(5) It must not have a noticeable effect on the internal load distribution.
(6) It must not have a noticeable effect on handling, performance, or flutter characteristics. Areas of special attention in clude deformations that may have altered the wing twist or stiffness properties.
(7) It must not weaken or damage the primary structure. No other permanent set may occur with repeated applications of limit load.
(8) It must not prevent compliance to any regulation.
25. I have permanent set in the vertical deflection of my airplane wing, at the wing tip, after limit load testing. The permanent set is approximately 5 percent of the maximum vertical deflection measured at limit load. Is this considered “detrimental” permanent deformation?
a. Not necessarily. Permanent set in wing tip deflection can result from permanent deformation or joint slippage. It is not necessarily true that a large permanent set implies that the structure has yielded. For example, a small deformation that occurs near the wing centerline can translate into a large deformation at the wing tip.
b. You should identify the location and size of the deformation that caused the permanent set. Wing tip deflection caused by bolted joint slippage is not necessarily an example of detrimental, permanent deformation.
c. You must evaluate the wing tip deflection and its root cause, joint slippage, or material deformation, by the qualitative criteria listed in the question above to determine its acceptability.
26. Is there any guidance available specifically for composite structures relative to strength and deformation?
Yes. See Policy Statement PS-ACE100-2001- 006, “Static Strength Substantiation of Composite Airplane Structure”.
23.307 Proof of structure (Original)
27. What should I consider for increases in maximum weight, in maximum landing weight, or in maximum zero fuel weight?
a. Due to changes in the operational requirements of an owner or operator, the need arises to modify and substantiate the structure for an increase in maximum weight, in maximum landing weight, or in maximum zero fuel weight. Any one of these changes affects the airplane basic loads and structural integrity and may affect the limitations and performance. Two examples follow:
(1) When increasing the original airplane maximum weight, special considerations are necessary. (See an acceptable method in paragraph 28.)
(2) When replacing a piston engine with a turbopropeller engine, one must consider that:
(a) Jet fuel weighs as much as 17 percent more than avgas, and
(b) Airplane total fuel quantity often must increase.
(c) Therefore, the maximum zero fuel weight changes. These kinds of modifications should be investigated to verify that either:
1 The critical loads have not increased, or 2 The loads that have increased are capable of being carried by the existing or modified structure.
28. What is one acceptable method for showing compliance for a weight increase?
a. Prepare a compliance checklist. It may be an advantage to the applicant to identify the airworthiness standards affected by the proposed weight increase. Both the applicant and the FAA Aircraft Certification Office personnel will have a better idea of what technical showings of compliance need to be made.
b. Identify the critical flight, landing, and ground loads. The loads may either be obtained from the existing type certificate data (if they are made available to the applicant by the holder of the type certificate) or the applicant may derive them. Derived loads should be verified to produce essentially the same results as those used for the original type certification work.
c. The airplane structural design airspeeds (see part 23, § 23.335) should be re evaluated to determine if the selected airspeeds are adequate at the increased design weights. Lateral gust conditions (see § 23.443) should reflect any changes in yaw moment-of-inertia resulting from revised mass distributions. An increase in airplane weight often causes an increase in wing loading. When pitch and roll inertia affect the airplane loads, due to increasing or redistributed mass, examine these also.
d. Confirm the support of all structures affected by load increases, however small. This may be accomplished after the critical loads are identified. Stress analyses, static tests, or a combination of both proof-of-strength methods may be used to prove the structures capable of sustaining ultimate loads (see §§ 23.307 and 23.641). The FAA encourages, but does not require, the applicant to conduct proof-of-strength tests to both limit and ultimate load conditions—and beyond (for the additional knowledge gained and growth capability). If static tests are used as the proof-of-strength method, the structure should be inspected for detrimental permanent set following the removal of the limit load(s). Any structure that shows detrimen tal permanent set requires some redesign and retest. When using analyses as the proof-o f-strength method, the material yield-point stress should not be exceeded when a limit load is analytically applied.
e. If an airplane was initially certified with maximum landing weight equal to maximum weight, and if an increase in the maximum weight is applied for, the applicant may take advantage of the five percent difference between landing weight and maximum 17
weight permitted by § 23.473(b). In that case, resubstantiation of the landing gear is not required for the first five percent of the weight increase as long as the airplane center of gravity (c.g.) remains within the original type certificate limits. (See subparagraph "k” below).
