(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.
Par 635 983
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.
984 Par 636
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.
Par 637 985
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