[Federal Register Volume 85, Number 224 (Thursday, November 19, 2020)]
[Proposed Rules]
[Pages 73644-73655]
[FR Doc No: 2020-23434]

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DEPARTMENT OF TRANSPORTATION

Federal Aviation Administration

14 CFR Part 33

[Docket No. FAA-2020-0894; Notice No. 33-19-01-SC]

Special Conditions: magniX USA, Inc., magni250 and magni500 Model
Engines

AGENCY: Federal Aviation Administration (FAA), DOT.

ACTION: Notice of proposed special conditions.

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SUMMARY: This action proposes special conditions for magniX USA, Inc.
(magniX), magni250 and magni500 model 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, 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 December 21, 2020.

ADDRESSES: Send comments identified by Docket No. FAA-2020-0894 using
any of the following methods:
     Federal eRegulations Portal: Go to http://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 (DOT), 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<PHONE_NUMBER>.
    Privacy: Except for Confidential Business Information (CBI) as
described in the following paragraph, and other information as
described in 14 CFR 11.35, the FAA will post all comments received,
without change, to http://www.regulations.gov/, including any personal
information you provide. The FAA will also post a report summarizing
each substantive verbal contact we received about this proposal.

Confidential Business Information

    Confidential Business Information (CBI) is commercial or financial
information that is both customarily and actually treated as private by
its owner. Under the Freedom of Information Act (FOIA) (5 U.S.C. 552),
CBI is exempt from public disclosure. If your comments responsive to
this Notice contain commercial or financial information that is
customarily treated as private, that you actually treat as private, and
that is relevant or responsive to this Notice, 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 they will
not be placed in the public docket of this Notice. Submissions
containing CBI should be sent to Gary Horan, AIR-6A1, Engine and
Propeller Standards Branch, Aircraft Certification Service, 1200
District Avenue, Burlington, Massachusetts 01803; telephone (781) 238-
7164<EMAIL_ADDRESS>Any commentary that the FAA receives which is
not specifically designated as CBI will be placed in the public docket
for this rulemaking.
    Docket: Background documents or comments received may be read at
http://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: Gary Horan, AIR-6A1, Engine and
Propeller Standards Branch, Aircraft Certification Service, 1200
District Avenue, Burlington, Massachusetts 01803; telephone (781) 238-
7164<EMAIL_ADDRESS>
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.

Background

    On June 4, 2019, magniX applied for a type certificate for its
magni250 and magni500 model electric engines. The FAA has not
previously type certificated an engine that uses electrical technology
for propulsion of the aircraft. 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.
    There is a growing interest within the aviation industry to utilize
electric propulsion technology. As a result, international agencies and
industry stakeholders formed a new committee under ASTM International
Committee F39 to identify the appropriate technical criteria for
aircraft engines using electrical technology that has not been
previously certificated for aircraft propulsion systems. ASTM
International, formerly known as American Society for Testing and
Materials, 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 standard and engine
information from magniX to develop special conditions to establish an
equivalent level of safety to that required by title 14, Code of
Federal Regulations (14 CFR) part 33.
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    \1\ https://www.astm.org/Standards/F3338.htm.
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Type Certification Basis

    Under the provisions of 14 CFR 21.17(a)(1), generally, magniX must
show that magni250 and magni500 model engines meet the applicable
provisions of part 33 in effect on the date of application for a type
certificate.
    If the Administrator finds that the applicable airworthiness
regulations (e.g., 14 CFR part 33) do not contain adequate or
appropriate safety standards for the magni250 and magni500 model
engines because of a novel or unusual design feature, special
conditions may be prescribed under the provisions of Sec.  21.16.
    Special conditions are initially applicable to the model for which
they are issued. Should the type certificate for that model be amended
later to include any other engine model that incorporates the same
novel or unusual design feature, these special conditions would also
apply to the other engine model under Sec.  21.101.
    In addition to the applicable airworthiness regulations and special
conditions, the magni250 and magni500 model engines must comply with
the noise certification requirements of 14 CFR part 36.
    The FAA issues special conditions, as defined in 14 CFR 11.19, in
accordance with Sec.  11.38, and they become part of the type
certification basis under Sec.  21.17(a)(2).

