23.562 Emergency landing dynamic conditions (Amendment 23-50)
165. Is there any policy available on paragraph (a) of this section as of January 1, 2007?
No.
166. Do I have to retest the seats in my airplane if I want to change the fabric dress covers, following the conduct of dynamic certification tests per the seat dynamic performance standards? No. You may substitute typical dress fa brics for each other, including leather, without retest. The change in friction due to dress cover alone is less significant than the variation possible due to clothing, etc. However, dress covers that exhibit low friction coefficients (i.e., hard plastics, which are not in use, to our knowledge) may require some resubstantiation.
167. Are there any certification issues at all for changing the dress covers or seat cushions on seats? Yes. Dress cover and cushion changes that affect the flammability of the dress cover and seat do require resubstantiation to the respective flammability requirements. Also see the next question about cushions.
168. Do I have to retest or analyze the seats in my airplane if I want to change the seat cushions, following the conduct of dynamic certification tests per the seat dynamic performance standards? Yes. The seat cushion static and dynamic res ponse characteristics can affect the lumbar column/pelvic load measured during the seat dynamic certification tests.
Thus, a change in the seat cushion would require re-substantiation. Analysis may be used per the guidance in AC 20-146.
169. In the acquisition of data for Head Injury Criteria (HIC), are 8-bit analog to digital converters adequate? No. In accordance with Society of Automotive Engineers (SAE) recommended practice J211 “Instrumentation for Impact Tests,” the converter should be at least 10-bit in order to obtain a sufficient level of resolution.
170. Can I run a vertical drop test as an acceptable alternate means for the determination of seat/floor deformation requirements found in § 23.562(b)(2)?
a. No. The seat/floor deformation requirement is more an assessment of the airplane seat’s tolerance to structural deformations than it is an assessment of potential airframe deformations. Intolerance of seat/floor interface to deformation is deemed a major cause of structural failure of the seat/floor structural attachment. If you want to pursue this type of test for information for research only (but not for certification), the following are the kind of items that you should consider.
(1) Drop tests are typically a simulation of the vertical impact velocity only; an actual airplane accident possesses both vertical and longitudinal impact velocity vectors. While the measured impact acceleration levels may provide some insight into the overall vertical impact characteristics of the airframe, the test condition does not simulate the combined vertical and longitudinal (and potentially lateral) loads on the airplane seat, its attachments, and interface structure.
(2) An actual airplane accident frequently occurs on an irregular impact surface that may also contain local obstructions, which can cause significant local loading, structural fractures, and deformations. A vertical drop test on a flat surface does not address the common and more critical impact case that may occur on irregular and obstructed surfaces.
(3) Large, significant, dynamic, structural deformations can occur during an impact test, or airplane accident that are not readily detectable for measure after the impact condition.
These results have been demonstrated during the general aviation airplane impact test series.
Consider this if you plan to measure the residual static floor deformation after the impact test.
(4) Choose accelerometers with an amplitude class that is adequate for the g-loads expected during your particular drop test.
b. The above paragraphs attempt to address the kinds of considerations that the 10 degree pitch and 10-degree roll in part 23, § 23.562, already account for.
171. Can I use static ultimate load tests as an alternate means of showing compliance to the dynamic seat test requirements of § 23.562? No. One of the key requirements of part 23, § 23.562 is to assure that an occupant’s vertical lumbar load is limited to 1,500 pounds or less during the 30 degree incline test. In addition, the HIC and upper torso restraint maximum loads can only be demonstrated by dynamic testing. The FAA is unaware of any analytical (except computer modeling techniques, discussed below) or static test procedure that will meet the dynamic seat test requirements.
172. Is there any policy available on paragraph (c) of this section as of January 1, 2007?
No.
173. Is there any policy available on paragraph (d) of this section as of January 1, 2007?
No.
174. Is analysis alone an acceptable alternative approach as stated in paragraph (e)?
a. Yes, per the guidance in AC 20-146.
b. The following two SAE papers provide helpful, related, information:
(1) SAE Paper No. 850853, “The Development of Dynamic Performance Standards for General Aviation Aircraft Seats,” Stephen J. Soltis and John W. Olcott.
(2) SAE Paper No. 851847, “Human Injury Criteria Relative to Civil Aircraft Seat and Restraint System,” Richard F. Chandler.
c. AC 23.562-1, Dynamic Testing of Part 23 Airplane Seat/Restraint Systems and Occupant Protection, contains useful guidance about this airworthiness standard topic.
