Table of Contents MAP Propulsion System Thermal Design Carol L. Mosier, NASA Goddard Space Flight Center, Greenbelt, MD Development of a Thermal Control Architecture for the Mars Exploration Rovers Keith S. Novak, Charles J. Phillips, Gajanana C. Birur, Eric T. Sunada and Conference on Thermophysics in Microgravity Michael T. Pauken, Jet Propulsion Laboratory, Pasadena, CA Active Heat Rejection System on Mars Exploration Rover - Design A01. Fundamentals of Two-Phase Flow and Heat Transfer in Changes from Mars Pathfinder Microgravity Gani B. Ganapathi, Gajanana C. Birur, Glenn T. Tsuyuki, Paul L. McGrath and Jack D. Patzold, Jet Propulsion Laboratory, Pasadena, CA Thermal Design Overview of NASA's Next Generation Space Tutorial on Quantification of Differences between Single- and Two- Telescope Component Two-Phase Flow and Heat Transfer Keith Parrish, Shaun Thomson and Stuart Glazer, NASA Goddard Space A.A.M. Delil, National Aerospace Laboratory NLR, Emmeloord, The Flight Center, Greenbelt, MD Netherlands Electric Field Effect on Bubble Detachment in Variable Gravity A04. Miscellaneous Topics on Thermophysics in Microgravity Environment Estelle Iacona, The Johns Hopkins University, Baltimore, MD; and Laboratoire EM2C du CNRS et de l'Ecole Centrale Paris, Chatenay-Malabry, Across-Gimbal and Miniaturized Cryogenic Loop Heat Pipes France; Cila Herman and Shinan Chang, The Johns Hopkins University, D. Bugby, B. Marland, C. Stouffer, and E. Kroliczek, Swales Aerospace, Baltimore, MD Beltsville, MD Superluminal Signals Test Cell for a Novel Planar MEMS Loop Heat Pipe Based on A.A. Stahlhofen, University of Koblenz, Koblenz, Germany Coherent Porous Silicon Debra Cytrynowicz, Mohammed Hamdan, Praveen Medis, H. Thurman A02. Thermal Control Technologies for Future Spacecraft Henderson, Frank M. Gerner, University of Cincinnati, Cincinnati, OH Zero-Gravity Test Results For Ultrasonic Sensing of Air-Liquid Solar Selective Coatings for High Temperature Applications Interface in a Vortex Separator Donald A. Jaworske, NASA Glenn Research Center, Cleveland, OH; R. Williams, Incipient Systems, Inc., Los Alamos, NM; I. Carron, C. Dean A. Shumway, Brigham Young University-Idaho, Rexburg, ID Kurwitz, and F. Best, Texas A&M University, College Station, TX; D. Bray, Sorption Heat Pipe - A New Thermal Control Device for Space Don Bray and Associates, Inc., College Station, TX Applications Experimental Investigation on Wetting of a Binary Volatile Sessile Leonard L. Vasiliev, Leonid L. Vasiliev, Luikov Heat and Mass Transfer Drop Institute, Minsk, Belarus K. Sefiane, University of Edinburgh, Edinburgh, UK; L. Tadrist, Comparison Between Acetone and Ammonia on the Thermal Laboratoire IUSTI, Marseille, France Performance of a Small-Scale Capillary Pumped Two-Phase Loop Roger R. Riehl, Edson Bazzo, Federal University of Santa Catarina, A05. Variable Emittance Technology for Spacecraft Thermal Florianopolis, Santa Catarina, Brazil Control Development Status of the Mechanically Pumped Two-Phase CO2 Cooling Loop for the AMS-2 TTCS Variable Emittance Materials Based on Conducting Polymers for A.A.M. Delil, A.A. Woering, National Aerospace Laboratory NLR, Spacecraft Thermal Control Emmeloord, The Netherlands; B. Verlaat, National Institute for Nuclear Prasanna Chandrasekhar, Brian J. Zay, Terrance McQueeney, David A. Physics and High Energy Physics NIKHEF, Amsterdam, The Netherlands Ross, Andre Lovis, Ashwin-Ushas Corporation, Lakewood, NJ; Rengasamy Ponappan, Air Force Research Laboratory, Wright-Patterson AFB, OH; A03. Thermal Control for Deep Space Missions Charlotte Gerhart, Air Force Research Laboratory, Kirtland AFB, NM; Theodore Swanson, Lonny Kauder, Donya Douglas, Wanda Peters, NASA 1 2 Goddard Space Flight Center, Greenbelt, MD;Gajanana C. Birur, Jet Park, MD Propulsion Laboratory, Pasadena, CA Bubble Behavior in Subcooled Pool Boiling of Water under Electrostatic Appliqué for Spacecraft Temperature Control Reduced Gravity William Biter, Stephen Hess, Sung Oh, Sensortex, Inc., Kennett Square, Koichi Suzuki, Motohiro Suzuki, Saika Takahash, Hirosi Kawamura, PA Tokyo University of Science, Chiba, Japan; Yoshiyuki Abe, The National All-Solid-State Electrochromic Variable Emittance Coatings for Institute of Advanced Industrial Science and Technology, Ibaragi, Japan Thermal Management in Space A Fundamental Study Regarding the Control Of Nucleate Boiling Nikolai Kislov and Howard Groger, Eclipse Energy Systems, Inc., St. in a Complex Magnetizable Fluid By an Petersburg, FL; Rengasamy Ponnappan, Air Force Research Laboratory, Applied Magnetic Field, in Microgravity Conditions Wright-Patterson AFB, OH Floriana D. Stoian, Gheorghe Pop, Virgil Stoica, Oana Marinica, Controlling Variable Emittance (MEMS) Coatings for Space "Politehnica" University of Timisoara, Timisoara, Romania; Doina Bica and Applications Ladislau, Romanian Academy-Timisoara Branch, Timisoara, Romania D. Farrar, W. Schneider, R. Osiander, J.L. Champion, A.G. Darrin, Johns Hopkins University Applied Physics Laboratory, Laurel, MD; D. Douglas and A08. High Temperature Heat Pipe Technology T.D. Swanson, NASA Goddard Space Flight Center, Greenbelt, MD SAFE-100 Module Fabrication and Test Peter J. Ring and Edwin D. Sayre, Advanced Methods and Materials, A06. Two-Phase Thermal Control Systems: Space Applications Sunnyvale, CA; J.Tom Sena, Los Alamos National Laboratory, Los Alamos, and Flight Experiments NM Thermal Vacuum Testing of a Novel Loop Heat Pipe Design for the SAFE-100 Module Processing Methodology Swift BAT Instrument Patrick G. Salvail and Robert W. Carter, NASA Marshall Space Flight Laura Ottenstein, Jentung Ku, NASA Goddard Space Flight Center, Center, Huntsville, AL Greenbelt, MD; David Feenan, Swales Aerospace, Beltsville, MD Transient Thermohydraulic Heat Pipe Modeling: Incorporating Development and Test Results of a Multi-Evaporator-Condenser THROHPUT into the CAESAR Environment Loop Heat Pipe Michael L. Hall, Los Alamos National Laboratory, Los Alamos, NM Yury F. Maydanik, Vladimir G. Pastukhov, Mariya A. Chernyshova, Ural SAFE Alkali Metal Heat Pipe Reliability Branch of the Russian Academy of Sciences, Ekaterinburg, Russia; and Robert S. Reid, Los Alamos National Laboratory, Los Alamos, NM Ad A.M. Delil, National Aerospace Laboratory NLR, Emmeloord, The Netherlands Development of a Two-Phase Capillary Pumped Heat Transport Conference on Commercial/Civil Next for Spacecraft Central Thermal Bus Generation Space Transportation Triem Hoang, TTH Research, Capitol Heights, MD; Michael Brown, Robert Baldauff, and Sheila Cummings, U.S. Naval Research Laboratory, Washington, DC B01. Cost Analysis Tutorial Flight Testing of the Capillary Pumped Loop 3 Experiment Cost Analysis Tutorial Laura Ottenstein, Dan Butler, Jentung Ku, NASA Goddard Space Flight Eric Shaw, NASA Marshall Space Flight Center, Huntsville, AL Center, Greenbelt, MD; Kwok Cheung and Robert Baldauff, Naval Research Laboratory, Washington, DC; Triem Hoang, TTH Research, Capitol Heights, B02. Market and Finance MD A Current Summary of RLV Activities in the U. S. Sam K. Mihara, Mihara Associates, Huntington Beach, CA A07. Boiling in Microgravity A Market 2 Come: On-Orbit Servicing of Satellites Joerg Kreisel, JOERG KREISEL INTERNATIONAL CONSULTANT Effect of Non-Ionic Surfactants on Nucleate Pool Boiling Space Technology and Intellectual Property: Funding Exploration J. P. Kizito, R. Balasubramanaim, M.J. Boggess, NASA Glenn Research Through Technology Commercialization Center, Cleveland, OH; K.J. Stebe, Johns Hopkins University, Baltimore, MD William N. Hulsey III, Hughes & Luce, LLP, Austin, TX Heater Size and Gravity Effects on Pool Boiling Heat Transfer Evolving Markets for Commercial, Civil, and Military Services Jungho Kim and Christopher Henry, University of Maryland, College Marshall H. Kaplan, Strategic Insight, Ltd., Rockville, MD

3 4 Geoffrey A. Landis, NASA Glenn Research Center, Cleveland, OH; Vincent Denis, International Space University, Strasbourg, France B03. Public Space Travel and Tourism Intelligent Launch and Range Operations Testbed Public Space Markets - What We Know and What We Don't Rodney D. Davis, Kevin R. Brown, Command and Control Technologies, Derek Webber, Futron Corporation, Bethesda, MD Titusville, FL Potential Effects of Government Regulations on Public Space Travel B07. Space Launch Initiative Systems Harvey A. Wichman, Claremont McKenna College, Claremont, CA Boeing Space Launch Initiative Program Overview Developing the Space Destinations – Near Term Possibilities Dallas Bienhoff, The Boeing Company, Huntington Beach, CA Paola Favata, University Frienza, Italy Selection of Lockheed Martin's Preferred TSTO Configurations Creating thee Future Space Tourism Movement for the space Launch Initiative Tsuyoshi Saotome, Crystal Space Place, Inc., Japan Joshua B. Hopkins, Lockheed Martin Astronautics, Denver, CO 2nd Generation RLV: Overview of Concept Development Process B04. Insuring the Enterprise and Results Risk Mitigation in the Space Industry Mark Benton, Jim Berry, Harry Benner Northrop Grumman Corporation, El. Patricia Maloney, The Aerospace Corporation, Los Angeles, CA Segundo, CA, The Mechanics of Space Insurance Overview of Orbital Sciences Corporation’s Orbital Space Plane Isabel Passoa-Lopes (OSP) for the Space Launch Initiative TBD Adrienne E. Wasko, Orbital Sciences Corporation, Dulles, VA Stephen C. Leonard, International Space Brokers, Rosslyn, VA Experiences In Securing Insurance For A Reusable Rocket: B08. Space Launch Initiative Technology Implications For Commercial RLVs Systems Engineering Approach to Technology Integration for Jeffrey K. Greason, XCOR Aerospace, Mojave, CA NASA’s 2nd Generation Reusable Launch Vehicle Sheryl Kittredge, Dale Thomas, Marc Verhage, Charles Smith, Leann B05. Spaceport Developments Thomas, NASA Marshall Space Flight Center, Huntsville, AL New Mexico's Commercial Space Program The Successful Development of an Automated Rendezvous and Louis Gomez, New Mexico Office of Space Commercialization, Santa Fe, Capture (AR&C) System for The National Aeronautics and Space NM Administration The Road to a Spaceport Fred D. Roe and Richard T. Howard, NASA Marshall Space Flight Jay T. Edwards, Oklahoma Space Industry Development Authority, Center, Huntsville, AL Oklahoma City, OK Advanced Checkout, Control, and Maintenance System (ACCMS) Functions and Requirements in the Advanced Spaceport Breakthrough Technologies Environment Cary Peaden, NASA Kennedy Space Center, KSC, Florida Roelof L. Schuiling, NASA Kennedy Space Center, FL Next Generation Launch Technologies Program Propulsion Activities Overview B06. Spaceport Technology Shayne Swint, NASA Marshall Space Flight Center, Huntsville, AL Launch System Testbed: An Innovative Approach for Design and Development of Future Launch Structures B09. Standards and Operations Bruce T. Vu, NASA Kennedy Space Center, FL; Max Kandula and Ravi Water Detection and Removal From Shuttle Tiles Margasahayam, Dynamics Inc., Kennedy Space Center, FL; Danielle M. Ford, Robert C. Youngquist, NASA Kennedy Space Center, FL Embry-Riddle Aeronautical University, Daytona Beach, FL Standards and Certification Processes for Reusable Space Electrochemical Evaluation of Alloys for Spaceport Design Transportation Development of NASA Technical Standards Luz Marina Calle and Louis G. MacDowell, NASA Kennedy Space Program Relative to Enhancing Engineering Capabilities Center, FL;Rubiela D. Vinje, Dynacs, Inc., Kennedy Space Center, FL Paul S. Gill, NASA Marshall Space Flight Center, Huntsville, AL; High Altitude Launch for a Practical SSTO William W. Vaughan, University of Alabama, Huntsville, AL

