Rule2026-01642
Special Conditions: Safran Electric & Power S.A. ENGINeTM US100 Series Electric Engines
Primary source
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Published
January 27, 2026
Effective
January 27, 2026
Issuing agencies
Transportation DepartmentFederal Aviation Administration
Abstract
This action amends the applicability of special conditions that were issued for the Safran Electric & Power S.A. Model ENGINe US100A1 electric engines. These engines have a novel or unusual design feature when compared to the state of technology envisioned in the airworthiness standards applicable to aircraft engines. The 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. This action amends the applicability of Special Conditions No. 33-23-01-SC, dated December 27, 2024, which 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, to include the ENGINe US100 series electric engines.
Full Text
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<title>Federal Register, Volume 91 Issue 17 (Tuesday, January 27, 2026)</title>
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[Federal Register Volume 91, Number 17 (Tuesday, January 27, 2026)]
[Rules and Regulations]
[Pages 3360-3372]
From the Federal Register Online via the Government Publishing Office [<a href="http://www.gpo.gov">www.gpo.gov</a>]
[FR Doc No: 2026-01642]
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DEPARTMENT OF TRANSPORTATION
Federal Aviation Administration
14 CFR Part 33
[Docket No. FAA-2023-0587; Special Conditions No. 33-030A-SC]
Special Conditions: Safran Electric & Power S.A. ENGINe\TM\ US100
Series Electric Engines
AGENCY: Federal Aviation Administration (FAA), DOT.
ACTION: Final special conditions, amendment.
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SUMMARY: This action amends the applicability of special conditions
that were issued for the Safran Electric & Power S.A. Model ENGINe
US100A1 electric engines. These engines have a novel or unusual design
feature when compared to the state of technology envisioned in the
airworthiness standards applicable to aircraft engines. The 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. This action
amends the applicability of Special Conditions No. 33-23-01-SC, dated
December 27, 2024, which 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,
to include the ENGINe US100 series electric engines.
DATES: Effective January 27, 2026.
FOR FURTHER INFORMATION CONTACT: Mark Bouyer, Engine & Propulsion
Standards Section, AIR-625, Technical Policy Branch, Policy & Standards
Division, Aircraft Certification Service, 1200 District Avenue,
Burlington, Massachusetts 01803; telephone (781) 238-7755;
<a href="/cdn-cgi/l/email-protection#a7cac6d5cc89c5c8d2dec2d5e7c1c6c689c0c8d1"><span class="__cf_email__" data-cfemail="3c515d4e57125e534945594e7c5a5d5d125b534a">[email protected]</span></a>.
SUPPLEMENTARY INFORMATION:
Background
On November 27, 2020, Safran applied for FAA validation of their
pending EASA type certificate application for the Model ENGINe US100A1
electric engine. The FAA published a Notice of Proposed Special
Conditions for public comment in the Federal Register on March 20, 2024
(89 FR 19763). On December 27, 2024, the FAA issued Final Special
Conditions No. 33-23-01-SC (89 FR 105432) for the US100A1 electric
engine.
Safran subsequently amended its EASA application to include the
models US100B1 and US100B2. Similarly, on January 15, 2025, Safran
amended their FAA application to include the B1 and B2 on the requested
type certificate for the ENGINe US100 electric engines. This amendment
to the special conditions updates the applicability of Special
Conditions No. 33-23-01-SC from Safran Electric & Power S.A. ``Model
ENGINe US100A1 Electric Engines'' to ``ENGINe US100 Series Electric
Engines''. The novel or unusual design of these additional models is
adequately addressed by the requirements of Special Conditions No. 33-
23-01-SC, and the requirements of these special conditions are
substantively unchanged. However, the
[[Page 3361]]
FAA is amending the Special Conditions number for these final special
conditions from Special Conditions No. 33-23-01-SC to Special
Conditions No. 33-030A-SC.
Type Certification Basis
Under the provisions of 14 CFR 21.17(a)(1) Safran must show that
ENGINe US100 series 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 ENGINe US100 series 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).
In addition to the applicable airworthiness regulations and special
conditions, the Safran ENGINe US100 series electric engines must comply
with the noise certification requirements of 14 CFR part 36.
Novel or Unusual Design Feature
The Safran ENGINe US100 series electric engines will incorporate
the following novel or unusual design feature:
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.
Safran's proposed aircraft engines will operate using electrical
power instead of air and fuel combustion to propel the aircraft. 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.
Safran's Proposed Electric Engines Require a Mix of Part 33 Standards
and Special Conditions
Although the electric aircraft engines Safran proposes 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 proposed Safran
ENGINe<SUP>TM</SUP> US100 series electric engines. 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
ENGINe US100A1, 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 ENGINe US100 series 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 ENGINe US100 series 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 requires 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 Safran to design and
construct the engines to minimize the development of an unsafe
condition between maintenance intervals, overhaul periods, and
mandatory actions described in the ICA to comply with Sec. 33.19. This
condition also requires Safran to develop maintenance instructions and
scheduling information.
