АС 29-2B 7/30/97 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 derogate 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.
(4) Lightning Protection.
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, because the early aircraft were constructed mostly of metal and had little, if any, dependence on advanced technology systems. Contemporary design transport category rotorcraft are utilizing the same advanced technology systems and materials as transport category airplanes. Because of this fact, a specific requirement has been added by Amendment 29-24 for the consideration of lightning strike protection of required systems, equipment, and installations. The addition of Paragraph (h) to § 29.1309 further defines the consideration required for the foreseeable operating condition of a lightning strike encounter on the rotorcraft.
(її) Procedures.
(A) Section 29.1309(h) requires, when showing compliance with § 29.1309(a) and (b), that the effects of lightning strikes on the rotorcraft be considered.
(1) 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.
(2) 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.
(3) The lightning strike models to be used for system justification should be as described in SAE AE4L Committee Report AE4L-87-3, Rev. B, dated January 1989 (or later version). The recommended reference for performing such analysis is Society of Automotive Engineers, Aerospace Recommended Practice 926A.
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(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 50 is not recommended for full scale low level pulse testing, due to the difficulty of obtaining valid results with poor signal-to-noise ratios.
(3) For a proven analysis technique, a safety factor of at least 2 will be necessary.
(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 § 29.1309(a); i.e., 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 29.610 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 § 29.1309(a). These avionic systems thus identified usually only require protection from indirect effects of lightning. If it is determined such is the case, then a method as outlined in Paragraph 621b(3), Lightning Strike Protection of Full Authority Digital Engine Controls, 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 if the malfunctioning of several "essential" systems in relation to other systems would result in a hazard to a Category B rotorcraft or preclude the continued safe flight and landing of a Category A rotorcraft. If groups of functions are so identified, sufficient lightning protection should be provided to prevent a hazardous malfunction situation on Category B rotorcraft or provide those conditions which prevent continued safe flight and landing on a Category A rotorcraft are extremely improbable. In performing this part of the analysis, attention should be given to the fact that many of the required equipment, systems, and installations may fail simultaneously with other required equipment, systems, and installations and result in a reduction of the capability of the rotorcraft but still not result in a catastrophe. An example of required equipment for which the simultaneous failure of all the required equipment is catastrophic is a failure which results in a total loss of attitude display for IFR certified rotorcraft operating in instrument meteorological conditions. Note that the analysis which is utilized to demonstrate that these failures are extremely improbable should have the encounter with a worst-case lightning strike as a given event; i.e., probability is unity. Additionally.
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for a Category A rotorcraft, an autorotation is not considered continued safe flight and landing.
c. Failure Analyses.
(1) Power and distribution systems should be analyzed to show compliance with § 29.1309.
One acceptable procedure for documenting the analysis is contained in Society of Automotive Engineers (SAL) Fault/Failure Analysis Procedure ARP 926A, revised November 15, 1979.
(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.
(¡i) 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/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 § 29.1309.
(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)(a) through c(1)(d) above. Examples of systems which may be critical are autopilots, hydraulic control systems, navigation and flight instruments on IFR approved rotorcraft, and bleed air systems.
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d. Reliability Analyses. Numerical Reliability Analyses may be developed, on an optional basis, as a continuation of the failure analysis procedure.
(1) Specific reliability numbers are not shown in § 29.1309. 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 service records or handbooks, such as MIL-HDBK-217C, (2) operator or manufacturer service records, such as airline records on sufficiently similar component designs, and (3) laboratory life tests.
(2) For the purpose of conducting or evaluating an analysis, the following terms and numerical values should apply:
(i) 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.
(ii) PROBABILITY CLASSIFICATIONS. Three probability classifications are defined below. 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.
(A) PROBABLE. Probable events may be expected to occur several times during the operational life of each rotorcraft. A probability on the order of 105 or greater.
(B) IMPROBABLE. Improbable 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. A probability on the order of 10 or less.
(C) EXTREMELY IMPROBABLE. 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 10-9 or less.
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NOTE:
If a quantitative analysis is used to help show compliance with Federal Aviation 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 three probability terms defined in Paragraph d(2)(ii) 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.
A new SAE document (ARP 4754), "Systems Integration Requirements" is being created. This document will completely change the present probability classifications from three to five. This change is necessary because the latest software standard DO178B has five levels of software qualification. While the caution that software levels cannot be expressed in reliability terms is still valid, there is need for some correlation between criticality levels (probability classification) and software levels. Since DO178B has five software qualification levels, it follows that corresponding criticality levels are appropriate. Additionally, another SAE document (ARP 4761) "Safety Assessment Guidelines for Civil Airborne System and Equipment", is in the early stage of writing. This document will address the safety assessment process and provide guidance for the entire system evaluation. Generally, the guidance for Failure Analysis of Paragraph d, is not required in its entirety for Category B rotorcraft. The only failure/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, ground and flight tests, and failure analyses, must be documented in sufficient detail to show compliance with § 29.1309 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.
f. Computer Software. The latest standard for qualification of software is DO178B; however, the use of DO178A 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.
