§ 7.260 (§ 29.561) without extreme distortion which might tend to rupture the fuel tank may be considered to comply with the requirements of this section
(7.261).
Puncture resistant "bladder" fuel cells that are adequately designed and also protected from the stated impact loads imposed on the fuselage may also satisfy the standards.
(8) For rotorcraft with retractable landing gear, alternative landing gear positions and the resulting effects on potential fuel release should be evaluated.
218A. § 29.561 (Amendment 29-29) EMERGENCY LANDING CONDITIONS - GENERAL.
a. Explanation. Amendment 29-29 adds or increases the design static load factors of § 29.561 in three different areas:
(1) The design static load factors for the cabin in § 29.561(b)(3) are increased in concert with the dynamic test requirements of new § 29.562.
(2) Design static load factors are added in § 29.561(c) for external items of mass located above and/or behind the crew and passenger compartment.
(3) The static load factors, which were formerly only referenced in § 29.561(d), are now included explicitly in § 29.561(d) for substantiation of internal fuel tanks which are below the passenger floor.
b. Procedures. The procedures of Paragraph 218, § 29.561, continue to apply except the new load factors of § 29.561 should be used. Penetration of any items of mass into the cabin or occupied areas should be prevented.
218B. § 29,561 (Amendment 29-38) EMERGENCY LANDING
CONDITIONS-GENERAL
a. Explanation. Amendment 29-38 adds a new rearward emergency load factor of
1.5g to both §§ 29.561(b)(3)(v) and 29.561(c)(5). The addition of the 1.5g rearward load factor in § 29.561(b)(3)(v) is to provide an aft ultimate load condition for substantiation of the restraints required for retention of both occupants and significant items of mass inside the cabin that could otherwise come loose and cause injuries in an emergency landing. The addition of the 1.5g rearward load factor to § 29.561(c)(5) is to provide an aft ultimate load condition for substantiation of the support structure for retention of significant items of mass above and forward of the occupied volumes) of the rotorcraft that could otherwise come loose and injure an occupant in an emergency landing. Amendment 29-38 also increases the forward, sideward, and downward emergency load factors of § 29.561(c)(2), (c)(3), and (c)(4), respectively, for retention of 334 Par 218
items of mass above and behind the occupied volume(s) that could otherwise come loose and injure an occupant in an emergency landing.
b. Procedures. The procedures of Paragraphs 218 and 218A continue to apply except the newly specified load factors must be used. A list of the significant items of mass to be considered should be compiled by the applicant and approved by the certifying authority.
219. § 29.562 EMERGENCY LANDING DYNAMIC CONDITIONS.
a. Explanation. Amendment 29-29 adds new requirements for the dynamic testing of all seats in rotorcraft.
b. Procedures. AC 20-137, Dynamic Evaluation of Seat Restraint Systems and Occupant Restraint for Rotorcraft (Normal and Transport)i provides procedures for complying with § 29. 562 using the 170-pound anthropomorphic test dummy specified in § 29.562(b). Those seats not occupied for takeoff and landing, and so placarded and identified in the rotorcraft flight manual (RFM), may be excluded from compliance.
220. § 29.563 (Amendment 29-12) STRUCTURAL DITCHING PROVISIONS.
a. Explanation. Amendment 29-12 included certification requirements for ditching approvals. The rotorcraft must be able to sustain an emergency landing in water as prescribed by § 29.801(e).
b. Procedures. Refer to Paragraph 337, § 29.801, for procedures.
220A. § 29.563 (Amendment 29-30) STRUCTURAL DITCHING PROVISIONS.
a. Explanation. Amendment 29-30 added specific structural conditions to be considered to support the overall ditching requirements of § 29.801. These conditions are to be applied to rotorcraft for which over-water operations and associated ditching approvals are requested.
(1) The forward speed landing conditions are specified as:
The rotorcraft should contact the most critical wave for reasonable, probable water conditions in the likely pitch, roll, and yaw attitudes.
