Hung D. Nguyen and Gynelle C. Steele

Glenn Research Center, Cleveland, Ohio

An Overview of SBIR Phase 2 In-Space Propulsion 

and Cryogenic Fluids Management

NASA/TM—2015-218829

July 2015
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Hung D. Nguyen and Gynelle C. Steele

Glenn Research Center, Cleveland, Ohio

An Overview of SBIR Phase 2 In-Space Propulsion 

and Cryogenic Fluids Management

NASA/TM—2015-218829

July 2015

National Aeronautics and

Space Administration

Glenn Research Center

Cleveland, Ohio 44135
Available from

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either expressed or implied, by the National Aeronautics and 

Space Administration.

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This report is available in electronic form at http://www.sti.nasa.gov/ and http://ntrs.nasa.gov/
Abstract 

Technological innovation is the overall focus of NASA’s Small Business Innovation Research 

(SBIR) program. The program invests in the development of innovative concepts and 

technologies to help NASA’s mission directorates address critical research and development 

needs for agency projects. 

This report highlights innovative SBIR Phase II projects from 2007-2012 specifically addressing 

Areas in In-Space Propulsion and Cryogenic Fluids Management which is one of six core 

competencies at NASA Glenn Research Center. There are nineteen technologies featured with 

emphasis on a wide spectrum of applications such as high-performance Hall thruster support 

system, thruster discharge power converter, high-performance combustion chamber, ion thruster 

design tool, green liquid monopropellant thruster, and much more. Each article in this booklet 

describes an innovation, technical objective, and highlights NASA commercial and industrial 

applications.

This report serves as an opportunity for NASA personnel including engineers, researchers, and 

program managers to learn of NASA SBIR’s capabilities that might be crosscutting into this 

technology area. As the result, it would cause collaborations and partnerships between the small 

companies and NASA Programs and Projects resulting in benefit to both SBIR companies and 

NASA. 

An Overview of SBIR Phase 2 

In-Space Propulsion and Cryogenic Fluids Management 

Hung D. Nguyen and Gynelle C. Steele 

National Aeronautics and Space Administration 

Glenn Research Center 

Cleveland, Ohio 44135 

NASA/TM—2015-218829 1
Low-Cost, High-Performance Hall 

Thruster Support System 

Power processing unit (PPU) for Hall thrusters

Colorado Power Electronics (CPE) has built an innovative modular PPU 

for Hall thrusters, including discharge, magnet, heater and keeper supplies, 

and an interface module. This high-performance PPU offers resonant circuit 

topologies, magnetics design, modularity, and a stable and sustained operation 

during severe Hall effect thruster current oscillations. Laboratory testing has 

demonstrated discharge module efficiency of 96 percent, which is considerably 

higher than current state of the art. 

The Phase II project developed an engineering model high-voltage Hall 

accelerator (HiVHAC) PPU that includes a digital control interface unit 

(DCIU). This will position CPE to manufacture a qualification model PPU 

as a Phase III project. The prototype digitally controlled flow controller with a 

PC interface developed in Phase I served as the foundation for a combination 

DCIU-flow module added to the PPU in Phase II. Thermal and vibration 

finite element analysis was performed on the reduced-mass chassis, and then 

a test brassboard PPU was built and tested. Additionally, the control loops of 

the PPU were analyzed and a stress analysis performed. The team designed 

and built an engineering model flight-like PPU that includes flight-like wire 

harnessing schemes, electromagnetic interference filtering, enhanced 

modularity, and the new DCIU-flow module.

Applications

NASA 

} PPU for Hall thrusters

} PPU for ion thrusters

Commercial

} Space-rated piezoelectric driver

} Resonant converters for 

high-power Hall thrusters

Phase II Objectives

} Develop an engineering model 

flight-like HiVHAC PPU

} Develop and integrate a DCIU that 

controls the PPU and xenon feed 

system

} Refine electrical and mechanical 

PPU design

} Improve manufacturability

} Develop electrical, thermal, and 

structural models

Benefits

} Offers discharge module efficiency 

of 96 percent

} Provides sustained operation during 

severe Hall effect thruster current 

oscillations

} Can operate a Hall thruster at a wide 

range of voltages

NASA photo

Firm Contact

Colorado Power Electronics 

Bryce Hesterman

bryce.hesterman@c-pwr.com

120 Commerce Drive, Unit 3

Fort Collins, CO 80524–4731

Phone: 970–482–0191

Proposal Number: 09-2 S3.04-8077

NASA/TM—2015-218829 2
Low-Mass, Low-Power Hall 

Thruster System 

For radioisotope electric propulsion (REP)

NASA is developing an electric propulsion system capable of producing 

20 mN thrust with input power up to 1,000 W and specific impulse ranging 

from 1,600 to 3,500 seconds. The key technical challenge is the target mass of 

1 kg for the thruster and 2 kg for the power processing unit (PPU). In Phase I, 

Busek Company, Inc., developed an overall subsystem design for the thruster/

cathode, PPU, and xenon feed system. This project demonstrated the feasibility 

of a low-mass power processing architecture that replaces four of the DC–DC 

converters of a typical PPU with a single multifunctional converter and a lowmass Hall thruster design employing permanent magnets.

In Phase II, the team developed an engineering prototype model of its lowmass BHT-600 Hall thruster system, with the primary focus on the low-mass 

PPU and thruster. The goal was to develop an electric propulsion thruster with 

the appropriate specific impulse and propellant throughput to enable radioisotope electric propulsion (REP). This is important because REP offers the 

benefits of nuclear electric propulsion without the need for an excessively large 

spacecraft and power system.

Applications

NASA 

} Orbiters around Pluto, Neptune, 

and Uranus

} Rendezvous and Centaurs, 

Kuiper belt objects, and primitive 

bodies in the outer solar system

} Extensive surveys of major 

asteroid groups

Commercial

} Low-power electric 

propulsion systems

} Low-power Hall effect 

thruster systems

} Commercial satellite 

manufacturers:

•	 Primary propulsion on low 

Earth orbit spacecraft

•	 Station keeping on 

geostationary satellites

Phase II Objectives

} Develop a low-mass BHT-600 Hall 

thruster system, with the primary 

focus on the low-mass PPU and 

thruster

} Design, fabricate, and demonstrate 

an engineering model version of a 

low-mass (2 kg) PPU and a low-mass 

(1 kg) version of the BHT-600 thruster

} Conduct an integrated system test 

and deliver a prototype PPU and 

thruster system 

Benefits

} Low mass

} Low power

NASA photo

Firm Contact

Busek Company, Inc.

