АС 27-1A 7/30/97 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 §§ 27.603 and 27.613. These criteria require application of the critical environmental limits such as temperature, humidity, and exposure to aircraft fluids (such as fuel, oils, and hydraulic fluids), to determine their effect on the performance of each composite material system. Temperature and humidity effects are commonly considered by coupon and component tests utilizing preconditioned test specimens for each material system selected. Material "A" & "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 tollowing:
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 g(1) above).
(її) 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.
(i¡i) 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 1120 Par 788
7/30/97 АС 27-1A 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. 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 Par 788 1121
АС 27-1A 7/30/97 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 tull-scale design criteria and parameters necessary to ensure that detrimental structural creep in full-scale structure does not occur in service, should be conducted early in certification and should be FAA/AUTHORITY-approved.
(vili) 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 § 27.609.
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 § 27.610.
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 § 27.610), 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.
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(5) The fifth area is the static strength evaluation requirements of §§ 27.305 and 27.307 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 the service period). Analytical reports and tests should consider all possible failure modes and should include the critical, allowable effects of:
Environment (reference Paragraphs 2 and 3.)
(i) Service Life (residual limit strength and stiffness demonstration.)
(ili) 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 is 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.
(vili) For static strength analysis laminae and laminate "A" and "B" basis allowables (determined in accordance with Paragraph (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.
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АС 27-1А 7/30/97 (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.
(6) The sixth area is the fatigue evaluation requirements of § 27.571. The fatigue evaluation method for the rotorcraft being certified should consider damage tolerance in accordance with AC 20-107A.
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)).
(її) 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 shoula 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.
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(vi) The method of generating S-N curves using approved raw data should be demonstrated, evaluated, and approved.
(vil) Any limited life items must be identified and placed in the Airworthiness Limitations section of the maintenance manual in accordance with § 27.571.
(vili) 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 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 § 27.629 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.
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 Vp, for a VSTOL aircraft) with the extent of damage for which residual strength and stiffness are demonstrated.
(iii) Where appropriate, crash impact dynamics considerations must 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 §§ 27.611, 27.1529, and FAR Part 27 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 Par 788 1125
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:
(1) The repair should be permanent.
(it) The repair should restore the structure to the required strength and stiffness.
The repair should restore all functional requirements.
(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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Appendix 1 APPENDIX 1
ROTORCRAFT ONE-ENGINE-INOPERATIVE POWER ASSURANCE
1. PURPOSE.
The purpose of this document is to establish an approach for an engine power assurance procedure which will assure that the required OE power level can be achieved.
2. GENERAL.
The data and methods described herein are intended to be utilized as a guide and not necessarily the only means of achieving the desired result.
3. APPLICABILITY The applicability of the document is intended to be primarily in support of the new 30-second and 2-minute OEI rotorcraft engine rating scheme.
4. PARTIAL POWER ASSURANCE (ENGINE "RUN-LINE").
a. Fundamental to the concept of limited-use one-engine inoperative (OEI) ratings is the requirement to be certain that the rated OEI power will indeed be available when needed. Conventional periodic power-assurance and topping checks are impractical with the limited-use rating concept because of the rapid expenditure of useful life during exposure at the engine speeds and temperatures consistent with limited-use ratings;
therefore, we require a means of assuring the power available, other than by actual demonstration on each service engine. The advent of more sophisticated controls and engine developments catering to the 30-second/2-minute OEl rating concepts can provide the means to determine: (1) that the thermodynamic/mechanical capability of the engine as tested at the prevailing ambient conditions, will permit reaching a specified power level at any other ambient condition and (2) the fuel system and the various limiters will not prevent the engine achieving OEI power on demand. Pending availability of these new methods, the "parallel run line check" approach is recommended.
b. The method commonly called the "parallel run-line check" that has been in use for two decades may require refinement for application to the new rating structure where the degree of extrapolation to the OE power level is more extensive and the slope of the individual engine characteristic is important. As in any power assurance method, success is strongly dependent on the validity of the data base, the maintenance of the engines and sensor/indicating systems, and the care taken during the conduct of the power check. In addition, trending of individual engine performance Page A1-1
АС 27-1A 7/30/97 Appendix 1 by the operator and associated analyses can be used to avoid unnecessary flight delays and engine removals.
