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
float design, maintaining the integrity of watertight compartments is a special concern— disbonding or delaminations that cause inter-compartment leakage may reduce the level of safety intended by part 23, § 23.751 and CAR 3.371.
207. Are there any requirements for maintenance instructions (relative to floats)? There are no requirements for maintenance instructions under the TSO general requirements (refer to part 21, Subpart O) or in both TSO-C27 and NAS 807 specifications for seaplane floats.
However, § 21.50(b) requires that the applicant prepare Instructions for Continued Airworthiness according to § 23.1529 for each supplemental type certificate (STC) applied for after January 28, 1981.
208. Is there additional guidance for this topic?
a. See AC 20-107A for additional guidance about this airworthiness standard topic.
b. All composite structures that are critical to flight safety should be designed to be damage tolerant. If impractical, the applicant is referred to § 23.573(a)(6), Damage tolerance and fatigue evaluation of structure. The manufacturer should substantiate scatter factors.
Consider the following items when demonstrating the damage tolerance capability of structures critical to safe flight:
(1) Introduce manufacturing defects and realistic impact damage up to the threshold of detectability.
(2) Substantiate the static ultimate load retention capability with impact damage up to the threshold of detectability.
(3) Introduce initially detectable damage.
(4) Apply a statistically significant number of flight-by-flight spectrum repeatedload cycles to validate one lifetime or an inspection interval for operational service use. An inspection program suitable for operation and maintenance application needs to be developed by considering the damage growth.
cycling.
(5) Limit load retention capability should be demonstrated after repeated-load
(6) Environmental accountability should be included in the above demonstrations.
(7) Residual strength capability to withstand critical limit flight loads should be demonstrated. Demonstrate this with the extent of detectable damage consistent with the results of the damage tolerance evaluations and the maximum disbonds of bonded joints permitted by design. (The damage tolerance evaluations incl ude damage sizes ranging from small detectable damages to larger damage sizes that are possible in service.) Consider the two damages separately.
(8) All tests should be on actual composite material being used. The alternative use of any other composite material should be substantiated.
(9) Repair procedures may be part of the substantiation program and can be published in the Continued Airworthiness Section of the Maintenance Manual. This is not a regulatory requirement, but rather a highly desirable FAA goal.
(10) All ultimate static tests on structures critical to flight safety should be conducted on full-scale component articles. Environmental effects should be considered.
(11) All critical conditions should be substantiated by tests, or by analysis supported by tests.
209. Is there any guidance on solar and thermal effects when designing composite airplane structures?
a. The thermal environmental analysis should be based on a parametric study of the following data to identify the highest structural temperature:
Hour Ambient Temperature (°F) Solar Radiation (Btu/ft2 /hr) 1100 111 330 1200 114 355 1300 119 355 1400 122 330 1500 123 291 1600 124 231 1700 123 160 NOTE: The above temperature values woul d not be exceeded 99.9 percent of the time, as derived from MIL-STD-210C statistical data. For the above data, the wind speed was 14 feet per second, and the relative humidity was 3 percent.
b. The effect of cooling airflow may be considered. We recommend that, after heat soak at the critical condition, the airplane taxi, take off, and climb to 1,000 feet above sea level. The airplane should accelerate in level flight to:
(1) The lesser of the design maneuvering speed, (VA); or the aircraft operating speed limit (in § 91.117(b)) if maneuver loads are critical;
Or,
(2) The lesser of the design cruise speed, (VC), or the aircraft operating speed limit (in § 91.117(b)) if gust loads are critical.
NOTE: In a commuter category airplane, the design speed for maximum gust intensity, (VB), applies instead of the design cruise speed, (VC). The aircraft operating speed limit in § 91.117(b) is 200 knots. This applies to major structure and may not be applicable to certain structures, such as flaps and landing gear doors that would be subject to limit loads at an earlier time in the flight profile.
For a small airplane, a maximum taxi speed of 10 m.p.h. is recommended. A four-minute taxi-time would be reasonable.