f. Until Amendment 23-48, part 23 did not require that a maximum zero fuel weight be established (see § 23.343, Design fuel loads). However, for airplane designs with wing fuel tanks, the minimum fuel condition may produce the highest wing-bending moment; it also affects the wing torsional moment. Evaluate these conditions during the compliance phase of the project (see § 23.301).
g. Verify weight distribution and c.g. design changes. Weight increases or relocated mass items, which change the overall mass distribution, also change the airplane c.g. at empty weight, maximum weight, and weights in between the two extremes. These kinds of design changes should be carefully investigated for their affects on the original weight versus the c.g. envelope. The designer should consider the effects of depletable payload items, like fuel; account for c.g. shifts; and calculate the influence these may have on the whole airplane.
h. Examine the effects of design changes to the airplane structural damping and speeds. Changes to the maximum weight, maxi mum zero fuel weight, airplane structural stiffness and the distribution of mass need to be examined with the effects of flutter in mind. Ground vibration survey tests permit the identification of airplane structure nodes, modes, and corresponding frequencies. From these, the airplane flutter characteristics can be analytically estimated. AC 23.629-1A, Means of Compliance with Section
23.629, “Flutter,” thoroughly discusses this subject.
i. Re-evaluate the fatigue strength or fail-safe strength estimates. A fatigue or fail-safe evaluation should be accomplished if the certification basis airworthiness standards include §§ 23.571 or 23.572, and they should be prudently considered in every design. This evaluation may indicate that cy clic tests should be run on a fatigue test specimen with the modifications incorporated.
j. The certifier should revise the Airplane Flight Manual and the Instructions for Continued Airworthiness. The revisions or supplements to these manuals should reflect any pertinent changes in weight and balance data, performance, flight procedures, maintenance procedures or practices, life-limited parts, etc. Note that the Maintenance Manual and the Instructions for Continued Airworthiness may also be affected.
k. Identify the critical flight, landing, and ground loads. The loads may be obtained from the existing type certificate data if the holder of the type certificate makes them available to the applicant. The designer (or modifier) can calculate the loads.
Verify calculated loads—the olderCARs often show acceptable methods to calculate loads that an airplane designer may use.
l. A decision about whether the maximum weight increase is “small” or “substantial” does not affect the application of guidance in this AC.
29. Is there any information available on wood airplane structures?
a. When designing aircraft that contain wooden structures, refer to “Design of Wood Aircraft Structures,” Army-Navy-Civil ANC-18 Bulletin. This bulletin was 18
prepared by the Forest Products Laboratory, Forest Service, United States Department of Agriculture, and ANC-23 Panel on Sandwich Construction for Aircraft, Subcommittee on Army-Navy-Civil Aircraft Design Criteria (Aircraft Committee Munitions Board). The following are general interest information items:
(1) In the continental United States, ANC-18 Bulletin, Section 2.1, indicates that 15 percent moisture is considered acceptable for wood used in airplane design. The moisture content expected in service would obviously depend on the geographic region of the earth where the aircraft is operated. However, where the relative humidity is expected to be greater than 90 percent for an extended time (the Tropics), 20 percent moisture content should be assumed.
NOTE: As moisture content increases, wood strength decreases.
(2) The FAA knows of no protective coating that will prevent wood from reaching an equilibrium condition in moisture content in ambient conditions.
(3) The FAA recommends that tests be conducted whenever the design is in question.
(4) Specific test requirements for strength due to moisture content, when proof of strength is shown by tests of Civil Air Regulations (CAR) 3.174-1(b), are not outlined and much is left to the FAA certification engineer. Test conditions should be reasonable and, without other data, the recommendations of ANC-18 Bulletin should be used (see Section 3.0111). Certification records of four previous successful airplane designs show that the moisture content of the test articles was not documented during tests.
(5) The FAA has no published methods or procedures about the effects of moisture content on the strength of wooden structures. Refer to “Design of Wood Aircraft Structures” (ANC-18 Bulletin) for methods and procedures that are acceptable to the procuring or certificating agency (see Section 1.0).