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
[[Page 3362]]
and usage in the engine installation manual, in accordance with Sec.
33.5, if cooling is required to satisfy the safety analysis described
in proposed 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 proposed design to comply with Sec. 33.23,
which requires the applicant to define, and the proposed design to
withstand, certain load limits for the engine mounting attachments and
related engine structure. These requirements would otherwise be
applicable only to reciprocating and turbine aircraft engines.
Accessory Attachments: Special condition no. 8 requires the
proposed 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 comply
with Sec. 33.27 to establish by test, validated analysis, or a
combination of both, that--
(a) the rotor overspeed must not result in a burst, rotor growth,
or damage that results in a hazardous engine effect;
(b) rotors must possess sufficient strength margin to prevent
burst; and
(c) 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 comply with Sec. 33.28 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, distinguishes between complex and simple electronic
hardware.\1\
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\1\ <a href="https://my.rtca.org/NC__Product?id=a1B36000001IcjTEAS">https://my.rtca.org/NC__Product?id=a1B36000001IcjTEAS</a>.
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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 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.\2\ Special condition
10(e) requires Safran to demonstrate by system or component tests in
accordance with special condition no. 27 any environmental limits that
cannot be adequately substantiated by the endurance demonstration,
validated analysis, or combination thereof.
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\2\ <a href="https://my.rtca.org/NC__Product?id=a1B36000001IcnSEAS">https://my.rtca.org/NC__Product?id=a1B36000001IcnSEAS</a>.
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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), 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.\3\ 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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\3\ <a href="https://www.faa.gov/documentLibrary/media/Advisory_Circular/AC_33_28-3.pdf">https://www.faa.gov/documentLibrary/media/Advisory_Circular/AC_33_28-3.pdf</a>.
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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.
[[Page 3363]]
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 proposed 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 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 ENGINe<SUP>TM</SUP> US100 series
electric engines. However, a stress analysis particular to these
electric engines is necessary to account for stresses resulting from
electric technology used in the electric engine series.
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 that 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.
[[Page 3364]]
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 condition also requires Safran to demonstrate
the unique lubrication attributes and functional capability of the
Safran ENGINe<SUP>TM</SUP> US100 series 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 ENGINe<SUP>TM</SUP>
US100 series electric engines. For example, electric engines do not
have a 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 ENGINe<SUP>TM</SUP> US100 series 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 ENGINe<SUP>TM</SUP> US100 series 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 ENGINe<SUP>TM</SUP> US100 series 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 ENGINe<SUP>TM</SUP> US100
series 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 ENGINe<SUP>TM</SUP> US100 series 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.
[[Page 3365]]
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 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 proposed 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 ENGINe<SUP>TM</SUP> US100 series 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 to 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 ENGINe US100 series 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 proposed 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
[[Page 3366]]
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. 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 ENGINe US100 series electric engines.
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 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 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
[[Page 3367]]
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
codified, the only electrical systems used on turbine engines were low-
voltage, electronic engine control systems (EEC) and the 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 it 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 electrical power. The term ``abnormal
conditions'' refers to various engine operating conditions such as:
<bullet> System or component characteristics outside of normal
statistical variation from circumstances such as systems degradation,
installation error, and engine response to fault conditions;
<bullet> Unusual environmental conditions from extreme temperature,
humidity, vibration, lightning, high-intensity radiated field (HIRF),
atmospheric neutron radiation; and
<bullet> 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
proposed 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 that can be
used at the aircraft safety assessment level.
The 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.
Applicability
As discussed above, these special conditions are applicable to
Safran ENGINe US100 series 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 ENGINe US100 series 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), 40113, 44701, 44702, 44704.
The Special Conditions
Accordingly, pursuant to the authority delegated to me by the
Administrator, the following special conditions are issued as part of
the type certification basis for Safran ENGINe US100 series 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
[[Page 3368]]
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 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--
(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's 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.
[[Page 3369]]
(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 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
[[Page 3370]]
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 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
[[Page 3371]]
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.
(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.
[[Page 3372]]
(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 in Kansas City, Missouri, on January 22, 2026.
Patrick R. Mullen,
Manager, Technical Policy Branch, Policy and Standards Division,
Aircraft Certification Service.
[FR Doc. 2026-01642 Filed 1-26-26; 8:45 am]
BILLING CODE 4910-13-P
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</html>Indexed from Federal Register on January 27, 2026.
This is legal information, not legal advice. Laws vary by jurisdiction and change frequently. Always verify current law with official sources and consult a licensed attorney in your jurisdiction for advice on your specific situation.