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CHAPTER 3

AIRWORTHINESS STANDARDS

TRANSPORT CATEGORY ROTORCRAFT

MISCELLANEOUS GUIDANCE (MG)

AC 29 MG 8. (Amendment 29-35) SUBSTANTIATION OF COMPOSITE

ROTORCRAFT STRUCTURE.

a. Reference FAR Sections. §§ 29.305, .307, .571, .603, .605, .609, .610, .611,

.613, .629, .923, .927, .931, .1529 and Appendix A.

b. Purpose. These substantiation procedures provide a more specialized

supplement to the general procedures outlined by AC 20-107A, “Composite Aircraft

Structure.” These procedures address 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, some variations and

deviations from the procedures described herein will be both necessary and acceptable.

Significant deviations from this material should be coordinated in advance with the

Rotorcraft Directorate.

c. 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 is 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.

d. 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 for metals is significantly different from fixed wing fatigue substantiation.

Since AC 20-107A, 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 20-107A 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 full-scale specimen fatigue tests for

rotorcraft programs rather than just one full-scale specimen test. Also, constant
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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 (reference Appendix 1 of this AC, “Fatigue Evaluation of

Transport Category Rotorcraft Structure (Including Flaw Tolerance)”). 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 specimens (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 Appendix 1 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., MIL-HDBK-5 for metals, MIL-HDBK-17B for

composites, or 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

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
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FAR’s 27 and 29 contain many sections (reference paragraph a.) to be considered in

substantiating composite rotorcraft structure, but this advisory material is needed to

supplement the general guidance of AC 20-107A by providing specific rotorcraft

guidance for obtaining consistent compliance with FAR sections applicable to rotorcraft.

e. Definitions. The following basic definitions are provided as a convenient

reading reference. MIL-HDBK-17, and other sources, contain more complete glossaries

of definitions.

(1) AUTOCLAVE. A closed apparatus usually equipped with variable

conditions of vacuum, pressure and temperature. Used for bonding, compressing or

curing materials.

(2) ALLOWABLES. Both A- basis and B- basis values statistically derived and

used for a particular composite design.

(3) BALANCED LAMINATE. A composite laminate in which all laminae at

angles other than 0° occur only in ± pairs (not necessarily adjacent).

(4) 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 of 95 percent.

(5) 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 of 95 percent.

(6) BOND. The adhesion of one surface to another, with or without the use of

an adhesive as a bonding agent.

(7) 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.

(8) 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 catalyst, and

with or without heat.

(9) DELAMINATION. The separation of the layers of material in a laminate.

(10) DISBOND. A lack of proper adhesion in a bonded joint. This may be local

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.
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(11) 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.

(12) 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.

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

(14) 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.

(15) GLASS TRANSITION TEMPERATURE. The approximate midpoint of the

temperature range over which the glass transition takes place.

(16) HYBRID. Any mixture of fiber types (i.e., graphite and glass).

(17) IMPREGNATE. An application of resin onto fibers or fabrics by several

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

(18) LAMINA. A single ply or layer in a laminate in which all fibers have the

same fiber orientation.

(19) LAMINATE. A product made by bonding together two or more layers or

laminae of material or materials.

(20) 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.

(21) MATERIAL SYSTEM CONSTITUENT. A single constituent (ingredient)

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

(22) MATERIAL SYSTEM. The combination of single constituents chosen

(e.g., fiber and resin).

(23) MATRIX. The essentially homogeneous material in which the fibers or

filaments of a composite are embedded. The resins used in most aircraft structure are

thermoset polymers.

(24) 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
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critical load case has an associated maximum structural temperature). This term is

synonymous with the term “maximum panel temperature.”

(25) POROSITY. A condition of trapped pockets of air, gas, or void within a

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

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

(26) 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 which will involve complete

curing. Prepreg is usually drapable, tacky and can be easily handled.

(27) RESIN. An organic material with indefinite and usually high molecular

weight and no sharp melting point.

(28) RESIN CONTENT. The amount of matrix present in a composite either by

percent weight or percent volume.

(29) 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 COCURING).

(30) SHELF LIFE. The length of time a material, substance, product, or

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

all applicable specification requirements and/or remain suitable for its intended function.

(31) 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, etc.). Strain level is generally measured in thousandths of an inch per

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

(32) SYMMETRICAL LAMINATE. A composite laminate in which the ply

orientation is symmetrical about the laminate midplane.

(33) TAPE. Hot melt impregnated fibers forming unidirectional pre-preg.

(34) 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.
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(35) 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).

(36) WARP. Yarns extended along the length of the fabric (in the 0° direction)

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

(37) WORK LIFE. The period during which a compound, after mixing with a

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

use.

f. RELATED REGULATORY AND GUIDANCE MATERIAL.

Document Title

(1) AC 20-95 “Fatigue Evaluation of Rotorcraft Structure”

(2) AC 20-107 “Composite Aircraft Structure”

(3) AC 21-26 “Quality Control for the Manufacture of

Composite Materials”

(4) MIL-HDBK-17 “Polymer Matrix Composites Volume 1:

Guidelines”

g. 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: (1) fabrication requirements; (2) basic constituent, pre-preg and laminate

material acceptance requirements and material property determination requirements;

(3) protection of structure; (4) lightning protection; (5) static strength evaluation;

(6) damage tolerance and fatigue evaluation; (7) dynamic loading and response

evaluation; and (8) special repair and continued airworthiness requirements. Original as

well as alternate or substitute material system constituents (e.g., fibers, resins, etc.),

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

(laminates, cocured assemblies, bonded assemblies, etc.) 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 strongly recommended that a FAA/AUTHORITY

certification team approach be used for composite structural substantiation. The team

should consist of FAA/AUTHORITY engineering, the MIDO inspector(s), the associated

Designated Engineering Representatives (DER’s), the associated Designated

Manufacturing Inspection Representatives (DMIR’s), and cognizant members of the

applicant’s organization. Personnel who are composites specialists (or are otherwise
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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.

(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.

