category effect on the rotorcraft or its crew and occupants. Level C qualification is less than Level B and is employed in some required flight instruments commensurate with failure category "Major." However, this lower level of qualification may be appropriate for either or both of the higher failure categories if the criteria can be met as discussed previously for level reduction. Level D qualification is less than Level C and is typically employed for required systems/equipment that do not exhibit the fault potential of the higher categories. Level E qualification is less than Level D and is employed in those systems/equipment that are not required by the regulations. Examples are entertainment systems, powered seats, etc.
(III) Although a rough correlation exists between software levels and criticality levels that in turn relate to a probability in numbers, these numbers cannot be applied to software to determine reliability. obtained by the processes delineated in DO-178B and at this time, have no correlation with probability.
f. High Intensity Radiated Fields (HIRF
(1) Explanation. A regulatory project is active to add requirements for the protection of aircraft electrical and electronic systems from the effects of the HIRF environment. This effort is the result of technological advances in airframe and electronic systems design and a concurrent increase in the levels of radiated power in the aircraft environment. These changes have raised vulnerability to the electromagnetic environment of the electrical and electronic systems which perform critical and essential functions. In current type certification programs involving advanced electrical and electronic systems the authority has adopted special conditions to provide an adequate level of safety.
The special conditions are directed toward the operation and operational capability of the installed electrical and electronic systems that perform critical functions. The applicant may demonstrate that these systems are not adversely affected when the aircraft is exposed to the HIRF environment, or as an alternative a laboratory test may be conducted, as discussed in the "Discussion" associated with each special condition. The laboratory tests would be conducted at a peak electromagnetic strength of 100 or 200 volts per meter, as appropriate, in a frequency range of 10 KHz to 18 GHz.
A definition of the HIRF environment is included in an FAA Aircraft Engineering Division Memorandum dated December 5, 1989, (Subject: High Energy Radiated Electromagnetic Fields (HERF) Interim Policy Guidelines on Certification Issues). SAE chartered a HIRF committee, SAE-4R, to define the HIRF environment for VFR and IFR aircraft, to prepare a Users Manual, and to submit a proposed advisory circular to the FAA. This committee has currently completed its task, except for the Users Manual and the rotorcraft VFR environment definition. The referenced FAA memo has been updated several times to reflect the latest SAE-4R recommended levels. The current revision of this memo is dated July 29, 1992.
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(ill) If the laboratory test alternative is selected the 100 volts/meter level is considered appropriate for a function that is critical during IFR operations and the 200 volts/meter level is considered appropriate for a function that is critical during VFR operations. This is because the minimum en route altitude for IFR flight is 1,000 feet or 500 feet (FAA/AUTHORITY or ICAO), and rotorcraft operating VFR can and do operate regularly at lower altitudes. The attitude system is an example of a system performing a critical function during IFR operation. A full authority digital engine control (FADEC) system is an example of a critical function during VFR and IFR operation.
(2) Procedures. It is recommended that the applicant present a plan to the cognizant FAA Aircraft Certification Office (ACO) for approval, outlining how the compliance with the HIRF requirements will be attained. This plan should also propose a pass/fail criteria for the operation of critical systems in the HIRF environment.
(i> A preliminary hazard analysis should be performed by the applicant for approval by the cognizant FAA ACO to identify electrical and/or electronic systems that perform critical functions. The term "critical functions" means those whose failure would prevent the continued safe flight and landing of the rotorcraft.
(ії) The systems performing critical functions that are identified by the preliminary hazard analysis are candidates for the application of HIRF requirements. A system may perform both critical and non-critical functions; however, the HIRF requirements only apply to critical functions. If redundant systems are used, all systems should be subjected to test/analysis for the HIRF requirements.
(iii) The latest revision of RTCA-DO-160, Section 20 is an appropriate reference for laboratory test procedures. In addition a separate advisory circular and users guide on the subject of HIRF is being drafted for the FAA by the SAE AE4R Subcommittee.
622.-631. RESERVED.
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SECTION 35. INSTRUMENTS: INSTALLATION
632. § 27,1321 (Amendment 27-13) ARRANGEMENT AND VISIBILITY.
a. Background. Part 27 contains specific requirements for instruments to allow any pilot to operate the rotorcraft safely within authorized limits and to indicate system conditions. The instruments should be arranged for use by any pilot and must be readily visible to the pilot.
b. Procedures.
