the rotorcraft. The representative bench test rig should have the ability to generate the torques, speeds, vibration frequency, and acceleration rate generated by the rotorcraft.
The power should have the same method/path of application as that used on the rotorcraft. The test rig should be configured with the same components used for conducting the endurance test on the rotorcraft except that the test components not affected by asymmetric power application may not have to be installed (i.e., if a combining gearbox is used it may not be necessary to have the main transmission installed on the bench test rig).
(3) The takeoff portion of the endurance test should be accomplished on the rotorcraft. When conducting the bench test for 30- second/2-minute OEI it is not necessary to repeat the takeoff portion of the endurance test; however, the simulated power failure and application of 30-second/2-minute OEI power by the remaining engine(s) should be accomplished after the input power has stabilized at takeoff power.
AC 27.927. § 27.927 (Amendment 27-12) ADDITIONAL TESTS.
a. Section 27.927(a):
(1) Explanation. This paragraph is the authority to require any special tests or investigations to establish that the rotor drive system is safe.
(2) Procedures. The certification engineer should review the design of the rotor drive system and its installation and intended operation for features or conditions that may not be adequately qualified in the tests prescribed by this part and, if necessary, additional qualification test programs should be developed and accomplished to ensure safe operation of the rotor drive system. Items of interest would include poorly defined load paths associated with redundant design features, flight deflections of structure and of mounting arrangements, and special or unusual operating procedures which may be anticipated or proposed by the applicant.
b. Section 27.927(b):
(1) Explanation. This paragraph prescribes testing to qualify the rotor drive system for the power excursions to be expected with governor-controlled engines wherein the power from the engine(s) changes automatically to maintain rotor speed at preselected values. At high collective flight control displacements, the normal rotor speed droop will result in the governor-controlled engine(s) automatically accelerating to maximum fuel flow or to any other power, speed, temperature, or torque limiting device, regardless of crew action or artificially established limitations reflected by instrument markings. This high power condition can occur typically during a normal landing when the crew applies high collective to cushion ground contact or, for multiengine rotorcraft, during any flight regime when an engine fails and the corresponding loss of power results in drooping the rotor speed. Special tests are prescribed by this section to provide assurance that the rotor drive system can safely sustain these conditions. The tests of this section should be conducted without intervening disassembly, and all rotor Page E - 19
drive system components should be in serviceable condition after the test. It is permissible but not required that these tests be performed on the same specimen of the rotor drive system used to show compliance with § 27.923.
(2) Procedures. Testing as prescribed by this section should be conducted on a ground-test rotorcraft conformed to the type design suggested for the endurance test of § 27.923. In most cases, testing to comply with § 27.927(b)(1) is accomplished as a continuation of the test of § 27.923 using the same test vehicle. For this test, the main rotor control (cyclic/collective) may be set to simulate vertical lift. The auxiliary rotor control (antitorque) may be set or adjusted to react main rotor torque. Rotation speed should be maximum normal for the test condition; i.e., all engines operating as for takeoff. Using the collective control, obtain torque as required to meet either § 27.927(b)(1)(i) or (ii). This will normally be 110 percent of takeoff torque or a lower value as limited by an approved, reliable device to simultaneously limit torque on all engines. If individual torque limiters are provided for each engine, rigging tolerances should be at maximum allowed mismatch for the type design. For the one-engine-inoperative (OEI) test of § 27.927(b)(2), rotor RPM droop, if any, may be allowed as would occur in service. Since this OEI test requires the remaining engine(s) to produce power not usually available under normal atmospheric conditions, supplemental power augmentation may be needed such as inlet air refrigeration, ramming, or overfueling the engine. Alternatively, bench testing with a transmission test rig may be appropriate providing close simulation of the drive system torsionals, shaft/coupling, misalignment, etc., is achieved. Overtesting (excessive torque) to compensate for inadequacies in the bench test may be negotiated with the FAA/AUTHORITY approval office. Note that compliance with § 27.903(b) requires that the remaining engine(s) be capable of safe, continued operations under the high power conditions of this test. This may require the engine manufacturer to conduct special testing or to produce suitable evidence that the stresses (and gas temperatures) associated with these governor-induced high power excursions do not compromise the airworthiness of the remaining engines or their capability to produce topping power automatically during the initial moments of flight after an engine failure.