5 6 Albuquerque, NM; R.R. Siergiej, B. Wernsman, & S.A. Derry, Bechtel Bettis 20th Symposium on Space Nuclear Power and Inc., West Mifflin, PA The Status of Thermophotovoltaic Energy Conversion Technology Propulsion at Lockheed Martin Corp. E.J. Brown, P.F. Baldasaro, S.R. Burger, L.R. Danielson, D.M. DePoy, C01. Advanced Concepts G.J. Nichols, W.F. Topper, Lockheed Martin Corp., Niskayuna, NY A Nuclear-Powered Laser-Accelerated Plasma Propulsion System Improved Thermophotovoltaic (TPV) Performance Using Terry Kammash, University of Michigan, Ann Arbor, MI Dielectric Photon Concentrations (DPC) Fusion Ship II- A Fast Manned Interplanetary Space Vehicle Using P.F. Baldasaro and P.M. Fourspring, Lockheed Martin Corp., Niskayuna, Inertial Electrostatic Fusion NY R.L. Burton, N. Richardson, Y. Shaban, G.H. Miley, University of Illinois, Radioisotopic Powered Thermophotovoltaic Energy Systems Urbana, IL; H. Momota and G.H. Miley, NPL Associates, Inc., Champaign, IL D.M. DePoy, E.J. Brown, P.F. Baldasaro, L.P. Rice, Lockheed Martin Ion Dynamic Capture Experiments with the High Performance Corp., Niskayuna, NY Antiproton Trap (HiPAT) James Martin, Suman Chakrabarti, William H Sims, J Boise Pearson, C04. Special Session: Voyager Revisited NASA Marshall Space Flight Center, Huntsville, AL; Raymond Lewis, R. Overview of Voyager Missions Lewis Company, Huntsville, AL; and Wallace E Fant, Cortez III, Huntsville, John Casani, Jet Propultion Laboratory, Pasadena, CA AL Overview Lincoln Experimental Satellite (LES) 8/9 Missions On the Performance Prediction and Scale Modelling of a Donald C. Maclellan, MIT Lincoln Laboratory, Lexington, MA Motorised Momentum Exchange Propulsion MHW-RTG (Multi-Hundred Watt Radioisotope Thermoelectric Tether Generator) Matthew P. Cartmell and David S. Neill, University of Glasgow, Glasgow, Charles E. Kelly, Lockheed Martin Corp., King of Prussia, PA Scotland; Spencer W. Ziegler, UMIST, Manchester, England Panel Discussion

C02. Fission Propulsion Systems for Science Missions C05. Mission / Systems Safety and Reliability Nuclear Propulsion Requirements for Science Missions Solid Rocket Propellant Fire Characterization Test Results L. Dudzinski , NASA Headquarters, Washington DC, WA L.W. Hunter, Y. Chang, H.N. Oguz and S.C. Walts, John Hopkins MITEE-B: A Compact Ultra Lightweight Bi-Modal Nuclear University, Laurel, MD Propulsion Engine for Robotic Planetary A Method for the Analysis of Impact Events Involving Nested Science Missions Safety Systems James Powell, George Maise, and John Paniagua, Plus Ultra James R. Coleman, Consulting, Cross, SC Technologies, Inc., Shoreham, NY; Stanley Borowski, NASA Glenn Research A Brief Discussion of Uncertainty as It Relates to Space Nuclear Center, Cleveland, OH Safety Analyses Design and Development of the MITEE-B Bi-Modal Nuclear James R. Coleman, Consulting, Cross, SC Propulsion Engine Nuclear Systems Initiative: Implementing a Managed Approach to John C. Paniagua, James R. Powell and George Maise, Plus Ultra Risk Communication Technologies, Inc., Stony Brook, NY Victoria P. Friedensen, NASA Headquarters, Washington, D.C.; Sandra L. SUSEE – An Ultra Lightweight Nuclear Electric Space Power Dawson, NASA Jet Propulsion Lab, Pasadena, CA System Based on Conventional Steam Cycle G. Maise, J. Powell, & J. Paniagua, Plus Ultra Technologies, Inc., Stoney C06. Fuels and Advanced Materials Brook, NY High Temperature Cermet Fuels – A Promising Candidate for

Space Reactors C03. Thermophotovoltaic Technology And Applications S.K. Bhattacharyya, Argonne National Laboratory, Argonne, IL

Review of the Historical Capabilities and Testing of Composite and 20% Efficient InGaAs/InPAs Thermophotovoltaic Cells Cermet Fuels in Los Alamos S.L. Murray, C.S. Murray, & F.D. Newman, Emcore Photovoltaics, 7 8 Robert J. Hanrahan, Jr., Robert L. Smith III, Jason Morgan, Los Alamos National Laboratory, Los Alamos, Antimatter Driven P-B11 Fusion Propulsion System NM Terry Kammash, University of Michigan, Ann Arbor, MI; James Martin, An Overview of Development and Testing of Uranium Tri-Carbide Thomas Godfroy, NASA Marshall Space Flight Center, Huntsville, AL Fuels Samim Anghaie and Travis Knight, University of Florida, Gainesville, FL C09. Low-Cost Space Exploration Missions - Lessons Learned Re-establishing Fabraication Capabilities for Space Nuclear Power Systems Implementing a New Line of Medium-Class, Unmanned Space Jeffrey Halfinger, Barry G. Miller and DeWayne L. Husser, BWX Exploration Missions Technologies, Lynchburg, VA Thomas H. Morgan and Susan M. Niebur, NASA Headquarters,

Washington, DC C07. Alkali-Metal Thermal-To-Electric Technology And Considerations and Lessons Learned in Implementing Effective, Applications - I Low-Cost, Unmanned Space Exploration Missions Development, Evaluation, and Design Applications of an AMTEC R. Brad Perry, NASA Langley Research Center, Hampton, VA; Dennis G. Converter Model Pelaccio, Science Applications International Corporation, Littleton, CO Cliff A. Spence, Michael Schuller, Tom R. Lalk, Texas A&M University, NASA's New Millennium Program: Validation of Advanced College Station, TX Technologies in Space Effect of Long Term, High Temperature Annealing on the Strength Christopher M. Stevens, Jet Propulsion Laboratory, Pasadena, CA of Beta"-Alumina Ceramics Lessons Learned in Sending an Ion-Drive System to Deep Space James R. Rasmussen, Advanced Modular Power Systems, Inc., Ann D.H. Lehman, Jet Propulsion Laboratory, Pasadena, CA Arbor, MI; Roger M. Williams and Adam K. Kisor, Jet Propulsion Laboratory, Pasadena, CA AMTEC Response to Changes in Resistive Loading C10. Fission Power Systems for Science Missions – I Robert W. Fletcher and Thomas K. Hunt, Advanced Modular Power Systems, Inc., Ann Arbor, MI Design Concept for a Nuclear Reactor-Powered Mars Rover Comparison of Measurement Techniques for Determining the John O. Elliott, Jet Propulsion Laboratory, Pasadena, CA; Ronald J. Thermal Emittance of Coupons at Elevated Temperatures Lipinski, Sandia National Laboratories, Albuquerque, NM; David I. Poston, Daniel P. Kramer, Roger G. Miller, Edwin I. Howell, BWXT of Ohio, Los Alamos National Laboratory, Los Alamos, NM Inc., Miamisburg, OH; Donald A. Jaworske, NASA Glenn Research Center, Reactor Shielding Calculations for the Mars Cryobot Lander Cleveland, OH; Kenneth E. Wilkes, Oak Ridge National Laboratory, Oak (MCL) and Mars Atomic Rover for Geographical Exploration Ridge, TN (MARGE) David I.Poston, Los Alamos National Laboratory, Los Alamos, NM C08. Fusion Space Systems And Applications – I High Efficiency Thermoelectrics in NEP Reactor Power Systems Daniel T. Allen, Saeid Ghamaty, and Norbert B. Elsner, Hi-Z Technology, Ablation Radiation Shields For Nuclear Fusion Rockets Inc., San Diego, CA Luis Coreano and Brice Cassenti, Rensselaer at Hartford and Pratt & Conceptual Design of a 100-kWe Space Nuclear Reactor Power Whitney, East Hartford, CT System with High-Power AMTEC First Results of the Gasdynamic Mirror Fusion Propulsion Mohamed S. El-Genk and Jean-Michel Tournier, University of New Experiment Mexico, Albuquerque, NM William J. Emrich, Jr., NASA Marshall Space Flight Center, Huntsville, AL C11. Fission Power Surface Systems for Human Missions A Design Study of a p-11B Gasdynamic Mirror Fusion Propulsion Power Needs For Planetary Surface Exploration System S. Hoffman and J. George, NASA Johnson Space Center, Houston, TX Chad Ohlandt, Terry Kammash, Kenneth G. Powell, University of Review of Power System Options for Human Planetary Michigan, Ann Arbor, MI Exploration

9 10 Robert Cataldo, NASA Glenn Research Center, Cleveland, OH hydraulic Simulator Robert W. Carter, Ray M. Guffee, Los Alamos National Laboratory, Los SUSSEE- An Ultra Lightweight Nuclear Electric Space Power Alamos, NM; Russell L. Rosmait, Pittsburgh State University, Pittsburgh, KS; System Based on Conventional Steam Cycle Pat Salvail, NASA Marshall Space Flight Center, Huntsville, AL George Maise and James Powell and John Paniagua, Plus Ultra Texhnologies Inc., Stony Brook, NY C14. Thermoelectric Technology and Applications StarTram - A Highway to Space That Everyone Can Afford James Powell, Plus Ultra Technologies, Stony Brook, NY Progress in the Development of High Efficiency Segmented Thermoelectric Unicouples C12. Space Nuclear Technology - General II T. Caillat and J.P. Fleurial and J. Snyder and J. Sakamoto, California SAFE Testing Nuclear Rockets Economically Institute of Technology, Pasadena, CA Steven D. Howe, Bryan Travis and David K. Zerkle, Los Alamos National Thermal Stability Characterization of Skutterudite Antimonides Laboratory, Los Alamos, NM and Phosphides Nuclear Thermal Rocket Exhaust Conditioning In Open Cycle and J. Sakamoto ,V. Shields ,T. Caillat and J. P. Fleurial, California Institute Closed Cycle Systems of Technology, Pasadena, CA Stanley V. Gunn, Rocketdyne-Retired, Chatsworth, CA Potential Improvements in Skutterudite Thermoelectric Properties Technologies to Improve Ion Propulsion System Life and due to Solid Solution Formation Efficiency V. Shields, T. Caillat, J. P. Fleurial, A. Zoltan, L. Zoltan and M. Ira Katz, John R. Brophy, John R. Anderson, James E. Polk, Jet Tuchscherer, California Institute of Technology, Pasadena, CA Propulsion Laboratory, Pasadena, CA Engineering Nanostructures for Efficient Thermoelectric Power Fusion Propulsion Through a Magnetic Nozzle and Open Divertor Conversion Craig H. Williams, Ian J. Dux, NASA Glenn Research Center, Cleveland, G. Chen, Massachusetts Institute of Technology, Cambridge, MA OH; Pavlos G. Mikellides, Arizona State University, Tempe, AZ; Ioannis G. Mikellides, Science Applications International Corporation, San Diego, CA; Richard A. Gerwin, Los Alamos National Laboratory, Los Alamos, NM C15. Space Nuclear Systems Development Testing - Lessons Learned

