AC 29-2C, Chg 4 7/6/2012 AC 29.927A. § 29.927 (Amendment 29-26) ADDITIONAL TESTS.
a. Section 29.927(c):
(1) Explanation.
(i) Amendment 29-26 revised the rotor drive system loss of lubrication test requirements for Category A rotorcraft in § 29.927(c). This requires testing to show that any failures that result in a loss of lubrication in any normal use lubrication system, unless the failures are extremely remote, will not prevent continued safe flight for at least 30 minutes after the flight crew recognizes the loss of lubricant failure.
(ii) The introductory phrase to this amendment to the regulation, “unless such failures are extremely remote” has caused confusion. The NPRM did not contain this expression and the only change documented in the preamble to the final rule (53 FR 34204) explains that the final rule was revised in response to a public comment that the proposed regulation could be interpreted to “preclude credit for auxiliary lubrication systems or to require consideration of lubricant failures to self-lubricated bearings.”
This was not intended and the final rule was “revised to eliminate this possible ambiguity.” The phrase, “unless such failures are extremely remote,” was introduced to resolve the public comment to convey that the applicant does not have to consider failures that may exist in the auxiliary lubrication system prior to performing the loss of lubrication testing. Under the current regul ation, the extremely remote language in the final rule means that testing to demonstrate at least 30 minutes continued flight capability (for Category A), following loss of lubrication in the normal lubrication system, is not required if the failures leading to that loss of lubrication condition are determined to be extremely remote. While this comp liance approach is allowed, it may not be achievable due, in part, to the unforeseen variables and complexity associated with predicting potential lubrication failure modes and their associated criticality and frequency of occurrence. This includes consi dering lubrication failures that may result from improper transmission maintenance and servicing. The expected compliance approach has been to assume a failure in the normal lubrication system leading to rapid loss of lubrication and to rely on an auxiliary lubrication system or the robustness of the transmission components to accomplish at least 30 minutes of operation (for Category A) at the prescribed conditions. With this approach, the normal and auxiliary systems must be independent in order to preclude common loss of lubrication failure points and possible cross contamination. Compli ance with § 29.1309 would only apply to any electrical and software design aspects of the normal and auxiliary lubrication systems.
The auxiliary lubrication system must also be designed, constructed, and functionally tested to show that it can perform its intended function.
(iii) The regulation is intended to apply to pressurized lubrication systems and has not been applied to splash lubricated gearboxes since historically their design has not been as critical or complex when compared to pressurized systems. The likelihood of loss of lubrication is significantly greater for transmissions that use pressure lubrication and external cooling. Th is is due to the increased complexity of the Page E –56
lubrication system and the external components that circulate oil outside the gearbox. A pressure lubrication system is more commonly used in the rotorcraft’s main transmission but may also be used in auxiliary transmissions or gearboxes.
(iv) The lubricating system has two primary functions. The first is to provide lubricating oil to contacting or rubbing surfaces and thus reduce friction losses. The second is to dissipate heat energy generated by friction of meshing gears and bearings, thus maintaining surface and material temperature. Accordingly, a loss of lubrication leads to increased friction between components and increased component surface temperatures. With increased component surface temperatures, component surface hardness can be lost resulting in the inability of the component to carry or transmit loads. Thermal expansion in transmission components can 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 internal and external failures. Failures include, but are not limited to, oil lines, fittings, seal plugs, sealing gaskets, valves, pumps, oil filters, oil coolers, accessory pads, etc. A leak caused by a crack in the transmission outer case need not be considered as a source of a loss of lubrication provided the outer case has been structurally substantiated to satisfy the requirements of §§ 29.307, 29.923(m), and 29.571.
(v) The intent of the rule change for Category A rotorcraft was to assure that these rotorcraft have significant continued flight capability after the loss of lubricant to any single transmission in order to optimize eventual landing opportunities. Extending the bench testing beyond 30 minutes, although not required, is considered highly desirable. Accomplishing this would further improve the capability of the rotorcraft to reach a suitable landing location in order to improve occupant safety when operating in remote geographic areas that include harsh environmental conditions. Indefinite flight with a lubrication system failure is not expected. However, it may be acceptable to include a time interval in the emergency procedures. That time interval should be reduced sufficiently when compared to the bench test demonstration to allow for an adequate safety margin.
(2) Procedures.
(i) Section 29.927(c)(1) prescribes a test to demonstrate that the effects of a loss of lubrication will not prevent continued safe powered operation for category A rotorcraft for at least 30 minutes after illumination of the low oil pressure warning device (required by § 29.1305). For category B roto rcraft, § 29.927(c)(2) prescribes the tests for safe operation under autorotative conditions must continue for at least 15 minutes.
