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[Federal Register Volume 89, Number 248 (Friday, December 27, 2024)]
[Rules and Regulations]
[Pages 105432-105446]
From the Federal Register Online via the Government Publishing Office [www.gpo.gov]
[FR Doc No: 2024-30855]

[[Page 105432]]

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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.

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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; 
[email protected].

SUPPLEMENTARY INFORMATION:

Background

    On November 27, 2020, Safran applied for FAA validation for a type 
certificate for their Model ENGINeUS 100A1 electric engine. The Safran 
Model ENGINeUS 100A1 electric engine will be used in a single-engine 
airplane that will be certified separately from the engine.
    The Safran Model ENGINeUS 100A1 electric engine is comprised of a 
direct-drive, radial-flux, permanent magnet motor, divided in two 
sections, each section having a three-phase motor, and one electric 
power inverter controlling each three-phase motor.

Type Certification Basis

    Under the provisions of 14 CFR 21.17(a)(1), generally, Safran must 
show that Model ENGINeUS 100A1 electric engines meet the applicable 
provisions of 14 CFR part 33 in effect on the date of application for a 
type certificate.
    If the Administrator finds that the applicable airworthiness 
regulations (e.g., part 33) do not contain adequate or appropriate 
safety standards for the Safran Model ENGINeUS 100A1 electric engines 
because of a novel or unusual design feature, special conditions may be 
prescribed under the provisions of Sec.  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 Sec.  21.101.
    The FAA issues special conditions, as defined in Sec.  11.19, in 
accordance with Sec.  11.38, and they become part of the type 
certification basis under Sec.  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.
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    \1\ https://www.astm.org/Standards/F3338.html.
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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

[[Page 105433]]

Safran aircraft engine is subject to operating conditions produced by 
chemical, thermal, and mechanical components working together, but the 
operating conditions are unlike those observed in internal combustion 
engine systems. Therefore, part 33 does not contain adequate or 
appropriate safety standards for the Safran Model ENGINeUS 100A1 
electric engine's novel or unusual design feature.
    Safran's aircraft engines will operate using electrical power 
instead of air and fuel combustion to propel the aircraft. These 
electric engines will be designed, manufactured, and controlled 
differently than turbine or reciprocating aircraft engines. They will 
be built with an electric motor, motor controller, and high-voltage 
electrical systems that draw energy from electrical storage or 
electrical energy generating systems. The electric motor is a device 
that converts electrical energy into mechanical energy by electric 
current flowing through windings (wire coils) in the motor, producing a 
magnetic field that interacts with permanent magnets mounted on the 
engine's main rotor. The controller is a system that consists of two 
main functional elements: the motor controller and an electric power 
inverter to drive the motor.\2\ The high-voltage electrical system is a 
combination of wires and connectors that integrate the motor and 
controller.
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    \2\ Sometimes the entire system is referred to as an inverter. 
Throughout this document, it is referred to as the controller.
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    In addition, the technology comprising these high-voltage and high-
current electronic components introduces potential hazards that do not 
exist in turbine and reciprocating aircraft engines. For example, high-
voltage transmission lines, electromagnetic shields, magnetic 
materials, and high-speed electrical switches are necessary to use the 
physical properties of an electric engine for propelling an aircraft. 
However, this technology also exposes the aircraft to potential 
failures that are not common to gas-powered turbine and reciprocating 
engines, technological differences which could adversely affect safety 
if not addressed through these special conditions.

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

    Although Safran's electric aircraft engines use novel or unusual 
design features that the FAA did not envisage during the development of 
its existing part 33 airworthiness standards, these engines share some 
basic similarities, in configuration and function, to engines that use 
the combustion of air and fuel, and therefore require similar 
provisions to prevent common hazards (e.g., fire, uncontained high 
energy debris, and loss of thrust control). However, the primary 
failure concerns and the probability of exposure to these common 
hazards are different for the Safran Model ENGINeUS 100A1 electric 
engine. This creates a need to develop special conditions to ensure the 
engine's safety and reliability.
    The requirements in part 33 ensure that the design and construction 
of aircraft engines, including the engine control systems, are proper 
for the type of aircraft engines considered for certification. However, 
part 33 does not fully address aircraft engines like the Safran Model 
ENGINeUS 100A1 electric engine, which operates using electrical 
technology as the primary means of propelling the aircraft. This 
necessitates the development of special conditions that provide 
adequate airworthiness standards for these aircraft engines.
    The requirements in part 33, subpart B, are applicable to 
reciprocating and turbine aircraft engines. Subparts C and D are 
applicable to reciprocating aircraft engines. Subparts E through G are 
applicable to turbine aircraft engines. As such, subparts B through G 
do not adequately address the use of aircraft engines that operate 
using electrical technology. Special conditions are needed to ensure a 
level of safety for electric engines that is commensurate with these 
subparts, as those regulatory requirements do not contain adequate or 
appropriate safety standards for electric aircraft engines that are 
used to propel aircraft.

