(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
TRANSPORT CATEGORY ROTORCRAFT
MISCELLANEOUS GUIDANCE (MG) AC 29 MG 11 FATIGUE TOLERANCE EVALUATION OF TRANSPORT CATEGORY
ROTORCRAFT METALLIC STRUCTURE
a. Purpose. This advisory material provides an acceptable means of compliance with the provisions of § 29.571, Amendment 28, of the Federal Aviation Regulations (FAR) dealing with the fatigue tolerance evaluation of transport category rotorcraft metallic structure. For safe- life evaluations, AC 27-1B, MG 11 (Fatigue Evaluation of Rotorcraft Structure) provides in depth guidance on acceptable means of compliance.
The fatigue evaluation procedures outlined in paragraph AC 29 MG 11 are for guidance purposes only and are neither mandatory nor regulatory in nature. Specific issues related to substantiation of composite rotorcraft structures are addressed in AC 29 MG 8. 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. It is recommended that major deviations from the procedures be coordinated with the certifying regulatory authority to assure compliance with the regulatory requirements.
b. Special Considerations . The structure of rotorcraft is subject to cyclic stresses in practically every regime of flight. In addition, since rotorcraft are highly maneuverable and capable of forward, rearward, sideward, vertical, and rotational flight, operating limitations due to fatigue are possible in practically all flight situations. Corrosion and other environmental damage are also common in rotorcraft operations. For these reasons, special attention should be focused on the fatigue evaluation of rotorcraft structure.
Background.
(1) Fatigue substantiation of rotorcraft dynamic components was first implemented in the 1950’s by means of Safe-Life Methodology. Many advances in design and analytical methods have been made in the state-of-the-art and in industry practices since that time. To date, Safe- Life Methodology, as described in AC 27-1B, MG 11, is considered successful in providing a high level of reliability, but could be improved by taking into account the strength-reducing effects of damage likely to occur in manufacturing, maintenance, and in service, including corrosion, accidental damage, or manufacturing/maintenance flaws. The introduction of composites led the manufacturers/certifying authorities to take into account the specific static and fatigue strength-reducing effects of aging, temperature, moisture absorption, impact damage, and recognition of a minimum quality standard. In parallel, crack growth methodology in Page MG 11 - 1
29-2C, Chg 1 2/12/03 metal structure has been successfully used for solving short-term airworthiness problems in rotorcraft, and as the certification basis for civil and military transport aircraft applications.
(2) Recognizing that advances in state-of-the-art and industry practice warranted changes to the existing fatigue requirements in Part 29, the regulatory requirements of § 29.571 were substantially revised. The revision to § 29.571 requires new guidance material containing compliance provisions related to the changes.
AC 29 MG 11 provides this material with respect to the flaw tolerance requirements for metallic structure and is supplemented by AC 27-1B MG 11. Guidance material specifically addressing composite rotorcraft structure is provided in Chapter 3, MG 8 of this AC (1) and is supplemented by AC 20-107A.
d. Introduction .
(1) Definitions. These definitions are provided to define the terms used in applying the requirements specified in FAR 29.571, Amendment 28, and may differ from other definitions.
Fatigue Tolerance. The capability of structure to continue functioning without catastrophic failure after being subjected to fatigue (repeated) loads expected during operation of the rotorcraft. Fatigue tolerance should be achieved by flaw tolerance design, or if impractical, safe-life design, or a combination.
(ii) Safe Life. The capability of as -manufactured structure as shown by tests, or analysis based on tests, not to initiate fatigue cracks during the service life of the rotorcraft or before an established replacement time.
(iii) Flaw Tolerance. The capability of rotorcraft structure to achieve fatigue tolerance accounting for the presence of flaws and damage which may occur in manufacturing and service use. Flaw tolerance can be achieved by either flaw tolerance safe-life or fail-safe designs. The term ‘Damage Tolerance’ is frequently used to describe the ability of a structure to tolerate the effects of flaws and damage;
however, the terminology of FAR 29.571, Amendment 28, is used in this AC to maintain consistency.
(iv) Flaw tolerant Safe Life. The capability of as -manufactured structure with expected flaws as shown by tests, or analysis based on tests, not to initiate fatigue cracks during the service life of the rotorcraft or before an established replacement time.
(v) Fail-Safe. The capability of structure remaining after a partial failure to withstand design limit loads without catastrophic failure within an inspection period.
(vi) Multiple Load Path. Structure providing two or more separate and distinct paths of structure that will carry limit load after complete failure of one of the members.