(ii) The forward velocity relative to wave surface should be in a range of O to 30 knots with a vertical descent rate of not less than 5 FPS relative to the mean water surface.
Par 218B 335
NOTE: A forward velocity of less than 30 knots may be used for multiengine rotorcraft if it can be demonstrated that the forward velocity selected would not be exceeded in a normal one-engine-out touchdown.
(iii) Rotor lift of not more than two-thirds of the design maximum weight may be used to act through the CG throughout the landing impact.
(2) For floats fixed or deployed before water contact, the auxiliary or emergency float conditions are specified in § 29.563(b)(i). Loads for a fully immersed float should be applied (unless it is shown that full immersion is unlikely). If full immersion is unlikely, loads resulting from restoring moments are specified for sidewind and unsymmetrical rotorcraft landing.
(3) Floats deployed after water contact are normally considered fully immersed during and after full inflation. An exception would be when the inflation interval is long enough that full immersion of the inflated floats does not occur; e.g., deceleration of the rotorcraft during water impact and natural buoyancy of the hull prevent full immersion loads on the fully inflated floats.
b. Procedures.
(1) The rotorcraft support structure, structure-float attachments, and floats should be substantiated for rational limit and ultimate ditching loads.
(2) The most severe wave heights for which approval is desired are to be considered. A minimum of Sea State 4 condition wave heights should be considered (reference Paragraph 337 (§ 29.801) of this AC for a description of Sea State 4 conditions).
(3) The landing structural design consideration should be based on water impact with a rotor lift of not more than two-thirds of the maximum design weight acting through the center of gravity under the following conditions:
Forward velocities of 0 to 30 knots (or a reduced maximum forward velocity if it can be demonstrated that a lower maximum velocity would not be exceeded in a normal one-engine-out landing).
(i¡) The rotorcraft pitch attitude that would reasonably be expected to occur in service. Autorotation flight tests or one-engine-inoperative flight tests, as applicable, should be used to confirm the attitude selected. This information should be included in the Type Inspection Report.
(iii) Likely roll and yaw attitudes.
(iv) Vertical descent velocity of 5 FPS or greater.
336 Par 220A
(4) Landing load factors and water load distribution may be determined by water drop tests or analysis based on tests.
(5) Auxiliary or emergency float loads should be determined by full immersion or the use of restoring moments required to react upsetting moments caused by sidewind, asymmetrical rotorcraft landing, water wave action, rotorcraft inertia, and probable structure damage and punctures considered under § 29.801. Auxiliary or emergency float loads may be determined by tests or analysis based on tests.
(6) Floats deployed after initial water contact are required to be substantiated by tests or analysis for the specified immersion loads (same as for (5) above and for the specified combined vertical and drag loads).
Par 220A 337
221.-229. RESERVED.
Intentionally Left Blank Par 221 338 (thru 346)
7/30/97 АС 29-2B SECTION 14. FATIGUE EVALUATION
230. § 29.571 FATIGUE EVALUATION OF FLIGHT STRUCTURE.
a. Explanation. An evaluation is required to assure structural reliability of the rotorcraft in flight. Advisory Circular 20-95 contains background information and acceptable means of compliance with the requirements. A safe life may be assigned or the structure may be fail sate as prescribed.
b. Procedures.
(1) The fatigue evaluation requires consideration of the following factors:
Identification of the structure/components to be considered.
(її) The stress during operating conditions.
(Ill) The operating spectrum or frequency of occurrence.
(iv) Fatigue strength, and/or fatigue crack propagation characteristics, residual strength of the cracked structure.
(2) Since the design limits, e.g., rotor RPM (maximum and minimum), airspeed, and blade angles (thrust, weight, etc.) affect the fatigue life of the rotor system, it is 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 flight test personnel to determine that the flight strain program 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. Flight test verification may be achieved through:
(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.)
(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.
(ill) Flying a chase aircraft which has a flight envelope appropriate to allow visual confirmation of the proposed and programmed flight maneuvers.
Par 230 347
(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.
Par 230 349
(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.
Par 241 361