AC 23-13A 

FATIGUE, FAIL-SAFE, AND DAMAGE 

TOLERANCE EVALUATION OF METALLIC 

STRUCTURE FOR NORMAL, UTILITY, 

ACROBATIC, AND COMMUTER CATEGORY 

AIRPLANES 

September 29, 2005 

For sale by the Superintendent of Documents, U.S. Government Printing Office, Washington, DC 20401

Initiated by: ACE-100

9/29/05 AC 23-13A

FOREWORD 

This Advisory Circular (AC) sets forth an acceptable means of showing compliance with Title 14 

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 

certification basis. This AC also clarifies the use of AC 20-128A in the evaluation of rotorburst 

structural hazards in small airplanes. Finally, this AC consolidates existing policy documents 

and certain technical reports into one document. 

s/ 

David R. Showers 

Acting Manager, Small Airplane Directorate 

Aircraft Certification Service 

i (and ii) 

9/29/05 AC 23-13A

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 What is incorporated in this AC and, thereby, superseded?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 REGULATIONS............................................ 4 

1-1. What is the goal of the small airplane fatigue requirements? ..................................................... 4 

1-2. How have the fatigue requirements 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 regulations 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 EVALUATION............................................................ 7 

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 evaluating 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 validate 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 safe-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-life using full-scale fatigue test results? .............................13 

2-14. How do I calculate the mean test life? ........................................................................................14 

2-15. What are the scatter factors for full-scale fatigue tests?..............................................................14 

2-16. What about statistical confidence intervals 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 substantiation 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 DESIGN ......................................................................................23 

3-1. What is a fail-safe design? ..........................................................................................................23 

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AC 23-13A 9/29/05

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 principal structural elements?...............................................................................26 

3-10. Can I use analysis to substantiate a fail-safe design?..................................................................27 

3-11. How do I select the critical areas?...............................................................................................27 

CHAPTER 4. DAMAGE TOLERANCE EVALUATION........................................................28 

4-1. Are there any differences between the part 23 and part 25 damage tolerance regulations?........28 

CHAPTER 5. ROTORBURST REQUIREMENTS ................................................................29 

5-1. What are the requirements for turbine or turbopropeller installations in small airplanes? .........29 

5-2. I thought there was not a discrete source damage requirement in part 23?.................................29 

CHAPTER 6. FLIGHT WITH KNOWN CRACKS.................................................................30 

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 allowing operational flight with known cracks?................30 

REFERENCES ......................................................................................................................34

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA...................................................A1-1 

APPENDIX 2. S/N CURVES FOR ALUMINUM COMPONENTS.........................................A2-1 

 
9/29/05 AC 23-13A

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 extensive 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. 

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AC 23-13A 9/29/05

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. 
9/29/05 AC 23-13A

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. 

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AC 23-13A 9/29/05

CHAPTER 1. SMALL AIRPLANE FATIGUE REGULATIONS 

1-1. What is the goal of the small 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 small airplanes appeared in the Civil 

Air Regulations (CAR), § 3.307. The CAR 3 paragraph 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 expanded 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 fatigue 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. 

4 
9/29/05 AC 23-13A

(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?

Required: Damage Tolerance Evaluation

14 CFR §§ 23.573(a), 23.575

See AC 20-107A

YES

Required: 1.) Damage Tolerance Evaluation

14 CFR §§ 23.574(a), 23.575

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

Are you seeking certification YES

in the commuter category?

NO

Required: For metallic structure in normal, utility, and acrobatic

categories, an applicant may choose any of the following:

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

NO

5 
AC 23-13A 9/29/05

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 regulations 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. Physical 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 evaluate 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. 
9/29/05 AC 23-13A

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 retired 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 fatigue 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 evaluating 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 
AC 23-13A 9/29/05

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 inspection 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 from 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. 
9/29/05 AC 23-13A

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 propose 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 applicant 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). Situations 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. 

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AC 23-13A 9/29/05

(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 training, 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 types 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 stress 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. 
9/29/05 AC 23-13A

b. Component testing is different from full-scale testing in that the component test omits 

substantial portions of adjacent structure that may impart local stresses on the tested component. 

Because of the omitted structure and other simplifications inherent in component testing, an 

applicant should validate that the stresses in the component test accurately or conservatively 

represent the stresses in the actual structure. One acceptable method of validating the internal 

loads is with strain gages on the airframe static test article. Strain gage measurements should be 

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. Consider 

applying unique conditions on the static test article to better represent fatigue loading conditions. 

Loads applied in the static test should conservatively represent actual flight load conditions. 

c. For fatigue evaluation by analysis, an applicant should validate the internal loads used 

in the analysis. In validating the internal loads, the applicant should validate the component’s 

total stress field, including magnitude and orientation. One acceptable method of validating the 

internal loads is with strain gages on the airframe static test article. Strain gage measurements 

should be 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. 

Consider applying unique conditions on the static test article to better represent fatigue loading 

conditions. Loads applied in the static test should conservatively represent actual flight load 

conditions. 

d. Most modifiers and applicants for supplemental type certificates (STCs) evaluate the 

effect of their modification using a comparative fatigue analysis. In this case, it is usually 

sufficient for an applicant to substantiate that the loads and stresses used in their analysis 

conservatively account for the effects of their modification. The substantiation should 

incorporate the validation of stress analysis methods used for static strength substantiation. 

2-10. How do I substantiate a fatigue safe-life with full-scale fatigue tests? 

Full-scale fatigue tests are the most reliable method for establishing the fatigue performance of a 

structure. A full-scale test may use a complete, or nearly complete, airframe with flight, ground, 

and pressurization loads applied to the entire structure. A full-scale test may also be a series of 

separate tests of the individual major subassemblies of the airframe, such as the complete wing 

assembly, the empennage assembly, and the pressurized cabin. In testing major subassemblies, 

you should simulate the loads applied from any omitted structure and any fretting that would 

have been present with the omitted structure. For example, in testing the pressurized cabin, 

accurately simulate the loads and deflections, including assembly preload, imparted on the cabin 

structure by the wing or empennage. 

2-11. What should I include in the test plan for my full-scale fatigue tests? 

Your test plan should include at least the following items:

a. Description of the structure you will test. Describe whether you will test a 

complete airframe or separate major assemblies. Your test plan should include an analysis of the 

loading effects of any omitted structure and how you will simulate the loads imparted on the test

article from these omitted structures. You may omit secondary structures from the full-scale test 

article provided the omitted structure does not contribute to the fatigue performance of the 

primary structure. The effects of secondary structure should be substantiated by analysis or 

previous test experience. Pay particular attention to the effects of fretting of any secondary 

structure you plan to omit. 

11 
AC 23-13A 9/29/05

b. Description of the test apparatus. Describe how the test article will be constrained 

and what effects the attachment to the test fixture will have on distributing internal loads in the 

test article. Describe how the test loads will be applied. The test set-up should be capable of 

accurately controlling the application of test loads. The set-up should record the time history of 

applied test loads. 

c. Validation of internal loads. See paragraphs 2-8 and 2-9 of this AC. 

d. Comparison of predicted internal loads and applied fatigue test loads. Make 

sure that the applied fatigue test loads produce internal loads that conservatively represent the 

predicted internal loads. You could show this with calibrated strain gages on the fatigue test 

article. 

e. Description of loading sequences. The most realistic method of applying fatigue 

test loads is a random application of loads on a flight-by-flight basis. However, you may apply 

test loads in ordered loading blocks. The block length should be no greater than the number of 

flight hours (unfactored) that can be repeated ten times during the expected life. The sequence of 

loads should be from low to high to low within each block and use at least six load levels to 

describe each block. You should apply the ground-air-ground cycles either individually or in 

frequent blocks. 

f. Description of truncated and clipped loads. 

(1) The highest load level applied in the test should not exceed limit load or the 

load level that will be equaled or exceeded ten times in the life of the part. The magnitude of 

these high loads should be reduced, or clipped, down to the load level that will not be equaled or

exceeded ten times in the expected life of the part. The intent of this precaution is to prevent the 

development of localized yielding. 

(2) Your test loads should extend to the lowest level that causes significant fatigue 

damage unless you otherwise account for such loads. To expedite testing, you may omit, or 

truncate, lower level loads and simulate the fatigue damage resulting from these loads by 

applying a limited number of loads at higher load levels. However, enough cycles must be 

applied to account for the effects of fretting. 

(3) Validate the load truncation level. You may do this with either coupon tests or 

analysis using data obtained from fatigue tests of similar structures. 

g. How, when, and where you will inspect the structure. 

(1) You should periodically inspect the test article to check for the development of 

detectable cracks. Use an analytical fatigue analysis of your structure to help determine the 

locations where cracks are likely to develop. Start inspections well before the predicted life of 

the structure. Later inspections should be frequent enough to allow an accurate estimate of the 

time of crack development. Your test plan should define the type of inspection and equipment 

required to perform the inspection. This will insure familiarity and proper calibration of the 

inspection equipment before the start of the test. 

(2) Consider a teardown inspection of the structure after completion of the test. A 

teardown inspection often reveals unpredicted fatigue problems and can be valuable in 

addressing continuing airworthiness issues in the future. 

h. Describe your definition of a completed test. The detection of a fatigue crack is 

not necessarily the end of a fatigue test. 
9/29/05 AC 23-13A

(1) When testing a complete wing assembly, if a crack develops on one side, you 

may want to continue the test to determine when a crack develops on the other side of the wing. 

You would then compute the safe-life based on two test results. Continuation of the test in this 

circumstance is acceptable provided the failed structure does not affect performing the test on the 

other wing. 

(2) Some applicants choose to repair the cracked structure and continue the test. 

Consider the effects of any repair on the distribution of internal loads and the associated effects 

on fatigue performance in the repaired area. We use the initial crack detection to determine the 

safe-life. However, applicants use the information gained from continuing the test to provide a 

greater understanding of the fatigue performance of the structure and as a database in addressing 

continuing airworthiness issues. You should obtain concurrence from the responsible Aircraft 

Certification Office (ACO) before installing any repair on the test article. 

(3) Decide before you start the fatigue test when you will stop if no detectable 

crack is found. Some applicants set a “design life goal.” Once the fatigue test article reaches 

this goal (“design life goal” multiplied by the appropriate scatter factor) without a detectable 

crack, they choose to stop the test at that point. Other applicants choose to let the test continue 

and set the safe-life as high as the test results will allow. 

i. Approval of test plan and conformity inspection. The responsible ACO should 

review and approve your test plan. The FAA must also perform a conformity inspection of the 

test article and test setup.

2-12. Do I need to complete the test before the FAA will issue the type certificate (TC)? 

a. The FAA recognizes the significant time often required to complete a full-scale 

fatigue test; therefore, the FAA may issue a type certificate before the applicant completes the 

full-scale fatigue test. In this instance, the safe-life specified in the limitations section of the 

Airplane Flight Manual (AFM) should be based on one of the following: 

(1) The safe-life shown by an on-going test at the time of TC issuance. The safelife should be computed as shown in the following paragraphs. 

(2) The safe-life substantiated by a fatigue analysis. Your analysis should follow 

the guidance on fatigue analysis included in paragraphs 2-24 through 2-26 of this AC. Your 

analysis may consider information gained from the on-going fatigue test but should not be a 

prediction or extrapolation based solely on the on-going test. 

b. Your certification plan, compliance checklist, and methods of compliance document 

should discuss how to determine the safe-life for TC issuance. In developing your certification 

plan, consider that a revision to the limitations section of the AFM may be all that is required to 

document a safe-life greater than that specified at the time of TC issuance. However, for a 

shorter safe-life, an airworthiness directive (AD) may be necessary. 