(ii) An acceptable means of demonstrating compliance with this rule is through the use of a bench test (transmission test rig). Since this is essentially a durability test of the transmission to operate with residual oil, typically the worst case failure (i.e., the undrainable oil or the oil remaining after a severe pressure leak, whichever results in a greater loss of oil in the transmission’s normal use lubrication system) is used as a critical entry point for the test; see paragraph a.(2)(iii).
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AC 29-2C, Chg 4 7/6/2012
(iii) The transmission should be stabilized at the torque associated with maximum continuous power (reacted as appropriate at the main mast and tail rotor output quills) at a normal main rotor mast speed, oil temperature that is at the highest limit for continuous operation, and oil pressure that is within the normal operating range.
A vertical load should be applied at the mast, equal to the gross weight of the rotorcraft at 1g. Once the transmission oil temperatur e is stabilized, simulate the worst case failure in the normal use lubrication system. Upon illumination of the low oil pressure warning device (required by § 29.1305), reduce the input torque for category A rotorcraft to the minimum torque necessary to sustain flight and continue the test for at least 30 minutes at the maximum gross weight and the most efficient flight conditions. To complete the test, apply an input torque to the transmission for approximately 25 seconds to simulate an autorotation. The last 10 seconds (of the 25 seconds) should be at the torque required for a minimum power landing. A successful demonstration may involve limited damage to the transmission, provided it is determined that the autorotative capabilities of the rotorcraft were not significantly impaired. For category B rotorcraft, upon illumi nation of the low oil pressure warning device, reduce the input torque to simulate an autorotation and continue transmission operation for 15 minutes. To complete the test, apply an input torque to the transmission for approximately 10 seconds to simulate a minimum power landing. A successful demonstration may involve limited damage to the transmission provided it is determined that the autorotative capabilities of the rotorcraft were not significantly impaired. If compliance with category A requirements is demonstrated, category B requirements will have been met.
b. Section 29.927(d):
(1) Explanation. The revision to paragraph (d) includes a requirement to demonstrate that overspeed conditions, which may result from an engine control device failure or other event such as a control misapplication, will not result in damage to the rotor drive system.
(2) Procedures. The overspeed endurance cycle and overspeed conditions to be demonstrated are defined in this section. The test may be conducted on a rotorcraft configured for the endurance tests prescribed by § 29.923. Turbine engines involved in the test may require fuel control re-rigging or operation on a manual fuel control system to meet test requirements. Fifty overs peed runs of 30 ±3 seconds must be run on the rotor drive system. The overspeed runs must be alternated with stabilizing runs of 1 to 5 minutes duration each at 60 to 80 percent of maximum continuous speed.
(i) The maximum speed to be demonstrated during the power on overspeed test is:
(A) The higher of:
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(1) The speed to be expected from an engine control device failure;
or,
(2) 105 percent of the maximum rotational speed to be expected in service, including transients.
(B) The maximum speed allowed by a speed limiting device if the device is installed independent of the engine controls and is shown to be reliable.
(ii) From the stabilizing run condition, increase power to achieve the maximum speed established from (i) above. Se t the collective for minimum blade pitch for smooth operation. The cyclic control s hould be positioned for vertical lift, and the anti-torque control should be set in flat pitch. Hold this condition for 30 seconds, then decelerate to the stabilizing run condition.
(iii) The acceleration and deceleration described above should be accomplished in 10 seconds or less except where it can be shown that the certified engine acceleration or deceleration rate exceeds 10 seconds. The time required for acceleration and deceleration may not be deducted from the 30 second overspeed period.
Note: Some equivalent safety findings have been issued based upon limiting the test speed to that permitted by an independent, reliable overspeed trip device. This has been done to avoid permanent damage to rotors, yokes, engines, etc., which are involved, but not under evaluation by this test.
c. Section 29.927(f):
(1) Explanation. Amendment 29-26 also added a new paragraph (f), which requires that the overtorque, lubrication system failure, and overspeed tests required by § 29.927(b), (c), and (d), respectively, be conducted without intervening disassembly during the individual test. After each test, a teardown inspection is performed, and except for the components used in the lubrication system failure test, the components are required to be in serviceable (return to service) condition.
(2) Procedures. None.
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AC 29-2C, Chg 4 5/1/2014 AC 29.931. § 29.931 (Amendment 29- 12) SHAFTING CRITICAL SPEEDS.
a. Explanation.
