Rule2024-29490
Special Conditions: BETA Technologies Inc. Model H500A Electric Engines
Primary source
Metadata and text below are from the Federal Register, a public-domain U.S. government work. Always verify the official published version before relying on it for any legal matter.
Published
December 17, 2024
Effective
January 16, 2025
Issuing agencies
Transportation DepartmentFederal Aviation Administration
Abstract
These special conditions are issued for BETA Technologies Inc. (BETA) Model H500A electric engines that operate using electrical technology installed on the aircraft, for use as an aircraft engine. These engines will have a novel or unusual design feature when compared to the state of technology envisioned in the airworthiness standards applicable to aircraft engines. This design feature is the use of an electric motor, motor controller, and high-voltage systems as the primary source of propulsion for an aircraft. The applicable airworthiness regulations do not contain adequate or appropriate safety standards for this design feature. These special conditions contain the additional safety standards that the Administrator considers necessary to establish a level of safety equivalent to that established by the existing airworthiness standards.
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<title>Federal Register, Volume 89 Issue 242 (Tuesday, December 17, 2024)</title>
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[Federal Register Volume 89, Number 242 (Tuesday, December 17, 2024)]
[Rules and Regulations]
[Pages 101854-101870]
From the Federal Register Online via the Government Publishing Office [<a href="http://www.gpo.gov">www.gpo.gov</a>]
[FR Doc No: 2024-29490]
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DEPARTMENT OF TRANSPORTATION
Federal Aviation Administration
14 CFR Part 33
[Docket No. FAA-2022-1641; Special Conditions No. 33-028-SC]
Special Conditions: BETA Technologies Inc. Model H500A Electric
Engines
AGENCY: Federal Aviation Administration (FAA), DOT.
ACTION: Final special conditions.
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SUMMARY: These special conditions are issued for BETA Technologies Inc.
(BETA) Model H500A electric engines that operate using electrical
technology installed on the aircraft, for use as an aircraft engine.
These engines will have a novel or unusual design feature when compared
to the state of technology envisioned in the airworthiness standards
applicable to aircraft engines. This design feature is the use of an
electric motor, motor controller, and high-voltage systems as the
primary source of propulsion for an aircraft. The applicable
airworthiness regulations do not contain adequate or appropriate safety
standards for this design feature. These special conditions contain the
additional safety standards that the Administrator considers necessary
to establish a level of safety equivalent to that established by the
existing airworthiness standards.
DATES: Effective January 16, 2025.
FOR FURTHER INFORMATION CONTACT: Mark Bouyer, Engine and Propulsion
Standards Section, AIR-625, Technical Policy Branch, Policy and
Standards Division, Aircraft Certification Service, 1200 District
Avenue, Burlington, Massachusetts 01803; telephone (781) 238-7755;
<a href="/cdn-cgi/l/email-protection#94f9f5e6ffbaf6fbe1edf1e6d4f2f5f5baf3fbe2"><span class="__cf_email__" data-cfemail="dfb2beadb4f1bdb0aaa6baad9fb9bebef1b8b0a9">[email protected]</span></a>.
SUPPLEMENTARY INFORMATION:
Background
On January 27, 2022, BETA applied for a type certificate for its
Model H500A electric engines. The BETA Model H500A electric engine
initially will be used as a ``pusher'' electric engine in a single-
engine airplane that will be certified separately from the engine. A
typical normal category general aviation aircraft locates the engine at
the front of the fuselage. In this configuration, the propeller
attached to the engine pulls the airplane along its flightpath. A
pusher engine is located at the rear of the fuselage, so the propeller
attached to the engine pushes the aircraft instead of pulling the
aircraft.
The BETA Model H500A 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. The magnets are arranged
in a Halbach magnet array, and the stator is a concentrated, tooth-
wound configuration. A stator is the stationary component in the
electric engine that surrounds the rotating hardware; for example: the
BETA propeller shaft, which consists of a bonded core with coils of
insulated wire, known as the windings. When alternating current is
applied to the coils of insulated wire in a stator, a rotating magnetic
field is created, which provides the motive force for the rotating
components.
[[Page 101855]]
Type Certification Basis
Under the provisions of 14 CFR 21.17(a)(1), generally, BETA must
show that Model H500A electric engines meet the applicable provisions
of 14 CFR part 33 in effect on the date of application for a type
certificate.
If the Administrator finds that the applicable airworthiness
regulations (e.g., part 33) do not contain adequate or appropriate
safety standards for the BETA Model H500A electric engines because of a
novel or unusual design feature, special conditions may be prescribed
under the provisions of Sec. 21.16.
Special conditions are initially applicable to the model for which
they are issued. Should the type certificate for that model be amended
later to include any other engine model that incorporates the same
novel or unusual design feature, these special conditions would also
apply to the other engine model under Sec. 21.101.
The FAA issues special conditions, as defined in Sec. 11.19, in
accordance with Sec. 11.38, and they become part of the type
certification basis under Sec. 21.17(a)(2).
Novel or Unusual Design Features
The BETA Model H500A electric engines will incorporate the
following novel or unusual design features:
An electric motor, motor controller, and high-voltage electrical
systems that are used as the primary source of propulsion for an
aircraft.
Discussion
Electric propulsion technology is substantially different from the
technology used in previously certificated turbine and reciprocating
engines. Therefore, these engines introduce new safety concerns that
need to be addressed in the certification basis.
BETA's Electric Engines Are Novel or Unusual
The BETA Model H500A electric 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. Therefore, part 33 does not contain adequate or
appropriate safety standards for the BETA Model H500A electric engine's
novel or unusual design feature.
BETA's aircraft engines will operate using electrical power instead
of air and fuel combustion to propel the aircraft. These electric
engines will be designed, manufactured, and controlled differently than
turbine or reciprocating aircraft engines. They will be built with an
electric motor, motor controller, and high-voltage electrical systems
that draw energy from electrical storage or electrical energy
generating systems. The electric motor is a device that converts
electrical energy into mechanical energy by electric current flowing
through windings (wire coils) in the motor, producing a magnetic field
that interacts with permanent magnets mounted on the engine's main
rotor. The controller is a system that consists of two main functional
elements: the motor controller and an electric power inverter to drive
the motor.\1\ The high-voltage electrical system is a combination of
wires and connectors that integrate the motor and controller.
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\1\ Sometimes the entire system is referred to as an inverter.
Throughout this document, it is referred to as the controller.
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In addition, the technology comprising these high-voltage and high-
current electronic components introduces potential hazards that do not
exist in turbine and reciprocating aircraft engines. For example, high-
voltage transmission lines, electromagnetic shields, magnetic
materials, and high-speed electrical switches are necessary to use the
physical properties of an electric engine for propelling an aircraft.
BETA's Electric Engines Require a Mix of Part 33 Standards and Special
Conditions
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 BETA Model
H500A, which operates using electrical technology as the primary means
of propelling the aircraft.
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.
The FAA proposed special conditions and received comments from many
commenters. Some comments resulted in changes to the special
conditions. These changes are explained in the Discussion of Comments.
FAA Special Conditions for the BETA Engine Design
Applicability: Special condition no. 1 requires BETA to comply with
part 33, except for those airworthiness standards specifically and
explicitly applicable only to reciprocating and turbine aircraft
engines.
Engine Ratings and Operating Limitations: Special condition no. 2,
in addition to compliance with Sec. 33.7(a), requires BETA to
establish engine operating limits related to the power, torque, speed,
and duty cycles specific to BETA Model H500A electric engines. The duty
or duty cycle is a statement of the load(s) to which the engine is
subjected, including, if applicable, starting, no-load and rest, and
de-energized periods, including their durations or cycles and sequence
in time. This special condition also requires BETA to declare cooling
fluid grade or specification, power supply requirements, and to
establish any additional ratings that are necessary to define the BETA
Model H500A electric engine capabilities required for safe operation of
the engine.
Materials: Special condition no. 3 requires BETA to comply with
Sec. 33.15, which sets requirements for the suitability and durability
of materials used in the engine, and which would otherwise be
applicable only to reciprocating and turbine aircraft engines.
Fire Protection: Special condition no. 4 requires BETA to comply
with Sec. 33.17, which sets requirements to protect the engine and
certain parts and components of the airplane against fire, and which
would otherwise be applicable only to reciprocating and turbine
aircraft engines. Additionally, this special condition requires BETA to
ensure that the high-voltage electrical wiring interconnect systems
that connect the controller to the motor are protected against arc
faults. An arc fault is a high-power discharge of electricity between
two or more conductors. This discharge generates heat, which can break
down the wire's insulation and trigger an electrical fire. Arc faults
can range in power from a few amps up to thousands of amps and are
highly variable in strength and duration.
Durability: Special condition no. 5 requires the design and
construction of BETA Model H500A electric engines to minimize the
development of an unsafe condition between maintenance intervals,
overhaul periods, and mandatory actions described in the
[[Page 101856]]
Instructions for Continued Airworthiness (ICA).
Engine Cooling: Special condition no. 6 requires BETA to comply
with Sec. 33.21, which requires the engine design and construction to
provide necessary cooling, and which would otherwise be applicable only
to reciprocating and turbine aircraft engines. Additionally, this
special condition requires BETA to document the cooling system
monitoring features and usage in the engine installation manual (see
Sec. 33.5) if cooling is required to satisfy the safety analysis
described in special condition no. 17. Loss of cooling to an aircraft
engine that operates using electrical technology can result in rapid
overheating and abrupt engine failure, with critical consequences to
safety.
Engine Mounting Attachments and Structure: Special condition no. 7
requires BETA and the design to comply with Sec. 33.23, which requires
the applicant to define, and the design to withstand, certain load
limits for the engine mounting attachments and related engine
structure. These requirements would otherwise be applicable only to
reciprocating and turbine aircraft engines.
Accessory Attachments: Special condition no. 8 requires the design
to comply with Sec. 33.25, which sets certain design, operational, and
maintenance requirements for the engine's accessory drive and mounting
attachments, and which would otherwise be applicable only to
reciprocating and turbine aircraft engines.
Rotor Overspeed: Special condition no. 9 requires BETA to establish
by test, validated analysis, or a combination of both, that--
(1) the rotor overspeed must not result in a burst, rotor growth,
or damage that results in a hazardous engine effect;
(2) rotors must possess sufficient strength margin to prevent
burst; and
(3) operating limits must not be exceeded in service.
The special condition associated with rotor overspeed is necessary
because of the differences between turbine engine technology and the
technology of these electric engines. Turbine rotor speed is driven by
expanding gas and aerodynamic loads on rotor blades. Therefore, the
rotor speed or overspeed results from interactions between
thermodynamic and aerodynamic engine properties. The speed of an
electric engine is directly controlled by electric current, and an
electromagnetic field created by the controller. Consequently, electric
engine rotor response to power demand and overspeed-protection systems
is quicker and more precise. Also, the failure modes that can lead to
overspeed between turbine engines and electric engines are vastly
different, and therefore this special condition is necessary.
