AC 29-2C, Chg 3 9/30/2008 [This page intentionally left blank.] [Section AC 29.1351 continued on next page.] Page F – 66.2
(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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AC 29-2C, Chg 4 5/1/2014
(2) After engine start, load the instrumented generator to its proposed amperage limit and begin recording temperatures.
(3) A total of 30 minutes should be spent on the ground prior to takeoff.
This is to account for taxi and holding times.
(4) After takeoff, climb at single-engine best-rate-of-climb speed using maximum continuous power, to the single-engine service ceiling. Above this, continue at twin-engine best-rate-of-climb speed, using maximum continuous power on both engines, to maximum altitude.
(5) Cruise at maximum altitude until all generator temperatures stabilize.
Temperatures shall be considered stabilized when they have peaked and have not increased for a period of 5 minutes.
(6) Descend, conduct an approach to include a go-around, hover until temperature stabilizes, then land and continue to record temperatures after shut-down until 5 minutes after all temperatures have peaked.
(7) Conduct cooling tests with the rotorcraft hovering at both the minimum and maximum hover altitudes.
(8) Correct all results for hot day conditions. Use the standard lapse rate of 3.6° F/1,000 feet for consideration of altitude. See paragraph AC 29.1309.b.(9)(ii)(A) for details on temperature correction.
(C) Manufacturer’s Limits. If at any time during the testing it appears the manufacturer’s limits are to be exceeded, the amperage load on the test generator/alternator should be reduced to prevent this from happening.
(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.
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).
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(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.
(ii) 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.
(3) Category A rotorcraft (See FAR 29.1351(d)(2) Amendment 29-40).
(i) 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.
(ii) 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;
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AC 29-2C, Chg 2 4/25/06 (E) Heading, airspeed and altitude information, including appropriate pitot head and static vent heating;
Note: Where the aircraft is to be approved for IFR, pitot heat and, where appropriate, static vent heating specified in paragraph (E) above (relating to required air data for airspeed information while operating in an emergency configuration) should be provided for the complete duration (at least 30 minutes) while operating on emergency power. A minimum 5-minute “landing” time for the landing light operation specified in paragraph (3)(i) above should be provided.
(F) Adequate flight controls;
(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 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, Page F – 72
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).
(i) An APU capable of continuous operation throughout an adequate flight envelope may be considered an acceptable means of supplying 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.
(ii) 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.
(i) 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.
(ii) 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 Page F – 73
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.
AC 29.1353. § 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. Additionally, wiring installation design should be documented sufficiently to maintain configuration control for manufacturing and to assure that the electromagnetic characteristics remain the same as the certification sample.
(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) 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 II 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:
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(i) Feeder wires from the rotorcraft’s generators and batteries should be routed separately from utilization system wiring.
(ii) 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.
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AC 29-2C, Chg 1 2/12/03
(i) Installation approval of batteries consists of two parts. One part is the approval of the battery itself to performance specifications that meet the helicopter’s requirements and the other part is the actual installation of the battery into the helicopter. The following methods of showing compliance can be demonstrated by a combination of tests and analysis.
(A) Approval of the battery for installation is dependent on the helicopter’s requirements for the installed equipment and operations.
(1) The battery’s capacity must be sufficient to supply the required current, at a voltage level that allows proper operation of the equipment/functions dependent on battery power. This capacity must be shown to be sufficient, at the proper voltage levels, for all operational and environmental conditions. Helicopters approved for IFR operations require sufficient battery power to operate a minimum set of required equipment for a time long enough to find a suitable place to land and effect a safe landing. The minimum time for IFR is 30 minutes. Additionally, for helicopters that employ electrical power for starting, the battery must have sufficient capacity to provide adequate engine starting on the ground, and also to have the capability to provide power for the air start function, when required. The power requirement for air start could have high-level safety considerations depending on the helicopter’s power distribution system and the number of engines.
(2) The battery should be selected with the knowledge of the type of charging that will be supplied to it from the helicopter. The type of charging will be a major factor to determine the number of cells that the battery should contain. If the battery is not adequately charged, the capacity may not be available, when required, to meet emergency power demands. Twenty cell NI -CAD batteries have been used for several years to prevent thermal runaway; however, aircraft bus level voltage may not be an adequate source to properly charge a 20 cell Ni-CAD battery. The nominal voltage of a Ni-CAD cell is 1.5 volts and if the number of battery cells is 20 then;
1.5 Volts/Cell X 20 Cells = 30 volts required to even match the nominal voltage of the installed battery. (This is actually inadequate to charge the Ni -CAD cell as a few tenth volts above that of the nominal voltage is required.) The maximum helicopter bus voltage available is typically 28.5 volts, which is deficient by 1.5 volts to just meet the nominal battery voltage. A compensating factor for this battery charging design is to require frequent battery maintenance activities. Future battery installation designs should not depend on the use of frequent battery maintenance as a compensating factor for designs that are inherently deficient in battery charging capability. Battery charging from the helicopter bus voltage is the prevalent method used to charge batteries, since no separate battery charger is needed. However, for the 20 cell Ni-CAD example, using the helicopter bus voltage as the battery-charging source can introduce battery capacity limitations. Several options are available to address these concerns.
One option would be to go to a 19-cell battery. Since Ni -CAD batteries were first introduced, much has been learned about their care and use, plus design improvements have been made that lessen the possibility of thermal runaways. Another approach could be to oversize the battery capacity recognizing that the battery will be in a "less Page F – 76
than" nominal capacity condition. This philosophy would also carry with it a Certification Maintenance Requirement as a limitation on the helicopter approval. Compliance could be shown by using a dedicated battery charger, that can profile the charging based on state of charge. This can be much more efficient and make it possible to have a better charged battery with less frequent maintenance, and increased battery life.
(3) One of the significant design issues associated with batteries is the internal impedance. Low internal impedance usually characterizes batteries capable of high discharge rates. Starting and fault clearing requirements are typically best satisfied by batteries with low internal impedance. The other extreme is batteries with high internal impedances that are generally more economical and perform low rate, long-time interval discharges well. This type of batteries can provide the emergency power for equipment efficiently. If both starting/fault clearing and emergency power requirements are to be provided using a single battery, a workable compromise on the internal impedance must be accomplished.
(4) The battery must be qualified to the helicopter’s defined environmental specifications associated with the battery installation location, as a minimum. High level environmental qualification may preclude the installation location dependency for the battery. In addition to the environmental qualification, the design/installation design should consider the risk of explosion presented by the natural production of gases. An explosive resistant case or container is an acceptable way of showing compliance to the rule requiring that the structure and essential systems do not suffer hazardous effects.
There may be other ways to show compliance, but use of the explosion-resistant case is the predominant way at this time.
(B) The part of the battery installation approval associated with installation into the helicopter should consider the adequacy of the associated battery interface components and the physical aspects of installing the battery.
(1) The adequacy of the associated components that are necessary to interface with the battery to provide functionality must be shown to perform their intended function for the installation environment and operations. These components should be qualified to meet the most severe environment of their installation and operate properly within all specified limits. Operational specified limits may be set by battery charging in-rush currents or possible fault currents.
(2) 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.
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