Page C - 97 AC 29.571A. §29.571 (Amendment 29-28) FATIGUE EVALUATION OF STRUCTURE.
a. Explanation . Amendment 29-28 adds a requirement to substantiate tolerance to flaws during the fatigue evaluation of structure. A flaw tolerant safe-life evaluation or a fail-safe (residual strength after flaw growth) evaluation is required by § 29.571(b) unless “the applicant establishes that these fatigue flaw tolerant methods for a particular structure cannot be achieved within the limitations of geometry, inspectability, and good design practices.”
b. Procedures .
(1) AC 29 MG 11 provides acceptabl e general procedures for complying with Amendment 29-28.
(2) Specific rotorcraft drive system gear fatigue evaluation procedures, which supplement Appendix 1, follow:
(i) Fatigue test evidence is necessary for the fatigue evaluation of gears.
The test evidence should be provided by rotating tests of complete gearbox specimens operating under power. The tests provide the basis for analysis leading to the establishment of safe-life.
(ii) The tests are conducted spec ifically for the purpose of gear tooth evaluation, and components subjected to the tests do not have to be considered serviceable on completion of the test. Excessive wear on bearings and shafts and marking (including spalling) of bearings and gear teeth are acceptable provided no fatigue damage is evident on the gear teeth. However, fatigue damage other than tooth fatigue should be considered for test validity and the integrity of the affected part confirmed as necessary.
(iii) The test conditions (torque versus number of cycles) should permit the setting of mean strength curve(s) to be associated with each primary gear in the drive train. The minimum test condition should encompass those power levels for which repeated application in service is expected under normal conditions. The S-n curve(s), for the material and type of gear, should be reduced by a factor of safety to take into account material and manufacturing variability. The factored curve will then be used in conjunction with the flight power spectrum to determine a life (limited or unlimited) for the gears in the primary drive system.
(iv) Special procedures, which do not affect fatigue evaluation of the gear teeth, may be allowed to facilitate completion of the test provided they have been justified and they do not affect life determination. These include periodic interruption for inspection, replacement of non-critical parts and the use of special lubricants, special cooling systems, and methods to prevent unrepresentative deflections at the test torque levels.
AC 29-2C, Chg 1 2/12/03 Page C - 98
(v) From evidence in relation to the strength of steel gears of conventional design, it is accepted that adequate fatigue strength can be demonstrated by the use of the above safety factor of 1.4 for a single test, 1.35 for two tests, 1.32 for three tests, and 1.3 for four or more tests. Where several tests are to be conducted, specimens should be selected from different manufacturing batches if practicable.
(vi) Demonstration of infinite lif e for gear teeth will normally require tests of a minimum of 107 cycles duration at factored power levels. Use of shorter duration tests should be justified.
Page D - 1 CHAPTER 2. PART 29
AIRWORTHINESS STANDARDS
TRANSPORT CATEGORY ROTORCRAFT
SUBPART D - DESIGN AND CONSTRUCTION
DESIGN AND CONSTRUCTION - GENERAL
AC 29.601. § 29.601 DESIGN .
a. Explanation .
(1) This rule requires that no design f eatures or details be used that experience has shown to be hazardous or unreliable.
(2) Further, the rule requires that the suitability of each questionable design detail and part must be established by tests.
b. Procedures .
(1) This rule is met partially by a review of service history of earlier model rotorcraft, or for a new model, review of service experience of models with similar design features. Specifically, this rule covers “features or details” such as the following:
(i) Seat track-to-seat interface fittings should have adequate locking devices to prevent both premature structural failure and premature unlatching.
(ii) Seat belt and harness shoul d be of a type and construction that service experience has shown to be easy to don, unlatch, and remove. They should also be of a type that is reliable, does not interfere with egress, and does not sustain unnecessary wear and tear under normal operations.
(iii) Metallic parts less than a certain thickness gauge and composite materials less than a certain number of plies should not be used. The minimum thickness and number of plies should be based to a large degree on service experience (normal wear and tear) with similar designs.
(2) The effects of service wear on t he loading of critical components should be considered. Flight testing, ground testing, and analyses may be used in these considerations.
(3) Tests are required for details and parts which the applicant chooses to use after questions have arisen concerning their suitability.
AC 29-2C, Chg 2 4/25/06 Page D - 2 AC 29.602 § 29.602 CRITICAL PARTS .
a. Explanation .
