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.
986 Par 637
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.
Par 637 987
(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.
988 Par 637
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.
Par 637 989
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.
Par 637 991
7/30/97 AC 29-2B ILS Glide Slope Critical Fault Point, Minimum Height for Operation of Autopilot 29:1 Minimum Operational Tolerance - Deviation Profile 50:1 Obstacle Clearance Plane → 15 ft Terrain .... Runway FIGURE 637-2 OPERATIONAL LIMITATION Par 637 992
638. § 29.1331 (Amendment 29-24) INSTRUMENTS USING A POWER SUPPLY.
a. Explanation. The rule concerns each flight instrument using a power supply that is installed in a Category A rotorcraft. A reference to Paragraph 618 will give a listing of the flight instruments that are specifically required for certification. The discussion included in this paragraph is directed toward electrical instruments since these are the type normally installed in Category A rotorcraft. It should be noted, however, that the rule is not restricted to electrical instruments. Further, the discussion provided here can be used to evaluate nonelectrical applications.
b. Procedures.
(1) This requirement must be considered when designing the electrical distribution system for Category A rotorcraft. It states that each required flight instrument must have two independent sources of power and a means of selecting either source. The flight instruments required for certification are listed in Paragraph 618, and independent power sources are discussed in Paragraph 654b(4).
(2) Some older flight instruments may not have integral visual means to indicate that adequate power is being supplied to the instrument as required by the rule.
For these instruments, external annunciation has been accepted that monitors the presence of adequate voltage at the power pin on the electrical disconnect that mates with the electrical connector on the back side of the instrument. The annunciator light should be located in close proximity to the indicator and placarded to identify its function. Note that the rule requires the voltage monitored to be within the approved limits for the instrument to be adequate. Since relay coils normally operate well outside the approved instrument voltage limits, the use of a relay contact that closes when the monitored voltage drops low enough to pull in or release a relay coil does not normally result in a satisfactory design to meet the regulatory requirement. Annunciator lights provided for this application are normally red.
(3) The power supply system requirements of this rule should be coordinated with the requirements of § 29.1355 (see Paragraph 654). Both rules concern equipment or systems that require two independent sources of electrical power.
Examples of faults in the distribution system to be considered include open feeders, shorted feeders, shorted busses, etc.
(4) Amendment 29-24 revised the regulation to further clarify the power adequacy indication requirement. The clarification provided was intended to make it easier to understand the meaning of adequate power in the event it was necessary to provide separate annunciation. The application of the rule in each form (before and after Amendment 29-24) should be the same.
Par 638 993
639. § 29.1333 (Amendment 29-24) INSTRUMENT SYSTEMS.
a. Explanation. Prior to Amendment 29-24, this requirement was titled Duplicate Instrument Systems," and its provisions were intended to be applied when duplicate flight instruments were required by any operating rule. Due to the increased complexity of instrumentation that is available and being used, it was considered appropriate to revise the provisions of this requirement to more appropriately consider the extreme range of operation environments to which rotorcraft are now routinely exposed. It is the intent of this rule to prevent degrading of the first pilot's instrument system, or the only pilot's instrument system in a single-pilot-approved rotorcraft, by not permitting peripheral systems to be connected to it. In addition, equipment must not be connected to operating systems for the second pilot's required instruments unless it is extremely improbable that failure of such additional equipment would affect that operating system.
Similar provisions are also included in Appendix B to Part 29, Airworthiness Criteria for Helicopter Instrument Flight.
b. Procedures.
(1) The provisions of the current rule are essentially self-explanatory.
(2) If the certification basis of the rotorcraft is prior to Amendment 29-24, the provisions are more precise; however, they only apply in the instance where duplicate instruments are required by the operating rules.
(3) If an IFR approval is part of the certification effort, then Part 29, Appendix B, applies, and the provisions of Paragraph VIll(b)(5) are essentially the same as the current rule. If the certification basis of the rotorcraft is prior to Amendment 29-24, and an IF approval is being added, the instrument systems should be carefully reviewed since their design may not have considered the provisions of the IFR rule.
640. § 29,1335 (Amendment 29-14) FLIGHT DIRECTOR SYSTEMS.
a. Explanation. This section prescribes the accepted display criteria for a rotorcraft three-cue flight director providing command guidance for pitch, roll, and power. 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.
994 Par 639
7130/97 AC 29-2B
(1) 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.
(2) 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/AUTHORITY with special symbology, special background effects, and special color coding, and has performed satisfactorily in service.
b. Procedures. 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 being confused with a glide slope display or angle of attack display.
One display uses a third cue, shaped like a small handle, to aid in identifying it as the collective pitch symbol.
(1) 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.