FAA Special Conditions for the Safran Engine Design

    Applicability: Special condition no. 1 requires Safran to comply 
with part 33, except for those airworthiness standards specifically and 
explicitly applicable only to reciprocating and turbine aircraft 
engines.
    Engine Ratings and Operating Limitations: Special condition no. 2, 
in addition to compliance with Sec.  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 
Sec.  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 Sec.  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 Sec.  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 
Sec.  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 Sec.  33.23, which 
requires the applicant to define, and the design to withstand, certain 
load limits for the engine mounting attachments and

[[Page 105434]]

related engine structure. These requirements would otherwise be 
applicable only to reciprocating and turbine aircraft engines.
    Accessory Attachments: Special condition no. 8 requires the design 
to comply with Sec.  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 Sec.  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 Sec.  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.
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    \3\ https://my.rtca.org/NC__Product?id=a1B36000001IcjTEAS.
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    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.
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    \4\ https://my.rtca.org/NC__Product?id=a1B36000001IcnSEAS.
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    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 Sec.  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.
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    \5\ https://www.faa.gov/documentLibrary/media/Advisory_Circular/AC_33_28-3.pdf.
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    The phrase ``in the full-up configuration'' used in special 
condition no. 10(f)(2) refers to a system without any fault conditions 
present. The electronic control system must, when in the full-up 
configuration, be single-fault tolerant, as determined by the 
Administrator, for electrical, electrically detectable, and electronic 
failures involving LOPC events.
    The term ``local'' in the context of ``local events'' used in 
special condition no. 10(f)(4) means failures or malfunctions leading 
to events in the intended aircraft installation such as fire, overheat, 
or failures leading to damage to engine control system components. 
These ``local events'' must not result in a hazardous engine effect due 
to engine control system failures or malfunctions.
    Special condition no. 10(g) requires Safran to conduct a safety 
assessment of the control system to support the safety analysis in 
special condition no. 17. This control system safety assessment 
provides engine response to failures, and rates of these failures that 
can be used at the aircraft-level safety assessment.
    Special condition no. 10(h) requires Safran to provide appropriate 
protection devices or systems to ensure that engine operating limits 
will not be exceeded in service.
    Special condition no. 10(i) is necessary to ensure that the 
controllers are self-sufficient and isolated from other aircraft 
systems. The aircraft-supplied data supports the analysis at the 
aircraft level to protect the aircraft from common mode failures that 
could lead to major propulsion power loss. The exception ``other than 
power command signals from the aircraft,'' noted in special condition 
no. 10(i), is based on the FAA's determination that the engine 
controller has no reasonable means to determine the validity of any in-
range signals from the electrical power system. In many cases, the 
engine control system can detect a faulty signal from the aircraft, but 
the engine control

[[Page 105435]]

system typically accepts the power command signal as a valid value.
    The term ``independent'' in the context of ``fully independent 
engine systems'' referenced in special condition no. 10(i) means the 
controllers should be self-sufficient and isolated from other aircraft 
systems or provide redundancy that enables the engine control system to 
accommodate aircraft data system failures. In the case of loss, 
interruption, or corruption of aircraft-supplied data, the engine must 
continue to function in a safe and acceptable manner without hazardous 
engine effects.
    The term ``accommodated,'' in the context of ``detected and 
accommodated,'' referenced in special condition 10(i)(2) is to assure 
that, upon detecting a fault, the system continues to function safely.
    Special condition no. 10(j) requires Safran to show that the loss 
of electric power from the aircraft will not cause the electric engine 
to malfunction in a manner hazardous to the aircraft. The total loss of 
electric power to the electric engine may result in an engine shutdown.
    Instrument Connection: Special condition no. 11 requires Safran to 
comply with Sec.  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 Sec.  33.29 
apply only to technologies used in reciprocating and turbine aircraft 
engines.
    Instrument connections (wires, wire insulation, potting, grounding, 
connector designs, etc.) must not introduce unsafe features or 
characteristics to the aircraft. Special condition no. 11 requires the 
safety analysis to include potential hazardous effects from failures of 
instrument connections to function properly. The outcome of this 
analysis might identify the need for design enhancements or additional 
ICA to ensure safety.
    Stress Analysis: Section 33.62 requires applicants to perform a 
stress analysis on each turbine engine. This regulation is explicitly 
applicable only to turbine engines and turbine engine components, and 
it is not appropriate for the Safran Model ENGINeUS 100A1 electric 
engines. However, a stress analysis particular to these electric 
engines is necessary to account for stresses resulting from electric 
technology used in the engine.
    Special condition no. 12 requires a mechanical, thermal, and 
electrical stress analysis to show that the engine has a sufficient 
design margin to prevent unacceptable operating characteristics. Also, 
the applicant must determine the maximum stresses in the engine by 
tests, validated analysis, or a combination thereof, and show that they 
do not exceed minimum material properties.
    Critical and Life-Limited Parts: Special condition no. 13 requires 
Safran to show whether rotating or moving components, bearings, shafts, 
static parts, and non-redundant mount components should be classified, 
designed, manufactured, and managed throughout their service life as 
critical or life-limited parts.
    The term ``low-cycle fatigue,'' referenced in special condition no. 
13(a)(2), is a decline in material strength from exposure to cyclic 
stress at levels beyond the stress threshold the material can sustain 
indefinitely. This threshold is known as the ``material endurance 
limit.'' Low-cycle fatigue typically causes a part to sustain plastic 
or permanent deformation during the cyclic loading and can lead to 
cracks, crack growth, and fracture. Engine parts that operate at high 
temperatures and high mechanical stresses simultaneously can experience 
low-cycle fatigue coupled with creep. Creep is the tendency of a 
metallic material to permanently move or deform when it is exposed to 
the extreme thermal conditions created by hot combustion gasses, and 
substantial physical loads such as high rotational speeds and maximum 
thrust. Conversely, high-cycle fatigue is caused by elastic 
deformation, small strains caused by alternating stress, and a much 
higher number of load cycles compared to the number of cycles that 
cause low-cycle fatigue.
    The engineering plan referenced in special condition no. 13(b)(1) 
informs the manufacturing and service management processes of essential 
information that ensures the life limit of a part is valid. The 
engineering plan provides methods for verifying the characteristics and 
qualities assumed in the design data using methods that are suitable 
for the part criticality. The engineering plan informs the 
manufacturing process of the attributes that affect the life of the 
part. The engineering plan, manufacturing plan, and service management 
plan are related in that assumptions made in the engineering plan are 
linked to how a part is manufactured and how that part is maintained in 
service. For example, environmental effects on life limited electric 
engine parts, such as humidity, might not be consistent with the 
assumptions used to design the part. Safran must ensure that the 
engineering plan is complete, available, and acceptable to the 
Administrator.
    The term ``manufacturing plan,'' referenced in special condition 
no. 13(b)(2), is the collection of data required to translate 
documented engineering design criteria into physical parts, and to 
verify that the parts comply with the properties established by the 
design data. Because engines are not intentionally tested to failure 
during a certification program, documents and processes used to execute 
production and quality systems required by Sec.  21.137 guarantee 
inherent expectations for performance and durability. These systems 
limit the potential manufacturing outcomes to parts that are 
consistently produced within design constraints.
    The manufacturing plan and service management plan ensure that 
essential information from the engineering plan, such as the design 
characteristics that safeguard the integrity of critical and life-
limited parts, is consistently produced and preserved over the lifetime 
of those parts. The manufacturing plan includes special processes and 
production controls to prevent inclusion of manufacturing-induced 
anomalies, which can degrade the part's structural integrity. Examples 
of manufacturing-induced anomalies are material contamination, 
unacceptable grain growth, heat-affected areas, and residual stresses.
    The service-management plan ensures the method and assumptions used 
in the engineering plan to determine the part's life remain valid by 
enabling corrections identified from in-service experience, such as 
service-induced anomalies and unforeseen environmental effects, to be 
incorporated into the design process. The service-management plan also 
becomes the ICA for maintenance, overhaul, and repairs of the part.
    Lubrication System: Special condition no. 14 requires Safran to 
ensure that the lubrication system is designed to function properly 
between scheduled maintenance intervals and to prevent contamination of 
the engine bearings. This special condition also requires Safran to 
demonstrate the unique lubrication attributes and functional capability 
of the Safran Model ENGINeUS 100A1 electric engine design.
    The corresponding part 33 regulations include provisions for 
lubrication systems used in reciprocating and turbine engines. The part 
33 requirements account for safety issues associated with specific 
reciprocating and turbine engine system configurations. These 
regulations are not appropriate for the Safran Model ENGINeUS 100A1 
electric engines. For example, electric engines do not have a

