various computer codes available that use computation fluid dynamics (CFD), such as, VS-Aero, MG-Aero, and TRANAIR.
a. The National Advisory Committee for Aeronautics (NACA) Technical Report 572 reports on the Anderson method. The wings under consideration covered a complete range of taper ratios and aspect ratios from 2 to 20. It does not say that 20 is a limit.
b. Reference 1 (page 228) states that NACA Technical Report 585 contains an exact method.
c. The “Fourier Series Method” is described in Reference 1 on pages 233 to 242.
This is the method in ANC-1 (which has tabular forms for ease of calculation). This method uses lifting line theory, which is good for conventional unswept wings with aspect ratios greater than 5 or 6 (see Reference 1, page 247).
d. Weissenger's “Method” (Reference 2) is applicable to straight or swept wings of low or high aspect ratio. This is a modified lifting theory method.
e. Schrenk's “Approximation” basically averages the lift forces obtained from an elliptical lift distribution with those obtained from a planform lift distribution. This approximation is accurate for wings that approach an elliptical planform (Reference 1, page 224). This method is contained in Civil Aeronautics Manual (CAM) 04, Appendix IV. The Limitations, in Section 6, state that it applies to the normal range of aspect ratios (from 5 to 12).
f. Reference 3 (page 14) lists NACA Technical Reports 572, 585, and 606. The first two of these technical reports are discussed in Items a and b above. Technical Report 606 is titled “Empirical Corrections to the Span Load Distribution at the Tip.”
Correction is only necessary if the wing is tapered less than 2:1 and has a blunt tip.
g. Reference 4 lists several references: 3.5 through 3.12 (see Enclosure 3).
h. In the paper titled “Application of Microcomputer Software to the Aerodynamic Design of a Motorglider” Technical Soaring, October 1993, the lifting line theory FORTRAN program in Reference 5 (pages 159-164) was used (on an Apple Macintosh Plus computer) to calculate the lift distribution of a 17-meter (55.76 foot) wing (aspect ratio unstated). A comparison wa s made of the lift distribution calculated by Schrenk's approximation.
i. Reference 6 contains a vortex lattice FORTRAN program.
j. Reference 7 is an aeroelastic supplement for the NASTRAN finite element program.
k. NACA TN 3030 titled, “A Method for Calculating the Subsonic Steady-state Loading on an Airplane with a Wing of Arbitrary Planform and Stiffness” includes both 6
the aeroelastic effects and the ability to base the span loading on linearized wind tunnel wing section data. It also allows for correc ting wind tunnel model elastics and jig twist.
FIGURE 1
SPANWISE WING-LOADING METHODOLOGIES
METHOD APPROACH TO SOLUTION LIMITATIONS ON
APPLICATION
COMMENTS ANC-1 Fourier Solution. Based on lifting line theory. Basic and additional lift.
Conventional unswept wings with aspect ratios greater than 5 to
6.
Exact Solution despite limitations of lifting line theory. Similar Methods:
LOTZ GLAUERT (Elements of Aerofoil and Airscrew Theory) ANDERSON (Theory of Wing Sections) SHRENK’S Approximate Conventional Accuracy approaches NACA TM- Solution. Based on unswept wings ANC-1. Small 948 lifting line theory.
Basic and additional lift.
with aspect ratios greater than 5 to
6.
inaccuracies overshadowed by original assumptions. Much easier to apply.
SHERMAN’S Successive Applicable to any Accuracy compares with NACA TN-732 approximations to straight wing ANC-1 but unsatisfactory match chord without high lift due to separate calculation distribution. Each devices. for each angle of attack.
angle of attack requires separate calculation.
PLANFORM Air load Untwisted high Unconservative for Taper (Often used for distribution directly taper ratio wings. Ratios < 0.25.
horizontal tail) proportional to chord variation.
Conservative for Taper Ratios > 0.35.
WEISSINGER Iteration process Applicable to Lifting surface method NACA TM based on modified straight or swept also presented as 1120 lifting line theory.
Basic and additional lift.
wings of low or high aspect ratio.
applicable to straight wings.
DeYOUNG & Based on simplified Applicable to Extension of Weissinger HARPER lifting surface straight or swept Method. Can account for NACA TR 921 theory. Basic and additional lift.
wings of low or high taper ratio.
compressibility M > 0.5 @ S.L.
DIEDERICH Matrix Iteration No planform Considers loading and NACA TR process utilizing limitations. divergence speed.
1000 aero-elastic matrix determined from aerodynamic and structural properties of wing.
NOTE: See NACA TN 606 for Empirical Tip Corrections, and RM L53B18 for Tip Tank Effects on Load Distribution.