175. How are the minimum safety standards determined for restricted category airplanes?
The FAA issues type certificates under the regulatory procedures of part 21, Certification Procedures for Products and Parts. The Small Airplane Directorate decides the appropriateness of part 23 airworthiness standards as they apply to the special purpose restricted category airplane. If a regulation is found inappropriate, the FAA excludes that regulation from the certification basis of that airplane.
NOTE: A finding of equivalent level of safe ty or the need for an exemption is unnecessary for that specific airworthiness standard.
176. How are the minimum safety standards determined for agricultural or special purpose airplanes?
a. The FAA has allowed the certification of restricted category aircraft since Civil Air Regulations (CAR) Part 8 was introduced (around 1950). The preamble to CAR 8 recognized that, for restricted category aircraft where the public was not endangered, it was unnecessary to provide an equivalent level of safety to the standard airworthiness requirements.
b. Policy: As of December 1, 1997, for emergency landing dynamic conditions, evaluate the airplane using at least the following considerations:
(1) The placement of the chemical hopper forward of the cockpit so that there is no large item of mass that threatens to collapse the cockpit should a crash occur.
(2) The elimination of any protruding knobs, handles, or other rigid structures from the cockpit that the pilot or crew member may contact in a crash. Approved Department of Transportation or Mil-Spec protective headgear is mandatory.
(3) The installation of a military type lap belt and shoulder harness having a 5,000 pound rating or greater or approved equivalent.
(4) Special purpose crew members who assist in the aerial application operation,
i.e., flaggers, loaders, may be carried in ferry flights provided certain conditions are met. First, each crew member must have a seat, a lap belt and shoulder harness comparable in strength to that of the pilots. Second, the crew seat must not be in the cockpit. Third, the crew seat must be located behind the pilot seat. Special purpose crew members that are carried for any other purpose will be afforded the same protection as that of the pilot.
c. Dynamic seat airworthiness standards used in part 23 were developed for normal, utility, and acrobatic category airplanes only. These standards were never intended for use on restricted category airplanes. When NASA, the general aviation industry, NTSB, and the FAA examined survivability envelopes, they did not consider restricted category airplanes in the database because there were notable differences in the crash scenarios.
d. The Small Airplane Directorate decides whether dynamic seat test requirements are appropriate or not, after reviewing crashworthiness design features of the specific make and model airplane. Service experience shows that agricultural airplane operators have a lower accident fatality rate than general aviation operators. Certain make and model agricultural airplane designs contain increasingly effective crashworthy features throughout recent development history, which are the obvious reasons for the fatality rate differences.
e. Policy: The FAA will not automatically exclude the emergency landing dynamic conditions requirement (§ 23.562) for any applicant who is seeking a type certificate for an agricultural airplane. Instead, the Small Airplane Directorate reviews the design for compensating features to the dynamic seat airworthiness standards and decides whether the dynamic seat test requirements are appropriate or not on a case-by-case basis.
f. For useful guidance, refer to AC 21.25-1, Issuance of Type Certificate: Restricted Category Agricultural Airplanes.
177. I would like to use ejection seats for emergency egress. Can I get my part 23 airplane certificated with ejection seats installed? Yes.
178. What are the criteria or minimum safety standards for approving ejection seats for use in an airplane certificated to part 23? Part 23 does not have any minimum safety standards for ejection seats, and the FAA does not have the expertise to evaluate ejection seat testing. However, the military has been develo ping ejection seat technology for 40 to 50 years.
Therefore, the applicable sections of the Air Force Systems Requirement Document (SRD) are used as the minimum safety standards. The SRD provides specific criteria for crew escape, and for ejection seat testing and verification. The Air Force is used as a consultant. We only approve the ejection seat data after the Air Force has reviewed the data and concurs that the ejection seat satisfies the SRD crew escape, testing, and evaluation requirements.
179. Do I have to show literal compliance to the emergency landing dynamic conditions of this rule (23.562) for ejection seats?
a. No. You may show literal compliance if you choose to, or you may pursue an Equivalent Level of Safety (ELOS) finding for compliance to this rule. An ELOS is allowed instead of literal compliance to a rule, by part 21, § 21.21(b)(1). You may use the Air Force SRD as the basis for this finding.
b. The lower limits of the ejection seat operating envelope must be zero altitude above ground level and zero airspeed if you choose to pursue an ELOS for this rule.
