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 st ructures. Paragraph f.(6) specifically addresses the advisory guidance applying to damage tolerance and fatigue evaluation as required by § 29.573, Amendm ent 29-54. Some information contained in AC 29-2C, MG 8 (Amendment 29-42) is repeated and updated, as appropr iate, to preserve the “building block” approach for analyses of com posite 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 roto rcraft. A uniform approach to composite

structural substantiation is des irable, but it is recognized th at in a continually developing technical area, which has diverse industr ial roots both in aerospace and in other industries, variations and deviations fr om 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 uni quely and is inherently subjected to severe cyclic stresses, special consideration is required for the

substantiation of all rotorcraft structur e, including composites. This special consideration is necessary to ensure that the level of safety intended by the current regulations are attained during the type certific ation process for all structure with special emphasis on composite struct ure because of its unique stru ctural characteristics, manufacturing quality and operat ional considerations, and failure mechanisms.

c. Background

.

(1) Historically, rotorcraft have requi red unique, conservative structural substantiation because of unique configuration effects, uniq ue loading considerations, severe fatigue spectrum effects, and the specialized comprehensive fatigue testing required by these effects. Ro torcraft structural static st rength substantiation for both metal and composite structure is essentially identical to that fo r fixed wing structure

once basic loads have been determined. Howe ver, rotorcraft st ructural fatigue substantiation is significantly different fr om fixed wing fatigue substantiation. Since AC 20-107, as developed, applies to both fi xed wing aircraft and ro torcraft, it, of necessity, was finalized in a broad generic fo rm. Accordingly, a need to supplementAC 29-2C 12/01/2011

2 AC 20-107 for rotorcra ft was recognized during type ce rtification pr ograms. One significant difference in traditional rotorcra ft 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 rat her than spectrum tests because of the high frequency loads common to rotorcraft operati ons. These rotorcraft fati gue 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 ma in or tail rotor system) using a monotonic (sinusoidal) fatigue spectrum based on maximum and minimum service stress values. Unless configuration differenc es or flight usage dat a 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 curv es 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, “Fati gue Tolerance Evaluation of Transport Category Rotorcraft Metallic Structure”). The three data points selected are a short specimen life (low-cycle fatigue), an interm ediate 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 re liability of the associated S-N curve. Order 8110.9, “Handbook on Vibration Subs tantiation and Fati gue Evaluation of Helicopter and Other Power Transmission S ystems” and AC 27-1B, AC 27 MG 11, “Fatigue Evaluation of Rotorc raft Structure”, contain comp rehensive discussions of the S-N curve generation process. The rotorcraft S-N curve process contrasts sharply with the fixed wing process of usi ng 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., Me tallic Materials Properties Development and Standardization (MMPDS), formerly the MIL-HDB K-5 for metals; Composites Materials Handbook-17 (CMH-17), formerly the MIL-HD BK-17 for composites; or AC 23-13, “Fatigue, Fail-Safe, and Damage Tolerance Ev aluation of Metallic St ructure for Normal, Utility, Acrobatic, and Commuter Airplanes”, which replaced AF S-120-73-2 for full-scale assemblies), a non-monotonic spectrum, and relati vely 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 ro tor 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 airc raft). In any case, the early 1950 period was well before a cl ear, detailed understanding of composite structural behavior (especially in the ar eas 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 fi eld. It also predated the initial issuance of AC 20-107. Currently, much composite des ign information is proprietary, either to government, industry or bot h, and many data gathering methods have not been completely standardized. Consequently, a signi ficant variation from laboratory toAC 29-2C 12/01/2011

3 laboratory in material property value dete rmination methods and results can exist. The early rotor blade designs (as well as current des igns) are by nature relatively low strain, tension structure designs. Also, by natur e, these designs ar e not damage or flaw critical. Thus, by circumstance as much as design, early composite rotor blade and other composite rotorcraft designs incorpor ated an acceptable fatigue tolerance level of safety. In the 1980’s, more test data, analytical knowledge, and analytical methodology became available to more completely subst antiate 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 guidanc e 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 s ources, 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 collisi ons, 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 equip ped 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 ± pai rs (not necessarily adjacent).AC 29-2C 12/01/2011

4 (9) Bond . The adhesion of one surface to anot her, 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 diffe rent materials in a single step. Examples include the curing of various com patible resin system pre-pregs, using the same cure cycle, to produce hybrid composit e structure or the curing of compatible composite materials and structural adhesives , using the same cure cycle, to produce

sandwich structure or skins wit h 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., usua lly a flat laminate) for evaluation of basic lamina or laminate properties or proper ties 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 (crosslinki ng) agents, with or without a catalyst, and

with or without heat.

