Dual-Mode Propulsion System 

Enabling CubeSat Exploration 

of the Solar System

NASA Innovative Advanced Concepts

2014 NIAC Symposium

Stanford University

February 4 – 6, 2014

Nathan Jerred, Troy Howe & Steven Howe

 Center for Space Nuclear Research 

Adarsh Rajguru

 University of Southern California 
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Inspiration

Limited budgets make large missions more difficult to fund

– achieve more science-per-dollar

• Develop technologies enabling reliable, compact exploration platforms

– long-lived and long-ranged mobile platform

off-the-shelf propulsion system

propulsion system for micro-satellite payloads

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• Target small launch vehicles 

– less than 1,000 kg to LEO

• Enable affordable deep space exploration

GOAL 

– deliver 6U CubeSat payload to Enceladus orbit (~15 kg) 

develop an appropriately sized propulsion system concept 

develop the mission architecture 

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1U CubeSat frame
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Concept

energy source

Radioisotope Decay

proposed radioisotope: 238PuO2 

heritage use within NASA 

high specific energy: 1.6 x 106 MJ/kg (thermal)

Chemical propellants: 10 MJ/kg (thermal) 

RTG: 9.6x104 MJ/kg (6%) vs Li-ion: 0.72 MJ/kg (electric) 

poor specific power: 0.392 W/g [238PuO2] 

238 PuO2 Pellet

DOE

• Fuel Containment

– fuel encapsulated in a tungsten-based matrix 

• provides high strength & toughness

– can provide great energy density

Tungsten-based CERMET

O’Brien et al.
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Approach

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Radioisotope-Based Core 

decay energy accumulated within central core, i.e, thermal capacitor 

direct propellant heating for 

propulsion

radioisotope thermal rocket 

(RTR) 

– thermal energy converted to 

electrical energy

• electric propulsion

• power generation

– accumulated energy is depleted 

through each impulse

artistic rendering of the concept in Earth orbit
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Concept

energy storage

Thermal Capacitor Qualifications

– high thermal storage, high thermal conductivity, high melting 

temperature

sensible heat  based on a material’s heat capacity

latent heat  based on energy needed to change a material’s phase

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Core Material

Silicon identified as a suitable material 

 ΔHfusion = 1.8 MJ/kg

 Tmelt = 1687 K

Allows for an operational 

temperature of around 1700 K

proposed cycle process

Radioisotope 

Heats 

Thermal 

Capacitor

Core 

Reaches 

Peak 

Temperature

Blowdown

of Gas 

Through 

Core

Electrical 

Conversion 

System 

Produces 

Power

Propellant 

Injected into 

Core 

Produces 

Thrust

Impulsebased 

Function 

Performed

Regeneration 

Cycle

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Concept

dual-mode

Direct Propellant Heating

provides thermal-based propulsion

heated propellant expelled through nozzle 

creating thrust

Energy Conversion

closed loop Brayton cycle 

10’s of kW per pulse

multiple pulses per day

dual Brayton engines

thermal propulsion 

flow schematic

electrical conversion 

flow schematic
Optimization

• Previous work performed on a Martian exploration probe

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Starr-CCM+ steady-state 

simulation of Mars Hopper 

concept 

22.5° symmetry StarrCCM+ steady-state 

model 

Thermal interactions within the core

details core sizing such as isotope loading, capacitor size, 

insulation, etc.

Thermal hydraulics between the core and flowing gas

details performance of propellant yielding chamber pressure 

& temperature aiding in nozzle development

details performance of working fluid for energy conversion 

to help determine power generation and efficiencies
Mission Design

instrumentation & communication

Proposed Instrument Package

Instrument Mass (g) Volume (cm3) Power (W)

Thermal Imaging Radiometer 1,000’s 1,000’s 10’s

Infrared Spectrometer 100’s 100’s -

High Resolution Camera (MAC) 100’s 100’s 1’s

Mass Spectrometer 100’s 10’s 1’s

6U payload frame

communication system

Communication System 

Item Symbol Units Value

Frequency f GHz 27.50

Transmission Power Pt Watts 25000

Transmission Antenna Dia. Dt m 1.5

Trans. Antenna Gain (net) Gt dBi 50.16

Prop. Path Length S km 1.27(10)9

Space Loss Ls dB -303.30

System Noise Temp. Ts K 84.10

Data Rate R Mbps 1 

SNR Eb/No - 5.94
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Payload Power

A compact Radioisotope Thermo-Photovoltaic (RTPV) power source 

~being studied at the CSNR

estimated RPS for 

5 We – 400g, “D cell” size

Solid-state energy conversion

continuous energy production

no moving parts & long mission lifetimes 

Encapsulation method allows for modification 

in the source size

can be sized to accommodate payload needs

Tungsten matrix shields harmful radiation

power sources can be housed among 

instrumentation 

provides high temperature strength

Emitter material or coating can be integrated 

with the CERMET fuel

increase conversion efficiency
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Trajectory

