Special Conditions: Safran Electric & Power S.A. Model ENGINe US100A1 Electric Engines, 19763-19775 [2024-05101]
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19763
Proposed Rules
Federal Register
Vol. 89, No. 55
Wednesday, March 20, 2024
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.
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
Federal Aviation
Administration (FAA), Department of
Transportation (DOT).
ACTION: Notice of proposed special
conditions.
AGENCY:
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Confidential Business Information
SUPPLEMENTARY 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.
Comments Invited
Background
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.
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.
FOR FURTHER INFORMATION CONTACT:
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
SUMMARY:
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.
Mark Bouyer, Engine and Propulsion
Standards Section, AIR–625, Technical
Policy Branch, Policy and Standards
Division, Aircraft Certification Service,
1200 District Avenue, Burlington,
Massachusetts 01803; telephone (781)
238–7755; mark.bouyer@faa.gov.
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.
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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
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design feature, special conditions may
be prescribed under the provisions of
§ 21.16.
Special conditions are initially
applicable to the model for which they
are issued. Should the type certificate
for that model be amended later to
include any other engine model that
incorporates the same novel or unusual
design feature, these special conditions
would also apply to the other engine
model under § 21.101.
The FAA issues special conditions, as
defined in § 11.19, in accordance with
§ 11.38, and they become part of the
type certification basis under
§ 21.17(a)(2).
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.
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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
1 https://www.astm.org/Standards/F3338.htm.
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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.
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.
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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 highvoltage electrical system is a
combination of wires and connectors
that integrate the motor and controller.
In addition, the technology
comprising these high-voltage and highcurrent electronic components
introduces potential hazards that do not
exist in turbine and reciprocating
aircraft engines. For example, highvoltage transmission lines,
electromagnetic shields, magnetic
materials, and high-speed electrical
switches are necessary to use the
physical properties of an electric engine
for propelling an aircraft. However, this
technology also exposes the aircraft to
potential failures that are not common
to gas-powered turbine and
reciprocating engines, technological
differences which could adversely affect
safety if not addressed through these
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
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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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 § 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, noload 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
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US100A1 engine capabilities required
for safe operation of the engine.
Materials: Proposed special condition
no. 3 would require Safran to comply
with § 33.15, which sets requirements
for the suitability and durability of
materials used in the engine, and which
would otherwise be applicable only to
reciprocating and turbine aircraft
engines.
Fire Protection: Proposed special
condition no. 4 would require Safran to
comply with § 33.17, which sets
requirements to protect the engine and
certain parts and components of the
airplane against fire, and which would
otherwise be applicable only to
reciprocating and turbine aircraft
engines. Additionally, this 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 § 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 § 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 § 33.23,
which requires the applicant to define,
and the proposed design to withstand,
certain load limits for the engine
mounting attachments and related
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19765
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
§ 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 § 33.28 at
amendment 33–15 to address the
evolution of the means of controlling
the fuel supplied to the engine, from
carburetors and hydro-mechanical
controls to electronic control systems.
These electronic control systems grew
in complexity over the years, and as a
result, the FAA amended § 33.28 at
amendment 33–26 to address these
increasing complexities. The controller
that forms the controlling system for
these electric engines is significantly
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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
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, highintensity 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.
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,
3 https://my.rtca.org/NC__Product?id=
a1B36000001IcjTEAS.
4 https://my.rtca.org/NC__
Product?id=a1B36000001IcnSEAS.
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‘‘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 § 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.
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 aircraftsupplied data supports the analysis at
the aircraft level to protect the aircraft
from common mode failures that could
lead to major propulsion power loss.
The exception, ‘‘other than power
command signals from the aircraft’’
noted in proposed special condition no.
5 https://www.faa.gov/documentLibrary/media/
Advisory_Circular/AC_33_28-3.pdf.
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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 § 33.29(a), (e),
and (g), which set certain requirements
for the connection and installation of
instruments to monitor engine
performance. The remaining
requirements in § 33.29 apply only to
technologies used in reciprocating and
turbine aircraft engines.