(15) Damage. A generic term for structural a nomalies caused by manufacturing (processing, fabrication, assembly or handli ng) or service usage. Trimming, fastener installation, or foreign object impact are pot ential 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 st ructure has sustained a given level of fatigue, corrosion, accide ntal or discrete source damage.

(17) Damage Tolerant Fail-Safe . The capability of structure remaining after a partial failure to withstand design limit l oads without catastrophic failure within an inspection period.

(18) Damage Tolerant Safe Life . Capability of struct ure with damage present to survive expected repeated loads of vari able magnitude without detectable damage growth and to maintain ultimate load capabil ity 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

5 (21) Detail . A non-generic structural element of a more complex structural member (e.g., specific design configured join ts, 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 ti me 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 c onditions (excluding mechanical loading), separately or in combination, t hat 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 delaminati ons, disbonds, splintering, or cracking.

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

one-dimensional in the macro-behavior sens e, used as a princi pal constituent in advanced composites because of it s 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 har d 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 ty pes (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 rela ted to manufacturing operations , processing or assembly, such as voids, gaps, porosity, inclusions, fi ber 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

6 (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 composit e structure provides for no flaw growth and thus provides damage tolerance of the maxi mum defects allowed during the certification process using the approved design fatigue spectrum.

(37) Material System . The combination of single c onstituents 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 resi ns, which are mainly thermoset polymers

in aircraft structure.

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

structural element due to serv ice parameters such as incide nt heat fluxes, temperature,

and air flow at the time of o ccurrence 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 lo ad paths with one or more of the members (or areas of a member) re latively unloaded until fa ilure 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 t he 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 impregnat ed with resin, usually partially cured, and ready for manufacturing use in a final produc t that will involve complete curing. Prepreg is usually drapable, ta cky, and can be easily handled.AC 29-2C 12/01/2011

7 (46) Principal St ructural 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 so me period of unrepaired use after a failure or partial failu re due to fatigue, accidental, or discrete source of damage.

(48) Resin . An organic material with indefinit e 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 alr eady-cured composite parts, duri ng which the only chemical or thermal reaction occurring is the curing of the adhesive itself. The joining together of one already-cured composite par t to an uncured composite pa rt, through the curing of the resin of the uncured part, is also cons idered for the purposes of this advisory circular to be a secondary bondi ng 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 require ments and remain suitable for its intended function.

(52) Strain Level . As used herein, is defined as t he 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 st ructural 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 se ction of the full stru cture (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 fo rming unidirectional pre-preg.

(56) Thermoplastic . A plastic that repeatedly can be softened by heating and hardened by cooling through a tem perature range characteristic of the plastic, and when in the softened stage, can be shaped by flow in to 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 c hemical change; (i.e., it is substantially infusible and insoluble after having been cured by heat or other means).AC 29-2C 12/01/2011

8 (58) Warp . Yarns extended along the length of t he 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 Vibr ation Substantiation and Fatigue Evaluation of Helicopt er and other Power Transmission Systems

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

AC 20-107 “Composite Aircraft Structure”

AC 21-26 “Quality Control for t he Manufacture of Composite

Materials”

CMH-17 “Composite Materials Handbook”

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

Rotorcraft Metallic Structure”

DOT/FAA/CT-86/39 Whitehead, R.S., Kan, H.P., Cordero, R., and Seather, R., “Certificati on Testing Methodology for Composite Structures ”, October 1986.

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

structures evaluation has been di vided into eight basic regulat ory areas to provide focus on relevant regulatory requireme nts. These eight areas are: fabrication requirements; basic constituent, pre-preg and laminate ma terial 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 spec ial 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 assembli es, bonded assemblies) should be qualified in accordance with the methodology presen ted 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 approv al in each of the

various areas outlined herein. It is reco mmended that a FAA/AUTHORITY certification team approach be used for composite structur al substantiation. The team shouldAC 29-2C 12/01/2011

9 consist of FAA/AUTHORITY and cognizant me mbers of the applic ant’s organization. Personnel who are composites specialists (or are otherwise knowledgeable in the subject) should be primary team member candidates.

Once selected, it is recomm ended that team meet ings be held periodically (possibly in conjunction with type boards) during certif ication to ensure the building block certification process is accomplished as intended. The team should assure that permanent documentation in the form of repor ts or other FAA/AUTHORITY acceptable

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

but is not limited to the structural subs tantiation reports (bot h analysis and test), manufacturing processes and quality cont rol, 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 practi ces for many of the areas id entified below are available

in CMH-17.