(1) Critical parts requirements apply to structural components, rotor drive systems, rotors, and mechanical control systems.
(2) The objective of identifying critical parts is to ensure that critical parts are controlled during design, manufacture, and throughout their service life so that the risk of failure in service is minimized by ensuring that the critical parts maintain the critical characteristics on which certification is based.
(3) Definitions with respect to § 29.602:
(i) The use of the word “could” in paragraph 29.602(a) of the rule means that this failure assessment should consider the effect of flight regime (i.e., forward flight, hover, etc.). The operational environment need not be considered.
(ii) With respect to this rule, the term “catastrophic” means the inability to conduct an autorotation to a safe landing, without exceptional piloting skills, assuming a suitable landing surface is available.
(iii) The use of the word “and” in paragraph 29.602(a) of the rule means the part must have both a catastrophic failure mode together with one or more critical characteristics.
(iv) With respect to this rule, t he term “part” means one piece, or two or more pieces permanently joined together.
(v) With respect to this rule, t he term “critical characteristic” means any dimension, tolerance, finish, material, or any manufacturing or inspection process, or other feature which cannot tolerate variation from type design requirements and, if nonconforming, would cause failure of the critical part.
(4) Many rotorcraft manufacturers alr eady have procedures in place within their companies for handling “critical parts.” These plans may be required by their dealings with other customers, frequently military (e.g., US DoD, UK MoD, Italian MoD).
Although these plans may have slightly different definitions of “critical parts” which have sometimes been called “Flight Safety Parts,” “Critical Parts,” “Vital Parts,” or “Identifiable Parts,” they have in the past been accepted as meeting the intent of this requirement and providing the expected level of safety. It is acceptable for these plans to use alternative names and terminology provided they meet the intent of this requirement.
b. Procedures . The rotorcraft manufacturer should establish a Critical Parts Plan that identifies and controls the critical characteristics. The policies and procedures which constitute that plan should be such as to ensure that--
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(1) All critical parts of the rotorcraft are identified by means of an appropriate failure assessment and a Critical Parts List is established.
(2) Documentation draws the attention of the personnel involved in the design, manufacture, maintenance, inspection, and overhaul of a critical part to the special nature of the part and details the relevant special instructions. For example all drawings, work sheets, inspection documents, etc., could be prominently annotated with the words "Critical Part" or equivalent and the Instructions for Continued Airworthiness and Overhaul Manuals (if applicable) should clearly identify critical parts and include the needed maintenance and overhaul instructions. The documentation should:
(i) Contain comprehensive instructions for the maintenance, inspection and overhaul of critical parts and emphasize the importance of these special procedures;
(ii) Indicate to operators and overhaulers that unauthorized repairs or modifications to critical parts may have hazardous consequences;
(iii) Emphasize the need for careful handling and protection against damage or corrosion during maintenance, overhaul, storage, and transportation and accurate recording and control of service life (if applicable);
(iv) Require notification of the manufacturer of any unusual wear or deterioration of critical parts and the return of affected parts for investigation when appropriate;
(3) Procedures should be established for identifying and controlling critical characteristics.
(4) To the extent needed for control of critical characteristics, procedures and processes for manufacturing critical parts (including test articles) are defined (for example material source, forging procedures, machining operations and sequence, inspection techniques, and acceptance and rejection criteria). Procedures for changing these manufacturing procedures should also be established.
( 5) Any changes to the manufacturing procedures, to the design of a critical part, to the approved operating environment, or to the design loading spectrum are evaluated to establish the effects, if any, on the fatigue evaluation of the part.
( 6) Materials review procedures for critical parts (i.e., procedures for determining the disposition of parts having manufacturing errors or material flaws) are in accordance with paragraphs (4) and (5) above.
( 7) Critical parts are identified as required, and relevant records relating to the identification are maintained such that it is possible to establish the manufacturing history of the individual parts or batches of parts.
AC 29-2C, Chg 2 4/25/06 Page D - 4 ( 8) The critical characteristics of critical parts produced in whole or in part by suppliers are maintained.
AC 29.603. § 29.603 (Amendm ent 29-17) MATERIALS.
a. Explanation . The rule requires that the suitability and durability of materials, the failure of which could adversely affect safety, must be determined by three-fold considerations:
(1) Considerations bas ed on experience or tests.
(2) Meeting approved specifications.
(3) Taking into account environmental conditions such as temperature and humidity.
b. Procedures .
