[Federal Register Volume 89, Number 46 (Thursday, March 7, 2024)]
[Proposed Rules]
[Pages 16474-16486]
From the Federal Register Online via the Government Publishing Office [www.gpo.gov]
[FR Doc No: 2024-04800]



[[Page 16474]]

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DEPARTMENT OF TRANSPORTATION

Federal Aviation Administration

14 CFR Part 33

[Docket No. FAA-2022-1641; Notice No. 33-22-01-SC]


Special Conditions: BETA Technologies Inc. Model H500A Electric 
Engines

AGENCY: Federal Aviation Administration (FAA), Department of 
Transportation (DOT).

ACTION: Notice of proposed special conditions.

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SUMMARY: This action proposes special conditions for 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 have a novel or unusual design feature when compared to 
the state of technology envisioned in the airworthiness standards 
applicable to aircraft engines. The design feature is the use of an 
electric motor, motor controller, and high-voltage systems as the 
primary source of propulsion for an aircraft. The applicable 
airworthiness regulations do not contain adequate or appropriate safety 
standards for this design feature. These proposed special conditions 
contain the additional safety standards that the Administrator 
considers necessary to establish a level of safety equivalent to that 
established by the existing airworthiness standards.

DATES: Send comments on or before April 8, 2024.

ADDRESSES: Send comments identified by Docket No. FAA-2022-1641 using 
any of the following methods:
     Federal eRegulations Portal: Go to https://www.regulations.gov/ and follow the online instructions for sending 
your comments electronically.
     Mail: Send comments to Docket Operations, M-30, U.S. 
Department of Transportation, 1200 New Jersey Avenue SE, Room W12-140, 
West Building, Ground Floor, Washington, DC 20590-0001.
     Hand Delivery or Courier: Take comments to Docket 
Operations in Room W12-140 of the West Building, Ground Floor at 1200 
New Jersey Avenue SE, Washington, DC, between 9 a.m. and 5 p.m., Monday 
through Friday, except Federal holidays.
     Fax: Fax comments to Docket Operations at 202-493-2251.
    Docket: Background documents or comments received may be read at 
https://www.regulations.gov/ at any time. Follow the online 
instructions for accessing the docket or go to Docket Operations in 
Room W12-140 of the West Building, Ground Floor at 1200 New Jersey 
Avenue SE, Washington, DC, between 9 a.m. and 5 p.m., Monday through 
Friday, except Federal holidays.

FOR FURTHER INFORMATION CONTACT: Mark Bouyer, Engine and Propulsion 
Standards Section, AIR-625, Technical Policy Branch, Policy and 
Standards Division, Aircraft Certification Service, 1200 District 
Avenue, Burlington, Massachusetts 01803; telephone (781) 238-7755; 
[email protected].

SUPPLEMENTARY INFORMATION:

Comments Invited

    The FAA invites interested people to take part in this rulemaking 
by sending written comments, data, or views. The most helpful comments 
reference a specific portion of the proposed special conditions, 
explain the reason for any recommended change, and include supporting 
data.
    The FAA will consider all comments received by the closing date for 
comments. The FAA may change these proposed special conditions based on 
the comments received.

Privacy

    Except for Confidential Business Information (CBI) as described in 
the following paragraph, and other information as described in title 
14, Code of Federal Regulations (14 CFR) 11.35, the FAA will post all 
comments received, without change, to https://www.regulations.gov/, 
including any personal information you provide. The FAA will also post 
a report summarizing each substantive verbal contact received about 
these special conditions.

Confidential Business Information

    Confidential Business Information is commercial or financial 
information that is both customarily and actually treated as private by 
its owner. Under the Freedom of Information Act (FOIA) (5 U.S.C. 552), 
CBI is exempt from public disclosure. If your comments responsive to 
this document contain commercial or financial information that is 
customarily treated as private, that you actually treat as private, and 
that is relevant or responsive to this document, it is important that 
you clearly designate the submitted comments as CBI. Please mark each 
page of your submission containing CBI as ``PROPIN.'' The FAA will 
treat such marked submissions as confidential under the FOIA, and the 
indicated comments will not be placed in the public docket of these 
proposed special conditions. Send submissions containing CBI to the 
individual listed in the For Further Information Contact section below. 
Comments the FAA receives, which are not specifically designated as 
CBI, will be placed in the public docket for these proposed special 
conditions.

Background

    On 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 
propeller shaft, that 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.

Type Certification Basis

    Under the provisions of 14 CFR 21.17(a)(1), generally, BETA must 
show that Model H500A 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 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

[[Page 16475]]

would also apply to the other engine model under Sec.  21.101.
    The FAA issues special conditions, as defined in Sec.  11.19, in 
accordance with Sec.  11.38, and they become part of the type 
certification basis under Sec.  21.17(a)(2).
    In addition to the applicable airworthiness regulations and special 
conditions, the BETA Model H500A engines must comply with the noise 
certification requirements of 14 CFR part 36.

Novel or Unusual Design Features

    The BETA Model H500A engines will incorporate the following novel 
or unusual design features:
    An electric motor, motor controller, and high-voltage electrical 
systems that are used as the primary source of propulsion for an 
aircraft.

Discussion

    Electric propulsion technology is substantially different from the 
technology used in previously certificated turbine and reciprocating 
engines. Therefore, these engines introduce new safety concerns that 
need to be addressed in the certification basis.
    A growing interest within the aviation industry involves electric 
propulsion technology. As a result, international agencies and industry 
stakeholders formed Committee F39 under ASTM International, formerly 
known as American Society for Testing and Materials, to identify the 
appropriate technical criteria for aircraft engines using electrical 
technology that has not been previously type certificated for aircraft 
propulsion systems. ASTM International is an international standards 
organization that develops and publishes voluntary consensus technical 
standards for a wide range of materials, products, systems, and 
services. ASTM International published ASTM F3338-18, ``Standard 
Specification for Design of Electric Propulsion Units for General 
Aviation Aircraft,'' in December 2018.\1\ The FAA used the technical 
criteria from the ASTM F3338-18, the published Special Conditions No. 
33-022-SC for the magniX USA, Inc. Model magni350 and magni650 engines, 
and information from the BETA Model H500A engine design to develop 
special conditions that establish an equivalent level of safety to that 
required by part 33.
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    \1\ https://www.astm.org/Standards/F3338.html.
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Part 33 Was Developed for Gas-Powered Turbine and Reciprocating Engines

    Aircraft engines make use of an energy source to drive mechanical 
systems that provide propulsion for the aircraft. Energy can be 
generated from various sources such as petroleum and natural gas. The 
turbine and reciprocating aircraft engines certificated under part 33 
use aviation fuel for an energy source. The reciprocating and turbine 
engine technology that was anticipated in the development of part 33 
converts oxygen and fuel to energy using an internal combustion system, 
which generates heat and mass flow of combustion products for turning 
shafts that are attached to propulsion devices such as propellers and 
ducted fans. Part 33 regulations set forth standards for these engines 
and mitigate potential hazards resulting from failures and 
malfunctions. The nature, progression, and severity of engine failures 
are tied closely to the technology that is used in the design and 
manufacture of aircraft engines. These technologies involve chemical, 
thermal, and mechanical systems. Therefore, the existing engine 
regulations in part 33 address certain chemical, thermal, and 
mechanically induced failures that are specific to air and fuel 
combustion systems operating with cyclically loaded, high-speed, high-
temperature, and highly stressed components.