(ii) Specific allowable defect limits on, for example, fiber waviness, warp

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

debonds, and delaminations, etc., 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 for defect detection

created, validated, and approved. Each critical engineering design should consider the

worse-case effects of the manufacturing process (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, reference paragraph g(5))

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

provided and FAA/AUTHORITY 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/AUTHORITY certification team members (both the

manufacturing and inspection district office (MIDO) and FAA/AUTHORITY 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

MIL-HDBK-17B, Volume I, Section 2.3.2. These methods can also be used, periodically

for qualification status renewals of existing material systems and suppliers.
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(B) Receiving Inspection and In Process Inspection. Applicants should be

able to demonstrate to FAA/AUTHORITY certification team members (both MIDO and

engineering), at any time, that their receiving and in-process quality control systems

provides products which continuously meet approved material and process

specifications. Quality systems should be designed with appropriate checks and

balances, such 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/AUTHORITY-witnessed.

(C) Material System Component Storage and Handling. Applicants

should be able to demonstrate to FAA/AUTHORITY 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, such that the necessary

statistical reliability and confidence levels for the items being inspected (which 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, etc.) 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/AUTHORITY-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.

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

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

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

presented and FAA/AUTHORITY-approved well in advance of the change’s production

effectivity.

(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
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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” & “B” basis allowable strength values and other basic

material properties (based on MIL-HDBK-17, or equivalent) 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 MIL-HDBK-17 has not completed development of

“B” basis allowables for inclusion in the handbook. 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 which is

approved and used should meet or exceed related MIL-HDBK-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, etc.) and constituents (e.g., fibers, resins, etc.) should be

identified, documented, and approved. These requirements, once approved, should be

placed in all appropriate procedures and specifications (such as those in

paragraph (g)(1) above).

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

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

approved methods or per FAA/AUTHORITY 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 humidity induced effects.

(iv) The 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. In actuality, the change from a hard polymeric material to a rubbery material

takes 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 (matrix) is the critical structural constituent in a

composite and since Tg exceedance is critical to structural integrity; Tg determination is

necessary. The Tg margin methodology of MIL-HDBK-17, Section 2.2.2.1, should be

implemented, i.e., the wet glass transition temperature (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 MIL-HDBK-17 techniques (e.g., consideration

of limited high temperature excursions) or equivalent methodologies based on tests

and/or experience should be established and approved early in the certification process.
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In no case should structural strength be degraded below limit load capability on a

maximum world wide high temperature day.

(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, etc.) and structural

discontinuities (e.g., joints, splices, etc.). 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 nucleation

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

scales 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 and approved such that specified impact damage considerations

are defined and properly limited. The effects of the approved strain levels should 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 applications have not been presented that desire 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 seen to be extremely important.

(vii) Composite and adhesive properties should be determined such 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
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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/AUTHORITY-approved.

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

should be developed using the coupon procedures presented in MIL-HDBK-17 or

equivalent. At least three batches of material samples should be used in material

allowable strength testing. Company standards should be prepared, evaluated and

FAA/AUTHORITY approved early in certification (as part of the building block process),

that reflect the material property determination considerations recommended in

MIL-HDBK-17 on a equal to or better than basis.

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

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

weathering, abrasion, fretting, hail, ultraviolet radiation, chemical effects, accidental

damage, etc.) 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

hazard analysis is conducted early in certification which 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 20-53A and FAA Report DOT/FAA/CT-86/8. For composite

structure projects involving rotorcraft certified 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 certification 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. Only conservative proven methods of static

analysis and failure criteria should be employed. The material stress-strain curve

should be clearly established, at least through the ultimate design load, for each

composite design. Composite structure should be statistically demonstrated,

incrementally, through a program of analysis, coupon tests, minor component ultimate

load tests and major component ultimate load tests. The static strength substantiation

program should 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
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the service period). Analytical reports and tests should consider all possible failure

modes and should include the critical, allowable effects of:

(i) Environment (reference paragraphs AC 29 MG 8 g(2) and g(3)).

(ii) Service Life (residual limit strength and stiffness demonstration).

(iii) Load path loss (fail-safe analysis and limit strength demonstration).

(iv) The standard fabrication process and its variability.

(v) Impact damage expected during service up to the established

threshold of detectability of the field inspection methods to be employed.

(vi) Point design and structural discontinuity considerations (e.g., stress

risers, joints, etc.).

(vii) Unless the ultimate strength of each critical bonded joint can be

reliably substantiated in production by NDI techniques (or other equivalent, approved

techniques), then limit load capability must be guaranteed by either of the following or a

combination thereof:

(A) The maximum disbond of each critical bonded joint which will carry

limit load is established by test, analysis, or both. Disbonds greater than these values

are typically prevented by design features.

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

tested to the critical limit load.

(viii) For static strength analysis laminae and laminate “A” and “B” basis

allowables (determined in accordance with paragraph g(2)) should be used subject to

the following conditions unless lower material properties are required by point design

considerations (e.g., stress risers, joints, etc.) stiffness requirements (e.g., flutter or

vibration margins), fatigue strength (including damage tolerance), or other overriding

considerations.

(A) When applied loads are distributed through a single load path or

single member within an assembly, the failure of which would result in the loss of the

structural integrity of the component involved or inability of the rotorcraft structure to

carry limit load, the part should be designed, analyzed, and tested using “A” basis

allowables.

(B) Redundant (fail-safe) structures in which the failure of individual

elements would result in applied loads being safely redistributed to other load carrying

members without exceeding the limit load capability of the rotorcraft structure may be

designed, analyzed, and tested using “B” basis allowables.
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(6) The sixth area is the fatigue evaluation requirements of § 29.571. The

fatigue evaluation method for the rotorcraft being certified should consider damage

tolerance in accordance with AC 20-107A.

(i) The safe-life method for composite structure as defined in

AC 20-107A is a flaw tolerant safe-life method (e.g., the test specimens consider

inherent production flaws and impact damage (reference paragraph (7)(ii)).

(ii) Large area disbonds, weak bonds, delaminations, or other defects

should be considered in tests or be prevented or be limited by appropriate flaw tolerant

special design features and by special manufacturing, maintenance, and inspection

procedures. Special attention should be assigned to all pure bond lines

(reference paragraph (5)).

(iii) Non-fail-safe or partially fail-safe dynamic component structure, which

may employ bond lines as the only load path, should be designed to relatively small

previously approved values of elastic, ultimate strain for the material system utilized,

and should be subjected to full-scale S-N curve testing. Six or more specimens are

recommended, as part of the substantiation process. Where practical, flight-by-flight

spectrum testing should be used.

(iv) All critical safety of flight composite structure must be designed to be

flaw (damage) tolerant. Environment degradation and in-service damage critical values

are typically included in the flaw tolerance evaluation. All other key factors, such as

material selection, manufacturing, and quality assurance controls, and in-service

inspection and maintenance, as noted previously, are also to be accounted for.

