Pressure Altitude Air data basis for Flight Speed establishing power Free Air Temperature plant inlet pressure (stagnation) and temperature Torque Power Turbine Speed Gas Generator Speed(s) Measured Gas Temperature Direct measurement of power output.
Primary thermodynamic and limiting parameters Fuel Flow Secondary trend monitoring and potential limiting parameter (її) The overall power check accuracy can be assessed on a suitable statistical basis using equations that link the measured parameters and inserting system accuracy distributions for each value. This approach will provide an overall assessment of power check accuracy and will highlight major contributors to error. The accuracy assessment at each parameter should include the following elements:
Sensor error Indicator error System error Reading error (i¡i) This assessment might show that while conventional instrument displays of air data are acceptable, servo driven digital displays are desired for engine parameters. Further, displays that provide a "snapshot" of engine readings at a given moment may be useful in avoiding variation in power level during the finite period needed to manually read and log the set of parameters.
(2) Installation Loss Definition.
(i) Installation loss definition is an extremely important aspect of any form of rotorcraft engine performance. Engines are certificated and sold with uninstalled performance guarantees and estimates as to the power output capabilities. Installation of the engine in the rotorcraft imposes power output penalties that must be accounted for in any sort of power assurance check procedure. Normal practice dictates that the engine manufacturer provides a computer program that accurately predicts the engine power output capability throughout the approved flight envelope. This computer program has the capability to correct the power output for the losses incurred by the rotorcraft installation.
(II) Losses that can reduce engine power available are as tollows:
(A) Air intake total pressure loss Page MG 9 - 6
(B) Air intake total temperature rise (C) Exhaust back pressure (D) Accessory power extraction (E) Compressor bleed air extraction (F) Off-optimum power turbine output speed effects
(iii) The above items and methods of dealing with them are clearly defined in SAE Aerospace Recommended Practice (ARP) 1702. Typically, these losses will not be a fixed percentage but will vary with engine operating conditions and environment.
(iv) Any calculations involving power assurance data should use the approved engine performance program, and the rotorcraft losses should be input on a discrete basis so that the interaction between losses and their independent variability is properly considered. This approach is clearly defined in ARP 1702. Accurate consideration of the losses should produce a Power Assurance Check that will preclude premature removal of acceptable engines or continued operation of inadequate power
g. Rotorcraft Flight Manual RFM).
(1) The Power Assurance Check data for the installed engine (engine data adjusted for inlet losses, exhaust losses, bleed extraction, power extraction, and off-optimum output shaft speed operation) should be presented in the RFM in an easily useable format. The data format may consist of charts of engine torque (at constant power turbine shaft speed) versus allowable values of gas generator speed and gas path temperature covering the range of ambient conditions for takeoff operations.
Associated limitations for the rotorcraft transmission and the engine should be noted.
(2) The RFM should also:
Include succinct statements of the reason for the Power Assurance Check and what must be done if the Power Assurance Check results are not acceptable.
(ii) Clearly state that Power Assurance Check either is a pre-takeoff or in-flight procedure, as required by operations, specifications and/or other approval authority documents.
(ili) Be kept simple, easy to use, and identify equipment operation limitations and requirements.
Page MG 9 - 7
Page MG 9 - 8 AC 27-1B FUEL
CONTROL FUEL FLOW
TRANSMITTER
FUEL PUMP FIGURE AC 27 MGIL1 BY PASS
TEE NEEDLE
FUEL PRESSURE VALVE
TRANSMITTER FUEL FUEL NOZZLE
PRESSURE GAGE
COCKPIT FUEL AIRCRAFT
FLOW GAGE FUEL TANK
INSTALLED ENGINE FUEL FLOW CAPACITY CHECK 9/30/99
CHAPTER 3
AIRWORTHINESS STANDARDS
NORMAL CATEGORY ROTORCRAFT
MISCELLANEOUS GUIDANCE (MG) AC 27 MG 10 ADVISORY MATERIAL FOR SUBSTANTIATION OF EMERGENCY
FLOTATION SYSTEM
a. Reference. FAR sections 27.521, 563(b), .751, 753(a)(1), (a)(2), .801(b), (d), .807 (d).
b. Explanation.