(1) At certain speeds, rotating shafts tend to vibrate violently in a transverse direction. These speeds are variously known as “critical speeds,” “whirling speeds,” or [Section AC 27.931 continued on next page.] Page E –58
“whipping speeds.” The vibration results from the unbalance of the rotating system and can be shown to reach destructive values with only minimal unbalance. The nature of this phenomena is that as shaft rotational speed increases, residual unbalance in the shaft gives rise to centrifugal forces. These forces cause the shaft to rotate in a bent or bowed configuration with the centrifugal force induced bending loads being balanced by coriolis and elastic forces in the shaft. As shaft rotational speed increases, the centrifugal forces increase to the point at which they exceed the elastic forces in the shaft, and divergence occurs. This point in the speed range is called the critical speed.
At shaft speeds above the critical speed, a 180° phase change occurs; the shaft’s mass center moves toward the center of rotation and the amplitude of vibration diminishes with further increases in shaft speed.
(2) The most prominent design option is to operate the shafting subcritical; i.e., below the first critical speed, with adequate margins from critical speed at the maximum allowable speed, including transients. However, another option, that of supercritical shaft operation; i.e., operating above the first or even higher critical speeds with adequate margins between any critical speed for the normal operating speed range.
This latter portion requires some form of fixed system damping to permit safe transition through the critical speed range and to avoid excessive nonsyncronous vibrations or instability in the critical speed mode at suboperating frequency.
(3) A review of typical design practices and drive system arrangements discloses several types of shaft support and loading:
(i) Main rotor/mast/transmission assemblies rigidly mounted to the airframe;
(ii) Main rotor/mast/transmission assemblies compliantly mounted to the airframe;
(iii) Main rotor supported through a bearing arrangement by a rigid nonrotating structure with a coaxial torque shaft driving the rotor;
(iv) Cross-shafting, interconnect shafting, tail rotor drive shafting which are generally supported by gearboxes at each end and by hanger bearings at semispan;
(v) Engine to transmission shafting which, for compliant pylons, incorporate flexible or geared coupling, to accommodate the misalignment and chucking; and
(vi) Tail rotor/mast/gearbox supported on the tailboom or near the upper extremity of a vertical fin.
(4) With regard to compliant pylon mountings, recent developments in vibration control have led to rotor isolation wherein the fuselage is isolated from the rotor and Page E - 59
transmission, resulting in improved vibration and system reliability. Rotor isolation systems typically entail the installation of isolation devices at the transmission-airframe interface. The crux of rotor isolation is providing adequate, low -frequency isolation without excessive relative displacement or loss of mechanical stability. Rotor isolation affects shaft critical speeds in the following ways:
(i) First, the transmission mounting configuration, system stiffness, and tuning requirements may result in different fore-and-aft and lateral natural frequencies, imposing additional analytical requirements. For compliant mounting, the response while transitioning through the fundamental or rocking modes is generally controlled by dampers or elastomeric elements.
(ii) Second, the relatively high displacements permitted by the isolation system, depending on configuration, may result in variations in shaft misalignment and length thus adding further complexity to the analytical prediction of critical speeds.
b. Procedures.
(1) Subcritical Shafting Designs. Three basic methods of qualification may be considered, with the required margins relative to the degree of assurance provided:
(i) Analytical.
(A) Simplistic model(s) as shown in figures AC 29.931-1 and AC 29.931-2; 35 percent margin shown above maximum operating speed.
(B) Detailed model, taking into account significant variations in shaft stiffness, mass distribution, cone adapters, support bearing stiffness, support structure;
20 percent margin shown above maximum operating speed.
(ii) Analytical supported by tests. Analysis supported by shake test (rotating or nonrotating) or by bench test, where appropriate adjustments are made for differences between the bench and the aircraft; 15 percent margin shown above maximum operating speed.
(iii) Whirl test on the aircraft.
(A) For all cases, it should be shown that, under maximum permissible unbalance and at the maximum operating speed, the shafting and support structure has acceptable clearance and does not have excessive vibration.
(B) For compliant pylon mountings, damping of the rigid body rocking modes, which are often transitioned during run-up to normal speed (and which are not critical flexing modes), may be verified by analysis, laboratory tests, or ground run-up with the rotor at maximum permissible unbalance. Damping on the order of 5 percent equivalent viscous damping is generally acceptable.
