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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 [www.gpo.gov]
[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; 
[email protected].

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\ https://standards.rtca.org/XanHrK.
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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\ https://my.rtca.org/NC__Product?id=a1B36000001IcnSEAS.
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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\ https://www.faa.gov/documentLibrary/media/Advisory_Circular/AC_33_28-3.pdf.
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    The phrase ``in the full-up configuration'' used in 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:
     System or component characteristics outside of normal 
statistical variation from circumstances such as systems degradation, 
installation error, and engine response to fault conditions;
     Unusual environmental conditions from extreme temperature, 
humidity, vibration, lightning, high-intensity radiated field (HIRF), 
atmospheric neutron radiation; and
     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:
     System or components outside of normal statistical 
variation due to degradation, or installation error
     Unusual environmental conditions such as extreme 
temperature, humidity, FOD impact, severe lightning, HIRF, or 
atmospheric radiation
     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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