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(vii) Active Multiple Load Path. Structure providing two or more load paths that are all loaded during operation to a similar load spectrum.
(viii) Passive Multiple Load Path. Structure providing load paths with one or more of the members (or areas of a member) relatively unloaded until failure of the other member or members.
(ix) Crack Arrest Feature. Structure that does not provide completely separate and distinct load paths but does provide features of design such as bonded and/or riveted straps, changes in geometry, or special processing techniques such as rolling or coining to retard or arrest crack growth.
(x) Slow Crack Growth Feature. Structure (single element or multiple element) that provides for slow crack growth by material selection, material processing, limitation of stress levels, geometrical design features, or by other methods.
(xi) Flaw. Intrinsic imperfec tions such as inclusions, forging laps, or porosity, and discrete damage such as gouges, scratches, nicks, corrosion, fretting, wear, or impact, that could be expected during manufacture or operation.
(xii) Design Limit Loads. The maximum loads to be expected in service, as defined by § 29.301(a).
(xiii) As Manufactured. Product and/or component that has passed the applicable quality control process and has been found to conform to the approved design within the allowable tolerances.
(xiv) Residual Strength. The strength retained for some period of unrepaired use after a failure or partial failure due to fatigue, corrosion, and accidental or discrete source of damage.
(xv) Principal Structural Element (PSE). A structural element that contributes significantly to the carrying of flight or ground loads and whose failure due to fatigue can lead to catastrophic failure of the rotorcraft.
(2) Rotorcraft Fatigue Tolerance. Fatigue tolerant design as substantiated by fail-safe flaw growth or flaw tolerant safe-life means outlined in § 29.571 and paragraph AC 29 MG 11g is required for all PSE’s, unless it entails such complications that an effective flaw tolerant structure cannot be achieved within the limitations of geometry, inspectability, or good design practice. Good design practice includes consideration of component complexity, component weight, methods of production, and component life cycle cost. Under these circumstances, a design that complies with safe-life criteria should be used. Typical examples of structure that might not be conducive to flaw tolerance designs are swashplates, main rotor shafts, push rods, small rotor head components (i.e., devices, bolts, etc.), landing gear, and gearbox Page MG 11 - 3
29-2C, Chg 1 2/12/03 internal parts, including bearings. In addition, the need for the use of inspection techniques and equipment or highly trained personnel--resources not available (for economic or other reasons) to the small operator or in remote areas of operation--should be carefully considered.
Test Background. Experience with the application of methods of fatigue evaluation 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. Flaw location and crack growth data based on test results and service history of similar parts, if available, should also be considered in establishing a recommended inspection program.
(4) Manufacturing and Maintenance Considerations. Assurance of structural adequacy also includes manufacturing, overhaul and repair, and service maintenance in accordance with design requirements, design specifications, maintenance procedures, overhaul and repair instructions, quality control to monitor compliance, and established manufacturing work processes including “frozen planning.” The fatigue tolerance substantiation should include an evaluation of the details of the specific work processes used on each component to determine the potential sensitivities.
(5) Fatigue Tolerance Considerations. In the fatigue tolerance evaluation, the following items should be considered:
(i) Identification of the structure to be considered in each evaluation (a failure mode and effects analysis or similar method should be used).
(ii) The stresses and strains (steady and oscillatory) associated with all representative steady and maneuvering operating conditions expected in service.
(iii) The frequency of occurrences of various flight conditions and the corresponding spectrum of loadings and stresses.
(iv) The fatigue strength, fatigue crack propagation characteristics of the materials used and of the structure, and the residual strength of the damaged structure.
(v) Inspectability, inspection methods, and detectable flaw sizes.
(vi) Variability of the measured stresses, the actual flight condition occurrences, and the fatigue strength material properties.
Flight Loads Measurement Program. See paragraph c. of AC 27- 1B MG 11.
f. Rotorcraft Usage Spectrum. See paragraph d of AC 27- 1B MG 11.
g. Fatigue Tolerance Evaluation.