Instrument connections (wires, wire
insulation, potting, grounding,
connector designs, etc.) must not
introduce unsafe features or
characteristics to the aircraft. 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
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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 nonredundant 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.’’ Lowcycle 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 lowcycle 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
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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
§ 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 servicemanagement plan ensure that essential
information from the engineering plan,
such as the design characteristics that
safeguard the integrity of critical and
life-limited parts, is consistently
produced, and preserved over the
lifetime of those parts. The
manufacturing plan includes special
processes and production controls to
prevent inclusion of manufacturinginduced anomalies, which can degrade
the part’s structural integrity. Examples
of manufacturing-induced anomalies are
material contamination, unacceptable
grain growth, heat-affected areas, and
residual stresses.
The service-management plan ensures
the method and assumptions used in the
engineering plan to determine the part’s
life remain valid by enabling corrections
identified from in-service experience,
such as service-induced anomalies and
unforeseen environmental effects, to be
incorporated into the design process.
The service-management plan also
becomes the ICA for maintenance,
overhaul, and repairs of the part.
Lubrication System: 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.
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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
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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 § 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 lifelimited or critical parts. Some lifelimited or critical engine parts can fail
suddenly in their primary failure mode,
from prolonged exposure to normal
engine environments such as
temperature, vibration, and stress, if
those engine parts are not removed from
service before the damage mechanisms
progress to a failure. Due to the
consequence of failure, these parts are
not allowed to be managed by oncondition or probabilistic means
because the probability of failure cannot
be sensibly estimated in numerical
terms. Therefore, the parts are managed
by compliance with integrity
requirements, such as mandatory
maintenance (life limits, inspections,
inspection techniques), to ensure the
qualities, features, and other attributes
that prevent the part from failing in its
primary failure mode are preserved
throughout its service life. For example,
if the number of engine cycles to failure
are predictable and can be associated
with specific design characteristics,
such as material properties, then the
applicant can manage the engine part
with life limits.
Complete or total power loss is not
assumed to be a minor engine event, as
it is in the turbine engine regulation
§ 33.75, to account for experience data
showing a potential for higher hazard
levels from power loss events in singleengine general aviation aircraft. The
criteria in these proposed special
conditions apply to an engine that
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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, §§ 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
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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 highvoltage 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 § 33.85 only apply to the
endurance test specified in § 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
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approach for electric engines. The
calibration tests in § 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
§ 33.8, from failing due to temperaturerelated 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
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Safran will use to certify the engine
must account for installation loads and
effects. These are aircraft-level effects
that could affect the engine
characteristics that are measured when
the engine is tested on a stand or in a
test cell. These effects could result from
elevated inlet cowl temperatures,
aircraft maneuvers, flowstream
distortion, and hard landings. For
example, an engine that is run in a sealevel, static test facility could
demonstrate more capability for some
operating characteristics than it will
have when operating on an aircraft in
certain flight conditions. Discoveries
like this during certification could affect
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 § 33.87 endurance test because it is
designed to address safety concerns in
combustion engines. For example,
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§ 33.91(b) requires the establishment of
temperature limits for components that
require temperature-controlling
provisions, and § 33.91(a) requires
additional testing of engine systems and
components where the endurance test
does not fully expose internal systems
and components to thermal conditions
that verify the desired operating limits.
Exceeding temperature limits is a safety
concern for electric engines. The FAA
proposes that the § 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 inflight. 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 § 33.75.
Teardown Inspection: Proposed
special condition no. 29 would require
Safran to perform a teardown or nonteardown evaluation, after the
endurance, durability, and overtorque
demonstrations, based on the criteria
proposed in special condition no. 29(a)
or (b).
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Proposed special condition no. 29(b)
includes restrictive criteria for ‘‘nonteardown 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 § 33.70. Therefore, the
procedure used to determine which
parts are life-limited, and how the life
limits are established, requires FAA
approval, as it does for § 33.70. Sections
33.55 and 33.93 do not contain similar
requirements because reciprocating and
turbine engines can be completely
disassembled for inspection.
Containment: 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.
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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 § 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
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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 § 33.28 was
codified, the only electrical systems
used on turbine engines were lowvoltage, electronic engine control
systems (EEC) and high-energy sparkignition systems. Electric aircraft
engines use high-voltage, high-current
electrical systems and components that
are physically located in the motor and
motor controller. Therefore, the existing
part 33 control system requirements do
not adequately address all the electrical
systems used in electric aircraft engines.