Novel or Unusual Design Features

    The magni250 and magni500 model engines will incorporate the
following novel or unusual design features:
    An electric motor, controller, and high-voltage systems that are
used as the primary source of propulsion for an aircraft.

Discussion

Part 33 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 certified 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 air 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 to design and manufacture 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.

magniX's Proposed Electric Engines Are Novel or Unusual

    The existing part 33 airworthiness standards for aircraft engines
date back to 1965. These airworthiness standards are based on fuel-
burning reciprocating and turbine engine technology. The magni250 and
magni500 model engines are not turbine or 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 aircraft engine is
also exposed to chemical, thermal, and mechanical operating conditions,
unlike those observed in internal combustion systems. Therefore, part
33 does not contain adequate or appropriate safety standards for the
magni250 and magni500 model engine's novel design feature.
    magniX'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, controller, and high-voltage systems
that draw energy from electrical storage or generating systems. The
electric motor is a device that converts electrical energy into
mechanical energy by electric current flowing through wire coils in the
motor producing a magnetic field that interacts with the magnets on the
rotating shaft. 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 system is a
combination of wires and the connectors that couple the motor and the
controller.
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    \2\ Sometimes this entire system is referred to as an inverter.
Throughout this document, it will be referred to as the controller.
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    In addition, the technology required to produce 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 essential to the electric engine. However,
this technology also exposes the aircraft to potential failures that
are not common to gas-powered turbine and reciprocating engines, which
could adversely affect safety.

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

    Although the electric aircraft engines proposed by magniX use novel
or unusual design features that are not addressed in the existing part
33 airworthiness standards, there are some basic similarities in
configuration and function that require similar provisions to prevent
hazards that are common to aircraft engines using air and fuel
combustion (e.g., fire, uncontained high-energy debris, and loss of
thrust control). However, the primary failure concerns and the
probability of exposure to common hazards are different for the
proposed electric aircraft engines. This creates a need to develop
special conditions to ensure the engine's safety and reliability.
    The requirements in part 33 ensure the design and construction of
aircraft engines, including the engine control systems, are proper for
the engine type design and operating limits. However, part 33 does not
fully address the use of aircraft engines like magniX's, which operate
using electrical technology as the primary means of propelling the
aircraft. This necessitates the development of special conditions to
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. This necessitates the development of
special conditions to ensure a level of safety commensurate with these
subparts, as those regulatory requirements do not contain adequate or
appropriate safety standards for aircraft engines that operate using
electrical technology to propel the aircraft.
    The special conditions that the FAA proposes for magniX's engine
design include:
    Applicability: Proposed special condition no. 1 would require
magniX to comply with 14 CFR 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 require magniX, in addition to compliance with 14
CFR 33.7(a), to establish engine operating limits related to the power,
torque, speed, and duty cycles specific to the magni250 and magni500
model 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.
    Materials: Proposed special condition no. 3 would require magniX to
comply with 14 CFR 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
magniX to comply with 14 CFR 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 magniX to ensure 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
proposed engine design and construction to ensure safe engine operation
between maintenance intervals, overhaul periods, and mandatory actions.
This proposed condition would also require magniX to develop
maintenance instructions and scheduling information.
    Engine Cooling: Proposed special condition no. 6 would require
magniX to comply with 14 CFR 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
magniX to document the cooling system monitoring features and usage in
the engine installation manual, in accordance with Sec.  33.5, if
cooling is required to satisfy the safety analysis described in
proposed special condition no. 17. Loss of adequate cooling to an
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 magniX and the proposed design to comply
with 14 CFR 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 14 CFR 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.
    Overspeed: Proposed special condition no. 9 would require magniX 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 speed is driven by hot air expansion and is impacted
by the aerodynamic loads on the rotor blades. Therefore, the speed or
overspeed is not directly controlled in turbine engines. The speed of
an electric engine is directly controlled by the electric field created
by the controller. The failure modes that can lead to overspeed between
turbine engines and these engines are vastly different, and therefore
this special condition is necessary.
    Engine Control Systems: Proposed special condition no. 10(b) would
require magniX 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.
    Proposed special condition no. 10(c) would require magniX to
develop and verify the software and complex electronic hardware used in
programmable logic devices, using proven methods that ensure it 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\ distinguish 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, magniX 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.\4\ Accordingly, proposed special condition
10(e) would require magniX to document the environmental limits to
which the system has been qualified in the engine installation
instructions.
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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 magniX to
evaluate various control system failures to assure that these failures
will not lead to unsafe conditions. The FAA issued Advisory Circular,
AC 33.28-3, Guidance Material For 14 CFR 33.28, Engine Control Systems,
on May 23, 2014.\5\ Paragraph 6-2 of this AC provides applicants with
guidance on defining an engine control system failure when showing
compliance with the requirements of 14 CFR 33.28. AC 33.28-3 also
includes objectives for the integrity requirements, criteria for a loss
of thrust (or power) control (LOTC/LOPC) event, and an acceptable LOTC/
LOPC rate. As with other topics within these proposed special
conditions, the failure rates that apply to electric engines were not
established when the FAA issued this AC.
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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 magniX to
conduct a safety assessment of the control system to support the safety
analysis in special condition no. 17. This control safety assessment
provides failures and rates