(1) When evaluating the location of the instruments for rotorcraft approved for VFR (IFR locations are discussed in Paragraph 775), the flight, navigation, and powerplant instruments should be placed such that the pilots) can easily see and read the instruments when seated normally. Additionally, the instruments should be located so that the necessity for the pilot's head to turn is minimized.
(2) On multiengine rotorcraft, the required powerplant instruments should be grouped so there will be no confusion regarding which engine an individual gage represents. This is usually accomplished by mounting the engine gages in a vertical grouping. Identical parameter gages are placed next to each other and positioned from left to right in the same position and sequence as the engine location in the airframe.
(3) An evaluation should be made to determine that vibration of the instrument panel does not exceed the tolerances of the instrument. The instrument manufacturer will usually provide data which indicate the level of vibration for which the instrument has been qualified. The flight test evaluation of the rotorcraft should explore and determine that the vibration of the instrument panel does not affect the readability of the instrument. To meet these two criteria, it has been necessary in some installations to "shock mount" or otherwise isolate the instrument panel.
(4) The flight test evaluation should also determine that the flags or malfunction indicators of the instruments are readily visible in all combinations of lighting for approved kinds of operations.
633. § 27.1322 (Amendment 27-11) WARNING, CAUTION, AND ADVISORY LIGHTS.
a. Explanation.
(1) Cockpit devices are color-coded to symbolically represent various functions and varying levels of importance for flightcrew operation. From early times, an attempt has been made to take full advantage of associations developed early in life as a result of continuous exposure to our daily environment.
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(2) Military design specifications were the first to reference color-coding in cockpit design requirements. In the mid-1940s, the CAA initiated the first color-coding requirements for civil cockpit design. Color-coding standards for cockpit visual signals soon followed. MIL-STD-411, May 31, 1957, identified three separate categories of light signals:
Warning Light - indicates the existence of a hazardous condition which may require immediate corrective action.
(#1) Caution Light - serves to alert the operator to an impending dangerous condition requiring attention but not necessarily immediate action.
(iii) Advisory Light - indicates safe or normal configuration, condition of performance, operation of essential equipment, or attracts attention for routine purposes.
(3) Examples of warning and caution signals are included in later versions of the military standard, and a few of those are shown below:
Warning Signals Caution Signals Cabin Pressure Failure Trim Failure Fire Fuel Low Fuel System Failure Generator Inoperative Landing Gear Unsafe Defrosting Failure
(4) Specific color designation for civil advisory lights was first addressed in Amendment 3 to the rotorcraft certification rules (Parts 27 and 29) on January 19, 1968, with adoption of new §§ 27.1322 and 29.1322.
(5) In a subsequent revision (Amendment 27-11), green lights were redesignated and additional colors introduced for flexibility in the requirement.
b. Procedures.
(1) Red shall be reserved for annunciation of emergency conditions requiring immediate corrective action. Typical examples include fire, transmission oil pressure, engine failure, and battery overheat. The use of red for annunciators which do not require immediate action must be avoided. Use of red when it is not needed tends to lessen the impact of a red annunciator and the needed pilot association for immediate action. In evaluating cockpit annunciators for acceptability, the FAA should ensure all annunciators which require immediate action are red and that only those requiring such action are red. If a master warning light is provided, it should be red, and it should be powered by the same signal that powers any of the individual red warning signals. An aural warning may accompany visual warning signals to enhance pilot response. Care Par 633 837
АС 27-1A 7130/97 should be taken that any aural signal is sufficiently distinct from other aural warnings, such as low rotor RPM to prevent confusion and to ensure proper crew response. A means to deactivate and reset the master warning (visual and aural) is required.
Resetting the master warning must not deactivate any individual warning signal.
(2) Amber shall be reserved for indicating malfunction or failure conditions which do not require immediate crew action to ensure safe flight. Typical examples include door unlatched, inverter failure, generator failure, fuel filter clogged, and parking brake engaged. Amber should generally be utilized for malfunction and failure conditions which do not require immediate action. The key word here is "require."
Obviously, a pilot should perform corrective action for malfunction or failure conditions in a timely manner as soon as other cockpit priorities allow. The time increment associated with "immediate action" may vary with the system involved, the flight regime, and the aircraft; however, 15 seconds is a representative value in evaluating this term.
This by no means indicates that any red annunciator can be ignored for 15 seconds.
