Luis R. Piñero, Robert J. Scheidegger, and Michael V. Aulisio

Glenn Research Center, Cleveland, Ohio

Arthur G. Birchenough

Vantage Partners LLC, Brook Park, Ohio 

High Input Voltage Discharge Supply for High 

Power Hall Thrusters Using Silicon Carbide Devices

NASA/TM—2014-216607

February 2014

IEPC–2013–388
NASA STI Program . . . in Profile

Since its founding, NASA has been dedicated to the 

advancement of aeronautics and space science. The 

NASA Scientific and Technical Information (STI) 

program plays a key part in helping NASA maintain 

this important role.

The NASA STI Program operates under the auspices 

of the Agency Chief Information Officer. It collects, 

organizes, provides for archiving, and disseminates 

NASA’s STI. The NASA STI program provides access 

to the NASA Aeronautics and Space Database and 

its public interface, the NASA Technical Reports 

Server, thus providing one of the largest collections 

of aeronautical and space science STI in the world. 

Results are published in both non-NASA channels 

and by NASA in the NASA STI Report Series, which 

includes the following report types:

• TECHNICAL PUBLICATION. Reports of 

completed research or a major significant phase 

of research that present the results of NASA 

programs and include extensive data or theoretical 

analysis. Includes compilations of significant 

scientific and technical data and information 

deemed to be of continuing reference value. 

NASA counterpart of peer-reviewed formal 

professional papers but has less stringent 

limitations on manuscript length and extent of 

graphic presentations.

• TECHNICAL MEMORANDUM. Scientific 

and technical findings that are preliminary or 

of specialized interest, e.g., quick release 

reports, working papers, and bibliographies that 

contain minimal annotation. Does not contain 

extensive analysis.

• CONTRACTOR REPORT. Scientific and 

technical findings by NASA-sponsored 

contractors and grantees.

• CONFERENCE PUBLICATION. Collected 

papers from scientific and technical 

conferences, symposia, seminars, or other 

meetings sponsored or cosponsored by NASA.

• SPECIAL PUBLICATION. Scientific, 

technical, or historical information from 

NASA programs, projects, and missions, often 

concerned with subjects having substantial 

public interest.

• TECHNICAL TRANSLATION. Englishlanguage translations of foreign scientific and 

technical material pertinent to NASA’s mission.

Specialized services also include creating custom 

thesauri, building customized databases, organizing 

and publishing research results.

For more information about the NASA STI 

program, see the following:

• Access the NASA STI program home page at 

http://www.sti.nasa.gov

• E-mail your question to help@sti.nasa.gov

• Fax your question to the NASA STI 

Information Desk at 443–757–5803

• Phone the NASA STI Information Desk at

443–757–5802

• Write to:

 STI Information Desk

 NASA Center for AeroSpace Information

 7115 Standard Drive

 Hanover, MD 21076–1320
Luis R. Piñero, Robert J. Scheidegger, and Michael V. Aulisio

Glenn Research Center, Cleveland, Ohio

Arthur G. Birchenough

Vantage Partners LLC, Brook Park, Ohio 

High Input Voltage Discharge Supply for High 

Power Hall Thrusters Using Silicon Carbide Devices

NASA/TM—2014-216607

February 2014

IEPC–2013–388

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135

Prepared for the

33rd International Electric Propulsion Conference (IEPC2013)

sponsored by the Electric Rocket Propulsion Society

Washington, D.C., October 6–10, 2013
Acknowledgments

The authors would like to acknowledge the dedication and hard work of a large group of GRC engineers, designers and technicians 

that have contributed to the success of this development including: Karin Bozak, Michael Capelety, Randy Clapper, Dragos Dinca, 

James Dolce, Justin Elchert, James Fleet, Michael Garrett, Marcelo Gonzalez, Gary Gorecki, Courtney Gras, Elmer Griebeler, 

Ahmad Hammoud, Kevin McCormick, Michael Pastel, Marcus Studmire, Deboshri Sadhukhan, Paul Trimarchi, and Brandon White.