(v) The fail-safe design features of the rotor heads and blade retention

systems, other critical primary composite structure, and point design features (e.g.,

bonded metal-to-composite joints) should be assessed and appropriate inspection

programs provided to prevent catastrophic failure from flaw/damage propagation.

(vi) The method of generating S-N curves using approved raw data should

be demonstrated, evaluated, and approved.

(vii) Any limited life items must be identified and placed in the

Airworthiness Limitations section of the maintenance manual in accordance with

§ 29.571.

(viii) Load spectra, load truncation methods and all other major aspects of

the fatigue evaluation are documented in test proposals and approved.

(ix) Flaw growth rates (from initial detectability to the established value for

residual strength) must be previously established and closely monitored during
AC 29-2C 9/30/99

Page MG 8 - 14

substantiation. This data should be used to establish special phased inspections and

maintenance intervals for critical structure, as required.

(7) The seventh major area is the dynamic loading and response requirements

of § 29.629 for vibration and resonance frequency determination and separation for

aeroelastic stability and stability margin determination for flutter critical flight structure.

Critical parts, locations, excitation modes, and separations are to be identified and

substantiated. This substantiation should consist of analysis supported by tests and

tests which account for repeated loading effects and environment exposure effects on

critical properties, such as stiffness, mass, and damping. Initial stiffness, residual

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 structure are identified

and properly substantiated.

(ii) All flutter-critical composite structure are identified and properly

substantiated. 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 FAR 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 must be shown to provide structure which continually meets the guidance of

paragraphs (1) through (7) of this AC paragraph. All certification based repair and

continued airworthiness standards, limits, and inspections must be clearly stated and

their provisions and limitations defined and documented to ensure continued

airworthiness. In general, no composite repair should be attempted which is out of

scope to repairs stated in an approved Structural Repair Manual (SRM) without an

engineering design approval by a qualified FAA/AUTHORITY representative (DER or

staff engineer). The following minimum criteria should be met in any acceptable

composite repair:

(i) The repair should be permanent.

(ii) The repair should restore the structure to the required strength and

stiffness.

(iii) The repair should restore all functional requirements.
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(iv) The repair should have a negligible weight penalty.

(v) The repair should be aerodynamically compatible.

(vi) The repair materials should be compatible in all essential aspects with

the parent materials.

In summary, primary composite structure is an especially critical structure that requires

a clearly defined, phased approval (building block) certification process. This process

should involve the entire project certification team from a project’s start to its finish so

that proper certification is continuously and ultimately achieved. Also, in some special

cases, involving new advanced state-of-the-art composite technology, an issue paper

may be necessary. However, in the majority of cases (using current composite

materials and design philosophy) the applicant’s acknowledged use of this advisory

material (as recorded in the type board minutes) should eliminate the need for a

separate issue paper.
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CHAPTER 3

AIRWORTHINESS STANDARDS

TRANSPORT CATEGORY ROTORCRAFT

MISCELLANEOUS GUIDANCE (MG)

AC 29 MG 11 FATIGUE EVALUATION OF TRANSPORT CATEGORY

ROTORCRAFT STRUCTURE (INCLUDING FLAW TOLERANCE)

a. PURPOSE. This advisory material provides an acceptable means of

compliance with the provisions of § 29.571 of the Federal Aviation Regulations (FAR)

dealing with the fatigue evaluation of transport category rotorcraft structure (AC 20-95,

Fatigue Evaluation of Rotorcraft Structure, May 18, 1976, applies to normal category

rotorcraft structure and older transport category rotorcraft). The fatigue evaluation

procedures outlined in paragraph AC 29 MG 11 are for guidance purposes only and are

neither mandatory nor regulatory in nature. Although a uniform approach to fatigue

evaluation is desirable, it is recognized that in such a complex problem, new design

features and methods of fabrication, new approaches to fatigue evaluation, and new

configurations may require variations and deviations from the procedures described

herein. It is recommended that major deviations from the procedures be coordinated

with the Rotorcraft Standards Staff, ASW-110, to assure national standardization.

b. SPECIAL CONSIDERATIONS. The structure of rotorcraft is subject to cyclic

stresses in practically every regime of flight. In addition, since rotorcraft are highly

maneuverable and capable of forward, rearward, sideward, vertical, and rotational flight,

operating limitations due to fatigue are possible in practically all flight situations.

Corrosion and other environmental damage are also common in rotorcraft operations.

For these reasons, special attention should be focused on the fatigue evaluation of

rotorcraft structure.

c. BACKGROUND.

(1) During recent years there have been significant state-of-the-art and industry

practice developments in the area of structural fatigue and fail-safe strength evaluations

of transport category rotorcraft. The advance in the state-of-the-art has resulted from

two primary programs: (1) the perfecting of production techniques for composite

construction, and (2) the damage tolerant design features required to meet the battle

damage requirements of military programs such as the Advanced Attack and Heavy Lift

Helicopter (AAH and HLH) programs.

(2) Recognizing that advances in state-of-the-art and industry practice

warranted changes to the existing fatigue requirements in Part 29, the regulatory

requirements of § 29.571 were substantially revised. The revision to § 29.571 requires

new guidance material containing compliance provisions related to the changes. Also,

paragraph AC 29 MG 11 supplements AC 20-95 for new transport category rotorcraft

and provides guidance material in the flaw tolerance area. General guidance material
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for flaw/damage tolerance of composite structures is provided in AC 20-107A,

Composite Aircraft Structure, April 25, 1984.

d. INTRODUCTION.

(1) Definitions.

(i) Fatigue Tolerance. The capability of structure to continue functioning

without catastrophic failure after being subjected to fatigue (repeated) loads expected

during operation of the rotorcraft. Fatigue tolerance may be achieved by safe-life

design, flaw tolerant safe-life design (enhanced safe life), fail-safe design considering

flaw growth, or a combination.

(A) Safe Life. The capability of pristine structure as shown by tests, or

analysis based on tests, not to sustain measurable cracks during the service life of the

rotorcraft or before an established replacement time. Special inspection intervals or

other special procedures are not usually prescribed for safe-life substantiations.

(Routine inspections for wear, fretting, corrosion, crack, and service damage as outlined

in § 29.1529 are, of course, appropriate.)