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(1) RTCA Document DO178A
(i) 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.
(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) Par 621 963
was that the present methods available for assigning "reliability" numbers to software do not yield credible results for 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.
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 pertorm their tasks accurately or completely.
(3) Adverse effects on occupants, including serious or potentially tatal 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 964 Par 621
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 FAA/AUTHORITY 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 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.
(ill) 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. The quality assurance of software is obtained by the processes delineated in DO178B and at this time, have no correlation with probability.
g. 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 airtrame 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 FAA/AUTHORITY has adopted special conditions to provide an adequate level of satety.
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 Par 621 965
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/AUTHORITY. 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.
(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/tall criteria for the operation of critical systems in the HIRF environment.
(1) 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.
(ii) 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.
(lI) The latest revision of RTCA document 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/AUTHORITY by the SAE AER Subcommittee.
622.-631. RESERVED.
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SECTION 35. INSTRUMENTS: INSTALLATION
632. § 29.1321 (Amendment 29-21) ARRANGEMENT AND VISIBILITY.
a. Background. This section is the first in a series which concerns the installation of instruments. Specific requirements for individual instruments are addressed in other paragraphs. The instruments should be arranged in a manner such that the pilot may avail himself of the information displayed by the instruments without undue distraction.
Additionally, for instrument flight, the rule requires that the attitude, altitude, airspeed, and compass indicators be grouped in the so-called standard "T."
b. Procedures.
(1) For rotorcraft certified for VFR operation, the flight, navigation, and powerplant instruments should be placed such that the pilot and copilot, if a required crewmember, can easily see and read these instruments when seated normally.
Additionally, the instruments should be located so that the necessity for the pilot to turn his head is minimized. The instruments which are necessary for safe operation including the airspeed indicator, gyroscopic direction indicator, gyroscopic bank-and-pitch indicator, slip-skid indicator, altimeter, rate-of-climb indicator, rotor tachometers, and the indicator most representative of engine power should be installed immediately in front of the pilot.
(2) The other powerplant instruments should be grouped together and visible to any appropriate crewmember. On multiengine rotorcraft, there should be no confusion regarding which engine an individual gauge represents. This is usually accomplished by mounting the engine gauges vertically in the center of the instrument panel.
Identical gauges are placed next to each other and positioned from left to right in the same position and sequence as the corresponding engine location in the airframe.
(3) An evaluation should determine that vibration of the instrument panel does not exceed the tolerances of the instruments. 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 instruments.
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 should be readily visible in all combinations of lighting for approved kinds of operations.
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(5) For IFR-certified rotorcraft, there is an additional requirement that the airspeed, altimeter, attitude, and compass instrument be located in a standard "T" configuration in front of the pilot. This configuration is:
airspeed - attitude - altitude compass (Paragraph 775 further addresses IFR panel arrangement.)
(6) Geometric variation from a perfect "T" has been permitted. Each installation should be evaluated for suitability based on criteria such as panel size, ease of scan, and readability of the individual elements in the overall presentation. Advisory Circular 25-11, Transport Category Airplane Electronic Display Systems, provides additional guidance for "glass cockpit" installation.
633. § 29.1322 (Amendment 29-12) WARNING, CAUTION, AND ADVISORY LIGHTS
a. Explanation.
(1) Cockpit devices are color coded to symbolically represent various functions and varying levels of importance for flight crew 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.
(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:
(i) Warning Light - indicates the existence of a hazardous condition which may require immediate corrective action.
(ii) Caution Light - serves to alert the operator to an impending dangerous condition requiring attention but not essential equipment, or attracts attention for routine purposes.
(iii) Advisory Light - indicates safe or normal configuration, condition of performance, or operation of essential equipment, or attracts attention for routine purposes.
(3) Examples of warning and caution signals were included in later versions of the military standard, and a few of those are shown below:
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Warning Signals Caution Signals Cabin Pressure Failure Trim Failure Fire Fuel Low Fuel System Failure Generator Inoperative Landing Gear Unsate 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 subsequent revision (Amendment 29-12), green lights were redesignated and additional colors introduced for flexibility in the requirement.
(6) 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.
(7) Other colors may be utilized as advisory lights in accordance with § 29.1322(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.
(8) The primary test for designation of color is:
(1) Red - Is immediate action required:
(ii) Amber - Is pilot action (other than immediate) required?
(li) Green - Is safe operation indicated, and is the indication sufficiently distinct to prevent confusion with the landing gear down indication?
(iv) Other advisory lights - Is the meaning clear and distinct enough to prevent confusion with other annunciators? Do the colors which are utilized differ sufficiently from the colors specified in Paragraphs (b)(1), (2), and (3) above?
(9) Annunciator lights should be visible during bright daylight conditions. This should include visibility in direct sunlight unless lights are located in such a manner that direct sunlight cannot impinge on them.
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