(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.
Definition of Environment. The SAE AE4L Committee report dated June 20, 1978, is an acceptable criteria to define the worst-case lightning strike which may be encountered by the rotorcraft in service. An additional explanation of the lightning environment may 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 Federal Aviation Regulations.
It presents a comprehensive test criteria to provide the essential information for the in-flight lightning protection of all types of fixed/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 March 5, 1990, FAA Advisory Circular AC 20-136, "Protection of Aircraft Electrical/Electronic Systems Against the Indirect Effects of Lightning," was issued.
Appendix 3 of that document contains an updated quantification of the severe natural lightning environment. Additionally, an AIR-100 policy letter dated August 25, 1993, addressing multiple burst lightning strikes, shoula be considered. It is recommended that for new designs and applications after March 5, 1990, this revised definition of the lightning be used.
(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 Par 621 817
АС 27-1A 7/30/97 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 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) Eull-Leve, 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/or 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 resistivities. 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 and/or electrical/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) 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 the recommended 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 nonrecoverable digital upset. One method to provide this consideration 818 Par 621
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. 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) For complete airframe verification testing, a minimum level of at least 4KA peak and a current rise time of 2KA/microsecond are recommended. It is often difficult to obtain valid results at lower levels due to poor signal-to-noise ratios.
When complete vehicle testing is accomplished at some lower level, or through some alternate test technique such as low level swept CW testing, consideration should be given to nonlinear airframe response, diffusion effects, and alterations in current paths caused by arcing and flashover.
(4) 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.
A level of 6 dB is recommended for those analyses which are confirmed by the use of reduced level, full-scale vehicle testing. This safety margin is the difference between the airframe installed system responses and the system component tolerance, not an adjustment to the quantification of the atmospheric environment. (The airframe system response to the worst-case lightning event should be at least 6 Bless than the FADEC system computer and components tolerance level. Number of dB is defined as 20 LOG10 (V1N2) and 20 LOG10 (11/l2) where V, and I, are the determined tolerance levels of the system components and V½ and 2 are the extrapolated airframe response.)
(5) 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. Approximately an additional 25 dB of protection has been deemed acceptable for a reasonably rigorous analysis performed on an airframe for which the response characteristics are known.
(E) Pass/Fail Criteria. The certification plan should address a pass/fail criteria for the testing and analysis to be performed. The following items should be satisfied to assure acceptable system performance:
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(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.
(G) After the system recovers, if the performance of the system has been degraded in a noncritical manner which would reduce the capability of the 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. Information should be made available 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 926A, revised November 15, 1979.
(ії) 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.
(ili) 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.
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(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/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.
(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)(1) through c(1) (iv) above. Examples of systems which may be critical are autopilots, hydraulic control systems, navigation and flight instruments on FR approved rotorcraft, and bleed air systems.
d. Documentation. All laboratory, ground, and 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/probability analysis it must also be documented in sufficient detail.
e. Computer Software. The latest standard for qualification of software is DO-178B; however, the use of DO-178A for a standard is not precluded at this time.
Because of this dual standard situation, at this time, use of both standards will be addressed for qualification of software that is used for airborne systems and equipment certification.
(1) RTCA Document DO-178A RTCA Document DO-178A, "Software Considerations in Airborne Systems and Equipment Certification," dated March 22, 1985, is a recommended standard to be used for the approval of system software. This document defines three levels of software; i.e., levels 1, 2, and 3. The level of the software is related to the consequence of a system malfunction caused by an error in the software. The criticality categories are:
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(A) Critical - Functions for which the occurrence of any failure condition or design error would prevent the continued safe flight and landing of the aircraft.
(B) Essential - Functions for which the occurrence of any failure condition or design error would reduce the capability of the aircraft or the ability of the crew to cope with adverse operating conditions.
(C) Nonessential - Functions for which failures or design errors could not significantly degrade aircraft capability or crew operational cue.
