AC 29-2C 12/01/2011 

1 

This new AC section is final and approved for inclusion in AC 29-2C. This AC 

section is effective and will be incorporated into the next published change or 

revision to AC 29-2. 

AC 29.573. § 29.573 (Amendment 29- 54) DAMAGE TOLERANCE AND FATIGUE

EVALUATION OF COMPOSITE ROTORCRAFT STRUCTURES 

 a. Purpose. This advisory material provides an acceptable means of compliance 

with the provisions of § 29.573, Amendment 29-54, Title 14 of the Code of Federal 

Regulations (CFR) dealing with the damage tolerance and fatigue evaluation of 

transport category composite rotorcraft structures. Paragraph f.(6) specifically 

addresses the advisory guidance applying to damage tolerance and fatigue evaluation 

as required by § 29.573, Amendment 29-54. Some information contained in AC 29-2C, 

MG 8 (Amendment 29-42) is repeated and updated, as appropriate, to preserve the 

“building block” approach for analyses of composite rotorcraft structure for compliance 

to § 29.573, Amendment 29-54 . (Supplemental guidance can be found in AC 20-107, 

“Composite Aircraft Structure.”) These procedures address the substantiation 

requirements for composite material system constituents, composite material systems, 

and composite structures common to rotorcraft. A uniform approach to composite 

structural substantiation is desirable, but it is recognized that in a continually developing 

technical area, which has diverse industrial roots both in aerospace and in other 

industries, variations and deviations from the procedures described here may be 

necessary. Deviations from this advisory material should be coordinated in advance 

with the Rotorcraft Directorate. 

 b. Special Considerations. Since rotorcraft structure is configured uniquely and is 

inherently subjected to severe cyclic stresses, special consideration is required for the 

substantiation of all rotorcraft structure, including composites. This special 

consideration is necessary to ensure that the level of safety intended by the current 

regulations are attained during the type certification process for all structure with special 

emphasis on composite structure because of its unique structural characteristics, 

manufacturing quality and operational considerations, and failure mechanisms. 

 c. Background. 

 (1) Historically, rotorcraft have required unique, conservative structural 

substantiation because of unique configuration effects, unique loading considerations, 

severe fatigue spectrum effects, and the specialized comprehensive fatigue testing 

required by these effects. Rotorcraft structural static strength substantiation for both 

metal and composite structure is essentially identical to that for fixed wing structure 

once basic loads have been determined. However, rotorcraft structural fatigue 

substantiation is significantly different from fixed wing fatigue substantiation. Since 

AC 20-107, as developed, applies to both fixed wing aircraft and rotorcraft, it, of 

necessity, was finalized in a broad generic form. Accordingly, a need to supplement 
AC 29-2C 12/01/2011 

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AC 20-107 for rotorcraft was recognized during type certification programs. One 

significant difference in traditional rotorcraft fatigue substantiation programs and fixed 

wing fatigue programs is the use of multiple component fatigue tests for rotorcraft 

programs rather than just one full-scale test. Also, constant amplitude, accelerated load 

tests are typically used rather than spectrum tests because of the high frequency loads 

common to rotorcraft operations. These rotorcraft fatigue tests have traditionally 

involved the generation of stress versus life or cycle (S-N) curves for each critical part 

(most of which are subjected to the cyclic loading of the main or tail rotor system) using 

a monotonic (sinusoidal) fatigue spectrum based on maximum and minimum service 

stress values. Unless configuration differences or flight usage data dictate otherwise, 

the monotonic fatigue spectrum’s period is typically based on six ground-air-ground 

(GAG) cycles for each flight hour of operation. The S-N curves for the substantiation of 

each detailed part are typically generated by plotting a curved line through three data 

points (see AC 29-2C, AC 29 MG 11, “Fatigue Tolerance Evaluation of Transport 

Category Rotorcraft Metallic Structure”). The three data points selected are a short 

specimen life (low-cycle fatigue), an intermediate specimen life and a long specimen life 

(high-cycle fatigue). Each raw data point is generated by monotonically fatigue testing 

at least two full-scale parts to failure or run out for each data point on the S-N curve. 

The raw data point values are then reduced by an acceptable statistical method to a 

single value for plotting to ensure proper reliability of the associated S-N curve. 

Order 8110.9, “Handbook on Vibration Substantiation and Fatigue Evaluation of 

Helicopter and Other Power Transmission Systems” and AC 27-1B, AC 27 MG 11, 

“Fatigue Evaluation of Rotorcraft Structure”, contain comprehensive discussions of the 

S-N curve generation process. The rotorcraft S-N curve process contrasts sharply with 

the fixed wing process of using a single full-scale fatigue article (usually an entire wing 

or airframe, which constitutes a single full-scale assembly data point), generic material 

or full-scale assembly S-N data (e.g., Metallic Materials Properties Development and 

Standardization (MMPDS), formerly the MIL-HDBK-5 for metals; Composites Materials 

Handbook-17 (CMH-17), formerly the MIL-HDBK-17 for composites; or AC 23-13, 

“Fatigue, Fail-Safe, and Damage Tolerance Evaluation of Metallic Structure for Normal, 

Utility, Acrobatic, and Commuter Airplanes”, which replaced AFS-120-73-2 for full-scale 

assemblies), a non-monotonic spectrum, and relatively large scatter factors to verify or 

determine the design fatigue life of the full-scale airplane. 

 (2) Also, rotorcraft have employed and mass-produced composite designs in 

primary structure (typically main and tail rotor blades) since the early 1950’s. This was 

10 or more years before composites were type certificated for primary fixed-wing 

structure in either military or civil aircraft applications (with some notable limited 

production exceptions, such as the Windecker fixed wing aircraft). In any case, the 

early 1950 period was well before a clear, detailed understanding of composite 

structural behavior (especially in the areas of macroscopic and microscopic failure 

mechanisms and modes) was relatively common and readily available in a usable 

format for the average engineer working in this field. It also predated the initial issuance 

of AC 20-107. Currently, much composite design information is proprietary, either to 

government, industry or both, and many data gathering methods have not been 

completely standardized. Consequently, a significant variation from laboratory to 
AC 29-2C 12/01/2011 

3 

laboratory in material property value determination methods and results can exist. The 

early rotor blade designs (as well as current designs) are by nature relatively low strain, 

tension structure designs. Also, by nature, these designs are not damage or flaw 

critical. Thus, by circumstance as much as design, early composite rotor blade and 

other composite rotorcraft designs incorporated an acceptable fatigue tolerance level of 

safety. In the 1980’s, more test data, analytical knowledge, and analytical methodology 

became available to more completely substantiate a composite design. Current 14 CFR 

parts 27 and 29 contain many sections to be considered in substantiating composite 

rotorcraft structure. This advisory material provides the current or updated information 

from AC 29-2C, MG 8, Amendment 29-42 to supplement the general guidance of 

AC 20-107 and provides compliance guidance for the requirements of § 29.573 

Amendment 29-54 for rotorcraft composite structure. 

 d. Definitions. The following basic definitions are provided as a convenient reading 

reference. CMH-17, and other sources, contain more complete glossaries of definitions. 

 (1) A-Basis Allowable. The “A” mechanical property value is the value above 

which at least 99 percent of the population of values is expected to fall, with a 

confidence level of 95 percent. 

 (2) Accidental Damage. Discrete damage, which may occur in service use or in 

manufacturing due to impacts or collisions, such as dents, scratches, gouges, 

abrasions, disbonds, splintering, and delaminations. 

 (3) Active Multiple Load Path. Structure providing two or more load paths that 

are all loaded during operation to a similar load spectrum. 

 (4) Allowables. Both A-basis and B-basis values statistically derived and used 

for a particular composite design. 

 (5) As-Manufactured. Product or component that has passed the applicable 

quality control process and has been found to conform to the approved design within 

the allowable tolerances. 

 (6) Autoclave. A closed apparatus usually equipped with variable conditions of 

vacuum, pressure, and temperature. It is used for bonding, compressing or curing 

materials. 

 (7) B-Basis Allowable. The “B” mechanical property value is the value above 

which at least 90 percent of the population of values is expected to fall, with a 

confidence level of 95 percent. 

 (8) Balanced Laminate. A composite laminate in which all laminae at angles 

other than 0° occur only in ± pairs (not necessarily adjacent). 
AC 29-2C 12/01/2011 

4 

 (9) Bond. The adhesion of one surface to another, with or without the use of an 

adhesive as a bonding agent. 

 (10) Catastrophic Failure. An event that could prevent continued safe flight and 

landing. 

 (11) Cocure. The process of curing several different materials in a single step. 

Examples include the curing of various compatible resin system pre-pregs, using the 

same cure cycle, to produce hybrid composite structure or the curing of compatible 

composite materials and structural adhesives, using the same cure cycle, to produce 

sandwich structure or skins with integrally molded fittings. 

 (12) Component. A major section of the airframe structure (e.g., wing, fin, body, 

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

 (13) 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). 