Bruce Pote

bpote@busek.com

11 Tech Circle

Natick, MA 01760–1023

Phone: 508–655–5565

Proposal Number: 09-2 S3.04-8131

NASA/TM—2015-218829 3
High-Efficiency Hall Thruster 

Discharge Power Converter 

For use in high-voltage Hall accelerator thrusters and other 

electric propulsion devices

Busek Company, Inc., is designing, building, and testing a new printed circuit 

board converter. The new converter consists of two series or parallel boards 

(slices) intended to power a high-voltage Hall accelerator (HiVHAC) thruster 

or other similarly sized electric propulsion devices. The converter accepts 80- to 

160-V input and generates 200- to 700-V isolated output while delivering 

continually adjustable 300-W to 3.5-kW power. Busek built and demonstrated 

one board that achieved nearly 94 percent efficiency the first time it was turned 

on, with projected efficiency exceeding 97 percent following timing software 

optimization. The board has a projected specific mass of 1.2 kg/kW, achieved 

through high-frequency switching.

In Phase II, Busek optimized to exceed 97 percent efficiency and built a second 

prototype in a form factor more appropriate for flight. This converter then was 

integrated with a set of upgraded existing boards for powering magnets and the 

cathode. The program culminated with integrating the entire power processing 

unit and testing it on a Busek thruster and on NASA’s HiVHAC thruster.

Applications

NASA 

} Lunar and Mars exploration

} Vesta–Ceres rendezvous 

(Dawn Mission)

} Comet 22P/Koppf rendezvous

} Asteroid 4660 Nereus sample 

return mission

Commercial

} Department of Defense 

space assets

} Defense Advanced Research 

Projects Agency Fast Access 

Spacecraft Testbed Program:

•	 All-electric, high-power space 

tug and geosynchronous Earth 

orbit servicing vehicle

} Commercial satellite 

manufacturers:

•	 Satellite servicing

•	 Orbit maintenance

•	 Orbit raising and lowering

•	 Inclination changes

•	 Repositioning

Phase II Objectives

} Demonstrate very high efficiency, low 

mass modular power processing unit 

(PPU) with input and output voltages 

appropriate for deep-space missions 

and NASA’s HiVHAC thruster

} Integrate new with existing converters 

operating at different input voltages

} Complete existing breadboard and 

build next version with the correct 

form factor

} Convert existing boards for magnets, 

cathode, etc.

} Integrate new and existing boards

} Test PPU on BHT-1500 thruster at 

Busek and on HiVHAC thruster at 

NASA Glenn

Benefits

} High efficiency Hall thruster

} Low mass

NASA photo

Firm Contact

Busek Company, Inc.

Thomas Jaquish

busek@busek.com

11 Tech Circle

Natick, MA 01760–1023

Phone: 508–655–5565

Proposal Number: 09-2 S3.04-8809

NASA/TM—2015-218829 4
Silicon Carbide (SiC) Power 

Processing Unit (PPU) for Hall 

Effect Thrusters 

A high-efficiency, radiation-hardened power supply

Arkansas Power Electronics International (APEI), Inc., is developing a highefficiency, radiation-hardened 3.8-kW SiC power supply for the PPU of Hall 

effect thrusters. This project specifically targets the design of a PPU for the 

high-voltage Hall accelerator (HiVHAC) thruster, with target specifications of 

80- to 160-V input, 200- to 700-V/5A output, efficiency greater than 

96 percent, and peak power density in excess of 2.5 kW/kg. The PPU under 

development uses SiC junction field-effect transistor power switches, 

components that APEI, Inc., has irradiated under total ionizing dose 

conditions to greater than 3 MRad with little to no change in device 

performance.

Applications

NASA 

} Satellite and spacecraft power 

management systems

} Satellite and spacecraft motors 

and actuators:

•	 Opening and rotating solar 

array panels, controlling 

robotic arms, aligning 

communications arrays, 

pointing cameras and 

instruments, etc.

} Extreme environment 

exploratory vehicles 

Commercial

} High-voltage hybrid electric 

vehicle battery packs

} Photovoltaic energy systems

} Energy exploration

} Industrial motor drives

} Electric vehicle motor drives

} Military systems

Phase II Objectives

} Determine the overall system 

requirements for the PPU power 

converter

} Develop the electrical design of the 

radiation-hardened SiC-based PPU 

power supply

} Fabricate and test PPU power supply 

prototype

Benefits

} More efficient than silicon

} Lower weight and volume

} Operates in high-temperature 

environments

NASA photo

Firm Contact

APEI, Inc.

Bradley Reese

breese@apei.net

535 W. Research Center Blvd.

Fayetteville, AR 72701–7175

Phone: 479–443–5759 ext. 8250

Proposal Number: 09-2 S3.04-8940

NASA/TM—2015-218829 5
Unibody Composite Pressurized 

Structure (UCPS) for In-Space 

Propulsion 

Single, all-composite structure reduces spacecraft 

mass and cost

Microcosm, Inc., in conjunction with the Scorpius Space Launch Company, 

is developing a UCPS for in-space propulsion. This innovative approach 

constitutes a clean break from traditional spacecraft design by combining what 

were traditionally separate primary and secondary support structures and metal 

propellant tanks into a single unit. The all-composite construction is stronger, 

significantly lighter, more robust and reliable, and capable of supporting much 

higher pressures and smaller volume than previous approaches. The single, 

all-composite structure includes linerless high-pressure propellant tank(s), 

composite bosses, flanges, longitudinal and circumferential stringers with 

integral shelves, holding mechanisms, and attach features to support all of the 

spacecraft equipment. These features will replace the separate mission-critical 

primary support structure, tanks, struts, straps, braces, clamps, and brackets 

traditionally required to hold subsystem parts in place. The new structure 

has nearly 0 coefficient of thermal expansion over a temperature range from 

cryogenic to more than 100 °C. 

The Phase II project refined the design of the UCPS integrated with a bladder; 

continued materials compatibility testing with hydrazine; fabricated two UCPS 

as positive expulsion tank (PET) structures; and completed expulsion, 

qualification, and burst testing on one of them.

Applications

NASA 

} Science mission directorate 

spacecraft

} High-pressure liquid and gaseous 

propellants for chemical and 

electric propulsion systems

Commercial

} Aerospace

} Automotive

} Electrical

} Medical

Phase II Objectives

} Refine the system objectives and 

requirements for next-generation 

propulsion systems and how the PET 

can best meet current and future 

needs

} Refine the design for the new UCPS 

structure employing the unibody 

integrated tank/structure approach 

that includes a bladder as a PET

} Conduct more extensive materials 

compatibility tests with UCPS 

structural material and hydrazine

} Fabricate and test a transparent bottle 

plus bladder to characterize bladder 

mechanics

} Fabricate and conduct water-only 

qualification tests of the PET

Benefits

} Allows much higher pressure chemical propulsion systems (in excess of 

2,000 psi) with less tank mass

} Uses all-composite structure 

combined with high-pressure, 

low-volume tanks

NASA photo

Firm Contact

Microcosm, Inc.