C13. Nuclear System Development and Testing Progress/Results Looking Back on Rover C. Paul Robinson, Sandia National Laboratories, Albuquerque, NM Test Facilities in Support of High Power Electric Propulsion Nuclear Thermal Rocket – An Established Space Propulsion Systems Technology Milton Klein, NASA/AEC (retired), Menlo Park, CA Melissa Van Dyke, Mike Houts, Thomas Godfroy, Ricky Dickens, James J. Martin, Patrick Salvail, and Robert Carter, NASA Marshall Space Flight Looking Backward, Looking Forward Center, Huntsville, AL Gary L. Bennett, Emmett, ID Thermal Stress Calculations for Heatpipe-Cooled Reactor Power Nuclear Thermal Rocket Ground Test Potential at NASA Glenn's Systems Plum Brook Station Richard J. Kapernick, Ray M. Guffee, Los Alamos National Laboratory, Brian P. Willis, Mark R. Woike, NASA Glenn Research Center/Plum Los Alamos, NM Brook Station, Sandusky, OH Direct-Drive Gas-Cooled Reactor Power System: Concept and Preliminary Testing C16. Alkali-Metal Thermal-To-Electric Technology and S.A. Wright, R.J. Lipinski, Sandia National Laboratories, Albuquerque, Applications – II NM; T.J. Godfroy, S.M. Bragg-Sitton, M.K. Van Dyke, NASA Marshall Space Flight Center, Huntsville, AL Recent Developments in Mixed Ionic and Electronic Conducting Mechanical Design and Fabrication of a SAFE-100 Heat Electrodes for the Alkali Metal Thermal Electric Converter Exchanger for Use in NASA's Advanced Propulsion Thermal- (AMTEC) 11 12 Robert W. Fletcher, Advanced Modular Power Systems, Inc., Ann Arbour, Stirling Convertor Performance Mapping Test Results for Future MI; Robert W. Fletcher and Johannes W. Schwank, University of Michigan, Radioisotope Power Systems Ann Arbor, MI Songgang Qiu, Allen A. Peterson, Franklyn Faultersack, Darin L. High Power AMTEC Converters for Deep-Space Nuclear Reactor Redinger, and John E. Augenblick, Stirling Technology Company, Kennewick, Power Systems WA Mohamed S. El-Genk and Jean-Michel Tournier, University of New Mexico, Albuquerque, NM C19. Space Nuclear Systems Technology Demonstration Missions Design Optimization of High-Power, Liquid Anode AMTEC Jean-Michel Tournier and Mohamed S. El-Genk, University of New Mexico, Albuquerque, NM Nuclear Electric Propulsion Design Factors For Deep Space Recent Developments in Mini-Electrode Test Cell Testing Robotic Missions J.R. Rasmussen and T.K. Hunt, Advanced Modular Power Systems, Inc., Joe Bonometti, Eric Stewart, Jeff Dilg, Larry Kos and Lee Mason, Ann Arbor, MI Marshall Space Flight Center, Huntsville, AL; Gary Langford, Glenn Research Center, Cleveland, OH The Impact of Mission Performance Requirements on the C17. Dynamic Energy Conversion Technology And Applications - I Development of an Early-Flight Space Fission Reactor Technology Development for a High Performance Brayton Cycle David Poston, Los Alamos National Laboratory, Los Alamos, NM Engine Hardware Based Technology Assessment in Support of Near-Term Paul Gill, Charles T. Kudija, Patrick E. Frye Missions Technology Development Program for an Advanced Potassium Mike Houts, Melissa Van Dyke, Tom Godfroy, James Martin, Shannon Rankine Power Conversion System Compatible with Several Space Bragg-Sitton, Ricky Dickens, Pat Salvail, Eric Williams, Roger Harper, Ivana Reactor Designs Hrbud, Robert Carter, NASA Marshall Space Flight Center, Huntsville, AL Bruce B. Bevard and Graydon L. Yoder, Oak Ridge National Laboratory, A 100-kWt NaK-Cooled Space Reactor Concept for an Early- Oak Ridge, TN Flight Mission Technology Concept for a Near-Term Closed Brayton Cycle Power David Poston, Los Alamos National Laboratory, Los Alamos, NM Conversion Unit John Foti, Dave Halsey, Tim Bauch and Glen Smith, Hamilton C20. High-Power Electric Propulsion – I Sundstrand, Rockford, IL A Closed Brayton Power Conversion Unit Concept for Nuclear Electric Propulsion for Deep Space Missions MPD Thruster Performance Analytic Models Claude Russell Joyner II, Bruce Fowler, John Matthews, Pratt-Whitney, James Gilland, Geoffrey Johnston, NASA Glenn Research Center, West Palm Beach, FL Cleveland, OH High Power Nuclear Electric Systems for Fast Outer Planet C18. Radioisotope Power Systems – I Missions Ben Donahue, The Boeing Company, Huntsville, AL, Micheal Cupples and Shaun Green, SAIC, Huntsville, AL Energy Conversion Options for Advanced Radioisotope Power Recent VASIMR Accomplishments Systems F. Chang-Diaz, Jared P. Squire,Timothy Glover, Andrew Petro, Mohamed S. El-Genk, University of New Mexico, Albuquerque, NM Verlin Jacobson, and Andrew Ilin, NASA Johnson Space Center, Houston, The Art of Dynamic System Testing - As Applied to a Stirling TX; Roger Bengtson, Boris Breizman, The University of Texas at Convertor Austin, Austin, TX; Wallace Baity, Richard Goulding and Mark Songgang Qiu, Stirling Technology Company, Kennewick, WA Carter, The Oak Ridge National Laboratory, Oak Ridge, TN; Oleg Continuing Develpoment for Free-Piston Stirling, Space Power Batischev, Massachusetts Institute of Technology, Cambridge, MA; Systems Allen A. Peterson, Songgang Qiu, Darin L. Redinger, John E. Augenblick, Roderick Boswell, Australian National University, Canberra, Australia; and Stephen L. Petersen, Stirling Technology Company, Kennewick, WA

13 14 MW-Class Thruster Experiments at NASA GRC Advanced Thermionic Converter Technology Program Michael R. LaPointe, Ohio Aerospace Institute, Cleveland, OH James R. Luke, New Mexico Institute of Mining and Technology, Albuquerque, NM C21. Dynamic Energy Conversion Technology And Applications-II Effects of Collector Temperature on the Performance of Grooved Electrode Thermionic Converters Yoichi Momozaki and Mohamed S. El-Genk, The University of New Advanced 35 W Free-Piston Stirling Engine for Space Power Mexico, Albuquerque, NM Applications Application of Electrometer Technology to Materials Evaluation J. Gary Wood and Neill Lane, Sunpower, Inc., Athens, OH for Future Planetary Spaceports Developments in Turbo-Brayton Power Converters C.I. Calle and E.E. Groop, NASA Kennedy Space Center, FL; C.R. Mark V. Zagarola, Chistopher J. Crowley, and Walter L. Swift, Creare Buhler, Swales Aerospace, Merritt Island, FL; J.G. Mantovani, Florida Institute Incorporated, Hanover, NH of Technology, Melbourne, FL; A.W. Nowicki, Dynacs, Inc., Kennedy Space Thermoacoustic Space Power Converter Center, FL Emanuel Tward, Michael Petach, TRW, Redondo Beach, CA; Scott Backhaus, Los Alamos National Laboratory, Los Alamos, NM C24. Dynamic Energy Conversion Technology And Applications - Reliability Assessment Approach for Stirling Convertors and III Generators Experimental Results from a 2 kW Brayton Power Conversion Ashwin R. Shah, Sest, Inc., Middleburg Heights, OH; Jeffrey G. Unit Schreiber, Edward Zampino and Timothy Best, NASA Glenn Research Center, Cleveland, OH David Hervol, Analex Corporation, Cleveland, OH; Lee Mason, Arthur Birchenough, NASA Glenn Research Center, Cleveland, OH C22. Potential Future Science Missions Initial Tests of a Thermoacoustic Space Power Engine Scott Backhaus, Los Alamos National Laboratory, Los Alamos, NM Mission Concept for a Nuclear Reactor-Powered Mars Cryobot NASA GRC Stirling Technology Development Overview Lander Lanny G. Thieme and Jeffrey G. Schreiber, NASA Glenn Research John O. Elliott, Jet Propulsion Laboratory, Pasadena, CA; Ronald J. Center, Cleveland, OH Lipinski, Sandia National Laboratories, Albuquerque, NM; David I. Poston, Los Alamos National Laboratory, Los Alamos, NM Component-Level Dynamic Modeling and Test Results for a A Fission Powered Mars Telecommunications Orbiter Mission Stirling Convertor Songgang Qiu and Allen A. Peterson, Stirling Technology Company, Concept Kennewick, WA Erik N. Nilsen, Jet Propulsion Laboratory, Pasadena, CA

Reaching the Outer Planets with Nuclear Electric Propulsion: C25. SP-100 Lessons Learned Trades and Sensitivities and the case for a Neptune System SP-100 Program Overview and Lessons Learned Explorer Stirling Bailey, Lockheed Martin, Los Gatos, CA Muriel A. Noca, Robert C. Moeller, Jet Propulsion Laboratory, Pasadena, Fuel Development and Testing for SP-100 CA James Stephen, GE Nuclear, Los Gatos, CA NEPTranS; A Shuttle-Tended NEP Interplanetary Transportation Component Fabrication Lessons Learned From The SP100 System Nuclear Space Power Program John O. Elliott, Roy Y. Nakagawa, Thomas R. Spilker, Jet Propulsion Laboratory, Pasadena, CA; Ronald J. Lipinski, Sandia National Laboratories, Edwin Sayre, Advanced Methods and Materials, Los Gatos, CA Albuquerque, NM; David I. Poston, Los Alamos National Laboratory, Los Conductively Coupled TE Cell Development for SP-100 Alamos, NM; Dean W. Moreland, NASA Johnson Space Center, Houston, TX Jaime Reyes and L.E. DeFillipo, Lockheed Martin, King of Prussia, PA

C23. Space Nuclear Technology - General C26. High-Power Electric Propulsion - II Technology for Space Reactor Applications Hydrodynamic Electrode Model for MPD Thruster F.W. Wiffen and S.J. Zinkle, Oak Ridge National Laboratory, Oak Ridge, Subrata Roy, Kettering University, Flint, MI TN Large Carbon-Carbon Grids for High Power, High Specific

15 16 Impulse Ion Thrusters C29. Fission Power Systems for Science Missions - II Jay Polk, John Brophy, California Institute of Technology, Pin-Type Gas Cooled Reactor for Nuclear Electric Propulsion Pasadena, CA; Vince Rawlin, George Williams, NASA Glenn Steven A. Wright and Ronald J. Lipinski, Sandia National Laboratories, Research Center, Cleveland, OH. Albuquerque, NM End-to-End Demonstrator of the Safe Affordable Fission Engine Liquid Metal Cooled Reactor for Space Power (SAFE) 30: Power Conversion and Ion Engine Operation Abraham Weitzberg, Potomac, MD Ivana Hrbud, Melissa Van Dyke, Mike Houts, NASA Marshall Space The Design of a Nb1Zr SAFE-400 Space Fission Reactor Flight Center, Huntsville, AL; Keith Goodfellow, Jet Propulsion Laboratory, David I. Poston, Los Alamos National Laboratory, Los Alamos, NM Pasadena, CA Design Development Analyses in Support of a Heatpipe-Brayton Numerical Simulations of the Pulsed Inductive Thruster Cycle Heat Exchanger Pavlos G. Mikellides, Arizona State University, Tempe, AZ Brian Steeve, Melissa Van Dyke, Alok Majumdar, Dalton Nguyen, Marshall Space Flight Center, Huntsville, AL; Melissa Corley, Stanford C27. Dynamic Energy Conversion Technology And Applications - University, Stanford, CA; Ray M. Guffee, Richard J. Kapernick, Los Alamos IV National Laboratory, Los Alamos, NM

