7/30/97 АС 29-2B
425. § 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 "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:
Main rotor/mast/transmission assemblies rigidly mounted to the airtrame;
(її) 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;
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(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 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.
(i¡) 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:
Analytical.
(A) Simplistic model(s) as shown in Figures 425-1 and 425-2;
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.
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(™1) 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 analysis, 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 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.
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АС 29-2B 7/30/97
(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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Wer M + 0.23m Wer = first critical speed, RAD/SEC = shaft spring rate, LB/IN = 3EI/L 3 = modulus of elasticity = moment on inertia M = mass of weight, LB-SECZ/IN m = mass of shaft, LB-SECZ/IN FIGURE 425-1. CANTILEVERED SHAFT, FIRST CRITICAL SPEED Wer E I 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 = a numerical constant: for first critical speed, a = (pi)z = 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 425-2. SHAFT BETWEEN SUPPORT BEARINGS, FIRST CRITICAL SPEED Par 425 675
426. § 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 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.
427. § 29.939 (Amendment 29-12) TURBINE ENGINE OPERATING CHARACTERISTICS.
a. Explanation. This section requires evaluation of engine operation, engine inlet airflow distortion, and engine/drive system torsional stability. A satisfactory rotorcraft design for all three items should be established by the manufacturer early in his development program since changes in design to satisfy these requirements are typically very expensive and will adversely impact other basic design features. The role of the certification engineer is to assure that the manufacturer's evaluation programs have been thorough and conclusive. The certification engineer should also determine, where applicable, that the FAA/AUTHORITY-approved Engine Installation Manual requirements are met.
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 reatures 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.
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(i) 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. This should include maximum airspeed-sideslip combinations, power recoveries, hover with wind from all azimuths and other maneuvers appropriate to the type. Recirculation of exhaust gases during hover can be critical for engine operation. Particular attention should be given to flight/operating conditions which 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 and rearward flight at high altitude has resulted in 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.
(2) Vibration.
(i) Explanation. Engine airflow patterns are deflected or distorted by the presence of airframe inlet hardware, cowling, fuselage panels, and, to a degree, in almost all flight regimes. Additional items such as airframe installed particle separators, deflectors for snow, ice, or sand protection, and obstructions forward of the engine inlet, such as a hoist kit, could affect the engine air flow patterns. The rotating elements of the engine, particularly the compressor blades, will be subjected to a cyclically varying air flow as these elements move into and out of areas of deflected airflow to the engine.
A corresponding aerodynamic load will be imposed on these engine elements. Since this loading is also cyclic, the possibility of critical frequency coupling with an engine component shall be investigated.
(®) Procedure. Typically, this evaluation would involve installation in the engine inlet of a special multiple probe, total pressure sensing system, and flight testing which largely follows that prescribed for evaluation of engine operating characteristics as described above. Data from these tests can be reduced to create a pressure map at the compressor inlet face which, in conjunction with compressor speeds, may be used to determine the frequencies and relative amplitudes of the cyclic air loading imposed on the engine compressor blades. The engine manufacturer either supplies the sensing probe or specifies its design and performance. Also, the engine manufacturer may evaluate the test results or publish acceptance criteria. A wave analysis may be involved in identifying higher order excitations. Engine exhaust ducts which include bends, noise suppressors, or other obstructions may require an evaluation similar to that discussed above for the engine inlet. The engine manufacturer should be consulted for instructions or approval of this aspect. High performance engines may also require an engine inlet temperature survey. Details of instrumentation and acceptance criteria should be provided by the engine manufacturer. Engines equipped with only centritugal compressors are less likely to encounter frequency coupling and Par 427 677
may not require this investigation. The engine manufacturer's recommendations should be followed in these cases.
(3) Torsional Stability.
(i) Explanation. Governor-controlled engines installed in rotorcraft are subject to a fuel control resonant feedback condition which could be divergent if not properly designed or compensated. This condition occurs when the response frequency of the governor on the engine is coincident with or close to a low order natural torsional frequency of the rotorcraft's rotor drive system. Typically, these frequencies appear in the 3 to 5 cycles per second (CPS) range. The manufacturer usually resolves torsional instability problems by introducing damping into the engine governor/fuel control. Provisions for this change must be supplied by or approved by the engine manufacturer. The final configuration may be a compromise between a lightly damped control, which will allow a positive but slow convergence of drive system torsional oscillations, and a highly damped control which exhibits excessive rotor speed droop or overspeed following rotorcraft collective control displacement.
