2024-30855. Special Conditions: Safran Electric & Power S.A. Model ENGINeUS 100A1 Electric Engines

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Department of Transportation

Federal Aviation Administration

1. 14 CFR Part 33
2. [Docket No. FAA-2023-0587; Special Conditions No. 33-23-01-SC]

AGENCY:

Federal Aviation Administration (FAA), DOT.

ACTION:

Final special conditions.

SUMMARY:

These special conditions are issued for the Safran Electric & Power S.A. (Safran) Model ENGINeUS 100A1 electric engines that operate using electrical technology installed on the aircraft for use as an aircraft engine. These engines will have a novel or unusual design feature when compared to the state of technology envisioned in the airworthiness standards applicable to aircraft engines. This design feature is the use of an electric motor, motor controller, and high-voltage systems as the primary source of propulsion for an aircraft. The applicable airworthiness regulations do not contain adequate or appropriate safety standards for this design feature. These special conditions contain the additional safety standards that the Administrator considers necessary to establish a level of safety equivalent to that established by the existing airworthiness standards.

DATES:

Effective January 27, 2025.

FOR FURTHER INFORMATION CONTACT:

Mark Bouyer, Engine and Propulsion Standards Section, AIR-625, Technical Policy Branch, Policy and Standards Division, Aircraft Certification Service, 1200 District Avenue, Burlington, Massachusetts 01803; telephone [phone]; [e-mail].

SUPPLEMENTARY INFORMATION:

Background

On November 27, 2020, Safran applied for FAA validation for a type certificate for their Model ENGINeUS 100A1 electric engine. The Safran Model ENGINeUS 100A1 electric engine will be used in a single-engine airplane that will be certified separately from the engine.

The Safran Model ENGINeUS 100A1 electric engine is comprised of a direct-drive, radial-flux, permanent magnet motor, divided in two sections, each section having a three-phase motor, and one electric power inverter controlling each three-phase motor.

Type Certification Basis

Under the provisions of 14 CFR 21.17(a)(1), generally, Safran must show that Model ENGINeUS 100A1 electric engines meet the applicable provisions of 14 CFR part 33 in effect on the date of application for a type certificate.

If the Administrator finds that the applicable airworthiness regulations ( e.g., part 33) do not contain adequate or appropriate safety standards for the Safran Model ENGINeUS 100A1 electric engines because of a novel or unusual design feature, special conditions may be prescribed under the provisions of § 21.16.

Special conditions are initially applicable to the model for which they are issued. Should the type certificate for that model be amended later to include any other engine model that incorporates the same novel or unusual design feature, these special conditions would also apply to the other engine model under § 21.101.

The FAA issues special conditions, as defined in § 11.19, in accordance with § 11.38, and they become part of the type certification basis under § 21.17(a)(2).

Novel or Unusual Design Features

The Safran Model ENGINeUS 100A1 electric engines will incorporate the following novel or unusual design features:

An electric motor, motor controller, and high-voltage electrical systems that are used as the primary source of propulsion for an aircraft.

Discussion

Electric propulsion technology is substantially different from the technology used in previously certificated turbine and reciprocating engines. Therefore, these engines introduce new safety concerns that need to be addressed in the certification basis.

A growing interest within the aviation industry involves electric propulsion technology. As a result, international agencies and industry stakeholders formed Committee F39 under ASTM International, formerly known as American Society for Testing and Materials, to identify the appropriate technical criteria for aircraft engines using electrical technology that has not been previously type certificated for aircraft propulsion systems. ASTM International is an international standards organization that develops and publishes voluntary consensus technical standards for a wide range of materials, products, systems, and services. ASTM International published ASTM F3338-18, “Standard Specification for Design of Electric Propulsion Units for General Aviation Aircraft,” in December 2018.^[1] The FAA used the technical criteria from the ASTM F3338-18, the published Special Conditions No. 33-022-SC for the magniX USA, Inc. Model magni350 and magni650 engines, and information from the Safran Model ENGINeUS 100A1 electric engine design to develop special conditions.

Part 33 Was Developed for Gas-Powered Turbine and Reciprocating Engines

Aircraft engines make use of an energy source to drive mechanical systems that provide propulsion for the aircraft. Energy can be generated from various sources such as petroleum and natural gas. The turbine and reciprocating aircraft engines certificated under part 33 use aviation fuel for an energy source. The reciprocating and turbine engine technology that was anticipated in the development of part 33 converts oxygen and fuel to energy using an internal combustion system, which generates heat and mass flow of combustion products for turning shafts that are attached to propulsion devices such as propellers and ducted fans. Part 33 regulations set forth standards for these engines and mitigate potential hazards resulting from failures and malfunctions. The nature, progression, and severity of engine failures are tied closely to the technology that is used in the design and manufacture of aircraft engines. These technologies involve chemical, thermal, and mechanical systems. Therefore, the existing engine regulations in part 33 address certain chemical, thermal, and mechanically induced failures that are specific to air and fuel combustion systems operating with cyclically loaded, high-speed, high-temperature, and highly stressed components.

Safran's Electric Engines Are Novel or Unusual

The existing part 33 airworthiness standards for aircraft engines date back to 1965. As discussed in the previous paragraphs, these airworthiness standards are based on fuel-burning reciprocating and turbine engine technology. The Safran Model ENGINeUS 100A1 electric engines are neither turbine nor reciprocating engines. These engines have a novel or unusual design feature, which is the use of electrical sources of energy instead of fuel to drive the mechanical systems that provide propulsion for aircraft. The ( print page 105433) Safran aircraft engine is subject to operating conditions produced by chemical, thermal, and mechanical components working together, but the operating conditions are unlike those observed in internal combustion engine systems. Therefore, part 33 does not contain adequate or appropriate safety standards for the Safran Model ENGINeUS 100A1 electric engine's novel or unusual design feature.

Safran's aircraft engines will operate using electrical power instead of air and fuel combustion to propel the aircraft. These electric engines will be designed, manufactured, and controlled differently than turbine or reciprocating aircraft engines. They will be built with an electric motor, motor controller, and high-voltage electrical systems that draw energy from electrical storage or electrical energy generating systems. The electric motor is a device that converts electrical energy into mechanical energy by electric current flowing through windings (wire coils) in the motor, producing a magnetic field that interacts with permanent magnets mounted on the engine's main rotor. The controller is a system that consists of two main functional elements: the motor controller and an electric power inverter to drive the motor.^[2] The high-voltage electrical system is a combination of wires and connectors that integrate the motor and controller.

In addition, the technology comprising these high-voltage and high-current electronic components introduces potential hazards that do not exist in turbine and reciprocating aircraft engines. For example, high-voltage transmission lines, electromagnetic shields, magnetic materials, and high-speed electrical switches are necessary to use the physical properties of an electric engine for propelling an aircraft. However, this technology also exposes the aircraft to potential failures that are not common to gas-powered turbine and reciprocating engines, technological differences which could adversely affect safety if not addressed through these special conditions.

Safran's Electric Engines Require a Mix of Part 33 Standards and Special Conditions

Although Safran's electric aircraft engines use novel or unusual design features that the FAA did not envisage during the development of its existing part 33 airworthiness standards, these engines share some basic similarities, in configuration and function, to engines that use the combustion of air and fuel, and therefore require similar provisions to prevent common hazards ( e.g., fire, uncontained high energy debris, and loss of thrust control). However, the primary failure concerns and the probability of exposure to these common hazards are different for the Safran Model ENGINeUS 100A1 electric engine. This creates a need to develop special conditions to ensure the engine's safety and reliability.

The requirements in part 33 ensure that the design and construction of aircraft engines, including the engine control systems, are proper for the type of aircraft engines considered for certification. However, part 33 does not fully address aircraft engines like the Safran Model ENGINeUS 100A1 electric engine, which operates using electrical technology as the primary means of propelling the aircraft. This necessitates the development of special conditions that provide adequate airworthiness standards for these aircraft engines.

The requirements in part 33, subpart B, are applicable to reciprocating and turbine aircraft engines. Subparts C and D are applicable to reciprocating aircraft engines. Subparts E through G are applicable to turbine aircraft engines. As such, subparts B through G do not adequately address the use of aircraft engines that operate using electrical technology. Special conditions are needed to ensure a level of safety for electric engines that is commensurate with these subparts, as those regulatory requirements do not contain adequate or appropriate safety standards for electric aircraft engines that are used to propel aircraft.

FAA Special Conditions for the Safran Engine Design

Applicability: Special condition no. 1 requires Safran to comply with part 33, except for those airworthiness standards specifically and explicitly applicable only to reciprocating and turbine aircraft engines.

Engine Ratings and Operating Limitations: Special condition no. 2, in addition to compliance with § 33.7(a), requires Safran to establish engine operating limits related to the power, torque, speed, and duty cycles specific to Safran Model ENGINeUS 100A1 electric engines. The duty or duty cycle is a statement of the load(s) to which the engine is subjected, including, if applicable, starting, no-load and rest, and de-energized periods, including their durations or cycles and sequence in time. This special condition also requires Safran to declare cooling fluid grade or specification, power supply requirements, and to establish any additional ratings that are necessary to define the Safran Model ENGINeUS 100A1 electric engine capabilities required for safe operation of the engine.

Materials: Special condition no. 3 requires Safran to comply with § 33.15, which sets requirements for the suitability and durability of materials used in the engine, and which would otherwise be applicable only to reciprocating and turbine aircraft engines.

Fire Protection: Special condition no. 4 would require Safran to comply with § 33.17, which sets requirements to protect the engine and certain parts and components of the airplane against fire, and which would otherwise be applicable only to reciprocating and turbine aircraft engines. Additionally, this special condition requires Safran to ensure that the high-voltage electrical wiring interconnect systems that connect the controller to the motor are protected against arc faults. An arc fault is a high-power discharge of electricity between two or more conductors. This discharge generates heat, which can break down the wire's insulation and trigger an electrical fire. Arc faults can range in power from a few amps up to thousands of amps and are highly variable in strength and duration.

