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
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.
642.-651. RESERVED.
998 (thru 1008) Par 641A
SECTION 36. ELECTRICAL SYSTEMS AND EQUIPMENT
652. § 29.1351 (Amendment 29-40) ELECTRICAL SYSTEMS AND EQUIPMENT - GENERAL.
a. Explanation. With the advent of more 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 by Military Specification MIL-E-7016F, and use of this standard is preferred since it has been received widespread acceptance. If other formats have been used and have been considered acceptable, their continued use is encouraged.
(2) Generating systems must be analyzed, inspected, or tested to assure conformance to the following criteria.
(i) For Category A, the generating system must perform as specified in § 29.1309(d) and (e).
(i¡) No probable malfunction in the generating system or in the generator drive system may result in the permanent loss of service to electric utilization systems, which are necessary to maintain controlled flight and to effect 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 malfunction or failure which is considered probable on the basis of past service experience with similar components in aircraft applications.
This definition should be extended to multiple malfunctions when:
(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.
(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.
Par 652 1009
(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,
(1) Adequate warning should be provided for any malfunction or failure requiring such corrective action.
(il) Controls should be so located as to permit such corrective action during any probable flight situation.
(ili) 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 1010 Par 652
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 fallure 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) Refer to Paragraph 777 of this advisory circular for detailed test procedures for DC electrical systems.
c. Operation with normal electrical power generating system inoperative. See FAR 29.1351(d).
(1) Definition: Normal electrical power generating system. The term normal electrical power generating system is intended to include all electrical power sources used for operation of the rotorcraft under any approved normal operating condition (VFR, IFR, Icing, etc.), not including batteries and emergency electrical power sources.
(2) All rotorcraft (See FAR 29.1351(d)(1) Amendment 29-40).
(i) FAR 29.1351(d)(1) requires, for all rotorcraft, continued safe VFR operation for a period of at least 5 minutes with the normal electrical power system inoperative. If loss of the normal electrical power generating system, followed by depletion of battery power, could prevent safe flight and landing, adequate warning of loss of the normal electrical power generating system should be provided for compliance with FAR 29.1309(c), and Flight Manual procedures compatible with the available battery endurance should be provided.
(її) One possible cause of loss of the normal electrical power generating system is engine failure. The requirement specifies consideration of engine flameout and restart attempts. A minimum battery endurance of 5 minutes is specified. To ensure safe operation under all conditions, however, the battery endurance should be not less than the time required for an autorotative descent to sea level from the maximum operating altitude. Where applicable, allowance should be made for the use of the batteries for attempts to restart the engines during the descent. It may be necessary to include limitations on the number of attempted starts or to provide a separate dedicated battery for such purposes.
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(3) Category A rotorcraft (See FAR 29.1351(d)(2) Amendment 29-40).
FAR 29.1351(d)(2) is applicable to Category A rotorcraft and requires that provision be made to ensure adequate electrical supplies to those systems which are necessary for continued safe flight and landing in the event of a failure of all normal generated electrical power. All components and wiring of the alternate supplies should be physically and electrically segregated from the normal system and should be such that no single failure, including the effects of fire, the cutting of a cable bundle, or the loss of a junction box or control panel will affect both normal and alternate supplies.
In considering the systems which should remain available following the loss of the normal electrical power generating system, consideration should be given to the role and flight conditions of the rotorcraft and the possible duration of flight time to reach a suitable landing site and make a safe landing.
(iii) The systems required by FAR 29.1351(d)(2) may differ between rotorcraft types and roles and should be agreed with the Authority. They should normally include:
(A) Attitude information;
(B) Radio communication and flight crew intercommunication;
(C) Navigation;
(D) Cockpit and instrument lighting;
(E) Heading, airspeed and altitude information, including appropriate pitot head heating;
(F) Adequate flight controis;
(G) Adequate engine instrumentation and control;
(H) Such warnings, cautions, and indications as are required for continued safe flight and landing;
(I) Any other services required for continued safe flight and landing;
e.g. fire extinguishing, emergency flotation equipment, landing light.