(B) Flaw Tolerant (Enhanced) Safe Life. The capability of flawed

structure as shown by tests or analysis based on tests to sustain, without measurable

flaw growth, the spectrum of operating loads expected during the service life of the

rotorcraft or during an established replacement time. Measurable flaw growth means

flaw growth beyond an acceptable threshold of detectability. Enhanced safe life may be

achieved by designing structure (single element or multiple element) that provides for

resistance to crack initiation from manufacturing or service flaws, by material selection,

by material processing, by limitation of stress levels, and by geometric design features.

The protection of structure against environmental damage and accidental mechanical

damage through the use of protective coatings may be used in determining initial flaw

types and sizes to be considered.”

(C) Fail-Safe Design Considering Flaw Growth. The capability of

rotorcraft structure to continue functioning without catastrophic failure after being

subjected to fatigue damage, corrosion, intrinsic flaws, or accidental damage expected

during fabrication and operation of the rotorcraft. The concept of fail-safety now

includes some discrete damage requirements and more explicitly requires consideration

of “life-remaining” after flaws occur until the flaws are detected using a prescribed

inspection plan or until the part is replaced. Fail-safe designs may include any of the

following features:

(1) 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.
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(i) Active Multiple Load Path. Structure providing two or more load paths

that are all loaded during operation to a similar load spectrum. The use of active

multiple load paths requires special attention to fatigue damage in the remaining

members after failure of a member.

(ii) 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.

(2) Flaw Arrest (Flaw Stopper) Feature. Structure that does not provide

completely separate and distinct load paths but does provide features of design such as

bonded and/or riveted straps, changes in geometry, or special processing techniques

such as rolling or coining to retard or arrest flaw growth.

(3) Slow Flaw Growth Feature. Structure (single element or multiple

element) that provides for slow flaw growth by material selection, material processing,

limitation of stress levels, geometrical design features, or by other methods.

NOTE: The detection of the crack or other flaw is an integral part of the flaw growth

method. The need to use complicated inspection techniques may not be practical in

some cases. See the discussion in paragraph AC 29 MG 11 g(1)(iii)(B).

(ii) Flaw. Structural imperfections in excess of type design allowances

for “as manufactured” or “pristine” structure.

(A) In Metallics.

(1) For Fail-Safe Crack Growth. Corner cracks for holes, semicircular

cracks for surfaces, realistic cracks for other locations.

(2) For Flaw Tolerance (Enhanced) Safe Life. Gouges, scratches,

corrosion, fretting, or wearing likely to occur during fabrication and operation of the

rotorcraft.

(B) In Nonmetallics. Flaws should be determined in accordance with

AC 20-107A guidance.

(iii) Limit Design Load. “The maximum loads to be expected in service,”

as defined by § 29.301(a), are considered as limit loads for new structure and as

ultimate loads for flaw tolerance residual strength demonstration purposes. The

residual strength after failure should equal or exceed these loads for flaw tolerance.

(iv) Damage Tolerance. No definition is provided in paragraph

AC 29 MG 11 for this expression because many definitions have been applied by the

civil and military communities. The expression has been used to describe design

features, structural systems (including design, manufacturing, and operating
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considerations), and substantiation techniques such as fracture mechanics analyses for

metallics. Paragraph AC 29 MG 11 uses expressions for those features of the general

damage tolerance concept used to provide and demonstrate fatigue tolerance of

rotorcraft structure including consideration of flaws. Tolerance to flaws rather than

tolerance to damage is specified in paragraph AC 29 MG 11 since flaws intrinsic to

certain manufacturing and fabrication processes are covered in addition to damage

from handling, corrosion, etc.

(2) Rotorcraft Flaw Tolerance. Flaw tolerant design as substantiated by

fail-safe flaw growth or flaw tolerant (enhanced) safe-life means outlined in § 29.571

and paragraph AC 29 MG 11 g is required, unless it entails such complications that an

effective flaw tolerant structure cannot be achieved within the limitations of geometry,

inspectability, or good design practice. Good design practice includes consideration of

component complexity, component weight, methods of production and component cost.

Under these circumstances, a design that complies with safe-life criteria should be

used. Typical examples of structure that might not be conducive to flaw tolerance

design are swashplates, main rotor shafts, push rods, small rotor head components

(i.e., devices, bolts, etc.), landing gear, and gearbox internal parts including bearings.

In addition, the need for the use of inspection techniques and equipment or highly

trained personnel--resources not available (for economic or other reasons) to the small

operator or in remote areas of operation--should be carefully considered (reference

paragraphs AC 29 MG 11 d(1)(i)(C)(3) and g(1)(iii)(B)).

(3) Test Background. Experience with the application of methods of fatigue

evaluation indicates that a relevant test background should exist in order to achieve the

design objective. Even under the flaw tolerance method discussed in

paragraph AC 29 MG 11g, it is the general practice within industry to conduct flaw

tolerance tests for design information and guidance purposes. Flaw location and flaw

growth data based on test results and service history of similar parts, if available,

should also be considered in establishing a recommended inspection program.

(4) Manufacturing Considerations. Assurance of structural adequacy also

includes manufacturing and fabrication in accordance with design requirements and

specifications, quality control to monitor compliance, and effective service inspection

procedures.

(5) Fatigue Tolerance Considerations. In the fatigue tolerance evaluation, the

following items should be considered:

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

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

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

representative steady and maneuvering operating conditions expected in service.
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Page MG 11 - 5

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

corresponding spectrum of loadings and stresses.

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

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

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

(vi) Variability of the measured stresses of

paragraph AC 29 MG 11 d(5)(ii), the actual flight condition occurrences of

paragraph AC 29 MG 11 d(5)(iii), and the fatigue strength material properties of

paragraph AC 29 MG 11 d(5)(iv).

e. FLIGHT STRAIN MEASUREMENT PROGRAM.

(1) General. Subsequent to design analysis, in which aircraft loads and

associated stresses are derived, the stress level and/or loads are to be verified by a

carefully controlled flight strain measurement program. (This guidance is similar to that

of AC 20-95.)

(2) Instrumentation.

(i) The instrumentation system used in the flight strain measurement

program should accurately measure and record the critical strains under test conditions

associated with normal operation and specific maneuvers. The location and distribution

of the strain gages should be based on a rational evaluation of the critical stress areas.

This may be accomplished by appropriate analytical means supplemented, when

deemed necessary, by strain sensitive coatings or photoelastic methods. The

distribution and number of strain gages should define the load spectrum adequately for

each part essential to the safe operation of the rotorcraft as identified in

§ 29.571(a)(1)(i). Other devices such as accelerometers may be used as appropriate.

