Advanced Space Power Systems-Teofilo-2008

8/9/2019 Advanced Space Power Systems-Teofilo-2008

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Slide 1

SPACE POWER SYSTEMS FOR THE 21ST CENTURY

Vincent L. Teofilo, Ph.D.

[email protected]

(408) 743-2275

Lockheed Martin Space Systems Company


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Outline

Satellite Power Systems

Solar Array-Battery Power

Radioisotope Power

Space Nuclear Reactors

Advanced Concepts

Fusion

Matter-Anti-Matter

Vacuum Energy

Warp-Drive Worm Holes


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Corona [1959-70]

10,000 kg of AgO/Zn batteries


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TELEVISION INFRARED OBSERVATION SATELLITE

AMSUA1

AVHRR

ESA

HIRS

IMP

IMU

REA

SAD

SAR ANTENNAS

THERMAL CONTROLPINWHEEL LOUVERS

SOA

VRA

UDA

SBA

AMSU-B

1100 W SOLAR ARRAY

40 Ah BATTERY MODULESTED

SBUV

AMSU-A2


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Hubble Space Telescope

90 Ah NiH2

GaAs/Ge


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Mars Global Surveyor


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Mars Climate Orbiter

Characteristic Requirement Capability

Mission Life 5 Years1 Yr Cruise+AB2 Years Mapping

2 Years Relay

7 Years +

Orbital AvePwr

300 W @ MarsPerihelion

350 W (1600W @ 1AM0Beginning of Life)

EnergyStorage

~ 10,000 DODCycles (60% MaxDOD @ AB End-Game)

(1) 16 A-Hr NiH2 Battery(70% Max DODCapability)

Bus VoltageRange

22 36Vdc @User Load I/F

24 36 Vdc @ User LoadI/F

Redundancy No MissionCritical SinglePoint Failures(MCSPF)

Single NiH2 Battery hasCredible but Low-RiskMCSPF


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60 Ah NiH2 SPV

Battery

Communications

Section

Bus Section

Gateway

Antenna

1 kW eSolar Array

Main Mission

Antenna Panel

Cross LinkAntenna

Iridium Satellite


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IKONOS Satellite


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Solar Power System

IN-T3052

IN-T3055

IN-T3049

Solar Array Drive

Assemblies

(North & South) Fuse Box AssemblyPower Regulation Unit

Battery (North) Battery (South)

Power to

Payload &

HK Loads

Power to

Arcjets

Power to Pyros &

Earth & Sun Sensor

Assemblies

Solar Array

Load Load

SolarArray

SolarArray

Shunt(Fullor

Partial)

SeriesRegulator

ChargeRegulator

DischargeRegulatoror Diode

Battery Load

SeriesRegulator

= Optional EPS Elements= Energy Source/Storage Elements

LoadLoad LoadLoad

SolarArray

SolarArray

Shunt(Fullor

Partial)

SeriesRegulator

ChargeRegulator

DischargeRegulatoror Diode

Battery Load

SeriesRegulator

= Optional EPS Elements= Energy Source/Storage Elements


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A2100 EPS ComponentAccommodation

DUAL Ku-band GIMBALLED(50 diameter) ANTENNA TRANSPONDER SUPORT STRUTS

ACCESS PANELS

SOLAR ARRAY

SHEAR-TIE STRUTS

NORTH BATTERY MODULES

SOUTH BATTERY MODULES

SOLAR ARRAY

PANELS (NORTH)

SOLAR ARRAYPANELS (SOUTH)

BASE PANELS TRANSITIONSTRUCTURE

TRANSPONDER PANEL(NORTH) includes

PRU & FBA

Ref: A. Salim, IECEC, 2000


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Space Solar Cell Development Projections

3rd generation of Lattice matched multi-junction solar cells willreach practical limit of 40-45%. IMM cells on kapton are beingdeveloped at 50-100 um thickness

Thin film nano-crystaline cells will be introduced for cost andmass efficiency intitiallly at 13% efficiency with potential forachieving 20% in 5-10 years using MJ thin film techniques

