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
except for wing torsional stiffness criteria that specify indicated airspeed (IAS) be used.
However, at 14,000 feet in altitude, the difference between EAS and IAS is small (about 3½ knots).
239. When did flight flutter tests become a requirement? Amendment 23-48 made flightflutter tests a requirement. Before February 9, 1996, the date of the amendment, freedom from flutter, control reversal, and divergence could be shown by either a rational analysis, by flightflutter tests, or by simplified flutter prevention criteria. For all new type certification projects for which application was made after February 9, 1996, flight flutter tests are required to demonstrate compliance to § 23.629.
240. Where can I get a copy of AEER 45? Airframe and Equipment Engineering Report No.
45, “Simplified Flutter Prevention Criteria for Personal Type Aircraft,” by Robert Rosenbaum, and Civil Aeronautics Manual 3, Supplement No. 11, dated March 28, 1952, may be used for simple airplane designs to show compliance with flutter. Copies may be obtained by mail from the following: Manager, Standards Office (ACE -110), Federal Aviation Administration, Small Airplane Directorate, DOT Building, 901 Locust, Room 301, Kansas City, MO 64106.
241. How can I achieve the objective of paragraph (f) of this section?
a. The objective is to prevent airplane flutter from occurring after the failure, malfunction, or disconnection of any single element in the primary flight-control system, in any tab-control system, or in any flutter damper. Balancing the control systems and then showing that the airplane is free from flutter can achieve this. Alternatively, the objective can also be shown by doing the following:
(1) Incorporating a structural fail-safe design throughout the entire flight-control system and then demonstrating that the airplane is free from flutter; or
(2) Incorporating a combination of structural fail-safe designs and balanced-control system.
b. If a hinge pin single failure would allow the pin to fall out of the hinge, create hinge design features to prevent the pin from separating.
242. Is the fail-safe design criterion of § 23.572(a)(2) an acceptable method of compliance for flutter? No. Section 23.629(f)(2) requires the failu re, malfunction, or disconnection of any single element be considered. Both control-surface balance and dual-load, path-tab-system designs have been judged as meeting the airworthiness requirement for irreversible systems.
243. Why isn’t the fail-safe design criterion of § 23.572(a)(2) an acceptable method of compliance for flutter? The fail-safe criterion in § 23.572(a)(2) imposes a static ultimate-load factor of 75 percent of the critical limit-load factor at design cruise speed, VC. This criterion is inadequate for flutter substantiation of dual-load path primary-control systems, or tab-control systems because of the lower speeds and lower structural loads imposed.
244. Is there any additional guidance available for this section? Yes. See AC 23.629-1B, Means of Compliance with Section 23.629, “Flutter,” for additional information about this airworthiness standard topic.
WINGS
23.641 Proof of strength (Original)
245. Is there any policy available on this section as of January 1, 2007? No.
CONTROL SURFACES
23.651 Proof of strength (Original)
246. Is there any policy available on this section as of January 1, 2007? No.
23.655 Installation (Amendment 23-45)
247. Is there any policy available on this section as of January 1, 2007? No.
23.657 Hinges (Amendment 23-48)
248. Is there any additional guidance available for hinges? See 23.393, Loads parallel to the hinge line, for additional guidance about hinges. Also see 23.651, Proof of strength, for information that may affect hinges.
23.659 Mass balance (Original)
249. Is there any policy available on this section as of January 1, 2007? No.
APPENDIX A TO PART 23—Simplified Design Load Criteria A23.1(a) References:
1. Clousing, Lawrence A. and Turner, William N.: Flight Measurements of Horizontal Tail Loads on a Typical Propeller-Driven Pursuit Airplane During Stalled Pull-Outs at High Speed. RMR (WR A - 81), May 1944.
2. Matheny, Coyce E.: Comparison Between Calculated and Measured Loads on Wing and Horizontal Tail in Pull-Up Maneuvers. ARR L5H11 (WR L-193), Oct.
1945.
3. Garvin, John B.: Flight Measurements of Aerodynamic Loads on the Horizontal Tail Surface of a Fighter-Type Airplane. TN 1483, Nov. 1947.
4. Sadoff, Melvin and Clousing, Lawrence A.: Measurements of the Pressure Distribution on the Horizontal-Tail Surfaces of a Typical Propeller-Driven Pursuit Airplane in Flight. III—Tail Loads in Pull-Up Push-Down Maneuvers. TN 1539, Feb. 1948.
5. NACA Flight Research Maneuvers Section: Flight Studies of the Horizontal-Tail Loads Experienced by a Modern Fighter Airplane in Abrupt Maneuvers. Report 792, 1944.
6. AC 23-15A, Small Airplane Certification Compliance Program The first paragraph of the Reference 1 CONCLUDING REMARKS reads:
“With the test airplane operated within maneuvering limits which were considered safe by design specifications in use at the time the airplane was designed, units loads were measured on the stabilizer which were not only considerably in excess of the design unit loads, but which occurred in a direction opposite to the design loads.”