(3) Arrested Growth Evaluation. This method is applicable when the flaw grows, but the growth is mechanically arrested or terminated before becoming critical (residual static strength reduced to limit load), as illustrated in Figure AC 29 MG 8-4.
Arrested Growth may occur due to design features such as a geometry change, reinforcement, thickness change, or a structural joint. This approach is appropriate for inspectable arrested growth damage. Structural details, elements, and sub- components of critical structural areas, components or full-scale structures, should be tested under repeated loads for validating an arrested growth approach to the flaw tolerance requirements. The number of cycles applied to validate an arrested growth concept should be statistically significant, and may be determined by load and/or life considerations. Residual strength testing or evaluation should be performed after repeated load cycling and demonstrate that the residual strength of the structure is equal to or greater than limit load considered as ultimate. Moreover, it should be shown that stiffness properties have not changed beyond acceptable levels. Inspection intervals should be established, considering the residual strength capability associated with the arrested growth damage. The intent of this is to ensure that structure is not exposed to an excessive period of time with static margins less than ultimate, providing a lower safety level than in the typical slow growth situation, as illustrated by Figure AC 29 MG 8-3. For any damage size that reduces load capability below ultimate, the component is either repaired to restore ultimate load capability or replaced.
The lower the residual strength caused by an arrested growth event, the shorter the inspection interval should be. Considerations of both inspectability and impact surveys (including probability of occurrence) for specific structure may be used to isolate the most critical threats to consider in setting a maintenance inspection interval. Knowledge of the residual strength for a given critical damage is also needed for such an evaluation. If it is known that the design is capable of handling large and clearly detectable damage, while maintaining a residual strength well above limit load, a less rigorous engineering approach may be applied in establishing the inspection interval.
(C) Safe-Life Evaluation. This method demonstrates that the structure, in an as-manufactured condition, is able to withstand repeated loads of variable magnitude without detectable cracks, disbonds, or delaminations for the life of the rotorcraft or within a specified retirement time. It is available for use only when both the Fail -Safe and Flaw Tolerant Safe-Life methods have been shown to be impractical due to considerations of geometry, inspectability, or good design practice. Further guidance for Safe-Life substantiation is provided in AC 27-1B, Chg 1, MG-11, “Fatigue Evaluation of Rotorcraft Structure”. The fatigue test articles should be fabricated and assembled in accordance with production specifications and processes so that they are representative of production structure including defects consistent with the limits established by manufacturing acceptance criteria.
(D) Combination of Safe Life and Fail Safe Evaluations. Generally it ma y be appropriate to establish both a retirement time and an inspection program for a given structure.
Page MG 8 - 25
AC 29-2C, Chg 2 4/25/06
(iv) Additional Considerations for FATIGUE AND FLAW TOLERANCE Evaluations.
(A) Experience with the application of methods of fatigue and flaw tolerance evaluations indicates that a relevant test background should exist in order to achieve the design objective. It is the general practice within industry to conduct flaw tolerance tests for design information and guidance purposes. It is crucial that the critical structure be identified and tested to the proper flight and ground loads. In the fatigue and flaw tolerance evaluation the following items must be considered:
(B) Identification of the structure to be considered in each evaluation (a failure mode and effects analysis or similar method should be used).
(1) Identification of Principal Structural Elements. Principal structural elements are those that contribute significantly to carrying flight and ground loads and whose failure could result in catastrophic failure of the rotorcraft. Typical examples of such elements are:
(i) Rotor blades and attachment fittings.
(ii) Rotor heads, including hubs, hinges, and some main rotor dampers.
(iii) Control system components subject to repeated loading, including control rods, servo structure, and swashplates.
(iv) Rotor supporting structure (lift path from airframe to rotor head).
(v) Fuselage, including stabilizers and auxiliary lifting surfaces.
(vi) Main fixed or retractable landing gear and fuselage attachment structure.
(2) Identification of Locations Within Principal Structural Elements to be Evaluated. The locations of damage to structure for damage tolerance evaluation can be determined by analysis or by fatigue test on complete structures or subcomponents.