c. Thermodynamic/mechanical capability can be addressed by test stand mapping of development engines over a range of ambient conditions to establish an adequate data base of engine characteristics. This will address characteristic slope variations between engines and establish correction factors necessary for extrapolation of data from a power assurance checkpoint to the 30-second OEI rating. Statistical verification and/or modification of the data base may be necessary during production by mapping of sample production engines. Performance data, at the 30-second OEI condition, taken during the supplementary block test and also during the "overhaul test" will demonstrate the capability of an engine and its control system near the end of an overhaul period to produce the required power. This will demonstrate capability with a deteriorated base-performance engine.
d. The question of fuel system limitations and other various limiters, which could prevent the engine from achieving OEI power on demand, may be addressed by use of more sophisticated control systems, for example, electronic controls utilizing several engine parameter limiters each with automatic datum reset capability. Such control systems can sense an engine failure and automatically reset the operating limiters upward from "normal" to "OEl" limits. Conventional flow and electronic bench testing can be used to verify the function and limit setting of the units when new or after overhaul or repair. The reset features can be extended in function to include a fixed magnitude pulldown type reset for use in verifying new and field production engine/control combination function ability. Pulldown type resets are currently in use today for verification of limiter settings on some engines and can be utilized in this application to avoid unneeded exposure of the engine to the rapid life expenditure conditions.
e. While the above is envisioned as the probable means in which assurance of capability will occur early in the application of such engines, there will be other means developed. One such means would be utilization of modern electronic engine condition or health monitors to display "go" or "no go" conditions relative to the ability of the engine and its control system to produce 30-second OEI power if required. In this application the device would be a "power assurance meter" and could be used with electronic, hydro-mechanical, and pneumo-mechanical control systems. It is entirely reasonable to expect that self-taught or self-programmed power assurance meters can be used that continually program the actual performance slope of the subject engine and extrapolate to the 30-second OEI with continuous engine monitoring.
Seli-programming occurs by sampling engine temperature, speed, torque, otner characteristics (such as tuel pressure), and ambient conditions, resulting in the reflection of an actual characteristic for the installed engine. The availability of this information permits treating engines individually, whether it is a new or deteriorated engine or one with either minimum or maximum slope, without the necessary compromises to "best" engines that necessarily occurs using the earlier statistical Page A1-2
Appendix 1 approach. The question of instantaneous fuel system capacity could be addressed by fuel pump/control systems incorporating bypass systems equipped with flow meters.
The health monitor or power assurance meter can continually integrate the fuel flow increment available in terms of power increment required in the event of OE and would include this intelligence in its pass-fail judgment criteria. Systems of this type would further be conducive to in-service ground checks by overt by-pass deactivation from low power settings to assure satisfactory mechanical function.
f. Power assurance for the limited-use OEl ratings depends on a complete understanding of the engine model's operating characteristics. Two approaches have been discussed, one where, with the aid of a sophisticated fuel control system, the engine "learns" its own characteristics, and the other where the performance extrapolation is compared with a known minimum standard. The establishment of the standard is obviously a vital part of the procedure, which depends to a large extent on the existence of a reliable data base. In a mature program this is relatively easy to maintain, since it is possible to use the new production engine acceptance data to establish engine-to-engine variation and also to test engines prior to overhaul to determine the effects of deterioration. Thus, an up-to-date minimum or worst-engine characteristic can be maintained and service engines would be compared with this minimum engine.
g. When the engine in question is a completely new design, or a remote derivative of an existing design, establishing the initial data base presents some problems which must be resolved. New production engines will eventually establish engine-to-engine variation, but initially an estimated worst variation must be assumed.
The rate of deterioration and its impact on the base standard must be accounted for from the first engine delivered, yet it may be some time before an acceptable number of engines can be tested after service.
h. A partial solution lies in the development and qualification cycle of the engine.
A typical new-design program requires several development engines, of which more than half can be expected to be used for endurance or accelerated endurance testing.