2-13. How do I calculate the fatigue safe-life using full-scale fatigue test results? 

a. You can calculate the safe-life from the results of a full-scale fatigue test using the 

following equation:

The Safe-Life = The mean test life ÷ The scatter factor Equation 1 

b. The following two paragraphs discuss the calculation of the mean test life and the 

purpose and derivation of the scatter factor. 

13 
AC 23-13A 9/29/05

2-14. How do I calculate the mean test life? 

a. The calculation of the mean test life might seem a simple issue. However, the 

statistical distribution of fatigue life follows a log-normal statistical distribution; therefore, we 

should not use the arithmetic mean. We should instead calculate the mean based on the 

logarithms of the test lives. You may use the following two equations to calculate the mean test 

life based on the logarithms of the individual specimen test lives. 

( ) ∑ ( =

=

n

i

i testlife

n

meantestlife

1

10 10 log 1 log )

)

 Equation 2 

(meantestlife meantest life 10 log =10 Equation 3 

b. The following example calculation shows the use of Equations 2 and 3. 

Given: Specimen 1 test life: 100,000 hours 

Specimen 2 test life: 200,000 hours 

Using Equation 2: 

( ) [log (100,000) log (200,000)] 2

1 log10 = 10 + 10 meantestlife

 [ ] 5.0 5.30103 5.150515

2

1 = + =

Using Equation 3: 

meantest life 10 141,421 hours 5.150515 = =

c. The mean test life using the logarithmic mean (141,421 hours) is less than you would 

calculate using the arithmetic mean (150,000 hours). 

2-15. What are the scatter factors for full-scale fatigue tests? 

a. The purpose of the scatter factor is to account for the statistical variability inherent in 

the fatigue performance of built-up structures. The equation for the minimum scatter factor, 

given below, assumes a log-normal statistical distribution of fatigue life. 

s

s

p n

n Z

FactorFST Scatter

1

10

+

= σ Equation 4 

where: FST denotes full-scale test. 

 Zp is the normal distribution variate for the specified probability of a 

 detectable crack-free safe-life. 

 σ is the standard deviation of the population of fatigue test lives. For 

 use in this equation, σ , is a percentage of the log of the mean test 

life for the built-up or monolithic structure. 

 ns is the number of fatigue specimens tested. 

b. Reference 3 specified a value for Zp = 3.511. This value for the normal distribution 

variate equates to a probability of 99.9777 percent. For any individual structural component in a 
9/29/05 AC 23-13A

small airplane, the FAA’s specified probability that the component will reach its safe-life without 

a detectable fatigue crack is 99.9777 percent. 

c. The value for the standard deviation of the fatigue test life is usually taken from

historic data. This is due to the high cost of testing sufficient numbers of full-scale test articles 

to develop an accurate value for the standard deviation. For built-up aluminum structures, 

Reference 3 specified a value for σ = 0.14. The following table summarizes the commonly 

accepted values for the standard deviation for various structural metals. 

Table 1. Standard Deviation of

Log Fatigue Life for Built-up Structures Constructed of 

Common Structural Metals

(Reference 5) 

Base Material of Built-Up 

or Monolithic Structure 

σ, Standard Deviation 

(as percent of log life) 

Aluminum alloys 14% 

Steels (ftu = 100 – 200 ksi) 17% 

Steels (ftu = 200 – 300 ksi) 25% 

Titanium alloys 20% 

d. These values for the standard deviation are for complete, built-up structures. Values 

for the standard deviation given in material handbooks, such as Reference 6, are from simple 

coupon tests and should not be used to determine full-scale fatigue test scatter factors. An 

applicant may propose alternate values for the standard deviation provided they have adequate 

substantiating data. Any alternate values for the standard deviation must be coordinated with the 

Small Airplane Directorate Standards Office. 

e. The table below summarizes the required full-scale test scatter factors for aluminum

structures. You may recognize the commonly known scatter factor of 4.0 from Reference 3. If 

your structure is made of a metal other than aluminum, you need to use Equation 4 to calculate 

the correct scatter factor.

Table 2. Full-Scale Fatigue Test Scatter Factors for Aluminum Structures

For Use with Safe-Life Structures 

Number of

Test Specimens 

Required 

Scatter Factor 

1 4.96

2 4.00

3 3.70

Probability of Detectable Crack-Free Safe-Life 

99.9777% (Zp = 3.511)

Standard Deviation of Log Fatigue Life, σ

0.14 (14%) for Aluminum Structures 4 3.54

15 
AC 23-13A 9/29/05

2-16. What about statistical confidence intervals on full-scale fatigue test results? 

a. Certification authorities, military specifications, and industry standards often require 

the use of confidence intervals in the analysis of test results. A common example is the 

requirement for material strength allowables to be given as an “A-basis,” defined as 99 percent 

probability of survival with 95 percent confidence. A common requirement for fatigue is 99.9 

percent probability of survival with 95 percent confidence. The confidence, or confidence 

interval, is a statistical method of estimating the possible range for the true mean of a population 

based on the test results of only a few specimens. The confidence interval gives a lower and 

upper boundary for the likely value of the population mean. In establishing design allowables, or 

design lives, we are interested in the lower boundary for the mean. 

b. In the statistical model we have used to determine the scatter factor, each individual 

fatigue critical component has a probability of reaching its safe-life without a detectable crack of 

99.9777 percent (with no confidence interval on the mean). This probability is equivalent to the 

following: 

(1) With two test results (scatter factor = 4.0): 99.9 percent probability with 91.3 

percent confidence (or 99.822 percent probability with 95 percent confidence). 

(2) With one test result (scatter factor = 4.96): 99.9 percent probability with 93.95 

percent confidence (or 99.868 percent probability with 95 percent confidence). 

c. The scatter factor model we are using provides probabilities nearly equivalent to 

alternate fatigue requirements that use confidence intervals. However, be aware that the CAAs 

of other countries often require the use of a somewhat larger scatter factor that provides 99.9 

percent probability with 95 percent confidence. 

2-17. What counts as one test specimen, and what counts as two test specimens? 

a. Structural components or details that occur more than once in a test article may count 

as two or more test specimens provided the details have the all of following: 

(1) Same structural function. 

(2) Same geometry. 

(3) Same materials.

(4) Same stress level, both flight and ground. 

(5) Same stress concentration factor. 

(6) Same flight and ground load spectrum. 

b. Testing of a complete wing, both the left and right side, generally counts as two test 

specimens. Similarly, certain frame, stringer, cutout, and other details that occur symmetrically 

in a complete pressurized cabin may count as two or more test specimens. 

c. Examples of testing that might count as only one test specimen include details in the 

vertical tail, the keel beam in a pressurized cabin, and non-symmetric details in the pressurized 

cabin, such as entrance and emergency exit doors. 

d. A 1996 policy letter further clarifies when a test article counts as only one test 

specimen. An applicant asked, “Does the fact that each fuselage fitting assembly consists of two 

parts, back to back, mean that the time to initial failure of the one assembly counts as 2 (two) 

results?” Our answer was, “If each part can independently sustain ultimate load for the 

assembly, then our answer is yes. If each part cannot sustain ultimate load, and safe-life 
9/29/05 AC 23-13A

certification is being done, then we consider the fitting assembly as having failed when one part 

has failed, and our answer is no.” 

2-18. Can I use a full-scale test scatter factor lower than that obtained from Equation 4? 

You should not use a full-scale test scatter factor less than that obtained from Equation 4 unless 

you substantiate compliance to fail-safe design requirements. Reference 1 allowed reduction of 

the scatter factor from 4.0 down to 3.0 with an inspection program based on the results of testing 

that today we recognize as crack growth testing. Industry practice and FAA regulation, policy, 

and guidance have evolved considerably since the guidance provided in Reference 3. Today, we 

do not allow a type certificate (TC), amended type certificate, or supplemental type certificate 

(STC) program to rely on inspections alone, without demonstrated residual strength, as a means 

of maintaining continued airworthiness. Use of scatter factors less than those obtained from

Equation 4 without the compensating features of fail-safe design features or significant residual 

strength capability does not provide adequate structural reliability. 

2-19. What about the ‘3-sigma’ method of computing the safe-life? 

a. Some industries and manufacturers use a method of safe-life computation known as 

the “3-sigma” method, wherein the safe-life is determined by subtracting three standard 

deviations from the mean fatigue test life. While this method is in common use, it will not 

provide the same probability of a detectable fatigue crack at the safe-life as our scatter factor 

model. 

(1) Our scatter factor model uses 3.511 standard deviations to obtain the required 

probability, compared to 3.0 standard deviations in the ‘3-sigma’ method. 

(2) Our scatter factor model incorporates an adjustment to account for the number of 

fatigue specimens tested. 

(3) Our scatter factor model incorporates values for the standard deviation of fatigue 

life based on historically accepted values for the standard deviation. 

b. To ensure standardization of fatigue safe-life evaluations, applicants should use the 

scatter factor model approach to compute the safe-life based on full-scale test results. 

2-20. How do you substantiate a fatigue safe-life with component tests? 

a. You can also substantiate the safe-life of a structure with component tests that use 

less than the complete structure or major subassembly. Examples of component tests include 

testing of the wing spar and testing of the wing-center section joint. When testing at the 

component level, take great care to ensure that the test stresses are valid and that you have 

included all critical details in your test. See paragraph 2-9 for a discussion of validation of test 

stresses. 

b. Review the design, stress analysis, static test, strain surveys, tests of similar structural 

configurations, and service experience to make sure that you have identified all structural details 

that are significant for fatigue. It is acceptable to use component specimens that include portions 

of immediately adjacent structural elements such as skins, webs, rib attachments and fasteners 

(without separately loading such elements) to simulate fretting action. The test specimen and 

surface conditions should be representative of the production article. Special attention should be 

given to areas of stress concentration such as holes, joints, changes in section, sharp corners, and 

rough surface finish. 

17 
AC 23-13A 9/29/05

c. Component testing is usually limited to simple and determinate structure that is free 

of stresses due to eccentricities, assembly preload, and so forth, unless enough adjacent structure 

is included to ensure valid test stresses. 

d. You should validate the test stresses, including the peak local stresses, by strain 

survey comparisons with complete structures for the test loading conditions. Validation of the 

test stresses is necessary to account for omitted structure and to gain an accurate knowledge of 

the stress state in the full-scale component. Additional guidance on validation of the internal 

loads for component testing is given in paragraphs 2-8 and 2-9 of this AC. 

2-21. What loads do I apply in a component test? 

There are two types of loadings used in component tests, as follows: 

a. Spectrum loading, similar to full-scale testing, intended to accurately simulate 

operational loadings and stress distributions. 

b. Constant amplitude, or S-N testing. This testing is intended to define the mean S-N 

curve for use in a cumulative damage analysis. Although the loads are constant amplitude, the 

stress distributions in the test specimen and the magnitudes of the applied loads should 

accurately simulate the stress distributions in the full-scale structure and duplicate the range of 

expected operational loadings. 

2-22. For component testing, how many specimens do I need to test? 

a. If you test using spectrum loading, you should test at least three specimens 

representative of the most critical location of each type of detail. 

b. If you are testing using constant amplitude, or S-N testing, test a sufficient number of 

specimens to define the mean S-N curve. Chapter 9 in Reference 6 provides guidance on 

methods of defining the mean S-N curve. 

2-23. For component testing, what scatter factor should I use? 

a. If your test uses spectrum loading, the scatter factor and safe-life are calculated 

similarly as for a full-scale test. The following equation defines the minimum scatter factor for 

component testing with spectrum loading. 

 CT FST Scatter Factor = 1.5× ScatterFactor Equation 5 

where: CT stands for component test 

 FST stands for full-scale test

b. The following example calculation shows the use of Equation 5. 