Durability: Special condition no. 5 requires the design and construction of Safran Model ENGINeUS 100A1 electric engines to minimize the development of an unsafe condition between maintenance intervals, overhaul periods, and mandatory actions described in the Instructions for Continued Airworthiness (ICA).

Engine Cooling: Special condition no. 6 requires Safran to comply with § 33.21, which requires the engine design and construction to provide necessary cooling, and which would otherwise be applicable only to reciprocating and turbine aircraft engines. Additionally, this special condition requires Safran to document the cooling system monitoring features and usage in the engine installation manual (see § 33.5) if cooling is required to satisfy the safety analysis described in special condition no. 17. Loss of cooling to an aircraft engine that operates using electrical technology can result in rapid overheating and abrupt engine failure, with critical consequences to safety.

Engine Mounting Attachments and Structure: Special condition no. 7 requires Safran and the design to comply with § 33.23, which requires the applicant to define, and the design to withstand, certain load limits for the engine mounting attachments and ( print page 105434) related engine structure. These requirements would otherwise be applicable only to reciprocating and turbine aircraft engines.

Accessory Attachments: Special condition no. 8 requires the design to comply with § 33.25, which sets certain design, operational, and maintenance requirements for the engine's accessory drive and mounting attachments, and which would otherwise be applicable only to reciprocating and turbine aircraft engines.

Rotor Overspeed: Special condition no. 9 requires Safran to establish by test, validated analysis, or a combination of both, that—

(1) the rotor overspeed must not result in a burst, rotor growth, or damage that results in a hazardous engine effect;

(2) rotors must possess sufficient strength margin to prevent burst; and

(3) operating limits must not be exceeded in service.

The special condition associated with rotor overspeed is necessary because of the differences between turbine engine technology and the technology of these electric engines. Turbine rotor speed is driven by expanding gas and aerodynamic loads on rotor blades. Therefore, the rotor speed or overspeed results from interactions between thermodynamic and aerodynamic engine properties. The speed of an electric engine is directly controlled by electric current, and an electromagnetic field created by the controller. Consequently, electric engine rotor response to power demand and overspeed-protection systems is quicker and more precise. Also, the failure modes that can lead to overspeed between turbine engines and electric engines are vastly different, and therefore this special condition is necessary.

Engine Control Systems: Special condition no. 10(b) requires Safran to ensure that these engines do not experience any unacceptable operating characteristics, such as unstable speed or torque control, or exceed any of their operating limitations.

The FAA originally issued § 33.28 at amendment 33-15 to address the evolution of the means of controlling the fuel supplied to the engine, from carburetors and hydro-mechanical controls to electronic control systems. These electronic control systems grew in complexity over the years, and as a result, the FAA amended § 33.28 at amendment 33-26 to address these increasing complexities. The controller that forms the controlling system for these electric engines is significantly simpler than the complex control systems used in modern turbine engines. The current regulations for engine control are inappropriate for electric engine control systems; therefore, the special condition no. 10(b) associated with controlling these engines is necessary.

Special condition no. 10(c) requires Safran to develop and verify the software and complex electronic hardware used in programmable logic devices, using proven methods that ensure that the devices can provide the accuracy, precision, functionality, and reliability commensurate with the hazard that is being mitigated by the logic. RTCA DO-254, “Design Assurance Guidance for Airborne Electronic Hardware,” dated April 19, 2000,^[3] distinguishes between complex and simple electronic hardware.

Special condition no. 10(d) requires data from assessments of all functional aspects of the control system to prevent errors that could exist in software programs that are not readily observable by inspection of the code. Also, Safran must use methods that will result in the expected quality that ensures the engine control system performs the intended functions throughout the declared operational envelope.

The environmental limits referred to in special condition no. 10(e) include temperature, vibration, high-intensity radiated fields (HIRF), and all others addressed in RTCA DO-160G, “Environmental Conditions and Test Procedures for Airborne Electronic/Electrical Equipment and Instruments,” dated December 8, 2010, which includes RTCA DO-160G, Change 1—“Environmental Conditions and Test Procedures for Airborne Equipment,” dated December 16, 2014, and “DO-357—User Guide: Supplement to DO-160G,” dated December 16, 2014.^[4] Special condition 10(e) requires Safran to demonstrate by system or component tests in special condition no. 27 any environmental limits that cannot be adequately substantiated by the endurance demonstration, validated analysis, or a combination thereof.

Special condition no. 10(f) requires Safran to evaluate various control system failures to ensure that such failures will not lead to unsafe engine conditions. The FAA issued Advisory Circular (AC) 33.28-3, “ Guidance Material for 14 CFR 33.28, Engine Control Systems,” on May 23, 2014 (AC 33.28-3), for reciprocating and turbine engines.^[5] This AC provides guidance for defining an engine control system failure when showing compliance with the requirements of § 33.28. AC 33.28-3 also includes objectives for control system integrity requirements, criteria for a loss of thrust control (LOTC) and loss of power control (LOPC) event, and an acceptable LOTC/LOPC rate. The electrical and electronic failures and failure rates did not account for electric engines when the FAA issued this AC, and therefore performance-based special conditions are established to allow fault accommodation criteria to be developed for electric engines.

The phrase “in the full-up configuration” used in special condition no. 10(f)(2) refers to a system without any fault conditions present. The electronic control system must, when in the full-up configuration, be single-fault tolerant, as determined by the Administrator, for electrical, electrically detectable, and electronic failures involving LOPC events.

The term “local” in the context of “local events” used in special condition no. 10(f)(4) means failures or malfunctions leading to events in the intended aircraft installation such as fire, overheat, or failures leading to damage to engine control system components. These “local events” must not result in a hazardous engine effect due to engine control system failures or malfunctions.

Special condition no. 10(g) requires Safran to conduct a safety assessment of the control system to support the safety analysis in special condition no. 17. This control system safety assessment provides engine response to failures, and rates of these failures that can be used at the aircraft-level safety assessment.

Special condition no. 10(h) requires Safran to provide appropriate protection devices or systems to ensure that engine operating limits will not be exceeded in service.

Special condition no. 10(i) is necessary to ensure that the controllers are self-sufficient and isolated from other aircraft systems. The aircraft-supplied data supports the analysis at the aircraft level to protect the aircraft from common mode failures that could lead to major propulsion power loss. The exception “other than power command signals from the aircraft,” noted in special condition no. 10(i), is based on the FAA's determination that the engine controller has no reasonable means to determine the validity of any in-range signals from the electrical power system. In many cases, the engine control system can detect a faulty signal from the aircraft, but the engine control ( print page 105435) system typically accepts the power command signal as a valid value.

The term “independent” in the context of “fully independent engine systems” referenced in special condition no. 10(i) means the controllers should be self-sufficient and isolated from other aircraft systems or provide redundancy that enables the engine control system to accommodate aircraft data system failures. In the case of loss, interruption, or corruption of aircraft-supplied data, the engine must continue to function in a safe and acceptable manner without hazardous engine effects.

The term “accommodated,” in the context of “detected and accommodated,” referenced in special condition 10(i)(2) is to assure that, upon detecting a fault, the system continues to function safely.

Special condition no. 10(j) requires Safran to show that the loss of electric power from the aircraft will not cause the electric engine to malfunction in a manner hazardous to the aircraft. The total loss of electric power to the electric engine may result in an engine shutdown.

Instrument Connection: Special condition no. 11 requires Safran to comply with § 33.29(a), (e), and (g), which set certain requirements for the connection and installation of instruments to monitor engine performance. The remaining requirements in § 33.29 apply only to technologies used in reciprocating and turbine aircraft engines.

Instrument connections (wires, wire insulation, potting, grounding, connector designs, etc.) must not introduce unsafe features or characteristics to the aircraft. Special condition no. 11 requires the safety analysis to include potential hazardous effects from failures of instrument connections to function properly. The outcome of this analysis might identify the need for design enhancements or additional ICA to ensure safety.

Stress Analysis: Section 33.62 requires applicants to perform a stress analysis on each turbine engine. This regulation is explicitly applicable only to turbine engines and turbine engine components, and it is not appropriate for the Safran Model ENGINeUS 100A1 electric engines. However, a stress analysis particular to these electric engines is necessary to account for stresses resulting from electric technology used in the engine.

Special condition no. 12 requires a mechanical, thermal, and electrical stress analysis to show that the engine has a sufficient design margin to prevent unacceptable operating characteristics. Also, the applicant must determine the maximum stresses in the engine by tests, validated analysis, or a combination thereof, and show that they do not exceed minimum material properties.

Critical and Life-Limited Parts: Special condition no. 13 requires Safran to show whether rotating or moving components, bearings, shafts, static parts, and non-redundant mount components should be classified, designed, manufactured, and managed throughout their service life as critical or life-limited parts.

The term “low-cycle fatigue,” referenced in special condition no. 13(a)(2), is a decline in material strength from exposure to cyclic stress at levels beyond the stress threshold the material can sustain indefinitely. This threshold is known as the “material endurance limit.” Low-cycle fatigue typically causes a part to sustain plastic or permanent deformation during the cyclic loading and can lead to cracks, crack growth, and fracture. Engine parts that operate at high temperatures and high mechanical stresses simultaneously can experience low-cycle fatigue coupled with creep. Creep is the tendency of a metallic material to permanently move or deform when it is exposed to the extreme thermal conditions created by hot combustion gasses, and substantial physical loads such as high rotational speeds and maximum thrust. Conversely, high-cycle fatigue is caused by elastic deformation, small strains caused by alternating stress, and a much higher number of load cycles compared to the number of cycles that cause low-cycle fatigue.

