AC 29-2C, Chg 4 2/27/2014 due to shifting cargo. In addition, w hen 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, non-accessible compartment. Also, the compartment weight limitations placard should be changed. AC 29-2, section 29.787 pertains to cargo and baggage compartments.
(vii) Safety Assessment. When installing any new equipm ent or modifying existing equipment, a safety assessment must be made to assure the FAA that all possible failure conditions that could occur from these changes have been adequately addressed to show compliance to the regulations.
(17) 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 (e .g. new model, manufacturer) is to be substituted, then an evaluation should be made to ensure the substitute equipment is also satisfactory. This evaluation would no rmally consist of comparing the attachment means, design features, failure modes, specifications, and operation of the two units.
The purpose of the evaluation is to ensure there are no differences that have an adverse effect on the airworthiness of the installation. Ot her 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 Regulations and References.
(1) Regulatory Sections. 14 CFR 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,
29.1589, part 91, and part 135.
(2) Other References Refer to the current version of each document.
(i) Helicopter Association International, Emergency Medical Services Recommended Guidelines.
(ii) National Highway Traffic Safety Administration, Air Ambulance Guidelines.
(iii) AC 23-2, Flammability Tests.
(iv) AC 43.13-2, Acceptable Methods, Techniques, and Practices-Aircraft Alterations.
(v) AC 135-14, Emergency Medical Services/Helicopter (EMS/H).
(vi) Oxygen Equipment for Aircraft, Society of Automotive Engineers Aerospace Information Report No. 825.
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(vii) MIL-D-19326G, Design and Installation of Liquid Oxygen Systems in Aircraft, General Specification for Military Specification.
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AC 29-2C, Chg 4 2/27/2014
BUILDUPIVENT VALVE IVA INTEGRAL ASSEMBLY (Most Cases) TO GAUGING SYSTEM LIQUID CONTAINER PRESSURE RELIEF PRES CLOSING VALVE PRES. BUILDUP COIL SHUTOFF VALVE
EVAPO- RATING ASSEMBLY
OVERBOARD VENT LINE RELIEF FLOW REGULATOR CONTROL
TO PATIENT FIGURE AC 29.MG 6-1 TYPICAL LIQUID OXYGEN SYSTEM Page MG 6 - 22
FILL, BUILDUP AND VENT VALVE NTEGRAL ASSEMBLY (Most Cases) TO GAUGING Z SVSTEM LIQUID CONTÂINER PRESSURE RELIEF PRES. CLOSING VALVE PRES BUILDUP COIL EMERGENCY SHUTOFF VALVE
EVAPO- RATING COIL ASSEMBLY OVERBOARD VENT LINE PRESSURE RELIEF • REGULATOR FIGURE AC 29.MG 6-2 TYPICAL LIQUID OXYGEN SYSTEM - USING
COMBINATION VALVE Page MG 6 - 23
CHAPTER 3
AIRWORTHINESS STANDARDS
TRANSPORT CATEGORY ROTORCRAFT
MISCELLANEOUS GUIDANCE (MG) AC 29 MG 7. STRUCTURAL CONDITION INDICATORS.
a. Related Sections. § 29.301 - Loads; § 29.305 - Strength and Deformation;
§ 29.571 - Fatigue Evaluation of Flight Structure; § 29.1301 - Function and Installation;
§ 29.1309 - Equipment, Systems and Installations; § 29.1321 - Arrangement and Visibility; § 29.1322 - Warning, Caution, and Advisory Lights; § 29.1355 - Distribution System; § 29.1503 - Airspeed Limitations: General; and § 29.1529 - Instructions for Continued Airworthiness.
b. Background.
(1) Structural condition indicators have been used on rotorcraft for several years in two main programs: as part of the basic type design and as part of airworthiness directive (AD) action. When approved as part of the basic type design, only limited “credit” has been given for the installation of structural condition indicators;
i.e., components provided with a structural condition indication system were required to be designed to § 29.571 “safe-life” criteria considering the structural condition indicator system inoperative. So- called “nonhazard” approvals were granted. When used as part of the mandatory actions of ADs, structural condition indicators have had a degree of “credit” recognized, primarily in the recognition of “fail-safety” provided by the indicator system.
(2) Since structural condition indicators have been used during both original type design and AD issuance, and since there is movement toward increased damage tolerance in rotorcraft design, policy concerning condition indicator use is considered appropriate.
c. At present, the use of structural condition indicators alone on new type designs is not considered an acceptable substitute for providing the necessary safe life for each component. However, areas which may be considered when approving these indicators for fail-safety credit are delineated in the following paragraphs.
d. What, how, when, where, and who of structural condition indicators.