Given: Three (3) component aluminum specimens 

 From Equation 4: Scatter FactorFST = 3.70 

 Using Equation 5: Scatter FactorCT = 1.5 x 3.70 = 5.55 

c. A scatter factor larger than that given by Equation 5 may be necessary if the 

component test specimen and test loading do not accurately simulate operational loadings, stress 

distributions, and the full-scale structure. An applicant may propose a lower scatter factor if the 

component test accurately simulates a full-scale test. Any reduced scatter factor must be 

coordinated with the Small Airplane Directorate Standards Office. 
9/29/05 AC 23-13A

d. For constant amplitude, or S-N testing, the component test results are used to develop 

the mean S-N curve for use in a cumulative damage analysis. The scatter factors, or statistical 

adjustments, applied to cumulative damage analysis are discussed in the following paragraphs. 

2-24. How do you substantiate a fatigue safe-life with analysis supported by test 

evidence? 

a. You may substantiate the fatigue safe-life of a structure by fatigue analysis supported 

by test evidence. The analysis method most commonly used is the Palmgren-Miner hypothesis, 

also known as Miner’s Linear Cumulative Damage Theory. 

b. The basic philosophy of Miner’s Theory is that the fatigue damage introduced by a 

given stress level is proportional to the number of applied cycles at that stress level divided by 

the total number of cycles to failure at that stress level. This ratio is called a cycle ratio and is 

used to measure damage. If you apply various levels of stress amplitude to the structure, the 

total damage is the sum of the cycle ratios. Failure is predicted to occur when the sum of the 

cycle ratios is equal to one, or the following: 

k

k

k

i i

i

N

n

N

n

N

n

N

n ∑ = = + + + =

1.0 ......

2

2

1

1

1

 Equation 6 

 where: ni is the fatigue damage due to the applied number of cycles 

at a certain stress level 

 Ni is the number of cycles to failure at that stress level. 

k is the number of intervals the stress spectrum is divided 

into. Typically, the stress spectrum is divided into 

 increments corresponding to 0.2g. 

c. Reference 7 provides a sample calculation using Miner’s Theory. The Small 

Airplane Directorate has developed a computer program based on Miner’s Theory. The 

Directorate makes this software available to airframe Designated Engineering Representatives 

(DERs). For additional information on the software, please see Reference 8. 

2-25. What are some concerns with substantiation by fatigue analysis? 

a. Accuracy of the cumulative damage hypothesis. Several references, including 

Reference 9, compare spectrum test results to cumulative damage predictions. These references 

show that fatigue failure may occur between damage summations between 0.5 to 2.0, depending 

on various loading parameters. In particular, severe usage, including acrobatic and agricultural 

usage, may significantly reduce the accuracy of the cumulative damage calculation. 

b. Adequacy of S-N data. 

(1) Reference 6, formerly MIL-HDBK-5, includes the following disclaimer 

regarding handbook S-N data: 

“The S-N or ε–N curves may not apply directly to the design of structures because they may not 

take into account the effect of specific concentration associated with reentrant corners, notches, 

holes, joints, rough surfaces, and other similar conditions, which are present in fabricated parts.” 

(2) Handbook S-N data may also not be directly applicable to your structure due to 

small differences in the material used to develop the handbook curves and the material used in 

your structure. Also, data for the exact stress concentration factors in your structure may not be 

available in the handbook data. 

19 
AC 23-13A 9/29/05

(3) Applicants may address these issues by adjusting the S-N curves to account for 

differences between the actual structure and the structure, or coupon, used to develop the S-N 

curves. Applicants should adjust the S-N curves for differences in geometry, bearing and bypass 

stress, including eccentric loads and orientation of principal stresses. You may use component 

testing or previous testing experience with similar structural details to select and adjust the most 

applicable S-N curves. In determining the applicability of S-N data, similar structures are those: 

(a) Constructed of the same material (same form, same heat treat). 

(b) Constructed with the same or similar manufacturing methods. 

(c) With the same or similar structural function and geometry. 

(d) With the same or similar stress concentration factor. 

(e) With the same or similar stress level and loading spectrum. 

(f) With the same or similar principal stresses, for example axial and bending 

loads. 

For complex details, including joints and fittings, you should not use S-N data without first 

adjusting the data for applicable component level or full-scale test results. 

c. Accuracy in determining the stress concentration factor, Kt. Identify all areas of 

stress concentration in the structure and determine the magnitude of the stress concentration 

factor, Kt, associated with each of these details. Accurate determination of Kt is necessary to 

make sure that you are using the correct S-N data in your fatigue analysis. You may determine 

stress concentration factors for some details, including open holes, radii, filets, and so forth, from

handbook references. Stress concentration factors for more complex details, including fittings 

and joints, may require detail analysis, including validated finite element methods. In this case, 

the determination of Kt should include consideration of eccentric loadings and the bearing and 

bypass stresses at each bolt or rivet in a joint. 

d. Accuracy in determining the stress level. Since you will not conduct full-scale 

fatigue tests or spectrum loading component level tests on your structure, the quality, detail, and 

completeness of your stress analysis is important. In fatigue substantiation by analysis, we rely 

completely on the stress analysis to identify all fatigue critical details. Applicants should follow 

the guidance in paragraphs 2-8 and 2-9 of this AC to validate the stresses used in a fatigue 

analysis. 

2-26. What scatter factors do I use in a fatigue analysis? 

The scatter factor you use in a fatigue analysis is larger than the scatter factors you would apply 

to full-scale and component level test results. This is due to the uncertainties in a fatigue 

analysis outlined in the previous paragraph. The scatter factor used in an analysis depends on the 

type of metal, statistical basis, and applicability of the S-N data used in the analysis.

a. For aluminum structures, you may use the following: 

(1) You may use the S-N data provided in Appendix 2. This S-N data is applicable to 

conventional built-up aluminum structure with no fittings (other than continuous splice fittings), 

no parts with high residual stresses, no unique structural features, and no stress concentrations 

greater than Kt = 4. This data is carried over from the guidance provided in Reference 3. Use of 

this S-N data, combined with a scatter factor of 8.0, has provided satisfactory service experience. 

(2) You may use proprietary, handbook, or material supplier mean S-N curves 

adjusted for any differences between the actual structure and the structure or coupon used to 
9/29/05 AC 23-13A

develop the base S-N curves. Test adjusted data is recommended for lugs, joints, and fittings. In 

this case, you may use a minimum scatter factor of 8.0. 

b. Adjustments to the typical scatter factor of 8.0 must be made for the following: 

(1) Structures made out of metals other than aluminum. 

(2) Structures, including aluminum, whose S-N curves have a statistical basis other 

than the mean (for example, curves based on 99 percent probability of survival). (Note: The S-N 

data provided in Reference 6 represent the mean S-N curve, with no adjustments for confidence

intervals.)

(3) Any adjustments to the scatter factor must be coordinated with the Small Airplane 

Directorate Standards Office. 

2-27. How can I extend the safe-life, or retirement life, of my airplane? 

a. Extensions of the safe-life are accomplished either as an amended type certificate 

(TC) by the TC holder or as a supplemental type certificate (STC) either by the TC holder or an 

STC applicant. In formulating a life extension program, applicants should consider the 

following: 

(1) Inspection programs alone are not sufficient to extend the safe-life of an airplane. 

(2) The safe-life of a structural component is usually set to the detail that has the 

shortest fatigue life. An applicant should address the possibility that other structural details or 

components may have nearly as short a fatigue life as the detail you are extending the life of. To 

avoid the development of unforeseen fatigue cracks in other parts of the structure, you should 

complete a fatigue evaluation of the entire structure. 

b. One method of extending the safe-life is a thorough re-evaluation of the fatigue 

performance of the structure. 

(1) A re-evaluation of the fatigue performance requires detailed knowledge of the 

data, including the loading spectra, stress levels, stress concentration factors, and S/N curves, 

used to substantiate the original type design. STC applicants are at a particular disadvantage in 

fatigue life extensions since they typically do not own or possess the type design data. 

(2) It may be difficult for you to substantiate a life extension through analysis alone. 

An analytical approach would require you to show that the type design analysis was overly 

conservative or that your analysis, including assumptions and data, is more accurate than that 

used in the type design substantiation. 

(3) A full-scale fatigue test of the airplane’s structure is the preferred method of 

extending the original safe-life, especially if you know that the original design was substantiated 

by analysis. In this case, follow the guidance provided in this AC for full-scale fatigue testing, 

including mission profile development, selection of load spectra, and validation of internal loads. 

Consider in-flight measurements of loads and stresses to further justify the life extension. 

(4) In your fatigue test, you should use an airframe with zero flight time. This is 

necessary since we do not know the usage history of any structural member that has accumulated 

flight time. It is also possible that an airframe that has accumulated flight time may have 

experienced an overload event that may have altered the fatigue performance of the structure. 

c. One life extension method that applicants have used successfully has been to provide

fail-safe design features to an existing safe-life design. In this approach, an applicant completes 

all of the following: 

21 
AC 23-13A 9/29/05

(1) Incorporates redundant structure into the existing design. For example, an 

external spar strap that provides the required residual strength capability.

(2) Inspects the structure for cracks. Any parts with existing cracks are replaced. 

(3) Oversizes existing bolt and rivet holes to remove any small cracks not detected 

during inspection. 

(4) Incorporates cold working of holes and stress concentrations to improve fatigue 

performance. 

(5) Develops an inspection program for the fail-safe and redundant structure. 

d. Life extensions must be approved by the responsible Aircraft Certification Office 

(ACO) and coordinated with the Small Airplane Directorate Standards Office. 
9/29/05 AC 23-13A

CHAPTER 3. FAIL-SAFE DESIGN

3-1. What is a fail-safe design? 

A fail-safe design is a design that retains its required residual strength after the failure or partial 

failure of a principal structural element. A fail-safe design typically consists of the fail-safe 

component or primary structural element, and a redundant or backup structural element. A failsafe design is, therefore, often said to be a redundant design or a multi-load path design. Failure 

in a multi-load path design is the inability to support the required residual strength loads or the 

inability to prevent excessive deformation that interferes with continued safe flight and landing. 

Follow the general guidance in sections 2-4, 2-7, 2-8, 2-9, and 2-11 of this AC for developing the 

loading spectra, mission profile, and test plan for evaluating your fail-safe design. 

3-2. What is the history of fail-safe design? 

a. Fail-safe design has been an option available to applicants since the first specific 

fatigue requirement in Amendment 3-2 to CAR 3. A fail-safe evaluation generally includes the 

following: 

(1) Establish the components you will make fail-safe. This includes identifying 

which primary structural elements will have redundant load paths. 

(2) Define the loading conditions, including the magnitude of external loads. This 

step includes determining the internal loads in the primary and redundant structure in both the 

pristine and damaged conditions. 

(3) Determine the extent of damage for which you will design the fail-safe structure. 

(4) Conduct structural tests and analyses to substantiate that you have achieved the 

fail-safe design objective. 

(5) Establish inspection programs aimed at early detection of fatigue damage. 

b. In the more than 45 years since CAR Amendment 3-2 first introduced the fail-safe 

option for addressing fatigue requirements, the redundant structure provided by fail-safe design 

has prevented otherwise catastrophic failures. However, there are well-documented cases in 

which a design thought to be fail-safe was not, and the design failed to prevent a catastrophic 

failure. Several references, including References 10 and 11, identify the following potential 

problems of designs mistakenly thought to be fail-safe: 

(1) Potential loss of fail-safe attributes with time due to normal fatigue wear-out. 

(2) Difficulty in making an accurate prediction and validation of failure modes. 

(3) Incorrect assumptions that a design is sufficiently and consistently selfannunciating. 

