AC 27-1B, Chg 4 3/31/2014 two is an acceptable margin for systems with catastrophic failure conditions, if this margin is verified by aircraft test or by analysis supported by aircraft tests. For other verification methods, the margin should be agreed upon by the cognizant certification office.
(C) Flight and engine controls are examples of "critical" functions. With these critical functions defined, an analysis and testing should be performed to show compliance with § 27.1309(a) (i.e., equipment, systems, and installations performing critical functions should be designed and installed to ensure they continue to perform those intended functions subject to the conditions encountered in a worst-case lightning strike). Section 27.610 contains some methods which may be utilized for less complex mechanical systems; however, a great deal of difficulty will be experienced in trying to use these criteria to demonstrate that a very complex avionic system complies with § 27.1309(a). These avionic systems usually only require protection from indirect effects of lightning, and therefore a method such as that outlined in paragraph b.(4) below is recommended. This method may be readily adapted to other avionic systems performing critical functions. Also, this identifies an acceptable quantification of the expected airborne environment. The next step involves expanding the fault/failure analysis to determine what probable malfunctions or failures of the installed equipment systems and installation may occur as a result of a lightning strike to the rotorcraft. For a twin-engine rotorcraft, these systems must be designed such that a failure or malfunction will not cause a hazard to the rotorcraft. For a single-engine rotorcraft, these equipment systems and installation should be designed to minimize hazardous effects of failures or malfunctions. Note that the analysis which determines what are probable malfunctions or failures should consider that the encounter with a worst-case lightning strike is a given event.
(4) Lightning Strike Protection of Full Authority Digital Engine Controls (FADEC)
(i) Explanation.
(A) The following discussion is written specifically for FADEC with an alternate technology backup fuel control installed on rotorcraft with part 27 category A engine isolation. The requirement for increased consideration of lightning strike encounter effects on avionic equipment and systems has been brought about by the increased use of avionics to perform functions, the failure or malfunction of which could result in a hazard to the rotorcraft. The susceptibility of current high technology avionic systems is increased by the use of large scale integration, very large scale integration, and complementary metallic oxide silicon technologies which exhibit a greatly reduced tolerance to large amplitude, low energy electrical transients as compared to conventional bipolar technology, and the reduced physical protection and electromagnetic shielding afforded aircraft avionic systems by the advanced technology composite airframe materials. Additionally, processor-based systems have the failure phenomenon of digital upset. A digital upset occurs when a system, perturbed by an Page F - 20
electrical transient, ceases proper operation in accordance with its embedded software while suffering no apparent component or device damage.
(B) Since elements of electrical and electronic engine subsystems are typically spread throughout much of the rotorcraft, transients caused by lightning are coupled into the subsystem interface cables and may damage the system or cause upset. Effective lightning protection must be designed and incorporated into these systems. Reliance upon redundancy as a means of protection against lightning effects is generally not adequate because lightning electromagnetic fields and structural IR voltages usually interact (to some extent) with all electrical wiring aboard a rotorcraft.
(C) The testing and analysis outlined in this discussion are methods by which the FAA/AUTHORITY may be assured that when the rotorcraft experiences "the foreseeable operating conditioni of a worst-case lightning strike encounter that the electronically controlled engines will continue to "perform their intended function" and therefore be in compliance with § 27.1309 as installed.
(D) The definition of what constitutes a full authority engine control is not at this time clearly defined. However, it has been accepted in past certification that any control which relies upon the electronics for the function on which Civil Certification or Military Qualification is based (e.g., rotor speed governing) is a full authority control, regardless of the backup control mode provided. If engine certification or qualification can be achieved without the electronic control which is subsequently added to achieve improved operational efficiency in the aircraft, the control is "supervisory." However, if the controls used in a multiengine rotorcraft have a common failure caused by a lightning strike which could result in simultaneous failures which would cause a reduction in power greater than the loss of one engine, this would also be considered "full authority."