Markus Rufer

mrufer@smad.com

4940 W. 147th Street

Hawthorne, CA 90250–6708

Phone: 310–219–2700

Proposal Number: 09-2 S3.04-9659

NASA/TM—2015-218829 6
Low-Cost, High-Performance 

Combustion Chamber

For use with cryogenic liquid oxygen/methane (LOX/CH4) 

propellant

Ultramet designed and fabricated a lightweight, high-temperature combustion 

chamber for use with cryogenic LOX/CH4 propellants that can deliver a specific 

impulse of ~355 seconds. This increase over the current 320-second baseline of 

nitrogen tetroxide/monomethylhydrazine (NTO/MMH) will result in a propellant mass decrease of 55 lbm for a typical lunar mission. The material system 

was based on Ultramet’s proven oxide-iridium/rhenium architecture, which has 

been hot-fire tested with stoichiometric oxygen/hydrogen for hours. Instead of 

rhenium, however, the structural material was a niobium or tantalum alloy that 

has excellent yield strength at both ambient and elevated temperatures. 

Phase I demonstrated alloys with yield strength-to-weight ratios more than 

three times that of rhenium, which will significantly reduce chamber weight. 

The starting materials were also two orders of magnitude less expensive than 

rhenium and were less expensive than the C103 niobium alloy commonly used 

in low-performance engines. 

Phase II focused on the design, fabrication, and hot-fire testing of a 12-lbf 

thrust class chamber with LOX/CH4, and a 100-lbf chamber for LOX/CH4. 

A 5-lbf

 chamber for NTO/MMH also was designed and fabricated.

Applications

NASA 

} Reaction control systems for 

lunar or Martian ascent/descent 

vehicles

} Lunar or Martian sample return 

vehicles

} Main engines for interplanetary 

spacecraft and spacecraft being 

placed into geostationary orbit 

} Apogee topping engines and 

attitude control systems for 

Earth-orbiting satellites

} Launch vehicles for attitude 

control

Commercial

} Apogee topping engines for 

commercial satellites

} Attitude control thrusters for 

launch vehicles 

} Primary propulsion as well as 

divert and attitude control system 

functions for ballistic missile 

defense and tactical missiles

} Drop-in replacements for 

iridium/rhenium engines 

currently being manufactured 

and flown

Phase II Objectives

} Optimize the alloy composition and 

deposition conditions and perform 

extensive mechanical characterization 

of the resulting alloys

} Design and fabricate a flight-weight 

12-lbf

 combustion chamber for 

hot-fire testing

} Perform hot-fire testing of the 12-lbf

chamber with LOX/CH4

} Design and fabricate a highperformance 100-lbf

 combustion 

chamber for use with LOX/CH4

} Design and fabricate a 5-lbf chamber 

for use with NTO/MMH

Benefits

} Reduces combustion chamber weight

} Lowers costs of materials

} Increases yield strength at ambient 

and elevated temperatures

NASA photo

Firm Contact

Ultramet 

Arthur J. Fortini

art.fortini@ultramet.com

12173 Montague Street

Pacoima CA 91331–2210

Phone: 818–899–0236 

Proposal Number: 09-2 X10.01-8636

NASA/TM—2015-218829 7
Wrapped Multilayer Insulation 

Thermal insulation for cryogenic piping

New NASA vehicles, such as Earth Departure Stage (EDS), Orion, landers, and 

orbiting fuel depots, need improved cryogenic propellant transfer and storage 

for long-duration missions. Current cryogen feed line multilayer insulation 

(MLI) performance is 10 times worse per area than tank MLI insulation. 

During each launch, cryogenic piping loses approximately 150,000 gallons 

(equivalent to $300,000) in boil-off during transfer, chill down, and ground 

hold. Quest Product Development Corp., teaming with Ball Aerospace, 

developed an innovative advanced insulation system, Wrapped MLI (wMLI), 

to provide improved thermal insulation for cryogenic feed lines.

wMLI is high-performance multilayer insulation designed for cryogenic piping. 

It uses Quest’s innovative discrete-spacer technology to control layer spacing/

density and reduce heat leak. The Phase I project successfully designed, built, 

and tested a wMLI prototype with a measured heat leak 3.6X lower than 

spiral-wrapped conventional MLI widely used for piping insulation. A wMLI 

prototype had a heat leak of 7.3 W/m2

, or 27 percent of the heat leak of 

conventional MLI (26.7 W/m2

). 

The Phase II project is further developing wMLI technology with custom, 

molded polymer spacers and advancing the product toward commercialization 

via a rigorous testing program, including developing advanced vacuuminsulated pipe for ground support equipment.

Applications

NASA 

} New NASA vehicles

} Orbiting fuel depots

} Vacuum-insulated pipe used to 

transfer cryogens

Commercial 

} Food, research, medical, and 

industrial applications:

•	 Transfers of cryogenic liquid 

into and from cryogenic 

dewars for liquid nitrogen 

(LN2

), hydrogen (LHe), and 

oxygen (LOX)

} Industrial:

•	 Handling LN2

, LOX, and 

liquefied natural gas (LNG)

•	 Handling piping, automatic 

filling equipment, dewar 

manifolds, and gas panels

} LNG:

•	 High-performance insulated 

cryogenic transfer piping 

to reduce LNG losses from 

vaporization during liquid 

transfer

} LN2 equipment:

•	 Semiconductor, electronics, 

and aerospace environmental 

temperature testing

•	 Special effects (fogging), 

biological freezing applications, inerting of food and 

beverage containers, container 

pressurization, and food 

freezing

Phase II Objectives

} Design and develop a custom, 

molded polymer spacer

} Further develop assembly and 

installation processes

} Develop and test wMLI for three 

different piping diameters

} Conduct testing to optimize spacer 

and wrap geometries

} Perform thermal testing on 12 

different wMLI test configurations

} Perform thermal testing on 

advanced ”clam-shell” netting MLI

} Design, develop, and test MLI in a 

vacuum-insulated pipe prototype for 

use in ground support equipment

Benefits

} Low heat leak (3.6X less than 

conventional MLI)

} Easy assembly

} Few layers

} Low cost

} Less mass

Firm Contact

Quest Product Development Corporation 

Scott A. Dye

sdye@quest-corp.com

6833 Joyce Street

Arvada, CO 80007–7570

Phone: 303–670–5088 ext. 12

Proposal Number: 09-2 X8.01-8258

NASA/TM—2015-218829 8
High-Power, High-Thrust Ion 

Thruster (HPHTion) 

For near-Earth applications

Advances in high-power photovoltaic technology have enabled the possibility 

of reasonably sized, high–specific power solar arrays. At high specific powers, 

power levels ranging from 50 to several hundred kilowatts are feasible. Ion 

thrusters offer long life and overall high efficiency (typically >70 percent 

efficiency). In Phase I, the team at ElectroDynamic Applications, Inc., built 

a 25-kW, 50-cm ion thruster discharge chamber and fabricated a laboratory 

model. This was in response to the need for a single, high-powered engine to 

fill the gulf between the 7-kW NASA’s Evolutionary Xenon Thruster (NEXT) 

system and a notional 25-kW engine.

