(VI) Equipment in Cargo and Baggage Compartments. When components are added to the compartment, rovisions should be made to protect the system components due to shifting cargo. In addition, when oxygen components are installed, the compartment should be placarded against the storage of oil or hydrocarbons. A smoke detector is recommended for a compartment if oxygen cylinders are installed in a closed, nonaccessible compartment. Also, the compartment weight limitations placard should be changed. Paragraph 336 of this AC pertains to cargo and baggage compartments.
(15) Equipment Substitution. The EMS modification that is presented for approval will contain specific items of equipment, and the approval will make reference to this equipment. If other equipment (new model, manufacturer, etc.) is to be substituted, then an evaluation should be made to assure the substitute equipment is also satistactory. This evaluation would normally consist of comparing the attachment means, design features, failure modes, specifications, and operation of the two units.
The purpose of the evaluation is to assure there are not differences that have an adverse effect on the airworthiness of the installation. Other differences would not be considered significant. A specific seat and litter design is approved as a part of the EMS configuration. Substitutions may be approved in accordance with the standards.
c. Related FAR Sections and References.
(1) FAR Sections. §§ 29.337, 29.471, 29.561, 29.773, 29.783, 29.785, 29.803,
29.805, 29.809, 29.811, 29.813, 29.815, 29.831, 29.853, 29.1301, 29.1309, 29.1353,
29.1357, 29.1411, 29.1431, 29.1557(d), 29.1583(d), 29.1585, and 29.1589.
(2) Other References.
(i) Helicopter Association International, Emergency Medical Services Recommended Guidelines, 1987, First Revision, 2 pages.
(ii) National Highway Traffic Safety Administration, Air Ambulance Guidelines, dated 1981.
(ill) FAR Part 91, General Operating and Flight Rules.
(iv) FAR Part 135, Air Taxi Operators and Commercial Operators.
(v) AC 67-1, Medical Information for Air Ambulance Operators, dated March 4, 1974.
(vi) AC 23-2, Flammability Tests, dated August 20, 1984.
(vii) Oxygen Equipment for Aircraft, Society of Automotive Engineers Aerospace Information Report No. 825B, Rev. 9/86.
Par 786 1257
(viii) Acceptable Methods, Techniques, and Practices--Aircraft Alterations, AC 43.13-2A, dated June 9, 1977.
(ix) Design and Installation of Liquid Oxygen Systems in Aircraft, General Specification for Military Specification MIL-D-19326G, dated October 1, 1985.
1258 Par 786
Par 786 INTEGRAL ASSEMBLY (Most Cases) 7/30/97
OVERBOARD B
VENT LINE PRESSURE BUILDUPIVENT
RELIEF VALVE TO
GAUGING SYSTEM
PRESSURE RELIEF PRES.
CLOSING LIQUID VALVE Oz
TO FLOW CONTAINER
PATIENT CONTRO REGULATOR ALV EVAPO- PRES.
BUILDUP
RATING COIL FILL COIL CHECK ¡VALVE ASSEMBLY IVALVE_ -
EMERGENCY
SHUTOFF VALVE FIGURE 786-1 TYPICAL LIQUID OXYGEN SYSTEM AC 29-2B 1259
INTEGRAL ASSEMBLY AC 29-2B (Most Cases)
OVERBOARD
VENT LINE
PRESSURE
RELIEF FILL, BUILDUP
AND VENT VALVE TO
GAUGING SYSTEM PRESSURE PRES.
RELIEF CLOSING LIQUID
VALVE CONTINER FLOW
CONTROL REGULATOR VALVE PRES.
EVAPO- BUILDUP
RATING COIL
COIL CHECK ASSEMBLY | VALVE
EMERGENCY
SHUTOFF VALVE FIGURE 786-2 TYPICAL LIQUID OXYGEN SYSTEM - USING COMBINATION VALVE 7/30/97 Par 786
787. RESERVED.
Intentionally Left Blank Par 787 1261
788. (Amendment 29-35) SUBSTANTIATION OF COMPOSITE ROTORCRAFT STRUÇTURE.
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 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 1262 Par 788
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 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 Par 788 1263
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) EIBER. 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) EIBER 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, (l.e., each Par 788 1265
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) SHELE LIEE. 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.
1266 Par 788
(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 Materialsi
(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 Par 788 1267
АС 29-2B 7/30/97 applicant's organization. Personnel who are composites specialists (or are otherwise knowledgeable in the subject) should be primary team member candidates. Once selected, it is recommended that team meetings be held periodically (possibly in conjunction with type boards) during certification to ensure the building block certification process is accomplished as intended.
(1) The first area is the fabrication requirements of § 29.605:
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.
(її) 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 detects) allowed by the reliability limitations of the approved inspection program.
(ill) 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 1268 Par 788