Tmin = Minimum temperature to which the installed equipment has been tested in the laboratory.
Ttest min = Minimum installed equipment temperature recorded during tests.
Tor = The low reference outside air temperature. It varies with altitude starting at the lowest sea level temperature at which rotorcraft operation is to be approved and decreases at 3.6° F/1,000-foot altitude.
T1 mar = Temperature margin between the minimum equipment temperature substantiated in the laboratory and the minimum installed equipment temperature. If the margin is zero or positive, the equipment passes. If the margin is negative, the equipment fails the test.
T1 mar = - (Tmin - (T test min + (Tort - Toat min))) NOTE: This equation assumes all temperatures are negative. It is necessary to place a (-) in front of the right side of the equation in order to convert the T1 mar value to the conventional positive answer for acceptance and a negative answer for rejection. Temperature in the zero to 32° F range can be handled by conversion to the centigrade scale.
Example #2: Assume the applicant is seeking a low temperature operating limit at sea level of -25° F. Assume the hydraulic control cylinder has been substantiated in the laboratory to operate at a cylinder temperature of -40° F.
The in-flight test data are:
Measured Altitude (ft. MSL) Cylinder Temp (Trest min) OAT = Toat max sea level -0° F -25° F 5,000 -9° F -45° F 10,000 -21° F -59° F 15,000 -32° F -65° F 20,000 -40° F -69° F (Tor must be calculated for each altitude test point) @ sea level, Tor = -25° F @ 5,000 ft., Tor = -25° F -5,000 ft., x 3.6°/1,000 ft. = -43° F @ 10,000 ft., Tor = -25° F - 10,000 ft. × 3.6°/1,000 ft. = -61° F @ 15,000 ft., Tor = -25° F - 15,000 ft. x 3.6°/1,000 ft. = -79° F* @ 20,000 ft., Tori = -25° F - 20,000 ft. x 3.6°/1,000 ft. = -97° F* 812 Par 621
7/30/97 • AC 27-1A According to § 27.1043(b), the lowest temperature to be considered is -69.7° F.
Then at sea level:
T1 mar = -(-40 -(0 + (-25 - (-25)))) = 40° F At 5,000 feet:
T1 mar = -(-40 - (-9 + (-43 - (-45)) = 33° F At 10,000 feet:
T1 mar = (40 - (-21 + (-61 - (-59))) = 17° F At 15,000 feet:
T1 mar = -(-40 - (-32 + (-69.7 - (-65)))) = 3.3° F At 20,000 feet:
T1 mar = (40 - (-40+ (-69.7 - (-69)))) =-0.7° F It can be seen that there is an acceptable margin at all altitudes up to and including 15,000 feet. However, at 20,000 feet, the margin is negative and the system fails.
(ii) Vibration tests. Normally, installed vibration tests are not necessary for equipment qualified in accordance with the latest revision of RTCA document DO-160. This paper categorizes vibration tests according to installed rotorcraft equipment location such as fuselage, engine compartment, instrument panel, equipment rack, etc. However, installed equipment vibration tests may be necessary when it appears the equipment location environment may exceed the laboratory-tested equipment vibration limits.
(il) Altitude tests. If the equipment has been laboratory tested to the maximum certified altitude of the rotorcraft, installed altitude tests are unnecessary.
The installed equipment must be either laboratory tested or tested in the rotorcraft to the maximum certified altitude of the rotorcraft.
(3) Lightning Protection.
(i) Background. During the original design and development of rotorcraft and the development of regulations concerning these aircraft, little attention was given to protection from the meteorological phenomenon of lightning. This was, in part, Par 621 813
because the early aircraft were constructed mostly of metal and had little, if any, dependence on advanced technology systems. Contemporary design normal category rotorcraft are utilizing the same advanced technology systems and materials as airplanes. Because of this fact, a specific requirement has been added by Amendment 27-21 for the consideration of lightning strike protection of required systems, equipment, and installations. The addition of Paragraph (d) to § 27.1309 further defines the consideration required for the foreseeable operating condition of a lightning strike encounter on the rotorcraft.
(її) Procedures.
(A) Section 27.1309(d) requires when showing compliance to § 27.1309(a) and (b) that the effects of lightning strikes on the rotorcraft must be considered. The first step in demonstrating compliance is to perform a fault/failure analysis (F/FA) to identify those functions for which the loss of function or malfunction may result in a catastrophe to the rotorcraft. An F/FA should be conducted on each system whose failure to function properly would prevent the continued safe flight and landing of the rotorcraft. These systems should be designed and installed to ensure that they can perform their intended function during and after exposure to lightning.
Additionally, evaluation must be performed to identify each system whose failure to function properly would reduce the capability of the rotorcraft or the ability of the flight crew to cope with adverse operating conditions. These systems should be designed and installed to ensure that they can perform their intended function after exposure to lightning.
The lightning strike models to be used for system justification should be as described in SAE AE4L Committee Report AE4L-87-3, Revision B, dated January 1989 or later version. The recommended reference for performing such analysis is Society of Automotive Engineers, Aerospace Recommended Practice 926A.
(B) Detailed means of compliance should be agreed with the authorities taking into account the effects on the rotorcraft and minimum considerations are as follows:
(1) Any combination of analysis and testing should be agreed with the authority.
(2) For test results, an extrapolation of the threat current parameters of more than a factor of 10 is not likely to be acceptable without an additional safety factor being applied.
(3) For a proven analysis technique, a safety factor of at least 2 will be necessary.
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(C) Flight and engine controls are examples of "critical" functions and with these critical functions defined, an analysis and/or testing should be performed to show compliance with § 27.1309(a); i.e., equipment, systems, and installations performing those identified functions equipment, systems, and installations performing those identified functions should be designed and installed to ensure that they continue to perform their intended functions considering the conditions of the rotorcraft experiencing a worst-case lightning strike encounter. 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). A method as outlined in Paragraph 621b(4), "Lightning Strike Protection of Full Authority Digital Engine Controls,i 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 F/FA 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.
(i> Explanation.
(A) The following discussion is written specifically for full authority digital engine controls (FADEC) with an alternate technology backup fuel control installed on rotorcraft with Part 29 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 (LSI), very large scale integration (VLSI), and complementary metallic oxide silicon (CMOS) 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 electrical transient, ceases proper operation in accordance with its embedded software while suffering no apparent component or device damage.
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АС 27-1A 7/30/97 (B) Since elements of electrical/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 condition" 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 OEI ratings are approved, cumulative loss of power from all engines must be limited to allow flight manual performance based on OEl ratings.
(її) 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 which describes the analytical procedures and/or 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 tests).
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(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