[[Page 105436]]

crankcase or lubrication oil sump. Electric engine bearings are sealed, 
so they do not require an oil circulation system. The lubrication 
system in these engines is also independent of the propeller pitch 
control system. Therefore, special condition no. 14 incorporates only 
certain requirements from the part 33 regulations.
    Power Response: Special condition no. 15 requires the design and 
construction of the Safran Model ENGINeUS 100A1 electric engines to 
enable an increase from the minimum--
    (1) power setting to the highest rated power without detrimental 
engine effects, and
    (2) within a time interval appropriate for the intended aircraft 
application.
    The engine control system governs the increase or decrease in power 
in combustion engines to prevent too much (or too little) fuel from 
being mixed with air before combustion. Due to the lag in rotor 
response time, improper fuel/air mixtures can result in engine surges, 
stalls, and exceedances above rated limits and durations. Failure of 
the combustion engine to provide thrust, maintain rotor speeds below 
rotor burst thresholds, and keep temperatures below limits can have 
engine effects detrimental to the aircraft. Similar detrimental effects 
are possible in the Safran Model ENGINeUS 100A1 electric engines, but 
the causes are different. Electric engines with reduced power response 
time can experience insufficient thrust to the aircraft, shaft over-
torque, and over-stressed rotating components, propellers, and critical 
propeller parts. Therefore, this special condition is necessary.
    Continued Rotation: Special condition no. 16 requires Safran to 
design the Model ENGINeUS 100A1 electric engines such that, if the main 
rotating systems continue to rotate after the engine is shut down while 
in-flight, this continued rotation will not result in any hazardous 
engine effects.
    The main rotating system of the Safran Model ENGINeUS 100A1 
electric engines consists of the rotors, shafts, magnets, bearings, and 
wire windings that convert electrical energy to shaft torque. For the 
initial aircraft application, this rotating system must continue to 
rotate after the power source to the engine is shut down. The safety 
concerns associated with this special condition are substantial 
asymmetric aerodynamic drag that can cause aircraft instability, loss 
of control, and reduced efficiency; and may result in a forced landing 
or inability to continue safe flight.
    Safety Analysis: Special condition no. 17 requires Safran to comply 
with Sec.  33.75(a)(1) and (a)(2), which require the applicant to 
conduct a safety analysis of the engine, and which would otherwise be 
applicable only to turbine aircraft engines. Additionally, this special 
condition requires Safran to assess its engine design to determine the 
likely consequences of failures that can reasonably be expected to 
occur. The failure of such elements, and associated prescribed 
integrity requirements, must be stated in the safety analysis.
    A primary failure mode is the manner in which a part is most likely 
going to fail. Engine parts that have a primary failure mode, a 
predictable life to the failure, and a failure consequence that results 
in a hazardous effect, are life-limited or critical parts. Some life-
limited or critical engine parts can fail suddenly in their primary 
failure mode, from prolonged exposure to normal engine environments 
such as temperature, vibration, and stress, if those engine parts are 
not removed from service before the damage mechanisms progress to a 
failure. Due to the consequence of failure, these parts are not allowed 
to be managed by on-condition or probabilistic means because the 
probability of failure cannot be sensibly estimated in numerical terms. 
Therefore, the parts are managed by compliance with integrity 
requirements, such as mandatory maintenance (life limits, inspections, 
inspection techniques), to ensure the qualities, features, and other 
attributes that prevent the part from failing in its primary failure 
mode are preserved throughout its service life. For example, if the 
number of engine cycles to failure are predictable and can be 
associated with specific design characteristics, such as material 
properties, then the applicant can manage the engine part with life 
limits.
    Complete or total power loss is not assumed to be a minor engine 
event, as it is in the turbine engine regulation Sec.  33.75, to 
account for experience data showing a potential for higher hazard 
levels from power loss events in single-engine general aviation 
aircraft. The criteria in these special conditions apply to an engine 
that continues to operate at partial power after a single electrical or 
electronic fault or failure. Total loss of power is classified at the 
aircraft level using special condition nos. 10(g) and 33(h).
    Ingestion: Special condition no. 18 requires Safran to ensure that 
these engines will not experience unacceptable power loss or hazardous 
engine effects from ingestion. The associated regulations for turbine 
engines, Sec. Sec.  33.76, 33.77, and 33.78, are based on potential 
performance impacts and damage from birds, ice, rain, and hail being 
ingested into a turbine engine that has an inlet duct, which directs 
air into the engine for combustion, cooling, and thrust. By contrast, 
the Safran Model ENGINeUS 100A1 electric engines are not configured 
with inlet ducts.
    An ``unacceptable'' power loss, as used in special condition no. 
18(b), is such that the power or thrust required for safe flight of the 
aircraft becomes unavailable to the pilot. The specific amount of power 
loss that is required for safe flight depends on the aircraft 
configuration, speed, altitude, attitude, atmospheric conditions, phase 
of flight, and other circumstances where the demand for thrust is 
critical to safe operation of the aircraft.
    Liquid and Gas Systems: Special condition no. 19 requires Safran to 
ensure that systems used for lubrication or cooling of engine 
components are designed and constructed to function properly. Also, if 
a system is not self-contained, the interfaces to that system would be 
required to be defined in the engine installation manual. Systems for 
the lubrication or cooling of engine components can include heat 
exchangers, pumps, fluids, tubing, connectors, electronic devices, 
temperature sensors and pressure switches, fasteners and brackets, 
bypass valves, and metallic chip detectors. These systems allow the 
electric engine to perform at extreme speeds and temperatures for 
durations up to the maintenance intervals without exceeding temperature 
limits or predicted deterioration rates.
    Vibration Demonstration: Special condition no. 20 requires Safran 
to ensure the engine--
    (1) is designed and constructed to function throughout its normal 
operating range of rotor speeds and engine output power without 
inducing excessive stress caused by engine vibration, and
    (2) design undergoes a vibration survey.
    The vibration demonstration is a survey that characterizes the 
vibratory attributes of the engine. It verifies that the stresses from 
vibration do not impose excessive force or result in natural frequency 
responses on the aircraft structure. The vibration demonstration also 
ensures internal vibrations will not cause engine components to fail. 
Excessive vibration force occurs at magnitudes and forcing functions or 
frequencies, which may result in damage to the aircraft. Stress margins 
to failure add conservatism to the highest values predicted by analysis 
for additional protection from failure