Note: If OE ratings are approved, cumulative loss of power from all engines must be limited to allow flight manual performance based on OE ratings.
(i1) Procedures. Although not a regulatory requirement, it is recommended that a formal written certification plan be used to assure regulatory compliance. The use of this plan is beneficial to both the applicant and the FAA/AUTHORITY because it identifies and defines an acceptable resolution to the critical issues early in the certification process. These are the usual steps to be followed when utilizing a certification plan:
(A) Prepare a certification plan that describes the analytical procedures and the qualification tests to be utilized to demonstrate protection effectiveness. Test plans should describe the rotorcraft and FADEC system to be utilized, test drawing(s) as required, the method of installation that simulates the production installation, the lightning zone(s) applicable, the lightning simulation method(s), test voltage or current waveforms to be used, diagnostic methods, and the appropriate schedules and location(s) of proposed test(s).
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AC 27-1B, Chg 4 3/31/2014 (B) Obtain FAA/AUTHORITY concurrence that the certification plan is adequate.
(C) Obtain FAA/AUTHORITY detail part conformity of the test articles and installation conformity of applicable portions of the test setup.
(D) Schedule FAA/AUTHORITY witnessing of the test.
(E) Submit a final test report describing all results and obtain FAA/AUTHORITY approval of the report.
, 1999, has beefinition of Environment. (iii) Definition of Environment. The SAE AE4L Committee report dated June 6, 1999, has been and remains acceptable criteria to define the worst-case lightning strike that may be encountered by the rotorcraft in service. An additional explanation of the lightning environment could be found in FAA Report DOT/FAA/CT-89/22, "Aircraft Lightning Protection Handbook." This handbook will assist aircraft design, manufacturing, and certification organizations in protecting aircraft against the direct and indirect effects of lightning strikes, in compliance with 14 CFR regulations. It presents a comprehensive test criteria to provide the essential information for the in-flight lightning protection of all types of fixed wing, rotary wing, and powered lift aircraft of conventional, composite, and mixed construction and their electrical and fuel systems. The handbook contains chapters on the natural phenomenon of lightning, the interaction between the aircraft and the electrically charged atmosphere, the mechanism of the lightning strike, and the interaction with the airframe, wiring, and fuel system. Further chapters cover details of designing for optimum protection; the physics behind the voltages, currents, and electromagnetic fields developed by the strike; and shielding techniques and damage analysis. The handbook ends with discussion of test and analytical techniques for determining the adequacy of a given protection scheme.
On September 7, 2011, FAA Advisory Circular AC 20-136B, "Aircraft Electrical and late Synew desins a pration ons this red seersedion previe ersions of
(iv) Certification Plan. The following subjects are not intended to provide a complete list of the items which should be included in the certification plan, but rather highlight some of the areas which should receive consideration. The certification plan should address the total protection which is required to allow the FADEC to continue to operate properly when the rotorcraft experiences a worst-case lightning strike encounter.
(A) Determination of Lightning Strike Attachments. Determine the locations on the rotorcraft where lightning strike attachment is likely to occur and the portions of the airframe through which currents may flow between attachments. The main and tail rotors are recognized as likely attachment points; however, consideration should be given to all possible attachment points. The swept stroke phenomenon may Page F - 22
not exist for all lightning strike encounters due to the fact that the rotorcraft may be airborne with little or no airspeed.
(B) Establish the Lightning Environment. Establish the components of the total lightning event to be considered. These are the currents and voltages which are described in the definition of the environment.
(C) Full-Level, Complete Vehicle Testing. In accordance with traditional FAA/AUTHORITY policy, the demonstration that the FADEC installed in a complete type design rotorcraft will continue to operate properly when exposed to a worst-case lightning strike is sufficient to demonstrate compliance with § 27.1309(a). Because of the difficulties involved in utilizing this type of an approach, it is generally not used.
