(ii) Annunciator colors conform to the colors specified in § 29.1322 (reference section 29.1322 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 § 29.779.
(v) Any disconnect of the autopilot should be annunciated.
c. Malfunction Evaluations. To preclude hazardous conditions that may result from any failure or malfunctioning of the autopilot the following failures should be evaluated.
This evaluation should also account for any hazards that also might be caused by inadvertent pilot action. The guidance in Appendix B 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.
(i) Recovery from malfunctions should simulate instrument conditions, or visual flight conditions, depending on the category of certification that is sought.
Justification should be provided for recognition (e.g., audio or visual warning, excess deviation alert, or acceleration cues in the case of a hardover). Continuous close monitoring of the flight attitude instruments by the pilot may not be relied upon as a reliable means for detecting low rate attitude deviations (typically < 3°/sec) and thus for determining the point at which slowover recognition occurs. In such cases, analysis should be employed to establish the recognition criteria for the particular helicopter and flight phase, and the acceptability of the recovery. For each flight regime, the maximum height loss recorded for all malfunction testing should be established. The applicant should ensure that sufficient data is generated to substantiate a height loss figure that can be used for an operational determination of a minimum use height, where appropriate.
(A) For cruise, the height loss is defined as the difference between the aircraft altitude at the time the failure is introduced, and the minimum altitude achieved in the recovery, taking into account appropriate pilot delays as discussed above.
(B) For a descent, the height loss should be determined as illustrated in Figure AC 29.1329-1. The evaluation should consider the maximum rate of descent approved for hands-off, IFR, operation.
(C) For approach without vertical guidance, the height loss should also be determined as illustrated in Figure AC 29.1329-1, but with consideration to the critical approach angle and nominal approach speed.
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. . .... ' · . ' · .... ....... .
Height Loss .... ·, . . ·, .... ·< .... ....... . ·, .... ·,
1. Failure Initiation
2. Failure Recogn ition by pilot
3. Initiation of Manual Recovery act ion by pilot ...........
·-....: .... ....... . ·, .... ·, _. ....
....... . . ....... ....
.... .... .... .... .... .... .... .... ....
AC 29-2C, Chg 4 5/1/2014 (D) For approach with vertical guidance, the height loss should be determined as discussed in paragraph (d) below.
(ii) The more critical of the following should be induced into the automatic pilot system.
(A) A signal about any axis equivalent to the cumulative effect of any single failure, including autotrim (if installed).
(B) The combined signals about all affected axes, if multiple axes failures can result from the malfunction of any single component.
- - FIGURE AC 29.1329-1 HEIGHT LOSS METHOD Limit Loads. The simulated fa ilure and the subsequent corrective action should n
(2) ot create loads in excess of structural limits or result in dangerous dynamic conditions or hazardous 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 b.(6) of Appendix B 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 Page F - 54
(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.
(3) Maneuvering Flight. Malfunctions should also be induced into the automatic pilot system similar to paragraph c.(1) above. When corrective action is taken, the resultant loads and speeds should not exceed the values contained in paragraph c.(2) above. Maneuvering flight tests should incl ude 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.
(i) 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 that 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.
(ii) 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. To aid in recovery of the rotorcraft, after a malfunction occurs, one pilot should be able to physically overpower and disengage the autopilot with ease, and the autopilot should remain disengaged until further pilot action to reengage. The control to disconnect the aut opilot 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 Page F - 55
AC 29-2C, Chg 4 5/1/2014 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) above 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 al titude loss experienced during these tests should be measured.
(6) External Interfaces. The autopilo t 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.
(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:
(i) The rotorcraft should be instrumented so the following information is recorded–– (A) The path of the rotorcraft with respect to the normal glide path;
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(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 AC 29.1329-2.
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AC 29-2C, Chg 4 5/1/2014 NOTE: Point of change of rotorcraft configuration may be Projected Automatic Deceleration/Hover Mode (if Provided) more than one point For instance: 1. Gain changes along the glide path. 2. The 200 ft. or middle marker transition. Point of Change to Flare Fault: Rotorcraft - Configuration No. 3 Deviation Protle No. 3 Runway
- Mode Point of Change of Rotorcraft Configuration Fault Rotorcraft Configuration No. 2 Deviation Profile No. 2 Fault: Rotorcraft Configuration No. 1 Deviation Profile No. 1 Terrain FIGURE AC 29.1329-2 DEVIATION PROFILE Page F - 58
(iii) 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.
(v) The minimum altitude at which the autopilot may be used should be determined as the altitude that results in the critical deviation profile becoming tangent with a minimum operational tolerance line. An example of this may be found in Figure AC 29.1329-3. The 29:1 sl ope 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 re cognized that this minimum altitude will vary with glide slope angle. Information regarding these variations should be obtained and presented.
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' ' . '. '.. ·g~ er _1cal F It. P . n ~ Mm f o Oper • 00 of· .a,~ . . ' .
' AC 29-2C, Chg 4 5/1/2014 FIGURE AC 29.1329-3 DEVIATION PROFILE
(vi) A malfunction of the autopilot during a coupled ILS approach should not place the aircraft in an attitude that would preclude conducting a satisfactory go-around or landing.
e. . The automatic p ilot system should be installed and adjusted so the syst Servo Authority em 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 autom atic pilot system will adequately control the aircraft should establish the lower limit. Thos e 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. Appr opriate aircraft loadings to produce the critical results should be used.