Overview of NASA Multi-Dimensional Stirling Convertor Code Development and Validation Effort Combined C & D Roy C. Tew and James E. Cairelli, NASA Glenn Research Center, CD1. Fission Propulsion Systems for Human Missions - I Cleveland, OH; Mounir B. Ibrahim, Cleveland State University, Cleveland, Revolutionary Concepts for Human Outer Planet Exploration OH; Terrence W. Simon, University of Minnesota, Minneapolis, (HOPE) MN; and David Gedeon, Gedeon Associates, Athens, OH Patrick A. Troutman, Fred Stillwagen, NASA Langley Research Center, Development of a Dynamic, End-to-End Free Piston Stirling Hampton, VA; Kristen Bethke, Princeton University, Princeton, NJ; Darrell L. Convertor Model Caldwell, Jr., Shawn A. Krizan, Analytical Mechanics Associates, Inc., Timothy F. Regan, Sest, Inc., Middleburg Heights, OH; Scott S. Gerber, Hampton, VA; Ram Manvi, Jet Propulsion Laboratory, Pasadena, CA; Chris Zin Technologies, Inc., Brookpark, OH; Mary Ellen Roth, NASA Glenn Strickland, Swales Aerospace, Hampton, VA Research Center, Cleveland, OH "Bimodal" Nuclear Thermal Rocket (BNTR) Propulsion for Radiator Concepts for Nuclear Powered Brayton Conversion Artificial Gravity HOPE Mission to Callisto Systems Stanley K. Borowski, Melissa L. McGuire, Lee M. Mason, NASA Glenn Devarakonda Angirasa, SEST, Inc., Cleveland, OH; Lee S. Mason and Research Center, Cleveland, OH; James H. Gilland, Ohio Aerospace Institute, Richard K. Shaltens, NASA Glenn Cleveland, OH; Thomas W. Packard, Analex Corporation, Cleveland, OH Research Center, Cleveland, OH Bimodal Nuclear Electric Propulsion for Human Missions to the Concepts for a Capillary-Pumped Heat Engine Solar System R.B. Williams, Los Alamos National Laboratory, Los Alamos, NM L. Dudzinski, NASA Glenn Research Center, Cleveland, OH High Power MPD Nuclear Electric Propulsion (NEP) for Artificial C28. Radioisotope Power Systems - II Gravity HOPE Missions to Callisto Radioisotope Power Systems for NASA Space Science Missions Melissa L. McGuire, Stanley K. Borowski, Lee M. Mason, NASA Glenn Therese M. Griebel, Raynor Taylor, NASA Headquarters, Washington, Research Center, Cleveland, OH; James Gilland, Ohio Aerospace Institute, DC; Bernard Edwards, NASA Goddard Space Flight Center, Greenbelt, MD; Cleveland, OH Steve Oleson, NASA Glenn Research Center, Cleveland, OH Coated Particle Fuel Compact-General Purpose Heat Source for CD2. Fission Propulsion Systems for Human Missions - II Advanced Radioisotope Power Systems Mohamed S. El-Genk and Jean-Michel Tournier, University of New High Power Nuclear Electric Propulsion (NEP) for Cargo and Mexico, Albuquerque, NM Propellant Transfer Missions in Cislunar Space Turotial: Safety and Launch Approval Robert D. Falck and Stanley K. Borowski, NASA Glenn Research Center, Lyle Rutger, US Department of Energy, Germantown, MD Cleveland, OH Use of Liquid Rocket Engine Technologies on a Hybrid Nuclear

17 18 Propulsion and Power System Update on an Electromagnetic Basis for Inertia, Gravitation, the Russell Joyner, Pratt and Whitney, West Palm Beach, FL Principle of Equivalence, Spin and Particle Mass Ratios Deep Space Propultion Requirements Development Bernard Haisch, California Institute for Physics and Astrophysics, San Melvin Bulman, GenCorp Aerojet, Sacramento, CA; Stanley Borowski, Mateo, CA; Alfonso Rueda, California State University, Long Beach, CA; L.J. NASA Glenn Research Center, Cleveland, OH Nickisch, Mission Research Corp., Monterey, CA; Jules Mollere, Henderson "24 Hours to the Moon" Using LOX-Augmented Nuclear Thermal State University, Arkadelphia, AR Rocket (LANTR) Propulsion It's All Gravity… Stanley K. Borowski, NASA Glenn Research Center, Cleveland, OH P.A. Murad, Vienna, VA Transient Interial Effects and Stationary Forces James F. Woodward, California State University, Fullerton, CA Conference on Human Space Exploration D04. Far-Term Propulsion Concepts – II D01. In-Space Habitats and Supportability The International Space Station Habitat New Experiments with Spinning Metallic Discs Patricia Mendoza Watson and Mike Engle, NASA Johnson Space Center, Konstantin Mazuruk and Richard N. Grugel, NASA Marshall Space Flight Houston, TX Center, Huntsville, AL Lessons Learned in Maintenance of the International Space Station Quantum Vehicle Propulsion William W. Robbins, Jr., NASA Johnson Space Center, Houston, TX Jerry E. Bayles, Gravitational Research, Medford, OR Machining in Microgravity Very Large Propulsive Effects Predicted for a 512 kV Rotator Graylan Vincent, University of Washington, Seattle, WA David Maker, Huntsville, AL; Glen A. Robertson, Gravi Atomic Panel Discussion Research, Madison, AL Panel Discussion D02. Science and Technology of Exploration - Humans, Machines, and Habitats D05. Radiation I - Methods in Radiation Analysis for Space Robots and Humans: Synergy in Planetary Exploration Exploration Vehicles Geoffrey A. Landis, NASA Glenn Research Center, Cleveland, OH Novel Amine-Functional Membrane for Metabolic CO2 Removal High-Speed Computational Applications for Space Radiation from Spacesuit Breathing Loop Shielding Analysis Chung-Yi A. Tsai and Xia Tang, United Technologies Research Center, John E. Nealy, Old Dominion University, Norfolk, VA; Brooke M. East Hartford, CT; Ipek Guray, Worcester Polytechnic Institute, Worcester, Anderson, Swales Aerospace, Hampton, VA; John W. Wilson and Garry D. MA; Tim Nalette and Catherine Thibaud-Erkey, Hamilton Sundstrand Space Qualls, NASA Langley Research Center, Hampton, VA Systems International, Windsor Locks, CT; C. Jeffrey Brinker, Sandia National Laboratories and The University of New Mexico, Albuquerque, NM; George Reliability Methods for Shield Design Process Xomerita, The University of New Mexico, Albuquerque, NM R.K. Tripathi and J.W. Wilson, NASA Langley Research Center, Because it is Hard: The Crucible of Space as a Source of Hampton, VA Innovation Development of Collaborative Engineering Environments for Michael D. White, Blank Rome Comisky & McCauley LLP, Washington, Spacecraft Design DC Robert C. Singleterry, Jr., F. McNeil Cheatwood, Garry D. Qualls, Panel Discussion Jaroslaw Sobieszczanski-Sobieski, John W. Wilson, NASA Langley Research Center, Hampton, VA; Bradley D. Johns, Kwok Y. Fan, Swales Aerospace, Hampton, VA; Todd A. Wareing, John McGhee, Shawn Pautz, Los Alamos D03. Far-Term Propulsion Concepts – I National Laboratory, Los Alamos, NM; Anil K. Prinja, Frederick Gleicher, University of New Mexico, Albuquerque, NM; Greg Failla, ICEM CFD A Critique of Theoretical Explanations of Gravity Shielding Engineering, Berkeley, CA Phenomena Panel Discussion R. Clive Woods, Iowa State University, Ames, IA 19 20 D06. Radiation II - Radiation Analysis for Gateway and Deep 1st Symposium on Space Colonization Space Missions

Immersive Shield Design of a Gateway Space Station Concept E01 Space Tourism – I Chris A. Sandridge, Brooke M. Anderson, NASA Langley Research Center, Hampton, VA; Aric R. Aumann, Analytical Services and Materials, Survey on the History of Space Tourism Inc., Hampton, VA Larry Ortega, Spaceflight Unlimited, Bellevue, NE Radiation Shielding Analysis for Deep Space Missions Benefit Estimation Model for Tourist Spaceflights Giovanni De Angelis, John E. Nealy, Old Dominion University, Norfolk, Robert A. Goehlich, Technical University Berlin, Berlin, Germany VA; Martha S. Clowdsley, College of William and Mary, Williamsburg, VA; Incredible Adventures Robert C. Singleterry, Ram K. Tripathi, John W. Wilson, NASA Langley Jane Reifert, Incredible Adventures, Sarasota, FL Research Center, Hampton, VA DARPA, Space Tourism and RASCAL Gateway/L1 Modeling and Radiation Analysis Preston Carter, DARPA Tactical Technology Office, Arlington, VA Brooke M. Anderson, Swales Aerospace, Hampton, VA; John E. Nealy, Old Dominion University, Hampton, VA; James R. Geffre, NASA Johnson E02. Space Tourism - II Space Center, Houston, TX; Garry D. Qualls, Pat Troutman, NASA Langley Research Center, Hampton, VA; Shawn A. Krizan, Analytical Mechanics Business Context of Space Tourism Associates, Inc., Hampton, VA Harrison H. Schmitt, Umiversity of Wisconsin-Madison, Albuquerque, Deep Space Environment and Shielding NM J.W. Wilson, NASA Langley Research Center, Hampton, VA; J.E. Nealy, The Pioneer Rocketplane XP Aircraft G. de Angelis, Old Dominion University, Norfolk, VA; M.S. Clowdsley, Mitchell Burnside Clapp, Pioneer Rocketplane, Solvang, CA College of William and Mary, Williamsburg, VA; F.F. Badavi, Christopher Space Hotels Newport University, Newport News, VA Chuck Lauer Space Stations For Dummies, Part 1: Accounting For Space D07. Radiation III - Radiation Ananlysis for LEO and Planetary Christopher Lee Martens, Mutual Space, Ltd., Crestline, CA Surface Missions E03. Space Colonization - Governmental, Political, Legal And Other Issues ISS as a Platform for Environmental Model Evaluation Human Mission from Planet Earth: Technology Assessment and Craig P. Hugger, Swales Aerospace, Hampton, VA; Garry D. Qualls and Social Forecasting the Future of Space John W. Wilson, NASA Langley Research Center, Hampton, VA; Frank A. Eligar Sadeh, University of North Dakota, Grand Forks, ND Cuccinotta and Mark R. Shavers, NASA Johnson Space Center, Houston, TX; Neil Zapp, Lockheed Martin, Houston, TX Energy Policies for Habitat and Terraforming of Mars Thomas Meyer, University of Colorado, Boulder, CO; Christopher P. Space Radiation Shielding Calculation Models for LEO Satellites Mckay, NASA Ames Research Centre, Moffett Field, CA Myung-Won Shin, Myung-Hyun Kim, Kyung Hee University, Gyeongki- do, Korea The International Space Development Authority Declan J. O'Donnell, United Societies in Space, Inc., Castle Rock, CO The Martian Radiation Environment Experiment (MARIE) on the A System Design for a Compact, Renewable and Energy Efficient, 2001 Mars Odyssey Spacecraft William Atwell, The Boeing Company, Houston, TX Oxygen, Waste Recycle and Food Supply for Manned Travel to Surface Environments for Exploration Mars M.S. Clowdsley, College of William and Mary, Williamsburg, VA; G. George F. Erickson, Nanotube Engineering, Los Alamos, NM DeAngelis, Old Dominion University, Norfolk, VA; F.F. Badavi, Christopher Newport University, Newport News, VA; J.W. Wilson, R.C. Singleterry, and E04. Space Colonization - An International Planet Earth Endeavor S.A. Thibeault, NASA Langley Research Center, Hampton, VA The Next Logical Step - Post ISS Douglas A. O'Handley, Santa Clara University, Santa Clara, CA Pathways to Colonization David V. Smitherman, Jr., NASA Marshall Space Flight Center, 21 22 Huntsville, AL E07. In-Situ Space Resources Utilization (ISRU) - II A View of Future Human Colonies on Mars Robert J. Gustafson, Eric E. Rice, Daniel J. Gramer, and Brant C. White, Operation, Modeling and Analysis of the Reverse Water Gas Shift Orbital Technologies Corporation, Madison, WI Process New Concepts for Permanently Manned Lunar Bases, Report of Jonathan E. Whitlow, Florida Institute of Technology, Melbourne, FL; the Lunar Base Design Workshop, held in Noordwijk, The Clyde F. Parrish, NASA Kennedy Space Center, FL Netherlands from 10-21 June 2002 Extraction of Water from the Martian Atmosphere Barbara Imhof, Institut fuer hochbau II, Vienna, Austria; Susmita Matthew A. Schneider and Adam P. Bruckner, University of Washington, Mohanty, Moonfront LLC, San Francisco, CA; Hans Jurgen Rombaut, Lunar Seattle, WA Architecture, Utrecht, The Netherlands; Paul J. van Susante, Colorado School Solar Energy for In-Situ Resource Utilization in Space of Mines, Golden, CO; Jim Volp, ESA/ESTEC, Noordwijk, The Netherlands Takashi Nakamura, John A. Case, and Connie L. Senior, Physical Sciences Inc., San Ramon, CA High Temperature Interaction Between H2, CH4, NH3 and E05. Mars Terraforming Ilmenite Giovanni De Maria, Bruno Brunetti, Giuseppe Trionfetti and Daniela Stages in the Terraforming of Mars: the Transition to Flowering Ferro, Università La Sapienza, Roma, Italy