[[Page 105437]]

caused by influences beyond those quantified in the analysis. The 
result of the additional design margin is improved engine reliability 
that meets prescribed thresholds based on the failure classification. 
The amount of margin needed to achieve the prescribed reliability rates 
depends on an applicant's experience with a product. The FAA considers 
the reliability rates when deciding how much vibration is 
``excessive.''
    Overtorque: Special condition no. 21 requires Safran to demonstrate 
that the engine is capable of continued operation without the need for 
maintenance if it experiences a certain amount of overtorque.
    Safran's electric engine converts electrical energy to shaft 
torque, which is used for propulsion. The electric motor, controller, 
and high-voltage systems control the engine torque. When the pilot 
commands power or thrust, the engine responds to the command and 
adjusts the shaft torque to meet the demand. During the transition from 
one power or thrust setting to another, a small delay, or latency, 
occurs in the engine response time. While the engine dwells in this 
time interval, it can continue to apply torque until the command to 
change the torque is applied by the engine control. The allowable 
amount of overtorque during operation depends on the engine's response 
to changes in the torque command throughout its operating range.
    Calibration Assurance: Special condition no. 22 requires Safran to 
subject the engine to calibration tests to establish its power 
characteristics and the conditions both before and after the endurance 
and durability demonstrations specified in special condition nos. 23 
and 26. The calibration test requirements specified in Sec.  33.85 only 
apply to the endurance test specified in Sec.  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 Sec.  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 
Sec.  33.8, from failing due to temperature-related causes when 
operating at the most extreme engine and environmental thermal 
conditions.
    Operation Demonstration: Special condition no. 25 requires the 
engine to demonstrate safe operating characteristics throughout its 
declared flight envelope and operating range. Engine operating 
characteristics define the range of functional and performance values 
the Safran Model ENGINeUS 100A1 electric engines can achieve without 
incurring hazardous effects. The characteristics are requisite 
capabilities of the type design that qualify the engine for 
installation into aircraft and that determine aircraft installation 
requirements. The primary engine operating characteristics are assessed 
by the tests and demonstrations that would be required by these special 
conditions. Some of these characteristics are shaft output torque, 
rotor speed, power consumption, and engine thrust response. The engine 
performance data Safran will use to certify the engine must account for 
installation loads and effects. These are aircraft-level effects that 
could affect the engine characteristics that are measured when the 
engine is tested on a stand or in a test cell. These effects could 
result from elevated inlet cowl temperatures, aircraft maneuvers, 
flowstream distortion, and hard landings. For example, an engine that 
is run in a sea-level, static test facility could demonstrate more 
capability for some operating characteristics than it will have when 
operating on an aircraft in certain flight conditions. Discoveries like 
this during certification could affect engine ratings and operating 
limits. Therefore, the installed performance defines the engine 
performance capabilities.
    Durability Demonstration: Special condition no. 26 requires Safran 
to subject the engine to a durability demonstration. The durability 
demonstration must show that the engine is designed and constructed to 
minimize the development of any unsafe condition between maintenance 
intervals or between engine replacement intervals if maintenance or 
overhaul is not defined. The durability demonstration also verifies 
that the ICA is adequate to ensure the engine, in its fully 
deteriorated state, continues to generate rated power or thrust, while 
retaining operating margins and sufficient efficiency, to support the 
aircraft safety objectives. The amount of deterioration an engine can 
experience is restricted by operating limitations and managed by the 
engine ICA. Section 33.90 specifies how maintenance intervals are 
established; it does not include provisions for an engine replacement. 
Electric engines and turbine engines deteriorate differently. 
Therefore, Safran will use different test effects to develop 
maintenance, overhaul, or engine replacement information for their 
electric engine.
    System and Component Tests: Special condition no. 27 requires 
Safran to show that the systems and components of the engine perform 
their intended functions in all declared engine environments and 
operating conditions.
    Sections 33.87 and 33.91, which are specifically applicable to 
turbine engines, have conditional criteria to decide if additional 
tests will be required after the engine tests. The criteria are not 
suitable for electric engines. Part 33 associates the need for 
additional testing with the outcome of the Sec.  33.87 endurance test 
because it is designed to address safety concerns in combustion 
engines. For example, Sec.  33.91(b) requires the establishment of 
temperature limits for components that require temperature-controlling 
provisions, and Sec.  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 Sec.  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.

