M. David Kankam

Glenn Research Center, Cleveland, Ohio

Malik E. Elbuluk

University of Akron, Akron, Ohio

A Survey of Power Electronics Applications

in Aerospace Technologies

NASA/TM—2001-211298

November 2001

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NASA/TM—2001-211298

November 2001

National Aeronautics and

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Prepared for the

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cosponsored by the ASME, IEEE, AIChE, ANS, SAE, and AIAA

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IECEC2001–AT–32

M. David Kankam

Glenn Research Center, Cleveland, Ohio

Malik E. Elbuluk

University of Akron, Akron, Ohio

A Survey of Power Electronics Applications

in Aerospace Technologies
Acknowledgments

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constructive comments. We acknowledge Marie DiNovo for her assistance in processing the diagrams.

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NASA/TM2001-211298 1 

A SURVEY OF POWER ELECTRONICS APPLICATIONS IN 

AEROSPACE TECHNOLOGIES 

M. David Kankam 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

Malik E. Elbuluk 

University of Akron 

Akron, Ohio 44325 

ABSTRACT 

 

The insertion of power electronics in aerospace 

technologies is becoming widespread. The application 

of semiconductor devices and electronic converters, as 

summarized in this paper, includes the International 

Space Station, satellite power system, and motor 

drives in ‘more electric’ technology applied to aircraft, 

starter/generators and reusable launch vehicles. 

Flywheels, servo systems embodying 

electromechanical actuation, and spacecraft on-board 

electric propulsion are discussed. Continued inroad by 

power electronics depends on resolving 

incompatibility of using variable frequency for 400Hzoperated aircraft equipment. Dual-use electronic 

modules should reduce system development cost. 

I. THE POWER ELECTRONIC SYSTEM [1–2] 

A power electronic system can comprise a 

modular power electronic subsystem (PESS) 

connected to a source and load at its input and output 

power ports, respectively. The third port of PESS is 

connected to the system control, as shown in Fig. 1. 

PESS has been described as a power device module, 

intelligent power module, smart power device, power 

control module and, more recently, power electronic 

building block (PEBB) of the Office of Naval 

Research-initiated program [1–2]. 

The next three sections discuss the commonly 

used power semiconductor devices and their resident 

electronic converters, and electric motor drive 

technologies which are enablers in the ‘more electric’ 

technology (MET) for aero-space vehicle upgrades. 

I–1. Power Semiconductor Devices

The semiconductor devices metal-oxidesemiconductor field effect transistor (MOSFET), 

insulated gate bipolar transistor (IGBT), moscontrolled thyristor (MCT) and gate-turn-off thyristor 

(GTO) represent the cornerstone of modern power 

electronic converters. They are fully controllable. 

Their applications depend on their power and 

frequency characteristics. 

Figure 1. Three-Port Power Electronic System 

Structure. 

The unipolar, voltage-controlled MOSFET is 

relatively very fast, requires only minimal snubbing 

due to its low switching losses, and is protected from 

breakdown by its inherent positive temperature 

coefficient of resistance. Its drain resistance increases 

with temperature. The device is used in applications of 

few MHz and few watts to few kilowatts, such as in 

voltage-source pulse-width-modulated inverter 

(PWMI) and a zero-voltage switching (ZVS) converter 

rated at 0.1 to 2kVA and 0.5 to 2MHz. 

The IGBT, a hybrid of a MOSFET and a turn/off 

bipolar junction transistor (BJT), is a MOS-gated 

device. When used in 0 to 300kVA, 1.5 to 5kHz 

voltage-source PWMI and 20 to 150kVA, 40kHz 

voltage-source converter drives, the IGBT requires 

minimum cooling, is more reliable, experiences 

reduced voltage spikes at turn-off, and exhibits 

improved thermal life. 

The GTO is turned-on/off by gate current pulses. 

Turn-off-induced spike in the anode voltage can cause 

hot spots in, and breakdown of the device. A snubber 

can protect the GTO the high switching losses of 
NASA/TM2001-211298 2 

which limit its use to 1.0 to 2.0kHz PWM. Example 

application areas are 0.75 to 1.5kVA, 500 to 600Hz 

voltage-source PWMI and 2 to 12MVA, 700 to 

1200Hz ZVS inverter. 