REFERENCES
1. D. J. Peery, Aircraft Structures, McGraw-Hill, (1950).
2. J. Weissinger, The Lift Distribution of Swept-Back Wings, NACA Technical Note 1120, (1947).
3. Anonymous, Basic Glider Criteria Handbook, Federal Aviation Agency, (1962).
4. Niu, M. C. Y., Airframe Structural Design, Conmilit Press, LTD., (1988).
5. Kuthe, A. M. & Chow, C. Y., Foundations of Aerodynamics, Fourth Edition, John Wiley & Sons, New York, (1986).
6. Margason, R. J. and Lamar, J. E., “Vortex Lattice FORTRAN Program for Estimating Subsonic Aerodynamic Characteristics of Complex Planforms,” NASA Technical Note D-6142, (February 1971).
7. MSC/NASTRAN Aeroelastic Supplement, MacNeal-Schwendler Corporation, 815 Colorado Boulevard, Los Angeles, CA 90041-1777.
8. DeYoung, J. and Harper, C. W., “Theoretical Symmetric Span Loading at Subsonic Speeds for Wings Having Arbitrary Planform.” NACA Report No. 921, (1948).
9. Sivells, J. C., “An Improved Approximate Method for Calculating Lift Distributions Due to Twist.” NACA Technical Note 2282, (1951).
10. DeYoung, J., “Theoretical Anti-symmetric Span Loading for Wings of Arbitrary Plan Form at Subsonic Speeds.” NACA Report No. 1056, (1951).
11. DeYoung, J., “Theoretical Symmetric Span Loading Due to Flap Deflection for Wings of Arbitrary Plan Form at Subsonic Speeds.” NACA Report No. 1071, (1952).
12. Falkner, V. M., “The Calculation of Aerodynamic Loading on Surfaces of any Shape.” ARC Report R&M No. 1910, (1943).
13. Multhopp, H., “Methods for Calculating the Lift Distribution of Wings (Subsonic lifting-surface theory).” ARC Report R&M No. 2884, (1955).
14. Watkins, C. E., Woolston, D. S., and Cummingham, H. J., “A Systematic Kernel Function Procedure for Determining Aerodynamic Forces on Oscillating or Steady Finite Wings at Subsonic Speeds.” NASA Technical Report R-48, (1959).
23.302 Canard or tandem wing configurations (Amendment 23-42)
16. Is there any policy available on this section as of January 1, 2007? No.
23.303 Factor of safety (Original)
17. Is there any policy available on paragraph (a) of this section as of January 1, 2007? The prescribed factor of safety of 1.5 is the minimum acceptable factor of safety.
Also see section 23.619.
23.305 Strength and deformation (Amendment 23-45)
18. Is it possible to return an article to service after experiencing ultimate loads from a test? Certain FAA Order 8110.4 practices, about returning articles to service that have experienced ultimate load tests, may be relaxed without compromising safety.
[Order 8110.4C, currently in effect, was issued several years after this example situation from Order 8110.4.] For instance, an engine mount assembly can be readily and completely inspected to determine that there is no structural damage (deformation, permanent set, material yielding). The previous FAA Order 8110.4 permits similar practice for limit load tested articles. Exercise judgment to determine the structures that can properly be inspected for damage.
19. How should I show that a part can withstand ultimate load without failure, and what is considered a failure on a static test specimen?
a. The interpretation of a structural failure of a static test specimen has varied greatly on past type certification programs. In the strictest interpretation, if one part fails beyond limit load but below ultimate load, the test is stopped—the part repaired—and the test rerun. The repair, in this case, becomes part of the type design. In a more liberal vein, a local failure up to ultimate load was accepted as long as the entire structure being tested was able to carry the ultimate load for 3 seconds. The applicant was not required to redesign or structurally “beef up” the locally failed part. In a third instance, the specimen was loaded to destruction with a continuously increasing load at a constant rate and with a continuous recording of the test results. The ultimate load was established as the load attained 3 seconds before the maximum load was recorded.
b. 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.
c. 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.
20. What is the definition of “acceptable local failures” in static structural tests?
a. Regulatory History
(1) Amendment 23-45, 14 CFR § 23.305(b) read as follows:
”The structure must be able to support ultimate loads without failure for at least three seconds.”
(2) Amendment 23-45, effective August 6, 1993, changed § 23.305(b) to read as follows:
”The structure must be able to support ultimate loads without failure for at least three seconds, except local failures or structural instabilities between limit and ultimate load are acceptable only if the structure can sustain the required ultimate load for at least three seconds.”
(3) The Notice of Proposed Rule Making (NPRM, Notice 90-18, Small Airplane Airworthiness Review Program Notice No. 4) for Amendment 23-45, provided the following information supporting the revision to § 23.305(b).