10%

15%

20%

25%

30%

35%

40%

45%

1971 1975 1980 1985 1990 1995 2000

Year

Silicon

GaAsMulti-Junction

2005

40-45% in 10-15 years

1st Generation

2nd Generation

3rd Generation

20152010 2020


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Solar Panel Characteristics

PV Cell Technology

BOL

Efficiency W/m2

Cell W/kg

Panel

W/kg1

Thin Film [Triple-Junction (TJ)

amorphous Si on 1 mm Poly] 7.5 93 440 352

Thin Film [CIGS on 1.5 mm Al] 13 169 627 502

High Efficiency Si 17.5 128 182 60

GaAs Triple Junction 28 245 291 88

GaAs XJ (projected~2009) 34 298 353 107

GaAs TJ on Kapton Substrate 28 245 353 174

1 - Panel with interconnects and substrate or support framing

Boeing 100 kW Array (SPW-2006)

ATK 10 kW Array (SPW-2006)LM A2100 Solar Array Wing Assembly


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Thin Film Solar PV

Evaporation of elements simultaneously or in aprescribed sequence,

Sputtering of metals followed by selenization with H2Se, Reactive sputtering of metals with Se vapor,

Printing of metals from ink precursors followed by selenization[requires no vacuum]


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HIGH-EFFICIENCY III-V Thin MJ CELLS(SpectroLab)

thin-cell coupon on a 95-mm radius cylinder.

Ref: D. Law et al., Lightweight, Flexible, High-efficiency Iii-vMultijunction Cells, WCPEC 2006

Triple-junction structures were deposited by metalorganic vapor phase epitaxy (MOVPE) in 4 wafer


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PRIMARYREFLECTOR

32.4 m

PCM

THERMIONIC

CONVERTER

THERMALRECEIVER

SECONDARY

CONCENTRATOR

THERMIONICCONVERTER

THERMAL STORAGE

PRIMARYREFLECTOR

32.4 m

PCM

THERMIONIC

CONVERTER

THERMALRECEIVER

SECONDARY

CONCENTRATOR

THERMIONICCONVERTER

THERMAL STORAGE

Solar Thermal Thermionic Power System

Able to Operate in Van Allen Belts

Volume ~1/2 of SA/Batt Power

Mass ~1/3 of SA/ Batt Power at 75 W/m2

Using a new diffracting concentrator lenswhich weighs 0.5 kg/m2

BeO Phase Change Material used to heat TIelements to provide power during Eclipse

Satellite applications at > 30 kWe

NASA MSFC Ground Demo [Clark- STAIF 2006]

MSFC Ground Test


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Space Battery Cell Comparisons

0

100

200

300

400

500

0 50 100 150 200 250 300 350

Li-SPE

Li-Ion

AgO/ZnSecondary

Ni/H2

Ni/Cd

Pb/PbO

Ni/MH

ENERGYDENSITY

(Wh/l)

SPECIFIC ENERGY (Wh/kg)

AgO/Zn Primary

LiSO2

LiMnO2

400

Li(CF)x

LiSOCl 2

Advanced Li Polymer

Current Li Ion batteries using liquid electrolytes provide 100-125- W/kg Future Li Ion batttery cells with solid electrolytes and nano-structured electrodes will provide

> 200 Wh/kg batteries in 10 years


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Potential Spacecraft Energy Sources

2.4 x 10108cal/s-cc

ZPF


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Range of Energy Source Applications


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Solar Array-Battery Systems vs. Solar Array -Fuel Cell Hybrid System

Fuel Cells are not competitive with eletrochemical energy storage due tolower charge-discharge efficiency

15

20

25

30

35

40

45

50

200 400 600 800 1000

Peak Power (W)

Mass(kg

Fuel Cell Hybrid

Batteries - 1 day

Batteries - 1 week

Batteries - 1 month

Batteries - 6 months

Batteries - 1 year

[Teofilo- IECEC 2006]


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>40 RTGs flown for space science missionsfrom 3 W to 300 W

Design for Launch Safety against Pu238

release is major cost driver

Uses 20 General Purpose Heat Source AssembliesTo generate 290 W

Currently using of Russian Pu238 but in future

INL to produce Pu

238

Radioisotope Power

GPHS RTG

56 kg20 GPHS


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Planetary Science Mission S/C EPS


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NASA Missions That Have Used RTGs