(4) Loss of type design strength due to inadequate crack detection. If a crack or 

other damage is not found by inspections or is not so obvious as to be detected before further 

flight, the capability of the structure may be below the design limit and ultimate load capability. 

(5) Inadequate residual life with obvious damage present. Redundant structure may 

not have sufficient safe-life to ensure that damaged structure will be found by inspections. 

23 
AC 23-13A 9/29/05

3-3. How can I assure that my design complies with the fail-safe design requirements? 

a. The historical guidance provided in Reference 3 described the importance of repeated 

inspections in maintaining the continued airworthiness of a fail-safe design. Detecting fatigue 

cracks before they become dangerous is the ultimate control to ensure the fail-safe characteristics 

of flight structure and pressurized cabin. For a fail-safe design, applicants should develop an 

inspection program capable of detecting fatigue cracks and other partial failures. Applicants 

should provide enough guidance information to aid operators in establishing the frequency and 

extent of the repeated inspections of the critical structures or critical areas. Include instructions 

for these inspections, including inspection schedule and inspection methods, in the information 

required for § 23.1529. 

b. You could develop an inspection program based on: 

(1) A service history based approach. For guidance on service history based 

inspections, see AC 91-60. 

(2) A damage tolerance approach. 

(3) A hybrid approach in which the inspection threshold equals the safe-life 

determined using Equation 1 and repeat inspection intervals based on crack growth analysis. 

3-4. What are some design considerations in achieving a fail-safe design? 

a. Use of multipath construction and providing crack stoppers to limit the growth of 

cracks. 

b. Use of composite (more than one element) duplicate structures so a fatigue failure 

occurring in one-half of the composite member will be confined to the failed half. Also, the 

remaining structure will still possess the load carrying ability required by § 23.571 and § 23.572. 

c. Use of backup structure wherein one member carries the entire load, with a second 

member available that can assume the extra load if the primary member fails. 

d. Selection of stress levels and materials with low notch sensitivity (particularly for 

components with high stress concentration) that provide a controlled slow rate of crack 

propagation combined with a high residual strength after initiation of cracks. 

e. Arrangement of design details to allow easy detection of failures in all critical 

structural elements before the failures can become dangerous or result in a loss of strength below 

that required by § 23.571 and § 23.572. You should also arrange design details to allow 

replacement or repair. 

Note: The last two items are examples of good design practice and improve fail-safe design 

concepts, but they cannot be used alone to achieve fail-safe design. 

3-5. What is a Principal Structural Element? 

A principal structural element (PSE) is an element that contributes significantly to carrying 

flight, ground, or cabin pressurization loads, and whose integrity is essential in maintaining the 

overall structural integrity of the airplane. (Note: Part 23 fatigue evaluation requirements do not 

apply to landing gear or unpressurized fuselage structure; however, ground loads are to be 

included to the extent that they affect wing, empennage, canard or pressurized cabin structure.) 

Typical examples of PSEs are as follows: 

a. Wing, horizontal stabilizer, vertical fin, canard, forward wing, winglets/tip fins: 
9/29/05 AC 23-13A

(1) Fixed surface, stabilator, or trimmable stabilizer attachment fittings; 

(2) Integrally stiffened plates; 

(3) Primary fittings; 

(4) Principal splices; 

(5) Skin or reinforcement around cutouts or discontinuities; 

(6) Skin-stringer combinations; 

(7) Spar caps; and 

(8) Spar webs. 

b. Pressurized cabin: 

(1) Circumferential frames and adjacent skin; 

(2) Pressure bulkheads; 

(3) Cockpit window posts; 

(4) Skin and any single frame or stiffener element around a cutout; 

(5) Skin or skin splices, or both, under circumferential loads; 

(6) Skin or skin splices, or both, under fore and aft loads; 

(7) Skin around a cutout; 

(8) Skin and stiffener combinations under fore and aft loads; 

(9) Door frames, skins, and latches; and 

(10) Window frames. 

3-6. How much fatigue damage should I consider in my fail-safe evaluation? 

Each particular design should be carefully assessed to establish appropriate damage criteria. As 

discussed in Reference 10, you should not rely on the detection of an “obvious partial failure.” 

Instead, you should consider fatigue damage in terms of the complete failure of the single 

element involved. Typical examples of the fatigue damage you should consider are outlined 

below: 

a. Skin cracks in splice joints and those emanating from the edge of structural openings 

or cutouts that can be readily detected by visual inspection of the area. 

b. Failure of one element where dual construction is used in components such as spar 

caps, window posts, window or door frames, and skin structure. 

c. The presence of a fatigue failure in at least the tension portion of the spar web or 

similar elements. 

d. Failure of one element of primary attachments, such as wing and empennage fixed 

surface, or stabilator attach fittings. 

e. Excessive loss of stiffness under load as evidenced by excessive deformation. 

f. Failure of skin, frames, stringers, and pressure bulkheads in pressurized cabins. 

25 
AC 23-13A 9/29/05

3-7. How do I address inaccessible areas? 

Designers should make every reasonable effort to ensure all principal structural elements can be 

inspected, as required by § 23.611. If inaccessible or blind areas are unavoidable, place 

emphasis on determining crack propagation and residual strength of the particular fatiguedamaged structure to ensure continued airworthiness of the structure with reasonable inspection 

methods and controls by the operator. Also consider providing additional fatigue strength to 

preclude fatigue cracking in the blind element or to conduct fatigue tests of the blind areas to 

establish that a high service life is provided. Pay particular attention to corrosion prevention in 

inaccessible areas. 

3-8. How do I account for dynamic effects? 

Apply the dynamic magnification factor of 1.15, required by §§ 23.571 and 23.572, to all loads, 

including pressure cabin loads. You may also make fail-safe cuts under load to show a different 

value for the factor, or you may show that the dynamic effects are negligible by comparison to 

dynamic test data from a similar structure. 

3-9. How do I test principal structural elements? 

a. The nature and extent of tests on either a complete structure or portions of the 

primary structure, or both, will depend upon previous experience with similar types of structures

and the crack propagation characteristics of the structure. Single elements or members, such as 

stringers and spar caps, should be completely severed and 1.15 times the critical fail-safe load 

applied after severing. If definite evidence is furnished that the dynamic failure effects are not 

present, the 1.15 factor may be eliminated or reduced in accordance with the effects noted. The 

required residual strength of the remaining structure is a static ultimate load factor of 75 percent 

of the critical limit load factor at VC, per §§ 23.571 and 23.572. 

b. You may apply the fail-safe loads to the structure before severing and omit the 1.15 

factor. In this case, the test specimen and test fixture must be carefully designed to ensure that 

the correct dynamic effects are obtained. If you are testing distributed members, such as a sheetstringer combination or an integrally stiffened tension skin, make a cut to represent an initial 

crack in the element under test. If there is no failure, you may increase the length of the cut with 

the fail-safe load applied until one of the following occurs: 

(1) The fail-safe damage has been simulated; or 

(2) The crack propagation rate decreases due to redistribution of load paths; or 

(3) Crack propagation stops due to a crack stopper. 

c. The simulated cracks should be as representative as possible of actual fatigue 

damage. If it is not practical to produce actual fatigue cracks, damage may be simulated by cuts 

made with a fine saw, sharp blades, or a guillotine. If saw-cuts in primary structure are used to 

simulate sharp fatigue cracks, sufficient evidence should be available from element tests to 

indicate equivalent residual strength. If it is necessary to simulate damage at joints or fittings, 

bolts may be removed to simulate the failure if this condition represents an actual failure. 

d. Failure is the inability of the redundant structure to support the required residual 

strength loads or the inability to prevent excessive deformation that would interfere with safe 

flight and landing. 
9/29/05 AC 23-13A

3-10. Can I use analysis to substantiate a fail-safe design? 

a. Sometimes you can substantiate fail-safe characteristics analytically. You may use 

an analytical approach when the structural configuration involved is similar to one already 

verified by fail-safe tests, whether conducted on a previously approved type design or on other 

similar areas of the design currently being evaluated. 

b. You may also use the analytical approach when conservative failures are assumed. 

However, the failure must be detected long before the critical crack length is approached, and the 

margins of safety resulting from the analysis must be considerably more than the fail-safe 

residual static strength level. In any such analysis, include the 1.15 dynamic magnification 

factor unless it can be shown that this factor is not required. 

3-11. How do I select the critical areas? 

Paragraph 3-5 of this AC lists typical single principal structural elements and detail design points 

that require fail-safe evaluation. The process of determining where you should simulate fail-safe 

damage in an element, such as a wing spar cap or fuselage frame, requires use of sound 

engineering judgment that considers various factors, such as the following: 

a. Conducting an analysis to locate areas of maximum stress and a low margin of 

safety. 

b. Conducting strain gage surveys on undamaged structure to establish points of high 

stress concentration as well as the magnitude of such concentration. 

c. Examining static test results to determine locations where excessive deformations 

occurred. 

d. Determining from repeated load tests where failure may have initiated or where the 

crack propagation rate is a maximum. 

e. Selecting location in an element (such as a spar cap) where the stresses in adjacent 

elements (such as the spar web or wing skin) would be the maximum with the spar cap failed. 

f. Selecting points in an element (such as a spar web or frame) in which high stress 

concentrations are present in the residual structure with the web failed. 

g. Assessing detail design areas that service experience records of similarly designed 

components indicate are prone to fatigue damage.

h. Locating areas susceptible to operational damage, such as foreign object damage, 

corrosion, and so forth. 

27 
AC 23-13A 9/29/05

CHAPTER 4. DAMAGE TOLERANCE EVALUATION 

The FAA has not developed guidance unique to part 23 damage tolerance evaluations. For 

damage tolerance evaluations, use the guidance provided in AC 25.571C, Damage Tolerance and 

Fatigue Evaluation of Structure. Follow the general guidance in sections 2-4, 2-7, 2-8, 2-9, and 

2-11 of this AC for developing the loading spectra, mission profile, and test plan for evaluating 

your damage tolerant design. 

4-1. Are there any differences between the part 23 and part 25 damage tolerance 

regulations? 

There is one significant regulatory difference between part 23 and part 25. Part 25 includes a 

discrete source damage requirement. There is no requirement for substantiation of discrete 

source damage capability in part 23. However, be aware that a discrete source damage analysis 

might be necessary to show compliance to rotorburst requirements. 
9/29/05 AC 23-13A

CHAPTER 5. ROTORBURST REQUIREMENTS 

5-1. What are the requirements for turbine or turbopropeller installations in small 

airplanes? 

a. Compliance with 14 CFR, part 23, § 23.903(b)(1) requires the applicant to take 

design precautions to minimize the hazards to the airplane, including the airframe, due to an 

engine rotor failure. A Small Airplane Directorate policy letter dated April 7, 1989, defines 

“minimize” in the context of § 23.903 as “to reduce to a minimum, decrease to the least practical 

amount. The 'least practical amount' obviously has to include consideration of what is attainable 

with current technology and materials.” 

b. Industry practice for part 23 turbine-powered airplanes is to use discrete source 

damage analysis to evaluate design precaution options. Applicants have used the discrete source 

damage evaluations to select the most effective, appropriate design precautions and to 

substantiate that the selected design precautions minimize the rotorburst hazard to the least 

practical amount. The design precautions selected by these applicants to show compliance to 

§ 23.903(b)(1) have provided sufficient residual strength capability to demonstrate compliance 

with the discrete source damage requirements in 14 CFR, part 25, § 25.571(e)(3). 

