SUBPART F - EQUIPMENT
ELECTRICAL SYSTEMS AND EQUIPMENT AC 27.1351. § 27.1351 (Amendment 27-13) 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) Electrical System Capacity. Rotorcraft electrical systems have grown in capacity, complexity, and impact on safety. This paragraph requires adequate electrical system capacity for safe operation of load circuits essential for safe operation at continuous rated power. If this capacity can be shown by electrical measurements, an electrical load analysis is not required. However, if the measurement method of compliance is chosen, data should be provided to show that the measured loads represent the worst case electrical loading for components or combination of components.
(i) Load circuits (systems) that are essential for safe operation are those systems necessary to maintain controlled flight and land safely and are generally those systems required to show compliance with the certification regulations. This includes most electrical utilization systems.
(ii) An electrical utilization system is a system of electrical equipment, devices, and connected wiring using electric energy to perform a specific aircraft function.
(iii) The specific utilization systems, which are necessary to maintain controlled flight and land safely, will vary with the type of rotorcraft and with the nature of operations. Examples of systems which may be essential are basic flight instruments, minimum navigation equipment, minimum radio communications, and flight control systems.
(2) Function.
(i) Generating systems should be analyzed, inspected, or tested to ensure that no probable malfunction in the generating system or in the generator drive system may cause loss of service to systems essential for safe operation. A probable malfunction is any single electrical or mechanical component malfunction or failure that Page F - 55
AC 27-1B, Chg 3 9/30/2008 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. Analyses should be performed on the electrical power system emphasizing the exclusion of single point failures and the possibility of latent failures. Test methods should be developed that uncover latent faults. To identify latent failure causing events, consider the functions and the results of the safety analysis, as well as any failure detection monitor that may be incorporated. Evaluate the system using various combinations of unforeseen internal signal patterns and states. Also, evaluate the non- monitored functions by selecting test conditions that use every signal path and decision point between the input and output. Refer to MG-2 for electrical system test methods. These analyses 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.
(ii) The generator drive system includes the prime movers (propulsion engines or other) and coupling devices such as gearboxes or constant speed drives.
(iii) Where crew corrective action is necessary:
(A) Adequate warning should be provided for any malfunction or failure requiring such corrective action;
(B) Controls should be located to permit such corrective action during any probable flight situation;
(C) If corrective action must be taken within a specified time to continue safe operation of the generating system, it should be demonstrated that such corrective action can be accomplished within the specified time during any probable flight situation; and (D) The procedure to be followed by the crew should be detailed in the Rotorcraft Flight Manual.
(iv) Chapter 11 of Advisory Circular 43.13-1A, “Acceptable Methods, Techniques, and Practices; Aircraft Inspection and Repair,” includes guidance on installation of electrical systems (routing, separation, tying, clamping, j-box installations, etc.). Special emphasis should be placed on wire routing during the rotorcraft compliance inspection. Control wires to the rotorcraft’s generators should be routed separately from generator output wiring. This should begin at the generator and continue to the voltage regulator. Additionally, wiring installation design should be Page F – 55.1
documented sufficiently to maintain configuration control for manufacturing and to assure that the electromagnetic characteristics remain the same as the certification sample.
(3) Generating System. When electrical power is needed for essential equipment, this paragraph requires at least one generator with adequate capacity for safe operation. Complete electrical failures have been caused by loss of voltage control in the voltage regulator. Overvoltage conditions can destroy electronic equipment. An acceptable method of overvoltage protection is the use of a separate overvoltage sensing relay to trip the generator off the line when overvoltage is detected. Another acceptable method is use of a voltage regulator with built-in overvoltage protection.
(4) Generator Ratings: 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 (continuous, 2 minute intermittent and 5 second transient 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) Loa d Bank. A load bank will usually be necessary to load the test unit to the amperage limit for which approval is requested.
(2) A mmeter. An ammeter should be provided with sufficient resolution to assure the amperage load is being maintained at the desired level.
(3) T emperature/Pressure Readouts. Readouts which are compatible with the temperature or pressure sensors installed in the test unit should be provided.
(4) C alibration Records. Calibration records should be available for all instrumentation.
(5) R ecordings. 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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AC 27-1B, Chg 3 9/30/2008 (C) Regulatory References. Sections 27.1301, 27.1309, 27.1351,
27.1521(f), 27.1041, 27.1043, 27.1045 (through Amendment 27-19).
