in line with the CS 29.965(d) requirement, the slosh and vibration test conditions shall be simultaneously applied to the test article.

Therefore the use of MIL-T-6396 should be restricted to paragraph 4.6.6 ‘Simultaneous Slosh and Vibration test’. Individual/separate performance of paragraph 4.6.7 ‘Vibrations test’ and paragraph 4.6.8 ‘Slosh Test’ of the referenced MIL Specification are not considered to be appropriate.

Application of the slosh effect during the test as prescribed in CS 29.965(d)(5):

CS 29.965(d)(5) prescribes the performance of the vibration test for 25h at 16 to 20 slosh cycles per minute (cpm).

MIL-T-6396 proposes two test durations in paragraph 4.6.6:

Option 1: Vibrate for 25h at 16 to 20 slosh cpm, which is identical to the CS 29.965 (d)(5) requirement.

or

Option 2: Vibrate for 25h at 10 to 16 slosh cpm with 15 hours of additional test at 10 to 16 slosh cpm.

While it is recognised that Option 2 is appropriate in terms of number of cycles to which the test article is finally submitted (extended testing duration to compensate for the reduction of rocking frequency), it potentially omits a major effect introduced by the higher rocking frequency which may induce more severe structural effects due to the fluid dynamics and subsequent shocks.

An applicant wishing to use Option 2 should demonstrate by analysis, test or a combination thereof, that the reduction of rocking frequency compared to Option 1 has no positive effect to the test results.

[Amdt No: 29/11]

ED Decision 2003/16/RM

(a)Each fuel tank must be supported so that tank loads are not concentrated on unsupported tank surfaces. In addition:

(1)There must be pads, if necessary, to prevent chafing between each tank and its supports;

(2)The padding must be non-absorbent or treated to prevent the absorption of fuel;

(3)If flexible tank liners are used, they must be supported so that they are not required to withstand fluid loads; and

(4)Each interior surface of tank compartments must be smooth and free of projections that could cause wear of the liner, unless:

(i)There are means for protection of the liner at those points; or

(ii)The construction of the liner itself provides such protection.

(b)Any spaces adjacent to tank surfaces must be adequately ventilated to avoid accumulation of fuel or fumes in those spaces due to minor leakage. If the tank is in a sealed compartment, ventilation may be limited to drain holes that prevent clogging and that prevent excessive pressure resulting from altitude changes. If flexible tank liners are installed, the venting arrangement for the spaces between the liner and its container must maintain the proper relationship to tank vent pressures for any expected flight condition.

(c)The location of each tank must meet the requirements of CS 29.1185(b) and (c).

(d)No rotorcraft skin immediately adjacent to a major air outlet from the engine compartment may act as the wall of an integral tank.

ED Decision 2003/16/RM

Each fuel tank or each group of fuel tanks with interconnected vent systems must have an expansion space of not less than 2% of the combined tank capacity. It must be impossible to fill the fuel tank expansion space inadvertently with the rotorcraft in the normal ground attitude.

ED Decision 2003/16/RM

(a)Each fuel tank must have a sump with a capacity of not less than the greater of:

(1)0.10% of the tank capacity; or

(2)0.24 litres (0.05 Imperial gallon/one sixteenth US gallon).

(b)The capacity prescribed in sub-paragraph (a) must be effective with the rotorcraft in any normal attitude, and must be located so that the sump contents cannot escape through the tank outlet opening.

(c)Each fuel tank must allow drainage of hazardous quantities of water from each part of the tank to the sump with the rotorcraft in any ground attitude to be expected in service.

(d)Each fuel tank sump must have a drain that allows complete drainage of the sump on the ground.

ED Decision 2003/16/RM

(a)Each fuel tank filler connection must prevent the entrance of fuel into any part of the rotorcraft other than the tank itself during normal operations and must be crash resistant during a survivable impact in accordance with CS 29.952(c). In addition:

(1)Each filler must be marked as prescribed in CS 29.1557(c)(1);

(2)Each recessed filler connection that can retain any appreciable quantity of fuel must have a drain that discharges clear of the entire rotorcraft; and

(3)Each filler cap must provide a fuel- tight seal under the fluid pressure expected in normal operation and in a survivable impact.

(b)Each filler cap or filler cap cover must warn when the cap is not fully locked or seated on the filler connection.

ED Decision 2003/16/RM

(a)Fuel tank vents. Each fuel tank must be vented from the top part of the expansion space so that venting is effective under normal flight conditions. In addition:

(1)The vents must be arranged to avoid stoppage by dirt or ice formation;

(2)The vent arrangement must prevent siphoning of fuel during normal operation;

(3)The venting capacity and vent pressure levels must maintain acceptable differences of pressure between the interior and exterior of the tank, during:

(i)Normal flight operation;

(ii)Maximum rate of ascent and descent; and

(iii)Refuelling and defuelling (where applicable);

(4)Airspaces of tanks with interconnected outlets must be interconnected;

(5)There may be no point in any vent line where moisture can accumulate with the rotorcraft in the ground attitude or the level flight attitude, unless drainage is provided;

(6)No vent or drainage provision may end at any point:

(i)Where the discharge of fuel from the vent outlet would constitute a fire hazard; or

(ii)From which fumes could enter personnel compartments; and

(7)The venting system must be designed to minimise spillage of fuel through the vents to an ignition source in the event of a rollover during landing, ground operations, or a survivable impact.

(b)Carburettor vapour vents. Each carburettor with vapour elimination connections must have a vent line to lead vapours back to one of the fuel tanks. In addition –

(1)Each vent system must have means to avoid stoppage by ice; and

(2)If there is more than one fuel tank, and it is necessary to use the tanks in a definite sequence, each vapour vent return line must lead back to the fuel tank used for take-off and landing.