of these failures that can be used at the aircraft safety assessment
level.
    Proposed special condition no. 10(h) requires magniX to provide
appropriate protection devices or systems to ensure that engine
operating limitations will not be exceeded in-service.
    Proposed special condition no. 10(i) is necessary to ensure 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 there are
no reasonable means for the engine controller to determine the validity
of any in-range signals from this system. In many cases, the engine
control system can detect a faulty signal from the aircraft. 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 it 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 unacceptable effects
on thrust or power, hazardous engine effects, or inability to comply
with the operation demonstrations in proposed special condition no. 25.
    The term ``accommodated'' in the context of ``detected and
accommodated'' referenced in proposed special condition 10(i)(2) is to
assure that once a fault has been detected, that the system continues
to function safely.
    Proposed special condition no. 10(j) would require magniX 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 magniX to comply with 14 CFR 33.29(a), (e), (f), and (g), which
set certain requirements for the connection and installation of
instruments to monitor engine performance. The remaining requirements
in section 33.29 apply only to technologies used in reciprocating and
turbine aircraft engines.
    Instrument connections (wires, wire insulation, potting, grounding,
connector designs) present opportunities for unsafe features to be
present on the aircraft. Proposed special condition no. 11 would
require the safety analysis to include potential hazardous effects from
failure of instrument connections to function properly. The outcome of
this analysis might identify the need for design enhancements or
additional Instructions for Continued Airworthiness (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
not appropriate for the magniX magni250 and magni500 model engines.
However, the FAA proposes that a stress analysis particular to these
electric engines is necessary.
    Proposed special condition no. 12 would require a mechanical,
thermal, and electrical stress analysis to show there is a sufficient
design margin to prevent unacceptable operating characteristics. Also,
the applicant must determine the maximum stresses in the engine by
tests, validated analysis, or a combination thereof, and show that they
do not exceed minimum material properties.
    Critical and Life-Limited Parts: Proposed special condition no. 13
would require magniX 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 engineering plan referenced in proposed special condition no.
13(b)(1) would require magniX to establish activities for managing
documents, practices, and procedures that govern key design criteria
essential to part airworthiness. The engineering plan would be required
to contain methods for verifying the characteristics and qualities
assumed in the design data using methods that are suitable for the part
criticality. The engineering plan flows information from engineering to
manufacturing about the criticality of key attributes that affect the
airworthiness of the part. The plan also includes a reporting system
that flows problematic issues that develop in engines while they
operate in service so the design process can address them. For example,
the effect of environmental influences on engine performance might not
be consistent with the assumptions used to design the part. The impact
of ice slab ingestion on engine parts might not be fully understood
until the engine ingests the specific ice quantities and shapes that
the airplane sheds. During the pre-certification activities, magniX
must ensure the engineering plan is complete, available, and acceptable
to the Administrator before the engine is certified.
    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 term ``manufacturing definition'' 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 verify that the parts comply with the properties established by the
design data. Since engines are not intentionally tested to failure
during a certification program, there are inherent expectations for
performance and durability guaranteed by the documents and processes
