[Federal Register Volume 89, Number 55 (Wednesday, March 20, 2024)]
[Proposed Rules]
[Pages 19763-19775]
From the Federal Register Online via the Government Publishing Office [www.gpo.gov]
[FR Doc No: 2024-05101]
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Proposed Rules
Federal Register
________________________________________________________________________
This section of the FEDERAL REGISTER contains notices to the public of
the proposed issuance of rules and regulations. The purpose of these
notices is to give interested persons an opportunity to participate in
the rule making prior to the adoption of the final rules.
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Federal Register / Vol. 89, No. 55 / Wednesday, March 20, 2024 /
Proposed Rules
[[Page 19763]]
DEPARTMENT OF TRANSPORTATION
Federal Aviation Administration
14 CFR Part 33
[Docket No. FAA-2023-0587; Notice No. 33-23-01-SC]
Special Conditions: Safran Electric & Power S.A. Model ENGINe
US100A1 Electric Engines
AGENCY: Federal Aviation Administration (FAA), Department of
Transportation (DOT).
ACTION: Notice of proposed special conditions.
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SUMMARY: This action proposes special conditions for the Safran
Electric & Power S.A. (Safran) Model ENGINeTM US100A1
electric engines that operate using electrical technology installed on
the aircraft for use as an aircraft engine. These engines have a novel
or unusual design feature when compared to the state of technology
envisioned in the airworthiness standards applicable to aircraft
engines. The design feature is the use of an electric motor, motor
controller, and high-voltage systems as the primary source of
propulsion for an aircraft. The applicable airworthiness regulations do
not contain adequate or appropriate safety standards for this design
feature. These proposed 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: Send comments on or before April 19, 2024.
ADDRESSES: Send comments identified by Docket No. FAA-2023-0587 using
any of the following methods:
Federal eRegulations Portal: Go to https://www.regulations.gov/ and follow the online instructions for sending
your comments electronically.
Mail: Send comments to Docket Operations, M-30, U.S.
Department of Transportation, 1200 New Jersey Avenue SE, Room W12-140,
West Building Ground Floor, Washington, DC 20590-0001.
Hand Delivery or Courier: Take comments to Docket
Operations in Room W12-140 of the West Building Ground Floor at 1200
New Jersey Avenue SE, Washington, DC, between 9 a.m. and 5 p.m., Monday
through Friday, except Federal holidays.
Fax: Fax comments to Docket Operations at 202-493-2251.
Docket: Background documents or comments received may be read at
https://www.regulations.gov/ at any time. Follow the online
instructions for accessing the docket or go to Docket Operations in
Room W12-140 of the West Building Ground Floor at 1200 New Jersey
Avenue SE, Washington, DC, between 9 a.m. and 5 p.m., Monday through
Friday, except Federal holidays.
FOR FURTHER INFORMATION CONTACT: Mark Bouyer, Engine and Propulsion
Standards Section, AIR-625, Technical Policy Branch, Policy and
Standards Division, Aircraft Certification Service, 1200 District
Avenue, Burlington, Massachusetts 01803; telephone (781) 238-7755;
[email protected].
SUPPLEMENTARY INFORMATION:
Comments Invited
The FAA invites interested people to take part in this rulemaking
by sending written comments, data, or views. The most helpful comments
reference a specific portion of the proposed special conditions,
explain the reason for any recommended change, and include supporting
data.
The FAA will consider all comments received by the closing date for
comments. The FAA may change these proposed special conditions based on
the comments received.
Privacy
Except for Confidential Business Information (CBI) as described in
the following paragraph, and other information as described in title
14, Code of Federal Regulations (14 CFR) 11.35, the FAA will post all
comments received, without change, to https://www.regulations.gov/,
including any personal information you provide. The FAA will also post
a report summarizing each substantive verbal contact received about
these special conditions.