Plants James M. Graham, University of Wisconsin, Madison, WI E08. Space Colonization Advanced Concepts, Infrastructures, and Terraforming Mars: Can We Feed Ourselves If We Go? Architectures – I James M. Graham and Kandis Elliott, University of Wisconsin, Madison, WI Visions and Possibilities for Future Exploration of Space Terraforming Mars: The Fluorine Bottleneck and the Importance Robert A. Cassanova , NASA Institute for Advanced Concepts, Atlanta, of Sample Return GA; Ronald E. Turner, ANSER Analytic Services, Inc., Arlington, VA; and Benton C. Clark, Lockheed Martin, Denver, CO Patricia L. Russell, Universities Space Research Association, Columbia, MD Artificial Biogeochemical Cycles for Mars The Initial Nine Space Settlements Penelope J. Boston, New Mexico Institute of Mining and Technology, Anita E. Gale and Richard P. Edwards, Space Settlement Design Socorro, NM Competitions, Nassau Bay, TX Colonization of Venus E06. In-Situ Space Resources Utilization (ISRU) - I Geoffrey A. Landis, NASA Glenn Research Center, Cleveland, OH ISRU Development Strategy and Recent Activities to Support Near Implications of Outside-the-Box Technologies on Future Space and Far Term Missions Exploration and Colonization Russell S. Baird, Gerald B. Sanders, and Thomas M. Simon, NASA Theodore C. Loder III, University of New Hampshire, Durham, NH Johnson Space Center, Houston, TX Evaluation of Private Sector Roles in Space Resource Development E09. Space Colonization - Economic Drivers And Justifications Elisabeth S. Lamassoure, Ramachandra Manvi, and Robert W. Easter, Jet Propulsion Laboratory, Pasadena, CA; Brad R. Blair, Javier Diaz, and Michael B. Duke, Colorado School of Mines, Golden, CO; Mark Oderman and Marc Space Colonization--Benefits for the World Vaucher, CSP Associates, Inc., Cambridge, MA W.H. Siegfried, The Boeing Company, Huntington Beach, CA Investigation into Uses for Lunar Regolith Space Colonization - Economic Drivers and Justifications Charles Horton, Carlos Gramajo, Lance Williams, Andenet Alemu, Alex Gordon Woodcock, Gray Research, Inc., Huntsville, AL Freundlich, and Alex Ignatiev; University of Houston, Houston, TX Triggering Events for the First Space Settlement Surface Mine Design and Planning for Lunar Regolith Production Anita E. Gale and Richard P. Edwards, Space Settlement Design Leslie Sour Gertsch and Richard E. Gertsch, Michigan Technological Competitions, Nassau Bay, TX University, Houghton, MI Prospects for Revolutionary Technology for Space Colonization Edward McCullough, The Boeing Company, Huntington Beach, CA

23 24 E10 In-Situ Space Resources Utilization (ISRU) – III Darel Preble, Space Solar Power Institute, Jonesboro, GA An Interplanetary Rapid Transit System Between Earth and Mars Kerry Nock, Angus McRonald, Paul Penzo, and Chris Wyszkowski, Solar Cells Using Lunar Resources Global Aerospace Corporation, Altadena, CA; Michael Duke, Robert King, Alex Freundlich, Charles Horton, Andenet Alemu, Carlos Gramajo, Lance Lee Johnson, Colorado School of Mines, Golden, CO; Mark Jacobs and Jerry Williams, Alex Ignatiev, University of Houston, Houston, TX Rauwolf, Science Applications International Corporation, Schaumburg, IL The Development of ISRU and ISSE Technologies Leveraging Canadian Mining Expertise Dale S. Boucher, Northern Centre for Advanced Technology Inc, Sudbury, Ontario, Canada; Jim Richard, Electric Vehicle Controllers Ltd., Hanmer, Ontario, Canada; Erick Dupuis, Canadian Space Agency, St-Hubert, Quebec, Canada ISRU Reactant, Fuel Cell Based Power Plant for Robotic and Human Mobile Exploration Applications Russell S. Baird, Gerald Sanders, Thomas Simon, and Kerri McCurdy, NASA Johnson Space Center, Houston, TX Optimized ISRU Propellants for Propulsion and Power Needs for Future Mars Colonization Eric E. Rice, Robert J. Gustafson, Daniel J. Gramer, Martin J. Chiaverini, Ronald R. Teeter, and Brant C. White, Orbital Technologies Corporation (ORBITEC), Madison, WI

E11 Space Colonization Advanced Concepts, Infrastructures, and Architectures - II

Tailored Force Fields for Space-Based Construction Narayanan M. Komerath, Sameh S. Wanis, Joseph Czechowski, Georgia Institute of Technology, Atlanta, GA Distributed Power Sources for Mars Colonization George H. Miley and Yasser Shaban, University of Illinois, Urbana, IL Architecture Studies for Commercial Production of Propellants From the Lunar Poles Michael B. Duke, Javier Diaz, Brad R. Blair, Colorado School of Mines, Golden, CO; Mark Oderman and Marc Vaucher, CSP Associates, Inc., Cambridge, MA Space Colonization Using Space-Elevators from Phobos Leonard M. Weinstein, NASA Langley Research Center, Hampton, VA

E12. Space Bases

Business Approach To Lunar Base Activation Harrison H. Schmitt, University of Wisconsin-Madison, Albuquerque, NM Space Base: Design Problem Identification Laurie Barlow, L.Barlow & Co., South Pasadena, CA AQUAPLEX: An Environmentally Aware Model Lunar Settlement

25 26 Electric Field Effect on Bubble Detachment in Variable Gravity Environment Tutorial on Quantification of Differences between 1,2 1 1 Single- and Two-Component Two-Phase Flow and Heat Estelle Iacona , Cila Herman and Shinan Chang Transfer 1Department of Mechanical Engineering, The Johns Hopkins University, 3400 N. Charles Street, Baltimore, MD 21218, USA A.A.M. Delil 2 Laboratoire EM2C du CNRS et de l’Ecole Centrale Paris, Grande National Aerospace Laboratory NLR Voie des Vignes, 92295 Chatenay-Malabry, France P.O. Box 153, 8300 AD Emmeloord, The Netherlands +01-410-516-446, FAX: +01-410-516-7254, [email protected] Phone +31 527 248 229, Fax +31 527 248 210, E-mail [email protected] Abstract. The subject of the present study, the process of bubble Abstract. Single-component two-phase systems are envisaged for detachment from an orifice in a plane surface, shows some resemblance aerospace thermal control applications: Mechanically Pumped Loops, to bubble departure in boiling. Because of the high heat transfer Vapour Pressure Driven Loops, Capillary Pumped Loops and Loop coefficients associated with phase change processes, boiling is utilized Heat Pipes. Thermal control applications are foreseen in different in many industrial operations and is an attractive solution to cooling gravity environments: Micro-g, reduced-g for Mars or Moon bases, 1-g problems in aerospace engineering. In terrestrial conditions, buoyancy during terrestrial testing, and hyper-g in rotating spacecraft, during is responsible for bubble removal from the surface. In space, the gravity combat aircraft manoeuvres and in systems for outer planets. In the level being orders of magnitude smaller than on earth, bubbles formed evaporator, adiabatic line and condenser sections of such single- during boiling remain attached at the surface. As a result, the amount of component two-phase systems, the fluid is a mixture of the working heat removed from the heated surface can decrease considerably. The liquid (for example ammonia, carbon dioxide, ethanol, or other use of electric fields is proposed to control bubble behavior and help refrigerants, etc.) and its saturated vapour. Results of two-phase two- bubble removal from the surface on which they form. The objective of component flow and heat transfer research (pertaining to liquid-gas the study is to investigate the behavior of individual air bubbles mixtures, e.g. water/air, or argon or helium) are often applied to support injected through an orifice into an electrically insulating liquid under research on flow and heat transfer in two-phase single-component the influence of a static electric field. Bubble cycle life were visualized systems. The first part of the tutorial updates the contents of two earlier in terrestrial conditions and for several reduced gravity levels. Bubble tutorials, discussing various aerospace-related two-phase flow and heat volume, dimensions and contact angle at detachment were measured transfer research. It deals with the different pressure gradient and analyzed for different parameters as gravity level and electric field constituents of the total pressure gradient, with flow regime mapping magnitude. Situations were considered with uniform or non-uniform (including evaporating and condensing flow trajectories in the flow electric field. Results show that these parameters significantly affect pattern maps), with adiabatic flow and flashing, and with thermal- bubble behavior, shape, volume and dimensions. gravitational scaling issues. The remaining part of the tutorial qualitatively and quantitatively determines the differences between single- and two-component systems: Two systems that physically look similar and close, but in essence are fully different. It was already elucidated earlier that, though there is a certain degree of commonality, the differences will be anything but negligible, in many cases. These differences (quantified by some examples) illustrates how careful one shall be in interpreting data resulting from two-phase two-component simulations or experiments, for the development of single-component two-phase thermal control systems for various gravity environments.

27 28 Superluminal Signals Solar Selective Coatings for High Temperature Applications Alfons A. Stahlhofen Donald A. Jaworske1 and Dean A. Shumway2 Univ. Koblenz, Inst. f. Physik, Universitätsstr. 1, D-56070 Koblenz 0049-261-287-2341; [email protected] 1NASA Glenn Research Center, 21000 Brookpark Road, Cleveland, OH 44135 Abstract. Photonic tunneling allows for superluminal signal velocities 2Brigham Young University—Idaho, Rexburg, ID 83460 predicted before in an analysis of quantum mechanical tunneling. Since 216-433-2312, Donald.A.Jaworske@grc.nasa.gov such velocities apparently violate causality, the experimental reports of superluminal photonic tunneling times triggered a lively debate raising Abstract. Solar selective coatings are envisioned for use on many open questions which are answered in the talk. The basics of minisatellites, for applications where solar energy is to be used to superluminal signals preserving causality are sketched from the views power heat engines or to provide thermal energy for remote regions in of both information theory and physics. The role of forerunners, the the interior of the spacecraft. These coatings are designed to have the problems associated with points on nonanalyticity in a signal and the combined properties of high solar absorptance and low infrared crucial role of frequency-band and time limitation of signals are emittance. These coatings must be durable at elevated temperatures. addressed. Applications of the results range from optical networking For thermal bus applications, the temperature during operation is likely over metamaterials also realizing superluminal causal signal velocities to be near 100°C. For heat engine applications, the temperature is to the radiative exchange of heat between nanostructures which is likely to be much greater. The objective of this work was to screen dominated at short distances by photonic tunneling. candidate solar selective coatings for their high temperature durability. Candidate solar selective coatings were composed of molecular mixtures of metal and dielectric, including: nickel and aluminum oxide, titanium and aluminum oxide, and platinum and aluminum oxide. To identify high temperature durability, the solar absorptance and infrared emittance of the candidate coatings were evaluated initially, and after heating to temperatures in the range of 400°C to 700°C. The titanium and aluminum oxide molecular mixture was found to be the most durable.

29 30 Sorption Heat Pipe –A New Thermal Control Device for Comparison Between Acetone and Ammonia on the Space Applications Thermal Performance of a Small-Scale Capillary Pumped Two-Phase Loop Leonard L. Vasiliev, Leonid L. Vasiliev Roger R. Riehl, Edson Bazzo Luikov Heat and Mass Transfer Institute, P. Brovka 15, 220072, Minsk, Belarus Federal University of Santa Catarina – Mechanical Engineering Tel/Fax: 375-17-284-21-33, [email protected] Department, Florianópolis, SC 88040-900, Brazil 55 48 331-9390, Fax: 55 48 331-7615, [email protected] Abstract. Sorption heat pipe is a novelty and combines the enhanced heat and mass transfer typical for conventional heat pipe with sorption Abstract. With the continuous increase of interest for satellite and phenomena in a sorbent structure. Sorption heat pipe can be used as a structures thermal control using capillary driven two-phase loops, this heat source / sink and be applied as a heat pipe. Sorption heat pipe is paper presents an experimental investigation of a small-scale capillary insensitive to some “g” acceleration and it is suggested for space and pumped loop (CPL). Tests were performed with an internally grooved ground application. extrusion with a hollow polyethylene porous structure as the capillary evaporator, using acetone and anhydrous ammonia as working fluids. For a range of power applied to the evaporator, the system presented reliable start-ups and continuous operation for several hours for each working fluid, without a tendency of temperature overshooting. Comparing the operation of the CPL using two different working fluids, the system showed faster startups, smaller levels of superheating and consequently better heat transport capability when anhydrous ammonia was used. The capillary evaporator temperature distribution was also experimentally investigated, which showed that the proposed configuration did not present a tendency of localized superheating or dryout, which may represent a reliable thermal control system to be used when limited area is available. Preliminary tests using a capillary evaporator with sintered nickel wick showed greater difference on the temperature distribution for low heat load, which is mainly due to the metallic wick thermal conductivity characteristics.