used to execute production and quality systems required by Sec.
21.137. These systems limit the potential manufacturing outcomes to
parts that are consistently produced within design constraints.
    The manufacturing plan and service management plan ensure essential
information from the engineering plan, such as the design
characteristics that ensure 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 has provisions for
enhanced detection and reporting of service-induced anomalies that can
cause the part to fail before it reaches its life limit or service
limit. Anomalies can develop in service from improper handling,
unforeseen operating conditions, and long-term environmental effects.
The service management plan ensures important information that might
affect the assumptions used to design a part is incorporated into the
design process to remove unforeseen potential unsafe features from the
engine.
    Lubrication System: Proposed special condition no. 14 would require
magniX to ensure the lubrication system is designed to function
properly between scheduled maintenance intervals and prevent
contamination of the engine bearings. This proposed condition would
also require magniX to demonstrate the unique lubrication attributes
and functional capability of the magni250 and magni500 model 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 magniX magni250 and magni500
model engines. For example, these engines do not have a crankcase or
lubrication oil sump. The 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 magni250 and magni500 model engines to
enable an increase (1) from the minimum power setting to the highest-
rated power without detrimental engine effects, and (2) from the
minimum obtainable power while in-flight and on the ground to the
highest-rated power within a time interval for safe operation of the
aircraft.
    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 engine to provide thrust, maintain rotor speeds below burst
thresholds, and temperatures below limits have the potential for
detrimental effects to the aircraft. Similar detrimental effects are
possible in the magni250 and magni500 model engines, but the causes are
different. Electric engines with reduced power response time can
experience insufficient thrust to the aircraft, shaft over-torque, and
over-stressed rotating components, propellers, and critical propeller
parts. Therefore, this special condition is necessary.
    Continued Rotation: Proposed special condition no. 16 would require
magniX to design the magni250 and magni500 model 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 magniX magni250 and magni500 model
engines consists of the rotors, shafts, magnets, bearings, and wire
windings that convert electrical energy to shaft torque. 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
result in a forced landing or inability to continue safe flight.
    Safety Analysis: Proposed special condition no. 17 would require
magniX to comply with 14 CFR 33.75(a)(1), (a)(2), and (a)(3), which
require the applicant to conduct a safety analysis of the engine, and
which would otherwise be applicable only to turbine aircraft engines.
Additionally, this proposed special condition would require magniX 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. 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.
    Ingestion: Proposed special condition no. 18 would require magniX
to ensure that these engines will not experience unacceptable power
loss or hazardous engine effects from ingestion. The associated
regulation for turbine engines, 14 CFR 33.76, is based on potential
damage from birds being ingested into the turbine engine that has an
inlet duct, which directs air into the engine for combustion, cooling,
and thrust. In contrast, these electric engines do not use an inlet for
those purposes.
    An ``unacceptable'' power loss, as used in proposed special
condition no. 18(a), is one in which 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 Systems: Proposed special condition no. 19 would require
magniX to ensure that liquid systems used for lubrication or cooling of
engine components are designed and constructed to function properly.
Also, if a liquid system is not self-contained, the interfaces to that
system would be required to be defined in the engine installation
manual. Liquid 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