5-2. I thought there was not a discrete source damage requirement in part 23? 

a. Part 23 does not include a discrete source damage requirement in the fatigue 

regulations. However, small airplane applicants have used, and the FAA has accepted, discrete 

source damage analysis as a quantitative method of substantiating that selected design 

precautions minimize the rotorburst hazard to the least practical amount. 

b. AC 20-128A, Design Considerations for Minimizing Hazards caused by Uncontained

Turbine Engine and Auxiliary Power Unit Rotor Failure, provides a means, but not the only 

acceptable means, of showing compliance to § 23.903(b)(1). The AC uses discrete source 

damage criteria to include airframe hazards in a safety assessment and risk analysis. The correct 

interpretation of the AC is that the discrete source damage evaluation is required for all 

categories of airplanes. If you choose to show compliance using AC 20-128A, include airframe

hazards, via discrete source damage analysis, in the rotorburst safety assessment and risk 

analysis. 

29 
AC 23-13A 9/29/05

CHAPTER 6. FLIGHT WITH KNOWN CRACKS 

6-1. What is the purpose of this policy? 

This policy is provided as guidance for Aircraft Certification Office (ACO) engineers when the 

FAA’s Flight Standards Service requests engineering support. The Flight Standards Service may 

request support from an ACO when addressing a known cracking situation in an individual 

airplane. 

6-2. What is the policy for operational flight with known cracks? 

It is the Small Airplane Directorate’s policy to not allow continued operational flight with known 

cracks in primary structure. An airplane with a known crack in the primary structure no longer 

meets its type design and may no longer possess its type design strength. 

6-3. Are there any exceptions to this policy? 

a. The exceptions to this policy, to be taken on a case-by-case basis, are the following: 

(1) Substantiation that the cracks are not in primary structure. 

(2) Substantiation of the ability of single load path structure with known cracks to 

carry ultimate loads. Only in unusual circumstances, such as the difficulty of an operator in 

obtaining replacement parts, will this be allowed. The time allowed for replacement or repair 

must be as short as practical. 

(3) Substantiation that the cracks are in fail-safe structure. The ability to sustain 

ultimate load with the maximum permissible crack must be substantiated. Temporary repairs, 

such as "stop drilling," should be specified. The conditions of a temporary repair, and the time

allowed before a permanent repair is required, may be less restrictive than for single load path 

structure. 

b. The Aircraft Certification Office (ACO) responsible for the airplane’s type certificate 

should make each specific approval for flight with known cracks. The ACO issuing the approval 

should coordinate the approval with the Small Airplane Directorate Standards Office. 

6-4. Which airplanes are covered by this policy? 

This policy applies to all normal, utility, acrobatic, or commuter category airplanes, regardless of 

certification basis. This policy provides acceptable methods for any airplane with a known 

crack, even those whose certification basis does not include specific fatigue requirements, to 

substantiate a limited time for continued operational flight. The policy does not expand current 

fatigue requirements to existing designs; it provides methods to substantiate that, in certain 

unusual circumstances, an airplane with a known crack may be acceptable for limited operational 

flight. 

6-5. What will the FAA consider before allowing operational flight with known cracks? 

Continued operational flight with known cracks in a structure whose failure may be catastrophic, 

in general, is not permissible. However, there are certain unusual circumstances under which 

flight is permissible for a limited period of time with known cracks in the structure. Any 

approval for flight with known cracks should meet all the following criteria. 
9/29/05 AC 23-13A

a. There must be an unusual need for allowing flight with known cracks. Unusual 

needs are those that arise from such things as the scarcity of parts or operator hardship. For 

example, unavailability of parts or tooling to support an immediate part replacement would 

qualify as an unusual need. Another would be the need for operational flight to a repair facility 

from an area where insufficient technical or maintenance resources are available to perform a 

repair. Deferral of corrective action should not be approached as either business as usual or as a 

matter of convenience to operators. 

b. The level of safety provided by the certification basis must not be compromised 

by the allowance to operate with a known crack. At all times during the expected crack 

growth phase, the airplane must meet its limit and ultimate load requirements with the crack 

present. Part modification, replacement, or repair must be implemented if the structure's 

strength, with the crack present, is anticipated to fall below these load requirements before the 

next inspection of the structure. Specifically, this includes the following: 

(1) Cracked structure must support limit load without detrimental permanent 

deformation (§ 23.305(a)). 

(2) Cracked structure must support ultimate load for three seconds without failure 

(§ 23.305(b)). 

(3) Airplanes certified as fail-safe should remain fail-safe with the damage present. 

This means that the cracked structure must be able to bear ultimate load and meet the fail-safe 

criterion specified in the airplane’s certification basis. 

(4) Any other certification requirement identified by the ACO as being potentially 

violated by repair deferral should be addressed. 

(5) Single failure requirements remain in effect with the damage present (e.g.,

§ 23.629 (g)(h)). 

c. There should be a high degree of confidence in the evaluation used to justify the 

particular case. Regardless of the certification basis of the airplane, a fracture mechanics 

assessment must be performed for the cracked structure before further flight. The evaluation 

should determine that the material, location, and type of damage is such that a reliable fracture 

mechanics assessment can be performed. The following items are especially critical to a 

thorough evaluation: 

(1) An investigation into the cause of the crack. This investigation should include 

a complete characterization of the crack location, size, and shape. The investigation should 

estimate the time of crack initiation and growth rate, using microscopic mapping techniques 

when possible. Identify the presence of any corrosion or manufacturing or material defect. 

Review the results of any previous inspections of the cracked part to find any indications of 

abnormalities. The investigation should also review the assumptions used in the cracked 

structure’s design. This should include a comparison of the load spectra used in the type 

certification fatigue evaluation and actual usage, validation of internal loads, and verification of 

material properties. 

(2) Static strength and fracture mechanics analysis. The fracture mechanics 

analysis must provide a high degree of confidence that the crack will not grow faster than the 

predicted rate with the expected usage until the next inspection, or until it is repaired. All 

residual strength substantiations should be shown by a conservative analysis, by test, or by test 

combined with analysis. Analysis is allowed only where experience shows that analysis is 

reliable. 

31 
AC 23-13A 9/29/05

(3) Inspection intervals. 

(a) This assessment must establish a viable inspection program for the cracked 

structure. The inspection program should provide a high level of confidence that the crack will 

not be allowed to grow to a size that would compromise the safety provided by the certification 

basis. The probability of detection and reliability of the inspection must be high. The inspection 

method should be suitable for the situation, and the cracked structure should be adequately 

accessible. The inspection program, along with any other limitations, must be placed in the 

mandatory section of the airplane’s maintenance manual for the period of time in which the 

cracked structure remains in operation. 

(b) Inspection intervals should be conservatively determined based on approved 

crack growth data and a fracture mechanics based evaluation. The inspection interval is 

established by calculating the crack growth time from the detected crack size (or maximum of 

detected size range for non-visual) for the Non-Destructive Inspection (NDI) method used to 

detect critical crack size and dividing that number of flight hours or flights by a conservative 

value. Inspection intervals should provide for at least two opportunities to detect crack growth 

before the crack size reduces the capability of the structure below that provided by the 

certification basis. 

(4) Temporary intervention. If possible, some intervention, such as stop drilling of 

the crack tip, should be implemented to slow crack growth (stop drilling is not considered a 

repair). 

(5) Single-Load path structures. Flight with known cracks is strongly discouraged 

for structure where applied loads are eventually distributed through a single member, the failure 

of which would result in the loss of the structural capability to carry the applied loads. If

cracking is found in this type of structure, the part must be replaced or repaired before further 

flight. Exceptions may be permitted only if all other criteria of this policy are met, stringent

inspection requirements are specified, and a conservative maximum permissible crack size able 

to sustain ultimate load is specified. A combination of a conservative fracture mechanics 

analysis correlated to test results, or conservative fracture mechanics analysis with added scatter 

factors, must be performed to allow continued operation with cracks. 

(6) Multiple-Load path structures. For redundant structures in which, with the 

failure of individual elements, the applied loads would be safely distributed to other load 

carrying members, adjacent structure must be inspected and be free of cracks or other damage. 

Various combinations of analysis and test, including that provided at the time of original 

certification, may be considered as ample substantiation. 

(7) Structures with multiple-site damage, corrosion and stress corrosion cracks. 

Continued operational flight with known cracks in instances of multiple-site damage, cracks in 

the presence of corrosion, or stress corrosion cracks should be avoided. 

d. The approval should provide a limited time for continued flight with known 

cracks. Under no circumstances can allowance for operation with known cracks be considered 

more than a temporary condition. The limited period must be based on the need for the deferral 

of the repair, replacement, or modification. Once the unusual need for the deferral has passed, 

the deferral must be eliminated, and the part repaired or replaced. 

e. The ACO approving flight with known cracks should coordinate each approval 

with the Small Airplane Directorate Standards Office. The approval to allow flight with 

known cracks must be clearly made, documented, and approved by the FAA before any flight 

with known cracks. The ACO responsible for the type certificate should make each specific 
9/29/05 AC 23-13A

approval for flight with known cracks in coordination with the Small Airplane Directorate 

Standards Office. 

33 
AC 23-13A 9/29/05

REFERENCES

1. FAA Report No. ACE-100-01, Fatigue Evaluation of Empennage, Forward Wing and 

Winglets/Tip Fins on Part 23 Airplanes, dated February 15, 1994. 

2. DOT/FAA/CT-91/20 General Aviation Aircraft – Normal Acceleration Data Analysis 

and Collection Project, dated February 1993. 

3. FAA Report No. AFS-120-73-2, Fatigue Evaluation of Wing and Associated Structure on 

Small Airplanes, dated May 1973. 

4. FAA Memorandum, Method of Establishing Flight Load Spectra for Safe-Life Fatigue 

Analysis, dated March 25, 2005. 

5. A Review of Fatigue Scatter Factors with Particular Reference to Civil Aircraft, M. B.

Benoy, Civil Aviation Authority of Australia, July 1978. 

6. DOT/FAA/AR-MMPDS-01, Metallic Materials Properties Development and 

Standardization, dated January 2003. 

7. FAA Memorandum, Example Cumulative Damage Calculation for Safe-Life Fatigue 

Analysis, dated February 7, 2005. 

8. User’s Instructions for Computer Program to Calculate Fatigue Safe-Life (Unfactored) 

for Small Airplanes, dated March 1996. 

9. Metal Fatigue, J. A. Pope (Editor), Chapman and Hall Ltd, dated 1959. 

10. A Critical Review of Strategies Used to Deal with Metal Fatigue, by Bob Eastin, Federal 

Aviation Administration, Presented at the International Committee on Aeronautical Fatigue 

(ICAF) 2003. 

11. Commercial Airplane Certification Process Study, An Evaluation of Selected Aircraft 

Certification, Operations, and Maintenance Process, Federal Aviation Administration, March 

2002. 

Note: You may obtain References 1, 3, 4, 7, and 8 from the Federal Aviation Administration, 

Small Airplane Directorate Standards Office, telephone number 816-329-4111. 
9/29/05 AC 23-13A 

Appendix 1

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA 

A1-1. What flight and ground load spectra are presented in this appendix? 

The following list summarizes the flight and ground load spectra presented in this appendix. 

Flight Spectra, General Usage, Single-Engine Unpressurized 

Description Graphic Data Tabulated Data Comments

Gust spectra for single-engine,

unpressurized operations, 

including basic instruction, 

personal, executive, and 

acrobatic usage. 

Figure A1-1 Table A1-1 These spectra should also be used for 

pressurized single-engine airplanes that

spend a significant amount of flight time at

low altitudes. 

Maneuver spectra for singleengine basic instruction usage.Figure A1-2 Table A1-2 

Maneuver spectra for singleengine personal usage. Figure A1-3 Table A1-3 An airplane in the personal usage category

has a single, reciprocating engine with 185

horsepower or less.

Maneuver spectra for singleengine executive usage. Figure A1-4 Table A1-4 An airplane in the executive usage category 

has a single, reciprocating engine with more

than 185 horsepower. The executive usage 

category also includes unpressurized, singleengine turboprop airplanes. 

