CONTENTS
Section PAGE
1. What is the purpose of this AC?1
2. What is incorporated in this AC and, thereby, superseded?1
3. What is the background for this AC and what time period does this AC cover?1
4. How should the information contained in this AC be applied?1
5. What regulations are discussed in this AC?1
6. What other publications are related to this AC?2
7. What are the definitions for the common terms used in this AC?2
CHAPTER 1. SMALL AIRPLANE FATIGUE REGULATIONS4
1-1. What is the goal of the sma ll airplane fatigue requirements?4
1-2. How have the fatigue requi rements changed over time?4
1-3. Can you summarize the current fatigue regulations for small airplanes?4
1-4. Which metallic components should I include in the fatigue evaluation?6
1-5. Are the small airplane fatigue regula tions different in other countries?6
1-6. How do I show compliance to the fatigue requirements for alterations, modifications, or
changes to the design?6
CHAPTER 2. SAFE-LIFE FATIGUE EVALUATION7
2-1. What is a safe-life fatigue evaluation?7
2-2. What is the definition of failure in a safe-life structure?7
2-3. What are the steps in evaluati ng the safe-life of a structure?7
2-4. What loading spectra should I use?8
2-5. How were the load spectra in Appendix 1 developed?9
2-6. Can I use other loading spectra in place of the Appendix 1 spectra?9
2-7. What should I consider in developing a mission profile?9
2-8. Does the FAA require a flight test to valid ate the loads I use in my fatigue evaluation?10
2-9. Does the FAA require validation of the stresses I use in my fatigue evaluation?10
2-10. How do I substantiate a fatigue saf e-life with full-scale fatigue tests?11
2-11. What should I include in the test plan for my full-scale fatigue tests?11
2-12. Do I need to complete the test before the FAA will issue the type certificate (TC)?13
2-13. How do I calculate the fatigue safe-lif e using full-scale fatigue test results?13
2-14. How do I calculate the mean test life?14
2-15. What are the scatter factor s for full-scale fatigue tests?14
2-16. What about statistical confidence inte rvals on full-scale fatigue test results?16
2-17. What counts as one test specimen, and what counts as two test specimens?16
2-18. Can I use a full-scale test scatter factor lower than that obtained from Equation 4?17
2-19. What about the '3-sigma' method of computing the safe-life?17
2-20. How do you substantiate a fatigue safe-life with component tests?17
2-21. What loads do I apply in a component test?18
2-22. For component testing, how many specimens do I need to test?18
2-23. For component testing, what scatter factor should I use?18
2-24. How do you substantiate a fatigue safe-life with analysis supported by test evidence?19
2-25. What are some concerns with s ubstantiation by fatigue analysis?19
2-26. What scatter factors do I use in a fatigue analysis?20
2-27. How can I extend the safe-life, or retirement life, of my airplane?21
CHAPTER 3. FAIL-SAFE DESIGN23
3-1. What is a fail-safe design?23
iii
3-2. What is the history of fail-safe design?23
3-3. How can I assure that my design complies with the fail-safe design requirements?24
3-4. What are some design considerations in achieving a fail-safe design?24
3-5. What is a Principal Structural Element?24
3-6. How much fatigue damage should I consider in my fail-safe evaluation?25
3-7. How do I address inaccessible areas?26
3-8. How do I account for dynamic effects?26
3-9. How do I test principa l structural elements?26
3-10. Can I use analysis to s ubstantiate a fail-safe design?27
3-11. How do I select the critical areas?27
CHAPTER 4. DAMAGE TOLERANCE EVALUATION28
4-1. Are there any differences between the part 23 and part 25 damage tolerance regulations?28
CHAPTER 5. ROTORBURST REQUIREMENTS29
5-1. What are the requirements for turbine or tu rbopropeller installations in small airplanes?29
5-2. I thought there was not a discrete s ource damage requirement in part 23?29
CHAPTER 6. FLIGHT WITH KNOWN CRACKS30
6-1. What is the purpose of this policy?30
6-2. What is the policy for operational flight with known cracks?30
6-3. Are there any exceptions to this policy?30
6-4 Which airplanes are covered by this policy?30
6-5 What will the FAA consider before allowi ng operational flight with known cracks?30