magniX to ensure (1) the engine 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) the engine design undergoes a vibration survey.
    The vibration demonstration is a survey that characterizes the
vibratory attributes of the engine and verifies 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 magniX
to demonstrate that the engine is capable of continued operation
without the need for maintenance if it experiences a certain amount of
overtorque.
    The electric engine proposed by magniX 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, there is a small delay, or
latency, in the engine response time. While the engine dwells in this
time interval, it can continue to apply torque until the command to
reduce the torque is applied by the engine control. The amount of
overtorque the FAA permits during operation depends on how well the
applicant demonstrates the engine's capability to remain operational
without the need for maintenance action. Therefore, this special
condition is necessary.
    Calibration Assurance: Proposed special condition no. 22 would
require magniX 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 magniX magni250 and magni500 model engines. Proposed special
condition no. 22 would allow magniX to demonstrate the endurance and
durability of the electric engine either together or independently,
whichever is most appropriate for the engine qualities being assessed.
Consequently, the proposed special condition applies the calibration
requirement to both the endurance and durability tests.
    Endurance Demonstration: Proposed special condition no. 23 would
require magniX 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, to
determine if the test is sufficient to show the engine design will not
exhibit unacceptable effects in-service, such as significant
performance deterioration, operability restrictions, engine power loss
or instability, when it is run for sustained periods at extreme
operating conditions.
    Temperature Limit: Proposed special condition no. 24 would require
magniX 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
from failing due to temperature-related causes when operating at the
most extreme 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 magniX magni250 and magni500 model engines can
achieve without incurring hazardous effects. They are requisite
capabilities of the type design that qualify the engine for
installation into aircraft and 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 magniX 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 in a test
cell. These effects could result from elevated inlet cowl temperatures,
extreme aircraft maneuvers, flowstream distortion, and hard landings.
An engine that is run in a test facility could demonstrate more
capability for some operating characteristics than it will when
operating on an aircraft and potentially decrease the engine ratings
and operating limits. Therefore, the installed performance defines the
engine performance capabilities.
    Durability Demonstration: Proposed special condition no. 26 would
require magniX to subject the engine to a durability demonstration. The
durability demonstration must show that each part of the engine is
designed and constructed to minimize the development of any unsafe
condition of the system between overhaul periods or between engine
replacement intervals if overhaul is not defined. Durability is the
ability of an engine, in the fully deteriorated state, to continue
generating rated power or thrust, retain adequate operating margins,
and retain sufficient efficiency that enables the aircraft to reach its
destination. The amount of deterioration an engine can experience is
restricted by operating limitations and managed by the 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, magniX will use
different test effects to establish overhaul periods or engine
replacement intervals if no maintenance is specified.
    System and Component Tests: Proposed special condition no. 27 would
require magniX 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) establishes a need for temperature limits and additional
testing where the endurance test does not fully expose internal
components to thermal conditions that verify the desired operating
limits. A safety concern for electric engines is extreme temperatures.
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. There
are also additional safety considerations that need to be addressed in
the test. Therefore, proposed special condition no. 27 would be a
performance-based requirement that allows magniX to determine how to
challenge the electric engine and to determine the appropriate
limitations that correspond to the technology.
    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 magniX engine installations are
not limited to vehicles 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 magniX to classify the hazard (major/minor)
and assign the appropriate quantitative criteria that meet the safety
objectives required by Sec.  33.75.
    Teardown Inspection: Proposed special condition no. 29 would
require magniX to perform either a teardown evaluation or a non-
teardown evaluation based on the criteria provided in proposed 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
magniX's are constructed in an integrated fashion that precludes the
possibility of tearing them down without destroying them. Sections
33.55 and 33.93 do not contain similar requirements because
reciprocating and turbine engines can be disassembled for inspection.
    Containment: Proposed special condition no. 30 would require the
engine to provide containment features that protect against likely
hazards from rotating components unless magniX can show, by test or
validated analysis, that the margin to rotor burst does not justify the
need for containment features. Rotating components in electric engines
are typically disks, shafts, bearings, seals, orbiting magnetic
components, and the assembled rotor core. However, if the margin to
rotor burst does not unconditionally rule out the possibility of a
rotor burst, then the condition would require magniX to assume a rotor
burst could occur and provide case features that will contain the
failed rotors. In addition, magniX must also determine the effects of
subsequent damage precipitated by the main rotor failure and
characterize any fragments that are released forward or aft of the
containment features. The fragment energy levels, trajectories, and
size must be documented in the installation manual because the aircraft
will need to account for the effects of a rotor failure in the aircraft
design. The intent of this special condition is to prevent hazardous
engine effects from structural failure of rotating components and the
rotating parts that are built into them.
    Operation with a Variable Pitch Propeller or Fan: Proposed special
condition no. 31 would require magniX to conduct functional
demonstrations, including feathering, negative torque, negative thrust,
and reverse thrust operations, as applicable, based on the propeller or
fan's variable pitch functions that are planned for use on these
electric engines, with a representative propeller. The tests prescribed
in Sec.  33.95, for engines operating with variable pitch propellers,
are based on the operating characteristics of turbine engines, which
include thrust response times, engine stall, propeller shaft overload,
loss of thrust control, and hardware fatigue. The electric engines
proposed by magniX have different operating characteristics that
substantially affect their susceptibility to these and other potential
failures. Since magniX's proposed electric engines may be installed
with a variable pitch propeller, the proposed special condition
associated with the operation with a variable pitch propeller or fan is
necessary.
    General Conduct of Tests: Proposed special condition no. 32 would
require magniX to (1) include scheduled maintenance in the engine ICA
before certification; (2) include any maintenance, in addition to the
scheduled maintenance, that was needed during the test to satisfy the
requirement; and (3) conduct any additional tests that the
Administrator finds necessary 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 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 the
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 magniX 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 in order to
comply with the certification requirements.
    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 the magniX magni250 and magni500 model engines. Should
magniX 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 magniX magni250 and magni500 model
engines. It is not a rule of general applicability.