”This proposal clarifies the FAA’s interpretation of failure during static ultimate load test. Using existing Section 23.305, the test is a failure if a part or component fails (e.g., a rivet) beyond limit load but below ultimate load during a static ultimate load test. Using a more liberal interpretation, a failure or structural instability between limit and ultimate load is acceptable as long as the entire structure demonstrates the capability to carry ultimate load for three seconds. This proposal clarifies this disparity, but is not intended to relieve the requirement for deflection shown in Section 11
23.301(c) or Section 23.305(a). The intent of this proposal was unopposed at the conference.”
(4) The ”conference” referred to in the NPRM is the Part 23 Regulatory Review Program public meeting held in St. Louis, Missouri, in October 1984. The minutes from the meeting (pg. 191, Regulatory Review Program, Committee II, Airframe, October 22, 1984) provide further insight into the intent of the revision to § 23.305(b).
“The idea, here, is to allow the test to continue between limit and ultimate.
In the event you had a rather minor local failure, such as a sheared rivet or two, the idea is, if the major component, assuming you are testing a wing, for example and you shear or rivet or two, there have been cases in which FAA personnel had discontinued the test and have told the applicant or the manufacturer that the test did not pass, because there was a failure. Now, this proposed change would permit the test to continue to ultimate and would consider to be an acceptable test, provided there were no major structural failures.”(Sic)
b. Policy
(1) The intent of § 23.305(b) is that the structure must support ultimate loads without failure. The revision at Am endment 23-45, clarified that minor, local failures, and instabilities may be acceptable. The intent of the revision was to provide relief in those instances where minor, insignificant failures might otherwise result in the test being declared a failure.
(2) The revision was not intended to make any failures acceptable as long as some part of the structure was able to sustain the required loads. The original intent of the regulation remains the same; any failures before or at ultimate load are not desirable.
The revision simply allows for the use of good engineering judgment, so that relatively minor local failures and structural instabilities may be discounted.
(3) Acceptable local failures or structural instabilities might include:
• A limited number of sheared rivets.
• Short cracks not extending a significant distance through the part or component.
• Localized panel buckling.
• Delaminations or disbonds over a small percentage of the part or joint area.
(4) The following types of failures are not acceptable even if some part of the structure is able to support the required load. If the structure develops these types of damage before or at ultimate load, the test should be considered a failure.
• Any failure that causes significant load shedding or redistribution of loads. Significant load shedding is indicative that a major load path or component has failed. Load versus deflection data and strain gage data can be observed during the test to verify that the design load paths are maintained.
• Large number of sheared fasteners.
• Widespread cracking.
• Extensive buckling
• Large areas of delaminations or disbonds.
21. Can I use structural analysis to show compliance to part 23, § 23.305? Yes.
Often, an engineer can perform structural analyses that will substantiate airplane designs and design changes. Contact an engineer who is familiar with the FAA certification process and the particular airworthiness standards. Among others, a Designated Engineering Representative (DER) can sometimes help in this endeavor. This is another way that allows a designer or a modifier to gain FAA approval for changes to the type design. See AC 183.29-1, Designated Engineer ing Representatives, current edition.
22. Can I employ static tests to show compliance to part 23, § 23.305?
a. Yes. The assessment of a structure at limit load is a visual check.
Deformations may be observed at limit load. However, those deformations should disappear when the load is removed. Also, any deformation that may occur at any load up to limit load should not interfere with safe operation. For example, when static testing a complete wing structure that includes installed control systems, ailerons, flaps, etc., the control systems and surfaces should perform their intended function during any deformation that may occur up to and including limit load.
b. The FAA CAUTIONS airplane designers and certifiers to watch out for the SPECIAL EXCEPTION to FAA LIMIT and ULTIMATE load regulatory failure conditions (Euler Column Buckling). COLU MN STRUCTURES, when they are used in a (primary structure) single-load-path design application, cannot be allowed to buckle under either FAA LIMIT or FAA ULTIMATE load conditions. Two common applications of column structures are wing struts and control system pushrods.
c. Settlement of structure due to the effects of riveting, fasteners, etc., does take place during limit load tests. When testing a pressurized fuselage, the pressure differential required by § 23.365 will introduce some settlement in the rivets and fasteners. The differential pressure required is 1.33 times the maximum relief valve setting. For altitudes that exceed 45,000 feet, previously issued part 23 special conditions required a differential pressure of 1.67 times the maximum relief valve setting. Under limit load, if it can be accomplished safely, visually inspect the test specimen. Accept sounds associated with the working of rivets, fasteners, and panels as the applied load increases to limit load. See 04.201, Civil Aeronautics Manual (CAM) 04, Deformations, revised July 1, 1944. (See 23.365 guida nce about pressure tests.)