NIMBUS B-1 1968 (Aborted) -- -- --III 1969 SNAP 19 (1) ~28 PbTe

APOLLO 11 1969 Heater Units -- --

12 1969 SNAP 27 (1) ~73 PbTe13 1970 (Aborted) -- -- --14 1971 SNAP 27 (1) ~73 PbTe15 1971 16 1972 17 1972

PIONEER 10 1972 SNAP 19 (4) ~40 PbTe/TAGS11 1973

VIKING 1 1975 SNAP 19 (2) ~35 PbTe/TAGS2 1975

VOYAGER 1 1977 MHW (3) ~150 SiGe

2 1977 GALILEO 1989 GPHS-RTG (2) ~285 SiGeULYSSES 1990 GPHS-RTG (2) ~285 SiGePATHFINDER 1996 Heater Units -- --CASSINI 1997 GPHS-RTG (3) ~285 SiGe

Power Level ThermoelectricsMissions Launch Year Type of RTG Per Unit (We) Used


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Radioisotope MissionsUsed safely on 24 missions since 1961

8 RTG Earth Orbit missions (Transit, Nimbus, LES) 7 RTG planetary missions (Pioneer, Voyager, Ulysses, Galileo, Cassini) 5 RTG moon missions (Apollo ALSEP) 2 RTG Mars missions (Viking 1&2) RHUs used on Apollo 11, Mars Pathfinder & MERs among others

MER (2003)


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Voyager Spacecraft


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Galileo Spacecraft


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Advanced Stirling RadioisotopeGenerator

DESIGN FEATURES

Power: 112 We (BOM)94 We (14 yrs)

Mass: 20 kg

System Efficiency: 30%

Dimensions: 88.9cm (length

26.7 (tip-tip)

Voltage: 28 + 0.2 Vdc

Fuel: 2 GPHS modules

Re-Programed: April 2006 to utilizeAdvanced Stirling


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Radioisotope Thermophotvoltaic

NASA Funded DevelopmentCreare Tested 100 We Engineering Test unitUses two GPHS modules producing

500WtDemonstrated Integrated System

Efficiency 17%Design Mass of 7.0 kg with Radiators for

specific energy of 14 W/kg


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Space Nuclear Reactors

Coolant In

Coolant Out

Fast Spectrum Reactors (>500keV)With high power density for compactness

External reflectors for mechanical simplicity

LiH Shield for low mass

Reflector

Shield

Fuel andModerator


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Space Nuclear Reactor History

500 We SNAP-10A 1965

SOVIET SPACENuclear Initiative

>30 reactors flownusing TE and TI(2)

1969-89 1.5 to 5 kWe

1st Space Nuclear Initiative1950-74

SP-100 Technolgy developed in 2nd Space Nuclear Initiative 1983-95

5 kWe TOPAZ


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Nuclear Thermal Propulsion

1 GWt nuclear reactor heats hydrogen to 2200-3000K

Generates 15,000 to 75,000 lb thrust in burns up to an hour

Can be configured to also provide steady state electrical power of

50-100kWe Bimodal or Trimodal designs

20 Reactor/Rockets designed, built and tested 1959-72 in Rover/Nerva Programfor ~$1.4B before termination in 1974 to fund Space Shuttle development


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STAR-C Thermionic Reactor

Solid Graphite Core with Thermionicelements Heated to 1800 K and rejecting heatto heat pipe radiator at 1000K 9000-1000 K

40 kWe unit ~1100 kg with PC and Radiator

Conversion Efficiency of 16% withTH = 1800K and TC= 1000 K

Develop High Efficiency Nano TechnogyTI Convertors with Lower TH and TC

STAR WARS STIMULATED SMALL REACTOR CONCEPTS


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Gas Cooled Reactor

Development Issue: He-Xe gas coolant replenishment over life


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Heat Pipe Reactor

Fission power is generated in uranium fuel-pins. The power is conducted toheatpipes which transfer the power to an ex-core power conversion system(orintermediate HX).

Heatpipes provide the efficiency of 2-phase liquid-metal heat transfer in apassive, simple, well-characterized volume.

HX can interface with any conversion system or heat pipes can go directly to

thermo-electric or thermophotvoltaic power conversion system.