The engineering plan referenced in special condition no. 13(b)(1) informs the manufacturing and service management processes of essential information that ensures the life limit of a part is valid. The engineering plan provides methods for verifying the characteristics and qualities assumed in the design data using methods that are suitable for the part criticality. The engineering plan informs the manufacturing process of the attributes that affect the life of the part. The engineering plan, manufacturing plan, and service management plan are related in that assumptions made in the engineering plan are linked to how a part is manufactured and how that part is maintained in service. For example, environmental effects on life limited electric engine parts, such as humidity, might not be consistent with the assumptions used to design the part. Safran must ensure that the engineering plan is complete, available, and acceptable to the Administrator.

The term “manufacturing plan,” referenced in special condition no. 13(b)(2), is the collection of data required to translate documented engineering design criteria into physical parts, and to verify that the parts comply with the properties established by the design data. Because engines are not intentionally tested to failure during a certification program, documents and processes used to execute production and quality systems required by § 21.137 guarantee inherent expectations for performance and durability. These systems limit the potential manufacturing outcomes to parts that are consistently produced within design constraints.

The manufacturing plan and service management plan ensure that essential information from the engineering plan, such as the design characteristics that safeguard the integrity of critical and life-limited parts, is consistently produced and preserved over the lifetime of those parts. The manufacturing plan includes special processes and production controls to prevent inclusion of manufacturing-induced anomalies, which can degrade the part's structural integrity. Examples of manufacturing-induced anomalies are material contamination, unacceptable grain growth, heat-affected areas, and residual stresses.

The service-management plan ensures the method and assumptions used in the engineering plan to determine the part's life remain valid by enabling corrections identified from in-service experience, such as service-induced anomalies and unforeseen environmental effects, to be incorporated into the design process. The service-management plan also becomes the ICA for maintenance, overhaul, and repairs of the part.

Lubrication System: Special condition no. 14 requires Safran to ensure that the lubrication system is designed to function properly between scheduled maintenance intervals and to prevent contamination of the engine bearings. This special condition also requires Safran to demonstrate the unique lubrication attributes and functional capability of the Safran Model ENGINeUS 100A1 electric engine design.

The corresponding part 33 regulations include provisions for lubrication systems used in reciprocating and turbine engines. The part 33 requirements account for safety issues associated with specific reciprocating and turbine engine system configurations. These regulations are not appropriate for the Safran Model ENGINeUS 100A1 electric engines. For example, electric engines do not have a ( print page 105436) crankcase or lubrication oil sump. Electric engine bearings are sealed, so they do not require an oil circulation system. The lubrication system in these engines is also independent of the propeller pitch control system. Therefore, special condition no. 14 incorporates only certain requirements from the part 33 regulations.

Power Response: Special condition no. 15 requires the design and construction of the Safran Model ENGINeUS 100A1 electric engines to enable an increase from the minimum—

(1) power setting to the highest rated power without detrimental engine effects, and

(2) within a time interval appropriate for the intended aircraft application.

The engine control system governs the increase or decrease in power in combustion engines to prevent too much (or too little) fuel from being mixed with air before combustion. Due to the lag in rotor response time, improper fuel/air mixtures can result in engine surges, stalls, and exceedances above rated limits and durations. Failure of the combustion engine to provide thrust, maintain rotor speeds below rotor burst thresholds, and keep temperatures below limits can have engine effects detrimental to the aircraft. Similar detrimental effects are possible in the Safran Model ENGINeUS 100A1 electric engines, but the causes are different. Electric engines with reduced power response time can experience insufficient thrust to the aircraft, shaft over-torque, and over-stressed rotating components, propellers, and critical propeller parts. Therefore, this special condition is necessary.

Continued Rotation: Special condition no. 16 requires Safran to design the Model ENGINeUS 100A1 electric engines such that, if the main rotating systems continue to rotate after the engine is shut down while in-flight, this continued rotation will not result in any hazardous engine effects.

The main rotating system of the Safran Model ENGINeUS 100A1 electric engines consists of the rotors, shafts, magnets, bearings, and wire windings that convert electrical energy to shaft torque. For the initial aircraft application, this rotating system must continue to rotate after the power source to the engine is shut down. The safety concerns associated with this special condition are substantial asymmetric aerodynamic drag that can cause aircraft instability, loss of control, and reduced efficiency; and may result in a forced landing or inability to continue safe flight.

Safety Analysis: Special condition no. 17 requires Safran to comply with § 33.75(a)(1) and (a)(2), which require the applicant to conduct a safety analysis of the engine, and which would otherwise be applicable only to turbine aircraft engines. Additionally, this special condition requires Safran to assess its engine design to determine the likely consequences of failures that can reasonably be expected to occur. The failure of such elements, and associated prescribed integrity requirements, must be stated in the safety analysis.

A primary failure mode is the manner in which a part is most likely going to fail. Engine parts that have a primary failure mode, a predictable life to the failure, and a failure consequence that results in a hazardous effect, are life-limited or critical parts. Some life-limited or critical engine parts can fail suddenly in their primary failure mode, from prolonged exposure to normal engine environments such as temperature, vibration, and stress, if those engine parts are not removed from service before the damage mechanisms progress to a failure. Due to the consequence of failure, these parts are not allowed to be managed by on-condition or probabilistic means because the probability of failure cannot be sensibly estimated in numerical terms. Therefore, the parts are managed by compliance with integrity requirements, such as mandatory maintenance (life limits, inspections, inspection techniques), to ensure the qualities, features, and other attributes that prevent the part from failing in its primary failure mode are preserved throughout its service life. For example, if the number of engine cycles to failure are predictable and can be associated with specific design characteristics, such as material properties, then the applicant can manage the engine part with life limits.

Complete or total power loss is not assumed to be a minor engine event, as it is in the turbine engine regulation § 33.75, to account for experience data showing a potential for higher hazard levels from power loss events in single-engine general aviation aircraft. The criteria in these special conditions apply to an engine that continues to operate at partial power after a single electrical or electronic fault or failure. Total loss of power is classified at the aircraft level using special condition nos. 10(g) and 33(h).

Ingestion: Special condition no. 18 requires Safran to ensure that these engines will not experience unacceptable power loss or hazardous engine effects from ingestion. The associated regulations for turbine engines, §§ 33.76, 33.77, and 33.78, are based on potential performance impacts and damage from birds, ice, rain, and hail being ingested into a turbine engine that has an inlet duct, which directs air into the engine for combustion, cooling, and thrust. By contrast, the Safran Model ENGINeUS 100A1 electric engines are not configured with inlet ducts.

An “unacceptable” power loss, as used in special condition no. 18(b), is such that the power or thrust required for safe flight of the aircraft becomes unavailable to the pilot. The specific amount of power loss that is required for safe flight depends on the aircraft configuration, speed, altitude, attitude, atmospheric conditions, phase of flight, and other circumstances where the demand for thrust is critical to safe operation of the aircraft.

Liquid and Gas Systems: Special condition no. 19 requires Safran to ensure that systems used for lubrication or cooling of engine components are designed and constructed to function properly. Also, if a system is not self-contained, the interfaces to that system would be required to be defined in the engine installation manual. Systems for the lubrication or cooling of engine components can include heat exchangers, pumps, fluids, tubing, connectors, electronic devices, temperature sensors and pressure switches, fasteners and brackets, bypass valves, and metallic chip detectors. These systems allow the electric engine to perform at extreme speeds and temperatures for durations up to the maintenance intervals without exceeding temperature limits or predicted deterioration rates.

Vibration Demonstration: Special condition no. 20 requires Safran to ensure the engine—

(1) is designed and constructed to function throughout its normal operating range of rotor speeds and engine output power without inducing excessive stress caused by engine vibration, and

(2) design undergoes a vibration survey.

The vibration demonstration is a survey that characterizes the vibratory attributes of the engine. It verifies that the stresses from vibration do not impose excessive force or result in natural frequency responses on the aircraft structure. The vibration demonstration also ensures internal vibrations will not cause engine components to fail. Excessive vibration force occurs at magnitudes and forcing functions or frequencies, which may result in damage to the aircraft. Stress margins to failure add conservatism to the highest values predicted by analysis for additional protection from failure ( print page 105437) caused by influences beyond those quantified in the analysis. The result of the additional design margin is improved engine reliability that meets prescribed thresholds based on the failure classification. The amount of margin needed to achieve the prescribed reliability rates depends on an applicant's experience with a product. The FAA considers the reliability rates when deciding how much vibration is “excessive.”

Overtorque: Special condition no. 21 requires Safran to demonstrate that the engine is capable of continued operation without the need for maintenance if it experiences a certain amount of overtorque.

Safran's electric engine converts electrical energy to shaft torque, which is used for propulsion. The electric motor, controller, and high-voltage systems control the engine torque. When the pilot commands power or thrust, the engine responds to the command and adjusts the shaft torque to meet the demand. During the transition from one power or thrust setting to another, a small delay, or latency, occurs in the engine response time. While the engine dwells in this time interval, it can continue to apply torque until the command to change the torque is applied by the engine control. The allowable amount of overtorque during operation depends on the engine's response to changes in the torque command throughout its operating range.

Calibration Assurance: Special condition no. 22 requires Safran to subject the engine to calibration tests to establish its power characteristics and the conditions both before and after the endurance and durability demonstrations specified in special condition nos. 23 and 26. The calibration test requirements specified in § 33.85 only apply to the endurance test specified in § 33.87, which is applicable only to turbine engines. The FAA determined that the methods used for accomplishing those tests for turbine engines are not appropriate for electric engines. The calibration tests in § 33.85 have provisions applicable to ratings that are not relevant to the Safran Model ENGINeUS 100A1 electric engines. Special condition no. 22 allows Safran to demonstrate the endurance and durability of the electric engine either together or independently, whichever is most appropriate for the engine qualities being assessed. Consequently, the special condition applies the calibration requirement to both the endurance and durability tests.