30. Are there any precautions I should take when placing tension pads for static tests? Yes. There was a fatal accident involvi ng an airplane that suffered a wing upperskin failure in buckling. When tension pads are used to apply loads during static tests, they may stiffen thin-skinned structure and bias test results non-conservatively. The following factors should be observed and documented in the test report regarding tension pad use:
a. Type: Round rubber pad with metal back, square rubber pad with metal back, round canvas pad, square canvas pad, etc.
b. Size and number: Percentage of lifting surface covered (outboard or inboard, fore or aft).
c. Location: Upper or lower lifting surface (the surface is a main tension or compression field); tension pad nearness to a spar, rib, or stringer structural element;
effects of fuel pressure loads.
31. Do you have any information about alternate means of compliance for seat structural requirements on agricultural airplanes? Yes. See 23.562, Emergency landing dynamic conditions (paragraph 174), for information and guidance about dynamic seat requirements.
32. Can I conduct flight tests as a substantiation method for secondary structure?
See question number 22.
33. What are the failure criteria for secondary structure? When secondary structure experiences forces that vary with angle of attack, or yaw conditions, demonstrate these structures to the same failure criteria as a primary structure by:
a. A support limit load without detrimental permanent deformation, and
b. A support ultimate load without failure for 3 seconds.
This can often be accomplished using a simple conservative analysis.
34. What are some of the methods? Structural analyses or static tests, or a combination of these, may be used to show compliance with both the limit and ultimate load conditions. The critical points on the flight envelope (V-n diagram) should be examined. The basic loads may be obtained from flight tests, wind tunnel tests, derived data from similar airplane designs, or by conservative analyses. Engineering judgment is involved. Some pertinent consid erations include the following:
a. Wind tunnel basic loads: It may be necessary to apply a conservative factor to ensure the confidence of the FAA project engineer in the full-scale loads. Factors to consider include Reynolds numbers, flow similarities between the tunnel-model and the full-scale airplane, and load measuring methods.
b. Three methods to obtain flight-test loads:
(1) When either compressibility or elasticity are negligible (or both), the first method is 1g flight-test data taken over the range of angle of attack that correspond to the airspeed critical points on the flight envelope (V-n diagram). This first method is also known as “scaling by dynamic pressure.” These data may then be corrected to the flight conditions on the V-n diagram (i.e., corrected for dynamic and static pressure).
(2) When compressibility effects are significant—and even if they are not, the second method is to repeat the first method at several higher load factors (within the limit load capabilities of the particular airplane design). One can measure loads in 1g level flight, and steady turning flight, at a series of constant airspeeds. This technique produces a family of load curves that can be extrapolated to limit load conditions. Obtain the desired Mach number by prudently choosing the airspeed and altitude.
(3) A third method is to fly roller coaster maneuvers (pull-ups and pushovers) to produce load factors above and below 1g conditions. This method demands accurate measuring techniques for loads and flight conditions during transient flight conditions. The transitory nature of the data sometimes gives more data problems than value.
c. When static pressure influences structural loads: The structure experiences a load increase due to the external air load when the external static pressure is different from the internal static pressure. This often happens on secondary structures like engine cowls, windshields, and windows.
d. Also see the question regarding the use of wind tunnel data under section
23.301.
35. What proof-of-strength factors should be considered for airplane design changes that may affect structure? (Also see 23.321, General (flight loads).)
a. When loads increase, the strength of structure is affected by shear, bending, and torsion—not just one of these conditions.
b. In some aircraft certifications, the structure was proven by tests only. Any modifications to these aircraft that alter the loads⎯or the load paths⎯must be assessed to determine if the change is significant enough to require retests. It is inappropriate to assume that additional strength resides in the structures beyond the values proven by tests.
c. For stressed-skin wings, if analyses are used to justify the strength between limit and ultimate load conditions, strain-gauge data coupled with panel-buckling stress data may be used to show that the strength extrapolation is reasonable and correct.
d. Consider the effects of stress concentration factors.
e. A small increase in design maximum weight may cause a severe reduction of fatigue life (see 23.572, Metallic wing, empennage, and associated structures).
f. Any load increase on the wing, tail, or landing gear structures, or passenger, cargo or equipment areas affects the fuselage.
g. Identify specific materials, dimensions, and processes used in the design (see part 21, § 21.31, and part 23, § 23.603).
h. A previously certificated airplane requires structural proof-of-strength substantiation to the Certification Basis regulations when:
(1) Someone puts a turbine engine on a previously certificated reciprocating-engine airplane; or when
(2) A turbine engine substitutes as a single-power-source for two or more reciprocating engines.