100 kWe Reactor/Shield Mass ~ 1000 kg

Combined with Stirling engine yields specific power of 50 W/kg


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Jupiter Icy Moons Orbiter Space Craft

2003 Notional Concept (JPL)

2001 A Space Odyssey (Universal Studios)


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Proposed JIMO Configuration

MISSION/

AVIONICS

MODULE

HIGH GAIN ANTENNA

3.00 METER

XENON TANK

11314 Kg

SP100 STYLE

130KWe REACTOR

DEPLOYABLEFLAT RADIATOR

90 M2

ATLAS HLV

5 M FAIRING

DEPLOYING BOOM

PMAD

4 BOXES

TE CONV

12 PLCS

NEXT THRUSTER

36 PLCS (18/SIDE)

STOWED RADIATOR

2 PLCS


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S b d S l P (SBSP)


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Space-based Solar Power (SBSP)

Renewed interest in an old idea giventodays perspectives

Peak oil, global environmentalconcerns, and growing world-wideenergy demand

Attractive option within future energyportfolio

Very few clean, safe, inexhaustible,reliable, and affordable alternatives

Significant changes in recent yearsenabling improved

SBSP economics Technology advancements, emerging

applications, and market pricingrealities

CPV selected for high efficiency (>40%)But impact of radiator for cooling cellsprovides net W/kg


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Progress in Magnetic Confinement Fusion

DT Terrestrial Controlled Fusion requires (1) periodic vacuumconfinement and thermal blanket structural Materialreplacement and (2) breeding sufficient tritium to replace thatconsumed. This makes it uneconomical for commercial powergeneration. Advanced fuels (d,3) and (p,B11) would eliminateSuch requirements but require much higher confinementconditions obtainable by magnetic fields > 30 Tesla.

Inertial Electrostatic Confinement IEC


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Inertial Electrostatic Confinement IEC

University Of Wisconsin IEC Experimental FacilitySee: http://fti.neep.wisc.edu/publist?which=fdm50

Fusion Reactions vs Grid Bias

Magnetically Channeled Spherical IEC propulsion Experiment


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Inertial Electrostatic Confinement Space Thruster

Plasma Jet Grid

Xe Propellant Estimated Performance

Thrust = 34 mN, Isp = 3000 s, Accelerating = 500 W, t ~

62-68% , Voltage = 600V

p-B11 Propellant Estimated Performance

Thrust= > 1 N , Isp >104s , V = 150kV , tt > 1 N> 1 N

Miley et al, Technolgy of Fusion Energy, 2008

Magneto-Inertial Fusion (MIF)


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Magneto-Inertial Fusion (MIF)

plasma density intermediate betweenconventional MFE and ICF

Avoid huge, steady-state magnets ofconventional MFE,

Avoid the problem of extremely highpulsed power (1000s TW) Compactreactor

Offers potentially a low-cost R&D path

Lawsons criteria excessenergy, n > 1014 s.m-3

Material Liner

Solid

Liquid

Gaseous

Target Plasma

Magnetic fieldused to insulate

the target plasma

from the liner

Magneto-Kinetic Compression MIF


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Magneto Kinetic Compression MIF

Flowing Liquid Metal Heat Exchanger/ Breeder

~ 20 m

BURN CHAMBER

(Dch ~ 5 cm)

10-30 m

1 m

Magnetic

Expansion ChamberAccelerator Source

Current experiment to createinitial FRC plasmoid forfusion breakeven experiment

Advantages over ITER tokamak Minimum B field at highest plasma pressure (~1)

Simple linear system reactor wall is a steel pipe

Variable output power ~ 10-100 MW not multi-GW

Burn chamber well separated from plasmoidformation/heating.

Direct electric power conversion with expansion offusion heated plasmoid (Brayton cycle - > 90%)

Low mass system directly applicable to spacepropulsion

Key physics and scaling have been demonstrated

Developmental cost orders of magnitude less -

Proof-of-Principle experiment ~ 3 M$ / year

Energy required to achievefusion conditions is transferredto FRC plasmoid from array ofaxially sequenced coils.