Page F - 60
f. Rotorcraft Flight Manual Information. The followi ng information should be placed in the rotorcraft flight manual:
(1) In the Operating Limitations Section: Airspeed, Maximum Height Loss following AFCS malfunction for each phase of flight, Minimum Use Height where appropriate, and other applicable operating limitations for use of the autopilot.
(2) In the Operating Procedures Section: Normal operation information.
(3) In the Emergency Operation Procedures Section:
(i) 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)(iii) and d(5)(iv) above, if this deviation exceeds 100 feet.
(ii) True profiles of deviations below the glide slope or projected flare path for the critical conditions tested (see paragraphs d.(5)(iv) above and Figure AC 29.1329-2) and the deviation profile indicating the lowest altitude at which the autopilot can be used(see paragraph d.(5)(v) above, 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.
AC 29.1329A. § 29.1329 (Amendment 29- 42) AUTOMATIC PILOT SYSTEM.
a. Explanation. Amendment 29-42 adds paragraph (f) to § 29.1329, which requires a means of indicating to the pilots the current mode of operation for those automatic pilot systems that can be coupled to the airborne navigation equipment.
b. Procedures. The policy material pertain ing to section 29.1329 of this AC remains in effect with the following additions:
(1) Mode annunciation must indicate the state of the system including mode change and disengagement. Mode annunciati on should be presented in a manner compatible with flightcrew procedures and tasks and be consistent with other flight deck systems’ mode annunciations. Mode selector switch position or status is not acceptable as the sole means of mode annunciation. Modes and mode changes should be depicted in a manner that achieves flightcrew attention and awareness.
Page F – 60.1
AC 29-2C, Chg 4 5/1/2014
(2) Mode annunciations must effectively and unambiguously indicate the active and armed modes of operation. The mode annunciation should convey explicitly, and as simply as possible, what the system is doing (for active modes), what it will be doing (for armed modes), and target information (such as selected speed, heading, and altitude) for satisfactory flightcrew awareness. The pilot must be alerted to any deviation from the pilot-commanded target.
Page F – 60.2
AC 29.1329B. § 29.1329 (AMENDMENT 29-59) AUTOMATIC PILOT AND FLIGHT
GUIDANCE SYSTEM
a. Explanation. Amendment 29-59 changed the section heading name to “Automatic pilot and flight guidance system” and combined the requirements for automatic pilot (14 CFR 29.1329) and flight director systems (14 CFR 29.1335) into one rule (14 CFR 29.1329). The amendment also removed 14 CFR 29.1335.
b. Procedures. This section of the AC is updated to include the guidance previously found in AC 29-2C Section 29.1335. Therefore, the guidance pertaining to section 29.1329 of this AC remains in effect with the following additions:
(1) Three-cue flight directors for rotorcraft use the usual pitch and roll command cues with the third cue displayed on the left side of the attitude director indicator (ADI). These instruments can be used in either the two-axes or three-axes modes. In either mode, the lateral command cue controls the roll attitude, and the vertical command cue controls the pitch attitude. The rotorcraft attitude, controlled by the cyclic control, is changed to satisfy the flight director commands. The third cue, when displayed, commands collective pitch position and is used when an airspeed or pitch attitude mode and a vertical mode (altitude hold, glide slope, etc.) are selected.
(2) The general convention for flight director design is that each command bar is a “fly to” command. The motion of the flight director indicator is such to command a corresponding sense of control system motion. This is true of flight director pitch and roll commands and should hold true for additional commands such as collective pitch.
(3) Some consideration should be given to the collective, or third cue, display.
For example, if the collective symbol is selected as the fixed index, the command cue and collective pitch control should move in opposite directions when collective pitch changes are made. This configuration would constitute a conventional “fly to” indicator.
If the collective symbol is selected for the movable index, the direction of motion of the collective symbol will coincide with the direction of collective pitch changes. In this case the moving collective symbol does not comply with the “fly to” convention; however, this configuration has been approved by the FAA/ with special symbology, special background effects, and special color coding and has performed satisfactorily in service.
(4) The recommended display for a three-cue flight director incorporates the standard pitch and roll command symbols: either pitch and roll bars or the “V” bar display. The third cue, or collective symbol, should be located on the left side of the ADI.
The shape of the moving cue and the background display should be unique to avoid it being confused with a glide slope display or angle of attack display. One display uses a Page F - 60.2a
third cue, shaped like a small handle, to aid in identifying it as the collective pitch symbol.
(5) The color of the pitch and roll command indicators, the aircraft symbol, the background marking of the third cue, and third cue itself should be consistent. The optimum color scheme uses the same color for the aircraft symbol and the collective symbol. This is usually fire orange. The command cues including the collective cue also should use the same color, usually yellow. The rationale for the different colors is that the aircraft symbol and the collective symbol (the same color) are moved toward their respective command cues. If the pitch command cue is above the center, the aircraft symbol is raised (nose pulled up) and, if the collective command cue is above the collective symbol, the collective pitch is raised, moving the collective symbol towards the command cue.
(6) If the ADI provides a monochromatic display, the collective pitch cue and its background markings must, per 14 CFR 29.1321, be distinctive to reduce the chance of them being confused with the glide slope indicator. This can be accomplished through the use of different shaped cues and background marks. A round cue with a chevron-shaped background marking has been satisfactory.
Page F - 60.2b