Engine Control Systems: Special condition no. 10(b) requires BETA
to ensure that these engines do not experience any unacceptable
operating characteristics, such as unstable speed or torque control, or
exceed any of their operating limitations.
The FAA originally issued Sec. 33.28 at amendment 33-15 to address
the evolution of the means of controlling the fuel supplied to the
engine, from carburetors and hydro-mechanical controls to electronic
control systems. These electronic control systems grew in complexity
over the years, and as a result, the FAA amended Sec. 33.28 at
amendment 33-26 to address these increasing complexities. The
controller that forms the controlling system for these electric engines
is significantly simpler than the complex control systems used in
modern turbine engines. The current regulations for engine control are
inappropriate for electric engine control systems; therefore, special
condition no. 10(b) associated with controlling these engines is
necessary.
Special condition no. 10(c) requires BETA 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,\2\ distinguishes between complex and simple electronic
hardware.
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\2\ <a href="https://standards.rtca.org/XanHrK">https://standards.rtca.org/XanHrK</a>.
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Special condition no. 10(d) requires data from assessments of all
functional aspects of the control system to prevent errors that could
exist in software programs that are not readily observable by
inspection of the code. Also, BETA must use methods that will result in
the expected quality that ensures the engine control system performs
the intended functions throughout the declared operational envelope.
The environmental limits referred to in special condition no. 10(e)
include temperature, vibration, high-intensity radiated fields (HIRF),
and all others addressed in RTCA DO-160G, ``Environmental Conditions
and Test Procedures for Airborne Electronic/Electrical Equipment and
Instruments,'' dated December 8, 2010, which includes RTCA DO-160G,
Change 1--``Environmental Conditions and Test Procedures for Airborne
Equipment,'' dated December, 16, 2014, and DO-357, ``User Guide:
Supplement to DO-160G,'' dated December 16, 2014.\3\ Special condition
10(e) requires BETA to demonstrate by system or component tests in
special condition no. 27 any environmental limits that cannot be
adequately substantiated by the endurance demonstration, validated
analysis, or a combination thereof.
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\3\ <a href="https://my.rtca.org/NC__Product?id=a1B36000001IcnSEAS">https://my.rtca.org/NC__Product?id=a1B36000001IcnSEAS</a>.
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Special condition no. 10(f) requires BETA to evaluate various
control system failures to ensure that such failures will not lead to
unsafe engine conditions. The FAA issued Advisory Circular (AC) 33.28-
3, ``Guidance Material for 14 CFR 33.28, Engine Control Systems,'' on
May 23, 2014 (AC 33.28-3), for reciprocating and turbine engines.\4\
This AC provides guidance for defining an engine control system failure
when showing compliance with the requirements of Sec. 33.28. AC 33.28-
3 also includes objectives for control system integrity requirements,
criteria for a loss of thrust control (LOTC) and loss of power control
(LOPC) event, and an acceptable LOTC/LOPC rate. The electrical and
electronic failures and failure rates did not account for electric
engines when the FAA issued this AC, and therefore performance-based
special conditions are established to allow fault accommodation
criteria to be developed for electric engines.
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\4\ <a href="https://www.faa.gov/documentLibrary/media/Advisory_Circular/AC_33_28-3.pdf">https://www.faa.gov/documentLibrary/media/Advisory_Circular/AC_33_28-3.pdf</a>.
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The phrase ``in the full-up configuration'' used in special
condition no. 10(f)(2) refers to a system without any fault conditions
present. The electronic control system must, when in the full-up
configuration, be single-fault tolerant, as determined by the
Administrator, for electrical, electrically detectable, and electronic
failures involving LOPC events.
The term ``local'' in the context of ``local events'' used in
special condition no. 10(f)(4) means failures or malfunctions leading
to events in the intended aircraft installation such as fire, overheat,
or failures leading to damage to engine control system components.
These ``local events'' must not result in a hazardous engine effect due
to engine control system failures or malfunctions.
Special condition no. 10(g) requires BETA to conduct a safety
assessment of the control system to support the safety analysis in
special condition no. 17. This control system safety assessment
[[Page 101857]]
provides engine response to failures, and rates of these failures that
can be used at the aircraft-level safety assessment.
Special condition no. 10(h) requires BETA to provide appropriate
protection devices or systems to ensure that engine operating limits
will not be exceeded in service.
Special condition no. 10(i) is necessary to ensure that the
controllers are self-sufficient and isolated from other aircraft
systems. The aircraft-supplied data supports the analysis at the
aircraft level to protect the aircraft from common mode failures that
could lead to major propulsion power loss. The exception ``other than
power command signals from the aircraft,'' noted in special condition
no. 10(i), is based on the FAA's determination that the engine
controller has no reasonable means to determine the validity of any in-
range signals from the electrical power system. In many cases, the
engine control system can detect a faulty signal from the aircraft, but
the engine control system typically accepts the power command signal as
a valid value.
The term ``independent'' in the context of ``fully independent
engine systems'' referenced in special condition no. 10(i) means the
controllers should be self-sufficient and isolated from other aircraft
systems or provide redundancy that enables the engine control system to
accommodate aircraft data system failures. In the case of loss,
interruption, or corruption of aircraft-supplied data, the engine must
continue to function in a safe and acceptable manner without hazardous
engine effects.
The term ``accommodated,'' in the context of ``detected and
accommodated,'' referenced in special condition 10(i)(2) is to assure
that, upon detecting a fault, the system continues to function safely.
Special condition no. 10(j) requires BETA to show that the loss of
electric power from the aircraft will not cause the electric engine to
malfunction in a manner hazardous to the aircraft. The total loss of
electric power to the electric engine may result in an engine shutdown.
Instrument Connection: Special condition no. 11 requires BETA to
comply with Sec. 33.29(a), (e), and (g), which set certain
requirements for the connection and installation of instruments to
monitor engine performance. The remaining requirements in Sec. 33.29
apply only to technologies used in reciprocating and turbine aircraft
engines.
Instrument connections (wires, wire insulation, potting, grounding,
connector designs, etc.) must not introduce unsafe features or
characteristics to the aircraft. Special condition no. 11 requires the
safety analysis to include potential hazardous effects from failures of
instrument connections to function properly. The outcome of this
analysis might identify the need for design enhancements or additional
ICA to ensure safety.
Stress Analysis: Section 33.62 requires applicants to perform a
stress analysis on each turbine engine. This regulation is explicitly
applicable only to turbine engines and turbine engine components, and
it is not appropriate for the BETA Model H500A electric engines.
However, a stress analysis particular to these electric engines is
necessary to account for stresses resulting from electric technology
used in the engine.
Special condition no. 12 requires a mechanical, thermal, and
electrical stress analysis to show that the engine has a sufficient
design margin to prevent unacceptable operating characteristics. Also,
the applicant must determine the maximum stresses in the engine by
tests, validated analysis, or a combination thereof, and show that they
do not exceed minimum material properties.
Critical and Life-Limited Parts: Special condition no. 13 requires
BETA to show whether rotating or moving components, bearings, shafts,
static parts, and non-redundant mount components should be classified,
designed, manufactured, and managed throughout their service life as
critical or life-limited parts.
The term ``low-cycle fatigue,'' referenced in special condition no.
13(a)(2), is a decline in material strength from exposure to cyclic
stress at levels beyond the stress threshold the material can sustain
indefinitely. This threshold is known as the ``material endurance
limit.'' Low-cycle fatigue typically causes a part to sustain plastic
or permanent deformation during the cyclic loading and can lead to
cracks, crack growth, and fracture. Engine parts that operate at high
temperatures and high mechanical stresses simultaneously can experience
low-cycle fatigue coupled with creep. Creep is the tendency of a
metallic material to permanently move or deform when it is exposed to
the extreme thermal conditions created by hot combustion gasses, and
substantial physical loads such as high rotational speeds and maximum
thrust. Conversely, high-cycle fatigue is caused by elastic
deformation, small strains caused by alternating stress, and a much
higher number of load cycles compared to the number of cycles that
cause low-cycle fatigue.
The engineering plan referenced in special condition no. 13(b)(1)
informs the manufacturing and service management processes of essential
information that ensures the life limit of a part is valid. The
engineering plan provides methods for verifying the characteristics and
qualities assumed in the design data using methods that are suitable
for the part criticality. The engineering plan informs the
manufacturing process of the attributes that affect the life of the
part. The engineering plan, manufacturing plan, and service management
plan are related in that assumptions made in the engineering plan are
linked to how a part is manufactured and how that part is maintained in
service. For example, environmental effects on life limited electric
engine parts, such as humidity, might not be consistent with the
assumptions used to design the part. BETA must ensure that the
engineering plan is complete, available, and acceptable to the
Administrator.
The term ``manufacturing plan,'' referenced in special condition
no. 13(b)(2), is the collection of data required to translate
documented engineering design criteria into physical parts, and to
verify that the parts comply with the properties established by the
design data. Because engines are not intentionally tested to failure
during a certification program, documents and processes used to execute
production and quality systems required by Sec. 21.137 guarantee
inherent expectations for performance and durability. These systems
limit the potential manufacturing outcomes to parts that are
consistently produced within design constraints.
The manufacturing plan and service management plan ensure that
essential information from the engineering plan, such as the design
characteristics that safeguard the integrity of critical and life-
limited parts, is consistently produced and preserved over the lifetime
of those parts. The manufacturing plan includes special processes and
production controls to prevent inclusion of manufacturing-induced
anomalies, which can degrade the part's structural integrity. Examples
of manufacturing-induced anomalies are material contamination,
unacceptable grain growth, heat-affected areas, and residual stresses.
The service-management plan ensures the method and assumptions used
in the engineering plan to determine the part's life remain valid by
enabling corrections identified from in-service experience, such as
service-induced anomalies and
[[Page 101858]]
unforeseen environmental effects, to be incorporated into the design
process. The service-management plan also becomes the ICA for
maintenance, overhaul, and repairs of the part.
Lubrication System: Special condition no. 14 requires BETA to
ensure that the lubrication system is designed to function properly
between scheduled maintenance intervals and to prevent contamination of
the engine bearings. This special condition also requires BETA to
demonstrate the unique lubrication attributes and functional capability
of the BETA Model H500A electric engine design.
The corresponding part 33 regulations include provisions for
lubrication systems used in reciprocating and turbine engines. The part
33 requirements account for safety issues associated with specific
reciprocating and turbine engine system configurations. These
regulations are not appropriate for the BETA Model H500A electric
engines. For example, electric engines do not have a crankcase or
lubrication oil sump. Electric engine bearings are sealed, so they do
not require an oil circulation system. The lubrication system in these
engines is also independent of the propeller pitch control system.
Therefore, special condition no. 14 incorporates only certain
requirements from the part 33 regulations.
Power Response: Special condition no. 15 requires the design and
construction of the BETA Model H500A electric engines to enable an
increase from the minimum--
(1) power setting to the highest rated power without detrimental
engine effects, and
(2) within a time interval appropriate for the intended aircraft
application.