31 32 Development Status of the Mechanically Pumped Two- MAP Propulsion System Thermal Design Phase CO2 Cooling Loop for the AMS-2 TTCS Carol L. Mosier A.A.M. Delil1, A.A.Woering1, and B. Verlaat2 NASA, Goddard Space Flight Center, Thermal Engineering Branch 1National Aerospace Laboratory NLR, Emmeloord, Netherlands Code 545, Greenbelt, MD 20771 2National Institute for Nuclear Physics and High Energy Physics 301-286-3168, Carol.Mosier@ gsfc.nasa.gov NIKHEF, Amsterdam,Netherlands +31 52 248229, Fax +31 52 248210, [email protected] Abstract. The propulsion system of the Microwave Anisotropy Probe (MAP) had stringent requirements that made the thermal design unique. To meet instrument stability requirements the system had to be Abstract. The Alpha Magnetic Spectrometer AMS is an international designed to keep temperatures of all components within acceptable experiment, led by Nobel prize laureate Samuel Ting (MIT), searching limits without heater cycling. Although the spacecraft remains at a o for anti-matter, dark matter and lost matter. It is a particle detector for fixed 22 sun angle at L2, the variations in solar constant, property high-energy cosmic rays, consisting of the sub-detectors: Tracker, Time degradation, and bus voltage range all significantly effect the of Flight (ToF) system, Veto Counters, Transition Radiation Detector temperature. Large portions of the fuel lines are external to the (TRD), Synchrotron Radiation Detector (SRD), Ring Imaging structure and all components are mounted to non-conductive composite Cherenkov Counter (RICH), Anti-Coincidence Counter, and structure. These two facts made the sensitivity to the MLI effective Electromagnetic Calorimeter. The demonstration experiment AMS-1 emissivity and bus temperature very high. Approximately two years has successfully flown in June ‘98 on the Space Shuttle Discovery prior to launch the propulsion system was redesigned to meet MAP (STS91). The paper focuses on TTCS issues. requirements. The new design utilized hardware that was already installed in order to meet schedule constraints. The spacecraft design and the thermal requirements were changed to compensate for inadequacies of the existing hardware. The propulsion system consists of fuel lines, fill and drain lines/valve, eight thrusters, a HXCM, and a propulsion tank. A voltage regulator was added to keep critical components within limits. Software was developed to control the operational heaters. Trim resistors were put in series with each operational heater circuits and the tank survival heater. A highly sophisticated test program which included “real time” model correlation was developed to determine trim resistors sizes. These trim resistors were installed during a chamber break and verified during thermal balance.

33 34 Development of a Thermal Control Architecture for the Active Heat Rejection System on Mars Exploration Mars Exploration Rovers Rover – Design Changes from Mars Pathfinder

Keith S. Novak, Charles J. Phillips, Gajanana C. Birur, Eric T. Sunada, Gani B. Ganapathi, Gajanana C. Birur, Glenn T. Tsuyuki, Paul L. and Michael T. Pauken McGrath and Jack D. Patzold

Jet Propulsion Laboratory, California Institute of Technology, Jet Propulsion Laboratory, California Institute of Technology. Pasadena, California 91109 Pasadena, California, Thermal and Propulsion Section, M/S 125-109, (818) 393-5841, [email protected] 4800 Oak Grove Drive, Pasadena, CA 91109 818-354-7449; fax: 818-393-6682; [email protected] Abstract. In June and July of 2003, the National Aeronautics and Space Administration (NASA) will launch two roving science vehicles Abstract. The active Heat Rejection System designed for Mars on their way to Mars. They will land on Mars in January and February Pathfinder was modified for the Mars Exploration Rover (Mars ’03) of 2004 and carry out 90-Sol missions. This paper addresses the mission and will be used to remove excess heat from the Rover thermal design architecture developed for the Mars Exploration Rover electronics during the cruise part of the mission. The Integrated Pump (MER) for operations on the Mars surface. The surface atmosphere Assembly design from MPF remained essentially intact; changes were temperature on Mars can vary from 0°C in the hottest part of the day to primarily made to reduce weight. However, the cooling loop was –100°C in the early morning, prior to sunrise. Heater energy usage at significantly redesigned to service totally different requirements for the night must be minimized in order to conserve battery energy. The MER rovers. In addition, the vent design was readdressed to alleviate desire to minimize nighttime heater energy led to a design in which all potentially excessive nutation as was induced on the MPF spacecraft in temperature sensitive electronics and the battery were placed inside a the process of dumping the CFC-11 overboard prior to well-insulated (carbon-opacified aerogel lined) Warm Electronics Box Entry/Descent/Landing. The current vent design was based on a better (WEB). In addition, radioisotope heater units (RHU’s) were mounted understanding of the flow characteristics during the blowdown process. on the battery and electronics inside the WEB. During the Martian day, This paper addresses some of the key design changes. This paper also the electronics inside the WEB dissipate a large amount of energy (over addresses lessons learned from the performance testing, and potential 740 W*hrs). This heat energy raises the internal temperatures inside changes to improve the HRS performance (e.g, temperature the WEB. Hardware items that have similar temperature limits were oscillations). conductively coupled together to share heat and concentrate thermal mass. Thermal mass helped to minimize temperature increases in the hot case (with maximum internal dissipation) and minimize temperature decreases in the cold case (with minimum internal dissipation). In order to prevent the battery from exceeding its maximum allowable flight temperature, paraffin-actuated passive thermal switches were placed between the battery and an external radiator. This paper discusses the design philosophies and system requirements that resulted in a successful Mars rover thermal design.

35 36 Thermal Design Overview of NASA’s Next Generation Across-Gimbal and Miniaturized Cryogenic Loop Heat Space Telescope Pipes

Keith Parrish, Shaun Thomson, Stuart Glazer D. Bugby, B. Marland, C. Stouffer, and E. Kroliczek

Thermal Engineering Branch, NASA Goddard Space Flight Center, Swales Aerospace, Beltsville, MD, 20705, USA Greenbelt, MD 20771, USA (301) 902-4385, [email protected] 301-286-3104, [email protected] Abstract. This paper describes the development status of three Abstract. Baseline configurations for NASA’s Next Generation Space advanced cryogenic loop heat pipes (CLHP) for solving important Telescope (NGST) include a multi-module science instrument package problems in cryogenic integration. The three devices described herein with near-infrared (near-IR) detectors passively cooled to below 30 K. are: (1) an across-gimbal CLHP; (2) a short transport length This integrated science instrument model (ISIM) will also house miniaturized CLHP; and (3) a long transport length miniaturized midinfrared (mid-IR) detectors that are cooled to 6-7 K with a CLHP. The across-gimbal CLHP, which is baselined for operation from mechanical cooler or stored cryogen. These complex cooling 80-100 K with nitrogen, provides a low weight, low torque, high requirements, combined with the NGST concept of a large deployed conductance solution for gimbaled cryogenic systems wishing to mount aperture optical telescope passively cooled to below 40 K, makes their cryocoolers off-gimbal. The short transport length miniaturized NGST one of the most unique and thermally challenging missions CLHP, which is baselined for operation near 35 K with neon, combines flown to date. This paper describes the current status and baseline localized thermal transport, flexibility, and thermal switching into one thermal/cryogenic systems design and analysis approach for the ISIM. device that can be directly mounted to a cryocooler cold head and a The extreme thermal challenges facing the ISIM are presented along cryogenic component just a short distance (10-20 cm) away. The long with supporting heat maps and analysis results. transport length miniaturized CLHP, which is also baselined for operation near 35 K with neon, adds to the capabilities of the short transport length miniaturized CLHP by increasing the transport length to over 250 cm to meet cryogenic heat transport device requirements of future NASA and DoD spacecraft.

37 38 Test Cell for a Novel Planar MEMS Loop Heat Pipe Zero-Gravity Test Results For Ultrasonic Sensing of Based on Coherent Porous Silicon Air-Liquid Interface in a Vortex Separator

Debra Cytrynowicz1, Mohammed Hamdan2, Praveen Medis1, 1R. Williams, 2I. Carron, 3D. Bray, 4C. Kurwitz, and 5F. Best H. Thurman Henderson1, Frank M. Gerner2 1Incipient Systems, Inc., P.O. Box 4778, Los Alamos, NM 87545 1Center for Microelectronic Sensors and MEMS, Department of 2Commercial Space Center for Engineering, TEES, Texas A&M Electrical Computer Engineering and Computer Science University, College Station, TX, 77843-3118 2Department of Mechanical, Industrial, and Nuclear Engineering, 3Don Bray and Associates, Inc., P.O. Box 10315 College Station, TX, University of Cincinnati, Cincinnati, Ohio, 45221 - 0030 77842-0315 (513) 556 – 4774, [email protected] 4Center for Space Power, TEES, Texas A&M University, Wisenbaker Building, Room 223, College Station, TX, 77843-3118 Abstract. Work towards the development of a novel, potentially high 5Nuclear Engineering Department, Texas A&M University, College power density MEMS loop heat pipe is in progress at the Center for Station, TX, 77843 Microelectronic Sensors and MEMS at the University of Cincinnati. 713-256-0552, [email protected] The design of the loop heat pipe evaporator is based upon the very novel coherent porous silicon technology, a technique in which vast Abstract. An ultrasonic pulse-echo method for detecting the location arrays of micrometer sized through-holes are photo-electrochemically of the air-liquid interface of a air-liquid vortex in microgravity was etched into a silicon wafer perpendicular to the (100) surface. The tested. A vortex was established in near-zero gravity conditions in a initial mathematical model, design, fabrication, and characterization of hydraulically-stirred phase separator. Near zero-gravity conditions the device in an open loop configuration were previously reported at were produced by flight tests of the experiment in parabolically flown this conference, STAIF 2002. This paper begins with a very brief NASA KC-135 aircraft, where 20-30 second periods of microgravity- explanation of the device and its theory of operation. Emphasis is like conditions are developed during each parabola. A single ultrasonic placed upon The design and construction of the device components by transducer located about the mid-plane of the vortex, and mounted on means of microelectronic and MEMS fabrication techniques. Recent the separator external wall, was used to sense the thickness of the layer modifications made in the photon-assisted electrochemical etch process of liquid inside the phase separator local to the ultrasonic transducer. that significantly increased the etch rate are explained. Some attention Comparison was made between the sensed liquid layer thickness and an is given to the mathematical model of the device with respect to the ideal right circular liquid annulus thickness calculated from the known generation of component dimensions through a summary of recent liquid inventory in the phase separator. The absolute error between the advances. The design and construction of the evacuated closed loop test sensed and ideal annulus thicknesses was found to be in the structure is summarized. neighborhood of 0.1 inches for the operational range of inventory of the separator.