The MCT is a MOS-gated device with a high 

turn-off current gain. With switching speed 

comparable to that of IGBT, the MCT shows promise 

for application in snubbered voltage-source inverters 

at 0 to 100kVA and 2kHz, zero-current switching 

(ZCS) at 150kVA and 40kHz, and ZVS converter at 

20 to 150kVA and 25kHz. The above devices are 

constituent parts of electronic converters which 

permeate aerospace power systems. 

I–2. Power Electronic Converters [3–4] 

Aerospace power systems have a considerable real 

estate of DC power usage. Over the past decade, AC 

power has emerged as a ‘driver’ for developing 

MET’s. This has increased the use of power electronic 

converters to condition and control power in the 

related systems. The high frequency converters of 

interest and their devices are noted in Fig. 2 [3]. 

Figure 2.Options for ‘More Electric’ Technology 

(MET).

The semiconductor devices in a hard-switching 

PWMI are stressed during turn-on/off modes. 

Capacitive and inductive stray effects compound the 

resulting switching losses. The PWMI AC voltage 

output is harmonic-rich, variable-frequency (VF), and 

commercially used in motor drives. The PWMI is 

characterized by limited power range and low 

switching frequencies, with attendant acoustic noise 

and reduced efficiency. However, it has low VoltAmpere rating, and is relatively easy to control. 

The snubbered converter employs switching 

similar to that of the PWMI. It incorporates a series 

inductive snubber to limit the inrush-current though its 

devices. A parallel capacitive snubber limits the device 

voltage, and reduces device stress. High switching 

losses are dissipated in the snubber. 

The zero-current switching (ZCS) converter 

uses an inductive snubber for device turn-off without 

current flow. Similarly, the zero-voltage switching 

(ZVS) converter employs a capacitive snubber for 

device turn-on, with an anti-parallel diode conducting. 

The converters have high switching frequency, lossless 

devices, high efficiency and reliability. 

The Resonant DC-Link (RDCL) in Fig. 3 and

Resonant AC Link converters overcome the 

limitations of PWMI. The RDCL converter has low 

switching losses, heat dissipation and acoustic noise, 

higher operating frequency and reliability, and reduced 

dv/dt and di/dt, resulting in low electromagnetic 

interference. Eliminating the Resonant Tank from 

Fig. 3 yields the conventional hard-switched, voltagesource PWMI [3]. 

Figure 3.The Conventional Resonant DC Link 

Inverter.

DC-DC converters have been used in aerospace 

power systems to provide the required voltage for the 

secondary distribution network. The basic topologies 

are the step-down (buck) and step-up (boost) 

converters for various load requirements [4]. For a 

given input voltage, the average output voltage is 

obtained by switching the electronic devices at 

constant frequency, while adjusting the device ‘ON’

duration. 

Electronic converters constitute the heart of motor 

drives which are essential for electromechanical 

actuation (EMA) in MET. 

I–3. Motor Drive Technologies [5–12] 

 The electric motor is the workhorse in a drive 

system. Drive characteristics depend on the motor, the 

power circuit, electronic devices and the controller. 

Power rating, operating speed range, environment, 
NASA/TM2001-211298 3 

fault tolerance, reliability, performance, thermal 

capability and cost affect motor selection for an 

application [5]. 

In the past, the excellent drive performance and 

low initial cost of DC machines made them the 

primary choice for servo applications. Their built-in 

commutators, high maintenance and spark-inducing 

brushes hinder DC machine use in drives. “Brushless”

motors have emerged from coupling DC, AC 

synchronous and induction motors with electronic 

controllers. The resulting maintenance- and spark-free 

brushless DC machine (BLDCM) or permanent 

magnet synchronous motor (PMSM), switched 

reluctance motor (SRM) and induction motor (IM) 

have higher torque/inertia ratio, peak torque 

capability, power density and reliability than the DC 

brush motor. 