Available from

NASA Center for Aerospace Information

7115 Standard Drive

Hanover, MD 21076–1320

National Technical Information Service

5301 Shawnee Road

Alexandria, VA 22312

Available electronically at http://www.sti.nasa.gov

Level of Review: This material has been technically reviewed by technical management. 
NASA/TM—2014-216607 1

High Input Voltage Discharge Supply for High Power 

Hall Thrusters Using Silicon Carbide Devices

Luis R. Piñero, Robert J. Scheidegger, and Michael V. Aulisio

National Aeronautics and Space Administration 

Glenn Research Center

Cleveland, Ohio 44135 

Arthur G. Birchenough 

Vantage Partners LLC

Brook Park, Ohio 44142

Abstract

A power processing unit for a 15 kW Hall thruster is under development at NASA Glenn Research 

Center. The unit produces up to 400 VDC with two parallel 7.5 kW discharge modules that operate from a 

300 VDC nominal input voltage. Silicon carbide MOSFETs and diodes were used in this design because 

they were the best choice to handle the high voltage stress while delivering high efficiency and low 

specific mass. Efficiencies in excess of 97 percent were demonstrated during integration testing with the 

NASA-300M 20 kW Hall thruster. Electromagnet, cathode keeper, and heater supplies were also 

developed and will be integrated with the discharge supply into a vacuum-rated brassboard power 

processing unit with full flight functionality. This design could be evolved into a flight unit for future 

missions that requires high power electric propulsion. 

Introduction

Recent advancements in solar cell and deployment mechanisms could lead to the development of highpower solar arrays (Ref. 1). Using this power for electric propulsion will result in a high performance 

spacecraft capable of satisfying the requirements of the most demanding missions. A team of engineers from

NASA Glenn Research Center (GRC) and the Jet Propulsion Laboratory (JPL) was chartered by NASA’s 

Space Technology Mission Directorate (STMD) Game Changing Development (GCD) Program to develop 

a high-power solar electric propulsion (SEP) system to revolutionize future missions that require moving 

cargo or humans beyond Earth’s orbit (Refs. 2 to 4). A critical part of an SEP system is the power 

processing unit (PPU) which processes power from solar arrays and converts it to the voltage and currents

required for thruster operation. PPUs require much attention as they are the most massive and complex part 

of an electric propulsion system. Previous efforts to develop high power PPUs for electric propulsion 

have been challenging because of the difficulties in controlling high currents, noise and conduction losses

(Refs. 5 to 7).

High voltage is advantageous for high power systems to reduce the current levels and minimize 

ohmic losses (Refs. 8 to 10). However, there is very limited availability of high-voltage, highperformance parts for space applications. Recent advances in solid-state power component technology 

have resulted in commercial availability of silicon carbide (SiC) transistors. These could enable the 

development of high power PPUs with higher input voltages by overcoming the performance limitations 

of state of the art silicon transistor technology. SiC has the potential of revolutionizing future PPU 

developments and enabling high-power SEP. 

GRC is leading the PPU development effort for the STMD/GCD Program. The first phase of this 

effort consisted of a comprehensive trade study of power semiconductors. This study led to the selection 

of SiC MOSFETs as the transistor for a high voltage, high power PPU because they had the highest 

potential to achieve high efficiency and low specific mass. The second phase of the effort focused on the 

development of a breadboard 15 kW discharge supply to demonstrate the capabilities and functionality of 
NASA/TM—2014-216607 2

SiC in a high power converter design. The third phase of the effort is currently developing an entire PPU 

that includes most electrical functionality required for flight and starts to address form and fit of a flight 

unit. The goal is to develop a brassboard PPU to technology readiness level (TRL) 5 with a 300 V 

nominal input, and capable of operating in vacuum at high efficiency. This PPU will be integrated with a 

thruster to demonstrate a high performance design that can evolve into a flight SEP system and scaled to 

higher power. 