(ii) The corresponding flight parameters (airspeed, rotor RPM, center of

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.

(iii) The instrumentation system should be adequately calibrated and

checked periodically throughout the flight strain measurement program to ensure

consistent and accurate results.

(3) 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.
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(4) Flight Regimes and Conditions to be Investigated.

(i) Typical flight and ground conditions to be investigated in the flight

strain measurement program are given in Attachment 1 to paragraph AC 29 MG 11.

(ii) 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 multiengined 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.

(iii) The severity of the maneuvers investigated during the flight strain

survey should be at least as severe as the maximum likely in service.

(iv) 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

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 fatigue evaluation.

f. FREQUENCY OF LOADING.

(1) Types of Operation.

(i) 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, rotorcraft 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:
9/30/99 AC 29-2C

Page MG 11 - 7

(A) Long flights to remote sites (low ground-air-ground cycles but high

cruise speeds).

(B) Typical, general types of operation.

(C) Short flights as used in logging operations.

(ii) One means is to substantiate for the most severe type of operation;

however, this method is not always economically feasible.

(iii) 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 rotorcraft.

(iv) Where one or more of the above 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 (i.e., logging with its high number of takeoffs/landings

per hour). A note to this effect should also appear in the rotorcraft airworthiness

limitations section of the maintenance manual prepared in accordance with § 29.1529.

(v) Should subsequent usage of the rotorcraft encompass a mission for

which the original structural substantiation did not account, 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 aircraft involved in the added mission provided--

(A) Proper part reidentification is established;

(B) A Rotorcraft Flight Manual (RFM) supplement outlining limitations is

approved;

(C) An airworthiness limitations section supplement is approved; or

(D) An appropriate combination of part reidentification, RFM supplement,

or airworthiness limitation section supplement is approved.

(2) 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 such that it is unlikely that

actual usage will subject the structure to damage beyond that associated with the
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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. 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 that govern Vc or

cruise speed. Where possible, it is desirable to conduct the flight strain-gage program

by simulating the usage as determined above, with continuous recording of stresses

and loads, thus obtaining directly the stress/load spectra for structural elements.

g. FATIGUE STRENGTH EVALUATION (INCLUDING ROTORCRAFT FLAW

TOLERANCE).

(1) General. A means should be established using the conventional safe-life

approach or another fatigue tolerant approach to control the airworthiness of principal

structural elements identified under § 29.571(a)(1)(i). While the conventional safe-life

approach is acceptable under certain circumstances as defined in § 29.571, the

enhanced safe-life and fail-safe flaw growth approaches are to be used unless shown

to be impractical as stated in § 29.571. A fatigue strength evaluation of structure

considering tolerance to flaws is intended to ensure that even when flaws are present

due to manufacturing of service operations, the structure will withstand service loads

without failure until the flawed parts are replaced or until the flaws (including resulting

fatigue cracks) are detected and appropriate action taken. Either of two types of fatigue

strength evaluation may be used for flaw tolerance substantiation: flaw tolerant

(enhanced) safe-life or fail-safe flaw growth methods. Flaw tolerant (enhanced)

safe-life includes the testing and analyses currently associated with safe-life

substantiation, plus consideration of flaws. Flaw growth methods include the testing

and analyses currently associated with damage tolerance assessment (DTA). Tests

are required to substantiate flaw propagation rates and residual strength. Either

method or a combination can be used to meet the requirements of § 29.571 for

tolerance to flaws. Flaw tolerance evaluation encompasses establishing the

components to be designed as flaw tolerant, defining the loading conditions and extent

of flaws for which the structure is to be designed, conducting structural tests and

analyses to substantiate that the design objectives have been achieved, and

establishing replacement times or establishing inspection programs as necessary to

assure detection of fatigue damage. On components predominantly loaded by

centrifugal force, care should be taken in selecting limit load to assure that it is the

maximum expected in service. Design features that should be used in attaining a flaw

tolerant structure are:

(i) Use of multi-path construction and the provision of crack stoppers to

limit the growth of cracks and to provide adequate residual strength.

(ii) Selection of materials and stress levels that preclude crack growth

from flaws or that provide a controlled slow rate of crack propagation combined with
9/30/99 AC 29-2C

Page MG 11 - 9

high residual strength after initiation of cracks. Tests are required to substantiate crack

propagation rates.

(iii) Design to permit detection of cracks and other flaws, including the use

of crack detection systems, in all critical structural elements before cracks can

propagate and become dangerous or result in appreciable strength loss and to permit

replacement or repair. Inspection means appropriate for flaw tolerant design follow:

(A) Routine Inspections. To support routine inspection programs, blind

areas should be avoided, where practical. Access panels and openings should be

considered early in design.

(B) Special Inspections. These inspections will generally result from test

results as well as the geometry of the design. Care should be given to special

inspection techniques to be used in the field. Inspection techniques requiring facilities

and resources beyond the capability of the small operator or not generally available in

remote-area operations traditionally associated with rotorcraft operations should not be

specified for field inspections. Conservative sizes for detectable cracks or other flaws

should be used. Sufficient interval inspections should be provided to detect cracks

before they grow from a detectable size to a size that reduces the remaining strength

below design limit strength. If special inspection techniques not commonly available

cannot be avoided, then use of enhanced safe-life substantiation techniques should be

used.

(C) Pressurized Chambers. This design feature may be used to detect

cracks that cause a chamber to lose its pressure (either positive or negative). The loss

of pressure can be indicated by gages, or dye may be used if it is shown to be a

dependable indicator. Care should be taken in the design of pressurized chambers and

their indicating systems to assure dependability. Undependable systems can cause an

inordinate number of false indications and maintenance problems.

(D) Vibration Generation. This characteristic should be considered both

from the aspect of vibrations giving indications of a failure and from the aspect of the

increased fatigue loading resulting from the vibrations.

(E) Noise Generation. If initial failure will result in a clear and

unmistakable noise that is sufficiently continuous and loud, this characteristic can be

used in achieving flaw tolerance without additional special inspections.

(F) Crack Detection Wire, Foil, etc. Detection wire may be used in areas

that are sufficiently well defined so that the wire can be properly located. This

technique is appropriate in areas otherwise difficult to inspect. The potential for false

readings and possible maintenance problems should be considered.