Endurance Demonstration: Special condition no. 23 requires Safran to perform an endurance demonstration test that is acceptable to the Administrator. The Administrator will evaluate the extent to which the test exposes the engine to failures that could occur when the engine is operated at up to its rated values, and determine if the test is sufficient to show that the engine design will not exhibit unacceptable effects in service, such as significant performance deterioration, operability restrictions, and engine power loss or instability, when it is run repetitively at rated limits and durations in conditions that represent extreme operating environments.

Temperature Limit: Special condition no. 24 requires Safran to ensure the engine can endure operation at its temperature limits plus an acceptable margin. An “acceptable margin,” as used in the special condition, is the amount of temperature above that required to prevent the least capable engine allowed by the type design, as determined by § 33.8, from failing due to temperature-related causes when operating at the most extreme engine and environmental thermal conditions.

Operation Demonstration: Special condition no. 25 requires the engine to demonstrate safe operating characteristics throughout its declared flight envelope and operating range. Engine operating characteristics define the range of functional and performance values the Safran Model ENGINeUS 100A1 electric engines can achieve without incurring hazardous effects. The characteristics are requisite capabilities of the type design that qualify the engine for installation into aircraft and that determine aircraft installation requirements. The primary engine operating characteristics are assessed by the tests and demonstrations that would be required by these special conditions. Some of these characteristics are shaft output torque, rotor speed, power consumption, and engine thrust response. The engine performance data Safran will use to certify the engine must account for installation loads and effects. These are aircraft-level effects that could affect the engine characteristics that are measured when the engine is tested on a stand or in a test cell. These effects could result from elevated inlet cowl temperatures, aircraft maneuvers, flowstream distortion, and hard landings. For example, an engine that is run in a sea-level, static test facility could demonstrate more capability for some operating characteristics than it will have when operating on an aircraft in certain flight conditions. Discoveries like this during certification could affect engine ratings and operating limits. Therefore, the installed performance defines the engine performance capabilities.

Durability Demonstration: Special condition no. 26 requires Safran to subject the engine to a durability demonstration. The durability demonstration must show that the engine is designed and constructed to minimize the development of any unsafe condition between maintenance intervals or between engine replacement intervals if maintenance or overhaul is not defined. The durability demonstration also verifies that the ICA is adequate to ensure the engine, in its fully deteriorated state, continues to generate rated power or thrust, while retaining operating margins and sufficient efficiency, to support the aircraft safety objectives. The amount of deterioration an engine can experience is restricted by operating limitations and managed by the engine ICA. Section 33.90 specifies how maintenance intervals are established; it does not include provisions for an engine replacement. Electric engines and turbine engines deteriorate differently. Therefore, Safran will use different test effects to develop maintenance, overhaul, or engine replacement information for their electric engine.

System and Component Tests: Special condition no. 27 requires Safran to show that the systems and components of the engine perform their intended functions in all declared engine environments and operating conditions.

Sections 33.87 and 33.91, which are specifically applicable to turbine engines, have conditional criteria to decide if additional tests will be required after the engine tests. The criteria are not suitable for electric engines. Part 33 associates the need for additional testing with the outcome of the § 33.87 endurance test because it is designed to address safety concerns in combustion engines. For example, § 33.91(b) requires the establishment of temperature limits for components that require temperature-controlling provisions, and § 33.91(a) requires additional testing of engine systems and components where the endurance test does not fully expose internal systems and components to thermal conditions that verify the desired operating limits. Exceeding temperature limits is a safety concern for electric engines. The FAA determined that the § 33.87 endurance test is not appropriate for testing the electronic components of electric engines because mechanical energy is generated differently by electronic systems than it is by the thermal conditions in turbine engines. ( print page 105438) Additional safety considerations also need to be addressed in the test. Therefore, special condition no. 27 is a performance-based requirement that allows Safran to determine when engine systems and component tests are necessary and to determine the appropriate limitations of those systems and components used in the Safran Model ENGINeUS 100A1 electric engine.

Rotor Locking Demonstration: Special condition no. 28 requires the engine to demonstrate reliable rotor locking performance and that no hazardous effects will occur if the engine uses a rotor locking device to prevent shaft rotation.

Some engine designs enable the pilot to prevent a propeller shaft or main rotor shaft from turning while the engine is running, or the aircraft is in-flight. This capability is needed for some installations that require the pilot to confirm the functionality of certain flight systems before takeoff. The Safran engine installations are not limited to aircraft that will not require rotor locking. Section 33.92 prescribes a test that may not include the appropriate criteria to demonstrate sufficient rotor locking capability for these engines. Therefore, this special condition is necessary.

The special condition does not define “reliable” rotor locking but allows Safran to classify the hazard as major or minor and assign the appropriate quantitative criteria that meet the safety objectives required by special condition no. 17 and the applicable portions of § 33.75.

Teardown Inspection: Special condition no. 29 requires Safran to perform a teardown or non-teardown evaluation after the endurance, durability, and overtorque demonstrations, based on the criteria in special condition no. 29(a) or (b).

Special condition no. 29(b) includes restrictive criteria for “non-teardown evaluations” to account for electric engines, sub-assemblies, and components that cannot be disassembled without destroying them. Some electrical and electronic components like Safran's are constructed in an integrated fashion that precludes the possibility of tearing them down without destroying them. The special condition indicates that, if a teardown cannot be performed in a non-destructive manner, then the inspection or replacement intervals must be established based on the endurance and durability demonstrations. The procedure for establishing maintenance should be agreed upon between the applicant and the FAA prior to running the relevant tests. Data from the endurance and durability tests may provide information that can be used to determine maintenance intervals and life limits for parts. However, if life limits are required, the lifing procedure is established by special condition no. 13, Critical and Life-Limited Parts, which corresponds to § 33.70. Therefore, the procedure used to determine which parts are life-limited, and how the life limits are established, requires FAA approval, as it does for § 33.70. Sections 33.55 and 33.93 do not contain similar requirements because reciprocating and turbine engines can be completely disassembled for inspection.

Containment: Special condition no. 30 requires the engine to have containment features that protect against likely hazards from rotating components, unless Safran can show the margin to rotor burst does not justify the need for containment features. Rotating components in electric engines are typically disks, shafts, bearings, seals, orbiting magnetic components, and the assembled rotor core. However, if the margin to rotor burst does not unconditionally rule out the possibility of a rotor burst, then the special condition requires Safran to assume a rotor burst could occur and design the stator case to contain the failed rotors, and any components attached to the rotor that are released during the failure. In addition, Safran must also determine the effects of subsequent damage precipitated by a main rotor failure and characterize any fragments that are released forward or aft of the containment features. Further, decisions about whether the Safran engine requires containment features, and the effects of any subsequent damage following a rotor burst, should be based on test or validated analysis. The fragment energy levels, trajectories, and size are typically documented in the installation manual because the aircraft will need to account for the effects of a rotor failure in the aircraft design. The intent of this special condition is to prevent hazardous engine effects from structural failure of rotating components and parts that are built into the rotor assembly.

General Conduct of Tests: Special condition no. 32 requires Safran to include scheduled maintenance in the engine ICA, include any maintenance, in addition to the scheduled maintenance that was needed during the test to satisfy the applicable test requirements, and conduct any additional tests that the Administrator finds necessary, as warranted by the test results.

For example, certification endurance test shortfalls might be caused by omitting some prescribed engine test conditions, or from accelerated deterioration of individual parts arising from the need to force the engine to operating conditions that drive the engine above the engine cycle values of the type design. If an engine part fails during a certification test, the entire engine might be subjected to penalty runs, with a replacement or newer part design installed on the engine, to meet the test requirements. Also, the maintenance performed to replace the part, so that the engine could complete the test, would be included in the engine ICA. In another example, if the applicant replaces a part before completing an engine certification test because of a test facility failure and can substantiate the part to the Administrator through bench testing, they might not need to substantiate the part design using penalty runs with the entire engine.

The term “excessive” is used to describe the frequency of unplanned engine maintenance, and the frequency of unplanned test stoppages, to address engine issues that prevent the engine from completing the tests in special condition nos. 32(b)(1) and (2), respectively. Excessive frequency is an objective assessment from the FAA's analysis of the amount of unplanned maintenance needed for an engine to complete a certification test. The FAA's assessment may include the reasons for the unplanned maintenance, such as the effects test facility equipment may have on the engine, the inability to simulate a realistic engine operating environment, and the extent to which an engine requires modifications to complete a certification test. In some cases, the applicant may be able to show that unplanned maintenance has no effect on the certification test results, or they might be able to attribute the problem to the facility or test-enabling equipment that is not part of the type design. In these cases, the ICA will not be affected. However, if Safran cannot reconcile the amount of unplanned service, then the FAA may consider the unplanned maintenance required during the certification test to be “excessive,” prompting the need to add the unplanned maintenance to mandatory ICA to comply with the certification requirements.

Engine electrical systems: The current requirements in part 33 for electronic engine control systems were developed to maintain an equivalent level of safety demonstrated by engines that operate with hydromechanical engine control systems. At the time § 33.28 was ( print page 105439) codified, the only electrical systems used on turbine engines were low-voltage, electronic engine control systems (EEC) and high-energy spark-ignition systems. Electric aircraft engines use high-voltage, high-current electrical systems and components that are physically located in the motor and motor controller. Therefore, the existing part 33 control system requirements do not adequately address all the electrical systems used in electric aircraft engines. Special condition no. 33 is established using the existing engine control systems requirement as a basis. It applies applicable airworthiness criteria from § 33.28 and incorporates airworthiness criteria that recognize and focus on the electrical power system used in the engine.