 (14) Cure. To change the properties of a thermosetting resin irreversibly by 

chemical reaction (i.e., condensation, ring closure, or addition). Cure may be 

accomplished by addition of curing (crosslinking) agents, with or without a catalyst, and 

with or without heat. 

 (15) Damage. A generic term for structural anomalies caused by manufacturing 

(processing, fabrication, assembly or handling) or service usage. Trimming, fastener 

installation, or foreign object impact are potential sources of damage, along with fatigue 

and environmental effects. 

 (16) 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. 

 (17) Damage Tolerant Fail-Safe. The capability of structure remaining after a 

partial failure to withstand design limit loads without catastrophic failure within an 

inspection period. 

 (18) Damage Tolerant Safe Life. Capability of structure with damage present to 

survive expected repeated loads of variable magnitude without detectable damage 

growth and to maintain ultimate load capability throughout service life of the rotorcraft. 

 (19) Delamination. The separation of the layers of material in a laminate. 

 (20) Design Limit Loads. The maximum loads to be expected in service, as 

defined by § 29.301(a). 
AC 29-2C 12/01/2011 

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 (21) 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). 

 (22) Disbond. A lack of proper adhesion in a bonded joint. This may be isolated 

or may cover a majority of the bond area. It may occur at any time in the cure or 

subsequent life of the bond area and may arise from a wide variety of causes. 

 (23) Element. A generic part of a more complex structural member (e.g., skin, 

stringers, shear panels, sandwich panels, joints, or splices). 

 (24) 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). 

 (25) Fatigue or Environmental Damage. Structural damage related to fatigue or 

environmental effects such as delaminations, disbonds, splintering, or cracking. 

 (25) Fiber. A single homogeneous strand of material, essentially 

one-dimensional in the macro-behavior sense, used as a principal constituent in 

advanced composites because of its high axial strength and modulus. 

 (26) Fiber Volume. The volume of fiber present in the composite. This is usually 

expressed as a percentage volume fraction or weight fraction of the composite. 

 (28) Fill. The 90° yarns in a fabric, also called the woof or weft. 

 (29) Glass Transition. The reversible change in an amorphous polymer or in 

amorphous regions of a partially crystalline polymer from (or to) a viscous or rubbery 

condition to (or from) a hard and relatively brittle one. 

 (30) Glass Transition Temperature. The approximate midpoint of the 

temperature range over which the glass transition takes place. 

 (31) Hybrid. Any mixture of fiber types (e.g., graphite and glass). 

 (32) Impregnate. An application of resin onto fibers or fabrics by several 

processes: hot melt, solution coat, or hand lay-up. 

 (33) Intrinsic or discrete manufacturing defects. Intrinsic or discrete 

imperfections or flaws related to manufacturing operations, processing or assembly, 

such as voids, gaps, porosity, inclusions, fiber dislocation, disbonds, and delaminations. 

 (34) Lamina. A single ply or layer in a laminate in which all fibers have the same 

fiber orientation. 
AC 29-2C 12/01/2011 

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 (35) Laminate. A product made by bonding together two or more layers or 

laminae of material or materials. 

 (36) Low Strain Level. As used herein, is defined as a principal, elastic axial 

gross strain level that for a given composite structure provides for no flaw growth and 

thus provides damage tolerance of the maximum defects allowed during the certification 

process using the approved design fatigue spectrum. 

 (37) Material System. The combination of single constituents chosen (e.g., fiber 

and resin). 

 (38) Material System Constituent. A single constituent (ingredient) chosen for a 

material system (e.g., a fiber, a resin). 

 (39) Matrix. The essentially homogeneous material in which the fibers or 

filaments of a composite are embedded in resins, which are mainly thermoset polymers 

in aircraft structure. 

 (40) Maximum Structural Temperature. The temperature of a part, panel or 

structural element due to service parameters such as incident heat fluxes, temperature, 

and air flow at the time of occurrence of any critical load case, (i.e., each critical load 

case has an associated maximum structural temperature). This term is synonymous 

with the term “maximum panel temperature.” 

 (41) Multiple Load Path. Structure providing two or more separate and distinct 

paths of structure that will carry limit load after complete failure of one of the members. 

 (42) Passive Multiple Load Path. Structure providing load paths with one or more 

of the members (or areas of a member) relatively unloaded until failure of the other 

member or members. 

 (43) 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. 

 (44) Porosity. A condition of trapped pockets of air, gas, or void within a solid 

material, usually expressed as a percentage of the total nonsolid volume to the total 

volume (solid + nonsolid) of a unit quantity of material. 

 (45) Pre-Preg, Preimpregnated. A combination of mat, fabric, nonwoven 

material, tape, or roving already impregnated with resin, usually partially cured, and 

ready for manufacturing use in a final product that will involve complete curing. Prepreg is usually drapable, tacky, and can be easily handled. 
AC 29-2C 12/01/2011 

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 (46) Principal Structural Element (PSE). A structural element that contributes 

significantly to the carrying of flight or ground loads and whose failure can lead to 

catastrophic failure of the rotorcraft. 

 (47) Residual Strength. The strength retained for some period of unrepaired use 

after a failure or partial failure due to fatigue, accidental, or discrete source of damage. 

 (48) Resin. An organic material with indefinite and usually high molecular weight 

and no sharp melting point. 

 (49) Resin Content. The amount of matrix present in a composite by either 

percent weight or percent volume. 

 (50) Secondary Bonding. The joining together, by the process of adhesive 

bonding, of two or more already-cured composite parts, during which the only chemical 

or thermal reaction occurring is the curing of the adhesive itself. The joining together of 

one already-cured composite part to an uncured composite part, through the curing of 

the resin of the uncured part, is also considered for the purposes of this advisory 

circular to be a secondary bonding operation. (See COCURE). 

 (51) Shelf Life. The lengths of time a material, substance, product, or reagent 

can be stored under specified environmental conditions and continue to meet all 

applicable specification requirements and remain suitable for its intended function. 

 (52) Strain Level. As used herein, is defined as the principal axial gross strain of 

a part or component due to the principal load or combinations of loads applied by a 

critical load case considered in the structural analysis (e.g., tension, bending, 

bending-tension). Strain level is generally measured in thousandths of an inch per unit 

inch of part or microinches/inch (e.g., .003 in/in equals 3000 microinches/inch). 

 (53) 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). 

 (54) Symmetrical Laminate. A composite laminate in which the ply orientation is 

symmetrical about the laminate midplane. 

 (55) Tape. Hot melt impregnated fibers forming unidirectional pre-preg. 

 (56) Thermoplastic. A plastic that repeatedly can be softened by heating and 

hardened by cooling through a temperature range characteristic of the plastic, and when 

in the softened stage, can be shaped by flow into articles by molding or extrusion. 

 (57) Thermoset (Or Chemset). A plastic that once set or molded cannot be 

re-set or remolded because it undergoes a chemical change; (i.e., it is substantially 

infusible and insoluble after having been cured by heat or other means). 
AC 29-2C 12/01/2011 

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 (58) Warp. Yarns extended along the length of the fabric (in the 0° direction) and 

being crossed by the fill yarns (90° fibers). 

 (59) Work Life. The period during which a compound, after mixing with a 

catalyst, solvent, or other compounding constituents, remains suitable for its intended 

use. 

 e. Related Regulatory and Guidance Material. 

Document Title

FAA Order 8110.9 Handbook on Vibration Substantiation and Fatigue 

Evaluation of Helicopter and other Power 

Transmission Systems 

AC 27-1B, MG 11 “Fatigue Evaluation of Rotorcraft Structure” 

AC 20-107 “Composite Aircraft Structure” 

AC 21-26 “Quality Control for the Manufacture of Composite 

Materials” 

CMH-17 “Composite Materials Handbook” 

AC 29-2C, MG 11 “Fatigue Tolerance Evaluation of Transport Category 

Rotorcraft Metallic Structure” 

DOT/FAA/CT-86/39 Whitehead, R.S., Kan, H.P., Cordero, R., and 

Seather, R., “Certification Testing Methodology for 

Composite Structures”, October 1986. 

 f. Procedures for Substantiation of Rotorcraft Composite Structure. The composite 

structures evaluation has been divided into eight basic regulatory areas to provide focus 

on relevant regulatory requirements. These eight areas are: fabrication requirements; 

basic constituent, pre-preg and laminate material acceptance requirements, and 

material property determination requirements; protection of structure; lightning 

protection; static strength evaluation; damage tolerance and fatigue evaluation; dynamic 

loading and response evaluation; and special repair and continued airworthiness 

requirements. 