(1) This section pertains to emergency flotation systems used to provide buoyancy for rotorcraft not specifically certificated for ditching but performing over-water operations. According to paragraph AC 27.801, ditching may be defined as an emergency landing on the water deliberately executed with the intent of abandoning the rotorcraft as soon as practical. Currently, ditching certification is not required by FAR 27; however, certification requirements are prescribed for applicants requesting ditching certification approval. If a rotorcraft operates over water during a Part 135 operation, the rotorcraft must comply with FAR 135.183, which may require floats.
(2) There are no airworthiness rules specifying the minimum standards for emergency flotation systems on rotorcraft not certificated for ditching requirements.
Equipment presented for evaluation must perform its intended function and not create a hazard for the rotorcraft or occupants. The objective in evaluating emergency flotation systems is safe flight and evacuation of the rotorcraft in emergency situations.
Adequate emergency flotation systems would aid in keeping rotorcraft sufficiently upright and in adequate trim to permit safe and orderly evacuation in an emergency water landing.
c. Procedures. The following guidance criteria is based on past certification policy and experience for emergency flotation systems. Demonstration of compliance to other criteria may produce acceptable results if adequately justified by rational analysis.
Model tests of the appropriate emergency water landing configuration may be conducted to demonstrate satisfactory flotation and trim characteristics where satisfactory correlation between model testing and flight testing has been established.
Model tests and other data from rotorcraft of similar configurations may be used to satisfy the water requirements where appropriate.
(1) Flotation Systems.
Normally inflated. The flotation systems which are normally inflated and intended for emergency use only, should be evaluated for:
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(A) Structural integrity when subjected to:
(1) Air loads throughout the approved flight envelope with floats installed,
2) Water loads during water entry, and
(3) Water loads after water entry at speeds likely to be experienced after water impact.
(B) Rotorcraft handling qualities throughout the approved flight envelope with floats installed.
(il) Normally deflated. Emergency flotation systems which are normally stowed in a deflated condition and inflated either in flight or after water contact during an emergency water landing should be evaluated for:
(A) Inflation.
(1) Proper Inflation. The inflation system design should minimize the probability of the floats not inflating properly or inflating asymmetrically. This may be accomplished by use of a single inflation agent container or multiple container system interconnected together. Redundant inflation activation systems will also normally be required. If the primary actuation system is electrical, a mechanical backup actuation system will usually provide the necessary reliability. A secondary electrical actuation system may also be acceptable if adequate electrical system independence and reliability can be documented.
(2) Inadvertent actuation. The inflation system should be safeguarded against spontaneous or inadvertent actuation for all flight conditions. It should be demonstrated that float inflation at any flight condition within the approved operating envelope will not result in a hazardous condition unless the safeguarding system can be shown to be reliable. Limitations to the approved envelope can be established so inadvertent actuation does not impose a hazard at the new envelope.
(3) Float actuation. The float activation means may be fully automatic or manual with a means to verify primary actuation system prior to each flight. If manually inflated, the float activation switch should be located on one of the primary flight controls. These activation means should be safeguarded against spontaneous or inadvertent actuation for all flight conditions.
(4) Flight Limitations. Maximum airspeeds for intentional in-flight actuation of the float system and for flight with the floats inflated should be established as limitations in the Rotorcraft Flight Manual (RFM) unless in-flight actuation is prohibited by the RFM.
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(5) Inflation time. For floats inflated automatically by water contact, inflation time from actuation to neutral buoyancy should be short enough to prevent the rotorcraft from becoming submerged to the point where egress is impeded.
(6) Pressure checking. A means should be provided for checking the pressure of the gas storage cylinders prior to each flight. A table or device showing acceptable gas cylinder pressure variation with ambient temperature and altitude (if applicable) should be provided.
Over inflation. A means should be provided to minimize the possibility of over inflation of float bags under any reasonably probable actuation conditions.