ED Decision 2003/16/RM

(a)There must be a fuel strainer for the fuel tank outlet or for the booster pump. This strainer must:

(1)For reciprocating engine powered rotorcraft, have 3 to 6 meshes per cm (8 to 16 meshes per inch); and

(2)For turbine engine powered rotorcraft, prevent the passage of any object that could restrict fuel flow or damage any fuel system component.

(b)The clear area of each fuel tank outlet strainer must be at least five times the area of the outlet line.

(c)The diameter of each strainer must be at least that of the fuel tank outlet.

(d)Each finger strainer must be accessible for inspection and cleaning.

ED Decision 2003/16/RM

(a)Each fuelling connection below the fuel level in each tank must have means to prevent the escape of hazardous quantities of fuel from that tank in case of malfunction of the fuel entry valve.

(b)For systems intended for pressure refuelling, a means in addition to the normal means for limiting the tank content must be installed to prevent damage to the tank in case of failure of the normal means.

(c)The rotorcraft pressure fuelling system (not fuel tanks and fuel tank vents) must withstand an ultimate load that is 2.0 times the load arising from the maximum pressure, including surge, that is likely to occur during fuelling. The maximum surge pressure must be established with any combination of tank valves being either intentionally or inadvertently closed.

(d)The rotorcraft defuelling system (not including fuel tanks and fuel tank vents) must withstand an ultimate load that is 2.0 times the load arising from the maximum permissible defuelling pressure (positive or negative) at the rotorcraft fuelling connection.

FUEL SYSTEM COMPONENTS

ED Decision 2003/16/RM

(a)Compliance with CS 29.955 must not be jeopardised by failure of:

(1)Any one pump except pumps that are approved and installed as parts of a type certificated engine; or

(2)Any component required for pump operation except the engine served by that pump.

(b)The following fuel pump installation requirements apply:

(1)When necessary to maintain the proper fuel pressure:

(i)A connection must be provided to transmit the carburettor air intake static pressure to the proper fuel pump relief valve connection; and

(ii)The gauge balance lines must be independently connected to the carburettor inlet pressure to avoid incorrect fuel pressure readings.

(2)The installation of fuel pumps having seals or diaphragms that may leak must have means for draining leaking fuel.

(3)Each drain line must discharge where it will not create a fire hazard.

ED Decision 2003/16/RM

(a)Each fuel line must be installed and supported to prevent excessive vibration and to withstand loads due to fuel pressure, valve actuation, and accelerated flight conditions.

(b)Each fuel line connected to components of the rotorcraft between which relative motion could exist must have provisions for flexibility.

(c)Each flexible connection in fuel lines that may be under pressure or subjected to axial loading must use flexible hose assemblies.

(d)Flexible hose must be approved.

(e)No flexible hose that might be adversely affected by high temperatures may be used where excessive temperatures will exist during operation or after engine shutdown.

ED Decision 2003/16/RM

In addition to meeting the requirements of CS 29.1189, each fuel valve must:

(a)Reserved.

(b)Be supported so that no loads resulting from their operation or from accelerated flight conditions are transmitted to the lines attached to the valve.

ED Decision 2003/16/RM

There must be a fuel strainer or filter between the fuel tank outlet and the inlet of the first fuel system component which is susceptible to fuel contamination, including but not limited to the fuel metering device or an engine positive displacement pump, whichever is nearer the fuel tank outlet. This fuel strainer or filter must:

(a)Be accessible for draining and cleaning and must incorporate a screen or element which is easily removable;

(b)Have a sediment trap and drain, except that it need not have a drain if the strainer or filter is easily removable for drain purposes;

(c)Be mounted so that its weight is not supported by the connecting lines or by the inlet or outlet connections of the strainer or filter itself, unless adequate strength margins under all loading conditions are provided in the lines and connections; and

(d)Provide a means to remove from the fuel any contaminant which would jeopardise the flow of fuel through rotorcraft or engine fuel system components required for proper rotorcraft or engine fuel system operation.

ED Decision 2003/16/RM

(a)There must be at least one accessible drain at the lowest point in each fuel system to completely drain the system with the rotorcraft in any ground attitude to be expected in service.

(b)Each drain required by sub-paragraph (a) including the drains prescribed in CS 29.971 must:

(1)Discharge clear of all parts of the rotorcraft;

(2)Have manual or automatic means to ensure positive closure in the off position; and

(3)Have a drain valve:

(i)That is readily accessible and which can be easily opened and closed; and

(ii)That is either located or protected to prevent fuel spillage in the event of a landing with landing gear retracted.

ED Decision 2003/16/RM

If a fuel jettisoning system is installed, the following apply:

(a)Fuel jettisoning must be safe during all flight regimes for which jettisoning is to be authorised.

(b)In showing compliance with sub-paragraph (a), it must be shown that:

(1)The fuel jettisoning system and its operation are free from fire hazard;

(2)No hazard results from fuel or fuel vapours which impinge on any part of the rotorcraft during fuel jettisoning; and

(3)Controllability of the rotorcraft remains satisfactory throughout the fuel jettisoning operation.

(c)Means must be provided to automatically prevent jettisoning fuel below the level required for an all-engine climb at maximum continuous power from sea-level to 1524 m (5000 ft) altitude and cruise thereafter for 30 minutes at maximum range engine power.

(d)The controls for any fuel jettisoning system must be designed to allow flight personnel (minimum crew) to safely interrupt fuel jettisoning during any part of the jettisoning operation.

(e)The fuel jettisoning system must be designed to comply with the powerplant installation requirements of CS 29.901(c).

(f)An auxiliary fuel jettisoning system which meets the requirements of sub-paragraphs (a), (b), (d) and (e) may be installed to jettison additional fuel provided it has separate and independent controls.

OIL SYSTEM

ED Decision 2003/16/RM

(a)Each engine must have an independent oil system that can supply it with an appropriate quantity of oil at a temperature not above that safe for continuous operation.