(D) Analytical Processes. A description should be given in the certification plan of the analytical process and certification tests to be utilized to demonstrate protection effectiveness. Typically, the certification plan will include a combination of analysis and tests. (Analytical techniques are most often utilized to predict the levels of lightning-induced transients in interconnecting wiring.) In most cases, successful analyses are based upon well-defined geometrical or electrical parameters such as structural dimensions and materials resistivity. When electrical characteristics of structural materials are not well established, development tests are often utilized to obtain this data which is subsequently utilized in an analysis. In more complex structures or electrical and electronic system installations, it is sometimes difficult or impossible to define the problem in terms that can be analyzed. In these cases, development or verification testing is often relied upon. The purpose of the certification plan is to show how developmental tests, analyses, and verification tests are combined to demonstrate protection design adequacy. In certain cases, previously verified designs can be incorporated and their adequacy confirmed by reference to previous verifications. Such reference should also be incorporated in the certification plan.
(1) The verification testing should be conducted on a system which simulates as closely as possible the installed configuration. As few items as possible of actual hardware should be simulated.
(2) Unless the applicant has opted to follow the guidance in AC 20-136B as a means of showing compliance with § 27.1309(d), then the following or other acceptable guidance should be followed. The use of various analytical processes usually requires that the system component tolerance is established. The SAE AE4L Committee Report No. AE4L-81-2 is one example of an acceptable reference to be used for the testing accomplished to determine these tolerances. The testing which is performed to determine the tolerance level of the control computer should include a consideration for the occurrence of a non-recoverable digital upset. One method to provide this consideration is to have the unit powered and the processor operating normally under software control (usually this should be the exact software for which approval is sought) when the test is performed. If strike testing is used, then several shots should be made to develop enough data to provide a reasonable confidence level.
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AC 27-1B, Chg 4 3/31/2014 It is an acceptable procedure for the engine manufacturer, while he is obtaining his type certificate, to accomplish this bench testing to determine the level of tolerance of the FADEC system components to lightning encounter indirect effects. This approach has the advantages that the bench tests are not necessarily required to be repeated when the engine is installed in a different airframe. This recommendation is not meant to add a requirement to the engine manufacturer but to propose a more efficient method of certification. If this tolerance was not determined by the engine manufacturer, the applicant installing the FADEC in a rotorcraft would be expected to furnish this data.
(3) As with any analytical method, it is prudent to include a margin of safety to account for the uncertainties involved in the analytical and testing processes.
Margins account for uncertainty in the compliance method. As confidence in the compliance method increases, the margin can decrease. A factor of two is an acceptable margin for systems with catastrophic failure conditions, if this margin is verified by aircraft test or by analysis supported by aircraft tests. For other verification methods, the margin should be approved by the cognizant certification office.
(4) When an analysis has no associated full-scale vehicle testing to confirm the analysis, the analysis should be very rigorous. Additionally, it should be expected in this situation that this analysis indicates a very large margin of protection.
Many factors must be considered in determining what constitutes an acceptably large margin. The specific additional margin required should be based on an assessment of the inherent uncertainty of a given analysis.
(E) Pass or Fail Criteria. The certification plan should address a pass or fail criteria for the testing and analysis to be performed. The following items should be satisfied to assure acceptable system performance:
(1) No immediate crew action must be required.
(2) Automatic control of the engine cannot be lost for any appreciable period of time. The engine must not be allowed to be out of control for a period of time which will result in a hazard in a worst-case flight condition. Obviously, any rapid, uncontrolled divergence is not acceptable.
(3) No crew action should be required to reset the system. This is not to imply that the system cannot be designed with a manual reset, but the manual reset cannot be used to show compliance to recover from a digital upset.
(4) The resumption of engine control after an upset must be reasonably within the range which existed before the upset.
(5) No critical data can be lost.