BETA's Proposed Electric Engines Are Novel or Unusual

    The existing part 33 airworthiness standards for aircraft engines 
date back to 1965. As discussed in the previous paragraphs, these 
airworthiness standards are based on fuel-burning reciprocating and 
turbine engine technology. The BETA Model H500A engines are neither 
turbine nor reciprocating engines. These engines have a novel or 
unusual design feature, which is the use of electrical sources of 
energy instead of fuel to drive the mechanical systems that provide 
propulsion for aircraft. The BETA aircraft engine is subject to 
operating conditions produced by chemical, thermal, and mechanical 
components working together, but the operating conditions are unlike 
those observed in internal combustion engine systems. Therefore, part 
33 does not contain adequate or appropriate safety standards for the 
BETA Model H500A engine's novel or unusual design feature.
    BETA's proposed aircraft engines will operate using electrical 
power instead of air and fuel combustion to propel the aircraft. These 
electric engines will be designed, manufactured, and controlled 
differently than turbine or reciprocating aircraft engines. They will 
be built with an electric motor, motor controller, and high-voltage 
electrical systems that draw energy from electrical storage or 
electrical energy generating systems. The electric motor is a device 
that converts electrical energy into mechanical energy by electric 
current flowing through windings (wire coils) in the motor, producing a 
magnetic field that interacts with permanent magnets mounted on the 
engine's main rotor. The controller is a system that consists of two 
main functional elements: the motor controller and an electric power 
inverter to drive the motor.\2\ The high-voltage electrical system is a 
combination of wires and connectors that integrate the motor and 
controller.
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    \2\ Sometimes the entire system is referred to as an inverter. 
Throughout this document, it is referred to as the controller.
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    In addition, the technology comprising these high-voltage and high-
current electronic components introduces potential hazards that do not 
exist in turbine and reciprocating aircraft engines. For example, high-
voltage transmission lines, electromagnetic shields, magnetic 
materials, and high-speed electrical switches are necessary to use the 
physical properties of an electric engine for propelling an aircraft. 
However, this technology also exposes the aircraft to potential 
failures that are not common to gas-powered turbine and reciprocating 
engines, technological differences which could adversely affect safety 
if not addressed through these proposed special conditions.

BETA's Proposed Electric Engines Require a Mix of Part 33 Standards and 
Special Conditions

    Although the electric aircraft engines BETA proposes use novel or 
unusual design features that the FAA did not envisage during the 
development of its existing part 33 airworthiness standards, these 
engines share some basic similarities, in configuration and function, 
to engines that use the combustion of air and fuel, and therefore 
require similar provisions to prevent common hazards (e.g., fire, 
uncontained high energy debris, and loss of thrust control). However, 
the primary failure concerns and the probability of exposure to these 
common hazards are different for the proposed BETA Model H500A electric 
engine. This creates a need to develop special conditions to ensure the 
engine's safety and reliability.

[[Page 16476]]

    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. This necessitates the development of 
special conditions that provide adequate airworthiness standards for 
these aircraft engines.
    The requirements in part 33, subpart B, are applicable to 
reciprocating and turbine aircraft engines. Subparts C and D are 
applicable to reciprocating aircraft engines. Subparts E through G are 
applicable to turbine aircraft engines. As such, subparts B through G 
do not adequately address the use of aircraft engines that operate 
using electrical technology. Special conditions are needed to ensure a 
level of safety for electric engines that is commensurate with these 
subparts, as those regulatory requirements do not contain adequate or 
appropriate safety standards for electric aircraft engines that are 
used to propel aircraft.

FAA Proposed Special Conditions for the BETA Engine Design

    Applicability: Proposed special condition no. 1 would require 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: Proposed special 
condition no. 2 would, in addition to compliance with Sec.  33.7(a), 
require BETA to establish engine operating limits related to the power, 
torque, speed, and duty cycles specific to BETA Model H500A 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 engine capabilities required for safe operation of the 
engine.
    Materials: Proposed special condition no. 3 would require 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: Proposed special condition no. 4 would require 
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 proposed special condition 
would require 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: Proposed special condition no. 5 would require the 
design and construction of BETA Model H500A engines to minimize the 
development of an unsafe condition between maintenance intervals, 
overhaul periods, and mandatory actions described in the Instructions 
for Continued Airworthiness (ICA).
    Engine Cooling: Proposed special condition no. 6 would require 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 proposed special condition would require 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 proposed special condition no. 17. 
Loss of cooling to an aircraft engine that operates using electrical 
technology can result in rapid overheating and abrupt engine failure, 
with critical consequences to safety.
    Engine Mounting Attachments and Structure: Proposed special 
condition no. 7 would require BETA and the proposed design to comply 
with Sec.  33.23, which requires the applicant to define, and the 
proposed design to withstand, certain load limits for the engine 
mounting attachments and related engine structure. These requirements 
would otherwise be applicable only to reciprocating and turbine 
aircraft engines.
    Accessory Attachments: Proposed special condition no. 8 would 
require the proposed design to comply with Sec.  33.25, which sets 
certain design, operational, and maintenance requirements for the 
engine's accessory drive and mounting attachments, and which would 
otherwise be applicable only to reciprocating and turbine aircraft 
engines.
    Rotor Overspeed: Proposed special condition no. 9 would require 
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 proposed special condition associated with rotor overspeed is 
necessary because of the differences between turbine engine technology 
and the technology of these electric engines. Turbine rotor speed is 
driven by expanding gas and aerodynamic loads on rotor blades. 
Therefore, the rotor speed or overspeed results from interactions 
between thermodynamic and aerodynamic engine properties. The speed of 
an electric engine is directly controlled by electric current, and an 
electromagnetic field created by the controller. Consequently, electric 
engine rotor response to power demand and overspeed-protection systems 
is quicker and more precise. Also, the failure modes that can lead to 
overspeed between turbine engines and electric engines are vastly 
different, and therefore this special condition is necessary.
    Engine Control Systems: Proposed special condition no. 10(b) would 
require 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, the 
proposed special condition no. 10(b) associated with controlling these 
engines is necessary.