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(1) General. A means should be established using the Safe- Life approach or a Flaw Tolerant approach to control the airworthiness of principle structural elements identified under § 29.571(a)(1)(i) While the safe- life approach is acceptable under certain circumstances as defined in § 29.571, the Flaw Tolerant safe Life and Fail-Safe (residual strength after crack/flaw growth) approaches are to be used unless shown to be impractical as stated in § 29.571. A Flaw Tolerance evaluation of structure is intended to ensure that even when expected flaws are present, the structure will withstand service loads without failure until the flawed parts are replaced or until the flaws (or resulting fatigue cracks) are detected and appropriate action taken. The Flaw Tolerant evaluation may be achieved by either Flaw Tolerant Safe Life, Fail Safe (residual strength after crack/flaw growth), or a combination thereof. Flaw Tolerant Safe Life includes the analyses and/or testing currently associated with safe-life substantiation, plus consideration of flaws to establish a replacement time, or a safe life greater than the service life of the rotorcraft. Crack growth methods include the analyses and/or testing currently associated with a Damage Tolerance Assessment (DTA) to establish an inspection program. Design features that should be used in attaining a fatigue tolerant structure are:
(i) Use of multi-path construction and the provision of crack stoppers to limit the growth of cracks and to provide adequate residual strength.
(ii) Selection of materials and stress levels that preclude crack growth or crack initiation from flaws or that provide a controlled slow rate of crack propagation combined with high residual strength after initiation of cracks. Tests are required to substantiate crack propagation rates.
(iii) Design to permit detection of cracks and other flaws, including the use of crack detection systems, in all critical structural elements before cracks can propagate and become dangerous or result in appreciable strength loss and to permit replacement or repair.
(iv) Use of multiple element structures may be provided so that damage or failure occurring in one element of the member will be confined to that element and the remaining structure will still possess adequate load-carrying ability until the failed element is discovered by inspection.
(v) Provisions to limit the probability of concurrent multiple damage, particularly after long service, should be provided. These provisions should ensure adequate independence of each failure mode of multi-path constructions. The use of full-scale fatigue test articles are recommended in this evaluation.
(vi) Identification of Principal Structural Elements. Typical examples of such elements are:
(A) Rotor blades, attachment fittings, and dynamic systems.
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29-2C, Chg 1 2/12/03 (B) Rotor heads, including hubs, hinges, some main rotor dampers, and support structures.
(C) Control system components subject to repeated loading, including control rods, servo structure, and swashplates.
(D) Rotor supporting structure (torque and lift path from airframe to rotor head).
(E) Fatigue critical fuselage structures, including stabilizers and auxiliary lifting surfaces.
(F) Main fixed or retractable landing gear and fuselage attachment structure.
(G) Gearboxes, driveshafts, and couplings.
(vii) Identification of Locations within Principal Structural Elements to be Evaluated. The locations to be considered for fatigue tolerance evaluation can be determined by analysis or by fatigue test on complete structures or subcomponents.
(A) The following should be considered:
(1) Strain gauge data on undamaged structure to establish points of high stress concentration as well as the magnitude of the concentration;
(2) Locations where analysis shows high stress or low margins of safety;
(3) Locations where permanent deformation occurred in static tests;
(4) Locations of potential fatigue damage identified by fatigue analysis;
(5) Locations where the stresses in adjacent elements will be at a maximum with an element in the location failed;
(6) Partial fracture locations in an element where high stress concentrations are present in the residual structure;
(7) Locations where detection would be difficult;
(8) Design details or similarly designed components that are prone to fatigue or other damage as shown by service experience; and,
(9) Components fabricated from materials of potentially low fracture toughness or high flaw growth rate.
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(B) In addition, the areas of probable damage from sources such as a severe corrosive and/or fretting environment, a wear and/or galling environment, or a high maintenance environment should be determined from a review of the design and past service experience.
(viii) Extent of Flaws. Each particular design should be assessed to establish appropriate damage criteria in relation to inspectability and flaw extension characteristics. In any flaw determination, it is possible to establish the extent of flaws in terms of detectability with the inspection techniques to be used, the associated single element failure or initially detectable flaw size, the residual strength capabilities of the structure, and the likely flaw extension rate (after either an element failure or a partial failure) considering the expected stress redistribution under the repeated loads expected in service and with the expected inspection frequency.
(ix) Provisions for Inspection. The designer should strive to ensure adequate inspectability of all structural parts to qualify them under the fail-safe crack growth provisions. In those cases where blind areas or surfaces exist, suitable design features should be provided to allow inspection techniques (either visual or nondestructive testing, as necessary) to assure adequate residual strength is achieved unless shown to be impractical due to limitations of geometry and good design practice.
In addition, the alternate safe-life approach to fatigue tolerance should be implemented if the inspection techniques are shown to be too complicated and impractical.
NOTE: Removed Figure AC 29 MG 11- 1, "S-N or P-N Curve Usage," AC 29-2C, dated 9/30/99.
