ASTM F3115/F3115M-23

This document is not an ASTM standard and is intended only to provide the user of an ASTM standard an indication of what changes have been made to the previous version. Because
it may not be technically possible to adequately depict all changes accurately, ASTM recommends that users consult prior editions as appropriate. In all cases only the current version
of the standard as published by ASTM is to be considered the official document.
Designation: F3115/F3115M − 22 F3115/F3115M − 23
Standard Specification for
Structural Durability for Small Aeroplanes
This standard is issued under the fixed designation F3115/F3115M; the number immediately following the designation indicates the year
of original adoption or, in the case of revision, the year of last revision. A number in parentheses indicates the year of last reapproval.
A superscript epsilon (´) indicates an editorial change since the last revision or reapproval.
1. Scope
1.1 This specification addresses the airworthiness requirements related to structural durability for the design of small aeroplanes.
The material was developed through open consensus of international experts in general aviation. This information was created by
focusing on Levels 1 through 4 Normal Category aeroplanes. The content may be more broadly applicable; it is the responsibility
of the applicant to substantiate broader applicability as a specific means of compliance.
1.2 An applicant intending to propose this information as Means of Compliance for a design approval must seek guidance from
their respective oversight authority (for example, published guidance from applicable Civil Aviation Authorities (CAAs), including
the guidance noted in Appendix X2, Guidance Material) concerning the acceptable use and application thereof. For information
on which oversight authorities have accepted this specification (in whole or in part) as an acceptable Means of Compliance to their
regulatory requirements (hereinafter referred to as “the Rules”), refer to ASTM Committee F44 webpage (www.astm.org/
COMMITTEE/F44.htm). Appendix X3Annex A1 maps the Means of Compliance of the ASTM standards to EASA CS-23,
amendment 5, or later, and FAA 14 CFR 23, amendment 64, or later, Structural Durability requirements of 23.2240.
1.3 Units—This document may present information in either SI units, English Engineering units, or both; the values stated in each
system are not necessarily exact equivalents; therefore, to ensure conformance with the standard, each system shall be used
independently of the other, and values from the two systems shall not be combined.
1.4 This standard does not purport to address all of the safety concerns, if any, associated with its use. It is the responsibility
of the user of this standard to establish appropriate safety, health, and environmental practices and determine the applicability of
regulatory limitations prior to use.
1.5 This international standard was developed in accordance with internationally recognized principles on standardization
established in the Decision on Principles for the Development of International Standards, Guides and Recommendations issued
by the World Trade Organization Technical Barriers to Trade (TBT) Committee.
2. Referenced Documents
2.1 ASTM Standards:
F3060 Terminology for Aircraft
F3066/F3066M Specification for Aircraft Powerplant Installation Hazard Mitigation
F3116/F3116M Specification for Design Loads and Conditions
F3174/F3174M Specification for Establishing Operating Limitations and Information for Aeroplanes
F3380 Practice for Structural Compliance of Very Light Aeroplanes
This specification is under the jurisdiction of ASTM Committee F44 on General Aviation Aircraft and is the direct responsibility of Subcommittee F44.30 on Structures.
Current edition approved Aug. 1, 2022Jan. 1, 2023. Published August 2022January 2023. Originally approved in 2015. Last previous edition approved in 20202022 as
F3115/F3115M-20.-22. DOI: 10.1520/F3115_F3115M-22.10.1520/F3115_F3115M-23.
For referenced ASTM standards, visit the ASTM website, www.astm.org, or contact ASTM Customer Service at service@astm.org. For Annual Book of ASTM Standards
volume information, refer to the standard’s Document Summary page on the ASTM website.
Copyright © ASTM International, 100 Barr Harbor Drive, PO Box C700, West Conshohocken, PA 19428-2959. United States
F3115/F3115M − 23
2.2 EASA Standard:
CS-23, Amendment 5 Certification Specifications for Normal Category Aeroplanes
2.3 FAA Standard:
14 CFR 23, Amendment 64 Airworthiness Standards: Normal Category Airplanes
3. Terminology
3.1 The following are a selection of relevant terms. See Terminology F3060 for more definitions and abbreviations.
3.2 Definitions:
3.2.1 catastrophic loss or catasrophic failure—failure condition that is expected to result in fatalities of the occupants, normally
with the loss of the aeroplane.
3.2.2 damage tolerance—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.
3.2.3 fail safe—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 structural element.
3.2.4 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 damage or complete fracture after a sufficient
number of fluctuations.
3.2.5 life (or load) enhancement factor—an additional load factor or test duration, or both, applied to structural repeated load tests,
relative to the intended design load and life values, used to account for material variability. It is used to develop the required level
of confidence in data.
3.2.6 residual strength—the strength capability of a structure after the structure has been damaged due to fatigue, corrosion,
accidental, or discrete source damage. The residual strength capability includes consideration of static strength, fracture, and
stiffness.
3.2.7 safe life—of a structure—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-induced damage.
3.2.8 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 and usage variability. A scatter factor can also be used in a
fatigue analysis to address the uncertainties inherent in a fatigue analysis.
3.2.9 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 given level of fatigue damage in a specimen at a given life (N), where N is expressed in the number of cyclic
applications of stress or strain.
4. Inspection Intervals for Continued Airworthiness
4.1 Evaluation for Aircraft Structure:
4.1.1 All aircraft structure must be designed to meet the following general durability criteria: avoid fatigue failure during normal
service.
