101854 Federal Register / Vol. 89, No. 242 / Tuesday, December 17, 2024 / Rules and Regulations 

rule that may affect family well-being. 

This final rule would not have any 

impact on the autonomy or integrity of 

the family as an institution. 

Accordingly, DOE has concluded that it 

is not necessary to prepare a Family 

Policymaking Assessment. 

L. Review Under the Treasury and 

General Government Appropriations 

Act, 2001 

Section 515 of the Treasury and 

General Government Appropriations 

Act, 2001 (44 U.S.C. 3516) provides for 

Federal agencies to review most 

disseminations of information to the 

public under guidelines established by 

each agency pursuant to general 

guidelines issued by OMB. OMB’s 

guidelines were published at 67 FR 

8452 (February 22, 2002), and DOE’s 

guidelines were published at 67 FR 

62446 (October 7, 2002). Pursuant to 

OMB Memorandum M–19–15, 

Improving Implementation of the 

Information Quality Act (April 24, 

2019), DOE published updated 

guidelines which are available at: 

www.energy.gov/sites/prod/files/2019/ 

12/f70/DOE%20Final%20

Updated%20IQA%20Guidelines%20

Dec%202019.pdf. 

DOE has reviewed this final rule and 

will ensure that information produced 

under this regulation remains consistent 

with the applicable OMB and DOE 

guidelines. 

M. Congressional Review 

As required by 5 U.S.C. 801, DOE will 

report to Congress on the promulgation 

of this rule prior to its effective date. 

The report will state that the rule does 

not, meet the criteria set forth in 5 

U.S.C. 804(2). 

V. Approval by the Office of the 

Secretary of Energy 

The Secretary of Energy has approved 

publication of this final rule. 

List of Subjects in 10 CFR Part 1008 

Administration practice and 

procedure, Freedom of information, 

Privacy, Reporting and recordkeeping 

requirements. 

Signing Authority 

This document of the Department of 

Energy was signed on [December 11, 

2024, by Ann Dunkin, Senior Agency 

Official for Privacy, pursuant to 

delegated authority from the Secretary 

of Energy. That document with the 

original signature and date is 

maintained by DOE. For administrative 

purposes only, and in compliance with 

requirements of the Office of the Federal 

Register, the undersigned DOE Federal 

Register Liaison Officer has been 

authorized to sign and submit the 

document in electronic format for 

publication, as an official document of 

the Department of Energy. This 

administrative process in no way alters 

the legal effect of this document upon 

publication in the Federal Register. 

Signed in Washington, DC, on December 

12, 2024. 

Treena V. Garrett, 

Federal Register Liaison Officer, U.S. 

Department of Energy. 

For the reasons set forth in the 

preamble, the Department of Energy 

amends part 1008 of chapter X of title 

10 of the Code of Federal Regulations as 

set forth below: 

PART 1008—RECORDS MAINTAINED 

ON INDIVIDUALS (PRIVACY ACT) 

■ 1. The authority citation for part 1008 

continues to read as follows: 

Authority: 42 U.S.C. 7101 et seq.; 50 

U.S.C. 2401 et seq.; 5 U.S.C. 552; 5 U.S.C. 

552a; 42 U.S.C. 7254; and 5 U.S.C. 301. 

Section 1008.22(c) also issued under 42 

U.S.C. 405 note. 

■ 2. Amend § 1008.12 by adding 

paragraph (b)(2)(ii)(Q) to read as 

follows: 

§ 1008.12 Exemptions. 

* * * * * 

(b) * * * 

(2) * * * 

(ii) * * * 

(Q) Nondiscrimination in Federally 

Assisted Program Files (DOE–42) 

* * * * * 

[FR Doc. 2024–29664 Filed 12–16–24; 8:45 am] 

BILLING CODE 6450–01–P 

DEPARTMENT OF TRANSPORTATION 

Federal Aviation Administration 

14 CFR Part 33 

[Docket No. FAA–2022–1641; Special 

Conditions No. 33–028–SC] 

Special Conditions: BETA 

Technologies Inc. Model H500A 

Electric Engines 

AGENCY: Federal Aviation 

Administration (FAA), DOT. 

ACTION: Final special conditions. 

SUMMARY: These special conditions are 

issued for BETA Technologies Inc. 

(BETA) Model H500A electric engines 

that operate using electrical technology 

installed on the aircraft, for use as an 

aircraft engine. These engines will have 

a novel or unusual design feature when 

compared to the state of technology 

envisioned in the airworthiness 

standards applicable to aircraft engines. 

This design feature is the use of an 

electric motor, motor controller, and 

high-voltage systems as the primary 

source of propulsion for an aircraft. The 

applicable airworthiness regulations do 

not contain adequate or appropriate 

safety standards for this design feature. 

These special conditions contain the 

additional safety standards that the 

Administrator considers necessary to 

establish a level of safety equivalent to 

that established by the existing 

airworthiness standards. 

DATES: Effective January 16, 2025. 

FOR FURTHER INFORMATION CONTACT: 

Mark Bouyer, Engine and Propulsion 

Standards Section, AIR–625, Technical 

Policy Branch, Policy and Standards 

Division, Aircraft Certification Service, 

1200 District Avenue, Burlington, 

Massachusetts 01803; telephone (781) 

238–7755; mark.bouyer@faa.gov. 

SUPPLEMENTARY INFORMATION: 

Background 

On January 27, 2022, BETA applied 

for a type certificate for its Model 

H500A electric engines. The BETA 

Model H500A electric engine initially 

will be used as a ‘‘pusher’’ electric 

engine in a single-engine airplane that 

will be certified separately from the 

engine. A typical normal category 

general aviation aircraft locates the 

engine at the front of the fuselage. In 

this configuration, the propeller 

attached to the engine pulls the airplane 

along its flightpath. A pusher engine is 

located at the rear of the fuselage, so the 

propeller attached to the engine pushes 

the aircraft instead of pulling the 

aircraft. 

The BETA Model H500A electric 

engine is comprised of a direct drive, 

radial-flux, permanent-magnet motor, 

divided in two sections, each section 

having a three-phase motor, and one 

electric power inverter controlling each 

three-phase motor. The magnets are 

arranged in a Halbach magnet array, and 

the stator is a concentrated, toothwound configuration. A stator is the 

stationary component in the electric 

engine that surrounds the rotating 

hardware; for example: the BETA 

propeller shaft, which consists of a 

bonded core with coils of insulated 

wire, known as the windings. When 

alternating current is applied to the 

coils of insulated wire in a stator, a 

rotating magnetic field is created, which 

provides the motive force for the 

rotating components. 


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1Sometimes the entire system is referred to as an 

inverter. Throughout this document, it is referred to 

as the controller. 

Type Certification Basis 

Under the provisions of 14 CFR 

21.17(a)(1), generally, BETA must show 

that Model H500A electric engines meet 

the applicable provisions of 14 CFR part 

33 in effect on the date of application 

for a type certificate. 

If the Administrator finds that the 

applicable airworthiness regulations 

(e.g., part 33) do not contain adequate or 

appropriate safety standards for the 

BETA Model H500A electric engines 

because of a novel or unusual design 

feature, special conditions may be 

prescribed under the provisions of 

§ 21.16. 

Special conditions are initially 

applicable to the model for which they 

are issued. Should the type certificate 

for that model be amended later to 

include any other engine model that 

incorporates the same novel or unusual 

design feature, these special conditions 

would also apply to the other engine 

model under § 21.101. 

The FAA issues special conditions, as 

defined in § 11.19, in accordance with 

§ 11.38, and they become part of the 

type certification basis under 

§ 21.17(a)(2). 

Novel or Unusual Design Features 

The BETA Model H500A electric 

engines will incorporate the following 

novel or unusual design features: 

An electric motor, motor controller, 

and high-voltage electrical systems that 

are used as the primary source of 

propulsion for an aircraft. 

Discussion 

Electric propulsion technology is 

substantially different from the 

technology used in previously 

certificated turbine and reciprocating 

engines. Therefore, these engines 

introduce new safety concerns that need 

to be addressed in the certification 

basis. 

BETA’s Electric Engines Are Novel or 

Unusual 

The BETA Model H500A electric 

engines have a novel or unusual design 

feature, which is the use of electrical 

sources of energy instead of fuel to drive 

the mechanical systems that provide 

propulsion for aircraft. Therefore, part 

33 does not contain adequate or 

appropriate safety standards for the 

BETA Model H500A electric engine’s 

novel or unusual design feature. 

BETA’s aircraft engines will operate 

using electrical power instead of air and 

fuel combustion to propel the aircraft. 

These electric engines will be designed, 

manufactured, and controlled 

differently than turbine or reciprocating 

aircraft engines. They will be built with 

an electric motor, motor controller, and 

high-voltage electrical systems that 

draw energy from electrical storage or 

electrical energy generating systems. 

The electric motor is a device that 

converts electrical energy into 

mechanical energy by electric current 

flowing through windings (wire coils) in 

the motor, producing a magnetic field 

that interacts with permanent magnets 

mounted on the engine’s main rotor. 

The controller is a system that consists 

of two main functional elements: the 

motor controller and an electric power 

inverter to drive the motor.1 The highvoltage electrical system is a 

combination of wires and connectors 

that integrate the motor and controller. 

In addition, the technology 

comprising these high-voltage and highcurrent electronic components 

introduces potential hazards that do not 

exist in turbine and reciprocating 

aircraft engines. For example, highvoltage transmission lines, 

electromagnetic shields, magnetic 

materials, and high-speed electrical 

switches are necessary to use the 

physical properties of an electric engine 

for propelling an aircraft. 

BETA’s Electric Engines Require a Mix 

of Part 33 Standards and Special 

Conditions 

The requirements in part 33 ensure 

that the design and construction of 

aircraft engines, including the engine 

control systems, are proper for the type 

of aircraft engines considered for 

certification. However, part 33 does not 

fully address aircraft engines like the 

BETA Model H500A, which operates 

using electrical technology as the 

primary means of propelling the aircraft. 

The requirements in part 33, subpart 

B, are applicable to reciprocating and 

turbine aircraft engines. Subparts C and 

D are applicable to reciprocating aircraft 

engines. Subparts E through G are 

applicable to turbine aircraft engines. As 

such, subparts B through G do not 

adequately address the use of aircraft 

engines that operate using electrical 

technology. Special conditions are 

needed to ensure a level of safety for 

electric engines that is commensurate 

with these subparts, as those regulatory 

requirements do not contain adequate or 

appropriate safety standards for electric 

aircraft engines that are used to propel 

aircraft. 

The FAA proposed special conditions 

and received comments from many 

commenters. Some comments resulted 

in changes to the special conditions. 

These changes are explained in the 

Discussion of Comments. 

FAA Special Conditions for the BETA 

Engine Design 

Applicability: Special condition no. 1 

requires BETA to comply with part 33, 

except for those airworthiness standards 

specifically and explicitly applicable 

only to reciprocating and turbine 

aircraft engines. 

Engine Ratings and Operating 

Limitations: Special condition no. 2, in 

addition to compliance with § 33.7(a), 

requires BETA to establish engine 

operating limits related to the power, 

torque, speed, and duty cycles specific 

to BETA Model H500A electric engines. 

The duty or duty cycle is a statement of 

the load(s) to which the engine is 

subjected, including, if applicable, 

starting, no-load and rest, and deenergized periods, including their 

durations or cycles and sequence in 

time. This special condition also 

requires BETA to declare cooling fluid 

grade or specification, power supply 

requirements, and to establish any 

additional ratings that are necessary to 

define the BETA Model H500A electric 

engine capabilities required for safe 

operation of the engine. 

Materials: Special condition no. 3 

requires BETA to comply with § 33.15, 

which sets requirements for the 

suitability and durability of materials 

used in the engine, and which would 

otherwise be applicable only to 

reciprocating and turbine aircraft 

engines. 

Fire Protection: Special condition no. 

4 requires BETA to comply with § 33.17, 

which sets requirements to protect the 

engine and certain parts and 

components of the airplane against fire, 

and which would otherwise be 

applicable only to reciprocating and 

turbine aircraft engines. Additionally, 

this special condition requires BETA to 

ensure that the high-voltage electrical 

wiring interconnect systems that 

connect the controller to the motor are 

protected against arc faults. An arc fault 

is a high-power discharge of electricity 

between two or more conductors. This 

discharge generates heat, which can 

break down the wire’s insulation and 

trigger an electrical fire. Arc faults can 

range in power from a few amps up to 

thousands of amps and are highly 

variable in strength and duration. 

Durability: Special condition no. 5 

requires the design and construction of 

BETA Model H500A electric engines to 

minimize the development of an unsafe 

condition between maintenance 

intervals, overhaul periods, and 

mandatory actions described in the 


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2https://standards.rtca.org/XanHrK. 

3https://my.rtca.org/NC__

Product?id=a1B36000001IcnSEAS. 

4https://www.faa.gov/documentLibrary/media/ 

Advisory_Circular/AC_33_28-3.pdf. 

Instructions for Continued 

Airworthiness (ICA). 