(1) Experience may be used to show a material’s resistance to wear and deterioration from environmental effects (environmental effects include both natural environmental effects such as exposure to sunlight, water, salt spray, etc., and installation environmental effects such as exposure to fuel, hydraulic fluids, deicing fluids, etc.). Installation environmental effects should consider both direct exposure contact and expected migration of potentially deleterious fluids and compounds.
Testing for environmental effects may use either coupon testing, full-scale testing, or a combination. A combination of testing and experience may also be used.
(i) MIL-HDBK’s-5, -17, and - 23 include consideration of some environmental effects and contain reference to additional methods of testing for environmental effects.
(ii) The use of AC 20-107A, Composite Aircraft Structure, is recommended for environmental and damage tolerance considerations of advanced composite materials. (Also see paragraph AC 29 MG 8.)
(iii) The effects of excessive w ear and delamination of elastomeric and self-lubricated bearings used in critical load carrying applications in relation to redistribution of loading should be considered.
(2) Where possible, materials that meet widely accepted specifications such as AISI, SAE, MIL, or AMS and alloys which have favorable experience or tests should be used. Where company-developed materials are used, approved specifications are required to ensure the developed properties are duplicated in each lot of material. Raw material quality control is defined in FAA Order N8020-11 which is scheduled to be integrated into a forthcoming advisory circular. Documented specification usage is necessary to maintain quality assurance of materials.
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(3) Section 29.613 concerns strengt h properties and design values. (See paragraph AC 29.613.)
AC 29.605. § 29.605 ( Amendment 29-17) FABRICATION METHODS.
a. Explanation . The basic requirement of this rule is that the methods of fabrication must produce sound structure and produce it consistently.
(1) A process specification is requir ed for fabrication processes requiring close control.
(2) A test program is explicitly requi red for each new aircraft fabrication method.
b. Procedures .
(1) The approved specifications required by this rule may either be established government/industry specifications such as MIL, AISI, ASTM, or SAE, or the specifications may be company-developed proprietary specifications. Sufficient data should be provided to the FAA/AUTHORITY aircraft engineering offices to show that the desired features are provided by the process specification. In addition, sufficient process controls, inspections, and tests should be coordinated with FAA/AUTHORITY manufacturing inspection personnel to ensure that continued quality of the process is provided.
(2) In addition to the examples given by the rule; i.e., gluing, spot welding, and heat treating process, specifications should also be prepared for types of welding other than spot welding, for platings of metals, for protective finishes (other than decorative), for sealing, and for unique fabrication methods such as those used for composite materials.
(3) The required test programs should c onsider static strength effects, fatigue strength effects, and environmental effects as appropriate to the processes.
(4) During the fabrication of advanced composite materials, the effects of fabrication anomalies (i.e., disbonds, voids, porosity) should be considered. Special nondestruct testing inspection techniques and procedures should be developed to cover fabrication with allowable anomalies and permitted repair procedures. (See also paragraph AC 29 MG 8.)
AC 29.607. § 29.607 (Amendment 29-5) FASTENERS .
a. Explanation . Section 29.607 of Amendment 29-5 requires dual locking removable fasteners in critical locations. A nonfriction locking device is specifically required in any bolt subject to rotation, as stated in the rules.
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b. Procedures . Advisory Circular 20-71, Dual Locking Devices or Fasteners, December 8, 1970, contains information, procedures, and means of complying with § 29.607 of Amendment 29-5.
AC 29.609. § 29.609 PROTECTION OF STRUCTURE .
a. Explanation . The structure should be suitably protected as specified in the rule to maintain its design strength. Ventilation and drainage provisions must be provided as specified in the rule. Overboard drains should be furnished for corrosive or waste liquids. Drains for flammable fluids are specified in other rules such as §§ 29.999 and
29.1187.
b. Procedures .
(1) The structure may be preserved, pai nted, or treated with chemical films to protect it from strength deterioration. An approved process specification should be used for these types of treatments.
(2) Parts may be plated or chemically tr eated, such as anodized, for protection.
An evaluation and substantiation may be required to assure the structure or parts are not adversely affected during, or as a result of, the plating or treatment process.
(§ 29.605 concerns approval of process specifications and fabrication methods.)
(3) Plating or material surface har dness or composition changes may require fatigue substantiation to assure the fatigue strength is not altered or is otherwise properly assessed. An approved process specification should be used for these types of treatments.
