(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.
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(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).
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(4) For rotorcraft that will be approved for IFR operation, the provisions of Appendix BVIII(b)(5) of Part 29 as discussed in Paragraph 775, should also be considered.
b. Procedures. The installation of the static system should consider the following:
(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 IFR 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/AUTHORITY 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 of this AC.
636. § 29.1327 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 § 29.1303(c) or the unit or system required for IFR operation by Appendix B VIlI(a) to Part 29. Both of these indicators provide the pilot with an aircraft heading which is referenced to the earth's magnetic field. The unit required by § 29.1303(c) is the indicator commonly referred to as a "whiskey compass." This 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 29 is usually a system of units which meets the requirements of TSO-C6c.
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b. Procedures. In showing compliance to § 29.1327 (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 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. The results of the investigation may be used to construct the calibration placard which is required by § 29.1547. 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., §§ 29.1301, 29.1309, 29.1555, etc.).
637. § 29.1329 (Amendment 29-24) AUTOMATIC PILOT SYSTEMS.
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 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 in § 29.1329 but is included under § 29.672. 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 782 of this AC discusses the use of automatic systems for Category I approaches, and Paragraph 775 of this AC discusses the evaluation of stability augmentation systems.
b. Procedures.
(1) General.
(i) 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 § 29.1329. 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.
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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 applicant is considering such a design, the applicant and the approving office should contact the Rotorcraft Standards Staff specialists for guidance.
(iii) 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 ensure adequate reliability. The question which often arises is: Should the preflight test function be interlocked so 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 tollowing items:
(i> 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 § 29.1322 (reference Paragraph 633 of this AC).
(ili) 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 § 29.779.
(v) Any disconnect of the autopilot should be annunciated.
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c. Malfunction Evaluations. To preclude hazardous conditions which may result from any failure or malfunctioning of the autopilot the following failures should be 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.
(i> A signal about any axis equivalent to the cumulative effect of any single failure, including autotrim (if installed).
(її) The combined signals about all affected axes, if multiple axes failures can result from the malfunction of any single component.
(2) Limit Loads. 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 deviation due to these 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 (i.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(2).
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.
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7/30/97 АС 29-2B An investigation should be made to determine the effects of an oscillatory signal of sufficient amplitude to saturate the servo amplifier of each device 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.
(ill) 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. To aid in recovery of the rotorcraft, after a malfunction occurs, one pilot should be able to physically overpower the autopilot and then disengage it with ease, and it should remain disengaged until further pilot action to reengage. 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 Paragraph c(2) 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.
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(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.
(2) The aircraft should be flown down the instrument landing system (ILS) in the configuration and at the approach speed specified by the applicant for approach.
Simulated autopilot malfunctions should be induced at critical points along the ILS, taking into consideration all possible variations in autopilot sensitivity and authority.
The malfunctions should be induced in each axis. While the pilot may know the purpose of the flight, the pilot should not be informed when a malfunction is about to be or has been applied except through aircraft action, control movement, or other acceptable warning devices.
(3) An engine failure during an automatic ILS approach should not cause a lateral deviation of the aircraft from the flight path at a rate greater than 3° per second or produce hazardous attitudes.
(4) If approval is sought for ILS approaches initiated with one engine inoperative, the automatic pilot should be capable of conducting the approach.
(5) Deviations from the ILS flight profile should be evaluated as follows:
The rotorcraft should be instrumented so the following information is recorded— (A) The path of the rotorcraft with respect to the normal glide path;
(B) The point along the glide path when the simulated malfunction is induced;
(C) The point where the pilot indicates recognition of the malfunction; and (D) The point along the path of the rotorcraft where recovery action is initiated.
(ii) Data obtained from the point of the indicated malfunction to the point where the rotorcraft has either again intersected the glide slope or is in level flight will define the deviation profile. When changes to the aircraft autopilot configuration are made during the approach and these changes alter the deviation profile, additional data should be obtained to define each of the applicable deviation profiles. An example of a deviation profile is shown in Figure 637-1.
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Recoveries from malfunctions should simulate under-the-hood instrument conditions with an appropriate time delay between pilot recognition of the fault and initiation of the recovery at all altitudes down to 80 percent of the minimum decision altitude for which the applicant requests approval.
(iv) Recoveries from malfunctions at altitudes between 80 percent of the minimum decision altitude for which the applicant requests approval and the minimum altitude for which the applicant requests approval to operate the autopilot may be visual with no time delay between pilot recognition of fault and initiation of recovery.
The minimum altitude at which the autopilot may be used should be determined as the altitude which results in the critical deviation profile becoming tangent with a minimum operational tolerance line. An example of this may be found in Figure 637-2. The 29:1 slope of the minimum operational tolerance line provides a 1 percent gradient factor of safety over the 50:1 obstacle clearance line. An additional factor of safety is provided by measuring the 29:1 slope from the horizontal at a point 15 feet above the runway threshold. It is recognized that this minimum altitude will vary with glide slope angle. Information regarding these variations should be obtained and presented.
(vi) A malfunction of the autopilot during a coupled ILS approach should not place the aircraft in an attitude which would preclude conducting a satisfactory go-around or landing.
e. Servo Authority. The automatic pilot system should be installed and adjusted so the system tolerances established during certification tests can be maintained in normal operation. This may be ensured by conducting flight tests at the extremes of the tolerances. Those tests conducted to determine that the automatic pilot system will adequately control the aircraft should establish the lower limit. Those tests to determine that the automatic pilot will not impose dangerous loads or deviation from the flight path should be conducted at the upper limit. Appropriate aircraft loadings to produce the critical results should be used.
f. Rotorcraft Flight Manual Information. The following information should be placed in the rotorcraft flight manual:
(1) In the Operating Limitations Section. Airspeed and other applicable operating limitations for use of the autopilot.
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AC 29-2B NOTE: Point of change of rotorcraft configuration may be more than one point. For instance:
1. Gain changes along the glide path.
2. The 200 ft. or middle marker transition.
Fault: Rotorcraft Configuration No. 1 Point of Change of Rotorcraft Configuration Fault: Rotorcraft Point of Change to Flare Configuration No. 2 Mode Projected Automatic Deceleration/Hover Deviation Mode (if Provided) Profile No. 1 Fault: Rotorcraft Deviation Profile No. 2 Configuration No. 3 Terrain Deviation Profile No. 3- Par 637 Runway 7/30/97 FIGURE 637-1 DEVIATION PROFILE
(2) In the Operating Procedures Section. The normal operation information.
(3) In the Emergency Operation Procedures Section.
(1) A statement of the downward flight path deviation in the cruise, climb, and descent configurations and the maneuvering flight configuration in accordance with Paragraphs d(5) (ili) and d(5) (iv) of this paragraph, if this deviation exceeds 100 feet.
(ii) True profiles of deviations below the glide slope or projected flare path for the critical conditions tested in accordance with Paragraphs d(5)(iv) (see Figure 627-1) and the deviation profile indicating the lowest altitude at which the autopilot can be used, as referenced in Paragraph d(5)(v), if applicable, and if this deviation exceeds 100 feet or excessive deviation for an ILS approach.
g. There should be a means of sequencing actions or interlocking engagement with sensor inputs to prevent autopilot initiated maneuvers that could result in hazardous operations due to:
(1) Engagement of the autopilot;
(2) Malfunctions of autopilot input or feedback signals that could result in unbounded output commands.
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