Engine Cooling: Special condition no. 

6 requires BETA to comply with § 33.21, 

which requires the engine design and 

construction to provide necessary 

cooling, and which would otherwise be 

applicable only to reciprocating and 

turbine aircraft engines. Additionally, 

this special condition requires BETA to 

document the cooling system 

monitoring features and usage in the 

engine installation manual (see § 33.5) if 

cooling is required to satisfy the safety 

analysis described in special condition 

no. 17. Loss of cooling to an aircraft 

engine that operates using electrical 

technology can result in rapid 

overheating and abrupt engine failure, 

with critical consequences to safety. 

Engine Mounting Attachments and 

Structure: Special condition no. 7 

requires BETA and the design to comply 

with § 33.23, which requires the 

applicant to define, and the design to 

withstand, certain load limits for the 

engine mounting attachments and 

related engine structure. These 

requirements would otherwise be 

applicable only to reciprocating and 

turbine aircraft engines. 

Accessory Attachments: Special 

condition no. 8 requires the design to 

comply with § 33.25, which sets certain 

design, operational, and maintenance 

requirements for the engine’s accessory 

drive and mounting attachments, and 

which would otherwise be applicable 

only to reciprocating and turbine 

aircraft engines. 

Rotor Overspeed: Special condition 

no. 9 requires BETA to establish by test, 

validated analysis, or a combination of 

both, that— 

(1) the rotor overspeed must not result 

in a burst, rotor growth, or damage that 

results in a hazardous engine effect; 

(2) rotors must possess sufficient 

strength margin to prevent burst; and 

(3) operating limits must not be 

exceeded in service. 

The special condition associated with 

rotor overspeed is necessary because of 

the differences between turbine engine 

technology and the technology of these 

electric engines. Turbine rotor speed is 

driven by expanding gas and 

aerodynamic loads on rotor blades. 

Therefore, the rotor speed or overspeed 

results from interactions between 

thermodynamic and aerodynamic 

engine properties. The speed of an 

electric engine is directly controlled by 

electric current, and an electromagnetic 

field created by the controller. 

Consequently, electric engine rotor 

response to power demand and 

overspeed-protection systems is quicker 

and more precise. Also, the failure 

modes that can lead to overspeed 

between turbine engines and electric 

engines are vastly different, and 

therefore this special condition is 

necessary. 

Engine Control Systems: Special 

condition no. 10(b) requires BETA to 

ensure that these engines do not 

experience any unacceptable operating 

characteristics, such as unstable speed 

or torque control, or exceed any of their 

operating limitations. 

The FAA originally issued § 33.28 at 

amendment 33–15 to address the 

evolution of the means of controlling 

the fuel supplied to the engine, from 

carburetors and hydro-mechanical 

controls to electronic control systems. 

These electronic control systems grew 

in complexity over the years, and as a 

result, the FAA amended § 33.28 at 

amendment 33–26 to address these 

increasing complexities. The controller 

that forms the controlling system for 

these electric engines is significantly 

simpler than the complex control 

systems used in modern turbine 

engines. The current regulations for 

engine control are inappropriate for 

electric engine control systems; 

therefore, special condition no. 10(b) 

associated with controlling these 

engines is necessary. 

Special condition no. 10(c) requires 

BETA to develop and verify the software 

and complex electronic hardware used 

in programmable logic devices, using 

proven methods that ensure that the 

devices can provide the accuracy, 

precision, functionality, and reliability 

commensurate with the hazard that is 

being mitigated by the logic. RTCA DO– 

254, ‘‘Design Assurance Guidance for 

Airborne Electronic Hardware,’’ dated 

April 19, 2000,2 distinguishes between 

complex and simple electronic 

hardware. 

Special condition no. 10(d) requires 

data from assessments of all functional 

aspects of the control system to prevent 

errors that could exist in software 

programs that are not readily observable 

by inspection of the code. Also, BETA 

must use methods that will result in the 

expected quality that ensures the engine 

control system performs the intended 

functions throughout the declared 

operational envelope. 

The environmental limits referred to 

in special condition no. 10(e) include 

temperature, vibration, high-intensity 

radiated fields (HIRF), and all others 

addressed in RTCA DO–160G, 

‘‘Environmental Conditions and Test 

Procedures for Airborne Electronic/ 

Electrical Equipment and Instruments,’’ 

dated December 8, 2010, which includes 

RTCA DO–160G, Change 1— 

‘‘Environmental Conditions and Test 

Procedures for Airborne Equipment,’’ 

dated December, 16, 2014, and DO–357, 

‘‘User Guide: Supplement to DO–160G,’’ 

dated December 16, 2014.3 Special 

condition 10(e) requires BETA to 

demonstrate by system or component 

tests in special condition no. 27 any 

environmental limits that cannot be 

adequately substantiated by the 

endurance demonstration, validated 

analysis, or a combination thereof. 

Special condition no. 10(f) requires 

BETA to evaluate various control system 

failures to ensure that such failures will 

not lead to unsafe engine conditions. 

The FAA issued Advisory Circular (AC) 

33.28–3, ‘‘Guidance Material for 14 CFR 

33.28, Engine Control Systems,’’ on May 

23, 2014 (AC 33.28–3), for reciprocating 

and turbine engines.4 This AC provides 

guidance for defining an engine control 

system failure when showing 

compliance with the requirements of 

§ 33.28. AC 33.28–3 also includes 

objectives for control system integrity 

requirements, criteria for a loss of thrust 

control (LOTC) and loss of power 

control (LOPC) event, and an acceptable 

LOTC/LOPC rate. The electrical and 

electronic failures and failure rates did 

not account for electric engines when 

the FAA issued this AC, and therefore 

performance-based special conditions 

are established to allow fault 

accommodation criteria to be developed 

for electric engines. 

The phrase ‘‘in the full-up 

configuration’’ used in special condition 

no. 10(f)(2) refers to a system without 

any fault conditions present. The 

electronic control system must, when in 

the full-up configuration, be single-fault 

tolerant, as determined by the 

Administrator, for electrical, electrically 

detectable, and electronic failures 

involving LOPC events. 

The term ‘‘local’’ in the context of 

‘‘local events’’ used in special condition 

no. 10(f)(4) means failures or 

malfunctions leading to events in the 

intended aircraft installation such as 

fire, overheat, or failures leading to 

damage to engine control system 

components. These ‘‘local events’’ must 

not result in a hazardous engine effect 

due to engine control system failures or 

malfunctions. 

Special condition no. 10(g) requires 

BETA to conduct a safety assessment of 

the control system to support the safety 

analysis in special condition no. 17. 

This control system safety assessment 


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provides engine response to failures, 

and rates of these failures that can be 

used at the aircraft-level safety 

assessment. 

Special condition no. 10(h) requires 

BETA to provide appropriate protection 

devices or systems to ensure that engine 

operating limits will not be exceeded in 

service. 

Special condition no. 10(i) is 

necessary to ensure that the controllers 

are self-sufficient and isolated from 

other aircraft systems. The aircraftsupplied data supports the analysis at 

the aircraft level to protect the aircraft 

from common mode failures that could 

lead to major propulsion power loss. 

The exception ‘‘other than power 

command signals from the aircraft,’’ 

noted in special condition no. 10(i), is 

based on the FAA’s determination that 

the engine controller has no reasonable 

means to determine the validity of any 

in-range signals from the electrical 

power system. In many cases, the engine 

control system can detect a faulty signal 

from the aircraft, but the engine control 

system typically accepts the power 

command signal as a valid value. 

The term ‘‘independent’’ in the 

context of ‘‘fully independent engine 

systems’’ referenced in special 

condition no. 10(i) means the 

controllers should be self-sufficient and 

isolated from other aircraft systems or 

provide redundancy that enables the 

engine control system to accommodate 

aircraft data system failures. In the case 

of loss, interruption, or corruption of 

aircraft-supplied data, the engine must 

continue to function in a safe and 

acceptable manner without hazardous 

engine effects. 

The term ‘‘accommodated,’’ in the 

context of ‘‘detected and 

accommodated,’’ referenced in special 

condition 10(i)(2) is to assure that, upon 

detecting a fault, the system continues 

to function safely. 

Special condition no. 10(j) requires 

BETA to show that the loss of electric 

power from the aircraft will not cause 

the electric engine to malfunction in a 

manner hazardous to the aircraft. The 

total loss of electric power to the electric 

engine may result in an engine 

shutdown. 

Instrument Connection: Special 

condition no. 11 requires BETA to 

comply with § 33.29(a), (e), and (g), 

which set certain requirements for the 

connection and installation of 

instruments to monitor engine 

performance. The remaining 

requirements in § 33.29 apply only to 

technologies used in reciprocating and 

turbine aircraft engines. 

Instrument connections (wires, wire 

insulation, potting, grounding, 

connector designs, etc.) must not 

introduce unsafe features or 

characteristics to the aircraft. Special 

condition no. 11 requires the safety 

analysis to include potential hazardous 

effects from failures of instrument 

connections to function properly. The 

outcome of this analysis might identify 

the need for design enhancements or 

additional ICA to ensure safety. 

Stress Analysis: Section 33.62 

requires applicants to perform a stress 

analysis on each turbine engine. This 

regulation is explicitly applicable only 

to turbine engines and turbine engine 

components, and it is not appropriate 

for the BETA Model H500A electric 

engines. However, a stress analysis 

particular to these electric engines is 

necessary to account for stresses 

resulting from electric technology used 

in the engine. 

Special condition no. 12 requires a 

mechanical, thermal, and electrical 

stress analysis to show that the engine 

has a sufficient design margin to prevent 

unacceptable operating characteristics. 

Also, the applicant must determine the 

maximum stresses in the engine by 

tests, validated analysis, or a 

combination thereof, and show that they 

do not exceed minimum material 

properties. 

Critical and Life-Limited Parts: 

Special condition no. 13 requires BETA 

to show whether rotating or moving 

components, bearings, shafts, static 

parts, and non-redundant mount 

components should be classified, 

designed, manufactured, and managed 

throughout their service life as critical 

or life-limited parts. 

The term ‘‘low-cycle fatigue,’’ 

referenced in special condition no. 

13(a)(2), is a decline in material strength 

from exposure to cyclic stress at levels 

beyond the stress threshold the material 

can sustain indefinitely. This threshold 

is known as the ‘‘material endurance 

limit.’’ Low-cycle fatigue typically 

causes a part to sustain plastic or 

permanent deformation during the 

cyclic loading and can lead to cracks, 

crack growth, and fracture. Engine parts 

that operate at high temperatures and 

high mechanical stresses 

simultaneously can experience lowcycle fatigue coupled with creep. Creep 

is the tendency of a metallic material to 

permanently move or deform when it is 

exposed to the extreme thermal 

conditions created by hot combustion 

gasses, and substantial physical loads 

such as high rotational speeds and 

maximum thrust. Conversely, high-cycle 

fatigue is caused by elastic deformation, 

small strains caused by alternating 

stress, and a much higher number of 

load cycles compared to the number of 

cycles that cause low-cycle fatigue. 

The engineering plan referenced in 

special condition no. 13(b)(1) informs 

the manufacturing and service 

management processes of essential 

information that ensures the life limit of 

a part is valid. The engineering plan 

provides methods for verifying the 

characteristics and qualities assumed in 

the design data using methods that are 

suitable for the part criticality. The 

engineering plan informs the 

manufacturing process of the attributes 

that affect the life of the part. The 

engineering plan, manufacturing plan, 

and service management plan are 

related in that assumptions made in the 

engineering plan are linked to how a 

part is manufactured and how that part 

is maintained in service. For example, 

environmental effects on life limited 

electric engine parts, such as humidity, 

might not be consistent with the 

assumptions used to design the part. 

BETA must ensure that the engineering 

plan is complete, available, and 

acceptable to the Administrator. 

The term ‘‘manufacturing plan,’’ 

referenced in special condition no. 

13(b)(2), is the collection of data 

required to translate documented 

engineering design criteria into physical 

parts, and to verify that the parts 

comply with the properties established 

by the design data. Because engines are 

not intentionally tested to failure during 

a certification program, documents and 

processes used to execute production 

and quality systems required by 

§ 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 manufacturinginduced anomalies, which can degrade 

the part’s structural integrity. Examples 

of manufacturing-induced anomalies are 

material contamination, unacceptable 

grain growth, heat-affected areas, and 

residual stresses. 

The service-management plan ensures 

the method and assumptions used in the 

engineering plan to determine the part’s 

life remain valid by enabling corrections 

identified from in-service experience, 

such as service-induced anomalies and 


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unforeseen environmental effects, to be 

incorporated into the design process. 

The service-management plan also 

becomes the ICA for maintenance, 

overhaul, and repairs of the part. 

Lubrication System: Special condition 

no. 14 requires BETA to ensure that the 

lubrication system is designed to 

function properly between scheduled 

maintenance intervals and to prevent 

contamination of the engine bearings. 

This special condition also requires 

BETA to demonstrate the unique 

lubrication attributes and functional 

capability of the BETA Model H500A 

electric engine design. 