Original as well as alternate or substitute material system constituents (e.g., fibers, 

resins), material systems (combinations of constituents and adhesives), and composite 

designs (e.g., laminates, cocured assemblies, bonded assemblies) should be qualified 

in accordance with the methodology presented in the following paragraphs. Each 

regulatory area will be addressed in turn. It is important to remember that proper 

certification of a composite structure is an incremental, building block process, which 

involves phased FAA/AUTHORITY involvement and incremental approval in each of the 

various areas outlined herein. It is recommended that a FAA/AUTHORITY certification 

team approach be used for composite structural substantiation. The team should 
AC 29-2C 12/01/2011 

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consist of FAA/AUTHORITY and cognizant members of the applicant’s organization. 

Personnel who are composites specialists (or are otherwise knowledgeable in the 

subject) should be primary team member candidates. 

Once selected, it is recommended that team meetings be held periodically (possibly in 

conjunction with type boards) during certification to ensure the building block 

certification process is accomplished as intended. The team should assure that 

permanent documentation in the form of reports or other FAA/AUTHORITY acceptable 

documents are included in the certification data package. The documentation includes 

but is not limited to the structural substantiation reports (both analysis and test), 

manufacturing processes and quality control, and Instructions for Continued 

Airworthiness (maintenance, overhaul, and repair manuals). The Airworthiness 

Limitations Section of the Instructions for Continued Airworthiness is approved by FAA 

engineering. Engineering practices for many of the areas identified below are available 

in CMH-17. 

 (1) The first area is the fabrication requirements of § 29.605: 

 (i) The quality control system should be developed considering the critical 

engineering, manufacturing, and quality requirements and a guidance standard such as 

AC 21-26, “Quality Control for the Manufacture of Composite Materials." This ensures 

that all special engineering, or manufacturing quality instructions for composites are 

presented, evaluated, documented, and approved, using drawings, process and 

manufacturing specifications, standards, or other equivalent means. This should be one 

of the early phases of a composite structure certification program, since this represents 

a major building block for sequential substantiation work. Some important concepts of 

AC 21-26 are included below. 

 (ii) Specific allowable defect limits (e.g., fiber waviness, warp defects, fill 

defects, porosity, hole edge effects, edge defects, resin content, large area disbonds, 

and delaminations) for a particular material system component, laminate design, 

detailed part, or assembly should be jointly established by engineering, manufacturing, 

and quality, and the associated inspection programs created, validated, and approved 

for defect detection. Each critical engineering design should consider the variability of 

the manufacturing process to determine the worse case effects (maximum waviness, 

disbonds, delaminations, and other critical defects) allowed by the reliability limitations 

of the approved inspection program. 

 (iii) If bonds or bond lines such as those typical of rotorcraft rotor blade 

structure are used, special inspection methods, special fabrication methods, or other 

approved verification methods (e.g., engineering proof tests - see paragraph f.(6)) 

should be provided to detect and limit disbonds or understrength bonds. 

 (iv) Structurally critical composite construction fabrication process and 

procurement specifications, for fabricating reproducible and reliable structure, must be 
AC 29-2C 12/01/2011 

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provided and FAA approved early during the certification process and should, as a 

minimum, cover the following: 

 (A) Vendor and Qualified Parts List (QPL) Control. Applicants should be 

able to demonstrate to FAA certification team members (both the manufacturing 

inspection district office (MIDO) and FAA engineering) at any time, that their quality 

control systems ensure on a continuous basis, that only qualified suppliers provide the 

basic material constituents or material systems (e.g., pre-pregs) that meet approved 

material specifications. Recommended guidelines for qualification of alternate material 

systems and suppliers are contained in CMH-17. These methods can also be used 

periodically for qualification status renewals of existing material systems and suppliers. 

 (B) Receiving Inspection and In-Process Inspection. Applicants should be 

able to demonstrate to FAA certification team members (both MIDO and engineering), 

at any time, that their receiving and in-process quality control systems provide products, 

which continuously meet approved material and process specifications. Quality 

systems should be designed with appropriate checks and balances, so that the 

necessary statistical reliability and confidence levels for the items being inspected (that 

are specified by engineering) are continuously maintained. This will require periodic 

standard inspections and engineering characterization tests on basic constituent and 

material system samples, which should be conducted, as a minimum, on a 

batch-to-batch basis. The periodic testing necessary to maintain the quality standard 

should be conducted by the applicants on conformed samples and should be FAA 

witnessed. 

 (C) Material System Component Storage and Handling. Applicants should 

be able to demonstrate to FAA certification team members (both MIDO and 

engineering), at any time, that their composite material system (or constituent) storage 

and handling procedures and specifications provide products, which continuously meet 

approved material and process specifications. Quality systems should be designed with 

appropriate checks and balances, so that the necessary statistical reliability and 

confidence levels for the items being inspected (that are specified by engineering) are 

continuously maintained. This should require, as a minimum, periodic inspections to 

ensure that proper records are kept on critical parameters (e.g., room temperature 

“bench” exposure, shelf life) and that periodic basic constituent and material system 

characterization tests are conducted, on a batch-to-batch basis. The periodic testing 

necessary to maintain the quality standard should be conducted by the applicants on 

conformed samples and should be FAA witnessed. 

 (D) Statistical Validation Level. It is necessary to maintain the minimum 

required statistical validation level of the quality control system, which should be 

specified for each critical item or constituent by the approved quality and engineering 

specifications. The statistical validation level should be defined and approved early in 

certification. Also, approval and proper usage should be continuously maintained 

during the entire procurement and manufacturing cycles. 
AC 29-2C 12/01/2011 

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 (v) Alternate fabrication and process specifications should be approved and 

must comply with § 29.605. Any alternate specifications should provide at least the 

same level of quality and safety as the original specification. Any changes should be 

presented for FAA approval well in advance of the effective date of the production 

change. 

 (2) The second area is the basic raw constituent, pre-preg and laminate material 

acceptance requirements, and material property determination requirements of 

§§ 29.603 and 29.613. These criteria require application of the critical environmental 

limits such as temperature, humidity, and exposure to aircraft fluids (such as fuel, oils, 

and hydraulic fluids), to determine their effect on the performance of each composite 

material system. Temperature and humidity effects are commonly considered by 

coupon and component tests utilizing preconditioned test specimens for each material 

system selected. Material “A” and “B” basis allowable strength values and other basic 

material properties (based on CMH-17 or equivalent procedures) are typically 

determined by small scale tests, such as coupon tests, for use in certification work. In 

the case of composites, determination of these basic constituent and material system 

properties will almost invariably involve the submittal, acceptance, and use of company 

standards. This is currently necessary because the FAA (new managers of CMH-17) 

has not completed development of “B” basis allowables for inclusion in CMH-17. Also, 

test methods vary somewhat from manufacturer to manufacturer; therefore, individual 

company results will exhibit some scatter in final material property values. Any 

company standard that is used should meet or exceed related CMH-17 requirements. 

Material structural acceptance criteria and property determination should, as a 

minimum, include the following: 

 (i) Property characterization requirements of all material systems (e.g., 

pre-pregs, adhesives) and constituents (e.g., fibers, resins) should be identified, 

documented, and approved. These requirements, once approved, should be placed in 

all appropriate procedures and specifications such as those in paragraph f.(1). 

 (ii) Moisture conditioning of test coupons, parts, subassemblies, or 

assemblies should be accomplished in accordance with CMH-17, other similar 

approved methods or per FAA approved programs. 

 (iii) The maximum and minimum temperatures expected in service (as derived 

from test measurements, thermal analyses on panels and other parts, experience, or a 

combination) should be determined and accounted for in static and fatigue strength 

(including damage tolerance) substantiation programs considering associated humidityinduced effects. 

 (iv) The wet glass transition temperature, Tg, is an important characteristic 

parameter of amorphous polymers, such as epoxies. It is the temperature below which 

the polymer behaves like a “glassy” solid and above which it behaves like a “rubbery” 

solid (i.e., it is the temperature at which there is a very rapid change in physical 

properties). The change from a hard polymeric material to a rubbery material takes 
AC 29-2C 12/01/2011 

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place over a narrow temperature range. A composite material will experience a drastic 

reduction in matrix-controlled mechanical material properties when loaded in this 

temperature range. Since the resin is the critical structural element in a composite 

matrix and the Tg is critical to structural integrity, a Tg determination is necessary. The 

Tg margin methodology of CMH-17 should be implemented (i.e., the Tg should be 50° F 

higher than the maximum structural temperature (see definition)). For any type of resin 

or adhesive, an acceptable temperature margin using CMH-17 techniques (e.g., 

consideration of limited high temperature excursions) or equivalent methodologies 

based on tests or experience, or both, should be established and approved early in the 

certification process. 

 (v) Local design values should be established by analysis and 

characterization tests and approved for specific structural configurations (point designs), 

which include the effects of stress risers (e.g., holes, notches) and structural 

discontinuities (e.g., joints, splices). Proper determination of these values for full-scale 

design and test should be considered one of the most critical building blocks in 

substantiating and evaluating a composite structure. These transitional load transfer 

areas typically produce the highest stresses (and strains) and serve as the initiation 

sites for many of the failures (including those due to the relatively low interlaminar 

strength of composites) that occur in service in a full-scale part or assembly. Small 

scale tests (such as coupon, element, and subcomponent tests), or equivalent approved 

testing programs, and analytical techniques should be carefully designed, prepared, and 

approved to evaluate potential “hot spots” and provide accurate simulations and 

representations of full-scale article stresses and strains in the critical transition areas. 