The engine control system governs the increase or decrease in power
in combustion engines to prevent too much (or too little) fuel from
being mixed with air before combustion. Due to the lag in rotor
response time, improper fuel/air mixtures can result in engine surges,
stalls, and exceedances above rated limits and durations. Failure of
the combustion engine to provide thrust, maintain rotor speeds below
rotor burst thresholds, and keep temperatures below limits can have
engine effects detrimental to the aircraft. Similar detrimental effects
are possible in the BETA Model H500A electric engines, but the causes
are different. Electric engines with reduced power response time can
experience insufficient thrust to the aircraft, shaft over-torque, and
over-stressed rotating components, propellers, and critical propeller
parts. Therefore, this special condition is necessary.
Continued Rotation: Special condition no. 16 requires BETA to
design the Model H500A electric engines such that, if the main rotating
systems continue to rotate after the engine is shut down while in-
flight, this continued rotation will not result in any hazardous engine
effects.
The main rotating system of the BETA Model H500A electric engines
consists of the rotors, shafts, magnets, bearings, and wire windings
that convert electrical energy to shaft torque. For the initial
aircraft application, this rotating system must continue to rotate
after the power source to the engine is shut down. The safety concerns
associated with this special condition are substantial asymmetric
aerodynamic drag that can cause aircraft instability, loss of control,
and reduced efficiency; and may result in a forced landing or inability
to continue safe flight.
Safety Analysis: Special condition no. 17 requires BETA to comply
with Sec. 33.75(a)(1) and (a)(2), which require the applicant to
conduct a safety analysis of the engine, and which would otherwise be
applicable only to turbine aircraft engines. Additionally, this special
condition requires BETA to assess its engine design to determine the
likely consequences of failures that can reasonably be expected to
occur. The failure of such elements, and associated prescribed
integrity requirements, must be stated in the safety analysis.
A primary failure mode is the manner in which a part is most likely
going to fail. Engine parts that have a primary failure mode, a
predictable life to the failure, and a failure consequence that results
in a hazardous effect, are life-limited or critical parts. Some life-
limited or critical engine parts can fail suddenly in their primary
failure mode, from prolonged exposure to normal engine environments
such as temperature, vibration, and stress, if those engine parts are
not removed from service before the damage mechanisms progress to a
failure. Due to the consequence of failure, these parts are not allowed
to be managed by on-condition or probabilistic means because the
probability of failure cannot be sensibly estimated in numerical terms.
Therefore, the parts are managed by compliance with integrity
requirements, such as mandatory maintenance (life limits, inspections,
inspection techniques), to ensure the qualities, features, and other
attributes that prevent the part from failing in its primary failure
mode are preserved throughout its service life. For example, if the
number of engine cycles to failure are predictable and can be
associated with specific design characteristics, such as material
properties, then the applicant can manage the engine part with life
limits.
Complete or total power loss is not assumed to be a minor engine
event, as it is in the turbine engine regulation Sec. 33.75, to
account for experience data showing a potential for higher hazard
levels from power loss events in single-engine general aviation
aircraft. The criteria in these special conditions apply to an engine
that continues to operate at partial power after a single electrical or
electronic fault or failure. Total loss of power is classified at the
aircraft level using special condition nos. 10(g) and 33(h).
Ingestion: Special condition no. 18 requires BETA 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 BETA electric engines are not configured with inlet ducts.
An ``unacceptable'' power loss, as used in special condition no.
18(b), is such that the power or thrust required for safe flight of the
aircraft becomes unavailable to the pilot. The specific amount of power
loss that is required for safe flight depends on the aircraft
configuration, speed, altitude, attitude, atmospheric conditions, phase
of flight, and other circumstances where the demand for thrust is
critical to safe operation of the aircraft.
Liquid and Gas Systems: Special condition no. 19 requires BETA to
ensure that systems used for lubrication or cooling of engine
components are designed and constructed to function properly. Also, if
a system is not self-contained, the interfaces to that system would be
required to be defined in the engine installation manual. Systems for
the lubrication or cooling of engine components can include heat
exchangers, pumps, fluids, tubing, connectors, electronic devices,
temperature sensors and pressure switches, fasteners and brackets,
bypass valves, and metallic chip detectors. These systems allow the
electric engine to perform at extreme speeds and temperatures for
durations up to the maintenance intervals without exceeding temperature
limits or predicted deterioration rates.
[[Page 101859]]
Vibration Demonstration: Special condition no. 20 requires BETA to
ensure the engine--
(1) is designed and constructed to function throughout its normal
operating range of rotor speeds and engine output power without
inducing excessive stress caused by engine vibration, and
(2) design undergoes a vibration survey.
The vibration demonstration is a survey that characterizes the
vibratory attributes of the engine. It verifies that the stresses from
vibration do not impose excessive force or result in natural frequency
responses on the aircraft structure. The vibration demonstration also
ensures internal vibrations will not cause engine components to fail.
Excessive vibration force occurs at magnitudes and forcing functions or
frequencies, which may result in damage to the aircraft. Stress margins
to failure add conservatism to the highest values predicted by analysis
for additional protection from failure caused by influences beyond
those quantified in the analysis. The result of the additional design
margin is improved engine reliability that meets prescribed thresholds
based on the failure classification. The amount of margin needed to
achieve the prescribed reliability rates depends on an applicant's
experience with a product. The FAA considers the reliability rates when
deciding how much vibration is ``excessive.''
Overtorque: Special condition no. 21 requires BETA to demonstrate
that the engine is capable of continued operation without the need for
maintenance if it experiences a certain amount of overtorque.
BETA's electric engine converts electrical energy to shaft torque,
which is used for propulsion. The electric motor, controller, and high-
voltage systems control the engine torque. When the pilot commands
power or thrust, the engine responds to the command and adjusts the
shaft torque to meet the demand. During the transition from one power
or thrust setting to another, a small delay, or latency, occurs in the
engine response time. While the engine dwells in this time interval, it
can continue to apply torque until the command to change the torque is
applied by the engine control. The allowable amount of overtorque
during operation depends on the engine's response to changes in the
torque command throughout its operating range.
Calibration Assurance: Special condition no. 22 requires BETA to
subject the engine to calibration tests to establish its power
characteristics and the conditions both before and after the endurance
and durability demonstrations specified in special condition nos. 23
and 26. The calibration test requirements specified in Sec. 33.85 only
apply to the endurance test specified in Sec. 33.87, which is
applicable only to turbine engines. The FAA determined that the methods
used for accomplishing those tests for turbine engines are not
appropriate for electric engines. The calibration tests in Sec. 33.85
have provisions applicable to ratings that are not relevant to the BETA
Model H500A electric engines. Special condition no. 22 allows BETA to
demonstrate the endurance and durability of the electric engine either
together or independently, whichever is most appropriate for the engine
qualities being assessed. Consequently, the special condition applies
the calibration requirement to both the endurance and durability tests.
Endurance Demonstration: Special condition no. 23 requires BETA to
perform an endurance demonstration test that is acceptable to the
Administrator. The Administrator will evaluate the extent to which the
test exposes the engine to failures that could occur when the engine is
operated at up to its rated values, and determine if the test is
sufficient to show that the engine design will not exhibit unacceptable
effects in service, such as significant performance deterioration,
operability restrictions, and engine power loss or instability, when it
is run repetitively at rated limits and durations in conditions that
represent extreme operating environments.
Temperature Limit: Special condition no. 24 requires BETA to ensure
the engine can endure operation at its temperature limits plus an
acceptable margin. An ``acceptable margin,'' as used in the special
condition, is the amount of temperature above that required to prevent
the least capable engine allowed by the type design, as determined by
Sec. 33.8, from failing due to temperature-related causes when
operating at the most extreme engine and environmental thermal
conditions.
Operation Demonstration: Special condition no. 25 requires the
engine to demonstrate safe operating characteristics throughout its
declared flight envelope and operating range. Engine operating
characteristics define the range of functional and performance values
the BETA Model H500A electric engines can achieve without incurring
hazardous effects. The characteristics are requisite capabilities of
the type design that qualify the engine for installation into aircraft
and that determine aircraft installation requirements. The primary
engine operating characteristics are assessed by the tests and
demonstrations that would be required by these special conditions. Some
of these characteristics are shaft output torque, rotor speed, power
consumption, and engine thrust response. The engine performance data
BETA will use to certify the engine must account for installation loads
and effects. These are aircraft-level effects that could affect the
engine characteristics that are measured when the engine is tested on a
stand or in a test cell. These effects could result from elevated inlet
cowl temperatures, aircraft maneuvers, flowstream distortion, and hard
landings. For example, an engine that is run in a sea-level, static
test facility could demonstrate more capability for some operating
characteristics than it will have when operating on an aircraft in
certain flight conditions. Discoveries like this during certification
could affect engine ratings and operating limits. Therefore, the
installed performance defines the engine performance capabilities.
Durability Demonstration: Special condition no. 26 requires BETA 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, BETA will use different test effects to develop maintenance,
overhaul, or engine replacement information for their electric engine.
System and Component Tests: Special condition no. 27 requires BETA
to show that the systems and components of the engine perform their
intended functions in all declared engine environments and operating
conditions.
Sections 33.87 and 33.91, which are specifically applicable to
turbine engines, have conditional criteria to
[[Page 101860]]
decide if additional tests will be required after the engine tests. The
criteria are not suitable for electric engines. Part 33 associates the
need for additional testing with the outcome of the Sec. 33.87
endurance test because it is designed to address safety concerns in
combustion engines. For example, Sec. 33.91(b) requires the
establishment of temperature limits for components that require
temperature-controlling provisions, and Sec. 33.91(a) requires
additional testing of engine systems and components where the endurance
test does not fully expose internal systems and components to thermal
conditions that verify the desired operating limits. Exceeding
temperature limits is a safety concern for electric engines. The FAA
determined that the Sec. 33.87 endurance test is not appropriate for
testing the electronic components of electric engines because
mechanical energy is generated differently by electronic systems than
it is by the thermal conditions in turbine engines. Additional safety
considerations also need to be addressed in the test. Therefore,
special condition no. 27 is a performance-based requirement that allows
BETA to determine when engine systems and component tests are necessary
and to determine the appropriate limitations of those systems and
components used in the BETA Model H500A electric engine.
Rotor Locking Demonstration: Special condition no. 28 requires the
engine to demonstrate reliable rotor locking performance and that no
hazardous effects will occur if the engine uses a rotor locking device
to prevent shaft rotation.
Some engine designs enable the pilot to prevent a propeller shaft
or main rotor shaft from turning while the engine is running, or the
aircraft is in-flight. This capability is needed for some installations
that require the pilot to confirm the functionality of certain flight
systems before takeoff. The BETA engine installations are not limited
to aircraft that will not require rotor locking. Section 33.92
prescribes a test that may not include the appropriate criteria to
demonstrate sufficient rotor locking capability for these engines.