d. For metallic structures only, the FAA Small Airplane Directorate allows tested structure to be used as part of an airplane in operational service. However, the structure must have been thoroughly examined, following static tests to any loads, and the structural deformations must have remained elastic. That is, the structure should be shown, by proper measurements, not to have experienced material stresses beyond the material yield-point stresses under the applied loads. Any deviations to the type design that are created due to structural tests should be dispositioned before the article is subjected to operational flights. If, during any static tests, portions of the structure become visibly damaged, the damaged items should be replaced before the structure is released for operational flight tests. On the other hand, any static test structure that has been tested and has yielded is not a candidate structure for later operational flight tests.
23. Can I conduct flight tests as the sole means to show compliance to part 23, § 23.305 for secondary structures?
a. No. In the past, there have been inst ances where flight tests to dive speed have been accepted as the only means of substantiating secondary structures.
Definition: Secondary structure is not a primary load-carrying member.
Failure of secondary structure neither reduces the airframe structural integrity nor prevents the airplane from continued safe flight and landing.
b. A dive-speed approval does not satisfy the requirement that structure should support ultimate load.
c. In other instances, dive-speed flight tests have not been accepted as the sole means of showing compliance with the airworthiness standards. In those cases, an applicant presents additional data. Some secondary structure modifications, or alterations, have been approved by structural analyses and tests without dive-speed tests;
these modifications include structures like windshields, windows, and radomes.
d. Sometimes, the FAA has accepted flight tests to design dive speed (not just to the never-exceed-speed, VNE) as the only means of “substantiating” secondary structures.
This substantiation does not satisfy the requirement for the structure to support ultimate load, nor does it apply the load factors of the V-n diagram to the structure. Flight tests to design dive speed should not be accepted as the sole means of substantiation; the applicant must present additional data to complete the “show compliance with Airworthiness Standards” requirements. Certain secondary structure modifications or alterations, which do not change the original external contour, have been FAA approved by structural substantiation alone. Examples of these types of secondary structures include windshields, windows, and radomes.
Caution: Compliance to flight tests should be prudently limited to an 80 percent design flight envelope (V-n diagram; limit loads) until structural tests to all ultimate load conditions are satisfactorily completed (or until all structural analyses to ultimate load conditions are satisfactorily completed).
Caution: Structural flight tests do not necessarily demonstrate ultimate load conditions for secondary structures. Limit dynamic pressure, limit maneuver load factor (and load), and limit landing impact load factor (and load) may be easily achieved during flight tests. Limit gust load factor may be difficult to achieve during flight tests. It can be dangerous to attempt ultimate load factor (and load) flight tests. Ultimate load flight tests are discouraged.
Caution: In rare instances, some secondary structures can be flight tested safely to ultimate load conditions well within the airplane flight envelope (V-n diagram), and well below the airplane design dive speed (VD). Landing gear doors are an example of this special case. The landing gear operational speeds, landing gear extended, VLE, and landing gear operating, VLO, can be considerably lower than the airplane design dive speed. Consequently, it is sometimes possible to flight test to an ultimate dynamic pressure for the landing gear that is safely below the limit dynamic pressure for the airplane design dive speed. Angle of attack may be a negligible factor in this case, and maximum airplane yaw angles may be accommodated within the airplane limit flight conditions.
24. What is your interpretation of “detrimental, permanent deformation,” as used in paragraph (a) of this section, on limit load testing?
a. Current policy does not specify a numeric value for “allowable permanent deformation”.
b. As stated previously, metallic structure should not be used in service if it has exceeded the deformation obtained by the 0.2 percent offset method, which is the common engineering practice for defining the yield point of metals. Yielded material is not airworthy. This addresses the airworthiness of materials, but it does not define what is considered “detrimental, permanent deformation.”
c. The definition of “detrimental, permanent deformation” is more qualitative in nature, but we can define a set of pass/fail criteria for the limit load tests. Therefore, every item on the following list of criteria must be met for any permanent deformations present after the removal of limit load, to meet the criteria that there is no “detrimental, permanent deformation.”
(1) It must not require repair.
(2) It must not be noticeable during routine maintenance and inspections.
(3) It must not prevent proper fit and function of primary structure, secondary structure, and flight controls. Removable primary structure, removable wings on gliders for example, must align properly when reinstalled. Doors, access panels and similar structure must fit properly, with correct fastener alignment. Control surfaces must function properly, with full range of travel and proper fairing.
(4) It must not have a noticeable effect on the external airload distribution.
Areas of special attention include deformations that may change the local airflow or that may have altered the wing twist.
(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