180. Are there other rules relevant to ejection seats that I may pursue an ELOS for? Yes.
You may also pursue an ELOS for the applicable portions of § 23.807 and § 23.1309.
181. The original certification basis of my airplane is before Amendment 23-36 and was manufactured on, or before, December 12, 1986. I wo uld like to install side-facing seats.
Do I have to comply with the emergency landing dynamic conditions of this rule (23.562)?
a. Yes. You must comply with the emergency landing dynamic conditions of this rule if you want approval by a Supplemental Type Certificate (STC), an amended Supplemental Type Certificate (ASTC), or an amended Type Certificate (ATC) for the installation of side-facing seats.
b. Earlier regulations were written with forward or aft-facing seats in mind, and they do not provide adequate standards for protection of occupants in side-facing seats. Therefore, in accordance with part 21, § 21.101, an applicant must comply with the current standards, including the emergency landing dynamic conditions contained in this rule (23.562), before any side-facing seat installation (to be occupied during takeoff or landing) can be approved.
c. An alternative to compliance with the current regulations is to prohibit occupation of side-facing seats during takeoff and landing through the use of a flight manual supplement and clearly visible placards. This t ype of installation should include a restraint system that keeps the occupant securely seated during turbulence or maneuvering.
d. This policy is applicable to airplanes whose original certification basis was before Amendment 23-36 to part 23, and the airplanes that were manufactured on, or before 62
December 12, 1986. This includes normal, utility, acrobatic, and commuter category airplanes with earlier amendments to part 23, Civil Air Regulations 3 (CAR-3), or airplanes with any previous airworthiness regulations as their original certification basis.
e. The one exception to this policy is an airplane that had side-facing seats as part of its original certification basis and with a certification basis that falls within the scope of paragraph d, above. In this instance, an application for an STC, ASTC, or ATC for side-facing seats does not need to meet the requirements of emergency landing dynamic conditions of this rule
(23.562).
f. Also see § 23.785 in AC “Systems and Equipment Guide for Certification of Part 23 Airplanes and Airplanes”.
FATIGUE EVALUATION
23.571 Metallic pressurized cabin structures (Amendment 23-48)
182. What is the history of this regulation?
a. History: Fatigue evaluation of pressurized cabins was first required for small airplanes by Amendment 3-2 of the Civil Air Regulations (CAR), Part 3, effective August 12, 1957, and it continued into the original part 23.
b. Safe-life requirements mandate that certain critical structural elements have a fatigue life determined during the airplane type certification process. Life-limited items are normally identified in the Type Certificate Data Sheet Note 3. As of Amendment 23-26, part 23, § 23.1529 requires time-limited items to be shown in the Instructions for Continued Airworthiness—Airworthiness Limitations Section.
c. Amendment 23-45, effective September 7, 1993, provides § 23.573(b), Damage tolerance and fatigue evaluation of structure, as an option to § 23.571(a) and (b).
d. Policy: The FAA allows an extension of the originally imposed safe-life limits only through a formal reinvestigation of the life-limited parts. Generally, the type certificate holder conducts a new fatigue certification program. This requires specific FAA attention since lifelimit approvals are normally beyond the scope of delegation authority granted by the FAA.
Replacing the safe-life structural elements is a viable alternative to recertification. However, this is not always the best course of action because some major structural elements, like the wing and fuselage pressure vessel assemblies, are not easily or economically replaceable.
183. Is there other guidance available on fatigue? Yes. AC 23-13, Fatigue and Fail-Safe Evaluation of Flight Structure and Pressurized Cabin for Part 23 Airplanes, contains useful guidance about this airworthiness standard topic.
184. What features should be given special consideration as it relates to fatigue and damage tolerance? Cutouts: Doors, windows, access holes, etc., in aircraft structure cause redistribution of axial and shear loads, pressure loads, and stiffness changes. Fatigue capabilities may be affected. Damage tolerance capabilities may be affected. Account for all such design changes.
23.572 Metallic wing, empennage, and associated (Amendment 23-48) structures
185. What is the history on paragraph (a) of this regulation?
a. Civil Air Regulations (CAR) 3, Amendment 3-2, effective August 12, 1957, first imposed a fatigue evaluation of pressurized cabin airplane designs.
b. Part 23 adopted an airworthiness standard to evaluate fatigue of an airplane wing and associated structure at Amendment 23-7, effective September 14, 1969.
c. Amendment 23-34, effective February 17, 1987, added commuter category airplanes to part 23. Empennage fatigue requirements were included for these airplanes. SFAR 41, which applied to part 23 derivative-model airplanes, always had such a requirement. Effective October 26, 1989, the FAA issued Amendment 23-38, which extended the fatigue requirement for empennage, canard surfaces, tandem wing, winglets, and tip fins to all part 23 airplanes.