Proposed special condition no. 33 is
established using the existing engine
control systems requirement as a basis.
It applies applicable airworthiness
criteria from § 33.28 and incorporates
airworthiness criteria that recognize and
focus on the electrical power system
used in the engine.
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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
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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
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
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19771
to the FAA in a form and manner
acceptable to the Administrator.
(2) Engine Ratings and Operating
Limits.
In addition to § 33.7(a), the engine
ratings and operating limits must be
established and included in the type
certificate data sheet based on:
(a) Shaft power, torque, rotational
speed, and temperature for:
(1) Rated takeoff power;
(2) Rated maximum continuous
power; and
(3) Rated maximum temporary power
and associated time limit.
(b) Duty cycle and the rating at that
duty cycle. The duty cycle must be
declared in the engine type certificate
data sheet.
(c) Cooling fluid grade or
specification.
(d) Power-supply requirements.
(e) Any other ratings or limitations
that are necessary for the safe operation
of the engine.
(3) Materials.
The engine design must comply with
§ 33.15.
(4) Fire protection.
The engine design must comply with
§ 33.17(b) through (g).
(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 § 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.
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(7) Engine mounting attachments and
structure.
The engine mounting attachments and
related engine structures must comply
with § 33.23.
(8) Accessory Attachments.
The engine must comply with § 33.25.
(9) Overspeed.
(a) A rotor overspeed must not result
in a burst, rotor growth, or damage that
results in a hazardous engine effect, as
defined in special condition no. 17(d)(2)
of these special conditions. Compliance
with this paragraph must be shown by
test, validated analysis, or a
combination of both. Applicable
assumed rotor speeds must be declared
and justified.
(b) Rotors must possess sufficient
strength with a margin to burst above
certified operating conditions and above
failure conditions leading to rotor
overspeed. The margin to burst must be
shown by test, validated analysis, or a
combination thereof.
(c) The engine must not exceed the
rotor speed operational limitations that
could affect rotor structural integrity.
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(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
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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.
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(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
§ 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.
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(13) Critical and life-limited parts.
(a) The applicant must show, by a
safety analysis or means acceptable to
the Administrator, whether rotating or
moving components, bearings, shafts,
static parts, and non-redundant mount
components should be classified,
designed, manufactured, and managed
throughout their service life as critical
or life-limited parts.
(1) Critical part means a part that
must meet prescribed integrity
specifications to avoid its primary
failure, which is likely to result in a
hazardous engine effect as defined in
special condition no. 17(d)(2) of these
special conditions.
(2) Life-limited parts may include but
are not limited to a rotor or major
structural static part, the failure of
which can result in a hazardous engine
effect, as defined in special condition
no. 17(d)(2) of these special conditions,
due to a low-cycle fatigue (LCF)
mechanism. A life limit is an
operational limitation that specifies the
maximum allowable number of flight
cycles that a part can endure before the
applicant must remove it from the
engine.
(b) In establishing the integrity of each
critical part or life-limited part, the
applicant must provide to the
Administrator the following three plans
for approval:
(1) an engineering plan, as defined in
§ 33.70(a);
(2) a manufacturing plan, as defined
in § 33.70(b); and
(3) a service-management plan, as
defined in § 33.70(c).
(14) Lubrication system.
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(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
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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
§ 33.75(a)(1) and (a)(2) using the failure
definitions in special condition no.
17(d) of these special conditions.
(b) The primary failure of certain
single elements cannot be sensibly
estimated in numerical terms. If the
failure of such elements is likely to
result in hazardous engine effects, then
compliance may be shown by reliance
on the prescribed integrity requirements
of § 33.15 and special condition nos. 9
and 13 of these special conditions, as
applicable. These instances must be
stated in the safety analysis.
(c) The applicant must comply with
§ 33.75(d) and (e) using the failure
definitions in special condition no.
17(d) of these special conditions, and
the ICA in § 33.4.
(d) Unless otherwise approved by the
Administrator, the following definitions
apply to the engine effects when
showing compliance with this
condition:
(1) A minor engine effect does not
prohibit the engine from performing its
intended functions in a manner
consistent with § 33.28(b)(1)(i),
(b)(1)(iii), and (b)(1)(iv), and the engine
complies with the operability
requirements of special condition no. 15
and special condition no. 25 of these
special conditions, as appropriate.