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(C) For tail rotor masts, the analysis should include fixed system structural response including tailboom, fixed control surfaces, and vertical fin. The frequency analysis will then contain both fixed system and rotating system modes. An energy analysis can then be used to identify whether the modes are predominantly fixed system or rotating system modes. Systems with up to 35 percent energy in the rotating system have been operated in the field without significant problems. For this type of shafting installations, it is advisable to avoid fixed system modes at multiples of shaft speed, particularly where highly nonisotropic mountings exist.
(2) Supercritical Shafting Design. Another facet occasionally encountered with shafting is the concept of normally operating at speeds above the critical speed, commonly referred to as “supercritical operation.” To function properly, suitable dampers must be installed to enable the shaft to pass safely through the lower critical speed up to the operating speed, and speed controls should be devised so as to avoid any tendency to operate continuously at any critical speed. Accurate balancing of the rotating components will also decrease the energy to be dissipated into the damping device during transition thereby increasing its serviceability and reliability. It should be noted that damper design and locations become more complex as selected operating speed increases through the third or fourth critical frequency. Multiple node points will exist where dampers will not be effective. Production specimen testing at high speed/high torque conditions should include checks for shaft straightness until experience verifies that shaft deflecting is not significant. For system utilizing squeeze film dampers at the support bearings, variations in oil pressure, flow restrictions, and the effects of bearing preload should be evaluated. The effects of shaft and unbalance and the proximity of the damper to bottoming under maximum unbalance should be evaluated.
(3) If the shafting configuration of the rotorcraft includes universal joints or misalignment couplings, a velocity differential will exist across the joint which creates sinusoidal torques and bending moments at both shafts at multiples of the rotation speed. To avoid amplification of these torques and bending moments, the design should preclude coincidence of critical speeds and multiples of normal speeds.
(4) Note that failure considerations required under § 29.901(d) may result in abnormal rotational speed and torque excursions. Resulting encounters with critical speeds should not create hazards.
(5) Order 8110.9, Handbook on Vibration Substantiation and Fatigue Evaluation of Helicopters and Other Power Transmission Systems, also addresses this subject. This document is distributed to section level and above in all Regional Aircraft Certification Offices.
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Wa M + 0.23m Wa = first critical speed, RAD/SEC K = shatt spring rate, LB/IN = 3EI/L3 = modulus of elasticity = moment on inertia M = mass of weight, LB-SEC?/IN m = mass of shaft, LB-SECHIN FIGURE AC 29.931-1. CANTILEVERED SHAFT, FIRST CRITICAL SPEED Wo = =^ EI u L4 Wer = first critical speed RAD/SEC u = mass per unit length E = Young's modulus L = length between supports | = inertia of shaft = a numerical constant. for first critical speed, a = (pi) = 9.87 The numerical constant (a) for higher critical whirl modes or other shaft support systems may be derived from standard texts on this subject.
FIGURE AC 29.931-2. SHAFT BETWEEN SUPPORT BEARINGS, FIRST CRITICAL SPEED Page E - 62
AC 29.935. § 29.935 SHAFTING JOINTS.
a. Explanation. This rule requires the design of shafting joints to include provisions for lubrication when such lubrication is necessary for operation.
b. Procedures. Review the design of t he rotor drive system for universal joints, slip joints (splines), and other shaft couplings. Lubrication access points (Zerk fittings) should be required unless the design incorporates alternate provisions for lubrication acceptable to the FAA/AUTHORITY.
AC 29.939. § 29.939 (Amendment 29-12) TURBINE ENGINE OPERATING CHARACTERISTICS.
a. Explanation. This section provides guidance for evaluation of engine operation, engine inlet airflow distortion, and engine to drive system torsional stability. A satisfactory rotorcraft design for all three items should be established by the manufacturer early in the development program since changes in design to satisfy these requirements are typically very expensive and will adversely impact other basic design features. Introduction of full authority digital engine control (FADEC) controls has increased the complexity in evaluating engine operating characteristics with the need to investigate FADEC degraded or failure modes. The certification test plan should address the engine control being used. In addition, where manufacturing tolerances could affect engine handling and rotor governing, tests should be performed considering the worst tolerances regarding the tests to be conducted. The results of these evaluations are used in part to verify that the requirements for the engine installation Instruction Manual mandated by § 33.5(a) are satisfied.
b. Procedures.
(1) Turbine engine operation.
