Rule2024-30855
Special Conditions: Safran Electric & Power S.A. Model ENGINeUS 100A1 Electric Engines
Primary source
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Published
December 27, 2024
Effective
January 27, 2025
Issuing agencies
Transportation DepartmentFederal Aviation Administration
Abstract
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.
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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 [<a href="http://www.gpo.gov">www.gpo.gov</a>]
[FR Doc No: 2024-30855]
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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;
<a href="/cdn-cgi/l/email-protection#325f5340591c505d474b5740725453531c555d44"><span class="__cf_email__" data-cfemail="335e5241581d515c464a5641735552521d545c45">[email protected]</span></a>.
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\ <a href="https://www.astm.org/Standards/F3338.html">https://www.astm.org/Standards/F3338.html</a>.
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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\ <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 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\ <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) 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\ <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.
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:
<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
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
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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.