(3) No puncture inflation. The ability of the floats to inflate without puncture when subjected to actual water pressure should be substantiated. A full scale rotorcraft immersion demonstration in a calm body of water is one acceptable method of substantiation. Other methods of substantiation may be acceptable depending upon the particular design of the flotation system.
(9) Flotation bag containment. Float installations should be evaluated to ascertain that emergency exits are not blocked by the inflated floats when the float bags are inflated to their maximum inflation pressure or their most adverse inflation pressure for emergency exits and the rotorcraft at its most critical weight and center of gravity configuration.
(B) Structural Integrity. The flotation bags should be evaluated for loads resulting from:
(1) Airloads during inflation and fully inflated during the most critical flight conditions and water loads with fully inflated floats during water impact for the rotorcraft desiring float deployment before water entry; or
(2) Water loads during inflation after water entry.
(C) Handling qualities. Rotorcraft handling qualities should be verified by tests or analysis to comply with the applicable regulations throughout the approved operating envelopes for:
(1) Deflated and stowed condition,
2) In-flight inflation condition
(3) Fully inflated condition, and
(4) Partially inflated condition, assuming the most critical float compartment fails to inflate.
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(2) The float system attachment hardware should be shown to be structurally adequate to withstand critical air loads and water loads during water entry when both deflated and stowed and fully inflated (unless in-flight inflation is prohibited). The appropriate vertical loads and drag loads determined from water entry conditions (or as limited by flight manual procedures) should be addressed. The effects of the vertical loads and the drag loads may be considered separately for the analysis.
(3) Flotation and Trim should be investigated for a range of sea states from zero to the maximum selected by the applicant and should be satisfactory in waves having height/length ratios of 1:12.5 for multi engine rotorcraft with Category A engine isolation and 1:10 for all other rotorcraft.
(i) Demonstrated to be satisfactory to at least sea state 4 water conditions.
(ii) Flotation tests should be investigated at the most critical rotorcraft loading condition.
(iii) Flotation time and trim requirements should be evaluated with a simulated, ruptured deflation of the most critical float compartment. Flotation characteristics should be satisfactory in this degraded mode to at least sea state 2 water conditions.
(iv) Probable rotorcraft door/window open or closed configurations and probable damage to the airframe/hull (i.e., failure of doors, windows, skin, etc.) should be considered when demonstrating compliance with the flotation and trim requirements.
(4) Float System Reliability. Reliability should be considered in the basic design to ensure approximately equal inflation of the floats to preclude excessive yaw, roll, or pitch in flight or in the water.
(i) Maintenance procedures should not degrade the flotation system (such as introducing contaminants which could affect normal operation, etc.).
(ii) The flotation system design should preclude inadvertent damage due to normal personnel traffic flow and excessive wear and tear. Protection covers should be evaluated for function and reliability.
(5) Buoyancy requirements for emergency flotation systems should be a minimum of 25 percent excess buoyancy at maximum internal gross weight. The weight of fresh water (density 62.42 Ib/ft3) displaced by fully submerged float or floats should be a minimum of 25 percent greater than the maximum certificated gross weight of the rotorcraft. Analysis may be used for buoyancy verification.
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(6) Sufficient watertight compartments should provide an acceptable margin of positive stability with any single main float compartment flooded or deflated. The location of the floats, the most critical compartment, the rotorcraft weight, mass moment of inertia, and center of gravity location are also important considerations for stability. Analyses, tests, or a combination thereof may be used to substantiate a positive margin of stability with the most critical compartment flooded or deflated.
(7) The inflatable bag type floats should be designed for the maximum pressure differential developed at the maximum design altitude. That is, the resulting pressure difference between an operational altitude and a take-off site elevation should be established and substantiated. This resulting pressure differential may become an operating limitation.
(8) The float landing load factors may be determined from the drop test of the float landing gear or the loads may be derived from landing gear drop test or loads may be determined from model or full scale water entry tests. The vertical loads are distributed over three fourths of the bag's projected area. Bag floats are not subject to the side loads. Rigid floats are to be designed for vertical, horizontal, and side loads distributed along the length of the float.