(b)The usable oil capacity of each system may not be less than the product of the endurance of the rotorcraft under critical operating conditions and the maximum allowable oil consumption of the engine under the same conditions, plus a suitable margin to ensure adequate circulation and cooling. Instead of a rational analysis of endurance and consumption, a usable oil capacity of 3.8 litres (0.83 Imperial gallon/1 US gallon) for each 151 litres (33.3 Imperial gallons/40 US gallons) of usable fuel may be used for reciprocating engine installations.

(c)Oil-fuel ratios lower than those prescribed in sub-paragraph (b) may be used if they are substantiated by data on the oil consumption of the engine.

(d)The ability of the engine oil cooling provisions to maintain the oil temperature at or below the maximum established value must be shown under the applicable requirements of CS 29.1041 to 29.1049.

ED Decision 2003/16/RM

(a)Installation. Each oil tank installation must meet the requirements of CS 29.967.

(b)Expansion space. Oil tank expansion space must be provided so that –

(1)Each oil tank used with a reciprocating engine has an expansion space of not less than the greater of 10% of the tank capacity or 1.9 litres (0.42 Imperial gallon/0.5 US gallon), and each oil tank used with a turbine engine has an expansion space of not less than 10% of the tank capacity;

(2)Each reserve oil tank not directly connected to any engine has an expansion space of not less than 2% of the tank capacity; and

(3)It is impossible to fill the expansion space inadvertently with the rotorcraft in the normal ground attitude.

(c)Filler connections. Each recessed oil tank filler connection that can retain any appreciable quantity of oil must have a drain that discharges clear of the entire rotorcraft. In addition –

(1)Each oil tank filler cap must provide an oil-tight seal under the pressure expected in operation;

(2)For Category A rotorcraft, each oil tank filler cap or filler cap cover must incorporate features that provide a warning when caps are not fully locked or seated on the filler connection; and

(3)Each oil filler must be marked under CS 29.1557(c)(2).

(d)Vent. Oil tanks must be vented as follows:

(1)Each oil tank must be vented from the top part of the expansion space so that venting is effective under all normal flight conditions.

(2)Oil tank vents must be arranged so that condensed water vapour that might freeze and obstruct the line cannot accumulate at any point.

(e)Outlet. There must be means to prevent entrance into the tank itself, or into the tank outlet, of any object that might obstruct the flow of oil through the system. No oil tank outlet may be enclosed by a screen or guard that would reduce the flow of oil below a safe value at any operating temperature. There must be a shutoff valve at the outlet of each oil tank used with a turbine engine unless the external portion of the oil system (including oil tank supports) is fireproof.

(f)Flexible liners. Each flexible oil tank liner must be approved or shown to be suitable for the particular installation.

ED Decision 2003/16/RM

Each oil tank must be designed and installed so that –

(a)It can withstand, without failure, any vibration, inertia, and fluid loads to which it may be subjected in operation; and

(b)It meets the requirements of CS 29.965, except that instead of the pressure specified in CS 29.965(b) –

(1)For pressurised tanks used with a turbine engine, the test pressure may not be less than 34 kPa (5 psi) plus the maximum operating pressure of the tank; and

(2)For all other tanks, the test pressure may not be less than 34 kPa (5 psi).

ED Decision 2003/16/RM

(a)Each oil line must meet the requirements of CS 29.993.

(b)Breather lines must be arranged so that –

(1)Condensed water vapour that might freeze and obstruct the line cannot accumulate at any point;

(2)The breather discharge will not constitute a fire hazard if foaming occurs, or cause emitted oil to strike the pilot’s windshield; and

(3)The breather does not discharge into the engine air induction system.

ED Decision 2003/16/RM

(a)Each turbine engine installation must incorporate an oil strainer or filter through which all of the engine oil flows and which meets the following requirements:

(1)Each oil strainer or filter that has a bypass must be constructed and installed so that oil will flow at the normal rate through the rest of the system with the strainer or filter completely blocked.

(2)The oil strainer or filter must have the capacity (with respect to operating limitations established for the engine) to ensure that engine oil system functioning is not impaired when the oil is contaminated to a degree (with respect to particle size and density) that is greater than that established for the engine under CS-E.

(3)The oil strainer or filter, unless it is installed at an oil tank outlet, must incorporate a means to indicate contamination before it reaches the capacity established in accordance with subparagraph (a)(2).

(4)The bypass of a strainer or filter must be constructed and installed so that the release of collected contaminants is minimised by appropriate location of the bypass to ensure that collected contaminants are not in the bypass flow path.

(5)An oil strainer or filter that has no bypass, except one that is installed at an oil tank outlet, must have a means to connect it to the warning system required in CS 29.1305(a)(18).

(b)Each oil strainer or filter in a powerplant installation using reciprocating engines must be constructed and installed so that oil will flow at the normal rate through the rest of the system with the strainer or filter element completely blocked.

ED Decision 2003/16/RM

A drain (or drains) must be provided to allow safe drainage of the oil system. Each drain must –

(a)Be accessible; and

(b)Have manual or automatic means for positive locking in the closed position.

ED Decision 2003/16/RM

(a)Each oil radiator must be able to withstand any vibration, inertia, and oil pressure loads to which it would be subjected in operation.

(b)Each oil radiator air duct must be located, or equipped, so that, in case of fire, and with the airflow as it would be with and without the engine operating, flames cannot directly strike the radiator.

ED Decision 2003/16/RM

(a)Each oil shutoff must meet the requirements of CS 29.1189.

(b)The closing of oil shutoffs may not prevent autorotation.

(c)Each oil valve must have positive stops or suitable index provisions in the ‘on’ and ‘off’ positions and must be supported so that no loads resulting from its operation or from accelerated flight conditions are transmitted to the lines attached to the valve.