The corresponding part 33 regulations 

include provisions for lubrication 

systems used in reciprocating and 

turbine engines. The part 33 

requirements account for safety issues 

associated with specific reciprocating 

and turbine engine system 

configurations. These regulations are 

not appropriate for the BETA Model 

H500A electric engines. For example, 

electric engines do not have a crankcase 

or lubrication oil sump. Electric engine 

bearings are sealed, so they do not 

require an oil circulation system. The 

lubrication system in these engines is 

also independent of the propeller pitch 

control system. Therefore, special 

condition no. 14 incorporates only 

certain requirements from the part 33 

regulations. 

Power Response: Special condition 

no. 15 requires the design and 

construction of the BETA Model H500A 

electric engines to enable an increase 

from the minimum— 

(1) power setting to the highest rated 

power without detrimental engine 

effects, and 

(2) within a time interval appropriate 

for the intended aircraft application. 

The engine control system governs the 

increase or decrease in power in 

combustion engines to prevent too 

much (or too little) fuel from being 

mixed with air before combustion. Due 

to the lag in rotor response time, 

improper fuel/air mixtures can result in 

engine surges, stalls, and exceedances 

above rated limits and durations. 

Failure of the combustion engine to 

provide thrust, maintain rotor speeds 

below rotor burst thresholds, and keep 

temperatures below limits can have 

engine effects detrimental to the aircraft. 

Similar detrimental effects are possible 

in the BETA Model H500A electric 

engines, but the causes are different. 

Electric engines with reduced power 

response time can experience 

insufficient thrust to the aircraft, shaft 

over-torque, and over-stressed rotating 

components, propellers, and critical 

propeller parts. Therefore, this special 

condition is necessary. 

Continued Rotation: Special condition 

no. 16 requires BETA to design the 

Model H500A electric engines such that, 

if the main rotating systems continue to 

rotate after the engine is shut down 

while in-flight, this continued rotation 

will not result in any hazardous engine 

effects. 

The main rotating system of the BETA 

Model H500A electric engines consists 

of the rotors, shafts, magnets, bearings, 

and wire windings that convert 

electrical energy to shaft torque. For the 

initial aircraft application, this rotating 

system must continue to rotate after the 

power source to the engine is shut 

down. The safety concerns associated 

with this special condition are 

substantial asymmetric aerodynamic 

drag that can cause aircraft instability, 

loss of control, and reduced efficiency; 

and may result in a forced landing or 

inability to continue safe flight. 

Safety Analysis: Special condition no. 

17 requires BETA to comply with 

§ 33.75(a)(1) and (a)(2), which require 

the applicant to conduct a safety 

analysis of the engine, and which would 

otherwise be applicable only to turbine 

aircraft engines. Additionally, this 

special condition requires BETA to 

assess its engine design to determine the 

likely consequences of failures that can 

reasonably be expected to occur. The 

failure of such elements, and associated 

prescribed integrity requirements, must 

be stated in the safety analysis. 

A primary failure mode is the manner 

in which a part is most likely going to 

fail. Engine parts that have a primary 

failure mode, a predictable life to the 

failure, and a failure consequence that 

results in a hazardous effect, are lifelimited or critical parts. Some lifelimited or critical engine parts can fail 

suddenly in their primary failure mode, 

from prolonged exposure to normal 

engine environments such as 

temperature, vibration, and stress, if 

those engine parts are not removed from 

service before the damage mechanisms 

progress to a failure. Due to the 

consequence of failure, these parts are 

not allowed to be managed by oncondition or probabilistic means 

because the probability of failure cannot 

be sensibly estimated in numerical 

terms. Therefore, the parts are managed 

by compliance with integrity 

requirements, such as mandatory 

maintenance (life limits, inspections, 

inspection techniques), to ensure the 

qualities, features, and other attributes 

that prevent the part from failing in its 

primary failure mode are preserved 

throughout its service life. For example, 

if the number of engine cycles to failure 

are predictable and can be associated 

with specific design characteristics, 

such as material properties, then the 

applicant can manage the engine part 

with life limits. 

Complete or total power loss is not 

assumed to be a minor engine event, as 

it is in the turbine engine regulation 

§ 33.75, to account for experience data 

showing a potential for higher hazard 

levels from power loss events in singleengine general aviation aircraft. The 

criteria in these special conditions 

apply to an engine that continues to 

operate at partial power after a single 

electrical or electronic fault or failure. 

Total loss of power is classified at the 

aircraft level using special condition 

nos. 10(g) and 33(h). 

Ingestion: Special condition no. 18 

requires BETA to ensure that these 

engines will not experience 

unacceptable power loss or hazardous 

engine effects from ingestion. The 

associated regulations for turbine 

engines, §§ 33.76, 33.77, and 33.78, are 

based on potential performance impacts 

and damage from birds, ice, rain, and 

hail being ingested into a turbine engine 

that has an inlet duct, which directs air 

into the engine for combustion, cooling, 

and thrust. By contrast, the BETA 

electric engines are not configured with 

inlet ducts. 

An ‘‘unacceptable’’ power loss, as 

used in special condition no. 18(b), is 

such that the power or thrust required 

for safe flight of the aircraft becomes 

unavailable to the pilot. The specific 

amount of power loss that is required 

for safe flight depends on the aircraft 

configuration, speed, altitude, attitude, 

atmospheric conditions, phase of flight, 

and other circumstances where the 

demand for thrust is critical to safe 

operation of the aircraft. 

Liquid and Gas Systems: Special 

condition no. 19 requires BETA to 

ensure that systems used for lubrication 

or cooling of engine components are 

designed and constructed to function 

properly. Also, if a system is not selfcontained, 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. 


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Vibration Demonstration: Special 

condition no. 20 requires BETA to 

ensure the engine— 

(1) is designed and constructed to 

function throughout its normal 

operating range of rotor speeds and 

engine output power without inducing 

excessive stress caused by engine 

vibration, and 

(2) design undergoes a vibration 

survey. 

The vibration demonstration is a 

survey that characterizes the vibratory 

attributes of the engine. It verifies that 

the stresses from vibration do not 

impose excessive force or result in 

natural frequency responses on the 

aircraft structure. The vibration 

demonstration also ensures internal 

vibrations will not cause engine 

components to fail. Excessive vibration 

force occurs at magnitudes and forcing 

functions or frequencies, which may 

result in damage to the aircraft. Stress 

margins to failure add conservatism to 

the highest values predicted by analysis 

for additional protection from failure 

caused by influences beyond those 

quantified in the analysis. The result of 

the additional design margin is 

improved engine reliability that meets 

prescribed thresholds based on the 

failure classification. The amount of 

margin needed to achieve the prescribed 

reliability rates depends on an 

applicant’s experience with a product. 

The FAA considers the reliability rates 

when deciding how much vibration is 

‘‘excessive.’’ 

Overtorque: Special condition no. 21 

requires BETA to demonstrate that the 

engine is capable of continued operation 

without the need for maintenance if it 

experiences a certain amount of 

overtorque. 

BETA’s electric engine converts 

electrical energy to shaft torque, which 

is used for propulsion. The electric 

motor, controller, and high-voltage 

systems control the engine torque. 

When the pilot commands power or 

thrust, the engine responds to the 

command and adjusts the shaft torque to 

meet the demand. During the transition 

from one power or thrust setting to 

another, a small delay, or latency, 

occurs in the engine response time. 

While the engine dwells in this time 

interval, it can continue to apply torque 

until the command to change the torque 

is applied by the engine control. The 

allowable amount of overtorque during 

operation depends on the engine’s 

response to changes in the torque 

command throughout its operating 

range. 

Calibration Assurance: Special 

condition no. 22 requires BETA to 

subject the engine to calibration tests to 

establish its power characteristics and 

the conditions both before and after the 

endurance and durability 

demonstrations specified in special 

condition nos. 23 and 26. The 

calibration test requirements specified 

in § 33.85 only apply to the endurance 

test specified in § 33.87, which is 

applicable only to turbine engines. The 

FAA determined that the methods used 

for accomplishing those tests for turbine 

engines are not appropriate for electric 

engines. The calibration tests in § 33.85 

have provisions applicable to ratings 

that are not relevant to the BETA Model 

H500A electric engines. Special 

condition no. 22 allows BETA to 

demonstrate the endurance and 

durability of the electric engine either 

together or independently, whichever is 

most appropriate for the engine qualities 

being assessed. Consequently, the 

special condition applies the calibration 

requirement to both the endurance and 

durability tests. 

Endurance Demonstration: Special 

condition no. 23 requires BETA to 

perform an endurance demonstration 

test that is acceptable to the 

Administrator. The Administrator will 

evaluate the extent to which the test 

exposes the engine to failures that could 

occur when the engine is operated at up 

to its rated values, and determine if the 

test is sufficient to show that the engine 

design will not exhibit unacceptable 

effects in service, such as significant 

performance deterioration, operability 

restrictions, and engine power loss or 

instability, when it is run repetitively at 

rated limits and durations in conditions 

that represent extreme operating 

environments. 

Temperature Limit: Special condition 

no. 24 requires BETA to ensure the 

engine can endure operation at its 

temperature limits plus an acceptable 

margin. An ‘‘acceptable margin,’’ as 

used in the special condition, is the 

amount of temperature above that 

required to prevent the least capable 

engine allowed by the type design, as 

determined by § 33.8, from failing due 

to temperature-related causes when 

operating at the most extreme engine 

and environmental thermal conditions. 

Operation Demonstration: Special 

condition no. 25 requires the engine to 

demonstrate safe operating 

characteristics throughout its declared 

flight envelope and operating range. 

Engine operating characteristics define 

the range of functional and performance 

values the BETA Model H500A electric 

engines can achieve without incurring 

hazardous effects. The characteristics 

are requisite capabilities of the type 

design that qualify the engine for 

installation into aircraft and that 

determine aircraft installation 

requirements. The primary engine 

operating characteristics are assessed by 

the tests and demonstrations that would 

be required by these special conditions. 

Some of these characteristics are shaft 

output torque, rotor speed, power 

consumption, and engine thrust 

response. The engine performance data 

BETA will use to certify the engine must 

account for installation loads and 

effects. These are aircraft-level effects 

that could affect the engine 

characteristics that are measured when 

the engine is tested on a stand or in a 

test cell. These effects could result from 

elevated inlet cowl temperatures, 

aircraft maneuvers, flowstream 

distortion, and hard landings. For 

example, an engine that is run in a sealevel, static test facility could 

demonstrate more capability for some 

operating characteristics than it will 

have when operating on an aircraft in 

certain flight conditions. Discoveries 

like this during certification could affect 

engine ratings and operating limits. 

Therefore, the installed performance 

defines the engine performance 

capabilities. 

Durability Demonstration: Special 

condition no. 26 requires BETA to 

subject the engine to a durability 

demonstration. The durability 

demonstration must show that the 

engine is designed and constructed to 

minimize the development of any 

unsafe condition between maintenance 

intervals or between engine replacement 

intervals if maintenance or overhaul is 

not defined. The durability 

demonstration also verifies that the ICA 

is adequate to ensure the engine, in its 

fully deteriorated state, continues to 

generate rated power or thrust, while 

retaining operating margins and 

sufficient efficiency, to support the 

aircraft safety objectives. The amount of 

deterioration an engine can experience 

is restricted by operating limitations and 

managed by the engine ICA. Section 

33.90 specifies how maintenance 

intervals are established; it does not 

include provisions for an engine 

replacement. Electric engines and 

turbine engines deteriorate differently; 

therefore, BETA will use different test 

effects to develop maintenance, 

overhaul, or engine replacement 

information for their electric engine. 

System and Component Tests: Special 

condition no. 27 requires BETA to show 

that the systems and components of the 

engine perform their intended functions 

in all declared engine environments and 

operating conditions. 

Sections 33.87 and 33.91, which are 

specifically applicable to turbine 

engines, have conditional criteria to 


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decide if additional tests will be 

required after the engine tests. The 

criteria are not suitable for electric 

engines. Part 33 associates the need for 

additional testing with the outcome of 

the § 33.87 endurance test because it is 

designed to address safety concerns in 

combustion engines. For example, 

§ 33.91(b) requires the establishment of 

temperature limits for components that 

require temperature-controlling 

provisions, and § 33.91(a) requires 

additional testing of engine systems and 

components where the endurance test 

does not fully expose internal systems 

and components to thermal conditions 

that verify the desired operating limits. 

Exceeding temperature limits is a safety 

concern for electric engines. The FAA 

determined that the § 33.87 endurance 

test is not appropriate for testing the 

electronic components of electric 

engines because mechanical energy is 

generated differently by electronic 

systems than it is by the thermal 

conditions in turbine engines. 

Additional safety considerations also 

need to be addressed in the test. 

Therefore, special condition no. 27 is a 

performance-based requirement that 

allows BETA to determine when engine 

systems and component tests are 

necessary and to determine the 

appropriate limitations of those systems 

and components used in the BETA 

Model H500A electric engine. 