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

(i) The quality control system shoul d 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 manufacturi ng quality instructions for composites are presented, evaluated, documented, and appr oved, using drawings, process and manufacturing specifications, standards, or ot her equivalent means. This should be one of the early phases of a com posite structure cert ification program, si nce this represents a major building block for sequential substant iation 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 defec ts, resin content, large area disbonds, and delaminations) for a particular mate rial system component, laminate design, detailed part, or assembly should be jointly established by engineering, manufacturing, and quality, and the associated inspection pr ograms created, validated, and approved for defect detection. Each critical engineerin g design should consider the variability of the manufacturing process to determine the worse case effects (maximum waviness, disbonds, delaminations, and other critical defec ts) 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 met hods, special fabrication methods, or other approved verification methods (e.g., engine ering proof tests - see paragraph f.(6)) should be provided to detect and limit disbonds or understrength bonds.

(iv) Structurally critical com posite construction fabrication process and procurement specifications, fo r fabricating reproducible and reliable structure, must beAC 29-2C 12/01/2011

10 provided and FAA approved ear ly 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 certificati on team members (bot h the manufacturing inspection district office (MIDO) and FAA engineering) at any time, that their quality control systems ensure on a continuous basis, t hat only qualified suppliers provide the basic material constituents or material syst ems (e.g., pre-pregs) that meet approved material specifications. Recommended guidelines for qualification of alternate material systems and suppliers are contained in CM H-17. These methods can also be used periodically for qualification status renewal s 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-p rocess 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 charac terization tests on basic constituent and material system samples, which should be conducted, as a minimum, on a batch-to-batch basis. The perio dic 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 certific ation team member s (both MIDO and engineering), at any time, that their compos ite material system (or constituent) storage and handling procedures and specif ications provide products, which continuously meet approved material and process specifications . Quality systems sh ould be designed with appropriate checks and balances, so that t he necessary statistical reliability and confidence levels for the items being inspecte d (that are specified by engineering) are continuously maintained. This should require , as a minimum, periodic inspections to ensure that proper records are kept on crit ical parameters (e.g., room temperature “bench” exposure, shelf life) and that peri odic basic constituent and material system characterization tests are conducted, on a ba tch-to-batch basis. The periodic testing necessary to maintain the quality standar d 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 validati on level of the quality contro l system, which should be specified for each critical item or cons tituent by the approv ed quality and engineering specifications. The statistical validation level should be defined an d approved early in certification. Also, approval and proper usage should be continuously maintained during the entire procurem ent and manufacturing cycles.AC 29-2C 12/01/2011

11 (v) Alternate fabrication and proce ss specifications should be approved and must comply with § 29.605. Any alternate s pecifications should provide at least the same level of quality and safety as the orig inal 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 app lication of the crit ical environmental limits such as temperature, hum idity, and exposure to aircraft fluids (such as fuel, oils, and hydraulic fluids), to determine their e ffect on the performanc e of each composite material system. Temperature and humidit y effects are commonly considered by coupon and component tests utilizing precondi tioned test specimens for each material system selected. Material “A” and “B” basis allowable strength va lues 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 a llowables for inclusion in CMH-17. Also, test methods vary somewhat from manufacturer to manufactu rer; 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 requir ements of all material systems (e.g., pre-pregs, adhesives) and constituents (e.g., fibers, resins) should be identified, documented, and approved. These requirement s, once approved, should be placed in all appropriate procedures and specificati ons such as those in paragraph f.(1).

(ii) Moisture conditioning of test coupons, parts, subassemblies, or assemblies should be accomplished in a ccordance with CMH-17, other similar approved methods or per FAA approved programs.

(iii) The maximum and minimum tem peratures expected in service (as derived from test measurements, t hermal analyses on panels and other parts, experience, or a combination) should be determined and acc ounted for in static and fatigue strength (including damage tolerance) substantiati on programs consideri ng associated humidityinduced effects.

(iv) The wet glass transition temperature, T g , 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 r ubbery material takesAC 29-2C 12/01/2011

12 place over a narrow temperatur e range. A composite materi al will experience a drastic

reduction in matrix-controlled mechanical mate rial properties when loaded in this temperature range. Since the resin is the cr itical structural element in a composite

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

T g margin methodology of CMH-17 should be implemented (i.e., the T g should be 50° F higher than the maximum structur al temperature (see definition)). For any type of resin or adhesive, an acceptable temperature ma rgin using CMH-17 techniques (e.g., consideration of limited high temperature excursions) or equivalent methodologies based on tests or experience, or both, should be establis hed and approved early in the certification process.