ED Decision 2003/16/RM

(a)The oil system for components of the rotor drive system that require continuous lubrication must be sufficiently independent of the lubrication systems of the engine(s) to ensure:

(1)Operation with any engine inoperative; and

(2)Safe autorotation.

(b)Pressure lubrication systems for transmissions and gearboxes must comply with the requirements of CS 29.1013, sub-paragraphs (c), (d) and (f) only, CS 29.1015, 29.1017, 29.1021, 29.1023 and 29.1337(d). In addition, the system must have:

(1)An oil strainer or filter through which all the lubricant flows, and must:

(i)Be designed to remove from the lubricant any contaminant which may damage transmission and drive system components or impede the flow of lubricant to a hazardous degree; and

(ii)Be equipped with a bypass constructed and installed so that:

(A)The lubricant will flow at the normal rate through the rest of the system with the strainer or filter completely blocked; and

(B)The release of collected contaminants is minimised by appropriate location of the bypass to ensure that collected contaminants are not in the bypass flow path;

(iii)Be equipped with a means to indicate collection of contaminants on the filter or strainer at or before opening of the bypass;

(2)For each lubricant tank or sump outlet supplying lubrication to rotor drive systems and rotor drive system components, a screen to prevent entrance into the lubrication system of any object that might obstruct the flow of lubricant from the outlet to the filter required by sub-paragraph (b)(1). The requirements of sub-paragraph (b)(1) do not apply to screens installed at lubricant tank or sump outlets.

(c)Splash type lubrication systems for rotor drive system gearboxes must comply with CS 29.1021 and 29.1337(d).

COOLING

ED Decision 2003/16/RM

(a)The powerplant and auxiliary power unit cooling provisions must be able to maintain the temperatures of powerplant components, engine fluids, and auxiliary power unit components and fluids within the temperature limits established for these components and fluids, under ground, water, and flight operating conditions for which certification is requested, and after normal engine or auxiliary power shut-down, or both.

(b)There must be cooling provisions to maintain the fluid temperatures in any power transmission within safe values under any critical surface (ground or water) and flight operating conditions.

(c)Except for ground-use-only auxiliary power units, compliance with sub-paragraphs (a) and (b) must be shown by flight tests in which the temperatures of selected powerplant component and auxiliary power unit component, engine, and transmission fluids are obtained under the conditions prescribed in those paragraphs.

ED Decision 2003/16/RM

(a)General. For the tests prescribed in CS 29.1041(c), the following apply:

(1)If the tests are conducted under conditions deviating from the maximum ambient atmospheric temperature specified in sub-paragraph (b), the recorded powerplant temperatures must be corrected under sub-paragraphs (c) and (d), unless a more rational correction method is applicable.

(2)No corrected temperature determined under sub-paragraph (a)(1) may exceed established limits.

(3)The fuel used during the cooling tests must be of the minimum grade approved for the engines, and the mixture settings must be those used in normal operation.

(4)The test procedures must be as prescribed in CS 29.1045 to 29.1049.

(5)For the purposes of the cooling tests, a temperature is ‘stabilised’ when its rate of change is less than 1°C (2°F) per minute.

(b)Maximum ambient atmospheric pressure. A maximum ambient atmospheric temperature corresponding to sea-level conditions of at least 38°C (100°F) must be established. The assumed temperature lapse rate is 2.0°C (3.6°F) per thousand feet of altitude above sea-level until a temperature of –56.5°C (–69.7°F) is reached, above which altitude the temperature is considered constant at –56.5°C (–69.7°F). However, for winterisation installations, the applicant may select a maximum ambient atmospheric temperature corresponding to sea-level conditions of less than 38°C (100°F).

(c)Correction factor (except cylinder barrels). Unless a more rational correction applies, temperatures of engine fluids and powerplant components (except cylinder barrels) for which temperature limits are established, must be corrected by adding to them the difference between the maximum ambient atmospheric temperature and the temperature of the ambient air at the time of the first occurrence of the maximum component or fluid temperature recorded during the cooling test.

(d)Correction factor for cylinder barrel temperatures. Cylinder barrel temperatures must be corrected by adding to them 0.7 times the difference between the maximum ambient atmospheric temperature and the temperature of the ambient air at the time of the first occurrence of the maximum cylinder barrel temperature recorded during the cooling test.

ED Decision 2003/16/RM

(a)Climb cooling tests must be conducted under this paragraph for:

(1)Category A rotorcraft; and

(2)Multi-engine Category B rotorcraft for which certification is requested under the Category A powerplant installation requirements, and under the requirements of CS 29.861(a) at the steady rate of climb or descent established under CS 29.67(b).

(b)The climb or descent cooling tests must be conducted with the engine inoperative that produces the most adverse cooling conditions for the remaining engines and powerplant components.

(c)Each operating engine must:

(1)For helicopters for which the use of 30-minute OEI power is requested, be at 30-minute OEI power for 30 minutes, and then at maximum continuous power (or at full throttle, when above the critical altitude);

(2)For helicopters for which the use of continuous OEI power is requested, be at continuous OEI power (or at full throttle when above the critical altitude); and

(3)For other rotorcraft, be at maximum continuous power (or at full throttle when above the critical altitude).

(d)After temperatures have stabilised in flight, the climb must be:

(1)Begun from an altitude not greater than the lower of:

(i)305 m (1000 ft) below the engine critical altitude; and

(ii)305 m (1000 ft) below the maximum altitude at which the rate of climb is 0.76 m/s (150 fpm); and

(2)Continued for at least 5 minutes after the occurrence of the highest temperature recorded, or until the rotorcraft reaches the maximum altitude for which certification is requested.

(e)For Category B rotorcraft without a positive rate of climb, the descent must begin at the all-engine-critical altitude and end at the higher of:

(1)The maximum altitude at which level flight can be maintained with one engine operative; and

(2)Sea-level.