Proper certification work in this area will ensure initial safety and continued 

airworthiness in full-scale production articles. 

 (vi) The design strain level for each major component and material system 

should be established so that specified impact damage considerations are defined and 

properly limited. The effects of the strain levels may be established for each composite 

material using small-scale characterization tests and the results should be used to 

establish or verify the maximum allowable design strain level for each full-scale article. 

The maximum allowable design strain values selected should also take into account the 

reliability and confidence levels established for the relevant portions of the quality 

control system. This methodology is necessary because the amount and size of flaws 

in the production article may restrict the allowable level of design strain. In a 

no-flaw-growth design, the maximum specified impact damage and manufacturing flaw 

size at the most critical location on the part will be a major factor in determining the 

maximum allowable elastic strain. This design approach is currently selected for nearly 

all civil and most military applications; since, under normal conditions, only visual 

inspections are required in the field (unless unusual external damage circumstances 

such as a hail storm occur) to maintain the initial level of airworthiness (safety). 

However, many military applications, because of their demanding missions, employ 

scheduled field non-destructive inspection (NDI) maintenance, (such as comparative 

ultrasonics) to ensure that flaw growth either does not occur, is controlled by approved 

structural repair, or by replacement of affected parts. To date, civil applicants have not 
AC 29-2C 12/01/2011 

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requested a flaw growth, phased NDI approach. Therefore, selection of the full-scale 

article’s design strain limit based on small-scale tests for a no flaw growth design is 

extremely important. 

 (vii) Composite and adhesive properties should be determined so that 

detrimental structural creep does not occur under the sustained loads and environments 

expected in service. Small-scale characterization tests (such as coupon, element, and 

subcomponent tests) and analysis, which verify and establish the full-scale design 

criteria and parameters necessary to ensure that detrimental structural creep in 

full-scale structure does not occur in service, should be conducted early in certification 

and should be FAA-approved. 

 (viii) Material allowable strength values for full-scale design and testing should 

be developed using the coupon procedures presented in CMH-17 or equivalent. The 

intent is to represent the material variability including the effects that can occur in 

multiple batches of material and process runs. At least three batches of material 

samples should be used in material allowable strength testing. Company standards 

should be prepared, evaluated and FAA-approved early in certification (as part of the 

building block process), that reflect the material property determination considerations 

recommended in CMH-17 on an equal to or better than basis. 

 (3) The third area is the protection of structure as required by § 29.609. 

Protection against thermal, humidity, and other environmental effects (e.g., weathering, 

abrasion, fretting, hail, ultraviolet radiation, chemical effects, accidental damage) should 

be provided, or the structural substantiation should consider the results of those effects 

for which total protection is impractical. Determination and approval of worst-case or 

most conservative operating limits, and damage scenarios should be accomplished. 

Appropriate flammability and fire-resistance requirements should also be considered in 

selecting and protecting composite structure. Usually, a threat analysis is conducted 

early in the certification process that identifies the various threats and threat levels for 

which protection must be provided. This data is then used to construct and submit for 

approval the methods-of-compliance necessary to provide proper structural protection. 

 (4) The fourth area is the lightning protection requirements of § 29.610. 

Protection should be provided and substantiated in accordance with analysis and with 

tests such as those of AC 20-53, “Protection of Aircraft Fuel Systems Against Fuel 

Vapor Ignition Caused by Lightning” and FAA Report DOT/FAA/CT-86/8. For 

composite structure projects involving rotorcraft certificated to earlier certification bases 

(which do not automatically include the lightning protection requirements of § 29.610), 

these requirements should be imposed as special conditions. The design should be 

reviewed early in the certification process to ensure proper protection is present. The 

substantiation test program should also be established, reviewed and approved early to 

ensure proper substantiation. 

 (5) The fifth area is the static strength evaluation requirements of §§ 29.305 and 

29.307 for composite structure. Structural static strength substantiation of a composite 
AC 29-2C 12/01/2011 

14 

design should consider all critical load cases and associated failure modes, including 

effects of environment, material and process variability, and defects or service damage 

that are not detectable or allowed by the quality control, manufacturing acceptance 

criteria, or maintenance documents of the end product. The static strength 

demonstration should include a program of component ultimate load tests, unless 

experience exists to demonstrate the adequacy of the analysis, supported by 

subcomponent tests or component tests to accepted lower load levels. The necessary 

experience to validate an analysis should include previous component ultimate load 

tests with similar designs, material systems, and load cases.

 (i) The effects of repeated loading and environmental exposure, both of 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 levels, or alternatively by existing data. Earlier 

discussions in this AC address the effects of environment on material properties (see 

paragraph f.(2)) and protection of structure (see paragraph f.(3)). Static strength tests 

should be conducted for substantiation of new structure. For the critical loading 

conditions, two approaches to account for prior repeated loading or environmental 

exposure for structural substantiation exist. 

 In the first approach, the large-scale static test should be conducted on 

structure with prior repeated loading and conditioned to simulate the 

environmental exposure and then tested in that environment. 

 The second approach relies upon coupon, element, and sub-component test 

data to assess the possible degradation of static strength after application of 

repeated loading and environmental exposure. The degradation 

characterized by these tests should then be accounted for in the static 

strength demonstration test (e.g., load enhancement), or in the analysis of 

these results (e.g., showing a positive margin of safety with allowables that 

include the degrading effects of environment and repeated load). 

In practice, the two approaches may be combined to get the desired result (e.g., a 

large-scale static test may be performed at a temperature with a load enhancement 

factor to account for moisture absorbed over the aircraft structure’s life). 

 (ii) The strength of the composite structure should be statistically established, 

incrementally, through a program of analysis and tests at the coupon, element, 

subcomponent, or component levels. As part of the evaluation, building block tests and 

analyses at the coupon, element, or subcomponent levels can be used to address the 

issues of variability, environment, structural discontinuity (e.g., joints, cut-outs or other 

stress risers), damage, manufacturing defects, and design or process-specific details. 

Figure AC 29.573-1 provides a conceptual schematic of tests included in the building 

block approach. The material stress-strain curve should be clearly established, at least 

through the ultimate design load, for each composite design. As shown in Figure AC 

29.573-1, the large quantity of tests needed to provide a statistical basis comes from the 
AC 29-2C 12/01/2011 

15 

lowest levels (coupons and elements) and the performance of structural details are 

validated in a lesser number of sub-component and component tests. The static 

strength substantiation program should also consider all critical loading conditions for all 

critical structure including residual strength and stiffness requirements after a 

predetermined length of service (e.g., end of life (EOL)), which takes into account 

damage and other degradation due to the service period. 
AC 29-2C 12/01/2011 

16 

Figure AC 29.573-1. Schematic diagram of building block tests.
AC 29-2C 12/01/2011 

17 

 (iii) Allowables should be evaluated and used as specified in § 29.613. These 

allowables may be generated at the lamina, laminate, or specific design feature level 

(e.g., filled hole, lap joint, stringer run-out), provided they accurately reflect the actual 

value and variability of the structural strength for the critical failure modes being 

considered, at each point design where margins need to be established.

 (iv) The static test articles should be fabricated and assembled in accordance 

with production specifications and processes so that they are representative of 

production structure including defects consistent with the limits established by 

manufacturing acceptance criteria. 

 (v) The material and processing variability of the composite structure should 

be considered in the static strength substantiation. This can be achieved by 

establishing sufficient process and quality controls to manufacture structure and reliably 

substantiate the required strength in tests and analyses, which support a building block 

approach. If sufficient process and quality controls cannot be achieved, it may be 

necessary to account for greater variability with special factors (§ 29.619) applied to the 

design. Such factors should be accounted for in the component static tests or analysis. 

 (vi) It should be shown that impact damage (or other minor discrete source 

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 

static strength capability can be shown by analysis supported by test evidence, or by a 

combination of tests at the coupon, element, subcomponent, and component levels. 

Later discussions in this AC address the issues associated with damage in excess of 

that considered in f.(5) and drops in residual strength below ultimate load capability (see 

paragraph f.(6) below). 

 (6) The sixth area is the damage tolerance and fatigue evaluation requirements 

of § 29.573. 

 (i) Background. The static strength determination required by §§ 29.305 and 

29.307 establishes the ultimate load capability for composite structures that are 

manufactured, operated, and maintained with established procedures and conditions. 