(G) Health Monitoring. Techniques such as vibration sensing and

analysis or real time oil analysis can be used to provide information on establishing

inspection time.
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(iv) Use of multiple element structures may be provided so that damage

or failure occurring in one element of the member will be confined to that element and

the remaining structure will still possess adequate load-carrying ability until the failed

element is discovered by inspection.

(v) Provisions to limit the probability of concurrent multiple damage,

particularly after long service, should be provided. These provisions should ensure

adequate independence of each failure mode of multi-path constructions. The use of

full-scale fatigue test articles are recommended in this evaluation. Examples of

concurrent multiple damage to be avoided are:

(A) Simultaneous failure or partial failure of multiple path discrete

elements working at similar stress levels.

(B) Failures or partial failures, in adjacent areas, due to redistribution of

loading following a failure of a single element.

(2) 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 rotor head).

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

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

structure.

(3) 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.

(i) The following should be considered:
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(A) Strain gage data on undamaged structure to establish points of high

stress concentration as well as the magnitude of the concentration;

(B) Locations where analysis shows high stress or low margins of safety;

(C) Locations where permanent deformation occurred in static tests;

(D) Locations of potential fatigue damage identified by fatigue analysis;

(E) Locations where the stresses in adjacent elements will be at a

maximum with an element in the location failed;

(F) Partial fracture locations in an element where high stress

concentrations are present in the residual structure;

(G) Locations where detection would be difficult;

(H) Design details that service experience of similarly designed

components indicates are prone to fatigue or other damage; and

(I) Components fabricated from materials of potentially low fracture

toughness or high flaw growth rate.

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

severe corrosive and/or fretting environment, a wear and/or galling environment, or a

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

past service experience.

(4) Extent of Flaws. Each particular design should be assessed to establish

appropriate damage criteria in relation to inspectability and flaw extension

characteristics. In any flaw determination, it is possible to establish the extent of flaws

in terms of detectability with the inspection techniques to be used, the associated single

element failure or initially detectable flaw size, the residual strength capabilities of the

structure, and the likely flaw extension rate (after either an element failure or a partial

failure) considering the expected stress redistribution under the repeated loads

expected in service and with the expected inspection frequency. Although

multiple-element design should be used where practical, and obvious partial failure

could be considered to be the extent of the flaw for residual strength assessment,

provided a positive determination is made that manufacturing or service flaws will not

initiate flaw growth within the life of the part or that the fatigue flaw growth will be

detectable by the available inspection techniques at a sufficiently early stage of the flaw

growth. In a swashplate or pin containing pressurized chamber, an obvious partial

failure might be detectable through the inability of the chamber to maintain pressure

after occurrence of the damage. Flaw tolerant (enhanced) safe-life evaluations from

flaws should consider flaws to be expected during manufacturing (including handling)

and during service. Special coatings against corrosion, flame plating or plasma plating
AC 29-2C 9/30/99

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against fretting corrosion, and energy absorption coating or shielding against damage

associated with maintenance may be used in determining the type and extent of flaws

to be considered in flaw tolerant (enhanced) safe-life evaluations. The following are

typical examples of the type of partial failures that may be considered in the flaw growth

fail-safe and enhanced safe-life evaluation. These may or may not be appropriate to

the design being considered. These examples have been sources of service difficulty

on prior/existing designs:

(i) Detectable skin cracks in the trailing edge sections of rotor blades.

(ii) Detectable failures of individual straps in “strap packs.”

(iii) Detectable skin cracks emanating from the edge of structural

openings or cutouts;

(iv) Detectable circumferential or longitudinal skin crack in the basic

fuselage or tail boom structure;

(v) Complete severance of interior frame elements or stiffeners in

addition to a detectable crack in the adjacent skin;

(vi) Presence of a detectable fatigue failure in at least the tension portion

of the spar web or similar element;

(vii) Detectable failure of a primary attachment, including blade

attachment fittings and control surface hinge and fittings; and

(viii) Fretting, corrosion, and galling conditions expected in service.

(5) Provisions for Inspection. The designer should strive to ensure adequate

inspectability of all structural parts to qualify them under the fail-safe flaw growth

provisions. In those cases where blind areas or surfaces exist, suitable design features

should be provided to allow inspection techniques (either visual or nondestructive

testing, as necessary) to assure adequate residual strength is achieved unless shown

to be impractical due to limitations of geometry and good design practice. In addition,

the alternate safe-life approach to fatigue tolerance should be implemented if the

inspection techniques are shown to be too complicated and impractical.

(6) Testing of Principal Structural Elements. The nature and extent of tests on

complete structures or on portions of the primary structure will depend upon applicable

previous design, construction, tests, and service experience in connection with similar

structures. For flaw tolerant safe-life testing, simulated flaws should be as

representative as possible of actual gouges, scratches, pitting, or fretting to be

expected in manufacture and service. For fail-safe testing considering crack

propagation, simulated cracks should be as representative as possible of actual fatigue

damage. Where it is not practical to produce actual fatigue cracks, flaws can be
9/30/99 AC 29-2C

Page MG 11 - 13

simulated by cuts made with a fine saw, sharp blade, guillotine, or other suitable

means. The validity of saw cuts, etc., should be verified by comparison to coupon tests

of a cracked specimen of the same material. In those cases where bolt failure, or its

equivalent, is to be simulated as part of a possible flaw configuration in joints or fittings,

bolts can be removed to provide that part of the simulation.

(7) Flaw Tolerance Demonstration (Flaw Tolerant (Enhanced) Safe-Life or

Fail-Safe Flaw Growth).

(i) It should be determined by analysis, supported by test evidence, that

the structure with the extent of damage established for residual strength evaluation can

withstand the specified design limit loads (considered as ultimate loads). Flaw tolerant

safe-life substantiation provides a safe period of operation of structure with flaws with

only routine inspections necessary. Safe crack growth (fracture mechanics)

substantiation for fail-safe designs, on the other hand, provides for limited operation

after crack initiation from flaws until the cracks can be safely detected. Since flaw

tolerant safe life does not provide for detailed crack growth beyond the threshold of

detectability, it will tend to apply to life-limited parts, particularly those hard to inspect.

Safe crack growth substantiation of fail-safe designs is applicable to readily inspectable

parts and may provide for replacement “on condition” rather than at a specified life.

(ii) The enhanced safe-life of metallic components will use analyses and

testing similar to that of basic safe-life except that “flawed” specimens will be tested

rather than “pristine” specimens.