For red annunciators, some type of immediate pilot response is expected. If immediate pilot action is not required, the FAA should recommend the use of an amber designation. If a master caution light is provided in addition to a master warning light, the master caution annunciator should be amber, and should be powered by the same signal that powers any of the individual amber caution signals. Reset considerations for the master caution are the same as those detailed above for the master warning.
(3) Green signifies a safe operating condition and more specifically has come to signify landing gear extended and locked. Extensive use of green annunciators throughout the cockpit should generally be avoided due to possible confusion with the special use of green for landing gear. If green annunciators are physically and functionally removed from the landing gear operation, they may be found acceptable for a variety of "safe operating" applications. One such application is "all green for approach" used in autopilot, flight director, and other navigation system displays.
(4) Other colors may be utilized as advisory lights in accordance with § 27.1322(d). Red and amber must not be used as advisory lights due to the possibility of introducing confusion into the cockpit. Obviously, yellow and pink annunciators should be avoided due to their similarity to amber and red. White and blue have been successfully utilized as advisory segments in past civil designs.
(5) The primary test for designation of color is:
(1) Red - Is immediate action required?
(II) Amber - Is pilot action (other than immediate) required?
(iii) Green - Is safe operation indicated, and is the indication sufficiently distinct to prevent confusion with the landing gear down indication?
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Other advisory lights - Is the meaning clear and distinct enough to prevent confusion with other annunciations? Do the colors which are utilized differ sufficiently from the colors specified in Paragraphs b(1), (2), and (3) above?
(6) 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.
(7) If dimming capability is provided, all annunciators, including master warning and caution, may be dimmable so long as the annunciation is clearly discernable 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.
(8) 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 falls, the light will illuminate and burn steadily.
(9) 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.
(10) Refer to Paragraph 779 of this AC, Annunciator Panels, for additional design information.
634. § 27.1323 (Amendment 27-13) 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 Par 633 839
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, etc. 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. 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.
(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.
635. § 27.1325 (Amendment 27-13) STATIC PRESSURE SYSTEMS.
a. Explanation.
(1) This section, in conjunction with § 27.1323, 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 § 27.1309(a).
(3) The standard for consideration of malfunction conditions is § 27.1309(b).
(4) For rotorcraft that will be approved for FR operation, the provisions of Appendix BVIII(b)(5) of Part 27 as discussed in Paragraph 775 should also be considered.
b. Procedures. The installation of the static system should consider the following:
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(1) Static lines should be initially routed upward immediately behind the static pressure port. This procedure will minimize the entry of moisture into the system when operating in rain or washing the rotorcraft.
(2) Drain(s) should be located at low points in the system. Line routing and clamping should allow for all moisture that does enter the system to be routed to the drain(s).
(3) If independent systems are provided, the placement of each system component should allow for maximum practicable separation of each system. As much as possible, one system should be on one side of the rotorcraft and the second system on the opposite side.
(4) Most static pressure ports that are provided for IF operation are heated.
Before any tests are conducted, a program to qualify the heater on the port should normally be agreed upon through discussions between the FAA and the applicant. It is suggested that the requirements of TSO C16, Airspeed Tubes (Heated), be used as a guide for these discussions. If the ports are not to be heated, a comprehensive analysis should be prepared, and limited testing should be conducted to verify the analysis.
(5) Other static system considerations are included in Paragraphs 634 and 775 ot this AC.
636. § 27.1327 (Amendment 27-13) MAGNETIC DIRECTION INDICATOR.
a. Background. This section contains specific requirements regarding installation and functioning of a magnetic direction indicator. The magnetic direction indicator (commonly referred to as a compass) described by this paragraph is the unit required by § 27.1303(c) or the unit or system required for IFR operation by Appendix BVIII (a) to Part 27. Both of these indicators provide the pilot with an aircraft heading which is referenced to the earth's magnetic field. The unit required by § 27.1303(c) is the indicator commonly referred to as a "whiskey compass." The unit was given this designation because early units were constructed using alcohol as the medium in which the compass ball floats. This unit is generally approved as meeting the requirements of TSO-C7c. The indicator required by Appendix B to Part 27 is usually a system of units which meets the requirements of TSO-C6c.