The damage tolerance and fatigue evaluation required by § 29.573 mandates 

procedures that allow the composite structure to retain the intended ultimate load 

capability when subjected to expected fatigue loads and conditions during its 

operational life. The requirements established for the damage tolerance and fatigue 

evaluation include component replacement times, inspection intervals, or other 

procedures as necessary to avoid catastrophic failure. These evaluations assume that 

the baseline ultimate strength capability might be compromised by damage caused by 

fatigue, environmental effects, intrinsic or discrete flaws, or accidental damage. This 

damage includes flaws or defects, which may occur in manufacturing or maintenance 

and which are used to set the ultimate strength capability and establish the 
AC 29-2C 12/01/2011 

18 

manufacturing acceptance criteria. The damage tolerance assessment establishes 

standards that allow the static strength capability to degrade below the ultimate strength 

capability assuming such damage occurs within the operational life of the structure. 

However, when this damage occurs, the remaining structure must withstand expected 

loads without failure or excessive structural deformations until the damage is detected 

and the component is either repaired to restore ultimate strength capability or retired. 

 (ii) General. The nature and extent of the required analysis or tests on 

complete structures and portions of the primary structure can be based on applicable 

previous fatigue or damage tolerant designs, construction, tests, and service experience 

on similar structures. In the absence of experience with similar designs, 

FAA/AUTHORITY 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. 

Rotorcraft structure provides a broad range of composite applications that are quite 

different in terms of functionality, geometry and inspectability. These include the rotors, 

the drive shafts, the fuselage, control system components (e.g., push-pull rods), and the 

control surfaces. When selecting the approach, attention should be given to the 

composite application under evaluation, the type of potential damage and degradation 

of the structural design details, the materials used and margin over flight loads. 

Whatever the approach selected, the following considerations will apply for tests and 

analysis: 

 (A) The test articles should be fabricated and assembled in accordance 

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

of production structure. 

 (B) The test articles should include material imperfections whose extent is 

not less than the limits established under the inspection and acceptance criteria used 

during the manufacturing process and consistent with the inspection techniques used in 

service (e.g., visual, ultrasonic, X-ray). The initial extent of these imperfections should 

be discussed and agreed with the FAA, taking into account experience in manufacturing 

and routine in-service inspections. Typical defects to be considered include but are not 

limited to the following: 

 (1) Disbonds and weak bonds (considered as disbonds). 

 (2) Delaminations, fiber waviness, porosity, voids. 

 (3) Scratches, gouges, and penetrations. 

 (4) Impact damage. 
AC 29-2C 12/01/2011 

19 

All of the damages identified in the preceding paragraph (B) above should be derived 

from the threat assessment described in the following paragraph (C). 

 (C) For each PSE, a threat assessment must be made of the probable 

locations, types, and sizes of damage considering fatigue, environmental effects, 

intrinsic and discrete flaws, and impact or other accidental damage. This determination 

must be submitted with accompanying rationale to the FAA/AUTHORITY for approval. 

This rationale may include experience with similar materials, designs, processes 

(manufacturing, maintenance, and overhaul), structural details, or structure, and may 

also include service failure evaluations, manufacturing records, overhauls and repair 

reports, field service reports, incident and accident investigations, service impact 

surveys, inspectability surveys, and engineering judgment. 

Consideration should also be given to factors that: 

 - Reduce scatter and deviations from nominal structures, such as frozen processes, 

Flight Critical Parts programs, and materials and manufacturing processes to mitigate 

intrinsic flaws (inclusions and defects). 

 - Preclude a type of damage by use of a specific design feature (material selection, 

surface treatment, protective coating, or shielding), a specific stress level (for fatigue 

damage), or a specific manufacturing inspection process (if it can be shown to be highly 

reliable, well-controlled, documented, and systematically required). 

The assessment should include: 

 A systematic evaluation of all the location, types, and sizes of damage and their 

estimated probability of occurrence. 

 A selection or elimination of this damage based on the above estimate. 

 A verification that the inspection method selected is capable of detecting the 

damage at the size and location determined. 

The types of damage to consider include: 

 (1) Intrinsic Flaws (imperfections), which are probable to exist in an asmanufactured structure based on the evaluation of the details and potential sensitivities 

of the specific manufacturing work processes used. The types of flaws to be considered 

include voids, disbonds, inclusions, foreign objects, resin-rich and resin-starved areas, 

and improper ply orientation or ply ending. The sizes of the intrinsic flaws considered 

should be based on the limits established under the manufacturing inspection and 

acceptance criteria and are expected to remain in service for the life of the structure. 

 (2) Impact Damage, which may occur during manufacturing and in 

service based on an evaluation of the threats by means of an impact survey and/or 
AC 29-2C 12/01/2011 

20 

service experiences. This type of damage can include dents, penetrations, gouges, 

abrasions, and scratches. A threat assessment is needed to identify impact damage 

severity and detectability for design and maintenance. A threat assessment usually 

includes damage data collected from service plus an impact survey. An impact survey 

consists of impact tests performed with configured structure, which is subjected to 

boundary conditions characteristic of the structure. Many different impact scenarios and 

locations are typically considered in the survey, which has a goal of identifying the most 

critical impacts (i.e., those causing the most serious damage but are least detectable). 

When simulating accidental impact damage, blunt or sharp impactors should be 

selected to represent the maximum criticality versus detectability, according to the load 

conditions (e.g., tension, compression or shear). Until sufficient service experience 

exists to make good engineering judgments on energy and impactor variables, impact 

surveys should consider a wide range of conceivable impacts, including runway or 

ground debris, hail, tool drops, and vehicle collisions. Service data collected over time, 

can better define impact surveys and design criteria for subsequent products, as well as 

establish more rational inspection intervals and maintenance practice. Refer to 

paragraph f.(6)(ii)(H) for various combinations of detectability and energy levels to be 

considered in the damage tolerance and fatigue evaluation. 

 (3) Discrete Source Damage. The structure should be able to 

withstand limit static loads (considered as ultimate loads) and fatigue loads, which are 

reasonably expected during a completion of a flight on which damage resulting from 

obvious discrete source occurs (e.g., hail damage, bird strike, uncontained engine 

failure, and uncontained high energy rotating machinery failure). The extent of damage 

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

relating to each discrete source. 

 (D) The use of composite secondary bonding in manufacturing or 

maintenance requires strict process and quality controls to achieve the reliability needed 

to use such technology in critical structures (see AC 21-26). Assuming good process 

and quality controls, service history has shown that additional damage tolerant design 

considerations are also needed to ensure the safety of structure with secondary bonds 

(i.e., random, but an unacceptable number of weak bonds discovered in service). 

Unless the ultimate strength of each critical bonded joint can be reliably substantiated in 

production by NDI techniques (or other equivalent, approved techniques), then the limit 

load capability should be ensured by any or a combination of the following: 

 (1) Consider isolated disbonds and weak bonds (represented by zero 

bond strength) in structural elements that use secondary bonding for primary load 

transfer. The associated disbond size should be up to the limitations provided by 

redundant design features (i.e., mechanical fasteners or a separate bonding detail). 

The structure containing such damage should be shown to carry limit load by tests, 

analyses, or some combination of both. For purposes of test or analysis demonstration, 

each disbond should be considered separately as a random occurrence (i.e., it is not 

necessary to demonstrate residual strength with all structural elements disbonded 

simultaneously). 
AC 29-2C 12/01/2011 

21 

 (2) Each critical bonded joint on each production article should be 

proof-tested to the critical limit load. 

 (3) Critical bonded joints that have high static margins of safety (e.g., 

some rotor blades) may be accepted based on satisfactory service history of like or 

similar components. 

 (E) The fatigue load spectrum developed for fatigue testing and analysis 

purposes should be representative of the anticipated service usage. Low amplitude 

load levels that can be shown not to contribute to fatigue damage may be omitted 

(truncated). Reducing maximum load levels (clipping) is generally not accepted.

 (F) Environmental effects (temperature and humidity representative of the 

expected service usage) on the static and fatigue behavior and damage growth should 

be considered. Unless tested in the environment, appropriate environmental knock 

down factors for the static and the fatigue test articles should be derived and applied in 

the evaluation. 

 (G) Variability in fatigue behavior should be covered by appropriate load or 

life scatter factors and these factors should take into account the number of specimens 

tested. 

 (H) The following Figure AC 29.573-2 illustrates the extent of the impact 

damage that needs to be considered in the damage tolerance and fatigue evaluation. 
AC 29-2C 12/01/2011 

22 

Detectability

Energy

°

+

+

+

Level of energy selected for compliance with 

§§ 29.305, 29.307 (Static Strength) and f(5)(vi) of this AC. 

Maximum level of energy selected for the Damage Tolerance 

evaluation. This level should not be exceeded in service.

Obvious Damage

Barely Detectable Impact 

Damage

Detectable damage

°

°

Undetectable damage

+

+

Zone 1 Zone 3

Zone 2

Figure AC 29.573-2. Characterization of Impact Damage.

 (1) Both the energy level associated with the static strength 

demonstration and the maximum energy level associated with the damage tolerance 

evaluation (depicted in Figure AC 29.573-2) are dependent on the part of the structure 

under evaluation and a threat assessment. 

 (2) Obvious impact damage is used to define the threshold from which 

damage is readily detectable and appropriate actions may be taken before the next 

flight. 