Therefore, this special condition is necessary.
The special condition does not define ``reliable'' rotor locking
but allows BETA to classify the hazard as major or minor and assign the
appropriate quantitative criteria that meet the safety objectives
required by special condition no. 17 and the applicable portions of
Sec. 33.75.
Teardown Inspection: Special condition no. 29 requires BETA to
perform a teardown or non-teardown evaluation after the endurance,
durability, and overtorque demonstrations, based on the criteria in
special condition no. 29(a) or (b).
Special condition no. 29(b) includes restrictive criteria for
``non-teardown evaluations'' to account for electric engines, sub-
assemblies, and components that cannot be disassembled without
destroying them. Some electrical and electronic components like BETA's
are constructed in an integrated fashion that precludes the possibility
of tearing them down without destroying them. The special condition
indicates that, if a teardown cannot be performed in a non-destructive
manner, then the inspection or replacement intervals must be
established based on the endurance and durability demonstrations. The
procedure for establishing maintenance should be agreed upon between
the applicant and the FAA prior to running the relevant tests. Data
from the endurance and durability tests may provide information that
can be used to determine maintenance intervals and life limits for
parts. However, if life limits are required, the lifing procedure is
established by special condition no. 13, Critical and Life-Limited
Parts, which corresponds to Sec. 33.70. Therefore, the procedure used
to determine which parts are life-limited, and how the life limits are
established, requires FAA approval, as it does for Sec. 33.70.
Sections 33.55 and 33.93 do not contain similar requirements because
reciprocating and turbine engines can be completely disassembled for
inspection.
Containment: Special condition no. 30 requires the engine to have
containment features that protect against likely hazards from rotating
components unless BETA can show the margin to rotor burst does not
justify the need for containment features. Rotating components in
electric engines are typically disks, shafts, bearings, seals, orbiting
magnetic components, and the assembled rotor core. However, if the
margin to rotor burst does not unconditionally rule out the possibility
of a rotor burst, then the special condition requires BETA 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, BETA 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 BETA engine
requires containment features, and the effects of any subsequent damage
following a rotor burst, should be based on test or validated analysis.
The fragment energy levels, trajectories, and size are typically
documented in the installation manual because the aircraft will need to
account for the effects of a rotor failure in the aircraft design. The
intent of this special condition is to prevent hazardous engine effects
from structural failure of rotating components and parts that are built
into the rotor assembly.
General Conduct of Tests: Special condition no. 32 requires BETA
to--
(1) Include any scheduled maintenance.
(2) Include any maintenance, in addition to the scheduled
maintenance, which 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 special condition nos. 32(b)(1) and (2),
respectively. Excessive frequency is an objective assessment from the
FAA's analysis of the amount of unplanned maintenance needed for an
engine to complete a certification test. The FAA's assessment may
include the reasons for the unplanned maintenance, such as the effects
test facility equipment may have on the engine, the inability to
simulate a realistic engine operating environment, and the extent to
which an engine requires modifications to
[[Page 101861]]
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 BETA
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. Special condition
no. 33 is established using the existing engine control systems
requirement as a basis. It applies applicable airworthiness criteria
from Sec. 33.28 and incorporates airworthiness criteria that recognize
and focus on the electrical power system used in the engine.
Special condition no. 33(b) ensures that all aspects of an
electrical system, including generation, distribution, and usage, do
not experience any unacceptable operating characteristics.
Special condition no. 33(c) requires the electrical power
distribution aspects of the electrical system to provide the safe
transfer of electrical energy throughout the electric engine.
The term ``abnormal conditions'' used in special condition no.
33(c)(2) is based on the term ``abnormal operation'' used in MIL-STD-
704F ``Aircraft Electric Power Characteristics'' which defines normal
operation and abnormal operation. MIL-STD-704F is a standard that
ensures compatibility between power sources that provide power to the
aircraft's electrical systems and airborne equipment that receive power
from the power source. This standard also establishes technical
criteria for aircraft electric power. The term ``abnormal conditions''
refers to various engine operating conditions such as:
<bullet> System or component characteristics outside of normal
statistical variation from circumstances such as systems degradation,
installation error, and engine response to fault conditions;
<bullet> Unusual environmental conditions from extreme temperature,
humidity, vibration, lightning, high-intensity radiated field (HIRF),
atmospheric neutron radiation; and
<bullet> Unusual and infrequent events such as landing on icy
runways, rejected take-offs or go-arounds, extended ground idling or
taxiing in a hot environment, and abrupt load changes from foreign
object damage or engine contamination.
The phrase ``safe transmission of electric energy'' used in special
condition no. 33(c)(3) refers to the transmission of electrical energy
in a manner that supports the operation of the electric engine(s) and
the aircraft safety objectives without detrimental effects such as
uncontrolled fire or structural failure due to severe overheating.
Special condition no. 33(d) requires the engine electrical system
to be designed such that the loss, malfunction, or interruption of the
electrical power source, or power conditions that exceed design limits,
will not result in a hazardous engine effect.
Special condition no. 33(e) requires BETA to identify and declare,
in the engine installation manual, the characteristics of any
electrical power supplied from the aircraft to the engine, or
electrical power supplied from the engine to the aircraft via energy
regeneration, and any other characteristics necessary for safe
operation of the engine.
Special condition no. 33(f) requires BETA to demonstrate that
systems and components will operate properly up to environmental
limits, using special conditions, when such limits cannot be adequately
substantiated by the endurance demonstration, validated analysis, or a
combination thereof. The environmental limits referred to in this
special condition include temperature, vibration, HIRF, and all others
addressed in RTCA DO-160G, ``Environmental Conditions and Test
Procedures for Airborne Electronic/Electrical Equipment and
Instruments.''
Special condition 33(g) requires BETA to evaluate various electric
engine system failures to ensure that these failures will not lead to
unsafe engine conditions. The evaluation includes single-fault
tolerance, ensures no single electrical or electronic fault or failure
would result in hazardous engine effects, and ensures that any failure
or malfunction leading to local events in the intended aircraft
application does not result in certain hazardous engine effects. The
special condition also implements integrity requirements, criteria for
LOTC/LOPC events, and an acceptable LOTC/LOPC rate.
Special condition 33(h) requires BETA to conduct a safety
assessment of the engine electrical system to support the safety
analysis in special condition no. 17. This safety assessment provides
engine response to failures, and rates of these failures, which can be
used at the aircraft safety assessment level.
Discussion of Comments
The FAA issued a notice of proposed special conditions (NPSC)
Docket No. FAA-2022-1641 for the BETA Model H500A electric engines,
which was published in the Federal Register on March 7, 2024 (89 FR
16474).
The FAA Received Comments From Eight Commenters
The FAA received comments from Transport Canada (TC), Transport
Canada Civil Aviation (TCCA), United Parcel Service Flight Forward
(UPSFF), Association for Uncrewed Vehicle Systems International
(AUVSI), magniX USA, Inc. (magniX), General Aviation Manufacturers
Association (GAMA), an individual, and an anonymous commenter.
The FAA received comments from TCCA.
TCCA indicated the discussion of proposed special condition
no.10(e), Environmental limits of engine cooling systems, in the
preamble states that the environmental limits referred to in this
special condition are addressed in RTCA DO-160G. However, TCCA
explained that some of the existing RTCA DO-160G test specifications,
methods, and categories may not be adequate for high voltage systems,
such as the high voltage components of this engine. Accordingly, TCCA
recommended adding the language ``or other appropriate industry
standards'' at the end of the discussion of special condition no. 10(e)
in the preamble.
The FAA does not agree with the recommended change. Although RTCA
DO-160G is not sufficient for the high voltage systems used in the BETA
Model H500A electric engine motor and inverter/controller, tests that
are appropriate for the BETA engine will be developed in accordance
with special condition nos. 1(b) and 1(c) using the testing techniques
in RTCA DO-160G and other aerospace environmental
[[Page 101862]]
documents. Independent tests are done for radiated and conducted
susceptibility and compared to the RTCA DO-160G HIRF spectrum for
susceptibility to ensure all electric engine radio frequency energy
emissions inherent to the engine design are addressed. If the equipment
under test passes the emission test in RTCA-DO-160 the susceptibility
spectrum is covered by RTCA DO-160G. The applicant can use the RTCA DO-
160G test. If not, the spectrum from the emission test would be
analyzed and could be adjusted for the applicant's design and applied
during the susceptibility test with FAA concurrence. No changes were
made to these special conditions as a result of this comment.
TCCA also indicated special condition no. 2, Engine ratings and
operating limits, should require that component life be considered when
establishing the engine operating limits. They explained, the engine
system or the electrical motor design may have components or parts that
require a life limit. For example, the insulation on the high voltage
system wiring may degrade with time and operating conditions. TCCA
requested the FAA add ``(f) Component life'' to special condition no.
2, Engine Ratings and Operating Limits, and explained that component
life should be considered when establishing the engine operating
limits, similar to Sec. 33.07(b)(7).
The FAA does not concur with TCCA's request. Component life is an
expected outcome of special conditions nos. 13 (Critical and life-
limited parts) and 17 (Safety analysis). Special condition no. 17
determines whether special condition no. 13 applies to the engine part.
Special condition no 13 determines the mandatory replacement times
(component life) and implements a maintenance program to manage these
parts composed of an engineering plan, manufacturing plan, and service
management plan. No changes were made to these special conditions as a
result of this comment.
TCCA requested the FAA confirm that special condition no. 33(a),
applicability for engine electrical systems, is not applicable to
energy storage systems (ESS) but it does include the interface between
the electric engine and the propulsion power source. TCCA further
explained this comment is a request for clarification, rather than
modification, of this special condition.
Special condition no. 33 does not apply to ESS but does apply to
the interface between the engine and ESS. No changes were made to these
special conditions as a result of this comment.
TCCA stated that proposed special condition no. 33(b), Electrical
systems, is written in a way that implies electrical load shedding is
mandatory even when not needed and explained electrical load shedding
should only be implemented if required. TCCA recommended adding ``if
required'' between parenthesis like the following: ``. . . , and
electrical load shedding (if required), . . .'' to special condition
no. 33(b).
The FAA concurs with TCCA's recommendation and has revised special
condition no. 33(b) accordingly. Load shedding is a capability of the
electric engine's power distribution system.
TCCA requested the FAA define the term ``abnormal condition,''
which is used in special condition 33(c)(2), Electrical power
distribution, and offered several potential interpretations of the
term. They also asked if an abnormal condition is any failure condition
not considered extremely improbable, and if it is equivalent to the
definition from MIL-STD-704F. The FAA's use of the term ``abnormal
conditions'' does not refer to internal malfunctions or failures. It
refers to operating conditions such as:
<bullet> System or components outside of normal statistical
variation due to degradation, or installation error
<bullet> Unusual environmental conditions such as extreme
temperature, humidity, FOD impact, severe lightning, HIRF, or
atmospheric radiation
<bullet> Infrequent scenarios such as landing on icy runways,
rejected take-offs or balked landings, extended ground idling, or
taxiing in hot environments.