c. Section 27.927(c):
(1) Explanation. This paragraph prescribes a test which is intended to demonstrate that in the event of a pressure failure of any pressurized lubrication system used on the rotor drive system, no failure or malfunction will occur in the rotor drive system that will impair the capability of the crew to execute an emergency descent and landing. The lubrication system failure modes of interest usually are limited to failur e of external lines, fittings, valves, coolers, etc., of pressure lubricated transmissions.
(2) Procedures. Conventionally, a bench test (transmission test rig) is used to demonstrate compliance with this rule. Since this is essentially a test of the capability of the residual oil in the transmission to provide limited lubrication, a critical entry condition for the test would be the critical eligible lubricant preheated to the transmission oil temperature limit. With the transmission operating at maximum normal speed, with Page E - 20
lubricant as described above, with nominal cruise torque applied (reacted as appropriate at main mast and tail rotor output quills), and with a vertical load at the mast equal to gross weight of the rotorcraft at 1g, disconnect or cause to leak an external oil plumbing device. Upon illumination of the low oil pressure warning (required by § 27.1305), reduce engine input torque to zero to simulate autorotation, and continue rotation for 15 minutes. Apply input torque to simulate a minimum power landing for approximately 15 seconds to complete the test. Successful demonstration may involve limited damage to the transmission provided it is determined that the autorotative capabilities of the rotorcraft were not significantly impaired.
AC 27.927A. § 27.927 (Amendment 27-23) ADDITIONAL TESTS.
a. Explanation. Amendment 27-23 changed § 27.927 by adding a requirement that the rotor drive system overtorque tests prescribed by § 27.927(b)(3) be conducted at the maximum rotational speed intended for the power condition of the test. The previous rule only specified the torque to be applied to the rotor drive system during the overtorque test.
b. Procedures. The changes to this section did not change the suggested method of compliance.
AC 27.931. § 27.931 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 “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) A design option would be to operate the shafting subcritical; i.e., below the first critical speed, with adequate margins between critical speed and 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 may be permitted. This latter option requires some form of system damping to permit safe Page E - 21
transition through the critical speed range and to avoid excessive nonsyncronous vibrations or instability when transitioning through the critical speed.
(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 and couplings at intervals along the tail rotor drive shaft.
(v) Engine to transmission shafting which, for compliant pylons, incorporates a flexible or geared coupling, to accommodate the misalignment and chucking.
(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 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.
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(1) Subcritical Shafting Designs. Three basic methods of qualification may be considered, with the required margins relative to the degree of assurance provided.
The margins are shown for guidance only.
(i) Analytical.
(A) Simplistic model(s) as shown in figures AC 27.931-1 and AC 27.931-2; 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 stiffnesses, 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 runup to normal speed (and which are not critical flexing modes), may be verified by analyses, laboratory tests, or ground runup with the rotor at maximum permissible unbalance. Damping on the order of 5 percent equivalent viscous damping is generally acceptable.
(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 installation, it is advisable to avoid fixed system modes at multiples of shaft speed, particularly where highly non-isotropic 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 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 Page E - 23
device during transition thereby increasing its serviceability and reliability. Note 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 systems 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) 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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We M + 0.23m = first critical speed, RAD/SEC K = shaft spring rate, LB/IN = 3EI/L3 E = modulus of elasticity = moment on inertia M = mass of weight, LB-SEC?/IN m = mass of shaft, LB-SECHIN FIGURE AC 27.931-1 CANTILEVERED SHAFT, FIRST CRITICAL SPEED DASTE E I We = first critical speed-RAD/SEC u = mass per unit length E = Young's modulus = length between supports 1 = inertia of shaft = a numerical constant: for first critical speed, a = (pi)2 = 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 27.931-2 SHAFT BETWEEN SUPPORT BEARINGS, FIRST CRITICAL SPEED Page E - 25
AC 27-1B, Chg 4 3/31/2014 AC 27.935. § 27.935 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 the 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 and shown valid by test or experience.