 (3) Barely Detectable Impact Damage (BDID) is the state of damage at 

the threshold of detectability for the approved inspection procedure. Barely Visible 

Impact Damage (BVID) is that threshold of visually detectable damage associated with 

a detailed visual inspection procedure. 

 (4) Detectable Damage is the state of damage that can be reliably 

detected at scheduled inspection intervals. Visible Impact Damage (VID) is that 

threshold associated with the type of damage that should be detectable during a 

detailed visual inspection. 

 (5) Three Zones are depicted by this figure: 
AC 29-2C 12/01/2011 

23 

 Zone 1: Since the damage is not detectable, Ultimate Load 

capability is required. The provisions of paragraph f.(5) provide a means of compliance. 

 Zone 2: Since the damage can be detected at a scheduled 

inspection, Limit Load (considered as Ultimate load) capability is the minimum 

requirement for this damage. 

 Zone 3: Since the damage is not detectable with the proposed inservice inspection procedures, ultimate load capability is required, unless an alternate 

procedure can show an equivalent level of safety. For example, residual strength lower 

than ultimate may be used in association with improved inspection procedures or with a 

probabilistic approach showing that the occurrence of energy levels is low enough so 

that an acceptable level of safety can be achieved. 

Of the three zones, only Zone 3 may have a residual strength requirement that can vary 

with alternate procedures or the probability of damage occurrence or both. In either 

case, any compromise for residual strength requirements less than the ultimate load 

requirement should only be considered when pursuing one of the options under the 

damage tolerant fail-safe means of compliance, as described in the following section, 

f.(6)(iii)(B). 

One example of the use of alternate procedures is for the rare damage threat from a 

high energy, blunt impact (e.g., service vehicle collision). Depending on the selected 

maintenance inspection scheme, such damage may fall under Zone 3. When 

considering such damage in the design of a part, it may be shown to be damage 

tolerant fail safe, even though the damage is not detectable, based on a very low 

probability of occurrence. As a result, the design may have sufficiently high residual 

strength (e.g., below Ultimate, but well above limit load capability to ensure safety 

without detection for long periods of time). If it is further determined that such impact 

events usually occur with the knowledge of maintenance or aircraft service personnel, 

then the alternate procedures may be added to the Instructions for Continued 

Airworthiness. For example, advanced inspection methods, which can detect damage 

from high-energy blunt impacts, may be used as alternate procedures to minimize the 

risk of catastrophic failure for such Zone 3 damage. 

 (iii) Means of compliance. For each PSE, inspections, replacement times, or 

other procedures must be established as necessary to avoid catastrophic failure. 

Compliance with the requirements of § 29.573(d) and (e) should be shown by one, or a 

combination of, the methods described subsequently. Generally, replacement times are 

established using Damage Tolerance Safe Life Evaluations and Inspection Intervals are 

established using Fail Safe Evaluations. From current state-of-the-art rotorcraft 

applications, it is widely accepted that composite materials have good flaw and damage 

tolerance capabilities and therefore the supplemental procedures may only be rarely 

necessary. Damage tolerance evaluations are best suited for composite structures, 

particularly those with structural redundancy and inherent resistance to damage growth. 

Damage resulting from anomalous or accidental events must be considered in the 
AC 29-2C 12/01/2011 

24 

damage tolerant evaluations. The damage tolerant evaluation for replacement times 

and inspection intervals is to be used unless it is established that neither can be 

achieved within the limitations of geometry, inspectability, or good design practice. In 

that case, supplemental procedures must be established and submitted to the FAA for 

approval. In any case, the FAA must approve the methodology used for compliance to 

§ 29.573. 

The substantiation method(s) should be chosen so that the structure is protected 

against catastrophic failure from each of the threats identified in paragraph f.(6)(ii)(C) of 

this AC by a specific procedure (inspection, replacement time, or other procedure). For 

example, a manufacturing-related void of a specific allowable size could be 

substantiated by means of a replacement time method with no scheduled inspection. 

An accidental impact in the same area could be substantiated by an inspection method 

with no specific replacement time. The result could be one structure with several 

different inspection requirements (location, method, and interval) and a fixed 

replacement time as well. This combination of procedures assures that each threat is 

covered. 

The fatigue substantiation should include sufficient coupon, element, sub-element, or 

component tests to establish the fatigue scatter, curve shapes, and the environmental 

effects. The substantiation should include full-scale, component, or sub-component 

fatigue testing but also may be accomplished by analysis supported by test evidence. 

When spectrum testing is used, the lowest load levels can be eliminated from the 

spectrum if they can be shown to be non-damaging. The substantiation should include 

a static strength evaluation to show that the required residual strength and adequate 

stiffness, accounting for the effects of environment, are retained for the life of the 

structure or the appropriate inspection interval. Damage as determined in paragraph 

f.(6)(ii) of this AC for the specific structure being substantiated should be imposed at 

each critical area of the structure. 

 (A) Damage Tolerant Safe-Life Evaluation. This is a “No-Growth” method 

in which the structure, with damage present, is able to withstand repeated loads of 

variable magnitude without detectable damage growth for the life of the rotorcraft or 

within a specified replacement time. This evaluation may be used to substantiate any 

type of damage that will remain in-service for the life of the part. 

No specific inspection requirements are generated from the test program in this method. 

However, compliance with routine inspections for cracking, delaminations, and service 

damage and other limitations prescribed in accordance with § 29.1529 are always 

required. Compliance using full-scale, component, or sub-component fatigue testing 

can be accomplished by either of the following methods: 

 (1) S-N Method. This method is based on determining the point where 

initiation of growth occurs for the damage present at critical locations in the structure. 

AC 27-1B, AC 27 MG 11, provides guidance that may be appropriate for this method. 

The method utilizes one or more full-scale, component, or sub-component test 
AC 29-2C 12/01/2011 

25 

specimens subjected to constant-amplitude or spectrum loading applied in a distribution 

on the structure that is representative of critical flight conditions. Any indication of 

growth of the imposed damage and defects, or structurally significant cracking, 

disbonding, splintering, or delaminating of the composite, defines the fatigue initiation 

characteristic of the structure in terms of applied load and cycles. Working S-N curves 

are established from the mean curve using strength or cycle reductions or both to 

account for fatigue scatter and environmental effects. Flight loads are compared to this 

working curve, and if any intercepts occur, a cumulative damage calculation is 

conducted to establish the component retirement time. Compliance with the ultimate 

load requirements should be demonstrated at the completion of the fatigue test. 

 (2) Life-Test Method. This method uses spectrum fatigue testing to 

verify the absence of damage growth over a large number of cycles that are equivalent 

to a lifetime of expected usage. The method uses one or more full-scale, component, or 

sub-component test specimens subjected to spectrum fatigue loading applied in a 

representative distribution of flight loads, including Ground-Air-Ground (GAG) loads. 

Fatigue test loads should be increased by factors for environment and fatigue strength 

scatter. The load may also be increased using an S-N curve approach to reduce the 

duration of the test. Any significant growth of the imposed damage, or structurally 

significant cracking, disbonding, splintering, or delamination of the composite during the 

test constitutes failure to achieve the desired lifetime. However, the equivalent life 

demonstrated at the time of inception of damage growth or cracking can be used as a 

retirement time for the component. Compliance with the ultimate load requirements 

should be demonstrated at the completion of the fatigue test. 

 (B) Damage Tolerant Fail-Safe (Residual Strength with Detectable 

Damage) Evaluation. This method establishes inspection intervals to ensure that the 

structure remaining after a partial failure is able to withstand design limit loads without 

failure or excessive structural deformations within a specified inspection interval. If the 

damage is detected in an inspection, the structure should be either replaced or repaired 

to restore ultimate load capability. Evaluation of Zone 3 damage should have 

sufficiently high residual strength and, if necessary, supplemental procedures should be 

established to minimize the risk of catastrophic failure. Full-scale, component, or subcomponent testing should be accomplished using one or more specimens subjected to 

constant amplitude or spectrum loading applied in a manner representative of flight load 

conditions. The test loads should be increased by factors that account for environment 

and fatigue strength scatter. The results of the testing can be used to manage the 

structure in one or a combination of the three methods described subsequently. 

 (1) No Growth Evaluation. This approach is appropriate for 

inspectable in-service damage, which does not grow in service (see Figure 

AC 29.573-3). Damage growth should be substantiated using either method described 

in f.(6)(iii)(B)(2) or f.(6)(iii)(B)(3). Structural details, elements, sub-components, and 

components of critical structural areas, or full-scale structures, should be tested under 

repeated loads for validating a no-growth approach to the damage tolerance 

requirements. The number of cycles applied to validate a no-growth concept should be 
AC 29-2C 12/01/2011 

26 

statistically significant, and may be determined by load or life considerations or both. 