The technical approach used in this development is to simplify the design wherever possible to avoid 

high recurring cost; use commonality at all levels including parts, circuits, assemblies, converters and 

modules; and finally, leverage as much experiences and lessons learned from other recent NASA 

development program like NSTAR, NEXT and HiVHAc (Refs. 10 to 13).

System Overview

A simplified block diagram of a typical Hall thruster system is shown in Figure 1. The most critical 

part of the PPU is the discharge supply because it processes 90 to 95 percent of the power into the thruster 

and regulates the high voltage for the main discharge to accelerate ions and produce thrust. The discharge 

supply also provides electrical isolation between the thruster and the solar arrays. The output filter in the 

discharge supply is critical for stable thruster operation because it provides current to the thruster to 

sustain discharge current oscillations. A PPU also provides power for the inner and outer electromagnets 

and the cathode keeper and heater. These are referred to as the auxiliary supplies. An input filter for the 

entire unit can be added to mitigate conducted emissions to the spacecraft power system and solar arrays.

Also included in the PPU are input and output telemetry circuits to accurately assess system performance 

and protection circuits to safe the system in case of a fault. Finally, the PPU must have an interface to 

communicate with the spacecraft to relay system telemetry and receive commands.

Specifications

The design specifications for the high power PPU are summarized in Table 1. The discharge supply is 

required to operate from an input voltage range of 270 to 330 V and generate a maximum of 15 kW for 

the thruster at a voltage range of 300 to 400 V. This supply must regulate both output voltage and current 

to control steady-state thruster operation and limit current during fault or start-up conditions. The 

Figure 1.—Hall thruster PPU block diagram. 
NASA/TM—2014-216607 3

TABLE 1.—PPU ELECTRICAL SPECIFICATIONS 

auxiliary supplies are required to operate from a 23 to 36 V input voltage range. This was selected so the 

thruster cathode can maintain operation during an eclipse by running the keeper supply from a standard 

28 V battery backed-up power system. The auxiliary supplies require a variety of outputs but they do not 

exceed 36 V and 10 A and are required to regulate output current. All supplies must have better than 

2 percent regulation and less than 5 percent output ripple. Finally, the cathode keeper must include a high 

voltage pulsed igniter to facilitate ignition. 

Semiconductor Trade Study

One of the biggest challenges in developing high power electronics with high input voltages for space 

is the availability of radiation-hardened, flight-qualified power semiconductors such as transistors and 

diodes. Silicon (Si) MOSFETs have very good performance only up to a few hundred volts. Beyond that, 

performance degrades quickly and options are extremely limited. In fiscal year 2011, a comprehensive 

trade study of power semiconductors was conducted at GRC to evaluate and assess available options. At 

the time, components using new semiconductor materials, like silicon carbide (SiC) and gallium nitride, 

were becoming commercially available so these were considered in addition to standard Si parts. A 

selection of components from several vendors were characterized, compared and modeled using test 

circuits that simulated the conditions they would experience in a typical converter. Performance 

parameters such as on-resistance, switching loss, gate drive requirements and switching times were 

evaluated. The results indicated that for high voltage applications, SiC MOSFETs provided superior 

electrical performance over traditional Si devices. Commercially available 1200 V SiC MOSFETs have 

very high voltage ratings and much lower on-resistance than Si devices even at lower voltage ratings.

They could facilitate building power converters with much lower specific mass. Another attractive feature 

is a very low gate capacitance that results in very high switching speeds for low switching losses.

However, there are a few issues with these parts. First, SiC MOSFET technology had not been evaluated 

for use in a radiation environment or used in a power converter application in space. Also, their gates 

require to be driven to more than 20 V for the devices to be completely on. None of the space rated, radhard gate drive chips available can operate at this voltage. 

