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9/8/09 AC 20-107B
(8) Representative structure must be included to gain valid test and analysis results.
Depending on aircraft loading (requiring investigation of various aircraft passenger and cargo configurations), structural dynamic considerations, and progressive failures, local strain rates and loading conditions may differ throughout the structure. Sensitivity of the structural behavior to reasonable impact orientation should also be considered for transport airplane and rotorcraft applications. This can be addressed by analysis supporte d by test evidence.
(9) Considering a need for comparative assessments with metal structure and a range of crash conditions, analysis with sufficient structural test evidence is often needed for transport and rotorcraft applications. Analysis requires extensive investigation of model sensitivity to modeling parameters (e.g., mesh optimization, representation of joints, element material input stress-strain data). Test also requires investigation of test equipment sensitivity appropriate to composites (e.g., filter frequencies with respect to expected pulse characteristics in the structure).
Model validation may be achieved using a building block approach, culminating in an adequately complex test (e.g., a drop test with sufficient structural details to properly evaluate the crashworthiness criteria).
b. Fire Protection, Flammability and Thermal Issues.
(1) Fire and exposure to temperatures that exceed maximum operating conditions require special considerations for composite airframe structure. (Refer to note below.) Requirements for flammability and fire protection of aircraft structure attempt to minimize the hazard to occupants in the event that flammable materials, fluids, or vapors ignite. The regulations associated with each aircraft product type (i.e., transport, small airplane, rotorcraft) should be used accordingly. Compliance may be shown by tests or analysis supported by test evidence. A composite design, including repair and alterations, should not decrease the existing level of safety relative to metallic structure. In addition, maintenance procedures should be available to evaluate the structural integrity of any composite aircraft structures exposed to fire and temperatures above the maximum operating conditions substantiated during design.
Note: Aircraft cabin interiors and baggage compartments have been areas of flammability concerns in protecting passenger safety.
This revision of the AC does not address composite materials used in aircraft interiors and baggage compartments.
Please consult other guidance material for acceptable means of compliance with flammability rules for interiors.
(2) Fire protection and flammability has traditionally been considered for engine mount structure, firewalls, and other powerplant structures that include composite elements. Additional issues critical to passenger safety have come with the expanded use of composites in wing and fuselage structures for transport airplanes. Existing regulations do not address the potential for the airframe structure itself to be flammable. Wing and fuselage applications should consider the effects of composite design and construction on the resulting passenger safety in the event of in-flight fires or emergency landing conditions, which combine with subsequent egress when a fuel-fed fire is possible.
(3) The results of fire protection and flammability testing with structural composite parts indicate dependence upon overall design and process details, as well as the origin of the fire and its extent. For example, the overa ll effects of composite fuselage structures exposed to fire may be significantly different when the fire originates within the cabin, where it can be controlled by limiting the structure’s contribution to spreading the fire, than when the fire occurs exterior to the fuselage after a crash landing, where fuel is likely to be the primary source for maintaining and spreading the fire. The threat in each case is different, and the approach to mitigation may also be different. In-flight fire safe ty addresses a fire originating within the airplane due to some fault, whereas post-crash fire safety addresses a fuel-fed pool fire external to the airplane.
Special conditions are anticipated for transport category airplanes with fuselage structure subjected to both in-flight and post-crash fire conditions. Transport wing structure will need to have special conditions for post-crash fire conditions.
(4) For an in-flight fire in transport category airplanes, it is critical that the fire not propagate or generate hazardous quantities of toxic by-products. In-flight fires have been catastrophic when they can grow in inaccessible areas. Composite fuselage structure could play a role different than traditional metal structure if the issue is not addressed.
(5) Metallic transport fuselage and wing structures have established a benchmark in fire protection that can be used to evaluate specific composite wing and fuselage structural details.
Exterior fire protection issues associated with composite transport structure must include the effects of an exterior pool fire following a survivable crash landing. Fuselage structure should provide sufficient time for passenger egress, without fire penetration or the release of gasses and/or materials that are either toxic to escaping passengers or could increase the fire severity.
Furthermore, these considerations must be extended to wing and fuel tank structure, which must also be prevented from collapse and release of fuel (including consideration of the influence of fuel load upon the structural behavior). For transport category airpla nes, the standards of § 25.856(b) provide the benchmark to establish the required level of safety.
(6) The exposure of composite structures to high temperatures needs to extend beyond the direct flammability and fire protection issues to other thermal issues. Many composite materials have glass transition temperatures, which mark the onset of reductions in strength and stiffness that are somewhat lower than the temperatures that can have a similar affect on equivalent metallic structure. The glass transition temperature of most composite materials is further reduced by moisture absorption. The reduced strength or stiffness of composites from high temperature exposures must be understood per the requirements of particular applications
(e.g., engine or other system failures). After a system failure and/or known fire, it may be difficult to detect the full extent of irreversible heat damage to an exposed composite structure.