(2) The engine effects in § 33.75(g)(2)
are hazardous engine effects with the
addition of:
(i) Electrocution of the crew,
passengers, operators, maintainers, or
others; and
(ii) Blockage of cooling systems that
could cause the engine effects described
in § 33.75(g)(2) and special condition
17(d)(2)(i) of these special conditions.
(3) Any other engine effect is a major
engine effect.
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(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
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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.
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(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,
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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.
(b) Reliable rotor unlocking
performance; and
(c) That no hazardous engine effects,
as specified in special condition no.
17(d)(2) of these special conditions, will
occur.
(29) Teardown inspection.
(a) Teardown evaluation.
(1) After the endurance and durability
demonstrations have been completed,
the engine must be completely
disassembled. Each engine component
and lubricant must be eligible for
continued operation in accordance with
the information submitted for showing
compliance with § 33.4.
(2) Each engine component, having an
adjustment setting and a functioning
characteristic that can be established
independent of installation on or in the
engine, must retain each setting and
functioning characteristic within the
established and recorded limits at the
beginning of the endurance and
durability demonstrations.
(b) Non-Teardown evaluation. If a
teardown cannot be performed for all
engine components in a non-destructive
manner, then the inspection or
replacement intervals for these
components and lubricants must be
established based on the endurance and
durability demonstrations and must be
documented in the ICA in accordance
with § 33.4.
(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.
(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.
(28) Rotor locking demonstration.
If shaft rotation is prevented by
locking the rotor(s), the engine must
demonstrate:
(a) Reliable rotor locking performance;
(31) Operation with variable pitch
propeller.
The applicant must conduct
functional demonstrations including
feathering, negative torque, negative
(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.
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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 § 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.
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(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.
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(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 steadystate 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.
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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
DEPARTMENT OF TRANSPORTATION
Federal Aviation Administration
14 CFR Parts 91, 125, 135, 137, and 145
[Docket No.: FAA–2024–0025; Notice No.
24–08A]
RIN 2120–AL20
Inspection Programs for Single-Engine
Turbine-Powered Airplanes and
Unmanned Aircraft; and Miscellaneous
Maintenance-Related Updates;
Correction
Federal Aviation
Administration (FAA), Department of
Transportation (DOT).
ACTION: Notice of proposed rulemaking
(NPRM); correction; extension of
comment period.
AGENCY:
The FAA is correcting an
NPRM published on January 31, 2024.
In that document, the FAA proposed to
amend its regulations to revise certain
aircraft maintenance inspection rules for
small, corporate-sized, and unmanned
aircraft. This document corrects errors
in the preamble of that document.
DATES: The comment period for the
proposed rule published January 31,
2024, at 89 FR 6056, is extended. The
comment period originally scheduled to
close on April 1, 2024, is extended to
close on May 1, 2024.
FOR FURTHER INFORMATION CONTACT: For
technical questions concerning this
action, contact Bryan B. Davis, Airmen
& Special Projects Branch, AFS–320,
Aircraft Maintenance Division, Flight
Standards Service, Federal Aviation
Administration, 800 Independence
Avenue SW, Washington, DC 20591;
telephone (202) 267–1675; email
Bryan.Davis@faa.gov.
SUPPLEMENTARY INFORMATION:
SUMMARY:
Background
On January 31, 2024, the FAA
published an NPRM titled, ‘‘Inspection
Programs for Single-Engine TurbinePowered Airplanes and Unmanned
Aircraft; and Miscellaneous
Maintenance-Related Updates’’ (89 FR
6056).
In that NPRM, the FAA proposed to
amend its regulations to revise certain
E:\FR\FM\20MRP1.SGM
20MRP1
]]>Agencies
[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]
========================================================================
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.
========================================================================
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.
-----------------------------------------------------------------------
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.
---------------------------------------------------------------------------
\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.
---------------------------------------------------------------------------
\2\ Sometimes the entire system is referred to as an inverter.
Throughout this document, it is referred to as the controller.
---------------------------------------------------------------------------
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.
---------------------------------------------------------------------------
\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