Rotor Locking Demonstration: Special 

condition no. 28 requires the engine to 

demonstrate reliable rotor locking 

performance and that no hazardous 

effects will occur if the engine uses a 

rotor locking device to prevent shaft 

rotation. 

Some engine designs enable the pilot 

to prevent a propeller shaft or main 

rotor shaft from turning while the 

engine is running, or the aircraft is inflight. This capability is needed for 

some installations that require the pilot 

to confirm the functionality of certain 

flight systems before takeoff. The BETA 

engine installations are not limited to 

aircraft that will not require rotor 

locking. Section 33.92 prescribes a test 

that may not include the appropriate 

criteria to demonstrate sufficient rotor 

locking capability for these engines. 

Therefore, this special condition is 

necessary. 

The special condition does not define 

‘‘reliable’’ rotor locking but allows 

BETA to classify the hazard as major or 

minor and assign the appropriate 

quantitative criteria that meet the safety 

objectives required by special condition 

no. 17 and the applicable portions of 

§ 33.75. 

Teardown Inspection: Special 

condition no. 29 requires BETA to 

perform a teardown or non-teardown 

evaluation after the endurance, 

durability, and overtorque 

demonstrations, based on the criteria in 

special condition no. 29(a) or (b). 

Special condition no. 29(b) includes 

restrictive criteria for ‘‘non-teardown 

evaluations’’ to account for electric 

engines, sub-assemblies, and 

components that cannot be 

disassembled without destroying them. 

Some electrical and electronic 

components like BETA’s are constructed 

in an integrated fashion that precludes 

the possibility of tearing them down 

without destroying them. The special 

condition indicates that, if a teardown 

cannot be performed in a nondestructive manner, then the inspection 

or replacement intervals must be 

established based on the endurance and 

durability demonstrations. The 

procedure for establishing maintenance 

should be agreed upon between the 

applicant and the FAA prior to running 

the relevant tests. Data from the 

endurance and durability tests may 

provide information that can be used to 

determine maintenance intervals and 

life limits for parts. However, if life 

limits are required, the lifing procedure 

is established by special condition no. 

13, Critical and Life-Limited Parts, 

which corresponds to § 33.70. 

Therefore, the procedure used to 

determine which parts are life-limited, 

and how the life limits are established, 

requires FAA approval, as it does for 

§ 33.70. Sections 33.55 and 33.93 do not 

contain similar requirements because 

reciprocating and turbine engines can be 

completely disassembled for inspection. 

Containment: Special condition no. 

30 requires the engine to have 

containment features that protect 

against likely hazards from rotating 

components unless BETA can show the 

margin to rotor burst does not justify the 

need for containment features. Rotating 

components in electric engines are 

typically disks, shafts, bearings, seals, 

orbiting magnetic components, and the 

assembled rotor core. However, if the 

margin to rotor burst does not 

unconditionally rule out the possibility 

of a rotor burst, then the special 

condition requires BETA to assume a 

rotor burst could occur and design the 

stator case to contain the failed rotors, 

and any components attached to the 

rotor that are released during the failure. 

In addition, BETA must also determine 

the effects of subsequent damage 

precipitated by a main rotor failure and 

characterize any fragments that are 

released forward or aft of the 

containment features. Further, decisions 

about whether the BETA engine requires 

containment features, and the effects of 

any subsequent damage following a 

rotor burst, should be based on test or 

validated analysis. The fragment energy 

levels, trajectories, and size are typically 

documented in the installation manual 

because the aircraft will need to account 

for the effects of a rotor failure in the 

aircraft design. The intent of this special 

condition is to prevent hazardous 

engine effects from structural failure of 

rotating components and parts that are 

built into the rotor assembly. 

General Conduct of Tests: Special 

condition no. 32 requires BETA to— 

(1) Include any scheduled 

maintenance. 

(2) Include any maintenance, in 

addition to the scheduled maintenance, 

which was needed during the test to 

satisfy the applicable test requirements; 

and 

(3) Conduct any additional tests that 

the Administrator finds necessary, as 

warranted by the test results. 

For example, certification endurance 

test shortfalls might be caused by 

omitting some prescribed engine test 

conditions, or from accelerated 

deterioration of individual parts arising 

from the need to force the engine to 

operating conditions that drive the 

engine above the engine cycle values of 

the type design. If an engine part fails 

during a certification test, the entire 

engine might be subjected to penalty 

runs, with a replacement or newer part 

design installed on the engine, to meet 

the test requirements. Also, the 

maintenance performed to replace the 

part, so that the engine could complete 

the test, would be included in the 

engine ICA. In another example, if the 

applicant replaces a part before 

completing an engine certification test 

because of a test facility failure and can 

substantiate the part to the 

Administrator through bench testing, 

they might not need to substantiate the 

part design using penalty runs with the 

entire engine. 

The term ‘‘excessive’’ is used to 

describe the frequency of unplanned 

engine maintenance, and the frequency 

of unplanned test stoppages, to address 

engine issues that prevent the engine 

from completing the tests in special 

condition nos. 32(b)(1) and (2), 

respectively. Excessive frequency is an 

objective assessment from the FAA’s 

analysis of the amount of unplanned 

maintenance needed for an engine to 

complete a certification test. The FAA’s 

assessment may include the reasons for 

the unplanned maintenance, such as the 

effects test facility equipment may have 

on the engine, the inability to simulate 

a realistic engine operating 

environment, and the extent to which 

an engine requires modifications to 


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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 § 33.28 was 

codified, the only electrical systems 

used on turbine engines were lowvoltage, electronic engine control 

systems (EEC) and high-energy sparkignition systems. Electric aircraft 

engines use high-voltage, high-current 

electrical systems and components that 

are physically located in the motor and 

motor controller. Therefore, the existing 

part 33 control system requirements do 

not adequately address all the electrical 

systems used in electric aircraft engines. 

Special condition no. 33 is established 

using the existing engine control 

systems requirement as a basis. It 

applies applicable airworthiness criteria 

from § 33.28 and incorporates 

airworthiness criteria that recognize and 

focus on the electrical power system 

used in the engine. 

Special condition no. 33(b) ensures 

that all aspects of an electrical system, 

including generation, distribution, and 

usage, do not experience any 

unacceptable operating characteristics. 

Special condition no. 33(c) requires 

the electrical power distribution aspects 

of the electrical system to provide the 

safe transfer of electrical energy 

throughout the electric engine. 

The term ‘‘abnormal conditions’’ used 

in special condition no. 33(c)(2) is based 

on the term ‘‘abnormal operation’’ used 

in MIL–STD–704F ‘‘Aircraft Electric 

Power Characteristics’’ which defines 

normal operation and abnormal 

operation. MIL–STD–704F is a standard 

that ensures compatibility between 

power sources that provide power to the 

aircraft’s electrical systems and airborne 

equipment that receive power from the 

power source. This standard also 

establishes technical criteria for aircraft 

electric power. The term ‘‘abnormal 

conditions’’ refers to various engine 

operating conditions such as: 

• System or component 

characteristics outside of normal 

statistical variation from circumstances 

such as systems degradation, 

installation error, and engine response 

to fault conditions; 

• Unusual environmental conditions 

from extreme temperature, humidity, 

vibration, lightning, high-intensity 

radiated field (HIRF), atmospheric 

neutron radiation; and 

• Unusual and infrequent events such 

as landing on icy runways, rejected 

take-offs or go-arounds, extended 

ground idling or taxiing in a hot 

environment, and abrupt load changes 

from foreign object damage or engine 

contamination. 

The phrase ‘‘safe transmission of 

electric energy’’ used in special 

condition no. 33(c)(3) refers to the 

transmission of electrical energy in a 

manner that supports the operation of 

the electric engine(s) and the aircraft 

safety objectives without detrimental 

effects such as uncontrolled fire or 

structural failure due to severe 

overheating. 

Special condition no. 33(d) requires 

the engine electrical system to be 

designed such that the loss, 

malfunction, or interruption of the 

electrical power source, or power 

conditions that exceed design limits, 

will not result in a hazardous engine 

effect. 

Special condition no. 33(e) requires 

BETA to identify and declare, in the 

engine installation manual, the 

characteristics of any electrical power 

supplied from the aircraft to the engine, 

or electrical power supplied from the 

engine to the aircraft via energy 

regeneration, and any other 

characteristics necessary for safe 

operation of the engine. 

Special condition no. 33(f) requires 

BETA to demonstrate that systems and 

components will operate properly up to 

environmental limits, using special 

conditions, when such limits cannot be 

adequately substantiated by the 

endurance demonstration, validated 

analysis, or a combination thereof. The 

environmental limits referred to in this 

special condition include temperature, 

vibration, HIRF, and all others 

addressed in RTCA DO–160G, 

‘‘Environmental Conditions and Test 

Procedures for Airborne Electronic/ 

Electrical Equipment and Instruments.’’ 

Special condition 33(g) requires BETA 

to evaluate various electric engine 

system failures to ensure that these 

failures will not lead to unsafe engine 

conditions. The evaluation includes 

single-fault tolerance, ensures no single 

electrical or electronic fault or failure 

would result in hazardous engine 

effects, and ensures that any failure or 

malfunction leading to local events in 

the intended aircraft application does 

not result in certain hazardous engine 

effects. The special condition also 

implements integrity requirements, 

criteria for LOTC/LOPC events, and an 

acceptable LOTC/LOPC rate. 

Special condition 33(h) requires 

BETA to conduct a safety assessment of 

the engine electrical system to support 

the safety analysis in special condition 

no. 17. This safety assessment provides 

engine response to failures, and rates of 

these failures, which can be used at the 

aircraft safety assessment level. 

Discussion of Comments 

The FAA issued a notice of proposed 

special conditions (NPSC) Docket No. 

FAA–2022–1641 for the BETA Model 

H500A electric engines, which was 

published in the Federal Register on 

March 7, 2024 (89 FR 16474). 

The FAA Received Comments From 

Eight Commenters 

The FAA received comments from 

Transport Canada (TC), Transport 

Canada Civil Aviation (TCCA), United 

Parcel Service Flight Forward (UPSFF), 

Association for Uncrewed Vehicle 

Systems International (AUVSI), magniX 

USA, Inc. (magniX), General Aviation 

Manufacturers Association (GAMA), an 

individual, and an anonymous 

commenter. 

The FAA received comments from 

TCCA. 

TCCA indicated the discussion of 

proposed special condition no.10(e), 

Environmental limits of engine cooling 

systems, in the preamble states that the 

environmental limits referred to in this 

special condition are addressed in 

RTCA DO–160G. However, TCCA 

explained that some of the existing 

RTCA DO–160G test specifications, 

methods, and categories may not be 

adequate for high voltage systems, such 

as the high voltage components of this 

engine. Accordingly, TCCA 

recommended adding the language ‘‘or 

other appropriate industry standards’’ at 

the end of the discussion of special 

condition no. 10(e) in the preamble. 

The FAA does not agree with the 

recommended change. Although RTCA 

DO–160G is not sufficient for the high 

voltage systems used in the BETA 

Model H500A electric engine motor and 

inverter/controller, tests that are 

appropriate for the BETA engine will be 

developed in accordance with special 

condition nos. 1(b) and 1(c) using the 

testing techniques in RTCA DO–160G 

and other aerospace environmental 


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documents. Independent tests are done 

for radiated and conducted 

susceptibility and compared to the 

RTCA DO–160G HIRF spectrum for 

susceptibility to ensure all electric 

engine radio frequency energy 

emissions inherent to the engine design 

are addressed. If the equipment under 

test passes the emission test in RTCA– 

DO–160 the susceptibility spectrum is 

covered by RTCA DO–160G. The 

applicant can use the RTCA DO–160G 

test. If not, the spectrum from the 

emission test would be analyzed and 

could be adjusted for the applicant’s 

design and applied during the 

susceptibility test with FAA 

concurrence. No changes were made to 

these special conditions as a result of 

this comment. 

TCCA also indicated special 

condition no. 2, Engine ratings and 

operating limits, should require that 

component life be considered when 

establishing the engine operating limits. 

They explained, the engine system or 

the electrical motor design may have 

components or parts that require a life 

limit. For example, the insulation on the 

high voltage system wiring may degrade 

with time and operating conditions. 

TCCA requested the FAA add ‘‘(f) 

Component life’’ to special condition 

no. 2, Engine Ratings and Operating 

Limits, and explained that component 

life should be considered when 

establishing the engine operating limits, 

similar to § 33.07(b)(7). 

The FAA does not concur with 

TCCA’s request. Component life is an 

expected outcome of special conditions 

nos. 13 (Critical and life-limited parts) 

and 17 (Safety analysis). Special 

condition no. 17 determines whether 

special condition no. 13 applies to the 

engine part. Special condition no 13 

determines the mandatory replacement 

times (component life) and implements 

a maintenance program to manage these 

parts composed of an engineering plan, 

manufacturing plan, and service 

management plan. No changes were 

made to these special conditions as a 

result of this comment. 

TCCA requested the FAA confirm that 

special condition no. 33(a), applicability 

for engine electrical systems, is not 

applicable to energy storage systems 

(ESS) but it does include the interface 

between the electric engine and the 

propulsion power source. TCCA further 

explained this comment is a request for 

clarification, rather than modification, 

of this special condition. 