TCCA also requested the FAA provide a definition for ``safe
transmission,'' which is used in special condition 33(c)(3).
The FAA concurs with TCCA's requests and has added definitions of
the terms ``abnormal condition'' and ``safe transmission'' to the
preamble discussion for special condition no. 33.
TCCA observed that proposed special condition nos. 33(e)(1) and
(e)(2), Electrical power characteristics, were linked with an ``or''
indicating that either condition could be applied, but not both. TCCA
stated both (e)(1) and (e)(2) are applicable, and therefore recommended
the FAA revise special condition no. 33(e) to replace the ``or'' with
an ``and.''
The FAA concurs with TCCA's recommendation and has revised special
condition no. 33(e) accordingly.
TCCA indicated that noise certification requirements are applicable
at the airframe level and not at the engine level. TCCA explained the
NPSC implies that an engine applicant demonstration of compliance to 14
CFR part 36 is part of the special conditions. However, TCCA stated
there is no definition of requirements within the special conditions
other than the preamble section titled the Type Certification Basis.
TCCA requested that the FAA remove the statement ``In addition to the
applicable airworthiness regulations and special conditions, the BETA
Model H500A electric engines must comply with the noise certification
requirements of 14 CFR part 36'' from the preamble. GAMA also commented
on this issue and stated the noise certification requirements do not
apply to engines and requested the FAA remove this statement from the
preamble.
The FAA concurs with TCCA's and GAMA's requests and has updated the
preamble of these special conditions accordingly.
TCCA suggested that the reference to ``consensus standards'' in
proposed special condition 1(b), Applicability, may not be necessary.
TCCA stated that consensus standards are not a means of compliance but
instead, they are derived/alternate requirements (i.e., ASTM) that are
formulated by industry to be used in lieu of published regulatory
guidance material. TCCA further suggested that the use of derived/
alternate requirements in lieu of the published standards is to be
accepted by the Administrator as being equivalent to the published
standards. Then, the means of compliance to the consensus standards are
to be accepted by the Administrator. TCCA recommended reducing the text
in special condition no. 1(b) to the following: ``(b) the applicant
must comply with this part using a means of compliance accepted by the
administrator.''
The FAA does not concur with TCCA's suggested change. The reference
to consensus standards provides clarification about potential sources
of information that may be used to determine a means of compliance. The
comment indicates a need to clarify how consensus standards are used.
For example, consensus standards developed by the standards development
organizations (SDOs) typically function as a method of compliance to 14
CFR requirements or special conditions. Published FAA guidance can
function either as a means of compliance, method of compliance, or
both. Special condition 1(b) permits consensus standards to be used for
showing compliance to certification requirements, but they are not a
[[Page 101863]]
requirement of that special condition. Therefore, special condition
1(b) supplements the performance-based special conditions by requiring
a means of compliance, which could include consensus standards
developed by SDOs. Further, special condition 1(b) is intended to be
equivalent to Sec. 23.2010(a), which also refers to consensus
standards as a potential means of compliance. No changes were made to
the special conditions as a result of this comment.
TCCA observed the BETA proposed special condition no. 17 does not
include a reference to Sec. 33.75(a)(3) which appears in the magniX
special conditions and recommended that the FAA explain this difference
in the discussion for that special condition in the preamble to avoid
ambiguity between the relative project requirements.
The FAA does not concur with TCCA's recommendation. The NPSC for
the magniX magni350 and magni650 model electric engines originally
proposed to incorporate Sec. 33.75(a)(3) into special condition no.
17. The FAA received a comment suggesting that Sec. 33.75(a)(3) may
not be needed for those engines. In the final special conditions
(Docket No. FAA-2020-0894, Special Conditions No. 33-022-SC), the FAA
agreed with the comment and removed the reference to Sec. 33.75(a)(3).
No changes were made to these special conditions as a result of this
comment.
The FAA received comments from TC.
TC disagreed with the text in proposed special condition nos. 17(a)
and 17(c) which say, ``The applicant must comply . . .'' TC stated that
the onus to show compliance with the applicable requirements with the
intent to obtain a type certificate is on the applicant and that the
elements that comply with the requirements themselves are those objects
of the type certificate, such as the engine and its systems. TC further
explained it is not clear to state that the applicant must comply,
where it is in fact the engine/systems which must comply with the
requirements. Instead, the applicant shows compliance. TC suggested
changing the phrase to read ``The applicant must show compliance . .
.''
TC's proposed change is not necessary. Section 21.20, ``Compliance
with Applicable Requirements'' contains an example that supports the
language used in special conditions nos. 17(a) and (c). Specifically,
Sec. 21.20(b) specifies the applicant must ``provide a statement
certifying that the applicant has complied with the applicable
requirements,'' which indicates the applicant complies with the
applicable requirements. . No changes were made as a result of this
comment.
TC observed the text in proposed special condition no. 17(d)(1),
Safety Analysis, does not include special condition no. 31, Operation
with Variable Pitch Propeller. TC recommended that the FAA either add a
reference to special condition no. 31 in special condition no. 17(d)(1)
because BETA's electric engine may be installed with a variable pitch
propeller or provide a rationale for not including it.
The FAA does not concur with TC's suggestion to add a reference to
special condition no. 31. Adding special condition no. 31 is not
necessary because the specific engine model BETA intends to certify is
not designed to use a variable pitch propeller. No changes were made to
the special conditions as a result of this comment.
TC indicated there is a similar electrical engine special condition
in the magniX special conditions (Special Conditions No. 33-022-SC)
that contains an ingestion requirement that does not appear in the BETA
special conditions. TC referred to special condition no. 18(d) in the
magniX special conditions, which requires ingestion sources that are
not evaluated must be declared in the engine installation manual. TC
recommended that the FAA either revise the BETA special conditions to
add this requirement or provide the rationale for not including it.
The FAA does not concur with TC's request to revise the BETA
special conditions to include special condition no. 18(d) from the
magniX special conditions. Special condition no. 18(d) was intended to
ensure ingestion sources that are not applicable to an electric engine
are enunciated in the engine documentation. The list of required
ingestion sources in BETA special condition nos. 18(a) and (b) are more
prescriptive compared to the ingestion requirements in the published
magniX special condition no. 18(a). Therefore, the FAA has determined
special condition no. 18(d) is not necessary to include in the BETA
special conditions because exceptions to the ingestion requirement
would be specified and managed using special condition no. 18(c), which
is similar to how exceptions are managed by the existing part 33
ingestion requirements. No changes were made to the special conditions
as a result of this comment.
TC noted that proposed special condition no. 33(c)(1) introduces
the term ``electrical power plant'' and recommended that the FAA update
the preamble to describe an electrical power plant.
The FAA disagrees with TC's recommendation to define ``electrical
power plant'' because the FAA revised special condition no. 33(c)(1) in
these final special conditions to change the term ``electrical power
plant'' to ``powerplant,'' as that term is defined in part 23, subpart
E, in Sec. 23.2400(a) powerplant installation, to include each
component necessary for propulsion, which affects propulsion safety, or
provides auxiliary power to the airplane, and in the installation
requirements in subpart E of parts 25, 27, and 29.
TC observed that the proposed system safety assessments in proposed
special condition no. 33(h), and proposed special condition no.10(g)
are different in that special condition no. 10(g) requires the rates of
hazardous and major faults to be declared in the engine installation
manual and special condition no. 33(h) does not. TC recommended that
the FAA either revise special condition no. 33(h) to match special
condition no. 10(g) or provide a rationale for why they are different.
The FAA agrees with TC's recommendation and has revised final
special condition no. 33(h) to match special condition no. 10(g).
The FAA received comments from GAMA.
GAMA recommended that the FAA align the special conditions for the
H500A electric engine with the electric engine requirements included in
the certification basis for special class powered lift aircraft that
certify an electric engine as part of the aircraft type certification.
GAMA stated that there are technical variations between the H500A
proposed special conditions and the electric engine airworthiness
criteria outlined in the Special Class Airworthiness Criteria for the
powered-lift and cited special condition no. 17(c) and special
condition no. 33(c) as examples of these technical differences. GAMA
further stated these variations could lead to two electric engines used
in the same aircraft having different requirements based solely on
whether the engine is certified as part of the aircraft or under part
33. AUVSI also commented on the importance of applying consistent
requirements across projects and requested the FAA substantiate any
inconsistencies introduced to the electric engine requirements.
There are no intended technical differences between the proposed
special class airworthiness criteria for the powered lift in draft
Advisory
[[Page 101864]]
Circular 21.17-4 (AC 21.17-4) and the BETA special conditions. For
example, the corresponding criteria to BETA special condition nos.
17(c) and 33(c) are PL.3375(f) and PL.3326(c) respectively. The engine
requirements are documented differently between the BETA special
conditions and powered-lift airworthiness criteria proposed in draft AC
21.17-4 because special conditions are written in accordance with the
requirements of Sec. 21.16, and the powered-lift airworthiness
criteria in draft AC 21.17-4 are not specific to one applicant. There
are also some minor differences in the documentation requirements
because engines are approved with the special class aircraft, so some
engine details may be included in the aircraft manuals. No changes were
made to the special conditions as a result of this comment.
GAMA indicated proposed special condition no. 9, Overspeed, lacks
clarity regarding whether ``rotor'' refers to an internal electric
engine component or an actual propulsive propeller. GAMA recommended
the FAA provide the necessary clarification to address this ambiguity.
The FAA agrees with GAMA's recommendation. The term ``rotor'' in
the proposed special conditions is intended to refer to an engine
component and not a propulsive propeller. A rotor in an electric engine
may consist of a circular disk and magnets fixed at the outer
circumference that rotates inside a stationary casing configured with
electrical windings (or coils), or a rotating cylindrical casing with
magnets fixed on the inside surface that rotates around a stationary
set of windings (or coils). Each configuration is attached to a
rotating shaft that drives a propulsive device, such as a propeller.
Project-specific decisions will be made regarding which engine parts
are applicable to the overspeed requirement. No changes were made to
the special conditions as a result of this comment.
GAMA stated that proposed special condition nos. 30(a) and (b),
Containment, utilize language tailored to an engine design featuring a
non-rotating stator situated outside the rotor. GAMA recommended the
FAA explore a rule version that is less design-specific. GAMA advised
against presuming that all rotating components possess a case,
particularly that the rotor is contained within the stator.
The FAA does not concur with GAMA's recommendation Special
condition 30(a) is intended to account for rotor designs with
exceedingly large margins to a rotor burst. The special condition does
not specify a particular rotor design. However, the amount of margin
needed to satisfy the requirement would be determined based on the
engine's design. Special condition 30(b) is intended to account for
rotors located inside a static stator case. No changes were made to the
special conditions as a result of this comment.
