AC 20-107A 4/25/84

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U.S. Department ADVISORY

of Transportation

Federal Aviation CIRCULAR

Administration

______________________________________________________________________________________________

Subject: COMPOSITE AIRCRAFT STRUCTURE Date: 4/25/84 AC No: 20-107A

Initiated by: AWS-103 Change:

__________________________________________

1. PURPOSE. This advisory circular sets forth an acceptable, but not the only, means of showing

compliance with the provisions of Federal Aviation Regulations (FAR), Parts 23, 25, 27, and 29

regarding airworthiness type certification requirements for composite aircraft structures, involving

fiber reinforced materials, e.g., carbon (graphite), boron, aramid (Kevlar), and glass reinforced

plastics. Guidance information is also presented on associated quality control and repair aspects.

2. CANCELLATION. AC 20-107, Composite Aircraft Structure dated July 10, 1978, is

canceled.

3. REGULATIONS AFFECTED. The material contained herein applies to normal, utility,

acrobatic, and transport category aircraft type certificated under Civil Aviation Regulations

(CARs) 3, 4b, 6, 7; and FARs 23, 25, 27, 29; and produced in compliance with FAR Part 21,

sections 21.125, or 21.143 as may be appropriate. The individual FARs applicable to each

paragraph are listed in Appendix 1 of this advisory circular.

4. GENERAL

a. The procedures outlined in this advisory circular provide guidance material for

composite structures and are considered acceptable to the FAA for showing compliance with

certification requirements of civil composite aircraft. This circular is published to aid in the

evaluation of certification programs for composite applications and reflects the current status of

composite technology. It is expected that this circular will be modified periodically to reflect

technology advances. The information contained herein is for guidance purposes and is not

mandatory nor regulatory in nature.

b. The extent of testing and/or analysis and the degree of environmental accountability

required will differ for each structure depending upon the expected service usage, the material

selected, the design margins, the failure criteria, the data base and experience with similar

structures, and on other factors affecting a particular structure. It is expected that these factors

will be considered when interpreting this advisory circular for use on a specific application.

c. Pertinent definitions are given in Appendix 2.
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5. MATERIAL AND FABRICATION DEVELOPMENT. To provide an adequate design data

base, environmental effects on the design properties of the material system should be established.

a. Environmental design criteria should be developed that identify the most critical

environmental exposures, including humidity and temperature, to which the material in the

application under evaluation may be exposed. This is not required where existing data demonstrate

that no significant environmental effects, including the effects of temperature and moisture, exist

for the material system and construction details, within the bounds of environmental exposure

being considered. Experimental evidence should be provided to demonstrate that the material

design values or allowables are attained with a high degree of confidence in the appropriate critical

environmental exposures to be expected in service. The effect of the service environment on static

strength, fatigue and stiffness properties should be determined for the material system through

tests; e.g., accelerated environmental tests, or from applicable service data. The effects of

environmental cycling (i.e., moisture and temperature) should be evaluated. Existing test data may

be used where it can be shown directly applicable to the material system.

b. The material system design values or allowables should be established on the laminate

level by either test of the laminate or by test of the lamina in conjunction with a test validated

analytical method.

c. For a specific structural configuration of an individual component (point design), design

values may be established which include the effects of appropriate design features (holes, joints,

etc.).

d. Impact damage is generally accommodated by limiting the design strain level.

6. PROOF OF STRUCTURE - STATIC. The static strength of the composite design should be

demonstrated through a program of component ultimate load tests in the appropriate environment,

unless experience with similar designs, material systems and loadings is available to demonstrate

the adequacy of the analysis supported by subcomponent tests, or limit load component tests.

a. The effects of repeated loading and environmental exposure which may result in

material property degradation should be addressed in the static strength evaluation. This can be

shown by analysis supported by test evidence, by tests at the coupon, element or subcomponent

level, or alternatively by relevant existing data.

b. Static strength structural substantiation tests should be conducted on new structure

unless the critical load conditions are associated with structure that has been subjected to a

repeated loading and environmental exposure. In this case either (1) the static test should be

conducted on structure with prior repeated loading and environmental exposure, or (2)

coupon/element/subcomponent test data should be provided to assess the possible degradation of

static strength after application of repeated loading and environmental exposure, and this

degradation accounted for in the static test or in the analysis of the results of the static test of the

new structure.
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c. The component static test may be performed in an ambient atmosphere if the effects of

the environment are reliably predicted by subcomponent and/or coupon tests and are accounted for

in the static test or in the analysis of the results of the static test.

d. The static test articles should be fabricated and assembled in accordance with

production specifications and processes so that the test articles are representative of production

structure.

e. When the material and processing variability of the composite structure is greater than

the variability of current metallic structures, the difference should be considered in the static

strength substantiation (1) by deriving proper allowables or design values for use in the analysis,

and the analysis of the results of supporting tests, or (2) by accounting for it in the static test when

static proof of structure is accomplished by component test.

f. Composite structures that have high static margins of safety (e.g., some rotorblades)

may be substantiated by analysis supported by subcomponent, element, and/or coupon testing.

g. It should be shown that impact damage that can be realistically expected from

manufacturing and service, but not more than the established threshold of detectability for the

selected inspection procedure, will not reduce the structural strength below ultimate load

capability. This can be shown by analysis supported by test evidence, or by tests at the coupon,

element or subcomponent level.