(6) After the system recovers, if the performance of the system has been degraded in a noncritical manner which would reduce the capability of the Page F - 24
rotorcraft or the ability of the pilot to cope with adverse operating conditions, then the crew must be alerted to this system degradation.
(v) System Installation Considerations. In most cases, the installation of the system components is a constituent part of the lightning protection. This is particularly true in the use of shielding techniques. If these installation features are required for adequate lightning protection, consideration should be given to ensure that their effectiveness is not derogated in service. ICAs must be provided to the parties who service and operate the rotorcraft to allow them to take actions necessary to ensure the continued effectiveness of the system lightning protection.
c. Failure Analyses.
(1) Power and distribution systems should be analyzed to show compliance with § 27.1309.
(i) One acceptable procedure for documenting the analysis is contained in Society of Automotive Engineers (SAE) Fault/Failure Analysis Procedure ARP 926B, revised June 1, 1997.
(ii) As a minimum, any analysis should consider the effect of failures of components and systems on the capability of the rotorcraft to perform its intended function without hazard.
(iii) The analysis should consider the indication of failure. Those latent failures that occur without indication should be considered in all possible sequences and combinations of additional failures until a positive indication of failure is provided.
(iv) The analysis should consider failure of indirectly related parts of installations which could induce failure in the system being analyzed; for example, the effect of hydraulic fluid sprayed on electrical components as a result of a ruptured hydraulic line. Another example is the result of a ruptured bleed air line and its effect on hydraulic, fuel, or electrical lines or cables.
(v) The Type Inspection Authorization (TIA) should call for specific simulated failures, evaluation of failure detection, failure warning, and performance of the remaining system on the ground and in flight to verify the critical aspects of the failure analysis. The applicant should provide a proposed detailed test procedure for incorporation in the TIA to accomplish this verification. The applicant's proposed tests simulating in-flight failures should be carefully reviewed by both the systems engineer and flight test pilot to assure the flight test crew will not be subjected to hazardous flight.
Where practicable those simulated failures that would be hazardous in flight should be evaluated by ground tests. Analyzed and tested systems (where functioning is required) exhibiting hazards or failing to perform their intended functions under any foreseeable operating conditions must be redesigned to comply with § 27.1309.
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AC 27-1B, Chg 4 3/31/2014
(2) Utilization systems that are required or critical as to performance of intended function or result in rotorcraft hazard upon failure should also be analyzed for failures by the procedures of paragraphs c(1)(i) through c(1) (iv) above. Examples of systems that may be critical are autopilots, hydraulic control systems, navigation and flight instruments on IF-approved rotorcraft, and bleed air systems.
d. Safety Assessment. The international safety community has contributed to several documents providing guidance for conducting safety assessments for civil airborne systems and equipment. "Safety assessment" is terminology commonly used in the aerospace industry and addresses all aspects of fault, failure, or reliability analysis. Depending on the complexity and assessed criticality of the systems or equipment, the safety assessment can vary from a simple analysis which shows that there are no catastrophic, hazardous or severe-major, or major failure conditions, to a Functional Hazard Assessment (FHA) with the various analyses that may be required.
Various analytical techniques may be used to assist the applicant and the FAA/AUTHORITY in conducting the safety assessment for the determination of compliance with the requirements. These analytical techniques are intended to supplement but not replace engineering and operational judgment. A safety assessment can be accomplished by several different methods. One method used in past certifications (Failure Analyses) is addressed in paragraph c. above and is acceptable for showing compliance. Another method is the safety assessment process that contains the FHA addressed herein and is a methodology of general application to show compliance for systems, and all or parts of it may be applied as necessary depending on the complexity of design. The concept of safety assessment is provided in SAE document ARP 4754A,i Guidelines for Development of Civil Aircraft and Systems, and that aspect of this document is endorsed as an acceptable way of showing compliance to the requirements for function and design assurance. SAE Document ARP 4761, "Guidelines and Methods for Conducting the Safety Assessment Process on Civil Airborne Systems and Equipmenti can be used as a guide to perform the different analyses that form the safety assessment. The starting point is a FHA that identifies the assessed criticality of aircraft or systems functions. Further analysis is directed to be performed on the identified sources of these functions (i.e., systems) or part of systems performing these function(s)). One of the most common directed analysis is the reliability analysis.