Special condition no. 33(b) ensures that all aspects of an electrical system, including generation, distribution, and usage, do not experience any unacceptable operating characteristics.

Special condition no. 33(c) requires the electrical power distribution aspects of the electrical system to provide the safe transfer of electrical energy throughout the electric engine.

The term “abnormal conditions” used in special condition no. 33(c)(2) is intended to be consistent with the definitions in MIL-STD-704F “Aircraft Electric Power Characteristics” which defines normal operation and abnormal operation. MIL-STD-704F is a standard that ensures compatibility between power sources that provide power to the aircraft's electrical systems and airborne equipment that receive power from the power source. This standard also establishes technical criteria for aircraft electric power. The term “abnormal conditions” refers to various engine operating conditions such as:

• System or component characteristics outside of normal statistical variation from circumstances such as systems degradation, installation error, and engine response to fault conditions;
• Unusual environmental conditions from extreme temperature, humidity, vibration, lightning, high-intensity radiated field (HIRF), atmospheric neutron radiation; and
• Unusual and infrequent events such as landing on icy runways, rejected take-offs or go-arounds, extended ground idling or taxiing in a hot environment, and abrupt load changes from foreign object damage or engine contamination.

The phrase “safe transmission of electric energy” used in special condition no. 33(c)(3) refers to the transmission of electrical energy in a manner that supports the operation of the electric engine(s) and the aircraft safety objectives without detrimental effects such as uncontrolled fire or structural failure due to severe overheating.

Special condition no. 33(d) requires the engine electrical system to be designed such that the loss, malfunction, or interruption of the electrical power source, or power conditions that exceed design limits, will not result in a hazardous engine effect.

Special condition no. 33(e) requires Safran to identify and declare, in the engine installation manual, the characteristics of any electrical power supplied from the aircraft to the engine, or electrical power supplied from the engine to the aircraft via energy regeneration, and any other characteristics necessary for safe operation of the engine.

Special condition no. 33(f) requires Safran to demonstrate that systems and components will operate properly up to environmental limits, using special conditions, when such limits cannot be adequately substantiated by the endurance demonstration, validated analysis, or a combination thereof. The environmental limits referred to in this special condition include temperature, vibration, HIRF, and others addressed in RTCA DO-160G, “Environmental Conditions and Test Procedures for Airborne Electronic/Electrical Equipment and Instruments.”

Special condition 33(g) requires Safran to evaluate various electric engine system failures to ensure that these failures will not lead to unsafe engine conditions. The evaluation includes single-fault tolerance, ensures no single electrical or electronic fault or failure would result in hazardous engine effects, and ensures that any failure or malfunction leading to local events in the intended aircraft application do not result in certain hazardous engine effects. The special condition also implements integrity requirements, criteria for LOTC/LOPC events, and an acceptable LOTC/LOPC rate.

Special condition 33(h) requires Safran to conduct a safety assessment of the engine electrical system to support the safety analysis in special condition no. 17. This safety assessment provides engine response to failures, and rates of these failures, which can be used at the aircraft safety assessment level.

Discussion of Comments

The FAA issued a notice of proposed special conditions (NPSC) Docket No. FAA-2023-0587 for the Safran Model ENGINeUS 100A1 electric engines, which was published in the Federal Register on March 20, 2024 (89 FR 19763).

The FAA received responses from four commenters, Airbus Helicopters (Airbus), Ampaire Inc. (Ampaire), Kite Magnetics Pty Ltd. (Kite Magnetics), and magniX USA, Inc. (magniX).

The FAA received one comment from Airbus that stated proposed special condition no. 4, Fire Protection, does not prescribe safety criteria for flammable cooling fluids and suggested that a fireproof wall, cooling fluid shut-off valve, fluid draining system, and fire detection system may be necessary because a potential ignition source (electrical failure) and flammable fluids share the same area in the aircraft.

The FAA does not concur with Airbus's comment that special condition no. 4 does not prescribe safety criteria for flammable cooling fluids. Special condition no. 4 incorporates § 33.17(b) through (g) into the Safran electric engine certification basis, which include provisions for flammable fluid. The FAA also revised special condition no 4. slightly to clarify that § 33.17(b) through (g) are required as part of that special condition.

The FAA received several comments from Ampaire.

Ampaire asked if the FAA determined that the definition of propeller options for part 33 electric propulsion systems are sufficiently covered by existing reciprocating and gas turbine regulations.

These special conditions are applicable to the Safran electric engine, which will be used with fixed pitch propellers. The existing requirements for reciprocating and gas turbine regulations are sufficient for the conventional fixed-pitch propellers and therefore no other propeller options are required. No changes were made as a result of this comment.

Ampaire regarded proposed special condition no. 10(f)(4) regarding engine control system failures as very similar to the corresponding part 33 regulation (§ 33.28(d)(4)), but noted that special condition is harder to understand without examples that describe the term “local events” such as those provided in the original part 33 regulation. Ampaire recommended adding the examples to special condition no. 10(f)(4) or including other more relevant examples.

The examples Ampaire requested are already in the preamble discussion for special condition no. 10(f)(4). The FAA did not intend to create a new definition of “local events.” As explained in the preamble, the term “local events” means failures or malfunctions leading to events in the intended aircraft installation such as fire, overheat, or ( print page 105440) failures leading to damage to engine control system components. No changes were made as a result of this comment.

Ampaire stated a system safety assessment is required by § 33.28 but there is no requirement in part 33 to add the rates of hazardous and major faults in the installation manual. Ampaire asked the FAA to explain why this requirement is included in special condition no. 10(g) for the Safran electric engine but not in part 33 for reciprocating and gas turbine engines.

The FAA added the requirement because electric engines enable a wide variety of new aircraft propulsion features, and the engine control system safety assessment is tied to these new propulsion features, which support aircraft that combine vertical takeoff and landing, multi-engine distributed-propulsion, propeller lift and tilt-wing functions, and zero velocity inflight maneuvering capabilities. The effects of an engine failure, such as power loss from an engine, and hazards to the aircraft are contingent on the aircraft design. Therefore, the hazards identified in the safety analysis, as well as the hazard level rates, are included in the engine installation manual to ensure any assumptions about aircraft capabilities that mitigate the effects of engine failures are taken into account when deciding if an engine can be installed in an aircraft. No changes were made as a result of this comment.

Ampaire asked the FAA to explain why the reference to special condition no. 31, Operation with variable pitch propeller, is included in the magniX special condition no. 17(d)(1), Safety analysis, but not the Safran proposed special condition no. 17(d)(1).

Safran's electric engine will be used with a fixed-pitch propeller, and therefore special condition no. 31 is not applicable to the Safran engine type design. No changes were made as a result of this comment.

Ampaire stated proposed special condition no. 23, Endurance demonstration, implies that endurance testing requires a demonstration of energy regeneration, but energy regeneration might not be a feature for some electric engines that operate normally at their limits. Ampaire suggested replacing the second sentence in special condition no. 23 with “The endurance demonstration must include dwellings and increases and decreases of the engine's power settings for sufficient durations that produce the extreme physical conditions the engine experiences at rated performance levels, operational limits, and at any other conditions or power settings including energy regeneration that are required to verify the limit capabilities of the engine.”

The FAA concurs with Ampaire's comment that energy regeneration might not be a feature for some electric engines that operate at their limits. The phrase “that produce the extreme physical conditions” in special condition no. 23 indicates the endurance test addresses engine properties where the extreme physical conditions can occur including conditions that cause the engine to operate at its limits of energy regeneration. As a result of this comment, the FAA changed special condition no. 23 in accordance with Ampaire's recommendation.

Ampaire requested the FAA revise special condition nos. 33(c)(1) and (d) for electrical power distribution and protection systems, respectively, by adding the conditional statement “due to a single fault” and explained electrical power distribution within the part 33 powerplant may take several faults to result in total loss. Ampaire also stated that electric power distribution outside the part 33 powerplant is the subject of part 23 aircraft certification.

The FAA does not concur with Ampaire's request to revise special condition nos. 33(c)(1) and (d) to provide protection from potential consequences resulting only from single electrical faults. Special condition nos. 33(g)(2) and (3), Electrical system failures, have safety criteria that already address single faults in all the engine electrical systems. The safety criteria in special condition no. 33(c)(1) and (d) are for loss of function in electrical power distribution systems, and the criteria apply regardless of the cause of system failures or malfunctions. Also, part 33 has provisions for electrical power supplied to electrical control systems, and therefore this special condition is within the scope of engine requirements. No changes were made as a result of this comment.

Ampaire asked the FAA to explain the regulatory significance of the term “detrimental” as it is used in proposed special condition no. 33, Engine electrical systems, and whether the term relates to hazard levels.

The FAA intends the term “detrimental” to have the same meaning as the meaning of the term as it is commonly used in the English language. The term is used extensively in existing FAA regulations and guidance. There is no intent to change how the term is used in these special conditions. Also, there is no correlation between the term “detrimental” and engine failure effect hazard levels. The term is intended to capture all engine effects that could result in an unsafe engine condition. No changes were made as a result of this comment.

The FAA received several comments from magniX.

MagniX noted proposed special condition nos. 1(b) and (c) state that a means of compliance, which may include consensus standards, must be “accepted by the Administrator” and “in a form and manner acceptable to the Administrator.” MagniX explained that these paragraphs are directly out of 14 CFR 23.2010, which contains performance-based language. MagniX also explained that part 33 and the Safran special conditions are prescriptive regulations, not performance-based. MagniX further indicated that requiring a performance-based process for establishing means of compliance with prescriptive regulations is unnecessary and overly burdensome to applicants and regulators. MagniX recommended the FAA not adopt proposed special condition nos. 1(b) and (c).