GAMA commented proposed special condition nos. 33(c)(1) and (c)(3),
Electrical power distribution for engine electrical systems, set forth
distinct criteria for the automatic measures needed when electrical-
energy generation encounters faults, which diverges from the
corresponding requirements in the special class airworthiness criteria
for powered-lift. GAMA indicated there are no evident variations in
electric engine configurations that warrant this inconsistency. GAMA
recommended that the FAA align these regulations to ensure that
electric engines certified as part of an aircraft or under part 33
adhere to uniform standards.
Proposed special condition nos. 33(c)(1) and (c)(3) are not the
same as the corresponding engine requirements in the powered-lift
airworthiness criteria used in another project. Proposed special
condition no. 33(c)(1) protects engine electrical systems from faulted
electrical energy generation or storage devices. Proposed special
condition no. 33(c)(3) prescribes a means of compliance (fault
isolation) to address (c)(1), but the means of compliance should be
tied to the safety assessment required in special condition no. 33(g),
which accounts for aircraft-level effects from faulted electrical-
energy generation or storage devices. The aircraft effects should not
be assumed in the engine requirements, and therefore the FAA revised
special condition no. 33(c)(3) to accommodate other potential
protection systems that might be more appropriate. Accordingly, final
special condition no. 33(c)(3) is changed to, ``The system must provide
mechanical or automatic means of isolating a faulted electrical energy
generation or storage device from leading to hazardous engine effects,
as defined in special condition no. 17(d)(2) of these special
conditions, or detrimental effects in the intended aircraft
application.''
The phrase, ``or detrimental engine effects in the intended
aircraft application'' was relocated to special condition no. 33(c)(3)
to maintain the connection with special condition no. 33(g).
GAMA commented proposed special condition no. 33(g), Electrical
system failures of engine electrical systems, extends beyond the
comparable part 33 regulation Sec. 33.28(d), which is originally
limited to the engine control system. GAMA suggested that expanding
this special condition to encompass the engine electrical system
instead of solely the engine control system entails subjecting
electrical components within the engine, such as windings, to failure
requirements historically not applied to engine mechanical components.
GAMA also stated that field experience indicates that component
failures are unpredictable based on wear and susceptible to random
failures. Electric engine components, like windings and insulation, are
better addressed using methods akin to those applied to traditional
engines to address mechanical failure aspects. GAMA recommended the FAA
revise this special condition to align with the existing regulatory
framework. The FAA does not concur with GAMA's recommendation. By their
nature, FAA special conditions are issued when the ``existing
regulatory framework'' is inadequate or insufficient. 14 CFR 21.16; see
also Amdt. 21-51. The existing requirements for engine control systems
were developed to address the failure characteristics of electrical
systems. For combustion engines, the only electrical system is the
engine control, but this is not the case for electric engines where
electrical systems extend beyond those addressed by Sec. 33.28(d).
Special condition no. 33(g) for the BETA electric engine provides the
same level of safety as Sec. 33.28(d) by applying the safety criteria
for electrical systems to all the electrical systems in the engine.
This includes the high-voltage systems used in the electric engine. No
changes were made to the special conditions as a result of this
comment.
The FAA received several comments from an individual commenter and
received similar comments from magniX (although these commenters
provided separate comments).
An individual and magniX commented proposed special condition nos.
1(b) and (c) state that a means of compliance, which may include
consensus standards, must be ``accepted by the Administrator'' and ``in
a form and manner acceptable to the Administrator.'' The individual and
magniX stated that these paragraphs are directly out of Sec. 23.2010,
which contains performance-based language. The individual and magniX
considered the BETA electric engine special conditions to be largely
prescriptive and not performance-based, which they stated would make
special condition
[[Page 101865]]
nos. 1(b) and (c) superfluous. The individual suggested these
requirements introduce a new regulatory layer to prescriptive
requirements and may lead to inadvertent consequences, while magniX
stated that requiring a performance-based process for establishing
means of compliance with prescriptive regulations is unnecessary and
overly burdensome to applicants and regulators. The individual and
magniX recommended the FAA not adopt proposed special condition nos.
1(b) and (c), and the individual also recommended holding public
consultations with stakeholders as was done when part 23 was being
reworked into a performance-based form.
The FAA does not concur with the individual's and magniX's
recommendation. While special conditions are rules of particular, not
general applicability, the FAA expects that special condition nos. 1(b)
and (c) support the FAA's transition to a performance-based approach
for developing new requirements. Although the BETA special conditions
are not prescriptive, they provide safety criteria that address hazards
presented by the new electric engine technology used in the BETA H500A
engine. Special condition nos. 1(b) and (c) will be used to incorporate
the additional details that apply to the BETA H500A engine design using
accepted means of compliance. No changes were made to these special
conditions as a result of this comment.
GAMA and magniX commented that special condition nos. 10(g), 15(b),
and 17(f) would require applicants to declare proprietary information
in the engine installation manual, these documentation requirements
establish a precedent beyond that required of their existing
reciprocating or turbine counterparts, and these requirements increase
the risk that sensitive information is disclosed. MagniX stated that
while it is understood this information is used during aircraft-level
certification efforts, traditional data sharing agreements sufficiently
provide the integrator with the required information while respecting
the proprietary nature of the data. MagniX also stated requiring
additional data in the engine installation manual overly constrains the
means whereby this information is shared when compared with established
means, introducing additional commercial risk. GAMA also stated
proposed special condition nos. 10(g), 15(b), and 17(f) are a
requirement for a manufacturer to disclose sensitive proprietary safety
analysis in the engine installation manual, a requirement not currently
imposed on part 33 engines. Additionally, GAMA stated the FAA has not
provided adequate justification for why an electric engine necessitates
this information in a manual. An individual provided a similar comment
regarding proposed special condition nos. 10(g) and 17(f), and stated
that historically such information was captured in other documents such
as the engine control systems interface control document and systems
safety assessment, that were only provided to the installer.
MagniX requested the FAA not adopt the documentation requirements
in proposed special conditions 10(g), 15(b), and 17(f), and proposed
that these data be provided to integrators through generic
``installation instructions'' in lieu of the engine installation
manual. GAMA also requested the FAA reconsider its approach and/or
provide justification for the added requirement of disclosing sensitive
proprietary safety analysis in the engine installation manual. An
individual requested the FAA preserve the engine OEM's flexibility to
document and protect proprietary data by changing ``installation
manual'' to the more generic ``installation instructions,'' which
consist of other documents such as interface control drawings,
technical memorandums, or other installer requested documentation. The
individual further stated that this change would harmonize the special
condition with Sec. 23.2400(e) which uses the verbiage of
``installation instructions,'' and this change could be promulgated to
other special condition paragraphs which refer to the engine
installation manual.
The FAA does not concur with magniX's and GAMA's comments that
special conditions 10(g), 15(b), and 17(f) require disclosing sensitive
information. The requirements cited in their comment do not require
disclosure of sensitive information. As discussed in the NPSC, the
documentation requirements in special conditions nos. 10(g), 15(b), and
17(f) are expected to ensure that the engine is used safely and
properly by constraining the installation of electric engines to only
aircraft types (configurations, flight capabilities, etc.) that were
used by the engine manufacturer to determine the engine ratings,
limits, performance characteristics, as well as the reliability and
criticality of engine systems and parts.
These documentation requirements are intended, and the FAA finds
necessary, to ensure enough information is included to safeguard
compatibility between the electric engine and aircraft, and to prevent
the engine from being used in an aircraft type that requires safety
features or performance characteristics that are not available from a
type certificated engine. For example, electric engines designed for
vertical lift in distributed propulsion tilt-wing aircraft provide
propulsion and act as flight control surfaces, and therefore these
engines have different performance requirements than those used in
conventional normal category airplanes. In addition, the FAA agrees
with the commenters' suggestion to remove the requirement that
specifies the information must be located in the engine installation
manual. These special conditions do not need to specify the document
that must have the information, but only that the information must be
provided to the installer in accordance with the engine installation
instructions under Sec. 33.5, ``Instruction manual for installing and
operating the engine.''. The proposed special conditions are modified
to incorporate this change.
The FAA received a comment from UPSFF.
UPSFF requested that the FAA align these special conditions with
the electric engine requirements included in the certification basis
for special class powered lift aircraft that certify an electric engine
as part of the aircraft type certification.
As stated previously, the engine requirements in the BETA special
conditions are documented differently from proposed powered lift
airworthiness criteria in draft AC 21.17-4 because special conditions
are written in accordance with the requirements of Sec. 21.16, and the
proposed powered-lift airworthiness criteria in draft AC 21.17-4 are
not specific to one applicant. Special conditions are project-specific
rules of particular applicability, and the special conditions for this
electric engine are based on certain novel or unusual design features.
Special conditions may evolve to a general standard as more experience
is gained with certifying the new technology (see Amdt. 21-51). No
changes were made to these special conditions as a result of this
comment.
The FAA received an anonymous comment. The commenter stated the
reference to Sec. 21.17(a) in the preamble of the NPSC seems
contradictory to the language in Sec. 21.17(b). The commenter
explained that since Sec. 21.17(b) applies to ``special classes of
aircraft, including the engines and propellers installed thereon (e.g.,
gliders, airships, and other nonconventional aircraft) . . .'' an
electric engine would be installed on a special class of aircraft as
described in Sec. 21.17(b) and referring to Sec. 21.17(a) seems to
contradict the language in paragraph (b) of that section.
[[Page 101866]]
The FAA does not concur with the comment that indicates the
reference to Sec. 21.17(a) is contradictory to the language in Sec.
21.17(b). Section 21.17(a) provides requirements for developing a
certification basis for an established aviation product, which includes
aircraft, engines, and propellers. The BETA electric engine is an
aircraft engine, which falls under Sec. 21.17(a), and therefore Sec.
21.17(a) is the appropriate reference for this project. Section
21.17(b) provides requirements for developing a certification basis for
special classes of aircraft, such as powered-lift. No changes were made
as a result of this comment.
The FAA also determined that the following changes were necessary.
The phrase, ``In addition'' is added to special condition no. 4, Fire
protection, to connect the introduction sentence to (a) and (b) and
avoid confusion. The FAA also revised the special conditions to use
consistent references to hazardous engine effects. Therefore, the
phrase ``as defined in special condition no. 17 of these special
conditions'' was added wherever ``hazardous engine effects'' is
mentioned.
The FAA recognizes energy regeneration might not be a feature for
some electric engines that operate at their limits, so special
condition no. 23 was changed to specify that ``The endurance
demonstration must include increases and decreases of the engine's
power settings, energy regeneration, and dwellings at the power
settings and energy regeneration for sufficient durations that produce
the extreme physical conditions the engine experiences at rated
performance levels, operational limits, and at any other conditions or
power settings, including energy regeneration that are required to
verify the limit capabilities of the engine.''
In addition, proposed special condition no. 31 was not adopted
because the specific engine model BETA intends to certify is not
designed to use a variable pitch propeller. Except as discussed above,
these special conditions are adopted as proposed.