(9) Design and/or support of the forward part of bag type floats should be evaluated for maximum design speeds to prevent collapse or significant distortion of the bag while in flight.
(10) Resistance to puncture and abrasion at attach/wear points is an important design consideration. Girt or attachment design loads should be sufficient to withstand the maximum imposed design loads.
(11) Occupant Egress and Survival. Each practicable design measure should be taken to minimize the probability that the behavior of the rotorcraft would cause immediate injury to the occupants or prevent evacuation of the rotorcraft after an emergency landing on water. Emergency exits should be located such that they are above the waterline and will not be blocked by the inflated or partially inflated floats, impeding evacuation of the rotorcraft. The flotation time and trim of the rotorcraft should allow the occupants to evacuate the rotorcraft. i.e., the rotorcraft should remain sufficiently upright and in adequate trim to permit safe and orderly evacuation of all personnel. For configurations which are considered to have critical occupant egress capabilities due to float proximity, an actual demonstration of egress may be required.
When a demonstration is required, it may be conducted on a full-scale rotorcraft actually immersed in a calm body of water or using any other rig/ground test facility shown to be representative. The demonstration should show that floats do not impede a satisfactory evacuation.
(12) Rotorcraft Flight Manual. The Rotorcraft Flight Manual should contain the information pertaining to the emergency flotation system. This material should include:
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(i) The information pertinent to the limitations applicable to the emergency float system and operating limitations for the emergency float system,
(ii) Procedures and limitations for flotation device inflation,
(iii) Procedures for use of emergency flotation equipment, and
(iv) Procedures for emergency water landing occupant evacuation.
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CHAPTER 3
AIRWORTHINESS STANDARDS
NORMAL CATEGORY ROTORCRAFT
MISCELLANEOUS GUIDANCE (MG) AC 27 MG 11 FATIGUE EVALUATION OF ROTORCRAFT STRUCTURE
a. Purpose. This revision to the advisory circular sets forth acceptable means of compliance with the provisions of Federal Aviation Regulations, §§ 27.571 and 29.571, dealing with the fatigue safe-life evaluation of metallic rotorcraft structure. The previous general guidance for fail-safe methodology is retained herein, as this approach also remains an option for compliance with Federal Aviation Regulation, § 27.571. General guidance and some background to fatigue evaluation issues are also provided.
Guidance for evaluation of composite structure may be found in Chapter 3, AC 27 MG 8, AC No. 20-107A, and AC 29-2C.
b. Background. The fatigue evaluation procedures outlined in this advisory circular are for guidance purposes only and are neither mandatory nor regulatory in nature. Although a uniform approach to fatigue evaluation is desirable, it is recognized that in such a complex problem, new design features and methods of fabrication, new approaches to fatigue evaluation, and new configurations may require variations and deviations from the procedures described herein. Engineering judgment should therefore be exercised for each particular application. The flight structure of the rotorcraft is subject to cyclic vibratory stresses in practically every regime of flight. In addition, since it is a highly maneuverable aircraft that is capable of forward, rearward, sideward, vertical, and rotational flight, operating limitations due to fatigue are possible in practically all flight situations. For these reasons, it is required that special attention be focused on the fatigue evaluation of the flight structure of the rotorcraft.
(1) Fatigue evaluation of the flight structure is intended to verify structural reliability. Assurance of structural reliability starts with design, including choice of materials for resistance to crack initiation and/or propagation, detail design to minimize stress concentration, and specification of surface finishes, fits, etc. Design analysis should include estimation of expected flight loads, and estimation of resistance to fatigue. Fatigue strength should be based on past full-scale fatigue tests and/or materials fatigue data with appropriate reductions for the variability in fatigue strength, size, shape, surface finish, and environments of the structure. In addition, design for fatigue should consider mode-of-failure analysis, areas susceptible to fatigue cracking, and methods to assure detectability of fatigue cracks; when fail-safe design is the chosen method. The residual strength of a cracked structure is an important consideration of fail-safe design.