Confidential Business Information
Confidential Business Information is commercial or financial
information that is both customarily and actually treated as private by
its owner. Under the Freedom of Information Act (FOIA) (5 U.S.C. 552),
CBI is exempt from public disclosure. If your comments responsive to
this notice of proposed special conditions contain commercial or
financial information that is customarily treated as private, that you
actually treat as private, and that is relevant or responsive to this
notice of proposed special conditions, it is important that you clearly
designate the submitted comments as CBI. Please mark each page of your
submission containing CBI as ``PROPIN.'' The FAA will treat such marked
submissions as confidential under the FOIA, and the indicated comments
will not be placed in the public docket of these proposed special
conditions. Send submissions containing CBI to the individual listed in
the For Further Information Contact section below. Comments the FAA
receives, which are not specifically designated as CBI, will be placed
in the public docket for these proposed special conditions.
Background
On November 27, 2020, Safran applied for FAA validation for a type
certificate for their Model ENGINeTM US100A1 electric
engine. The Safran Model ENGINeTM US100A1 electric engine
will be used in a single-engine airplane that will be certified
separately from the engine.
The Safran Model ENGINeTM US100A1 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 ENGINeTM US100A1 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 ENGINeTM US100A1
engines because of a novel or unusual
[[Page 19764]]
design feature, special conditions may be prescribed under the
provisions of Sec. 21.16.
Special conditions are initially applicable to the model for which
they are issued. Should the type certificate for that model be amended
later to include any other engine model that incorporates the same
novel or unusual design feature, these special conditions would also
apply to the other engine model under Sec. 21.101.
The FAA issues special conditions, as defined in Sec. 11.19, in
accordance with Sec. 11.38, and they become part of the type
certification basis under Sec. 21.17(a)(2).
In addition to the applicable airworthiness regulations and special
conditions, the Safran Model ENGINeTM US100A1 engines must
comply with the noise certification requirements of 14 CFR part 36.
Novel or Unusual Design Features
The Safran Model ENGINeTM US100A1 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 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 ENGINeTM US100A1
engine design to develop special conditions that establish an
equivalent level of safety to that required by part 33.
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\1\ https://www.astm.org/Standards/F3338.htm.
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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 Proposed 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 ENGINeTM US100A1
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 Safran aircraft engine is
subjected 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 ENGINeTM US100A1 engine's
novel or unusual design feature.
Safran's proposed 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 proposed special conditions.
Safran's Proposed Electric Engines Require a Mix of Part 33 Standards
and Special Conditions
Although the electric aircraft engines Safran proposes use novel or
unusual design features that the FAA did not envisage during the
development of its existing part 33 airworthiness standards, these
engines share some basic similarities, in configuration and function,
to engines that use the combustion of air and fuel, and therefore
require similar provisions to prevent common hazards (e.g., fire,
uncontained high energy debris, and
[[Page 19765]]
loss of thrust control). However, the primary failure concerns and the
probability of exposure to these common hazards are different for the
proposed Safran Model ENGINeTM US100A1 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
ENGINeTM US100A1, which operates using electrical technology
as the primary means of propelling the aircraft. This necessitates the
development of special conditions that provide adequate airworthiness
standards for these aircraft engines.
The requirements in part 33, subpart B, are applicable to
reciprocating and turbine aircraft engines. Subparts C and D are
applicable to reciprocating aircraft engines. Subparts E through G are
applicable to turbine aircraft engines. As such, subparts B through G
do not adequately address the use of aircraft engines that operate
using electrical technology. Special conditions are needed to ensure a
level of safety for electric engines that is commensurate with these
subparts, as those regulatory requirements do not contain adequate or
appropriate safety standards for electric aircraft engines that are
used to propel aircraft.
FAA Proposed Special Conditions for the Safran Engine Design
Applicability: Proposed special condition no. 1 would require
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: Proposed special
condition no. 2 would, in addition to compliance with Sec. 33.7(a),
require Safran to establish engine operating limits related to the
power, torque, speed, and duty cycles specific to Safran Model
ENGINeTM US100A1 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
ENGINeTM US100A1 engine capabilities required for safe
operation of the engine.