(v) Local design values should be established by analysis and characterization tests and approved for specific st ructural configurations (point designs),

which include the effects of stress ri sers (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. Thes e transitional load transfer areas typically produce the highest stresses (and strains) and serve as the initiation sites for many of the failures (including t hose 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 s ubcomponent 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 ar ticle 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 im pact 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 desi gn strain level for each full-scale article. The maximum allowable design strain values selected should also take into account the reliability and confidence levels establishe d for the relevant portions of the quality control system. This methodology is necessa ry 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 spec ified 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 des ign approach is currently selected for nearly all civil and most military applications; since, under normal conditions, only visual inspections are required in the field (unl ess unusual external damage circumstances such as a hail storm occur) to maintain the initial level of airw orthiness (safety). However, many military app lications, because of thei r 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 repl acement of affected parts. To date, civil applicants have notAC 29-2C 12/01/2011

13 requested a flaw growth, phased NDI approach. Therefore, selection of the full-scale article’s design strain limit based on small-sca le tests for a no flaw growth design is extremely important.

(vii) Composite and adhesive pro perties should be determined so that detrimental structural creep does not occur under the sustained lo ads and environments expected in service. Small-scale charac terization tests (such as coupon, element, and subcomponent tests) and analysis, which ve rify and establish the full-scale design criteria and parameters necessary to ensur e that detrimental structural creep in

full-scale structure does not occur in servic e, should be conducted early in certification and should be FAA-approved.

(viii) Material allowable strength val ues for full-scale design and testing should be developed using the coupon procedures pres ented in CMH-17 or equivalent. The intent is to represent the ma terial variability including t he effects that can occur in multiple batches of material and process r uns. At least three batches of material samples should be used in material allow able strength testing. Company standards should be prepared, eval uated 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, ultrav iolet radiation, chemical effe cts, 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 compos ite structure. Usually, a threat analysis is conducted early in the certification proc ess that identifies the various threats and threat levels for which protection must be provided. This dat a 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, “Protecti on of Aircraft Fuel Systems Against Fuel Vapor Ignition Caused by Lightning” and FAA Report DOT/FAA/CT-86/8. For composite structure projects involving rotorc raft certificated to earlie r certification bases (which do not automatically include the li ghtning protection require ments of § 29.610), these requirements should be imposed as spec ial conditions. The design should be reviewed early in the certification process to ensure proper protection is present. The substantiation test program should also be es tablished, 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 compositeAC 29-2C 12/01/2011

14 design should consider all critical load cases and associated failure modes, including effects of environment, material and proce ss variability, and defects or service damage

that are not detectable or a llowed by the quality control, manufacturing acceptance criteria, or maintenance documents of the end produc t. The static strength demonstration should include a program of component ultimate load tests, unless experience exists to demonstrate t he 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, ma terial 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 s upported 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 fo r prior repeated loading or environmental exposure for structural substantiation exist.

 In the first approach, the large-sca le static test should be conducted on structure with prior repeated loadi ng 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 fo r in the static strength demonstration test (e.g., load enhancement), or in the analysis of these results (e.g., showing a positive ma rgin 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 ma y 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 t he composite structure should be statistically established, incrementally, through a pr ogram of analysis and tests at the coupon, element, subcomponent, or component levels. As part of the evaluation, build ing block tests and analyses at the coupon, element, or subcom ponent levels can be used to address the issues of variability, environmen t, 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 te sts included in the building block approach. The material stress-strain cu rve should be clearly established, at least through the ultimate design load, for each com posite design. As shown in Figure AC 29.573-1, the large quant ity of tests needed to provide a st atistical basis comes from theAC 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 compo nent tests. The static strength substantiation program should also consider all cr itical loading conditions for all critical structure including residual st rength and stiffness requirements after a predetermined length of service (e.g., end of life (EOL)), which takes into account damage and other degradat ion due to the service period.AC 29-2C 12/01/2011

16 Figure AC 29.573-1. Sc hematic 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, la minate, or specific design feature level (e.g., filled hole, lap joint, st ringer run-out), provided they a ccurately reflect the actual value and variability of the st ructural strength for the cr itical failure modes being considered, at each point design where margins need to be established.

(iv) The static test articles shoul d be fabricated and assembled in accordance with production specifications and processe s so that they ar e representative of production structure including defects consistent with the limits established by manufacturing acceptance criteria.

