Advisory Circular (AC) No. 500-009 Issue 1
Composite Aircraft Structure
|File No.||5009-6-500||AC No.||500-009|
|RDIMS No.||528317-V2||Issue No.||01|
|Issuing Branch||Aircraft Certification||Effective Date||2004-12-01|
- 1.0 Introduction
- 2.0 References
- 3.0 Background
- 4.0 Bonded Joints
- 5.0 Compliance Documentation - Primary Structures
- 6.0 Damage Tolerance
- 7.0 Levels of Damage
- 8.0 Secondary Structures
- 9.0 Failure Criteria
- 10.0 Quality Control
- 11.0 Repairs
- 12.0 Appendix - Proof Of Structures
- 13.0 Headquarters Contact
- Appendix A
The purpose of this Advisory Circular (AC) is to provide guidance to applicants who wish to submit compliance documentation for structures made from composite materials. As a supplement to Federal Aviation Administration Advisory Circular (FAA AC) 20-107A, related interpretative material is included as well as a suggested format for the necessary compliance documentation. Although the approach set forth in this document is an acceptable one, it is not to be considered the only acceptable way of showing compliance. The applicant may elect to follow an alternate method, which must be acceptable to Transport Canada.
1.2 Guidance Applicability
This document is applicable to all Transport Canada personnel, delegates and industry.
1.3 Description of Changes
This document, formerly AMA No. 500/8B, is reissued as an AC. With the exception of minor editorial changes and updated references, the content is unaltered.
This document does not have a terminating action. It will however, be reviewed periodically for suitability of content.
2.1 Reference Documents
It is intended that the following reference materials be used in conjunction with this document:
- Chapter 523 of the Airworthiness Manual (AWM) - Normal, Utility, Aerobatic and Commuter Category Aeroplanes;
- Chapter 525 of the AWM - Transport Category Aeroplanes;
- Chapter 527 of the AWM - Normal Category Rotorcraft;
- Chapter 529 of the AWM - Transport Category Rotorcraft;
- Federal Aviation Administration Advisory Circular (FAA AC) 20-107A - Composite Aircraft Structure;
- FAA AC 21-26 (incorporating Change 1) - Quality Control for the Manufacture of Composite Structures;
- FAA AC 23-3 - Structural Substantiation of Secondary Structures;
- U.S. Military Handbook (MIL-HDBK)-5E - Metallic Materials and Elements for Aerospace Vehicle Structures Handbook;
- MIL-HDBK-17/1D - Composite Materials Handbook Volume 1 - Polymer Matrix Composites Guidelines for Characterization of Structural Materials; and
- U.S. Department of Transportation, Federal Aviation Administration (FAA), Technical Center Report No. DOT/FAA/CT-85/6 - Fiber Composite Analysis and Design, Volume I.
2.2 Cancelled Document
As of the effective date of this document, AMA No. 500/8B dated 8 November 1999 is cancelled.
Since its publication, FAA AC 20-107A has become the standard reference document for applicants seeking guidance in the field of composite structures. This AC intends to provide further information that would assist applicants in interpreting the FAA document and suggests a specific format for the required compliance documentation. Also, guidance material is included which should be followed by applicants seeking to show compliance with subsections 523.605(a), 525.605(a), 527.605(a) or 529.605(a) of the AWM, for processes used to fabricate bonded joints. While the suggested approaches will provide most of the additional information needed, it is recommended that interested applicants consult the Aircraft Certification Branch of Transport Canada at an early stage. In this way agreement may be reached on the acceptability of the proposed method of showing compliance with the applicable airworthiness standards.
For some applicants, the lack of adequate published design data for composite materials represents a major obstacle to the successful demonstration of compliance with airworthiness standards. Nevertheless, the importance of the material and fabrication development aspects of a composites program cannot be overemphasized. Without comprehensive design data, including design allowables for all expected environmental conditions, it will not be possible to obtain approval for composite primary structure. Similarly, for point design features or damaged structure, conservative but statistically unquantified design values are normally developed.
Since the publication of FAA AC 20-107A, revised statistical methods have been published, setting out the required procedure for obtaining composite material allowables from test data. These methods are to be found in MIL-HDBK-17/1D, Vol. I, and supersede all previously used procedures. Specifically, the statistical analysis given in MIL-HDBK-5E, Chapter 9, is only to be used for metallic materials.