Materials: Proposed special condition no. 3 would require 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: Proposed 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 proposed special condition
would require 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: Proposed special condition no. 5 would require the
design and construction of Safran Model ENGINeTM US100A1
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: Proposed special condition no. 6 would require
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 proposed special condition would require 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 proposed special condition no. 17.
Loss of cooling to an aircraft engine that operates using electrical
technology can result in rapid overheating and abrupt engine failure,
with critical consequences to safety.
Engine Mounting Attachments and Structure: Proposed special
condition no. 7 would require Safran and the proposed design to comply
with Sec. 33.23, which requires the applicant to define, and the
proposed design to withstand, certain load limits for the engine
mounting attachments and related engine structure. These requirements
would otherwise be applicable only to reciprocating and turbine
aircraft engines.
Accessory Attachments: Proposed special condition no. 8 would
require the proposed design to comply with Sec. 33.25, which sets
certain design, operational, and maintenance requirements for the
engine's accessory drive and mounting attachments, and which would
otherwise be applicable only to reciprocating and turbine aircraft
engines.
Rotor Overspeed: Proposed special condition no. 9 would require
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 proposed 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: Proposed special condition no. 10(b) would
require 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
[[Page 19766]]
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 proposed special
condition no. 10(b) associated with controlling these engines is
necessary.
Proposed special condition no. 10(c) would require Safran to
develop and verify the software and complex electronic hardware used in
programmable logic devices, using proven methods that ensure that the
devices can provide the accuracy, precision, functionality, and
reliability commensurate with the hazard that is being mitigated by the
logic. RTCA DO-254, ``Design Assurance Guidance for Airborne Electronic
Hardware'' dated April 19, 2000, distinguishes between complex and
simple electronic hardware.\3\
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\3\ https://my.rtca.org/NC__Product?id=a1B36000001IcjTEAS.
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Proposed special condition no. 10(d) would require 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 proposed special condition
no. 10(e) include temperature, vibration, high-intensity radiated
fields (HIRF), and others addressed in RTCA DO-160G, ``Environmental
Conditions and Test Procedures for Airborne Electronic/Electrical
Equipment and Instruments'' dated December 08, 2010, which includes
``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\ Proposed special
condition 10(e) would require Safran to demonstrate by systems or
component tests in special condition no. 27 any environmental limits
that cannot be adequately substantiated by the endurance demonstration,
validated analysis, or a combination thereof.
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\4\ https://my.rtca.org/NC__Product?id=a1B36000001IcnSEAS.
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Proposed special condition no. 10(f) would require Safran to
evaluate various control system failures to assure that such failures
will not lead to unsafe engine conditions. The FAA issued advisory
circular (AC), AC 33.28-3, ``Guidance Material for 14 CFR 33.28, Engine
Control Systems'' on May 23, 2014, for reciprocating and turbine
engines.\5\ Paragraph 6-2 of 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
(or power) control (LOTC/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 proposed to allow fault
accommodation criteria to be developed for electric engines.
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\5\ https://www.faa.gov/documentLibrary/media/Advisory_Circular/AC_33_28-3.pdf.
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The phrase ``in the full-up configuration'' used in proposed
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
proposed 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.
Proposed special condition no. 10(g) would require Safran to
conduct a safety assessment of the control system to support the safety
analysis in proposed 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.
Proposed 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.
Proposed special condition no. 10(i) is necessary to ensure that
the controllers are self-sufficient and isolated from other aircraft
systems. The aircraft-supplied data supports the analysis at the
aircraft level to protect the aircraft from common mode failures that
could lead to major propulsion power loss. The exception, ``other than
power command signals from the aircraft'' noted in proposed special
condition no. 10(i), is based on the FAA's determination that the
engine controller has no reasonable means to determine the validity of
any in-range signals from the electrical power system. In many cases,
the engine control system can detect a faulty signal from the aircraft,
but the engine control system typically accepts the power command
signal as a valid value.