The Phase II project matured the laboratory model into a protoengineering 

model ion thruster. This involved the evolution of the discharge chamber to 

a high-performance thruster by performance testing and characterization via 

simulated and full beam extraction testing. Through such testing, the team 

optimized the design and built a protoengineering model thruster. Coupled 

with gridded ion thruster technology, this technology can enable a wide range 

of missions, including ambitious near-Earth NASA missions, Department of 

Defense missions, and commercial satellite activities.

Applications

NASA 

} Cargo propulsion requirements 

with power levels extending from 

30 kW to 300 kW

} Space science endeavors 

Commercial

} Satellites

} Orbit transfer needs

} Propulsion systems

Phase II Objectives

} Evaluate prototype discharge chamber 

with simulated beam extraction

} Characterize performance of 

discharge chamber with full beam 

extraction

} Design and fabricate stainless steel 

protoengineering model

} Complete performance 

characterization of 

protoengineering model

Benefits

} Long life

} High thrust-to-power ratio 

(~50 mN/kW)

} Low system complexity

NASA photo

Firm Contact

ElectroDynamic Applications, Inc. 

Peter Y. Peterson

info@edapplications.com

3600 Green Court, Suite 300

Ann Arbor, MI 48105–1570

Phone: 734–734–1434

Proposal Number: 09-2 T3.01-9881

NASA/TM—2015-218829 9
Normalized Radial Position

Normalized Axial Position

Tech-X Corp. chart

Next-Generation Ion Thruster 

Design Tool 

To support future space missions

Computational tools that accurately predict the performance of electric 

propulsion devices are highly desirable and beneficial to NASA and the 

broader electric propulsion community. The current state of the art in electric 

propulsion modeling relies heavily on empirical data and numerous 

computational “knobs.” 

In Phase I of this project, Tech-X Corporation developed the most detailed 

ion engine discharge chamber model that currently exists. This kinetic model 

simulates all particles in the discharge chamber along with a physically correct 

simulation of the electric fields. In addition, kinetic erosion models are 

included for modeling the ion-impingement effects on thruster component 

erosion. 

In Phase II, Tech-X developed a user-friendly computer program for NASA 

and other governmental and industry customers. Tech-X has implemented a 

number of advanced numerical routines to bring the computational time down 

to a commercially acceptable level. NASA now has a highly sophisticated, 

user-friendly ion engine discharge chamber modeling tool.

Applications

NASA 

} 40-cm-diameter NASA’s 

Evolutionary Xenon Thruster 

(NEXT)

} Hall thrusters, such as the 

high-voltage Hall accelerator 

(HiVHAC) thruster 

Commercial

} Electric propulsion

} Military and commercial 

satellites: 

•	 Satellite station keeping 

•	 Orbit-changing maneuvers 

in space

} Ion source and plasma 

processing

Phase II Objectives

} Reduce computational time of new 

modeling tool

} Make new modeling tool user friendly 

and commercially attractive

Benefits

} Reduces time and expense in 

designing new/different sizes of 

ion engines 

} Analyzes existing ion engine 

performance

} Extends the operating range of the 

NEXT ion engine to higher power 

levels

Firm Contact

Tech-X Corporation

Peter Stoltz

pstoltz@txcorp.com

5621 Arapahoe Avenue

Boulder, CO 80303–1379

Phone: 303–996–7527

Proposal Number: 09-2 T3.01-9893

NASA/TM—2015-218829 10
Nanowire Photovoltaic Devices 

For ultrahigh-efficiency, radiation-tolerant space solar cells

Firefly Technologies, in collaboration with the Rochester Institute of 

Technology and the University of Wisconsin–Madison, developed synthesis 

methods for highly strained nanowires. Two synthesis routes resulted in 

successful nanowire epitaxy: direct nucleation and growth on the substrate 

and a novel selective-epitaxy route based on nanolithography using diblock 

copolymers. The indium-arsenide (InAs) nanowires are implemented in situ

within the epitaxy environment—a significant innovation relative to conventional semiconductor nanowire generation using ex situ gold nanoparticles. The 

introduction of these nanoscale features may enable an intermediate band solar 

cell while simultaneously increasing the effective absorption volume that can 

otherwise limit short-circuit current generated by thin quantized layers. The 

use of nanowires for photovoltaics decouples the absorption process from the 

current extraction process by virtue of the high aspect ratio. 

While no functional solar cells resulted from this effort, considerable fundamental understanding of the nanowire epitaxy kinetics and nanopatterning 

process was developed. This approach could, in principle, be an enabling 

technology for heterointegration of dissimilar materials. The technology also is 

applicable to virtual substrates. Incorporating nanowires onto a recrystallized 

germanium/metal foil substrate would potentially solve the problem of grain 

boundary shunting of generated carriers by restricting the cross-sectional area 

of the nanowire (tens of nanometers in diameter) to sizes smaller than the 

recrystallized grains (0.5 to 1 μm2

).

Applications

NASA 

} Space-based applications

} Virtual substrates 

Commercial

} Space-based power generation

} Photovoltaics:

•	 Integration of III-V 

nanostructures on low-cost 

silicon substrates

Phase II Objectives

} Demonstrate controllable 

gallium-arsenide (GaAS)-embedded 

InAs nanowire growth and 

characterization

} Demonstrate improved cell 

performance from nanowire-infused 

cells compared to standard GaAs cells

Benefits

} Offers solar cell efficiency of more 

than 40 percent

} Reduces costs in terms of photovoltaic 

array size, array weight, and launch 

costs

} Enables more optimal bandgap and 

material combinations for novel 

devices

NASA photo

Firm Contact

Firefly Technologies

David Forbes

dvfsps@rit.edu

111 Lomb Memorial Drive

Rochester, NY 14623–5608

Phone: 330–421–2104

Proposal Number: 09-2 T3.01-9972

NASA/TM—2015-218829 11
Solid model 

of full-scale 

prototype 

thruster

Lab 

model 

propellant 

feed 

system

NASA photos

20-mN Variable Specific Impulse 

(Isp) Colloid Thruster 

For colloid-based, multimode propulsion using ionic liquids

Busek Company, Inc., has designed and manufactured an electrospray emitter 

capable of generating 20 mN in a compact package (7x7x1.7 in). The thruster 

consists of nine porous-surface emitters operating in parallel from a common 

propellant supply. Each emitter is capable of supporting over 70,000 electrospray emission sites with the plume from each emitter being accelerated 

through a single aperture, eliminating the need for individual emission site 

alignment to an extraction grid. The total number of emission sites during 

operation is expected to approach 700,000.