39 40 Experimental Investigation on Wetting of a Binary Variable Emittance Materials Based on Conducting Volatile Sessile Drop Polymers for Spacecraft Thermal Control

K. Sefiane1a and L. Tadrist2 Prasanna Chandrasekhar1, Brian J. Zay1, Terrance McQueeney1, David A. Ross1, Andre Lovas1, Rengasamy Ponappan2, Charlotte Gerhart3, 1aSchool of Chemical Engineering ,University of Edinburgh Kings Theodore Swanson4, Lonny Kauder4 , Donya Douglas4, Wanda Peters4 Building, Mayfield road, Edinburgh, EH9 3JL, UK and Gajanana C. Birur5 2 Laboratoire IUSTI, UMR CNRS 6595, Technopôle de Château- Gombert - 5, rue Enrico Fermi,v13453 Marseille CEDEX 13, FRANCE 1Ashwin-Ushas Corporation, Inc. 500 James St., Unit 7, Lakewood, NJ 00 44 (0) 131 650 4873, [email protected] 08701 2Air Force Research Lab (AFRL), Wright-Patterson AFB, OH Abstract. In this paper we study the evaporation of a drop on rough 3Air Force Research Lab (AFRL), Kirtland AFB, NM 87117; substrates in PTFE and Aluminium, the drop of few millimetres size is 4NASA-Goddard Space Flight Center, Greenbelt, MD evaporating in an environment of a controlled pressure. An 5Jet Propulsion Laboratory, Pasadena, CA experimental set up is built to investigate wetting behaviour of an 732-901-9096; [email protected] evaporating binary drop on the two substrates with different roughness. The dynamic contact angle, the drop volume and the base width are Abstract. Ashwin-Ushas has developed a unique, patented Variable measured using a non-intrusive optical technique. The investigations Emittance technology based on the infrared (IR) electrochromism of are carried out for pure water and ethanol substances. The evaporation unique Conducting Polymers. This has features of: very thin (< 0.5 rate and the contact angle for each substance are found to vary in a mm), flexible, light weight (1.6 kg/m2) variable area (1 cm2 to 1 m2 ), monotonous fashion. For the binary mixtures typical behaviours of the entirely solid-state, extremely physically durable construction; measured parameters were found, revealing several stages in the switching times < 5 s at room temperature, < 3 min at –35oC; evaporation process. The wetting angle for the mixture is closely cyclabilities > 104 cycles; low power consumption (< 40 mW/ cm2); related to the drop volume as it evaporates. and most importantly, space environment durability (space vacuum and –40 oC to + 75 oC, Solar Wind, gamma radiation to 7.6 Mrad). A The evaporation rate seems to indicate that the more volatile demonstrator spaceflight is tentatively planned on NASA-Goddard’s component evaporates entirely in the first stage while in the last stage ST5 mission. This paper describes the features and current status of the the evaporation of the less volatile component prevails. technology, including results from the most recent tests. It is shown that the technology is the most promising among proposed new Variable Emittance technologies, and possibly one of the only technologies applicable to microspacecraft, besides also being applicable to large spacecraft, space based radars, and future interplanetary missions.

41 42 Electrostatic Appliqué for Spacecraft Temperature All-Solid-State Electrochromic Variable Emittance Control Coatings for Thermal Management in Space

William Biter, Stephen Hess, Sung Oh Nikolai Kislov1, Howard Groger1, and Rengasamy Ponnappan2

Sensortex, Inc. 515 Schoolhouse Road, Kennett Square, PA 19348 1Eclipse Energy Systems, Inc., 2345 Anvil Street North, St. Petersburg, (610) 444-2383, [email protected] FL 33710, USA 2Propulsion Directorate, Air Force Research Laboratory, Wright- Abstract. The electrostatically controlled radiator (ESR) uses Patterson AFB, OH 45433, USA electrostatic hold-down of a high emissivity composite film to control 1(727)344-7300, [email protected] spacecraft skin temperature. It functions as a thermal switch and changes the mode of heat transfer between the spacecraft skin and the Abstract. The thermal state of a satellite in space environment can be radiator film from conduction to radiation and has demonstrated large controlled through varying the rate of energy absorption (radiation from changes in apparent emissivity. The present device operates at high DC the sun, albedo, and earth shine), the rate of heat generation (by internal voltages and is designed with a rigid backing. An improved version, electrical circuits) and the rate of energy dissipation by thermal termed a micro-ESR, is being fabricated as an appliqué. Since the size emission from the satellite surface as infrared (IR) radiation. Small has been reduced, much lower operating voltages are possible. In light-weight satellites have reduced thermal mass and are rapidly addition, the system is conformal, allowing it to be applied to complex affected by large temperature variations resulting from changes in surfaces. This paper discusses the results from vacuum testing of the orbital conditions. In order to compensate for temperature variations, existing ESR devices. It also describes the process to form the active thermal control systems are used such as heat pipes, thermal appliqué. louvers, and heaters. These systems may not be appropriate for small satellites due to restricted payload weight and volume. All-solid-state inorganic electrochromic (EC) variable emittance coatings (VECs) are being considered as a promising technology for thermal control in the space environment.

This article demonstrates the feasibility of “all-solid-state” inorganic EC VECs for thermal management in space. VECs were built on glass substrates, flexible polyimide (KaptonTM) films, and on high resistance silicon wafers. The best VECs were found to modulate mid-infrared emittance from 0.15 to 0.46 on a Kapton film and from 0.24 to 0.48 on a silicon wafer that performs the dual role as both IR window and substrate. The results of thermal estimations for a two-dimensional plate showed that EC systems with emittance modulation ratio high/ low = 3 can be practically used to advantage in providing enhanced thermal control for lightweight structures in space. In addition, EC variable solar reflectance coatings (VSRCs) were built on polyethylene terephthalate (PET) film. The average reflection of the VSRC in a visual spectral range was 63% in the bleached condition and 20% in the colored condition. Thermal analysis shows that a combination of VECs and VSRCs described in this work provides an enhanced range of thermal control for satellites and small space vehicles in the space environment.

43 44 Controlling Variable Emittance (MEMS) Coatings for Thermal Vacuum Testing of a Novel Loop Heat Pipe Space Applications Design for the Swift BAT Instrument

D. Farrar1, W. Schneider1, R. Osiander1, J.L. Champion1, A.G. Darrin1 Laura Ottenstein1, Jentung Ku1, David Feenan2 D.Douglas2, T.D.Swanson2 1Thermal Engineering Branch, NASA/Goddard Space Flight Center, 1 Johns Hopkins University Applied Physics Laboratory Code 545, Greenbelt, MD 20771 Laurel, MD 20723 2Swales Aerospace, 5050 Powder Mill Road, Beltsville, MD 20705 2 NASA Goddard Space Flight Center (301) 286-4141, [email protected] Greenbelt, MD 20771 [email protected] Abstract. An advanced thermal control system for the Burst Alert Telescope on the Swift satellite has been designed and an engineering Abstract. Small spacecraft, including micro and nanosats, as they are test unit (ETU) has been built and tested in a thermal vacuum chamber. envisioned for future missions, will require an alternative means to The ETU assembly consists of a propylene loop heat pipe (LHP), two achieve thermal control due to their small power and mass budgets. constant conductance heat pipes, a variable conductance heat pipe One of the proposed alternatives is Variable Emittance (Vari-E) (VCHP), which is used for rough temperature control of the system, Coatings for spacecraft radiators. Space Technology-5 (ST-5) is a and a radiator. The entire assembly was tested in a thermal vacuum technology demonstration mission through NASA Goddard Space chamber at NASA/GSFC in early 2002. Tests were performed with Flight Center (GSFC) that will utilize Vari-E Coatings. This mission thermal mass to represent the instrument and with electrical resistance involves a constellation of three (3) satellites in a highly elliptical orbit heaters providing the heat to be transferred. Start-up and heat transfer with a perigee altitude of ~200 km and an apogee of ~38,000 km. Such of over 300 W was demonstrated with both steady and variable an environment will expose the spacecraft to a wide swing in the condenser sink temperatures. Radiator sink temperatures ranged from a thermal and radiation environment of the earth’s atmosphere. There are high of approximately 273 K to a low of approximately 83 K, and the three (3) different technologies associated with this mission. The three system was held at a constant operating temperature of 278 K technologies are electrophoretic, electrochromic, and Micro throughout most of the testing. A novel LHP temperature control ElectroMechanical Systems (MEMS). The ultimate goal is to make use methodology using both temperature-controlled electrical resistance of Vari-E coatings, in order to achieve various levels of thermal heaters and a small VCHP was demonstrated. This paper describes the control. The focus of this paper is to highlight the Vari-E Coating system and the tests performed, and includes a discussion of the test MEMS instrument, with an emphasis on the Electronic Control Unit results. responsible for operating the MEMS device. The Test & Evaluation approach, along with the results, is specific for application on ST-5, yet the information provides a guideline for future experiments and/or thermal applications on the exterior structure of a spacecraft.

45 46 Development and Test Results of a Development of a Two-Phase Capillary Pumped Heat Multi-Evaporator-Condenser Loop Heat Pipe Transport for Spacecraft Central Thermal Bus

Yury F. Maydanik1, Vladimir G. Pastukhov1, Triem Hoang1, Michael Brown2, Robert Baldauff2, Mariya A. Chernyshova 1, and Ad A.M. Delil 2 and Sheila Cummings2

1Institute of Thermal Physics, Ural Branch of the Russian Academy of 1TTH Research, Capitol Heights, MD 20743 Sciences, Amundsena St., 106, 620016 Ekaterinburg, Russia 2U.S. Naval Research Laboratory, Washington, DC 20375 2National Aerospace Laboratory NLR, Voorsterweg 31, 8316 PR (301) 641-2954; [email protected] Markneesse, the Netherlands (+7-3432) 678-791, [email protected] Abstract. Thermal requirements of future spacecraft and satellites will certainly outgrow the capability of conventional heat pipes in terms of Abstract. Results are presented of the development and tests of a 1 m heat transport, heat density, and temperature control. Emerging passive long ammonia ramified loop heat pipe, with two cylindrical evaporators heat transport technologies such as Capillary Pumped Loop (CPL) and (24 mm in diameter with an active zone length of 150 mm) and two Loop Heat Pipe (LHP) have demonstrated in both ground testing and condensers (length 200 mm, diameter 24 mm), made as pipe-in-pipe micro-gravity flight experiments that they have the potential to replace heat exchangers. Tests of the device at different orientations in 1-g heat pipes as primary heat transport devices in next generation thermal have shown that it can efficiently operate at symmetrical and non- control technology. Like heat pipes, CPLs and LHPs are completely symmetrical heat load distributions between the evaporators, and also passive systems which have no mechanical moving part to wear out or at different temperatures of the condensers cooling. The maximum total to introduce unwanted vibration to the spacecraft. However, the heat transport capacity is 1100-1400 W. Shutting down the active cooling of transport capabilities of CPLs and LHPs are at least one order of one condenser results in an abrupt decrease in the maximum transport magnitude higher than those of heat pipes. Despite sharing many capability of the device. operational characteristics, CPLs and LHPs do have differences. CPLs require a lengthy and tedious start-up procedure to prime the wicks before heat is applied to the evaporator plate. Even with the start-up procedure, start-ups are not always successful. LHPs, on the other hand, do not require a wick pre-conditioning process. But the LHP effective thermal conductance is not as high as that of a CPL. Temperature control of a LHP is not easily achieved. A novel concept, which combined a CPL and a LHP into one loop, was proposed to take advantage of selective features of each system without inheriting their shortcomings. The resultant loop was called Advanced Loop Heat Pipe (A-LHP). A proof-of-concept testbed was put together and tested at the Naval Research Laboratory. Test results showed that the A-LHP performed like a CPL without start-up problems associated with CPLs.

47 48 Flight Testing of the Capillary Pumped Loop 3 Effect of Non-ionic Surfactants on Nucleate Pool Boiling Experiment J.P. Kizito1, R. Balasubramanaim1, M.J. Boggess1, K.J. Stebe2 1 1 1 2 Laura Ottenstein , Dan Butler , Jentung Ku , Kwok Cheung , Robert 1 Baldauff2, and Triem Hoang3 National Center for Microgravity Research, Mail Stop 110-3, NASA Glenn Research Center, Cleveland, OH 44135 2 1Thermal Engineering Branch, NASA/Goddard Space Flight Center, Department of Chemical Engineering, Johns Hopkins University, Greenbelt, MD 20771 Baltimore, MD 21218 2Naval Research Laboratory, Washington, DC, 20375 (216) 433-2275; [email protected] 3TTH Research, Capitol Heights, MD 20743 (301) 286-4141, [email protected] Abstract. The ultimate aim of this project is to understand the influence of surfactants on boiling heat transfer in microgravity, and Abstract. The Capillary Pumped Loop 3 (CAPL 3) experiment was a identify the regimes which promote improved heat flux and the multiple evaporator capillary pumped loop experiment that flew in the mechanisms for it. The ground-based experiments we have performed Space Shuttle payload bay in December 2001 (STS-108). The main examine boiling curves for a flat substrate that consists of a thin layer objective of CAPL 3 was to demonstrate in micro-gravity a multiple of platinum that is deposited on a smooth silicon wafer, in ultra- evaporator capillary pumped loop system, capable of reliable start-up, purified water. These experiments establish a 1 g baseline, to which reliable continuous operation, and heat load sharing, with hardware for microgravity boiling curves can be compared. We use the non-ionic a deployable radiator. Tests performed on orbit included start-ups, surfactants Triton X-100 and Surfynol 104 in this study. Boiling data power cycles, low power tests (100 W total), high power tests (up to are obtained for aqueous solutions of these surfactants at various 1447 W total), heat load sharing, variable/fixed conductance transition concentrations and are compared to the data for pure water. tests, and saturation temperature change tests. The majority of the tests were completed successfully, although the experiment did exhibit an unexpected sensitivity to shuttle maneuvers. This paper describes the experiment, the tests performed during the mission, and the test results.