However, tests will be necessary when the basis for analytical prediction is not reliable, such as for complex components. If less than the complete structure is tested, care should be taken to ensure that the internal loads and boundary conditions are valid.
The following should be considered:
(i) Strain gauge data on undamaged structure to establish points of high stress concentration as well as the magnitude of the concentration.
(ii) L ocations where analysis shows high stress or low margins of safety.
(iii) Locations where permanent deformation occurred in static tests.
Page MG 8 - 26
(iv) Locations of potential fatigue damage identified by fatigue analysis.
(v) Locations where the stresses in adjacent elements will be at a maximum with an element in the location failed.
(vi) Partial failure locations in an element where high stress concentrations are present in the remaining structure.
(vii) Locations where detection would be difficult.
(viii)Design details that are prone to fatigue or other damage indicated by service experience of similarly designed components.
(3) In addition, the areas of probable damage from sources such as a severe corrosive or fretting environment, a wear or galling environment, or a high maintenance environment should be determined from a review of the design and past service experience.
(C) The stresses and strains (steady and oscillatory) associated with all representative steady and maneuvering operating conditions expected in service.
(D) The frequency of occurrences of various flight conditions and the corresponding spectrum of loadings and stresses.
(E) The fatigue strength, fatigue crack propagation characteristics of the materials used and of the structure, and the residual strength of the damaged structure.
(F) Inspectability, inspection methods, and detectable flaw sizes.
(G) Variability of the measured stresses of paragraph g(6)(iv)(C) above, the actual flight condition occurrences of paragraph g(6)(iv)(D) above, and the fatigue strength material properties of paragraph g(6)(iv)(E).
(v) FLIGHT STRAIN MEASUREMENT PROGRAM.
(A) General. Subsequent to desi gn analysis, in which aircraft loads and associated stresses are derived, the stress level and loads are to be verified by a carefully controlled flight strain measurement program. (This guidance is similar to that of AC 27-1B, MG 11, Chg 1.)
(B) Instrumentation.
(1) The instrumentation system used in the flight strain measurement program should accurately measure and record the critical strains under test conditions associated with normal operation and specific maneuvers. The location and distribution of the strain gauges should be based on a rational evaluation of the critical stress areas.
Page MG 8 - 27
AC 29-2C, Chg 4 5/1/2014 This may be accomplished by appropriate analytical means supplemented, when deemed necessary, by strain sensitive coatings or photoelastic methods. The distribution and number of strain gauges should define the load spectrum adequately for each part essential to the safe operation of the rotorcraft as identified in § 29.571(a)(1)(i). Other devices such as accelerometers may be used as appropriate.
(2) The corresponding flight parameters (airspeed, rotor RPM, center of gravity accelerations, etc.) should also be recorded simultaneously by appropriate methods. This is necessary to correlate the loads and stresses with the maneuver or operating conditions at which they occurred.
(3) The instrumentation system should be adequately calibrated and checked periodically throughout the flight strain measurement program to ensure consistent and accurate results.
(C) Parts to be Strain-Gauged. Fa tigue critical portions of the rotor systems, control systems, landing gear, fuselage, and supporting structure for rotors, transmissions, and engine are to be strain-gauged. For rotorcraft of unusual or unique design, special consideration might be necessary to ensure that all the essential parts are evaluated.
(D) Flight Regimes and Conditions to be investigated.
(1) Typical flight and ground conditions to be investigated in the flight strain measurement program are given in paragraph c. and d. of section AC 29 MG 11 of this AC.
(2) The determination of flight conditions to be investigated in the flight strain measurement program should be based on the anticipated use of the rotorcraft and, if available, on past service records for similar designs. In any event, the flight conditions considered appropriate for the design and application should be representative of the actual operation in accordance with the rotorcraft flight manual. In the case of multiengine rotorcraft, the flight conditions concerning partial engine-out operation should be considered in addition to complete power-off operation. The flight conditions to be investigated should be submitted in connection with the flight evaluation program.
(3) The severity of the maneuvers investigated during the flight strain survey should be at least as severe as the maximum likely in service.
