(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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AC 29-2C, Chg 1 2/12/03
(3) 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.
(4) 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.”
(ii) 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.
(A) 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.
(B) 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 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 Appendix B and appropriate Category A requirements.
(C) Replacement batteries must show compliance to the requirements of FAR 29 for the same areas of concern identified in paragraph b(2)(i)(B).
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6/23/3 AC 29-2C, Chg 9 AC 29.1353A. § 29.1353 (AMENDMENT 29-59) ENERGY STORAGE SYSTEMS.
a. Explanation. Amendment 29-59 revised § 29.1353 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 § 29.1351. Additionally, electrical equipment, controls and wiring installation design should be documented sufficiently to maintain configuration control for manufacturing and to assure 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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(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 § 29.1353(c). This design scheme must not expose structure or adjacent equipment to corrosive fluids, as required by § 29.1353(c), or allow explosive or toxic gasses to accumulate to hazardous levels, as required by § 29.1353(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 § 29.1353(d).
b. Procedures.
(1) General. Chapter 11 of Advisory Circular 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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(A) The energy storage system capacity must be sufficient to supply the required current at a voltage level that allows proper operation of the equipment or functions dependent on the energy storage system. This capacity must be shown to be sufficient at the proper voltage levels for all operational and environmental conditions, as required by § 29.1351(a)(1). Any energy storage system required for continued safe flight and landing of the rotorcraft must have monitoring features and the means to indicate to the pilot the status of all critical system parameters (see § 29.1353(e)).
Accordingly, warning, caution, or advisory indication must be provided appropriate to any hazard condition. Section 91.167 requires rotorcrafts approved for IFR operations have sufficient back-up 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 if primary power generation fails. The minimum time for IFR is 30 minutes. Additionally, for rotorcraft that employ electrical power for starting, to meet the requirements in § 29.1351, the energy storage system must have sufficient capacity to provide adequate engine starting on the ground, and also 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 rotorcraft power distribution system and the number of engines.
(B) The energy storage system must be qualified for the specific location and rotorcraft environment. Acceptable environmental test procedures may be found in RTCA DO-160, which provides test procedures to meet a wide range of environmental conditions that may exist on a wide variety of aircraft. Environmental tests should meet or exceed the expected installation environment. In addition to the environmental qualification, the equipment and installation design should consider the risk of explosion presented by the production of gases that occur during normal operation or as a result of any failed condition. An explosive resistant case or safe venting may be an acceptable way of showing compliance if it is shown that structure and essential systems do not suffer hazardous effects that jeopardize continued safe flight and landing.
(C)Energy storage systems shall have protective features to prevent unsafe conditions during operation. The protective features shall automatically control all critical operational parameters of the system in order to prevent any unsafe condition. The term “automatically” is intended to preclude pilot action as a mitigating feature and is not intended to require a specific mitigating design feature. Development assurance levels (DAL) for hardware and software that implement monitoring and control functions should be commensurate with the effect associated with system failure. An acceptable certification process for airborne software is found in AC 20-115 and, for complex hardware, in AC 20-152.
(D)The occurrence of self-sustaining, uncontrolled increases in temperature or pressure must be prevented for normal energy storage system operation or as a result of a failure of any part of the system installation (see § 29.1353(d)).
System architecture should ensure that a component failure does not propagate to adjacent components. A safe design may require a compartmentalized structure to prevent cascading failures. Any required monitoring and control functions implemented in hardware or software should have a DAL commensurate with the hazard associated Page F - 78.3
with the loss of function. Installation design features to mitigate explosive increases in temperature or pressure may be required if energy storage system DALs are not suitable for the inherent hazard.
(E) Fluids or gases that may escape from an energy storage system or its installation may not damage surrounding structure or any adjacent systems, equipment, or electrical wiring of the rotorcraft (see § 29.1353(c)). Installation design data should include features to contain, vent, or drain these fluids or gases away from the rotorcraft. Flight tests may be required to evaluate the effectiveness of venting or draining throughout the normal operating envelope.
(F) Explosive or toxic gases that may be emitted by an energy storage system should not accumulate in hazardous quantities within the rotorcraft (see § 29.1353(b)). Installation design data should include features to contain or vent these gases away from the rotorcraft. Flight tests may be required to evaluate the effectiveness of venting or draining throughout the normal operating envelope.
(G) The energy storage system and its installation must prevent any hazardous effects on the rotorcraft structure or other installed systems caused by the maximum amount of heat generated during normal operation or by a failure of any part of the energy storage system or its installation (see § 29.1353(d)). Shielding, cooling, or venting provisions may be required to ensure that temperatures do not exceed allowed temperatures of any adjacent equipment, electrical wiring, hydraulic lines, or structure.
(H) Energy storage system temperatures and pressures must be maintained within a safe operating range during all foreseeable charging and discharging conditions, and during any failure of the control or monitoring functions (see § 29.1353(d)). Energy storage system installation design features must mitigate all hazards associated with monitoring and control failures (see § 29.1353(e)). The monitoring and control function implemented in hardware or software should have a DAL commensurate with the hazard associated with the loss of function.
(I) The instructions for continued airworthiness (ICA) must include maintenance requirements to ensure that required energy storage system maintenance functions are performed at appropriate intervals specified for each energy storage system component, to meet the requirements in Appendix A to Part 29. This is necessary to ensure system performance will not degrade below specified energy levels sufficient to power rotorcraft systems, for intended applications. Similarly, the ICA must also include procedures for the maintenance of energy storage system components in spares storage to prevent the replacement of components with stored components that have experienced degraded functional capability or other damage due to prolonged storage. Replacement energy storage system components must be of the same manufacturer and part number as initially approved, unless additional installation approval is obtained (see § 21.9(a)). Precautions should be included in the ICA maintenance instructions to prevent mishandling of the energy storage system components that could result in unintentional impact damage caused by dropping or other destructive means that could result in personal injury or property damage.
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