(ii) Procedures. A ground and flight test program should be devised to evaluate the torsional response of the engine and drive system combination presented by the applicant. Instrumentation to record drive system torsionals should be applied to all major branches of the drive system. Engine parameters such as torque, RPM fuel manifold/nozzle pressure, compressor discharge pressure, and governor lever position should be recorded simultaneously with drive system parameters. The test program should include ground tie-down operation and flight operation across a range of engine power and rotor speeds while injecting control inputs as close to the first order drive system natural frequency as possible. Mechanical methods of making these inputs are not usually necessary if the desired frequency is in the 3 to 5 CPS range and the instrumentation readout confirms that the drive system was actually excited torsionally at its natural frequency. Control inputs should include collective, antitorque, and throttle. Also, cyclic inputs may be important on tandem rotor rotorcraft. The acceptance criteria may be dependent on several items. Among these are rotor and drive system fatigue loading, engine power response characteristics, limitations established by the engine manufacturer, etc. The acceptance criteria are usually stated as a percent damping (minimum). Typically, 1 percent of critical equivalent viscous damping (or greater) is acceptable. In effect, this means that the free vibration response to a control input damps to ½ amplitude in 11 cycles or less.
428.-445. RESERVED.
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SECTION 26. FUEL SYSTEM
446. § 29.951 (Amendment 29-12) GENERAL.
a. Explanation.
(1) The term "fuel system" means a system which includes all components required to deliver fuel from the tank(s) to the engine(s). This includes, but is not limited to, all components provided to contain, convey, drain, filter, shutoff, pump, jettison, meter, and distribute fuel to the engines.
(2) Paragraph (a) of this section is a general statement of the performance requirements for fuel systems and constitutes authority to require fuel systems to be adequate notwithstanding compliance with detail requirements listed in §§ 29.953 through 29.999 of this subpart.
(3) Paragraph (b) of this section requires fuel systems to be designed so that air will not enter the system under any operating conditions by either arranging the system so that no fuel pump can draw fuel from more than one tank or by other acceptable means.
(4) Paragraph (c) of this section sets forth a fuel system performance requirement intended to ensure that ice to be expected in fuel when operating in cold weather will not prevent the fuel system from supplying adequate fuel to the engines.
Although fuel system filters and strainers are the items in the fuel system most susceptible to clogging from ice particles in the fuel, this paragraph requires that the entire fuel system be shown to be capable of delivering fuel, initially contaminated with ice, to the engine(s).
b. Procedures.
(1) For Paragraph (a), the applicant should show compliance with the fuel system requirements of this subpart, except that if unusual fuel system arrangements or requirements exist which are not adequately addressed by these subparts, this paragraph may be used as authority to require special tests, analysis, or system performance needed for proper engine functioning.
(2) For Paragraph (b), review the fuel system design with special attention to fuel tank selector valves, crossfeed systems, and multiple tank outlet arrangements to ensure that no allowable fuel system configuration will permit air to enter the system.
For questionable situations, the applicant should conduct ground or flight tests, as necessary, to verify compliance with this section.
(3) Paragraph (c) provides for sustained satisfactory operation of the fuel system with cold fuel initially contaminated with water. Since ice in the fuel system is Par 446 697
АС 29-28 7/30/97 not considered to be an emergency condition but, rather, is an expected service encounter, compliance would not involve the imposition of special rotorcraft limitations.
Flight manual instructions such as land as soon as practicable, reduce altitude to some value less than otherwise permitted, reduce power, turn on boost pumps, etc., are not appropriate in demonstrating compliance. Some methods of fuel system ice protection which have been used to show compliance follow.
(i) Fuel heater. Usually these devices are fuel-to-engine oil heat exchangers and are normally located to protect the fuel filter from blockage by ice in the fuel. The adequacy of these devices should be established. Usually this involves generation of a heat balance between heat gained by fuel and heat lost by oil using performance data provided by the manufacturers of the fuel-oil heater, the oil cooler, the heat rejected by the engine to the oil, etc. A minimum oil temperature associated with the adequacy of the fuel heater may need to be established, marked on the oil temperature gauge, and verified to be maintained during critical flight conditions. Other unprotected parts of the fuel system remain to be evaluated and substantiated for compliance with this requirement.
(ії) Oversized fuel filter. This method may only substantiate the fuel filter and, as with the fuel heater method, is incomplete without evaluation of the remainder of the fuel system. An icing test of the filter should be accomplished. Fuel preparation procedures and method of testing should follow the applicable portion of SAE Aerospace Recommended Practice (ARP) No. 1401. A satisfactory configuration is achieved when a filter is demonstrated to have the capacity to continue to provide the filtration function, without bypassing, when subjected to fuel contaminated by ice to the degree required by this rule. Usually, a delta pressure caution signal for the filter is needed to alert the flightcrew that progressive filter blockage is in progress. The caution device setting should be established by test which demonstrates that after illumination of the caution signal sufficient filter capacity exists to enable completion of the flight. Fuel pressure should not fall below established limits because of ice accumulation on the filter.
(¡¡i) Anti-ice additives. This method utilizes the properties of ethylene glycol to reduce the freezing temperature of water in the fuel. It has the advantage over other methods of protecting all components in the fuel system from ice blockage.
Compliance with the rule by this method involves the following.