[[Page 16477]]

    Proposed special condition no. 10(c) would require 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,\3\ distinguishes between complex and 
simple electronic hardware.
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    \3\ https://my.rtca.org/NC__Product?id=a1B36000001IcjTEAS.
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    Proposed special condition no. 10(d) would require data from 
assessments of all functional aspects of the control system to prevent 
errors that could exist in software programs that are not readily 
observable by inspection of the code. Also, 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 proposed special condition 
no. 10(e) include temperature, vibration, high-intensity radiated 
fields (HIRF), and others addressed in RTCA DO-160G, ``Environmental 
Conditions and Test Procedures for Airborne Electronic/Electrical 
Equipment and Instruments'' dated December 08, 2010, which includes 
``DO-160G Change 1--Environmental Conditions and Test Procedures for 
Airborne Equipment'' dated December, 16, 2014, and ``DO-357--User 
Guide: Supplement to DO-160G'' dated December 16, 2014.\4\ Proposed 
special condition 10(e) would require BETA to demonstrate by system or 
component tests in proposed special condition no. 27 any environmental 
limits that cannot be adequately substantiated by the endurance 
demonstration, validated analysis, or a combination thereof.
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    \4\ https://my.rtca.org/NC__Product?id=a1B36000001IcnSEAS.
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    Proposed special condition no. 10(f) would require BETA to evaluate 
various control system failures to assure that such failures will not 
lead to unsafe engine conditions. The FAA issued Advisory Circular (AC) 
AC 33.28-3, ``Guidance Material for 14 CFR Sec.  33.28, Engine Control 
Systems,'' on May 23, 2014, for reciprocating and turbine engines.\5\ 
Paragraph 6-2 of this AC provides guidance for defining an engine 
control system failure when showing compliance with the requirements of 
Sec.  33.28. AC 33.28-3 also includes objectives for control system 
integrity requirements, criteria for a loss of thrust (or power) 
control (LOTC/LOPC) event, and an acceptable LOTC/LOPC rate. The 
electrical and electronic failures and failure rates did not account 
for electric engines when the FAA issued this AC, and therefore 
performance-based special conditions are proposed to allow fault 
accommodation criteria to be developed for electric engines.
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    \5\ https://www.faa.gov/documentLibrary/media/Advisory_Circular/AC_33_28-3.pdf.
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    The phrase ``in the full-up configuration'' used in proposed 
special condition no. 10(f)(2) refers to a system without any fault 
conditions present. The electronic control system must, when in the 
full-up configuration, be single-fault tolerant, as determined by the 
Administrator, for electrical, electrically detectable, and electronic 
failures involving LOPC events.
    The term ``local'' in the context of ``local events'' used in 
proposed special condition no. 10(f)(4) means failures or malfunctions 
leading to events in the intended aircraft installation such as fire, 
overheat, or failures leading to damage to engine control system 
components. These local events must not result in a hazardous engine 
effect due to engine control system failures or malfunctions.
    Proposed special condition no. 10(g) would require BETA to conduct 
a safety assessment of the control system to support the safety 
analysis in proposed special condition no. 17. This control system 
safety assessment provides engine response to failures, and rates of 
these failures that can be used at the aircraft-level safety 
assessment.
    Proposed special condition no. 10(h) requires BETA to provide 
appropriate protection devices or systems to ensure that engine 
operating limits will not be exceeded in service.
    Proposed special condition no. 10(i) is necessary to ensure that 
the controllers are self-sufficient and isolated from other aircraft 
systems. The aircraft-supplied data supports the analysis at the 
aircraft level to protect the aircraft from common mode failures that 
could lead to major propulsion power loss. The exception ``other than 
power command signals from the aircraft,'' noted in proposed special 
condition no. 10(i), is based on the FAA's determination that the 
engine controller has no reasonable means to determine the validity of 
any in-range signals from the electrical power system. In many cases, 
the engine control system can detect a faulty signal from the aircraft, 
but the engine control system typically accepts the power command 
signal as a valid value.
    The term ``independent'' in the context of ``fully independent 
engine systems'' referenced in proposed special condition no. 10(i) 
means the controllers should be self-sufficient and isolated from other 
aircraft systems or provide redundancy that enables the engine control 
system to accommodate aircraft data system failures. In the case of 
loss, interruption, or corruption of aircraft-supplied data, the engine 
must continue to function in a safe and acceptable manner without 
hazardous engine effects.
    The term ``accommodated,'' in the context of ``detected and 
accommodated,'' referenced in proposed special condition 10(i)(2) is to 
assure that, upon detecting a fault, the system continues to function 
safely.
    Proposed special condition no. 10(j) would require 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: Proposed special condition no. 11 would 
require 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. Proposed special condition no. 11 
would require the safety analysis to include potential hazardous 
effects from failures of instrument connections to function properly. 
The outcome of this analysis might identify the need for design 
enhancements or additional ICA to ensure safety.
    Stress Analysis: Section 33.62 requires applicants to perform a 
stress analysis on each turbine engine. This regulation is explicitly 
applicable only to turbine engines and turbine engine components, and 
it is not appropriate for the BETA Model H500A engines. However, the 
FAA proposes that a stress analysis particular to these electric 
engines is necessary to account for stresses resulting from electric 
technology used in the engine.
    Proposed special condition no. 12 would require a mechanical, 
thermal, and electrical stress analysis to show that the engine has a 
sufficient design margin to prevent unacceptable operating 
characteristics. Also, the

[[Page 16478]]

applicant must determine the maximum stresses in the engine by tests, 
validated analysis, or a combination thereof, and show that they do not 
exceed minimum material properties.
    Critical and Life-Limited Parts: Proposed special condition no. 13 
would require 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 proposed special 
condition no. 13(a)(2), is a decline in material strength from exposure 
to cyclic stress at levels beyond the stress threshold the material can 
sustain indefinitely. This threshold is known as the ``material 
endurance limit.'' Low-cycle fatigue typically causes a part to sustain 
plastic or permanent deformation during the cyclic loading and can lead 
to cracks, crack growth, and fracture. Engine parts that operate at 
high temperatures and high mechanical stresses simultaneously can 
experience low-cycle fatigue coupled with creep. Creep is the tendency 
of a metallic material to permanently move or deform when it is exposed 
to the extreme thermal conditions created by hot combustion gasses, and 
substantial physical loads such as high rotational speeds and maximum 
thrust. Conversely, high-cycle fatigue is caused by elastic 
deformation, small strains caused by alternating stress, and a much 
higher number of load cycles compared to the number of cycles that 
cause low-cycle fatigue.
    The engineering plan referenced in proposed special condition no. 
13(b)(1) informs the manufacturing and service management processes of 
essential information that ensures the life limit of a part is valid. 
The engineering plan provides methods for verifying the characteristics 
and qualities assumed in the design data using methods that are 
suitable for the part criticality. The engineering plan informs the 
manufacturing process of 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 proposed special 
condition no. 13(b)(2), is the collection of data required to translate 
documented engineering design criteria into physical parts, and to 
verify that the parts comply with the properties established by the 
design data. Because engines are not intentionally tested to failure 
during a certification program, documents and processes used to execute 
production and quality systems required by Sec.  21.137 guarantee 
inherent expectations for performance and durability. These systems 
limit the potential manufacturing outcomes to parts that are 
consistently produced within design constraints.
    The manufacturing plan and service management plan ensure that 
essential information from the engineering plan, such as the design 
characteristics that safeguard the integrity of critical and life-
limited parts, is consistently produced and preserved over the lifetime 
of those parts. The manufacturing plan includes special processes and 
production controls to prevent inclusion of manufacturing-induced 
anomalies, which can degrade the part's structural integrity. Examples 
of manufacturing-induced anomalies are material contamination, 
unacceptable grain growth, heat-affected areas, and residual stresses.
    The service-management plan ensures the method and assumptions used 
in the engineering plan to determine the part's life remain valid by 
enabling corrections identified from in-service experience, such as 
service-induced anomalies and unforeseen environmental effects, to be 
incorporated into the design process. The service-management plan also 
becomes the ICA for maintenance, overhaul, and repairs of the part.
    Lubrication System: Proposed special condition no. 14 would require 
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 proposed special condition 
would also require BETA to demonstrate the unique lubrication 
attributes and functional capability of the BETA Model H500A 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 engines. For 
example, electric engines do not have a crankcase or lubrication oil 
sump. Electric engine bearings are sealed, so they do not require an 
oil circulation system. The lubrication system in these engines is also 
independent of the propeller pitch control system. Therefore, proposed 
special condition no. 14 incorporates only certain requirements from 
the part 33 regulations.
    Power Response: Proposed special condition no. 15 would require the 
design and construction of the BETA Model H500A 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 engines, but the causes are 
different. Electric engines with reduced power response time can 
experience insufficient thrust to the aircraft, shaft over-torque, and 
over-stressed rotating components, propellers, and critical propeller 
parts. Therefore, this proposed special condition is necessary.
    Continued Rotation: Proposed special condition no. 16 would require 
BETA to design the Model H500A 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 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 proposed special condition are substantial asymmetric 
aerodynamic drag that can cause aircraft instability, loss of control, 
and reduced efficiency; and may result in a forced landing or inability 
to continue safe flight.
    Safety Analysis: Proposed special condition no. 17 would require 
BETA to comply with Sec.  33.75(a)(1) and (a)(2), which require the 
applicant to conduct