(x) Testing of Principal Structural Elements. 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. Test on complete structures are recommended as baseline for design validation and demonstration of compliance with the requirements. The nature and extent of tests on complete structures or on portions of the primary structure will depend upon:
(A) applicable previous design, construction and tests.
(B) service experience in connection with similar structures.
(C) extent and validation of the structural analysis (FEM models, design manuals, stress or strain measurements, …).
(D) conservatism and margin of safety of the validation, included assumptions on the loading conditions.
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29-2C, Chg 1 2/12/03
(2) Flaw Tolerant Safe Life Demonstration
(i) Flaw tolerant safe-life substantiations provide a safe period of operation of structure with expected flaws with only routine inspections necessary.
(ii) Designs, processes, materials and stress levels that preclude the occurrence of flaws, or that provide a tolerance to crack initiation from flaws, can be used to achieve a Flaw-Tolerant Safe Life structure.
(iii) Flaw-Tolerant Safe Life uses analyses and/or testing similar to that of Safe Life, except that structures with expected flaws are tested rather than asmanufactured structures. Tests will be necessary when the basis for analytical prediction may not be reliable, such as for complex components.
(iv) Evaluation of Flaw Types and Sizes (“Threat Assessment”).
(A) For each zone of the component, a systematic evaluation of the types and sizes of flaws should be conducted, based on the processes and practices used in design, manufacturing, maintenance, and overhaul, and service experience.
(B) Flaw types. The types of flaws considered should include intrinsic/discrete flaws, impacts, scratches, corrosion, fretting and wear. Consideration should also be given to factors that reduce scatter and deviations from nominal structures, such as "frozen processes", Flight Critical Parts programs, material selection to mitigate intrinsic flaws (inclusions and defects), and procedures to reduce manufacturing deviations. Implementation of a specific manufacturing inspection process can justify the elimination of some flaws from consideration if that process can be shown to be highly reliable, well-controlled and documented, and systematically required. Where surface treatments, protective coatings, or shields are used, these should be considered when establishing the likely location and type of the flaw.
(C) Flaw size. The flaw sizes to be considered should be representative of those which are likely to be encountered during the structure’s service life resulting from the manufacturing, maintenance, and service environment. An analysis may be used combining the distribution of likely flaw sizes, the criticality of location and orientation, and the likelihood of remaining in place for a significant period of time.
(v) Determination of Retirement Time.
(A) Flaw Characterization. S/N curve shapes and scatter factors should be defined for the specific materials selected for the Flaw Tolerant Safe Life structure.
This may be accomplished by testing a number of small specimens (“coupons”) or by reference to existing characterizations of the selected material. In addition, the effect of the flaws identified in paragraph g(2)(iv) above, on the S/N curve shape and basic fatigue strength of the selected material should be determined. This may be accomplished by testing a number of coupons incorporating the representative flaws, or Page MG 11 - 8
by reference to existing characterizations of the same material and flaws. A coupon program incorporating representative manufacturing processes and relevant design features may also be appropriate to define “equivalent” flaws that produce the same strength-reducing effect as the representative flaws, but can be more easily applied and controlled.
(B) Flawed Full-Scale Specimens. In order to determine the mean fatigue strength considering flaws, a number of flawed full-scale specimens should be tested.
Representative flaws should be imposed on these specimens in each critical location on the structure where flaws are likely to occur. Equivalent flaws may be used if they have the same or a more severe strength-reducing effect, as determined by a flawed coupon test program or by experience with similar applications.
(C) Mean Strength Determination. Conventional stress vs. number of cycles (S/N) or spectrum safe-life fatigue testing may be performed on flawed full-scale specimens to establish the mean fatigue strength. AC 27- 1B MG 11 provides general guidance in the conduct of safe-life fatigue testing and the establishment of mean fatigue curves from the test data. The strength of flawed structure may also be derived from un-flawed structure test results by imposing the reductions in strength determined in paragraph g(2)(v)(A) above, or derived from experience with similar structure.
(D) Working S/N curve. Reduction factors should be applied to the mean curve in deriving a working S/N curve. These factors should include consideration of the number of specimens tested, variability (scatter), previous test data on the same materials or similar structures, as well as service experience. To preclude a dual penalty situation, reduction by two standard deviations rather than three (conventional safe life) may be used if justified by appropriate design features such as multiple elements or unmistakable flaw indications or by material properties that provide benign types of failure modes. Where new materials or designs are being evaluated, it is recommended that a larger reduction factor be used until additional test data justifying a change are available.
testing followed by analysis. (E) Replacement Time. Utilize the working S/N curve and loading spectrum of paragraph AC 29 MG 11f in substantiating a replacement time for the flawed structure by cumulative damage means. The replacement time established should be included in the airworthiness limitations section of the document established under § 29.1529.