(4) All flight conditions considered appropriate for the particular design are to be investigated over the complete rotor speed, airspeed, center of gravity, altitude, and weight ranges to determine the most critical stress levels associated with each flight condition. The temperature effe cts on loads as affected by elastomeric components are to be investigated. To a ccount for data scatter and to determine the stress levels present, a sufficient amount of data points should be obtained at each Page MG 8 - 28
flight condition. Consideration can be given to the use of scatter factors in determining the sufficiency of data points. In some instances, the critical weight, center of gravity, and altitude ranges for the various maneuvers can be based on past experience with similar design. This procedure is acceptable where adequate flight tests are performed to substantiate such selections. The combinations of flight parameters that produce the most critical stress levels should be used in the fatigue evaluation.
(vi) FREQUENCY OF LOADING.
(A) Types of Operation.
(1) The probable types of operation (transport, utility, etc.) for the rotorcraft should be established. The type of operation can have a major influence on the loading environment. In the past, rotorcraft have been substantiated for the most critical general types of operation with some consideration of special, occasional types of operation. To assure that the most critical types of operation are considered, each major rotorcraft structural component should be substantiated for the most critical types of operation as established by the manufacturer. The types of operation shown below should be considered and, if applicable, used in the substantiation:
(i) Long flights to remote sites (low ground-air-ground cycles but high cruising speeds).
(ii) Typical, general types of operation.
(iii) Short flights as used in logging operations.
(2) One means is to substantiate for the most severe type of operation;
however, this method is not always economically feasible.
(3) A second means is to quantify the influence of mission type on fatigue damage by adding to or replacing hour limitations by flight cycle limitations (if properly defined and easily identifiable by the crew, for example: one landing, one load transportation). A special type of flight hour limitation replacement using factorization of flight hours for multiple types of operations may be feasible if continuing manufacturer’s technical support is provided and documented; i.e., the manufacturer either provides the factorization analyses or checks them on a continuing basis for each rotorcraft.
(4) Where one or more of the above operations are not among the general uses intended for the rotorcraft, the rotorcraft flight manual should state in the limitations section that the intended use of the rotorcraft does not include certain missions or repeated maneuvers (i.e., logging with its high number of takeoffs/landings per hour). A note to this effect should also appear in the rotorcraft airworthiness limitations section of the maintenance manual prepared in accordance with § 29.1529.
Page MG 8 - 29
AC 29-2C, Chg 2 4/25/06
(5) Should subsequent usage of the rotorcraft encompass a mission for which the original structural substantiation did not account, the effects of this new mission environment on the frequency of loading and structural substantiation should be addressed and where practicable, in the interest of safety, a reassessment made. If this reassessment indicates the necessity for revised retirement times, those new times may be limited to aircraft involved in the added mission provided.
(i) Proper part re-identification is established;
(ii) a Rotorcraft Flight Manual (RFM) supplement outlining limitations is approved;
(iii) an airworthiness limitations section supplement is approved; or
(iv) an appropriate combination of part re-identification, RFM supplement, or airworthiness limitation section supplement is approved.
(B) Loading Spectrum. The spectrum allocating percentage of time or frequencies of occurrence to flight conditions or maneuvers is to be based on the expected usage of the rotorcraft. This spectrum is to be such that it is unlikely that actual usage will subject the structure to damage beyond that associated with the spectrum. Considerations to be included in developing this spectrum should include prior knowledge based on flight history recorder data, design limitations established in compliance with § 29.309, and recommended operating conditions and limitations specified in the rotorcraft flight manual. The distribution of times at various forward flight speeds should reflect not only the relation of these speeds to VNE but also the recommended operating conditions in the rotorcraft flight manual that govern Vc or cruise speed. Where possible, it is desirable to conduct the flight strain- gauge program by simulating the usage as determined above, with continuous recording of stresses and loads, thus obtaining directly the stress and load spectra for structural elements.