[[Page 16479]]

a safety analysis of the engine, and which would otherwise be 
applicable only to turbine aircraft engines. Additionally, this 
proposed special condition would require 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 proposed special conditions apply to an 
engine that continues to operate at partial power after a single 
electrical or electronic fault or failure. Total loss of power is 
classified at the aircraft level using proposed special condition nos. 
10(g) and 33(h).
    Ingestion: Proposed special condition no. 18 would require 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 proposed special 
condition no. 18(b), is such that the power or thrust required for safe 
flight of the aircraft becomes unavailable to the pilot. The specific 
amount of power loss that is required for safe flight depends on the 
aircraft configuration, speed, altitude, attitude, atmospheric 
conditions, phase of flight, and other circumstances where the demand 
for thrust is critical to safe operation of the aircraft.
    Liquid and Gas Systems: Proposed special condition no. 19 would 
require 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.
    Vibration Demonstration: Proposed special condition no. 20 would 
require 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: Proposed special condition no. 21 would require 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 proposed electric engine converts electrical energy to shaft 
torque, which is used for propulsion. The electric motor, controller, 
and high-voltage systems control the engine torque. When the pilot 
commands power or thrust, the engine responds to the command and 
adjusts the shaft torque to meet the demand. During the transition from 
one power or thrust setting to another, a small delay, or latency, 
occurs in the engine response time. While the engine dwells in this 
time interval, it can continue to apply torque until the command to 
change the torque is applied by the engine control. The allowable 
amount of overtorque during operation depends on the engine's response 
to changes in the torque command throughout its operating range.
    Calibration Assurance: Proposed special condition no. 22 would 
require 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 proposed special 
condition nos. 23 and 26. The calibration test requirements specified 
in Sec.  33.85 only apply to the endurance test specified in Sec.  
33.87, which is applicable only to turbine engines. The FAA proposes 
that the methods used for accomplishing those tests for turbine engines 
is not the best approach for electric engines. The calibration tests in 
Sec.  33.85 have provisions applicable to ratings that are not relevant 
to the BETA Model H500A engines. Proposed special condition no. 22 
would allow 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 
proposed special condition applies the calibration requirement to both 
the endurance and durability tests.

[[Page 16480]]

    Endurance Demonstration: Proposed special condition no. 23 would 
require 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: Proposed special condition no. 24 would require 
BETA to ensure the engine can endure operation at its temperature 
limits plus an acceptable margin. An ``acceptable margin,'' as used in 
the proposed special condition, is the amount of temperature above that 
required to prevent the least capable engine allowed by the type 
design, as determined by Sec.  33.8, from failing due to temperature-
related causes when operating at the most extreme engine and 
environmental thermal conditions.
    Operation Demonstration: Proposed special condition no. 25 would 
require the engine to demonstrate safe operating characteristics 
throughout its declared flight envelope and operating range. Engine 
operating characteristics define the range of functional and 
performance values the BETA Model H500A 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 proposed engine ratings and 
operating limits. Therefore, the installed performance defines the 
engine performance capabilities.
    Durability Demonstration: Proposed special condition no. 26 would 
require 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: Proposed special condition no. 27 would 
require BETA to show that the systems and components of the engine 
would perform their intended functions in all declared engine 
environments and operating conditions.
    Sections 33.87 and 33.91, which are specifically applicable to 
turbine engines, have conditional criteria to decide if additional 
tests will be required after the engine tests. The criteria are not 
suitable for electric engines. Part 33 associates the need for 
additional testing with the outcome of the Sec.  33.87 endurance test 
because it is designed to address safety concerns in combustion 
engines. For example, Sec.  33.91(b) requires the establishment of 
temperature limits for components that require temperature-controlling 
provisions, and Sec.  33.91(a) requires additional testing of engine 
systems and components where the endurance test does not fully expose 
internal systems and components to thermal conditions that verify the 
desired operating limits. Exceeding temperature limits is a safety 
concern for electric engines. The FAA proposes that the Sec.  33.87 
endurance test might not be the best way to achieve the highest thermal 
conditions for all the electronic components of electric engines 
because heat is generated differently in electronic systems than it is 
in turbine engines. Additional safety considerations also need to be 
addressed in the test. Therefore, proposed special condition no. 27 
would be a performance-based requirement that allows 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: Proposed special condition no. 28 
would require the engine to demonstrate reliable rotor locking 
performance and that no hazardous effects will occur if the engine uses 
a rotor locking device to prevent shaft rotation.
    Some engine designs enable the pilot to prevent a propeller shaft 
or main rotor shaft from turning while the engine is running, or the 
aircraft is in-flight. This capability is needed for some installations 
that require the pilot to confirm functionality of certain flight 
systems before takeoff. The proposed 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 proposed special condition does not define ``reliable'' rotor 
locking but would allow 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: Proposed special condition no. 29 would 
require BETA to perform a teardown or non-teardown evaluation after the 
endurance, durability, and overtorque demonstrations, based on the 
criteria proposed in special condition no. 29(a) or (b).
    Proposed special condition no. 29(b) includes restrictive criteria 
for ``non-teardown evaluations'' to account for electric engines, sub-
assemblies, and components that cannot be disassembled without 
destroying them. Some electrical and electronic components like BETA's 
are constructed in an integrated fashion that precludes the possibility 
of tearing them down without destroying them. The proposed special 
condition indicates that, if a teardown cannot be performed in a non-
destructive manner, then the inspection or replacement intervals must 
be established based on the endurance and durability demonstrations. 
The