(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.
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(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.
(3) The battery may be connected to the bus during the generator/alternator cooling test. The generator/alternator temperatures should be recorded at intervals sufficiently close to show the rate of temperature increase and stabilization. The temperature may be considered stabilized when it peaks and has not increased in the last 5 minutes. The 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 conducted at maximum altitude in the cruise configuration. Generator/alternator 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 amperage 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 shutdown.
(B) Multiengine Procedures. Conduct a generator cooling test in accordance with the following procedures:
(1) All ground operational and in-ground-effect hover tests should be conducted in ambient winds of 5 knots or less.
(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.
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(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 shutdown 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 27.1309b(2)(i) 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 obt ained 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.
(5) Instruments. Voltage and current supplied by each generator are parameters which define system operation. Some systems are provided with voltmeters and ammeters to display these parameters to the crew. These instruments may be multifunctional with switches to select the functions displayed. Some designs have annunciated safe operation of each generator with lights and have no voltmeter and ammeter. If additional limitations, such as maximum loading of portions of the systems, are necessary to account for fault condition, that information should be made available to appropriate personnel (crew, owner, modifier, etc.) to ensure the limits are not exceeded.
(6) An external power source supplying reverse polarity or reverse phase sequence to the rotorcraft electrical system could seriously damage the system. This paragraph requires a means to prevent such an occurrence. This can be accomplished by use of a standard polarized receptacle and protective relays.
AC 27.1353. § 27.1353 (Amendment 27-14) STORAGE BATTERY DESIGN AND INSTALLATION.
a. Explanation. Batteries must not be designed and/or installed to create a hazard to the rotorcraft under any operating conditions.
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b. Procedures.
(1) Battery General. As part of the electrical system evaluation, the battery installation should be reviewed to ensure 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 ensure that fluids are drained clear of the rotorcraft. Some aircraft rely on the installation of a sump jar to dispose of corrosive fluids.
(2) Battery Installation. 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.
(i) Approval of the battery for installation is dependent on the helicopter’s requirements for the installed equipment and operations.
(A) 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.
(B) 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 Page F - 59
AC 27-1B, Chg 1 2/12/03 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 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.
(C) 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.
(D) 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.
(ii) 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.
(A) 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 Page F - 60
operate properly within all specified limits. Operational specified limits may be set by battery charging in-rush currents or possible fault currents.
(B) A battery charge control system may be used that automatically controls the battery charge to prevent battery overheating. Unless otherwise specified by the battery manufacturer, temperatures above 140° F are considered overheat for NI-CAD batteries. The system is acceptable if the charge rate is automatically adjusted by controlling the charging current as a function of battery temperature, and in an over-temperature condition, the charge current is automatically reduced to a safe value.
Zero to 10 amperes has been considered safe for batteries rated at less than 34 amp-hours, and zero to 15 amperes has been considered safe for batteries rated at 34 amp-hours or more. The actual number chosen should be substantiated. Means and/or procedures should be provided for the crew to monitor the charger performance or the battery condition. If there is an automatic disconnect of the charger from the batteries and associated bus on an over-temperature condition, provisions should be provided in the cockpit to warn of a disconnect.
(C) If a temperature monitoring system is used, 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.
(D) Battery failure sensing and warning systems have also been used to show compliance with this rule.
(3) Battery replacement with any battery other than the one that was installed and accepted as part of the original type design must consider several aspects of battery design and installation to assure that the type design is not negatively impacted.
(i) The replacement battery should provide at least the minimum usable capacity as required by the current load analysis. This requirement of minimum usable capacity should be valid over the entire range of environmental conditions. Different types of batteries have different inherent characteristics. Nickel cadmium batteries have a voltage discharge curve that allows most of the amp-hour capacity to be usable from a delivered voltage standpoint, where only about 66% of the amp-hour capacity for lead acid batteries is usable due to the delivered voltage becoming too low to use. If a replacement battery has a similar discharge rate, and the capacity is similar throughout the original battery’s environmental range, then no further testing should be needed.