(v) The material and processing variabi lity of the composit e structure should be considered in the static strength s ubstantiation. This can be achieved by establishing sufficient process and quality cont rols to manufacture structure and reliably substantiate the required str ength in tests and analyses, which support a building block approach. If sufficient process and quality co ntrols cannot be achieved, it may be necessary to account for greater variability wi th 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 impac t damage (or other minor discrete source damage) that can be realistically expected from manufacturing and service, but not

more than the established threshold of det ectability for the selected inspection procedure, will not reduce the st ructural strength below ultima te load capability. This static strength capability can be shown by anal ysis supported by test evidence, or by a combination of tests at the coupon, el ement, 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 streng th determination required by §§ 29.305 and 29.307 establishes the ultimate load capabil ity for composite structures that are manufactured, operated, and maintained with established procedures and conditions. The damage tolerance and fatigue evaluat ion required by § 29.573 mandates procedures that allow the co mposite structure to retain the intended ultimate load capability when subjected to expected fa tigue loads and conditions during its operational life. The requirements estab lished for the damage tolerance and fatigue evaluation include component replacement ti mes, inspection intervals, or other procedures as necessary to avoid catastrophic failure. These evaluations assume that the baseline ultimate strength capability mi ght be compromised by damage caused by fatigue, environmental effects, intrinsic or di screte flaws, or accidental damage. This damage includes flaws or defects, which ma y occur in manufacturing or maintenance

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

18 manufacturing acceptance criteria. T he damage tolerance assessment establishes standards that allow the static strength capab ility to degrade below the ultimate strength capability assuming such damage occurs within the operational life of the structure. However, when this damage o ccurs, the remaining structur e must withstand expected

loads without failure or ex cessive structural deformati ons until the damage is detected and the component is either repaired to restor e ultimate strength capability or retired.

(ii) General. The nature and extent of the required ana lysis or tests on complete structures and portions of the primary structure can be based on applicable previous fatigue or damage tolerant designs, c onstruction, 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 perform ed. The following considerations are unique to the use of composite materi al systems and should be observed for the

method of substantiation se lected by the applicant.

Rotorcraft structure provides a broad range of composite applicatio ns that are quite

different in terms of functional ity, 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 appr oach, attention should be given to the composite application under ev aluation, the type of pot ential damage and degradation of the structural design details, the mate rials 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 includ e material imperfections whose extent is

not less than the limits estab lished 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 thes e imperfections should be discussed and agreed with the FAA, taking into account experience in manufacturing and routine in-service inspections. Typical defects to be consider ed include but are not limited to the following:

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

(2 ) Delaminations, fiber wavi ness, porosity, voids.

(3 ) Scratches, gouges, and penetrations.

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

19 All of the damages identifi ed in the preceding paragraph (B) above should be derived

from the threat assessment descri bed in the following paragraph (C).

(C) For each PSE, a threat a ssessment must be made of the probable locations, types, and sizes of damage consid ering fatigue, environmental effects, intrinsic and discrete flaws, and impact or ot her accidental damage. This determination must be submitted with accompanying ratio nale to the FAA/AUTHORITY for approval. This rationale may include experience with similar materials, designs, processes (manufacturing, maintenance, and overhaul), st ructural 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 surv eys, 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 mate rials and manufacturing processes to mitigate intrinsic flaws (inclusions and defects).

- Preclude a type of damage by use of a sp ecific design feature (material selection, surface treatment, protective coating, or sh ielding), a specific stress level (for fatigue damage), or a specific manufactu ring inspection process (if it can be shown to be highly reliable, well-controlled, document ed, and systematically required).

The assessment should include:

 A systematic evaluation of all the locati on, 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 me thod 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 deta ils 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 es tablished under the manufacturing inspection and acceptance criteria and are expected to remain in service for the lif e 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/orAC 29-2C 12/01/2011

20 service experiences. This type of dam age can include dents, penetrations, gouges, abrasions, and scratches. A threat asse ssment is needed to identify impact damage severity and detectability for design and ma intenance. A threat assessment usually includes damage data collected fr om service plus an impact su rvey. An impact survey consists of impact tests performed with conf igured structure, which is subjected to boundary conditions characteristic of the stru cture. Many differe nt impact scenarios and locations are typically considered in the surv ey, which has a goal of identifying the most critical impacts (i.e., thos e causing the most serious dam age but are least detectable). When simulating accidental impact damage, blunt or sharp impactors should be selected to represent the maxi mum criticality versus detect ability, according to the load conditions (e.g., tension, compression or s hear). Until sufficient service experience exists to make good engine ering judgments on energy and impactor variables, impact surveys should consider a wide range of c onceivable impacts, including runway or ground debris, hail, tool drops, and vehicle coll isions. Service data collected over time, can better define impact survey s and design criteria for subs equent products, as well as establish more rational inspection interval s 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 co mpletion of a flight on wh ich damage resulting from obvious discrete source occurs (e.g., hail damage, bird st rike, 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 struct ures (see AC 21-26). Assuming good process and quality controls, service history has s hown 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 equiva lent, approved techniques), then the limit load capability should be ensur ed by any or a combi nation 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 s hould be up to the limitations provided by redundant design features (i.e., mechanical fa steners 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 sa tisfactory service history of like or similar components.