Residual strength testing or evaluations should be performed after repeated load cycling 

demonstrating that the residual strength of the structure is equal to or greater than limit 

load considered as ultimate. Moreover, it should be shown that stiffness properties 

have not changed beyond acceptable levels. Inspection intervals should be 

established, considering the residual strength capability associated with the assumed 

damage. The intent of this is to assure that structure is not exposed to an excessive 

period of time with static margins less than ultimate, providing a lower safety level than 

in the typical slow growth situation, as illustrated by the Figure AC 29.573-3. Once the 

damage is detected, the component is either repaired to restore ultimate load capability 

or replaced. 

Residual strength

Time

No-growth approach **

Slow-growth approach *

Ultimate loads

Limit loads

Damage initiation 

or occurrence

* Repair to Restore Ultimate Strength 

** No growth without repair is not acceptable

Figure AC 29.573-3. Residual Strength vs. Time.

The lower the residual strength of a structure after an accidental damage event, the 

shorter the inspection interval should be. Considerations of both inspectability and 

impact surveys (including probability of occurrence) for specific structure may be used 

to isolate the most critical threats to consider in setting a maintenance inspection 

interval. Knowledge of the residual strength for a given critical damage is also needed 

for such an evaluation. If it is known that the design is capable of handling large and 

clearly detectable damage, while maintaining a residual strength well above limit load, a 

less rigorous engineering approach may be applied in establishing the inspection 

interval. 

 (2) Slow Growth Evaluation. This method is applicable when the 

damage grows in the test and the growth rate is shown to be slow, stable, and 

predictable, as illustrated in Figure AC 29.573-4. An inspection program should be 

developed consisting of the frequency, extent, and methods of inspection for inclusion in 

the maintenance plan. Inspection intervals should be established so that the damage 
AC 29-2C 12/01/2011 

27 

will have a very high probability of detection between the time it becomes initially 

inspectable and the time at which the extent of the damage reduces the residual static 

strength to limit load (considered as ultimate), including the effects of environment. For 

any damage size that reduces the load capability below ultimate, the component is 

either repaired to restore ultimate load capability or replaced. Should functional 

impairment (such as unacceptable loss of stiffness) occur before the damage becomes 

otherwise critical, this should be accounted for in the development of the inspection 

program. 

Damage Size

Static Strength

time

Ultimate

Limit

Slow Growth

Arrested Growth

Slow Growth

Inspection Interval Basis

Critical

Size

Arrested Growth

Slow Growth

No Growth

Threshold of 

Detectability

Figure AC 29.573-4. Illustration of Residual Strength and Damage Size Relationships

 for Fail-Safe Substantiation.

 (3) Arrested Growth Evaluation. This method is applicable when the 

damage grows, but the growth is mechanically arrested or terminated before becoming 

critical (residual static strength reduced to limit load), as illustrated in Figure 

AC 29.573-4. Arrested Growth may occur due to design features such as a geometry 

change, reinforcement, thickness change, or a structural joint. This approach is 

appropriate for inspectable arrested growth damage. Structural details, elements, and 

sub-components of critical structural areas, or full-scale structures, should be tested 

under repeated loads for validating an arrested growth approach to the flaw tolerance 

requirements. The number of cycles applied to validate an arrested growth concept 

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

or both. Residual strength testing or evaluation should be performed after repeated 

load cycling and a demonstration that the residual strength of the structure is equal to or 

greater than limit load considered as ultimate. Moreover, it should be shown that 

stiffness properties have not changed beyond acceptable levels. Inspection intervals 
AC 29-2C 12/01/2011 

28 

should be established, considering the residual strength capability associated with the 

arrested growth damage. The intent of this is to assure that structure is not exposed to 

an excessive period of time with static margins less than ultimate, providing a lower 

safety level than in the typical slow growth situation, as illustrated by Figure 

AC 29.573-3. For any damage size that reduces the load capability below ultimate, the 

component is either repaired to restore ultimate load capability or replaced. 

The lower the residual strength of a structure after an arrested growth event, the shorter 

the inspection interval should be. Considerations of both inspectability and impact 

surveys (including probability of occurrence) for specific structure may be used to 

isolate the most critical threats to consider in setting a maintenance inspection interval. 

Knowledge of the residual strength for a given critical damage is also needed for such 

an evaluation. If it is known that the design is capable of handling large and clearly 

detectable damage, while maintaining a residual strength well above limit load, a less 

rigorous engineering approach may be applied in establishing the inspection interval. 

 (C) Combination of Damage Tolerant Safe Life and Fail Safe Evaluations. 

Generally, it may be appropriate to establish both a replacement time and an inspection 

program for a given structure as calculated by the Damage Tolerant Safe Life and Fail 

Safe Evaluations. 

 (D) Other Procedures. Other procedures are allowed according to 

§ 29.573(d). Such alternative procedures must still provide the same degree of damage 

tolerance to the same identified threats as the replacement time or inspection interval 

methods. 

One possible alternate approach is the use of indirect damage detection methods 

instead of the specific mandated inspection procedures that are determined in the Fail 

Safe Evaluations of f.(6)(iii)(B). These indirect detection methods should be 

documented and shown to have the same degree of reliability, repeatability, and margin 

provided by a conventional inspection approach. These methods could include: (1) 

establishing measurable vibration or blade out-of-track conditions and limits, (2) defining 

indirect inspections, which would detect damage, and (3) in-flight detecting of damage 

by means of monitoring and warning devices. 

 (E) Supplemental Procedures. If the damage tolerant evaluations as 

described previously cannot be achieved within the limitations of geometry, 

inspectability, or good design practice, a fatigue evaluation using supplemental 

procedures may be proposed to the FAA/AUTHORITY per § 29.573(e). The applicant 

must establish that the damage tolerance criteria are impracticable and cannot be 

satisfied for the specific PSE, locations, and threats considered. In addition, the types 

of damage considered in the evaluations must be identified. Finally, supplemental 

procedures must be established to minimize the risk of catastrophic failure with the 

damages considered. 

 (iv) Additional considerations for damage tolerance and fatigue evaluations. 
AC 29-2C 12/01/2011 

29 

 (A) Experience with the application of methods of fatigue and damage 

tolerance evaluations indicates that a relevant test background should exist in order to 

achieve the design objective. It is the general practice within industry to conduct 

damage tolerance tests for design information and guidance purposes. It is crucial that 

the critical structure be identified and tested to the proper flight and ground loads. 

 (B) Identification of the structure to be considered in each evaluation (a 

failure mode and effects analysis or similar method should be used). 

 (1) Identification of Principal Structural Elements. Principal structural 

elements are those that contribute significantly to carrying flight and ground loads and 

whose failure could result in catastrophic failure of the rotorcraft. Typical examples of 

such elements are: 

 (i) Rotor blades and attachment fittings. 

 (ii) Rotor heads, including hubs, hinges, and some main rotor 

dampers. 

 (iii) Control system components subject to repeated loading, 

including control rods, servo structure, and swashplates. 

 (iv) Rotor supporting structure (lift path from airframe to rotorhead). 

 (v) Fuselage, including stabilizers and auxiliary lifting surfaces, 

airframe provisions for engine and transmission mountings. 

 (vi) Main fixed or retractable landing gear and fuselage attachment 

structure. 

 (2) Identification of Locations Within Principal Structural Elements to be 

Evaluated. The locations of damage to structure for damage tolerance evaluation can 

be determined by analysis or by fatigue test on complete structures or subcomponents. 

However, tests will be necessary when the basis for analytical prediction is not reliable, 

such as for complex components. If less than the complete structure is tested, care 

should be taken to ensure that the internal loads and boundary conditions are valid. 

The following should be considered: 

 (i) strain gauge data on undamaged structure to establish points of 

high stress concentration as well as the magnitude of the concentration; 

 (ii) locations where analysis shows high stress or low margins of 

safety; 

 (iii) locations where permanent deformation occurred in static tests; 
AC 29-2C 12/01/2011 

30 

 (iv) locations of potential fatigue damage identified by fatigue 

analysis; 

 (v) locations where the stresses in adjacent elements will be at a 

maximum with an element in the location failed; 

 (vi) partial fracture locations in an element where high stress 

concentrations are present in the residual structure; 

 (vii) locations where detection would be difficult; and 

 (viii) design details that service experience of similarly designed 

components indicates are prone to fatigue or other damage. 

 (3) In addition, the areas of probable damage from sources such as a 

severe corrosive or fretting environment, a wear or galling environment, or a high 

maintenance environment should be determined from a review of the design and past 

service experience. 

 (C) The stresses and strains (steady and oscillatory) associated with all 

representative steady and maneuvering operating conditions expected in service. 

 (D) The frequency of occurrences of various flight conditions and the 

corresponding spectrum of loadings and stresses. 

 (E) The fatigue strength, fatigue crack propagation characteristics of the 

materials used and of the structure, and the residual strength of the damaged structure. 

 (F) Inspectability, inspection methods, and detectable flaw sizes. 

