critical impacts (i.e., those causing the most serious damage but are least detectable).
When simulating accidental impact damage, blunt or sharp impactors should be selected to represent the maximum criticality versus detectability, according to the load conditions (e.g., tension, compression, or shear). Until sufficient service experience exists to make good engineering judgments on energy and impactor variables, impact surveys should consider a wide range of conceivable impacts, including runway or ground debris, hail, tool drops, and vehicle collisions. Service data collected over time can better define impact surveys and design criteria for subsequent products, as well as establish more rational inspection intervals and maintenance practice.
(6) Three Zones are defined by Figure AC 27 MG 8-2:
• Zone 1: Since the damage is not detectable, Ultimate Load capability is required.
The provisions of paragraph g(5) above provide a means of compliance.
• Zone 2: Since the damage can be detected at scheduled inspection, Limit Load (considered as Ultimate) capability is the minimum requirement for this damage.
• Zone 3: Since the damage is not detectable with the proposed in- service inspection procedures, ultimate load capability is required, unless an alternate procedure can show an equivalent level of safety. For example, residual strength lower than ultimate may be used in association with an improved inspection procedure.
(iii) FATIGUE TOLERANCE EVALUATION – MEANS OF COMPLIANCE.
One, or a combination of, the methods below should show compliance with the requirements of this section. The Flaw Tolerant Safe- Life Evaluation or the Fail-Safe Evaluation are to be used unless it can be shown that neither can be achieved within the limitations of geometry, inspectability, or good design practice. In that case, the Safe-Life Evaluation should be used. From current state-of-the-art with rotorcraft applications, it is widely admitted that composite materials have good flaw or damage tolerance capabilities and therefore the safe-life option is rarely necessary. Flaw Tolerance evaluations are best suited for most composite structures, particularly those with structural redundancy and inherent resistance to damage growth. Damage resulting from anomalous or accidental events must be considered in the Flaw Tolerant Safe-Life and Fail-Safe evaluations.
The fatigue substantiation should include sufficient coupon, element, sub-element, or component tests to establish the fatigue scatter, curve shapes, and the environmental effects. The substantiation should include full -scale testing but also may be accomplished by analysis supported by test evidence. When spectrum testing is used, the lowest load levels can be eliminated from the spectrum if they can be shown to be non-damaging. The substantiation should include a static strength evaluation to show that the required residual strength and adequate stiffness, accounting for the effects of environment, are retained for the life of the structure or the appropriate inspection Page MG 8 - 21
AC 27-1B, Chg 2 4/25/06 interval. Flaws and damage as determined in paragraph g(6)(ii) above for the specific structure being substantiated should be imposed at each critical area of the structure.
(A) Flaw Tolerant Safe-Life Evaluation. This is a “No -Growth” method in that it demonstrates that the structure, with flaws present, is able to withstand repeated loads of variable magnitude without detectable flaw growth for the life of the rotorcraft or within a specified replacement time. This fatigue evaluation may be used to substantiate any type of damage that will remain in-service for the life of the structure.
No specific inspection requirements are generated from the test program in this method.
However, routine inspections for cracking, delaminations, and service damage as outlined in § 27.1529 are always required. Compliance using full-scale, component, or sub-component fatigue testing can be accomplished by either of the following methods:
(1) `S-N Method. This method is based on determining the point where initiation of growth occurs for the flaws present at critical locations in the structure.
AC 27-1B, Chg 1, MG-11, provides guidance that can be appropriate for this method in composites. The method utilizes one or more full -scale, component, or sub-component test specimens subjected to constant-amplitude or spectrum loading applied in a distribution on the structure that is representative of critical flight conditions. Any indication of growth of the imposed flaws and defects, or structurally significant cracking, disbonding, splintering or delamination of the composite, defines the fatigue initiation characteristic of the structure in terms of applied load and cycles. Working S -N curves are established from the mean curve using strength or cycle reductions to account for fatigue scatter and environmental effects. Flight loads are compared to this working curve, and if any intercepts occur, a cumulative damage calculation is conducted to establish the component retirement time. Compliance with the ultimate load requirements should be demonstrated at the completion of the fatigue test.
(2) Life Test Method. This method uses spectrum fatigue testing to verify the absence of flaw growth over a large number of cycles that are equivalent to a lifetime of expected usage. The method uses one or more full-scale, component, or sub-component test specimens subjected to spectrum fatigue loading applied in a representative distribution of flight loads, including Ground-Air-Ground (GAG) loads.