(A) In order to determine the mean fatigue strength and the variability in

fatigue strength considering flaws, it is necessary to test a number of specimens in

establishing stress versus number of cycles (S-N) curves. Both full-scale and coupon

specimens may be used to account for the variability in fatigue strength. A reduction

factor should be applied to the mean curve in arriving at a working S-N curve. This

factor should include consideration of the number of specimens tested, the variability of

the fatigue results and, where available, previous test data on the same materials or

similar components, as well as service experience.

(B) Where new materials or designs are being evaluated, it is

recommended that a larger reduction factor be used until additional test data justifying a

change are available. The mean and reduced S-N data acceptable to the Administrator

on specimens with stress concentration factor, as applicable. A reduced S-N curve and

the loading spectrum of paragraph AC 29 MG 11 f(2) should be used in determining

replacement times.

(C) Figure AC 29 MG 11-1 represents the method of constructing a

typical S-N curve from the fatigue test data. Four to six full-scale specimens have

commonly been used in past safe-life programs to determine mean S-N curves. The

applicant may propose a specific number of specimens that may be evaluated with
AC 29-2C 9/30/99

Page MG 11 - 14

respect to the proposed methodology in determining the acceptability of the structural

substantiation program.
9/30/99 AC 29-2C

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(D) Additional coupon tests have been used, when necessary, to identify

more completely test scatter due to material properties variability. When the

agreed-upon number of specimens is tested at various acceptable load levels, a

reduction by three mean standard deviations, sigma (σ), necessary to account for

material properties, should be made to establish a reduced S-N curve for determination

of retirement times. This reduction may take into account the fact that flawed

specimens were tested to preclude a dual penalty situation. Reduction by two mean

standard deviations rather than three, may be used if justified by appropriate design

features such as multiple elements or unmistakable flaw indications or by material

properties that provide benign types of failure modes.

(E) “Run-outs” (specimens that do not fail) may be treated in one of three

ways: deleted from data (conservative), considered as a failure at maximum test cycles

(conservative), or considered in a rational manner to develop a more realistic “σ.” To

further account for variability in fatigue strength not totally accounted for by this

reduction in S-N data, the lesser of this reduction, 80 percent of mean, or bottom of

scatter, may be used as a limit, where justified, depending on material properties.

(iii) The procedures in Paragraph 7a of AC 20-107A provide criteria for

substantiating safe flaw growth for composite structure.

(iv) Safe crack growth (fracture mechanics) substantiation should show

that the damage growth rate under the repeated loads expected in service (between the

time at which the damage becomes initially detectable and the time at which the extent

of damage reaches the value for residual strength evaluation) provides a practical basis

for development of the inspection programs and procedures described in

paragraph AC 29 MG 11g(8). For multiengine load paths, a minimum of three

inspection intervals is recommended between the initially detectable damage time and

the time when residual strength is reduced to design limit load by crack growth. For

single element structures, a minimum of four inspection intervals is recommended. The

repeated loads should be defined in the loading, temperature, and humidity spectra.

The loading conditions should take into account the effects of structural flexibility and

rate of loading where it is significant.

(v) For flaw tolerance to achieve an improvement in safety over safe-life

for composite structures, the procedures of Paragraph 7a of AC 20-107A are

recommended. For flaw tolerance to achieve and improvement in safety over simple

safe-life for metallic structures, the following testing criteria are recommended to

supplement flaw tolerant safe-life analysis using flawed specimen S-N curves or crack

growth analyses using appropriate stress intensity factors and da/dn data from coupon

tests:

(A) Flaw Tolerant (Enhanced) Safe-Life Testing Criteria. Test full-scale

specimens with flaws or a mix of full-scale and coupon specimens with flaws to obtain

S-N data. Plot the data as shown in figure AC 29 MG 11-1. Utilize the S-N data and
AC 29-2C 9/30/99

Page MG 11 - 16

loading spectrum of paragraph AC 29 MG 11 f(2) in substantiating a crack-free life or in

arriving at a replacement time by cumulative damage analysis means. Where practical

(as allowed by the number of damaging cycles in the loading spectrum), spectrum

testing may be used for fatigue substantiation in lieu of S-N testing followed by analysis.

The replacement time established should be included in the airworthiness limitations

section of the document established under § 29.1529.

(B) Fail-Safe Crack Growth (Fracture Mechanics) Substantiation. Test

two or more specimens to obtain crack propagation data using either a realistic load

spectrum or an accelerated load (spectrum or single) associated with the use of

propagation theory and data after cracks have been initiated. Unless a more rational

method with an equivalent level-of-safety is applied for, the following methods of setting

inspection intervals should be applied. In all cases, the inspection methods and

intervals should adequately consider variables such as inspectability, type of inspection,

crack growth behavior, and other scheduled maintenance considerations.

(1) For single element (load path) structure, plot the data and set the

inspection as shown in figure AC 29 MG 11-2.

(i) Set the initial inspection at L1 /3.

(ii) Set the repetitive inspection intervals at L2 /4.
9/30/99 AC 29-2C

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AC 29-2C 9/30/99

Page MG 11 - 18

(2) For multi-element load path structure:

(i) Test all paths simultaneously with a 0.05-inch thick crack (or size

detectable by manufacturing quality control procedures) in the critical element at the

start of tests.

(ii) Note when field detectable cracking occurs.

(iii) Note when one element fails.

(iv) Note when the residual strength of the remaining elements decreases

to limit load due to crack growth.

(v) From figure AC 29 MG 11-3, set initial inspection at L1 /3.

(vi) Set repetitive inspection intervals at (L2 + LR)/3.

NOTE: If partial failure of the critical element is not detectable, then L2 becomes zero.
9/30/99 AC 29-2C

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AC 29-2C 9/30/99

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(C) Other test and inspection programs may be used rather than those of

paragraph AC 29 MG 11 g(7)(v)(A) and paragraph AC 29 MG 11 g(7)(v)(B), if shown to

have comparable or better probability of assuring that a catastrophic fatigue failure will

not occur.

(vi) The flaw tolerance characteristics can also be shown analytically by

reliable or conservative methods such as the following:

(A) By demonstrating quantitative relationships with structure already

verified as damage tolerant;

(B) By demonstrating that the damage would be detected before it

reaches the value for residual strength evaluation; or

(C) By demonstrating that the repeated loads and limit load stresses do

not exceed those of previously verified designs of similar configuration, materials, and

inspectability.