Thermal Propulsion Mode

provides high thrust & moderate Isp 

impulse function allows for phasing 

maneuvers (perigee pumping) to achieve 

Earth escape

Electric Propulsion Mode

provides high Isp for interplanetary travel 

allows for shorter transit times 

four 2.2 kW Hall Effect thrusters 

graphic of Earth-based phasing maneuvers

Earth – Jupiter – Saturn

• Earth Departure: Jan-18-2018

Saturn Arrival: July-11-2023

Total Duration: 5.48 years

Injection ΔV (Earthescaped): 6.42 km/s

Jupiter Gravity Assist: 0.12 km/s

Post Injection ΔV (Saturncapture): 0.54 km/s

Total ΔV required by propulsion system = 6.96 km/s

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AeroJet BPT-2000 Hall 

Effect thruster

possible trajectory to Saturn
Performance

PRELIMINARY RESULTS

based on Earth escape

Operation

 timeblowdown: 360 s

 timeheatup: 0.2 day

 timeescape: 80 days 

 ΔVescape: 3.8 km/s

 temp: 1700 K

Core

 30 cm length

 20 cm diameter

 2 mm dia. flow channels

 18.1 kg Silicon

 7.16 kg 238PuO2

 89.8 kWt core power

 27 kWe electric power

 31.1 kg TOTAL

Performance

 propellant: Hydrogen

 Isp : 694 s

 blowdowns: 400 

 ΔVper burn: 0.0095 km/s

 thrustper burn: 26.39 N

 prop. massper burn: 1.39 kg 

 prop. mass: ~558 kg TOTAL

6U frame

6’

artistic rendering of proposed concept 

6’
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Future Work

Phase I

Further Optimization

concept design

mission design

Simulations

thermal hydraulics

trajectory optimization

Technology Comparison

chemical propulsion

Mission Comparison

Enceladus Orbiter via Decadal 

Survey, etc.

Phase II

Detailed Concept Optimization

Technology Demonstration

energy storage within thermal 

capacitor

energy conversion with closed-loop 

Brayton 

thrust demonstration with propellant & 

nozzle

available experimentation 

hardware
Summary

A reliable interplanetary propulsion 

system based on radioisotopic energy is 

possible

provides versatility to use available 

power for propulsion or electrical power 

production

Extends the realm of CubeSat-based 

exploration and experimentation

Enables a broader range of researchers 

and research institutions

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artistic rendering of concept in Enceladus orbit

The development of a low mass, low cost propulsion system is achievable

smaller launch vehicles  cheaper launch costs

Enables ‘Public Access’ To Outer Planet Exploration!!
Acknowledgments

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Funding Provided By

NIAC Phase I Award  Fiscal Year 2014

Co-Investigators

Troy Howe & Adarsh Rajguru

Steven Howe  Advisor 

NIAC Group  Thank you for program support

artistic rendering of 

proposed concept 
THANK YOU!!

Questions??
References

O’Brien R. C., Ambrosi R. M., Bannister, N. P., et al., Spark Plasma Sintering of simulated radioisotope materials within tungsten cermets, Journal of Nuclear 

Materials, 2009, 393(1), 108-113.

O’Brien R. C., Ambrosi R. M., Bannister, N. P., et al., Safe radioisotope thermoelectric generators and heat sources for space applications, Journal of Nuclear 

Materials, 2008, 377(3), 506-521.

Mattarolo, G., “Development and Modeling of a Thermophotovoltaic System,” Thesis, Electrical Engineering and Computer Science Dept., University of 

Kassel. Kassel, Germany (2007).

Carl M. Stoots. “Emissivity Tuned Emitter For RTPV Power Sources.” Nuclear and Emerging Technologies For Space,The Woodlands, 

TX,03/21/2012,03/23/2012. (2012).

Larson W. J., Wertz J.R., Space Mission Analysis and Design, 3rd Edition, Microcosm Press, 2005. 

Curtis H. D., Orbital Mechanics for Engineering Students, Elsevier Aerospace Engineering Series, 2005, pg 268-273.

Gaskell, D. R. Intro. to the Thermodynamics of Materials, 4th Ed. (2003) 587.

Kelley, K. K. “The Specific Heats at Low Temperatures Of Crystalline Boric Oxide, Boron Carbide And Silicon Carbide”. Journal of the American Chemical 

Society. 63 (1941) 1137-9.

Kantor, K., P. B. Krasovitskaya, R. M. Kisil, O. M. Fiz. “Determining The Enthalpy And Specific Heat Of Beryllium In The Range 600-2200” Phys. Metals 

and Metallog. 10 (6) (1960) 42-4. Mcl-905/1, Ad-261792.

Booker, J. Paine, R. M. Stonehouse, A. J. Wright. “Investigation Of Intermetallic Compounds For Very High Temperature Applications”. Air Development 

Division (1961) 1-133. Wadd Tr 60-889, Ad 265625.

