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.
Par 621 961
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.
962 Par 621
(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.
966 (thru 976) Par 621
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.
Par 632 977
(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:
978 Par 632
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.
Par 633 979
(10) If dimming capability is provided, all annunciators, including master warning and caution, may be dimmable so long as the annunciation is clearly discernible for night operation at the lower lighting level. Undimmed annunciations have been found unacceptable for night operation due to disruption of cockpit vision at the high intensity. The dimming circuit should automatically revert to the high intensity setting when power is removed. Automatic dimming/brightening through the use of a photo cell is also acceptable, as are circuits which enable a dimming switch through a position light or other cockpit lighting controls.
(11) The use of flashing lights should be minimized. If a flashing feature is used, it should be controllable through pilot action so that flashing annunciation does not persist indefinitely. The indicator should be so designed that if it is energized and the flasher device fails, the light will illuminate and burn steadily.
(12) The activation of caution and warning lights should readily attract the attention of the appropriate crewmember while performing duties under both normal and high workload conditions.
(13) Refer to Paragraph 779 of this AC, Annunciator Panels, for additional design information.
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/AUTHORITY should assure 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 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 should be taken that any aural signal is sufficiently distinct from other aural warnings, such as low rotor RPM, to prevent confusion and to assure 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 assure 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 980 Par 633
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/AUTHORITY 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.
634. § 29.1323 (Amendment 29-3) AIRSPEED INDICATING SYSTEM.
a. Explanation.
(1) The accuracy of all flight test data concerned with the velocity of the rotorcraft is dependent on the calibration of the airspeed indicating system. For this reason, the airspeed system position error should be determined very early in the program.
(2) Since air density varies with altitude, the speed reading will only be correct under standard sea level conditions. However, in an actual installation, the indicator reading, even under standard sea level conditions, may differ from the calibrated airspeed because the static system does not sense true static pressure. This error in detection of static pressure is called position error. It is caused by the pressure field built up around the rotorcraft in flight. This pressure field will vary in intensity with dynamic pressure making the position error a function of calibrated airspeed. Since airspeed information is presented to the crew in terms of indicated airspeed, it is necessary to determine the position error for the rotorcraft to be flown safely.
b. Procedures.
(1) There are different methods to determine position error such as trailing bomb, airspeed course, boom system, and so forth. Each method has its own advantages and disadvantages, but will yield satisfactory results if done correctly. The airspeed system should be calibrated throughout the airspeed range of the rotorcraft and under the various flight conditions of cruise, climb, and autorotation standard. In addition, the effects of gross weight and center of gravity should be investigated.
(2) It may also be necessary to recalibrate the system with a change in external configuration if such a change may affect the airflow near the pitot or static sources.
(3) Additional information regarding position error is included in Paragraph 775b(10) and should be considered if pursuing an IFR approval.
Par 633 981
(4) Static system installation information is included in Paragraph 635.
Technical Standard Order (TSO) C16, Airspeed Tubes (Heated), gives minimum performance standards for pitot tubes, and pitot tubes qualified to this TSO normally allow for a satisfactory aircraft installation.
(5) The calibration requirements of the standard seem to be self-explanatory and are not discussed further in this paragraph.
634A. § 29.1323 (Amendment 29-24) AIRSPEED INDICATING SYSTEM.
a. Explanation. Amendment 29-24 to the regulations provides the requirements for Category A and Category B and defines the maximum allowable error for both.
b. Procedures. All of the policy material pertaining to this section remains in effect. In addition, calibration should be determined in level flight speeds of 20 knots and greater, and over an appropriate range of speeds for flight conditions of climb and autorotation; and takeoff. The takeoff calibration should be repeatable with respect to field lengths defined in the flight manual and avoidance of height-speed limiting envelope defined in § 29.79. Calibration errors, excluding instrument errors, may not exceed the following:
(1) Category A - Three percent or 5 knots, whichever is greater, in level flight at speeds above 80 percent of takeoff safety speed; and 10 knots in climb at speeds from 10 knots below takeoff safety speed to 10 knots above Vy.
(2) Category B - Three percent or 5 knots, whichever is greater, in level flight at speeds above 80 percent of the climb-out speed attained at 50 feet when complying with § 29.63.
635. § 29.1325 (Amendment 29-24) STATIC PRESSURE SYSTEMS.
a. Explanation.
(1) This section, in conjunction with § 29.1323, provides minimum performance standards for static pressure systems. The standard provides some relief when considering the icing environmental condition in that it allows the use of an alternate static port to account for the icing condition.
(2) The standard for the consideration of environmental conditions is § 29.1309(a).
(3) The standard for consideration of malfunction conditions is § 29.1309(b).
982 Par 634