(f)The climb or descent must be conducted at an airspeed representing a normal operational practice for the configuration being tested. However, if the cooling provisions are sensitive to rotorcraft speed, the most critical airspeed must be used, but need not exceed the speeds established under CS 29.67(a)(2) or 29.67(b). The climb cooling test may be conducted in conjunction with the take-off cooling test of CS 29.1047.

ED Decision 2003/16/RM

(a)Category A. For each Category A rotorcraft, cooling must be shown during take-off and subsequent climb as follows:

(1)Each temperature must be stabilised while hovering in ground effect with:

(i)The power necessary for hovering;

(ii)The appropriate cowl flap and shutter settings; and

(iii)The maximum weight.

(2)After the temperatures have stabilised, a climb must be started at the lowest practicable altitude and must be conducted with one engine inoperative.

(3)The operating engines must be at the greatest power for which approval is sought (or at full throttle when above the critical altitude) for the same period as this power is used in determining the take-off climbout path under CS 29.59.

(4)At the end of the time interval prescribed in sub-paragraph (b)(3), the power must be changed to that used in meeting CS 29.67(a)(2) and the climb must be continued for:

(i)30 minutes, if 30-minute OEI power is used; or

(ii)At least 5 minutes after the occurrence of the highest temperature recorded, if continuous OEI power or maximum continuous power is used.

(5)The speeds must be those used in determining the take-off flight path under CS 29.59.

(b)Category B. For each Category B rotorcraft, cooling must be shown during take-off and subsequent climb as follows:

(1)Each temperature must be stabilised while hovering in ground effect with:

(i)The power necessary for hovering;

(ii)The appropriate cowl flap and shutter settings; and

(iii)The maximum weight.

(2)After the temperatures have stabilised, a climb must be started at the lowest practicable altitude with take-off power.

(3)Take-off power must be used for the same time interval as take-off power is used in determining the take-off flight path under CS 29.63.

(4)At the end of the time interval prescribed in sub-paragraph (a)(3), the power must be reduced to maximum continuous power and the climb must be continued for at least 5 minutes after the occurrence of the highest temperature recorded.

(5)The cooling test must be conducted at an airspeed corresponding to normal operating practice for the configuration being tested. However, if the cooling provisions are sensitive to rotorcraft speed, the most critical airspeed must be used, but need not exceed the speed for best rate of climb with maximum continuous power.

ED Decision 2023/001/R

The hovering cooling provisions must be shown –

(a)At maximum weight or at the greatest weight at which the rotorcraft can hover (if less), at sea-level, with the power required to hover but not more than maximum continuous power, in the ground effect in still air, until at least 5 minutes after the occurrence of the highest temperature recorded; and

(b)With maximum continuous power, maximum weight, and at the altitude resulting in zero rate of climb for this configuration, until at least 5 minutes after the occurrence of the highest temperature recorded.

For rotorcraft for which a 30-minute power rating is claimed, the hovering cooling provisions must be shown:

(a)At maximum weight or at the greatest weight at which the rotorcraft can hover (if less), at sea level, with the power required to hover but not more than 30-minute power rating, in the ground effect in still air, until:

at least 5 minutes after the occurrence of the highest temperature recorded, or

the continuous time limit of the 30-minute power rating if the highest temperature recorded is not stabilised before.

(b)With 30-minute power rating, maximum weight, and at the altitude resulting in zero rate of climb for this configuration, until:

at least 5 minutes after the occurrence of the highest temperature recorded, or

the continuous time limit of the 30-minute power rating if the highest temperature recorded is not stabilised before.

[Amdt No: 29/11]

INDUCTION SYSTEM

ED Decision 2003/16/RM

(a)The air induction system for each engine and auxiliary power unit must supply the air required by that engine and auxiliary power unit under the operating conditions for which certification is requested.

(b)Each engine and auxiliary power unit air induction system must provide air for proper fuel metering and mixture distribution with the induction system valves in any position.

(c)No air intake may open within the engine accessory section or within other areas of any powerplant compartment where emergence of backfire flame would constitute a fire hazard.

(d)Each reciprocating engine must have an alternate air source.

(e)Each alternate air intake must be located to prevent the entrance of rain, ice, or other foreign matter.

(f)For turbine engine powered rotorcraft and rotorcraft incorporating auxiliary power units:

(1)There must be means to prevent hazardous quantities of fuel leakage or overflow from drains, vents, or other components of flammable fluid systems from entering the engine or auxiliary power unit intake system; and

(2)The air inlet ducts must be located or protected so as to minimise the ingestion of foreign matter during take-off, landing, and taxying.

ED Decision 2003/16/RM

(a)Reciprocating engines. Each reciprocating engine air induction system must have means to prevent and eliminate icing. Unless this is done by other means, it must be shown that, in air free of visible moisture at a temperature of –1°C (30°F) and with the engines at 60% of maximum continuous power –

(1)Each rotorcraft with sea-level engines using conventional venturi carburettors has a preheater that can provide a heat rise of 50°C (90°F);

(2)Each rotorcraft with sea-level engines using carburettors tending to prevent icing has a preheater that can provide a heat rise of 39°C (70°F);

(3)Each rotorcraft with altitude engines using conventional venturi carburettors has a preheater that can provide a heat rise of 67°C (120°F); and

(4)Each rotorcraft with altitude engines using carburettors tending to prevent icing has a preheater that can provide a heat rise of 56°C (100°F).

(b)Turbine engines:

(1)It must be shown that each turbine engine and its air inlet system can operate throughout the flight power range of the engine (including idling):

(i)Without accumulating ice on engine or inlet system components that would adversely affect engine operation or cause a serious loss of power under the icing conditions specified in Appendix C; and

(ii)In snow, both falling and blowing, without adverse effect on engine operation, within the limitations established for the rotorcraft.