As a result, composite structures exposed to high temperatures may require special inspections, tests, and analysis for proper disposition of heat damage. All appropriate damage threats and degradation mechanisms need to be identified and integrated into the damage tolerance and maintenance evaluation accordingly. Reliable inspections and test measurements of the extent of damage that exists in a part exposed to unknown levels of high temperatures should be documented. Particular attention should be given to defining the maximum damages that likely could remain undetected by the selected inspection procedures.
c. Lightning Protection. Lightning protection design features are needed for composite aircraft structures. Current carbon fiber composites are approximately 1,000 times less electrically conductive than standard aluminum materials, and composite resins and adhesives are traditionally non-conductive. Glass and aramid fiber composites are non-conductive. A lightning strike to composite structures can result in structural failure or large area damage, and it can induce high lightning current and voltage on metal hydraulic tubes, fuel system tubes, and electrical wiring if proper conductive lightning protection is not provided. Aircraft lightning protection design guidance can be found in the FAA Technical Report “Aircraft Lightning Protection Handbook” (DOT/FAA/CT-89/22). The lightning protection effectiveness for composite structures should be demonstrated by tests or analysis supported by tests. Such tests are typically performed on panels, coupons, subassemblies, or coupons representative of the aircraft structure, or tests on full aircraft. The lightning test waveforms and lightning attachment zones are defined in the SAE reports referenced in AC 20-155. Any structural damage observed in standard lightning tests should be limited to Category 1, 2 or 3, depending on the level of detection. This damage is characterized and integrated into damage tolerance analyses and tests as appropriate. Small simple airplanes certif ied under 14 CFR part 23 for VFR use only may be certified based on engineering assessment, according to AC 23-15. The effects of composite structural repairs and maintenance on the lightning protection system should be evaluated.
Repairs should be designed to maintain lightning protection.
(1) Lightning Protection for Structural Integrity.
(a) The composite structural design should incorporate the lightning protection when appropriate for the anticipated lightning attachment. The extent of lightning protection features depends on the lightning attachment zone designated for that area of the aircraft. Lightning protection features may include, but are not limited to, metal wires or mesh added to the outside surface of the composite structure where direct lightning attachment is expected.
(b) When lightning strikes an aircraft, very high currents flow through the airframe.
Proper electrical bonding must be incorporated between structural parts, which is most difficult for moveable parts (i.e., ailerons, rudders, and elevators). The electrical bonding features must be sized to conduct the lightning currents or they can vaporize, sending the high currents through unintended paths such as control cables, control rods, or hydraulic tubes. Guidance for certification of lightning protection of aircraft structures can be found in SAE Aerospace Recommended Practice (ARP) 5577, referenced in Transport Airplane Directorate Policy Statement ANM-111-05-004.
(2) Lightning Protection for Fuel Systems.
(a) Special consideration must be given to the fuel system lightning protection for an aircraft with integral fuel tanks in a composite structure. Composite structure with integral fuel systems must incorporate specific lightning protection features on the external composite surfaces, on joints, on fasteners, and for structural supports for fuel system plumbing and components to eliminate structural penetration, arcing, sparks or other ignition sources. AC 20-53 provides certification guidance for aircraft fuel system lightning protection.
(b) Transport airplane regulations for fuel system ignition prevention in § 25.981 require lightning protection that is failure tolerant. As a result, redundant and robust lightning protection for composite structure joints and fasteners in fuel tank structure is needed to ensure proper protection in preventing ignition sources.
(3) Lightning Protection for Electrical and Electronic Systems.
(a) Lightning strike protection of composite structures is needed to avoid inducing high lightning voltages and currents on the wiring for electrical and electronic systems whose upset or damage could affect safe aircraft operation. The consequences from a lightning strike of unprotected composite structures can be catastrophic for electrical and electronic systems that perform highly critical functions, such as fly-by-wire flight controls or engine controls.
(b) Electrical shields over system wiring and robust circuit design of electrical and electronic equipment both provide some protection against system upset or damage due to lightning. Since most composite materials provide poor shielding, at best, metal foil or mesh is typically added to the composite structure to provide additional shielding for wiring and equipment. Electrical bonding between composite structure parts and panels should be provided for the shielding to be effective. AC 20-136 provides certification guidance for aircraft electrical and electronic system lightning protection.
s/ David Hempe Manager, Aircraft Engineering Division Aircraft Certification Service Page 27
---------------------- ---------------------- ---------------------- ---------------------- ---------------------- ------------------------- ------------------------- ------------------------- ------------------------- ------------------------- 9/8/09 AC 20-107B Appendix 1. Applicable Regulations and Relevant Guidance
1. Applicable Regulations. A list of applicable regulations is provided for subjects covered in this AC. In most cases, these regulations apply regardless of the type of materials used in aircraft structures.