Maneuver spectra for singleengine acrobatic usage. Figure A1-5 Table A1-5 Applicable to typical acrobatic category 

airplanes, nz = +6g / -3g 

Flight Spectra, General Usage, Twin-Engine Unpressurized 

Description Graphic Data Tabulated Data Comments

Gust spectra for twin-engine, 

unpressurized operations, 

including instruction, and 

general usage.

Figure A1-6 Table A1-6 These spectra should also be used for 

pressurized twin-engine airplanes that may 

spend a significant amount of flight time at

low altitudes, 7,000 feet and below. An 

example of this type of operation is short

flight duration commuter and air taxi

operations. 

Maneuver spectra for twinengine instruction usage. Figure A1-7 Table A1-7 

Maneuver spectra for twinengine general usage. Figure A1-8 Table A1-8 These spectra should also be used for 

pressurized and unpressurized twin-engine 

airplanes that may be operated in short flight 

duration commuter and air taxi operations.

A1-1 
AC 23-13A 9/29/05 

Appendix 1

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

Flight Spectra, General Usage, Single-Engine and Twin-Engine Pressurized

Description Graphic Data Tabulated Data Comments

Gust spectra for pressurized 

usage. 

Figure A1-9 Table A1-9 

Maneuver spectra for 

pressurized usage. 

Figure A1-10 Table A1-10 

Flight Spectra, Special Usage – Agricultural (Aerial Application)

Description Graphic Data Tabulated Data Comments

Gust spectra for agricultural or

aerial application usage. 

Figure A1-11 Table A1-11 

Maneuver spectra for 

agricultural or aerial 

application usage. 

Figure A1-12 Table A1-12 

Flight Spectra, Special Usage – Survey (Pipeline Patrol)

Description Graphic Data Tabulated Data Comments

Gust spectra for low-level 

survey or pipeline patrol

usage. 

Figure A1-13 Table A1-13 

Maneuver spectra for lowlevel survey or pipeline patrol

usage. 

Figure A1-14 Table A1-14 

Ground Spectra 

Description Graphic Data Tabulated Data Comments

Landing Impact Figure A1-15 

Taxi Figure A1-16

 

A1-2 
9/29/05 AC 23-13A 

Appendix 1 

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

A1-2. How do I compute the gust limit load factor used in the gust load spectra? 

a. The equation for computing the gust load limit load factor has evolved through the 

regulatory history of the Civil Air Regulation Part 3 (CAR 3) and 14 CFR Part 23. These 

changes to the gust load formula incorporate the improving knowledge of airplane response to 

atmospheric turbulence. The gust load spectra presented in this AC are normalized to the gust 

load formula in use before Amendment 23-7 to 14 CFR Part 23. Regardless of the certification 

basis of your airplane, you must calculate a gust load limit load factor (for use in developing 

fatigue loading spectra only) using the same equation used to derive the gust spectra. 

b. Formula for computing the incremental gust load limit load factor: 

 

S

W

U K V m

anLLF

498

=

where U = 30.0, nominal gust velocity in feet per second 

 K = 4

1

2

1 ⎟

⎠

⎞ ⎜

⎝

⎛

S

W for W/S < 16 lbs / ft

2 

 

=

4

3

2.67 1.33

⎟

⎠

⎞ ⎜

⎝

⎛

−

S

W

 for W/S > 16 lbs / ft

2 

 W / S = Wing loading at maximum weight, lbs / ft2 

 V = Airplane structural design cruise speed, VC, KEAS 

 m = Wing lift curve slope, CLα, rad-1

A1-3. How are gust and maneuver load spectra ‘normalized’? 

a. We normalized the gust spectra to the incremental gust limit load factor computed in 

Paragraph A1-2. We expressed the gust spectra in terms of the gust load factor ratio: 

IncrementalDesignLimitGustLoad FactoratMaximumGrossWeight

IncrementalGustLoad Factorat OperatingWeight

a

a

nLLF

n =

b. We normalized the maneuver spectra to the incremental maneuver limit load factor. 

We expressed the maneuver spectra in terms of the maneuver load factor ratio: 