[[Page 105438]]

Additional safety considerations also need to be addressed in the test. 
Therefore, special condition no. 27 is a performance-based requirement 
that allows Safran to determine when engine systems and component tests 
are necessary and to determine the appropriate limitations of those 
systems and components used in the Safran Model ENGINeUS 100A1 electric 
engine.
    Rotor Locking Demonstration: Special condition no. 28 requires the 
engine to demonstrate reliable rotor locking performance and that no 
hazardous effects will occur if the engine uses a rotor locking device 
to prevent shaft rotation.
    Some engine designs enable the pilot to prevent a propeller shaft 
or main rotor shaft from turning while the engine is running, or the 
aircraft is in-flight. This capability is needed for some installations 
that require the pilot to confirm the functionality of certain flight 
systems before takeoff. The Safran engine installations are not limited 
to aircraft that will not require rotor locking. Section 33.92 
prescribes a test that may not include the appropriate criteria to 
demonstrate sufficient rotor locking capability for these engines. 
Therefore, this special condition is necessary.
    The special condition does not define ``reliable'' rotor locking 
but allows Safran to classify the hazard as major or minor and assign 
the appropriate quantitative criteria that meet the safety objectives 
required by special condition no. 17 and the applicable portions of 
Sec.  33.75.
    Teardown Inspection: Special condition no. 29 requires Safran to 
perform a teardown or non-teardown evaluation after the endurance, 
durability, and overtorque demonstrations, based on the criteria in 
special condition no. 29(a) or (b).
    Special condition no. 29(b) includes restrictive criteria for 
``non-teardown evaluations'' to account for electric engines, sub-
assemblies, and components that cannot be disassembled without 
destroying them. Some electrical and electronic components like 
Safran's are constructed in an integrated fashion that precludes the 
possibility of tearing them down without destroying them. The special 
condition indicates that, if a teardown cannot be performed in a non-
destructive manner, then the inspection or replacement intervals must 
be established based on the endurance and durability demonstrations. 
The procedure for establishing maintenance should be agreed upon 
between the applicant and the FAA prior to running the relevant tests. 
Data from the endurance and durability tests may provide information 
that can be used to determine maintenance intervals and life limits for 
parts. However, if life limits are required, the lifing procedure is 
established by special condition no. 13, Critical and Life-Limited 
Parts, which corresponds to Sec.  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 Sec.  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 Sec.  33.28 was

[[Page 105439]]

codified, the only electrical systems used on turbine engines were low-
voltage, electronic engine control systems (EEC) and high-energy spark-
ignition systems. Electric aircraft engines use high-voltage, high-
current electrical systems and components that are physically located 
in the motor and motor controller. Therefore, the existing part 33 
control system requirements do not adequately address all the 
electrical systems used in electric aircraft engines. Special condition 
no. 33 is established using the existing engine control systems 
requirement as a basis. It applies applicable airworthiness criteria 
from Sec.  33.28 and incorporates airworthiness criteria that recognize 
and focus on the electrical power system used in the engine.
    Special condition no. 33(b) ensures that all aspects of an 
electrical system, including generation, distribution, and usage, do 
not experience any unacceptable operating characteristics.
    Special condition no. 33(c) requires the electrical power 
distribution aspects of the electrical system to provide the safe 
transfer of electrical energy throughout the electric engine.
    The term ``abnormal conditions'' used in special condition no. 
33(c)(2) is intended to be consistent with the definitions in MIL-STD-
704F ``Aircraft Electric Power Characteristics'' which defines normal 
operation and abnormal operation. MIL-STD-704F is a standard that 
ensures compatibility between power sources that provide power to the 
aircraft's electrical systems and airborne equipment that receive power 
from the power source. This standard also establishes technical 
criteria for aircraft electric power. The term ``abnormal conditions'' 
refers to various engine operating conditions such as:
     System or component characteristics outside of normal 
statistical variation from circumstances such as systems degradation, 
installation error, and engine response to fault conditions;
     Unusual environmental conditions from extreme temperature, 
humidity, vibration, lightning, high-intensity radiated field (HIRF), 
atmospheric neutron radiation; and
     Unusual and infrequent events such as landing on icy 
runways, rejected take-offs or go-arounds, extended ground idling or 
taxiing in a hot environment, and abrupt load changes from foreign 
object damage or engine contamination.
    The phrase ``safe transmission of electric energy'' used in special 
condition no. 33(c)(3) refers to the transmission of electrical energy 
in a manner that supports the operation of the electric engine(s) and 
the aircraft safety objectives without detrimental effects such as 
uncontrolled fire or structural failure due to severe overheating.
    Special condition no. 33(d) requires the engine electrical system 
to be designed such that the loss, malfunction, or interruption of the 
electrical power source, or power conditions that exceed design limits, 
will not result in a hazardous engine effect.
    Special condition no. 33(e) requires Safran to identify and 
declare, in the engine installation manual, the characteristics of any 
electrical power supplied from the aircraft to the engine, or 
electrical power supplied from the engine to the aircraft via energy 
regeneration, and any other characteristics necessary for safe 
operation of the engine.
    Special condition no. 33(f) requires Safran to demonstrate that 
systems and components will operate properly up to environmental 
limits, using special conditions, when such limits cannot be adequately 
substantiated by the endurance demonstration, validated analysis, or a 
combination thereof. The environmental limits referred to in this 
special condition include temperature, vibration, HIRF, and others 
addressed in RTCA DO-160G, ``Environmental Conditions and Test 
Procedures for Airborne Electronic/Electrical Equipment and 
Instruments.''
    Special condition 33(g) requires Safran to evaluate various 
electric engine system failures to ensure that these failures will not 
lead to unsafe engine conditions. The evaluation includes single-fault 
tolerance, ensures no single electrical or electronic fault or failure 
would result in hazardous engine effects, and ensures that any failure 
or malfunction leading to local events in the intended aircraft 
application do not result in certain hazardous engine effects. The 
special condition also implements integrity requirements, criteria for 
LOTC/LOPC events, and an acceptable LOTC/LOPC rate.
    Special condition 33(h) requires Safran to conduct a safety 
assessment of the engine electrical system to support the safety 
analysis in special condition no. 17. This safety assessment provides 
engine response to failures, and rates of these failures, which can be 
used at the aircraft safety assessment level.