(2) If the attitude director indicator (ADI) provides a monochromatic display, the collective pitch cue and its background markings must be distinctive to reduce the chance of 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.
Par 640 995
641. § 29.1337 (Amendment 29-13) POWERPLANT INSTRUMENTS - (Paragraph (b) - FUEL GAUGE CALIBRATION).
a. Explanation. Section 29.1337(b) requires, in part, a means to indicate to the flightcrew the quantity of useable fuel in each tank in flight. Since the flight attitude of a rotorcraft may vary significantly with CG (center of gravity) and airspeed, a standard attitude for calibration of the fuel quantity gauge is needed. In addition, guidelines for gauge accuracy and comments regarding other fuel quantity gauging aspects are offered.
b. Procedures.
(1) Determine the rotorcraft pitch attitudes for most forward and most aft CG at a median gross weight and at an airspeed of 0.9 VNE or 0.9 VH, whichever is less. The mean attitude of the extremes defined above, further adjusted for lateral CG effects, if necessary, define the rotorcraft attitude for fuel gauge calibration.
(2) After establishing the calibration attitude, the requirements of 8 29.133/ (b) can be accomplished. The aircraft should be placed in the calibration attitude. Add fuel to the filler neck spillover level. Defuel the aircraft in increments corresponding to fuel gauge increment markings or at least 10 increments until gauge zero is obtained.
Precautions should be taken during this step to be sure that the fuel transmitter is sensing fuel level and not simply reflecting a physical "STOP" or end point in the system range. The fuel remaining in the tank below the "ZERO" mark must not be less than that amount determined by flight testing under § 29.959. (Otherwise, the zero point must be adjusted upward.) The gauging system accuracy is acceptable when it meets a tolerance of +2 percent of the total useable fuel plus 14 percent of the remaining usable fuel at any gauge reading, provided that the gauge indicates zero fuel with unusable fuel in accordance with § 29.959 in the tank. (For a 100-gallon tank this formula would allow a t6-gallon error at the full level, 14-gallon error at 50-gallon level, converging to a t2-gallon error at low fuel with the further provision that the zero mark accurately reflects unusable fuel.)
(3) Certain other aspects of a fuel gauging system need attention in order to minimize fuel exhaustion incidents:
Gauge reading with the aircraft at ground attitude is frequently used by the crew in calculating range, weight and balance, and actual gross weight.
Significant gauge errors in either direction during this reading can introduce hazards to the operation of the aircraft. If a calibration at this attitude indicates an unconservative error in excess of 6 percent of the gauge reading, corrective information should be applied adjacent to the fuel quantity gauge or be made available to the crew in other handbook data.
996 Par 641
(i¡) Flight during hover with maximum rearward wind may introduce significantly different fuel gauge readings. A check should be made to assure that the gauge is either accurate or at least does not read high (unconservative) in this attitude.
(4) Fuel gauging system transmitters which are strictly volumetric measuring devices (float-actuated variable rheostats) introduce a gauge readout error of about 5 percent if calibrated with a fuel temperature of 0° C and subjected to -55° C fuel or +55° C fuel. This error may be minimized by calibrating the gauge with fuel temperature in the middle of the useful range; i.e., 15° C.
(5) Capacitance transmitters have become the standard for most modern fuel systems. These transmitters ordinarily need no temperature compensation since the fuel volume and the fuel dielectric constant vary inversely as temperature changes.
The basic capacitance transmitter does not compensate for the different dielectric values to be expected with different type fuels. An add-on capacitance located so as to be submerged in fuel at all times can be devised to automatically compensate for other fuels.
641A. § 29.1337 (Amendment 29-26) POWERPLANT INSTRUMENTS.
a. Explanation. Amendment 29-26 adds § 29.1337(e) that requires certain rotor drive system transmissions and gearboxes to be equipped with chip detector systems.
These detectors will sense and signal the presence of ferromagnetic particles to the flight crew. The rule also requires a means to permit the crewmembers to check, in flight, the function of each detector's electrical circuit and signal. This amendment will improve the level of safety available with the installation of chip detector systems.
b. Procedures. All of the policy material pertaining to this section remains in effect. Additionally, the following information is added about chip detectors. The chip detectors should:
(1) Indicate the presence of ferromagnetic particles in the transmission or gearbox;
(2) Be easily removable for inspection of the magnetic poles for metallic chips;
and,
(3) Prevent the loss of lubricant in the event of failure of the retention device for the removable portion of the chip detector (debris monitor).
(4) Provide a test system to allow the crew to check, in flight, the function of each detector and wiring. The test circuit should test, at least, as much of the circuitry as reasonably possible. Where detectors are used that have a test feature in the form Par 641 997