AC 29.1335. § 29.1335 (Amendment 29-14) FLIGHT DIRECTOR SYSTEMS.
a. Explanation. This section prescri bes the accepted display criteria for a rotorcraft three-cue flight director providing command guidance for pitch, roll, and power. Three-cue flight directors for ro torcraft 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 a ttitude, 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.
(1) The general convention for flight director design is that each command bar is a “fly to” command. The motion of the fli ght director indicator is such to command a corresponding sense of control system motion. Th is 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 cons titute 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 displa y 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 Page F – 62
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.
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AC 29.1335A. § 29.1335 (AMENDMENT 29-59) FLIGHT DIRECTOR SYSTEMS.
Explanation. Modern designs include both automatic pilot and flight director systems, referring to them now as automatic pilot and flight guidance system. Having these systems in separate rules that use different terminology has resulted in some confusion. Therefore, amendment 29-59 removed Section 29.1335 and incorporated its provisions into Section 29.1329.
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AC 29.1337. § 29.1337 (Amendment 29-13) POWERPLANT INSTRUMENTS - (Paragraph (b) - FUEL QUANTITY INDICATOR).
a. Explanation. Section 29.1337(b) requires, in part, a means to indicate to the flight crew the quantity of usable fuel during flight in each tank. When two or more tanks are interconnected so that a failure of the system could cause fuel to become trapped in a fuel tank, the fuel quantity indicating system must provide the flight crew with the ability to determine the total effective amount of remaining usable fuel. Since the flight attitude of a rotorcraft may vary significantly with center of gravity (CG) 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 § 29.1337(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 bel ow 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 ±4 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 ±6-gallon error at the full level, ±4-gallon error at 50-gallon level, converging to a ±2-gallon error at low fuel with the further provision that the zero mark accurately reflects unusable fuel.)
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AC 29-2C, Chg 4 5/1/2014
(3) Certain other aspects of a fuel gauging system need attention in order to minimize fuel exhaustion incidents:
(i) 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 th is 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.
(ii) 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.
(iii) Consistent with the requirements of § 29.1337(b)(2), a separate fuel quantity indication is necessary for any interconnected fuel tank that has a flow control device, such as a fuel transfer pump or flapper valve, which could fail and trap fuel.
This requirement also applies to auxiliary fuel tanks. A sight gauge that is readable by the flight crew in flight may be acceptable for use with auxiliary fuel tanks.
(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 (32°F) and subjected to -55°C (-67°F) fuel or +55°C (131°F) fuel. This error may be minimized by calibrating the gauge with fuel temperature in the middle of the useful range (i.e., 15°C (59°F)).
(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.
AC 29.1337A. § 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.
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b. Procedures. All of the policy material perta ining 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 te st, at least, as much of the circuitry as reasonably possible. Where detectors are used that have a test feature in the form of an extra pin, all of the circuit, exclusive of the detector may be tested. Some chip detectors have a fuzz burner capability to eliminate nuisance indication of non-relevant conducting materials that result from oil contamination and very small wear particles.
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AC 29-2C, Chg 3 9/30/2008
SUBPART F - EQUIPMENT
ELECTRICAL SYSTEMS AND EQUIPMENT AC 29.1351. § 29.1351 (Amendment 29- 40) ELECTRICAL SYSTEMS AND EQUIPMENT -GENERAL.
a. Explanation. With the advent of mo re sophisticated rotorcraft and operations under more critical conditions, such as IFR and icing, it is essential that the electrical system be very carefully analyzed and evaluated to assure proper operation under any foreseeable operating condition, and that hazards do not result from any malfunctions or failures.
b. Procedures.
(1) An acceptable method of preparing an electrical load analysis is given in Military Specification MIL-E-7016F, and use of this standard is preferred since it has received widespread acceptance. If other formats have been used and have been considered acceptable, their continued use is acceptable.
(2) Generating systems must be analyzed, inspected, or tested to assure conformance to the following criteria. Analysis should be performed on the electrical power system emphasizing the exclusion of single point failures and possibilities of latent failures. Test methods should be developed that uncover latent faults. Ref MG-2 for electrical system test methods.
(i) Analyses should be performed on the electrical power system with an emphasis on excluding single point and latent failures. Also, evaluate the nonmonitored functions by selecting test conditions that use every signal path and decision point between the input and output. Test methods should be developed that uncover latent faults. Refer to MG-2 fo r electrical system test methods.
(ii) For Category A, the generating system must perform as specified in § 29.1309(d) and (e).
(iii) No probable malfunction in the generating system or in the generator drive system may result in loss of service to electric utilization systems which are necessary to maintain controlled flight and to affect a safe landing, unless the aircraft is equipped with an independent source of electrical power capable of supplying continuous emergency service to these utilization systems. A probable malfunction is any single electrical or mechanical component malfunction or failure that is likely to occur based on past service experience. This past service experience can include malfunction of components of previously approved rotorcraft, other aircraft, or qualitative analysis of similar components in rotorcraft applications. These analyses should be extended to multiple malfunctions when:
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(A) The first malfunction would not be detected during normal operation of the system, including periodic checks established at intervals which are consistent with the degree of hazard involved; or (B) The first malfunction would inevitably lead to other malfunctions.
(3) The generator drive system includes the prime movers (propulsion engines or other) and coupling devices such as gear boxes or constant speed drives.
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(4) An electric utilization system is a system of electric equipment, devices, and connected wiring which utilizes electric energy to perform a specific aircraft function.
(5) The specific electric utilization systems, which are necessary to maintain controlled flight and effect a safe landing, will vary with the type of aircraft and with the nature of the operation in which the aircraft is utilized. Examples of systems which may be in this category are as follows: basic flight instruments, minimum navigation equipment, minimum radio communications, and control system boost.