The term ``independent'' in the context of ``fully independent
engine systems'' referenced in proposed 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 proposed special condition 10(i)(2) is to
assure that, upon detecting a fault, the system continues to function
safely.
Proposed special condition no. 10(j) would require 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: Proposed special condition no. 11 would
require 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. Proposed special condition no. 11
would require 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 ENGINeTM US100A1
engines. However, the FAA
[[Page 19767]]
proposes that a stress analysis particular to these electric engines is
necessary to account for stresses resulting from electric technology
used in the engine.
Proposed special condition no. 12 would require 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: Proposed special condition no. 13
would require 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 proposed 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 proposed 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 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 proposed 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: Proposed special condition no. 14 would require
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 proposed special condition
would also require Safran to demonstrate the unique lubrication
attributes and functional capability of the Safran Model
ENGINeTM US100A1 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
ENGINeTM US100A1 engines. For example, electric engines do
not have a crankcase or lubrication oil sump. Electric engine bearings
are sealed, so they do not require an oil circulation system. The
lubrication system in these engines is also independent of the
propeller pitch control system. Therefore, proposed special condition
no. 14 incorporates only certain requirements from the part 33
regulations.
Power Response: Proposed special condition no. 15 would require the
design and construction of the Safran Model ENGINeTM US100A1
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 ENGINeTM US100A1 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 proposed special condition is
necessary.
Continued Rotation: Proposed special condition no. 16 would require
Safran to design the Model ENGINeTM US100A1 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 ENGINeTM
US100A1 engines consists of the rotors, shafts, magnets, bearings, and
wire windings that convert electrical energy to shaft torque. For the
initial aircraft
[[Page 19768]]
application, this rotating system must continue to rotate after the
power source to the engine is shut down. The safety concerns associated
with this proposed 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: Proposed special condition no. 17 would require
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 proposed special condition would require 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 proposed 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 proposed special condition nos.
10(g) and 33(h).
Ingestion: Proposed special condition no. 18 would require 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 electric engines are not configured
with inlet ducts.
An ``unacceptable'' power loss, as used in proposed 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: Proposed special condition no. 19 would
require 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: Proposed special condition no. 20 would
require Safran to ensure the engine--
(1) is designed and constructed to function throughout its normal
operating range of rotor speeds and engine output power without
inducing excessive stress caused by engine vibration, and
(2) design undergoes a vibration survey.
The vibration demonstration is a survey that characterizes the
vibratory attributes of the engine. It verifies that the stresses from
vibration do not impose excessive force or result in natural frequency
responses on the aircraft structure. The vibration demonstration also
ensures internal vibrations will not cause engine components to fail.
Excessive vibration force occurs at magnitudes and forcing functions or
frequencies, which may result in damage to the aircraft. Stress margins
to failure add conservatism to the highest values predicted by analysis
for additional protection from failure caused by influences beyond
those quantified in the analysis. The result of the additional design
margin is improved engine reliability that meets prescribed thresholds
based on the failure classification. The amount of margin needed to
achieve the prescribed reliability rates depends on an applicant's
experience with a product. The FAA considers the reliability rates when
deciding how much vibration is excessive.
Overtorque: Proposed special condition no. 21 would require 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 proposed 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: Proposed special condition no. 22 would
require Safran to subject the engine to calibration tests to establish
its power characteristics and the conditions both before and after the
endurance and durability demonstrations specified in proposed special
condition nos. 23 and 26. The calibration test requirements specified
in Sec. 33.85 only apply to the endurance test specified in Sec.