(2) Assurance of structural reliability also includes manufacture and fabrication in accordance with design requirements and specifications, quality control to monitor compliance, and effective service inspection procedures.
AC Page MG 11 - 1
AC 27-1B, Chg 1 2/12/03
(3) Fatigue evaluation of the structure, measurement of flight loads and stresses, and evaluations of fatigue strength and/or fatigue crack propagation are the subjects of this advisory circular. There is some question whether a completely reliable method for the prediction of time to fatigue crack initiation and fracture exists. Nevertheless, one engineering approach to the subject is to use theiLinear Cumulative Damage Hypothesis." This hypothesis states that every cycle of stress above an "endurance limit produces fatigue damage proportional to the ratio of cycles accumulated at the stress to fatigue "life" at that stress.
(4) Laboratory tests of this hypothesis indicate that it is reasonably valid when the loading spectrum consists of stresses that are, in effect, random. Despite the lack of an adequate theory connecting this hypothesis with more basic properties of materials, it has been successfully used in a number of applications to calculate a safelife retirement time.
(5) In addition, fatigue evaluation generally requires a method of accounting for the effect of steady loads and stresses on fatigue. Where the manufacturer does not provide other substantiating data, a Goodman diagram may be used to account for these effects.
(6) In any rational fatigue evaluation, the following factors should be considered:
Identification of the structure to be considered in the fatigue evaluation. Those elements of the rotorcraft structure that may be critical in fatigue should be identified. Typical elements include:
(A) Rotor blades and attachment fittings.
(B) Rotor heads, including hubs, hinges, dampers.
(C) Rotor drive components, including gearboxes and transmission shafts.
(D) Control system components, including control rods, servos, swashplates.
(E) Rotor supporting structure.
(F) Primary flight and ground load paths of the fuselage, including landing gear, lift frames, stabilizers and auxiliary lifting surfaces.
(ii) The loads and stresses associated with steady and maneuvering operating conditions expected in service.
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(III) The frequency of occurrences associated with various flight conditions and the corresponding spectrum of loading and stresses.
(iv) The fatigue characteristics of the structure, including fatigue strength, and as necessary, crack propagation and residual strength characteristics.
c. Flight Load Measurement Program.
(1) General. Subsequent to design analysis, in which aircraft loads and associated stresses are derived, the stress level and/or loads are to be verified by a carefully controlled flight load measurement program. The flight load measurement program shall demonstrate maximum and minimum loads for the entire flight envelope.
It shall also gather steady and cyclic load/strain data for use in the fatigue evaluation required by §§ 27.571 and 29.571. The parameters to be measured are primarily load calibrated strains supported by local strain measurement, accelerations, and deflections as necessary.
(2) Instrumentation.
The instrumentation system used in the flight parameter measurement program should accurately measure and record the critical parameters under operational test conditions. For critical maneuvers, the instrumentation should be capable of recording data sequences for several related channels for stationary and rotating channels. The location and distribution of the strain gauges should be based on a rational evaluation of the critical stress areas. Appropriate analytical methods should be used, such as Finite Element Modeling (FEM) and may be supplemented by other techniques including strain sensitive coatings, photoelastic methods, and thermography. Manual calculations based on the historical precedent of similar structure can also be very helpful. As much as possible, the instrumentation plan should standardize the gauge locations for each component so that all testing and other experience can be related to the same common set of measurement parameters. The gauge sensitivity, frequency response, location, distribution, and number of strain gauges must provide the strain/load spectrum and strain/load distribution for each part essential to the safe operation of the rotorcraft.
(її) The corresponding flight and ground operation parameters (airspeed, rotor RPM, center of gravity accelerations, etc.) should be recorded simultaneously and, where appropriate, as time histories. This is necessary to correlate the loads and stresses with the maneuver or operating condition during which they occurred. If the number of data parameters required exceeds the system capability, enough "carry-over" data channels should be included to reasonably relate all data for the specific maneuver when several flights are necessary.
(iii) The instrumentation system should be adequately calibrated and checked frequently during the strain survey. Ideally this would occur at the beginning AC Page MG 11 - 3