The permanent magnet in the rotor of a BLDCM 

produces an armature current-independent field. The

commutatorless BLDCM has laterally stiff rotor which 

permits higher speed, especially in servo applications. 

The intra-stator placement of the rotor improves heat 

conduction which increases electric loading, and yields 

a higher torque/amp, better effective power factor and 

higher efficiency. BLDCM-based drives are popular, 

due to their performance and price improvements. 

Disadvantages include the need for shaft position 

sensing and more complex electronic controller. 

The IM has been the traditional workhorse for 

fixed and variable speed drive applications. It is 

rugged, relatively inexpensive and almost 

maintenance-free. The rotor slip-dependent torque 

production worsens the performance and decreases the 

efficiency of the motor. In the fractional and low 

integral hp range requiring dynamic performance, high 

efficiency and a wide speed range, the complexity of 

induction motor drive makes the BLDCM favorable. 

The gear-less, direct-application SRM drives are 

widely accepted. Their applicability to aircraft engine 

starter/generator has been demonstrated [6–7]. The 

motor design and converter topologies of the drive 

have undergone significant research and development 

in over two decades [8–12]. The series connection of 

the converter phase-leg switches to the motor phase 

winding prevents shoot-through fault by the converter 

switches. The motor is economical, compact in 

construction, and has high torque-to-inertia ratio, high 

torque output at low-to-moderate speeds, and faster 

response in servo systems. Its drawbacks are higher 

torque ripple and acoustic noise, complex control, and 

the need for an absolute rotor position sensor for the 

controller to establish the phase current pulses. 

II. AEROSPACE POWER SYSTEMS [2] 

The advent of MET for aerospace systems has focused 

attention on AC-, and hybrid DC- and AC-based 

power management and distribution (PMAD) 

systems. The schematics in Figs. 4 and 5 show a 

commonality in the use of electronic converters, 

photovoltaic (PV) solar arrays and batteries [2]. In 

aerospace systems, PEBB-related integration issues 

are the level of power and frequency range, 

application- and mission-dependent extreme 

temperature range, weight and size, electromagnetic 

interference and performance. Resolution of these 

issues is expected to promote expeditious insertion of 

electronic modules in aerospace technologies. 

III. POWER ELECTRONICS APPLICATION IN 

AEROSPACE TECHNOLOGIES 

The International Space Station (ISS), satellite 

power systems, MET, starter/generator system, 

reusable launch vehicles, flywheel technology and onboard electric propulsion are discussed to highlight the 

important role of power electronics in these systems. 

Figure 4.DC-PMAD-Based Electric Power System. 

Figure 5.AC-PMAD-Based Electric Power System. 

III–1. Space Station Power System [13–14] 

A single channel diagram, Fig. 6, of the ISS electric 

power system (EPS) shows a DC network of PV solar 
NASA/TM2001-211298 4 

arrays, batteries, power converters, switches and user 

loads [13]. The networks of 120V American and 28V 

Russian can exchange bi-directional power flow via 

American-to-Russian Converter (ARCU) and 

Russian-to-American Converter (RACU) units. The 

primary distribution system (PDS) comprises the PV 

arrays, batteries and the network up to the DC-DC

converter units (DDCU’s), for 160V to 120V stepdown to the secondary distribution system (SDS). The 

sequential shunt unit (SSU) regulates the voltage 

output of the PV array. The DDCU’s isolate the PDS 

and SDS from each other, and condition the source 

power for the SDS. The batteries store energy during 

insolation periods, and supply load power during 

orbital eclipse. The battery charge/discharge units 

(BCDU’s) isolate the battery from the primary bus. 

Remote power controller modules (RPCM’s), or 

switchgears, distribute power to the load converters. 

The BCDU’s, DDCU’s and the load converter contain 

semiconductor switches in their circuitry, thus 

underscoring the importance of power electronics in 

the ISS EPS. 

Figure 6.Single Channel Diagram of ISS Power 

System. 

In their modular form, power electronics are 

expected to facilitate the control of PMAD and 

diagnosis of system malfunctions, to yield reliability 

improvement [14]. Some level of electronic 

modularity has already been built into some satellite 

EPS, to permit adaptability of the EPS to various and 

future programs, with minimal re-design. 