(ii) The different software levels may be related to the criticality categories. Level 1 software, the most error free software, is usually required for critical functions. However, Level 1 software may sometimes be reduced by system architecture techniques such as the use of redundant (dissimilar) software performing the same function. Level 2 software is required for essential functions. It should be noted that those systems, equipment, and installations, with functioning required by 14 CFR subchapter C, are by this definition, essential functions. The criticality of the function should be determined by the use of a fault/failure or hazard analysis. The Society of Automotive Engineers Aeronautical Recommended Practice Document Nos. 926A and 1834 are the recommended references for performing these analyses.
CAVEAT: The user of DO-178A is cautioned by a caveat in Chapter 3 that for a certain class of systems, the techniques in DO-178A, Level 1 software are not by themselves sufficient consideration for reliance on system software to preclude a catastrophic event. Additional considerations are required with this class of system for software verification and validation (V&V) in addition to those required for DO-178A Level 1. This class of systems is one which has been called, "full flight regime critical."
An example of such a system is a fly-by-wire flight control. This system must perform its intended function through the full flight regime to provide for the continued safe flight and landing of the rotorcraft. For this system, software and system level validation beyond the scope of DO-178A are required. Also, DO-178A cautions the user against the assignment of probabilities of residual software errors. The conclusion of Special Committee No. 152 (The RTCA committee that wrote DO-178A) was that the present methods available for assigning "reliability" numbers to software do not yield credible results fo certification purposes.
(2) RTCA Document DO-178B
(i) RTCA Document DO-178B, "Software Considerations in Airborne Systems and Equipment Certification," dated December 1, 1992, is the latest standard and is recommended to be used for qualification and subsequent approval of system software. This document defines five levels of software; i.e., levels A, B, C, D, and E.
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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:
(A) Catastrophic - failure conditions that would prevent continued safe flight and landing.
(B) Hazardous 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:
(1) A large reduction in safety margins or functional capabilities.
(2) Physical distress or higher workload such that the flight crew could not be relied on to perform their tasks accurately or completely.
(3) Adverse effects on occupants, including serious or potentially fatal injuries, to a small number of those occupants.
(C) 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.
(D) 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.
(E) No Effect - failure conditions that do not affect the operational capability of the aircraft or increase crew workload.
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/software, redundancy of systems using dissimilar software, hardware/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/severe major critical Par 621 823
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 discussed previously for level reduction. Level D qualification is less than Level C and is typically employed for required systems/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/equipment that are not required by the regulations. Examples are entertainment systems, powered seats, etc.
(III) 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-178B and at this time, have no correlation with probability.
f. High Intensity Radiated Fields (HIRF
(1) Explanation. A regulatory project is active to add requirements for the protection of aircraft electrical and electronic systems from the effects of the HIRF environment. This effort is the result of technological advances in airframe and electronic systems design and a concurrent increase in the levels of radiated power in the aircraft environment. These changes have raised vulnerability to the electromagnetic environment of the electrical and electronic systems which perform critical and essential functions. In current type certification programs involving advanced electrical and electronic systems the authority has adopted special conditions to provide an adequate level of safety.
The special conditions are directed toward the operation and operational capability of the installed electrical and electronic systems that perform critical functions. The applicant may demonstrate that these systems are not adversely affected when the aircraft is exposed to the HIRF environment, or as an alternative a laboratory test may be conducted, as discussed in the "Discussion" associated with each special condition. The laboratory tests would be conducted at a peak electromagnetic strength of 100 or 200 volts per meter, as appropriate, in a frequency range of 10 KHz to 18 GHz.
A definition of the HIRF environment is included in an FAA Aircraft Engineering Division Memorandum dated December 5, 1989, (Subject: High Energy Radiated Electromagnetic Fields (HERF) Interim Policy Guidelines on Certification Issues). SAE chartered a HIRF committee, SAE-4R, to define the HIRF environment for VFR and IFR aircraft, to prepare a Users Manual, and to submit a proposed advisory circular to the FAA. This committee has currently completed its task, except for the Users Manual and the rotorcraft VFR environment definition. The referenced FAA memo has been updated several times to reflect the latest SAE-4R recommended levels. The current revision of this memo is dated July 29, 1992.