The FAA does not concur with magniX's recommendation. The FAA considers special condition nos. 1(b) and (c) to be essential for achieving an equivalent level of safety to the level of safety provided by the part 33 engine requirements. The Safran electric engine criteria are a combination of part 33 requirements and special conditions to the requirements in part 33. Special conditions are developed under the provisions of § 21.16, which are issued when the applicable regulations do not contain adequate or appropriate safety standards. Special condition nos. 1(b) and (c) will be used to incorporate the additional details that apply to the Safran engine design using accepted means of compliance. No changes were made as a result of this comment.

MagniX stated proposed special condition nos. 10(g), 15(b), and 17(f) would require applicants to declare proprietary information in the engine installation manual, these documentation requirements establish a precedent beyond that required of their existing reciprocating or turbine engine counterparts, and these requirements increase the risk that sensitive information is disclosed. MagniX explained that while it is understood this information is used during aircraft-level certification efforts, traditional data sharing agreements sufficiently provide the integrator with the required information while respecting the proprietary nature of the data. MagniX also stated requiring these additional data in the engine installation manual overly constrains the means of ( print page 105441) compliance and introduces commercial risk. MagniX recommended the FAA not adopt the requirement to include these specific disclosures in the engine installation manual. MagniX proposed that these data be provided to integrators through generic “installation instructions” in lieu of the engine installation manual and explained this will allow specific proprietary disclosures in other installation documents such as interface control drawings, technical memorandums, or other installer requested documentation.

Special condition nos. 10(g), 15(b), and 17(f) do not require the disclosure of sensitive information. As discussed in the NPSC, the documentation requirements in special conditions nos. 10(g), 15(b), and 17(f) are expected to ensure that the engine is used safely and properly by constraining the installation of electric engines to only aircraft types (configurations, flight capabilities, etc.) that were used by the engine manufacturer to determine the engine ratings, limits, performance characteristics, as well as the reliability and criticality of engine systems and parts.

These documentation requirements are intended, and the FAA finds necessary, to ensure enough information is included to safeguard compatibility between the electric engine and aircraft, and to prevent the engine from being used in an aircraft type that requires safety features or performance characteristics that are not available from an engine that was type-certificated for an aircraft that does not require the same safety features or performance characteristics. The FAA modified the proposed special conditions to clarify the requirement by specifying the information identified in special condition nos. 6 “Engine cooling,” 10 “Engine control systems,” 15 “Power response,” 17 “Safety analysis,” 18 “Ingestion,” 19 “Liquid and gas systems,” 30 “Containment,” and 33 “Engine electrical systems” must be documented and provided to the installer as part of the requirements in § 33.5.

The FAA received several comments from Kite Magnetics.

Kite Magnetics stated that special condition no. 14 for the lubrication system of the Safran Model ENGINeUS 100A1 electric engine should focus specifically on the unique lubrication attributes and inherent functional capabilities of the Safran electric engine design, rather than apply requirements for the entire lubrication system. Kite Magnetics suggested changing special condition no. 14 to apply component-level requirements that would be better suited for the unique attributes of electric engines such as the Safran Model ENGINe US 100A1, promote clarity and relevance of the special condition to critical aspects of the lubrication system pertinent to electric engines, and avoid unnecessary requirements that do not apply to this engine type.

The FAA does not concur with Kite Magnetics' comment that the special conditions for an electric engine lubrication system should be established at the component level. These special conditions are engine-level requirements; however, the means of compliance to the special conditions can involve component-level assessments using special condition no. 27, System and component tests, which can focus on the unique lubrication attributes and inherent functional capabilities of the Safran electric engine design. No changes were made as a result of this comment.

Kite Magnetics stated the language “Any system or device that provides, uses, conditions, or distributes electrical power, and is part of the engine type design” in proposed special condition no. 33(a) could imply that energy storage systems (ESS) are part of the engine electrical system. Kite Magnetics explained that ESS fall under the category of systems that provide electrical power and may be perceived as part of the engine's electrical system. However, Kite Magnetics noted that an ESS is a distinct system that supports the engine's electrical power needs, but it is not inherently integrated into the engine's core electrical system design. Kite Magnetics requested confirmation that special condition 33(a) does not apply to ESS. Kite Magnetics did not request changes to this special condition.

The FAA confirms special condition 33(a) does not apply to ESS. No changes were made as a result of this comment.

Kite Magnetics requested clarification regarding the components and devices that are considered part of the engine's electrical power distribution system, as outlined in proposed special condition no. 33(c). Kite Magnetics explained this request is intended to ensure a clear understanding of the scope and components included within the electrical power distribution system. Kite Magnetics did not request changes to this special condition.

The FAA confirms special condition no. 33(c) applies only to the electrical power distribution systems that are part of Safran's electric engine type design. However, the partition between the engine and aircraft electrical power distribution systems must be clearly described and documented with the data provided for showing compliance to § 33.5(a). No changes were made as a result of this comment.

The FAA also determined that the following changes are necessary.

The phrase “In addition” is added to special condition no. 4, Fire protection, to connect the introduction sentence to (a) and (b) and avoid confusion.

The phrase “as defined in special condition no. 17 of these special conditions” is also added where the term “hazardous engine effects” is mentioned in these special conditions.

The applicability of special condition no. 33(b) “Electrical systems” to electrical load shedding is clarified to affect the electrical system only when required.

The term “electrical power plant” is changed to “powerplant” in special condition no. 33(c)(1), which is a term used in part 23, subpart E.

Definitions of the terms “abnormal condition” used in special condition no. 33(c)(2) and “safe transmission” used in special condition no. 33(c)(3) are included in the preamble discussion for special condition no. 33.

Special condition no. 33 was modified to provide flexibility in electric engine protection system designs. Special condition no. 33(c)(3) is changed to, “The system must provide mechanical or automatic means of isolating a faulted electrical-energy generation or storage device from leading to hazardous engine effects, as defined in special condition no. 17(d)(2) of these special conditions, or detrimental effects in the intended aircraft application.” The phrase, “or detrimental engine effects in the intended aircraft application” is also relocated to special condition no. 33(c)(3) to maintain the connection with special condition no. 33(g).

Special condition nos. 33(e)(1) and (e)(2) are both required and therefore “or” is replaced with “and” in special condition no. 33(e)(1), “Electrical power characteristics.”

The documentation requirement in special condition no. 10(g) is also applied to special condition no. 33 (h) “Engine Electrical Systems—System Safety Assessment.”

The FAA did not adopt proposed special condition no. 31 “Operation with a variable pitch propeller” because the Safran Model ENGINeUS 100A1 electric engine will not use a variable pitch propeller.

Except as discussed above, these special conditions are adopted as proposed. ( print page 105442)

Applicability

As discussed above, these special conditions are applicable to Safran Model ENGINeUS 100A1 electric engines. Should Safran apply at a later date for a change to the type certificate to include another model on the same type certificate, incorporating the same novel or unusual design feature, these special conditions would apply to that model as well.

Conclusion

This action affects only Safran Model ENGINeUS 100A1 electric engines. It is not a rule of general applicability.

List of Subjects in 14 CFR Part 33

• Aircraft
• Aviation safety
• Reporting and recordkeeping requirements

Authority Citation

The authority citation for these special conditions is as follows:

Authority: 49 U.S.C. 106(f), 106(g), 40113, 44701, 44702, 44704.

The Special Conditions

Accordingly, pursuant to the authority delegated to me by the Administrator, the following special conditions are issued as part of the type certification basis for Safran Model ENGINeUS 100A1 electric engines. The applicant must also comply with the certification procedures set forth in part 21.

(1) Applicability

(a) Unless otherwise noted in these special conditions, the engine design must comply with the airworthiness standards for aircraft engines set forth in part 33, except for those airworthiness standards that are specifically and explicitly applicable only to reciprocating and turbine aircraft engines or as specified herein.

(b) The applicant must comply with this part using a means of compliance, which may include consensus standards, accepted by the Administrator.

(c) The applicant requesting acceptance of a means of compliance must provide the means of compliance to the FAA in a form and manner acceptable to the Administrator.

(2) Engine Ratings and Operating Limits

In addition to § 33.7(a), the engine ratings and operating limits must be established and included in the type certificate data sheet based on:

(a) Shaft power, torque, rotational speed, and temperature for:

(1) Rated takeoff power;

(2) Rated maximum continuous power; and

(3) Rated maximum temporary power and associated time limit.

(b) Duty cycle and the rating at that duty cycle. The duty cycle must be declared in the engine type certificate data sheet.

(c) Cooling fluid grade or specification.

(d) Power-supply requirements.

(e) Any other ratings or limitations that are necessary for the safe operation of the engine.

(3) Materials

The engine design must comply with § 33.15.

(4) Fire Protection

The engine design must comply with § 33.17(b) through (g). In addition—

(a) The design and construction of the engine and the materials used must minimize the probability of the occurrence and spread of fire during normal operation and failure conditions and must minimize the effect of such a fire.

(b) High-voltage electrical wiring interconnect systems must be protected against arc faults that can lead to hazardous engine effects as defined in special condition no. 17(d)(2) of these special conditions. Any non-protected electrical wiring interconnects must be analyzed to show that arc faults do not cause a hazardous engine effect.

(5) Durability

The engine design and construction must minimize the development of an unsafe condition of the engine between maintenance intervals, overhaul periods, or mandatory actions described in the applicable ICA.

(6) Engine Cooling

The engine design and construction must comply with § 33.21. In addition, if cooling is required to satisfy the safety analysis as described in special condition no. 17 of these special conditions, the cooling system monitoring features and usage must be documented in the and provided to the installer as part of the requirements in § 33.5.

(7) Engine Mounting Attachments and Structure

The engine mounting attachments and related engine structures must comply with § 33.23.

(8) Accessory Attachments

The engine must comply with § 33.25.