Interplanetary Fusion Propulsion


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Interplanetary Fusion Propulsion

Reactions: d + t = He(3.52 Mev) + n (14 MeV)d + He3 = He4 + p [ 18.3 MeV]

Lawson Criteria: n > 1014 sec exceeded throughsteady state and pulsed magnetic confinement

Effective Exhaust Velocity, Ve (m/s)

Chem

Fission

Fusion

0.1 (kW/kg)

1.0

10.0

100.0

103

104

105

106

101

102

103

2000 2010 2020 2030 2040

PHYSICS DEMO

Tech Development

Ground DEMO

FLIGHT DEMO

MANNED MISSION

Interstellar Anti-matter Propulsion


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Interstellar Anti matter Propulsion

Fuel Mass for round-trip to Alpha Centauri at 0.1 c: 0.01 Msc / 0.1 (conversion eff.) = 0.1 Msc

For 100 ton Space Ship 10 tons of anti-protons cost $1021

~ 3 ng of p- produced /yr in 2002

Dark Energy and Matter


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Dark Energy and Matter

General Relativity Equation: r/r = 4G ( + 3p)/3Normal matter creates gravitational attraction and slows down cosmological expansion. Dark matter alsocauses attraction. For vacuum the situation is opposite: positive vacuum energy anti-gravitates, i.e.creates gravitational repulsion and its density is constant.

QUANTUM VACUUM ENERGY FIELD


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Q

Spectral Energy Density Planck ( 1911)

(, ) = 82 [h/( h/ 1) + h/2 ]c3

Heisenberg Uncertainty Principle: x v < hMin Energy at T = 0 : h/2 for each EM harmonic mode

up to Planck Frequency p = 1.9 x 1043 Hz

p = 22c /G2h = 10109 J/cc Einstein, Nernst study zero-point energy (1913-16)

Casimir identifies Quantum Zero point Field (ZPF) assource of force between parallel plates (1948)

Thought experiment by Forward shows principle oftapping zero-point energy (1984).

USAF Study identified ZPF experiments for furtherstudy (1996).

Casimir Force definitively measured by Lamoreaux(1997) and to 1% by Mohideen (2003)

Newtons Laws derived from Maxwells Equations byHaisch, Rueda and Puthoff (Phys. Rev 1994) and(Anal.Physik- 2005)

Casimir Force

Quantum Vacuum Plasma Thruster (QVPT)?


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( )

The local quantum vacuum density is related to local matter density?

- [ H. White STAIF 2007]

A QVPT, in principle, can be likened to a conventional plasma thruster that

uses crossed E and B fields to induce a plasma drift in the propellant whichfor the vacuum field are p+-p- and e+-e- pairs.

The difference arises in the fact that a QVPT uses quantum vacuumfluctuations as the fuel source mitigating the need to carry propellant.

This suggests much higher ISP is available for QVPT systems limited only bysupply power storage densities.

Shawyer EM Drive for Chinese QVPT

vaclocalm

vac

localm

vaclocalvac

_

_

_ ==

Energy for Warp Drive


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By creating the energy equivalent of negative mass the space in front of aspace ship can be contracted and that behind expanded with energy:

E = -3.5 x 1030 kg x vs/c

or x c2 = ( 3 x 106m/s)2

E = -3.15 x 1042 Joules

for each factor of the speed of light

Vacuum energy of p = 10109 J/cc may be more then enough to provide themass equivalence energy to Warp Space.

However, current vacuum energy extraction concepts may be limited to 50 W/kg

Advanced Radioisotope Power Systems will be needed for Lunarexploration using Stirling or TPV energy conversion

Nuclear reactors may be need for manned Lunar or Mars bases after ~2020

Nuclear Thermal Propulsion development interrupted in 1974 may becontinued for more efficient manned space transportation

Fusion power/propulsion may supplant nuclear fission due to greater

safety and implemeted early for replacing HCTs

Worm holes may be the only possible method to explore the universe

Future R&D in Advanced Space Power Sources


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Development of advanced thin film PV and Li Ion batteries

Development of Solar Thermal Thermionic Power system for high

power applications

Development of nuclear reactors and for Lunar Bases and SpaceCraft power/propulsion [2006-2030]

Research and Develop Space Fusion Power/PropulsionTechnology [2006-2040]

Research Physics and Technology for Interstellar Space Travel

[2006- 2???]

See: AIAA Frontiers in Propulsion Science [http://www.aiaa.org/content.cfm?pageid=360&id=1743 ]

Advanced Space Power Systems-Teofilo-2008

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