(iv) Observation of telemetered flight data to assure desired control deflections, rates, and aircraft attitudes.
(v) Some combinations of items b(2)(i) through b(2) (iv) above.
(3) Assessing the operation spectrum and the flight loads or strain measurement program will involve airframe, propulsion, and flight test personnel.
(4) Variation in the operating or loading spectrum among models, and variations in the spectrum for a particular model rotorcraft, should be evaluated.
AC 20-95, Paragraph 7, entitled "Loading Spectrum," contains the statement that Table 1 (of the circular) contains typical percent of occurrences for various flight conditions for a single-piston-engine powered rotorcraft used in utility operations. In addition, the table should be used only as a guide and should be modified as necessary for each particular rotorcraft design.
(5) The difference in loading spectrum for different models that may be anticipated is illustrated by comparing the percentage of time assigned to level flight conditions, specifically 0.8 Vu to 1.0 Vu for three different rotorcraft designs where V. is the maximum airspeed at maximum continuous power in level flight. The first was obtained from Table 1, AC 20-95 which applies to a single-piston-engine powered small rotorcraft used in utility operations. The second was obtained from data for a single-turbine-engine powered seven-place small business and utility rotorcraft. The third was obtained from data for a twin-engine-powered 13 passenger transport rotorcraft. It should be noted that the level flight percentage of occurrences shown in the table below for the turbine utility business and twin turbine transport rotorcraft are only examples of a particular design. The high percentage of time shown in this flight regime could be unconservative for some designs, especially if the stresses under these design conditions produce an infinite fatigue life for the particular component.
The fatigue spectrum percentage of occurrences in AC 20-95 may be modified according to the intended operational usage of the rotorcraft. However, a conservative application should be considered.
Table 230-1 Comparison Percent of Time in Level Flight Turbine Piston Utility Twin Turbine Utility Business Transport
0.8 VNE 25% 0.8 VH 16% 0.8 VH 15%
1.0 VH 15% 0.9 V 21% 0.9 VH 20%
1.0 VNE 3% 1.0 VH 24% 1.0 VH 38% Total 43% 61% 73% 348 Par 230
This variation illustrates the "tailoring" of the loading spectrum for the type of rotorcraft and the anticipated usage.
(6) External cargo operations are a unique and demanding operation. A "logging" operator may use 50 maximum power applications per flight hour to move logs from a cutting site to a hauling site. Power is used to accelerate, decelerate, or hover prior to load release. Lifting loads over an obstruction or natural barrier is another example of very frequent high power applications for takeoff and for hovering over the release area. Similar types of operations require flight loads data to assess the effects on fatigue critical components.
(7) Frequently the applicant may request approval of a gross weight for an external cargo configuration that exceeds the standard configuration gross weight. The external cargo VNE is typically significantly lower than the standard configuration VNE possibly due to adverse effects on flight loads at the increased weight.
(8) The impact of the external cargo operation on standard configuration limits should be assessed to determine whether or not the component service lives will be affected. The assessment may be done by calculating an "external cargo configuration" service life for each critical component. The lowest service life obtained from standard configuration flight loads data and loading spectrum, or from external cargo configuration flight loads data and loading spectrum is generally the approved service life. This procedure avoids prorating the operating time between the two types of operations. This procedure is necessary since the regulatory maintenance and operating rules do not require recording time in service for the different types of operations.
(9) The applicant should plan to conduct a flight loads survey program for both a standard configuration and an external cargo configuration, if appropriate. This procedure will avoid delays associated with reinstallation and calibration of equipment.
230A. §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 practicesi
b. Procedures.
(1) Appendix 1 (formerly AC 29-571-1, December 11, 1992, Draft) provides acceptable general procedures for complying with Amendment 29-28.
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(2) Specific rotorcraft drive system gear fatigue evaluation procedures, which supplement Appendix 1, follow:
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.
(її) The tests are conducted specifically 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 curves) 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 curves), 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.
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.
Demonstration of infinite life 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.
350 Par 230A
231.-239. RESERVED.
Intentionally Left Blank Par 231 351 (thru 359)
SECTION 15. DESIGN AND CONSTRUCTION - GENERAL
240. § 29.601 DESIGN.
a. Explanation.
(1) This rule requires that no design features 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:
Seat track-to-seat interface fittings should have adequate locking devices to prevent both premature structural failure and premature unlatching.
Seat belt and harness should 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 the 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.
241. § 29.603 (Amendment 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:
360 Par 240
(1) Considerations based 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.
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 788 of this document.)
(i) The effects of excessive wear 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.
(3) Section 29.613 concerns strength properties and design values. (See Paragraph 247 of this document.)
242. § 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.
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(1) A process specification is required for fabrication processes requiring close control.
(2) A test program is explicitly required 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 consider 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 788 of this document.)
243. § 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.
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.
362 Par 242
244. § 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, painted, 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 treated, 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 hardness 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 (l.e., transmission support structure and engine mount structure) should be designed to prevent entry and collection of corrosive fluids or vapor, including water.
(1) 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.
(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.
Par 244 363
245. § 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 621 of this AC addresses the 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.
(4) This type of strike effect on the rotorcraft 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.
364 Par 245
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 which 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 tests).
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 concurrence that the certification plan is adequate.
(iii) Obtain FAA/AUTHORITY detail part 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 witnessing 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.
(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.
Par 245 365
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.
(її) 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, April 1987, Determination of Electrical Properties of Grounding, Bonding, and Fastening Techniques for Composite Materials, may assist the applicant.
(4) Euel Systems. Refer to Report DOT/FAA/CT-83/3 referenced in Paragraph 245a. 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 449 of this AC.
d. Aircraft Design Criteria for Lightning and Static Electricity Protection.
(1) Lightning Protection.
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 that 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-
(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.
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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 fitted 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 external 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.
(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.
(i) General. All items, which by the accumulation and discharge of static charges may cause a danger of electrical shock, ignition of flammable vapors or interference with essential equipment (e.g. radio communications and navigational aids) should be adequately bonded to the main aircraft grounding system.
Intermittent Contact. The design should ensure that random intermittent contact between metallic and/or metallized parts (such as could cause Par 245 367