(1) Functional Hazard Assessment. The FHA is a top down assessment that, on an aircraft level, starts at flight operations and identifies the criticality levels associated with the malfunction or failure of functions. The source of these functions can then be identified after completion of the FHA. There are five criticality levels defined in paragraph f. below, as they pertain to part 27. These level definitions are used in the FHA to decide which level matches the malfunctions or failures under consideration. The bottom-up analysis is then directed to the identified function source.
Reliability analysis is the most common of the bottom-up analysis.
(2) Reliability Analyses. Numerical Reliability Analyses may be developed, on an optional basis, as a continuation of the failure analysis procedure.
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(i) Specific reliability numbers are not shown in § 27.1309. The necessary degree of reliability is a function of the criticality of the system under consideration.
Acceptable sources of component failure rates are (1) military aircraft component failure or fault reports or handbooks, such as the latest revision of MIL-HDBK-217; (2) operator or manufacturer component malfunction or defect records, such as airline component defect records on sufficiently similar component designs; and (3) laboratory life tests.
(A) FLIGHT TIME (Block Time). The time from the moment the rotorcraft first moves under its own power for the purpose of flight until the moment it comes to rest at the next point of landing.
(B) PROBABILITY CLASSIFICATIONS. Five new probability classifications are defined below, although three probability classifications have been used in past certifications. This is not considered a problem as long as the conversion below is applied, and it is recognized that only the divisions within the same range are different. Quantitative ranges are also provided as a common point of reference if numerical probabilities are used. The quantitative ranges given for these classifications are considered to overlap due to the inexact nature of probability estimates. When assessing the acceptability of a failure condition using a quantitative analysis, the numerical ranges given below should normally be interpreted to be the allowable risk for an hour of flight time based on a flight of mean duration for the rotorcraft type.
However, when assessing a function which is used only at a specific time during a flight, the probability of the failure condition should be calculated for the specific time period and expressed as the risk for the flight condition, takeoff, landing, etc., as appropriate.
(1) FREQUENT. (This is the lower part of the range 105 or greater previously applied to the term "PROBABLE".) Frequent events may be expected to occur several times during the operational life of each rotorcraft that is based on a probability on the order of 103 or greater.
(2) REASONABLY PROBABLE. (This is the upper part of the range 10" or greater previously applied to the term "PROBABLE".) Probable events may be expected to occur several times during the operational life of each rotorcraft that is based on a probability on the order of between 103 to 105.
(3) REMOTE. ( The term "REMOTE" is not related to the structural use of the term.) (This is the lower part of the range 109 to 10-5 previously applied to the term "IMPROBABLE.") Remote events are not expected to occur during the total operational life of a random single rotorcraft of a particular type, but may occur during the total operational life of all rotorcraft of a particular type that is based on a probability on the order of between 105 to 107.
(4) EXTREMELY REMOTE. (The term "REMOTE" is not related to the structural use of the term.) (This is the upper part of the range 10° to 10-5 previously applied to the term "IMPROBABLE.") Extremely remote events are not Page F - 27
AC 27-1B, Chg 4 3/31/2014 expected to occur during the total operational life of a random single rotorcraft of a particular type, but may occur during the total operational life of all rotorcraft of a particular type that is based on a probability on the order of between 107 to 10-9.
(5) EXTREMELY IMPROBABLE. (Remains the same for both the three or the five classifications.) Extremely improbable events are so unlikely that they need not be considered to ever occur, unless engineering judgment would require their consideration. A probability on the order of 10 or less is assigned to this classification.
