501329main TA02 InSpaceProp DRAFT Nov2010 A

8/11/2019 501329main TA02 InSpaceProp DRAFT Nov2010 A

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National Aeronautics and Space Administration

November 2010

DRAFT I n -S pAce p RopulSIon S ySTemS R oADmApT A 02

Mike Meyer,Co-chair Les Johnson,Co-chair Bryan PalaszewskiDan Goebel

Harold WhiteDavid Coote


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Table of Contents

FE S TA02-11. G O TA02-21.1. Technical Approach TA02-21.2. Benets TA02-51.3. Applicability/Traceability to NASA Strategic Goals, AMPM, DRMs, DRAs TA02-51.4. Top Technical Challenges TA02-5

2. D P D TA02-62.1. Chemical Propulsion TA02-72.2. Nonchemical Propulsion TA02-102.3. Advanced Propulsion Technologies TA02-162.4. Supporting Technologies TA02-19

3. I O T A TA02-204. P B O N N TA02-20 A TA02-21R M TA02-22 A TA02-22


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FOREWORDNASAs integrated technology roadmap, including both technology pull and technology push strategies,considers a wide range of pathways to advance the nations current capabilities. Te present state of this effortis documented in NASAs DRAF Space echnology Roadmap, an integrated set of fourteen technologyarea roadmaps, recommending the overall technology investment strategy and prioritization of NASAs spacetechnology activities. Tis document presents the DRAF echnology Area 02 input: In-Space PropulsionSystems. NASA developed this DRAF Space echnology Roadmap for use by the National Research Council(NRC) as an initial point of departure. Trough an open process of community engagement, the NRC willgather input, integrate it within the Space echnology Roadmap and provide NASA with recommendationson potential future technology investments. Because it is difficult to predict the wide range of future advancespossible in these areas, NASA plans updates to its integrated technology roadmap on a regular basis.


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safely (mission enabling), getting there quickly(reduced transit times), getting a lot of mass there(increased payload mass), and getting there cheap-ly (lower cost). Te simple act of getting thererequires the employment of an in-space propul-sion system, and the other metrics are modiers

to this fundamental action.Development of technologies within this A will result in technical solutions with improve-ments in thrust levels, Isp, power, specic mass(or specic power), volume, system mass, systemcomplexity, operational complexity, commonali-ty with other spacecraft systems, manufacturabil-ity, durability, and of course, cost. Tese types ofimprovements will yield decreased transit times,increased payload mass, safer spacecraft, and de-creased costs. In some instances, developmentof technologies within this A will result in mis-sion-enabling breakthroughs that will revolution-ize space exploration. Tere is no single propul-sion technology that will benet all missions ormission types. Te requirements for in-space pro-pulsion vary widely due according to their intend-ed application. Te technologies described herein

will support everything from small satellites androbotic deep space exploration to space stationsand human missions to Mars.

Figure 1 is a graphical representation of the In-Space Propulsion echnology Area BreakdownStructure ( ABS). Te ABS is divided into fourbasic groups: (1) Chemical Propulsion, (2) Non-chemical Propulsion, (3) Advanced Propulsionechnologies, and (4) Supporting echnologies,based on the physics of the propulsion system andhow it derives thrust as well as its technical matu-rity. Tere may be credible meritous in-space pro-pulsion concepts not foreseen or captured in thisdocument that may be shown to be benecial tofuture mission applications. Care should be taken

when implementing future investment strategiesto provide a conduit through which these con-cepts can be competitively engaged to encouragecontinued innovation.

Figure 2 is the roadmap for the development ofadvanced in-space propulsion technologies show-ing their traceability to potential future missions.Te roadmap makes use of the following set ofdenitions and ground rules. Te term missionpull denes a technology or a performance char-acteristic necessary to meet a planned NASA mis-sion requirement. Any other relationship betweena technology and a mission (an alternate propul-sion system, for example) is categorized as tech-nology push. Also, a distinction is drawn be-

EXECUTIVE SUMMARYIn-space propulsion begins where the launch ve-

hicle upper stage leaves off, performing the func-tions of primary propulsion, reaction control,station keeping, precision pointing, and orbit-al maneuvering. Te main engines used in space

provide the primary propulsive force for orbittransfer, planetary trajectories and extra planetarylanding and ascent. Te reaction control and or-bital maneuvering systems provide the propulsiveforce for orbit maintenance, position control, sta-tion keeping, and spacecraft attitude control. Advanced in-space propulsion technologies will

enable much more effective exploration of ourSolar System and will permit mission designersto plan missions to y anytime, anywhere, andcomplete a host of science objectives at the desti-nations with greater reliability and safety. Witha wide range of possible missions and candidatepropulsion technologies, the question of whichtechnologies are best for future missions is a dif-cult one. A portfolio of propulsion technologiesshould be developed to provide optimum solu-tions for a diverse set of missions and destinations. A large fraction of the rocket engines in use to-

day are chemical rockets; that is, they obtain theenergy needed to generate thrust by chemical re-actions to create a hot gas that is expanded to pro-duce thrust. A signicant limitation of chemicalpropulsion is that it has a relatively low specif-ic impulse (Isp, thrust per mass ow rate of pro-pellant). A signicant improvement (>30%) inIsp can be obtained by using cryogenic propel-lants, such as liquid oxygen and liquid hydrogen,for example. Historically, these propellants havenot been applied beyond upper stages. Further-more, numerous concepts for advanced propul-sion technologies, such as electric propulsion, arecommonly used for station keeping on commer-cial communications satellites and for prime pro-pulsion on some scientic missions because theyhave signicantly higher values of specic im-pulse. However, they generally have very smallvalues of thrust and therefore must be operated

for long durations to provide the total impulse re-quired by a mission. Several of these technologiesoffer performance that is signicantly better thanthat achievable with chemical propulsion. Tisroadmap describes the portfolio of in-space pro-pulsion technologies that could meet future spacescience and exploration needs.

In-space propulsion represents technologies thatcan signicantly improve a number of critical met-rics. Space exploration is about getting somewhere


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tween an in-space demonstration of a technologyversus an in-space validation. A space demonstra-tion refers to the spaceight of a scaled version ofa particular technology or of a critical technolo-gy subsystem; a space validation would serve as aqualication ight for future mission implemen-tation. A successful validation ight would not re-quire any additional space testing of a particulartechnology before it can be adopted for a scienceor exploration mission. Te graphical roadmapprovides suggested technology pursuits within thefour basic categories, and ties these efforts to theportfolio of known and potential future NASA/non-NASA missions.

1. GENERAL OVERVIEW

1.1. Technical ApproachFor both human and robotic exploration, tra-

versing the solar system is a struggle against timeand distance. Te most distant planets are 4.56billion kilometers from the Sun and to reach themin any reasonable time requires much more capa-ble propulsion systems than conventional chemi-cal rockets. Rapid inner solar system missions with

exible launch dates are difficult, requiring pro-pulsion systems that are beyond today's currentstate of the art. Te logistics, and therefore the to-tal system mass required to support sustained hu-man exploration beyond Earth to destinationssuch as the Moon, Mars or Near Earth Objects,are daunting unless more efficient in-space pro-pulsion technologies are developed and elded. With the exception of electric propulsion sys-

tems used for commercial communications satel-lite orbit positioning and station-keeping, and ahandful of lunar and deep space science missions,all of the rocket engines in use today are chemi-cal rockets; that is, they obtain the energy neededto generate thrust by combining reactive chemi-cals to create a hot gas that is expanded to producethrust. A signicant limitation of chemical pro-pulsion is that it has a relatively low specic im-pulse (thrust per unit of mass ow rate of propel-lant). Numerous concepts for advanced in-spacepropulsion technologies have been developed overthe past 50 years. While generally providing sig-nicantly higher specic impulse compared tochemical engines, they typically generate muchlower values of thrust. Trust to weight ratios

Figure 1. In-Space Propulsion Technology Area Breakdown Structure


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Figure 2: In Space Propulsion Technology Area Strategic Roadmap (TASR)

Technology Roadmap: In Space PropulsionESMD

SOMD

SMD

Non-NASA Users

FlightDemonstration

Chemical Propulsion:Liquid & Gelled

Solid & Hybrid Cold/Warm Gas & Micro

Non-Chemical Propulsion:

Low-TRL AdvancedPropulsion:

ONGOING RESEARCH

Supporting Technologies

Mission Implementation Flight Demo TRL 5/6 Significant Demo (TRL 1,000 m2

Solar SailOrbital Demo

Space DemoRefuelingfrom tanker >2,500 m 2

Solar Sail Space Demo

MomentumExchangeTether BoostStation Demo

CRYOSTAT

InterstellarProbe

CrewedOrbitMars

MarsOrbitMission

CrewedMars

Surface

MarsCargo

CrewedNEO

NEO MissionCrewedGEO

TSSMFlagship

LWS(L1 Diamond)

xPRMNEO

Cargo transportin LEO

Cargo transportbeyond GEO

Transfer Missions toGEO and beyond

ISSPropellantlessReboost

COMETMission

GeoStorm(with NOAA)Evolved CommonUpper StageDARPA

LOKI Mission

Common Upper Stage Engine

Hybrid Nozzle& TVC Eval

Adv. CompositeNozzle Demo

LOX/LCH4Engine Components

LOX/LCH4 Comb.Instability Test

LO2/LCH42,000 lbf decent engine

LO2/LCH410,000 lbf decent engine

LO2/LCH4100,000 lbf decent engineVariable Thrust/

Multi-Burn Motor

Deep ThrottlingCryo Engine

High Energy DensityGreen Propellants

Colloid Electrospray

MPD Demo

VASIMR VF-500 VASIMR VF-1000

NTR 25klbf Engine test

Advanced Fission/Fusion Research

< 1 g/m 2 SystemLevel Sail Areal Density

NTR Fuel Development

NTR FuelDownselection

Tether/PayloadCatch Mechanism

High Specific StrengthConductive Tethers

Tape TetherDeployer

SynthesizeMetallicHydrogen

ISHMEP PPUCFM & XFER Adv Heat Rejection

BeamedEnergy Orbital(microsat)

BeamedEnergySuborbital

FusionThrustObserved

FusionGround50 MW

FusionGroundBreakeven

Engineering Breadboardto produce desired effect

Investigate if effects of interest can be observed

Solar Thermal: Isp~1200 s,85% EfficiencySolar Thermal:Isp~800 s,

70% EfficiencyNTR Fuel,Coating Experiments

ColloidMicroPropulsion

< 10 g/m 2 SystemLevel Sail Areal Density

NEXT IonThruster

Miniature Ion Thruster

50 kW HallThruster

VASIMRVF-200

SEPStage100 kW Hall

Thruster MEMS 10 WElectrospray

MEMS 100 WElectrospray

Ionic RCS

Multiple TetheredSatellites for

Ionospheric StudiesGrace II

DARPA

HEOMission

2000 kmBridgeto Space

Drag-freeEarth-observationLEO-to-GEO

ReusableBoost Station

ExpolanetFinder

ExpolanetFinder

PS00326Oct1

1 MWPIT

70% Eff. PIT

1 MWLi MPD100,000 hr life200 kW

Li MPD 250-500 kWLi MPD

1 MWLi MPD

NTR 5klbf Engine test

NTP5 klbf Demo

>40,000 m 2Solar SailSpace Demo

Adv. Expander Engine Demo

Research Propulsive Applications of (1) Quantum Vacuum,(2) Inertial Frames, (3) Spacetim e Metrics

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greater than unity are required to launch from thesurface of the Earth, and chemical propulsion iscurrently the only propulsion technology capa-ble of producing the magnitude of thrust neces-sary to overcome Earths gravity. However, oncein space, more efficient propulsion systems can be

used to reduce total mission propellant mass re-quirements. Advanced In-Space Propulsion technologies will

enable much more effective exploration of our So-lar System and will permit mission designers toplan missions to y anytime, anywhere, and com-plete a host of science objectives at their destina-tions. A wide range of possible missions and can-didate chemical and advanced in-space propulsiontechnologies with diverse characteristics offersthe opportunity to better match propulsion sys-tems for future missions. Developing a portfolioof in-space propulsion technologies will allow op-timized propulsion solutions for a diverse set ofmissions and destinations. Te portfolio of con-cepts and technologies described in this roadmapare designed to address these future space scienceand exploration needs.1.2. Benets

In-space propulsion is a category of technolo-gy where developments can benet a number ofcritical Figures of Merit (metrics) for space explo-ration. Space exploration is about getting somewhere safely (mission enabling), getting therequickly (reduced transit times), getting a lot of

mass there (increased payload mass), and gettingthere cheaply (lower cost). Te simple act of "get-ting" there requires the employment of an in-space propulsion system, and the other metricsare modiers to this fundamental action. Simplyput, without a propulsion system, there would beno mission.

Development of technologies within this A will result in technical solutions with improve-

ments in thrust levels, Isp, power, specic mass(or specic power), volume, system mass, systemcomplexity, operational complexity, commonali-ty with other spacecraft systems, manufacturabil-ity, durability, and of course, cost. Tese types ofimprovements will yield decreased transit times,

increased payload mass, safer spacecraft, and de-creased costs. In some instances, development oftechnologies within this A will result in missionenabling breakthroughs that will revolutionizespace exploration.1.3. Applicability/Traceability to NASA

Strategic Goals, AMPM, DRMs, DRAsTe In-Space Propulsion Roadmap team used

the NASA strategic goals and missions detailed inthe following reference materials in the develop-ment of this report: Human Exploration Frame-

work eam products to extract reference missions with dates, the SMD Decadal Surveys, past De-sign Reference Missions, Design Reference Ar-chitectures, historical mission studies, In-SpacePropulsion echnology concept studies, and in-ternal ISS utilization studies. Appendix B con-tains references for reports and studies identify-ing missions used for categorizing pull and pushtechnology designations.1.4. Top Technical Challenges

Te major technical challenges for In-SpacePropulsion Systems echnology Area (ISPS A)

were identied and prioritized through teamconsensus based on perceived mission need orpotential impact on future in-space transporta-tion systems. Tese challenges were then catego-rized into near- (present to 2016), mid- (20172022), and far-term (20232028) time frames,representing the point at which RL 6 is expect-ed to be achieved. It is likely that support of thesetechnologies would need to begin well before thelisted time horizon.

RL-6 readiness dates were determined by con-Rank Description

1 Power Processing Units (PPUs) for ion, Hall, and other electric propulsion systems N

2 Long-term in-space cryogenic propellant storage and transfer M

3 High power (e.g. 50300 kW) class Solar Electric Propulsion scalable to MW class Nuclear Electric Propulsion M

4 Advanced in-space cryogenic engines and supporting components M

5 Developing and demonstrating MEMS-fabricated electrospray thrusters N

6 Demonstrating large (over 1000 m 2) solar sail equipped vehicle in space N

7 Nuclear Thermal Propulsion (NTP) components and systems F

8 Advanced space storable propellants M

9 Long-life (>1 year) electrodynamic tether propulsion system in LEO N

10 Advanced In-Space Propulsion Technologies (TRL


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sidering stated mission pull (for example, HEFor Decadal Surveys stating mission need dates,etc.), the state-of-the-art for specic technologiesthat could be matured to the point of quickly en-abling missions of interest to potential users (tech-nology push), and the need for a breadth of tech-

nology-enabled capabilities across all timeframes.2. DETAILED PORTFOLIO DISCUSSION

Te roadmap for this technical area is dividedinto four basic groups: (1) Chemical Propulsion,(2) Nonchemical Propulsion, (3) Advanced Pro-pulsion echnologies, and (4) Supporting ech-nologies. Te rst two categories are groupedaccording to the governing physics. ChemicalPropulsion includes propulsion systems that op-erate through chemical reactions to heat and ex-pand a propellant (or use a uid dynamic expan-sion, as in a cold gas) to provide thrust. Propulsionsystems that use electrostatic, electromagnet-ic, eld interactions, photon interactions, or ex-ternally supplied energy to accelerate a space-craft are grouped together under the section titledNonchemical Propulsion. Te third section, Ad-vanced Propulsion echnologies, is meant to cap-ture technologies and physics concepts that are ata lower RL level (< RL3). Te fourth section,Supporting echnologies, identies the pertinenttechnical areas that are strongly coupled to, butare not part of, in-space propulsion, such that fo-cused research within these related areas will al-low signicant improvements in performance forsome in-space propulsion technical areas. In ad-dition, development of some advanced forms ofchemical propulsion will have modeling challeng-es to better understand and predict dynamic insta-bility during combustion, and electric propulsiontechnologies require the enhancement and vali-dation of complicated life models to shorten life-qualication testing.

Development of technologies within this A willresult in technical solutions with improvementsin thrust levels, specic impulse, power, specicmass, system complexity, operational complexity,

commonality with other spacecraft systems, man-ufacturability, and durability. Te benets to bederived from each technology in the ABS will beidentied with one of the icons as described in thefollowing table on the right. Within each section of the following tables there

are three columns. Te left-most column providesa summary description of a particular technolo-gy, explaining its governing physics and methodof operation. Te middle column identies at a

high-level the technical challenges that must beovercome to raise its maturity. Te right-most col-umn for Sections 2.1, 2.2, and 2.4 describes thesignicant milestones to be reached for a giventechnology to attain RL-6. In Section 2.3 thiscolumn describes the milestones required for at-

taining RL >3.Tis roadmap makes use of the following set ofdenitions and ground rules. Te term "missionpull" denes a technology or a performance char-acteristic necessary to meet a planned NASA mis-sion requirement. Any other relationship betweena technology and a mission (an alternate propul-sion system for example) is categorized as ech-nology Push. Also, a distinction is drawn be-tween an in-space demonstration of a technologyversus an in-space validation. A space demonstra-tion refers to the spaceight of a scaled version ofa particular technology or of a critical technolo-gy subsystem; a space validation would serve as aqualication ight for future mission implemen-tation. A successful validation ight would not re-quire any additional space testing of a particulartechnology before it can be adopted for a scienceor exploration mission.

Te graphical Roadmap representation (Fig. 2,p. 3/4) provides suggested technology pursuits

within the four basic categories, and ties these ef-forts to the portfolio of known and potential fu-ture NASA/non-NASA missions. Most of thenear-term content on the graphic is based on ac-tual plans while the out years can be considered tohave larger uncertainties bars on the placement ofitems within the timeline.

Improvement Results

Icon Designator Description

T Decreased transit times

M Increased payload mass

C Reduces cost /system complexity/ im-proved system reliability

E Enable missions to new science and ex-ploration targets

R Provide potential propulsion break-throughs that will revolutionize spaceexploration


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2.1. Chemical PropulsionChemical Propulsion involves the chemical re-

action of propellants to move or control a space-craft. Chemical propulsion system functionsinclude primary propulsion, reaction control, sta-tion keeping, precision pointing, and orbital ma-

neuvering. Te main engines provide the primarypropulsive force for orbit transfer, planetary trajectories and extra planetary landing and ascent.Te reaction control and orbital maneuvering sys-tems provide the propulsive force for orbit main-tenance, position control, station keeping andspacecraft attitude control.

2.1 CHEMICAL PROPULSION

Description and State of the Art Technical Challenges Milestones to TRL 6

2.1.1 Liquid Storable

2.1.1.1 Monopropellants

Hydrazine thrusters use a catalytic decomposi-tion reaction to generate high temperature gas forthrust. Hydrazine is SOA. Spacecraft reaction controlsystem (RCS) performance is near I sp 228 sec. Landerengines have higher Isp (238 secs). Freezing point is3 C.

Catalyst life, inability for cold starts. In-creased thrust and Isp performance withpumped systems. Reduction of freezingpoint from 3 C needed without compro-mising the performance.

Evaluate alternate propellants such asNOFB, and AF315E. Develop thrustersto operate in pulse and continuousoperation with new propellant. Qual-ify propellants, components (valves,lters, regulators etc).

2.1.1.2 Bipropellants

Bipropellant thrusters use the chemical reaction,typically hypergolic, to generate high temperaturegas that is expanded to generate thrust. Nitrogen Tetroxide (NTO)/Hydrazine (N 2H4) is SOA with Isp 326secs for xed thrust (450 N) planetary main engine.

Increased thrust with improved packagingfor landers & orbit insertion. Throttle capa-bility for planetary landers. Pumped sys-tems desirable for planetary spacecraft vs.pressure fed systems. Mixture-ratio controland propellant gauging to reduce residuals& improve performance.

Develop and qualify pumped bi-pro-pellant system. Develop and qualifythrottleable bi-propellant valve/sys-tem. Recapture XLR-132 NTO/MMHpump-fed engine technology.

2.1.1.3 High-Energy Propellants

Bipropellant thrusters use the chemical reactions togenerate high temperature gas that is expanded togenerate thrust. One of the two may be cryogenicuid and may also require spark ignition systems.LO2 /N2H4 is a hypergolic option has comparable per-formance to LO 2 /LCH4. Higher thrust levels neededfor SMD missions.

No cryogenic engines have own otherthan RL10 for


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2.1.2.2 LO2, LH2

SOA is MMH/NTO at TRL 9 for Reaction Control Sys-tems (RCS) and orbital maneuvering propulsion,which are integrated. Development of LOX/LH 2 RCS(liquid propellants) allows integration with an up-per stage that uses LOX/LH 2. An O2 /H2 RCS typicallyinvolves taking low-pressure propellants from themain tanks, pumping to higher pressure, turning liq-uid to a gas, and then storing in a gas accumulator. The TRL is 4-5 with engines having been tested, dat-ing back to 1970 for early shuttle designs. Throttle-able LO2 /LH2 main engines for planetary descent areat TRL 4 with recent CECE engine testing. LOX/LH 2 for primary in-space propulsion trades well for somemission applications.

For integrated RCS, reducing system com-plexity and dry mass. For throttle-ablemain engine, developing deep throttlingcapability with good performance. Cryo-genic uid management issues must alsobe addressed.

Develop components (pumps, heatexchangers, accumulators) for the O 2 /H2 feed system and perform integrat-ed system level tests. Develop andtest LO 2 /LH2 throttleable main enginefor crewed descent.

2.1.3 Gelled & Metalized-Gelled Propellants

Gelled and metallized fuels are a class of thixotropic(shear thinning) fuels which improved the perfor-mance of rocket and airbreathing systems in severalways: increased rocket specic impulse, increasedfuel density, reduced spill radius in an accidentalspill, lower volatility during low pressure accidentalpropellant res, reduced fuel sloshing, and lowerleak potential from damaged fuel tanks (due tohigher propellant viscosity). Military systems havesought gelled fuels for all of these reasons. NASAsystems have studied gelled fuels analytically andexperimentally for lunar and Mars missions, upperstages, interplanetary robotic missions, and launchvehicle applications. Increased fuel density and in-creased engine specic impulse are the primarybenets. Missile ight tests, 1999, 2001, with Earth-storable propellants: Inhibited Red Fuming NitricAcid for the oxidizer, and gelled-MMH/Carbon forthe fuel.

Gelled cryogenic propellants have onlybeen tested in laboratory experiments andhave not yet own in a space representa-tive environment. One potential issue tobe addressed would be boil-off and a cor-responding shift in gellant-loading in thefuel. Cryogenic uid management issuesmust also be addressed. Storable NTO/MMH/Aluminum, Oxygen/RP-1/Aluminum,and Cryogenic Oxygen/Hydrogen/Aluminum are the primary candidates tobe investigated. The primary challengesare with gelling the fuels with the alumi-num particles.

Recapture gelled hydrogen/cryogen-ic fuel work from 1970s. Cryogenicuid management issues must alsobe addressed. Large scale (500-1000lbs thrust) RP-1/Aluminum, and Hy-drogen/Aluminum engine and com-ponent testing must be conducted.

2.1.4 Solid Rocket Propulsion Systems

Solid propellants are usually pre-mixed oxidizer andfuel that are then formed into a particular shape,so that when the surface is ignited the surface areaburns at a predetermined & tailored burn rate to

generate the thrust and duration required for themission. I sp values are normally less than 300 secs.For space-based solids, Hydroxyl-Terminated Poly-butadiene (HTPB) propellant has been used exclu-sively in apogee kick-motors & upper stages. Thrustvectoring is controlled by gimballing or gaseous/liquid injection.

Increase Isp by using nano-particles toincrease surface burning area. Improvedurability of thrust vector control systemto withstand long-term exposure to space

environment. Improve pointing accuracyin thrust vector control. Maintaining pres-sure in chamber to ensure ignition (covers,purge systems) is required for use in space.Long-term storage issues in space may notyet be well understood.

Develop and hot-re test formula-tions with nano-particles includedin solid propellant mix. Develop andqualify propellant for long-duration

space applications.

2.1.5 Hybrid

Hybrid rockets utilize a solid fuel and liquid oxidizer. They are potentially safer and have a higher I sp thansolid rockets, and are less complex and cheaper thanliquid rockets. Hybrid rockets are generally larger involume than solid rockets due to the lower densityof its propellants. Also, for most hybrid fuels the re-gression rate is much less than solids, which requiresmore burning surface for an equivalent thrust. Forsimple grain shapes, they are restrained to highlength/diameter ratios, which translate to long, thinmotors. Hybrid motors have been demonstrated atthe 250K thrust level. Recent funded investmentsin hybrid technology development has resulted insignicant progress reducing technology risk, withlong burn duration rings at 20K thrust level.

Further fundamental fuel/oxidizer/addi-tives/ propellant-web design investiga-tions combined with burn rate additives,paraffin, different oxidizer ows, and mul-tiport multilayer congurations need to beconducted. re-start/multiple ring use for upper stage is need-ed. For in-space propulsion, in general,higher mass-fraction and higher I sp rangesare critical design issues, and optimizationtrade studies are required.

System studies on large upper stagesand apogee kick-motors must beconducted over the range of tech-nology options. The most promisingcandidates (Aluminum-loading, highregression fuel, paraffin, additives,vortex, multiport/multi-layer congu-ration, etc.) should be tested at largerthrust levels to determine propulsionscaling and combustion efficiency.Subscale testing and analysis is need-ed to prepare candidate system(s) forenhanced component demonstra-tions at 250,000 lbs thrust. For futuremissions/applications, requirements,system studies and development ofcomponents will be needed.


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2.1.6 Cold Gas/Warm Gas

Cold Gas systems have been ying in space since the1950's with thrust levels from fractions of a pound to10s of pounds. Warm gas systems havebeen used in ight systems for pressurization, butnot for main propulsion. The principal advantageof the warm gas version of a cold gas system is a ~50% reduced tank volume. Gas propulsion systemsare typically used for small delta-V or when smalltotal impulse is required. Generally inexpensive andvery reliable, inert gases are inherently non-toxicand most of the residual risk lies with the high-pres-sure storage tanks, although good design providesample margin for safety. Cold gas systems are TRL 9;warm gas TRL5/6.

Getting a ight demonstration of a warmgas system is the next challenge. Thrusterdevelopment and development of a com-bination isolation valve/regulator is animportant improvement for packaging thesystem, and such a combination compo-nent could be used for other systems (suchas pressurization).

Denition of a specic mission, orthrust class to drive the required ap-plications engineering is importantfor a technology demonstration. Thruster development and devel-opment of a combination isolationvalve/regulator is required. Other al-ternative catalyst options need to beevaluated.

2.1.7 Micropropulsion

2.1.7.1 Solids

Solid motor microthrusters are simply miniature ver-sions of large solid-booster rockets. There are manysolid motor options that are ight ready for micro-propulsion applications.

Determine minimal size and thermal scal-ing for given applications.

Off-the-shelf designs are already at TRL > 6, and ready for ight demon-stration. Denition of a specic mis-sion and thrust class to drive applica-tions engineering is needed.

2.1.7.2 Cold Gas/Warm Gas

Micropropulsion cold/warm gas thrusters are min-iature versions of these devices described earlier. There are many off-the shelf small cold gas thrust-ers that are ight ready for micropropulsion appli-cations, some with ight heritage. Smaller thrustersystems based on liquied gas (e.g. butane) havebeen developed for inspector spacecraft and cube-sat application, and MEMS based thrusters havebeen developed. Most of these thrusters are await-ing ight demonstration.

Very few technical challenges exist otherthan ight demonstration.

These types of thrusters are already at TRL>6 and require ight demonstra-tion.

2.1.7.3 Hydrazine or Hydrogen Peroxide Monopropellant

Microthrusters using monopropellants such as N2H4(and others) are very small engines that produce lowthrust levels and minimum impulse bits for reactioncontrol systems (RCS). Hydrazine micro-thrusters areused for primary propulsion on small sats (~100 kgclass). SOA Performance for an MR-103H(TRL 9) isthrust: 1.07 N, I sp: 220 s, Power 6.5 W, Min. Ibit: 5000uNs, Mass: 195 g. A TRL 4-5 Hydrazine milli-Newton Thruster (HmNT) is under development that pro-duces thrust: 129 mN, I sp: 150 s, Power (valve): 8.25W, Min. Ibit: 50 uNs, Mass: 40 g. Under development(TRL 2-3) are MEMS fabricated versions, but no per-formance data is available yet.

Challenges include the development ofsmall catalyzer-beds, small high-speedow control valves and thermal controltechniques.

Successful testing and performancemeasurements need to be made onthese devices to elevate the TRL to 5.Qualication and long-duration test-ing need to be conducted to reach TRL 6.


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2.2. Nonchemical PropulsionNonchemical Propulsion serves the same set of

functions as chemical propulsion, but without us-ing chemical reactants. Example technologies in-clude: systems that accelerate reaction mass elec-trostatically and/or electromagnetically (Electric

Propulsion), systems the energize propellant ther-mally (Solar or Nuclear Termal Propulsion), andthose that interact with the space environment toobtain thrust electromagnetically (Solar Sail andether Propulsion).

2.2 Nonchemical Propulsion


2.2.1 Electric Propulsion

2.2.1.1 Elec trothermal

2.2.1.1.1 Resistojets

Resistojets use an electrically heated element incontact with the propellant to increase the enthal-py prior to expansion through a nozzle. Additionalheat may be added chemically, with hydrazine pro-pellant for example. Resistojets are a mature (TRL9) technology with hundreds of thrusters in opera-tion on commercial communications satellites forstation keeping, orbit insertion, attitude control,and de-orbiting. Off-the-shelf resistojets withpower levels ranging from 467-885 W are available.Low power (


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2.2.1.2 Electrostatic

2.2.1.2.1 Ion Thrusters

Ion thrusters employ a variety of plasma generationtechniques to ionize a large fraction of the propel-

lant. High voltage grids then extract the ions fromthe plasma and electrostaticly accelerate them tohigh velocity at voltages up to and exceeding 10kV. Ion thrusters feature the highest efficiency (60to >80%) and very high specic impulse (2000 toover 10,000 sec) compared to other thruster types.Over 130 ion thrusters have own in space on over30 spacecraft in both primary propulsion and sat-ellite station keeping applications. The propellantpresently used is xenon for its high atomic mass,easy storage on spacecraft and lack of contami-nation issues, although other propellants can beused. Flight thrusters operate at power levels from100 W to 4.5 kW. Various ion thrusters are at TRL 9(13cm XIPS, 25cm XIPS, NSTAR, T5 Kaufman Thrust-er, RIT10, 10 ECR, and ETS-8). The 7.2 kW NEXT ionthruster is already at TRL6 and requires ight dem-onstration or mission application.

Ion thruster performance and life isdetermined by the grids. Thrusters

operate at voltages of 750 V - 10,000V, and voltage breakdown of closelyspace multi-aperture grids is an im-portant issue. Improve-ments in low-erosion grid materials and longer lifecathodes are needed for future deepspace missions. Improvements in ef-ciency based on better plasma gen-erator design is needed. Improvedmodeling & model-based design &life predictions are also needed forfuture ion thruster development.

Ion thrusters require development to pro-duce higher I sp and longer life by utilizing

advanced grid materials. High power ionthrusters developed at JPL and GRC in the20 to 30 kW range are at TRL4 and requireenvironmental testing and life qualicationto achieve TRL6.

2.2.1.2.2 Hall Thrusters

Hall thrusters are electrostatic thrusters that utilizea cross-eld discharge described by the Hall effectto generate the plasma. An electric eld perpen-dicular to the applied magnetic eld acceleratesions to high exhaust velocities, while the transversemagnetic eld inhibits electron motion that wouldtend to short out the electric eld. Hall thrusterefficiency and specic impulse is somewhat lessthan that achievable in ion thrusters, but the thrustat a given power is higher and the device is muchsimpler. Over 240 xenon Hall thrusters have ownin space since 1971 with a 100% success rate. Com-mercially developed ight Hall thrusters operatebetween 0.2 and 4.5 kW with 50% efficiency, thrustdensities of 1 mN/cm 2, and Isp of 12002000 secs.Hall thrusters have been demonstrated from 0.1to 100 kW with efficiencies of 50-70%. Recent re-search has demonstrated operation with alterna-tive propellants and I sp increases to 30008000secs.

Scaling to high-power and achiev-ing sufficient lifetime are centralchallenges. Scaling to higher power(>10 kW) normally results in in-creased specic mass (kg/kW), butprovides longer lifetime due togreater amounts of wall materialinherent in larger designs. A majorchallenge is to capitalize on recentbreakthroughs on reducing wall ero-sion rates to realize very long life andthroughput (>1000 kg) and increaseIsp . Life validation of high-power,long-life thrusters requires develop-ment of physics-based models of theplasma & erosion processes.

Hall thruster power level must progress fromthrusters capable of 10s of kW of power tosystems of multiple thrusters capable of theorder of 1 MW [ref. HEFT study]. Key mile-stones for high power Hall thrusters aredemonstration of long-life technology onlarge thrusters (10's to 100's of kW), devel-opment of 100 kW or multi-100kW thrusterswith demonstration of performance andlife, and development of associated powerprocessing units (PPUs). The10-20-kW classthrusters developed by AFRL must be lev-eraged to achieve TRL6 within 3-5 years asa steppingstone to higher power thrusters.Larger thrusters operating at power levelsof 50 kW and higher require performancedemonstration at Isp from 2000 to 3000 sec,environmental testing and life qualicationto achieve TRL6.

2.2.1.3 Electromagnetic

2.2.1.3.1 Pulsed Inductive Thruster

The Pulsed Inductive Thruster (PIT) is an electro-magnetic plasma accelerator that creates its plas-ma by inductive breakdown of a layer of gaseouspropellant transiently puffed onto the surface of aninduction coil. Energy stored in a bank of capaci-tors switched into the coil produces an azimuthalelectric eld generates a at ring of current thatprovides a piston against which the rising mag-netic eld acts, entraining and ionizing the balanceof the propellant and ejecting it along the thrusteraxis. The PIT has demonstrated efficiency of greaterthan 50%, and an I sp of 2000-9000 secs in a singlepulse. New pulsed inductive concepts have operat-ed at much lower stored energy (100 J versus 2-4 KJof the high-power PIT) and provides a >3x smallerthruster operating at 20-40x less energy than thelarger variant.

Demonstration of the life of propel-lant valves and solid-state switches,and continuous operation at xedperformance. For ISRU, continuousoperation with H 2O without loss inperformance is needed. Similar chal-lenges exist for the scaled-down ver-sions of the PIT. Sustained power lev-els of 200 kW e at efficiencies of 70%or higher with I sp of 3000-10000 secsare required.

The key milestones to TRL 6 include demon-stration of switch and valve l ife >1010 pulses(>3yrs @100pps, >6yrs @50pps), and demon-stration of >70% thrust efficiency and I sp inthe range of 4,000 to 10,000 sec during con-tinuous operation. Continuous operationwith ammonia and/or water for in situ pro-pellant utilization must be demonstratedand the life veried.


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2.2.1.3.2 Magnetoplasmadynamic Thruster

Magnetoplasmadynamic (MPD) thrusters employthe interaction of high currents with either appliedmagnetic elds or the self-induced magnetic eldto accelerate ionized propellant. MPD thrustersoffer high efficiency and very high power process-ing capability in a small volume, and have beendemonstrated at steady state power levels up to 1MWe. There are three variants on the MPD thrust-er: steady state self-eld engines, steady stateapplied-eld engines, and quasi-steady thrusters.MPD thrusters show that they can achieve efficien-cies over 50% at I sps greater than 10,000 secs andthruster power levels of multi-MW e. State-of-the-art laboratory model (TRL 3-4) lithium applied eldthrusters have demonstrated efficiencies greaterthan 50% at 4,000 secs in a 200 kW e device andmodeling indicates they can achieve over 60% ef-ciency at power levels of 250 kW e and above.

Challenges are component lifetime,thermal management, and perfor-mance limitations due to the onsetphenomenon. The cathode mustoperate at high temperatures forsufficient emission current densities.Reduction in the cathode operatingtemperature by lowering the workfunction, passive radiative and/or ac-tive cooling must be demonstrated.Approaches to improving thrusterperformance by increasing the onsetcurrent must be understood theo-retically and experimentally veried.

MPD thruster power levels must progressfrom thrusters capable of 100s of kW to theorder of 1 MW or higher. Near-term mile-stones to achieve TRL6 at 200 to 250 kW e are demonstration of lithium applied-eldthruster performance at 60% efficiency atIsp of 6,000 sec and long life cathodes withanode thermal management and thrusterlife validated in long duration (10,000 hrs)wear tests. Mid-term milestones for TRL6thrusters at 1 MW include demonstrationof lithium self-eld thruster performance at>50% efficiency, long cathode life and an-ode thermal management, technologies forraising the onset current, and verication ofthruster life. In the far term, very high powerhydrogen-fueled thrusters must be devel-oped and demonstrated.

2.2.1.3.3 Variable Specic Impulse Magnetoplasma Rocket

The Variable Specic Impulse MagnetoplasmaRocket (VASIMR) is a high power electric space pro-pulsion engine capable of Isp/thrust modulation atconstant input power. Plasma is produced by heli-con discharge with energy added by ion cyclotronresonant heating (ICRH). Axial momentum is ob-tained by adiabatic expansion of plasma in a mag-netic nozzle. Thrust/specic impulse ratio controlachieved by partitioning RF power and propellantow between helicon and ICRH systems. The mostadvanced VASIMR prototype to date is the VX-200device that uses 35 kW helicon source to generateargon plasma, 170 kW ICRH section to heat p lasma,and an adiabatic magnetic nozzle to produce exitplume. The VX-200 leverages commercially devel-oped solid-state RF generators with efficienciesof 98 % & specic mass less than 1 kg/kW. It usesa cryogen-free low-temperature supercon-ductingmagnet to produce elds approaching 1 Tesla re-quired for helicon and ICRH sections. The VX-200has been operated at power levels of up to 200 kW& with performance data estimates of 5,000 sec I

sp

and 60% efficiency for 5 sec pulses.

The cur rent technical challenges arethe cryogen-free low-temperaturesuperconducting magnet, the mag-netic nozzle performance and life-time (from sputtering) and the hightemperature heat rejection systemrequired to extend the pulse length.Future technical challenges includespecic mass, heat rejection andthermal control, interactions of thedivergent plume with the spacecraft,and life qualication.

Subsystem: continue cryo-cooler testingand characterization for use with supercon-ducting magnets, trade 60kW pumped uidsystem (uses ISS PVTCS pump with Down-therm-J) performance against titanium-H 2OLoop Heat Pipe system (interdependencywith TA14), complete verication plan for200kW Nautel amplifer.System: Test VF-200 in a large vacuum cham-ber (e.g., Space Power Facility) to collectplume characteristics data for additionalmodelling validation. Measure & quantifyRF nature of VF-200 plume to ensure com-pliance with NASA spaceight hardwarerequirements. Perform testing of the VF-200 system to meet qualication & accep-tance environmental testing requirements.Demonstrate operation of VF-200 on ISSat 200kW power level for 15-min ring du-ration, ~5 N thrust, I sp of 4000-6000 secs,and efficiency of ~50-55%. Collect on-orbitplume characteristics data.

2.2.1.4 Micropropulsion

2.2.1.4.1 Microresistojets

See description on resistojets earlier. Microresisto- jets are smaller versions of conventional resistojetsand generally have an overall lower TRL because ofa lack of ight hardware development and demon-stration. Small thruster systems based on liqui-ed gas (e.g., butane) have been developed forInspector spacecraft and Cubesat applications andare ready to be ight demonstrated. MEMS basedresistojet thrusters have also been developed pri-marily for small-sat applications.

Conventional butane-based microre-sistojet thrusters require ight demo.For MEMS based designs, there areissues associated with microfabri-cation techniques, wall losses, andthermal design of the smaller thrust-er volumes. Additionally, the smallnozzle dimensions generally reducethe efficiency of these devices anddesigning MEMS based miniatureow-control valves has proven dif-cult.

Butane thrusters are at TRL 5-6 and ready foright demonstration. For MEMS based de-vices, demonstration is needed of the scala-biity to very small sizes, microfabrication ofthruster components, and verication ofthruster performance and life.

2.2.1.4.2 Microcavity Discharge, Teon

Microcavity discharge thrusters are similar to ar-cjects (see description above), except that the di-mensions are much smaller and the TRLs are lower(no ight development to date). MEMS fabricationhas been used along with smaller more conven-tional techniques. In these exceedingly small scaleddown versions of arc-jets, the arc is concentrated ina very small volume at lower power to maintain thelife of the single thruster, or congured in arraysone thruster can be used if another fails.

Challenges are MEMS-fabricationtechniques, reliability of arc elec-trodes, array design, and thermal is-sues.

Demonstrate scalability to very small sizes,microfabrication of thruster components,and verication of thruster performanceand life.


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2.2.1.4.3 Micropulse Plasma

Pulsed plasma thrusters (PPTs) use capacitively orinductively coupled plasma discharge to acceler-ate conductive gas, liquid, or solid at high powers(kW to MW in discharge) for very short pulses (80%).

This high efficiency removes many thermal issues,and the lack of a plasma discharge or connementrequirements simplies the construction and volt-age holdoff.

Challenges are MEMS fabricationof large tip arrays, microuidics fordistributing propellant to all thetips, cleanliness, power-processing,thruster testing, and lifetime. To date,the challenges of demonstratingmicrofabcation and microuidics tofeed the arrays with propellant in thelaboratory have prevented a space-qualiable protoype demonstration.

Milestones to TRL 6 include ight of the ST-7/LISA Pathnder that will demonstrate col-loid thruster performance, demonstrationof a MEMS fabricated prototype array, andperformance measurements, ight dem-onstration and successful life test of MEMSfabricated arrays.


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2.2.2 Solar Sail Propulsion

Solar sails are large, lightweight reective struc-tures that produce thrust by reecting solar pho-tons and thus transferring much of their momen-tum to the sail. The state-of-the-art solar sailswere produced for the NASA In Space Propulsion20 meter Ground System Demonstrations (GSD) in2005. The JAXA funded Interplanetary Kite-craftAccelerated by Radiation Of the Sun (IKAROS),launched in May, 2010, deployed its sail in June andhas since demonstrated both photon accelerationand attitude control. IKAROS has a square sail thatis approximately 14 meters long on each side, 7.5micrometers thick and used a spin deploymentmethod to deploy its sail.

System level integration and testof the component technologies for>1,000-m 2 sail using existing mate-rials and technologies are needed.Gravity offload and deployment testsof the large sail are needed.

Due to the constraints of gravity, solar sailpropulsion performance will not be totallydemonstrated to TRL 6 on the ground. Aspace ight demonstration will be requiredto fully achieve TRL 6.

2.2.3 Thermal Propulsion

2.2.3.1 Solar Thermal

Solar Thermal Propulsion (STP) heats the propel-lant with concentrated sunlight inside an absorbercavity and provides a very high specic impulse(~5001200 seconds). The solar energy is concen-trated and focused inside either a direct gain orthermal storage type engine conguration. Thesolar concentrator is may be rigid, segmented orinatable. The engine would be fabricated fromultra high temperature materials and operated asa heat exchanger with the propellant. A variety ofpropellants have been considered (e.g., hydrogen,methane, ammonia). The thrust level for this sys-tem would be on the order of 1.0 lbf. The LGarde ight experiment in 1996 demonstrat-ed the deployment of a large inatable concentra-tor (TRL6). The 30-day LH2 storage with controlledboil-off was demonstrated in 1998-1999 (TRL5).Various engine concepts have been made andfabricated from Rhenium, Tungsten/Rhenium, andRhenium coated graphite (TRL4). A new Ebeammanufacturing process has been demonstrated tofabricate complex STP engine designs. In addition,a new ultra-high temperature material (tri-carbide)has the potential to allow greater I sp than 1000 sec-onds.

Challenges are optical concentratoraccuracy and performance (from 50-60% to 85-90%), system/stage pack-aging, sun pointing (sub-arcsec accu-racy in at, 1 cm by 1 cm packages),inatable deployment, controlledcryogenic boil-off, and engine per-formance. An integrated overall sys-tem test has never been performed.STP is limited by payload shroudvolume when considering liquid hy-drogen LH 2. Options to overcomethis hurdle include the use of hightemperature carbides with melt-ing point ~4200K. At temperaturesabove 3200K and pressures ~25 psia.

Perform experiments with inatable or rigidconcentrators to increase the optical perfor-mance to 85-90% efficiency. Demonstratethe pointing capabilities of dual concentra-tors to keep the focuses on target duringall expected spacecraft orientations withthe sun. Demonstrate the propulsion per-formance of a high temperature carbidethruster manufactured with modern materi-al processing techniques, which allow com-plex geometries to be fabricated. Groundtest the performance of more dense propel-lants (e.g., methane, ammonia). In addition,demonstrate the utilization of deployableLH2 tanks.

2.2.3.2 Nuclear Thermal

Nuclear Thermal Rockets are a high thrust, highIsp propulsion technology. The state of the artground demonstrated engine, Nuclear Engine forRocket Vehicle Applications (NERVA) demonstratedthrusts (in the 1970's) comparable to chemical pro-pulsion (7,500 to 250,000 lbs of thrust with specicimpulses of 800 to 900 seconds, double that ofchemical rockets). Vehicles with solid-core NTR en-gines have been considered for human missions toMars as their high Isp allows reductions of the initialmass in low earth orbit (IMLEO) from 12 to 7 heavylift launch vehicle with 200 metric ton payloads.

The NERVA program matured thetechnology to a TRL 6 level in the1970s. The current challenge is tocapture the engineering and tech-nical knowledge base of the NERVAprogram. New fuel elements withlonger life is another technical chal-lenge to be addressed by future ef-forts.

Complete fuel tests select primary & back-up fuel/element design; design Ground &Flight Technology Demonstration engines;complete borehole gas injection tests - de-tailed design of ground test facility begins Authority to Proceed (ATP); complete small(5 klbf) ground technology demonstrationengines tests; conduct 5 klbf NTP Flight Technology demonstrator (FTD) mission;develop engine scale-up design for crewedMars missions.


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2.2.4 Tether Propulsion

2.2.4.1 Electrodynamic

An electrodynamic tether (EDT) can work as athruster because a magnetic eld exerts a force on a

current-carrying wire. This force is perpendicular tothe wire and to the eld vector. It is this interactionthat allows relatively short electrodynamic tethersto use solar power to push against a planetarymagnetic eld to achieve propulsion without ex-penditure of propellant. The groundwork has beenlaid for this type of propulsion. Important achieve-ments include retrieval of a tether in space (TSS-1, 1992), successful deployment of a 20-km-longtether in space (SEDS-1, 1993), and operation of anEDT in space with tether current driven in both di-rections (PMG, 1993). The Propulsive Small Expend-able Deployer System (ProSEDS) experiment wasto use an EDT to achieve ~0.4 N drag thrust, thusdeorbiting the stage. JAXA demonstrated in spacethe operation of a 300-m uninsulated tape tetheranode (T-Rex, 2010). Long-life, Micro-Meteoroid/Orbital Debris (MMOD) survivable tethers and al-ternative technologies that could replace hollow

cathode plasma contactors have been tested in thelaboratory.

The critical EDT subsystems havebeen tested both in space and in the

laboratory to TRL 5/6. To eld thistechnology, a full-scale (multi-kilo-meter, multi-kW) space validation atthe system level is required. The pri-mary challenge for EDT propulsionis at the system level, not at the sub-system level. Improvements to bet-ter preclude arcing and mechanismdeployment design can be pursued,but are not necessary for ight dem-onstration.

A space ight validation of an EDT propul-sion system will be required to fully achieve

TRL 6.

2.2.4.2 Momentum Exchange

A spinning, or Momentum Exchange Tether (MET)system can be used to boost payloads into higherorbits with a Hohmann-type delta-V or as an "el-evator" between two different orbits. A tether sys-tem would be anchored to a relatively large mass inLEO awaiting rendezvous with a payload deliveredto orbit by a launch vehicle. The uplifted payloadmeets with the tether facility, which then beginsa slow spin-up using electrodynamic tethers (forpropellantless operation) or another low thrust,high Isp thruster. At the proper moment and tethersystem orientation, the payload is released into atransfer orbit potentially to geostationary trans-fer orbit (GTO) or Earth Escape. Most of the achieve-ments cited for EDT propulsion establish the stateof the art for MET systems. The missing system-lev-el validations are spin-up, multi-spacecraft inter-action, orbital transfer dynamics, and rendezvousand docking with the tether tip.

There are numerous subsystem andcomponent-level engineering chal-lenges remaining before this capabil-ity reaches TRL-6: The developmentof long-life, survivable conductingand nonconducting tethers, deploy-ers, no expellant anodes and cath-odes, and a robust tether dynamicsmodeling capability are chief amongthem.

Very long (up to 100+ km), high strength-to-weight conducting & nonconducting teth-ers survivable in the LEO micrometeoroid,orbital debris, atomic oxygen & EUV envi-ronment.Reliable wire and/or tape tether deployerswith reel-in and reel-out capabilityAlternative anodes and cathodes with mini-mal or no expellants (Field Emitter ArrayCathodes, Solid Expellants, etc.)Flywheel energy storage (crossover to spacepower team)High-res, long-tether dynamics modelingto allow precise rendezvous between tethertip and payload.High-res, long-tether dynamics modeling toallow precision spin-up (if required, Rotova-tor only) orbital transfer of payloads.Rendezvous & docking of payload & tethertip for rapid payload transfer betweenlaunch vehicle (orbital or suborbital) & teth-er tip.


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2.3. Advanced Propulsion Technologies Advanced Propulsion echnologies are those

that use chemical or nonchemical physics to pro-duce thrust, but are generally considered to be oflower technical maturity ( RL< 3) than those de-scribed in Sections 2.1.1 and 2.1.2. Gravity Assist

is often used in conjunction with In-Space Propul-sion to provide the required mission v, but doesnot directly inuence or impact In-Space Propul-sion technologies discussed here. AeroGravity As-sist (AGA) is covered in A09.

2.3 Advanced (TRL


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3. INTERDEPENDENCY WITHOTHER TECHNOLOGY AREAS

Te ISPS A team evaluated the technical con-tent of the other fourteen technology area roadmaps for relationships with the content of theISPS A roadmap. In completing this evaluation,

three types of interdependencies were identiedin which either ISPS A technology development was related to technology development in anotherroad map, as illustrated in table below. In some cas-es, propulsion technology development is plannedin both road maps and close synergies could wereidentied. In other cases, ISPS A technology de-velopment depends on successful technology de-velopment in another A. For example, solar pow-er array and power management technologies are

developed by the Space Power and Energy Stor-age road map to support high power solar elec-tric propulsion. In the third type of dependency,another A road map is supported by technolo-gy developed by ISPS A. An example of this caseis development of throttleable cryogenic propul-

sion in the IPS A road map supporting the Entry,Descent and Landing A road map. For all threetypes of relationships, signicant benet will berealized through coordination of the implement-ing technology development projects.

4. POSSIBLE BENEFITS TOOTHER NATIONAL NEEDS

More capable and efficient in-space propul-sion will benet NASA, national defense, and the

2.4 Supporting Technologies


2.4.2 Propellant Storage, Transfer & Gauging

Cryogenic Fluid Management (CFM) broadlydescribes the suite of technologies that canbe used to enable the efficient in-spaceuse of cryogens despite their propensity toabsorb environmental heat, their complexthermodynamic and uid dynamic behav-ior in low gravity and the uncertainty of theposition of the liquid-vapor interface sincethe propellants are not settled. In additionto propulsion, this technology can supportpower reactant storage and ECLSS needs. The State of the Art is dened by commer-cial upper stages with a capability of up to9 hours in space (propellant boil-off rateson the order of 30%/day) and which requirepropellant settling (via thruster ring to ac-celerate the vehicle) before performing criti-cal functions. Many of the CFM technologyelements have been matured to a TRL ofnear 5 through ground testing.

Enable long duration (months to years) du-ration in space missions with cryogens andefficient in-space transfer of propellants fortanker or depot architectures. Specic chal-lenges to be addressed by in space demon-stration before reaching TRL 6 include:- Safe and efficient venting of unsettled pro-pellant tanks- Zero boil-off (ZBO) storage of cryogenicpropellants for long duration missions-Accurate micro-g propellant gauging anduid acquisition-Automated cryogenic uid couplings & pro-pellant transfer-Performance issues due to integration ofcomponents

1) Conduct ground tests in representativethermal vacuum environments of high del-ity CFM components (partial crossover with TA14 Thermal Management) and systems3) Complete short duration cryogenic ightexperiment(s) to obtain data to maturemodels of critical uid/thermal physics.4) Conduct CFM ight demonstration - Dem-onstrate in space long duration (>6 months)storage of LO 2 (ZBO) and LH2 (


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commercial space industry virtually any orga-nization that builds or uses space satellites. Spe-cic technologies with multi-user applicabilityinclude: Metalized Propellants for higher perfor-mance and safer missile systems; Long-DurationCryogenic Propulsion for high-energy orbit trans-fer; Electric Propulsion for longer-life communi-cations and Earth observing satellites; Solar Sailsfor sustained observation of the Earths polar re-gions; research toward Nuclear Termal Propul-sion will lead to smaller and more efficient reactordesigns; ethers will enable lower-cost access tospace and orbit transfer; and the development ofnew technologies from robust research and tech-

nology development will enable new missions andapplications for all potential users.

ACRONYMS AHMS Advanced Health Management System AMPM Agency Mission Planning Model ARC Ames Research Center ATP Authority to ProceedCFM Cryogenic Fluid ManagementClF3 Chlorine Tri uorideClF5 Chlorine Penta uorideDRM Design Reference MissionECLS Environmental Control and Life Support

ECR Electron Cyclotron ResonanceEDT Electrodynamic TetherEHS Environmental Health SystemGRC Glenn Research CenterGTO Geostationary Transfer OrbitHEDM High Energy Density MaterialsHmNT Hydrazine milli-Newton ThrusterHTPB Hydroxyl-Terminated PolybutadieneICRH Ion Cyclotron Resonant Heating

IKAROS Interplanetary Kite-Craft Accelerated by Radiation Of the Sun

IMLEO Initial Mass in Low-Earth OrbitISHM Integrated System Health ManagementISPSTA In-Space Propulsion Systems

Technology AreaISRU In-Situ Resource UtilizationISS International Space StationITOC Ion Thruster On a ChipJAXA Japanese Aerospace Exploration AgencyJSC Johnson Space CenterKSC Kennedy Space CenterLST Life Support TechnologiesMET Momentum Exchange TetherMMOD Micro-Meteoroid/Orbital DebrisMPD MagnetoplasmadynamicMMH MonomethylhydrazineMSFC Marshall Space Flight CenterNERVA Nuclear Engine for Rocket Vehicle

ApplicationsOF2 Oxygen Di uoridePIT Pulsed Inductive ThrusterPPUs Power Processing UnitsProSEDS Propulsive Small Expendable

Deployer SystemRCS Reaction Control SystemSDI Strategic Defense InitiativeSOA HydrazineTRL Technology Readiness LevelTSS-1 Tether In SpaceVASIMR Variable Speci c Impulse

Magnetoplasma RocketVAT Vacuum Arc ThrusterZBO Zero Boil-Off


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