33.87, which is applicable only to turbine engines. The FAA proposes
that the methods used for accomplishing those tests for turbine engines
is not the best
[[Page 19769]]
approach for electric engines. The calibration tests in Sec. 33.85
have provisions applicable to ratings that are not relevant to the
Safran Model ENGINeTM US100A1 engines. Proposed special
condition no. 22 would allow 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 proposed special condition applies the calibration
requirement to both the endurance and durability tests.
Endurance Demonstration: Proposed special condition no. 23 would
require 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: Proposed special condition no. 24 would require
Safran to ensure the engine can endure operation at its temperature
limits plus an acceptable margin. An ``acceptable margin,'' as used in
the proposed 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: Proposed special condition no. 25 would
require 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 ENGINeTM US100A1 engines
can achieve without incurring hazardous effects. The characteristics
are requisite capabilities of the type design that qualify the engine
for installation into aircraft and that determine aircraft installation
requirements. The primary engine operating characteristics are assessed
by the tests and demonstrations that would be required by these special
conditions. Some of these characteristics are shaft output torque,
rotor speed, power consumption, and engine thrust response. The engine
performance data Safran will use to certify the engine must account for
installation loads and effects. These are aircraft-level effects that
could affect the engine characteristics that are measured when the
engine is tested on a stand or in a test cell. These effects could
result from elevated inlet cowl temperatures, aircraft maneuvers,
flowstream distortion, and hard landings. For example, an engine that
is run in a sea-level, static test facility could demonstrate more
capability for some operating characteristics than it will have when
operating on an aircraft in certain flight conditions. Discoveries like
this during certification could affect proposed engine ratings and
operating limits. Therefore, the installed performance defines the
engine performance capabilities.
Durability Demonstration: Proposed special condition no. 26 would
require 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: Proposed special condition no. 27 would
require Safran to show that the systems and components of the engine
would 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 proposes that the Sec. 33.87
endurance test might not be the best way to achieve the highest thermal
conditions for all the electronic components of electric engines
because heat is generated differently in electronic systems than it is
in turbine engines. Additional safety considerations also need to be
addressed in the test. Therefore, proposed special condition no. 27
would be 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 ENGINeTM US100A1 electric engine.
Rotor Locking Demonstration: Proposed special condition no. 28
would require the engine to demonstrate reliable rotor locking
performance and that no hazardous effects will occur if the engine uses
a rotor locking device to prevent shaft rotation.
Some engine designs enable the pilot to prevent a propeller shaft
or main rotor shaft from turning while the engine is running, or the
aircraft is in-flight. This capability is needed for some installations
that require the pilot to confirm functionality of certain flight
systems before takeoff. The proposed 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 proposed special condition does not define ``reliable'' rotor
locking but would allow 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: Proposed special condition no. 29 would
require Safran to perform a teardown or non-teardown evaluation, after
the endurance, durability, and overtorque demonstrations, based on the
criteria proposed in special condition no. 29(a) or (b).
[[Page 19770]]
Proposed 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 proposed
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: Proposed special condition no. 30 would require 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 condition would require
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.
Operation with a Variable Pitch Propeller: Proposed special
condition no. 31 would require Safran to conduct functional
demonstrations, including feathering, negative torque, negative thrust,
and reverse thrust operations, as applicable, based on the propeller's
or fan's variable pitch functions that are planned for use on these
electric engines, using a representative propeller. The requirements of
Sec. 33.95 prescribe tests based on the operating characteristics of
turbine engines equipped with variable pitch propellers, which include
thrust response times, engine stall, propeller shaft overload, loss of
thrust control, and hardware fatigue. The electric engines Safran
proposes have different operating characteristics that substantially
affect their susceptibility to these and other potential failures
typical of turbine engines. Because Safran's proposed electric engines
may be installed with a variable pitch propeller, the proposed special
condition is necessary.
General Conduct of Tests: Proposed special condition no. 32 would
require Safran to--
(1) include scheduled maintenance in the engine ICA;
(2) include any maintenance, in addition to the scheduled
maintenance, that was needed during the test to satisfy the applicable
test requirements; and
(3) 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 proposed 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
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. Proposed 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.