b. Procedures. In showing compliance to § 27.1327(a), generally the magnetic indicator and its respective components will be tested to an appropriate standard such as RTCA DO 160B for use in a rotorcraft. If the unit functioned properly as described in the TSO during this testing, then no additional evaluation is generally required concerning vibration immunity. To determine the immunity of the indicator (system) from magnetic effects and its installed accuracy, a ground and flight test should be Par 635 841
performed. This test should turn the rotorcraft a full 360° heading change in 45° increments. The indicator should not have an error in excess of 10° on any of the 45° increments. When performing these tests, the electrical equipment and systems should be functioning normally, and the effect of windshield heating (if installed) should be investigated. During this investigation, it is permissible for the error caused by the effect of the operation of electrical equipment on the indicator required by § 27.1303(c) to exceed 10° if a gyroscopic heading indicator is installed. This gyroscopic heading indicator may either be a directional gyro or a slaved compass system (gyro-stabilized magnetic direction indicator). The results of the investigation may be used to construct the calibration placard which is required by § 27.1547. It should be noted that a calibration placard has not been traditionally required for slaved compass systems.
Also, it should be emphasized that other aspects of the functioning and installation of these indicators should comply with the other general requirements (i.e., §§ 27.1301,
27.1309, 27.1555, etc.).
637. § 27.1329 (Amendment 27-21) AUTOMATIC PILOT SYSTEM.
a. Explanation. The automatic flight control systems used on most modern rotorcraft often perform two different and distinct functions when viewed from a regulatory compliance aspect. These two functions are an augmentation of the stability of the rotorcraft and a pilot aid in maintaining attitude, altitude, and airspeed, or in radio navigation tasks. The first function of stability is not covered by § 27.1329 but is included under § 27.672. The second function as a pilot aid is the automatic pilot function covered by this section. The following procedure discusses only those parts or systems which are installed as a pilot aid. Paragraph 775 of this AC discusses the evaluation of stability augmentation systems.
b. Procedures.
(1) General.
The automatic pilot system should be evaluated to demonstrate that it can perform its intended function of flying the rotorcraft and that it complies with the installation, operation, and malfunction requirements of § 27.1329. In demonstrating malfunctions of the autopilot system, generally servo actuator hardovers are the most critical malfunction. If this is the case and the autopilot system utilizes the same servos and servo amplifiers as the stability augmentation system (SAS) and the autopilot function cannot produce a more severe hardover than the SAS, then no additional consideration is required for this malfunction. An evaluation using the guidance in Paragraph 775 of this AC would be sufficient.
(ii) There have been autopilots approved which require the use of a monitor since they cannot meet the hardover malfunction requirements. These approvals have involved a finding of equivalent safety which is beyond the scope of this AC. Such findings of equivalent safety are made on a case-by-case basis. If an 842 Par 636
applicant is considering such a design, the applicant and the approving office should contact the Rotorcraft Standards Staff specialists for guidance.
(i) The rule specifies that unless there is automatic synchronization, there should be some method to indicate the alignment of the actuating device to the pilot. The intent of this requirement is to provide a means such that the pilot does not inadvertently engage the system into a hardover condition. One method of achieving this has been the use of servo force meters. These meters monitor the current into the servo motor and indicate to the pilot if a signal is being sent to the servo prior to system engagement.
(iv) Various autopilot systems have used a preflight test to assure adequate reliability. The question which often arises is: Should the preflight test function be interlocked so that the autopilot cannot be engaged if the preflight test has not been accomplished? The guidance used in the past to answer this question is: If the preflight test is simple and rapid enough that the pilot may reasonably be expected to perform such a test, then it is not required to be interlocked. If, however, the preflight test is very complicated and lengthy and a pilot who was pressed by a schedule might skip such a test, then this preflight test should be interlocked.
(v) Most of the autopilots which have been approved utilize series actuators or servos such as those required for a SAS. However, this does not preclude the approval of an autopilot which uses outer loop parallel actuation. This type of autopilot may be particularly helpful in a VFR aircraft.
(2) Cockpit controls. Evaluation of the cockpit controls should include the following items:
Location of the automatic pilot system controls are such that their operation is properly labeled and is readily accessible to the pilots).
(ii) Annunciator colors conform to the colors specified in § 27.1322 (reference Paragraph 633 of this AC).
(iii) A determination is made that the controls, control labels, and placards are readable and discernible under all expected cockpit lighting conditions.
(iv) Motion and effect of the autopilot cockpit controls should conform with the requirements of § 27.779.
