Radioisotope Propulsion System - American Nuclear …anstd.ans.org/wp-content/uploads/2015/07/5098_Jerred-et...RADIOISOTOPE-BASED PROPULSION SYSTEM ENABLING EXPLORATION WITH SMALL

RADIOISOTOPE-BASED PROPULSION SYSTEM ENABLING EXPLORATION WITH SMALL PAYLOADS

Nuclear and Emerging Technologies for Space 2015 Albuquerque, New Mexico

February 24th, 2015

A Project Overview: NASA Innovative Advanced Concepts (NIAC) Phase I Award FY 2014

Nathan Jerred, Troy Howe & Dr. Steven Howe Center for Space Nuclear Research

Adarsh Rajguru University of Southern California

• 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

Inspiration

1U CubeSat frame

Project GOAL – deliver 6U CubeSat payload to Enceladus orbit (~15 kg) – develop an appropriately sized propulsion system concept – develop the mission architecture

• Target small launch vehicles – less than 1,000 kg to LEO

• Enable affordable deep space exploration

• 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]

O’Brien et al.

Concept energy source

Tungsten-based CERMET

238 PuO2 Pellet

DOE

• Fuel Containment – fuel encapsulated in a tungsten-based matrix

• provides high strength & toughness – can provide great energy density

• Radioisotope-Based Core – decay energy accumulated within central core, i.e. thermal

capacitor

Approach

– 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

• 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 induce

a material’s phase change

Allows for an operational temperature of around 1700 K

Concept energy storage

• Core Material – Silicon identified as a suitable material ΔHfusion = 1.8 MJ/kg kth = 148 W/m-K Tmelt = 1687 K

Radioisotope Heats

Thermal Capacitor

Core Reaches Peak

Temperature

Blowdown of Gas

Through Core

Electrical Conversion

System Produces

Power

Propellant Injected into

Core Produces

Thrust

Impulse-based

Function Performed

Energy Depleted from Core

Regeneration Cycle

• 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 induce

a material’s phase change

Concept energy storage

Thermal Capacitor

Radioisotope Power Source

Insulation

Support Structure

Flow Channels

Qin Component Mass Thermal Budget

RPS [kg] 6.44 Qin [Wt] 1,180 238PuO2 [kg] 3

Qresidual [Wt] 385 Silicon [kg] 15.6

Supports [kg] 2.1 Estored [MJ] 30

Insulation [kg] 108

Qloss [Wt] 795 TOTAL [kg] 132

• Thermal interactions within the subsystem – details such as isotope loading, capacitor size, insulation, etc. – determine heat-up time of the thermal subsystem

– 385 Wt for melting the core – heat-up: 20.5 hrs

Thermal Subsystem Modeling

full thermal subsystem

heat-up time modeling

Silicon & W/PuO2 Core

• Direct Propellant Heating – provides thermal-based propulsion – heated propellant expelled through

nozzle creating thrust • Energy Conversion

– closed loop Brayton cycle – 10’s kWe per day

thermal propulsion flow schematic

dual Brayton engines

Concept dual-mode

electrical conversion flow schematic

• Dual -Brayton Engines – 2 x 12.5 kWe

• 25 kWe per pulse – 6 min pulse per day – ~30 % efficiency

• Mass – Turbine/Compressor: 54.1 kg – Heat Rejection: 16.9 kg – Alternator: 15.9 kg – Housing/Pipping: 10 kg

Total: 97 kg

dual Brayton engines

Conversion Subsystem Brayton Engine

electrical conversion flow schematic

heat rejection subsystem dissipation of thermal energy

• Heat Rejection *able to dissipate thermal energy between pulses – Lithium ideal absorber material (Cp = 3.58 kJ/kg-K & Tmelt = 453 K)

• will need a containment canister *enables a low mass, low footprint thermal rejection system & overall conversion subsystem

Conversion Subsystem Heat Rejection

electrical conversion flow schematic

heat rejection subsystem dissipation of thermal energy

Lithium Absorber

Tin = 796 K

Tout = 325 K

dissipation time: ≈ 21hrs

Enceladus Science Traceability Matrix Science Objectives Science Investigation Instrument Payload

1. Physical conditions at the plume source

Topography & stratigraphy; Thermal output; vent shape; surface strength; surface roughness; subsurface structure of tiger stripes;

cavern size; subsurface lake; particle size distribution and speed; ice temperature

MAC, thermal imaging radiometer, dust analyzer, MS

2. Chemistry of the plume source

Chemical inventory of plume gas and dust species; chemical equilibria; isotopic ratios