REFERENCES34
APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRAA1-1
APPENDIX 2. S/N CURVES FOR ALUMINUM COMPONENTSA2-1
1. What is the purpose of this AC?
This advisory circular (AC) provides information and guidance concerning acceptable means, but not the only means, of compliance with Title 14 of the Code of Federal Regulations (14 CFR) part 23. This guidance is applicable to fatigue, fail-safe, and damage tolerance evaluations of metallic structure in normal, utility, acrobatic, and commuter category airplanes. This AC provides information on approval of continued operational flight with known cracks in the structure of small airplanes, regardless of their certification basis. This AC also clarifies the use of AC 20-128A in the evaluation of rotorburst structural hazards in small airplanes. This AC consolidates existing policy documents, and certain technical reports, into a single document.
This AC incorporates, and supersedes, FAA Report No. AFS-120-73-2, Fatigue Evaluation of Wing and Associated Structure on Small Airplanes, dated May 1973. Material in this AC is neither mandatory nor regulatory in nature and does not constitute a regulation.
2. What is incorporated in this AC and, thereby, superseded?
The following AC has been incorporated into this revision in whole or in part and is hereby superseded:
AC 23-13, Fatigue and Fail-Safe Evaluation of Flight Structure and Pressurized Cabin for Part 23 Airplanes, dated April 15, 1993.
3. What is the background for this AC and what time period does this AC cover?
The AC is current through Amendment 23-55, effective March 1, 2002. This material spans approximately 38 years of Federal Aviation Administration (FAA) aviation history.
4. How should the information contained in this AC be applied?
a. This material does not have any legal status and should be treated accordingly. However, to ensure standardization in the certification process, this guidance should be considered during all small airplane Type Certification (TC) and Supplemental Type Certification (STC) activities.
The user of this AC should check the latest amendments to the applicable 14 CFR part to verify that the regulations referenced in this AC have not been superseded or that additional regulations have not been added.
b. New approaches to fatigue evaluation, design features, methods of fabrication, and new airplane configurations may require variations and deviations from the guidance described in this AC. Engineering judgment, guided by the extens ive literature on the subject, should be exercised for each particular application. An applicant should evaluate the primary structure of the pressurized cabin, wing, empennage, and associated structures for any original design and for any design changes that affect the loading spectra, internal stresses, or stress concentrations or that change the construction methods or materials. Changes to the design that may be minor from a static strength standpoint can have a major effect on fatigue characteristics.
5. What regulations are discussed in this AC?
a. The requirements for fatigue evaluation included in 14 CFR, part 23, §§ 23.571, 23.572,
23.573, 23.574, and 23.627.
b. The requirements for inspections and Instructions for Continued Airworthiness included in 14 CFR, part 23, §§ 23.575, 23.611, and 23.1529.
c. The requirements for minimization of rotorburst hazards included in 14 CFR, part 23, § 23.903.
6. What other publications are related to this AC?
a. AC 23-19, Airframe Guide for Certification of Part 23 Airplanes, dated January 27, 2003.
b. AC 25.571-1C, Damage Tolerance and Fatigue Evaluation of Structure, dated April 29, 1998.
c. AC 25.613-1, Material Strength Properties and Material Design Values, dated August 6, 2003.
d. AC 20-128A, Design Considerations for Minimizing Hazards caused by Uncontained Turbine Engine and Auxiliary Power Unit Rotor Failure, dated March 25, 1997.
e. AC 20-107A, Composite Aircraft Structure, dated April 25, 1984.
f. AC 91-56A, Continuing Structural Integrity Program for Large Transport Category Airplanes, dated April 29, 1998.
g. AC 91-60, The Continued Airworthiness of Older Airplanes, dated June 13, 1983.