Applicability
As discussed above, these special conditions are applicable to BETA
Model H500A electric engines. Should BETA 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 BETA Model H500A electric engines. It is
not a rule of general applicability and affects only the applicant who
applied to the FAA for approval of these features on the airplane.
List of Subjects in 14 CFR Part 33
Aircraft, Aviation safety, Reporting and recordkeeping
requirements.
Authority Citation
The authority citation for these special conditions is as follows:
Authority: 49 U.S.C. 106(f), 106(g), 40113, 44701, 44702,
44704.
The Special Conditions
[ssquf] Accordingly, pursuant to the authority delegated to me by
the Administrator, the following special conditions are issued as part
of the type certification basis for BETA Technologies Inc. Model H500A
electric engines. The applicant must also comply with the certification
procedures set forth in part 21.
(1) Applicability
(a) Unless otherwise noted in these special conditions, the engine
design must comply with the airworthiness standards for aircraft
engines set forth in part 33, except for those airworthiness standards
that are specifically and explicitly applicable only to reciprocating
and turbine aircraft engines or as specified herein.
(b) The applicant must comply with this part using a means of
compliance, which may include consensus standards, accepted by the
Administrator.
(c) The applicant requesting acceptance of a means of compliance
must provide the means of compliance to the FAA in a form and manner
acceptable to the Administrator.
(2) Engine Ratings and Operating Limits
In addition to Sec. 33.7(a), the engine ratings and operating
limits must be established and included in the type certificate data
sheet based on:
(a) Shaft power, torque, rotational speed, and temperature for:
(1) Rated takeoff power;
(2) Rated maximum continuous power; and
(3) Rated maximum temporary power and associated time limit.
(b) Duty cycle and the rating at that duty cycle. The duty cycle
must be declared in the engine type certificate data sheet.
(c) Cooling fluid grade or specification.
(d) Power-supply requirements.
(e) Any other ratings or limitations that are necessary for the
safe operation of the engine.
(3) Materials
The engine design must comply with Sec. 33.15.
(4) Fire Protection
The engine design must comply with Sec. 33.17(b) through (g). In
addition--
(a) The design and construction of the engine and the materials
used must minimize the probability of the occurrence and spread of fire
during normal operation and failure conditions and must minimize the
effect of such a fire.
(b) High-voltage electrical wiring interconnect systems must be
protected against arc faults that can lead to hazardous engine effects
as defined in special condition no. 17(d)(2) of these special
conditions. Any non-protected electrical wiring interconnects must be
analyzed to show that arc faults do not cause a hazardous engine
effect.
(5) Durability
The engine design and construction must minimize the development of
an unsafe condition of the engine between maintenance intervals,
overhaul periods, or mandatory actions described in the applicable ICA.
(6) Engine Cooling
The engine design and construction must comply with Sec. 33.21. In
addition, if cooling is required to satisfy the safety analysis as
described in special condition no. 17 of these special conditions, the
cooling system monitoring features and usage must be documented and
provided to the installer as part of the requirements in Sec. 33.5.
(7) Engine Mounting Attachments and Structure
The engine mounting attachments and related engine structures must
comply with Sec. 33.23.
(8) Accessory Attachments
The engine must comply with Sec. 33.25.
(9) Overspeed
(a) A rotor overspeed must not result in a burst, rotor growth, or
damage that results in a hazardous engine effect, as defined in special
condition no. 17(d)(2) of these special conditions. Compliance with
this paragraph must be shown by test, validated analysis, or a
combination of both. Applicable assumed rotor speeds must be declared
and justified.
(b) Rotors must possess sufficient strength with a margin to burst
above
[[Page 101867]]
certified operating conditions and above failure conditions leading to
rotor overspeed. The margin to burst must be shown by test, validated
analysis, or a combination thereof.
(c) The engine must not exceed the rotor speed operational
limitations that could affect rotor structural integrity.
(10) Engine Control Systems
(a) Applicability. The requirements of this special condition apply
to any system or device that is part of the engine type design that
controls, limits, monitors, or protects engine operation, and is
necessary for the continued airworthiness of the engine.
(b) Engine control. The engine control system must ensure that the
engine does not experience any unacceptable operating characteristics
or exceed its operating limits, including in failure conditions where
the fault or failure results in a change from one control mode to
another, from one channel to another, or from the primary system to the
back-up system, if applicable.
(c) Design Assurance. The software and complex electronic hardware,
including programmable logic devices, must be--
(1) Designed and developed using a structured and systematic
approach that provides a level of assurance for the logic commensurate
with the hazard associated with the failure or malfunction of the
systems in which the devices are located; and
(2) Substantiated by a verification methodology acceptable to the
Administrator.
(d) Validation. All functional aspects of the control system must
be substantiated by test, analysis, or a combination thereof, to show
that the engine control system performs the intended functions
throughout the declared operational envelope.
(e) Environmental Limits. Environmental limits that cannot be
adequately substantiated by endurance demonstration, validated
analysis, or a combination thereof must be demonstrated by the system
and component tests in special condition no. 27 of these special
conditions.
(f) Engine control system failures. The engine control system
must--
(1) Have a maximum rate of loss of power control (LOPC) that is
suitable for the intended aircraft application. The estimated LOPC rate
must be documented and provided to the installer as part of the
requirements in Sec. 33.5;
(2) When in the full-up configuration, be single-fault tolerant, as
determined by the Administrator, for electrical, electrically
detectable, and electronic failures involving LOPC events;
(3) Not have any single failure that results in hazardous engine
effects as defined in special condition no. 17(d)(2) of these special
conditions; and
(4) Ensure failures or malfunctions that lead to local events in
the aircraft do not result in hazardous engine effects, as defined in
special condition no. 17(d)(2) of these special conditions, due to
engine control system failures or malfunctions.
(g) System safety assessment. The applicant must perform a system
safety assessment. This assessment must identify faults or failures
that affect normal operation, together with the predicted frequency of
occurrence of these faults or failures. The intended aircraft
application must be taken into account to assure that the assessment of
the engine control system safety is valid. The rates of hazardous and
major faults must be documented and provided to the installer as part
of the requirements in Sec. 33.5.
(h) Protection systems. The engine control devices and systems'
design and function, together with engine instruments, operating
instructions, and maintenance instructions, must ensure that engine
operating limits that can lead to a hazard will not be exceeded in
service.
(i) Aircraft supplied data. Any single failure leading to loss,
interruption, or corruption of aircraft-supplied data (other than
power-command signals from the aircraft), or aircraft-supplied data
shared between engine systems within a single engine or between fully
independent engine systems, must--
(1) Not result in a hazardous engine effect, as defined in special
condition no. 17(d)(2) of these special conditions, for any engine
installed on the aircraft; and
(2) Be able to be detected and accommodated by the control system.
(j) Engine control system electrical power.
(1) The engine control system must be designed such that the loss,
malfunction, or interruption of the control system electrical power
source will not result in a hazardous engine effect, unacceptable
transmission of erroneous data, or continued engine operation in the
absence of the control function. Hazardous engine effects are defined
in special condition no. 17(d)(2) of these special conditions. The
engine control system must be capable of resuming normal operation when
aircraft-supplied power returns to within the declared limits.
(2) The applicant must identify, document, and provide to the
installer as part of the requirements in Sec. 33.5, the
characteristics of any electrical power supplied from the aircraft to
the engine control system, including transient and steady-state voltage
limits, and any other characteristics necessary for safe operation of
the engine.
(11) Instrument Connection
The applicant must comply with Sec. 33.29(a), (e), and (g).
(a) In addition, as part of the system safety assessment of special
condition nos. 10(g) and 33(h) of these special conditions, the
applicant must assess the possibility and subsequent effect of
incorrect fit of instruments, sensors, or connectors. Where
practicable, the applicant must take design precautions to prevent
incorrect configuration of the system.
(b) The applicant must provide instrumentation enabling the flight
crew to monitor the functioning of the engine cooling system unless
evidence shows that:
(1) Other existing instrumentation provides adequate warning of
failure or impending failure;
(2) Failure of the cooling system would not lead to hazardous
engine effects before detection; or
(3) The probability of failure of the cooling system is extremely
remote.
(12) Stress Analysis
(a) A mechanical and thermal stress analysis, as well as an
analysis of the stress caused by electromagnetic forces, must show a
sufficient design margin to prevent unacceptable operating
characteristics and hazardous engine effects as defined in special
condition no. 17(d)(2) of these special conditions.
(b) Maximum stresses in the engine must be determined by test,
validated analysis, or a combination thereof, and must be shown not to
exceed minimum material properties.
(13) Critical and Life-Limited Parts
(a) The applicant must show, by a safety analysis or means
acceptable to the Administrator, whether rotating or moving components,
bearings, shafts, static parts, and non-redundant mount components
should be classified, designed, manufactured, and managed throughout
their service life as critical or life-limited parts.
(1) Critical part means a part that must meet prescribed integrity
specifications to avoid its primary failure, which is likely to result
in a hazardous engine effect as defined in special condition no.
17(d)(2) of these special conditions.
(2) Life-limited parts may include but are not limited to a rotor
or major
[[Page 101868]]
structural static part, the failure of which can result in a hazardous
engine effect, as defined in special condition no. 17(d)(2) of these
special conditions, due to a low-cycle fatigue (LCF) mechanism. A life
limit is an operational limitation that specifies the maximum allowable
number of flight cycles that a part can endure before the applicant
must remove it from the engine.
(b) In establishing the integrity of each critical part or life-
limited part, the applicant must provide to the Administrator the
following three plans for approval:
(1) an engineering plan, as defined in Sec. 33.70(a);
(2) a manufacturing plan, as defined in Sec. 33.70(b); and
(3) a service-management plan, as defined in Sec. 33.70(c).
(14) Lubrication System
(a) The lubrication system must be designed and constructed to
function properly between scheduled maintenance intervals in all flight
attitudes and atmospheric conditions in which the engine is expected to
operate.
(b) The lubrication system must be designed to prevent
contamination of the engine bearings and lubrication system components.
(c) The applicant must demonstrate by test, validated analysis, or
a combination thereof, the unique lubrication attributes and functional
capability of (a) and (b).
(15) Power Response
(a) The design and construction of the engine, including its
control system, must enable an increase--
(1) From the minimum power setting to the highest rated power
without detrimental engine effects;
(2) From the minimum obtainable power while in-flight and while on
the ground to the highest rated power within a time interval determined
to be appropriate for the intended aircraft application; and
(3) From the minimum torque to the highest rated torque without
detrimental engine effects in the intended aircraft application.
(b) The results of (a)(1), (a)(2), and (a)(3) of this special
condition must be documented and provided to the installer as part of
the requirements in Sec. 33.5.
(16) Continued Rotation
If the design allows any of the engine main rotating systems to
continue to rotate after the engine is shut down while in-flight, this
continued rotation must not result in any hazardous engine effects, as
defined in special condition no. 17(d)(2) of these special conditions.