Notes: If a quantitative analysis is used to help show compliance with regulations for equipment which is installed and required only for a specific operating condition for which the rotorcraft is thereby approved, credit may not be taken for the fact that the operating condition does not always exist. Except for this limitation, appropriate statistical randomness of environmental or operational conditions may be considered in the analysis. (However, the particular condition and probability of that condition should be agreed to with the FAA/AUTHORITY.)
The five probability terms defined in paragraph d. (2)(ii)(B) above are intended to relate to an identified failure condition resulting from or contributed to by the improper operation or loss of a function or functions. These terms do not define the reliability of specific components or systems.
Generally, the guidance for Failure Analysis of paragraph c. above, is not required in its entirety for non-IFR rated rotorcraft. The only failure or reliability requirement is that no single failure can result in a hazard to the rotorcraft. This can usually be accomplished by a systems safety assessment that may or may not, depending on complexity and configuration, require a numerical reliability analysis.
e. Documentation. All laboratory tests, ground tests, flight tests, and failure analyses must be documented in sufficient detail to show compliance with § 27.1309 and included in the type design file. Section 21.31(a) provides the regulatory basis for requiring this documentation. If the applicant elects to use a numerical reliability and probability analysis it must also be documented in sufficient detail.
f. Software.
(1) RTCA Document DO-178C, "Software Considerations in Airborne Systems and Equipment Certification," dated December 13, 2011, is the latest standard and is recommended to be used for qualification and subsequent approval of airborne software. See AC 20-115C for guidance on using DO-178C and earlier standards.
DO-178C defines five levels of software (i.e., levels A, B, C, D, and E). The level of software is related to the criticality of the function that may be adversely affected by an error in the software. The criticality categories are as follows:
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(i) Catastrophic - failure conditions that would prevent continued safe flight and landing.
(ii) Hazardous or Severe-Major - failure conditions that would reduce the capability of the aircraft or the ability of the crew to cope with adverse operating conditions to the extent that there would be:
(A) A large reduction in safety margins or functional capabilities.
(B) Physical distress or higher workload such that the flight crew could not be relied on to perform their tasks accurately or completely.
(C) Adverse effects on occupants, including serious or potentially fatal injuries, to a small number of those occupants.
(iii) Major - failure conditions that would reduce the capability of the aircraft or the ability of the crew to cope with adverse operating conditions to the extent that there would be, for example: a significant reduction in safety margins or functional capabilities, a significant increase in crew workload or in conditions impairing crew efficiency, or discomfort to occupants, possibly including injuries.
(iv) Minor - failure conditions that would not significantly reduce aircraft safety, and would involve crew actions that are well within their capabilities. Minor failure conditions may include, for example: a slight reduction in safety margins or functional capabilities, a slight increase in crew workload, such as, routine flight plan changes, or some inconvenience to occupants.
(v) No Effect - failure conditions that do not affect the operational capability of the aircraft or increase crew workload.
(2) The software levels are usually related to the criticality categories. Level A qualified software is the most error-free software and is usually required for functions that could exhibit catastrophic failures. However, additional considerations may moderate this direct relationship and allow some lower level of software qualification for higher function criticality categories. Some of the moderating factors may be architecture of the system or software, redundancy of systems using dissimilar software, hardware or software monitors, and independent function contributions. It is recommended that these practices are carefully employed and prior approval of the methodology should be obtained before the design is pursued. Typically, Level A is required for flight controls where the catastrophic criteria applies. Level B qualification is less than Level A and is employed in some flight controls and required flight instruments, where their malfunction would result in the hazardous or severe major critical category effect on the rotorcraft or its crew and occupants. Level C qualification is less than Level B and is employed in some required flight instruments commensurate with failure category "Major." However, this lower level of qualification may be appropriate for either or both of the higher failure categories if the criteria can be met as Page F - 29
AC 27-1B, Chg 4 3/31/2014 discussed previously for level reduction. Level D qualification is less than Level C and is typically employed for required systems or equipment that do not exhibit the fault potential of the higher categories. Level E qualification is less than Level D and is employed in those systems or equipment that are not required by the regulations and whose malfunction results in no safety effect on the rotorcraft or its crew and occupants.