Furthermore, by the time certification is completed and production deliveries have commenced, these engines will normally have amassed several thousand hours of running usually to a schedule far more rigorous than normal service. The information gathered during these tests will provide the necessary data base for the assessment of in-service engines, and it can be progressively enlarged, and the derived data refined, as further production and service data are obtained.
ENGINE CONSIDERATIONS. This section describes the potential causes of an engine not delivering specification OEI power levels in spite of passing a parallel run-line power assurance check. Possible solutions are discussed in the context of one time use 30-second and 2-minute ratings.
a. Fuel Flow.
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Appendix 1
(1) An engine may not achieve maximum power available or emergency rating because insufficient fuel is supplied. This condition has a number of possible causes:
(i) Low acceleration schedule (Il) Low maximum fuel stop (III) Low fuel pump output
(iv) Restrictions between the fuel control and the combustor
(2) The proposed emergency ratings (OEI) may preclude the use of a topping check to uncover the above problems; therefore the following procedures are advanced which can be used either separately or in combination with other approved methods to assure that the required fuel flow is available.
(3) During engine acceleration the fuel flow rate is considerably higher when compared with the normal steady state condition. This fact can be used to verify the availability of OEl fuel flow. The verification can be done by a direct measurement of fuel flow during an acceleration or derived indirectly from the engine acceleration rate.
It is envisaged that the determination of fuel flow by these procedures should be done by some automatic means.
(4) Figure 1 is a bypass technique in which some of the fuel controls output is routed away from the engine and back to tank. This forces the fuel control onto the acceleration schedule in order to maintain gas generator speed. The design of the system should ensure that with the bypass flowing the fuel control outlet pressure and flow at the OE ratings are simulated. The bypass system can be either permanently installed and operated in flight, (Failure Malfunction Effects Analysis must be provided), or as an item of ground test equipment. The quantity of fuel bypassed should be equivalent to the worst case difference between fuel flow at the 30-second rating and typical power assurance power levels. However, trend monitoring and service history may provide the basis of an alternative to periodic measurement.
b. Limiters. A means must be provided to assure that a lower than required (for OEI power) limiter setting does not exist. Limiters that could prohibit reaching OEI power are as follows:
(1) Ng - Maximum Compressor Speed Limiter or Governor
(2) Measured gas temperature limiter
(3) Output shaft torque limiter Page A1-4
Appendix 1
(4) Np limiter or power turbine governor-power turbine governors can be verified at lower than OEI power conditions.
(5) Fuel flow limiter or maximum fuel flow stop-Fuel flow limiting has been addressed in previous paragraphs.
c. Eailure Modes and Effects Analysis.
Failure modes and effects analysis, along with limited demonstration and suitable engine health monitoring procedures, may provide the basis of an acceptable solution to possible unexpected power limiting due to engine condition. It should be shown in the analysis that there is no probable event or combination of events which can cause a latent problem leading to inadequate fuel flow at high powers. The analysis should include all components of the fuel system such as: pump(s), control system (mechanical, hydromechanical, electronic, etc.) pipework, filters, fuel nozzle(s), and electrical interfaces. It should also address the probable effects of accumulated running time, dirty fuel, and hostile environment.
d. High Corrected Gas Producer Speed.
(1) The proposed OEl ratings will cause the engine to run at high corrected gas producer speeds (Ng/). At high Ng, performance characteristics of components, especially in the compressor, can change significantly and to an extent which would change the extrapolation of low speed run line data.
(2) In operation, the effects of the accretion of dirt, FOD, component deterioration, and erosion of blading may also cause changes in the high-speed performance of an engine.
(3) The above effects must be considered when developing power assurance procedures and data.
e. Special Devices.
(1) The satisfactory operation of devices or systems whose functioning is required in order to achieve the OEI powers should be verified. Devices or systems, which in normal operations are not exercised through the range of travel needed to achieve the OE powers, may require special checks to assure adequate capability.
(2) Special devices that are required only in order to achieve the OEl powers (for example, solenoids to provide additional cooling flow to hot-section components or a water/anti-freeze mixture into the compressor), should be subjected to periodic checks and have a demonstrated high reliability.
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