4.1.2 Additionally, aircraft structure whose failure could result in catastrophic loss of the aeroplane within the operational life of
the aeroplane are to be evaluated using 4.2 through 4.7.
4.1.2.1 For metallic structure, this must include pressurized cabin, wing, empennage, and associated structure.
Available from European Aviation Safety Agency (EASA), Postfach 10 12 53, D-50452 Cologne, Germany, http://easa.europa.eu
Available from Federal Aviation Administration (FAA), 800 Independence Ave., SW, Washington, DC 20591, http://www.faa.gov.
F3115/F3115M − 23
4.1.2.2 For composite structure, this must include each wing (including canards, tandem wings, and winglets), empennage, their
carry-through and attaching structure, moveable control surfaces and their attaching structure, fuselage, and pressure cabin.
4.1.3 Evaluation of a particular structure is not required if it is shown that the structure, operating stress level, materials, and
expected uses are comparable, from a fatigue standpoint, to a similar design (including design details) that has had extensive
satisfactory service experience. For composite structure, additional similarity must be shown for the cure process, quality process,
and ply layup.
4.2 Load Considerations:
4.2.1 The evaluations (except for evaluations defined in 4.3.2.2, 4.3.3, and 4.4.2.5) must:
4.2.1.1 Include expected loading spectra (for example, taxi, ground-air-ground cycles, maneuver, gust, aerodynamic configuration
changes such as deploying flaps or slats);
4.2.1.2 Account for any significant effects due to the mutual influence of aerodynamic surfaces; and
4.2.1.3 Consider any significant effects from the following: propeller slipstream loading, vibration, and buffeting.
4.2.2 Severe usage limitations must be either identified and addressed in accordance with 4.7.2 or included in the expected loading
spectra. Examples of severe usage include pipeline and utility patrol, instruction, and short duration commuter and air taxi flights.
4.3 Metallic Structure Evaluations:
4.3.1 Metallic components must be evaluated using one of the following methods:
4.3.1.1 For Level 1, 2, or 3 aeroplanes, the methods described in 4.3.2 or 4.3.3 or 4.3.4.
4.3.1.2 For Level 4, the method described in 4.3.4.
4.3.1.3 If certification for operation above 12 500 m [41 000 ft] is requested, the damage tolerance evaluation of 4.3.4 must be
conducted for the fuselage pressure boundary.
4.3.1.4 For aeroplanes evaluated to the damage tolerance criteria of 4.3.4: if it is shown that damage-tolerance criteria is
impractical for a particular structure, the structure must be evaluated in accordance with the methods described in 4.3.2.
4.3.2 Fatigue Strength Evaluation:
4.3.2.1 The structure must be shown by test, or by analysis supported by test evidence, to be able to withstand the repeated loads
of variable magnitude (or equivalent) expected in service. Using this data, a safe life limit must be established for the structure,
defined by the cyclic mean fatigue life divided by appropriate scatter factors. Exceptions are noted in 4.3.2.2.
4.3.2.2 In lieu of a cyclic test, a safe life limit can be established for a complete airframe based on a strength evaluation of critical
structure for very light aeroplanes. This method and limitations are presented in Practice F3380.
4.3.3 Fail Safe Strength Evaluation:
4.3.3.1 The structure must be shown by analysis, test, or both, that catastrophic failure of the structure is not probable after failure
of a principal structural element. Modes of damage should be identified in the analysis to show the extent of damage evaluated.
4.3.3.2 The damage extent must be shown to be obvious, such that it would be detectable before the next flight, and is large enough
to be detected by obvious visual indications during walk around or by indirect means such as cabin pressure loss, cabin noise, or
fuel leakage.
4.3.3.3 Additional procedures must be used to prevent loss of fail-safe capability with damaged structural components until
damage is found. This would include aircraft maintenance manual instructions for detecting failures in fail-safe structure within
50 flights.
F3115/F3115M − 23
4.3.3.4 The evaluation must show that the remaining structure is able to withstand the residual strength loads in 4.5 multiplied by
a factor of 0.75, where the factor is applied to critical limit flight loads, not pressure loads or 1 g loads. In addition, a factor of
1.15 must be applied unless the dynamic effects of failure under static load are otherwise considered. For a pressurized cabin, the
normal operating pressures combined with the expected external aerodynamic pressures must be applied simultaneously with the
flight loading conditions.
4.3.4 Damage Tolerance Evaluation:
4.3.4.1 An evaluation of the strength, detail design, and fabrication must show that catastrophic failure due to fatigue, corrosion,
manufacturing defects, or accidental damage will be avoided throughout the operational life of the aeroplane. The evaluation must
be analysis supported by test evidence and, if available, service experience.
4.3.4.2 Damage at multiple sites due to fatigue must be included where the design is such that this type of damage can be expected
to occur.
4.3.4.3 Modes of damage should be identified in the analysis to show the extent of damage evaluated.
4.3.4.4 The extent of damage for residual strength evaluation at any time within the operational life of the aeroplane must be
consistent with the initial detectability and subsequent growth under repeated loads.
4.3.4.5 The residual strength evaluation must show that the remaining structure is able to withstand the residual strength loads in
4.5 with the extent of detectable damage consistent with the results of the damage tolerance evaluations.
4.4 Composite Structure Evaluations:
4.4.1 Composite components must be evaluated as follows:
4.4.1.1 For Level 1, 2 and 3 non-aerobatic aeroplanes, the methods described in 4.4.2 or 4.4.3.
4.4.1.2 For Level 4 or any aer
...