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.
(1) 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 belmouth 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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(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.
(1) 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).
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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.
(E) 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.
1) 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 Np 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/Np 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/Np droop must not result in a condition that requires exceptional piloting skill, alertness, or strength to maintain safe flight.
(E) 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.
(ji) Collective decrease inputs.
(A) Collective decrease inputs will verity:
(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 Np should be envisaged.
(4) Compressor stall investigation.
1) 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:
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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.
(I) 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
concern is the effects of degradations, discrepancies or failures in the basic helicopter systems on engine governing laws and power available. Therefore, a review of the FMEA before beginning engine failure or degraded mode evaluations should be performed.
(3) It is recommended to have special test equipment installed on the test engine in order to simulate failures to the FADEC or digital engine control. Test risk mitigation should include the evaluation of possible malfunctions of this special test equipment.
(4) As a minimum, the following typical failures or degraded modes, including engine control reversion that is based on safety analysis and ground test results, should be evaluated in flight:
(i) Failure causing the engine to increase power.
(I) Failure causing the engine to decrease power.
(II) Failure fixing the engine power at a static value.
(iv) Failure introducing oscillations.
(v) Failure introducing lower acceleration or deceleration.
(vi) Failure causing the N2/NR to increase or decrease.
(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.
(il) 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 Page E - 27
AC 27-1B, Chg 4 3/31/2014 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 centrifugal compressors are less likely to encounter frequency coupling and 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 rotor drive system. Typically, these frequencies appear in the 3 to 5 CPS range. The manufacturer usually resolves torsional instability problems by introducing damping into the engine governor or 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) Procedure. 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 and power turbine speed 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.
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SUBPART E - POWERPLANT
FUEL SYSTEM AC 27.951. § 27.951 (through Amendment 27-9) FUEL SYSTEM--GENERAL.
a. Explanation.
(1) The term "fuel system" means a system which includes all components required to deliver fuel 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 the fuel system to be adequate notwithstanding compliance with detail requirements listed in §§ 27.953 through 27.999 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 water and cooled to critical icing conditions, 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 fuel system configuration will allow air to enter the system. For questionable situations, the applicant should conduct ground tests and flight tests as necessary to verify compliance with this section.
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(3) Paragraph (c) provides for sustained satisfactory operation of the fuel system, with initially ice-contaminated fuel. Since ice in the fuel system is 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.
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.
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 flight crew 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.
(ili) 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 materials can Page E - 30
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 tuel 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 anti-ice additive PFA-55MB MIL-I-27686 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 AC 27.951b(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 further 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 fuel 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, § 25.997(b). 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.
AC 27.952. § 27.952 (Amendment 27-30) FUEL SYSTEM CRASH RESISTANCE.
a. Explanation.
(1) Section 27.952 (added by Amendment 27-30) 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 27.952 accomplishes this for survivable impacts by:
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AC 27-1B, Chg 1 2/12/03
(1) Providing comprehensive criteria to minimize fuel leaks and potential ignition sources;
(її) 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 § 27.561 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 27.952 applies to all fuel systems (including auxiliary propulsion unit (APU) systems).
(3) Some similarities exist among the fire protection requirements of §§ 27.863,
27.1337(a)(2), and 27.952. 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 § 27.952. If a design does not use bladders, the application should be treated as a new and unusual design feature and 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 § 27.952 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. (canceled 6-7-97 without replacement)
(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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(5) Military Specification, MIL-H-25579 (USAF).
(6) Military Specification, MIL-H-38360.
(7) U.S. Army Publication USARTL-TR-79-22E, "Aircraft Crash Survival Design Guide, Volume V---Aircraft Postcrash Survival", dated January 1989.
NOTE: Section 4, "Postcrash Fire Protection" of Volume V of the Design Guide is the modern update to MIL-STD-1290. Section 4 contains a comprehensive design guide for military CRFS designs that may be useful for civil CRFS designs.
c. Conceptual Definitions.
(1) Survivable Impact. An impact (crash) where human tolerance acceleration limits are not exceeded in any of the principal rotorcraft axes, where the structure and structural volume surrounding occupants are sufficiently intact during and after impact to constitute a livable volume and permit survival, and where an item of mass does not become unrestrained and create an occupant hazard. "Livable volume" relates to the ability of an airframe to maintain a protective shell around occupants during a crash and to minimize threats, such as accelerations, applied to the occupiable portion of the aircraft during otherwise survivable impacts. In lieu of a more rational, approved criteria, the load factors of § 27.952(b)(1) constitute the structural human survivability accelerations limits.
(2) Postcrash Fire (PCF). A fire occurring immediately after and as a direct result of an impact. The fire is either the result of fuel released from a leaking fuel system reaching an existing or a crash-induced ignition source, a crash-induced ignition source internal to an undamaged or damaged fuel system, or a combination. PCF's have an intensity range from the minimum of a small local flame to the maximum of an instantaneous massive fire or fireball (explosion).
(3) Fuel Tank or Cell. A reservoir that contains fuel and may consist of a hard shell (of a composite, metal, or hybrid construction) with either a laced-in, snapped in, or otherwise attached semirigid or flexible rubber matrix bladder (or liner), spray-on bladder, or no bladder. The hard shell may be either the airframe (integral tank) or a separate rigid tank attached to the airframe. The device has inlets and outlets for fuel transter and internal pressure control.
(4) Ignition Source. An ignition source that when wet with fuel or in contact with fuel vapor would cause a PCF.
(5) Major Fuel System Component. A fuel system part with enough mass, installation location hazard or a combination to be structurally considered in a crash.
Structural consideration is required when crash-induced relative motion can occur between the part and its surrounding structure from inertial impact forces, airframe deformation forces, or for other reasons.
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