Fatigue test loads should be increased by factors for environment and fatigue strength scatter. The load may also be increased using an S-N curve approach to reduce the duration of the test. Please reference "Certification Testing Methodology for Composite Structure", Report No. DOT/FAA/CT-86/39 for a discussion of the S-N approach. Any significant growth of the imposed flaws and defects, or structurally significant cracking, disbonding, splintering, or delamination of the composite during the test constitutes failure to achieve the desired lifetime. However, the equivalent life demonstrated at the time of inception of flaw growth or cracking can be used as a retirement time for the component. Compliance with the ultimate load requirements should be demonstrated at the completion of the fatigue test.
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(B) Fail-Safe (Residual Strength after Flaw Growth) Evaluation. This method demonstrates that the structure following a partial failure still has a sufficient residual strength capability within a specified inspection interval or the established retirement life of the component. If a retirement life is established, an ultimate design load capability is generally required while, if an inspection interval is determined, a limit load capability is the minimum acceptable residual strength capability that needs to be demonstrated. Full-scale, component, or sub-component testing should be accomplished using one or more specimens subjected to constant amplitude or spectrum loading applied in a manner representative of flight load conditions. The test loads should be increased by factors that account for environment and fatigue strength scatter. The results of the testing can be used to manage the structure in one of the three methods described below or a combination thereof.
(1) Fail-Safe, No Growth Evaluation. This approach is appropriate for inspectable in-service accidental damage. Structural details, elements, and subcomponents of critical structural areas, components, or full-scale structures, should be tested under repeated loads for validating a no-growth approach to the flaw tolerance requirements. The number of cycles applied to validate a no- growth concept should be statistically significant, and may be determined by load 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 assumed damage. The intent of this is to assure 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 in the Figure AC 27 MG 8-3. Once the damage is detected, the component is either repaired to restore ultimate load capability or replaced.
Da mage init iation Residual strength or occurrence Ultimate loads Limit loads ** No growth without repair is not acceptable Time No-growth approach ** Slow-growth approach * * Repair to Restore Ultimate St rength Figure AC 27 MG 8-3: Residual Strength vs. Time Page MG 8 - 23
AC 27-1B, Chg 2 4/25/06 The lower the residual strength caused by an accidental damage 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.
(2) Slow Growth Evaluation. This method is applicable when the flaw grows in the test and the growth rate is shown to be slow, stable, and predictable, as illustrated in Figure AC 27 MG 8-4. An inspection program should be developed consisting of the frequency, extent, and methods of inspection for inclusion in the maintenance plan. Inspection intervals should be established such that the damage will have a very high probability of detection between the time it becomes initially inspectable and the time at which the extent of the damage reduces the residual static strength to limit load (considered as ultimate), including the effects of environment. For any damage size that reduces the load capability below ultimate, the component is either repaired to restore ultimate load capability or replaced. Should functional impairment (such as unacceptable loss of stiffness) occur before the damage becomes otherwise critical, this should be accounted for in the development of the inspection program.
Damage Size Static Strength time Ultimate Limit Slow Growth Arrested Growth Slow Growth Critical Size Arrested Growth Slow Growth No Gr owth Threshold of Detectability Inspection Interval Basis Figure AC 27 MG 8-4: Illustration of Residual Strength and Damage Size Relationships for Fail-Safe Substantiation.
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(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 27 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 27 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 may be appropriate to establish both a retirement time and an inspection program for a given structure.
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AC 27-1B, 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) Locations where analysis shows high stress or low margins of safety.
(iii) Locations where permanent deformation occurred in static tests.
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(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 design 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.
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AC 27-1B, Chg 4 3/31/2014 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 § 27.571(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. Fatigue 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 paragraphs c. and d. of section AC 27 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 effects on loads as affected by elastomeric components are to be investigated. To account 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 § 27.1529.
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AC 27-1B, 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. T his 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 § 27.309, 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 §§ 27.241, 27.251, and 27.629 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 §§ 27.562 and 27.952.
(8) The eighth area is the special repair and continued airworthiness requirements of §§ 27.611, 27.1529, and 14 CFR Part 27 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).
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AC 27-1B, Chg 2 4/25/06 Page MG 8 - 32