IncrementalDesignLimitManeuverLoad FactoratMaximumGrossWeight

IncrementalManeuverLoad Factorat OperatingWeight

a

a

nLLF

n =

A1-3 
AC 23-13A 9/29/05 

Appendix 1

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

Acceleration Fraction, an / anLLF

Cum

ulative Frequency of Exceedanc

e per Nauti

cal 

Mile

1.0

3

2

1.5

5

4

6

8

10-6

10-5

10-4

10-3

10-2

10-1

100

-1.2 -1.0 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 1.0 1.2

Figure A1-1

Gust Spectra

For Single-Engine Unpressurized Operations

Including Basic Instruction, Personal, Executive, and Acrobatic Usage

See Table A1-1 for

Tabulated Spectra

A1-4 
9/29/05 AC 23-13A 

Appendix 1 

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

Table A1-1 

Gust Spectra 

For Single-Engine Unpressurized Operations 

Including Basic Instruction, Personal, Executive, and Acrobatic Usage

Acceleration 

 Fraction 

Cumulative Frequency of 

Exceedance per Nautical Mile

Acceleration

 Fraction 

Cumulative Frequency of 

Exceedance per Nautical Mile 

0.10 7.99040E-01 -0.10 8.86903E-01

0.15 2.39762E-01 -0.15 2.90400E-01

0.20 8.26537E-02 -0.20 7.93185E-02

0.25 2.59957E-02 -0.25 2.50615E-02

0.30 9.43467E-03 -0.30 8.64457E-03

0.35 3.74392E-03 -0.35 3.13716E-03

0.40 1.56672E-03 -0.40 1.17560E-03

0.45 6.76810E-04 -0.45 4.49619E-04

0.50 2.96910E-04 -0.50 1.74053E-04

0.55 1.31101E-04 -0.55 6.77344E-05

0.60 5.79043E-05 -0.60 2.63806E-05

0.65 2.55611E-05 -0.65 1.02760E-05

0.70 1.12764E-05 -0.70 4.00187E-06

0.75 4.97072E-06 -0.75 1.55782E-06

0.80 2.18708E-06 -0.80 6.05622E-07

0.85 9.58226E-07 -0.85 2.34685E-07

0.90 4.15769E-07 -0.90 9.01733E-08

0.95 1.76305E-07 -0.95 3.38733E-08

1.00 7.05919E-08 -1.00 1.19395E-08

A1-5 
AC 23-13A 9/29/05 

Appendix 1

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

Acceleration Fraction, an / anLLF

Cumul

ative Frequenc

y of Exce

e

dance per Nauti

c

al Mil

e

1.0

3

2

1.5

5

4

6

8

10-6

10-5

10-4

10-3

10-2

10-1

100

-1.2 -1.0 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 1.0 1.2

Figure A1-2

 Maneuver Spectra

For Single-Engine Basic Instruction Usage

See Table A1-2 for

Tabulated Spectra

A1-6 
9/29/05 AC 23-13A 

Appendix 1 

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

Table A1-2 

Maneuver Spectra 

For Single-Engine Basic Instruction Usage

Acceleration 

 Fraction 

Cumulative Frequency of 

Exceedance per Nautical Mile

Acceleration

 Fraction 

Cumulative Frequency of 

Exceedance per Nautical Mile 

0.10 2.37466E-01 -0.10 2.22202E-01

0.15 1.20023E-01 -0.15 7.79188E-02

0.20 6.20605E-02 -0.20 2.89635E-02

0.25 3.45611E-02 -0.25 1.18361E-02

0.30 1.99278E-02 -0.30 4.73564E-03

0.35 1.16586E-02 -0.35 1.81743E-03

0.40 6.84325E-03 -0.40 6.69379E-04

0.45 4.01154E-03 -0.45 2.39525E-04

0.50 2.34706E-03 -0.50 8.45918E-05

0.55 1.36820E-03 -0.55 2.98155E-05

0.60 7.93694E-04 -0.60 1.04723E-05

0.65 4.59287E-04 -0.65 3.66958E-06

0.70 2.65584E-04 -0.70 1.28565E-06

0.75 1.53197E-04 -0.75 4.50315E-07

0.80 8.79528E-05 -0.80 1.57610E-07

0.85 5.01008E-05 -0.85 5.50454E-08

0.90 2.81951E-05 -0.90 1.91064E-08

0.95 1.55016E-05 -0.95 6.51330E-09

1.00 8.15702E-06 -1.00 2.10063E-09

A1-7 
AC 23-13A 9/29/05 

Appendix 1

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

Acceleration Fraction, an / anLLF

Cumulative Frequenc

y of Exc

eedanc

e per Nauti

cal 

Mile

1.0

3

2

1.5

5

4

6

8

10-6

10-5

10-4

10-3

10-2

10-1

100

-1.2 -1.0 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 1.0 1.2

Figure A1-3

 Maneuver Spectra

For Single-Engine Personal Usage

See Table A1-3 for

Tabulated Spectra

A1-8 
9/29/05 AC 23-13A 

Appendix 1 

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

Table A1-3 

Maneuver Spectra 

For Single-Engine Personal Usage

Acceleration 

 Fraction 

Cumulative Frequency of 

Exceedance per Nautical Mile

Acceleration

 Fraction 

Cumulative Frequency of 

Exceedance per Nautical Mile 

0.10 1.28156E-01 -0.10 9.71117E-02

0.15 6.27614E-02 -0.15 3.13880E-02

0.20 3.09500E-02 -0.20 1.09533E-02

0.25 1.58120E-02 -0.25 4.71523E-03

0.30 8.12433E-03 -0.30 2.07230E-03

0.35 4.12127E-03 -0.35 8.91311E-04

0.40 2.04241E-03 -0.40 3.75403E-04

0.45 9.97803E-04 -0.45 1.55129E-04

0.50 4.83819E-04 -0.50 6.29010E-05

0.55 2.33291E-04 -0.55 2.50877E-05

0.60 1.12320E-04 -0.60 9.84679E-06

0.65 5.40555E-05 -0.65 3.80374E-06

0.70 2.59889E-05 -0.70 1.44615E-06

0.75 1.24766E-05 -0.75 5.41094E-07

0.80 5.97051E-06 -0.80 1.99191E-07

0.85 2.83754E-06 -0.85 7.20825E-08

0.90 1.32888E-06 -0.90 2.55778E-08

0.95 6.02384E-07 -0.95 8.83264E-09

1.00 2.52545E-07 -1.00 2.89847E-09

A1-9 
AC 23-13A 9/29/05 

Appendix 1

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

Acceleration Fraction, an / anLLF

Cumulative Frequenc

y of Exc

eedanc

e per Nauti

cal 

Mile

1.0

3

2

1.5

5

4

6

8

10-6

10-5

10-4

10-3

10-2

10-1

100

-1.2 -1.0 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 1.0 1.2

Figure A1-4

 Maneuver Spectra

For Single-Engine Executive Usage

See Table A1-4 for

Tabulated Spectra

A1-10 
9/29/05 AC 23-13A 

Appendix 1 

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

Table A1-4 

Maneuver Spectra 

For Single-Engine Executive Usage

Acceleration 

 Fraction 

Cumulative Frequency of 

Exceedance per Nautical Mile

Acceleration

 Fraction 

Cumulative Frequency of 

Exceedance per Nautical Mile 

0.10 5.10093E-02 -0.10 5.54495E-02

0.15 1.97508E-02 -0.15 1.88747E-02

0.20 7.38092E-03 -0.20 6.05428E-03

0.25 3.13527E-03 -0.25 2.32888E-03

0.30 1.40483E-03 -0.30 9.20494E-04

0.35 6.60426E-04 -0.35 3.59308E-04

0.40 3.12393E-04 -0.40 1.35529E-04

0.45 1.48798E-04 -0.45 4.99636E-05

0.50 7.13657E-05 -0.50 1.81837E-05

0.55 3.43582E-05 -0.55 6.53422E-06

0.60 1.65390E-05 -0.60 2.31853E-06

0.65 7.96002E-06 -0.65 8.12393E-07

0.70 3.82975E-06 -0.70 2.81096E-07

0.75 1.84127E-06 -0.75 9.60362E-08

0.80 8.83928E-07 -0.80 3.23840E-08

0.85 4.23025E-07 -0.85 1.07639E-08

0.90 2.01127E-07 -0.90 3.51182E-09

0.95 9.42952E-08 -0.95 1.10946E-09

1.00 4.28623E-08 -1.00 3.23514E-10

A1-11 
AC 23-13A 9/29/05 

Appendix 1

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

Acceleration Fraction, an / anLLF

Cumulative Frequency of 

E

xceed

ance per Flight Hour

1.0

3

2

1.5

5

4

6

8

10-1

100

101

102

103

104

105

-1.2 -1.0 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 1.0 1.2

Figure A1-5

 Maneuver Spectra

For Single-Engine Acrobatic Usage

See Table A1-5 for

Tabulated Spectra

A1-12 
9/29/05 AC 23-13A 

Appendix 1 

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

Table A1-5 

Maneuver Spectra 

For Single-Engine Acrobatic Usage

Acceleration 

 Fraction 

Cumulative Frequency of 

Exceedance per Flight Hour

Acceleration

 Fraction 

Cumulative Frequency of 

Exceedance per Flight Hour 

0.10 1.8078 x 10+4 -0.125 8.7670 x 10+3

0.20 1.4840 x 10+4 -0.250 5.5320 x 10+3

0.30 1.0334 x 10+4 -0.375 2.6830 x 10+3

0.40 6.9640 x 10+3 -0.500 1.1400 x 10+3

0.50 3.8520 x 10+3 -0.625 3.7300 x 10+2

0.60 1.7490 x 10+3 -0.750 2.0000 x 10+1

0.70 5.9100 x 10+2

0.80 1.3900 x 10+2

0.90 2.0000 x 10+1

1.00 1.0000 x 10+0

A1-13 
AC 23-13A 9/29/05 

Appendix 1

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

Acceleration Fraction, an / anLLF

Cum

ulative Frequency of Exceedanc

e per Nautical 

Mile

1.0

3

2

1.5

5

4

6

8

10-6

10-5

10-4

10-3

10-2

10-1

100

-1.2 -1.0 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 1.0 1.2

Figure A1-6

Gust Spectra

For Twin-Engine Unpressurized Operations

General and Instructional Usage

See Table A1-6 for

Tabulated Spectra

A1-14 
9/29/05 AC 23-13A 

Appendix 1 

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

Table A1-6 

Gust Spectra 

For Twin-Engine Unpressurized Operations 

General and Instructional Usage

Acceleration 

 Fraction 

Cumulative Frequency of 

Exceedance per Nautical Mile

Acceleration

 Fraction 

Cumulative Frequency of 

Exceedance per Nautical Mile 

0.10 4.44668E-01 -0.10 4.06116E-01

0.15 1.61492E-01 -0.15 1.34138E-01

0.20 7.99116E-02 -0.20 7.30813E-02

0.25 3.17301E-02 -0.25 2.83576E-02

0.30 1.30374E-02 -0.30 1.18229E-02

0.35 5.87427E-03 -0.35 5.35232E-03

0.40 2.76767E-03 -0.40 2.55578E-03

0.45 1.30141E-03 -0.45 1.26417E-03

0.50 8.28867E-04 -0.50 6.41583E-04

0.55 4.71092E-04 -0.55 3.29008E-04

0.60 2.58561E-04 -0.60 1.70215E-04

0.65 1.42661E-04 -0.65 8.72554E-05

0.70 7.91060E-05 -0.70 4.41745E-05

0.75 4.51160E-05 -0.75 2.30083E-05

0.80 2.52331E-05 -0.80 1.16222E-05

0.85 1.40821E-05 -0.85 6.12841E-06

0.90 7.83366E-06 -0.90 2.98732E-06

0.95 4.24939E-06 -0.95 1.50864E-06

1.00 2.18891E-06 -1.00 7.03079E-07

A1-15 
AC 23-13A 9/29/05 

Appendix 1

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

Acceleration Fraction, an / anLLF

Cum

ulativ

e Frequency of Exc

eedanc

e per Nautical 

Mile

1.0

3

2

1.5

5

4

6

8

10-6

10-5

10-4

10-3

10-2

10-1

100

-1.2 -1.0 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 1.0 1.2

Figure A1-7

 Maneuver Spectra

For Twin-Engine Basic Instruction Usage

See Table A1-7 for

Tabulated Spectra

A1-16 
9/29/05 AC 23-13A 

Appendix 1 

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

Table A1-7 

Maneuver Spectra 

For Twin-Engine Basic Instruction Usage

Acceleration 

 Fraction 

Cumulative Frequency of 

Exceedance per Nautical Mile

Acceleration

 Fraction 

Cumulative Frequency of 

Exceedance per Nautical Mile 

0.10 4.73568E-01 -0.10 4.78815E-01

0.15 1.90555E-01 -0.15 1.38794E-01

0.20 8.52330E-02 -0.20 4.55853E-02

0.25 4.27211E-02 -0.25 1.65114E-02

0.30 2.27318E-02 -0.30 5.66718E-03

0.35 1.24765E-02 -0.35 1.80926E-03

0.40 6.94308E-03 -0.40 5.30806E-04

0.45 3.87445E-03 -0.45 1.41855E-04

0.50 2.15159E-03 -0.50 3.45231E-05

0.55 1.18268E-03 -0.55 7.68064E-06

0.60 6.41120E-04 -0.60 1.56206E-06

0.65 3.41647E-04 -0.65 2.90578E-07

0.70 1.79389E-04 -0.70 4.94860E-08

0.75 9.35425E-05 -0.75 7.76661E-09

0.80 4.85944E-05 -0.80 1.17677E-09

0.85 2.50637E-05 -0.85 2.26425E-10

0.90 1.27451E-05 -0.90 1.01278E-10

0.95 6.29624E-06 -0.95 1.84240E-12

1.00 2.92019E-06 -1.00 1.89361E-13

A1-17 
AC 23-13A 9/29/05 

Appendix 1

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

Acceleration Fraction, an / anLLF

Cum

ulative Frequency of Exceedanc

e per Nautical 

Mile

1.0

3

2

1.5

5

4

6

8

10-6

10-5

10-4

10-3

10-2

10-1

100

-1.2 -1.0 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 1.0 1.2

Figure A1-8

 Maneuver Spectra

For Twin-Engine Unpressurized General Usage

See Table A1-8 for

Tabulated Spectra

A1-18 
9/29/05 AC 23-13A 

Appendix 1 

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

Table A1-8 

Maneuver Spectra 

For Twin-Engine Unpressurized General Usage

Acceleration 

 Fraction 

Cumulative Frequency of 

Exceedance per Nautical Mile

Acceleration

 Fraction 

Cumulative Frequency of 

Exceedance per Nautical Mile 

0.10 6.89490E-02 -0.10 5.91059E-02

0.15 2.45115E-02 -0.15 1.49523E-02

0.20 9.65055E-03 -0.20 3.47737E-03

0.25 3.95338E-03 -0.25 1.23851E-03

0.30 1.73800E-03 -0.30 4.62353E-04

0.35 7.95244E-04 -0.35 1.52520E-04

0.40 3.75183E-04 -0.40 4.60500E-05