(7) The seventh major area is the dynamic loading and response requirements of §§ 29.241, 29.251, and 29.629 for vibration and resonance frequency determination and separation for aeroelastic stability and stability margin determination for dynamically critical flight structure. Critical parts, locations, excitation modes, and separations are to be identified and substantiated. This substantiation should consist of analysis supported by tests and tests that account for repeated loading effects and environment exposure effects on critical properties, such as stiffness, mass, and damping. Initial stiffness, residual stiffness, proper critical frequency design, and structural damping are provided as necessary to prevent vibration, resonance, and flutter problems.
(i) All vibration and resonance critical composite structure are identified and properly substantiated.
(ii) All flutter -critical composite structures are identified and properly substantiated. This structure must be shown by analysis to be flutter free to 1.1 V NE (or Page MG 8 - 30
any other critical operating limit, such as VD, for a VSTOL aircraft) with the extent of damage for which residual strength and stiffness are demonstrated.
(iii) Where appropriate, crash impact dynamics considerations should be taken into account to ensure proper crash resistance and a proper level of occupant safety for an otherwise survivable impact. Please reference §§ 29.562 and 29.952.
(8) The eighth area is the special repair and continued airworthiness requirements of §§ 29.611, 29.1529, and 14 CFR Part 29 Appendix A for composite structures. When repair and continued airworthiness procedures are provided in service documents (including approved sections of the maintenance manual or instructions for continued airworthiness) the resulting repairs and maintenance provisions must be shown to provide structure that continually meets the guidance of paragraphs (1) through (7) of this AC paragraph. All certification based repair and continued airworthiness standards, limits, and inspections must be clearly stated and their provisions and limitations defined and documented to ensure continued airworthiness. No composite structural repair should be attempted that is beyond the scope of the applicable approved Structural Repair Manual (SRM) without an engineering design approval by a qualified FAA/AUTHORITY representative (DER or staff engineer).
Page MG 8 - 31
CHAPTER 3
AIRWORTHINESS STANDARDS
TRANSPORT CATEGORY ROTORCRAFT
MISCELLANEOUS GUIDANCE (MG) AC 29 MG 9 ROTORCRAFT ONE -ENGINE-INOPERATIVE POWER ASSURANCE
a. Purpose. The purpose of this document is to establish an approach for an engine power assurance procedure which will assure that the required OEI power level can be achieved.
b. General. The data and methods described herein are intended to be utilized as a guide and not necessarily the only means of achieving the desired result.
c. Applicability. The applicability of the document is intended to be prim arily in support of the new 30-second and 2-minute OEI rotorcraft engine rating scheme.
d. Partial Power Assurance (Engine “Run-Line”).
(1) Fundamental to the concept of limited-use one-engine-inoperative (OEI) ratings is the requirement to be certain that the rated OEI power will indeed be available when needed. Conventional periodic power -assurance and topping checks are impractical with the limited-use rating concept because of the rapid expenditure of useful life during exposure at the engine speeds and temperatures consistent with limited-use ratings; therefore, we require a means of assuring the power available, other than by actual demonstration on each service engine. The advent of more sophisticated controls and engine developments catering to the 30-second/2-minute OEI rating concepts can provide the means to determine: (1) that the thermodynamic/mechanical capability of the engine as tested at the prevailing ambient conditions, will permit reaching a specified power level at any other ambient condition and (2) the fuel system and the various limiters will not prevent the engine achieving OEI power on demand. Pending availability of these new methods, the “parallel run line check” approach is recommended.
(2) The method commonly called the “parallel run-line check” that has been in use for two decades may require refinement for application to the new rating structure where the degree of extrapolation to the OEI power level is more extensive and the slope of the individual engine characteristic is important. As in any power assurance method, success is strongly dependent on the validity of the data base, the maintenance of the engines and sensor/indicating systems, and the care taken during the conduct of the power check. In addition, trending of individual engine performance by the operator and associated analyses can be used to avoid unnecessary flight delays and engine removals.
Page MG 9 - 1
(3) Thermodynamic/mechanical capability can be addressed by test stand mapping of development engines over a range of ambient conditions to establish an adequate data base of engine characteristics. This will address characteristic slope variations between engines and establish correction factors necessary for extrapolation of data from a power assurance checkpoint to the 30-second OEI rating. Statistical verification and/or modification of the data base may be necessary during production by mapping of sample production engines. Performance data, at the 30 -second OEI condition, taken during the supplementary block test and also during the “overhaul test” will demonstrate the capability of an engine and its control system near the end of an overhaul period to produce the required power. This will demonstrate capability with a deteriorated base-performance engine.