List of Subjects in 14 CFR Part 33

    Aircraft, Aviation safety, Reporting and recordkeeping
requirements.

Authority Citation

    The authority citation for these special conditions is as follows:

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

The Proposed Special Conditions

    Accordingly, the Federal Aviation Administration (FAA) proposes the
following special conditions as part of the type certification basis
for magniX USA, Inc., magni250 and magni500 model engines. The
applicant must also comply with the certification procedures set forth
in 14 CFR part 21.

1. Applicability

    Unless otherwise noted in these special conditions, the design must
comply with the airworthiness standards for aircraft engines set forth
in 14 CFR part 33, except those airworthiness standards specifically
and explicitly applicable only to reciprocating and turbine aircraft
engines.

2. Engine Ratings and Operating Limits

    In addition to Sec.  33.7(a), the design must comply with the
following:
    Ratings and operating limitations must be established and included
in the type certificate data sheet based on:
    (a) Power, torque, speed, and time for:
    (1) Rated maximum continuous power; and
    (2) Rated maximum temporary power and associated time limit.
    (b) The duty cycle and the rating at that duty cycle. The
manufacturer must declare the duty cycle or cycles in the engine
certificate data sheet.

3. Materials

    The engine design must comply with 14 CFR 33.15.

4. Fire Protection

    The engine design must comply with 14 CFR 33.17.
    In addition, high-voltage electrical wiring interconnect systems
must be protected against arc-faults. Any non-protected electrical
wiring interconnects must be analyzed to show that arc-faults do not
cause a hazardous engine effect.

5. Durability

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

6. Engine Cooling

    The engine design and construction must comply with 14 CFR 33.21.
In addition, if cooling is required to satisfy the safety analysis as
described in special condition no. 17, 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 structure must
comply with 14 CFR 33.23.

