AC 29.1331. § 29.1331 (Amendment 29-24) INSTRUMENTS USING A POWER SUPPLY.
a. Explanation. The rule concerns each flight instrument usin g a power supply that is installed in a Category A rotorcraft. A reference to paragraph AC 29.1303 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 non-electrical 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 AC 29.1303, and independent power sources are discussed in paragraph AC 29.1355b(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 AC 29.1355). 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.
Page F – 60.3
AC 29-2C, Chg 4 5/1/2014 [This page intentionally blank.] Page F – 60.4
AC 29.1333. § 29.1333 (Amendment 29-24) INSTRUMENT SYSTEMS.
a. Explanation.
(1) Electromechanical Displays. Prio r to Amendment 29-24, this requirement was intended to apply to duplicate flight instruments required by any operating rule.
Due to the increased complexity of instrumentation that was available and being used, it was considered appropriate to amend the provisions of this requirement by adopting §
29.1333 in Amendment 29-24 to more appropriately consider the extreme range of operational environments to which rotorcraft were being routinely exposed. It is the intent of § 29.1333 to prevent degrading 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 VIII.(b)(5)(i) through (iii) of Appendix B to Part 29, Airworthiness Criteria for Helicopter Instrument Flight.
(2) Advanced Display Systems. The increased use of microprocessor technology in avionic systems has resulted in the use of computer-generated graphics to replace conventional electromechanical instruments. These displays may replace individual instruments or may integrate several flight critical parameters into single displays. For display of redundant informati on, “crosstalk” between the pilot and copilot displays and supporting systems has been allowed to provide detection and annunciation of faults or “miscompare” of critical flight information. A level of safety finding equivalent to that level of safety provided by § 29.1333 may be possible through the implementation of integration technology that will assure that failure of one system does not and can not adversely affect the other system.
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 VIII(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 IFR approval is being added, the instrument systems should be carefully reviewed since their design may not have considered the provisions of the IFR rule.
Page F – 61
AC 29.1333A. § 29.1333 (AMENDMENT 29-59) INSTRUMENT SYSTEMS.
Explanation. Amendment 29-59 revised § 29.1333 by revising paragraph (a), which requires isolating the pilot instrument system from any other operating systems, to only apply to pneumatic systems. This change allows other types of pilot instrument systems to be connected to other systems on the rotorcraft and allows the use of modern technology to monitor and display highly integrated information.
Page F - 61.1
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.
Page F – 63
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.
Page F - 63.1
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.)
Page F – 63.2
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.
Page F – 64
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.
Page F – 65
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:
Page F – 66
(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.
[The remainder of this page intentionally left blank.] {Section AC 29.1351 continued on next page.] Page F – 66.1