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”.
23.611 Accessibility (Amendment 23-48)
218. Is there any policy available on this section as of January 1, 2007? No.
23.613 Material strength properties and design values (Amendment 23-45)
219. Is there any policy available on this section as of January 1, 2007? No.
220. How should the structural test article be configured?
a. Ideally, the structural test article (a whole wing, an empennage, a fuselage, etc.)
would contain all elements that are made of specification guaranteed-minimum-strength materials. Also, each element’s physical dimensions (geometry) would be at the nominal size, plus or minus specified tolerances, to conservatively represent the least strength or least stiff part that could be used according to approved design data (drawings, specifications, stress, or structural analyses).
b. Parts (elements) are manufactured and delivered to nominal sizes within tolerances.
This means that they will either deliver minimal performance or more than promised.
221. How do I determine the strength of the airplane’s structure?
a. There are, fundamentally, four actions an airplane designer can take to determine the strength of the airplane’s structure:
(1) The designer can analyze the airplane structure to both limit and ultimate load conditions, using guaranteed minimum-strength-material properties and conservative geometric characteristics;
(2) The designer can test the airplane structure to limit loads and then analyze the airplane structure to ultimate loads;
(3) The designer can test the airplane structure to limit loads and, later, to ultimate conditions; and
(4) The designer can test the airplane structure to beyond ultimate load conditions.
b. This last option is usually chosen to determine excess strength or growth capabilities.
It also exceeds the minimum FAA airworthiness standards for which compliance should be shown. Certain airworthiness standards require one of these methods instead of the others.
c. An applicant should substantiate that the strength properties of components used in structural tests are such that subsequent components used in airplanes presented for certification will have strengths equal to or exceeding the demonstrated strength of the tested components.
d. If the applicant chooses to demonstrate strength capability by tests of structural components, the applicant should substantiate that the strength of the tested component conservatively represents the strength of subsequent production components. Substantiating data might include quality control data, material and process specifications, material certifications, coupon sampling tests, or other appropriate information.
e. An applicant may also apply material correction factors to the applied test loads to account for material variability. Applicants should use material correction factors for ultimate load tests of single load path critical flight structure and for fail-safe tests of dual load path critical flight structure with one load path failed. Applicants do not need to use material correction factors for limit load tests or for ultimate load tests of fail-safe designs where loads from one failed component are distributed to and carried by two or more remaining components.
222. Is there any guidance available specifically for composite structures relative to material strength properties and design values? Yes. See Policy Statement PS-ACE100 2001-006, “Static Strength Substantiation of Composite Airplane Structure”.
223. Are there any other useful references for this section? Yes. See AC 20-33B, Technical Information Regarding Civil Aeronautics Manuals (CAM’s) 1, 3, 4a, 4b, 5, 6, 7, 8, 9, 13 and 14;
CAM 3, paragraphs 3.174-1 and 3.301-1; and CAM 4a, paragraph 4a.230.
224. Are there any other part 23 rules related to the material correction factors?
a. Yes. See §§ 23.305(a) and (b); 23.307(a); 23.603(a); 23.613(c); and, before Amendment 23-45, § 23.615(a) and (c).
b. Section 23.305, paragraphs (a) and (b), Strength and deformation requirements, § 23.307(a), Proof of structure standards, and § 23.603(a)(1), Materials and workmanship regulations has one objective. The objective is that the lowest strength conforming airframe produced to a set of FAA-approved type design data will comply with the requirements of § 23.305.
23.615 Design properties (Amendment 23-45) [Removed]
23.619 Special factors (Amendment 23-7)
225. Can you summarize the various special factors and where they are located? Yes. This table (see Figure 9) summarizes various special factors and where they are located. Read the appropriate airworthiness standards to determine when these special factors replace the factor of safety and when they multiply the factor of safety.
FIGURE 9
SUMMARY OF VARIOUS SPECIAL FACTORS § 23.303 § 23.365 § 23.621 § 23.623 § 23.625 Factor of Safety
1.5 Burst pressure factor
1.33,
1.67 with Special Conditions Casting factor
1-1.25
1.25-1.5
1.5-2.0
2.0-higher Bearing factor FS = 6.67 FS ≥ 3.33
2.0≤FS≤3.33 Fitting factor
1.15
1.33
226. What is the difference between a Factor of Safety and a Margin of Safety, and how is the Factor of Safety determined relative to § 23.303 – Figure 9?
Traditionally, a Factor of Safety (FS) is defined as the maximum allowable load (P) divided by the maximum applied load (p);
FS = P/p A Margin of Safety (MS) is defined as the Factor of Safety minus one;
MS = FS – 1 = (P/p) –1 The Factor of Safety as discussed in Figure 9, relative to § 23.303 refers to the multiplying factor used to determine the ultimate loads from the limit loads (§ 23.301(a)). In other words:
Ultimate Load = [Limit Load] * [FS] = [Limit Load] * [1.5 (minimum)]
227. Is there any guidance available specifically for composite structures relative to special factors? Yes. See Policy Statement PS-ACE100-2001- 006, “Static Strength Substantiation of Composite Airplane Structure”.