[[Page 16481]]

procedure for establishing maintenance should be agreed upon between 
the applicant and the FAA prior to running the relevant tests. Data 
from the endurance and durability tests may provide information that 
can be used to determine maintenance intervals and life limits for 
parts. However, if life limits are required, the lifing procedure is 
established by special condition no. 13, Critical and Life-Limited 
Parts, which corresponds to Sec.  33.70. Therefore, the procedure used 
to determine which parts are life-limited, and how the life limits are 
established, requires FAA approval, as it does for Sec.  33.70. 
Sections 33.55 and 33.93 do not contain similar requirements because 
reciprocating and turbine engines can be completely disassembled for 
inspection.
    Containment: Proposed special condition no. 30 would require the 
engine to have containment features that protect against likely hazards 
from rotating components, unless 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 proposed special condition 
would require 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 proposed special condition is 
to prevent hazardous engine effects from structural failure of rotating 
components and parts that are built into the rotor assembly.
    Operation with a Variable Pitch Propeller: Proposed special 
condition no. 31 would require BETA to conduct functional 
demonstrations, including feathering, negative torque, negative thrust, 
and reverse thrust operations, as applicable, based on the propeller's 
or fan's variable pitch functions that are planned for use on these 
electric engines, using a representative propeller. The requirements of 
Sec.  33.95 prescribe tests based on the operating characteristics of 
turbine engines equipped with variable pitch propellers, which include 
thrust response times, engine stall, propeller shaft overload, loss of 
thrust control, and hardware fatigue. The electric engines BETA 
proposes have different operating characteristics that substantially 
affect their susceptibility to these and other potential failures 
typical of turbine engines. Because BETA's proposed electric engines 
may be installed with a variable pitch propeller, the proposed special 
condition is necessary.
    General Conduct of Tests: Proposed special condition no. 32 would 
require BETA to--
    (1) include scheduled maintenance in the engine ICA;
    (2) include any maintenance, in addition to the scheduled 
maintenance, that was needed during the test to satisfy the applicable 
test requirements; and
    (3) conduct any additional tests that the Administrator finds 
necessary, as warranted by the test results.
    For example, certification endurance test shortfalls might be 
caused by omitting some prescribed engine test conditions, or from 
accelerated deterioration of individual parts arising from the need to 
force the engine to operating conditions that drive the engine above 
the engine cycle values of the type design. If an engine part fails 
during a certification test, the entire engine might be subjected to 
penalty runs, with a replacement or newer part design installed on the 
engine, to meet the test requirements. Also, the maintenance performed 
to replace the part, so that the engine could complete the test, would 
be included in the engine ICA. In another example, if the applicant 
replaces a part before completing an engine certification test because 
of a test facility failure and can substantiate the part to the 
Administrator through bench testing, they might not need to 
substantiate the part design using penalty runs with the entire engine.
    The term ``excessive'' is used to describe the frequency of 
unplanned engine maintenance, and the frequency of unplanned test 
stoppages, to address engine issues that prevent the engine from 
completing the tests in proposed special condition nos. 32(b)(1) and 
(2), respectively. Excessive frequency is an objective assessment from 
the FAA's analysis of the amount of unplanned maintenance needed for an 
engine to complete a certification test. The FAA's assessment may 
include the reasons for the unplanned maintenance, such as the effects 
test facility equipment may have on the engine, the inability to 
simulate a realistic engine operating environment, and the extent to 
which an engine requires modifications to complete a certification 
test. In some cases, the applicant may be able to show that unplanned 
maintenance has no effect on the certification test results, or they 
might be able to attribute the problem to the facility or test-enabling 
equipment that is not part of the type design. In these cases, the ICA 
will not be affected. However, if 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. Proposed special 
condition no. 33 is established using the existing engine control 
systems requirement as a basis. It applies applicable airworthiness 
criteria from Sec.  33.28 and incorporates airworthiness criteria that 
recognize and focus on the electrical power system used in the engine.
    Proposed special condition no. 33(b) would ensure that all aspects 
of an electrical system, including generation, distribution, and usage, 
do not experience any unacceptable operating characteristics.
    Proposed special condition no. 33(c) would require the electrical 
power distribution aspects of the electrical system to provide the safe 
transfer of electrical energy throughout the electric engine.
    Proposed special condition no. 33(d) would require the engine 
electrical system to be designed such that the loss, malfunction, or 
interruption of the electrical power source, or power conditions that 
exceed design limits,

[[Page 16482]]

will not result in a hazardous engine effect.
    Proposed 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.
    Proposed 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 
proposed special condition include temperature, vibration, HIRF, and 
others addressed in RTCA DO-160G, ``Environmental Conditions and Test 
Procedures for Airborne Electronic/Electrical Equipment and 
Instruments.''
    Proposed special condition 33(g) would require BETA to evaluate 
various electric engine system failures to ensure that these failures 
will not lead to unsafe engine conditions. The evaluation would include 
single-fault tolerance, would ensure no single electrical or electronic 
fault or failure would result in hazardous engine effects, and ensure 
that any failure or malfunction leading to local events in the intended 
aircraft application do not result in certain hazardous engine effects. 
The special condition would also implement integrity requirements, 
criteria for LOTC/LOPC events, and an acceptable LOTC/LOPC rate.
    Proposed special condition 33(h) would require 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, that can be 
used at the aircraft safety assessment level.
    These proposed special conditions contain the additional safety 
standards that the Administrator considers necessary to establish a 
level of safety equivalent to that established by the existing 
airworthiness standards for reciprocating and turbine aircraft engines.

Applicability

    As discussed above, these proposed special conditions are 
applicable to BETA Model H500A 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 engines. It is not a rule 
of general applicability.

List of Subjects in 14 CFR Part 33

    Aircraft, Aviation safety, Reporting and recordkeeping 
requirements.

Authority Citation

    The authority citation for these special conditions is as follows:

    Authority:  49 U.S.C. 106(f), 106(g), 40113, 44701, 44702, 
44704.

The Proposed Special Conditions

0
 Accordingly, the Federal Aviation Administration (FAA) proposes the 
following special conditions as part of the type certification basis 
for BETA Technologies Inc. Model H500A engines. The applicant must also 
comply with the certification procedures set forth in title 14, Code of 
Federal Regulations (14 CFR) part 21.

(1) Applicability

    (a) Unless otherwise noted in these special conditions, the engine 
design must comply with the airworthiness standards for aircraft 
engines set forth in 14 CFR part 33, except for those airworthiness 
standards that are specifically and explicitly applicable only to 
reciprocating and turbine aircraft engines or as specified herein.
    (b) The applicant must comply with this part using a means of 
compliance, which may include consensus standards, accepted by the 
Administrator.
    (c) The applicant requesting acceptance of a means of compliance 
must provide the means of compliance to the FAA in a form and manner 
acceptable to the Administrator.

(2) Engine Ratings and Operating Limits

    In addition to Sec.  33.7(a), the engine ratings and operating 
limits must be established and included in the type certificate data 
sheet based on:
    (a) Shaft power, torque, rotational speed, and temperature for:
    (1) Rated takeoff power;
    (2) Rated maximum continuous power; and
    (3) Rated maximum temporary power and associated time limit.
    (b) Duty cycle and the rating at that duty cycle. The duty cycle 
must be declared in the engine type certificate data sheet.
    (c) Cooling fluid grade or specification.
    (d) Power-supply requirements.
    (e) Any other ratings or limitations that are necessary for the 
safe operation of the engine.