The study also evaluated high-power Si and SiC rectifier diodes. Parts were primarily compared 

based on forward drop, reverse recovery characteristics and voltage and current ratings. A high-voltage 

SiC Schottky diode was the best choice for this application. The high voltage rating of 1200 V could 

enable building a discharge supply with minimum rectifier losses. Although their forward voltage drop 

and conduction losses were slightly higher than Si diodes, their lack of reverse recovery losses and 

“softer” turn-off resulted in better overall performance. 
NASA/TM—2014-216607 4

After selecting the best components for the application, the trade study evaluated various topologies 

and the optimum power level for individual converter modules. Three criteria drove the selection of the 

converter topology. First, topologies capable of processing high power so the PPU requires fewer 

discharge modules and have lower specific mass. Second, topologies with low voltage stress since a high 

input voltage was required for this design. Lastly, topologies with better transformer utilization were 

favored to minimize winding complexity and parasitic elements like leakage inductance that could be 

detrimental to converter performance. These led to the selection of the full-bridge converter topology.

Also, a single full-bridge rectifier was selected for the output stage because of the large margin between 

the SiC rectifier diodes and the 400 V output required for this design and to minimize conduction losses. 

After having selected components and topologies, the remaining question for the trade study was the 

power level for a single discharge module. Gate drive circuit capability, magnetics size and switching 

frequency influenced this part of the study. A module size of 7.5 kW was selected for this design to take 

full advantage of the SiC components. Also, a modest switching frequency of 30 kHz was chosen to 

minimize switching losses without the need of complex resonant switching circuits. Higher discharge 

power will be attained by paralleling modules. 

PPU Design

PPU development was initiated by design, fabrication and testing of proof-of-concept breadboards of 

the main components of the PPU including the discharge and auxiliary supplies; input filters; a master 

control circuit; and a system controller. These elements are discussed below.

Discharge Supply

In fiscal year 2012, the PPU team at GRC started the development of a breadboard discharge supply 

with a 300 V nominal input using SiC components. The topology selected for this design was a fullbridge utilizing SiC MOSFETs and single bridge rectifier with SiC Schottky diodes as recommended by 

the trade study. A simplified schematic of the topology is shown in Figure 2. Peak-current pulse-widthmodulation (PWM) was selected because it is simple to implement, protects against over-current faults 

and has a fast response to input and output voltage variations. An innovative gate drive circuit with 

transformer isolation and the ability to adjust the on and off gate voltages was developed so the switching 

performance of the SiC MOSFETs could be optimized. The main power transformer was designed on a 

ferrite core and using interleaving techniques on the primary and secondary windings to minimize leakage 

inductance. The final version of the transformer had less than 2.0 μH of leakage inductance which is 

small considering the size of the transformer.

Two 7.5 kW discharge modules were built and connected in parallel. A master control circuit 

processes the feedback to regulate both output voltage and current. It also provides clock signals and 

common controls. A photograph of the breadboard discharge supply is shown on Figure 3. 

Auxiliary Supply

The auxiliary supplies provide thruster power for the inner and outer electromagnets and cathode 

keeper and heater. These are required to process between 100 and 400 W from a nominal input of 28 V. 

Although somewhat oversized for this application, a full-bridge topology was also chosen for these 

supplies to maintain commonality with the discharge supply, good transformer utilization, and scalability 

to higher power. Scalability can be advantageous as the system development matures. Since the Hall 

thruster being developed for this effort is also in early stages of development, separate electromagnet 

supplies were included to provide flexibility and better control. However, once the thruster design is 

matured, both electromagnets could be powered from a common supply. Implementing this should be 

relatively easy with the full-bridge topology. A center-tapped full-wave rectifier was chosen for reduced 

conduction losses at the lower output voltages. A simplified schematic of the power supply is shown in 
NASA/TM—2014-216607 5

Figure 2.—Discharge module simplified schematic. 