Amendment 23-45, effective September 7, 1993, added the option of a damage tolerance evaluation, as defined in § 23.573(b). Amendment 23-48, effective March 11, 1996, made damage tolerance evaluation mandatory for commuter category airplanes.
d. The narrative on spar-component fatigue tests, combined with the safe-life and fail safe design philosophies, displays the FAA’s intentions before 1971.
186. What testing is typically required?
a. Component testing, while acceptable under certain conditions, presents a problem of determining which structure to test and how to ensure that the correct testing conditions are applied. The FAA intends for the airplane de signer to show that the wing, wing carry-through structure, and attaching structures comply with the fatigue requirements (§ 23.572). These examinations may exclude the control surfaces and their attachments.
b. These examinations normally include the main spar, the secondary spar, stringers, torque box skin, and at least the main internal ribs. While a main spar-component test could adequately substantiate the spar, the remaining structures should be proven by additional component tests, or analyzed as either safe-life or fail-safe structures. If fail-safe compliance is chosen, determine if the specified loads can be supported with a failed element. Also determine the number and kind of inspections to find the damage before catastrophic failure.
c. When testing the main spar, simulate load transfer through skin attachment units and the associated fretting. Further, consider and simulate any significant eccentricities and rib-to spar-cap loads.
d. Exercise good judgment to ensure that elements and aspects of primary importance to safety receive the most emphasis.
187. Which design fatigue life governs when the airplane is certificated for use in normal and utility categories? When a safe-life limit is established for airplane designs certified in both normal and utility categories, show the lower of the two lives in the TCDS. Show the following note in the TCDS, the Airplane Flight Manual, and the Airworthiness Limitations Section of the Instructions for Continued Airworthiness (when they exist): “Since the airplane is type certificated under both normal and utility categories, the lower fatigue life has been listed in the TCDS.”
188. Is fatigue an issue for wood structure? Research indicates wood is not sensitive to fatigue if the stress levels are low. Emphasis should be placed on ultimate load tests and environmental protection of the wood to prevent deterioration due to dry rot and other related environmental factors.
189. What effect will an increase in gross weight or an increase in airplane speed have?
a. A gross weight increase or an increase in airplane speed frequently demands an increase in horsepower. An engine horsepower increase may significantly relocate the engine center of gravity or increase engine weight with respect to the original installation, or both. A different engine mount (stiffness change) may accompany a different engine installation. When engine c.g., engine weight, or the engine mount is changed, one should consider the benefits of a pre-modification and post-modification ground vibration survey. The purpose of the survey is to identify any coupling or resonant characteristics changed by a different engine installation (i.e., nodes, modes, and frequencies).
b. An in-flight vibration monitoring program or a flight-strain survey could assess the effects of fatigue on airplane empennage structure, due to propeller slipstream impingement.
These in-flight tests could also provide insights on whether increased horsepower forced vibrations affect airplane critical vibration environments that were previously benign.
c. One should also compare the in-flight torsional and bending peak stresses, and some selected panel strain-gauge readings, with the original airplane design data. The objective is to verify that the new stress levels will not adversely affect the fatigue life of the empennage.
d. Often, a modifier neither has, nor can they get, access to the original airplane design data. Consequently, other means of showing co mpliance to the airworthiness standards may be used. One evaluation technique is to compare th e new fatigue stresses to the material endurance level, i.e., the S-N curves. A ground vibration survey and a flight-strain survey conducted before and after the modification can provide some data to perform a comparative analysis. If changes in the stresses are large enough to affect the empennage fatigue life, the modifier should determine appropriate structural design changes to include in the modification. Good engineering judgment should be exercised here. The service history of a similar type airplane incorporating a like modification may be used to identify potential fatigue crack locations and to serve as a guide when preparing detailed structure inspection methods and frequencies.
e. Fatigue critical structure is defined as structure whose failure would cause catastrophic loss of the whole airplane. Critical structure would include the spar, the primary fittings, the pressurized fuselage skin-stringer combinations, and the frames.