(1) Indication of what?
(i) Previous structural condition indicators have primarily been used for crack detection. Several types of through- the-thickness crack detection systems are currently in use. Two types which detect changes in pressure in an instrumented chamber due to gas movement through a cracked wall are known as the blade Page MG 7 - 1
inspection method (BIM) system and the integral spar inspection system (ISIS). These systems can only detect full-depth cracks which are large enough to allow loss (of gain) of pressure from the instrumented chamber. This presents a limitation since full -depth cracks may be fast growing before detection. Another through-t he-thickness crack method is a pressurized, dyed fluid or oil system to detect through cracks in specially designed bolts (NASA patent), spindles, pins, or other closed chamber mechanical equipment.
(ii) Surface cracks can be found by systems such as surface-mounted crack detection wires. These systems would allow a greater safe crack growth period for assuring safe landing after detection than the through-crack-detection systems, but they have been used little in operations because of significant limitations; e.g., complexity of installation, durability problems, limited areas of coverage, and strain level limitations.
(iii) Some aircraft have had mast moment indicators or other load indicators to help prevent the pilot from inadvertently applying a high load to the instrumented system or to help the pilot reduce the load by control movements. These load indicators only indirectly give indications of structural condition; therefore, only limited “credit” is allowed for this use. “Credit” is limite d in that the fatigue life substantiations of § 29.571 should consider a reasonable number of excursions into the higher ranges established for the load indicator, and special inspections, rework, or replacement instructions should be provided for any strength degradation associated with high range excursions.
(2) How indicated?
(i) Current BIM systems use two types of indicators. The visual blade inspection method (VBIM) uses a gauge mounted on the blade which must be read visually by maintenance personnel while the aircraft is parked. The cockpit blade inspection method (CBIM) uses lights mounted in the cockpit which may be monitored by the crew. Other pressurized chambers have used dyes or oils to improve visual inspection effectiveness. Mast moment indicators and other load indicators use instruments with marked ranges and needles.
(ii) No specific types of load indicators are required by the FAA/AUTHORITY but the type used should be evaluated for accuracy, readability, and overall effectiveness. Paragraphs AC 29 MG 7(e) and (f) cover, in more detail, the use of structural condition indicators.
(3) When indicated? Structural condition indicators are used before flight, during flight, and for normal maintenance inspections. Paragraphs AC 29 MG 7 (e) and
(f) contain guidance for cockpit-mounted instruments which are monitored during flight.
Indicators used for normal maintenance inspections are the preferred type since they can be scheduled to allow the most effective use of available maintenance personnel of well-equipped maintenance facilities and of parts available.
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(4) Where indicated? Indications on the component are provided by VBIM systems and by systems utilizing dye or colored oil leakage. Cockpit -mounted lights and gauges may be used for certain critical structures which require frequent, but simple, checks. Maintenance panel locations (cabin, equipment bay, etc.) are the preferred locations for use in routine maintenance.
(5) Who reads indicators? The flightcrew, of necessity, monitors indicators mounted in the cockpit for use during flight. Gauges with ranges of values representing mast bending moments or other structural loads are monitored by the flightcrew, as necessary, to reduce or to prevent control operations from imposing excessive loads or to prevent too many high load applications. Maintenance personnel are generally responsible for reading component-mounted indicators and for monitoring indicators which are mounted on maintenance panels. The before- flight checks may be conducted by maintenance personnel or by flightcrew in certain cases (i.e., cockpit-mounted gauges or “push-to-test” checks).
e. Actions required by indicators .
(1) On-ground indications. Indications noted on the ground should be followed by a functional check of the indication system as provided for by its design. If indications persist after the system has been checked and found to be functional, further inspection of the affected component(s) should be conducted for damage assessment.
Any damage found as a result of the detailed inspections should be repaired or replaced as appropriate.
(2) In-flight indications.
(i) Indications used for in-flight monitoring have in the past been used for two main reasons: to provide a structural load display (such as mast bending moment) and to help resolve a service problem (CBIM systems have been used to supplement conventional inspection methods in blind areas).
(ii) Structural load display systems should not be used instead of correcting deficient designs. Structural load display systems are appropriate for use in locating control positions, such as the cyclic stick, under transient conditions such as slope landings and hover in sidewinds, but structural load display systems are not considered appropriate for routine operations such as climbout or cruise with constant attention required by the flightcrew. If the load indicator provides a needed tool to the pilot in limited types of operations and does not significantly add to pilot workload otherwise, its use can be considered.
