causes the associated panel segment to illuminate. At this time all panel segments connected to the test circuit will illuminate.
(iii) This configuration becomes a special problem when one or more of the panel segments are red. A red light calls for immediate action by the crew, and the crew does not have adequate information for immediate action when many false panel segments are illuminated.
(iv) If the design review indicates a problem, a redesign of the panel to eliminate the condition is considered to be the best solution and is highly encouraged.
(2) Procedures.
(i) An alternative to panel redesign might be the following:
(A) Review the annunciator panel design and note which segments are red.
(B) Determine if cross reference information is available in the cockpit to allow elimination from consideration of any of the red segments (e.g., red low fuel pressure light and low fuel pressure gauge). Normal operation of the gauge would be a reason to assume the light did not cause the problem.
(C) Where a cross reference is available, further design review of that function is not necessary; however, it may be appropriate to include procedural information in the emergency procedures section of the rotorcraft flight manual.
(D) If cross references are not available for red segments, additional isolation should be incorporated into the annunciator design for those functions.
(ii) If cross referencing is not practical, the following approach is encouraged.
(A) Review the annunciator panel design and note which segments are red.
(B) Determine if isolation diodes are checked during the application of battery or external power before starting the engines (e.g., red low oil pressure light). If isolation diode is shorted, all panel segments will light as soon as battery or external power is applied.
(C) When the isolation diode can be checked before starting engines, further design review is not necessary.
(D) If diodes are not automatically checked before starting, then additional isolation, should be considered.
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(3) Annunciator Panel Arrangement. The annunciator panels should be arranged in a logical manner to reduce the crew’s time required to locate faults and to increase their efficiency in following aircraft flight manual procedures. For example, engine annunciators on multiengine rotorcraft should be physically located on the panel to coincide with engine location (left or right) so that properly operating engines are not inadvertently shut down due to crew confusion over which engine has malfunctioned.
AC 27.1323. § 27.1323 (Amendment 27-13) AIRSPEED INDICATING SYSTEM.
a. Explanation.
(1) The accuracy of all flight test data concerned with the velocity of the rotorcraft is dependent on the calibration of the airspeed indicating system. For this reason, the airspeed system position error should be determined very early in the program.
(2) Since air density varies with altitude, the speed reading will only be correct under standard sea level conditions. However, in an actual installation, the indicator reading, even under standard sea level conditions, may differ from the calibrated airspeed because the static system does not sense true static pressure. This error in detection of static pressure is called position error. It is caused by the pressure field built up around the rotorcraft in flight. This pressure field will vary in intensity with dynamic pressure making the position error a function of calibrated airspeed. Since airspeed information is presented to the crew in terms of indicated airspeed, it is necessary to determine the position error for the rotorcraft to be flown safely.
b. Procedures.
(1) There are different methods to determine position error such as trailing bomb, airspeed course, boom system, etc. Each method has its own advantages and disadvantages, but will yield satisfactory results if done correctly. The airspeed system should be calibrated throughout the airspeed range of the rotorcraft and under the various flight conditions of cruise, climb, and autorotation. In addition, the effects of gross weight and center of gravity should be investigated.
(2) It may also be necessary to recalibrate the system with a change in external configuration if such a change may affect the airflow near the pitot or static sources.
(3) Additional information regarding position error is included in paragraph AC 27 Appendix B b(10) and should be considered if pursuing an IFR approval.
(4) Static system installation information is included in paragraph AC 27.1325.
Technical Standard Order (TSO) C16, Airspeed Tubes (Heated), gives minimum Page F - 39
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.
AC 27.1325. § 27.1325 (Amendment 27-13) STATIC PRESSURE SYSTEMS.
a. Explanation.
(1) This section, in conjunction with § 27.1323, provides minimum performance standards for static pressure systems. The standard provides some relief when considering the icing environmental condition in that it allows the use of an alternate static port to account for the icing condition.
(2) The standard for the consideration of environmental conditions is § 27.1309(a).
(3) The standard for consideration of malfunction conditions is § 27.1309(b).
(4) For rotorcraft that will be approved for IFR operation, the provisions of Appendix B VIII(b)(5) of Part 27 as discussed in paragraph AC 27 Appendix B 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 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.
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(5) Other static system considerations are included in paragraphs AC 27.1323 and AC 27 Appendix B.
AC 27.1327. § 27.1327 (Amendment 27-13) MAGNETIC DIRECTION INDICATOR.
a. Background. This section contains specific requirements regarding installation and functioning of a magnetic direction indicator. The magnetic direction indicator (commonly referred to as a compass) described by this paragraph is the unit required by § 27.1303(c) or the unit or system required for IFR operation by Appendix B VIII(a) to Part 27. Both of these indicators provide the pilot with an aircraft heading whi ch is referenced to the earth’s magnetic field. The unit required by § 27.1303(c) is the indicator commonly referred to as a “whiskey compass.” The unit was given this designation because early units were constructed using alcohol as the medium in which the compass ball floats. This unit is generally approved as meeting the requirements of TSO-C7c. The indicator required by Appendix B to Part 27 is usually a system of units which meets the requirements of TSO-C6c.
b. Procedures. In showing compliance to § 27.1327(a), generally the magnetic indicator and its respective components will be tested to an appropriate standard such as RTCA DO 160B for use in a rotorcraft. If the unit functioned properly as described in the TSO during this testing, then no additional evaluation is generally required concerning vibration immunity. To determine the immunity of the indicator (system) from magnetic effects and its installed accuracy, a ground and flight test should be performed. This test should turn the rotorcraft a full 360° heading change in 45° increments. The indicator should not have an error in excess of 10° on any of the 45° increments. When performing these tests, the electrical equipment and systems should be functioning normally, and the effect of windshield heating (if installed) should be investigated. During this investigation, it is permissible for the error caused by the effect of the operation of electrical equipment on the indicator required by § 27.1303(c) to exceed 10° if a gyroscopic heading indicator is installed. This gyroscopic heading indicator may either be a directional gyro or a slaved compass system (gyro-stabilized magnetic direction indicator). The results of the investigation may be used to construct the calibration placard which is required by § 27.1547. It should be noted that a calibration placard has not been traditionally required for slaved compass systems.