h. Aging Airplane Inspection and Maintenance Baseline Checklist for Airplanes Without a Type Specific Checklist (Best Practices Guide for Maintaining Aging General Aviation Airplanes). Electronic version available at:
http://www.faa.gov/aircraft/air_cert/design_approvals/small_airplanes/cos/aging_aircraft/media/ aging_aircraft_best_practices.pdf
7. What are the definitions for the common terms used in this AC?
This Advisory Circular uses the definitions listed below. Some of these terms are used throughout and will only be defined here. Sometimes, the term definitions may be repeated in the relevant document section as an aid to the reader.
a. Principal Structural Element (PSE). A PSE is an element that contributes significantly to the carrying of flight, ground, or pressurization loads, and whose integrity is essential in maintaining the overall structural integrity of the airplane.
b. Primary Structure. Primary structure is that structure which carries flight, ground, or pressurization loads, and whose failure would reduce the structural integrity of the airplane.
c. Fatigue. The process of progressive localized permanent structural change occurring in a material subjected to conditions that produce fluctuating stresses and strains at some point or points, which may result in cracks or complete fracture after a sufficient number of fluctuations.
d. Safe-Life. The safe-life of a structure is that number of events, such as flights, landings, or flight hours, during which there is a low probability that the strength will degrade below its design ultimate value due to fatigue cracking.
e. S-N or ε–N. Stress–Life (S-N) or Strain–Life (ε -N) curves depict the magnitude of applied stress (S) or strain (ε) necessary to develop a fatigue crack in a specimen at a given life (N), where N is expressed in the number of cyclic applications of stress or strain.
f. Scatter Factor. The scatter factor, or life reduction factor, is a statistically derived divisor applied to fatigue test results to account for the variation in fatigue performance of built-up or monolithic structures. A scatter factor can also be used in a fatigue analysis to address the uncertainties inherent in a fatigue analysis.
g. Fail-Safe. Fail-safe is the attribute of the structure that permits it to retain its required residual strength for a period of unrepaired use after the failure or partial failure of a principal structural element.
h. Damage Tolerance. Damage tolerance is the attribute of the structure that permits it to retain its required residual strength for a period of use after the structure has sustained a given level of fatigue, corrosion, accidental, or discrete source damage.
i. Discrete Source Damage. Damage to the structure caused by a discrete source, such as fragments from uncontained engine failures.
j. Residual Strength. The strength capability of a structure after the structure has been damaged due to fatigue, corrosion, or discrete source damage. The residual strength capability includes consideration of static strength, fracture, and stiffness.
CHAPTER 1. SMALL AIRPLANE FATIGUE REGULATIONS
1-1. What is the goal of the sma ll airplane fatigue requirements?
The goal of the fatigue requirements is to maintain the design strength capability, both limit and ultimate load, throughout the operational life of the airplane.
1-2. How have the fatigue requirements changed over time?
a. Before 1957 . The only fatigue requirements for sm all airplanes appeared in the Civil Air Regulations (CAR), § 3.307. The CAR 3 paragr aph required the airplane designer to choose design details that avoided stress concentrations. This airworthiness requirement applied to all structural design details on the whole airplane. When the CAR requirements were recodified to the 14 CFR parts, this requirement was carried over as § 23.627.
b. 1957. Amendment 3-2 added a specific requirement to CAR 3 for the fatigue assessment of pressurized fuselage structure. The requirement gave applicants the option of a safe-life evaluation or incorporating fail-safe design features. CAR § 3.270 was recodified as 14 CFR, part 23, § 23.571.
c. 1969. Amendment 23-7, effective September 14, 1969, added fatigue requirements for the wing, wing carry-through, and attaching structure in § 23.572.
d. 1979. Special Federal Aviation Regulation No. 41 (SFAR 41), effective September 17, 1979, required a fatigue substantiation of “those parts of the wing, wing carrythrough, vertical fin, horizontal stabilizer, and attaching structure whose failure would be catastrophic.”
e. 1989. Amendment 23-38, effective October 26, 1989, added the empennage to the § 23.572 fatigue requirements.
f. 1993. Amendment 23-45 introduced § 23.573 requiring damage tolerance and fatigue evaluation for composite structures. This amendment also added damage tolerance as an option for metallic pressurized cabin in § 23.571 and for wing, empennage, and associated structure in § 23.572. Amendment 23-45 also expande d the fatigue assessment of wing structure to include canards, tandem wings, and winglets/tip fins.
g. 1996.