 (G) Variability of the measured stresses of paragraph f.(6)(iv)(C), the 

actual flight condition occurrences of paragraph f.(6)(iv)(D), and the fatigue strength 

material properties of paragraph f.(6)(iv)(E). 

 (v) Flight strain measurement program. 

 (A) General. Subsequent to design analysis, in which aircraft loads and 

associated stresses are derived, the stress level or loads are to be verified by a 

carefully controlled flight strain measurement program. (This guidance is similar to that 

of AC 27-1B, MG 11.) 

 (B) Instrumentation. 

 (1) The instrumentation system used in the flight strain measurement 

program should accurately measure and record the critical strains under test conditions 
AC 29-2C 12/01/2011 

31 

associated with normal operation and specific maneuvers. The location and distribution 

of the strain gauges should be based on a rational evaluation of the critical stress areas. 

This may be accomplished by appropriate analytical means supplemented, when 

necessary, by strain sensitive coatings or photoelastic methods. The distribution and 

number of strain gauges should cover the load spectrum adequately for each part 

essential to the safe operation of the rotorcraft as identified in § 29.573(d)(1). Other 

devices such as accelerometers may be used as appropriate. 

 (2) The corresponding flight parameters (airspeed, rotor RPM, centerof-gravity accelerations, etc.) should also be recorded simultaneously by appropriate 

methods. This is necessary to correlate the loads and stresses with the maneuver or 

operating conditions at which they occurred. 

 (3) The instrumentation system should be adequately calibrated and 

checked periodically throughout the flight strain measurement program to ensure 

consistent and accurate results. 

 (C) Parts to be Strain-Gauged. Fatigue critical portions of the rotor 

systems, control systems, landing gear, fuselage, and supporting structure for rotors, 

transmissions, and engine are to be strain-gauged. For rotorcraft of unusual or unique 

design, special consideration might be necessary to ensure that all the essential parts 

are evaluated. 

 (D) Flight Regimes and Conditions to be investigated. 

 (1) Typical flight and ground conditions to be investigated in the flight 

strain measurement program are given in paragraphs c. and d. of AC 27-1B, MG 11. 

 (2) The determination of flight conditions to be investigated in the flight 

strain measurement program should be based on the anticipated use of the rotorcraft 

and, if available, on past service records for similar designs. In any event, the flight 

conditions considered appropriate for the design and application should be 

representative of the actual operation in accordance with the rotorcraft flight manual. In 

the case of multiengine rotorcraft, the flight conditions concerning partial engine-out 

operation should be considered in addition to complete power-off operation. The flight 

conditions to be investigated should be submitted in connection with the flight evaluation 

program. 

 (3) The severity of the maneuvers investigated during the flight strain 

survey should be at least as severe as the maneuvers likely to occur in service. 

 (4) All flight conditions considered appropriate for the particular design 

are to be investigated over the complete rotor speed, airspeed, center of gravity, 

altitude, and weight ranges to determine the most critical stress levels associated with 

each flight condition. The temperature effects on loads as affected by elastomeric 

components are to be investigated. To account for data scatter and to determine the 
AC 29-2C 12/01/2011 

32 

stress levels present, a sufficient amount of data points should be obtained at each 

flight condition. Consideration can be given to the use of scatter factors in determining 

the sufficiency of data points. In some instances, the critical weight, center of gravity, 

and altitude ranges for the various maneuvers can be based on past experience with 

similar design. This procedure is acceptable where adequate flight tests are performed 

to substantiate such selections. The combinations of flight parameters that produce the 

most critical stress levels should be used in the evaluation. 

 (vi) Frequency of loading. 

 (A) Types of Operation. 

 (1) The probable types of operation (transport, utility, etc.) for the 

rotorcraft should be established. The type of operation can have a major influence on 

the loading environment. In the past, rotorcrafts have been substantiated for the most 

critical general types of operation with some consideration of special, occasional types 

of operation. To assure that the most critical types of operation are considered, each 

major rotorcraft structural component should be substantiated for the most critical types 

of operation as established by the manufacturer. The types of operation shown below 

should be considered and, if applicable, used in the substantiation: 

 (i) Long flights to remote sites (low ground-air-ground cycles but 

high cruising speeds). 

 (ii) Typical, general types of operation. 

 (iii) Short flights as used in logging operations. 

 (2) One means is to substantiate for the most severe type of operation; 

however, this method is not always economically feasible. 

 (3) A second means is to quantify the influence of mission type on 

fatigue damage by adding to or replacing hour limitations by flight cycle limitations (if 

properly defined and easily identifiable by the crew, for example: one landing, one load 

transportation). A special type of flight hour limitation replacement using factorization of 

flight hours for multiple types of operations may be feasible if continuing manufacturers’ 

technical support is provided and documented (i.e., the manufacturer either provides the 

factorization analyses or checks them on a continuing basis for each type of rotorcraft 

operation). 

 (4) Where one or more operations are not among the general uses 

intended for the rotorcraft, the rotorcraft flight manual should state in the limitations 

section that the intended use of the rotorcraft does not include certain missions or 

repeated maneuvers (e.g., logging with its high number of takeoffs and landings per 

hour). A note to this effect should also appear in the rotorcraft airworthiness limitations 
AC 29-2C 12/01/2011 

33 

section of the maintenance manual prepared in accordance with §§ 29.573 

and 29.1529. 

 (5) Should subsequent usage of the rotorcraft encompass a mission 

outside the original structural substantiation, the effects of this new mission environment 

on the frequency of loading and structural substantiation should be addressed and 

where practicable, in the interest of safety, a reassessment made. If this reassessment 

indicates the necessity for revised retirement times, those new times may be limited to 

specific rotorcraft model involved in the added mission provided: 

 (i) changes are adopted through the airworthiness directives 

process and proper part re-identification is established; or 

 (ii) a Rotorcraft Flight Manual (RFM) supplement outlining the 

limitations is approved; or 

 (iii) an airworthiness limitations section (ALS) supplement is 

approved; or 

 (iv) an appropriate combination of part re-identification, RFM 

supplement, or airworthiness limitation section supplement is approved. 

 (B) Loading Spectrum. The spectrum allocating percentage of time or 

frequencies of occurrence to flight conditions or maneuvers is to be based on the 

expected usage of the rotorcraft. This spectrum is to be established so that it is unlikely 

that actual usage will subject the structure to damage beyond that associated with the 

spectrum. Considerations to be included in developing this spectrum should include 

prior knowledge based on flight history recorder data, design limitations established in 

compliance with § 29.309, and recommended operating conditions and limitations 

specified in the rotorcraft flight manual or instructions for continued airworthiness (ICA). 

The distribution of times at various forward flight speeds should reflect not only the 

relation of these speeds to VNE but also the recommended operating conditions in the 

rotorcraft flight manual or ICA that govern Vc or cruise speed. It is desirable to conduct 

the flight strain-gauge program by simulating the usage as determined previously, with 

continuous recording of stresses and loads, thus obtaining directly the stress or load 

spectra for structural elements. 

 (7) The seventh major area is the dynamic loading and response requirements of 

§§ 29.241, 29.251, and 29.629 for vibration and resonance frequency determination and 

separation for aeroelastic stability and stability margin determination for dynamically 

critical flight structure. Critical parts, locations, excitation modes, and separations 

should be identified and substantiated. This substantiation should consist of analysis 

supported by tests, including tests that account for repeated loading effects and 

environment exposure effects on critical properties, such as stiffness, mass, and 

damping. This must be accomplished to assure that the initial stiffness, residual 
AC 29-2C 12/01/2011 

34 

stiffness, proper critical frequency design, and structural damping are provided as 

necessary to prevent vibration, resonance, and flutter problems. 

 (i) All vibration and resonance critical composite structures must be identified 

and properly evaluated. 

 (ii) All flutter-critical composite structures must be identified and properly 

evaluated. This structure must be shown by analysis to be flutter free to 1.1 VNE (or any 

other critical operating limit, such as VD, for a VSTOL aircraft) with the extent of damage 

for which residual strength and stiffness are demonstrated. 

 (iii) Where appropriate, crash impact dynamics considerations should be 

taken into account to ensure proper crash resistance and a proper level of occupant 

safety for an otherwise survivable impact. 

 (8) The eighth area is the special repair and continued airworthiness

requirements of §§ 29.611, 29.1529, and 14 CFR part 29 Appendix A, for composite 

structures. When repair and continued airworthiness procedures are provided in 

service documents (including approved sections of the maintenance manual or 

instructions for continued airworthiness), the resulting repairs and maintenance 

provisions should be shown to provide structure, which continually meets the guidance 

of paragraphs (1) through (7) of this AC. All certification-based repair and continued 

airworthiness standards, limits, and inspections must be clearly stated, and their 

provisions and limitations clearly documented to ensure continued airworthiness. No 

composite structural repair should be attempted that is beyond the scope of the 

applicable approved Structural Repair Manual (SRM) without an engineering design 

approval by a qualified FAA/AUTHORITY designated representative.