MS, dust analyzer

3. Presence of biological activity Organic molecules inventory to high masses MS, dust analyzer

4. Plume dynamics and mass loss rates

Plume structure, ejection rates; particle size vertical structure; particle velocities; time variability (density, particle size, velocity;

composition) MAC, MS, dust analyzer

5. Origin of south-polar surface features

Topography & stratigraphy, temperature distribution of active features

MAC, thermal imaging radiometer

6. Internal structure Static gravity, potential Love numbers, magnetic field Radio science, magnetometer, imaging

7. Presence, physics, and chemistry of the ocean

Potential Love numbers, magnetic induction, plume chemistry Radio science, magnetometer, MS, dust analyzer

8. Tidal dissipation rates and mechanisms

Long-wavelength global thermal emission, bolometric albedos MAC, thermal imaging radiometer

9. Chemical clues to Enceladus’ origin and evolution

Isotopic and elemental analysis of plume gases and dust grains MS, dust analyzer

10. Nature and origin of geological features and geologic

history Geology, topography, stratigraphy MAC, radio science

11. Plasma and neutral clouds Spatial distribution, composition, and time variability of neutral

clouds, correlation with plume activity MS, MAC to monitor plume activity

12. E-ring Variation, composition, and relation to Enceladus activity Dust analyzer, MAC to monitor plume activity and E-ring

structure

13. Modification of the surfaces of Enceladus and the other

satellites

Relative ages, surface texture on meter and centimeter scales, exogenic coatings, exogenic impact and ion environment; molecular

lifetimes Dust analyzer, thermal imaging radiometer, MAC, MS

14. Surfaces and interiors of Rhea, Dione and Tethys

Geology and evolution of surfaces of neighboring satellites, shapes, gravity fields

MAC, magnetometer, radio science, MS

15. Nature of potential landing sites

Topography, surface texture, thermal inertia MAC, thermal imaging, radiometer

Mission Design

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 5’s High Resolution Camera (MAC) 100’s 100’s 1’s Mass Spectrometer 100’s 10’s 1’s

6U limits 15,000 6,000 5-10

instrumentation & communication

6U payload frame

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 3 SNR Eb/No - 5.94

communication system

• 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

• 6,000 hrs operation • can provide ΔVtotal ≈ 5.67 km/s

Mission Design Trajectory

graphic of Earth-based phasing maneuvers

Earth – Venus – Saturn

• Injection ΔV (Earthesc): 4.35 km/s

• Venus Flyby: 0.47 km/s

• Earth1 Flyby: 0.19 km/s

• Earth2 Flyby: 0.50 km/s

• Saturn Insertion ΔV : 1.88 km/s

• Total Propulsion System ΔV Need: 6.23 km/s AeroJet BPT-2000 Hall Effect thruster

potential Saturn trajectory

Saturn Orbit

Earth Orbit

proposed concept

System Performance

6’

Thermal Propulsion (Earthescape)

timeblowdown: 360 s Mi: 1000 kg

timeheatup: ~1 day ΔVescape: 3.2 km/s

timeescape: 310 days temp: 1700 K

Performance: Thermal-Hybrid Scenario

propellant: Hydrogen ΔVburn: 0.016 km/s

Isp: 694 s ΔVtotal: 3.075 km/s

burns: 280 propburn: 1.3 kg

total H2 mass: 363 kg *final orbit period 6.33 days

propellant: Xenon ΔVburn: 0.125 km/s

Isp: 87.8 s ΔVtotal: 0.125 km/s

burns: 1 propburn: 88 kg

total Xe mass: 88 kg

thermal propulsion flow schematic

6U frame

6’

System Size

4.88 m

1.58 m Subsystem Mass [kg] Thermal Propulsion

Propellant (Earthescape) 451.65

Structure 100

Thermal Propulsion Propellant

(Saturn Capture) 100

Electric Propulsion Propellant 50

Instrumentation 15

Communication Subsystem 46.22

Thermal Subsystem & Conversion Subsystem 229.02

Total 991.89

• Star48B ≈ 2100 kg – Upper stage rocket motor

• Enceladus Orbiter ≈ 3600 kg – Decadal Survey Enceladus

Mission

Possible Launch Vehicle • Minotaur –C 3210 XL

– 1,278 kg to LEO (400 km)

NIAC Phase I: FY 2014 • Concept Design • Mission Profile • Modeling

– thermal subsystem – conversion subsystem – trajectory optimization - EMTG

• Technology Comparison – chemical propulsion

• Mission Comparison – Enceladus Orbiter via Decadal

Survey, etc.