f. Instead of a fatigue or fail-safe strength investigation, § 23.572(a) permits alternate compliance. It permits compliance by showing that the structure, operating stress level, materials, and expected use are comparable, from a fatigue standpoint, to a similar design that had extensive satisfactory service experience.
g. FAA Report No. ACE-100-01, “Fatigue Evaluation of Empennage, Forward Wing, and Winglets/Tip Fins,” contains comprehensive guidance on this subject. It is available from the National Technical Information Service, Springfield, VA 22161.
h. See 23.571(a), Metallic pressurized cabin structures, for FAA guidance about safelife airplane fatigue limitations imposed during the type certification process of civil aircraft.
i. Fail-safe strength: No policy av ailable as of January 1, 2007.
j. Damage tolerance: Effective September 7, 1993, Amendment 23-45 provided damage tolerance, § 23.573(b), as an option to § 23.572(a)(1) and (a)(2).
190. Is there any policy available on paragraph (b) of this section as of January 1, 2007?
No.
191. What other guidance is available for fatigue? AC 23-13, Fatigue and Fail-Safe Evaluation of Flight Structure and Pressurized Cabin for Part 23 Airplanes, contains comprehensive guidance about this airworthiness standard topic. Please refer to it for additional information. AC 23-13A will soon be released. Therefore, until AC 23-13 is revised to convey the following information, it is presented here as a courtesy: Report AFS-120-73-2, “Fatigue Evaluation of Wing and Associated Structure on Small Airplanes,” contains fatigue load spectra for various part 23 airplane usage categories. It also contains detailed procedures for the fatigue strength (safe life) investigation of § 23.572(a)(1).
192. Should a pilot make airplane maintenance record entries about airplane operations that relate to the established life-limits of the airplane according to instructions in a Pilot Operating Handbook? A type certificate applicant should show compliance with § 23.572 (fatigue life limits) using certification procedures for airplane design. The FAA does not allow a type certificate applicant to impose a requirement for the pilot to record aerobatic flight time in the airplane maintenance record. It is not possible for the FAA to validate these kinds of airplane record entries. Maintenance records, in the 1997 version of part 91, require the following registered owner or operator entries for the airframe, each engine, each propeller, and each rotor:
a. Total time in service; and
b. Current status of life-limited parts.
23.573 Damage tolerance and fatigue evaluation of (Amendment 23-48) structure
193. Is there any policy available on paragraph (a) of this section as of January 1, 2007?
Since Amendment 23-45 (effective September 7, 1993), the damage tolerance and fatigue evaluation of composite structure has been based on the applicable requirements of part 23, § 23.573. AC 20-107A, Composite Aircraft Structure, which existed before Amendment 23-45, contains acceptable means of showing compliance with these requirements. One consideration that is different for composite materials than for metallic structures is that of impact damage.
For composite structures, impact damage resulting from events such as dropped tools or hail impacts is difficult to detect, but it may cause degradation of static or fatigue strength. Another difference for composite structures is that, in addition to tensile loads, compressive loads may also drive damage growth. When demonstrating compliance with the growth rate or no-growth rate of a damage requirement (§ 23.573(a)(2)), it is important to consider compressive loads that may drive the growth of disbonds or delaminations in composite structures. For movable control surfaces, include any structure that, if it failed, would cause loss of the airplane. Also see Policy Statement PS-ACE100-2001-006, “Static Strength Substantiation of Composite Airplane Structure”.
194. Is there any policy available on paragraph (b) of this section as of January 1, 2007?
There is no guidance for metallic structures damage tolerance assessment for part 23 airplanes.
Now, a Transport category airplane (part 25), AC 25.571-1C, Damage-Tolerance and Fatigue Evaluation of Structure, dated April 29, 1998, provides the only information available from the FAA.
23.574 Metallic damage tolerance and fatigue evaluation of commuter category airplanes (Amendment 23-48)
195. Is there any policy available on this section as of January 1, 2007? No.
23.575 Inspections and other procedures (Amendment 23-48)
196. Is there any policy available on this section as of January 1, 2007? No.
SUBPART D—DESIGN AND CONSTRUCTION
23.601 General (Original)
197. Are there any special considerations for testing?
a. The manner of testing can bias the test results. Several years ago, the wings failed on an FAA-approved airplane whose pilot was performing aerobatic maneuvers with a full load of passengers. When comparing the static test resu lts with the pattern of failure of the wing, investigators cast suspicion that the tension pads used to apply wing-bending loads likely stabilized the upper wing skin during tests. The stabilizing effect kept the upper wing skin from buckling during the tests. More static tests verified the theory; the retested wing failed below ultimate load.
b. When approving any static tests set up for thin-skinned structure, consider the effects of installing tension pads because they may contribute a stabilizing effect on the structure and bias the test results.
c. See 23.307, Proof of structure, for secondary structure guidance.