7. PROOF OF STRUCTURE - FATIGUE/DAMAGE TOLERANCE. The evaluation of

composite structure should be based on the applicable requirements of FAR 23.571, 23.572,

25.571, 27.571, and 29.571. The nature and extent of analysis or tests on complete structures

and/or portions of the primary structure will depend upon applicable previous fatigue/damage

tolerant designs, construction, tests, and service experience on similar structures. In the absence of

experience with similar designs, FAA-approved structural development tests of components,

subcomponents, and elements should be performed. The following considerations are unique to the

use of composite material systems and should be observed for the method of substantiation selected

by the applicant. When selecting the damage tolerance or safe life approach, attention should be

given to geometry, inspectability, good design practice, and the type of damage/degradation of the

structure under consideration.

a. Damage Tolerance (Fail-Safe) Evaluation.

(1) Structural details, elements, and subcomponents of critical structural areas

should be tested under repeated loads to define the sensitivity of the structure to damage growth.

This testing can form the basis for validating a no-growth approach to the damage tolerance

requirements. The testing should assess the effect of the environment on the flaw growth

characteristics and the no-growth validation. The environment used should be appropriate to the

expected service usage. The repeated loading should be representative of anticipated service usage.

The repeated load testing should include damage levels (including impact damage)

typical of those
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that may occur during fabrication, assembly, and in-service, consistent with the inspection

techniques employed. The damage tolerance test articles should be fabricated and assembled in

accordance with production specifications and processes so that the test articles are representative

of production structure.

(2) The extent of initially detectable damage should be established and be

consistent with the inspection techniques employed during manufacture and in service.

Flaw/damage growth data should be obtained by repeated load cycling of intrinsic flaws or

mechanically introduced damage. The number of cycles applied to validate a no-growth concept

should be statistically significant, and may be determined by load and/or life considerations. The

growth or no growth evaluation should be performed by analysis supported by test evidence or by

tests at the coupon, element, or subcomponent level.

(3) The extent of damage for residual strength assessments should be established.

Residual strength evaluation by component or subcomponent testing or by analysis supported by

test evidence should be performed considering that damage. The evaluation should demonstrate

that the residual strength of the structure is equal to or greater than the strength required for the

specified design loads (considered as ultimate). It should be shown that stiffness properties have

not changed beyond acceptable levels. For the no-growth concept residual strength testing should

be performed after repeated load cycling.

(4) An inspection program should be developed consisting of frequency, extent ,

and methods of inspection for inclusion in the maintenance plan. Inspection intervals should be

established such that the damage will be detected between the time it initially becomes detectable

and the time at which the extent of damage reaches the limits for required residual strength

capability. For the case of no-growth design concept, inspection intervals should be established as

part of the maintenance program. In selecting such intervals the residual strength level associated

with the assumed damages should be considered.

(5) The structure should be able to withstand static loads (considered as ultimate

loads) which are reasonably expected during a completion of the flight on which damage resulting

from obvious discrete sources occur (i.e., uncontained engine failures, etc.). The extent of damage

should be based on a rational assessment of service mission and potential damage relating to each

discrete source.

(6) The effects of temperature, humidity, and other environmental factors which

may result in material property degradation should be addressed in the damage tolerance

evaluation.

b. Fatigue (Safe-Life) Evaluation. Fatigue substantiation should be accomplished by

component fatigue tests or by analysis supported by test evidence, accounting for the effects of the

appropriate environment. The test articles should be fabricated and assembled in accordance with

production specifications and processes so that the test articles are representative of
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production structure. Sufficient component, subcomponent, element or coupon tests should be

performed to establish the fatigue scatter and the environmental effects. Component,

subcomponent, and/or element tests may be used to evaluate the fatigue response of structure with

impact damage levels typical of those that may occur during fabrication, assembly, and in service,

consistent with the inspection procedures employed. The component fatigue test may be performed

with an as-manufactured test article if the effects of impact damage are reliably predicted by

subcomponent and/or element tests and are accounted for in the fatigue test or in analysis of the

results of the fatigue test. It should be demonstrated during the fatigue tests that the stiffness

properties have not changed beyond acceptable levels. Replacement lives should be established

based on the test results. An appropriate inspection program should be provided.