(3) Materials

    The engine design must comply with Sec.  33.15.

(4) Fire Protection

    The engine design must comply with Sec.  33.17(b) through (g).
    (a) The design and construction of the engine and the materials 
used must minimize the probability of the occurrence and spread of fire 
during normal operation and failure conditions and must minimize the 
effect of such a fire.
    (b) High-voltage electrical wiring interconnect systems must be 
protected against arc faults that can lead to hazardous engine effects 
as defined in special condition no. 17(d)(2) of these special 
conditions. Any non-protected electrical wiring interconnects must be 
analyzed to show that arc faults do not cause a hazardous engine 
effect.

(5) Durability

    The engine design and construction must minimize the development of 
an unsafe condition of the engine between maintenance intervals, 
overhaul periods, or mandatory actions described in the applicable ICA.

(6) Engine Cooling

    The engine design and construction must comply with Sec.  33.21. In 
addition, if cooling is required to satisfy the safety analysis as 
described in special condition no. 17 of these special conditions, the 
cooling system monitoring features and usage must be documented in the 
engine installation manual.

(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 16483]]

certified operating conditions and above failure conditions leading to 
rotor overspeed. The margin to burst must be shown by test, validated 
analysis, or a combination thereof.
    (c) The engine must not exceed the rotor speed operational 
limitations that could affect rotor structural integrity.

(10) Engine Control Systems

    (a) Applicability. The requirements of this special condition apply 
to any system or device that is part of the engine type design that 
controls, limits, monitors, or protects engine operation, and is 
necessary for the continued airworthiness of the engine.
    (b) Engine control. The engine control system must ensure that the 
engine does not experience any unacceptable operating characteristics 
or exceed its operating limits, including in failure conditions where 
the fault or failure results in a change from one control mode to 
another, from one channel to another, or from the primary system to the 
back-up system, if applicable.
    (c) Design Assurance. The software and complex electronic hardware, 
including programmable logic devices, must be--
    (1) Designed and developed using a structured and systematic 
approach that provides a level of assurance for the logic commensurate 
with the hazard associated with the failure or malfunction of the 
systems in which the devices are located; and
    (2) Substantiated by a verification methodology acceptable to the 
Administrator.
    (d) Validation. All functional aspects of the control system must 
be substantiated by test, analysis, or a combination thereof, to show 
that the engine control system performs the intended functions 
throughout the declared operational envelope.
    (e) Environmental Limits. Environmental limits that cannot be 
adequately substantiated by endurance demonstration, validated 
analysis, or a combination thereof must be demonstrated by the system 
and component tests in special condition no. 27 of these special 
conditions.
    (f) Engine control system failures. The engine control system 
must--
    (1) Have a maximum rate of loss of power control (LOPC) that is 
suitable for the intended aircraft application. The estimated LOPC rate 
must be specified in the engine installation manual;
    (2) When in the full-up configuration, be single-fault tolerant, as 
determined by the Administrator, for electrical, electrically 
detectable, and electronic failures involving LOPC events;
    (3) Not have any single failure that results in hazardous engine 
effects as defined in special condition no. 17(d)(2) of these special 
conditions; and
    (4) Ensure failures or malfunctions that lead to local events in 
the aircraft do not result in hazardous engine effects, as defined in 
special condition no. 17(d)(2) of these special conditions, due to 
engine control system failures or malfunctions.
    (g) System safety assessment. The applicant must perform a system 
safety assessment. This assessment must identify faults or failures 
that affect normal operation, together with the predicted frequency of 
occurrence of these faults or failures. The intended aircraft 
application must be taken into account to assure that the assessment of 
the engine control system safety is valid. The rates of hazardous and 
major faults must be declared in the engine installation manual.
    (h) Protection systems. The engine control devices and systems' 
design and function, together with engine instruments, operating 
instructions, and maintenance instructions, must ensure that engine 
operating limits that can lead to a hazard will not be exceeded in 
service.
    (i) Aircraft supplied data. Any single failure leading to loss, 
interruption, or corruption of aircraft-supplied data (other than 
power-command signals from the aircraft), or aircraft-supplied data 
shared between engine systems within a single engine or between fully 
independent engine systems, must--
    (1) Not result in a hazardous engine effect, as defined in special 
condition no. 17(d)(2) of these special conditions, for any engine 
installed on the aircraft; and
    (2) Be able to be detected and accommodated by the control system.
    (j) Engine control system electrical power.
    (1) The engine control system must be designed such that the loss, 
malfunction, or interruption of the control system electrical power 
source will not result in a hazardous engine effect, unacceptable 
transmission of erroneous data, or continued engine operation in the 
absence of the control function. Hazardous engine effects are defined 
in special condition no. 17(d)(2) of these special conditions. The 
engine control system must be capable of resuming normal operation when 
aircraft-supplied power returns to within the declared limits.
    (2) The applicant must identify and declare, in the engine 
installation manual, the characteristics of any electrical power 
supplied from the aircraft to the engine control system, including 
transient and steady-state voltage limits, and any other 
characteristics necessary for safe operation of the engine.

(11) Instrument Connection

    The applicant must comply with Sec.  33.29(a), (e), and (g).
    (a) In addition, as part of the system safety assessment of special 
condition nos. 10(g) and 33(h) of these special conditions, the 
applicant must assess the possibility and subsequent effect of 
incorrect fit of instruments, sensors, or connectors. Where 
practicable, the applicant must take design precautions to prevent 
incorrect configuration of the system.
    (b) The applicant must provide instrumentation enabling the flight 
crew to monitor the functioning of the engine cooling system unless 
evidence shows that:
    (1) Other existing instrumentation provides adequate warning of 
failure or impending failure;
    (2) Failure of the cooling system would not lead to hazardous 
engine effects before detection; or
    (3) The probability of failure of the cooling system is extremely 
remote.

(12) Stress Analysis

    (a) A mechanical and thermal stress analysis, as well as an 
analysis of the stress caused by electromagnetic forces, must show a 
sufficient design margin to prevent unacceptable operating 
characteristics and hazardous engine effects as defined in special 
condition no. 17(d)(2) of these special conditions.
    (b) Maximum stresses in the engine must be determined by test, 
validated analysis, or a combination thereof, and must be shown not to 
exceed minimum material properties.

(13) Critical and Life-Limited Parts

    (a) The applicant must show, by a safety analysis or means 
acceptable to the Administrator, whether rotating or moving components, 
bearings, shafts, static parts, and non-redundant mount components 
should be classified, designed, manufactured, and managed throughout 
their service life as critical or life-limited parts.
    (1) Critical part means a part that must meet prescribed integrity 
specifications to avoid its primary failure, which is likely to result 
in a hazardous engine effect as defined in special condition no. 
17(d)(2) of these special conditions.
    (2) Life-limited parts may include but are not limited to a rotor 
or major structural static part, the failure of which can result in a 
hazardous engine effect, as defined in special condition

[[Page 16484]]

no. 17(d)(2) of these special conditions, due to a low-cycle fatigue 
(LCF) mechanism. A life limit is an operational limitation that 
specifies the maximum allowable number of flight cycles that a part can 
endure before the applicant must remove it from the engine.
    (b) In establishing the integrity of each critical part or life-
limited part, the applicant must provide to the Administrator the 
following three plans for approval:
    (1) an engineering plan, as defined in Sec.  33.70 (a);
    (2) a manufacturing plan, as defined in Sec.  33.70 (b); and
    (3) a service-management plan, as defined in Sec.  33.70 (c).