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(ill) If the laboratory test alternative is selected the 100 volts/meter level is considered appropriate for a function that is critical during IFR operations and the 200 volts/meter level is considered appropriate for a function that is critical during VFR operations. This is because the minimum en route altitude for IFR flight is 1,000 feet or 500 feet (FAA/AUTHORITY or ICAO), and rotorcraft operating VFR can and do operate regularly at lower altitudes. The attitude system is an example of a system performing a critical function during IFR operation. A full authority digital engine control (FADEC) system is an example of a critical function during VFR and IFR operation.
(2) Procedures. It is recommended that the applicant present a plan to the cognizant FAA Aircraft Certification Office (ACO) for approval, outlining how the compliance with the HIRF requirements will be attained. This plan should also propose a pass/fail criteria for the operation of critical systems in the HIRF environment.
(i> A preliminary hazard analysis should be performed by the applicant for approval by the cognizant FAA ACO to identify electrical and/or electronic systems that perform critical functions. The term "critical functions" means those whose failure would prevent the continued safe flight and landing of the rotorcraft.
(ії) The systems performing critical functions that are identified by the preliminary hazard analysis are candidates for the application of HIRF requirements. A system may perform both critical and non-critical functions; however, the HIRF requirements only apply to critical functions. If redundant systems are used, all systems should be subjected to test/analysis for the HIRF requirements.
(iii) The latest revision of RTCA-DO-160, Section 20 is an appropriate reference for laboratory test procedures. In addition a separate advisory circular and users guide on the subject of HIRF is being drafted for the FAA by the SAE AE4R Subcommittee.
622.-631. RESERVED.
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SECTION 35. INSTRUMENTS: INSTALLATION
632. § 27,1321 (Amendment 27-13) ARRANGEMENT AND VISIBILITY.
a. Background. Part 27 contains specific requirements for instruments to allow any pilot to operate the rotorcraft safely within authorized limits and to indicate system conditions. The instruments should be arranged for use by any pilot and must be readily visible to the pilot.
b. Procedures.
(1) When evaluating the location of the instruments for rotorcraft approved for VFR (IFR locations are discussed in Paragraph 775), the flight, navigation, and powerplant instruments should be placed such that the pilots) can easily see and read the instruments when seated normally. Additionally, the instruments should be located so that the necessity for the pilot's head to turn is minimized.
(2) On multiengine rotorcraft, the required powerplant instruments should be grouped so there will be no confusion regarding which engine an individual gage represents. This is usually accomplished by mounting the engine gages in a vertical grouping. Identical parameter gages are placed next to each other and positioned from left to right in the same position and sequence as the engine location in the airframe.
(3) An evaluation should be made to determine that vibration of the instrument panel does not exceed the tolerances of the instrument. The instrument manufacturer will usually provide data which indicate the level of vibration for which the instrument has been qualified. The flight test evaluation of the rotorcraft should explore and determine that the vibration of the instrument panel does not affect the readability of the instrument. To meet these two criteria, it has been necessary in some installations to "shock mount" or otherwise isolate the instrument panel.
(4) The flight test evaluation should also determine that the flags or malfunction indicators of the instruments are readily visible in all combinations of lighting for approved kinds of operations.
633. § 27.1322 (Amendment 27-11) WARNING, CAUTION, AND ADVISORY LIGHTS.
a. Explanation.
(1) Cockpit devices are color-coded to symbolically represent various functions and varying levels of importance for flightcrew operation. From early times, an attempt has been made to take full advantage of associations developed early in life as a result of continuous exposure to our daily environment.
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(2) Military design specifications were the first to reference color-coding in cockpit design requirements. In the mid-1940s, the CAA initiated the first color-coding requirements for civil cockpit design. Color-coding standards for cockpit visual signals soon followed. MIL-STD-411, May 31, 1957, identified three separate categories of light signals:
Warning Light - indicates the existence of a hazardous condition which may require immediate corrective action.