Figure 3.—Breadboard discharge power supply. Figure 4.—Simplified schematic of auxiliary supply. 

Figure 4. A common PWM control circuit and printed circuit board as used in the discharge module was 

used for the auxiliary supplies. Similar gate-drive circuit as in the discharge module was also used with 

some changes due to lower gate voltage and power requirements for the 100 V Si MOSFETs used in this 

design. A higher switching frequency of 60 kHz was selected to allow reduction in transformer and filter 

masses. The main power transformer was also built on a ferrite core with interleaved windings to 

minimize leakage inductance. There are two versions of the power transformer with identical primary 

windings and only a minor difference in secondary turns. One version is for the electromagnet supplies 

and the other for the heater and keeper supplies as they require about twice the output voltage as the 

magnet supplies. The only other difference between the auxiliary supplies is the output inductor and 

pulsed igniter used on the cathode keeper supply. 

A

B

C

D

300 V

300 to

400 V

Module #1

Module #2

Master Controls #1

Controls #2

≤ 36 V

A

B

C

D

28 V
NASA/TM—2014-216607 6

Master Controller

The master controller includes several internal control functions for the discharge modules, the 

auxiliary supplies, and the overall PPU. It generates 30 and 60 kHz clock signals for PWM used by the

power converters. The discharge supply output voltage and current feedback circuitry also resides in the 

master controller. It also contains various fault protection circuits, interlocks and fault flags for thruster 

anomalies. All controls, output setpoints, and telemetry signals are generated or routed through the master 

controller on the way to the system control board (SCB) described below. 

Input Filter

The PPU has two independent inputs. The 300 V input provides power for the discharge supply and 

the 28 V input powers the auxiliary and housekeeping supplies. Each one has its own input filter 

consisting of a differential low-pass stage and a common-mode inductor. The filters are damped to 

mitigate overshoots during transient conditions. The filters have cut-off frequencies around 5 kHz so they 

provide enough attenuation at the switching frequency. The 28 V input filter PCB also includes a 

commercial off the shelf power supply for housekeeping power to provide bias voltages for internal 

circuits. Input voltage and current signals are generated to be included with telemetry. 

System Control Board 

The SCB controls the PPU and the propulsion system. The SCB communicates with a user computer 

through a transformer isolated Mil-Std-1553B interface. It provides users control over power supply and 

feed system outputs and displays telemetry and fault flags from the PPU master controller and propellant 

feed system. The SCB uses a field programmable gate array (FPGA) and additional digital circuitry to 

generate control signals in response to user commands and to “translate” telemetry for the operator. The 

SCB is being developed by JPL. 

Test Results

Both the discharge module and auxiliary supplies were tested to verify that the hardware meets design 

specifications including input and output range, loop response, efficiency, regulation and overcurrent 

protection. Testing was conducted using a resistive load and digital multi-meters to measure input and 

output voltages and currents. 

Breadboard Discharge Module Performance

A 7.5 kW discharge module was tested to validate electrical performance. The module is capable of 

delivering full power at output voltages from 300 to 400 V for input voltages from 270 to 330 V. It has 

the capability of operating up to 500 V at nominal input or at lower output voltages at reduced power. 

Figure 5a illustrates the efficiency of the discharge module as a function of output power for a variety of 

output voltages at a nominal 300 V input. Figure 5b illustrates efficiency as a function of output power for 

a 300 V output for the entire range of input voltages. Efficiency at full power conditions was between 96 

and 97 percent and was as high as 97.5 percent at 4 kW output power. Line and load regulation is better 

than 1.0 percent for the entire range of inputs and outputs. The output voltage ripple was lower than 

1 percent at full power conditions while running into a resistive load. 

The control loop of the discharge module was compensated to ensure stability while operating a Hall 

thruster. Figure 6 shows the gain and phase plot for the discharge module voltage loop. The bandwidth of

approximately 1 kHz, gain margin is approximately 12 dB, and a phase margin is about 90°, provide 

adequate stability margin and response for this application. 
NASA/TM—2014-216607 7

a) 300 V input b) 300 V output

Figure 5.—Breadboard discharge module efficiency. 