(17) Safety Analysis
(a) The applicant must comply with Sec. 33.75(a)(1) and (a)(2)
using the failure definitions in special condition no. 17(d) of these
special conditions.
(b) The primary failure of certain single elements cannot be
sensibly estimated in numerical terms. If the failure of such elements
is likely to result in hazardous engine effects, then compliance may be
shown by reliance on the prescribed integrity requirements of Sec.
33.15 and special condition nos. 9 and 13 of these special conditions,
as applicable. These instances must be stated in the safety analysis.
(c) The applicant must comply with Sec. 33.75(d) and (e) using the
failure definitions in special condition no. 17(d) of these special
conditions, and the ICA in Sec. 33.4.
(d) Unless otherwise approved by the Administrator, the following
definitions apply to the engine effects when showing compliance with
this condition:
(1) A minor engine effect does not prohibit the engine from
performing its intended functions in a manner consistent with Sec.
33.28(b)(1)(i), (b)(1)(iii), and (b)(1)(iv), and the engine complies
with the operability requirements of special condition no. 15 and
special condition no. 25 of these special conditions, as appropriate.
(2) The engine effects in Sec. 33.75(g)(2) are hazardous engine
effects with the addition of:
(i) Electrocution of the crew, passengers, operators, maintainers,
or others; and
(ii) Blockage of cooling systems that could cause the engine
effects described in Sec. 33.75(g)(2) and special condition
17(d)(2)(i) of these special conditions.
(3) Any other engine effect is a major engine effect.
(e) The intended aircraft application must be taken into account
when performing the safety analysis.
(f) The results of the safety analysis, and the assumptions about
the aircraft application used in the safety analysis, must be
documented and provided to the installer as part of the requirements in
Sec. 33.5.
(18) Ingestion
(a) Rain, ice, and hail ingestion must not result in an abnormal
operation such as shutdown, power loss, erratic operation, or power
oscillations throughout the engine operating range.
(b) Ingestion from other likely sources (birds, induction system
ice, foreign objects--ice slabs) must not result in hazardous engine
effects defined by special condition no. 17(d)(2) of these special
conditions, or unacceptable power loss.
(c) If the design of the engine relies on features, attachments, or
systems that the installer may supply, for the prevention of
unacceptable power loss or hazardous engine effects, as defined in
special condition no. 17(d)(2) of these special conditions, following
potential ingestion, then the features, attachments, or systems must be
documented and provided to the installer as part of the requirements in
Sec. 33.5.
(19) Liquid and Gas Systems
(a) Each system used for lubrication or cooling of engine
components must be designed and constructed to function properly in all
flight attitudes and atmospheric conditions in which the engine is
expected to operate.
(b) If a system used for lubrication or cooling of engine
components is not self-contained, the interfaces to that system must be
defined, documented and provided to the installer as part of the
requirements in Sec. 33.5.
(c) The applicant must establish by test, validated analysis, or a
combination of both that all static parts subject to significant
pressure loads will not:
(1) Exhibit permanent distortion beyond serviceable limits, or
exhibit leakage that could create a hazardous condition when subjected
to normal and maximum working pressure with margin;
(2) Exhibit fracture or burst when subjected to the greater of
maximum possible pressures with margin.
(d) Compliance with special condition no. 19(c) of these special
conditions must take into account:
(1) The operating temperature of the part;
(2) Any other significant static loads in addition to pressure
loads;
(3) Minimum properties representative of both the material and the
processes used in the construction of the part; and
(4) Any adverse physical geometry conditions allowed by the type
design, such as minimum material and minimum radii.
(e) Approved coolants and lubricants must be listed, documented,
and provided to the installer as part of the requirements in Sec.
33.5.
(20) Vibration Demonstration
(a) The engine must be designed and constructed to function
throughout its normal operating range of rotor speeds and engine output
power, including
[[Page 101869]]
defined exceedances, without inducing excessive stress in any of the
engine parts because of vibration and without imparting excessive
vibration forces to the aircraft structure.
(b) Each engine design must undergo a vibration survey to establish
that the vibration characteristics of those components subject to
induced vibration are acceptable throughout the declared flight
envelope and engine operating range for the specific installation
configuration. The possible sources of the induced vibration that the
survey must assess are mechanical, aerodynamic, acoustical, internally
induced electromagnetic, installation induced effects that can affect
the engine vibration characteristics, and likely environmental effects.
This survey must be shown by test, validated analysis, or a combination
thereof.
(21) Overtorque
When approval is sought for a transient maximum engine overtorque,
the applicant must demonstrate by test, validated analysis, or a
combination thereof, that the engine can continue operation after
operating at the maximum engine overtorque condition without
maintenance action. Upon conclusion of overtorque tests conducted to
show compliance with this special condition, or any other tests that
are conducted in combination with the overtorque test, each engine part
or individual groups of components must meet the requirements of
special condition no. 29 of these special conditions.
(22) Calibration Assurance
Each engine must be subjected to calibration tests to establish its
power characteristics, and the conditions both before and after the
endurance and durability demonstrations specified in special conditions
nos. 23 and 26 of these special conditions.
(23) Endurance Demonstration
The applicant must subject the engine to an endurance
demonstration, acceptable to the Administrator, to demonstrate the
engine's limit capabilities. The endurance demonstration must include
increases and decreases of the engine's power settings, energy
regeneration, and dwellings at the power settings and energy
regeneration for sufficient durations that produce the extreme physical
conditions the engine experiences at rated performance levels,
operational limits, and at any other conditions or power settings,
including energy regeneration 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, which would impact temperature
limits.
(25) Operation Demonstration
The engine design must demonstrate safe operating characteristics,
including but not limited to power cycling, starting, acceleration, and
overspeeding throughout its declared flight envelope and operating
range. The declared engine operational characteristics must account for
installation loads and effects.
(26) Durability Demonstration
The engine must be subjected to a durability demonstration to show
that each part of the engine has been designed and constructed to
minimize any unsafe condition of the system between overhaul periods,
or between engine replacement intervals if the overhaul is not defined.
This test must simulate the conditions in which the engine is expected
to operate in service, including typical start-stop cycles, to
establish when the initial maintenance is required.
(27) System and Component Tests
The applicant must show that systems and components that cannot be
adequately substantiated in accordance with the endurance demonstration
or other demonstrations will perform their intended functions in all
declared environmental and operating conditions.
(28) Rotor Locking Demonstration
If shaft rotation is prevented by locking the rotor(s), the engine
must demonstrate:
(a) Reliable rotor locking performance;
(b) Reliable rotor unlocking performance; and
(c) That no hazardous engine effects, as specified in special
condition no. 17(d)(2) of these special conditions, will occur.
(29) Teardown Inspection
(a) Teardown evaluation.
(1) After the endurance and durability demonstrations have been
completed, the engine must be completely disassembled. Each engine
component and lubricant must be eligible for continued operation in
accordance with the information submitted for showing compliance with
Sec. 33.4.
(2) Each engine component, having an adjustment setting and a
functioning characteristic that can be established independent of
installation on or in the engine, must retain each setting and
functioning characteristic within the established and recorded limits
at the beginning of the endurance and durability demonstrations.
(b) Non-Teardown evaluation. If a teardown cannot be performed for
all engine components in a non-destructive manner, then the inspection
or replacement intervals for these components and lubricants must be
established based on the endurance and durability demonstrations and
must be documented in the ICA in accordance with Sec. 33.4.
(30) Containment
The engine must be designed and constructed to protect against
likely hazards from rotating components as follows--
(a) The design of the stator case surrounding rotating components
must provide for the containment of the rotating components in the
event of failure, unless the applicant shows that the margin to rotor
burst precludes the possibility of a rotor burst.
(b) If the margin to burst shows that the stator case must have
containment features in the event of failure, then the stator case must
provide for the containment of the failed rotating components. The
applicant must define by test, validated analysis, or a combination
thereof, and document and provide to the installer as part of the
requirements in Sec. 33.5, the energy level, trajectory, and size of
fragments released from damage caused by the main-rotor failure, and
that pass forward or aft of the surrounding stator case.
(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
[[Page 101870]]
(4) Replacement of an engine part is found necessary during the
tests, or due to the teardown inspection findings.
(c) Upon completion of all demonstrations and testing specified in
these special conditions, the engine and its components must be--
(1) Within serviceable limits;
(2) Safe for continued operation; and
(3) Capable of operating at declared ratings while remaining within
limits.
(33) Engine Electrical Systems
(a) Applicability. Any system or device that provides, uses,
conditions, or distributes electrical power, and is part of the engine
type design, must provide for the continued airworthiness of the
engine, and must maintain electric engine ratings.
(b) Electrical systems. The electrical system must ensure the safe
generation and transmission of power, and electrical load shedding if
required, and that the engine does not experience any unacceptable
operating characteristics or exceed its operating limits.
(c) Electrical power distribution.
(1) The engine electrical power distribution system must be
designed to provide the safe transfer of electrical energy throughout
the powerplant. The system must be designed to provide electrical power
so that the loss, malfunction, or interruption of the electrical power
source will not result in a hazardous engine effect, as defined in
special condition no. 17(d)(2) of these special conditions.
(2) The system must be designed and maintained to withstand normal
and abnormal conditions during all ground and flight operations.
(3) The system must provide mechanical or automatic means of
isolating a faulted electrical energy generation or storage device from
leading to hazardous engine effects, as defined in special condition
no. 17(d)(2) of these special conditions, or detrimental effects in the
intended aircraft application.
(d) Protection systems. The engine electrical system must be
designed such that the loss, malfunction, interruption of the
electrical power source, or power conditions that exceed design limits,
will not result in a hazardous engine effect, as defined in special
condition no. 17(d)(2) of these special conditions.
(e) Electrical power characteristics. The applicant must identify,
declare, document, and provide to the installer as part of the
requirements in Sec. 33.5, the characteristics of any electrical power
supplied from--
(1) the aircraft to the engine electrical system, for starting and
operating the engine, including transient and steady-state voltage
limits, and
(2) the engine to the aircraft via energy regeneration, and any
other characteristics necessary for safe operation of the engine.
(f) Environmental limits. Environmental limits that cannot
adequately be substantiated by endurance demonstration, validated
analysis, or a combination thereof must be demonstrated by the system
and component tests in special condition no. 27 of these special
conditions.
(g) Electrical system failures. The engine electrical system must--
(1) Have a maximum rate of 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. The rates of hazardous and major faults
must be declared, documented, and provided to the installer as part of
the requirements in Sec. 33.5.
Issued in Kansas City, Missouri, on December 10, 2024.
Patrick R. Mullen,
Manager, Technical Policy Branch, Policy and Standards Division,
Aircraft Certification Service.
[FR Doc. 2024-29490 Filed 12-16-24; 8:45 am]
BILLING CODE 4910-13-P
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</html>Indexed from Federal Register on December 17, 2024.
This is legal information, not legal advice. Laws vary by jurisdiction and change frequently. Always verify current law with official sources and consult a licensed attorney in your jurisdiction for advice on your specific situation.