(4) To prevent water accumulation, drain holes should be placed at possible dams such as bulkheads, and at low points in the fuselage and in the stabilizing surfaces.
(5) Control tubes and tubes used as primary mount structures (i.e., transmission support structure and engine mount structure) should be designed to prevent entry and collection of corrosive fluids or vapor, including water.
(i) A closed insert in each tube end may be used.
(ii) A sealant applied around the tube ends and around each rivet head may be used.
(6) Overboard drains should discharge clear of the entire rotorcraft. Dyed water discharged in flight, may be used to assure fluids are properly drained.
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(7) Welded tubes should be flushed and sealed after welding in accordance with an approved process specification.
(8) Refer to AC 43-4, “Corrosion Control for Aircraft,“ for further procedures.
AC 29.610. § 29.610 (Amendment 29-40) LIGHTNING AND STATIC ELECTRICITY PROTECTION .
a. Background . During the initial development and promulgation of the standards concerning the airworthiness of rotorcraft, it was not necessary to specify design features that would protect the rotorcraft from the meteorological phenomenon of lightning. This was due, in part, to the fact that rotorcraft were primarily operated in a VFR and nonicing environment. Also, a prudent pilot avoided thunderstorms where the possibility of encountering severe weather and a lightning strike was much greater. The construction, design, and operating environment of civil rotorcraft have changed markedly within the past two decades. Many rotorcraft are now authorized to fly IFR in all types of weather environment. One transport design has been approved for flight into known icing conditions. Additionally, many rotorcraft now use the same advanced technologies in structures and systems as do airplanes. Because of these facts, a specific rule on lightning protection of rotorcraft was adopted in Amendment 29-24. For further information, see the preamble of Amendment 29-24 (49 FR 44437; 11/6/84), Proposal 2-14. Section 29.610 is similar to § 25.581 which applies to the protection of structures of transport airplanes. However, the standard provides for specific protection of the aircraft structures as well as the systems of the rotorcraft. In addition, the protection of fuel systems from the effects of lightning is found and referenced in Report DOT/FAA/CT-83/3, the applicable version of Users Manual for AC 20-53, Protection of Airplane Fuel Systems Against Fuel Vapor Ignition Due to Lightning.
b. Explanation .
(1) The regulation requires that the rotorcraft must be protected against the catastrophic effects of lightning. This means that a lightning strike encounter should not prevent the continued safe flight and landing of the rotorcraft.
(2) Paragraph AC 29.1309 addresses t he protection required for systems.
Protection of the rotorcraft structures may be accomplished in a similar fashion.
(3) The structural components of the rotorcraft should be designed in such a manner that the lightning current may be safely diverted or conducted through the rotorcraft without damaging any critical structure or without causing damage to noncritical structure, the failure of which would preclude the continued safe flight and landing of the rotorcraft. A radome or fin cap which explodes due to a lightning strike and results in catastrophic damage to main or tail rotors is a scenario of lightning damage to a noncritical structure which has catastrophic results.
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(4) This type of strike effect on the ro torcraft is generally referred to as direct effects. Direct effects are damage which includes the burning, eroding, blasting, or structural deformation produced by the high currents of the lightning flash passing through the rotorcraft structure.
c. Procedures .
(1) Certification Plan . Although not a regulatory requirement, it is recommended that a formal written certification plan be used to ensure regulatory compliance. The use of this plan is beneficial to both the applicant and the FAA/AUTHORITY because it identifies and defines an acceptable resolution to the critical issues early in the certification process. These are the usual steps to be followed when utilizing a certification plan:
(i) Prepare a certification plan wh ich describes the analytical procedures and/or the qualification tests to be utilized to demonstrate protection effectiveness. Test proposals should describe the rotorcraft and system to be utilized, test drawing(s) as required, the method of installation that simulates the production installation, the lightning zone(s) applicable, the lightning simulation method(s), test voltage or current waveforms to be used, diagnostic methods, and the appropriate schedules and location(s) of proposed test(s).
NOTE: The recommended reference for quantification of the lightning environment, the determination of the aircraft lightning strike zones, and the determination of appropriate test methods is SAE AE4L Committee Report, dated June 20, 1978, Lightning Test Waveforms and Techniques for Aerospace Vehicles and Hardware. Additionally, information may also be found in the NASA publication No. RP-1008, Lightning Protection of Aircraft.
(ii) Obtain FAA/AUTHORITY concu rrence that the certification plan is adequate.