Figure 6.—Discharge supply voltage loop response. 

Breadboard Discharge Supply/Thruster Integration

The 15 kW breadboard discharge supply was tested with the NASA-300M 20 kW Hall thruster in 

VF-5 at NASA GRC. The purpose of the test was to verify stable operation throughout the entire range of 

operating conditions. One of the most important issues to validate is the ability of the power supply to 

drive the large discharge oscillations that can occur during thruster operation. If the discharge output filter 

is not properly designed, it could result in large voltage oscillations and unstable operation of the 

discharge supply. 

During the integration test, the thruster was operated from 3 to 15 kW at 300 and 400 V output from a 

nominal 300 V input. Efficiencies in excess of 97 percent were measured. The two discharge modules 

shared current within 0.1 A of each other at all operating conditions. Measured current oscillations had a 

frequency of approximately 40 to 50 kHz and a peak-to-peak magnitude of 5 to 10 A depending on 

thruster operating conditions. The worst-case voltage oscillations had a peak-to-peak magnitude of less 

than 5 V at full power. Figure 7 illustrates these oscillations at full power conditions. 0.94 0.96 0.98

0 1 2 3 4 5 6 7 8 9

Efficiency

Output Power, kW

200V

Output

0.94 0.96 0.98

0 1 2 3 4 5 6 7 8 9

Efficiency

Output Power, kW

330V Input

300V Input

270V Input

0.01 0.1 1 10

Frequency, kHz

60

40

20

0

-20

-40

-60

-80

-100

100

80

Phase, degrees

40

20

0

-20

-40

-60

-80

-100

-120

-140

-160 Gain, dB

Phase 

Margin

~90 ̊

Gain

Margin

~12 dB
NASA/TM—2014-216607 8

Figure 7.—Discharge oscillations with NASA 300-M thruster. 

Breadboard Auxiliary Supply Performance 

An auxiliary supply was tested to validate electrical performance. It was tested in the electromagnet 

configuration because the magnet supplies have to operate continuously during steady-state thruster 

operation. The supply is capable of delivering up to 200 W of power at an output current of 10 A for an 

input voltage range of 23 to 36 V. Efficiencies in excess of 92 percent were measured when operating at 

full power. Line and load regulation was better than 1 percent and output current ripple was lower than 

3 percent for the entire operating range.

Brassboard Development

Brassboard PPU Design

After demonstrating functionality with the breadboard power supplies, the effort focused on design 

and fabrication of a brassboard version of the PPU. The goal was to develop a unit capable of operating in 

vacuum while maintaining the high efficiency demonstrated with the breadboard power supplies. This 

would require to address the thermal design of the unit and provide an iteration on the electrical design to 

include additional circuitry for telemetry and fault protection. This design could then be transferred to a 

contractor to evolve it into a flight design with full form, fit, and function. 

The brassboard PPU consists of interlocking modules that supply a variety of functions. The height 

and depth of the modules were fixed and the width was adjusted to accommodate various modules. 

Figure 8 illustrates a drawing of the PPU that consists of two discharge modules, an auxiliary module, an 

input filter module, and a system control module. 

The PPU was designed to reject waste heat through its baseplate. The thermal design approach 

consists of mounting large power components like MOSFETs, rectifier diodes, transformers and inductors 

on heat sinks that mount directly onto the baseplate. This maintains the parts with the most power 

dissipation at the lowest possible temperature within the unit. PCBs rely on internal thermal planes to 

reject waste heat from components and into a heat sink also attached to the baseplate. Thermal models 

were developed to validate the thermal design approach and verify that critical components like magnetics 

and the SiC semiconductors would operate below their temperature limits. Figure 9 shows the 

implementation of this design approach for the discharge module. The module consists of subassemblies 

for the discharge full-bridge inverter, power transformer, output rectifier and output inductor. 