[[Page 19771]]
Proposed special condition no. 33(b) would ensure that all aspects
of an electrical system, including generation, distribution, and usage,
do not experience any unacceptable operating characteristics.
Proposed special condition no. 33(c) would require the electrical
power distribution aspects of the electrical system to provide the safe
transfer of electrical energy throughout the electric engine.
Proposed special condition no. 33(d) would require 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.
Proposed 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.
Proposed special condition no. 33(f) requires Safran to demonstrate
that systems and components will operate properly up to environmental
limits, using special conditions, when such limits cannot be adequately
substantiated by the endurance demonstration, validated analysis, or a
combination thereof. The environmental limits referred to in this
proposed special condition include temperature, vibration, HIRF, and
others addressed in RTCA DO-160G, ``Environmental Conditions and Test
Procedures for Airborne Electronic/Electrical Equipment and
Instruments.''
Proposed special condition 33(g) would require Safran to evaluate
various electric engine system failures to ensure that these failures
will not lead to unsafe engine conditions. The evaluation would include
single-fault tolerance, would ensure no single electrical or electronic
fault or failure would result in hazardous engine effects, and ensure
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 would also implement integrity requirements,
criteria for LOTC/LOPC events, and an acceptable LOTC/LOPC rate.
Proposed special condition 33(h) would require Safran to conduct a
safety assessment of the engine electrical system to support the safety
analysis in special condition no. 17. This safety assessment provides
engine response to failures, and rates of these failures, that can be
used at the aircraft safety assessment level.
These proposed 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 for reciprocating and turbine aircraft engines.
Applicability
As discussed above, these proposed special conditions are
applicable to Safran Model ENGINeTM US100A1 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 ENGINeTM US100A1
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 Proposed Special Conditions
0
Accordingly, the Federal Aviation Administration (FAA) proposes the
following special conditions as part of the type certification basis
for Safran Electric & Power S.A. Model ENGINeTM US100A1
engines. The applicant must also comply with the certification
procedures set forth in title 14, Code of Federal Regulations (14 CFR)
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 14 CFR 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).
(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
Instructions for Continued Airworthiness (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
engine installation manual.
[[Page 19772]]
(7) Engine mounting attachments and structure.
The engine mounting attachments and related engine structures must
comply with Sec. 33.23.
(8) Accessory Attachments.
The engine must comply with Sec. 33.25.
(9) Overspeed.
(a) A rotor overspeed must not result in a burst, rotor growth, or
damage that results in a hazardous engine effect, as defined in special
condition no. 17(d)(2) of these special conditions. Compliance with
this paragraph must be shown by test, validated analysis, or a
combination of both. Applicable assumed rotor speeds must be declared
and justified.
(b) Rotors must possess sufficient strength with a margin to burst
above certified operating conditions and above failure conditions
leading to rotor overspeed. The margin to burst must be shown by test,
validated analysis, or a combination thereof.
(c) The engine must not exceed the rotor speed operational
limitations that could affect rotor structural integrity.
(10) Engine Control Systems.
(a) Applicability. The requirements of this special condition apply
to any system or device that is part of the engine type design that
controls, limits, monitors, or protects engine operation, and is
necessary for the continued airworthiness of the engine.
(b) Engine control. The engine control system must ensure that the
engine does not experience any unacceptable operating characteristics
or exceed its operating limits, including in failure conditions where
the fault or failure results in a change from one control mode to
another, from one channel to another, or from the primary system to the
back-up system, if applicable.
(c) Design Assurance. The software and complex electronic hardware,
including programmable logic devices, must be--
(1) Designed and developed using a structured and systematic
approach that provides a level of assurance for the logic commensurate
with the hazard associated with the failure or malfunction of the
systems in which the devices are located; and
(2) Substantiated by a verification methodology acceptable to the
Administrator.