(6) Where crew corrective action is necessary,
(i) Adequate warning should be provided for any malfunction or failure requiring such corrective action.
(ii) Controls should be so located as to permit such corrective action during any probable flight situation.
(iii) If corrective action must be taken within a specified time interval for continued safe operation of the generating system, it should be demonstrated that such corrective action can be accomplished within the specified time interval during any probable flight situation. For Category A rotorcraft, compliance with § 29.903(b)(2) must be considered.
(iv) The procedure to be followed by the crew should be detailed in the Rotorcraft Flight Manual.
(7) Voltage and current supplied by each generator are considered essential parameters for definition of system operation and most systems are provided with voltmeters and ammeters to display these parameters to the crew. Some recent designs have annunciated safe operation of each generator with lights and have eliminated the voltmeter and ammeter. For these systems, in addition to distribution system design precautions, parameters such as over and under voltage, reverse current sensing, feeder ground faults, and over and under frequency (AC generators only) are being monitored and provided as inputs to the generator annunciators. For systems not incorporating voltmeters and ammeters, and with automatic protective switching and annunciator lights, the pilot should be provided as a minimum, with sufficient information to determine the type of fault, and to identify portions of the system that have been lost.
If additional limitations such as maximum loading of portions of the system are necessary to account for fault conditions, that information should be made available to appropriate personnel (crew, owner, modifier, etc.) to assure the limits are not exceeded.
(8) For rotorcraft with a certification basis of FAR 29 after Amendment 29-14 (effective on July 18, 1977), the electrical wire and cable insulation and other materials used to show compliance with § 29.1351(d) must be self-extinguishing when tested in accordance with Part 25 Appendix F. This means the wire must be tested at an angle Page F – 67
of 60° in accordance with the applicable portions of Appendix F of FAR 25 which contain acceptable test procedures and define burn length.
(9) An area where a possibly hazardous malfunction of an electrical power source might occur is the supply of cooling air to the electrical generators. The hazard exists because the failure of a generator bearing usually produces metallic sparks and hot surfaces which are a potential ignition source. Consideration should be given to this failure. One method is if the generator is rated explosion proof, then the intake and output of cooling air into the engine compartment should not cause a hazard. If the generator is not rated explosion-proof and a failed bearing test cannot conclusively demonstrate that failure of the generator will not produce an ignition source, then cooling air should be ducted into and out of the generator from outside the aircraft. The ducting material should be sufficient to contain the failed generator fragments.
(10) Generator ratings are often the result of installation temperature limitations. The determination of these limitations, if any, is by testing the actual installation. The procedures for performing generator cooling tests are as follows:
(i) Test Requirements.
(A) General. The applicant should contact the generator (alternator) manufacturer and obtain the maximum limits for the unit to be tested. This will normally be in terms of temperatures at various locations within the unit (stator, bearings, diodes, heat sinks, brushes, etc.) or in terms of pressure drop across the generator. The manufacturer should either supply an instrumented unit or give complete details for instrumenting the test unit.
(B) Instrumentation.
(1) Load Bank. A load bank will usually be necessary to load the t est unit to the amperage limit for which approval is requested.
(2) Ammeter. An ammeter should be provided with sufficient resolution to assure the amperage load is being maintained at the desired level.
(3) Temperature/Pressure Readouts. Readouts which are compatible with the temperature or pressure sensors installed in the test unit should be provided.
(4) Calibration Records. Calibration records should be available for all instrumentation.
(5) Recordings. Permanent recordings should be provided for time, temperatures, current and/or pressure. The recording device should have provisions for placing event marks on the recording medium.
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(C) Regulatory References. Sections 29.1301, 29.1309, 29.1351,
29.1363(b), 29.1521(e), 29.1041, 29.1043, 29.1045, 29.1047, and 29.1049.
(D) Miscellaneous. The results obtained from the tests should be corrected for hot day conditions using a standard lapse rate (3.6° F/1,000 feet). The tests are conducted to determine the maximum generator capacity that does not result in surpassing the limits given from the manufacturer. This is for a continuous rating, any credit for short time over current ratings must also be verified by the same methods, particularly for short time ratings longer than 5 seconds.
(ii) Test Procedures.
(A) Single Engine Procedure.
(1) The cooling test is to be conducted during ground operation, climb-out, cruise, approach, and hover flight regimes.
(2) All ground operational and in-ground effect hover tests should be conducted in ambient winds of 5 knots or less. Wind direction relative to the aircraft should be from the most critical direction.
(3) The battery may be connected to the bus during the generator/alternator cooling test. The gener ator/alternator temperatures should be recorded at intervals sufficiently close to show the rate of temperature increase and stabilization. The temperature may be cons idered stabilized when it peaks and has not increased in the last 5 minutes. T he climb-out speed and power setting should correspond to the best rate of climb speed, using maximum continuous power or any other normal conditions of climb that would cause the generator/alternator temperatures to be critical. The cruise test should be c onducted at maximum altitude in the cruise configuration. Generator/a lternator cooling should be conducted at rated output consistent with the RPM at which it is operating. For instance, during the ground tests the engine RPM may be lower than that necessary to sustain maximum rated amperage output. In this case the maximum am perage output of the generator/alternator corresponding to the lower RPM should be assured.
(4) The test sequence should begin with about 30 minutes of ground operation to account for taxi and holding times, and end 5 minutes after all temperatures have peaked after engine shut down.
(B) Multi-engine Procedures. Conduct a generator cooling test in accordance with the following procedure:
(1) All ground operational and in-ground effect hover tests should be conducted in ambient winds of 5 knots or less. Wind direction relative to the aircraft should be from the most critical direction.
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