Special condition no. 33 does not 

apply to ESS but does apply to the 

interface between the engine and ESS. 

No changes were made to these special 

conditions as a result of this comment. 

TCCA stated that proposed special 

condition no. 33(b), Electrical systems, 

is written in a way that implies 

electrical load shedding is mandatory 

even when not needed and explained 

electrical load shedding should only be 

implemented if required. TCCA 

recommended adding ‘‘if required’’ 

between parenthesis like the following: 

‘‘. . . , and electrical load shedding (if 

required), . . .’’ to special condition no. 

33(b). 

The FAA concurs with TCCA’s 

recommendation and has revised 

special condition no. 33(b) accordingly. 

Load shedding is a capability of the 

electric engine’s power distribution 

system. 

TCCA requested the FAA define the 

term ‘‘abnormal condition,’’ which is 

used in special condition 33(c)(2), 

Electrical power distribution, and 

offered several potential interpretations 

of the term. They also asked if an 

abnormal condition is any failure 

condition not considered extremely 

improbable, and if it is equivalent to the 

definition from MIL–STD–704F. The 

FAA’s use of the term ‘‘abnormal 

conditions’’ does not refer to internal 

malfunctions or failures. It refers to 

operating conditions such as: 

• System or components outside of 

normal statistical variation due to 

degradation, or installation error 

• Unusual environmental conditions 

such as extreme temperature, humidity, 

FOD impact, severe lightning, HIRF, or 

atmospheric radiation 

• Infrequent scenarios such as 

landing on icy runways, rejected takeoffs or balked landings, extended 

ground idling, or taxiing in hot 

environments. 

TCCA also requested the FAA provide 

a definition for ‘‘safe transmission,’’ 

which is used in special condition 

33(c)(3). 

The FAA concurs with TCCA’s 

requests and has added definitions of 

the terms ‘‘abnormal condition’’ and 

‘‘safe transmission’’ to the preamble 

discussion for special condition no. 33. 

TCCA observed that proposed special 

condition nos. 33(e)(1) and (e)(2), 

Electrical power characteristics, were 

linked with an ‘‘or’’ indicating that 

either condition could be applied, but 

not both. TCCA stated both (e)(1) and 

(e)(2) are applicable, and therefore 

recommended the FAA revise special 

condition no. 33(e) to replace the ‘‘or’’ 

with an ‘‘and.’’ 

The FAA concurs with TCCA’s 

recommendation and has revised 

special condition no. 33(e) accordingly. 

TCCA indicated that noise 

certification requirements are applicable 

at the airframe level and not at the 

engine level. TCCA explained the NPSC 

implies that an engine applicant 

demonstration of compliance to 14 CFR 

part 36 is part of the special conditions. 

However, TCCA stated there is no 

definition of requirements within the 

special conditions other than the 

preamble section titled the Type 

Certification Basis. TCCA requested that 

the FAA remove the statement ‘‘In 

addition to the applicable airworthiness 

regulations and special conditions, the 

BETA Model H500A electric engines 

must comply with the noise certification 

requirements of 14 CFR part 36’’ from 

the preamble. GAMA also commented 

on this issue and stated the noise 

certification requirements do not apply 

to engines and requested the FAA 

remove this statement from the 

preamble. 

The FAA concurs with TCCA’s and 

GAMA’s requests and has updated the 

preamble of these special conditions 

accordingly. 

TCCA suggested that the reference to 

‘‘consensus standards’’ in proposed 

special condition 1(b), Applicability, 

may not be necessary. TCCA stated that 

consensus standards are not a means of 

compliance but instead, they are 

derived/alternate requirements (i.e., 

ASTM) that are formulated by industry 

to be used in lieu of published 

regulatory guidance material. TCCA 

further suggested that the use of 

derived/alternate requirements in lieu of 

the published standards is to be 

accepted by the Administrator as being 

equivalent to the published standards. 

Then, the means of compliance to the 

consensus standards are to be accepted 

by the Administrator. TCCA 

recommended reducing the text in 

special condition no. 1(b) to the 

following: ‘‘(b) the applicant must 

comply with this part using a means of 

compliance accepted by the 

administrator.’’ 

The FAA does not concur with 

TCCA’s suggested change. The reference 

to consensus standards provides 

clarification about potential sources of 

information that may be used to 

determine a means of compliance. The 

comment indicates a need to clarify how 

consensus standards are used. For 

example, consensus standards 

developed by the standards 

development organizations (SDOs) 

typically function as a method of 

compliance to 14 CFR requirements or 

special conditions. Published FAA 

guidance can function either as a means 

of compliance, method of compliance, 

or both. Special condition 1(b) permits 

consensus standards to be used for 

showing compliance to certification 

requirements, but they are not a 


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requirement of that special condition. 

Therefore, special condition 1(b) 

supplements the performance-based 

special conditions by requiring a means 

of compliance, which could include 

consensus standards developed by 

SDOs. Further, special condition 1(b) is 

intended to be equivalent to 

§ 23.2010(a), which also refers to 

consensus standards as a potential 

means of compliance. No changes were 

made to the special conditions as a 

result of this comment. 

TCCA observed the BETA proposed 

special condition no. 17 does not 

include a reference to § 33.75(a)(3) 

which appears in the magniX special 

conditions and recommended that the 

FAA explain this difference in the 

discussion for that special condition in 

the preamble to avoid ambiguity 

between the relative project 

requirements. 

The FAA does not concur with 

TCCA’s recommendation. The NPSC for 

the magniX magni350 and magni650 

model electric engines originally 

proposed to incorporate § 33.75(a)(3) 

into special condition no. 17. The FAA 

received a comment suggesting that 

§ 33.75(a)(3) may not be needed for 

those engines. In the final special 

conditions (Docket No. FAA–2020– 

0894, Special Conditions No. 33–022– 

SC), the FAA agreed with the comment 

and removed the reference to 

§ 33.75(a)(3). No changes were made to 

these special conditions as a result of 

this comment. 

The FAA received comments from 

TC. 

TC disagreed with the text in 

proposed special condition nos. 17(a) 

and 17(c) which say, ‘‘The applicant 

must comply . . .’’ TC stated that the 

onus to show compliance with the 

applicable requirements with the intent 

to obtain a type certificate is on the 

applicant and that the elements that 

comply with the requirements 

themselves are those objects of the type 

certificate, such as the engine and its 

systems. TC further explained it is not 

clear to state that the applicant must 

comply, where it is in fact the engine/ 

systems which must comply with the 

requirements. Instead, the applicant 

shows compliance. TC suggested 

changing the phrase to read ‘‘The 

applicant must show compliance . . .’’ 

TC’s proposed change is not 

necessary. Section 21.20, ‘‘Compliance 

with Applicable Requirements’’ 

contains an example that supports the 

language used in special conditions nos. 

17(a) and (c). Specifically, § 21.20(b) 

specifies the applicant must ‘‘provide a 

statement certifying that the applicant 

has complied with the applicable 

requirements,’’ which indicates the 

applicant complies with the applicable 

requirements. . No changes were made 

as a result of this comment. 

TC observed the text in proposed 

special condition no. 17(d)(1), Safety 

Analysis, does not include special 

condition no. 31, Operation with 

Variable Pitch Propeller. TC 

recommended that the FAA either add 

a reference to special condition no. 31 

in special condition no. 17(d)(1) because 

BETA’s electric engine may be installed 

with a variable pitch propeller or 

provide a rationale for not including it. 

The FAA does not concur with TC’s 

suggestion to add a reference to special 

condition no. 31. Adding special 

condition no. 31 is not necessary 

because the specific engine model BETA 

intends to certify is not designed to use 

a variable pitch propeller. No changes 

were made to the special conditions as 

a result of this comment. 

TC indicated there is a similar 

electrical engine special condition in 

the magniX special conditions (Special 

Conditions No. 33–022–SC) that 

contains an ingestion requirement that 

does not appear in the BETA special 

conditions. TC referred to special 

condition no. 18(d) in the magniX 

special conditions, which requires 

ingestion sources that are not evaluated 

must be declared in the engine 

installation manual. TC recommended 

that the FAA either revise the BETA 

special conditions to add this 

requirement or provide the rationale for 

not including it. 

The FAA does not concur with TC’s 

request to revise the BETA special 

conditions to include special condition 

no. 18(d) from the magniX special 

conditions. Special condition no. 18(d) 

was intended to ensure ingestion 

sources that are not applicable to an 

electric engine are enunciated in the 

engine documentation. The list of 

required ingestion sources in BETA 

special condition nos. 18(a) and (b) are 

more prescriptive compared to the 

ingestion requirements in the published 

magniX special condition no. 18(a). 

Therefore, the FAA has determined 

special condition no. 18(d) is not 

necessary to include in the BETA 

special conditions because exceptions to 

the ingestion requirement would be 

specified and managed using special 

condition no. 18(c), which is similar to 

how exceptions are managed by the 

existing part 33 ingestion requirements. 

No changes were made to the special 

conditions as a result of this comment. 

TC noted that proposed special 

condition no. 33(c)(1) introduces the 

term ‘‘electrical power plant’’ and 

recommended that the FAA update the 

preamble to describe an electrical power 

plant. 

The FAA disagrees with TC’s 

recommendation to define ‘‘electrical 

power plant’’ because the FAA revised 

special condition no. 33(c)(1) in these 

final special conditions to change the 

term ‘‘electrical power plant’’ to 

‘‘powerplant,’’ as that term is defined in 

part 23, subpart E, in § 23.2400(a) 

powerplant installation, to include each 

component necessary for propulsion, 

which affects propulsion safety, or 

provides auxiliary power to the 

airplane, and in the installation 

requirements in subpart E of parts 25, 

27, and 29. 

TC observed that the proposed system 

safety assessments in proposed special 

condition no. 33(h), and proposed 

special condition no.10(g) are different 

in that special condition no. 10(g) 

requires the rates of hazardous and 

major faults to be declared in the engine 

installation manual and special 

condition no. 33(h) does not. TC 

recommended that the FAA either 

revise special condition no. 33(h) to 

match special condition no. 10(g) or 

provide a rationale for why they are 

different. 

The FAA agrees with TC’s 

recommendation and has revised final 

special condition no. 33(h) to match 

special condition no. 10(g). 

The FAA received comments from 

GAMA. 

GAMA recommended that the FAA 

align the special conditions for the 

H500A electric engine with the electric 

engine requirements included in the 

certification basis for special class 

powered lift aircraft that certify an 

electric engine as part of the aircraft 

type certification. GAMA stated that 

there are technical variations between 

the H500A proposed special conditions 

and the electric engine airworthiness 

criteria outlined in the Special Class 

Airworthiness Criteria for the poweredlift and cited special condition no. 17(c) 

and special condition no. 33(c) as 

examples of these technical differences. 

GAMA further stated these variations 

could lead to two electric engines used 

in the same aircraft having different 

requirements based solely on whether 

the engine is certified as part of the 

aircraft or under part 33. AUVSI also 

commented on the importance of 

applying consistent requirements across 

projects and requested the FAA 

substantiate any inconsistencies 

introduced to the electric engine 

requirements. 

There are no intended technical 

differences between the proposed 

special class airworthiness criteria for 

the powered lift in draft Advisory 


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Circular 21.17–4 (AC 21.17–4) and the 

BETA special conditions. For example, 

the corresponding criteria to BETA 

special condition nos. 17(c) and 33(c) 

are PL.3375(f) and PL.3326(c) 

respectively. The engine requirements 

are documented differently between the 

BETA special conditions and poweredlift airworthiness criteria proposed in 

draft AC 21.17–4 because special 

conditions are written in accordance 

with the requirements of § 21.16, and 

the powered-lift airworthiness criteria 

in draft AC 21.17–4 are not specific to 

one applicant. There are also some 

minor differences in the documentation 

requirements because engines are 

approved with the special class aircraft, 

so some engine details may be included 

in the aircraft manuals. No changes 

were made to the special conditions as 

a result of this comment. 

GAMA indicated proposed special 

condition no. 9, Overspeed, lacks clarity 

regarding whether ‘‘rotor’’ refers to an 

internal electric engine component or an 

actual propulsive propeller. GAMA 

recommended the FAA provide the 

necessary clarification to address this 

ambiguity. 

The FAA agrees with GAMA’s 

recommendation. The term ‘‘rotor’’ in 

the proposed special conditions is 

intended to refer to an engine 

component and not a propulsive 

propeller. A rotor in an electric engine 

may consist of a circular disk and 

magnets fixed at the outer 

circumference that rotates inside a 

stationary casing configured with 

electrical windings (or coils), or a 

rotating cylindrical casing with magnets 

fixed on the inside surface that rotates 

around a stationary set of windings (or 

coils). Each configuration is attached to 

a rotating shaft that drives a propulsive 

device, such as a propeller. Projectspecific decisions will be made 

regarding which engine parts are 

applicable to the overspeed 

requirement. No changes were made to 

the special conditions as a result of this 

comment. 