(d) Validation. All functional aspects of the control system must
be substantiated by test, analysis, or a combination thereof, to show
that the engine control system performs the intended functions
throughout the declared operational envelope.
(e) Environmental Limits. Environmental limits that cannot be
adequately substantiated by endurance demonstration, validated
analysis, or a combination thereof must be demonstrated by the system
and component tests in special condition no. 27 of these special
conditions.
(f) Engine control system failures. The engine control system
must--
(1) Have a maximum rate of loss of power control (LOPC) that is
suitable for the intended aircraft application. The estimated LOPC rate
must be specified in the engine installation manual;
(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 declared in the engine installation manual.
(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 and declare, in the engine
installation manual, 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.
[[Page 19773]]
(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 included in the engine installation manual.
(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 in the engine installation manual.
(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 in the engine installation manual.
(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 in the engine installation manual.
(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
[[Page 19774]]
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 listed in the engine
installation manual.
(20) Vibration demonstration.
(a) The engine must be designed and constructed to function
throughout its normal operating range of rotor speeds and engine output
power, including defined exceedances, without inducing excessive stress
in any of the engine parts because of vibration and without imparting
excessive vibration forces to the aircraft structure.
(b) Each engine design must undergo a vibration survey to establish
that the vibration characteristics of those components subject to
induced vibration are acceptable throughout the declared flight
envelope and engine operating range for the specific installation
configuration. The possible sources of the induced vibration that the
survey must assess are mechanical, aerodynamic, acoustical, internally
induced electromagnetic, installation induced effects that can affect
the engine vibration characteristics, and likely environmental effects.
This survey must be shown by test, validated analysis, or a combination
thereof.
(21) Overtorque.
When approval is sought for a transient maximum engine overtorque,
the applicant must demonstrate by test, validated analysis, or a
combination thereof, that the engine can continue operation after
operating at the maximum engine overtorque condition without
maintenance action. Upon conclusion of overtorque tests conducted to
show compliance with this special condition, or any other tests that
are conducted in combination with the overtorque test, each engine part
or individual groups of components must meet the requirements of
special condition no. 29 of these special conditions.
(22) Calibration assurance.
Each engine must be subjected to calibration tests to establish its
power characteristics, and the conditions both before and after the
endurance and durability demonstrations specified in special conditions
nos. 23 and 26 of these special conditions.
(23) Endurance demonstration.
The applicant must subject the engine to an endurance
demonstration, acceptable to the Administrator, to demonstrate the
engine's limit capabilities. The endurance demonstration must include
increases and decreases of the engine's power settings, energy
regeneration, and dwellings at the power settings or 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 that
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, that 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 in the engine installation manual, 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) Operation with variable pitch propeller.
The applicant must conduct functional demonstrations including
feathering, negative torque, negative
[[Page 19775]]
thrust, and reverse thrust operations, as applicable, with a
representative propeller. These demonstrations may be conducted in a
manner acceptable to the Administrator as part of the endurance,
durability, and operation demonstrations.
(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, 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 electrical power plant. 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, or detrimental engine effects in the intended aircraft
application.
(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
affecting the safe transmission of electric energy to the electric
engine.
(d) Protection systems. The engine electrical system must be
designed such that the loss, malfunction, interruption of the
electrical power source, or power conditions that exceed design limits
will not result in a hazardous engine effect, as defined in special
condition no. 17(d)(2) of these special conditions.
(e) Electrical power characteristics. The applicant must identify
and declare, in the engine installation manual, 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, or
(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 loss of power control (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.
Issued in Kansas City, Missouri, on March 6, 2024.
Patrick R. Mullen,
Manager, Technical Policy Branch, Policy and Standards Division,
Aircraft Certification Service.
[FR Doc. 2024-05101 Filed 3-19-24; 8:45 am]
BILLING CODE 4910-13-P