necessary that flight conditions be conducted at limits that are appropriate for the particular rotorcraft and at the correct combination of these limits. It will be the responsibility of engineering and flight test personnel to determine that the flight strain program proposal includes conditions of flight at the various combinations of rotor RPM, airspeed, thrust, etc., that will be representative of the limits used in service. The flight test personnel should assure that the severity of the maneuvers to be investigated is such that actual service use will not be more severe. Verification that proposed maneuvers are suitable may be achieved by:
(i) Flying a representative set of maneuvers with the applicant's pilot in the test aircraft at noncritical combinations of weight, CG, and speed. (An FAA/AUTHORITY letter for specific test authorization would ordinarily be required.) If the procedure is used, the applicant should provide adequate preliminary flight strain data from development or other tests to confirm a cleared (non-critical) flight envelope for conduct of these representative maneuvers.
(ii) Flying a representative set of maneuvers with the applicant's pilot in a similar (certified) model to assess and agree upon the required maneuvers, control deflections, and aircraft rates. The required maneuvers or conditions will be specified in the flight strain program plan.
(iii) Flying a chase aircraft which has a flight envelope appropriate to allow visual confirmation of the proposed and programmed flight maneuvers.
(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 small 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 as shown in Table 230-1. (V. is the maximum airspeed at maximum continuous power in level flight.) The first column was obtained from Table 1, AC 20-95 which applies to a 286 Par 230
single-piston-engine powered small rotorcraft used in utility operations. The second column is appropriate tor a single-turbine-engine powered seven-place small business and utility rotorcraft. The third column is appropriate for a twin-engine-powered 13 passenger transport rotorcraft. It should be noted that the level flight percentage of occurrences shown in Table 230-1 for the turbine utility business and turbine transport rotorcraft are examples of particular designs. The high percentage of time shown in this level 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 shall be modified according to the intended operational usage of the rotorcraft. However, a conservative application should be considered. This variation illustrates the "tailoring" of the loading spectrum for the type of rotorcraft and the anticipated usage.
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 VH 21% 0.9 VH 20%
1.0 VNE 3% 1.0 VH 24% 1.0 VH 38% Total 43% 61% 73%
(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) The impact of the external cargo operation on standard configuration limits should be assessed to determine whether or not the component service lives, inspections, etc., 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 or from frequent ground-air-ground cycles is generally the approved service life or replacement time. Since the regulatory maintenance and operating rules do not require recording time in service for the different types of operations, this procedure could be used if an "operational cycles" equation for equivalent flight hours is not approved (see (8) below).
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(8) The Airworthiness Limitations Section of the maintenance manual shall contain the required information derived from complying with the standard. If an "operational cycles" equation for "equivalent flight hours" is approved under the standard, the equation is included in this approved section of the manual.
(9) The applicant should plan to conduct a flight loads survey program for both a standard configuration and an external cargo configuration, if applicable. The ground-air-ground cycle is inherent in these conditions. This procedure will avoid delays associated with reinstallation and calibration of equipment.
230A. $27.571 (Amendment 27-33) FATIGUE EVALUATION OF FLIGHT STRUCTURE FOR CATEGORY A CERTIFICATION.
a. Explanation. Amendment 27-33 added Appendix C to specify the requirements for Category A certification of normal category rotorcraft. The requirement for fatigue tolerance evaluation will require test evidence to support the analysis.
b. Procedures. For Category A certification, the tests specified in AC 29-2A, Section 230A are required for fatigue tolerance evaluation. AC 29-2A, Section 230A is repeated in this section.
(1) 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.
(2) 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.
(3) 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 test conditions, should at a minimum, encompass those power levels for which repeated application inservice is expected under normal circumstances. 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.
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(4) 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, etc., 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.
(5) 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.
(6) The 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.
231.-240. RESERVED.
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SECTION 15, DESIGN AND CONSTRUCTION - GENERAL
241. $ 27.601 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. These 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 and 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 (normal wear and tear) experience 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.
242. § 27.603 (Amendment 27-16) 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:
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(1) Considerations based on experience or tests.
(2) By meeting approved specifications.
(3) By taking into account environmental conditions such as temperature and humidity.
b. Procedures.
(1) 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.
(2) Environmental conditions may be taken into account by service experience, coupon testing, full-scale testing, or a combination of testing and experience, MIL-HDBK's -5, -17, and -23 include some environmental effects and contain reference to additional methods of testing for environmental effects.
(3) Section 27.613 concerns strength properties and design values. (See Paragraph 248 of this document.)
243. § 27.605 (Amendment 27-16) 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 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, ASIM, 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.
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(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.
244. § 27.607 (Amendment 27-4) FASTENERS.
a. Explanation. Section 27.607 of Amendment 27-4 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 contains information, procedures, and means of complying with § 27.607 of Amendment 27-4.
245. § 27.609 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 §§ 27.999 and
27.1193.
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 ensure the structure or parts are not adversely affected during, or as a result of, the plating or treatment process.