GAMA stated that proposed special 

condition nos. 30(a) and (b), 

Containment, utilize language tailored 

to an engine design featuring a nonrotating stator situated outside the rotor. 

GAMA recommended the FAA explore 

a rule version that is less designspecific. GAMA advised against 

presuming that all rotating components 

possess a case, particularly that the rotor 

is contained within the stator. 

The FAA does not concur with 

GAMA’s recommendation Special 

condition 30(a) is intended to account 

for rotor designs with exceedingly large 

margins to a rotor burst. The special 

condition does not specify a particular 

rotor design. However, the amount of 

margin needed to satisfy the 

requirement would be determined based 

on the engine’s design. Special 

condition 30(b) is intended to account 

for rotors located inside a static stator 

case. No changes were made to the 

special conditions as a result of this 

comment. 

GAMA commented proposed special 

condition nos. 33(c)(1) and (c)(3), 

Electrical power distribution for engine 

electrical systems, set forth distinct 

criteria for the automatic measures 

needed when electrical-energy 

generation encounters faults, which 

diverges from the corresponding 

requirements in the special class 

airworthiness criteria for powered-lift. 

GAMA indicated there are no evident 

variations in electric engine 

configurations that warrant this 

inconsistency. GAMA recommended 

that the FAA align these regulations to 

ensure that electric engines certified as 

part of an aircraft or under part 33 

adhere to uniform standards. 

Proposed special condition nos. 

33(c)(1) and (c)(3) are not the same as 

the corresponding engine requirements 

in the powered-lift airworthiness 

criteria used in another project. 

Proposed special condition no. 33(c)(1) 

protects engine electrical systems from 

faulted electrical energy generation or 

storage devices. Proposed special 

condition no. 33(c)(3) prescribes a 

means of compliance (fault isolation) to 

address (c)(1), but the means of 

compliance should be tied to the safety 

assessment required in special 

condition no. 33(g), which accounts for 

aircraft-level effects from faulted 

electrical-energy generation or storage 

devices. The aircraft effects should not 

be assumed in the engine requirements, 

and therefore the FAA revised special 

condition no. 33(c)(3) to accommodate 

other potential protection systems that 

might be more appropriate. 

Accordingly, final special condition no. 

33(c)(3) is changed to, ‘‘The system 

must provide mechanical or automatic 

means of isolating a faulted electrical 

energy generation or storage device from 

leading to hazardous engine effects, as 

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

of these special conditions, or 

detrimental effects in the intended 

aircraft application.’’ 

The phrase, ‘‘or detrimental engine 

effects in the intended aircraft 

application’’ was relocated to special 

condition no. 33(c)(3) to maintain the 

connection with special condition no. 

33(g). 

GAMA commented proposed special 

condition no. 33(g), Electrical system 

failures of engine electrical systems, 

extends beyond the comparable part 33 

regulation § 33.28(d), which is originally 

limited to the engine control system. 

GAMA suggested that expanding this 

special condition to encompass the 

engine electrical system instead of 

solely the engine control system entails 

subjecting electrical components within 

the engine, such as windings, to failure 

requirements historically not applied to 

engine mechanical components. GAMA 

also stated that field experience 

indicates that component failures are 

unpredictable based on wear and 

susceptible to random failures. Electric 

engine components, like windings and 

insulation, are better addressed using 

methods akin to those applied to 

traditional engines to address 

mechanical failure aspects. GAMA 

recommended the FAA revise this 

special condition to align with the 

existing regulatory framework. The FAA 

does not concur with GAMA’s 

recommendation. By their nature, FAA 

special conditions are issued when the 

‘‘existing regulatory framework’’ is 

inadequate or insufficient. 14 CFR 

21.16; see also Amdt. 21–51. The 

existing requirements for engine control 

systems were developed to address the 

failure characteristics of electrical 

systems. For combustion engines, the 

only electrical system is the engine 

control, but this is not the case for 

electric engines where electrical systems 

extend beyond those addressed by 

§ 33.28(d). Special condition no. 33(g) 

for the BETA electric engine provides 

the same level of safety as § 33.28(d) by 

applying the safety criteria for electrical 

systems to all the electrical systems in 

the engine. This includes the highvoltage systems used in the electric 

engine. No changes were made to the 

special conditions as a result of this 

comment. 

The FAA received several comments 

from an individual commenter and 

received similar comments from magniX 

(although these commenters provided 

separate comments). 

An individual and magniX 

commented proposed special condition 

nos. 1(b) and (c) state that a means of 

compliance, which may include 

consensus standards, must be ‘‘accepted 

by the Administrator’’ and ‘‘in a form 

and manner acceptable to the 

Administrator.’’ The individual and 

magniX stated that these paragraphs are 

directly out of § 23.2010, which 

contains performance-based language. 

The individual and magniX considered 

the BETA electric engine special 

conditions to be largely prescriptive and 

not performance-based, which they 

stated would make special condition 


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nos. 1(b) and (c) superfluous. The 

individual suggested these requirements 

introduce a new regulatory layer to 

prescriptive requirements and may lead 

to inadvertent consequences, while 

magniX stated that requiring a 

performance-based process for 

establishing means of compliance with 

prescriptive regulations is unnecessary 

and overly burdensome to applicants 

and regulators. The individual and 

magniX recommended the FAA not 

adopt proposed special condition nos. 

1(b) and (c), and the individual also 

recommended holding public 

consultations with stakeholders as was 

done when part 23 was being reworked 

into a performance-based form. 

The FAA does not concur with the 

individual’s and magniX’s 

recommendation. While special 

conditions are rules of particular, not 

general applicability, the FAA expects 

that special condition nos. 1(b) and (c) 

support the FAA’s transition to a 

performance-based approach for 

developing new requirements. Although 

the BETA special conditions are not 

prescriptive, they provide safety criteria 

that address hazards presented by the 

new electric engine technology used in 

the BETA H500A engine. Special 

condition nos. 1(b) and (c) will be used 

to incorporate the additional details that 

apply to the BETA H500A engine design 

using accepted means of compliance. 

No changes were made to these special 

conditions as a result of this comment. 

GAMA and magniX commented that 

special condition nos. 10(g), 15(b), and 

17(f) would require applicants to 

declare proprietary information in the 

engine installation manual, these 

documentation requirements establish a 

precedent beyond that required of their 

existing reciprocating or turbine 

counterparts, and these requirements 

increase the risk that sensitive 

information is disclosed. MagniX stated 

that while it is understood this 

information is used during aircraft-level 

certification efforts, traditional data 

sharing agreements sufficiently provide 

the integrator with the required 

information while respecting the 

proprietary nature of the data. MagniX 

also stated requiring additional data in 

the engine installation manual overly 

constrains the means whereby this 

information is shared when compared 

with established means, introducing 

additional commercial risk. GAMA also 

stated proposed special condition nos. 

10(g), 15(b), and 17(f) are a requirement 

for a manufacturer to disclose sensitive 

proprietary safety analysis in the engine 

installation manual, a requirement not 

currently imposed on part 33 engines. 

Additionally, GAMA stated the FAA has 

not provided adequate justification for 

why an electric engine necessitates this 

information in a manual. An individual 

provided a similar comment regarding 

proposed special condition nos. 10(g) 

and 17(f), and stated that historically 

such information was captured in other 

documents such as the engine control 

systems interface control document and 

systems safety assessment, that were 

only provided to the installer. 

MagniX requested the FAA not adopt 

the documentation requirements in 

proposed special conditions 10(g), 15(b), 

and 17(f), and proposed that these data 

be provided to integrators through 

generic ‘‘installation instructions’’ in 

lieu of the engine installation manual. 

GAMA also requested the FAA 

reconsider its approach and/or provide 

justification for the added requirement 

of disclosing sensitive proprietary safety 

analysis in the engine installation 

manual. An individual requested the 

FAA preserve the engine OEM’s 

flexibility to document and protect 

proprietary data by changing 

‘‘installation manual’’ to the more 

generic ‘‘installation instructions,’’ 

which consist of other documents such 

as interface control drawings, technical 

memorandums, or other installer 

requested documentation. The 

individual further stated that this 

change would harmonize the special 

condition with § 23.2400(e) which uses 

the verbiage of ‘‘installation 

instructions,’’ and this change could be 

promulgated to other special condition 

paragraphs which refer to the engine 

installation manual. 

The FAA does not concur with 

magniX’s and GAMA’s comments that 

special conditions 10(g), 15(b), and 17(f) 

require disclosing sensitive information. 

The requirements cited in their 

comment do not require disclosure of 

sensitive information. As discussed in 

the NPSC, the documentation 

requirements in special conditions nos. 

10(g), 15(b), and 17(f) are expected to 

ensure that the engine is used safely and 

properly by constraining the installation 

of electric engines to only aircraft types 

(configurations, flight capabilities, etc.) 

that were used by the engine 

manufacturer to determine the engine 

ratings, limits, performance 

characteristics, as well as the reliability 

and criticality of engine systems and 

parts. 

These documentation requirements 

are intended, and the FAA finds 

necessary, to ensure enough information 

is included to safeguard compatibility 

between the electric engine and aircraft, 

and to prevent the engine from being 

used in an aircraft type that requires 

safety features or performance 

characteristics that are not available 

from a type certificated engine. For 

example, electric engines designed for 

vertical lift in distributed propulsion 

tilt-wing aircraft provide propulsion and 

act as flight control surfaces, and 

therefore these engines have different 

performance requirements than those 

used in conventional normal category 

airplanes. In addition, the FAA agrees 

with the commenters’ suggestion to 

remove the requirement that specifies 

the information must be located in the 

engine installation manual. These 

special conditions do not need to 

specify the document that must have the 

information, but only that the 

information must be provided to the 

installer in accordance with the engine 

installation instructions under § 33.5, 

‘‘Instruction manual for installing and 

operating the engine.’’. The proposed 

special conditions are modified to 

incorporate this change. 

The FAA received a comment from 

UPSFF. 

UPSFF requested that the FAA align 

these special conditions with the 

electric engine requirements included in 

the certification basis for special class 

powered lift aircraft that certify an 

electric engine as part of the aircraft 

type certification. 

As stated previously, the engine 

requirements in the BETA special 

conditions are documented differently 

from proposed powered lift 

airworthiness criteria in draft AC 21.17– 

4 because special conditions are written 

in accordance with the requirements of 

§ 21.16, and the proposed powered-lift 

airworthiness criteria in draft AC 21.17– 

4 are not specific to one applicant. 

Special conditions are project-specific 

rules of particular applicability, and the 

special conditions for this electric 

engine are based on certain novel or 

unusual design features. Special 

conditions may evolve to a general 

standard as more experience is gained 

with certifying the new technology (see 

Amdt. 21–51). No changes were made to 

these special conditions as a result of 

this comment. 

The FAA received an anonymous 

comment. The commenter stated the 

reference to § 21.17(a) in the preamble 

of the NPSC seems contradictory to the 

language in § 21.17(b). The commenter 

explained that since § 21.17(b) applies 

to ‘‘special classes of aircraft, including 

the engines and propellers installed 

thereon (e.g., gliders, airships, and other 

nonconventional aircraft) . . .’’ an 

electric engine would be installed on a 

special class of aircraft as described in 

§ 21.17(b) and referring to § 21.17(a) 

seems to contradict the language in 

paragraph (b) of that section. 


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The FAA does not concur with the 

comment that indicates the reference to 

§ 21.17(a) is contradictory to the 

language in § 21.17(b). Section 21.17(a) 

provides requirements for developing a 

certification basis for an established 

aviation product, which includes 

aircraft, engines, and propellers. The 

BETA electric engine is an aircraft 

engine, which falls under § 21.17(a), and 

therefore § 21.17(a) is the appropriate 

reference for this project. Section 

21.17(b) provides requirements for 

developing a certification basis for 

special classes of aircraft, such as 

powered-lift. No changes were made as 

a result of this comment. 

The FAA also determined that the 

following changes were necessary. The 

phrase, ‘‘In addition’’ is added to special 

condition no. 4, Fire protection, to 

connect the introduction sentence to (a) 

and (b) and avoid confusion. The FAA 

also revised the special conditions to 

use consistent references to hazardous 

engine effects. Therefore, the phrase ‘‘as 

defined in special condition no. 17 of 

these special conditions’’ was added 

wherever ‘‘hazardous engine effects’’ is 

mentioned. 

The FAA recognizes energy 

regeneration might not be a feature for 

some electric engines that operate at 

their limits, so special condition no. 23 

was changed to specify that ‘‘The 

endurance demonstration must include 

increases and decreases of the engine’s 

power settings, energy regeneration, and 

dwellings at the power settings and 

energy regeneration for sufficient 

durations that produce the extreme 

physical conditions the engine 

experiences at rated performance levels, 

operational limits, and at any other 

conditions or power settings, including 

energy regeneration that are required to 

verify the limit capabilities of the 

engine.’’ 

In addition, proposed special 

condition no. 31 was not adopted 

because the specific engine model BETA 

intends to certify is not designed to use 

a variable pitch propeller. Except as 

discussed above, these special 

conditions are adopted as proposed. 