Also, it should be emphasized that other aspects of the functioning and installation of these indicators should comply with the other general requirements (i.e., §§ 27.1301,
27.1309, 27.1555, etc.).
AC 27.1329. § 27.1329 (Amendment 27-21) AUTOMATIC PILOT SYSTEM.
a. Explanation. The automatic flight control systems used on most modern rotorcraft often perform two different and distinct functions when viewed from a regulatory compliance aspect. These two functions are an augmentation of the stability of the rotorcraft and a pilot aid in maintaining attitude, altitude, and airspeed, or in radio navigation tasks. The first function of stability is not covered by § 27.1329 but is included under § 27.672. The second function as a pilot aid is the automatic pilot Page F - 41
AC 27-1B, Chg 4 3/31/2014 function covered by this section. The following procedure discusses only those parts or systems which are installed as a pilot aid. Appendix B 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 § 27.1329. In demonstrating malfunctions of the autopilot system, generally servo actuator hardovers are the most critical malfunction. If this is the case and the autopilot system utilizes the same servos and servo amplifiers as the stability augmentation system (SAS) and the autopilot function cannot produce a more severe hardover than the SAS, then no additional consideration is required for this malfunction. An evaluation using the guidance in Appendix B of this AC would be sufficient.
(ii) There have been autopilots approved which require the use of a monitor since they cannot meet the hardover malfunction requirements. These approvals have involved a finding of equivalent safety which is beyond the scope of this guidance.
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 assure adequate reliability. The question that often arises is: Should the preflight test function be interlocked so that the autopilot cannot be engaged if the preflight test has not been accomplished? The guidance used in the past to answer this question is: If the preflight test is simple and rapid enough that the pilot may reasonably be expected to perform such a test, then it is not required to be interlocked. If, however, the preflight test is very complicated and lengthy and a pilot who was pressed by a schedule might skip such a test, then this preflight test should be interlocked.
(v) Most of the autopilots which have been approved utilize series actuators or servos such as those required for a SAS. However, this does not preclude the approval of an autopilot which uses outer loop parallel actuation. This type of autopilot may be particularly helpful in a VFR aircraft.
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(2) Cockpit controls. Evaluation of the cockpit controls should include the following items:
(i) Location of the automatic pilot system controls are such that their operation is properly labeled and is readily accessible to the pilot(s).
(ii) Annunciator colors conform to the colors specified in § 27.1322 (reference section 27.1322 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 § 27.779.
(v) Annunciation should be provided if the autopilot disconnects for any reason other than pilot action.
c. Malfunction Evaluations. To preclude hazardous conditions which may result from any failure or malfunctioning of the autopilot the following failures should be evaluated.
This evaluation should also account for any hazards which 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.
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AC 27-1B, Chg 4 3/31/2014 (B) For a descent, the height loss should be determined as illustrated in Figure AC 27.1329-1. 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 27.1329-1, but with consideration to the critical approach angle and nominal approach speed.
(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.
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Minimum Use Height Actual Flight Path Intended flisht path Height Lass Path of rotorcraft skids/wheels as a 1. Failure Initiation result of failure. 2. Failure Recognition by pilot
3. Initiation of Manual Recovery action by pilot delay Height Loss Normal/Fault-Free Skid/Wheel Path Tangential Skid/Wheel Height FIGURE AC 27.1329-1 HEIGHT LOSS METHOD Page F - 45
AC 27-1B, Chg 4 3/31/2014
(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 hazardous deviations from the flight path. Additional guidance regarding the method of determining pilot recognition times and reasonable flight path deviations due to those simulated failures is contained in paragraph 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 (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 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.
(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 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.
(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.
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(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, the 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 autopilot should be easily available to the pilot who is now resisting the malfunctioning force of the autopilot. It is recommended that the disconnect button be placed on the cyclic control. It should be red and conspicuously marked “Autopilot Disconnect.” The pilot should be able to return the rotorcraft to its normal flight attitude under full manual control without exceeding the loads or speed limits defined in this paragraph and without engaging in any dangerous maneuvers during recovery. The maximum servo authority used for these tests should not exceed those values shown to be within the structural limits for which the rotorcraft was designed. The maximum altitude loss experienced during these tests should be measured.
(6) External Interfaces. The autopilot system should have appropriate interlocks to its engagement to ensure it does not operate improperly as a result of information furnished by an external device or system. An example of this is the navigation receivers and the compass system. If for a particular mode of operation the autopilot uses signals from these systems, the autopilot should be interlocked from operating in those modes if invalid information is being received from that system.
d. Automatic Pilot Instrument Approach Approval.
(1) Throughout an approach, no signal or combination of signals simulating the cumulative effect of any single failure or malfunction in the automatic pilot system, except vertical gyro mechanical failures, should provide hazardous deviations from flight path or any degree of loss of control.
(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.
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AC 27-1B, Chg 4 3/31/2014
(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.
(B) The point along the glide path when the simulated malfunction is induced.
(C) The point where the pilot indicates recognition of the malfunction.
(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 found in Figure AC 27.1329-2.
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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 27.1329-2 DEVIATION PROFILE Page F - 49