Examples are entertainment systems, powered seats, etc.
(3) Although a rough correlation exists between software levels and criticality levels that in turn relate to a probability in numbers, these numbers cannot be applied to software to determine reliability obtained by the processes delineated in DO-178C and at this time, have no correlation with probability.
g. Airborne Electronic Hardware (AEH). For airborne complex custom microcoded electronic components assessed as level A, B, or C in the safety assessment, RTCA Document DO-254, "Design Assurance Guidance for Airborne Electronic Hardware," dated April 19, 2000, provides certification guidance for qualification and subsequent approval of these components. Per DO-254, the component will need to be identified as either "simple" or "complex" and tagged to one of the five failure condition classifications. Based on these identifiers, DO-254 provides the guidance to show compliance to § 27.1309. Although RTCA/DO-254 applies specifically to complex custom micro-coded components, applicants are highly encouraged to apply DO-254 to other hardware components up through to the line replaceable unit and systems level.
AC 27.1309A. § 27.1309 (Amendment 27-46) EQUIPMENT, SYSTEMS, AND INSTALLATIONS.
a. Explanation. Amendment 27-46 removed § 27.1309(d) and incorporated a dedicated rule § 27.1316 to better address Electrical and Electronic System Lightning Protection.
b. Procedures. If the certification basis includes § 27.1309 (d), applicants should utilize the previous section of this guidance (section 27.1309 of this AC) to show compliance with protection from lightning effects. If the certification basis includes § 27.1316, applicants should show compliance with protection from lightning effects by referring to the applicable section of this AC (section 27.1316, Electrical and Electronic System Lightning Protection). The paragraphs other than b. (3) of section 27.1309 of this AC are still applicable. This should be coordinated with the appropriate certification office early in the discussions so the path to show compliance with electrical and electronic system lightning protection is clear to the applicant and authority.
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AC 27.1309B. § 27.1309 (AMENDMENT 27-51) EQUIPMENT, SYSTEMS, AND INSTALLATIONS.
a. Explanation. Amendment 27-51 revised § 27.1309 to address advances in technology and eliminate the distinction between multi-engine and single-engine rotorcraft. The rule now includes the same language as § 29.1309. The rule also:
• Revised § 27.1309 (a) to clarify the requirements for performing a failure analysis.
• Replaced § 27.1309 (b) and (c) with language consistent with contemporary industry practices.
b. Types of Equipment. Section 27.1309 covers, but is not limited to, electrical, pneumatic, and hydraulic power sources, distribution, and corresponding utilization systems.
(1) Applicability. Section 27.1309 is intended as a general requirement that is applicable to any equipment or system as installed, in addition to specific systems requirements, except as indicated below.
(i) General. If a specific part 27 requirement predefines systems safety aspects (e.g., redundancy level or criticality) for a specific type of equipment, system, or installation, then the specific part 27 requirement will take precedence. This precedence does not preclude accomplishment of a system safety assessment, if necessary. For example, § 27.695 predefines a required level of redundancy and an implied level of system reliability. However, a system safety assessment approach may still be required to show that the implied system reliability is met and to address assessment of the failure modes.
(ii) Section 27, Subparts B, C, and D. Section 27.1309 does not apply to the design features required by Subparts B, C, and D of part 27 for aspects such as the performance, flight characteristics, structural loads, and structural strength. However, the requirements in § 27.1309 do apply to any equipment and system that the proposed design uses to comply with the requirements of Subparts B, C, and D (e.g., Health Usage Monitoring System certified for maintenance credit and Stability Augmentation System).
(iii) Section 27, Subpart E.
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