(E) The fatigue load spectrum de veloped for fatigue testing and analysis purposes should be representative of the anticipated service usage. Low amplitude

load levels that can be shown not to c ontribute to fatigue damage may be omitted (truncated). Reducing maximum load leve ls (clipping) is generally not accepted.

(F) Environmental effects (temper ature and humidity re presentative of the expected service usage) on the static and fatigue behavior and damage growth should

be considered. Unless tested in the env ironment, 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 ta ke 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 t he damage tolerance and fatigue evaluation.AC 29-2C 12/01/2011

22 Detectability

Energ y 

## ° +

+ +

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 servic e

. Obvious Damage

Barely Detectable Impact
Damag e Detectable damag e



## °

° Undetectable damage 

## +

+

Zone 1 Zone 3 Zone 2

Figure AC 29.573-2. Characte rization of Impact Damage.

(1 ) Both the energy level associ ated 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 (B DID) is the state of damage at the threshold of detectability for the approv ed inspection procedure. Barely Visible Impact Damage (BVID) is that threshold of visually det ectable damage associated with a detailed visual inspection procedure.

(4 ) Detectable Damage is the st ate 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 dam age is not detect able, Ultimate Load capability is required. The pr ovisions of paragraph f.(5) prov ide 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 sa fety. For example, residual strength lower than ultimate may be used in association wi th 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 resi dual strength requirement that can vary with alternate procedures or the probability of damage occurrenc e or both. In either case, any compromise for re sidual 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 complianc e, as described in the following section, f.(6)(iii)(B).

One example of the use of alternate procedur es is for the rare damage threat from a high energy, blunt impact (e.g., service v ehicle 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 abov e limit load capability to ensure safety without detection for long periods of time). If it is furthe r 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 in spection 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 eac h PSE, inspections, replacement times, or other procedures must be established as nece ssary 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 subs equently. Generally, replacement times are established using Damage Tolerance Safe Li fe Evaluations and Inspection Intervals are established using Fail Safe Evaluations. From current state- of-the-art rotorcraft applications, it is widely accepted that co mposite materials have good flaw and damage tolerance capabilities and ther efore the supplemental procedures may only be rarely necessary. Damage tolerance evaluations are best suited for composite structures, particularly those with struct ural redundancy and inherent resistance to damage growth. Damage resulting from anomalous or accident al events must be considered in theAC 29-2C 12/01/2011

24 damage tolerant evaluations. The damage tole rant evaluation for replacement times and inspection intervals is to be used unle ss 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 chos en so that the structure is protected against catastrophic failure from each of the th reats identified in paragraph f.(6)(ii)(C) of this AC by a specific procedure (inspection, replacement time, or ot her procedure). For example, a manufacturing-re lated 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 coul d be substantiated by an inspection method with no specific replacement time. The result could be one structure with several different inspection requirements (locat ion, method, and interval) and a fixed replacement time as well. This combinatio n of procedures assures that each threat is

covered.

The fatigue substantiation should include su fficient coupon, element, sub-element, or component tests to establish the fatigue sca tter, curve shapes, and the environmental

effects. The substantiation should incl ude 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-da maging. The substantiation should include a static strength evaluation to show t hat the required resi dual strength and adequate stiffness, accounting for the effects of environment, are retained for the life of the structure or the appropriate inspection in terval. Damage as determined in paragraph f.(6)(ii) of this AC for the specific structure being s ubstantiated 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 presen t, is able to withstand repeated loads of variable magnitude without detectable damage gr owth 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 require ments are generated from the test program in this method. However, compliance with routine inspection s for cracking, delaminations, and service damage and other limitations prescribed in accordanc e 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 fo r the damage present at critical locations in the structure. AC 27-1B, AC 27 MG 11, provides guidance t hat may be appropriate for this method.

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

25 specimens subjected to const ant-amplitude or spectrum load ing applied in a distribution on the structure that is repres entative of critical flight c onditions. Any indication of growth of the imposed damage and defects, or structurally significant cracking, disbonding, splintering, or delaminating of t he 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 effect s. Flight loads are compared to this working curve, and if any intercepts occu r, a cumulative damage calculation is conducted to establish the component retire ment time. Compliance with the ultimate load requirements should be de monstrated at the comple tion of the fatigue test.