(#1) Caution Light - serves to alert the operator to an impending dangerous condition requiring attention but not necessarily immediate action.
(iii) Advisory Light - indicates safe or normal configuration, condition of performance, operation of essential equipment, or attracts attention for routine purposes.
(3) Examples of warning and caution signals are included in later versions of the military standard, and a few of those are shown below:
Warning Signals Caution Signals Cabin Pressure Failure Trim Failure Fire Fuel Low Fuel System Failure Generator Inoperative Landing Gear Unsafe Defrosting Failure
(4) Specific color designation for civil advisory lights was first addressed in Amendment 3 to the rotorcraft certification rules (Parts 27 and 29) on January 19, 1968, with adoption of new §§ 27.1322 and 29.1322.
(5) In a subsequent revision (Amendment 27-11), green lights were redesignated and additional colors introduced for flexibility in the requirement.
b. Procedures.
(1) Red shall be reserved for annunciation of emergency conditions requiring immediate corrective action. Typical examples include fire, transmission oil pressure, engine failure, and battery overheat. The use of red for annunciators which do not require immediate action must be avoided. Use of red when it is not needed tends to lessen the impact of a red annunciator and the needed pilot association for immediate action. In evaluating cockpit annunciators for acceptability, the FAA should ensure all annunciators which require immediate action are red and that only those requiring such action are red. If a master warning light is provided, it should be red, and it should be powered by the same signal that powers any of the individual red warning signals. An aural warning may accompany visual warning signals to enhance pilot response. Care Par 633 837
АС 27-1A 7130/97 should be taken that any aural signal is sufficiently distinct from other aural warnings, such as low rotor RPM to prevent confusion and to ensure proper crew response. A means to deactivate and reset the master warning (visual and aural) is required.
Resetting the master warning must not deactivate any individual warning signal.
(2) Amber shall be reserved for indicating malfunction or failure conditions which do not require immediate crew action to ensure safe flight. Typical examples include door unlatched, inverter failure, generator failure, fuel filter clogged, and parking brake engaged. Amber should generally be utilized for malfunction and failure conditions which do not require immediate action. The key word here is "require."
Obviously, a pilot should perform corrective action for malfunction or failure conditions in a timely manner as soon as other cockpit priorities allow. The time increment associated with "immediate action" may vary with the system involved, the flight regime, and the aircraft; however, 15 seconds is a representative value in evaluating this term.
This by no means indicates that any red annunciator can be ignored for 15 seconds.
For red annunciators, some type of immediate pilot response is expected. If immediate pilot action is not required, the FAA should recommend the use of an amber designation. If a master caution light is provided in addition to a master warning light, the master caution annunciator should be amber, and should be powered by the same signal that powers any of the individual amber caution signals. Reset considerations for the master caution are the same as those detailed above for the master warning.
(3) Green signifies a safe operating condition and more specifically has come to signify landing gear extended and locked. Extensive use of green annunciators throughout the cockpit should generally be avoided due to possible confusion with the special use of green for landing gear. If green annunciators are physically and functionally removed from the landing gear operation, they may be found acceptable for a variety of "safe operating" applications. One such application is "all green for approach" used in autopilot, flight director, and other navigation system displays.
(4) Other colors may be utilized as advisory lights in accordance with § 27.1322(d). Red and amber must not be used as advisory lights due to the possibility of introducing confusion into the cockpit. Obviously, yellow and pink annunciators should be avoided due to their similarity to amber and red. White and blue have been successfully utilized as advisory segments in past civil designs.
(5) The primary test for designation of color is:
(1) Red - Is immediate action required?
(II) Amber - Is pilot action (other than immediate) required?
(iii) Green - Is safe operation indicated, and is the indication sufficiently distinct to prevent confusion with the landing gear down indication?
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