(3) Fail-Safe Demonstration
(i) Safe Crack Growth. Safe crack growth (fracture mechanics) substantiation should show that the damage growth rate under the repeated loads expected in service (between the time at which the damage becomes initially detectable and the time at which the extent of damage reaches the value for residual strength evaluation) provides a practical basis for development of the inspection programs.
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29-2C, Chg 1 2/12/03
(ii) Inspection methods The designer should strive to ensure adequate inspectability of all structural parts to qualify them under the fail-safe flaw growth provisions. Inspection means that are appropriate for safe crack growth design follow:
(A) Routine Inspections. To support routine inspection programs, blind areas should be avoided, where practical.
(B) Special Inspections. These inspections will generally result from test results as well as the geometry of the design. Care should be given to special inspection techniques to be used in the field. Inspection techniques requiring facilities and resources beyond the capability of the small operator or not generally available in remote-area operations traditionally associated with rotorcraft operations should not be specified for field inspections. Conservative sizes for detectable cracks or other flaws should be used. Sufficient inspection intervals should be provided to detect cracks before they grow from a detectable size to a size that reduces the remaining strength below design limit strength.
(C) Pressurized Chambers. This design feature may be used to detect cracks that cause a chamber to lose its pressure (either positive or negative). Gauges can indicate the loss of pressure, or dye may be used if it is shown to be a dependable indicator.
(D) Vibration Generation. This characteristic should be considered both from the aspect of vibrations giving indications of a failure, and from the aspect of the increased fatigue loading resulting from the vibrations.
(E) Noise Generation. If initial failure will result in a clear and unmistakable noise that is sufficiently continuous and loud, this characteristic can be used in achieving flaw tolerance without additional special inspections.
(F) Crack Detection Wire, Foil, etc. Detection wire may be used in areas that are sufficiently well defined so that the wire can be properly located. This technique is appropriate in areas otherwise difficult to inspect.
(iii) Multiple Elements . Use of multiple element structures may be provided so that damage or failure occurring in one element of the member will be confined to that element and the remaining structure will still possess adequate load-carrying ability until the failed element is discovered by inspection. These provisions should be designed to provide adequate independence of each failure mode of multi-path constructions. The use of full -scale fatigue test articles is recommended in this evaluation. Examples of concurrent multiple damage to be avoided are:
(A) Simultaneous failure or partial failure of multiple path discrete elements working at similar stress levels.
(B) Failures or partial failures, in adjacent areas, due to redistribution of loading, following a failure of a single element.
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(iv) Partial Failures. The following are typical examples of the type of partial failures that may be considered in the flaw growth fail-safe evaluation:
(A) Detectable skin cracks in the trailing edge sections of rotor blades.
(B) Detectable failures of individual straps in “strap packs.”
(C) Detectable skin cracks emanating from the edge of structural openings or cutouts.
(D) Detectable circumferential or longitudinal skin crack in the basic fuselage or tail boom structure.
(E) Complete severance of interior frame elements or stiffeners in addition to a detectable crack in the adjacent skin.
(F) Presence of a detectable fatigue failure in at least the tension portion of the spar web or similar element.
(G) Detectable failure of a primary attachment, including blade attachment fittings and control surface hinge and fittings.
(H) Fretting, corrosion, and galling conditions expected in service.
(v) Initial inspection time. The initial crack size to be used for evaluati on of the initial inspection time should be the flaw size controlled by manufacturing quality and damage that can go undetected for the life of the part.
(vi) Evaluation of the inspection intervals. For multiple load paths, a minimum of three inspection intervals is recommended between the initially detectable damage time and the time when residual strength is reduced to design limit load by crack growth. For single element structures, a minimum of four inspection intervals is recommended. The repeated loads should be defined in the loading, temperature, and humidity spectra. The loading conditions should take into account the effects of structural flexibility and rate of loading where it is significant.
Tests of two or more specimens should be used to obtain crack propagation data using either a realistic load spectrum or an accelerated load (spectrum or single) associated with the use of propagation theory and data after cracks have been initiated. As far as applicable, crack growth tests should be carried out considering a fatigue crack naturally induced in testing. In the other cases, crack initiation points should be made by suitable methods, reference paragraph g(3)(viii). Unless a more rational method with an equivalent level-of-safety is applied for, the following methods of setting inspection intervals should be applied. In all cases, the inspection methods and intervals should Page MG 11 - 11