(iv) Emergency Power Source Duration and Integrity (A) Time Limited Power Source. Where an emergency power source provided to comply with FAR 29.1351(d)(2) is time limited (e.g., battery), the required duration will depend on the type and role of the rotorcraft. Unless it can be 1012 Par 652
shown that a lesser time is adequate, such a power source should have an endurance of at least half the rotorcraft endurance, or the Flight Manual limitations section should define aircraft endurance. However, an endurance of less than 30 minutes would not normally be acceptable. The endurance, with any associated procedures, should be specified in the Flight Manual. The endurance time should be determined by calculation or test, due allowance being made for-
(1) Delays in flight crew recognition of failures and completion of the appropriate drills where flight crew action is necessary. This should be assumed to be 5 minutes provided that the failure warning system has clear and unambiguous attention-getting characteristics and where such a delay is compatible with the crew's primary attention being given to the corresponding emergency procedures and/or other possible related failures such as engine fire, fumes in the cockpit, etc. A delay of less than 5 minutes may be acceptable if justified by simple procedures or an adequate degree of automation.
(2) The minimum voltage acceptable for the required loads, the battery state of charge, the minimum capacity permitted during service life and the battery efficiency at the discharge rates and temperatures likely to be experienced.
Unless otherwise agreed for the purpose of this calculation, a battery capacity at normal ambient conditions of 80 percent of the nameplate rated capacity, at the 1 hour rate, and a 90 percent state of charge, may be assumed (i.e., 72 percent of nominal demonstrated rated capacity at +20° C).
(3) For those rotorcraft where the battery is also used for engine or APU starting on the ground, it should be shown that following engine starts, the charge rate of the battery is such that the battery is maintained in a state of charge that will ensure adequate emergency power source duration should a failure of generated power occur shortly after takeoff.
NOTE: This could, for example, be achieved by ensuring that, following battery-powered starting, the battery charging current has fallen to a specified level prior to take-off.
(B) Non-time-limited Power Source. Where an emergency power source is provided by a non-time-limited source (e.g., standby generator driven by APU, transmission, pneumatic or hydraulic motor), due account should be taken of any limitation imposed by rotorcraft speed, altitude, etc., which may affect the capabilities of that power source. In considering the power source, account should be taken of the following:
(1) Auxiliary Power Unit (APU).
An APU capable of continuous operation throughout an adequate flight envelope may be considered an acceptable means of supplying Par 652 1013
electrical power to the required services provided that its air start capability is adequate and may be guaranteed. Where, however, the APU is dependent for its starting current on a battery source which is supplying critical loads, such starting loads may prejudice the time duration of the flight if APU start is not achieved.
(i) It may be necessary, therefore, to include limitations on the number of attempted starts or to provide a separate battery for APU starting, if this method of supplying electrical power is adopted. Consideration should also be given to the equipment, services and duration required prior to the APU generator coming on-line. Common failures which could affect the operation of all engines and the APU
(e.g., fuel supply) should be taken into consideration.
(2) Transmission-driven Generator.
A transmission-driven generator may be utilized to provide an emergency electrical power source, but due consideration should be given to ensuring that the means of bringing the generator into use are not dependent on a source which may have been lost as a result of the original failure.
(i) The continuity of electrical power to those services, which should remain operative without crew action prior to the generator being brought into operation, may necessitate the use of a battery, unless the operation of the emergency power source is automatic and immediate in the event of failure of the normal electrical power generating system.
(3) Pneumatic or Hydraulic Motor Driven Power Source. A pneumatic- or hydraulic-motor-driven electrical power source may be utilized subject to the same constraints on activation as the transmission-driven generator (See 3.4.2(b)).
Care should be taken in ensuring that the operation of the pneumatic or hydraulic system is not prejudiced by faults leading to, or resulting from, the original failure, including the loss of, or inability to restart engines.
653. § 29.1353 (Amendment 29-14) ELECTRICAL EQUIPMENT AND INSTALLATIONS.
a. Explanation.
(1) Electrical equipment, controls, and wiring must be installed so that operation of any one unit or system of units will not adversely affect the simultaneous operation of any other electrical unit or system essential to safe operation.