(14) Lubrication System

    (a) The lubrication system must be designed and constructed to 
function properly between scheduled maintenance intervals in all flight 
attitudes and atmospheric conditions in which the engine is expected to 
operate.
    (b) The lubrication system must be designed to prevent 
contamination of the engine bearings and lubrication system components.
    (c) The applicant must demonstrate by test, validated analysis, or 
a combination thereof, the unique lubrication attributes and functional 
capability of (a) and (b).

(15) Power Response

    (a) The design and construction of the engine, including its 
control system, must enable an increase--
    (1) From the minimum power setting to the highest rated power 
without detrimental engine effects;
    (2) From the minimum obtainable power while in-flight and while on 
the ground to the highest rated power within a time interval determined 
to be appropriate for the intended aircraft application; and
    (3) From the minimum torque to the highest rated torque without 
detrimental engine effects in the intended aircraft application.
    (b) The results of (a)(1), (a)(2), and (a)(3) of this special 
condition must be included in the engine installation manual.

(16) Continued Rotation

    If the design allows any of the engine main rotating systems to 
continue to rotate after the engine is shut down while in-flight, this 
continued rotation must not result in any hazardous engine effects, as 
defined in special condition no. 17(d)(2) of these special conditions.

(17) Safety Analysis

    (a) The applicant must comply with Sec.  33.75(a)(1) and (a)(2) 
using the failure definitions in special condition no. 17(d) of these 
special conditions.
    (b) The primary failure of certain single elements cannot be 
sensibly estimated in numerical terms. If the failure of such elements 
is likely to result in hazardous engine effects, then compliance may be 
shown by reliance on the prescribed integrity requirements of Sec.  
33.15 and special condition nos. 9 and 13 of these special conditions, 
as applicable. These instances must be stated in the safety analysis.
    (c) The applicant must comply with Sec.  33.75(d) and (e) using the 
failure definitions in special condition no. 17(d) of these special 
conditions, and the ICA in Sec.  33.4.
    (d) Unless otherwise approved by the Administrator, the following 
definitions apply to the engine effects when showing compliance with 
this condition:
    (1) A minor engine effect does not prohibit the engine from 
performing its intended functions in a manner consistent with Sec.  
33.28(b)(1)(i), (b)(1)(iii), and (b)(1)(iv), and the engine complies 
with the operability requirements of special condition no. 15 and 
special condition no. 25 of these special conditions, as appropriate.
    (2) The engine effects in Sec.  33.75(g)(2) are hazardous engine 
effects with the addition of:
    (i) Electrocution of the crew, passengers, operators, maintainers, 
or others; and
    (ii) Blockage of cooling systems that could cause the engine 
effects described in Sec.  33.75(g)(2) and special condition 
17(d)(2)(i) of these special conditions.
    (3) Any other engine effect is a major engine effect.
    (e) The intended aircraft application must be taken into account 
when performing the safety analysis.
    (f) The results of the safety analysis, and the assumptions about 
the aircraft application used in the safety analysis, must be 
documented in the engine installation manual.

(18) Ingestion

    (a) Rain, ice, and hail ingestion must not result in an abnormal 
operation such as shutdown, power loss, erratic operation, or power 
oscillations throughout the engine operating range.
    (b) Ingestion from other likely sources (birds, induction system 
ice, foreign objects--ice slabs) must not result in hazardous engine 
effects defined by special condition no. 17(d)(2) of these special 
conditions, or unacceptable power loss.
    (c) If the design of the engine relies on features, attachments, or 
systems that the installer may supply, for the prevention of 
unacceptable power loss or hazardous engine effects, as defined in 
special condition no. 17(d)(2) of these special conditions, following 
potential ingestion, then the features, attachments, or systems must be 
documented in the engine installation manual.

(19) Liquid and Gas Systems

    (a) Each system used for lubrication or cooling of engine 
components must be designed and constructed to function properly in all 
flight attitudes and atmospheric conditions in which the engine is 
expected to operate.
    (b) If a system used for lubrication or cooling of engine 
components is not self-contained, the interfaces to that system must be 
defined in the engine installation manual.
    (c) The applicant must establish by test, validated analysis, or a 
combination of both that all static parts subject to significant 
pressure loads will not:
    (1) Exhibit permanent distortion beyond serviceable limits, or 
exhibit leakage that could create a hazardous condition when subjected 
to normal and maximum working pressure with margin;
    (2) Exhibit fracture or burst when subjected to the greater of 
maximum possible pressures with margin.
    (d) Compliance with special condition no. 19(c) of these special 
conditions must take into account:
    (1) The operating temperature of the part;
    (2) Any other significant static loads in addition to pressure 
loads;
    (3) Minimum properties representative of both the material and the 
processes used in the construction of the part; and
    (4) Any adverse physical geometry conditions allowed by the type 
design, such as minimum material and minimum radii.
    (e) Approved coolants and lubricants must be listed in the engine 
installation manual.

(20) Vibration Demonstration

    (a) The engine must be designed and constructed to function 
throughout its normal operating range of rotor speeds and engine output 
power, including defined exceedances, without inducing excessive stress 
in any of the engine parts because of vibration and without imparting 
excessive vibration forces to the aircraft structure.
    (b) Each engine design must undergo a vibration survey to establish 
that the

[[Page 16485]]

vibration characteristics of those components subject to induced 
vibration are acceptable throughout the declared flight envelope and 
engine operating range for the specific installation configuration. The 
possible sources of the induced vibration that the survey must assess 
are mechanical, aerodynamic, acoustical, internally induced 
electromagnetic, installation induced effects that can affect the 
engine vibration characteristics, and likely environmental effects. 
This survey must be shown by test, validated analysis, or a combination 
thereof.

(21) Overtorque

    When approval is sought for a transient maximum engine overtorque, 
the applicant must demonstrate by test, validated analysis, or a 
combination thereof, that the engine can continue operation after 
operating at the maximum engine overtorque condition without 
maintenance action. Upon conclusion of overtorque tests conducted to 
show compliance with this special condition, or any other tests that 
are conducted in combination with the overtorque test, each engine part 
or individual groups of components must meet the requirements of 
special condition no. 29 of these special conditions.

(22) Calibration Assurance

    Each engine must be subjected to calibration tests to establish its 
power characteristics, and the conditions both before and after the 
endurance and durability demonstrations specified in special conditions 
nos. 23 and 26 of these special conditions.

(23) Endurance Demonstration

    The applicant must subject the engine to an endurance 
demonstration, acceptable to the Administrator, to demonstrate the 
engine's limit capabilities. The endurance demonstration must include 
increases and decreases of the engine's power settings, energy 
regeneration, and dwellings at the power settings or energy 
regeneration for sufficient durations that produce the extreme physical 
conditions the engine experiences at rated performance levels, 
operational limits, and at any other conditions or power settings that 
are required to verify the limit capabilities of the engine.