Possible Future Work: NIAC Phase II • Detailed Concept Optimization • Alternative communication systems • Modeling

– expand & refine modeling – CFD thermal hydraulic

interactions between gas & core • Demonstration of Enabling

Technologies

Current & Future Work

decadal survey – enceladus orbiter concept

– energy storage within PCM, i.e. thermal capacitor

– energy conversion with open- & closed-loop Brayton cycle

– demonstrate passive heat rejection with absorber

Europa?

• Extends the realm of micro-satellite exploration and experimentation & engages a broader range of researchers

• Concept can be adapted for the mission – low mass power production unit for orbit station keeping – low mass, high power for deep planetary surface drilling & sampling – strictly radioisotope electric propulsion (REP) system – a planetary surface exploration probe using thermal propulsion to maneuver

Summary

• A reliable interplanetary propulsion system based on radioisotopic energy is possible

– provides versatility to use available power as needed

artistic rendering of concept in Enceladus orbit

Acknowledgments & Thanks

• Funding Provided By: NASA NIAC Phase I Award

• NIAC staff & NIAC general counsel • R.C. O’Brien

– fuel encapsulation • J. Breedlove

– turbo-machinery much insight gained through discussions with all

of CSNR staff & many others artistic rendering of proposed

concept

THANK YOU!! Questions??

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).

Goel A., B. Franz, K. J. Schillo, S. Reddy, S. Howe. Design of a Flight Demonstration Experiment for Radioisotope Thermophotovoltaic (RTPV) Power System. Proceedings of Nuclear and Emerging Technologies for Space. Paper #5002. Albuquerque, NM (2015) 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/

References

Appendix

• 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

• 6,000 hrs operation • can provide ΔVtotal ≈ 5.67 km/s

Mission Design Trajectory

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 (Earthesc): 6.42 km/s

• Jupiter Gravity Assist: 0.12 km/s

• Post Injection ΔV (Saturncap): 0.54 km/s

• Total ΔV required by propulsion system = 6.96 km/s possible trajectory to Saturn AeroJet BPT-2000 Hall

Effect thruster

• 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 induce

a material’s phase change

Allows for an operational temperature of around 1700 K

Concept energy storage

• Core Material – Silicon identified as a suitable material ΔHfusion = 1.8 MJ/kg kth = 148 W/m-K Tmelt = 1687 K

Sensible Heat

Radioisotope Heats

Thermal Capacitor

Core Reaches Peak

Temperature

Blowdown of Gas

Through Core

Electrical Conversion

System Produces

Power

Propellant Injected into

Core Produces

Thrust

Impulse-based

Function Performed

Energy Depleted from Core

Regeneration Cycle

Thermal Subsystem Design

Silicon Core *15.6 kg

Heat Source *6.44 kg Flow Channels

Cannister

Primary Insulation ~ZrO2

*108 kg

Secondary Insulation ~Carbon-Aerogel

Support Structure ~Tantalum Rods

*2.1 kg

Shell

Total Subsystem Mass: 132 kg

Thermal Subsystem Design

Qloss

Qin

PuO2 3 kg ≈ 0.392 W/g Qin = 1.18 kWt Qloss = 795 Wt

Total Subsystem Mass: 132 kg

Alternative Isotopes Qloss

Qin

238PuO2 System T1/2 = 87 yr Specific Power: 0.392 W/g Isotope Mass: 3 kg Encapsulation Mass: 3.44 kg Silicon Mass: 15.6 kg TOTAL: 22 kg

Qin = 1.18 kWt

244Cm2O3 System T1/2 = 18.1 yr Specific Power: 2.269 W/g Isotope Mass: 0.52 kg Encapsulation Mass: 0.72 kg Silicon Mass: 15.6 kg TOTAL: 16.8 kg *will need increased shielding?

241AmO2 System T1/2 = 433 yr Specific Power: 0.094 W/g Isotope Mass: 12.5 kg Encapsulation Mass: 17.3 kg Silicon Mass: 15.6 kg TOTAL: 45.4 kg

Mission Design instrument package

proposed instrument package for Enceladus

*designed for about a 6U CubeSat payload

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

Earth Escape options

• Thermal Mode – Hyrogen propellant 100% – long final elliptical orbit?

• Thermal-Hybrid Mode – large final ΔV to quickly escape – heavy gas: Xenon

– increases prop. mass – small solid rocket motor

– increases mass < 100 kg? • Chemical

– sized to achieve Earth escape in a single burn – solid rocket kick-motor ATK Star 30 BP [500 kg] – replaces hydrogen tank smaller volume

Summary

• The development of a low mass, low cost propulsion system is achievable – smaller launch vehicles cheaper launch costs

• 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

artistic rendering of concept in Enceladus orbit

Enables ‘Public Access’ To Outer Planet Exploration!!