This Phase II project optimized and characterized the thruster fabricated 

during the Phase I effort. Additional porous emitters also were fabricated for 

full-scale testing. Propellant is supplied to the thruster via existing feed-system 

and microvalve technology previously developed by Busek, under the NASA 

Space Technology 7’s Disturbance Reduction System (ST7-DRS) mission and 

via follow-on electric propulsion programs. This project investigated methods 

for extending thruster life beyond the previously demonstrated 450 hours. The 

life-extending capabilities will be demonstrated on a subscale version of the 

thruster.

Applications

NASA 

} Missions requiring exceptional 

thrust/power where power is 

limited

} Missions benefiting from variable Isp (200 to 2,000+ s) and 

variable thrust (10x throttling) 

Commercial

} Colloid-based multimode 

propulsion using ionic liquids, 

where there is a lack of suitable 

thrusters at the millinewton level 

or greater

} Replacement for lower 

thrust plasma-based electric 

propulsion devices, which suffer 

from decreased efficiency at 

low power due to unfavorable 

surface-to-volume scaling

} Spacecraft requiring both dwell/

station keeping and rapid 

maneuvers from a single 

propulsion system

Phase II Objectives

} Validate Busek’s concept for scaling 

electrospray thrust to the 20-mN level 

through self-organized electrospray 

emission from porous surfaces

} Develop a thruster package that 

is compact, robust, and does not 

require complicated manufacturing 

techniques

} Optimize the design for best performance and improve the total impulse 

achievable through increased thruster 

lifetime

Benefits

} Offers a compact, robust package

} Potential future use of shared 

propellant tanks for high-thrust 

chemical propulsion as well as high 

Isp electric propulsion

} Enables new classes of missions 

benefiting from variable Isp of 200 to 

5,000+ s and variable thrusts up to 

and exceeding 20 mN

Firm Contact

Busek Company, Inc.

Nathaniel Demmons

nate@busek.com

11 Tech Circle

Natick, MA 01760–1023

Phone: 508–655–5565

Proposal Number: 10-2 S3.04-8878

NASA/TM—2015-218829 12
NASA photo

Green Liquid Monopropellant 

Thruster 

For in-space propulsion in challenging environmental 

conditions

Physical Sciences, Inc. (PSI), and Orbital Technologies Corporation 

(ORBITEC) are developing a unique chemical propulsion system for 

next-generation NASA science spacecraft and missions. The system is compact, 

lightweight, and can operate with high reliability over extended periods of 

time and under a wide range of thermal environments. The system uses a new 

storable, low-toxicity liquid monopropellant as its working fluid. In Phase I, 

the team demonstrated experimentally the critical ignition and combustion 

processes for the propellant and used the data to develop thruster design 

concepts. 

In Phase II, the team developed and demonstrated in the laboratory a proofof-concept prototype thruster. A Phase III project is envisioned to develop a 

full-scale protoflight propulsion system applicable to a class of NASA missions.

Applications

NASA 

This technology is ideal for 

propulsion systems in challenging 

environmental conditions, with long 

operational life, and with high duty 

cycles:

} Missions to other planets, their 

moons, and other small bodies:

•	 Sampling their atmospheres 

•	 Descent and landing on their 

surfaces

•	 Returning soil samples 

from their surfaces in ascent 

modules

•	 Rendezvous and docking with 

orbiting mother ships

} Missions to the Earth’s Moon, 

Venus, Mars and its moons, the 

moons of Jupiter, and asteroids

Military

} Air Force:

•	 Liquid thrusters for in-space 

propulsion

} National Reconnaissance 

Organization:

•	 Fast-response, long-life 

maneuvering propulsion 

systems

} Army:

•	 Green monopropellant–based, 

high-pressure gas generators 

for pressurizing gelled 

propellants

} Other Department of Defense:

•	 Liquid engines for propelling 

highly maneuverable, 

throttleable tactical missiles

Phase II Objectives

} Demonstrate operation of a prototype 

liquid monopropellant thruster 

incorporating the new fast-response, 

high-reliability, and long-life ignition 

system

} Develop ignition system designs to 

optimize performance

} Develop a modular, small-scale test 

thruster incorporating the ignition 

system to allow rapid changeability of 

components representative of different 

ignition/combustor designs

} Characterize the operation and performance of the small-scale thruster to 

demonstrate propulsive performance 

and ignition system start/stop/restart

} Develop functional prototype of a 

large-scale thruster and demonstrate 

its operation

} Develop complete propulsion system 

design incorporating the proposed 

in-space thruster technology

Benefits

} Compact and lightweight

} Can operate reliably over extended 

periods of time and under a wide 

range of thermal environments

Firm Contact

Physical Sciences, Inc.

Prakash B. Joshi

joshi@psicorp.com

20 New England Business Center

Andover, MA 01810–1077

Phone: 978–689–0003 ext. 8202

Proposal Number: 10-2 S3.04-8888

NASA/TM—2015-218829 13
NASA photo

Improved Rhenium Thrust 

Chambers 

Reduces chamber fabrication costs by 30 percent or more

Radiation-cooled bipropellant thrust chambers are being considered for ascent/

descent engines and reaction control systems on various NASA missions and 

spacecraft, such as the Mars Sample Return and Orion Multi-Purpose Crew 

Vehicle (MPCV). Currently, iridium (Ir)-lined rhenium (Re) combustion 

chambers are the state of the art for in-space engines. NASA’s Advanced 

Materials Bipropellant Rocket (AMBR) engine, a 150-lbf

 Ir-Re chamber 

produced by Plasma Processes and Aerojet Rocketdyne, recently set a hydrazine 

specific impulse record of 333.5 seconds. 

To withstand the high loads during terrestrial launch, Re chambers with 

improved mechanical properties are needed. Recent electrochemical forming 

(EL-FormTM) results have shown considerable promise for improving Re’s 

mechanical properties by producing a multilayered deposit composed of a 

tailored microstructure (i.e., Engineered Re). 