Anode Voltage

5 V/div(ac)

Anode Current

5 A/div(ac)

20 μs/div
NASA/TM—2014-216607 9

Figure 8.—Brassboard PPU. 

Figure 9.—Brassboard discharge module. 

Brassboard DDU Design

In a direct-drive system, electrical power from a power source, like solar arrays, is fed directly into 

the discharge in a Hall thruster. This eliminates the need for a discharge supply making the system 

potentially more efficient and lightweight. The direct-drive unit (DDU) contains the power electronics 

necessary to operate a direct-drive system (Ref. 14). 

Another goal of this development effort is to develop a DDU. It will be developed by leveraging PPU 

modules like the auxiliary module, the input module, and the SCB. However, in place of the two PPU 

discharge modules is a single DDU discharge module that houses a discharge filter to control ripple 

current from the thruster and a high-voltage switch to isolate the thruster from the power source and 

extinguish short circuit faults. In this design, a mechanical relay will be used as an isolation switch but in 

the future a hybrid switch that combines a relay and semiconductor devices could be implemented to 

improve reliability and transient response.

JPL has been conducting experiments with a 10 kW solar array to investigate technical issues 

associated with direct-drive. Some of the biggest challenges of direct-drive are the effect of discharge 

current oscillations on the various spacecraft systems, how to handle ignitions, transients and faults, and 

how to force current sharing when operating multiple thrusters. One outcome of their effort is the 

development of techniques and algorithms to control operation of Hall thrusters in a direct-drive system.

The algorithms will be included in the SCB for the DDU to control the entire Hall propulsion system. 

Inverter Power

Transformer

Output

Inductor

Rectifier
NASA/TM—2014-216607 10

SiC Component Radiation Testing

A parallel effort was initiated to evaluate the radiation tolerance of the SiC power devices being used 

to develop the PPU and other parts used in the design that are not space qualified. Personnel from NASA 

GRC and Goddard Spaceflight Center (GSFC) are conducting dose rate testing at GSFC facilities and

heavy ion radiation testing at Texas A&M University. The most critical parts for this application are the 

SiC MOSFETs and diodes that are being used for the power converters. SiC MOSFETs from various 

vendors are being evaluated as well as SiC JFET and super junction transistors (SJT). Also, two critical 

components of the gate drive circuit developed for this effort are being evaluated. 

Future Work Plan

Once the brassboard PPU and DDU assembly is completed, they will be subjected to a thorough 

performance test using a resistive load. Then the PPU will be integrated with a Hall thruster at GRC and 

the DDU will be tested at JPL. Lastly, the PPU will be tested while operating in vacuum. Thermal data 

will be collected to validate thermal models and investigate the thermal capability of the design. The PPU 

will then be redesigned to optimize thermal performance, mass, volume and structural design for vibration 

environment. 

Conclusion

NASA GRC is leading the development of a 15 kW Hall thruster PPU with a high input voltage. A 

comprehensive trade study of power semiconductors concluded that SiC was the superior choice for this 

application because it overcomes the performance limitations of state-of-the-art Si components. The 

discharge supply consists of two parallel 7.5 kW modules that use SiC MOSFETs and diodes to process 

power from a nominal input voltage of 300 V into an output of 400 V. The breadboard discharge supply 

demonstrated efficiencies in excess of 97 percent and was successfully integrated with the NASA-300M 

20 kW Hall thruster. Power supplies for the electromagnet, cathode keeper and heater functions were also 

developed and tested. A brassboard PPU with most functionality required for flight and capable of 

operating in vacuum is being developed using a modular approach and common design elements to 

mature the PPU to TRL 5. A brassboard DDU is also being developed using common PPU modules. 