(9) Overspeed

(a) A rotor overspeed must not result in a burst, rotor growth, or damage that results in a hazardous engine effect, as defined in special condition no. 17(d)(2) of these special conditions. Compliance with this paragraph must be shown by test, validated analysis, or a combination of both. Applicable assumed rotor speeds must be declared and justified.

(b) Rotors must possess sufficient strength with a margin to burst above certified operating conditions and above failure conditions leading to rotor overspeed. The margin to burst must be shown by test, validated analysis, or a combination thereof.

(c) The engine must not exceed the rotor speed operational limitations that could affect rotor structural integrity.

(10) Engine Control Systems

(a) Applicability. The requirements of this special condition apply to any system or device that is part of the engine type design that controls, limits, monitors, or protects engine operation, and is necessary for the continued airworthiness of the engine.

(b) Engine control. The engine control system must ensure that the engine does not experience any unacceptable operating characteristics or exceed its operating limits, including in failure conditions where the fault or failure results in a change from one control mode to another, from one channel to another, or from the primary system to the back-up system, if applicable.

(c) Design Assurance. The software and complex electronic hardware, including programmable logic devices, must be—

(1) Designed and developed using a structured and systematic approach that provides a level of assurance for the logic commensurate with the hazard associated with the failure or malfunction of the systems in which the devices are located; and

(2) Substantiated by a verification methodology acceptable to the Administrator.

(d) Validation. All functional aspects of the control system must be substantiated by test, analysis, or a combination thereof, to show that the engine control system performs the intended functions throughout the declared operational envelope.

(e) Environmental Limits. Environmental limits that cannot be adequately substantiated by endurance demonstration, validated analysis, or a combination thereof must be demonstrated by the system and component tests in special condition no. 27 of these special conditions.

(f) Engine control system failures. The engine control system must— ( print page 105443)

(1) Have a maximum rate of loss of power control (LOPC) that is suitable for the intended aircraft application. The estimated LOPC rate must be documented and provided to the installer as part of the requirements in § 33.5;

(2) When in the full-up configuration, be single-fault tolerant, as determined by the Administrator, for electrical, electrically detectable, and electronic failures involving LOPC events;

(3) Not have any single failure that results in hazardous engine effects as defined in special condition no. 17(d)(2) of these special conditions; and

(4) Ensure failures or malfunctions that lead to local events in the aircraft do not result in hazardous engine effects, as defined in special condition no. 17(d)(2) of these special conditions, due to engine control system failures or malfunctions.

(g) System safety assessment. The applicant must perform a system safety assessment. This assessment must identify faults or failures that affect normal operation, together with the predicted frequency of occurrence of these faults or failures. The intended aircraft application must be taken into account to assure that the assessment of the engine control system safety is valid. The rates of hazardous and major faults must be documented and provided to the installer as part of the requirements in § 33.5.

(h) Protection systems. The engine control devices and systems' design and function, together with engine instruments, operating instructions, and maintenance instructions, must ensure that engine operating limits that can lead to a hazard will not be exceeded in service.

(i) Aircraft supplied data. Any single failure leading to loss, interruption, or corruption of aircraft-supplied data (other than power-command signals from the aircraft), or aircraft-supplied data shared between engine systems within a single engine or between fully independent engine systems, must—

(1) Not result in a hazardous engine effect, as defined in special condition no. 17(d)(2) of these special conditions, for any engine installed on the aircraft; and

(2) Be able to be detected and accommodated by the control system.

(j) Engine control system electrical power.

(1) The engine control system must be designed such that the loss, malfunction, or interruption of the control system electrical power source will not result in a hazardous engine effect, unacceptable transmission of erroneous data, or continued engine operation in the absence of the control function. Hazardous engine effects are defined in special condition no. 17(d)(2) of these special conditions. The engine control system must be capable of resuming normal operation when aircraft-supplied power returns to within the declared limits.

(2) The applicant must identify, document, and provide to the installer as part of the requirements in § 33.5, the characteristics of any electrical power supplied from the aircraft to the engine control system, including transient and steady-state voltage limits, and any other characteristics necessary for safe operation of the engine.

(11) Instrument Connection

The applicant must comply with § 33.29(a), (e), and (g).

(a) In addition, as part of the system safety assessment of special condition nos. 10(g) and 33(h) of these special conditions, the applicant must assess the possibility and subsequent effect of incorrect fit of instruments, sensors, or connectors. Where practicable, the applicant must take design precautions to prevent incorrect configuration of the system.

(b) The applicant must provide instrumentation enabling the flight crew to monitor the functioning of the engine cooling system unless evidence shows that:

(1) Other existing instrumentation provides adequate warning of failure or impending failure;

(2) Failure of the cooling system would not lead to hazardous engine effects before detection; or

(3) The probability of failure of the cooling system is extremely remote.

(12) Stress Analysis

(a) A mechanical and thermal stress analysis, as well as an analysis of the stress caused by electromagnetic forces, must show a sufficient design margin to prevent unacceptable operating characteristics and hazardous engine effects as defined in special condition no. 17(d)(2) of these special conditions.

(b) Maximum stresses in the engine must be determined by test, validated analysis, or a combination thereof, and must be shown not to exceed minimum material properties.

(13) Critical and Life-Limited Parts

(a) The applicant must show, by a safety analysis or means acceptable to the Administrator, whether rotating or moving components, bearings, shafts, static parts, and non-redundant mount components should be classified, designed, manufactured, and managed throughout their service life as critical or life-limited parts.

(1) Critical part means a part that must meet prescribed integrity specifications to avoid its primary failure, which is likely to result in a hazardous engine effect as defined in special condition no. 17(d)(2) of these special conditions.

(2) Life-limited parts may include but are not limited to a rotor or major structural static part, the failure of which can result in a hazardous engine effect, as defined in special condition no. 17(d)(2) of these special conditions, due to a low-cycle fatigue (LCF) mechanism. A life limit is an operational limitation that specifies the maximum allowable number of flight cycles that a part can endure before the applicant must remove it from the engine.

(b) In establishing the integrity of each critical part or life-limited part, the applicant must provide to the Administrator the following three plans for approval:

(1) an engineering plan, as defined in § 33.70(a);

(2) a manufacturing plan, as defined in § 33.70(b); and

(3) a service-management plan, as defined in § 33.70(c).

(14) Lubrication System

(a) The lubrication system must be designed and constructed to function properly between scheduled maintenance intervals in all flight attitudes and atmospheric conditions in which the engine is expected to operate.

(b) The lubrication system must be designed to prevent contamination of the engine bearings and lubrication system components.

(c) The applicant must demonstrate by test, validated analysis, or a combination thereof, the unique lubrication attributes and functional capability of (a) and (b).

(15) Power Response

(a) The design and construction of the engine, including its control system, must enable an increase—

(1) From the minimum power setting to the highest rated power without detrimental engine effects;

(2) From the minimum obtainable power while in-flight and while on the ground to the highest rated power within a time interval determined to be appropriate for the intended aircraft application; and

(3) From the minimum torque to the highest rated torque without detrimental engine effects in the intended aircraft application.

(b) The results of (a)(1), (a)(2), and (a)(3) of this special condition must be ( print page 105444) documented and provided to the installer as part of the requirements in § 33.5.

(16) Continued Rotation

If the design allows any of the engine main rotating systems to continue to rotate after the engine is shut down while in-flight, this continued rotation must not result in any hazardous engine effects, as defined in special condition no. 17(d)(2) of these special conditions.

(17) Safety Analysis

(a) The applicant must comply with § 33.75(a)(1) and (a)(2) using the failure definitions in special condition no. 17(d) of these special conditions.

(b) The primary failure of certain single elements cannot be sensibly estimated in numerical terms. If the failure of such elements is likely to result in hazardous engine effects, then compliance may be shown by reliance on the prescribed integrity requirements of § 33.15 and special condition nos. 9 and 13 of these special conditions, as applicable. These instances must be stated in the safety analysis.

(c) The applicant must comply with § 33.75(d) and (e) using the failure definitions in special condition no. 17(d) of these special conditions, and the ICA in § 33.4.

(d) Unless otherwise approved by the Administrator, the following definitions apply to the engine effects when showing compliance with this condition:

(1) A minor engine effect does not prohibit the engine from performing its intended functions in a manner consistent with § 33.28(b)(1)(i), (b)(1)(iii), and (b)(1)(iv), and the engine complies with the operability requirements of special condition no. 15 and special condition no. 25 of these special conditions, as appropriate.

(2) The engine effects in § 33.75(g)(2) are hazardous engine effects with the addition of:

(i) Electrocution of the crew, passengers, operators, maintainers, or others; and

(ii) Blockage of cooling systems that could cause the engine effects described in § 33.75(g)(2) and special condition 17(d)(2)(i) of these special conditions.

(3) Any other engine effect is a major engine effect.

(e) The intended aircraft application must be taken into account when performing the safety analysis.

(f) The results of the safety analysis, and the assumptions about the aircraft application used in the safety analysis, must be documented and provided to the installer as part of the requirements in § 33.5.

(18) Ingestion

(a) Rain, ice, and hail ingestion must not result in an abnormal operation such as shutdown, power loss, erratic operation, or power oscillations throughout the engine operating range.

(b) Ingestion from other likely sources (birds, induction system ice, foreign objects—ice slabs) must not result in hazardous engine effects defined by special condition no. 17(d)(2) of these special conditions, or unacceptable power loss.

(c) If the design of the engine relies on features, attachments, or systems that the installer may supply, for the prevention of unacceptable power loss or hazardous engine effects, as defined in special condition no. 17(d)(2) of these special conditions, following potential ingestion, then the features, attachments, or systems must be documented and provided to the installer as part of the requirements in § 33.5.