(2)Each turbine engine must idle for 30 minutes on the ground, with the air bleed available for engine icing protection at its critical condition, without adverse effect, in an atmosphere that is at a temperature between -9°C and –1°C (between 15°F and 30°F) and has a liquid water content not less than 0.3 grams per cubic meter in the form of drops having a mean effective diameter not less than 20 microns, followed by momentary operation at take-off power or thrust. During the 30 minutes of idle operation, the engine may be run up periodically to a moderate power or thrust setting in a manner acceptable to the Agency.

(c)Supercharged reciprocating engines. For each engine having a supercharger to pressurise the air before it enters the carburettor, the heat rise in the air caused by that supercharging at any altitude may be utilised in determining compliance with subparagraph (a) if the heat rise utilised is that which will be available, automatically, for the applicable altitude and operation condition because of supercharging.

ED Decision 2023/001/R

This AMC is primarily applicable to rotorcraft equipped with air intake external screens (or any other air intake prone to the same kind of icing which may exist downstream), and has been developed based on in-service experience.

In icing conditions, as defined in CS-29 Appendix C, when the outside air temperature (OAT) is quite cold, typically below -5°C, the water droplets freeze at the helicopter air intake external screen that, once clogged, acts as passive protection by preventing subsequent super-cooled droplets to enter the engine duct and plenum. The air, then, enters the engine intake through screen areas where water droplets do not accrete, or through an air intake by-pass, if necessary.

For warmer temperatures, typically between -5°C and 0°C, a critical temperature can exist at which the water droplets do not freeze completely and immediately on the external screen and therefore icing conditions may exist downstream in the engine air intake ducts or engine internal screen.

Furthermore, ice accretions behind the air intake screen can then be released during an engine acceleration or a rotorcraft descent in a warmer atmosphere and thus may lead to engine damage, surge or in-flight shutdown.

In the case where the engine is also protected by its own screen, then the engine screen can then become clogged by ice. This may also lead to high pressure drop or distortion across the engine screen, resulting into engine surge, engine damage or engine shutdown.

The purpose of this AMC is to provide specific and complementary guidance for showing compliance with CS 29.1093(b)(1)(i) in the determination of this critical temperature, but does not provide any other guidance to demonstrate full compliance with CS 29.1093(b)(1)(i) to cope with icing conditions as detailed in Appendix C to CS-29.

Analysis only should not be considered in the determination of the critical temperature due to the level of accuracy required for such an assessment. Its determination should be validated during combined rotorcraft (air intake / engine) icing tests in a wind tunnel or a similar test facility where the temperature can be controlled accurately showing whether icing conditions downstream the air intake screen are an issue or not. Typically, an accuracy of 0.5°C could be envisaged.

If the above-mentioned testing is done without the engine, it should be first demonstrated that the engine flow is correctly simulated, and the engine thermal impact adequately considered and validated on air intake. In a second step, the repercussion of any ice accretion should be assessed at engine level both in terms of airflow distortion and engine ingestion and duly validated by appropriate means. It has to be noted that this alternative approach without the engine may lead to difficulties in interpreting the results at engine level.

During these tests, the engine should be run at critical power in the icing conditions defined in CS-29 Appendix C depending on the claimed certification (inadvertent icing encounter or full icing certification). The critical power could be determined following a critical point analysis (other methodologies might be acceptable) to assess the engine operability with regard to the feared events such as airflow distortion or engine ice ingestion.

To determine the temperature at which the water does not freeze on the external screen, the test temperature may be decreased by accurate steps (typically a value of 0.5°C is suggested) from 0°C until accretion downstream the external air intake screen, if any, is maximised. If no ice is observed after 15 minutes of water injection, the test point is believed to be performed at a too warm temperature and can be stopped.

When decreasing the temperature step by step, if no ice accretion is observed downstream the helicopter external screen — typically for temperatures below -5°C the external screen catches the majority of the super-cooled droplets — it means that the above-described phenomenon does not occur.

Some other method can be proposed to reduce the test point number.

The test should demonstrate that, at the determined critical temperature, the maximum potential ice accretions downstream the rotorcraft screen do not have an adverse effect on the engine both in the full range of claimed operation and when the rotorcraft then descends in an atmosphere with a positive OAT.

As an example, the following test procedure may be considered:

—A 1st run: at the end of the test (in fact, when reaching the highest measured pressure drop in the air intake), perform three consecutive engine quick decelerations (from maximum power to idle) / accelerations (from idle to maximum power).

—A 2nd run: at the end of the test (in fact, when reaching the highest measured pressure drop in the air intake), simulate a quick descent in atmosphere with a positive OAT considering a tunnel warm-up procedure.

Quick accelerations / decelerations are to be understood as the maximum acceleration / deceleration rates that can be performed by a pilot during flight operation. The intent is to simulate a real-life engine behaviour which affects the flow/ice ingestion accordingly. For example, values close to one second from minimum to maximum power have been considered in the past for such testing.

As specified in CS 29.1093(b)(1)(i), these tests shall demonstrate that the engine operation is not adversely affected by icing conditions.

Whenever an applicant is willing to use previous icing wind tunnel tests, an analysis might be an acceptable means of compliance provided that this analysis is adequately validated and covers as a minimum the changes in configurations (air intakes, engines, engine installations, etc.), engine operability (airflow, ingestion capabilities, surge margins, etc.) and thermal environment of the air intake.

For rotorcraft certified in full icing conditions, in order to determine the rotorcraft performance in icing conditions, this test point should be used to identify the engine installation losses for flight into known icing conditions, in particular if the engine is also equipped with its own screen.

[Amdt No: 29/11]

ED Decision 2003/16/RM

Each carburettor air preheater must be designed and constructed to:

(a)Ensure ventilation of the preheater when the engine is operated in cold air;

(b)Allow inspection of the exhaust manifold parts that it surrounds; and

(c)Allow inspection of critical parts of the preheater itself.