(§ 27.605 concerns approval of process specifications and fabrication methods.)
(3) Plating or material surface hardness or composition changes may require fatigue substantiation to ensure 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.
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(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.
(il) 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 ensure fluids are properly drained.
(7) Drains or vents which handle corrosive fumes (such as battery case vent line) may incorporate a container with an agent to neutralize the fumes prior to venting overboard.
(8) Welded tubes should be flushed and sealed after welding in accordance with an approved process specification.
(9) Refer to AC 43-4, "Corrosion Control for Aircraft," for further procedures.
246. § 27,610 (Amendment 27-21) LIGHTNING 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 27-21. For further information, see the preamble of Amendment 27-21 (49 FR 44433; 11/6/84), Proposal 2-14. Section 27.610 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. Section 27.610 is the standard referenced in the requirement for lightning protection of systems in § 27.1309(d) (see Paragraph 621 of this AC). In addition, the protection of fuel systems from the effects of lightning is found and referenced in Report DOT/FAA/CT-83/3, User Manual for AC 20-53A, Protection of Airplane Fuel Systems Against Fuel Vapor Ignition Due to Lightning, dated April 12, 1985.
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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 operation of the rotorcraft.
(2) Paragraph 621 of this AC addresses the protection required for systems.
The 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.
c. Procedures.
(1) Certification Plan. Although not a regulatory requirement, it is recommended that a formal written certification plan be used to assure 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 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, 304 Par 246
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.
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.
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.
(i¡) 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.
(ili) 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) Fuel Systems. Refer to Report DOT/FAA/CT-83/3 referenced in Paragraph 246a. For additional information on the lightning protection requirements for Par 246 305
fuel systems for rotorcraft with a certification basis which includes Amendment 27-23 refer to Paragraph 449 of this AC.
247. § 27.611 INSPECTION PROVISIONS.
a. Explanation. The rotorcraft must have access panels or openings that will allow for proper maintenance and/or adjustment of the rotorcraft systems.
(1) The rule states: There must be means to allow close examination of each part that requires recurring inspection, adjustment for proper alignment and functioning, or lubrication."
(2) "Structural" or load-carrying access panels may be used to comply with the rule. Structural panels should have stencils or permanent labels (§ 27.1541(a)(2)) stating the panels must be installed prior to ground or flight operation.
(3) Holes or "nonstructural" access panels should be used whenever possible.
b. Procedures.
(1) The determination of compliance can be accomplished in conjunction with the following activities:
(i) Reviewing type design drawings.
(ii) Conformity inspections accomplished during certification testing.
(ili) Be evaluated during the control system proof and operation tests (§§ 27.681 and 27.683).
(iv) During type inspection tests and functioning and reliability testing.
(2) Equipment requiring frequent inspections (at less than 25-hour intervals), lubrication, or adjustments should be accessible through "nonstructural" doors. Areas or items requiring daily attention should be accessible through "nonstructural" doors since properly rated maintenance personnel are required to "open and close" or reinstall structural panels, and special design features, such as multiple pins and latches, are generally necessary for structural doors.
248. § 27.613 (Amendment 27-16) MATERIAL STRENGTH PROPERTIES AND DESIGN VALUES.
a. Explanation. The rule requires the use of materials that have a known minimum strength value. The structure must not be understrength and must be designed to minimize fatigue failure.
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(1) Material design values in certain specified documents may be used. The FAA/AUTHORITY may approve other material design values thus allowing the applicant greater flexibility in selection of materials by proving their strength properties and design values as stated in § 27.613(d).
(2) Other materials that may be new or are not included in the specified documents may be tested and design values established as provided by § 27.613(a) and (d).
(3) Section 27.613(d) requires the selection of materials that will retain design values and properties in the type of service environment and for the length of service time intended for the structure.
(4) Section 27.613(c) is an objective rule concerning minimizing fatigue failures and § 27.571 concerns quantitative fatigue substantiation requirements.
b. Procedures.
(1) The properties and design values in the documents noted in the rule may be used.
(2) MIL-HDBK-5, Metallic Materials and Elements for Flight Vehicle Structure, Chapter 9, contains procedures for establishing design values of additional materials.
Uniform means of presenting the data are also contained in this chapter.
(3) Design values and properties must include effects of the service environment and service time. An example is exposure at elevated temperatures on the ultimate tensile strength of 7079-T6 aluminum alloys as found in Figure 3.7.4.1.1(c) of MIL-HDBK-5.
(4) The probability of disastrous fatigue failures must be minimized. This may be accomplished by using design features usually identified as fail-safe features, such as the following, which were obtained from Advisory Circular 20-95.
Selection of materials with stress levels to provide a controlled slow rate of crack propagation combined with high residual strength after initiation of cracks (lightly loaded structures).
(її) Use of multipath construction and the provision of crack stoppers to limit the growth of cracks.
(iii) Use of composite (multielement) duplicate structures so that a fatigue crack or failure occurring in one element of the composite (multielement) member will Par 248 307