The Engineered Re processing techniques were optimized, and detailed characterization and mechanical properties tests were performed. The most promising 

techniques were selected and used to produce an Engineered Re AMBR-sized 

combustion chamber for testing at Aerojet Rocketdyne.

Applications

NASA 

} Ascent/descent engines and reaction 

control systems for missions: 

•	 Mars Sample Return

•	 Orion MPCV

} In-space propulsion components for:

•	 Apogee insertion

•	 Attitude control

•	 Orbit maintenance

•	 Repositioning of satellites/

spacecraft

•	 Reaction control systems

•	 Descent/Ascent engines

•	 Nuclear power/propulsion

•	 Microgravity containment 

crucibles and cartridges

Commercial

} Defense

} Material research and development

} Nuclear power

} Aerospace

} Propulsion

} Automotive

} Electronics

} Crystal growth

} Medical

} Net-shaped fabrication of refractory 

and platinum group metals for:

•	 Rocket nozzles

•	 Crucibles

•	 Heat pipes

•	 Propulsion subcomponents

} Advanced coating systems for:

•	 X-ray targets

•	 Sputtering targets

•	 Turbines

•	 Rocket engines

•	 Wear and thermal/electrical 

insulation

Phase II Objectives

} Optimize the Engineered Re processing techniques to produce an 

AMBR-sized chamber that has room 

temperature yield strength of 40 ksi 

and 10 percent elongation

} Perform detailed characterization 

of the pinning layers and tailor the 

architecture of the Engineered Re to 

achieve the desired properties

} Perform extensive materials properties 

testing and select the most promising 

Engineered Re fabrication method

} Optimize the multicomponent 

processing technique to produce 

Engineered Re deposits on multiple 

AMBR-sized mandrels

} Demonstrate repeatability of the 

optimized Engineered Re fabrication 

technique

} Produce an AMBR-sized combustion 

chamber that incorporates the 

Engineered Re structure and test

Benefits

} Reduces chamber fabrication costs by 

30 percent or more

EL-Form is a registered trademark of Plasma Processes, LLC.

Firm Contact

Plasma Processes

John Scott O’Dell

scottodell@plasmapros.com

4914 Moores Mill Road

Huntsville, AL 35811–1558

Phone: 256–851–7653 ext. 104

Proposal Number: 10-2 S3.04-9213

NASA/TM—2015-218829 14
NASA photo

Nontoxic Hydroxylammonium 

Nitrate (HAN) Monopropellant 

Propulsion 

To replace hydrazine with a green monopropellant

Nontoxic monopropellants have been developed that provide better performance than toxic hydrazine. Formulations based on HAN have superior 

performance as compared to hydrazine with enhanced specific impulse (Isp), 

higher density and volumetric impulse, lower melting point, and much lower 

toxicity. However, HAN-based monopropellants require higher chamber 

temperatures (2,083 K vs. 883 K) to combust. Current hydrazine-based 

combustion chamber technology (Inconel® or niobium C103 and silicide 

coating) and catalyst (Shell 405) are inadequate. In Phase I, state-of-the-art 

iridium-lined rhenium chambers and innovative new foam catalysts were 

demonstrated in pulse and 10-second firings. Phase II developed and tested a 

flight-weight thruster for an environmentally green monopropellant.

Applications

NASA 

} Mars Ascent Vehicle

} Lunar landers

} Reaction control systems

} In-space propulsion:

•	 Attitude control

•	 Orbit maintenance

•	 Repositioning of satellites/

spacecraft

•	 Descent/Ascent engines

} Nuclear power/propulsion

} Microgravity containment 

crucibles and cartridges

Commercial

} Rocket nozzles for satellites and 

military

} Crucibles, heat pipes, propulsion 

subcomponents, X-ray targets, 

sputtering targets, turbines, 

rotors, furnaces, power 

generation, jet engine restarters, 

catalysts, etc.

Phase II Objectives

} Develop a nontoxic HAN-based 

monopropellant thruster to replace 

hydrazine

} Pursue improvements to Phase I foam 

catalyst and injector

} Demonstrate ignition of nontoxic 

AF-M315E with Phase II design

} Test life and response time of the 

thrust chamber assembly

} Analyze posttest thruster and catalyst

} Fabricate multiple thrust chambers for 

commercial partners 

} Commercialize green thruster

Benefits

} Enhanced Isp of 261 seconds (12 

percent greater than hydrazine)

} Higher density and volumetric impulse 

(60 percent greater density impulse 

than hydrazine)

} Lower melting point than hydrazine

} Low toxicity (no self-contained 

breathing apparatus required)

Inconel is a registered trademark of Special Metals Corp.

Firm Contact

Plasma Processes

Timothy N. McKechnie

timmck@plasmapros.com

4914 Moores Mill Road

Huntsville, AL 35811–1558

Phone: 256–851–7653 ext. 116

Proposal Number: 10-2 X2.02-9554

NASA/TM—2015-218829 15
NASA photo

Extreme-Environment SiliconCarbide (SiC) Wireless Sensor Suite 

For real-time monitoring of nuclear thermal propulsion 

(NTP) engines

A number of critical telemetry measurements must be monitored under 

continuous field operation, including temperature data across the reactor 

chamber and the nozzle, pressure data, neutron flux density, and flow rate of the 

propellant. Real-time monitoring of these data in nuclear thermal engines would 

greatly improve operational safety and performance, reduce operational costs, 

and significantly impact maintenance costs and reliability. Even though some 

extreme-environment sensors have become available recently, it is still impossible 

to directly and accurately measure the critical operational parameters of NTP 

engines due to the lack of extreme-environment electronics for those sensors. 

Data from extreme-environment sensors is delivered via wire-line to an external 

actively cooled electronics box, where it is processed. This approach presents 

significant drawbacks, such as the need for complex shielded wiring harnesses 

that not only are heavy but also limit sensor location and signal quality (i.e., 

signal-to-noise ratio). Additionally, these systems suffer from reliability issues 

due to wiring connections. 

In this Phase II project, Arkansas Power Electronics International (APEI), Inc., 

delivered a set of SiC-based integrated wireless sensor-transmitter suites for 

extreme-temperature operation (450 °C) in NTP engines. These sensor suites 

allow for real-time monitoring of critical engine components, reducing the risk 

of catastrophic failure and decreasing the inherent risk associated with NTP 

operation. The final wireless sensor systems are fully integrated into an 

autonomous drop-in solution for advanced sensing systems, including 

wireless energy harvesting.