23.605 Fabrication methods (Amendment 23-23)
210. Is there any guidance available specifically for composite structures relative to fabrication methods?
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”
211. Are there any known processes that may have detrimental effects on a material?
a. The FAA found the process of paint removal by the Plastic Media Blasting method to show varying degrees of decrease in fatigue life and an increase in crack growth rate. Reports submitted by Battelle, Columbus Laboratories, indicate variations were experienced, depending on the material, material thickness, and number of plastic bead blastings. The FAA is concerned about the detrimental effects that plastic bead blasting could have on the static and fatigue strength of the base metal(s) when removing the paint or protective coating.
b. To approve a process specification for Plastic Media Blasting paint removal, it is necessary to establish that the method is not damaging to the aircraft; therefore, as a minimum, the following parameters should be considered:
(1) Material variations, e.g., 7075-T6 alclad, 7075-T6 bare, 2024-T3 alclad, 2024 T3 bare, etc.;
(2) Paint, primer, and number of coatings to be removed;
(3) Plastic media size and type;
(4) Plastic media hardness;
(5) Nozzle pressure (maximum);
(6) Distance (nozzle to component);
(7) Angular nozzle displacement;
(8) Plastic media (mesh) flow level and nozzle diameter; and
(9) Dwell time.
c. Plastic media vendor suppliers should demonstrate that their materials could be supplied to the same standards (shown above).
d. Coupon tests for static strength and fatigue properties (including crack initiation and propagation) should demonstrate the compatibility of the material and the non-plastic media blasted material.
e. Blasting equipment users should demonstrate that the equipment would give the proper pressures and precise metering of the plastic media flow rate. The media separator should be capable of removing foreign particles from the reclaimed media using methods, such as the following:
(1) Vibrating screens;
(2) Magnetic separation;
(3) Electrostatic separation;
(4) Floatation method;
(5) Liquid gravity settling;
(6) Wet classification;
(7) Dry screening;
(8) Air separators;
(9) Gravity separators;
(10) Fixed chamber separators; and
(11) Mechanical separators or cyclone classifiers.
212. Is wood spar construction with Weldwood Plastic Resin Glue approved? Weldwood Plastic Resin Glue has been approved for wood spar construction in a few airplanes.
23.607 Fasteners (Amendment 23-48)
213. Is it appropriate to use self-locking nuts without a secondary locking mechanism?
a. Self-locking nuts, alone, should not be used in any system when movement of the joint may result in motion of the nut or bolt head relative to the surface against which it is bearing. Joint seizure (bearing, uniball, or bushing) does not have to be considered by this regulation when determining the relative motion of the parts in question, although it is advisable to do so. Suitable protection and material pr operties of the joint are required by part 23, §§ 23.609 and 23.613.
b, Self-locking castellated nuts, with cotter pins or lockwire, may be used in any system.
c. Self-locking nuts should not be used with bolts or screws on turbine-engine airplanes in locations where the loose nut, bolt, washer, or screw could fall or be drawn into the engine airintake scoop.
d. Self-locking nuts should not be used with bolts, screws, or studs to attach accesspanels or doors, or to assemble any parts that are routinely disassembled before or after each flight. This advice does not intend to exclude self-locking nut plates in these named applications. Nut plate designs permit the fastener to float, which is a desirable feature that is not provided by a non-floating fastener device.
214. Is there any intended difference between the terminology used in paragraph (a), “retaining devices,” and that used in paragraph (b), “locking devices”? No. The lack of consistency in the terminology between these two paragraphs was not intentional. The terminology “retaining devices” (in paragraph (a)), is identical in meaning and purpose to “locking devices” (in paragraph (b)), in this instance.
215. Is there any related guidance material on this subject? Yes. See AC 20-71, Dual- Locking Devices on Fasteners, for guidance about removable-fastener dual-locking devices for rotorcraft and transport category aircraft.
23.609 Protection of structure (Original)
216. Is there any policy available on this section as of January 1, 2007? No.
217. Is there any guidance available specifically for composite structures relative to protection of structure? Yes. See Policy Statement PS- ACE100-2001-006, “Static Strength Substantiation of Composite Airplane Structure”.