AC Paragraphs
1. Purpose of this AC
2. To Whom this AC Applies
3. Cancellation
4. Related Regulations and Guidance
5. General
6. Material and Fabrication Development
7. Proof of Structure – Static
8. Proof of Structure – Fatigue and
Damage Tolerance
9. Proof of Structure – Flutter
10. Continued Airworthiness
11. Additional Considerations
a. Crashworthiness Part 23 603 605 609 613 619 305 307 573
629 Part 25 Not Applicable
Not Applicable
Not Applicable
Not Applicable
Not Applicable
603 605 609 613 619 305 307 571
629 Part 27 Part 29 603 603
605 605
609 609 613 613
619 619
305 305
307 307
571 571 629 629
1529 1529 1529 1529 App. G App. H App. A App. A 561 561 561 561
562 562 562 562
601 601 601 601 631 631 721 721 783 783 783 783 785 785 785 785
787 787 787 787
807 789 801 801
965 801 807 803
967 809 965 809 963 963 967 967 981 1413 Page A1-1
b. Fire Protection, Flammability, 609 609 609 609 and Thermal Issues 787 863 861 861 863 865 863 863 865 867 1183 903 867 903 1185 967 954 967 1191 1013 1121 1121 1193 1121 1182 1181 1194 1183 1183 1182 1185 1189 1183 1189 1191 1185 1191 1193 1189 1193 1359 1191 1194 1365 1193
c. Lightning Protection 867 581 610 610 954 954 954 954 981 1309 1316 1309 1309 --------------------------------------- Notes:
(1) This list may not be all inclusive and there may be differences between regulatory authorities.
(2) Special conditions may be issued for novel and unusual design features (e.g., new composite materials systems).
2. Guidance. FAA issues guidance providing supportive information of showing compliance with regulatory requirements. Guidance may include the AC and policy statements (PS). In general, an AC presents information concerning acceptable means, but not the only means, of complying with regulations. The guidance listed below is deemed supportive to the purposes of this AC. These FAA documents can be located via website:
http://www.faa.gov/regulations_policies/.
a. ACs.
(1) AC 20-53, Protection of Airplane Fuel Systems Against Fuel Vapor Ignition Due to Lightning [6/06]
(2) AC 20-135, Powerplant Installation and Propulsion System Component Fire Protection Test Methods, Standards, and Criteria [2/90]
(3) AC 20-136, Protection of Aircraft Electrical/Electronic Systems Against the Indirect Effects of Lightning [12/06]
(4) AC 20-155, SAE Documents to Support Aircraft Lightning Protection Certification [4/06]
(5) AC 21-26, Quality Control for the Manufacture of Composite Structures [6/89]
(6) AC 21-31, Quality Control for the Manufacture of Non-Metallic Compartment Interior Components [11/91] Page A1-2
(7) AC 23-15, Small Airplane Certification Compliance Program [12/03]
(8) AC 23-20, Acceptance Guidance on Material Procurement and Process Specifications for Polymer Matrix Composite Systems [9/03]
(9) AC 25.571-1, Damage Tolerance and Fatigue Evaluation of Structure [4/98]
(10) AC 29 MG 8, Substantiation of Composite Rotorcraft Structure [4/06]
(11) AC 35.37-1, Guidance Material for Fatigue Limit Tests and Composite Blade Fatigue Substantiation [9/01]
(12) AC 145-6, Repair Stations for Composite and Bonded Aircraft Structure [11/96]
b. Policy Statements
(1) Static Strength Substantiation of Composite Airplane Structure [PS-ACE100-2001-006, December 2001]
(2) Final Policy for Flammability Testing per 14 CFR Part 23, Sections 23.853,
23.855 and 23.1359 [PS-ACE100-2001-002, January 2002]
(3) Material Qualification and Equivalency for Polymer Matrix Composite Material Systems [PS-ACE100-2002-006, September 2003]
(4) Bonded Joints and Structures - Technical Issues and Certification Considerations [PS-ACE100-2005-10038, September 2005]
(5) Policy Statement on Acceptance of SAE International Aerospace Recommended Practice 5577 as an Acceptable Method of Compliance to the Lightning Direct Effects requirements of § 25.581 [ANM-111-05-004, April 2006] Page A1-3
Appendix 2. Definitions
1. Allowables: Material values that are determined from test data at the laminate or lamina level on a probability basis (e.g., A or B basis values, with 99% probability and 95% confidence, or 90% probability and 95% confidence, respectively). The amount of data required to derive these values is governed by the statistical significance (or basis) needed.