0.45 1.81414E-04 -0.45 1.37289E-05

0.50 8.94887E-05 -0.50 4.06783E-06

0.55 4.53754E-05 -0.55 1.20075E-06

0.60 2.40712E-05 -0.60 3.54448E-07

0.65 1.37295E-05 -0.65 1.04628E-07

0.70 8.69829E-06 -0.70 3.08837E-08

0.75 6.19922E-06 -0.75 9.11504E-09

0.80 4.92938E-06 -0.80 2.68914E-09

0.85 6.64545E-07 -0.85 7.92276E-10

0.90 3.36409E-07 -0.90 2.32341E-10

0.95 1.69605E-07 -0.95 6.70536E-11

1.00 8.48118E-08 -1.00 1.82622E-11

A1-19 
AC 23-13A 9/29/05 

Appendix 1

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

Acceleration Fraction, an / anLLF

Cum

ulative Frequency of Exceedanc

e per Na

utic

al Mil

e

1.0

3

2

1.5

5

4

6

8

10-6

10-5

10-4

10-3

10-2

10-1

100

-1.2 -1.0 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 1.0 1.2

Figure A1-9

 Gust Spectra

For Single-Engine and Twin-Engine

Pressurized Usage

See Table A1-9 for

Tabulated Spectra

A1-20 
9/29/05 AC 23-13A 

Appendix 1 

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

Table A1-9 

Gust Spectra 

For Single-Engine and Twin-Engine Pressurized Usage

Acceleration 

 Fraction 

Cumulative Frequency of 

Exceedance per Nautical Mile

Acceleration

 Fraction 

Cumulative Frequency of 

Exceedance per Nautical Mile 

0.10 2.23784E-01 -0.10 2.54285E-01

0.15 8.49264E-02 -0.15 9.72598E-02

0.20 3.88721E-02 -0.20 3.64316E-02

0.25 1.40381E-02 -0.25 1.44418E-02

0.30 5.65101E-03 -0.30 6.36368E-03

0.35 2.42471E-03 -0.35 2.99910E-03

0.40 1.08187E-03 -0.40 1.47318E-03

0.45 4.95557E-04 -0.45 7.42192E-04

0.50 2.33345E-04 -0.50 3.79530E-04

0.55 1.12397E-04 -0.55 1.95259E-04

0.60 5.47654E-05 -0.60 1.00525E-04

0.65 2.68277E-05 -0.65 5.17244E-05

0.70 1.31310E-05 -0.70 2.65940E-05

0.75 6.41930E-06 -0.75 1.36445E-05

0.80 3.13035E-06 -0.80 6.97417E-06

0.85 1.51868E-06 -0.85 3.53834E-06

0.90 7.28911E-07 -0.90 1.76854E-06

0.95 3.41903E-07 -0.95 8.56921E-07

1.00 1.52259E-07 -1.00 3.87347E-07

A1-21 
AC 23-13A 9/29/05 

Appendix 1

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

Acceleration Fraction, an / anLLF

Cum

ulative Frequency of Exceedanc

e per Nauti

cal Mil

e

1.0

3

2

1.5

5

4

6

8

10-6

10-5

10-4

10-3

10-2

10-1

100

-1.2 -1.0 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 1.0 1.2

Figure A1-10

 Maneuver Spectra

For Single-Engine and Twin-Engine

Pressurized Usage

See Table A1-10 for

Tabulated Spectra

A1-22 
9/29/05 AC 23-13A 

Appendix 1 

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

Table A1-10 

Maneuver Spectra 

For Single-Engine and Twin-Engine Pressurized Usage

Acceleration 

 Fraction 

Cumulative Frequency of 

Exceedance per Nautical Mile

Acceleration

 Fraction 

Cumulative Frequency of 

Exceedance per Nautical Mile 

0.10 5.61695E-02 -0.10 3.78784E-02

0.15 2.02968E-02 -0.15 9.82270E-03

0.20 6.63342E-03 -0.20 1.36144E-03

0.25 2.36553E-03 -0.25 2.41970E-04

0.30 9.20697E-04 -0.30 5.03001E-05

0.35 3.74938E-04 -0.35 1.16053E-05

0.40 1.56632E-04 -0.40 2.79594E-06

0.45 6.60459E-05 -0.45 6.79568E-07

0.50 2.78816E-05 -0.50 1.71237E-07

0.55 1.17951E-05 -0.55 4.91413E-08

0.60 5.01701E-06 -0.60 1.98149E-08

0.65 2.15974E-06 -0.65 1.27710E-08

0.70 9.55665E-07 -0.70 2.69318E-09

0.75 4.48349E-07 -0.75 2.72578E-10

0.80 2.34443E-07 -0.80 1.74971E-10

0.85 1.44293E-07 -0.85 3.53231E-11

0.90 1.06299E-07 -0.90 1.78106E-12

0.95 1.09189E-08 -0.95 4.28533E-13

1.00 4.17046E-09 -1.00 1.03669E-13

A1-23 
AC 23-13A 9/29/05 

Appendix 1

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

Acceleration Fraction, an / anLLF

Cum

ulative Frequency of Exceedanc

e per Na

utic

al Mil

e

1.0

3

2

1.5

5

4

6

8

10-6

10-5

10-4

10-3

10-2

10-1

100

-1.2 -1.0 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 1.0 1.2

Figure A1-11

 Gust Spectra

For Agricultural Operations

See Table A1-11 for

Tabulated Spectra

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

X

A1-24 
9/29/05 AC 23-13A 

Appendix 1 

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

Table A1-11 

Gust Spectra 

For Agricultural Operations

Acceleration 

 Fraction 

Cumulative Frequency of 

Exceedance per Nautical Mile

Acceleration

 Fraction 

Cumulative Frequency of 

Exceedance per Nautical Mile 

0.10 3.77707E-01 -0.10 3.11916E-01

0.15 1.61542E-01 -0.15 1.31643E-01

0.20 7.22319E-02 -0.20 5.16164E-02

0.25 3.20268E-02 -0.25 2.15305E-02

0.30 1.34855E-02 -0.30 8.16495E-03

0.35 5.36510E-03 -0.35 3.21070E-03

0.40 2.13189E-03 -0.40 1.26787E-03

0.45 8.47618E-04 -0.45 5.01627E-04

0.50 3.36518E-04 -0.50 1.98343E-04

0.55 1.33203E-04 -0.55 7.78616E-05

0.60 5.22586E-05 -0.60 3.02250E-05

0.65 2.04295E-05 -0.65 1.15962E-05

0.70 7.94281E-06 -0.70 4.39690E-06

0.75 3.12936E-06 -0.75 1.64734E-06

0.80 1.23810E-06 -0.80 6.09684E-07

0.85 4.86588E-07 -0.85 2.22658E-07

0.90 1.89600E-07 -0.90 8.00213E-08

0.95 7.28972E-08 -0.95 2.80746E-08

1.00 2.72943E-08 -1.00 9.38019E-09

A1-25 
AC 23-13A 9/29/05 

Appendix 1

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

Acceleration Fraction, an / anLLF

Cum

ulative Frequency of Exceedanc

e per Na

utic

al Mil

e

1.0

3

2

1.5

5

4

6

8

10-6

10-5

10-4

10-3

10-2

10-1

100

-1.2 -1.0 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 1.0 1.2

Figure A1-12

 Maneuver Spectra

For Agricultural Operations

See Table A1-12 for

Tabulated Spectra

A1-26 
9/29/05 AC 23-13A 

Appendix 1 

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

Table A1-12 

Maneuver Spectra 

For Agricultural Operations

Acceleration 

 Fraction 

Cumulative Frequency of 

Exceedance per Nautical Mile

Acceleration

 Fraction 

Cumulative Frequency of 

Exceedance per Nautical Mile 

0.10 1.95680E+00 -0.10 1.06968E+00

0.15 1.62395E+00 -0.15 5.94801E-01

0.20 1.30823E+00 -0.20 3.13178E-01

0.25 9.97985E-01 -0.25 1.61990E-01

0.30 6.88328E-01 -0.30 7.11840E-02

0.35 4.44979E-01 -0.35 2.71675E-02

0.40 2.73441E-01 -0.40 8.78560E-03

0.45 1.61710E-01 -0.45 2.43350E-03

0.50 8.97378E-02 -0.50 6.28618E-04

0.55 4.80394E-02 -0.55 1.42839E-04

0.60 2.48058E-02 -0.60 2.96876E-05

0.65 1.23385E-02 -0.65 5.42758E-06

0.70 5.88998E-03 -0.70 9.62766E-07

0.75 2.76188E-03 -0.75 2.06299E-07

0.80 1.26089E-03 -0.80 9.21867E-08

0.85 5.56561E-04 -0.85 1.41585E-09

0.90 2.38497E-04 -0.90 1.11517E-09

0.95 9.89316E-05 -0.95 1.79203E-11

1.00 3.75956E-05 -1.00 1.43870E-12

A1-27 
AC 23-13A 9/29/05 

Appendix 1

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

Acceleration Fraction, an / anLLF

Cumul

ativ

e

Frequen

cy of Exceedance per Nauti

cal Mil

e

1.0

3

2

1.5

5

4

6

8

10-6

10-5

10-4

10-3

10-2

10-1

100

-1.2 -1.0 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 1.0 1.2

See Table A1-13 for

Tabulated Spectra

Figure A1-13

 Gust Spectra For Survey (Pipeline Patrol) Usage

A1-28 
9/29/05 AC 23-13A 

Appendix 1 

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

Table A1-13 

Gust Spectra 

For Survey (Pipeline Patrol) Operations

Acceleration 

 Fraction 

Cumulative Frequency of 

Exceedance per Nautical Mile

Acceleration

 Fraction 

Cumulative Frequency of 

Exceedance per Nautical Mile 

0.10 4.82042E+00 -0.10 4.81125E+00

0.15 2.07621E+00 -0.15 2.11771E+00

0.20 7.96900E-01 -0.20 8.25646E-01

0.25 2.46602E-01 -0.25 2.88959E-01

0.30 8.89211E-02 -0.30 1.17488E-01

0.35 3.52779E-02 -0.35 5.26061E-02

0.40 1.48846E-02 -0.40 2.51299E-02

0.45 6.53546E-03 -0.45 1.25490E-02

0.50 2.94288E-03 -0.50 6.46049E-03

0.55 1.34544E-03 -0.55 3.39526E-03

0.60 6.20257E-04 -0.60 1.80803E-03

0.65 2.87768E-04 -0.65 9.69846E-04

0.70 1.34389E-04 -0.70 5.21987E-04

0.75 6.29908E-05 -0.75 2.80579E-04

0.80 2.95677E-05 -0.80 1.50280E-04

0.85 1.38490E-05 -0.85 7.99172E-05

0.90 6.43049E-06 -0.90 4.18344E-05

0.95 2.92107E-06 -0.95 2.12119E-05

1.00 1.26226E-06 -1.00 1.00431E-05

A1-29 
AC 23-13A 9/29/05 

Appendix 1

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

Acceleration Fraction, an / anLLF

Cum

ulative Frequency of Exceedanc

e per Na

utic

al Mil

e

1.0

3

2

1.5

5

4

6

8

10-6

10-5

10-4

10-3

10-2

10-1

100

-1.2 -1.0 -0.8 -0.6 -0.4 -0.2 0 0.2 0.4 0.6 0.8 1.0 1.2

See Table A1-14 for

Tabulated Spectra

Figure A1-14

 Maneuver Spectra

For Survey (Pipeline Patrol) Usage

A1-30 
9/29/05 AC 23-13A 

Appendix 1 

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

Table A1-14 

Maneuver Spectra 

For Survey (Pipeline Patrol) Operations

Acceleration 

 Fraction 

Cumulative Frequency of 

Exceedance per Nautical Mile

Acceleration

 Fraction 

Cumulative Frequency of 

Exceedance per Nautical Mile 

0.10 7.42090E-01 -0.10 1.11391E-01

0.15 4.85640E-01 -0.15 4.71406E-02

0.20 2.81225E-01 -0.20 1.35410E-02

0.25 1.53241E-01 -0.25 5.45945E-03

0.30 7.81332E-02 -0.30 2.72563E-03

0.35 3.70863E-02 -0.35 1.41442E-03

0.40 1.63236E-02 -0.40 7.52533E-04

0.45 6.64508E-03 -0.45 4.06358E-04

0.50 2.49947E-03 -0.50 2.21451E-04

0.55 8.71435E-04 -0.55 1.20991E-04

0.60 2.89471E-04 -0.60 6.60759E-05

0.65 9.39729E-05 -0.65 3.60584E-05

0.70 3.01380E-05 -0.70 1.96504E-05

0.75 9.66349E-06 -0.75 1.06815E-05

0.80 3.09532E-06 -0.80 5.77900E-06

0.85 9.89409E-07 -0.85 3.09921E-06

0.90 3.13870E-07 -0.90 1.63439E-06

0.95 9.71706E-08 -0.95 8.33698E-07

1.00 2.76574E-08 -1.00 3.96027E-07

A1-31 
AC 23-13A 9/29/05 

Appendix 1

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

Sink Rate, ft / sec

C

umulative Frequency per 10,00

0 L

andings

1.0

3

2

1.5

5

4

6

8

100

101

102

103

104

105

106

0 2 4 6 8

Curve Description

 1 Single-Engine Basic Flight Instruction

 2 Single-Engine Unpressurized Operations,

Including Personal, Executive and

Aerobatic Usage

 3 Twin-Engine Unpressurized Usage,

Single-Engine and Twin-Engine

Pressurized Usage

Note: Curves reproduced from Figure 9,

Reference 3.

Figure A1-15

 Landing Impact Spectra

1 3 5 7 9

1

2 3

A1-32 
9/29/05 AC 23-13A 

Appendix 1

APPENDIX 1. FLIGHT AND GROUND LOAD SPECTRA (CONTINUED)

Normal Load Factor, nz

, g’s

Occ

urrences per 1,000 Lan

dings

1.0

3

2

1.5

5

4

6

8

100

101

102

103

104

105

106

1 .0 1 .2 1 .4 1 .6

Curve Description

 1 All Single-Engine and Twin-Engine

Operations, Excluding Aerial Application

 2 Agricultural Usage (Aerial Application)

Note: Curves reproduced from Figure 10,

Reference 3.

Figure A1-16

 Taxi Spectra

1 .1 1 .3 1 .5

1

2

+- +- +- +- +- +- +-

A1-33 (and 34) 

9/29/05 AC 23-13A 

Appendix 2

APPENDIX 2. S/N CURVES FOR ALUMINUM COMPONENTS 

i sk ~TLA S, ssert S gnit anr etl A

N ~ 

N

umber of Cycles

103 104 105 106 107

100

8060504030

8

2010

654321

FIGURE A2-1

ENDURANCE OF COMPLETE WINGS AND TAILPLANES

(ALUMINUM ALLOY MATERIALS)

Note: Curves reproduce

d from 

Figure 11, Ref

erence 3.

0

15

10

7.5

5

2.5

30

20

Mean Stress, SMEAN, ksi

A2- 1
AC 23-13A 9/29/05 

Appendix 2

APPENDIX 2. S/N CURVES FOR ALUMINUM COMPONENTS (CONTINUED)

i sk ~TLA S, ssert S gnit anr etl A

Mean Stress,

SMEAN, ksi

FIGURE A2-2

TYPICAL CONSTANT LIFE CURVES

FOR COMPLETE WINGS AND TAILPLANES

(ALUMINUM ALLOY MATERIALS)

Not

e: C

urv

e

s reproduce

d fro

m Figure 12, Reference 3.

0 20 15 10 5 25 30

0

201510

5

1 x 104

1 x 105

3 x 104

3 x 105

1 x 106

3 x 106

N

Number of 

Cycles

A2-2