8. PROOF OF STRUCTURE - FLUTTER. The effects of repeated loading and environmental

exposure on stiffness, mass and damping properties should be considered in the verification of

integrity against flutter and other aeroelastic mechanisms. These effects may be determined by

analysis supported by test evidence, or by tests at the coupon, element or subcomponent level.

9. ADDITIONAL CONSIDERATIONS.

a. Impact Dynamics. The present approach in airframe design is to assure that occupants

have every reasonable chance of escaping serious injury under realistic and survivable impact

conditions. Evaluation may be by test or by analysis supported by test evidence. Test evidence

includes but is not limited to element or subcomponent tests and service experience. Analytical

comparison to conventional structure may be used where shown to be applicable.

b. Flammability.

(1) The existing requirements for flammability and fire protection of aircraft

structure attempt to minimize the hazard to the occupants in the event ignition of flammable fluids

or vapors occur. In addition, components exposed to heat, flames or sparks should withstand these

effects. The use of composite structure should not decrease this existing level of safety.

Compliance may be shown by analysis supported by test evidence that aircraft interior material

subjected to these hazards can withstand fire and heat as required in FAR 25.

(2) Certain aircraft structure is required to be fire resistant. The following test is

considered acceptable for demonstrating compliance for aircraft exterior structure and engine

compartment materials that are to be fire resistant. A comparison test should be made between the

specimen and an aluminum alloy sheet of the thickness normally used for the intended installation.

The structure and materials should be tested by subjecting a specimen sheet 24 inches by 24 inches

positioned perpendicular to a 2000° F plus or minus 150° F flame produced by a modified oil

burner consuming two gallons of kerosene per hour. The burner should be positioned so that the

time required for the flame to penetrate the aluminum alloy sample would be approximately five

minutes. The test specimen should be positioned at the same distance from the burner flame as the

aluminum alloy sheet. The specimen will be considered satisfactory if it resists flame penetration

for a time period equal to or greater than the aluminum alloy sheet.
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c. Lightning Protection.

(1) Some composites are susceptible to lightning damage, and do not dissipate Pstatic electrical charges or provide electromagnetic shielding. Therefore it should be demonstrated

by analysis support by test evidence that the structure can dissipate P-static electrical charges,

provides electromagnetic protection where required and provides an acceptable means of diverting

the resulting electrical current (as a result of a lightning strike) so as not to endanger the aircraft.

(2) Consideration should be given possible deterioration and undetected damage

to the lightning protection system.

d. Protection of Structure. Weathering, abrasion, erosion, ultraviolet radiation, and

chemical environment (glycol, hydraulic fluid, fuel, cleaning agents, etc.) may cause deterioration

in a composite structure. Suitable protection against and/or consideration of degradation in

material properties should be provided for and demonstrated by test.

e. Quality Control. The overall plan required by the certifying agency should involve all

relevant disciplines, i.e., engineering, manufacturing and quality control. This quality control plan

should be responsive to special engineering requirements that arise in individual parts or areas as a

result of potential failure modes, damage tolerance and flaw growth requirements, loadings,

inspectability, and local sensitivities to manufacture and assembly.

f. Production Specifications. Specifications covering material, material processing, and

fabrication procedures should be developed to ensure a basis for fabricating reproducible and

reliable structure. The discrepancies permitted by the specifications should be substantiated by

analysis supported by test evidence, or tests at the coupon, element or subcomponent level.

g. Inspection and Maintenance. Maintenance manuals developed by manufacturers

should include appropriate inspection, maintenance and repair procedures for composite structures.

h. Substantiation of Repair. When repair procedures are provided in FAA approved

documents or the maintenance manual, it should be demonstrated by analysis and/or test that

methods and techniques of repair will restore the structure to an airworthy condition.

M.C. Beard

Director of Airworthiness
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APPENDIX 1. APPLICABLE FARs AND RELATED ADVISORY CIRCULARS

 Text Paragraphs FAR 23 FAR 25 FAR 27 FAR 29

1. PURPOSE Not Applicable

2. CANCELLATION Not Applicable

3. REGULATIONS AFFECTED Not Applicable

4. GENERAL Not Applicable

5. MATERIAL AND FABRICATION DEVELOPMENT

.603 .603 .603 .603

.613 .613 .613 .613

.615 .615

6. PROOF OF STRUCTURE - STATIC .305 .305 .305 .305

.307(a) .307(a) .307(a) .307(a)