(ii) Replacement batteries must be compatible with type design requirements of the helicopter (reference paragraph b(2)(i)(A) above). For IFR and CAT A certified helicopters that depend on fault clearing features, an analysis may be Page F - 61
AC 27-1B , Chg 4 3/31/2014 needed to show that the electrical system still complies to the original certification requirements. Subsequent analysis validation by testing may be needed in those cases where the system that provides fault clearing is difficult to evaluate by analysis alone. In these cases, the effects of a change in battery internal impedance on the fault clearing features is uncertain and may not be adequately determined by analysis.
There are high rate discharge types of batteries and slow-rate discharge batteries that have different internal impedances and their behavior in a fault clearing circuit is also different. The regulatory basis for analysis and testing for fault clearing is contained in Appendices B and C.
(iii) Replacement batteries must show compliance to the requirements of FAR 27 for the same areas of concern identified in paragraph b(2)(i)(B).
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AC 27-1B, Chg 9 AC 27.1353A. § 27.1353 (AMENDMENT 27-51) ENERGY STORAGE SYSTEMS.
a. Explanation. Amendment 27-51 revised § 27.1353 to generalize the regulation to apply to any energy storage system. Additionally, the amendment changed the section title from “Storage battery design and installation” to “Energy storage systems.” This AC material is updated to address lithium batteries. The guidance pertaining to lead-acid and nickel-cadmium batteries has not changed.
(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 to meet the requirements of § 27.1351. Additionally, electrical equipment, controls and wiring installation design should be documented sufficiently to maintain configuration control for manufacturing and to ensure that characteristics such as electromagnetic and thermal attributes remain the same as the certification sample.
(2) The applicant should submit the results of qualification testing. Environmental qualification of electrical equipment may be accomplished using the appropriate revision of RTCA DO-160. Electrical equipment should be tested as installed to show there are no adverse environmental effects on previously installed equipment and that newly installed equipment is not adversely affected by previously installed equipment.
Potential adverse effects include, but are not limited to, electromagnetic interference or overheating.
(3) Installed energy storage systems serve many required and non-required rotorcraft functions. Energy storage systems vary in size, capacity, chemistry, discharge rates, recharge rates, function, function criticality, and reliability. They supply power for the primary electrical system architecture, backup power for required functions, and backup power for non-required functions. Throughout the life of the aircraft, these systems and their components will be maintained, replaced, and modified. Replacement systems or components should be identical or all differences assessed and justified.
Replacement or system modification data should show that any changes to rotorcraft systems and wiring remain in compliance and that functionality and reliability meet or exceed current requirements. Capacity requirements should result from a careful and accurate system’s load analysis. System management requirements such as electrical load shedding should safely integrate into the overall cockpit management scheme and not adversely affect pilot workload for any normal or emergency operation.
(4) Energy storage systems may present hazards when installed on rotorcraft.
Energy storage may be achieved by various methods and each may present different hazards that must be addressed for installation. Hazards associated with installation of any energy storage system should be identified and evaluated to determine any effects on continued safe flight and landing. If required, methods should be developed to mitigate those effects.
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AC 27-1B, Chg 9
(5) Energy storage systems that have the potential to emit corrosive fluids, explosive gasses, or toxic gasses will require a design scheme to contain, drain, or vent these fluids or gasses, to meet the requirements of § 27.1353(c). This design scheme must not expose structure or adjacent equipment to corrosive fluids, as required by § 27.1353(c), or allow explosive or toxic gasses to accumulate to hazardous levels, as required by § 27.1353(b).
(6) Energy storage systems that are susceptible to thermal runaway should include a sense and control capability to prevent such an event or a design to safely contain or vent heat and pressure. Designs for energy storage systems with potential for thermal runaway must prevent an uncontrolled cascading failure that results in a hazardous condition, to meet the requirements of § 27.1353(d).
b. Procedures.
(1) General. Chapter 11 of AC 43.13-1B, 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:
(i) Feeder wires from the rotorcraft’s generators and energy storage systems 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) Energy Storage System Installation.
(i) Installation approval of an energy storage system consists of two parts.
One part is the approval of the energy storage system performance specifications that meet the rotorcraft requirements and the other part is the actual installation of the energy storage system into the rotorcraft. The following methods of showing compliance can be demonstrated by a combination of tests and analysis.
(ii) Approval of the energy storage system for installation is dependent on rotorcraft requirements for the installed equipment and operations.
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