105432 Federal Register / Vol. 89, No. 248 / Friday, December 27, 2024 / Rules and Regulations 

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

DEPARTMENT OF TRANSPORTATION 

Federal Aviation Administration 

14 CFR Part 33 

[Docket No. FAA–2023–0587; Special 

Conditions No. 33–23–01–SC] 

Special Conditions: Safran Electric & 

Power S.A. Model ENGINeUS 100A1 

Electric Engines 

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 (781) 

238–7755; mark.bouyer@faa.gov. 

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 directdrive, 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 


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2Sometimes the entire system is referred to as an 

inverter. Throughout this document, it is referred to 

as the controller. 

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 highvoltage electrical system is a 

combination of wires and connectors 

that integrate the motor and controller. 

In addition, the technology 

comprising these high-voltage and highcurrent electronic components 

introduces potential hazards that do not 

exist in turbine and reciprocating 

aircraft engines. For example, highvoltage 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 


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3https://my.rtca.org/NC__

Product?id=a1B36000001IcjTEAS. 

4https://my.rtca.org/NC__

Product?id=a1B36000001IcnSEAS. 

5https://www.faa.gov/documentLibrary/media/ 

Advisory_Circular/AC_33_28-3.pdf. 

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 aircraftsupplied 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 


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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 lowcycle 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 manufacturinginduced 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 


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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 lifelimited or critical parts. Some lifelimited 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 oncondition 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 singleengine 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 selfcontained, 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 


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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 sealevel, 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. 


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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 inflight. 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 nondestructive 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 


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codified, the only electrical systems 

used on turbine engines were lowvoltage, electronic engine control 

systems (EEC) and high-energy sparkignition 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 shutoff 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 


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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 distributedpropulsion, 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 performancebased 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 aircraftlevel 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 


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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 typecertificated 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 componentlevel 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 enginelevel 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. 


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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— 


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(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 


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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 


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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. 


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105446 Federal Register / Vol. 89, No. 248 / Friday, December 27, 2024 / Rules and Regulations 

(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 steadystate 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. 

[FR Doc. 2024–30855 Filed 12–26–24; 8:45 am] 

BILLING CODE 4910–13–P 

DEPARTMENT OF TRANSPORTATION 

Federal Aviation Administration 

14 CFR Part 61 

[Docket No. FAA–2024–2580; Amdt. No. 61– 

158A] 

RIN 2120–AM06 

Regulatory Updates to BasicMed; 

Correction 

AGENCY: Federal Aviation 

Administration (FAA), U.S. Department 

of Transportation (DOT). 

ACTION: Correcting amendment. 

SUMMARY: On November 18, 2024, the 

Federal Aviation Administration (FAA) 

published a final rule titled ‘‘Regulatory 

Updates to BasicMed.’’ That document 

implemented the provisions of sections 

815 and 828 of the FAA Reauthorization 

Act of 2024. The final rule’s revision to 

the part 61 authority citation that cites 

the 2024 FAA Reauthorization Act was 

erroneously overwritten by another 

rulemaking action revising part 61. This 

document corrects the amendatory 

instruction and revises the part 61 

authority citation to be inclusive of the 

intended updates promulgated in the 

BasicMed final rule. 

DATES: This correction is effective 

December 27, 2024. 

FOR FURTHER INFORMATION CONTACT: 

Bradley C. Zeigler, Federal Aviation 

Administration, 800 Independence 

Avenue SW, Washington, DC 20591; 

telephone: (202) 267–1100; email: 

Bradley.C.Zeigler@faa.gov. 

SUPPLEMENTARY INFORMATION: On 

October 1, 2024, the FAA published the 

final rule Removal of Expiration Date on 

a Flight Instructor Certificate; 

Additional Qualification Requirements 

To Train Initial Flight Instructor 

Applicants; and Other Provisions in the 

Federal Register (89 FR 80020) 

(hereinafter the CFI rule). That final rule 

removed the expiration date on flight 

instructor certificates to align with other 

airman certificates. Additional 

amendments included updating renewal 

requirements to recent experience 

requirements, introducing a new 

method for establishing recent flight 

instructor experience, and allowing 

instructors with a lapse of no more than 

three months to reinstate privileges via 

an approved flight instructor refresher 

course instead of a practical test. The 

final rule also added two new methods 

for flight instructors to qualify to train 

initial applicants. 

As part of the amendments made in 

the CFI rule, the FAA revised the 

authority citation for part 61 by 

removing 49 U.S.C. 106(g) as that 

paragraph had been removed and 

reserved as part of the FAA 

Reauthorization Act of 2024. The final 

rule added section 820 of Public Law 

118–63 to the authority citation. Section 

820 of the FAA Reauthorization Act of 

2024 required the FAA to promulgate 

the CFI final rule within 18 months after 

the date of enactment of the Act, to at 

a minimum, update part 61, Code of 

Federal Regulations, to remove the 

expiration date on the flight instructor 

certificate and replace the requirement 

that a flight instructor renews their 

flight instructor certificate with 

appropriate recent experience 

requirements. The CFI rule became 

effective on December 1, 2024. 

Subsequently, on November 18, 2024, 

the FAA published the final rule 

Regulatory Updates to BasicMed (89 FR 

90572) (hereinafter the BasicMed rule). 

That final rule implemented the 

provisions of sections 815 and 828 of 

the FAA Reauthorization Act of 2024. In 

the BasicMed rule, the FAA revised the 

authority citation for part 61 by adding 

sections 815 and 828 of Public Law 

118–63 to the authority citation for part 

61. The BasicMed final rule became 

effective on the date of publication, 

November 18, 2024. 

However, when the CFI rule became 

effective on December 1, 2024, and the 

authority citation for part 61 was 

revised in accordance with the 

instructions contained in the CFI rule, 

those revisions inadvertently removed 

the revisions made in the BasicMed 

final rule which published after, but 

became effective before, the CFI rule. 

Therefore, this correction revises the 

authority citation for part 61 to include 

both the revisions from the CFI rule and 

the BasicMed rule. 

List of Subjects in 14 CFR Part 61 

Aircraft, Airmen, Aviation safety, 

Reporting and recordkeeping 

requirements. 

PART 61—CERTIFICATION: PILOTS, 

FLIGHT INSTRUCTORS, AND GROUND 

INSTRUCTORS 

■ 1. The authority citation for part 61 is 

revised to read as follows: 

Authority: 49 U.S.C. 106(f), 40113, 44701– 

44703, 44707, 44709–44711, 44729, 44903, 

45102–45103, 45301–45302; Sec. 2307, Pub. 

L. 114–190, 130 Stat. 615 (49 U.S.C. 44703 

note); sec. 318, Pub. L. 115–254, 132 Stat. 

3186 (49 U.S.C. 44703 note); sec. 820, Pub. 

L. 118–63, 138 Stat. 1330 (49 U.S.C. 44939 

note); secs. 815 and 828, Pub. L. 118–63, 138 

Stat. 1328, 1336 (49 U.S.C. 44703 note). 


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