7. PROOF OF STRUCTURE - FATIGUE/DAMAGE TOLERANCE

.571 .571 .571 .571

.572 AC 20-95 AC 20-95

8. PROOF OF STRUCTURE - FLUTTER .629 .629 .629 .629

9. ADDITIONAL CONSIDERATIONS

a. Impact Dynamics .561 .561 .561 .561

.601 .601

.785 .721 .601 .601

 .783(c)(g) .785 .783(d)

.787(e) .785 .787(c) .785

.807(b)(4) .787(a)(b) .801 .787(c)

.967(e) .789 .807(b)(4)

 .801 .965 .801

 .809 .803(c)(1)

 .963(d)
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APPENDIX 1. APPLICABLE FARs AND RELATED ADVISORY CIRCULARS

Text Paragraphs FAR 23 FAR 25 FAR 27 FAR 29

ADDITIONAL CONSIDERATIONS

a. Impact Dynamics (cont’d)

* .1413 .809

.963(b)

.967(f)

b. Flammability .609(a) .609(a) .609(a) .609(a)

.787(d) .853 .853 .853

.853 .855 .855 .855

.859 .859 .859 .859

.865 .863 .861 .861

.1121(c) .865 .1183 .863

.1182 .867 .1185 .903(c)

.1183 .903(c) .1191 .967(e)

 .1189(b)(2) .967(e) .1193(d)(e)

.1191 .1121(c) .1194 .1013(e)

 .1193(c)(d)(e) .1181 .1121(c)

.1182 .1183

.1183 .1185

.1185 .1189(a)(2)

.1189(a)(2) .1191

.1191 .1193(c)(d)(e)

 .1193(c)(d)(e) .1194

* Special Conditions have been issued in the past on wide body airplanes concerning emergency

 wheels up landing.
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APPENDIX 1. APPLICABLE FARs AND RELATED ADVISORY CIRCULARS

Text Paragraphs FAR 23 FAR 25 FAR 27 FAR 29

ADDITIONAL CONSIDERATIONS

c. Lightning Protection .609 .581 .609 .609

.867 .609

d. Protection of Structure .609 .609 .609 .609

e. Quality Control **   ** 

f. Production Specifications

.603 .603 .603 .603

.605 .605 .605 .605

** A new Advisory Circular on Quality Control for Composites is under development.
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APPENDIX 2. DEFINITIONS

Design values - material, structural element, and structural detail properties that have been 

determined from test data and chosen to assure a high degree of confidence in the integrity

of the completed structure [reference FAR 25.613(b)].

Allowables - material values that are determined from test data at the laminate or lamina level on 

a probability basis, e.g., A or B base values [reference FAR 25.615(a)].

Laminate level design values or allowables - established from multi-ply laminate test data and/or 

from test data at the lamina level and then established at the laminate level by test 

validated analytical methods.

Lamina level material properties - established from test data for a single ply or multi-ply singledirection oriented lamina layup.

Point design - an element or detail of a specific design which is not considered generically 

applicable to other structure for the purpose of substantiation, e.g., lugs and major joints. 

Such a design element or detail can be qualified by test or by a combination of test and 

analysis.

Environment - external, non-accidental conditions (excluding mechanical leading), separately or 

in combination, that can be expected in service and which may affect the structure (e.g., 

temperature, moisture, UV radiation, and fuel).

Degradation - the alteration of material properties (e.g., strength, modulus, coefficient of 

expansion) which may result from deviations in manufacturing or from repeated loading 

and/or environmental exposure.

Discrepancy - a manufacturing anomaly allowed and detected by the planned inspection 

procedure. They can be created by processing, fabrication or assembly procedures.

Flaw - a manufacturing anomaly created by processing, fabrication or assembly procedures.

Damage - a structural anomaly caused by manufacturing (processing, fabrication, assembly or 

handling) or service usage. Usually caused by trimming, fastener installation or foreign 

object contact.

Impact damage - a structural anomaly created by foreign object impact.

Coupon - a small test specimen (e.g., usually a flat laminate) for evaluation of basic lamina or 

laminate properties or properties of generic structural features (e.g., bonded or 

mechanically fastened joints).
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APPENDIX 2. DEFINITIONS

Element - a generic element of a more complex structural member (e.g., skin, stringers, shear 

panels, sandwich panels, joints, or splices).

Detail - a non-generic structural element of a more complex structural member (e.g. specific 

design configurated joints, splices, stringers, stringer runouts, or major access holes).

Subcomponent - a major three-dimensional structure which can provide complete structural 

representation of a section of the full structure (e.g. stub-box, section of a spar, wing 

panel, wing rib, body panel, or frames).

Component - a major section of the airframe structure (e.g., wing, body, fin, horizontal 

stabilizer) which can be tested as a complete unit to qualify the structure.