Discussion of Comments

    The FAA issued a notice of proposed special conditions (NPSC) 
Docket No. FAA-2023-0587 for the Safran Model ENGINeUS 100A1 electric 
engines, which was published in the Federal Register on March 20, 2024 
(89 FR 19763).
    The FAA received responses from four commenters, Airbus Helicopters 
(Airbus), Ampaire Inc. (Ampaire), Kite Magnetics Pty Ltd. (Kite 
Magnetics), and magniX USA, Inc. (magniX).
    The FAA received one comment from Airbus that stated proposed 
special condition no. 4, Fire Protection, does not prescribe safety 
criteria for flammable cooling fluids and suggested that a fireproof 
wall, cooling fluid shut-off valve, fluid draining system, and fire 
detection system may be necessary because a potential ignition source 
(electrical failure) and flammable fluids share the same area in the 
aircraft.
    The FAA does not concur with Airbus's comment that special 
condition no. 4 does not prescribe safety criteria for flammable 
cooling fluids. Special condition no. 4 incorporates Sec.  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 Sec.  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 (Sec.  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

[[Page 105440]]

failures leading to damage to engine control system components. No 
changes were made as a result of this comment.
    Ampaire stated a system safety assessment is required by Sec.  
33.28 but there is no requirement in part 33 to add the rates of 
hazardous and major faults in the installation manual. Ampaire asked 
the FAA to explain why this requirement is included in special 
condition no. 10(g) for the Safran electric engine but not in part 33 
for reciprocating and gas turbine engines.
    The FAA added the requirement because electric engines enable a 
wide variety of new aircraft propulsion features, and the engine 
control system safety assessment is tied to these new propulsion 
features, which support aircraft that combine vertical takeoff and 
landing, multi-engine distributed-propulsion, propeller lift and tilt-
wing functions, and zero velocity inflight maneuvering capabilities. 
The effects of an engine failure, such as power loss from an engine, 
and hazards to the aircraft are contingent on the aircraft design. 
Therefore, the hazards identified in the safety analysis, as well as 
the hazard level rates, are included in the engine installation manual 
to ensure any assumptions about aircraft capabilities that mitigate the 
effects of engine failures are taken into account when deciding if an 
engine can be installed in an aircraft. No changes were made as a 
result of this comment.
    Ampaire asked the FAA to explain why the reference to special 
condition no. 31, Operation with variable pitch propeller, is included 
in the magniX special condition no. 17(d)(1), Safety analysis, but not 
the Safran proposed special condition no. 17(d)(1).
    Safran's electric engine will be used with a fixed-pitch propeller, 
and therefore special condition no. 31 is not applicable to the Safran 
engine type design. No changes were made as a result of this comment.
    Ampaire stated proposed special condition no. 23, Endurance 
demonstration, implies that endurance testing requires a demonstration 
of energy regeneration, but energy regeneration might not be a feature 
for some electric engines that operate normally at their limits. 
Ampaire suggested replacing the second sentence in special condition 
no. 23 with ``The endurance demonstration must include dwellings and 
increases and decreases of the engine's power settings for sufficient 
durations that produce the extreme physical conditions the engine 
experiences at rated performance levels, operational limits, and at any 
other conditions or power settings including energy regeneration that 
are required to verify the limit capabilities of the engine.''
    The FAA concurs with Ampaire's comment that energy regeneration 
might not be a feature for some electric engines that operate at their 
limits. The phrase ``that produce the extreme physical conditions'' in 
special condition no. 23 indicates the endurance test addresses engine 
properties where the extreme physical conditions can occur including 
conditions that cause the engine to operate at its limits of energy 
regeneration. As a result of this comment, the FAA changed special 
condition no. 23 in accordance with Ampaire's recommendation.
    Ampaire requested the FAA revise special condition nos. 33(c)(1) 
and (d) for electrical power distribution and protection systems, 
respectively, by adding the conditional statement ``due to a single 
fault'' and explained electrical power distribution within the part 33 
powerplant may take several faults to result in total loss. Ampaire 
also stated that electric power distribution outside the part 33 
powerplant is the subject of part 23 aircraft certification.
    The FAA does not concur with Ampaire's request to revise special 
condition nos. 33(c)(1) and (d) to provide protection from potential 
consequences resulting only from single electrical faults. Special 
condition nos. 33(g)(2) and (3), Electrical system failures, have 
safety criteria that already address single faults in all the engine 
electrical systems. The safety criteria in special condition no. 
33(c)(1) and (d) are for loss of function in electrical power 
distribution systems, and the criteria apply regardless of the cause of 
system failures or malfunctions. Also, part 33 has provisions for 
electrical power supplied to electrical control systems, and therefore 
this special condition is within the scope of engine requirements. No 
changes were made as a result of this comment.
    Ampaire asked the FAA to explain the regulatory significance of the 
term ``detrimental'' as it is used in proposed special condition no. 
33, Engine electrical systems, and whether the term relates to hazard 
levels.
    The FAA intends the term ``detrimental'' to have the same meaning 
as the meaning of the term as it is commonly used in the English 
language. The term is used extensively in existing FAA regulations and 
guidance. There is no intent to change how the term is used in these 
special conditions. Also, there is no correlation between the term 
``detrimental'' and engine failure effect hazard levels. The term is 
intended to capture all engine effects that could result in an unsafe 
engine condition. No changes were made as a result of this comment.
    The FAA received several comments from magniX.
    MagniX noted proposed special condition nos. 1(b) and (c) state 
that a means of compliance, which may include consensus standards, must 
be ``accepted by the Administrator'' and ``in a form and manner 
acceptable to the Administrator.'' MagniX explained that these 
paragraphs are directly out of 14 CFR 23.2010, which contains 
performance-based language. MagniX also explained that part 33 and the 
Safran special conditions are prescriptive regulations, not 
performance-based. MagniX further indicated that requiring a 