III–2. Satellite Power Systems [15] 

A partial block diagram of a satellite modular EPS 

is shown in Fig. 7. It depicts only the north portion of 

a ‘Dual Bus’ power system for a geosynchronous 

satellite [15]. The primary-side elements of the EPS 

are the PV arrays, battery and power control unit 

(PCU). On the secondary side, the PCU embodies the 

SSU, battery charge and discharge converter modules 

(BCCM, BDCM), and low voltage converter module 

(LVCM) of redundant, main bus-connected DDCUs 

which feed the spacecraft loads via the power 

distribution unit (PDU). The operation of the satellite 

Figure 7.Block Diagram of Satellite Power System. 

EPS is similar to that of the ISS, regarding sunlight 

and eclipse portions of a mission. The built-in 

modularity makes it possible to vary battery voltage 

and output power levels, by adding and removing 

converter module(s). This feature renders the EPS 

configurable for future missions. Also, the modularity 

facilitates power electronics packaging, equipment 

deployment into space, and needed on-orbit EPS 

modifications. 

The use of power electronics-based motion 

control systems in selected aerospace systems is 

discussed next. 

III–3. Motor Drive Applications [5, 14, 16–33] 

 The key elements in electric actuation (EA) for 

MET are the electric motor, its power electronics, the 

control system and the actuator load(s) [5]. Lower 

costs and advances in power electronics and highspeed electric machines have fuelled the interest of 

technologists, developers and researchers in industry 

[16–19], Government Agencies [20–23], and academia 

[24–27], in aerospace motion control systems. A key 

premise of the MET is to replace the traditionally 

mounted auxiliary drives and bleed air extraction with 

integral engine starter/generators (S/G’s), electricaldriven actuators and engine-gearbox-driven fuel 

pumps. The replacement eliminates hydraulic, 

pneumatic and mechanical power, and minimizes 

and/or eliminates their associated costs, as well as high 

pre-flight operation, maintenance and refurbishment of 

hydrazine-driven auxiliary power units (APU’s). 
NASA/TM2001-211298 5 

III–3.1. More Electric Technology for Aircraft 

Figure 8 shows a conceptual diagram of the Air 

Force’s ‘more electric aircraft’ (MEA) subsystems 

[20]. The hydraulic-driven flight control actuators, the 

engine-gearbox driven fuel pump and air-driven 

environmental control system (ECS) are electrically 

powered by electric motor drives. A S/G supplies 

electric power to a fault-tolerant PMAD subsystem 

which feeds power to the EA, engine starting, braking, 

ECS, fuel pump and anti-icing. Uninterrupted power 

from an integrated APU and battery system provides 

redundancy and engine start-up. 

Figure 8.Concept of More Electric Aircraft.

The candidate electric motor drives for the MET 

are the IM, BLDCM and SRM drives. For each 

selected drive in the MET, depending on the 

application, the EA and its electronic controller must 

suitably match the safety and reliability of hydraulic 

actuation. This may require motor drive redundancy, 

for assured flight and landing. Thus, the built-in 

redundancy in its independent motor windings due to 

magnetic isolation, and in power switching circuits by 

electrical isolation, makes the SRM an attractive 

choice for fault-tolerant EA. This ‘limp home’ ability 

of the SRM has been one of the key factors in its 

selection by the Air Force in their MEA development 

[21–22]. 

III–3.2. Starter/Generators 

 A general variable speed constant frequency 

(VSCF) S/G system is shown in Fig. 9 [23]. The 

machine may be any of the three candidates in Section 

III–3. During motoring, constant frequency (CF) 

electrical power from the main AC bus is converted to 

VF by the bi-directional power converter, and fed to 

the machine to start the load such as an aircraft engine. 

In the generating mode, the variable speed load 

provides mechanical power to run the machine the 

variable frequency of which is converted to a constant 

frequency for the main bus. The control system 

Figure 9.Typical VSCF Starter/Generator System. 

receives inputs from the VSCF sources, and provides 

gating signals for the converter to maintain proper 

interface between VF and CF requirements. 