8. Accessory Attachments

    The engine must comply with 14 CFR 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). Compliance with this paragraph must be shown by
test, validated analysis, or a combination of both. Applicable assumed
speeds must be declared and justified.
    (b) Rotors must possess sufficient strength with a margin to burst
above certified operating conditions and above failure conditions
leading to rotor overspeed. The margin to burst must be shown by tests,
validated analysis, or a combination of both.
    (c) The engine must not exceed the speed operational limitations
that could affect rotor structural integrity.

10. Engine Control Systems

    (a) Applicability.
    The requirements of this paragraph apply to any system or device
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 the engine does not
experience any unacceptable operating characteristics or exceed any of
its operating limitations.
    (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 tests, 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 demonstrations, validated analysis, or a combination thereof,
must be demonstrated by the system and component tests in special
condition no. 27.
    (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 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 result in hazardous engine
effects; and
    (4) Not have any likely failure or malfunction that lead to local
events in the intended aircraft installation.
    (g) 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.
    (h) Protection systems.
    The design and function of the engine control devices and systems,
together with engine instruments, operating instructions and
maintenance instructions, must ensure that engine operating limitations
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), 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.
    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, as defined in
special condition no. 17(d)(2), the unacceptable transmission of
erroneous data, or continued engine operation in the absence of the
control function.

11. Instrument Connection

    The applicant must comply with 14 CFR 33.29(a), (e), (f), and (g).
In addition, as part of the system safety assessment of special
condition no. 10(g), 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.

12. Stress Analysis

    (a) A mechanical, thermal, and electrical stress analysis must show
there is a sufficient design margin to prevent unacceptable operating
characteristics.
    (b) Maximum stresses in the engine must be determined by tests,
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 part means a rotor and major structural static
part whose failure can result in a hazardous engine effect due to a
low-cycle fatigue (LCF) mechanism or any LCF driven mechanism coupled
with creep. 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) The applicant must establish the integrity of each critical
part or life-limited part by providing the following three plans to the
Administrator for approval:
    (1) An engineering plan that establishes and maintains that the
combination of loads, material properties, environmental influences,
and operating conditions, including the effects of engine parts
influencing these parameters, are sufficiently well-known and
predictable by validated analysis, test, or service experience. The
engineering plan must ensure each critical part or life-limited part is
withdrawn from service at an approved life before hazardous engine
effects can occur. The engineering plan must establish activities to be
executed both pre- and post-certification. magniX must perform
appropriate damage tolerance assessments to address the potential for
failure from material, manufacturing, and service-induced anomalies
within the approved life of the part. The approved life must be
published in the mandatory ICA.
    (2) A manufacturing plan that identifies the specific manufacturing
definition (drawings, procedures, specifications, etc.) necessary to
consistently produce critical or life-limited parts with the attributes
required by the engineering plan.
    (3) A service management plan that defines in-service processes for
maintenance and repair of critical or life-limited parts that maintain
attributes consistent with those required by the engineering plan.
These processes must become part of the mandatory ICA.

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 by particle debris.
    (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

    The design and construction of the engine must enable an increase--
    (a) From the minimum power setting to the highest-rated power
without detrimental engine effects; and
    (b) From the minimum obtainable power while in-flight and while on
the ground to the highest-rated power within a time interval for safe
operation of the aircraft.

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
specified in special condition no. 17(d)(2).

17. Safety Analysis

    (a) The applicant must comply with Sec.  33.75(a)(1), (a)(2), and
(a)(3) using the failure definitions in special condition no. 17(d).
    (b) If the failure of such elements is likely to result in
hazardous engine effects, then the applicant may show compliance by
reliance on the prescribed integrity requirements of Sec.  33.15,
special condition no. 9, or special condition no. 13, as determined by
analysis. The failure of such elements and associated prescribed
integrity requirements must be stated in the safety analysis.
    (c) The applicant must comply with 14 CFR 33.75(d) and (e) using
the failure definitions in special condition no. 17(d) of this special
condition.
    (d) Unless otherwise approved by the Administrator, the following
definitions apply to the engine effects when showing compliance with
this condition:
    (1) An engine failure in which the only consequence is the
inability to dispatch the aircraft will be regarded as a minor engine
effect.
    (2) The engine effects in Sec.  33.75(g)(2) are hazardous engine
effects with the addition of:
    Electrocution of crew, passengers, operators, maintainers, or
others.
    (3) Any other engine effect is a major engine effect.