Experience with bonded joints has shown that some bonding processes will occasionally produce joints that appear to be cohesive but which actually have very little strength. This so-called "weak bond" phenomenon, which has defied attempts at detection by non-destructive inspection (NDI) methods, has caused the FAA to adopt a fail-safe approach to bonded joints. It has thus become common practice to use mechanical fasteners as a means of ensuring that, even if a weak bond is produced, a joint will have the ability to carry limit loads. Transport Canada's position has been that newly manufactured structures should always be capable of carrying ultimate loads. Also, for fabrication methods in general, another fundamental requirement is that they must produce consistently sound structures with the requirements of subsection 523.605(a) of the AWM.
Proof of composite structures is, as for metallic materials, divided into two sections: static strength and fatigue/damage tolerance. The means of compliance set out in FAA AC 20-107A, for static proof of strength and fatigue (damage tolerance), are reproduced in "flow-chart" form in Appendix A, Figures 1 and 2. It should be noted that, of the two possibilities shown for assessing the behaviour of damaged structures, the "no-growth" approach is the one that is normally chosen.
4.0 Bonded Joints
Only bonding processes that are well understood and that can be demonstrated to comply with subsections 523.605(a), 525.605(a), 527.605(a) or 529.605(a) of the AWM should be used in primary aircraft structures. Given a reliable process, the quality control procedures in place should be well documented in terms of their limitations. It should be known which possible deviations from the process specification may remain undetected and what would be the consequences of such deviations in terms of strength. Transport Canada does not believe that it is appropriate to rely on either mechanical fasteners or yet-to-be-developed NDI methods as a substitute for compliance with the requirements of subsections 523.605(a), 525.605(a), 527.605(a) or 529.605(a) of the AWM. Therefore, in showing compliance with these subsections the applicant should, in order to ensure the ultimate strength of each bonded joint critical to safe flight, determine:
- Which types of undetectable bonding process deviation are likely to cause a bonded joint to be understrength; and
- That the probability of any such undetectable deviation occurring and resulting in the joint being understrength is extremely remote.
For any bonded joint, the failure of which would result in catastrophic loss of the aircraft, the limit load capacity should be substantiated by one of the following methods:
- The maximum disbonds of each bonded joint should be consistent with the capability to withstand the loads in paragraph 6.(c) of this AC, determined by analysis, tests or both. Disbonds of each bonded joint greater than this should be prevented by design features; or
- Proof testing should be conducted on each production article that will apply the critical limit design load to each critical bonded joint; or
- Repeatable and reliable NDI techniques should be established that ensure the strength of each joint.
5.0 Compliance Documentation - Primary Structures
The following format would be acceptable for compliance documentation to be submitted for approval of designs using composite materials in primary aircraft structures:
Environmental Conditions Report
This report substantiates the climatic conditions that define the temperature and humidity design envelope. End-of-lifetime moisture content and maximum transient panel temperatures are documented.
Damage Scenario Report
This report should present a full description of the anticipated fabrication and in-service damage to the composite structures and the manner in which this damage will be accounted for in the structural test program.
Certification Test Plan
This report documents the way in which the test articles, damaged in accordance with the damage scenario, will be tested. Methods of accounting for environmental effects and material variability are described and substantiated.
The applicant may also choose to combine the Environmental Conditions and Damage Scenario Reports, and the Certification Test Plan, into a single "Certification/Approval Plan".
Certification/Approval Test Results
This is a summary of the results of the testing described in the Certification Test Plan.
Material and Design Allowables Report
This report serves primarily to satisfy the requirements of sections 523.613, 523.615, 523.619, 525.613, 525.615, 525.619, 527.613, 527.615, 527.619, 529.613, 529.615, and 529.619 of the AWM. In general, all material allowables must have been obtained from an acceptable material and fabrication development program.
Stress Analysis Report
The stress analysis report documents compliance with the requirements of sections 523.305, 523.307, 525.305, 525.307, 527.305, 527.307, 529.305, 529.307 of the AWM, and complements the certifications testing. Adequate margins of safety are demonstrated for static and damage-tolerance strengths under all critical loading conditions. Test data should be available to validate any analytical strength envelopes or failure criteria that are used.