 

III–3.3. Reusable Launch Vehicles 

EMA has been under consideration for replacing 

hydraulic systems used on reusable launch vehicles for 

thrust vector control (TVC) gimbaling of engines and 

aerosurface control [28–29]. The projected benefits are 

as stated in Section III–3. In the early 90’s, the then 

NASA Lewis Research Center (currently GRC) and 

General Dynamics Space Systems Division cooperated 

on demonstrating EMA technology readiness to meet 

the hydraulic TVC requirements for the Atlas 

Expendable Launch Vehicle [28]. Concurrently, a 

study by the NASA Kennedy Space Center indicated 

that the MET would save nearly 66 percent of manhours needed for hydraulic TVC processing of the 

Shuttle Solid Rocket Booster [30]. The EMA in 

Ref. [28] embodied IM drive with a field-oriented 

control [5, 31–32] for independent control of torque 

and flux, and a Pulse Population Modulation technique 

[28] for independent control of voltage and frequency. 

A more advanced motor drive is in use for ongoing Government-Industry development of flywheel 

technology. 

III–3.4. Flywheel Technology 

United States Government Agencies, Industry and 

academia are jointly developing advanced flywheel 

technologies to provide high performance and 

reliability, and reduced losses in a high-speed, light 

weight flywheel energy storage system (FESS), peak 

power and load leveling in spacecraft and aircraft 

applications [14]. The FESS embodies advanced 

composite materials-based rotor, low-loss magnetic 

bearings, high-speed motor-generator set and 

electronic converter drive. The above efforts and 

concurrent component technology developments 

represent an advancement of prior work by the 

collaborators [33]. NASA GRC has been leading an 

effort for a combined flywheel energy storage and 
NASA/TM2001-211298 6 

attitude control system, namely, integrated power and 

control system, to enable the development of a low 

cost, lightweight and higher specific energy spacecraft. 

Additionally to conditioning power and enabling 

bi-directional power and stored energy flow in 

flywheel systems, motor drives feature in electric 

upgrade of aircraft pumps, for the same MET benefits. 

III–3.5. Servo System Applications 

The MET proposes the use of VF motor drives to 

operate hydraulic and fuel pumps on aircraft. Using 

high-density motor drives can eliminate the usual size 

and weight limitations of drives. However, issues of 

VF incompatibility with 400Hz-operated aircraft 

equipment such as fuel and hydraulic pumps [19], 

attendant increase in motor weight to achieve the 

required torque at high frequencies, and potentially 

high upgrade cost must be resolved. 

 By comparison with CF power, VF motor 

controllers can reduce transient inrush current at motor 

start. Furthermore, a variable speed motor-driven fuel 

pump can provide only the required amount of fuel. 

Also, such a fuel pump can improve aircraft 

performance by reducing engine gearbox weight and 

enabling direct integration with the aircraft electronic 

propulsion and flight control [16]. 

Additionally to their use in aerospace power 

systems, electronic converters play an important 

function in the on-board electric propulsion of 

spacecraft. 

III–4. On-Board Electric Propulsion [34–37] 

 Power electronics are constituent parts of the 

power processing unit (PPU) of spacecraft electric 

propulsion which is credited with reducing launch 

vehicle requirements, notably for north-south station 

keeping of commercial geosynchronous satellites [34]. 

A PPU comprises one or more electronic converters. It 

provides electric power for the spacecraft thruster, and 

commands and telemetry interface to the electric 

propulsion system, as shown in Fig. 10 [35]. The 

converters may be current-controlled and voltage-fed, 

to rapidly supply constant current to offset thruster 

voltage variations, typically during a start-up period. 

 Small-sized PPUs with a minimal number of 

lightweight, highly efficient, soft-switching converters 

yield increased payload and power [36]. Such PPUs 

can generate high voltage start pulse to ignite as many 

as four arcjet thrusters for north/south station keeping 

orbit maneuvers [37], thus reducing propulsion system 

mass. 

Figure 10.Electric Propulsion System.