(1) Amendment 23-48 introduced § 23.574, which requires damage tolerance for metallic structure on commuter category airplanes. This amendment became effective on March 11, 1996.
(2) Amendment 23-48 also introduced § 23.575, which clarifies that inspections and any other procedures necessary to prevent catastrophic failure must be included in the Limitations Section of the Instructions for Continued Airworthiness (ICA).
1-3. Can you summarize the current fatigu e regulations for small airplanes?
a. For metallic pressurized cabin, wing, empennage, and associated structure in normal, utility, and acrobatic category airplanes, you have the option of the following fatigue evaluation methods:
(1) A fatigue strength evaluation, more commonly known as a safe-life evaluation. See § 23.571(a) for pressurized cabin and § 23.572(a)(1) for wing, empennage, and associated structure.
(2) A fail-safe strength evaluation. See § 23.571(b) for pressurized cabin and § 23.572(a)(2) for wing, empennage, and associated structure.
(3) A damage tolerance evaluation. See § 23.573(b) and § 23.571(c) for pressurized cabin and § 23.572(a)(3) for wing, empennage, and associated structure.
b. For wings, empennage, and associated structure, you may be able to show compliance to the fatigue requirements by comparing your design to an existing design. This method of substantiation includes showing that the structure, operating stress level, materials, stress concentrations, and expected uses are equivalent, from a fatigue standpoint, to a similar design that has had extensive satisfactory service experience. See § 23.572(a).
c. For metallic commuter category airplanes, you must use a damage tolerance evaluation. If you can show that damage tolerance is impractical for a particular structure, you may use a fatigue strength, or safe-life evaluation. See § 23.574.
d. For all categories of small airplanes constructed with composite materials, you must use a damage tolerance evaluation. AC 20-107A provides guidance for composite structures.
e. The following flowchart may help clarify which method of fatigue evaluation is required for your application.
FIGURE 1. WHICH FATIGUE REGULATIONS APPLY TO MY NEW TYPE DESIGN?
(Amendment 23-48, effective February 9, 1996.)
Is your structure constructed of composite materials?
NO Required: Damage Tolerance Evaluation YES 14 CFR §§ 23.573(a), 23.575 See AC 20-107A Required: 1.) Damage Tolerance Evaluation
14 CFR §§ 23.574(a), 23.575
Are you seeking certification YES in the commuter category?
or,
2.) If applicant can substantiate that the damage tolerance approach is impractical for the application, they may use a safe-life approach. 14 CFR §§ 23.574(b), 23.575 Required: For metallic structure in normal, utility, and acrobatic categories, an applicant may choose any of the following:
NO
1.) Safe-Life Evaluation For pressurized cabin, 14 CFR §§ 23.571(a), 23.575 For wing, empennage, and associated structure, 14 CFR §§ 23.572(a)(1), 23.575
2.) Fail-Safe Evaluation For pressurized cabin, 14 CFR §§ 23.571(b), 23.575 For wing, empennage, and associated structure, 14 CFR §§ 23.572(a)(2), 23.575
3.) Damage Tolerance Evaluation For pressurized cabin, 14 CFR §§ 23.571(c),
23.573(b), and 23.575 For wing, empennage, and associated structure, 14 CFR §§ 23.572(a)(3), 23.573(b), and 23.575 5
f. The most recent amendment to part 23 that revised the fatigue requirements was Amendment 23-48, effective February 9, 1996. Always check the latest amendment to part 23 to verify that the regulations referenced in this AC have not been superseded or that additional regulations have not been added.
1-4. Which metallic components should I include in the fatigue evaluation?