Applications

NASA 

} Health monitoring systems of 

NASA space exploration vehicles:

•	 Spacecraft

•	 Rockets

•	 Aircraft 

Commercial

} Health monitoring of turbine 

engines for both military and 

commercial aircraft:

•	 This technology enables 

nearly continuous onboard 

situational awareness of the 

vehicle health state for use by 

the flight crew, ground crew, 

and maintenance depot.

} Power generation, including both 

nuclear power generation and gas 

turbine power generation:

•	 By introducing hightemperature sensors and 

wireless transmitters into the 

gas turbine units (specifically 

within the blades where 

temperatures range from 450 

to 1,200 °C), very accurate 

turbine conditions can be 

determined in real time.

Phase II Objectives

} Develop an integrated SiC wireless 

sensor suite capable of in situ 

measurements of critical 

characteristics of NTP engine

} Compose SiC wireless sensor suite of: 

•	 Extreme-environment sensors

•	 Dedicated high-temperature 

(450 °C) SiC electronics that 

provide power and signal conditioning capabilities as well as radio 

frequency modulation and wireless 

data 

transmission capabilities

•	 An onboard energy harvesting 

system as a power source

Benefits

} Allows for real-time monitoring 

of critical engine components

} Reduces the risk of catastrophic 

failure

} Decreases risk associated with NTP 

operation

} Eliminates (or reduces) the need for 

thermal shielding and active cooling 

systems, reducing the size, weight, 

and the complexity of control systems

Firm Contact

APEI, Inc.

Jie Yang

jyang@apei.net

535 West Research Center Blvd., Suite 209

Fayetteville, AR 72701–6959

Phone: 479–443–5759 ext. 8280

Proposal Number: 10-2 X2.03-9748

NASA/TM—2015-218829 16
NASA photo

Data Concentrator 

For modular and distributed control of propulsion systems

Orbital Research, Inc., developed, built, and tested three high-temperature 

components for use in the design of a data concentrator module in distributed 

turbine engine control. The concentrator receives analog and digital signals 

related to turbine engine control and communicates with a full authority digital 

engine control (FADEC) or high-level command processor. This data concentrator follows the Distributed Engine Controls Working Group (DECWG) 

roadmap for turbine engine distributed controls communication development 

that operates at temperatures at least up to 225 °C.

In Phase I, Orbital Research developed detailed specifications for each 

component needed for the system and defined the total system specifications. 

This entailed a combination of system design, compiling existing component 

specifications, laboratory testing, and simulation. The results showed the 

feasibility of the data concentrator.

Phase II of this project focused on three key objectives. The first objective 

was to update the data concentrator design modifications from DECWG and 

prime contractors. Secondly, the project defined requirements for the three new 

high-temperature, application-specific integrated circuits (ASICs): one-time 

programmable (OTP), transient voltage suppression (TVS), and 3.3V. Finally, 

the project validated each design by testing over temperature and under load.

Applications

NASA 

} NASA space programs will 

benefit from the development of 

the high-temperature electronic 

(HTE) component chips, data 

concentrator, and multichip 

modules.

Commercial

} Next-generation military and 

civilian aircraft turbine engines:

•	 Rotorcraft

•	 Unmanned aircraft systems

•	 Land vehicles

} Downhole drilling and 

geothermal drilling controls

} Ground testing rocket and 

turbine engines

} Prognostic/Integrated system 

health management (PHM/

ISHM)

} Chemical, nuclear, refinery, and 

process plant instrumentation

} Powertrain controls for gas 

or diesel internal combustion 

engines (e.g., improved waste 

gate turbobooster)

Phase II Objectives

} Update data concentrator 

modifications from DECWG and 

prime contractor

} Define requirements for the OTP, TVS, 

and 3.3V linear regulator

} Validate each design by testing over 

temperature and under load

Benefits

} Enables replacement of the large 

bundles of noise-sensitive analog 

cable interfaces between sensors/

actuators and the FADEC with a digital 

data bus, thus reducing weight

} Optimizes actuator performance 

through closed-loop control

} Lowers maintenance costs through 

condition-based maintenance (CBM) 

and PHM capabilities

} Lowers costs for FADEC upgrades

Firm Contact

Orbital Research, Inc.

Mike Willett

willett@orbitalresearch.com

4415 Euclid Avenue, Suite 500

Cleveland, OH 44103–3757

Phone: 216–649–0399

Proposal Number: 10-2 A2.07-8712

NASA/TM—2015-218829 17
Busek Co., Inc. photo

Magnesium Hall Thruster 

For solar system exploration

This Phase II project is developing a magnesium (Mg) Hall effect thruster 

system that would open the door for in situ resource utilization (ISRU)-based 

solar system exploration. Magnesium is light and easy to ionize. For a Mars–

Earth transfer, the propellant mass savings with respect to a xenon Hall effect 

thruster (HET) system are enormous. Magnesium also can be combusted in 

a rocket with carbon dioxide (CO2

) or water (H2O), enabling a multimode 

propulsion system with propellant sharing and ISRU. In the near term, CO2

and H2

O would be collected in situ on Mars or the moon. In the far term, Mg 

itself would be collected from Martian and lunar regolith. 

In Phase I, an integrated, medium-power (1- to 3-kW) Mg HET system was 

developed and tested. Controlled, steady operation at constant voltage and 

power was demonstrated. Preliminary measurements indicate a specific impulse 

(Isp) greater than 4,000 s was achieved at a discharge potential of 400 V. The 

feasibility of delivering fluidized Mg powder to a medium- or high-power 

thruster also was demonstrated. 

Phase II of the project evaluated the performance of an integrated, highpower Mg Hall thruster system in a relevant space environment. Researchers 

improved the medium power thruster system and characterized it in detail. 

Researchers also designed and built a high-power (8- to 20-kW) Mg HET. A 

fluidized powder feed system supporting the high-power thruster was built and 

delivered to Busek Company, Inc.

Applications

NASA 

} NASA Flagship, Frontier, and 

Discovery class missions: 

•	 Extremely high Isp at 

voltages typical of low-cost, 

flight-qualified power 

processors

} Missions to asteroids, comets, 

and the outer planets

} Sample return missions

} Lunar and Martian missions

} Manned missions for transporting fuel and cargo

} In situ propellant use at the 

Moon and Mars

} Multimode Mg-based propulsion 

system:

•	 Featuring a Mg rocket and 

a Hall thruster with full or 

partial propellant sharing

Commercial

} Satellite orbit maintenance, orbit 

raising, and repositioning

} One-half of a multimode 

propulsion system that also 

contains a Mg-based rocket 

Phase II Objectives

} Evaluate the performance of an 

integrated, laboratory model Mg Hall 

thruster system in a relevant space 

environment

} Develop a Mg Hall thruster system

} Achieve Isp of 4,000 to 6,000 s from 

commercial orbital transportation 

services–derived 400- to 600-V 

power processing unit (PPU) 

Benefits

} Higher Isp and less life-limiting 

erosion than a xenon HET:

•	 Isp for a high-efficiency Mg Hall 

thruster driven by a 400-V PPU 

may exceed 5,000 s

} Low propellant cost and low-pressure 

propellant storage

} System efficiency in excess of 50 

percent

Firm Contact

Busek Company, Inc.