Applicability 

As discussed above, these special 

conditions are applicable to BETA 

Model H500A electric engines. Should 

BETA apply at a later date for a change 

to the type certificate to include another 

model on the same type certificate, 

incorporating the same novel or unusual 

design feature, these special conditions 

would apply to that model as well. 

Conclusion 

This action affects only BETA Model 

H500A electric engines. It is not a rule 

of general applicability and affects only 

the applicant who applied to the FAA 

for approval of these features on the 

airplane. 

List of Subjects in 14 CFR Part 33 

Aircraft, Aviation safety, Reporting 

and recordkeeping requirements. 

Authority Citation 

The authority citation for these 

special conditions is as follows: 

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

44701, 44702, 44704. 

The Special Conditions 

D Accordingly, pursuant to the 

authority delegated to me by the 

Administrator, the following special 

conditions are issued as part of the type 

certification basis for BETA 

Technologies Inc. Model H500A electric 

engines. The applicant must also 

comply with the certification 

procedures set forth in part 21. 

(1) Applicability 

(a) Unless otherwise noted in these 

special conditions, the engine design 

must comply with the airworthiness 

standards for aircraft engines set forth in 

part 33, except for those airworthiness 

standards that are specifically and 

explicitly applicable only to 

reciprocating and turbine aircraft 

engines or as specified herein. 

(b) The applicant must comply with 

this part using a means of compliance, 

which may include consensus 

standards, accepted by the 

Administrator. 

(c) The applicant requesting 

acceptance of a means of compliance 

must provide the means of compliance 

to the FAA in a form and manner 

acceptable to the Administrator. 

(2) Engine Ratings and Operating Limits 

In addition to § 33.7(a), the engine 

ratings and operating limits must be 

established and included in the type 

certificate data sheet based on: 

(a) Shaft power, torque, rotational 

speed, and temperature for: 

(1) Rated takeoff power; 

(2) Rated maximum continuous 

power; and 

(3) Rated maximum temporary power 

and associated time limit. 

(b) Duty cycle and the rating at that 

duty cycle. The duty cycle must be 

declared in the engine type certificate 

data sheet. 

(c) Cooling fluid grade or 

specification. 

(d) Power-supply requirements. 

(e) Any other ratings or limitations 

that are necessary for the safe operation 

of the engine. 

(3) Materials 

The engine design must comply with 

§ 33.15. 

(4) Fire Protection 

The engine design must comply with 

§ 33.17(b) through (g). In addition— 

(a) The design and construction of the 

engine and the materials used must 

minimize the probability of the 

occurrence and spread of fire during 

normal operation and failure conditions 

and must minimize the effect of such a 

fire. 

(b) High-voltage electrical wiring 

interconnect systems must be protected 

against arc faults that can lead to 

hazardous engine effects as defined in 

special condition no. 17(d)(2) of these 

special conditions. Any non-protected 

electrical wiring interconnects must be 

analyzed to show that arc faults do not 

cause a hazardous engine effect. 

(5) Durability 

The engine design and construction 

must minimize the development of an 

unsafe condition of the engine between 

maintenance intervals, overhaul 

periods, or mandatory actions described 

in the applicable ICA. 

(6) Engine Cooling 

The engine design and construction 

must comply with § 33.21. In addition, 

if cooling is required to satisfy the safety 

analysis as described in special 

condition no. 17 of these special 

conditions, the cooling system 

monitoring features and usage must be 

documented and provided to the 

installer as part of the requirements in 

§ 33.5. 

(7) Engine Mounting Attachments and 

Structure 

The engine mounting attachments and 

related engine structures must comply 

with § 33.23. 

(8) Accessory Attachments 

The engine must comply with § 33.25. 

(9) Overspeed 

(a) A rotor overspeed must not result 

in a burst, rotor growth, or damage that 

results in a hazardous engine effect, as 

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

of these special conditions. Compliance 

with this paragraph must be shown by 

test, validated analysis, or a 

combination of both. Applicable 

assumed rotor speeds must be declared 

and justified. 

(b) Rotors must possess sufficient 

strength with a margin to burst above 


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certified operating conditions and above 

failure conditions leading to rotor 

overspeed. The margin to burst must be 

shown by test, validated analysis, or a 

combination thereof. 

(c) The engine must not exceed the 

rotor speed operational limitations that 

could affect rotor structural integrity. 

(10) Engine Control Systems 

(a) Applicability. The requirements of 

this special condition apply to any 

system or device that is part of the 

engine type design that controls, limits, 

monitors, or protects engine operation, 

and is necessary for the continued 

airworthiness of the engine. 

(b) Engine control. The engine control 

system must ensure that the engine does 

not experience any unacceptable 

operating characteristics or exceed its 

operating limits, including in failure 

conditions where the fault or failure 

results in a change from one control 

mode to another, from one channel to 

another, or from the primary system to 

the back-up system, if applicable. 

(c) Design Assurance. The software 

and complex electronic hardware, 

including programmable logic devices, 

must be— 

(1) Designed and developed using a 

structured and systematic approach that 

provides a level of assurance for the 

logic commensurate with the hazard 

associated with the failure or 

malfunction of the systems in which the 

devices are located; and 

(2) Substantiated by a verification 

methodology acceptable to the 

Administrator. 

(d) Validation. All functional aspects 

of the control system must be 

substantiated by test, analysis, or a 

combination thereof, to show that the 

engine control system performs the 

intended functions throughout the 

declared operational envelope. 

(e) Environmental Limits. 

Environmental limits that cannot be 

adequately substantiated by endurance 

demonstration, validated analysis, or a 

combination thereof must be 

demonstrated by the system and 

component tests in special condition no. 

27 of these special conditions. 

(f) Engine control system failures. The 

engine control system must— 

(1) Have a maximum rate of loss of 

power control (LOPC) that is suitable for 

the intended aircraft application. The 

estimated LOPC rate must be 

documented and provided to the 

installer as part of the requirements in 

§ 33.5; 

(2) When in the full-up configuration, 

be single-fault tolerant, as determined 

by the Administrator, for electrical, 

electrically detectable, and electronic 

failures involving LOPC events; 

(3) Not have any single failure that 

results in hazardous engine effects as 

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

of these special conditions; and 

(4) Ensure failures or malfunctions 

that lead to local events in the aircraft 

do not result in hazardous engine 

effects, as defined in special condition 

no. 17(d)(2) of these special conditions, 

due to engine control system failures or 

malfunctions. 

(g) System safety assessment. The 

applicant must perform a system safety 

assessment. This assessment must 

identify faults or failures that affect 

normal operation, together with the 

predicted frequency of occurrence of 

these faults or failures. The intended 

aircraft application must be taken into 

account to assure that the assessment of 

the engine control system safety is valid. 

The rates of hazardous and major faults 

must be documented and provided to 

the installer as part of the requirements 

in § 33.5. 

(h) Protection systems. The engine 

control devices and systems’ design and 

function, together with engine 

instruments, operating instructions, and 

maintenance instructions, must ensure 

that engine operating limits that can 

lead to a hazard will not be exceeded in 

service. 

(i) Aircraft supplied data. Any single 

failure leading to loss, interruption, or 

corruption of aircraft-supplied data 

(other than power-command signals 

from the aircraft), or aircraft-supplied 

data shared between engine systems 

within a single engine or between fully 

independent engine systems, must— 

(1) Not result in a hazardous engine 

effect, as defined in special condition 

no. 17(d)(2) of these special conditions, 

for any engine installed on the aircraft; 

and 

(2) Be able to be detected and 

accommodated by the control system. 

(j) Engine control system electrical 

power. 

(1) The engine control system must be 

designed such that the loss, 

malfunction, or interruption of the 

control system electrical power source 

will not result in a hazardous engine 

effect, unacceptable transmission of 

erroneous data, or continued engine 

operation in the absence of the control 

function. Hazardous engine effects are 

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

of these special conditions. The engine 

control system must be capable of 

resuming normal operation when 

aircraft-supplied power returns to 

within the declared limits. 

(2) The applicant must identify, 

document, and provide to the installer 

as part of the requirements in § 33.5, the 

characteristics of any electrical power 

supplied from the aircraft to the engine 

control system, including transient and 

steady-state voltage limits, and any 

other characteristics necessary for safe 

operation of the engine. 

(11) Instrument Connection 

The applicant must comply with 

§ 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 


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structural static part, the failure of 

which can result in a hazardous engine 

effect, as defined in special condition 

no. 17(d)(2) of these special conditions, 

due to a low-cycle fatigue (LCF) 

mechanism. A life limit is an 

operational limitation that specifies the 

maximum allowable number of flight 

cycles that a part can endure before the 

applicant must remove it from the 

engine. 

(b) In establishing the integrity of each 

critical part or life-limited part, the 

applicant must provide to the 

Administrator the following three plans 

for approval: 

(1) an engineering plan, as defined in 

§ 33.70(a); 

(2) a manufacturing plan, as defined 

in § 33.70(b); and 

(3) a service-management plan, as 

defined in § 33.70(c). 

(14) Lubrication System 

(a) The lubrication system must be 

designed and constructed to function 

properly between scheduled 

maintenance intervals in all flight 

attitudes and atmospheric conditions in 

which the engine is expected to operate. 

(b) The lubrication system must be 

designed to prevent contamination of 

the engine bearings and lubrication 

system components. 

(c) The applicant must demonstrate 

by test, validated analysis, or a 

combination thereof, the unique 

lubrication attributes and functional 

capability of (a) and (b). 

(15) Power Response 

(a) The design and construction of the 

engine, including its control system, 

must enable an increase— 

(1) From the minimum power setting 

to the highest rated power without 

detrimental engine effects; 

(2) From the minimum obtainable 

power while in-flight and while on the 

ground to the highest rated power 

within a time interval determined to be 

appropriate for the intended aircraft 

application; and 

(3) From the minimum torque to the 

highest rated torque without detrimental 

engine effects in the intended aircraft 

application. 

(b) The results of (a)(1), (a)(2), and 

(a)(3) of this special condition must be 

documented and provided to the 

installer as part of the requirements in 

§ 33.5. 

(16) Continued Rotation 

If the design allows any of the engine 

main rotating systems to continue to 

rotate after the engine is shut down 

while in-flight, this continued rotation 

must not result in any hazardous engine 

effects, as defined in special condition 

no. 17(d)(2) of these special conditions. 

(17) Safety Analysis 

(a) The applicant must comply with 

§ 33.75(a)(1) and (a)(2) using the failure 

definitions in special condition no. 

17(d) of these special conditions. 

(b) The primary failure of certain 

single elements cannot be sensibly 

estimated in numerical terms. If the 

failure of such elements is likely to 

result in hazardous engine effects, then 

compliance may be shown by reliance 

on the prescribed integrity requirements 

of § 33.15 and special condition nos. 9 

and 13 of these special conditions, as 

applicable. These instances must be 

stated in the safety analysis. 

(c) The applicant must comply with 

§ 33.75(d) and (e) using the failure 

definitions in special condition no. 

17(d) of these special conditions, and 

the ICA in § 33.4. 

(d) Unless otherwise approved by the 

Administrator, the following definitions 

apply to the engine effects when 

showing compliance with this 

condition: 

(1) A minor engine effect does not 

prohibit the engine from performing its 

intended functions in a manner 

consistent with § 33.28(b)(1)(i), 

(b)(1)(iii), and (b)(1)(iv), and the engine 

complies with the operability 

requirements of special condition no. 15 

and special condition no. 25 of these 

special conditions, as appropriate. 

(2) The engine effects in § 33.75(g)(2) 

are hazardous engine effects with the 

addition of: 

(i) Electrocution of the crew, 

passengers, operators, maintainers, or 

others; and 

(ii) Blockage of cooling systems that 

could cause the engine effects described 

in § 33.75(g)(2) and special condition 

17(d)(2)(i) of these special conditions. 

(3) Any other engine effect is a major 

engine effect. 

(e) The intended aircraft application 

must be taken into account when 

performing the safety analysis. 

(f) The results of the safety analysis, 

and the assumptions about the aircraft 

application used in the safety analysis, 

must be documented and provided to 

the installer as part of the requirements 

in § 33.5. 

(18) Ingestion 

(a) Rain, ice, and hail ingestion must 

not result in an abnormal operation 

such as shutdown, power loss, erratic 

operation, or power oscillations 

throughout the engine operating range. 

(b) Ingestion from other likely sources 

(birds, induction system ice, foreign 

objects—ice slabs) must not result in 

hazardous engine effects defined by 

special condition no. 17(d)(2) of these 

special conditions, or unacceptable 

power loss. 

(c) If the design of the engine relies on 

features, attachments, or systems that 

the installer may supply, for the 

prevention of unacceptable power loss 

or hazardous engine effects, as defined 

in special condition no. 17(d)(2) of these 

special conditions, following potential 

ingestion, then the features, 

attachments, or systems must be 

documented and provided to the 

installer as part of the requirements in 

§ 33.5. 