Although the fatigue regulations discuss “those parts of the airframe structure whose failure would be catastrophic,” a 1998 policy letter clarifies that you are required to include only certain sections of metallic airplanes in your fatigue evaluation. For airplanes certificated in any small airplane category, these sections include the pressurized cabin, wing (includes tandem wings, canards, winglets and tip fins), empennage, and associated structures. You do not need to include other portions of the airplane, including the landing gear and engine mounts.
1-5. Are the small airplane fatigue re gulations different in other countries?
a. The FAA fatigue regulations for small airplanes are the same as those of many foreign countries. The FAA and the CAA (Civil Aviation Authority) of these countries have worked together to harmonize the part 23 regulatory wording.
b. Even though the regulations may be identical, some differences in policy, guidance, and acceptable means of compliance may exist between the FAA and the CAA of foreign countries. These differences include:
(1) The CAA of some countries may require fatigue evaluation of the engine mounts and landing gear.
(2) The CAA of some countries may require different scatter factors in a safe-life evaluation.
1-6. How do I show compliance to the fatigue requirements for alterations, modifications, or changes to the design?
a. The fatigue performance of the structure should be evaluated for any alteration, modification, or change to the type design. P hysical changes to the design that may affect fatigue performance include changes that increase the external or internal loads, alter the distribution of internal loads, change the stress concentrations, or change the materials or method of construction. You should also ev aluate the effect of any change to the airplane’s operational characteristics that invalidate the assumptions used in the original fatigue substantiation.
Changes to the operational characteristics that may be important for fatigue include higher design airspeeds or higher average speed. They also include changes to the maximum allowable weight and center of gravity envelope, changes to the average weight and center of gravity location, and engine or propeller changes.
b. Modifiers and applicants for supplemental type certificates (STCs) often do not have access to the Type Certificate holder’s original fatigue substantiation data. In these cases, applicants often complete a “comparative fatigue analysis” in which the fatigue performance of the modified and unmodified airplane are compared. Applicants use the results of the comparative analysis to determine the appropriate changes to the safe-life or inspection program of the modified airplane.
CHAPTER 2. SAFE-LIFE FATIGUE EVALUATION
2-1. What is a safe-life fatigue evaluation?
A safe-life fatigue evaluation is the assessment of a structure to ensure that the structure is able to withstand, without catastrophic failure, the repeated loads of variable magnitude expected in service throughout its operational life. The “safe-life” is a point in the airplane’s operational life, expressed in hours or flight cycles, at which the structure must be taken out of service or replaced. The structure is reti red or replaced at the safe-life to prevent the structure from developing fatigue cracks. An applicant may determine the safe-life of a structure using the following methods:
a. full-scale testing,
b. component testing, or
c. analysis supported by test evidence.
2-2. What is the definition of a fati gue failure in a safe-life structure?
For a safe-life structure, fatigue failure is the development of a detectable crack. A detectable crack is one that can be detected by common inspection methods, or the inspection methods required in the maintenance instructions. Since a safe-life evaluation usually does not include demonstration of crack growth rates or residual strength capability, we assume that the development of a detectable crack may result in catastrophic failure of the structure.
2-3. What are the steps in evaluati ng the safe-life of a structure?
We can summarize the essential elements of any safe-life evaluation as follows:
a. Estimate or measure the expected loading spectra for the structure. The loading spectra should include all loads you expect the structure to experience in normal operation. This includes gust loads, maneuver loads, and pressurization loads. Although you are not required to include the landing gear and engine mounts in your fatigue evaluation, you need to include appropriate taxi loads, landing impact loads, engine loads, vibration, buffet, and propeller slipstream effects in your loading spectra.
b. Conduct a structural analysis, which includes consideration of the effects of stress concentrations. Your structural analysis should include determination of gross stress levels and determination of the stress concentration factors, Kt, for all details in principal structural elements. Your analysis should consider, in detail, structural joints and fittings, paying particular attention to eccentrically loaded joints and the bearing and bypass stresses in the joint. The structural analysis should identify locations where fretting may occur. Identify areas where corrosion may develop or extreme thermal environments may affect fatigue performance.