(4) The question of fuel system limitations and other various limiters, which could prevent the engine from achieving OEI power on demand, may be addressed by use of more sophisticated control systems, for example, electronic controls utilizing several engine parameter limiters each with automatic datum reset capability. Such control systems can sense an engine failure and automatically reset the operating limiters upward from “normal” to “OEI” limits. Conventional flow and electronic bench testing can be used to verify the function and limit setting of the units when new or after overhaul or repair. The reset features can be extended in function to include a fixed magnitude pulldown type reset for use in verifying new and field production engine/control combination function ability. Pulldown type resets are currently in use today for verification of limiter settings on some engines and can be utilized in this application to avoid unneeded exposure of the engine to the rapid life expenditure conditions.
(5) While the above is envisioned as the probable means in which assurance of capability will occur early in the application of such engines, there will be other means developed. One such means would be utilization of modern electronic engine condition or health monitors to display “go” or “no go” conditions relative to the ability of the engine and its control system to produce 30-second OEI power if required. In this application the device would be a “power assurance meter” and could be used with electronic, hydro-mechanical, and pneumo-mechanical control systems. It is entirely reasonable to expect that self-taught or self-programmed power assurance meters can be used that continually program the actual performance slope of the subject engine and extrapolate to the 30-second OEI with continuous engine monitoring.
Self-programming occurs by sampling engine temperature, speed, torque, other characteristics (such as fuel pressure), and ambient conditions, resulting in the reflection of an actual characteristic for the installed engine. The availability of this information permits treating engines individually, whether it is a new or deteriorated engine or one with either minimum or maximum slope, without the necessary compromises to “best” engines that necessarily occurs using the earlier statistical approach. The question of instantaneous fuel system capacity could be addressed by fuel pump/control systems incorporating bypass systems equipped with flow meters.
The health monitor or power assurance meter can continually integrate the fuel flow increment available in terms of power increment required in the event of OEI and would Page MG 9 - 2
include this intelligence in its pass-fail judgment criteria. Systems of this type would further be conducive to in-service ground checks by overt by-pass deactivation from low power settings to assure satisfactory mechanical function.
(6) Power assurance for the limited-use OEI ratings depends on a complete understanding of the engine model’s operating characteristics. Two approaches have been discussed, one where, with the aid of a sophisticated fuel control system, the engine “learns” its own characteristics, and the other where the performance extrapolation is compared with a known minimum standard. The establish ment of the standard is obviously a vital part of the procedure, which depends to a large extent on the existence of a reliable data base. In a mature program this is relatively easy to maintain, since it is possible to use the new production engine acceptance data to establish engine-to-engine variation and also to test engines prior to overhaul to determine the effects of deterioration. Thus, an up- to-date minimum or worst-engine characteristic can be maintained and service engines would be compared with this minimum engine.
(7) When the engine in question is a completely new design, or a remote derivative of an existing design, establishing the initial data base presents some problems which must be resolved. New production engines will eventually es tablish engine-to-engine variation, but initially an estimated worst variation must be assumed.
The rate of deterioration and its impact on the base standard must be accounted for from the first engine delivered, yet it may be some time before an acceptable number of engines can be tested after service.
(8) A partial solution lies in the development and qualification cycle of the engine. A typical new -design program requires several development engines, of which more than half can be expected to be used for endurance or accelerated endurance testing. Furthermore, by the time certification is completed and production deliveries have commenced, these engines will normally have amassed several thousand hours of running usually to a schedule far more rigorous than normal service. The information gathered during these tests will provide the necessary data base for the assessment of in-service engines, and it can be progressively enlarged, and the derived data refined, as further production and service data are obtained.
e. Engine Considerations. This section describes the potential causes of an engine not delivering specification OEI power levels in spite of passing a parallel run-line power assurance check. Possible solutions are discussed in the contex t of one time use 30-second and 2-minute ratings.