198. Is there any guidance available specifically for composite structures?
Yes. See Policy Statement PS-ACE100-2001-006, “Static Strength Substantiation of Composite Airplane Structure”.
23.603 Materials and workmanship (Amendment 23-23)
199. Is there any guidance available specifically for composite structures relative to materials and workmanship?
Yes. The following advisory circular and policie s on composites are related to this section:
• Advisory Circular 23-20, “Acceptance Guidance on Material Procurement and Process Specifications for Polymer Matrix Composite Systems”
• Policy Statement PS-ACE100-2001-006, “Static Strength Substantiation of Composite Airplane Structure”
• Policy Statement PS-ACE100-2002-006, “Material Qualification and Equivalency for Polymer Matrix Composite Material Systems”
• Policy Statement PS-ACE100-2004-10030, “Substantiation of Secondary Composite Structures”
• Policy Statement PS-ACE100-2005-10038, “Bonded Joints and Structures – Technical Issues and Certification Considerations”
200. Is there any policy available for glass fiber fabric? See AC 20-44, Glass Fiber Fabric for Aircraft Covering, for more guidance about this airworthiness standard topic.
201. Is there any policy available for floats constructed from composite materials? Both Technical Standard Order TSO-C27 and National Aircraft Standards NAS 807, “Twin Seaplane 69
Floats,” and AC 20-107A, Composite Aircraft Structure, contain guidance and criteria for floats constructed from composite materials. While NAS 807 does not specifically address composite materials, the standards in Section 3, Material and Workmanship, and in Section 4.2, Strength, are general enough to apply to materials other than the conventional aluminum materials normally employed.
202. Are there any special considerations for composite materials (relative to floats)?
Composite materials require special considerations for handling and storage that are not commonly required for metallic materials. These factors may affect material and process specifications.
203. Do any special factors apply for composite materials (relative to floats)? In addition to the 1.5 factor of safety, another factor could be applied for material variability when substantiated by tests. Environm ental effects include moisture, saltwater exposure, ultraviolet light, temperature variations, material composition, and geometric dimensions. These factors can be required under Section 4.2.1, Material Strength Properties, and Section 4.2.3.1, Special Factors, of NAS 807.
204. What kind of defect testing should I consider for composite materials (relative to floats)? Component static strength tests should include defects such as disbonds and voids.
They should also reflect impact damage, up to the threshold of detectability, for the inspection system in use during manufacturing and operations. Both defects and impact damage should be located in critical areas that are expected as a result of production assembly bonding processes, and in operational service conditions. The designer should identify the nature and size of such defects, and damage. The manufacturer should have an inspection system functioning during production and operational service. This system should ensure that defects do not reduce structural strength below ultimate load. Critical areas should include bonding of support strut attachment fittings to the basic structure. An alternative approach would be the use of mechanical fasteners, in critical areas, where bonding defects would cause critical loss of strength.
205. What are some considerations regarding corrosion when dealing with composite materials (relative to floats)? Galvanic corrosion may occur when unprotected metal is in contact with graphite composite material in a corrosive environment. Affected metal parts will need suitable protection. Cadmium plated metals will corrode. Use fasteners made of corrosion resistant materials (e.g., titanium or corrosion resistant steel). Also, the bonds of aluminum parts in contact with composite material (including fiberglass) may seriously degrade over time due to moisture absorption. Special treatment (e.g ., phosphoric anodizing) of the aluminum is necessary to maintain the original strength of the bond.
206. Are there any fatigue requirements for composite materials (relative to floats)?
Fatigue requirements appearing in NAS 807, Section 4.2.1, are minimal; they are also identical to the Civil Air Regulations (CAR) 3.307 general requirement for airframe construction. The FAA has generally not required fatigue testing or analysis to meet this requirement. Float service history has generally been satisfactory and more stringent fatigue requirements have not been applied to part 23 airplane landing gear, therefore, AC 20-107A, Section 6, Proof of Structure - Fatigue, does not apply. However, the manufacturer should develop instructions for appropriate tests or inspections to detect problems of hidden damage or delaminations. The manufacturer should also develop repair instructions for the composite material structures. For 70