c. Conduct a fatigue test of any structure that cannot be related to a test background to establish the structure’s response to the typical loading spectrum expected in service. Full-scale fatigue testing is the most reliable method of determining the response of the structure to expected operational loadings. Fatigue substantiation by analysis is also acceptable. However, the stress-life (S-N) or strain-life (ε–N) data used in the analysis must be based on fatigue tests of similar structures or adjusted to account for differences between the actual structure and the tested structure.
d. Determine reliable replacement times by interpreting the loading history, variable load analysis, fatigue test data, service experience, and fatigue analysis. The reliability of the 7
replacement times is usually achieved through the use of fatigue scatter factors or other statistical adjustments.
e. Provide data for inspection and maintenance instructions and guidance information to the operators. This data should include inspectio n methods, inspection thresholds and intervals.
2-4. What loading spectra should I use?
a. Appendix 1 presents gust and maneuver flight load spectra in both graphic and tabulated form for the following types of airplanes and usage:
(1) Single-Engine Executive Usage (non-pressurized, engine size greater than 185 hp)
(2) Single-Engine Personal Usage (non-pressurized, engine size less than or equal to 185 hp)
(3) Single-Engine Instructional Usage (non-pressurized)
(4) Single-Engine Acrobatic Usage (non-pressurized)
(5) Twin-Engine General Usage (non-pressurized)
(6) Twin-Engine Instructional Usage (non-pressurized)
(7) Pressurized Usage
(8) Special Usage (including survey and aerial application)
b. For other types of loads:
(1) Appendix 1 also provides ground load spectra, including landing impact and taxi loads.
(2) The loading spectrum for pressurized cabins should include the repeated application of the normal operating differential pressure, and the superimposed effects of the gust and maneuver flight loads and external aerodynamic pressure.
(3) If the wing center section skin panels are affected by cabin pressurization, the wing loading spectrum should include the effects of pressurization.
(4) Reference 1 provides additional guidance on load spectra development and fatigue evaluation for the empennage, and for configurations with canards (or forward wings) and winglets (or tip fins).
c. Appendix 1 also discusses the gust load formula you should use to develop the gust load spectra for your airplane.
d. Your loading spectra should include the ground-air-ground cycle (GAG). The GAG cycle represents the range of the maximum and minimum loads expected to occur on a per flight basis. Typically, the minimum load results fr om landing or taxi conditions and the maximum load from the gust or maneuver spectra. For typical airplanes certificated in the normal category, two-thirds to three-fourths of the total fatigue damage on the wing may be caused by the GAG cycle.
e. While positive and negative load cycles are considered to occur randomly in service, the high positive and negative loads of a given type of repeated loading tend to occur at the same time. For testing and analysis purposes, the high positive load cycles are typically combined with the high negative load cycles of the same frequency.
2-5. How were the load spectra in Appendix 1 developed?
a. The FAA developed the flight load spectra in Appendix 1 based on a statistical analysis of the data presented in Reference 2. We used the ground load spectra directly from Reference 3 for the Appendix 1 ground load spectra.
b. The Appendix 1 flight load spectra include an increment (1.5 standard deviations) added to the average measured load frequency. The increment accounts for the variability in loading spectra experienced from individual airplane to airplane. The magnitude of the increment was selected to maintain the probability that a component will reach its safe-life without a detectable fatigue crack established by the scatter factor (see paragraph 2-15).
Reference 4 provides additional details on the development of the Appendix 1 flight load spectra.
2-6. Can I use other loading spectra in place of the Appendix 1 spectra?