(v) Annunciation should be provided if the autopilot disconnects for any reason other than pilot action.
c. Malfunction Evaluations. To preclude hazardous conditions which may result from any failure or malfunctioning of the autopilot the following failures should be Par 637 843
evaluated. This evaluation should also account for any hazards which also might be caused by inadvertent pilot action. The guidance in Paragraph 775 of this AC should be used to determine the appropriate reaction times of the human pilot to an autopilot malfunction.
(1) Climb, cruise, and descent flight regimes. The more critical of the following should be induced into the automatic pilot system.
A signal about any axis equivalent to the cumulative effect of any single failure, including autotrim (if installed).
(ii) The combined signals about all affected axes, if multiple axes failures can result from the malfunction of any single component.
(2) The simulated failure and the subsequent corrective action should not create loads in excess of structural limits or result in dangerous dynamic conditions or deviations from the flight path. Additional guidance regarding the method of determining pilot recognition times and reasonable flight path deviations due to those simulated failures is contained in Paragraph 775b(6) of this AC. Resultant flight loads outside the envelope of zero to 2g will be acceptable provided adequate analysis and flight test measurements are conducted to establish that no resultant aircraft load is beyond limit loads for the structure, including a critical assessment and consideration of the effects of structural loading parameter variations (l.e., center of gravity, load distribution, control system variations, maneuvering gradients, etc.). Analysis alone may be used to establish that limit loads are not exceeded where the aircraft loads are in the linear range of loading (i.e., aerodynamic coefficients for the flight condition are adequately established and no significant nonlinear air loadings exist). If significant nonlinear effects could exist, flight load survey measurements may be necessary to substantiate that the limit loads are not exceeded. The power for climb should be the most critical of: (1) that used in the performance climb demonstrations; (2) that used in the longitudinal stability tests; or (3) that actually used for operational climb speeds.
The altitude loss should be measured.
(3) Maneuvering Flight. Malfunctions should also be induced into the automatic pilot system similar to Paragraph c(1). When corrective action is taken, the resultant loads and speeds should not exceed the values contained in Paragraph c(1).
Maneuvering flight tests should include turns with the malfunction induced when maximum bank angles for normal operation of the system have been established and in the critical aircraft configuration and/or stages of flight likely to be encountered when using the automatic pilot. The altitude loss should be measured.
(4) Oscillatory Tests.
An investigation should be made to determine the effects of an oscillatory signal of sufficient amplitude to saturate the servo amplifier of each device 844 Par 637
that can move a control. The investigation should cover the range of frequencies which can be induced by a malfunction of the automatic pilot system and systems functionally connected to it, including an open circuit in a feedback loop.
(її) The results of this investigation should show that the peak loads imposed on the parts of the aircraft by the application of the oscillatory signal are within the limit loads for these parts.
(iii) The investigation may be accomplished largely through analysis with sufficient flight data to verify the analytical studies or largely through flight tests with analytical studies extending the flight data to the conditions which impose the highest percentage of limit load to the parts.
(iv) When flight tests are conducted in which the signal frequency is continuously swept through a range, the rate of frequency change should be slow enough to permit determining the amplitude of response of any part under steady frequency oscillation at any critical frequency within the test range.
(5) Recovery of Flight Control. Recovery of the rotorcraft should be accomplished by the pilot by first overpowering the malfunctioning autopilot and then disconnecting it. The control to disconnect the autopilot should be easily available to the pilot who is now resisting the malfunctioning force of the autopilot. It is recommended that the disconnect button be placed on the cyclic control. It should be red and conspicuously marked "Autopilot Disconnect." The pilot should be able to return the rotorcraft to its normal flight attitude under full manual control without exceeding the loads or speed limits defined in this paragraph and without engaging in any dangerous maneuvers during recovery. The maximum servo authority used for these tests should not exceed those values shown to be within the structural limits for which the rotorcraft was designed. The maximum altitude loss experienced during these tests should be measured.
(6) External Interfaces. The autopilot system should have appropriate interlocks to its engagement to ensure it does not operate improperly as a result of information furnished by an external device or system. An example of this is the navigation receivers and the compass system. If for a particular mode of operation the autopilot uses signals from these systems, the autopilot should be interlocked from operating in those modes if invalid information is being received from that system.
d. Automatic Pilot Instrument Approach Approval.
(1) Throughout an approach, no signal or combination of signals simulating the cumulative effect of any single failure or malfunction in the automatic pilot system, except vertical gyro mechanical failures, should provide hazardous deviations from flight path or any degree of loss of control.
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