18. Ingestion

    (a) Ingestion from likely sources (foreign objects, birds, ice,
rain, hail) must not result in unacceptable power loss, or in hazardous
engine effects as defined by special condition no. 17(d)(2).
    (b) If the design of the engine relies on features, attachments, or
systems that may be supplied by the installer for the prevention of
unacceptable power loss or hazardous engine effects following potential
ingestion, then the features, attachments, or systems must be
documented in the engine installation manual.

19. Liquid Systems

    (a) Each liquid 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 liquid 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.

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 proposed engine design must undergo a vibration survey to
establish that the vibration characteristics of those components that
may be 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, or electromagnetic. 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 tests, validated analysis, or a
combination thereof, that the engine is capable of continued operation
after operating at the maximum engine overtorque condition without
maintenance action.

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.

23. Endurance Demonstration

    The applicant must subject the engine to an endurance demonstration
acceptable to the Administrator to demonstrate the limit capabilities
of the engine. The endurance demonstration elevates and decreases the
engine's power settings, and dwells at the power settings for 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 at each rated condition
to the Administrator. The demonstration must be repeated for all
declared duty cycles and associated ratings.

25. Operation Demonstration

    The engine design must demonstrate safe operating characteristics,
including but not limited to, power cycling, 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 the development of any unsafe condition of the system between
overhaul periods, or between engine replacement intervals if 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.

27. System and Component Tests

    The applicant must show that systems and components will perform
their intended functions in all declared environmental and operating
conditions.

28. Rotor Locking Demonstration

    If shaft rotation is prevented by a means to lock the rotor(s), the
engine must demonstrate reliable rotor locking performance and that no
hazardous effects will occur.

29. Teardown Inspection

    The applicant must comply with either (a) or (b) as follows:
    (a) Teardown evaluation.
    (1) After the endurance and durability demonstrations have been
completed, the engine must be completely disassembled. Each engine
component must be within service limits and eligible for continued
operation in accordance with the information submitted for showing
compliance with Sec.  33.4, Instructions for Continued Airworthiness.
    (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 limits that were established and
recorded at the beginning of the endurance and durability
demonstrations.
    (b) Non-Teardown evaluation.
    If a teardown is not performed for all engine components, then the
life limits for these components must be established based on the
endurance and durability demonstrations.

30. Containment

    The engine must provide containment features that protect against
likely hazards from rotating components as follows--
    (a) The design of the case surrounding rotating components must
provide for the containment of the rotating components in the event of
failure unless the applicant shows that the rotor has a margin to burst
that would justify no need for containment features.
    (b) If the margin to burst shows the case must have containment
features in the event of failure, the case must provide for the
containment of the failed rotating components. The applicant must
define by test, validated analysis, or combination thereof, and
document in the installation manual the energy level, trajectory, and
size of any fragments released from damage caused by the main rotor
failure that pass forward or aft of the surrounding case.

31. Operation With a Variable Pitch Propeller or Fan

    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 as part of the endurance and durability 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 contained in
the proposed ICA.
    (b) The applicant must subject the engine or its parts to
maintenance and additional tests that the Administrator finds necessary
if--
    (1) The frequency of the service is excessive;
    (2) The number of stops due to engine malfunction is excessive;
    (3) Major repairs are needed; or
    (4) Replacement of a part is found necessary during the tests or as
the result of findings from the teardown inspection.
    (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.

    Issued in Burlington, Massachusetts, on October 19, 2020.
Robert J. Ganley,
Engine and Propeller Standards Branch, Policy and Innovation Division,
Aircraft Certification Service.
[FR Doc. 2020-23434 Filed 11-18-20; 8:45 am]
BILLING CODE 4910-13-P