Pankratz, L. B. K. K. Kelley. Thermodynamic Data for Magnesium Oxide U S Bur Mines. Report. 1-5 (1963); Bm-Ri-6295. 

Kandyba, K., V. V. Kantor, P. B. Krasovitskaya, R. M. Fomichev, E. N. Dokl “Determination Of Enthalpy And Thermal Capacity Of Beryllium Oxide In The 

Temperature Range From 1200 – 2820” Aec-Tr-4310. (1960) 1-4.

Hedge, J. C., J. W. Kopec, C. Kostenko, J. I. Lang. Thermal Properties Of Refractory Alloys. Aeronautical Systems Division. (1963) 1-128; ( Asd-Tdr-63-597, 

Ad 424375 ) 

X-123CdTe (X-Ray & Gamma-Ray Detector System) – http://www.amptek.com/x123cdte.html 

Argus Infrared Spectrometer – http://www.thoth.ca/spectrometers.htm 

NanoCam C1U (High Resolution Camera) – http://gomspace.com/index.php?p=products-c1u 

Low Voltage Gated Electrostatic Mass Spectrometer (LVGEMS) – http://www.techbriefs.com/component/content/article/16137 

http://www.universetoday.com/106288/indias-mars-orbiter-mission-mom-requires-extra-thruster-firing-after-premature-engine-shutdown/ 

http://trajbrowser.arc.nasa.gov/ 
Appendix
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Concept

core material

Thermal Capacitor Qualifications

high thermal storage, high thermal conductivity, high density, 

high melting temperature

Silicon identified as a 

suitable material

Energy Storage Potential: 

ΔHfusion =1.8 MJ/kg (1700 K)

Allows for an 

operational temperature 

of 1700 K

Specific Heat verse temperature of possible core materials
Mission Design

communication

link budget between spacecraft and DSN

Item Symbol Units Value

Frequency f GHz 27.5

Transmitter Power Pt Watts 2500

Transmitter Line Loss Lt dB -1.10

Transmit Antenna Diameter Dt m 1.5

Transmit Antenna Gain (net) Gt dBi 50.16

Propagation Path Length S km 1.27(10)9

Space Loss Ls dB -303.30

Propagation & Polarization Loss La dB -0.06

Receive Antenna Diameter Dr m 34.00

Receive Antenna Pointing Loss Lpr dB -6.87

Receive Antenna Gain Gr dBi 77.22

System Noise Temperature Ts K 84.10

Data Rate R Mbps 1 

SNR Eb/No - 5.94

table describing communication budget

*designed for 9.47 AU distance
Communication

frequency bands

L-Band

S-Band

X-Band

AM Radio

CB Radio

FM Radio

TV Stations

Garage Openers

Cell Phones

Infrared 

Visible Light

Ultraviolet

X-Rays

kHz

MHz

G

Gamma Rays

Hz

THz

Frequency [Hz]

Frequency 

Band

Frequency Range

Uplink Downlink

GHz GHz

UHF 0.2 - 0.45 0.2 - 0.45

L 1.635 - 1.66 1.535 - 1.56

S 2.65 - 2.69 2.5 - 2.54

C 5.9 - 6.4 3.7 - 4.2

X 7.9 - 8.4 7.25 - 7.75

Ku 14.0 - 14.5 12.5 - 12.75

Ka 27.5 - 31 17.7 - 19.7

SHF / EHF 43.5 - 45.5 19.7 - 20.7

Spacecraft

DSN HEA
Communication 

link budget design
Mission Design

instrument package

proposed instrument package for Enceladus

Instrument TRL Mass (kg) Volume 

(cm3)

Peak Loading 

Power (W)

X-123CdTe (X-Ray & Gamma-Ray 

Detector System) 7 0.18 175 2.5

Argus Infrared Spectrometer 9 0.23 180 -

NanoCam C1U (High Resolution 

Camera) 8 0.166 501 0.66

Low Voltage Gated Electrostatic Mass

Spectrometer (LVGEMS) 7 0.25 32 0.5

table describing possible instrument package

*designed for about a 6U CubeSat payload
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Trajectory

Electric Propulsion

4 ion thrusters  redundancy

Xenon propellant 

allows for shorter transit times

Earth – Jupiter - Saturn

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Earth Departure: Jan-18-2018

Saturn Arrival: July-11-2023

Total Duration: 5.48 years

Launch Vehicle Injection C3 required: 80.7 km^2/s^2

Declination of the launching asymptote: 4 degrees. 

Injection delta V(Earth escape delta V requirement):6.42 

km/s

Jupiter Flyby: Jan-08-2020 

Gravity Assist: 123 m/s (0.12 km/s) 

Post injection delta V (Saturn capture): 0.54 km/s

Total delta V required by propulsion system = 6.96 km/s 

possible heliocentric trajectory to Saturn