36. I want to get an STC for a gross weight increase, but I am not the Original Equipment Manufacturer (OEM), so I do not have access to the original Type Certificate data. Do I have to assume that the margin of safety of all structural components is zero? Yes. Assume that the margin is zero unless you can confirm two things. First, you corroborate that the proof of strength was demonstrated by a combination of analysis and tests. Second, you confirm that these tests showed a margin of safety greater than zero in an "as delivered" condition.
37. Is there any guidance available specifically for composite relative to proof of structure?
Yes. See Policy Statement PS-ACE100-2001- 006, “Static Strength Substantiation of Composite Airplane Structure”.
FLIGHT LOADS
23.321 General (Amendment 23-45)
38. Is there any policy available on paragraph (a) of this section as of January 1, 2007? No.
39. Is there any policy available on paragraph (b) of this section as of January 1, 2007? No. However, you can see 23.307 for guidan ce about airplane weight increases.
See 23.471 for guidance about ramp weight and take-off weight.
40. Is there any policy available on paragraph (c) of this section as of January 1, 2007? No. However, for information a bout canard and tandem wing airplane configurations, see 23.421, balancing loads for horizontal stabilizing and balancing surfaces. Also, laminar flow aerodynamics information may be found in 23.21, proof of compliance.
41. Can you show a graphical relational image of the airworthiness standards that concern airplane loads? Yes. See Figure 2 and Fi gure 3 on the next two pages.
LO/OE/₺ UNSYMMETRICAL §23.347 Vertical surfaces Horizontal stabilizing and balancing surfaces Vertical surfaces Ailerons Rolling Conditions Engine torque Side load on engine mount Gyroscopic and aerodynamic loads combined Unsymmetrical loads due to engine failure ACE-100, E. Gabriel, May 1996 FLIGHT LOADS- Part 23[ xxx| Amendment 45 or winglets Pressurized cabin loads Canard or tandem wing configurations Outboard fins Speed control devices Wing flaps Wing flaps High lift devices High lift devices SYMMETRICAL §23.331 Clean Airplane Horizontal stabilizing and balancing surfaces Balancing horizontal tail load Limit load factor Pitching Wing Flaps Slipstream Effects Rear lift truss
GUST MANEUVER ENGINE OTHER
FIGURE 3
FLIGHT LOADS Flight Loads - Part 23 Amendment 45 UNSYMMETRICAL § 23.347 Vertical Surfaces Lateral gust: + 50 fps @ Vc [§ 23.443(a)] Commuter category [§ 23.443(b)] Gusts normal to plane of symmetry Vc, Vo clean airplane high lift devices Horizontal Stabilizing and Balancing Surfaces [§ 23.445] [§ 23.427] Loads from gusts combined with yawing and slipstream effects, clean airplane and with high lift devices Vertical Surfaces [§ 23.441] - @ V, Yaw, sideslip. and [§ 23.445(d)] rudder deflection Outboard Fins or Ailerons [§ 23.445] Abrupt maximum control movement Winglets @VA. Control deflection requirements @Vc and Vp Rolling Conditions [§ 23.349] - Wing and wing bracing Airload Distribution Condition (See § 23.333) 100%/70% to 75% Normal, Utility, Commuter 100%/60% A and F Wing loads due to aileron deflections § 23.445 Engine Torque [§ 23.361] - Combined with symmetrical limit loads @ VA Side Load on Engine Mount [§ 23.363] Gyroscopic and Aerodynamic Loads [§ 23.371] - Pitching and yawing, applies only to turbine installations Unsymmetrical Loads Due to Engine Failure [§ 23.367] - Turboprops only Pressurized Cabin Loads, combined with flight loads [§ 23.365] Canard or Tandem Wing Configurations [$ 23.302] Category Acrobatic [§ 23.373] Speed Control Devices SYMMETRICAL § 23.331 High Lift Devices Clean Airplane [§ 23.345] Discrete Vertical Gusts [§ 23.333(c), § 23.341] + 25 ft/sec vertical 25 ft/sec head-on + 50 fps @ Vc and + 25 fps @ Vo Wing Flaps + 66 fps @ VB [§ 