performance-based process for establishing means of compliance with 
prescriptive regulations is unnecessary and overly burdensome to 
applicants and regulators. MagniX recommended the FAA not adopt 
proposed special condition nos. 1(b) and (c).
    The FAA does not concur with magniX's recommendation. The FAA 
considers special condition nos. 1(b) and (c) to be essential for 
achieving an equivalent level of safety to the level of safety provided 
by the part 33 engine requirements. The Safran electric engine criteria 
are a combination of part 33 requirements and special conditions to the 
requirements in part 33. Special conditions are developed under the 
provisions of Sec.  21.16, which are issued when the applicable 
regulations do not contain adequate or appropriate safety standards. 
Special condition nos. 1(b) and (c) will be used to incorporate the 
additional details that apply to the Safran engine design using 
accepted means of compliance. No changes were made as a result of this 
comment.
    MagniX stated proposed special condition nos. 10(g), 15(b), and 
17(f) would require applicants to declare proprietary information in 
the engine installation manual, these documentation requirements 
establish a precedent beyond that required of their existing 
reciprocating or turbine engine counterparts, and these requirements 
increase the risk that sensitive information is disclosed. MagniX 
explained that while it is understood this information is used during 
aircraft-level certification efforts, traditional data sharing 
agreements sufficiently provide the integrator with the required 
information while respecting the proprietary nature of the data. MagniX 
also stated requiring these additional data in the engine installation 
manual overly constrains the means of

[[Page 105441]]

compliance and introduces commercial risk. MagniX recommended the FAA 
not adopt the requirement to include these specific disclosures in the 
engine installation manual. MagniX proposed that these data be provided 
to integrators through generic ``installation instructions'' in lieu of 
the engine installation manual and explained this will allow specific 
proprietary disclosures in other installation documents such as 
interface control drawings, technical memorandums, or other installer 
requested documentation.
    Special condition nos. 10(g), 15(b), and 17(f) do not require the 
disclosure of sensitive information. As discussed in the NPSC, the 
documentation requirements in special conditions nos. 10(g), 15(b), and 
17(f) are expected to ensure that the engine is used safely and 
properly by constraining the installation of electric engines to only 
aircraft types (configurations, flight capabilities, etc.) that were 
used by the engine manufacturer to determine the engine ratings, 
limits, performance characteristics, as well as the reliability and 
criticality of engine systems and parts.
    These documentation requirements are intended, and the FAA finds 
necessary, to ensure enough information is included to safeguard 
compatibility between the electric engine and aircraft, and to prevent 
the engine from being used in an aircraft type that requires safety 
features or performance characteristics that are not available from an 
engine that was type-certificated for an aircraft that does not require 
the same safety features or performance characteristics. The FAA 
modified the proposed special conditions to clarify the requirement by 
specifying the information identified in special condition nos. 6 
``Engine cooling,'' 10 ``Engine control systems,'' 15 ``Power 
response,'' 17 ``Safety analysis,'' 18 ``Ingestion,'' 19 ``Liquid and 
gas systems,'' 30 ``Containment,'' and 33 ``Engine electrical systems'' 
must be documented and provided to the installer as part of the 
requirements in Sec.  33.5.
    The FAA received several comments from Kite Magnetics.
    Kite Magnetics stated that special condition no. 14 for the 
lubrication system of the Safran Model ENGINeUS 100A1 electric engine 
should focus specifically on the unique lubrication attributes and 
inherent functional capabilities of the Safran electric engine design, 
rather than apply requirements for the entire lubrication system. Kite 
Magnetics suggested changing special condition no. 14 to apply 
component-level requirements that would be better suited for the unique 
attributes of electric engines such as the Safran Model ENGINe US 
100A1, promote clarity and relevance of the special condition to 
critical aspects of the lubrication system pertinent to electric 
engines, and avoid unnecessary requirements that do not apply to this 
engine type.
    The FAA does not concur with Kite Magnetics' comment that the 
special conditions for an electric engine lubrication system should be 
established at the component level. These special conditions are 
engine-level requirements; however, the means of compliance to the 
special conditions can involve component-level assessments using 
special condition no. 27, System and component tests, which can focus 
on the unique lubrication attributes and inherent functional 
capabilities of the Safran electric engine design. No changes were made 
as a result of this comment.
    Kite Magnetics stated the language ``Any system or device that 
provides, uses, conditions, or distributes electrical power, and is 
part of the engine type design'' in proposed special condition no. 
33(a) could imply that energy storage systems (ESS) are part of the 
engine electrical system. Kite Magnetics explained that ESS fall under 
the category of systems that provide electrical power and may be 
perceived as part of the engine's electrical system. However, Kite 
Magnetics noted that an ESS is a distinct system that supports the 
engine's electrical power needs, but it is not inherently integrated 
into the engine's core electrical system design. Kite Magnetics 
requested confirmation that special condition 33(a) does not apply to 
ESS. Kite Magnetics did not request changes to this special condition.
    The FAA confirms special condition 33(a) does not apply to ESS. No 
changes were made as a result of this comment.
    Kite Magnetics requested clarification regarding the components and 
devices that are considered part of the engine's electrical power 
distribution system, as outlined in proposed special condition no. 
33(c). Kite Magnetics explained this request is intended to ensure a 
clear understanding of the scope and components included within the 
electrical power distribution system. Kite Magnetics did not request 
changes to this special condition.
    The FAA confirms special condition no. 33(c) applies only to the 
electrical power distribution systems that are part of Safran's 
electric engine type design. However, the partition between the engine 
and aircraft electrical power distribution systems must be clearly 
described and documented with the data provided for showing compliance 
to Sec.  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.