(2 ) Life-Test Method . This method uses spectrum fatigue testing to verify the absence of damage growth over a la rge 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 Gr ound-Air-Ground (GAG) loads. Fatigue test loads should be increased by fact ors 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 grow th of the imposed damage , or structurally significant cracking, disbonding, splintering, or delamination of t he composite during the

test constitutes failure to achieve the desir ed lifetime. Howeve r, the equivalent life demonstrated at the time of inception of damage growth or cracking can be used as a retirement time for the component. Compli ance with the ultimate load requirements

should be demonstrated at the co mpletion 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 parti al failure is able to withsta nd design limit loads without

failure or excessive structural deformations wit hin a specified inspection interval. If the damage is detected in an inspection, the struct ure should be either replaced or repaired to restore ultimate load capability. Evaluation of Zone 3 damage should have sufficiently high residual str ength and, if necessary, supplem ental procedures should be

established to minimize the risk of catast rophic failure. Full-scale, component, or subcomponent testing should be accomplished usi ng one or more specimens subjected to constant amplitude or spectrum loading appli ed in a manner represent ative of flight load

conditions. The test loads should be increas ed 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 subs tantiated using eith er 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 structur es, should be tested under repeated loads for validating a no-gro wth approach to the damage tolerance requirements. The number of cycles applied to validate a no-growth concept should beAC 29-2C 12/01/2011

26 statistically significant, and may be determined by load or life consider ations or both. Residual strength testing or evaluations s hould be performed after repeated load cycling demonstrating that the residual strength of the stru cture is equal to or greater than limit load considered as ultimate. Moreover, it should be shown that st iffness properties have not changed beyond acceptable levels . Inspection intervals should be established, considering the residual str ength capability associat ed with the assumed damage. The intent of this is to assure t hat structure is not exposed to an excessive period of time with static marg ins less than ultimate, providin g a lower safety level than in the typical slow growth situation, as illu strated 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. Resi dual Strength vs. Time.

The lower the residual strength of a stru cture after an accident al damage event, the shorter the inspection interval should be. Considerations of both inspectability and impact surveys (including probability of occu rrence) for specific structure may be used

to isolate the most critical threats to c onsider in setting a maintenance inspection interval. Knowledge of the residual strengt h 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 appl ied 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 fr equency, extent, and methods of inspection for inclusion in the maintenance plan. Inspection intervals should be established so that the damageAC 29-2C 12/01/2011

27 will have a very high probability of detecti on 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 fo r 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 Resi dual 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 mechanica lly arrested or term inated 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 strengt h testing or evaluation s hould be performed after repeated load cycling and a demonstration that the residual strengt h of the structure is equal to or greater than limit load considered as ulti mate. Moreover, it should be shown that stiffness properties have not changed beyond acceptable levels. Inspection intervalsAC 29-2C 12/01/2011

28 should be established, considering the resi dual strength capability a ssociated 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 ma rgins less than ultimate, providing a lower safety level than in the typical slow gr owth situation, as illustrated by Figure AC 29.573-3. For any damage si ze that reduces the load c apability below ultimate, the component is either repaired to restor e ultimate load capability or replaced.

The lower the residual strength of a structur e after an arrested growth event, the shorter the inspection interval should be. Consi derations of both inspectability and impact surveys (including probability of occurrence) for specific structure may be used to

isolate the most critical thr eats to consider in setting a ma intenance 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 ma y be applied in establishing the inspection interval.

(C) Combination of Damage Toler ant Safe Life and Fail Safe Evaluations. Generally, it may be appropriate to establis h 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 t he replacement time or inspection interval

methods.

One possible alternate approach is the us e 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). T hese 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/AU THORITY per § 29.573(e). The applicant must establish that the damage tolerance criteria are imprac ticable and cannot be satisfied for the specific PSE, locations, and threats considered. In addition, the types of damage considered in the evaluations mu st 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 applic ation of methods of fatigue and damage

tolerance evaluations indicates that a relev ant test background should exist in order to achieve the design objective. It is the gener al 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 pr oper flight and ground loads.

(B) Identificat ion of the structure to be considered in each evaluation (a failure mode and effects analysis or similar method should be used).

(1 ) Identification of Principa l Structural Elements. Principal structural elements are those that contribute signific antly to carrying flight and ground loads and whose failure could result in catastrophic failu re 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 s ubject to repeated loading, including control rods, servo structure, and swashplates.

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

(v ) Fuselage, including stabilizers and auxiliary lifting surfaces, airframe provisions for engi ne and transmission mountings.

(vi ) Main fixed or retractable l anding 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 t han the complete structure is tested, care

should be taken to ensure that the inter nal loads and boundary conditions are valid. The following should be considered:

(i ) strain gauge data on undamaged struct ure 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 deforma tion 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 env ironment, or a high

maintenance environment s hould 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 m aneuvering 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, inspecti on methods, and detectable flaw sizes.