AC 27.939. § 27.939 (Amendment 27-11) 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. Smooth, 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 these 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.
Page E - 26
(ii) Procedure. 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. The 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, recirculation 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 operating 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, causes loss of control, causes severe torsional shock through the rotor drive system, or otherwise damages the rotorcraft) is unacceptable. The flight test 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. Operating 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:
(A) Governing stability (e.g., absence of engine parameters oscillations).
Page E – 26.1
AC 27-1B, Chg 4 3/31/2014 (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) Effect of sideslip.
(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 – 26.2
should be reduced when altitude increases) and the rotorcraft category (speed of collective input should be reduced when rotorcraft max weight increases), the duration of the maneuver should be from 1 to 3 seconds at low altitude to approximately 3 to 5 seconds at higher altitude.
(C) Some collective inputs should be made from desynchronized conditions (NR > N2) but should be carefully done as those maneuvers generally lead to larger reductions of NR and high free wheel or engine constraints.
(D) On multi-engine helicopters, collective inputs at lower speeds should be done with one engine at idle position to check the behavior of the other engines in case of an engine failure.
(E) Droop of N2/NR out of the green arc is permitted during the test, provided it is acceptable to the engine or airframe manufacturer for the purpose of the test. The goal is to check for generally satisfactory acceleration (no surge, no extreme N2/NR droop, no significant overshoot). Any N2/NR droop must not result in a condition that requires exceptional piloting skill, alertness, or strength to maintain safe flight.
(F) During those collective inputs, NR below the minimum NR should be avoided by lowering the collective. The reasons for poor engine acceleration should be determined.
(ii) Collective decrease inputs.
(A) Collective decrease inputs will verify:
(a) Combustion stability (i.e., that there is no flame-out).
(b) The engine behavior in transients and in particular the maximum free turbine speed, minimum oil pressure, etc.
(c) The rotorcraft handling qualities (variation of pitch and roll and yaw, cross coupling, etc.) during the transients.
(B) Rate of collective lowering should be built up from 5 seconds down to under 3 seconds. Regarding N2 overspeed, a quick stop type maneuver that leads to an initial acceleration of the NR should be envisaged.
(4) Compressor stall investigation.
(i) The turbine engine installation should not suffer from compressor stalls anywhere in the flight envelope. Previous tests as described above include some of the required areas, but additional testing is necessary. In stabilized conditions, the collective pitch is moved in slow and small oscillations in:
Page E – 26.3
AC 27-1B, Chg 4 3/31/2014 (A) Level flight with or without sideslip at different power settings.
(B) Pull up or in turn.
(C) Autorotation.
(D) Descent at different rates.
(ii) The maneuvers described in paragraphs b.(1)(ii)(A)(4)(i)(A) through (D) above check the engine susceptibility to air flow distortions. The test technique involves oscillations of the collective at various rates around the identified thresholds of the compressor control bleed valves.
(B) Degraded modes: Governor failures.
(1) The specific degraded mode testing for a hydromechanically controlled engine is different than for a digitally controlled engine. However, the basic principles for both types of engine controls include:
(i) Review of the safety analysis, especially the FMEA.
(ii) Determine the failures to be checked in flight, based on prior analysis or testing.
(iii) For each failure to be checked in flight:
(A) Evaluate the behavior of the engine when the failure occurs.
(B) Determine the acceptability of the procedure of identification of the degraded engine.
(C) Assess the human factor aspects of the machine interface (cautions, warning, etc.).
(D) During the use of the degraded engine, consider the workload impact related to engine response in adjusting NR or power for approach and landing.
(E) Evaluate the adequacy of the RFM procedures.
(2) Helicopters equipped with FADECs are becoming increasingly integrated into the helicopter systems for engine control sensor inputs. For example, N2/NR values available could depend on parameters such as airspeed, altitude, or temperature as provided by the basic helicopter systems and sent to the FADEC. A Page E – 26.4