23.457(a)] (commuter category only) Horizontal Stabilizing and Balancing Surfaces [§ 23.425] Clean airplane and with high lift devices Balancing Horizontal Tail Load [§$ 23.331, 23.421] Limit Load Factor [§ 23.337] Normal or Commuter Category n = 3.8* Utility Category n = 4.4 Acrobatic Category n = 6.0 *May reduce for W > 4,118 Ibs. High Lift Devices [§ 23.345] Pitching: Checked and Unchecked n = 2.0 g Applies to horizontal stabilizing and Wing Flaps balancing surfaces [§ 23.423] [§ 23.457(a)] Abrupt maximum control input @ VA Wing Flaps Slipstream Effects, n = 1.0 [$ 23.457(b)] Rear Lift Truss, reverse air flow [§ 23.369] ANTONE
23.331 Symmetrical flight controls (Amendment 23-42)
42. Is there any policy available on this section as of January 1, 2007? No.
23.333 Flight envelope (Amendment 23-34)
43. Do I have to revise the velocity-load factor (V-n) diagram for Supplemental Type Certificates (STC) involving engine changes or modifications? It is not an automatic requirement. This question usually arises because some foreign certifying agencies require the development of a new V-n diagram if the original minimum design cruising speed (VC) was defined by the wing loading (W/S) [reference: part 23, § 23.335(a)(1) and (2)], instead of the maximum level flight speed (VH) [reference: part 23, § 23.335(a)(3)]. However, we do not require a new V-n diagram unless the applicant wishes to increase VC.
44. Is there any policy available on paragraph (b) or (c) of this section as of January 1, 2007? No.
23.335 Design airspeeds (Amendment 23-48)
45. When applying design airspeed criteria, how may the designer establish a minimum design cruising speed?
a. The designer may establish the minimum design cruising speed, VC min, according to the following:
W33 ≤ V ≤ 0 9 . V , for normal, utility and commuter category C min H S and W ≤ V ≤ 0 9 . V , for acrobatic category36 C min H S 26
b. Both minimum and maximum design cruise speed values, VC min and VC max, may be chosen with the following understanding:
VC ≤ 0.9VH
46. Why would I want to define VC as equal to 0.9 VH? Use this definition if you are designing an airplane that is capable of a sustained speed (VC) higher than that obtained by using the wing loading (W/S) formula.
47. Is there any policy available on paragraph (b) of this section as of January 1, 2007? No.
48. What is the design maneuvering speed VA?
a. The design maneuvering speed is a value chosen by the applicant. It may not be less than Vs√ n and need not be greater than Vc, but it could be greater if the applicant chose the higher value. The loads resulting from full control surface deflections at VA are used to design the empennage and ailerons in part 23, §§ 23.423, 23.441, and 23.455.
b. VA should not be interpreted as a speed that would permit the pilot unrestricted flight-control movement without exceeding airplane structural limits, nor should it be interpreted as a gust penetration speed. Only if VA = Vs √n will the airplane stall in a nose-up pitching maneuver at, or near, limit load factor. For airplanes where VA>VS√n, the pilot would have to check the maneuver; otherwise the airplane would exceed the limit load factor.
c. Amendment 23-45 added the operating maneuvering speed, VO, in § 23.1507.
VO is established not greater than VS√n, and it is a speed where the airplane will stall in a nose-up pitching maneuver before exceeding the airplane structural limits.
49. Is there any policy available on paragraph (d) of this section as of January 1, 2007? No.
23.337 Limit maneuvering load factors (Amendment 23-48)
50. Is there any policy available on this section as of January 1, 2007? No.
51. Are there any related sections? Yes. See Section 23.423, Maneuvering loads.
23.341 Gust load factors (Amendment 23-48)
52. Is there any policy available on this section as of January 1, 2007? No.