2. Anisotropic: Not isotropic; having mechanical and/or physical properties which vary with direction relative to natural reference axes inherent in the material.
3. Arrested Growth Approach: A method that requires demonstration that the structure, with defined flaws present, is able to withstand appropriate repeated loads with flaw growth which is either mechanically arrested or terminated before becoming critical (residual static strength reduced to limit load). This is to be associated with appropriate inspection intervals and damage detectability.
4. Category of Damage: Five categories of damage have been defined based on residual strength capability, required load level, detectability, inspection interval, damage threat and whether (or not) the event creating damage is self evident.
5. 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.
6. 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).
7. Critical Structure: A load bearing structure/element whose integrity is essential in maintaining the overall flight safety of the aircraft. This definition was adopted for this AC because there are differences in the definitions of primary structure, secondary structure, and principle structural elements (PSE) when considering the different categories of aircraft. For example, PSE are critical structures for Transport Category Aircraft.
8. Damage: A structural anomaly caused by manufacturing (processing, fabrication, assembly or handling) or service usage.
9. Debond: Same as Disbond.
10. 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.
11. Delamination: The separation of the layers of material in a laminate. This may be local or may cover a large area of the laminate. It may occur at any time in the cure or subsequent life of the laminate and may arise from a wide variety of causes.
Page A2-1
12. Design Values: Material, structural elements, 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. These values are most often based on allowables adjusted to account for actual structural conditions, and used in analysis to compute margins-of-safety.
13. Detail: A non-generic structural element of a more complex structural member (e.g., specific design configured joints, splices, stringers, stringer runouts, or major access holes).
14. Disbond: An area within a bonded interface between two adherends in which an adhesion failure or separation has occurred. It may occur at any time during the life of the substructure and may arise from a wide variety of causes. Also, colloquially, an area of separation between two laminae in the finished laminate (in this case, the term “delamination” is normally preferred).
15. Discrepancy: A manufacturing anomaly allowed and detected by the planned inspection procedure. They can be created by proces sing, fabrication or assembly procedures.
16. Element: A generic element of a more complex structural member (e.g., skin, stringers, shear panels, sandwich panels, joints, or splices).
17. Environment: External, non-accidental conditions (excluding mechanical loading), separately or in combination, that can be expected in service and which may affect the structure
(e.g., temperature, moisture, UV radiation, and fuel).
18. Factor(s):
a. Life (or Load) Enhancement Factor: An additional load factor and/or test duration 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.
b. Life Scatter Factor: Same as Life/Load Enhancement Factor.
c. Overload Factor: A load factor applied to a specific structure test which is used to address parameters (e.g., environment, a short test pyramid, etc.) not directly addressed in that test. This factor is usually developed from lower pyramid testing addressing such parameters.
19. Heterogeneous: Descriptive term for a material consisting of dissimilar constituents separately identifiable; a medium consisting of regions of unlike properties separated by internal boundaries.
20. Impact Damage: A structural anomaly created by foreign object impact.
21. Intrinsic Flaw: Defect inherent in the composite material or resulting from the production process.
Page A2-2
22. Manufacturing Defect: An anomaly or flaw occurring during manufacturing that can cause varying levels of degradation in structural strength, stiffness and dimensional stability. Those manufacturing defects (or permissible manufacturing variability) allowed by the quality control, manufacturing acceptance criteria are expected to meet appropriate structural requirements for the life of the aircraft part. Other manufactur ing defects that escape detection in manufacturing quality control should be included in a damage threat assessment and must meet damage tolerance requirements until detected and repaired.
23. No-Growth Approach: A method that requires demonstration that the structure, with defined flaws present, is able to withstand appropriate repeated loads without detrimental flaw growth for the life of the structure.
24. Primary Structure: The structure which carries flight, ground, or pressurization loads, and whose failure would reduce the structural integrity of the airplane.
25. 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.
26. Slow Growth Approach: A method that requires demonstration that the structure, with defined flaws present, is able to withstand appropriate repeated loads with slow, stable, and predictable flaw growth for the life of the structure, or beyond appropriate inspection intervals associated with appropriate damage detectability.
27. Structural Bonding: A structural joint created by the process of adhesive bonding, comprising of one or more previously-cured composite or metal parts (referred to as adherends).