[[Page 105442]]

Applicability

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

Conclusion

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

List of Subjects in 14 CFR Part 33

    Aircraft, Aviation safety, Reporting and recordkeeping 
requirements.

Authority Citation

    The authority citation for these special conditions is as follows:

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

The Special Conditions

0
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 Sec.  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 Sec.  33.15.

(4) Fire Protection

    The engine design must comply with Sec.  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 Sec.  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 Sec.  
33.5.

 (7) Engine Mounting Attachments and Structure

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

 (8) Accessory Attachments

    The engine must comply with Sec.  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--

[[Page 105443]]

    (1) Have a maximum rate of loss of power control (LOPC) that is 
suitable for the intended aircraft application. The estimated LOPC rate 
must be documented and provided to the installer as part of the 
requirements in Sec.  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 Sec.  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 Sec.  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 Sec.  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 Sec.  33.70(a);
    (2) a manufacturing plan, as defined in Sec.  33.70(b); and
    (3) a service-management plan, as defined in Sec.  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

[[Page 105444]]

documented and provided to the installer as part of the requirements in 
Sec.  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 Sec.  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 Sec.  
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 Sec.  33.75(d) and (e) using the 
failure definitions in special condition no. 17(d) of these special 
conditions, and the ICA in Sec.  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 Sec.  
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 Sec.  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 Sec.  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 
Sec.  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 
Sec.  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 Sec.  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 Sec.  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

[[Page 105445]]

settings, energy regeneration, and dwellings at the power settings and 
energy regeneration for sufficient durations that produce the extreme 
physical conditions the engine experiences at rated performance levels, 
operational limits, and at any other conditions or power settings, 
including energy regeneration, which are required to verify the limit 
capabilities of the engine.

 (24) Temperature Limit

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

 (25) Operation Demonstration

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

 (26) Durability Demonstration

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

 (27) System and Component Tests

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

 (28) Rotor Locking Demonstration

    If shaft rotation is prevented by locking the rotor(s), the engine 
must demonstrate:
    (a) Reliable rotor locking performance;
    (b) Reliable rotor unlocking performance; and
    (c) That no hazardous engine effects, as specified in special 
condition no. 17(d)(2) of these special conditions, will occur.

 (29) Teardown Inspection

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

[[Page 105446]]

    (e) Electrical power characteristics. The applicant must identify, 
declare, document, and provide to the installer as part of the 
requirements in Sec.  33.5, the characteristics of any electrical power 
supplied from--
    (1) the aircraft to the engine electrical system, for starting and 
operating the engine, including transient and steady-state voltage 
limits, and
    (2) the engine to the aircraft via energy regeneration, and any 
other characteristics necessary for safe operation of the engine.
    (f) Environmental limits. Environmental limits that cannot 
adequately be substantiated by endurance demonstration, validated 
analysis, or a combination thereof must be demonstrated by the system 
and component tests in special condition no. 27 of these special 
conditions.
    (g) Electrical system failures. The engine electrical system must--
    (1) Have a maximum rate of LOPC that is suitable for the intended 
aircraft application;
    (2) When in the full-up configuration, be single-fault tolerant, as 
determined by the Administrator, for electrical, electrically 
detectable, and electronic failures involving LOPC events;
    (3) Not have any single failure that results in hazardous engine 
effects; and
    (4) Ensure failures or malfunctions that lead to local events in 
the intended aircraft application do not result in hazardous engine 
effects, as defined in special condition no. 17(d)(2) of these special 
conditions, due to electrical system failures or malfunctions.
    (h) System safety assessment. The applicant must perform a system 
safety assessment. This assessment must identify faults or failures 
that affect normal operation, together with the predicted frequency of 
occurrence of these faults or failures. The intended aircraft 
application must be taken into account to assure the assessment of the 
engine system safety is valid. The rates of hazardous and major faults 
must be declared, documented, and provided to the installer as part of 
the requirements in Sec.  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
```