(iii) Obtain FAA/AUTHORITY detail par t conformity of the test articles and installation conformity of applicable portions of the test setup. Obtain FAA/AUTHORITY approval of the test proposal. A comprehensive test proposal may be used.
(iv) Schedule FAA/AUTHORITY witnessi ng of the test or tests proposed.
(v) Submit a test report describing all results and obtain FAA/AUTHORITY approval of each report prepared.
(2) Test Conditions . Refer to SAE AE4L Committee Report, dated June 20, 1978, and the NASA publication noted in paragraph c(1)(i) to determine the appropriate test parameters.
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(3) Aircraft Design Features and Criteria . MIL-B-5087B, Amendment 2, or later amendment contains valuable information to assist the designer. Figure 6 in the specification contains fault current versus bond resistance information. Refer to the NASA publication noted above also.
(i) Aluminum wire screen or mesh applied to the control or stabilizing surface and electrically bonded at each joint or juncture has been successful in conducting the current without serious damage.
(ii) Metal skin surfaces combined with surface wire screen or mesh have been successful. Also, successful use of surface treatment has been reported. For composites, treatments such as the following have been used: flame spray coatings, aluminized glass, metal foil, metallized fabrics, and conductive paint.
(iii) Ball or roller bearings may be used to conduct the current at rotating joints. However, increased friction or possible seizure of the bearing may occur. The potential for this should be evaluated. Inspection and replacement criteria for possible damage should be addressed in the manual for continued airworthiness. Bearings are especially susceptible to pitting and internal arcing.
(iv) Report DOT/FAA/CT-86/8, Ap ril 1987, Determination of Electrical Properties of Grounding, Bonding, and Fastening Techniques for Composite Materials, may assist the applicant.
(4) Fuel Systems . Refer to Report DOT/FAA/CT-83/3 referenced in paragraph AC 29.610a. For additional information on the lightning protection requirements for fuel systems for rotorcraft with a certification basis which includes Amendment 29-26 refer to paragraph AC 29.954.
d. Aircraft Design Criteria for Light ning and Static Electricity Protection.
(1) Lightning Protection .
(i) General . The rotorcraft structure should be provided with means to conduct lightning so that the rotorcraft and its occupants will not be endangered.
(ii) Rotors and Control Systems .
(A) It should be established t hat an adequate primary bonding path exists between the rotors and the airframe, such that a lightning strike on a rotor will not result in damage to or seizure of gearbox or swashplate bearings, control jacks, etc.
(B) Each hinge and bearing of rotor blades and control surfaces should either-
Page D - 10 ( 1 ) Be of a type that is capable of withstanding a lightning discharge without damage or seizure leading to loss of function, or
( 2 ) Be provided with at least one primary bonding conductor. Where bonding conductors are provided, they should be as flexible and short as possible and should be installed so that there is no danger of the conductor jamming the hinge or bearing, particularly if partially disrupted by a lightning strike.
(iii) External Non-metallic Parts .
(A) Where non-metallic parts are fi tted externally to the rotorcraft (e.g., rotors, radomes, composite skin panels) and are subjected to lightning, they should be protected against the following risks:
( 1 ) The disruption of the materials because of rapid expansion of gases within them (e.g., water vapor);
( 2 ) The rapid build-up of pressure in voids or in the enclosure provided by the parts resulting in mechanical disruption of the parts themselves or of the structure enclosed by them;
( 3 ) Fire caused by the ignition of the materials themselves or of the materials contained within the enclosures.
(B) Materials used for exter nal non-metallic parts should have low water-absorption characteristics, should not occlude gases, and should be of high dielectric strength in order to encourage surface flashover rather than puncture.
(C) Rotors and other external parts of nonmetallic construction should be provided with effective lightning diverters and/or primary conductors, which are capable of safely carrying lightning discharges, unless it can be shown that damage due to lightning discharge will not endanger the rotorcraft or its occupants. Bonding straps/leads are not required for small gaps between metallic structure and diverters in non-conducting panels in order to comply with the lightning protection criteria.
However, an electrical bonding path may be required to achieve static electricity protection.
(D) In some cases (e.g. radomes and rotors), confirmatory tests may be required to check the adequacy of the lightning protection provided.
(2) Characteristics of Lightning Discharges . The data contained in FAA AC 20-53 should be used for the purpose of assessing the adequacy of lightning discharge protection of rotorcraft.
(3) Protection Against the Accumulation of Static Charges.