23.621 Casting factors (Amendment 23-45)
228. Is there any policy available on paragraph (a), (b) or (e) of this section as of January 1, 2007? No.
229. Can you summarize the inspection requirements for a given casting factor as dictated by paragraphs (c) and (d) of this section? Yes. The following logic tree (Figure 10) permits a reader to quickly determine the inspection requirements associated with a chosen casting factor.
FIGURE 10
CASTING FACTOR INSPECTION REQUIREMENTS
CRITICAL NON-CRITICAL
CASTING FACTORS
1.25 - 1.50 > 1.50 > 2.0 1.00 1.25 - 1.50 1.50 - 2.0 > 2.0 3 sample test* 3 sample test* 100% visual 100% visual 100% visual 100% visual 100% radiographic 100% radiographic100% approved NDI 100% mag. part.
or dye pent.
100% mag. part.
or dye pent.
100% mag. part.
or dye pent.
or equivalent or equivalent 100% visual * Ultimate load corresponding to a casting factor of 1.25 (U = L x 1.5 x 1.25 = 1.875L);
deformation requirements of § 23.305 at a Load = 1.15 x L, where U = ultimate and L = limit load.
23.623 Bearing factors (Amendment 23-7)
230. Is there any policy available on this section as of January 1, 2007? No.
23.625 Fitting factors (Amendment 23-7)
231. Is there any policy available on paragraph (a), (b) or (c) of this section as of January 1, 2007? No.
232. Should the 1.33 fitting factor be applied to the dynamic emergency landing conditions as well as the static emergency landing conditions (see part 23, §§ 23.562 and 23.561, respectively)? The seat attachment fittings must be included in the dynamic test and, therefore, there is no need to apply a fitting factor.
23.627 Fatigue strength (Original)
233. Is an actual fatigue evaluation required? Regarding the fatigue evaluation (under part 23, § 23.627), the FAA has interpreted this standard as requiring only that the manufacturer exercise good design practice when avoiding severe stress concentrations, but not requiring a fatigue evaluation in itself.
234. What is meant by the “fatigue limit” (also known as the endurance limit)? When the results of a fatigue test are plotted on an S-n diagram (stress versus number-of-cycles to failure), the fatigue limit is the constant stress level reached at a high number of cycles. Below that stress level, failure is not expected to occur. Aluminum alloys may not show a clearly defined fatigue limit. Guidance provided in FAA report AFS-120-73-2, Fatigue Evaluation of Wing and Associated Structure on Small Airplanes, defines the fatigue endurance limit as the stress at 3 x 107 cycles.
235. Is there additional guidance available for this section? Yes. See AC 23-13, Fatigue and Fail-Safe Evaluation of Flight Structure and Pressurized Cabin for Part 23 Airplanes, for additional guidance about this airworthiness standard topic.
23.629 Flutter (Amendment 23-48)
236. When did the regulations start requiring flutter clearance to 1.2 VD for the rational analysis? Flutter clearance, using rational analyses, has been required to 1.2 VD since Amendment 23-7 became effective on September 14, 1969. Clearance, for true or equivalent airspeed, depends on whether appropriate density correction factors are included in the analysis.
Before Amendment 23-7, including Civil Air Regulations (CAR) 3, it was required to show flutter-free operation to 1.0 VD only.
237. When was the first authorization of the use of AEER 45 as a means of meeting flutter prevention requirements? Authorization to use AEER 45 (corrected February 1952) as a means of meeting the flutter prevention requirements of CAR 3.311 first appeared in Civil Aeronautics Manual (CAM) 3.311-1 on March 13, 1952. The simplified criteria do not specify applicable airspeed or altitude limits.
NOTE: Existing copies of AEER 45 are undated and do not indicate if they are corrected.
The correction appears on page 8 in the equation at paragraph 3(a). The corrected equation constant is 48.
238. What is history of the airspeed and altitude (if any) limitations for the use of AEER 45?
a. Under CAM 3.311-1, no airspeed or altitude limits were given.
b. On December 1, 1978, Amendment 23-23 established an airspeed limit less than 260 knots equivalent airspeed (EAS) at altitudes below 14,000 feet and less than Mach 0.6 at altitudes at and above 14,000 feet. See below.
Amendment 23-23 VD < 260 EAS below 14,000 feet and MD < Mach 0.6 at and above 14,000 feet
c. On January 8, 1979, AC 23.629-1, Means of Compliance with Section 23.629, “Flutter,” set new lower airspeed limits for AEER 45 to a design dive speed less than 200 m.p.h.
EAS at altitudes below 14,000 feet. See below.
AC 23.629-1 VD < 200 m.p.h. EAS below 14,000 feet
d. On October 23, 1985, FAA published revised AC 23.629-1A and changed the airspeed limits for AEER 45 to a design dive speed less than 260 knots EAS at altitudes below 14,000 feet. See below.
AC 23.629-1A VD < 260 EAS below 14,000 feet
e. On September 7, 1993, Amendment 23-45 again changed the Mach number for AEER 45 to a design dive Mach number less than Mach 0.5 to closely agree with the calculated Mach number at 14,000 feet and 260 knots EAS. See below.
Amendment 23-45 VD < 260 EAS and MD < Mach 0.5 at and above 14,000 feet
f. Note that part 23 never established an altitude limitation on the applicability of AEER
45. Speed units, although not addressed in AEER 45, have historically been taken as EAS— 82