(24) Temperature Limit

    The engine design must demonstrate its capability to endure 
operation at its temperature limits plus an acceptable margin. The 
applicant must quantify and justify the margin to the Administrator. 
The demonstration must be repeated for all declared duty cycles and 
ratings, and operating environments, that would impact temperature 
limits.

(25) Operation Demonstration

    The engine design must demonstrate safe operating characteristics, 
including but not limited to power cycling, starting, acceleration, and 
overspeeding throughout its declared flight envelope and operating 
range. The declared engine operational characteristics must account for 
installation loads and effects.

(26) Durability Demonstration

    The engine must be subjected to a durability demonstration to show 
that each part of the engine has been designed and constructed to 
minimize any unsafe condition of the system between overhaul periods, 
or between engine replacement intervals if the overhaul is not defined. 
This test must simulate the conditions in which the engine is expected 
to operate in service, including typical start-stop cycles, to 
establish when the initial maintenance is required.

(27) System and Component Tests

    The applicant must show that systems and components that cannot be 
adequately substantiated in accordance with the endurance demonstration 
or other demonstrations will perform their intended functions in all 
declared environmental and operating conditions.

(28) Rotor Locking Demonstration

    If shaft rotation is prevented by locking the rotor(s), the engine 
must demonstrate:
    (a) Reliable rotor locking performance;
    (b) Reliable rotor unlocking performance; and
    (c) That no hazardous engine effects, as specified in special 
condition no. 17(d)(2) of these special conditions, will occur.

(29) Teardown Inspection

    (a) Teardown evaluation.
    (1) After the endurance and durability demonstrations have been 
completed, the engine must be completely disassembled. Each engine 
component and lubricant must be eligible for continued operation in 
accordance with the information submitted for showing compliance with 
Sec.  33.4.
    (2) Each engine component, having an adjustment setting and a 
functioning characteristic that can be established independent of 
installation on or in the engine, must retain each setting and 
functioning characteristic within the established and recorded limits 
at the beginning of the endurance and durability demonstrations.
    (b) Non-Teardown evaluation. If a teardown cannot be performed for 
all engine components in a non-destructive manner, then the inspection 
or replacement intervals for these components and lubricants must be 
established based on the endurance and durability demonstrations and 
must be documented in the ICA in accordance with Sec.  33.4.

(30) Containment

    The engine must be designed and constructed to protect against 
likely hazards from rotating components as follows--
    (a) The design of the stator case surrounding rotating components 
must provide for the containment of the rotating components in the 
event of failure, unless the applicant shows that the margin to rotor 
burst precludes the possibility of a rotor burst.
    (b) If the margin to burst shows that the stator case must have 
containment features in the event of failure, then the stator case must 
provide for the containment of the failed rotating components. The 
applicant must define by test, validated analysis, or a combination 
thereof, and document, in the engine installation manual, the energy 
level, trajectory, and size of fragments released from damage caused by 
the main-rotor failure, and that pass forward or aft of the surrounding 
stator case.

(31) Operation With Variable Pitch Propeller

    The applicant must conduct functional demonstrations including 
feathering, negative torque, negative thrust, and reverse thrust 
operations, as applicable, with a representative propeller. These 
demonstrations may be conducted in a manner acceptable to the 
Administrator as part of the endurance, durability, and operation 
demonstrations.

(32) General Conduct of Tests

    (a) Maintenance of the engine may be made during the tests in 
accordance with the service and maintenance instructions submitted in 
compliance with Sec.  33.4.
    (b) The applicant must subject the engine or its parts to any 
additional tests that the Administrator finds necessary if--
    (1) The frequency of engine service is excessive;
    (2) The number of stops due to engine malfunction is excessive;

[[Page 16486]]

    (3) Major engine repairs are needed; or
    (4) Replacement of an engine part is found necessary during the 
tests, or due to the teardown inspection findings.
    (c) Upon completion of all demonstrations and testing specified in 
these special conditions, the engine and its components must be--
    (1) Within serviceable limits;
    (2) Safe for continued operation; and
    (3) Capable of operating at declared ratings while remaining within 
limits.

(33) Engine Electrical Systems

    (a) Applicability. Any system or device that provides, uses, 
conditions, or distributes electrical power, and is part of the engine 
type design, must provide for the continued airworthiness of the 
engine, and must maintain electric engine ratings.
    (b) Electrical systems. The electrical system must ensure the safe 
generation and transmission of power, and electrical load shedding, and 
that the engine does not experience any unacceptable operating 
characteristics or exceed its operating limits.
    (c) Electrical power distribution.
    (1) The engine electrical power distribution system must be 
designed to provide the safe transfer of electrical energy throughout 
the electrical power plant. The system must be designed to provide 
electrical power so that the loss, malfunction, or interruption of the 
electrical power source will not result in a hazardous engine effect, 
as defined in special condition no. 17(d)(2) of these special 
conditions or detrimental engine effects in the intended aircraft 
application.
    (2) The system must be designed and maintained to withstand normal 
and abnormal conditions during all ground and flight operations.
    (3) The system must provide mechanical or automatic means of 
isolating a faulted electrical energy generation or storage device from 
affecting the safe transmission of electric energy to the electric 
engine.
    (d) Protection systems. The engine electrical system must be 
designed such that the loss, malfunction, interruption of the 
electrical power source, or power conditions that exceed design limits, 
will not result in a hazardous engine effect, as defined in special 
condition no. 17(d)(2) of these special conditions.
    (e) Electrical power characteristics. The applicant must identify 
and declare, in the engine installation manual, the characteristics of 
any electrical power supplied from--
    (1) the aircraft to the engine electrical system, for starting and 
operating the engine, including transient and steady-state voltage 
limits, or
    (2) the engine to the aircraft via energy regeneration, and any 
other characteristics necessary for safe operation of the engine.
    (f) Environmental limits. Environmental limits that cannot 
adequately be substantiated by endurance demonstration, validated 
analysis, or a combination thereof must be demonstrated by the system 
and component tests in special condition no. 27 of these special 
conditions.
    (g) Electrical system failures. The engine electrical system must--
    (1) Have a maximum rate of loss of power control (LOPC) that is 
suitable for the intended aircraft application;
    (2) When in the full-up configuration, be single-fault tolerant, as 
determined by the Administrator, for electrical, electrically 
detectable, and electronic failures involving LOPC events;
    (3) Not have any single failure that results in hazardous engine 
effects; and
    (4) Ensure failures or malfunctions that lead to local events in 
the intended aircraft application do not result in hazardous engine 
effects, as defined in special condition no. 17(d)(2) of these special 
conditions, due to electrical system failures or malfunctions.
    (h) System safety assessment. The applicant must perform a system 
safety assessment. This assessment must identify faults or failures 
that affect normal operation, together with the predicted frequency of 
occurrence of these faults or failures. The intended aircraft 
application must be taken into account to assure the assessment of the 
engine system safety is valid.

    Issued in Kansas City, Missouri, on March 1, 2024.
Patrick R. Mullen,
Manager, Technical Policy Branch, Policy and Standards Division, 
Aircraft Certification Service.
[FR Doc. 2024-04800 Filed 3-6-24; 8:45 am]
BILLING CODE 4910-13-P