(A) Eligible additives. PFA-55MB (Phillips Petroleum Co.) and additives per specification MIL-I-27868, Revision D, or earlier. Later versions of this specification do not require glycerin, which may be needed to protect fuel tank coatings.
(B) Compatibility. Both engine fuel system and aircraft fuel system should be verified to be chemically compatible with the additive at the maximum concentration to be expected in the fuel system. Usually, information on eligible system 698 Par 446
materials can be obtained from the engine manufacturer for the engine fuel system and from the additive manufacturer for aircraft fuel system materials.
(C) Adding or blending the additive to the fuel. These additives do not mix well with the fuel and indiscriminate dumping of additive into the tank will not only fail to protect the system from ice accumulation but likely will damage nonmetallic components in the system. Some fuels may have additive premixed in the fuel. If other fuels are to be eligible, a method for blending additive into the fuel during refueling must be devised and demonstrated to be effective.
(D) Placards should be added near the fuel filler opening to note that fuel must contain the approved anti-ice additive within the minimum and maximum allowed concentration.
(E) The FAA/AUTHORITY-approved flight manual should contain necessary information to attain satisfactory blending of the additive and procedures to allow the operator to check the blend in the fuel tank.
(iv) Fuel system protection (other than filters). If the fuel heater method or oversize filter method (paragraphs b(3)(i) and b(3)(ii)) is proposed, the remainder of the fuel system should be shown to be free from obstruction by fuel ice. This may be shown by testing the system with ice contaminated fuel (prepared as suggested for filter tests) or, in many cases, by selecting fuel system components which by test or by previous experience are known to be free of ice collection tendencies. Tank outlet screens (or tank-mounted pump inlet screens) may be the significant fuel system feature for turther evaluation. In some instances, fuel turbulence due to pump motions may be sufficient to keep the screen clear of ice. In other instances, small screen bypass openings (approximately one-fourth inch in diameter) located outside the predominant tuel flow path have been found satisfactory.
NOTE: Advisory Circular (AC) 20-29 contains information regarding compliance with the fuel ice protection requirements of Part 25. The information in this AC is largely valid except for references to the quantity of water to be expected in fuel and the amount of additive required to ensure freedom from fuel ice hazards.
447. § 29.952 (Amendment 29-35) FUEL SYSTEM CRASH RESISTANCE.
a. Explanation.
(1) Section 29.952 provides safety standards that minimize postcrash fire (PCF) in a survivable impact. The rule contains comprehensive crash resistant fuel system (CRFS) design and test criteria that significantly minimize fuel leaks, creation of potential ignition sources, and the occurrence of PCF. Section 29.952 accomplishes this for survivable impacts by- Par 446 699
(i) Providing comprehensive criteria to minimize fuel leaks and potential ignition sources;
(il) Requiring increased crash load factors for fuel cells in and behind occupied areas to ensure the static, ultimate strength necessary for impact energy absorption, structural integrity, fuel containment, and occupant safety;
(iii) Maintaining the load factors of § 29.561 for fuel cells in other areas (particularly underfloor cells) to ensure leak-tight fuel cell deformation in energy absorbing underfloor structure without unduly crushing or penetrating the occupiable volume; and
(iv) Requiring a 50 ft. dynamic vertical impact (drop) test to measure fuel tank structural and fuel containment integrity.
(2) Section 29.952 applies to all fuel systems (including auxiliary propulsion unit (APU) systems).
(3) Some similarities exist among the fire protection requirements of §$ 29.863,
29.1337(a)(2), and 29.952. The requirements in each standard are not mutually exclusive. Overlapping requirements should be certified simultaneously.
(4) The use of bladders is not mandated as this would unduly dictate design.
However, in the majority of cases, their use is necessary to meet the test requirements of § 29.952. If a design does not use bladders, the application should be treated as a new and unusual design feature that should be thoroughly coordinated with the Airworthiness Authority for technical policy to insure adequate safety. Experience has shown that bladders with wall thicknesses from 0.03 to 0.018 inches typically meet the § 29.952 test requirements.
b. Related Material. Documents shown below may be obtained from The Naval Publications and Forms Center, 5801 Tabor Avenue, Philadelphia, Pennsylvania 19120-5094, ATTN: Customer Service (NPODS).
(1) Military Specification, MIL-T-27422B, Amendment 1, April 13, 1971, Tank, Fuel, Crash-resistant Aircraft.
(2) Military Standard, MIL-STD-1290 (AV), January 25, 1974, Light Fixed and Rotary Wing Aircraft Crashworthiness.
(3) Military Standard, MIL-H-83796, August 1, 1974, Hose Assembly, Rubber, Lightweight, Medium Pressure, General Specification for.
(4) Military Specification, MIL-V-27393 (USAF), July 12, 1960, Valve, Safety, Fuel Cell Fitting, Crash Resistant, General Specification, for.
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