ED Decision 2003/16/RM

(a)Each induction system duct upstream of the first stage of the engine supercharger and of the auxiliary power unit compressor must have a drain to prevent the hazardous accumulation of fuel and moisture in the ground attitude. No drain may discharge where it might cause a fire hazard.

(b)Each duct must be strong enough to prevent induction system failure from normal backfire conditions.

(c)Each duct connected to components between which relative motion could exist must have means for flexibility.

(d)Each duct within any fire zone for which a fire-extinguishing system is required must be at least:

(1)Fireproof, if it passes through any firewall; or

(2)Fire resistant, for other ducts, except that ducts for auxiliary power units must be fireproof within the auxiliary power unit fire zone.

(e)Each auxiliary power unit induction system duct must be fireproof for a sufficient distance upstream of the auxiliary power unit compartment to prevent hot gas reverse flow from burning through auxiliary power unit ducts and entering any other compartment or area of the rotorcraft in which a hazard would be created resulting from the entry of hot gases. The materials used to form the remainder of the induction system duct and plenum chamber of the auxiliary power unit must be capable of resisting the maximum heat conditions likely to occur.

(f)Each auxiliary power unit induction system duct must be constructed of materials that will not absorb or trap hazardous quantities of flammable fluids that could be ignited in the event of a surge or reverse flow condition.

ED Decision 2003/16/RM

If induction system screens are used:

(a)Each screen must be upstream of the carburettor;

(b)No screen may be in any part of the induction system that is the only passage through which air can reach the engine, unless it can be deiced by heated air;

(c)No screen may be deiced by alcohol alone; and

(d)It must be impossible for fuel to strike any screen.

ED Decision 2003/16/RM

Each inter-cooler and after-cooler must be able to withstand the vibration, inertia, and air pressure loads to which it would be subjected in operation.

ED Decision 2003/16/RM

It must be shown under CS 29.1043 that each installation using two-stage superchargers has means to maintain the air temperature, at the carburettor inlet, at or below the maximum established value.

EXHAUST SYSTEM

ED Decision 2003/16/RM

For powerplant and auxiliary power unit installations the following apply:

(a)Each exhaust system must ensure safe disposal of exhaust gases without fire hazard or carbon monoxide contamination in any personnel compartment.

(b)Each exhaust system part with a surface hot enough to ignite flammable fluids or vapours must be located or shielded so that leakage from any system carrying flammable fluids or vapours will not result in a fire caused by impingement of the fluids or vapours on any part of the exhaust system including shields for the exhaust system.

(c)Each component upon which hot exhaust gases could impinge, or that could be subjected to high temperatures from exhaust system parts, must be fireproof. Each exhaust system component must be separated by a fireproof shield from adjacent parts of the rotorcraft that are outside the engine and auxiliary power unit compartments.

(d)No exhaust gases may discharge so as to cause a fire hazard with respect to any flammable fluid vent or drain.

(e)No exhaust gases may discharge where they will cause a glare seriously affecting pilot vision at night.

(f)Each exhaust system component must be ventilated to prevent points of excessively high temperature.

(g)Each exhaust shroud must be ventilated or insulated to avoid, during normal operation, a temperature high enough to ignite any flammable fluids or vapours outside the shroud.

(h)If significant traps exist, each turbine engine exhaust system must have drains discharging clear of the rotorcraft, in any normal ground and flight attitudes, to prevent fuel accumulation after the failure of an attempted engine start.

ED Decision 2003/16/RM

(a)Exhaust piping must be heat and corrosion resistant, and must have provisions to prevent failure due to expansion by operating temperatures.

(b)Exhaust piping must be supported to withstand any vibration and inertia loads to which it would be subjected in operation.

(c)Exhaust piping connected to components between which relative motion could exist must have provisions for flexibility.

ED Decision 2003/16/RM

For reciprocating engine powered rotorcraft the following apply:

(a)Each exhaust heat exchanger must be constructed and installed to withstand the vibration, inertia, and other loads to which it would be subjected in operation. In addition:

(1)Each exchanger must be suitable for continued operation at high temperatures and resistant to corrosion from exhaust gases;

(2)There must be means for inspecting the critical parts of each exchanger;

(3)Each exchanger must have cooling provisions wherever it is subject to contact with exhaust gases; and

(4)No exhaust heat exchanger or muff may have stagnant areas or liquid traps that would increase the probability of ignition of flammable fluids or vapours that might be present in case of the failure or malfunction of components carrying flammable fluids.

(b)If an exhaust heat exchanger is used for heating ventilating air used by personnel –

(1)There must be a secondary heat exchanger between the primary exhaust gas heat exchanger and the ventilating air system; or

(2)Other means must be used to prevent harmful contamination of the ventilating air.

POWERPLANT CONTROLS AND ACCESSORIES

ED Decision 2003/16/RM

(a)Powerplant controls must be located and arranged under CS 29.777 and marked under CS 29.1555.

(b)Each control must be located so that it cannot be inadvertently operated by persons entering, leaving or moving normally in the cockpit.

(c)Each flexible powerplant control must be approved.

(d)Each control must be able to maintain any set position without:

(1)Constant attention; or

(2)Tendency to creep due to control loads or vibration.

(e)Each control must be able to withstand operating loads without excessive deflection.

(f)Controls of powerplant valves required for safety must have:

(1)For manual valves, positive stops or in the case of fuel valves suitable index provisions, in the open and closed position; and

(2)For power-assisted valves, a means to indicate to the flight crew when the valve:

(i)Is in the fully open or fully closed position; or

(ii)Is moving between the fully open and fully closed position.

ED Decision 2003/16/RM

Means must be provided on the flight deck for starting, stopping, and emergency shutdown of each installed auxiliary power unit.