(i) Explanation. Smoot h, stable operation of turbine engines is essential to safety and control of rotorcraft. This can be adversely affected by rotorcraft maneuvers, turbulence, high altitude, temperature, airspeed, and installation features such as the engine air inlet duct, exhaust duct, and the location with respect to other airframe items which induce or influence air flow through the engine. Powerplant control displacement rate can also be a factor, although most modern engines incorporate internal protection for this aspect. The engine’s tolerance to th ese factors is reflected as the “stall margin” which is established by the engine manufacturer through design and test. However, this stall margin is applicable only to an engine with a specified inlet and exhaust and at specified altitude, temperature, and effective airspeed. Typically, the specified engine inlet duct is a symmetrical bellmouth and the exhaust is a short straight duct of specified diameter and length. The stall margin, even under the above test conditions, usually varies with engine power, acceleration or deceleration, compressor air bleed, and accessory power extraction.
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AC 29-2C, Chg 4 5/1/2014
(ii) Procedures. The official flight test plan should include requirements to investigate the engine operating characteristics for stall, surge, flameout, acceleration and deceleration response, and transient response (within approved limits) throughout the operating range of the rotorcraft. T he results must show that no adverse characteristics are present, to a hazardous degree, during normal and emergency operation within the range of operating limitations of the rotorcraft and engine. Test configurations should encompass the critical engine power settings and design or rigging tolerances expected to be seen in service. In addition, normal and degraded engine operating control modes should be evaluated where applicable. Test conditions should include maximum airspeed-sideslip combinations, power recoveries, hover with wind from all directions including tailwinds and other maneuvers appropriate to the certificated operating envelope of the rotorcraft. In addition, reci rculation of exhaust gases during hover or rearward flight can be critical for engine operation and should be evaluated. Particular attention should be given to flight and operat ing conditions that can be judged critical from review of data on engine inlet pressure and temperature distribution patterns and engine stall margin data if available. High altitude has typically been critical for these tests unless other critical areas of the flight envelope are identified. In addition, during rearward flight at high altitude, results may indicate unacceptable thermal distortions in the inlet due to reingestion. Stall, surge, or flameout which may be hazardous (i.e., causes loss of engine function, loss of control, severe torsional shock through the rotor drive system, or otherwise damages the rotorcraft) is unacceptable. The flight te st program should include:
(A) Normal operation (hover, forward stabilized flight, collective inputs, rearward flight).
(1) Checks in hover. Hover with wind from all azimuths to the allowable wind limits, including tailwinds, should be evaluated. This evaluation should include maneuvers with rapid power changes. Recirculation of exhaust gases can be critical. Particular attention should be given to flight and operating conditions that can be judged critical from review of data on engine inlet pressure and temperature distribution patterns and engine stall margin data. Operat ing at high density altitudes, especially engine operation during low-speed rearward flight, is more likely to result in unacceptable engine inlet thermal distortion due to reingestion of exhaust gas.
Behavior of the compressor control bleed air valves, if any, must be carefully checked by collective oscillations around the shut off or open bleed air valves operating points.
To mitigate the possible risk of power loss, a safe test build up should be considered, such as a build up in actual wind conditions before progressing to rearward or sideward flight.
(2) Checks in forward stabilized flight.
(i) Behavior of the engine must be checked during level flight, climb, and descent at various power settings in order to verify the:
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oscillations).
(A) Governing stability (e.g., absence of engine parameters (B) Variation of rotor speed with the requested power.
(C) Engine matching on multi-engine helicopters.
(D) Accuracy of the engine parameters available to the pilot.
(E) Opening and closing thresholds of the bleed valve, if any.
( F ) E f f e c t o f s i d e s l i p .
(ii) Stabilized flight conditions should be long enough to allow the engine to reach its thermal stabilization (typically 2 or 3 minutes) with the aim of:
(A) Assessing that the engine cycle (TOT and turbine inlet temperature) as installed is consistent with the engine certificated type design.
(B) Evaluating installation effects.
(3) Checks of engine behavior during collective inputs. Before conducting the below collective increase inputs and collective decrease inputs tests, the tested engines should be precisely defined, especially the gas generator turbine nozzle and free turbine nozzle sections (key point for the stall characteristics). Additionally, the acceleration controllers’ settings (key points for acceleration and deceleration) should be identified.
(i) Collective increase inputs.
(A) The collective increase inputs will verify:
(a) The transient behavior of the engine, the acceleration, and the minimum NR speed achieved.
(b) The values that the engines could reach in transients depending of the collective inputs.
(c) The effects of the collective anticipation, if any.
(d) The rotorcraft handling qualities (amplitude of variation of pitch, roll and yaw, cross coupling, etc.) during the transients.
(B) The flight test techniques consist of increasing collective from the limit of desynchronization (N2/NR needles are still matched) up to the collective pitch corresponding to the MCP. Depending on the altitude (speed of collective input Page E – 64.1