28. Subcomponent: A major three-dimensional structure which can provide completed structural representation of a section of the full structure (e.g., stub-box, section of a spar, wing panel, body panel with frames).
29. Weak Bond: A bond line with mechanical properties lower than expected, but without any possibility to detect that by normal NDI procedures. Such situation is mainly due to a poor chemical bonding.
Page A2-3
Appendix 3. Change of Compos ite Material and/or Process
1. It is necessary to re-certify composite structures, which during production, incorporate substitutions of, or changes to, the materials and/or processes from those originally substantiated at the time of initial certification. For example, the original material supplier may either change its product, or cease production. Manufacturers may also find it necessary to modify their production processes to improve efficiency or correct product deficiencies. In either case, care must be taken to ensure that modifications and/or changes are adequately investigated to ensure the continued adequacy of already certified composite structure. This appendix covers such material and/or process changes, but does not address other changes to design (e.g., geometry, loading). The definition of the materials and pr ocesses used is required in the specifications by 14 CFR 21.31. Changes to the material and proce ss specifications are often major changes in type design and must be addressed as such under 14 CFR part 21, subpart D.
2. The qualification and structural substantiation of new or modified materials and/or processes used to produce parts of a previously certified aircraft product requires:
a. The identification of the key material and/or process parameters governing performances;
b. The definition of the appropriate tests able to measure these parameters; and
c. The definition of pass/fail criteria for these tests.
3. “Qualification” procedures developed by every manufacturer include specifications covering:
a. Physical and chemical properties,
b. Mechanical properties (coupon level), and
c. Reproducibility (by test ing several batches).
4. Specifications and manufacturing quality procedures are designed to control specific materials and processes to achieve stable and repeatable structure for that combination of materials and processes. However, the interchangeability of alternate materials and processes for a structural application cannot be assumed if one only considers the properties outlined in those specifications (as it could be for materials that are much less process dependent, e.g., some metallic material forms). A structure fabricated using new or modified materials and/or processes, which meets the “qualification” tests required for the original material and process specifications, does not necessarily produce components that meet all the original engineering requirements for the previously certified structure.
5. Until improvements in identifying the complex relations between key material parameters that govern composite processing occurs, there will be a need for extensive and diverse testing that directly interrogates material performance using a range of representative specimens of increasing complexity in building block tests. Furthermore, failure modes may vary from one material and/or process to another, and analytical models are sometimes insufficiently precise to Page A3-1
reliably predict failure without sufficient empirical data. Therefore, a step-by-step test verification with more complex specimens may be required.
6. Classification of Material or Process Change.
a. Any of the following situations requires further investigation of possible changes to a given composite structure:
(1) Case A: A change in one or both of the basic constituents, resin, or fiber (including sizing or surface treatment alone) would yield an alternate material. Other changes that result in an alternate material include changes in fabric weave style, fiber aerial weight, and resin content.
(2) Case B: Same basic constituents, but any change of the resin impregnation method.
Such changes include: (i) prepregging process (e.g., solvent bath to hot melt coating), (ii) tow size (3k, 6k, 12k) for tape material forms with the same fiber areal weight, (iii) prepregging machine at the same suppliers, (iv) supplier change for a same material (licensed supplier).
(3) Case C: Same material, but modification of the processing route (if the modification to the processing route governs eventual composite mechanical properties). Example process changes of significance include: (i) curing cycle, (ii) bond surface preparation, (iii) changes in the resin transfer molding process used in fabricating parts from dry fiber forms, (iv) tooling, (v) lay-up method, (vi) environmental parameters of the material lay-up room, and (vii) major assembly procedures.
b. For each of the above cases, a distinction should be made between those changes intended to be a replica of the former material/process combination (Case B and some of Case C) and those which are “truly new material” (Case A and some of Case C). So, two classes are proposed:
(1) “Identical materials/processes” in cases intended to create a replica structure.
(2) “Alternative materials/processes” in cases intended to create truly new structure.
c. Within the “identical materials/processes” class, a subclassification can be made between a change of the prepregging machine alone at the supplier and licensed production elsewhere.
For the time being, a change to a new fiber produced under a licensed process and reputed to be a replica of the former one, will be dealt with as an “alternative material/process.”
d. Some minor changes within the class representing identical materials/processes may not interact with structural performances (e.g., prepreg release papers, some bagging materials, etc.)
and should not be submitted as part of the recertification. However, the manufacturers (or the supplier) should develop a proper system for screening those changes, with adequate proficiency at all relevant decision levels. Other minor material changes that fall under Case B may warrant sampling tests to show equivalency only at lower levels of building block substantiation.
e. Case C changes that may yield major changes in material and structural performance Page A3-2