ED Decision 2003/16/RM

(a)There must be a separate power control for each engine.

(b)Power controls must be arranged to allow ready synchronisation of all engines by:

(1)Separate control of each engine; and

(2)Simultaneous control of all engines.

(c)Each power control must provide a positive and immediately responsive means of controlling its engine.

(d)Each fluid injection control other than fuel system control must be in the corresponding power control. However, the injection system pump may have a separate control.

(e)If a power control incorporates a fuel shutoff feature, the control must have a means to prevent the inadvertent movement of the control into the shutoff position. The means must –

(1)Have a positive lock or stop at the idle position; and

(2)Require a separate and distinct operation to place the control in the shutoff position.

(f)For rotorcraft to be certificated for a 30-second OEI power rating, a means must be provided to automatically activate and control the 30-second OEI power and prevent any engine from exceeding the installed engine limits associated with the 30-second OEI power rating approved for the rotorcraft.

ED Decision 2023/001/R

(a)For each engine, means must be provided in the cockpit so as to:

(1)control, either directly by the crew or by the crew via a system (such as the FADEC), each ignition circuit;

(2)readily allow the crew to conduct the flight and manage both ground start and in-flight restart;

(3)check the health condition of each ignition circuit; and

(4)maintain an isolation between each engine control.

(b)There must be means to quickly shut off all ignition by the grouping of switches or by a master ignition control.

(c)Each group of ignition switches, except ignition switches for turbine engines for which continuous ignition is not required, and each master ignition control, must have a means to prevent its inadvertent operation.

[Amdt No: 29/11]

ED Decision 2023/001/R

(a)Compliance with CS 29.1145(a) is considered to be demonstrated by providing for each engine one of the following design solutions:

(1)Independent ignition controls should be provided for each ignition circuit, or

(2)A single ignition control acting on two ignition switches should be provided to control each ignition circuit via a dual-channel FADEC.

(i)Each switch should be connected to one channel of the FADEC.

(ii)The FADEC should ensure the following functions:

(A)Ability to control automatically and independently each ignition circuit of the engine

(B)Ability to perform a health monitoring of each ignition circuit for the aircraft to meet the safety objectives of CS-29

(b)The check of the health condition of each ignition circuit could be achieved in automatic or initiated test or by procedure without any difference.

[Amdt No: 29/11]

ED Decision 2003/16/RM

(a)If there are mixture controls, each engine must have a separate control, and the controls must be arranged to allow:

(1)Separate control of each engine; and

(2)Simultaneous control of all engines.

(b)Each intermediate position of the mixture controls that corresponds to a normal operating setting must be identifiable by feel and sight.

ED Decision 2003/16/RM

(a)It must be impossible to apply the rotor brake inadvertently in flight.

(b)There must be means to warn the crew if the rotor brake has not been completely released before take-off.

ED Decision 2003/16/RM

There must be a separate carburettor air temperature control for each engine.

ED Decision 2003/16/RM

Each supercharger control must be accessible to:

(a)The pilots; or

(b)(If there is a separate flight engineer station with a control panel) the flight engineer.

ED Decision 2003/16/RM

(a)Each engine-mounted accessory must:

(1)Be approved for mounting on the engine involved;

(2)Use the provisions on the engine for mounting; and

(3)Be sealed in such a way as to prevent contamination of the engine oil system and accessory system.

(b)Electrical equipment subject to arcing or sparking must be installed, to minimise the probability of igniting flammable fluids or vapours.

(c)If continued rotation of an engine-driven cabin supercharger or any remote accessory driven by the engine will be a hazard if they malfunction, there must be means to prevent their hazardous rotation without interfering with the continued operation of the engine.

(d)Unless other means are provided, torque limiting means must be provided for accessory drives located on any component of the transmission and rotor drive system to prevent damage to these components from excessive accessory load.

ED Decision 2003/16/RM

(a)Each battery ignition system must be supplemented with a generator that is automatically available as an alternate source of electrical energy to allow continued engine operation if any battery becomes depleted.

(b)The capacity of batteries and generators must be large enough to meet the simultaneous demands of the engine ignition system and the greatest demands of any electrical system components that draw from the same source.

(c)The design of the engine ignition system must account for:

(1)The condition of an inoperative generator;

(2)The condition of a completely depleted battery with the generator running at its normal operating speed; and

(3)The condition of a completely depleted battery with the generator operating at idling speed, if there is only one battery.

(d)Magneto ground wiring (for separate ignition circuits) that lies on the engine side of any firewall must be installed, located, or protected, to minimise the probability of the simultaneous failure of two or more wires as a result of mechanical damage, electrical fault or other cause.

(e)No ground wire for any engine may be routed through a fire zone of another engine unless each part of that wire within that zone is fireproof.

(f)Each ignition system must be independent of any electrical circuit that is not used for assisting, controlling, or analysing the operation of that system.

(g)There must be means to warn appropriate crew members if the malfunctioning of any part of the electrical system is causing the continuous discharge of any battery necessary for engine ignition.

POWERPLANT FIRE PROTECTION

ED Decision 2003/16/RM

(a)Designated fire zones are:

(1)The engine power section of reciprocating engines;

(2)The engine accessory section of reciprocating engines;

(3)Any complete powerplant compartment in which there is no isolation between the engine power section and the engine accessory section, for reciprocating engines;

(4)Any auxiliary power unit compartment;

(5)Any fuel-burning heater and other combustion equipment installation described in CS 29.859;

(6)The compressor and accessory sections of turbine engines; and

(7)The combustor, turbine, and tailpipe sections of turbine engine installations except sections that do not contain lines and components carrying flammable fluids or gases and are isolated from the designated fire zone prescribed in sub-paragraph (a)(6) by a firewall that meets CS 29.1191.

(b)Each designated fire zone must meet the requirements of CS 29.1183 to 29.1203.