(G) Variability of the measur ed stresses of paragraph f.(6)(iv)(C), the actual flight condition occurrences of par agraph 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 stra in 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 conditionsAC 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 cr itical 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 roto rcraft as identified in § 29.573(d)(1). Other devices such as accelerometers may be used as appropriate.

(2 ) The corresponding flight parameter s (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 s hould be adequately calibrated and checked periodically throughout the flight strain meas urement program to ensure

consistent and accurate results.

(C) Parts to be Strain-Gauged. Fatigue critical portions of the rotor systems, control systems, landing gear, fuse lage, 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 fo r similar designs. In any event, the flight conditions considered appropriate fo r the design and application should be representative of the actual oper ation in accordance with the rotorcraft flight manual. In the case of multiengine roto rcraft, the flight condition s concerning partial engine-out operation should be considered in addition to co mplete power-off operation. The flight conditions to be investigated should be submi tted in connection with the flight evaluation

program.

(3 ) The severity of the maneuvers inve stigated during the flight strain survey should be at least as severe as t he 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 st ress levels associated with each flight condition. The temperature effe cts on loads as affected by elastomeric components are to be investigated. To a ccount for data scatter and to determine theAC 29-2C 12/01/2011

32 stress levels present, a sufficient amount of data points should be obtained at each flight condition. Considerat ion 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 maneuver s can be based on past experience with similar design. This procedure is acceptabl e where adequate flight tests are performed

to substantiate such selections. The combin ations of flight param eters that produce the

most critical stress levels s hould be used in the evaluation.

(vi) Frequency of loading.

(A) Types of Operation.

(1 ) The probable types of operation (t ransport, utility, etc.) for the rotorcraft should be established. The ty pe of operation can have a major influence on the loading environment. In t he 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 manufactu rer. 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 lim itations by flight cycle limitations (if properly defined and easily identifiable by t he crew, for example: one landing, one load transportation). A special type of flight hour limitation replacement us ing 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 c ontinuing 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 shoul d state in the limitations

section that the intended use of the rotorc raft does not include certain missions or repeated maneuvers (e.g., logging with its hi gh number of takeoffs and landings per hour). A note to this effect should also appear in the rotorcraft airworthiness limitationsAC 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 safe ty, a reassessment made. If this reassessment indicates the necessity for re vised retirement times, thos e new times may be limited to specific rotorcraft model invo lved in the added mission provided:

(i ) changes are adopted through the airworthiness directives process and proper part re-identific ation 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 limita tion section supplement is approved.

(B) Loading Spectrum. The spec trum allocating percentage of time or frequencies of occurrence to flight condition s or maneuvers is to be based on the expected usage of the rotorc raft. This spectrum is to be es tablished 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 recor der data, design limitations established in compliance with § 29.309, and recommended operating conditions and limitations specified in the rotorcraft flight manual or instructions for c ontinued airworthiness (ICA).

The distribution of times at various forwar d flight speeds should reflect not only the relation of these speeds to V NE but also the recommended oper ating conditions in the rotorcraft flight manual or ICA that govern V
c 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, t hus obtaining directly the stress or load spectra for structural elements.

(7) The seventh major area is the dynamic lo ading and response requirements of §§ 29.241, 29.251, and 29.629 for vibration and resonance frequency determination and separation for aeroelastic st ability and stability margin determination for dynamically critical flight structure. Critical parts, locations, exci tation modes, and separations

should be identified and subst antiated. This substantiation should consist of analysis supported by tests, including tests that account for repeated loading effects and environment exposure effects on critical pr operties, such as stiffness, mass, and damping. This must be accomplished to assu re that the initial stiffness, residualAC 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 st ructures must be identified and properly

evaluated. This structure must be show n by analysis to be flutter free to 1.1 V NE (or any other critical operating limit, such as V D , for a VSTOL aircraft) with the extent of damage for which residual strength and stiffness are demonstrated.

(iii) Where appropriate, crash impa ct 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 repai r and continued airworthiness requirements of §§ 29.611, 29.1529, and 14 CF R part 29 Appendix A, for composite structures. When repair and continued airw orthiness procedures are provided in service documents (including approved sect ions of the maintenance manual or

instructions for continued airworthine ss), the resulting repairs and maintenance provisions should be shown to provide struct ure, which continually meets the guidance of paragraphs (1) through (7) of this AC. All certification-bas ed repair and continued airworthiness standards, limits, and inspec tions must be clearly stated, and their provisions and limitations cl early documented to ensure continued airworthiness. No composite structural repair should be a ttempted that is bey ond the scope of the applicable approved Structural Repair Manu al (SRM) without an engineering design approval by a qualified FAA/AUTHOR ITY designated representative.