(iii) In the past, certain service problems have been solved by adding in-flight indicators such as CBIM systems. When retrofit of the affected structure is impossible or impractical, and when conventional inspection techniques are shown to be inadequate by themselves, CBIM or similar systems may be the only practical Page MG 7 - 3
solution, despite the increase in pilot workload and the potential for problems caused by overreaction by the pilot to a structural fault indication. When used for correction of service difficulties, the structural condition indicator system should be accompanied by clear, concise crew directions to prevent possible catastrophic overreaction. Load reduction measures such as rotor speed changes, airspeed reductions, altitude changes, etc., should be clearly provided, if needed. Crack propagation time from indication should be sufficient to allow continued safe flight to a safe landing area. For new designs, CBIM or similar systems which add to the pilot’s workload are considered inappropriate. Proper redesign to provide the needed safe life, fail -safety, and inspectability is considered the appropriate action.
f. Complementary considerations of structural condition indicator use.
(1) Two basic programs are commonly used for approval of structural condition indicators. Basic type certification procedures are used for mast moment indicators and similar systems, and AD’s (with appropriate type design changes) are used for CBIM systems which require pilot attention and corrective action when an indication of a structural fault is detected.
(2) The fatigue substantiation required by §§ 27.571 and 29.571 should consider a conservative number of excursions into the high load range monitored by a structural condition indicator such as a mast moment indicator. Static strength should not be adversely affected by a single excursion into the high load range monitored by the indicator.
(3) Complementary design provisions should accompany the use of a structural condition indicator system. Redundancy of load paths and inspection systems and indicator system failure analyses should be provided, as necessary, to meet the requirements of § 29.1309. The life remaining after the indicator system detec ts a structural failure should be calculated (with test verification), and compatible inspection and/or overhaul programs should be provided.
(4) The FAA/AUTHORITY approval of a structural condition indicator system requires evaluation by the airframe, systems and equipment, and flight test specialists.
The airframe specialist has the responsibility to review effects of structural condition indicator system use on aircraft loads, strength and deformation, and structural fatigue evaluation as well as the instructions for continued airworthiness. The systems and equipment specialist needs to evaluate the system for function and installation as well as the reliability requirements of § 29.1309. Flight test evaluation of the instruments’ arrangement and visibility, effect on crew workload, and possible changes for RFM is also needed. Care should be exercised to assure that responsibilities are not given to the flightcrew which would be more appropriately handled by a redesign or by the maintenance personnel. Early coordination between all specialists is necessary to prevent delays from last minute design changes.
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CHAPTER 3
AIRWORTHINESS STANDARDS
TRANSPORT CATEGORY ROTORCRAFT
MISCELLANEOUS GUIDANCE (MG) AC 29 MG 8. (Amendment 29-42) SUBSTANTIATION OF COMPOSITE ROTORCRAFT STRUCTURE.
a. Reference. 14 CFR §§ 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 procedur es 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 FAA/AUTHORITY.
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 st rength substantiation for both metal and composite structure is essentially identical to that for fixed wing structure once basic loads have been determined. However, rotorcraft st ructural fatigue substantiation for composites is significantly different from fixed wing fatigue substantiation. Since AC 20-107A, as dev eloped, 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 component fatigue tests for rotorcraft programs rather than just one full-scale test. Also, constant amplitude, Page MG 8 - 1
AC 29-2C, Chg 2 4/25/06 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 AC 29-2C, Chg 1, MG 11, “Fatigue Tolerance Evaluation of Transport Category Rotorcraft Metallic 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 AC 27-1B, Chg 1, MG 11, “Fatigue Evaluation of Rotorcraft Structure,” contain comprehensive discussions of the S-N curve generation process. The rotorcraft S-N curve process contrasts sharply with the fixed wing process of using a single full-scale fatigue article (usually an entire wing or airframe, which constitutes a single full-scale assembly data point), generic material or full-scale assembly S-N data (e.g., MIL-HDBK-5 for metals, MIL-HDBK-17 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 natur e, these designs are not damage or flaw critical. Thus by circumstance as much as design, early composite rotor blade and other composite rotorcraft designs incorporated an acceptable fatigue tolerance level of safety. In the 1980’s, more test data, analytical knowledge, and analytical methodology became available to more completely substantiate a composite design. Current 14 CFR Page MG 8 - 2
Parts 27 and 29 contain many sections (reference paragraph a above) 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 14 CFR 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.
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AC 29-2C, Chg 2 4/25/06
(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.
(11) FIBER. A s ingle 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 (e.g., 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.
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