James J. Szabo

jszabo@busek.com

11 Tech Circle

Natick, MA 01760–1023

Phone: 508–655–5565

Proposal Number: 10-2 T3.01-9872

NASA/TM—2015-218829 18
Aspen Aerogels photo

Thin Aerogel as a Spacer in 

Multilayer Insulation 

For cryogenic space applications

Cryogenic fluid management is a critical technical area that is needed for 

future space exploration. A key challenge is the storability of liquid hydrogen 

(LH2

), liquid methane (LCH4), and liquid oxygen (LOX) propellants for 

long-duration missions. The storage tanks must be well-insulated to prevent 

overpressurization and venting, which can lead to unacceptable propellant 

losses for long-duration missions to Mars and beyond. 

Aspen Aerogels had validated the key process step to enable the fabrication of 

thin, low-density aerogel materials. The multilayer aerogel insulation (MLAI) 

system prototypes were prepared using sheets of aerogel materials with superior 

thermal performance exceeding current state-of-the-art insulation for space 

applications. The exceptional properties of this system include a new breakthrough in high-vacuum cryogenic thermal insulation, providing a durable 

material with excellent thermal performance at a reduced cost when compared 

to longstanding state-of-the-art multilayer insulation systems. During the 

Phase II project, further refinement and qualification/system-level testing of the 

MLAI system will be performed for use in cryogenic storage applications.

Aspen has been in discussions with United Launch Alliance, LLC; NASA’s 

Kennedy Space Center; and Yetispace, Inc., to test the MLAI system on 

real-world tanks such as Vibro-Acoustic Test Article (VATA) or the Cryogenic 

Orbital Testbed (CRYOTE).

Applications

NASA 

} Insulation for cryotanks and 

cryogen transfer pipelines for 

ground processing 

} Cryogen storage insulation for 

in-space applications

} Satellite thermal management

} Extravehicular activity (EVA) 

suits 

} Internal insulation on future 

generations of reusable launch 

vehicles

Commercial

} Durable and reliable insulation 

systems for any cryogenic, 

high-vacuum, or thin and 

flexible applications: 

•	 Appliances

•	 Airliner fuselage

•	 Liquid natural gas fuel storage 

tanks and transfer lines

•	 Apparel

Phase II Objectives

} Refine low-density and thin aerogel 

formulations

} Optimize thin aerogel scale-up 

process and MLAI system fabrication

} Assess MLAI prototype performance:

•	 Cryostat 500 (size 8-in diameter) 

testing

} Large-scale testing of optimum 

MLAI system:

•	 Cryostat 100 (2 x 4-ft size)

•	 System-level testing at Ball 

Aerospace & Technologies Corp. 

Benefits

} More durable and robust multilayer 

insulation system that is easy to install

} Improved process for manufacturing 

thin, low-density aerogel materials

} Superior thermal insulation for 

cryogenic applications

Firm Contact

Aspen Aerogels 

Nancy Moroz

contracts@aerogel.com

30 Forbes Road, Building B

Northborough, MA 01532–2501

Phone: 508–691–1161

Proposal Number: 11-2 X10.01-9295

NASA/TM—2015-218829 19
NASA photo

Iodine Hall Thruster 

For space exploration

Iodine enables dramatic mass and cost savings for lunar and Mars cargo 

missions, including Earth escape and near-Earth space maneuvers. The 

demonstrated throttling ability of iodine is important for a singular thruster 

that might be called upon to propel a spacecraft from Earth to Mars or Venus. 

The ability to throttle efficiently is even more important for missions beyond 

Mars.

In the Phase I project, Busek Company, Inc., tested an existing Hall thruster, 

the BHT-8000, on iodine propellant. The thruster was fed by a high-flow 

iodine feed system and supported by an existing Busek hollow cathode flowing 

xenon gas. The Phase I propellant feed system was evolved from a previously 

demonstrated laboratory feed system. Throttling of the thruster between 2 and 

11 kW at 200 to 600 V was demonstrated. Testing showed that the efficiency 

of iodine fueled BHT-8000 is the same as with xenon, with iodine delivering a 

slightly higher thrust-to-power (T/P) ratio.

In Phase II, a complete iodine-fueled system was developed, including the 

thruster, hollow cathode, and iodine propellant feed system. The nominal 

power of the Phase II system is 8 kW; however, it can be deeply throttled as 

well as clustered to much higher power levels. The technology also can be scaled 

to greater than 100 kW per thruster to support megawatt-class missions. The 

target thruster efficiency for the full-scale system is 65 percent at high specific 

impulse (Isp) (~3,000 s) and 60 percent at high thrust (Isp ~2,000 s).

Applications

NASA 

} Orbit raising and interplanetary 

transfers:

•	 Exploration and science 

missions to near-Earth objects, 

asteroids, comets, and planets

} Reboosting the International 

Space Station (ISS), which 

is currently accomplished by 

chemical propulsion

Commercial

} Orbiting spacecraft:

•	 Orbit raising

•	 Orbit circularization

•	 Inclination changes

•	 Station keeping 

•	 Repositioning

} High-power electric upper stage 

for a commercial launch vehicle 

Phase II Objectives

} Design and build fully integrated 

high-power iodine thruster system 

consisting of a thruster, cathode, and 

propellant feed system 

} Measure performance of integrated 

system

} Assess materials compatibility 

requirements and system issues

} Collect data to assess spacecraft 

interactions

Benefits

} High-purity iodine is available 

commercially in large quantities and 

at much lower cost than xenon.

} Iodine stores at two to three times 

greater density than xenon and at 

approximately one-thousandth of 

the pressure and may be stored in 

low-mass, low-cost propellant tanks.

} Passive, long-term storage of a fully 

fueled system is feasible, including 

storage in conformal tanks that may 

be used to shield internal components 

against some types of space radiation.

Firm Contact

Busek Company, Inc.

James Szabo

jszabo@busek.com

11 Tech Circle

Natick, MA 01760–1023

Phone: 508–655–5565

Proposal Number: 11-2 X2.03-8838

NASA/TM—2015-218829 20