(19) Liquid and Gas Systems

(a) Each system used for lubrication or cooling of engine components must be designed and constructed to function properly in all flight attitudes and atmospheric conditions in which the engine is expected to operate.

(b) If a system used for lubrication or cooling of engine components is not self-contained, the interfaces to that system must be defined, documented, and provided to the installer as part of the requirements in § 33.5.

(c) The applicant must establish by test, validated analysis, or a combination of both that all static parts subject to significant pressure loads will not:

(1) Exhibit permanent distortion beyond serviceable limits, or exhibit leakage that could create a hazardous condition when subjected to normal and maximum working pressure with margin;

(2) Exhibit fracture or burst when subjected to the greater of maximum possible pressures with margin.

(d) Compliance with special condition no. 19(c) of these special conditions must take into account:

(1) The operating temperature of the part;

(2) Any other significant static loads in addition to pressure loads;

(3) Minimum properties representative of both the material and the processes used in the construction of the part; and

(4) Any adverse physical geometry conditions allowed by the type design, such as minimum material and minimum radii.

(e) Approved coolants and lubricants must be documented and provided to the installer as part of the requirements in § 33.5.

(20) Vibration Demonstration

(a) The engine must be designed and constructed to function throughout its normal operating range of rotor speeds and engine output power, including defined exceedances, without inducing excessive stress in any of the engine parts because of vibration and without imparting excessive vibration forces to the aircraft structure.

(b) Each engine design must undergo a vibration survey to establish that the vibration characteristics of those components subject to induced vibration are acceptable throughout the declared flight envelope and engine operating range for the specific installation configuration. The possible sources of the induced vibration that the survey must assess are mechanical, aerodynamic, acoustical, internally induced electromagnetic, installation induced effects that can affect the engine vibration characteristics, and likely environmental effects. This survey must be shown by test, validated analysis, or a combination thereof.

(21) Overtorque

When approval is sought for a transient maximum engine overtorque, the applicant must demonstrate by test, validated analysis, or a combination thereof, that the engine can continue operation after operating at the maximum engine overtorque condition without maintenance action. Upon conclusion of overtorque tests conducted to show compliance with this special condition, or any other tests that are conducted in combination with the overtorque test, each engine part or individual groups of components must meet the requirements of special condition no. 29 of these special conditions.

(22) Calibration Assurance

Each engine must be subjected to calibration tests to establish its power characteristics, and the conditions both before and after the endurance and durability demonstrations specified in special conditions nos. 23 and 26 of these special conditions.

(23) Endurance Demonstration

The applicant must subject the engine to an endurance demonstration, acceptable to the Administrator, to demonstrate the engine's limit capabilities. The endurance demonstration must include increases and decreases of the engine's power ( print page 105445) settings, energy regeneration, and dwellings at the power settings and energy regeneration for sufficient durations that produce the extreme physical conditions the engine experiences at rated performance levels, operational limits, and at any other conditions or power settings, including energy regeneration, which are required to verify the limit capabilities of the engine.

(24) Temperature Limit

The engine design must demonstrate its capability to endure operation at its temperature limits plus an acceptable margin. The applicant must quantify and justify the margin to the Administrator. The demonstration must be repeated for all declared duty cycles and ratings, and operating environments, which would impact temperature limits.

(25) Operation Demonstration

The engine design must demonstrate safe operating characteristics, including but not limited to power cycling, starting, acceleration, and overspeeding throughout its declared flight envelope and operating range. The declared engine operational characteristics must account for installation loads and effects.

(26) Durability Demonstration

The engine must be subjected to a durability demonstration to show that each part of the engine has been designed and constructed to minimize any unsafe condition of the system between overhaul periods, or between engine replacement intervals if the overhaul is not defined. This test must simulate the conditions in which the engine is expected to operate in service, including typical start-stop cycles, to establish when the initial maintenance is required.

(27) System and Component Tests

The applicant must show that systems and components that cannot be adequately substantiated in accordance with the endurance demonstration or other demonstrations will perform their intended functions in all declared environmental and operating conditions.

(28) Rotor Locking Demonstration

If shaft rotation is prevented by locking the rotor(s), the engine must demonstrate:

(a) Reliable rotor locking performance;

(b) Reliable rotor unlocking performance; and

(c) That no hazardous engine effects, as specified in special condition no. 17(d)(2) of these special conditions, will occur.

(29) Teardown Inspection

(a) Teardown evaluation.

(1) After the endurance and durability demonstrations have been completed, the engine must be completely disassembled. Each engine component and lubricant must be eligible for continued operation in accordance with the information submitted for showing compliance with § 33.4.

(2) Each engine component, having an adjustment setting and a functioning characteristic that can be established independent of installation on or in the engine, must retain each setting and functioning characteristic within the established and recorded limits at the beginning of the endurance and durability demonstrations.

(b) Non-Teardown evaluation. If a teardown cannot be performed for all engine components in a non-destructive manner, then the inspection or replacement intervals for these components and lubricants must be established based on the endurance and durability demonstrations and must be documented in the ICA in accordance with § 33.4.

(30) Containment

The engine must be designed and constructed to protect against likely hazards from rotating components as follows—

(a) The design of the stator case surrounding rotating components must provide for the containment of the rotating components in the event of failure, unless the applicant shows that the margin to rotor burst precludes the possibility of a rotor burst.

(b) If the margin to burst shows that the stator case must have containment features in the event of failure, then the stator case must provide for the containment of the failed rotating components. The applicant must define by test, validated analysis, or a combination thereof, and document and provide to the installer as part of the requirements in § 33.5, the energy level, trajectory, and size of fragments released from damage caused by the main-rotor failure, and that pass forward or aft of the surrounding stator case.

(31) [RESERVED]

(32) General Conduct of Tests

(a) Maintenance of the engine may be made during the tests in accordance with the service and maintenance instructions submitted in compliance with § 33.4.

(b) The applicant must subject the engine or its parts to any additional tests that the Administrator finds necessary if—

(1) The frequency of engine service is excessive;

(2) The number of stops due to engine malfunction is excessive;

(3) Major engine repairs are needed; or

(4) Replacement of an engine part is found necessary during the tests, or due to the teardown inspection findings.

(c) Upon completion of all demonstrations and testing specified in these special conditions, the engine and its components must be—

(1) Within serviceable limits;

(2) Safe for continued operation; and

(3) Capable of operating at declared ratings while remaining within limits.

(33) Engine Electrical Systems

(a) Applicability. Any system or device that provides, uses, conditions, or distributes electrical power, and is part of the engine type design, must provide for the continued airworthiness of the engine, and must maintain electric engine ratings.

(b) Electrical systems. The electrical system must ensure the safe generation and transmission of power, and electrical load shedding if load shedding is required, and that the engine does not experience any unacceptable operating characteristics or exceed its operating limits.

(c) Electrical power distribution.

(1) The engine electrical power distribution system must be designed to provide the safe transfer of electrical energy throughout the powerplant. The system must be designed to provide electrical power so that the loss, malfunction, or interruption of the electrical power source will not result in a hazardous engine effect, as defined in special condition no. 17(d)(2) of these special conditions.

(2) The system must be designed and maintained to withstand normal and abnormal conditions during all ground and flight operations.

(3) The system must provide mechanical or automatic means of isolating a faulted electrical energy generation or storage device from leading to hazardous engine effects, as defined in special condition no. 17(d)(2) of these special conditions, or detrimental effects in the intended aircraft application.

(d) Protection systems. The engine electrical system must be designed such that the loss, malfunction, interruption of the electrical power source, or power conditions that exceed design limits, will not result in a hazardous engine effect, as defined in special condition no. 17(d)(2) of these special conditions. ( print page 105446)

(e) Electrical power characteristics. The applicant must identify, declare, document, and provide to the installer as part of the requirements in § 33.5, the characteristics of any electrical power supplied from—

(1) the aircraft to the engine electrical system, for starting and operating the engine, including transient and steady-state voltage limits, and

(2) the engine to the aircraft via energy regeneration, and any other characteristics necessary for safe operation of the engine.

(f) Environmental limits. Environmental limits that cannot adequately be substantiated by endurance demonstration, validated analysis, or a combination thereof must be demonstrated by the system and component tests in special condition no. 27 of these special conditions.

(g) Electrical system failures. The engine electrical system must—

(1) Have a maximum rate of LOPC that is suitable for the intended aircraft application;

(2) When in the full-up configuration, be single-fault tolerant, as determined by the Administrator, for electrical, electrically detectable, and electronic failures involving LOPC events;

(3) Not have any single failure that results in hazardous engine effects; and

(4) Ensure failures or malfunctions that lead to local events in the intended aircraft application do not result in hazardous engine effects, as defined in special condition no. 17(d)(2) of these special conditions, due to electrical system failures or malfunctions.

(h) System safety assessment. The applicant must perform a system safety assessment. This assessment must identify faults or failures that affect normal operation, together with the predicted frequency of occurrence of these faults or failures. The intended aircraft application must be taken into account to assure the assessment of the engine system safety is valid. The rates of hazardous and major faults must be declared, documented, and provided to the installer as part of the requirements in § 33.5.

Issued in Kansas City, Missouri, on December 19, 2024.

Patrick R. Mullen,

Manager, Technical Policy Branch, Policy and Standards Division, Aircraft Certification Service.

Footnotes

1.   https://www.astm.org/​Standards/​F3338.html.

2.  Sometimes the entire system is referred to as an inverter. Throughout this document, it is referred to as the controller.

3.   https://my.rtca.org/​NC_​_​Product?​id=​a1B36000001IcjTEAS.

4.   https://my.rtca.org/​NC_​_​Product?​id=​a1B36000001IcnSEAS.

5.   https://www.faa.gov/​documentLibrary/​media/​Advisory_​Circular/​AC_​33_​28-3.pdf.