6.0 Damage Tolerance
When applied, damage tolerance criteria are intended to provide a level of safety equivalent to that of conventional, metal structures. The strength of composite materials is considered more likely to be adversely affected by manufacturing defects, discrete-source impact damage and degradation due to environmental effects. Therefore, design standards and the associated means of substantiation should provide for these effects.
The effects of material variability and environmental conditions should be accounted for in the following evaluations:
- Demonstration by test, or by analysis supported by tests, that the structures is capable of carrying ultimate load with damage up to the threshold of detectability, considering the inspection methods employed.
- Demonstration by test, or by analysis supported by tests, of growth rate or no-growth rate of damage that may occur from fatigue, corrosion, manufacturing flaws or impact damage, under repeated loads expected in service.
- Demonstration by residual strength test, or by analysis supported by residual strength test demonstrating ability to withstand critical limit flight loads, considered as ultimate loads, with the extent of detectable damage consistent with the results of the damage tolerance evaluations. The damage growth, between initial detectability and the value selected for residual strength demonstration, factored to obtain inspection intervals, must allow development of an inspection program suitable for application by operation and maintenance personnel.
7.0 Levels of Damage
Typically three levels of damage are considered:
Damage (including manufacturing impact and defects and service impact) that is below the threshold of detectability for the chosen inspection procedure. This level is applicable to ultimate load static testing, which may include the effects of repeated-load cycling and environmental exposure.
Detectable minor damage that may occur during fabrication, assembly and in-service. This type of damage is applicable to damage-tolerance substantiation and would normally be introduced into the test article at the beginning of repeated-load testing.
A pre-determined level of impact damage that is visible on the outer surface. If it can be shown that this damage will not grow in service, it is acceptable to choose a level so that the damage may not be detected during any one routine inspection. Depending on the approach taken, it may be possible to combine damage levels 2 and 3 for the purpose of compliance analysis and compliance qualification tests.
8.0 Secondary Structures
For secondary structures it is only necessary to address the applicable static strength requirements. For small aeroplanes, guidance material on this subject is contained in FAA AC 23-3. Compliance with strength requirements may be shown by conservative analysis in most cases.
9.0 Failure Criteria
Theoretical failure criteria exist which, when applied to the analysis of composite laminates will generate "margins of safety". An inadequacy of some failure criteria is that the mode of failure is not predicted. Because of this, there is a degree of uncertainty associated with margins of safety calculated on the basis of failure indices, generated using such theoretical methods. In the case of metallic materials it has not been customary to derive margins of safety without knowledge of the failure mode. Static strength substantiation of composite materials, obtained by an analysis that cannot predict the mode of failure, would not be acceptable. The FAA publication entitled "Fiber Composite Analysis and Design: Composite, Volume I" (DOT/FAA/AR-85/6), section 2.3.5, contains a discussion of this subject from an airworthiness point of view. This text emphasises the importance of physically realistic failure criteria and the possibility that such criteria could be simpler than some purely analytical approaches.
10.0 Quality Control
The fabrication methods used to manufacture all aircraft must produce consistently sound structures. Composite structures, unlike metal, involve the simultaneous fabrication of both the material and the structures and therefore require a higher level of quality control. Guidance on this subject is contained in FAA AC 21-26, which should be consulted at an early stage by all applicants.
Normally, repair procedures, including damage size limitations, are developed as part of the substantiation program and are published in the Continued Airworthiness section of the Maintenance Manual.
12.0 Appendix - Proof Of Structures
The mean of compliance set out in FAA AC 20-107A, for static proof of strength are reproduced in "flow-chart" form in Figure 1. For fatigue (damage tolerance) the FAA AC 20-107A method is similarly reproduced in Figure 2.
Static strength substantiation begins with the introduction of "Level 1" damage and then proceeds along one of two test paths depending on whether or not the structure needs to be exposed to repeated load and environmental cycling. For damage tolerance substantiation a choice has to be made between the "growth" and "no-growth" approaches. The objective is to demonstrate the required level of residual strength with damage of a predetermined size. FAA AC 20-107A provides a more comprehensive discussion of these means of substantiation.
13.0 Headquarters Contact
For more information please contact:
Policy Standards Coordinator (AARDH/P)
Phone: (613) 990-3923
Facsimile: (613) 996-9178
Original signed by Maher Khouzam
Chief, Regulatory Standards
Aircraft Certification Branch
- Date modified: