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National Aeronautics and Space Administration

DRAFT In-Space Systems Roadmap Technology 02

Mike Meyer, Co-chair Les Johnson, Co-chair Bryan Palaszewski Dan Goebel Harold White David Coote

November • 2010 DRAFT This page is intentionally left blank

DRAFT Table of Contents Foreword Executive Summary TA02-1 1. General Overview TA02-2 1.1. Technical Approach TA02-2 1.2. Benefits TA02-5 1.3. Applicability/Traceability to NASA Strategic Goals, AMPM, DRMs, DRAs TA02-5 1.4. Top Technical Challenges TA02-5 2. Detailed Portfolio Discussion TA02-6 2.1. Chemical Propulsion TA02-7 2.2. Nonchemical Propulsion TA02-10 2.3. Advanced Propulsion Technologies TA02-16 2.4. Supporting Technologies TA02-19 3. Interdependency with Other Technology TA02-20 4. Possible Benefits to Other National Needs TA02-20 Acronyms TA02-21 Reference Missions TA02-22 Acknowledgements TA02-22

DRAFT Foreword NASA’s integrated technology roadmap, including both technology pull and technology push strategies, considers a wide range of pathways to advance the nation’s current capabilities. The present state of this effort is documented in NASA’s DRAFT Space Technology Roadmap, an integrated set of fourteen technology area roadmaps, recommending the overall technology investment strategy and prioritization of NASA’s space technology activities. This document presents the DRAFT Technology Area 02 input: In-Space Propulsion Systems. NASA developed this DRAFT Space Technology Roadmap for use by the National Research Council (NRC) as an initial point of departure. Through an open process of community engagement, the NRC will gather input, integrate it within the Space Technology Roadmap and provide NASA with recommendations on potential future technology investments. Because it is difficult to predict the wide range of future advances possible in these areas, NASA plans updates to its integrated technology roadmap on a regular basis.

DRAFT Executive Summary safely (mission enabling), getting there quickly In-space propulsion begins where the launch ve- (reduced transit times), getting a lot of there hicle upper stage leaves off, performing the func- (increased payload mass), and getting there cheap- tions of primary propulsion, reaction control, ly (lower cost). The simple act of “getting” there station keeping, precision pointing, and - requires the employment of an in-space propul- al maneuvering. The main engines used in space sion system, and the other metrics are modifiers provide the primary propulsive for orbit to this fundamental action. transfer, planetary trajectories and extra planetary Development of technologies within this TA landing and ascent. The reaction control and or- will result in technical solutions with improve- bital maneuvering systems provide the propulsive ments in levels, Isp, , specific mass force for orbit maintenance, position control, sta- (or specific power), volume, system mass, system tion keeping, and . complexity, operational complexity, commonali- Advanced in-space propulsion technologies will ty with other spacecraft systems, manufacturabil- enable much more effective exploration of our ity, durability, and of course, cost. These types of and will permit mission designers improvements will yield decreased transit times, to plan missions to “fly anytime, anywhere, and increased payload mass, safer spacecraft, and de- complete a host of science objectives at the desti- creased costs. In some instances, development nations” with greater reliability and safety. With of technologies within this TA will result in mis- a wide range of possible missions and candidate sion-enabling breakthroughs that will revolution- propulsion technologies, the question of which ize . There is no single propul- technologies are “best” for future missions is a dif- sion technology that will benefit all missions or ficult one. A portfolio of propulsion technologies mission types. The requirements for in-space pro- should be developed to provide optimum solu- pulsion vary widely due according to their intend- tions for a diverse set of missions and destinations. ed application. The technologies described herein A large fraction of the engines in use to- will support everything from small and day are chemical ; that is, they obtain the robotic to space stations needed to generate thrust by chemical re- and human missions to Mars. actions to create a hot gas that is expanded to pro- Figure 1 is a graphical representation of the In- duce thrust. A significant limitation of chemical Space Propulsion Technology Area Breakdown propulsion is that it has a relatively low specif- Structure (TABS). The TABS is divided into four basic groups: (1) Chemical Propulsion, (2) Non- ic (Isp, thrust per of pro- pellant). A significant improvement (>30%) in chemical Propulsion, (3) Advanced Propulsion Technologies, and (4) Supporting Technologies, Isp can be obtained by using cryogenic propel- lants, such as liquid oxygen and liquid , based on the physics of the propulsion system and for example. Historically, these have how it derives thrust as well as its technical matu- not been applied beyond upper stages. Further- rity. There may be credible meritous in-space pro- more, numerous concepts for advanced propul- pulsion concepts not foreseen or captured in this sion technologies, such as electric propulsion, are document that may be shown to be beneficial to commonly used for station keeping on commer- future mission applications. Care should be taken cial communications satellites and for prime pro- when implementing future investment strategies pulsion on some scientific missions because they to provide a conduit through which these con- have significantly higher values of specific im- cepts can be competitively engaged to encourage pulse. However, they generally have very small continued innovation. values of thrust and therefore must be operated Figure 2 is the roadmap for the development of for long durations to provide the total impulse re- advanced in-space propulsion technologies show- quired by a mission. Several of these technologies ing their traceability to potential future missions. offer performance that is significantly better than The roadmap makes use of the following set of that achievable with chemical propulsion. This definitions and ground rules. The term “mission roadmap describes the portfolio of in-space pro- pull” defines a technology or a performance char- pulsion technologies that could meet future space acteristic necessary to meet a planned NASA mis- science and exploration needs. sion requirement. Any other relationship between In-space propulsion represents technologies that a technology and a mission (an alternate propul- can significantly improve a number of critical met- sion system, for example) is categorized as “tech- rics. Space exploration is about getting somewhere nology push.” Also, a distinction is drawn be- DRAFT TA02-1 Figure 1. In-Space Propulsion Technology Area Breakdown Structure tween an in-space demonstration of a technology flexible launch dates are difficult, requiring pro- versus an in-space validation. A space demonstra- pulsion systems that are beyond today's current tion refers to the of a scaled version of state of the art. The logistics, and therefore the to- a particular technology or of a critical technolo- tal system mass required to support sustained hu- gy subsystem; a space validation would serve as a man exploration beyond to destinations qualification flight for future mission implemen- such as the , Mars or Near Earth Objects, tation. A successful validation flight would not re- are daunting unless more efficient in-space pro- quire any additional space testing of a particular pulsion technologies are developed and fielded. technology before it can be adopted for a science With the exception of electric propulsion sys- or exploration mission. The graphical roadmap tems used for commercial communications satel- provides suggested technology pursuits within the lite orbit positioning and station-keeping, and a four basic categories, and ties these efforts to the handful of lunar and deep space science missions, portfolio of known and potential future NASA/ all of the rocket engines in use today are chemi- non-NASA missions. cal rockets; that is, they obtain the energy needed to generate thrust by combining reactive chemi- 1. General Overview cals to create a hot gas that is expanded to produce thrust. A significant limitation of chemical pro- 1.1. Technical Approach pulsion is that it has a relatively low specific im- For both human and robotic exploration, tra- pulse (thrust per unit of mass flow rate of propel- versing the solar system is a struggle against time lant). Numerous concepts for advanced in-space and distance. The most distant planets are 4.5–6 propulsion technologies have been developed over billion kilometers from the and to reach them the past 50 years. While generally providing sig- in any reasonable time requires much more capa- nificantly higher compared to ble propulsion systems than conventional chemi- chemical engines, they typically generate much cal rockets. Rapid inner solar system missions with lower values of thrust. Thrust to weight ratios

TA02-2 DRAFT Figure 2: In Space Propulsion Technology Area Strategic Roadmap (TASR)

Technology Roadmap: In Space Propulsion 2010 2015 2020 2025 2030 Beyond Cargo transport xPRM Cargo transport Crewed HEO Crewed NEO Mission Mars Crewed Mars Crewed in LEO NEO ISS beyond GEO GEO Mission NEO Cargo Orbit Orbit Mars ESMD Propellantless 2000 km Mars Mission Surface Reboost Bridge SOMD Transfer Missions to to Space Multiple Tethered GEO and beyond Mission Satellites for Expolanet -free Expolanet Interstellar TSSM LWS SMD GeoStorm Ionospheric Studies Flagship Finder Earth- Finder Probe Evolved Common Grace II (L1 Diamond) observation DARPA Upper Stage (with NOAA) LEO-to-GEO LOKI Mission DARPA Reusable Non-NASA Users Boost Station T-Rex flight Orbital System Demo FTD >1,000 m2 Space Demo Fusion Flight Falcon NTP ≥1MW Flight (JAXA) CRYOSTAT SEP Refueling 2 50 MW, Propulsion Orbital Sat Orbital Demo from tanker >2,500 m2 NTP >40,000 m 25 klbf Demo ISS Exchange Solar Sail 1 kg/kW Deflection IKAROS flight Demo EP&P Solar Sail Tether Boost 5 klbf Demonstration 2 DARPA Space Demo Demo Demonstration (200 m JAXA) ST7 Testbed Space Demo Station Demo Demo LOX/LCH4 Chemical Propulsion: Common Upper Engine Components LO2/LCH4 Stage Engine 2,000 lbf Adv. Expander Adv. Composite decent engine Liquid & Gelled LOX/LCH4 Comb. Demo Ionic RCS Engine Demo Instability Test Solid & Hybrid Hybrid Nozzle LO2/LCH4 Deep Throttling LO2/LCH4 & TVC Eval 10,000 lbf Cryo Engine 100,000 lbf Variable Thrust/ decent engine decent engine Cold/Warm Gas & Micro Multi-Burn Motor High Energy Density Green Propellants SEP Stage Non-Chemical Propulsion: NEXT 50 kW Hall 100 kW Hall MPD Demo Thruster Thruster Thruster MEMS 100 W MEMS 10 W Electrospray 1 MW Colloid Colloid Electrospray MicroPropulsion Electrospray Miniature Li MPD Electric 200 kW 100,000 hr Li MPD 250-500 kW 1 MW 70% Eff. PIT Li MPD 1 MW Li MPD VASIMR PIT VF-200 VASIMR VF-500 VASIMR VF-1000

< 10 g/m2 System < 1 g/m2 System Solar Sail Level Sail Areal Density Level Sail Areal Density NTR Fuel Solar Thermal: Isp~1200 s, Solar Thermal: Downselection 85% Efficiency NTR 5klbf NTR 25klbf Isp~800 s, Engine test Engine test Thermal NTR Fuel, 70% Efficiency NTR Fuel Development Coating Experiments High Specific Strength Tether/Payload Tape Tether Tether Conductive Tethers Catch Mechanism Deployer Beamed Fusion Synthesize Beamed Fusion Advanced Fission/Fusion Research Fusion Low-TRL Advanced Energy Ground Metallic Energy Orbital Thrust Ground Propulsion: Suborbital Breakeven Hydrogen (microsat) Observed 50 MW Research Propulsive Applications of (1) Quantum , Investigate if effects of Engineering Breadboard Breakthrough Physics (2) Inertial Frames, (3) Spacetime Metrics ONGOING RESEARCH interest can be observed to produce desired effect

Supporting Technologies CFM & XFER EP PPU ISHM Adv Heat Rejection

Mission Implementation Flight Demo TRL 5/6 Significant Demo (TRL <5) Mission Pull Technology Push

PS–00326 Oct 10

DRAFT TA02–3/4 This page is intentionally left blank greater than unity are required to launch from the ments in thrust levels, Isp, power, specific mass surface of the Earth, and chemical propulsion is (or specific power), volume, system mass, system currently the only propulsion technology capa- complexity, operational complexity, commonali- ble of producing the magnitude of thrust neces- ty with other spacecraft systems, manufacturabil- sary to overcome Earth’s . However, once ity, durability, and of course, cost. These types of in space, more efficient propulsion systems can be improvements will yield decreased transit times, used to reduce total mission mass re- increased payload mass, safer spacecraft, and de- quirements. creased costs. In some instances, development of Advanced In-Space Propulsion technologies will technologies within this TA will result in mission enable much more effective exploration of our So- enabling breakthroughs that will revolutionize lar System and will permit mission designers to space exploration. plan missions to fly anytime, anywhere, and com- 1.3. Applicability/Traceability to NASA plete a host of science objectives at their destina- Strategic Goals, AMPM, DRMs, DRAs tions. A wide range of possible missions and can- The In-Space Propulsion Roadmap team used didate chemical and advanced in-space propulsion the NASA strategic goals and missions detailed in technologies with diverse characteristics offers the following reference materials in the develop- the opportunity to better match propulsion sys- ment of this report: Human Exploration Frame- tems for future missions. Developing a portfolio work Team products to extract reference missions of in-space propulsion technologies will allow op- with dates, the SMD Decadal Surveys, past De- timized propulsion solutions for a diverse set of sign Reference Missions, Design Reference Ar- missions and destinations. The portfolio of con- chitectures, historical mission studies, In-Space cepts and technologies described in this roadmap Propulsion Technology concept studies, and in- are designed to address these future space science ternal ISS utilization studies. Appendix B con- and exploration needs. tains references for reports and studies identify- 1.2. Benefits ing missions used for categorizing pull and push In-space propulsion is a category of technolo- technology designations. gy where developments can benefit a number of 1.4. Top Technical Challenges critical Figures of Merit (metrics) for space explo- The major technical challenges for In-Space ration. Space exploration is about getting some- Propulsion Systems Technology Area (ISPSTA) where safely (mission enabling), getting there were identified and prioritized through team quickly (reduced transit times), getting a lot of consensus based on perceived mission need or mass there (increased payload mass), and getting potential impact on future in-space transporta- there cheaply (lower cost). The simple act of "get- tion systems. These challenges were then catego- ting" there requires the employment of an in- rized into near- (present to 2016), mid- (2017– space propulsion system, and the other metrics 2022), and far-term (2023–2028) time frames, are modifiers to this fundamental action. Simply representing the point at which TRL 6 is expect- put, without a propulsion system, there would be ed to be achieved. It is likely that support of these no mission. technologies would need to begin well before the Development of technologies within this TA listed time horizon. will result in technical solutions with improve- TRL-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. 50–300 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 N 6 Demonstrating large (over 1000 m2) solar sail equipped 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 <3) to enable a robust technology portfolio for future missions. F

DRAFT TA02-5 sidering stated mission pull (for example, HEFT high-level the technical challenges that must be or Decadal Surveys stating mission need dates, overcome to raise its maturity. The right-most col- etc.), the state-of-the-art for specific technologies umn for Sections 2.1, 2.2, and 2.4 describes the that could be matured to the point of quickly en- significant milestones to be reached for a given abling missions of interest to potential users (tech- technology to attain TRL-6. In Section 2.3 this nology push), and the need for a breadth of tech- column describes the milestones required for at- nology-enabled capabilities across all timeframes. taining TRL >3. This roadmap makes use of the following set of 2. Detailed Portfolio Discussion definitions and ground rules. The term "mission The roadmap for this technical area is divided pull" defines a technology or a performance char- into four basic groups: (1) Chemical Propulsion, acteristic necessary to meet a planned NASA mis- (2) Nonchemical Propulsion, (3) Advanced Pro- sion requirement. Any other relationship between pulsion Technologies, and (4) Supporting Tech- a technology and a mission (an alternate propul- nologies. The first two categories are grouped sion system for example) is categorized as “Tech- according to the governing physics. Chemical nology Push.” Also, a distinction is drawn be- Propulsion includes propulsion systems that op- tween an in-space demonstration of a technology erate through chemical reactions to heat and ex- versus an in-space validation. A space demonstra- pand a propellant (or use a fluid dynamic expan- tion refers to the spaceflight of a scaled version of sion, as in a cold gas) to provide thrust. Propulsion a particular technology or of a critical technolo- systems that use electrostatic, electromagnet- gy subsystem; a space validation would serve as a ic, field interactions, interactions, or ex- qualification flight for future mission implemen- ternally supplied energy to accelerate a space- tation. A successful validation flight would not re- craft are grouped together under the section titled quire any additional space testing of a particular Nonchemical Propulsion. The third section, Ad- technology before it can be adopted for a science vanced Propulsion Technologies, is meant to cap- or exploration mission. ture technologies and physics concepts that are at The graphical Roadmap representation (Fig. 2, a lower TRL level (

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 thrusters use a catalytic decomposi- Catalyst life, inability for cold starts. In- Evaluate alternate propellants such as tion reaction to generate high temperature gas for creased thrust and Isp performance with NOFB, and AF315E. Develop thrusters thrust. Hydrazine is SOA. Spacecraft reaction control pumped systems. Reduction of freezing to operate in pulse and continuous

system (RCS) performance is near Isp 228 sec. point from 3 °C needed without compro- operation with new propellant. Qual- engines have higher Isp (238 secs). Freezing point is mising the performance. ify propellants, components (valves, 3 °C. filters, regulators etc). 2.1.1.2 Bipropellants Bipropellant thrusters use the chemical reaction, Increased thrust with improved packaging Develop and qualify pumped bi-pro- typically hypergolic, to generate high temperature for landers & . Throttle capa- pellant system. Develop and qualify gas that is expanded to generate thrust. bility for planetary landers. Pumped sys- throttleable bi-propellant valve/sys-

Tetroxide (NTO)/Hydrazine (N2H4) is SOA with Isp 326 tems desirable for planetary spacecraft vs. . Recapture XLR-132 NTO/MMH secs for fixed thrust (450 N) planetary main engine. pressure fed systems. Mixture-ratio control pump-fed engine technology. and propellant gauging to reduce residuals & improve performance. 2.1.1.3 High-Energy Propellants

Bipropellant thrusters use the chemical reactions to No cryogenic engines have flown other Develop and qualify pump fed LO2/ generate high temperature gas that is expanded to than RL10 for <24 hrs. Cryogenic storage MMH engine. Demonstrate operation generate thrust. One of the two may be cryogenic and operation for long duration space and performance and qualify long- fluid and may also require spark ignition systems. based missions have not been demon- term storage of LO2 for space applica-

LO2/N2H4 is a hypergolic option has comparable per- strated. Valve leakage, boiloff manage- tion (see Section 2.4.2). formance to LO2/LCH4. Higher thrust levels needed ment and thermal environment significant for SMD missions. challenge. 2.1.1.4 High-Energy Oxidizers High-energy oxidizers such as fluorinated com- Fluorinated propellants have safety issues The stage development for this tech- pounds include chlorine trifluoride (ClF3), chlorime (high reactivity), but the upper stage pro- nology was designed for SDI, etc. pentafluoride (ClF5) and & oxygen difluoride (OF2). cessing methods to isolate ground support Recapturing the handling and upper These oxidizers have a long history of testing with personnel from the oxidizers have been stage ground processing methods is most recent testing in the 1980s under the Strategic developed. These processing methods needed. Defense Initiative (SDI). Stages for interceptors were have not been exercised since the 1980s. created for flight testing using hydrazine/ClF5. 2.1.2 Liquid Cryogenic 2.1.2.1 LO2, CH4 SOA is MMH/NTO at TRL 9 for level integration and test of the Perform system-level integration and System (RCS) and orbital maneuvering propulsion, component technologies are needed. test of the component technologies. which are integrated. LOX/Methane is proposed Improvement in the main engine injector Some component improvements are to enable higher performance, space storability, performance and stability. Development required such as to improve the main

pressure-fed and pump-fed options, common LO2 of flight-weight compact exciter, and dem- engine injector performance and sta- and LCH4 components (lower cost), application to In- onstrating the ability to deliver the correct bility. Test a regeneratively cooled Situ Resource Utilization (ISRU) for Mars, and higher quality of propellant for repeatable engine main engine. density for improved packaging. LOX/Methane is performance are needed. TRL 4-5 in that Cryogenic Fluid Management (CFM), feed systems, RCS, main engine, & components have been tested in vacuum environments.

DRAFT TA02-7 2.1 CHEMICAL PROPULSION Description and State of the Art Technical Challenges Milestones to TRL 6

2.1.2.2 LO2, LH2 SOA is MMH/NTO at TRL 9 for Reaction Control Sys- For integrated RCS, reducing system com- Develop components (pumps, heat

tems (RCS) and orbital maneuvering propulsion, plexity and dry mass. For throttle-able exchangers, accumulators) for the O2/ which are integrated. Development of LOX/LH2 RCS main engine, developing deep throttling H2 feed system and perform integrat- (liquid propellants) allows integration with an up- capability with good performance. Cryo- ed system level tests. Develop and

per stage that uses LOX/LH2. An O2/H2 RCS typically genic fluid management issues must also test LO2/LH2 throttleable main engine involves taking low-pressure propellants from the be addressed. for crewed descent. main 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 are at TRL 4 with recent CECE engine testing. LOX/LH2 for primary in-space propulsion trades well for some mission applications. 2.1.3 Gelled & Metalized-Gelled Propellants Gelled and metallized fuels are a class of thixotropic Gelled cryogenic propellants have only Recapture gelled hydrogen/cryogen- (shear thinning) fuels which improved the perfor- been tested in laboratory experiments and ic fuel work from 1970s. Cryogenic mance of rocket and airbreathing systems in several have not yet flown in a space representa- fluid management issues must also ways: increased rocket specific impulse, increased tive environment. One potential issue to be addressed. Large scale (500-1000 fuel density, reduced spill radius in an accidental be addressed would be boil-off and a cor- lbs thrust) RP-1/Aluminum, and Hy- spill, lower volatility during low pressure accidental responding shift in gellant-loading in the drogen/Aluminum engine and com- propellant fires, reduced fuel sloshing, and lower fuel. Cryogenic fluid management issues ponent testing must be conducted. leak potential from damaged fuel tanks (due to must also be addressed. Storable NTO/ higher propellant viscosity). Military systems have MMH/Aluminum, Oxygen/RP-1/Aluminum, sought gelled fuels for all of these reasons. NASA and Cryogenic Oxygen/Hydrogen/ systems have studied gelled fuels analytically and Aluminum are the primary candidates to experimentally for lunar and Mars missions, upper be investigated. The primary challenges stages, interplanetary robotic missions, and launch are with gelling the fuels with the alumi- vehicle applications. Increased fuel density and in- num particles. creased engine specific impulse are the primary benefits. Missile flight tests, 1999, 2001, with Earth- storable propellants: Inhibited Red Fuming Nitric Acid for the oxidizer, and gelled-MMH/Carbon for the fuel. 2.1.4 Solid Rocket Propulsion Systems Solid propellants are usually pre-mixed oxidizer and Increase Isp by using nano-particles to Develop and hot-fire test formula- fuel that are then formed into a particular shape, increase surface burning area. Improve tions with nano-particles included so that when the surface is ignited the surface area durability of thrust vector control system in solid propellant mix. Develop and burns at a predetermined & tailored burn rate to to withstand long-term exposure to space qualify propellant for long-duration generate the thrust and duration required for the environment. Improve pointing accuracy space applications.

mission. Isp values are normally less than 300 secs. in thrust vector control. Maintaining pres- For space-based solids, Hydroxyl-Terminated Poly- sure in chamber to ensure ignition (covers, butadiene (HTPB) propellant has been used exclu- purge systems) is required for use in space. sively in apogee kick-motors & upper stages. Thrust Long-term storage issues in space may not vectoring is controlled by gimballing or gaseous/ yet be well understood. liquid injection. 2.1.5 Hybrid Hybrid rockets utilize a solid fuel and liquid oxidizer. Further fundamental fuel/oxidizer/addi- System studies on large upper stages

They are potentially safer and have a higher Isp than tives/ propellant-web design investiga- and apogee kick-motors must be solid rockets, and are less complex and cheaper than tions combined with burn rate additives, conducted over the range of tech- liquid rockets. Hybrid rockets are generally larger in paraffin, different oxidizer flows, and mul- nology options. The most promising volume than solid rockets due to the lower density tiport multilayer configurations need to be candidates (Aluminum-loading, high of its propellants. Also, for most hybrid fuels the re- conducted. re-start/ regression fuel, paraffin, additives, gression rate is much less than solids, which requires multiple firing use for upper stage is need- vortex, multiport/multi-layer configu- more burning surface for an equivalent thrust. For ed. For in-space propulsion, in general, ration, etc.) should be tested at larger

simple grain shapes, they are restrained to high higher mass-fraction and higher Isp ranges thrust levels to determine propulsion length/diameter ratios, which translate to long, thin are critical design issues, and optimization scaling and combustion efficiency. motors. Hybrid motors have been demonstrated at trade studies are required. Subscale testing and analysis is need- the 250K thrust level. Recent funded investments ed to prepare candidate system(s) for in hybrid technology development has resulted in enhanced component demonstra- significant progress reducing technology risk, with tions at 250,000 lbs thrust. For future long burn duration firings at 20K thrust level. missions/applications, requirements, system studies and development of components will be needed.

TA02-8 DRAFT 2.1 CHEMICAL PROPULSION Description and State of the Art Technical Challenges Milestones to TRL 6 2.1.6 Cold Gas/Warm Gas Cold Gas systems have been flying in space since the Getting a flight demonstration of a warm Definition of a specific mission, or 1950's with thrust levels from fractions of a to gas system is the challenge. Thruster thrust class to drive the required ap- 10s of pounds. Warm gas systems have development and development of a com- plications engineering is important been used in flight systems for pressurization, but bination isolation valve/regulator is an for a technology demonstration. not for main propulsion. The principal advantage important improvement for packaging the Thruster development and devel- of the warm gas version of a cold gas system is a ~ system, and such a combination compo- opment of a combination isolation 50% reduced tank volume. Gas propulsion systems nent could be used for other systems (such valve/regulator is required. Other al- are typically used for small delta-V or when small as pressurization). ternative catalyst options need to be total impulse is required. Generally inexpensive and evaluated. very reliable, inert gases are inherently non-toxic and most of the residual risk lies with the high-pres- sure storage tanks, although good design provides ample margin for safety. Cold gas systems are TRL 9; warm gas TRL5/6.

2.1.7 Micropropulsion 2.1.7.1 Solids Solid motor microthrusters are simply miniature ver- Determine minimal size and thermal scal- Off-the-shelf designs are already at sions of large solid-booster rockets. There are many ing for given applications. TRL > 6, and ready for flight demon- solid motor options that are flight ready for micro- stration. Definition of a specific mis- propulsion applications. 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- Very few technical challenges exist other These types of thrusters are already at iature versions of these devices described earlier. than flight demonstration. TRL>6 and require flight demonstra- There are many off-the shelf small cold gas thrust- tion. ers that are flight ready for micropropulsion appli- cations, some with flight heritage. Smaller thruster systems based on “liquified gas” (e.g. butane) have been developed for inspector spacecraft and cube- sat application, and MEMS based thrusters have been developed. Most of these thrusters are await- ing flight demonstration. 2.1.7.3 Hydrazine or Hydrogen Peroxide Monopropellant Microthrusters using monopropellants such as N2H4 Challenges include the development of Successful testing and performance (and others) are very small engines that produce low small catalyzer-beds, small high-speed measurements need to be made on thrust levels and minimum impulse bits for reaction flow control valves and thermal control these devices to elevate the TRL to 5. control systems (RCS). Hydrazine micro-thrusters are techniques. Qualification and long-duration test- used for primary propulsion on small sats (~100 kg ing need to be conducted to reach class). SOA Performance for an MR-103H(TRL 9) is TRL 6. thrust: 1.07 N, Isp: 220 s, Power 6.5 W, Min. Ibit: 5000 uNs, Mass: 195 g. A TRL 4-5 Hydrazine milli- Thruster (HmNT) is under development that pro- duces thrust: 129 mN, Isp: 150 s, Power (valve): 8.25 W, Min. Ibit: 50 uNs, Mass: 40 g. Under development (TRL 2-3) are MEMS fabricated versions, but no per- formance data is available yet.

DRAFT TA02-9 2.2. Nonchemical Propulsion Nonchemical 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 Thermal Propulsion), and those that interact with the space environment to obtain thrust electromagnetically (Solar Sail and Tether Propulsion).

2.2 Nonchemical Propulsion Description and State of the Art Technical Challenges Milestones to TRL 6 2.2.1 Electric Propulsion 2.2.1.1 Electrothermal 2.2.1.1.1 Resistojets Resistojets use an electrically heated element in Challenges are in scaling technol- Demonstrate scalability to very small sizes, contact with the propellant to increase the enthal- ogy to much smaller sizes and power microfabrication of thruster components, py prior to expansion through a nozzle. Additional levels for use on microspacecraft, in- and verification of thruster performance heat may be added chemically, with hydrazine pro- cluding microfabrication techniques, and life. pellant for example. Resistojets are a mature (TRL high temp materials, low-leak rate 9) technology with hundreds of thrusters in opera- microfabricated valves for small gas- tion on commercial communications satellites for fed systems, achieving high perfor- station keeping, orbit insertion, attitude control, mance with low Reynolds number and de-orbiting. Off-the-shelf resistojets with , and lifetime of high-temp power levels ranging from 467-885 W are available. components. Low power (<50 W) resistojets that operate on xe- non, nitrogen or butane have been developed and flown on over 20 spacecraft. A multipropellant re- sistojet designed to operate on waste gases from the ISS was developed, but never flown. Applica- tions for resistojets are attitude control or proximity operations near small bodies that might benefit from longer life and higher performance. 2.2.1.1.2 Arcjets Arcjets use an electric arc to heat the propellant Minor product improvements are be- No immediate applications that require ad- prior to expansion through a nozzle. Additional ing made on existing products, but vanced arcjets. heat may be added chemically, with hydrazine there is little mission pull for more propellant for example. Arcjets are a mature (TRL advanced arcjets. 9) technology with hundreds of thrusters in opera- tion on commercial communications satellites, pri- marily for station keeping. Off-the shelf hydrazine arcjet systems have power levels of 1670 to 2000 W. Lower power hydrazine arcjets (~500 W) have achieved TRL 5-6. arcjets at 30 kW were flight-qualified (TRL 7). Laboratory model hydro- gen arcjets have power levels ranging from 1 to 100 kW, but did not progress beyond ~TRL 4.

TA02-10 DRAFT 2.2 Nonchemical Propulsion Description and State of the Art Technical Challenges Milestones to TRL 6 2.2.1.2 Electrostatic 2.2.1.2.1 Ion Thrusters Ion thrusters employ a variety of generation Ion thruster performance and life is Ion thrusters require development to pro- techniques to ionize a large fraction of the propel- determined by the grids. Thrusters duce higher Isp and longer life by utilizing lant. High voltage grids then extract the from operate at voltages of 750 V - 10,000 advanced grid materials. High power ion the plasma and electrostaticly accelerate them to V, and voltage breakdown of closely thrusters developed at JPL and GRC in the high at voltages up to and exceeding 10 space multi-aperture grids is an im- 20 to 30 kW range are at TRL4 and require kV. Ion thrusters feature the highest efficiency (60 portant issue. Improve-ments in low- environmental testing and life qualification to >80%) and very high specific impulse (2000 to erosion grid materials and longer life to achieve TRL6. over 10,000 sec) compared to other thruster types. are needed for future deep Over 130 ion thrusters have flown in space on over space missions. Improvements in ef- 30 spacecraft in both primary propulsion and sat- ficiency based on better plasma gen- ellite station keeping applications. The propellant erator design is needed. Improved presently used is xenon for its high atomic mass, modeling & model-based design & easy storage on spacecraft and lack of contami- life predictions are also needed for nation issues, although other propellants can be future ion thruster development. used. Flight thrusters operate at power levels from 100 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 ion thruster is already at TRL6 and requires flight dem- onstration or mission application. 2.2.1.2.2 Hall Thrusters Hall thrusters are electrostatic thrusters that utilize Scaling to high-power and achiev- Hall thruster power level must progress from a cross-field discharge described by the Hall effect ing sufficient lifetime are central thrusters capable of 10’s of kW of power to to generate the plasma. An perpen- challenges. Scaling to higher power systems of multiple thrusters capable of the dicular to the applied accelerates (>10 kW) normally results in in- order of 1 MW [ref. HEFT study]. Key mile- ions to high exhaust , while the transverse creased specific mass (kg/kW), but stones for high power Hall thrusters are magnetic field inhibits motion that would provides longer lifetime due to demonstration of long-life technology on tend to short out the electric field. Hall thruster greater amounts of wall material large thrusters (10's to 100's of kW), devel- efficiency and specific impulse is somewhat less inherent in larger designs. A major opment of 100 kW or multi-100kW thrusters than that achievable in ion thrusters, but the thrust challenge is to capitalize on recent with demonstration of performance and at a given power is higher and the device is much breakthroughs on reducing wall ero- life, and development of associated power simpler. Over 240 xenon Hall thrusters have flown sion rates to realize very long life and processing units (PPUs). The10-20-kW class in space since 1971 with a 100% success rate. Com- throughput (>1000 kg) and increase thrusters developed by AFRL must be lev- mercially developed flight Hall thrusters operate Isp. Life validation of high-power, eraged to achieve TRL6 within 3-5 years as between 0.2 and 4.5 kW with 50% efficiency, thrust long-life thrusters requires develop- a steppingstone to higher power thrusters. densities of 1 mN/cm2, and Isp of 1200–2000 secs. ment of physics-based models of the Larger thrusters operating at power levels Hall thrusters have been demonstrated from 0.1 plasma & erosion processes. of 50 kW and higher require performance to 100 kW with efficiencies of 50-70%. Recent re- demonstration at Isp from 2000 to 3000 sec, search has demonstrated operation with alterna- environmental testing and life qualification tive propellants and Isp increases to 3000–8000 to achieve TRL6. secs. 2.2.1.3 Electromagnetic 2.2.1.3.1 Pulsed Inductive Thruster The Pulsed Inductive Thruster (PIT) is an electro- Demonstration of the life of propel- The key milestones to TRL 6 include demon- magnetic plasma accelerator that creates its plas- lant valves and solid-state switches, stration of switch and valve life >1010 pulses ma by inductive breakdown of a layer of gaseous and continuous operation at fixed (>3yrs @100pps, >6yrs @50pps), and demon- propellant transiently puffed onto the surface of an performance. For ISRU, continuous stration of >70% thrust efficiency and Isp in induction coil. Energy stored in a bank of capaci- operation with H2O without loss in the range of 4,000 to 10,000 sec during con- tors switched into the coil produces an azimuthal performance is needed. Similar chal- tinuous operation. Continuous operation electric field generates a flat ring of current that lenges exist for the scaled-down ver- with ammonia and/or water for in situ pro- provides a piston against which the rising mag- sions of the PIT. Sustained power lev- pellant utilization must be demonstrated netic field acts, entraining and ionizing the balance els of 200 kWe at efficiencies of 70% and the life verified. of the propellant and ejecting it along the thruster or higher with Isp of 3000-10000 secs axis. The PIT has demonstrated efficiency of greater are required. than 50%, and an Isp of 2000-9000 secs in a single pulse. New pulsed inductive concepts have operat- ed at much lower stored energy (100 J versus 2-4 KJ of the high-power PIT) and provides a >3x smaller thruster operating at 20-40x less energy than the larger variant.

DRAFT TA02-11 2.2 Nonchemical Propulsion Description and State of the Art Technical Challenges Milestones to TRL 6 2.2.1.3.2 Magnetoplasmadynamic Thruster Magnetoplasmadynamic (MPD) thrusters employ Challenges are component lifetime, MPD thruster power levels must progress the interaction of high currents with either applied thermal management, and perfor- from thrusters capable of 100’s of kW to the magnetic fields or the self-induced magnetic field mance limitations due to the “onset” order of 1 MW or higher. Near-term mile-

to accelerate ionized propellant. MPD thrusters phenomenon. The must stones to achieve TRL6 at 200 to 250 kWe offer high efficiency and very high power process- operate at high temperatures for are demonstration of applied-field ing capability in a small volume, and have been sufficient emission current densities. thruster performance at 60% efficiency at

demonstrated at steady state power levels up to 1 Reduction in the cathode operating Isp of 6,000 sec and long life cathodes with MWe. There are three variants on the MPD thrust- temperature by lowering the work thermal management and thruster er: steady state self-field engines, steady state function, passive radiative and/or ac- life validated in long duration (10,000 hrs) applied-field engines, and quasi-steady thrusters. tive cooling must be demonstrated. wear tests. Mid-term milestones for TRL6 MPD thrusters show that they can achieve efficien- Approaches to improving thruster thrusters at 1 MW include demonstration

cies over 50% at Isp’s greater than 10,000 secs and performance by increasing the onset of lithium self-field thruster performance at thruster power levels of multi-MWe. State-of-the- current must be understood theo- >50% efficiency, long cathode life and an- art laboratory model (TRL 3-4) lithium applied field retically and experimentally verified. ode thermal management, technologies for thrusters have demonstrated efficiencies greater raising the onset current, and verification of

than 50% at 4,000 secs in a 200 kWe device and thruster life. In the far term, very high power modeling indicates they can achieve over 60% ef- hydrogen-fueled thrusters must be devel-

ficiency at power levels of 250 kWe and above. oped and demonstrated. 2.2.1.3.3 Variable Specific Impulse Magnetoplasma Rocket The Variable Specific Impulse Magnetoplasma The current technical challenges are Subsystem: continue cryo-cooler testing Rocket (VASIMR) is a high power electric space pro- the cryogen-free low-temperature and characterization for use with supercon- pulsion engine capable of Isp/thrust modulation at superconducting magnet, the mag- ducting magnets, trade 60kW pumped fluid constant input power. Plasma is produced by heli- netic nozzle performance and life- system (uses ISS PVTCS pump with Down-

con discharge with energy added by ion cyclotron time (from sputtering) and the high therm-J) performance against titanium-H2O resonant heating (ICRH). Axial momentum is ob- temperature heat rejection system Loop Heat Pipe system (interdependency tained by adiabatic expansion of plasma in a mag- required to extend the pulse length. with TA14), complete verification plan for netic nozzle. Thrust/specific impulse ratio control Future technical challenges include 200kW Nautel amplifer. achieved by partitioning RF power and propellant specific mass, heat rejection and System: Test VF-200 in a large vacuum cham- flow between helicon and ICRH systems. The most thermal control, interactions of the ber (e.g., Space Power Facility) to collect advanced VASIMR prototype to date is the VX-200 divergent plume with the spacecraft, plume characteristics data for additional device that uses 35 kW helicon source to generate and life qualification. modelling validation. Measure & quantify plasma, 170 kW ICRH section to heat plasma, RF nature of VF-200 plume to ensure com- and an adiabatic magnetic nozzle to produce exit pliance with NASA spaceflight hardware plume. The VX-200 leverages commercially devel- requirements. Perform testing of the VF- oped solid-state RF generators with efficiencies 200 system to meet qualification & accep- of 98 % & specific mass less than 1 kg/kW. It uses tance environmental testing requirements. a cryogen-free low-temperature supercon-ducting Demonstrate operation of VF-200 on ISS magnet to produce fields approaching 1 Tesla re- at 200kW power level for 15-min firing du-

quired for helicon and ICRH sections. The VX-200 ration, ~5 N thrust, Isp of 4000-6000 secs, has been operated at power levels of up to 200 kW and efficiency of ~50-55%. Collect on-orbit

& with performance data estimates of 5,000 sec Isp plume characteristics data. and 60% efficiency for 5 sec pulses. 2.2.1.4 Micropropulsion 2.2.1.4.1 Microresistojets See description on resistojets earlier. Microresisto- Conventional butane-based microre- Butane thrusters are at TRL 5-6 and ready for jets are smaller versions of conventional resistojets sistojet thrusters require flight demo. flight demonstration. For MEMS based de- and generally have an overall lower TRL because of For MEMS based designs, there are vices, demonstration is needed of the scala- a lack of flight hardware development and demon- issues associated with microfabri- biity to very small sizes, microfabrication of stration. Small thruster systems based on “liqui- cation techniques, wall losses, and thruster components, and verification of fied gas” (e.g., butane) have been developed for thermal design of the smaller thrust- thruster performance and life. Inspector spacecraft and Cubesat applications and er volumes. Additionally, the small are ready to be flight demonstrated. MEMS based nozzle dimensions generally reduce resistojet thrusters have also been developed pri- the efficiency of these devices and marily for small-sat applications. designing MEMS based miniature flow-control valves has proven dif- ficult. 2.2.1.4.2 Microcavity Discharge, Teflon Microcavity discharge thrusters are similar to ar- Challenges are MEMS-fabrication Demonstrate scalability to very small sizes, cjects (see description above), except that the di- techniques, reliability of arc elec- microfabrication of thruster components, mensions are much smaller and the TRLs are lower trodes, array design, and thermal is- and verification of thruster performance (no flight development to date). MEMS fabrication sues. and life. has been used along with smaller more conven- tional techniques. In these exceedingly small scaled down versions of arc-jets, the arc is concentrated in a very small volume at lower power to maintain the life of the single thruster, or configured in arrays one thruster can be used if another fails.

TA02-12 DRAFT 2.2 Nonchemical Propulsion Description and State of the Art Technical Challenges Milestones to TRL 6 2.2.1.4.3 Micropulse Plasma Pulsed plasma thrusters (PPTs) use capacitively or Miniaturization of PPT technology Demonstrate scalability to very small sizes, inductively coupled plasma discharge to acceler- must focus on mass reduction of the microfabrication of thruster components, ate conductive gas, liquid, or solid at high powers supporting power electronics (in and verification of thruster performance (kW to MW in discharge) for very short pulses (<1 particular on lighter capacitors and and life. msec). Typical PPTs ablate teflon using a surface switches). The thruster itself faces discharge to provide the propellant. Pulsing allows challenges during miniaturization in average power consumption to be very low but the tailoring of discharge energy to requires energy storage and switching. Small PPTs the decreased fuel rod cross section- were first flown onboard Russian Zond 2 spacecraft al area. If the discharge energy be- in 1964 and several other missions. PPTs offer very comes too small, carbon neutrals in small impulse bits, solid propellant storage, modu- the plasma arc can return to the fuel larity, and proven operation. Recent advances in surface & result in charring, which ul- PPT miniaturization have been achieved. The Vac- timately can lead to shorting of the uum Arc Thruster (VAT) is another ablative pulse thruster electrodes. plasma type that uses metal electrodes and an arc discharge. A VAT module has been developed for Cubesat. 2.2.1.4.4 Miniature Ion/Hall These are scaled down versions of ion and Hall In addition to the challenges of re- Develop microfabrication techniques for the thrusters described earlier. Miniature ion and Hall ducing the size of engines (associ- thruster and subsystem components, scal- engines have recently been developed for forma- ated with lower efficiencies due to ability to very small sizes, and demonstrate tion flying applications of future space telescopes higher surface losses and voltage subsystem performance, life and integration (the 100-W NASA MiXI ion thruster) and for small breakdown in small gaps), other re- with spacecraft in demonstration missions. satellites orbit maneuvering (the 200-W = Minia- quired subsystems have to be equally ture Hall thruster). Use of inert, noncontaminating miniaturized including the feed sys- xenon propellant near sensitive optical surfaces, as tem and the power processing unit well as ability to smoothly modulate thrust ampli- (PPU). Off-the shelf solutions feed tude have made miniature electric engines attrac- systems and power supplies used tive. Ion Thruster On a Chip (ITOC) concepts have in conventional thrusters could be been investigated that included subcomponent adapted for micro-ion/Hall engines developments including microfabricated accel- based on present performance. A erator grids and field emission based cathodes for key challenge to the thruster design discharge generation and neutralization. A 4-cm is to scale appropriately the applied diameter RF ion thruster was developed in Europe magnetic fields in the smaller sized and is slated for upcoming ESA earth-observing ion and Hall thrusters, and the devel- missions, and miniature ion thrusters based on opment of miniature hollow cathode Electron Cyclotron Resonance (ECR) plasma pro- technology. duction at frequencies are under devel- opment in Japan and in US universities. 2.2.1.4.5 MEMS Electrospray Electrospray thrusters use a conductive fluid Challenges are MEMS fabrication Milestones to TRL 6 include flight of the ST-7/ and electrostatic fields to extract and acceler- of large tip arrays, microfluidics for LISA Pathfinder that will demonstrate col- ate charged droplets, clusters of molecules, and/ distributing propellant to all the loid thruster performance, demonstration or individual molecules or ions. The ion-emission tips, cleanliness, power-processing, of a MEMS fabricated prototype array, and pointed-tips are on the order of microns which thruster testing, and lifetime. To date, performance measurements, flight dem- lends to MEMS based fabrication to produce large the challenges of demonstrating onstration and successful life test of MEMS arrays of emitters. The SOA in electrospray thrust- microfabcation and microfluidics to fabricated arrays. ers is a 5-30 uN precision thruster (TRL6) planned feed the arrays with propellant in the for stabilization and solar pressure compensation, laboratory have prevented a space- which is awaiting a flight demonstration on ST7/ qualifiable protoype demonstration. Interferometer Space Pathfinder. Other designs are under development. Using MEMS techniques, closely packed emitter tips are used to decrease the system volume while increas- ing the number of emitters (and hence the thrust). MEMS fabricated electrospray thrusters in develop- ment in US laboratories and in Europe are at TRL 3, and have demonstrated high efficiency (>80%). This high efficiency removes many thermal issues, and the lack of a plasma discharge or confinement requirements simplifies the construction and volt- age holdoff.

DRAFT TA02-13 2.2 Nonchemical Propulsion Description and State of the Art Technical Challenges Milestones to TRL 6 2.2.2 Solar Sail Propulsion Solar sails are large, lightweight reflective struc- System level integration and test Due to the constraints of gravity, solar sail tures that produce thrust by reflecting solar pho- of the component technologies for propulsion performance will not be totally tons and thus transferring much of their momen- >1,000-m2 sail using existing mate- demonstrated to TRL 6 on the ground. A tum to the sail. The state-of-the-art solar sails rials and technologies are needed. space flight demonstration will be required were produced for the NASA In Space Propulsion Gravity offload and deployment tests to fully achieve TRL 6. 20 meter Ground System Demonstrations (GSD) in of the large sail are needed. 2005. The JAXA funded Interplanetary Kite-craft Accelerated by Of the Sun (IKAROS), launched in May, 2010, deployed its sail in June and has since demonstrated both photon acceleration and attitude control. IKAROS has a square sail that is approximately 14 meters long on each side, 7.5 micrometers thick and used a spin deployment method to deploy its sail. 2.2.3 Thermal Propulsion 2.2.3.1 Solar Thermal Solar Thermal Propulsion (STP) heats the propel- Challenges are optical concentrator Perform experiments with inflatable or rigid lant with concentrated inside an absorber accuracy and performance (from 50- concentrators to increase the optical perfor- cavity and provides a very high specific impulse 60% to 85-90%), system/stage pack- mance to 85-90% efficiency. Demonstrate (~500–1200 seconds). The is concen- aging, sun pointing (sub-arcsec accu- the pointing capabilities of dual concentra- trated and focused inside either a direct gain or racy in flat, 1 cm by 1 cm packages), tors to keep the focuses on target during thermal storage type engine configuration. The inflatable deployment, controlled all expected spacecraft orientations with solar concentrator is may be rigid, segmented or cryogenic boil-off, and engine per- the sun. Demonstrate the propulsion per- inflatable. The engine would be fabricated from formance. An integrated overall sys- formance of a high temperature carbide ultra high temperature materials and operated as tem test has never been performed. thruster manufactured with modern materi- a heat exchanger with the propellant. A variety of STP is limited by payload shroud al processing techniques, which allow com- propellants have been considered (e.g., hydrogen, volume when considering liquid hy- plex geometries to be fabricated. Ground

methane, ammonia). The thrust level for this sys- drogen LH2. Options to overcome test the performance of more dense propel- tem would be on the order of 1.0 lbf. this hurdle include the use of high lants (e.g., methane, ammonia). In addition, The L’Garde flight experiment in 1996 demonstrat- temperature carbides with melt- demonstrate the utilization of deployable

ed the deployment of a large inflatable concentra- ing point ~4200K. At temperatures LH2 tanks. tor (TRL6). The 30-day LH2 storage with controlled above 3200K and pressures ~25 psia. boil-off was demonstrated in 1998-1999 (TRL5). Various engine concepts have been made and fabricated from Rhenium, Tungsten/Rhenium, and Rhenium coated (TRL4). A new Ebeam manufacturing process has been demonstrated to fabricate complex STP engine designs. In addition, a new ultra-high temperature material (tri-carbide)

has the potential to allow greater Isp than 1000 sec- onds.

2.2.3.2 Nuclear Thermal Nuclear Thermal Rockets are a high thrust, high The NERVA program matured the Complete fuel tests – select primary & back- Isp propulsion technology. The state of the art technology to a TRL 6 level in the up fuel/element design; design Ground & ground demonstrated engine, Nuclear Engine for 1970s. The current challenge is to Flight Technology Demonstration engines; Rocket Vehicle Applications (NERVA) demonstrated capture the engineering and tech- complete borehole gas injection tests - de- (in the 1970's) comparable to chemical pro- nical knowledge base of the NERVA tailed design of ground test facility begins – pulsion (7,500 to 250,000 lbs of thrust with specific program. New fuel elements with “Authority to Proceed” (ATP); complete small impulses of 800 to 900 seconds, double that of longer life is another technical chal- (5 klbf) ground technology demonstration chemical rockets). with solid-core NTR en- lenge to be addressed by future ef- engines tests; conduct 5 klbf NTP Flight gines have been considered for human missions to forts. Technology demonstrator (FTD) mission; Mars as their high Isp allows reductions of the initial develop engine scale-up design for crewed mass in (IMLEO) from 12 to 7 heavy Mars missions. lift with 200 metric ton payloads.

TA02-14 DRAFT 2.2 Nonchemical Propulsion Description and State of the Art Technical Challenges Milestones to TRL 6 2.2.4 Tether Propulsion 2.2.4.1 Electrodynamic An electrodynamic tether (EDT) can work as a The critical EDT subsystems have A space flight validation of an EDT propul- thruster because a magnetic field exerts a force on a been tested both in space and in the sion system will be required to fully achieve current-carrying wire. This force is perpendicular to laboratory to TRL 5/6. To field this TRL 6. the wire and to the field vector. It is this interaction technology, a full-scale (multi-kilo- that allows relatively short electrodynamic tethers meter, multi-kW) space validation at to use to “push” against a planetary the system level is required. The pri- magnetic field to achieve propulsion without - ex mary challenge for EDT propulsion penditure of propellant. The groundwork has been is at the system level, not at the sub- laid for this type of propulsion. Important achieve- system level. Improvements to bet- ments include retrieval of a tether in space (TSS- ter preclude arcing and mechanism 1, 1992), successful deployment of a 20-km-long deployment design can be pursued, tether in space (SEDS-1, 1993), and operation of an but are not necessary for flight dem- EDT in space with tether current driven in both di- onstration. rections (PMG, 1993). The Propulsive Small Expend- able Deployer System (ProSEDS) experiment was to use an EDT to achieve ~0.4 N drag thrust, thus deorbiting the stage. JAXA demonstrated in space the operation of a 300-m uninsulated tape tether anode (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 the laboratory. 2.2.4.2 Momentum Exchange A spinning, or (MET) There are numerous subsystem and Very long (up to 100+ km), high strength-to- system can be used to boost payloads into higher component-level engineering chal- weight conducting & nonconducting teth- with a Hohmann-type delta-V or as an "el- lenges remaining before this capabil- ers survivable in the LEO micrometeoroid, evator" between two different orbits. A tether sys- ity reaches TRL-6: The development orbital debris, atomic oxygen & EUV envi- tem would be anchored to a relatively large mass in of long-life, survivable conducting ronment. LEO awaiting rendezvous with a payload delivered and nonconducting tethers, deploy- Reliable wire and/or tape tether deployers to orbit by a launch vehicle. The uplifted payload ers, no expellant and cath- with reel-in and reel-out capability meets with the tether facility, which then begins odes, and a robust tether dynamics Alternative anodes and cathodes with mini- a slow spin-up using electrodynamic tethers (for modeling capability are chief among mal or no expellants (Field Emitter Array propellantless operation) or another low thrust, them. Cathodes, Solid Expellants, etc.) high Isp thruster. At the proper moment and tether Flywheel energy storage (crossover to space system orientation, the payload is released into a power team) transfer orbit – potentially to geostationary trans- High-res, long-tether dynamics modeling fer orbit (GTO) or Earth Escape. Most of the achieve- to allow precise rendezvous between tether ments cited for EDT propulsion establish the state tip and payload. of the art for MET systems. The missing system-lev- High-res, long-tether dynamics modeling to el validations are spin-up, multi-spacecraft inter- allow precision spin-up (if required, Rotova- action, orbital transfer dynamics, and rendezvous tor only) orbital transfer of payloads. and docking with the tether tip. Rendezvous & docking of payload & tether tip for rapid payload transfer between launch vehicle (orbital or suborbital) & teth- er tip.

DRAFT TA02-15 2.3. Advanced Propulsion Technologies Advanced Propulsion Technologies are those that use chemical or nonchemical physics to pro- duce thrust, but are generally considered to be of lower technical maturity (TRL< 3) than those de- scribed in Sections 2.1.1 and 2.1.2. is often used in conjunction with In-Space Propul- sion to provide the required mission Δv, but does not directly influence or impact In-Space Propul- sion technologies discussed here. AeroGravity As- sist (AGA) is covered in TA09.

2.3 Advanced (TRL <3) Propulsion Technologies Description and State of the Art Technical Challenges TRL Maturation 2.3.1 Beamed Energy Propulsion Beamed energy propulsion uses laser or microwave Development of MW Free Electron Demonstrate mode using energy from a ground or space based energy source . liquid, gaseous, or Delrin ablation propel- and beams it to an orbital vehicle which uses it to heat a Development of novel optics/track- lant for in space maneuvers. propellant, with the advantage being high exit velocity ing and pointing systems for orbit of exhaust products over traditional chemical propul- transfers. sion. Earth-to-Orbit technology has Propellant feeds or ablative propel- been investigated both analytically and experimentally lants will also need to be technically as a first step to orbital transfer. In space applications addressed. to be demonstrated are orbit transfer and earth escape. Development of efficient capture Other in-space applications could be to de-orbit orbital and transformation of beamed en- debris by way of ablation. ergy into propulsive energy (e.g., heat exchangers, direct plasma breakdown in propellant).

2.3.2 Propulsion Consists of a number of thin, long, and conducting Quantification of thrust magnitudes Validate physics models. wires that are kept in a high positive potential by an with on-orbit data. Develop system level performance mod- onboard electron gun. The positively charged wires Demonstration of noninterfering els. repel protons, thus deflecting their paths centrifugal deployment of multiple Develop control laws for attitude control and extracting momentum from them. Simultaneously wires from a single spacecraft. using multiple wire anodes. they also attract from the solar wind plasma. Validation of current collection and Perform subscale space flight validation A way to deploy the wires is to rotate the spacecraft electrostatic propulsion from the (outside of the ). and have the keep them stretched. By solar wind. fine-tuning the electrical potentials of individual wires Validation of electrostatic attitude and thus the solar wind force individually, the attitude control in the solar wind. of the spacecraft can also be controlled. Deployment of multikilometer length wires in space has been demon- strated (see electrodynamic tether propulsion). Elec- tron guns have also been flown in space. Other techni- cal approaches to achieve electrostatic propulsion from the solar wind include the superconducting magsail and Mini-Magnetospheric Plasma Propulsion (M2P2), but none of these have yet been demonstrated; all pro- pulsive effects have been only predicted in theory and modeling. 2.3.3 Fusion Propulsion Fusion propulsion involves using fusion reactions to Creation of a sustained fusion reac- Develop plasma thruster concept capa- produce the energy required for the spacecraft propul- tion that can drive a plasma thruster ble of efficiently converting high-energy, sion. This can be accomplished either indirectly (with a with a specific mass low enough charged particle fusion products into pro- fusion reactor producing electrical power that is in turn (alpha < 4) to be competitive with pellant energy. utilized in an electric thruster), or directly, by using the advanced fission is the primary Demonstrate plasma thruster concept on thermal/kinetic energy resulting from the fusion reac- challenge. Production of a positive the ground in space-like simulated envi- tions to accelerate a propellant. This is accomplished energy output with Deuterium-Tri- ronment. either by creating a hot, thermal plasma that is then tium reactions has yet to be demon- Perform testing and validation of engine expelled through a magnetic nozzle to provide thrust strated even in ground-based Toka- technology. (in the same manner as in a plasma thrusters) or using mak reactor concepts. Production high-energy, charged particle, fusion products to cre- of a thermal plasma suitable for an ate the hot, thermal plasma in the thrust chamber. The electric thruster from high-energy physics and related technologies are is still under inves- fusion products (such as would tigation at the laboratory scale level. A gain (energy come from an aneutronic fusion re- out of the reaction to energy into the reaction) of ap- actor) is needed. proximately 1 has been achieved, but for useful fusion propulsion, a gain of 100 to 1000 is needed.

TA02-16 DRAFT 2.3 Advanced (TRL <3) Propulsion Technologies Description and State of the Art Technical Challenges TRL Maturation 2.3.4 High Energy Density Materials 2.3.4.1 Metallic Hydrogen Metallic hydrogen is a theoretically dense energetic Upgrading existing experimental Demonstrate synthesis of metallic hydro- material (not yet produced on earth). The TRL level is equipment is required for synthesis gen in lab. not at level 1 as the characteristics are based on theo- and characterization of small quan- Evaluate characteristics of metallic hydro- retical calculations. The estimated density at ambient tities of metallic hydrogen. Scal- gen in lab. conditions is 7 g/cc, 10 times LH2. Above a critical tem- ing up production by many orders Develop production scaling techniques. perature, possibly 1000 K, metallic hydrogen will be- of magnitude is required. Engine Develop engine components and test come unstable and recombine to the molecular phase, components must be developed various diluents. releasing the energy of recombination, 216 MJ/kg (for that are compatible with metallic Perform propellant tankage develop- reference: H2 + O2 in the SSME releases 10 MJ/kg, LO2/ hydrogen. Test engines must be ment. RP1 releases 6 MJ/kg). Ongoing experiments are using developed to verify expected op- Perform tests of various engine sizes and diamond anvil cells and short pulse laser technologies erations and performance with a diluents. to follow the hydrogen melt line toward the conditions variety of diluents and mixture ra- for the metallic state. Expected Isp values are in the tios. Potential need for tankage that 500-2000 secs range. operates at millions of psi. 2.3.4.2 Atomic Boron/Carbon/Hydrogen trapped in solid cryogens (neon, etc.) at 0.2 to 2 Storage of atoms at 10, 15, or 50 Formulate storage methods for weight percent. Atomic hydrogen, boron, and carbon weight percent is needed for effec- high density. fuels are very high energy density, free-radical propel- tive propulsion. Develop engine designs for recombining lants. Atomic hydrogen may deliver an Isp of 600 to propellants. 1,500 secs. There has been great progress in the im- Perform testing and validation of engine provement of atom storage density over the last several designs. decades. Lab studies have demonstrated 0.2 & 2 weight percent atomic hydrogen in a solid hydrogen matrix. If the atom storage were to reach 10–15 percent, which would produce a specific impulse (Isp) of 600–750 secs. 2.3.4.3 High Nitrogen Compounds (N4+, N5+) These are the most powerful explosives created in his- The propellants are highly shock Perform inhibitor research to facilitate tory. Work was conducted under the High Energy Den- sensitive. Challenges include fab- safe scaling. sity Materials (HEDM) Program. Gram quantities for- rication, transportation, ground Develop high-speed deflagration/deto- mulated in laboratory (1999). Theoretical studies have processing, and personnel safety nation engine technology. shown that these materials may have in-space propul- to name a few. Presently, there are Perform testing and validation of engine sion applications. no integrated vehicle designs that technology. can make use of this possible pro- pellant. 2.3.5 Antimatter Propulsion Antimatter propulsion is based on conversion of a large The next step is experimental dem- Develop and demonstrate proof-of- percentage (up to ~75%) of fuel mass into propulsive onstration of these propulsion con- concept experiment to verify producing, energy by annihilation of atomic particles with their cepts. This requires a significant controlling, and exhausting annihilation/ antiparticles. Creation, manipulation, storage and anni- source of antimatter available for reaction products. hilation of picogram amounts of antimatter is routine at engineering research. Current pro- Develop and demonstrate thruster con- high-energy physics laboratories such as Fermi Lab. In duction rates of antiprotons are not cept. addition, very small amounts of positrons are routinely sufficient for any known propulsion Develop and demonstrate full-scale en- created, manipulated, stored and annihilated at various applications, but could be scaled gine in vacuum chamber. university labs and in hospitals. Low energy storage of up by several orders of magnitude. Develop and demonstrate long-term small amounts of positrons and antiprotons has been Portable storage of antiprotons storage and transportation systems for demonstrated at research institutions. Many antimatter needs to be developed. antiprotons. propulsion concepts have been explored analytically over the years.

DRAFT TA02-17 2.3 Advanced (TRL <3) Propulsion Technologies Description and State of the Art Technical Challenges TRL Maturation 2.3.6 Advanced Fission 2.3.6.1 Gas Core In gas-core rocket, radiant energy is transferred from An open-cycle engine relies on flow Demonstrate effective solutions to tech- high-temperature fissioning plasma to hydrogen pro- dynamics to control fuel loss. With nical challenges in simulated non-nuclear pellant. Propellant temperature can be significantly both open and closed cycle con- laboratory environment. higher than engine structural temperature. In some de- cepts, cooling engine walls is a ma- Demonstrate breadboard system in non- signs, propellant stream is seeded with submicron par- jor engineering problem. It is not nuclear environment. ticles (~20% weight fraction) to enhance heat transfer. known whether practical means Demonstrate breadboard cavity reactor Both open-cycle and closed-cycle configurations have can be devised to contain nuclear with sustained fissioning gas core. been proposed. Radioactive fuel loss and its effect on fuel, adequately cool the cham- Demonstrate breadboard propulsion sys- performance is a major problem with open-cycle con- ber, and achieve useful thrust to tem driven with fissioning gas core. cept. Cavity reactors for critical gas core assemblies are weight characteristics. Experiments Perform testing and validation of engine known to be achievable. To date, small scale, non-nu- to demonstrate effective contain- technology. clear gas flow tests have been performed. Evaluation of ment of simulated fuel particulate critical physics and engineering aspects for integrated and effective transfer of heat from open cycle systems have been limited to computa- simulated fuel particulate cloud tional studies. Closed cycle nuclear bulb concept, to flowing hydrogen are needed. whereby energy from hot reactant gas is transferred Closed-cycle concept avoids fuel to hydrogen propellant via radiation transfer through loss problem but requires develop- a transparent wall, has been successfully studied in a ment of adequate transparent wall laboratory environment to limited extent. By increasing material and effective coupling of temperature of fissioning gas, it is possible to achieve radiation heat transfer between hot full and formation of a plasma. Fissioning reactant gas and hydrogen propel- plasma offers superior performance potential but will lant. Open-cycle fuel loss must be require some type of electromagnetic containment limited to less than one percent of system much like those encountered for magnetic fu- the total flow.

sion systems. Isp values of 1500-2500 secs have been estimated. 2.3.6.2 Fission Fragment In a fission fragment engine, the thrust results from the The first challenge is to perform Develop and test concept to show that expulsion of high velocity (~5% light speed) nuclear fis- more detailed system studies to large portion of fission fragments are di- sion fragments produced from a high power nuclear better understand the system-level rected before interacting with surround- reactor through a magnetic nozzle. Several concepts technologies required for this ap- ings. have been proposed to directly use this fission energy proach to be viable. Because the Demonstrate critical reactor configura- as part of an extremely high specific impulse rocket en- fission fragments are very highly tion.

gine (Isp ~50,000+ secs). In one concept, thin wires of a charged, they have a very short Develop and demonstrate magnetic field fissionable material are brought into a critical state and mean free path in most materials. configuration to maximize thrust. the fission fragments collected as they exit the wire. A As such, they are very difficult to somewhat similar concept utilizes uranium foils where- collect before they interact with in the fission fragments are collected as they exit the their surroundings, and thus lose foil. A third concept recently proposed uses a "dusty" their energy before they can be plasma core wherein the uranium dust particles are expelled through a nozzle. In addi- suspended in a container and brought into a critical tion, their extremely high energy state. A magnetic field is used to contain the fuel and requires extremely high magnetic at the same time extract the thrust producing fission fields in order to direct them out of fragments. the nozzle to produce useful thrust. 2.3.6.3 External Pulsed Plasma Propulsion (EPPP) EPPP was first studied in 1950s and early-1960s for ARPA Some recent ideas have pointed Develop practical concept that relies on and then NASA. Known as Project Orion, the system to approaches that could mitigate decoupling of initiator/driver mechanism employed small nuclear bombs to provide thrust via a nuclear proliferation and radiation (e.g., laser, plasma guns) from fissile/fu- large pusher plate at the rear of the spacecraft. More issues. sionable fuel/propellant. Develop accu- advanced versions of EPPP have been considered since rate computational analyses and proof- then involving fissile fuel pellets compressed via laser of-principle demonstrations to assess or particle beams, and other concepts involving com- operation of key processes (e.g., driver, bined fission and fusion reactions. EPPP can achieve fuel/driver material interaction, fuel en-

both high average accelerations (1–2 g) and high Isp ergy release/thermalization process and (~10,000–100,000 s), making it well suited for rapid in- thrust production). Demonstration of terplanetary spaceflight. integrated subscale system operation in secure facilities, leading to TRL 6.

TA02-18 DRAFT 2.3 Advanced (TRL <3) Propulsion Technologies Description and State of the Art Technical Challenges TRL Maturation 2.3.7 Breakthrough Propulsion Breakthrough Propulsion Physics is specifically look- Challenges in this area are to de- A small sustained investment is needed to ing for propulsion breakthroughs from physics. It is velop theoretical models and high identify & support affordable, near-term, not looking for further technological refinements of fidelity laboratory experiments & credible research that will make incre- existing methods. It is an area of fundamental scientific for model verification /validation mental progress toward these propulsion research that seeks to explore and develop a deeper (coupling of gravity & electromag- goals. Prioritizing and pursuing focused understanding of the nature of space-time, gravitation, netism, vacuum fluctuation energy, research to: (1) establish if an idea has inertial frames, quantum vacuum, and other funda- warp drives & wormholes, & super- propulsion applications, (2) investigate if mental physical phenomenon with the pinnacle objec- luminal quantum effects). the effect of interest can be observed in tive of developing advanced propulsion applications the laboratory, & (3) begin engineering and systems that will revolutionize how we explore breadboard development to produce space. Past research efforts have yielded a number of the desired effect in a manner useful for publications in peer-reviewed literature detailing ap- spaceflight applications. plied theoretical models and laboratory investigation Once a concept has progressed through results/conclusions. Fundamental scientific research in these wickets, it should be ready to mi- this area is a high risk/high payoff venture. Individual grate beyond the TRL 3 level and could investigations may yield good science, but not always be recategorized as a game-changing result in a propulsion physics breakthrough. technology.

2.4. Supporting Technologies mal Management Systems TA, whereas micro- Supporting Technologies are those technolo- gravity fluid dynamics and the integration of the gies that support an in-space propulsion system thermal control and fluid management technolo- or subsystem but which are not directly propul- gies are covered within this road map. There is also sive. The supporting technology areas given signif- a need for future propulsion systems to be more icant consideration by the ISPSTA team included serviceable/maintainable as system life and reuse pervasive technologies (Integrated System Health increase, and those requirements have been treat- Management, Materials and Structures, Heat Re- ed as embedded in the individual technologies jection and Power) and cryogenic fluid manage- rather than a separate supporting technology area. ment (CFM) for propellants. For the pervasive technologies, technology gaps for propulsion ap- plication are identified in the preceding sections, embedded in the text of the individual propulsion technology supported. In each case, the technol- ogies are directly or significantly addressed by the following TAs respectively: Robotics, Tele-robot- ics, and Autonomous Systems; Materials, Struc- tures/Mechanical Systems, and Manufacturing; Thermal Management Systems; and Space Pow- er and Energy Storage Systems. For cryogenic flu- id management and transfer, the thermal control components are addressed in detail by the Ther-

2.4 Supporting Technologies Description and State of the Art Technical Challenges Milestones to TRL 6 2.4.1 Engine Health Monitoring and Safety Integrated System Health Management (ISHM) as applied to propulsion relies on the automation of interpretation, reasoning, and decision making based on data collected during the processing and operation of propulsion system to enable the following basic functionality: anomaly detection, diagnostics, prediction of future anomalies (Prognostics), and enabling intuitive and rapid integrated awareness about configuration and condition of every element in a system. Only a few health management systems have been implemented in operational mode; all are considered to provide low functional capability, and do not truly represent knowledge systems. The Main En- gine Advanced Health Management System (AHMS) monitors a selected number of engine sensors, detects threshold violations based on experimental models, and can activate redline conditions. See TA04 road map for more details.

DRAFT TA02-19 2.4 Supporting Technologies Description and State of the Art Technical Challenges Milestones to TRL 6 2.4.2 Propellant Storage, Transfer & Gauging Cryogenic Fluid Management (CFM) broadly Enable long duration (months to years) du- 1) Conduct ground tests in representative describes the suite of technologies that can ration in space missions with cryogens and thermal vacuum environments of high fidel- be used to enable the efficient in-space efficient in-space transfer of propellants for ity CFM components (partial crossover with use of cryogens despite their propensity to tanker or depot architectures. Specific chal- TA14 Thermal Management) and systems absorb environmental heat, their complex lenges to be addressed by in space demon- 3) Complete short duration cryogenic flight thermodynamic and fluid dynamic behav- stration before reaching TRL 6 include: experiment(s) to obtain data to mature ior in low gravity and the uncertainty of the - Safe and efficient venting of unsettled pro- models of critical fluid/thermal physics. position of the liquid-vapor interface since pellant tanks 4) Conduct CFM flight demonstration - Dem- the propellants are not settled. In addition - Zero boil-off (ZBO) storage of cryogenic onstrate in space long duration (>6 months)

to propulsion, this technology can support propellants for long duration missions storage of LO2 (ZBO) and LH2 (<0.5% loss/ power reactant storage and ECLSS needs. -Accurate micro-g propellant gauging and month). Demonstrate low loss in-space

The State of the Art is defined by commer- fluid acquisition transfer of LO2 and LH2 including automated cial upper stages with a capability of up to -Automated cryogenic fluid couplings & pro- fluid couplers. Demonstrate micrograv- 9 hours in space (propellant boil-off rates pellant transfer ity venting, liquid acquisition, and quantity

on the order of 30%/day) and which require -Performance issues due to integration of gauging of LO2 and LH2. (note that methane propellant settling (via thruster firing to ac- components is stored at similar temperature to LO2 and celerate the vehicle) before performing criti- therefore can use similar technology). cal functions. Many of the CFM technology elements have been matured to a TRL of near 5 through ground testing. 2.4.3 Materials & Manufacturing Technologies Structures and materials play a critical role in all in-space propulsion systems for both human and robotic missions. The reader is referred to the Technology Area 12, Materials, Structures Mechanical Systems and Manufacturing technology, road map for a State of the Art descrip- tion and specific technology development recommendations. In some cases, material, structural, or manufacturing advances are required to enable propulsion technology advances, in other cases the structural/material advances are enhancing, in still others (for example composite tanks) advancements can result in a significant propulsion system improvement of their own. In general in-space propulsion desires reduced cost, lighter weight, wider operating temperatures, and robustness against in-space and propulsion system environmen- tal conditions. 2.4.4 Heat Rejection Heat rejection is a key supporting capability for several in space propulsion systems. Some examples include rejection of the waste heat generated due to inefficiencies in electric propulsion devices and rejection of the heat removed from a cryogenic propellant storage sys- tem by a cryocooler. The reader is referred to the Technology Area 14, Thermal Management Systems technology, road map for a State of the Art description and specific technology development recommendations. In general the key heat rejection system metrics for in-space propulsion are cost, weight, operating temperature, and environmental durability (e.g. radiation, MMOD). 2.4.5 Power Power systems play an integral role in all in-space propulsion systems for both human and robotic missions. The reader is referred to the Technology Area 3, Space Power and Energy Storage Systems technology, roadmap for state-of-the-art description & specific technology development recommendations. In some cases, power is only required for basic functions of instrumentation and controls and technol- ogy advances are not required, but in other cases the propulsion energy of the technology is derived from electrical power generation, management and distribution, & power system technology advances are critical for the advancement of the propulsion system (e.g., high power solar or nuclear electric propulsion). In general the key power system metrics for in-space propulsion are cost, reliability, specific power, operating ranges (power and environmental), & maximum power generation.

3. Interdependency with developed by the Space Power and Energy Stor- Other Technology Areas age road map to support high power solar elec- The ISPSTA team evaluated the technical con- tric propulsion. In the third type of dependency, tent of the other fourteen technology area road another TA road map is supported by technolo- maps for relationships with the content of the gy developed by ISPSTA. An example of this case ISPSTA roadmap. In completing this evaluation, is development of throttleable cryogenic propul- three types of interdependencies were identified sion in the IPSTA road map supporting the Entry, in which either ISPSTA technology development Descent and Landing TA road map. For all three was related to technology development in another types of relationships, significant benefit will be road map, as illustrated in table below. In some cas- realized through coordination of the implement- es, propulsion technology development is planned ing technology development projects. in both road maps and close synergies could were identified. In other cases, ISPSTA technology de- 4. Possible Benefits to velopment depends on successful technology de- Other National Needs velopment in another TA. For example, solar pow- More capable and efficient in-space propul- er array and power management technologies are sion will benefit NASA, national defense, and the TA02-20 DRAFT commercial space industry – virtually any orga- IKAROS Interplanetary Kite-Craft nization that builds or uses space satellites. Spe- Accelerated by Radiation Of the Sun cific technologies with multi-user applicability IMLEO Initial Mass in Low-Earth Orbit include: Metalized Propellants for higher perfor- ISHM Integrated System Health Management mance and safer missile systems; Long-Duration ISPSTA In-Space Propulsion Systems Cryogenic Propulsion for high-energy orbit trans- Technology Area fer; Electric Propulsion for longer-life communi- ISRU In-Situ Resource Utilization cations and Earth observing satellites; Solar Sails ISS International Space Station for sustained observation of the Earth’s polar re- ITOC Ion Thruster On a Chip gions; research toward Nuclear Thermal Propul- JAXA Japanese Aerospace Exploration Agency sion will lead to smaller and more efficient reactor JSC Johnson Space Center designs; Tethers will enable lower-cost access to KSC Kennedy Space Center space and orbit transfer; and the development of LST Life Support Technologies new technologies from robust research and tech- MET Momentum Exchange Tether nology development will enable new missions and MMOD Micro-Meteoroid/Orbital Debris applications for all potential users. MPD Magnetoplasmadynamic MMH Monomethylhydrazine Acronyms MSFC Marshall Space Flight Center AHMS Advanced Health Management System NERVA Nuclear Engine for Rocket Vehicle Applications AMPM Agency Mission Planning Model OF2 Oxygen Difluoride ARC PIT Pulsed Inductive Thruster ATP Authority to Proceed PPUs Power Processing Units CFM Cryogenic Fluid Management ProSEDS Propulsive Small Expendable ClF3 Chlorine Trifluoride Deployer System ClF5 Chlorine Pentafluoride RCS Reaction Control System DRM Design Reference Mission SDI Strategic Defense Initiative ECLS Environmental Control and Life Support SOA Hydrazine ECR Electron Cyclotron Resonance TRL Technology Readiness Level EDT Electrodynamic Tether TSS-1 Tether In Space EHS Environmental Health System VASIMR Variable Specific Impulse GRC Magnetoplasma Rocket GTO Geosta¬tionary Transfer Orbit VAT HEDM High Energy Density Materials ZBO Zero Boil-Off HmNT Hydrazine milli-Newton Thruster HTPB Hydroxyl-Terminated Polybutadiene ICRH Ion Cyclotron Resonant Heating

DRAFT TA02-21 Reference Missions Frisbee, Robert H.: “Advanced Space Propulsion for the 21st Century,” Journal of Propulsion and Power, Vol. 19, No. 6, November– December 2003, Jet Propulsion Laboratory. Hoffman, Stephen J.; and Kaplan, David I. (eds): NASA Special Publication 6107, Human Exploration of Mars: The Reference Mission of the NASA Mars Exploration Study Team, Lyndon B. Johnson Space Center, , TX, July 1997. Human Exploration of Mars Design Reference Architecture 5.0, NASA SP—2009–566, July 2009. Human Exploration Framework Team (HEFT) DRM Review - Phase 1 Closeout, Steering Council, , September 2. 2010. Johnson, L.; Farris, B.; Eberle, B.; Woodcock, G.; and Negast, B.: “Integrated In-Space Transportation Plan,” NASA/CR—2002– 212050, October 2002. “NASA Exploration Team (NEXT), Design Reference Missions Summary,” FOR INTERNAL NASA USE ONLY, National Aeronautics and Space Administration, Draft, July 11, 2002. Balint, Tibor S.: “Design Reference Mission Set for RPS Enabled Missions in Support of NASA’s SSE Roadmap,” Jet Propulsion Laboratory, California Institute of Technology, IEEE Aerospace Conference (IEEEAC) Paper #1461, Version 2, updated December 5, 2006.

Acknowledgements The draft NASA technology area roadmaps were developed with the support and guidance from the Office of the Chief Technologist. In addi- tion to the primary author’s, major contributors for the TA02 roadmap included: the OCT TA02 Roadmapping POC, Jill Prince; the NASA Center Chief Technologist and NASA Mission Director- ate reviewers, and the following individuals Ron Reeve, David Jones, Kay Glover, Nancy Miecz- kowski.

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NASA Headquarters Washington, DC 20546 www..gov TA02-24 DRAFT