The Appendix 1 spectra are the FAA’s approved load spectra for use in small airplane fatigue evaluations. Applicants may propos e alternate spectra in certain situations. If an applicant proposes alternate spectra, the Small Airplane Directorate Standards Office must approve these spectra. Any spectrum proposed by an appli cant should maintain the probability that a component will reach its safe-life without a detectable fatigue crack established by the scatter factor (see paragraph 2-15 and Reference 4). Si tuations in which applicants might consider alternate spectra include the following:
a. The applicant may develop gust spectra for pressurized usage based on power spectral density (PSD) methods. It is acceptable in this case to use PSD mission analysis techniques to account for typical operating mission profiles.
b. The applicant has access to a large database of measured spectra data on a similar configuration or on an airplane with similar usage. This type of data is particularly useful in situations involving severe or unique usage, such as acrobatic and agricultural usage.
2-7. What should I consider in developing a mission profile?
In developing the mission profile for your airplane, you should use the following assumptions:
a. Flight time
(1) Single-Engine and Twin-Engine (non-pressurized) 0.65 hours
(2) Pressurized 1.1 hours
(3) Single-Engine special usage (low level survey) 2.0 hours
(4) Twin-Engine special usage 3.0 hours
b. Airplane Speed
(1) The speed for determining miles flown should not be less than 0.9 VNO or 0.9 VMO.
(2) For special usage, the speed for determining miles flown may be 100 knots or
0.9 VA, whichever is less.
c. Gross weight and load distribution
(1) You should base estimates of the gross weight and distribution of disposable load on conservative estimates of typical operating conditions.
(2) One acceptable means of compliance, used by some small airplane applicants, is to assume the weight condition that gives the highest 1-g stress.
d. The mission profiles of certain airplanes should consider the unique aspects of the usage. Some examples follow:
(1) Instructional usage . Takeoff and landing traini ng, or “touch-and-go” training, is a significant portion of the student pilot training curriculum. Incorporate “touch-and-go” training into the mission profiles of any airplane used for instructional purposes. Each “touchand-go” should be treated as a short duration flight, 6 to 10 minutes in length.
(2) Mixed usage . Some airplanes are certificated in more than one category. A common example is an airplane certificated in both the acrobatic and utility/normal categories.
In general, the safe-life should be based on the usage that results in the shorter life. Some applicants have estimated a mix of missions, the percent of time spent operating in either category, as a way of accounting for the different types of usage. In using this approach, it is not acceptable to use a mix of missions within a single flight. For purposes of fatigue evaluation, a single flight is operated within a single category and within a single type of usage.
(3) Severe usage. Some airplanes may be used more severely than you had assumed in the fatigue evaluation. Common examples of this include airplanes considered to be in the single-engine personal or executive usage category used for pipeline and utility patrol, for instruction, or for short duration commuter and air taxi flights. Some applicants address this issue by placing a statement in the limitations section of the Airplane Flight Manual (AFM) that certain types of usage require a re-evaluation of the structure by the Type Certificate (TC) holder. The AFM lists these t ypes of usage and requests the owner to contact the TC holder regarding the usage.
2-8. Does the FAA require a flight test to validate the loads I use in my fatigue evaluation?
In most cases, the FAA will not require flight test validation of external loads or the internal stresses used in a fatigue evaluation. Typically, it is acceptable for an applicant to substantiate that the predicted external loads conservatively represent actual flight load conditions. However, it is in the applicant’s best interest to accurately know the internal loads, or stresses, at critical fatigue details. A one percent increase in stre ss level can shorten the expected fatigue life by more than five percent. Due to the sensitivity of fatigue life to stress level, an applicant should consider flight-testing to measure strains at or near fatigue critical structural details of the principal structural elements.
2-9. Does the FAA require validation of the stresses I use in my fatigue evaluation?
In determining whether validation or substantiation of the stresses used in a fatigue evaluation is necessary, the applicant and the FAA should consider the following:
a. For full-scale fatigue testing, an applicant should substantiate that the loads applied in the test either accurately or conservatively represent actual flight load conditions. Substantiation may be shown with comparisons of applied and predicted external loads. Substantiation should include demonstration that the applied external loads produce conservative stresses in the fatigue critical components. Strain gages should be used on the full-scale fatigue test and compared and correlated against the static test article or a validated analysis of the structure. Caution should be taken when using strain gages in areas of high stress gradients.