(8) Inspection. Detection of flaws before they become dangerous is the

ultimate control in ensuring the flaw tolerance characteristics of the structure.

Therefore, the applicant should provide sufficient guidance information to assist

operators in establishing the frequency, extent, and methods of inspection of the critical

structure, and this kind of information should, under § 29.571(a)(2), be included in the

maintenance manual required by § 29.1529. Due to the inherent, complex interactions

of the many parameters affecting flaw tolerance, such as operating practices,

environmental effects, load sequence on flaw growth, and variations in inspection

methods, related operational experience should be taken into account in establishing

inspection procedures. Comparative analysis can be used to guide the changes from

successful past practice, when necessary. Therefore, maintenance and inspection

requirements should recognize the dependence on experience and should be specified

in a document that provides for revision as a result of operational experience, such as

the one containing the operator’s FAA/AUTHORITY-approved structural inspection

program developed through the Maintenance Review Board (MRB) procedures for FAR

Part 121 operators.

h. COMBINATION OF REPLACEMENT TIME AND FLAW GROWTH

EVALUATION. It may be possible to extend the replacement time of safe-life

components that exhibit limited flaw tolerance capability by using a combination of the

safe-life and flaw growth characteristics as described elsewhere in

paragraph AC 29 MG 11 and by assigning both a replacement time and inspection

period to these components. The replacement time may then be based on the

combined probability of not initiating a fatigue crack at or before the replacement time

and the probability that the crack, if initiated, will be detected prior to catastrophic failure

or loss of limit load (or maximum attainable load, whichever is less) carrying capability.

The probability of detection should be based on consideration of the inspection

effectiveness, the inspection intervals, and the fatigue life remaining after an obvious
9/30/99 AC 29-2C

Page MG 11 - 21

partial failure. A lower strength reduction factor commensurate with this probability of

detection may then be used in the determination of the replacement time.
AC 29-2C 9/30/99

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Attachment 1 to AC 29 MG 11

FLIGHT STRAIN PROGRAM CONDITIONS TO BE INVESTIGATED

1. GROUND CONDITIONS.

a. Normal start.

b. Rapid increases of RPM on ground to maximum power-on RPM of main rotor.

c. Taxiing with full cyclic control.

d. Landing run (if applicable).

e. Braking (if applicable).

f. Normal shutdown.

g. Special ground checks (if applicable).

2. IN-GROUND-EFFECT (IGE) MANEUVERS.

a. Hovering.

(1) Steady with rotor at maximum side of RPM tolerance.

(2) Steady with rotor at minimum side of RPM tolerance.

(3) 90° right turn.

(4) 90° left turn.

(5) Control reversal.

(i) Longitudinal.

(ii) Lateral.

(iii) Rudder.

(6) Sideward flight.

(i) Right.

(ii) Left.

(iii) Rearward flight.

b. Maneuvering.

(1) Jump takeoff.

(2) Normal takeoff and accelerate to climb airspeed.

(3) Normal approach and landing.

(i) Multiengine.

(ii) One-engine-inoperative.

(4) Full autorotational landing.
9/30/99 AC 29-2C

Page MG 11 - 23

Attachment 1 to AC 29 MG 11

FLIGHT STRAIN PROGRAM CONDITIONS TO BE INVESTIGATED

(continued)

3. FORWARD FLIGHT-POWER ON.

a. Level flight.

(1) 40 percent VH.

(i) Minimum side of main rotor RPM tolerance (RPM +)

(ii) Maximum side of main rotor RPM tolerance (RPM -)

(2) 60 percent VH.

(i) (RPM +)

(ii) (RPM -)

(3) 80 percent VH.

(i) (RPM +)

(ii) (RPM -)

(4) VH.

(i) (RPM +)

(ii) (RPM -)

(5) VNE.

(i) (RPM +)

(ii) (RPM -)

b. Maneuvers.

(1) Full power climbs.

(i) All engines operative.

(ii) One-engine-inoperative.

(2) Cyclic pull-ups.

(i) 60 percent VH.

(ii) 90 percent VH.

(3) Normal acceleration from climb airspeed to 90 percent VH.

(4) Turns.

(i) Right at 60 percent VH and 90 percent VH.

(ii) Left at 60 percent VH and 90 percent VH.

(5) Control reversals at 90 percent VH.

(i) Longitudinal.

(ii) Lateral.

(iii) Rudder.

(6) Deceleration from 90 percent VH to descent airspeed.

(7) Part power descent.

(i) All engines.

(ii) One engine out.
AC 29-2C 9/30/99

Page MG 11 - 24

Attachment 1 to AC 29 MG 11

FLIGHT STRAIN PROGRAM CONDITIONS TO BE INVESTIGATED

(continued)

4. POWER TRANSITIONS.

a. All engines operating to one engine out.

(1) In full power climb.

(2) At 90 percent VH.

b. One engine out to all engines operating in powered descent.

c. All engines operating to autorotation.

(1) At 60 percent VH.

(2) At maximum forward transition speed.

d. Stabilized autorotation to all engines operating at normal autorotation airspeed.

5. AUTOROTATION.

a. Stabilized.

(1) At 70 percent VNE.

(2) At NE.

b. Turns at 70 percent and 100 percent VNE.

(1) Right.

(2) Left.

c. Cyclic pull-up.

d. Control reversals.

(1) Longitudinal.

(2) Lateral.

(3) Rudder.

The flight and ground conditions of this table may be used as practical in

developing operating load spectra for the range of projected aircraft usage such

as:

(1) Short missions with high-power cycles and ground-air-ground (GAG) cycles

(such as used in logging operations).

(2) Medium range missions of 1 to 2 hours that include three or more GAG or

LO-HI power cycles per mission.
9/30/99 AC 29-2C

Page MG 11 - 25

Attachment 1 to AC 29 MG 11

FLIGHT STRAIN PROGRAM CONDITIONS TO BE INVESTIGATED

(continued)

(3) Long range missions that consider the longest practical range of the aircraft

and consider two or more GAG (or high power) cycles per mission (such as used

for remote offshore drilling platforms).

If aircraft usages that produce a large number of high-power cycles (approximately

50 per hour) or high GAG cycles (approximately 25 per hour) such as logging

operations are not included in the fatigue tolerance substantiation, this should be

noted in the RFM and Maintenance Manual airworthiness limitations section

(reference paragraph AC 29 MG 11 f(1)).