(1) Fuel Flow.
(i) An engine may not achieve maximum power available or emergency rating because insufficient fuel is supplied. This condition has a number of possible causes:
Page MG 9 - 3
(A) Low acceleration schedule (B) Low maximum fuel stop (C) Low fuel pump output (D) Restrictions between the fuel control and the combustor
(ii) The proposed emergency ratings (OEI) may preclude the use of a topping check to uncover the above problems; therefore the following procedures are advanced which can be used either separately or in combination with other approved methods to assure that the required fuel flow is available.
(iii) During engine acceleration the fuel flow rate is considerably higher when compared with the normal steady state condition. This fact can be used to verify the availability of OEI fuel flow. The verification can be done by a direct measurement of fuel flow during an acceleration or derived indirectly from the engine acceleration rate. It is envisaged that the determination of fuel flow by these procedures should be done by some automatic means.
(iv) Figure AC 29 MG 9-1 is a bypass technique in which some of the fuel controls output is routed away from the engine and back to tank. This forces the fuel control onto the acceleration schedule in order to maintain gas generator speed. The design of the system should ensure that with the bypass flowing the fuel control outlet pressure and flow at the OEI ratings are simulated. The bypass system can be either permanently installed and operated in flight, (Failure Malfunction Effects Analysis must be provided), or as an item of ground test equipment. The quantity of fuel bypassed should be equivalent to the worst case difference between fuel flow at the 30-second rating and typical power assurance power levels. However, trend monitoring and service history may provide the basis of an alternative to periodic measurement.
(2) Limiters. A means must be provided to assure that a lower than required (for OEI power) limiter setting does not exist. Limiters that could prohibit reaching OEI power are as follows:
(i) Ng Limiter - (Maximum Compressor Speed Limiter or Governor)
(ii) Measured gas temperature limiter.
(iii) Output shaft torque limiter.
(iv) Np limiter or power turbine governor - (Power turbine governors can be verified at lower than OEI power conditions.)
(v) Fuel flow limiter or maximum fuel flow stop - (Fuel flow limiting has been addressed in previous paragraphs.)
Page MG 9 - 4
(3) Failure Modes and Effects Analysis. Failure modes and effects analysis, along with limited demonstration and suitable engine health monitoring procedures, may provide the basis of an acceptable solution to possible unexpected power limiting due to engine condition. It should be shown in the analysis that there is no probable event or combination of events which can cause a latent problem leading to inadequate fuel flow at high powers. The analysis should include all components of the fuel system such as:
pump(s), control system (mechanical, hydromechanical, electronic, etc.) pipework, filters, fuel nozzle(s), and electrical interfaces. It should also address the probable effects of accumulated running time, dirty fuel, and hostile environment.
(4) High Corrected Gas Producer Speed.
(i) The proposed OEI ratings will cause the engine to run at high corrected gas producer speeds (Ng/√θ). At high Ng/√θ, performance characteristics of components, especially in the compressor, can change significantly and to an extent which would change the extrapolation of low speed run line data.
(ii) In operation, the effects of the accretion of dirt, FOD, component deterioration, and erosion of blading may also cause changes in the high-speed performance of an engine.
(iii) The above effects must be considered when developing power assurance procedures and data.
(5) Special Devices.
(i) The satisfactory operation of devices or systems whose functioning is required in order to achieve the OEI powers should be verified. Devices or systems, which in normal operations are not exercised through the range of travel needed to achieve the OEI powers, may require special checks to assure adequate capability.
(ii) Special devices that are required only in order to achieve the OEI powers (for example, solenoids to provide additional cooling flow to hot-section components or a water/anti-freeze mixture into the compressor), should be subjected to periodic checks and have a demonstrated high reliability.
f. Airframe Considerations.
(1) Instrumentation Accuracy.
(i) The accuracy of any power assurance check is strongly dependent on the air data and engine parameters. SAE ARP 1217 (May 1979) provides guidance on the desired measurement accuracy for parameters used for engine health and diagnostic monitoring. The parameters to be considered with their respective functions include:
Page MG 9 - 5