(2) Results of qualification testing should be available to ensure that the installation of equipment or a system will not result in adverse interference being introduced into the rotorcraft electrical equipment. A good reference for interference testing is the applicable version of Radio Technical Commission of Aeronautics (RTCA) 1014 Par 652
Document No. DO-160, "Environmental Conditions and Test Procedures for Airborne Equipment."
(3) The DO-160 type tests would normally be accomplished by the equipment manufacturer. The airframe manufacturer's test are normally more subjective and are oriented more toward watching for unwanted meter movement, noise in the interphone systems, and so forth. The combination of the equipment manufacturer's tests, supplemented by the airframe manufacturer's installation tests, should be adequate to assure compliance with this regulation.
b. Procedures.
(1) General. Chapter Il of Advisory Circular 43.13-1A, "Acceptable Methods, Techniques, and Practices: Aircraft Inspection and Repair," includes considerable guidance regarding the installation of electrical systems (routing, separation, typing, clamping, j-box installations, etc.). The following areas are overlooked in many cases and special emphasis should be placed on them during the compliance inspection of the rotorcraft:
Feeder wires from the rotorcraft's generators and batteries should be routed separately from utilization system wiring.
(ії) Generator field wiring should be routed separately from generator output wiring. This should begin at the generator and continue to the voltage regulator.
(2) Battery Installation.
As part of the electrical system evaluation, the battery installation should be reviewed to assure the battery is vented and drained. If there is some doubt regarding the ability of the drain to satisfactorily dispose of corrosive fluids, TIA tests should be conducted to resolve the issue. Normally this is done by expelling a dye solution through the drain system during different phases of flight to assure that fluids are drained clear of the rotorcraft. Some aircraft rely on the installation of a sump jar to dispose of corrosive fluids.
(ii) In nickel cadmium batteries are used for engine starts and compliance with § 29.1353(c)(6) is achieved through the use of a temperature monitoring system, the temperature sensor should be located in a position that will most accurately reflect the internal battery temperature without causing adverse effects to the sensor. The location normally used is near the center of the battery. If the sensor is placed between two cells, the indication should be very close to the actual temperature within the cell. If the sensor is placed in a cell strap, there will normally be a period of time just after a heavy current drain (e.g., engine start) when the sensor shows a temperature that is hotter than the actual cell temperature.
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Other aspects of the battery installation can be resolved by reviewing § 29.1353(c), AC 43.13-1A, and AC 43.13-2A, "Acceptable Methods, Techniques, and Practices Aircraft Alterations."
654. § 29.1355 (Amendment 29-14) DISTRIBUTION SYSTEM.
a. Explanation. None.
b. Procedures.
(1) When determining compliance with the portion of the rule that concerns supplying essential circuits in the event of reasonably probable faults or open circuits, the effects of tripped circuit breakers or blown fuses should be considered.
(2) Various means may be used to ensure an energy supply. Examples include duplicate electrical equipment, throwover switching, and multichannel or loop circuits separately routed.
(3) Essential load circuits are those circuits whose functioning is required to show regulatory compliance with the certification basis. In addition to those circuits specifically required by the regulations, this definition also includes those circuits required by general rules such as § 29.1309.
(4) An independent power source includes not only the electrical power source
(e.g., generator) but other items such as a regulator or a reverse current cutout that are necessary to make the electrical power source deliver power to a distribution bus.
When a regulatory requirement exists for two independent power sources, the required items should not be shared.
(5) Electrical system faults may occur that will result in a portion of the system (feeders, buses, etc.) being lost. Where portions of the electrical system may be switched from one power source to another to compensate for a fault, it is important that the transfer action not result in the loss of the replacement source. Circuit design should be such as to assure this will not happen.
654A. § 29.1355 (Amendment 29-24) DISTRIBUTION SYSTEM.
a. Explanation. Amendment 29-24 provides clarification for availability of the remaining electrical power source after a failure of one of two independent power sources.
b. Procedures. All of the policy material pertaining to this section remains in effect with the addition that an automatic or manually selectable means is required to maintain operation of the equipment or system for which the two independent power sources were required.
1016 Par 653