Once fabrication is completed, the PPU and DDU will undergo performance testing on resistive load and 

with a Hall thruster. The technology could then continue its evolution toward a flight ready design. This

game changing technology has the potential of revolutionizing future PPU developments and enable high 

power SEP. 

References

1. Murphy, D., “MegaFlex – The Scaling Potential of UltraFlex Technology,” AIAA-2012-1581, 53rd

Structures, Structural Dynamics and Materials Conference, Honolulu, HI, April 23-26, 2012. 

2. Brophy, J.R., et al., “300-kW Solar Electric Propulsion System Configuration for Human 

Exploration of Near-Earth Asteroids,” AIAA 2011-5514, 47th Joint Propulsion Conference, San 

Diego, CA, July 31-Aug 3, 2011. 

3. Snyder, J.S., et al., “Integration Power and Propulsion System Design for Near-Earth Object Robotic 

Exploration,” AIAA 2011-5662, 47th Joint Propulsion Conference, San Diego, CA, July 31- August 

3, 2011. 

4. Mercer, C.R., et al., “Benefits of Power and Propulsion Technology for a Piloted Electric Vehicle to 

an Asteroid,” AIAA 2011-7252, AIAA SPACE 2011 Conference and Exposition, Long Beach, CA, 

Sep. 27-29, 2011. 

5. Piñero, L.R. and Bowers, G.E., “Multi-Kilowatt Power Module for High-Power Hall Thrusters,” 

AIAA 2004-3973, 40th Joint Propulsion Conference, Fort Lauderdale, FL, July 11-14, 2004. 
NASA/TM—2014-216607 11

6. Hamley, J.A., Piñero, L.R., and Hill, G.M., “10 kW Power Electronics for Hydrogen Arcjets”, 

JANNAF Propulsion Meeting, Indianapolis, IN, February 24-28, 1992. 

7. Wong, S.P., et al., “30 kWe Class High Efficiency Arcjet Power Conditioner: Volume 1,” AFALTR-88-049, Aug. 1988. 

8. Aintablian, H.O., Kirkhan, H., and Timmerman, P., “High Power, High Voltage Electric Power 

System for Electric Propulsion,” AIAA 2006-4134, 4th International Energy Conversion Engineering 

Conference and Exhibit, San Diego, CA, June 26-29, 2006. 

9. Kerslake, T.W., “Effect of Voltage Level on Power System Design for Solar Electric Propulsion 

Missions,” NASA/TM—2003-212304, April 2003. 

10. Piñero, L.R., Kamhawi, H., and Drummond, G., “Integration Testing of a Modular Discharge Supply 

for NASA’s High Voltage Hall Accelerator Thruster,” IEPC-2009-275, 31st International Electric 

Propulsion Conference, Sep. 20-24, 2009. 

11. Piñero, L.R., Hopson, M., Todd, P., and Wong, B., “Performance of the NEXT Engineering Model 

Power Processing Unit,” AIAA 2007-5214, 43th Joint Propulsion Conference, Cincinnati, OH, July 

8-11, 2007. 

12. Piñero, L.R. and Benson, S.W., “NEXT Engineering Model PPU Development, Progress and Plans,” 

AIAA 2011-5659, 47th Joint Propulsion Conference, San Diego, CA, July 31-Aug. 3, 2011. 

13. Hamley, J.A., Cardwell, G.I., McDowell, J., and Matranga, M., “The Design and Performance 

Characteristics of the NSTAR PPU and DCIU”, AIAA 98-3938, 34th Joint Propulsion Conference, 

Cleveland, OH, July 13-15, 1998. 

14. Snyder, J.S., Brophy, J.R., Hofer, R.R., Goebel, D.M. and Katz, I., “Experimental Investigation of a 

Direct-Drive Hall Thruster and Solar Array System at Power Levels up to 10 kW”, AIAA 2012-

4068, 48th Joint Propulsion Conference, Atlanta, GA, July 30-Aug. 1, 2012.