(19) Liquid and Gas Systems 

(a) Each system used for lubrication or 

cooling of engine components must be 

designed and constructed to function 

properly in all flight attitudes and 

atmospheric conditions in which the 

engine is expected to operate. 

(b) If a system used for lubrication or 

cooling of engine components is not 

self-contained, the interfaces to that 

system must be defined, documented 

and provided to the installer as part of 

the requirements in § 33.5. 

(c) The applicant must establish by 

test, validated analysis, or a 

combination of both that all static parts 

subject to significant pressure loads will 

not: 

(1) Exhibit permanent distortion 

beyond serviceable limits, or exhibit 

leakage that could create a hazardous 

condition when subjected to normal and 

maximum working pressure with 

margin; 

(2) Exhibit fracture or burst when 

subjected to the greater of maximum 

possible pressures with margin. 

(d) Compliance with special condition 

no. 19(c) of these special conditions 

must take into account: 

(1) The operating temperature of the 

part; 

(2) Any other significant static loads 

in addition to pressure loads; 

(3) Minimum properties 

representative of both the material and 

the processes used in the construction 

of the part; and 

(4) Any adverse physical geometry 

conditions allowed by the type design, 

such as minimum material and 

minimum radii. 

(e) Approved coolants and lubricants 

must be listed, documented, and 

provided to the installer as part of the 

requirements in § 33.5. 

(20) Vibration Demonstration 

(a) The engine must be designed and 

constructed to function throughout its 

normal operating range of rotor speeds 

and engine output power, including 


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defined exceedances, without inducing 

excessive stress in any of the engine 

parts because of vibration and without 

imparting excessive vibration forces to 

the aircraft structure. 

(b) Each engine design must undergo 

a vibration survey to establish that the 

vibration characteristics of those 

components subject to induced 

vibration are acceptable throughout the 

declared flight envelope and engine 

operating range for the specific 

installation configuration. The possible 

sources of the induced vibration that the 

survey must assess are mechanical, 

aerodynamic, acoustical, internally 

induced electromagnetic, installation 

induced effects that can affect the 

engine vibration characteristics, and 

likely environmental effects. This 

survey must be shown by test, validated 

analysis, or a combination thereof. 

(21) Overtorque 

When approval is sought for a 

transient maximum engine overtorque, 

the applicant must demonstrate by test, 

validated analysis, or a combination 

thereof, that the engine can continue 

operation after operating at the 

maximum engine overtorque condition 

without maintenance action. Upon 

conclusion of overtorque tests 

conducted to show compliance with 

this special condition, or any other tests 

that are conducted in combination with 

the overtorque test, each engine part or 

individual groups of components must 

meet the requirements of special 

condition no. 29 of these special 

conditions. 

(22) Calibration Assurance 

Each engine must be subjected to 

calibration tests to establish its power 

characteristics, and the conditions both 

before and after the endurance and 

durability demonstrations specified in 

special conditions nos. 23 and 26 of 

these special conditions. 

(23) Endurance Demonstration 

The applicant must subject the engine 

to an endurance demonstration, 

acceptable to the Administrator, to 

demonstrate the engine’s limit 

capabilities. The endurance 

demonstration must include increases 

and decreases of the engine’s power 

settings, energy regeneration, and 

dwellings at the power settings and 

energy regeneration for sufficient 

durations that produce the extreme 

physical conditions the engine 

experiences at rated performance levels, 

operational limits, and at any other 

conditions or power settings, including 

energy regeneration that are required to 

verify the limit capabilities of the 

engine. 

(24) Temperature Limit 

The engine design must demonstrate 

its capability to endure operation at its 

temperature limits plus an acceptable 

margin. The applicant must quantify 

and justify the margin to the 

Administrator. The demonstration must 

be repeated for all declared duty cycles 

and ratings, and operating 

environments, which would impact 

temperature limits. 

(25) Operation Demonstration 

The engine design must demonstrate 

safe operating characteristics, including 

but not limited to power cycling, 

starting, acceleration, and overspeeding 

throughout its declared flight envelope 

and operating range. The declared 

engine operational characteristics must 

account for installation loads and 

effects. 

(26) Durability Demonstration 

The engine must be subjected to a 

durability demonstration to show that 

each part of the engine has been 

designed and constructed to minimize 

any unsafe condition of the system 

between overhaul periods, or between 

engine replacement intervals if the 

overhaul is not defined. This test must 

simulate the conditions in which the 

engine is expected to operate in service, 

including typical start-stop cycles, to 

establish when the initial maintenance 

is required. 

(27) System and Component Tests 

The applicant must show that systems 

and components that cannot be 

adequately substantiated in accordance 

with the endurance demonstration or 

other demonstrations will perform their 

intended functions in all declared 

environmental and operating 

conditions. 

(28) Rotor Locking Demonstration 

If shaft rotation is prevented by 

locking the rotor(s), the engine must 

demonstrate: 

(a) Reliable rotor locking performance; 

(b) Reliable rotor unlocking 

performance; and 

(c) That no hazardous engine effects, 

as specified in special condition no. 

17(d)(2) of these special conditions, will 

occur. 

(29) Teardown Inspection 

(a) Teardown evaluation. 

(1) After the endurance and durability 

demonstrations have been completed, 

the engine must be completely 

disassembled. Each engine component 

and lubricant must be eligible for 

continued operation in accordance with 

the information submitted for showing 

compliance with § 33.4. 

(2) Each engine component, having an 

adjustment setting and a functioning 

characteristic that can be established 

independent of installation on or in the 

engine, must retain each setting and 

functioning characteristic within the 

established and recorded limits at the 

beginning of the endurance and 

durability demonstrations. 

(b) Non-Teardown evaluation. If a 

teardown cannot be performed for all 

engine components in a non-destructive 

manner, then the inspection or 

replacement intervals for these 

components and lubricants must be 

established based on the endurance and 

durability demonstrations and must be 

documented in the ICA in accordance 

with § 33.4. 

(30) Containment 

The engine must be designed and 

constructed to protect against likely 

hazards from rotating components as 

follows— 

(a) The design of the stator case 

surrounding rotating components must 

provide for the containment of the 

rotating components in the event of 

failure, unless the applicant shows that 

the margin to rotor burst precludes the 

possibility of a rotor burst. 

(b) If the margin to burst shows that 

the stator case must have containment 

features in the event of failure, then the 

stator case must provide for the 

containment of the failed rotating 

components. The applicant must define 

by test, validated analysis, or a 

combination thereof, and document and 

provide to the installer as part of the 

requirements in § 33.5, the energy level, 

trajectory, and size of fragments released 

from damage caused by the main-rotor 

failure, and that pass forward or aft of 

the surrounding stator case. 

(32) General Conduct of Tests 

(a) Maintenance of the engine may be 

made during the tests in accordance 

with the service and maintenance 

instructions submitted in compliance 

with § 33.4. 

(b) The applicant must subject the 

engine or its parts to any additional tests 

that the Administrator finds necessary 

if— 

(1) The frequency of engine service is 

excessive; 

(2) The number of stops due to engine 

malfunction is excessive; 

(3) Major engine repairs are needed; 

or 


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(4) Replacement of an engine part is 

found necessary during the tests, or due 

to the teardown inspection findings. 

(c) Upon completion of all 

demonstrations and testing specified in 

these special conditions, the engine and 

its components must be— 

(1) Within serviceable limits; 

(2) Safe for continued operation; and 

(3) Capable of operating at declared 

ratings while remaining within limits. 

(33) Engine Electrical Systems 

(a) Applicability. Any system or 

device that provides, uses, conditions, 

or distributes electrical power, and is 

part of the engine type design, must 

provide for the continued airworthiness 

of the engine, and must maintain 

electric engine ratings. 

(b) Electrical systems. The electrical 

system must ensure the safe generation 

and transmission of power, and 

electrical load shedding if required, and 

that the engine does not experience any 

unacceptable operating characteristics 

or exceed its operating limits. 

(c) Electrical power distribution. 

(1) The engine electrical power 

distribution system must be designed to 

provide the safe transfer of electrical 

energy throughout the powerplant. The 

system must be designed to provide 

electrical power so that the loss, 

malfunction, or interruption of the 

electrical power source will not result in 

a hazardous engine effect, as defined in 

special condition no. 17(d)(2) of these 

special conditions. 

(2) The system must be designed and 

maintained to withstand normal and 

abnormal conditions during all ground 

and flight operations. 

(3) The system must provide 

mechanical or automatic means of 

isolating a faulted electrical energy 

generation or storage device from 

leading to hazardous engine effects, as 

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

of these special conditions, or 

detrimental effects in the intended 

aircraft application. 

(d) Protection systems. The engine 

electrical system must be designed such 

that the loss, malfunction, interruption 

of the electrical power source, or power 

conditions that exceed design limits, 

will not result in a hazardous engine 

effect, as defined in special condition 

no. 17(d)(2) of these special conditions. 

(e) Electrical power characteristics. 

The applicant must identify, declare, 

document, and provide to the installer 

as part of the requirements in § 33.5, the 

characteristics of any electrical power 

supplied from— 

(1) the aircraft to the engine electrical 

system, for starting and operating the 

engine, including transient and steadystate voltage limits, and 

(2) the engine to the aircraft via 

energy regeneration, and any other 

characteristics necessary for safe 

operation of the engine. 

(f) Environmental limits. 

Environmental limits that cannot 

adequately be substantiated by 

endurance demonstration, validated 

analysis, or a combination thereof must 

be demonstrated by the system and 

component tests in special condition no. 

27 of these special conditions. 

(g) Electrical system failures. The 

engine electrical system must— 

(1) Have a maximum rate of loss of 

power control (LOPC) that is suitable for 

the intended aircraft application; 

(2) When in the full-up configuration, 

be single-fault tolerant, as determined 

by the Administrator, for electrical, 

electrically detectable, and electronic 

failures involving LOPC events; 

(3) Not have any single failure that 

results in hazardous engine effects; and 

(4) Ensure failures or malfunctions 

that lead to local events in the intended 

aircraft application do not result in 

hazardous engine effects, as defined in 

special condition no. 17(d)(2) of these 

special conditions, due to electrical 

system failures or malfunctions. 

(h) System safety assessment. The 

applicant must perform a system safety 

assessment. This assessment must 

identify faults or failures that affect 

normal operation, together with the 

predicted frequency of occurrence of 

these faults or failures. The intended 

aircraft application must be taken into 

account to assure the assessment of the 

engine system safety is valid. The rates 

of hazardous and major faults must be 

declared, documented, and provided to 

the installer as part of the requirements 

in § 33.5. 

Issued in Kansas City, Missouri, on 

December 10, 2024. 

Patrick R. Mullen, 

Manager, Technical Policy Branch, Policy and 

Standards Division, Aircraft Certification 

Service. 

[FR Doc. 2024–29490 Filed 12–16–24; 8:45 am] 

BILLING CODE 4910–13–P 

DEPARTMENT OF TRANSPORTATION 

Federal Aviation Administration 

14 CFR Part 129 

[Docket No.: FAA–2024–0176; Amdt. No. 

129–55] 

RIN 2120–AL93 

Foreign Air Operator Certificates 

Issued by a Regional Safety Oversight 

Organization 

AGENCY: Federal Aviation 

Administration (FAA), Department of 

Transportation (DOT). 

ACTION: Final rule. 

SUMMARY: This amendment will allow 

the FAA to review and, if acceptable to 

the Administrator, recognize as valid air 

operator certificates issued by a 

Regional Safety Oversight Organization 

to foreign air carriers when the State of 

the Operator is a member of that 

Regional Safety Oversight Organization, 

for purposes of evaluating foreign 

applicants for operating specifications. 

DATES: Effective January 16, 2025. 

ADDRESSES: For information on where to 

obtain copies of rulemaking documents 

and other information related to this 

final rule, see ‘‘Additional Information’’ 

in the SUPPLEMENTARY INFORMATION

section of this document. 

FOR FURTHER INFORMATION CONTACT: Tim 

Shaver, International Program Division/ 

International Operations Branch, 

Federal Aviation Administration, 800 

Independence Avenue SW, Washington, 

DC, 20591; telephone (202) 267–1704; 

email tim.shaver@faa.gov. 

SUPPLEMENTARY INFORMATION: 

I. Authority for This Rulemaking 

The FAA’s authority to issue rules on 

aviation safety is found in title 49 of the 

United States Code. Subtitle I, section 

106, describes the authority of the FAA 

Administrator. Subtitle VII, Aviation 

Programs, describes in more detail the 

scope of the FAA’s authority. 

This rulemaking is issued under the 

authority described in subtitle VII, part 

A, subpart III, section 44701(a)(5). 

Under that section, the FAA is charged 

with promoting safe flight of civil 

aircraft in air commerce by prescribing 

regulations and minimum standards for 

practices, methods, and procedures the 

Administrator finds necessary to ensure 

safety in air commerce. This regulation 

is within the scope of that authority. 

Amending the regulations for 

applications for operations 

specifications under part 129 submitted 

by foreign air carriers or foreign persons, 

and the related standards for denial of 


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