 Several challenges must be overcome for 

continued penetration of power electronics into 

aerospace systems. 

IV. FUTURE TRENDS [14,19,24,38] 

 VF is currently used in turbo-prop and business 

jets. The cost and savings attractions of VF are 

tempered by the potential high cost of VF-upgrade for 

conventional, 400Hz-operating equipment on aircraft. 

Reference [19] points out judicious use of power 

electronics, via a hybrid hydraulic/pneumatic/motor 

drive design, to circumvent the VF–400Hz aircraft 

equipment incompatibility. 

 Continued improvements in power electronic 

devices and their switching schemes, advances in 

magnetic materials and capacitors, and better design of 

motors and electronic controls are expected to 

ameliorate weight, size and reliability issues of MET 

application to aerospace systems. 

 The need for bi-directional converters for battery 

charge/discharge functions and fixed frequency power 

and voltage, and expected varying requirements of 

multiple loads in aerospace systems suggest future use 

of hybrid AC and DC multi-converters with multivoltage levels [24]. 

 Increasing use of power electronic modules will 

require consideration of device ratings, bidirectionality or otherwise of power flow, power 

density requirements and degree of integration, when 

developing aerospace systems. Hardware commonality 

will promote dual-use application of the modules, and 

decrease system development cost [14]. For instance, 

NASA-planned development of 2 to 3kW power 

processor/thruster is expected to provide modular 

elements for various mission requirements [38]. 

V. CONCLUSIONS 

This paper presents a survey of power electronics 

applications in aerospace technologies. It encompasses 
NASA/TM2001-211298 7 

the International Space Station, satellite and aircraft 

power systems, flywheel technology, spacecraft onboard propulsion, and the ‘more electric’ technology 

(MET) insertion in spacecraft, aircraft and launch 

vehicles. 

 Power electronic converters are central to the 

performance of aerospace power systems and 

spacecraft on-board electric propulsion. Resolution of 

incompatibility between conventional, 400Hz 

operating equipment and the variable frequency of 

MET should promote increased penetration of power 

electronics into aerospace systems. Future multivoltage needs and varied load requirements will 

necessitate the use of multi-voltage level converters. 

The use of electronic modules with dual-use options 

and hardware commonality for aircraft and spacecraft 

should reduce development cost and maximize system 

re-use, while improving system reliability and 

performance. 

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No. 3, May/June 1990, pp. 434–440. 
NASA/TM2001-211298 8 

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This publication is available from the NASA Center for AeroSpace Information, 301–621–0390.

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NASA TM—2001-211298

IECEC2001–AT–32

E–13102

WU–910–30–11–00

Aerospace power systems; Electromechanical actuation; Flywheels; International space station; 14

Modularity; More electric aircraft; Motor drives; Power converters; Electronics; Reusable launch

vehicles; Semiconductor devices; Servomotor; Spacecraft propulsion; Starter/generator

Unclassified -Unlimited

Subject Categories: 15, 20, 33, and 44 Distribution: Nonstandard

The insertion of power electronics in aerospace technologies is becoming widespread. The application of semiconductor

devices and electronic converters, as summarized in this paper, includes the International Space Station, satellite power

system, and motor drives in ‘more electric’ technology applied to aircraft, starter/generators and reusable launch vehicles.

Flywheels, servo systems embodying electromechanical actuation, and spacecraft on-board electric propulsion are

discussed. Continued inroad by power electronics depends on resolving incompatibility of using variable frequency for

400 Hz-operated aircraft equipment. Dual-use electronic modules should reduce system development cost.

Prepared for the 36th Intersociety Energy Conversion Engineering Conference cosponsored by the ASME, IEEE, AIChE,

ANS, SAE, and AIAA, Savannah, Georgia, July 29–August 2, 2001. M. David Kankam, NASA Glenn Research Center,

and Malik E. Elbuluk, Department of Electrical Engineering, University of Akron, Akron, Ohio 44325. Responsible

person, M. David Kankam, organization code 5450, 216–433–6143.

M. David Kankam and Malik E. Elbuluk

A Survey of Power Electronics Applications in Aerospace Technologies