49 50 Heater Size and Gravity Effects on Pool Boiling Bubble Behavior in Subcooled Pool Boiling of Water Heat Transfer under Reduced Gravity

Jungho Kim and Christopher Henry Koichi Suzuki1, Motohiro Suzuki2, Saika Takahash2, Hirosi Kawamura1 and Yoshiyuki Abe3 University of Maryland, Dept. of Mechanical Engineering, College Park, MD 20742 1Department of Mechanical Engineering, Tokyo University of Science, 301-405-5437, [email protected] Yamasaki 2641, Noda, Chiba, 278-8510 Japan 2Graduate School of Science and Technology, Tokyo University of Abstract. Pool boiling heat transfer measurements from heaters of Science various sizes were obtained in low gravity aboard the KC-135. Visual 3 The National Institute of Advanced Industrial Science and observations indicate boiling in earth and high gravity can be Technology, Umezono 1-1-1, Tsukuba, Ibaragi, 305-8568 Japan dominated by surface tension effects exhibiting behavior similar to 1+81 (4) 7124 1501, [email protected] microgravity boiling. Surface tension dominated boiling occurred in earth and high gravity when the bubble departure diameter was greater Abstract. Subcooled pool boiling of water was conducted in reduced than the heater length. The Bond number at bubble departure was not gravity performed by a parabolic flight of aircraft and a drop-shaft constant across gravity levels and heater length scales, and was found facility. A small stainless steel plate was physically burned out in the to depend additionally on the wall superheat and the bulk fluid subcooled water by AC electric power during the parabolic flight. subcooling. Boiling bubbles grew with increasing heating power but did not detached from the heating surface. The burnout heat fluxes obtained were 200 ~ 400 percent higher than the existing theories. In the ground experiment, boiling bubbles were attached to the heating surface with a flat plate placed over the heating surface, and the experiment was performed by the same heating procedure as practiced under the reduced gravity. Same burnout heat fluxes as under the reduced gravity were obtained by adjusting the plate clearance to the heating surface. As the heating time extended longer than the reduced gravity duration, the burnout heat fluxes decreased gradually and became constant. Contact area of bubbles with heating surface was observed using a transparent heating surface in microgravity performed by a drop-shaft facility. The contact area of bubbles increased significantly at the start of microgravity. It is suggested by the experimental results that the boiling bubbles expand rapidly in the high heat flux region and the rapid evaporation of liquid layer remained between the bubbles and the heating surface raises up the critical heat flux higher than the existing theories in microgravity.

51 52 A Fundamental Study Regarding the Control of SAFE-100 Module Fabrication and Test Nucleate Boiling in a Complex Magnetizable Fluid by 1 1 2 an Applied Magnetic Field, in Microgravity Conditions Peter J. Ring , Edwin D. Sayre , and J. Tom Sena

1 Floriana D. Stoian1a,1c, Gheorghe Pop1a,1c, Doina Bica2,1c, Advanced Methods and Material, Sunnyvale, CA 94086 2 Virgil Stoica1a,1c, Oana Marinicã1b,1c, and Ladislau Vékás2,1c Los Alamos National Laboratory, Los Alamos, NM 87545 (408) 739-7772, [email protected] 1aDepartment of Thermal Machines & Transport, 1bInstitute for Complex Fluids and Abstract. Reliable, long-life, low-cost heat pipes can enable safe, affordable 1cNational Center for Engineering of Systems with Complex Fluids, space fission power and propulsion systems. Advanced versions of these ”Politehnica” University of Timisoara, 1 Mihai Viteazu Bv., systems can in turn allow rapid access to any point in the solar system. Stainless steel heat pipe modules are being built at Advanced Methods and Timisoara, RO-1900, Romania, Materials for use in a non-nuclear thermal hydraulic simulation of the SAFE- telephone/fax: + 40 256 221547, email: [email protected] 100 reactor. SAFE-100 is a near-term, low-cost space fission system 2 Laboratory of Magnetic Liquids, Center of Fundamental and demonstration. The heat pipes were designed to remove thermal power from Advanced Technical Researches, the SAFE-100 core, and transfer this power to an electrical power conversion Romanian Academy - Timisoara Branch, 24 Mihai Viteazu Bv., system. These heat pipe modules are being delivered to NASA Marshall Space Timisoara, RO-1900, Romania Flight Center to be filled and tested in a prototypical configuration during CY2003. The construction and test of a SAFE-100 module prototype is Abstract. A new type of working fluid to be used for boiling or described. multiphase heat transfer in microgravity conditions is proposed. The Heat pipes are readily manufactured using established procedures and advantages of using a complex magnetizable fluid for boiling or their use poses few unresolved technical issues. This paper describes multiphase heat transfer in microgravity are presented, based on the the building and test of the first module for the SAFE 100 reactor theoretical analysis of the body forces exerted by an applied magnetic prototype. This effort, building on the success of the SAFE 30 program, field in such a fluid, in comparison with the similar phenomena involves cooperation between Advanced Methods and Materials, determined by the body forces exerted by an applied electric field in a NASA Marshall Space Flight Center, and Los Alamos National dielectric fluid. The experimental set-up designed to study the effect of Laboratory. The successful test of this first module clearly the applied magnetic field on the bubble dynamics, to be proposed for a demonstrates that reactor module technology can be quickly transferred microgravity experiment, and currently in preparation for to industry for use in a fast paced flight program. The fabrication experimentation under terrestrial conditions, are presented. procedures used to produce these modules were adapted from those used to make alkali metal heat pipes that flew aboard the space shuttle Endeavor in 1996 and operated in a fast neutron flux in the EBR-II Reactor.

Currently, a total of nineteen modules are on order from Advanced Methods and Materials for the SAFE-100 prototype. As a part of the SAFE-100 effort, NASA Marshall has developed an in-house alkali metal handling capability. This capability allows test articles to be charged and their contents evaluated in a controlled environment. The remaining 18 heat pipe modules will be filled with sodium at MSFC and qualification tested during FY2003. The modules will then be assembled into the SAFE 100 core configuration and attached to a Brayton-cycle heat pipe-to-gas heat exchanger. The overall system will then be tested so that each module rejects 1.63 kW at 973 K. 53 54 Alkali Metal Handling Practices at NASA MSFC Transient Thermohydraulic Heat Pipe Modeling: Incorporating THROHPUT into the CÆSAR Patrick G. Salvail and Robert R. Carter Environment NASA Marshall Space Flight Center, Huntsville, AL 35812 Michael L. Hall (256) 544-1818, FAX: (256) 544-5877, [email protected] Los Alamos National Laboratory, P.O. Box 1663, MS-D409 Abstract. NASA Marshall Space Flight Center (MSFC) is NASA’s Los Alamos, NM 87545 principle propulsion development center. Research and development is 505-665-4312, [email protected] coordinated and carried out on not only the existing transportation systems, but also those that may be flown in the near future. Heat pipe Abstract. The THROHPUT code, which models transient cooled fast fission cores are among several concepts being considered thermohydraulic heat pipe behavior, is being incorporated into the for the Nuclear Systems Initiative. Marshall Space Flight Center has CÆSAR computational physics development environment. The developed a capability to handle high-purity alkali metals for use in CÆSAR environment provides many beneficial features for enhanced heat pipes or liquid metal heat transfer loops. This capability is a low model development, including levelized design, unit testing, Design by budget prototype of an alkali metal handling system that would allow ContractTM (Meyer, 1997), and literate programming (Knuth, 1992), in the production of flight qualified heat pipe modules or alkali metal a parallel, object-based manner. The original THROHPUT code was loops. The processing approach used to introduce pure alkali metal into developed as a doctoral thesis research code; the current emphasis is on heat pipe modules and other test articles are described in this paper. making a robust, verifiable, documented, component-based production package. Results from the original code are included. As a part of the SAFE-100 effort NASA Marshall has developed an in- house alkali metal handling capability. This capability allows test articles to be charged and their contents evaluated in a controlled environment. The alkali metal handling system is modular and flexible to accommodate small samples and multiple heat pipe modules. It may be modified to allow the fill of modules having arbitrary length or even a pre-assembled core. This equipment is not strictly limited to heat pipes it may also be used to support the development of other candidate systems such as potassium or lithium liquid metal heat transfer loops.

A stricter fabrication and quality control regimen would be justified for space flight qualified components. To keep prototype development cost below $0.5M, sodium purification was done by bulk vacuum distillation. The manufacture of alkali metal heat pipes or a liquid metal loop for a flight system would justify the development of a more elaborate handling capability. An advanced fill system for flight qualified units would vacuum distill alkali metal into individual heat pipe modules and might be made to fill modules for an entire core simultaneously. Detailed assays of non-metallic and gas residual impurity levels might also be performed in situ.

55 56 SAFE Alkali Metal Heat Pipe Reliability “A Market 2 Come: On-Orbit Servicing of Satellites”

Robert S. Reid1, 2 Joerg Kreisel 1Los Alamos National Laboratory, Los Alamos, NM 87545 2NASA Marshall Space Flight Center, Huntsville, AL 35812 JOERG KREISEL International Consultant (505) 667-2626, [email protected] Melanieweg 25, D-52072 Aachen, Germany [email protected]

Abstract. Near term deployment of a fission system is key to the Abstract. On-orbit servicing of space assets has been discussed success of the Nuclear Systems Initiative. To achieve an early flight, a frequently for many years and is known especially since the successful chosen design must be quickly converted into flight hardware. The Hubble repair mission conducted by astronauts in 1993. A major subset design must be simple with few integration issues. Integrating of on-orbit servicing is unmanned satellite servicing missions using components into a working unit can be very challenging, and becomes robotics and automation technologies (OOS). To date various concepts exponentially harder as system complexity increases. The design must for on-orbit servicing of satellites have been developed and take maximum advantage of non-nuclear testing, to allow fast-paced investigated, mostly driven by the technological challenges involved. and realistic, full power, non-nuclear heated, testing of the actual flight Particularly over the last 2 years this topic is gaining momentum, e.g. in unit. The system must be highly reliable and redundant with minimum Japan, Germany, Canada and the US, and two companies are currently (preferably zero) single-point failures. The system should also have no aiming at providing OOS exclusively. fundamental limitations that would preclude it from achieving long life. This paper elaborates on the big picture of on-orbit servicing and its Alkali metal heat pipes are among the best understood and tested of issues. Special attention is paid to the market drivers and potential components for first generation space fission reactors. The use of impact on the overall space industry from a global perspective. passive heat pipes for primary heat transport yields significant benefits Background of the findings presented are DLR-initiated high-level to space fission system testability, safety, and reliability yet reduces studies conducted by an international consortium led by the author and complexity, development cost, and production cost over other primary findings elaborated during the DLR-CSA international workshop "OOS heat transport options. Inherently redundant heat pipe arrays are 2002 - Defining a Way Forward" (Germany, November 2002). immune to single point pressure boundary failure. The breach of one or more heat pipes has minimal consequences to the overall system. A heat pipe array provides a generic interface that gives flexibility in choice of power conversion system, whether it is Brayton, Stirling, or static. A heat pipe array permits use of multiple power conversion units to achieve even higher levels of redundancy.

A flight reactor will require production of a hundred or more heat pipes with assured reliability over a number of years. To date, alkali metal heat pipes have been built mostly in low budget development environments with little formal quality assurance. Despite this, heat pipe test samples suggest that high reliability can be achieved with the care justified for space flight qualification. Fabrication procedures have been established that, if consistently applied, ensure long-term trouble- free heat pipe operation.

57 58 Space Technology and Intellectual Property: Funding Evolving Markets for Commercial, Civil, and Military Exploration Through Technology Commercialization Services

William N. Hulsey III Marshall H. Kaplan

Hughes & Luce, LLP, 111 Congress Ave., Austin, TX 78701 Strategic Insight, Ltd., 2440 Research Blvd., Suite 400 512-482-6856, [email protected] Rockville, MD 20850 301-977-1676; [email protected] Abstract. Future space missions will require revolutionary new technologies, and many of these technologies will dramatically improve people's lives here on Earth. Moreover, at least a portion of revenues Abstract. Recent commercial failures in the LEO market, declining derived from commercializing technologies first derived from continuing budgets for research, and other political factors have made it difficult space exploration can be redirected specifically to further early-stage for entrepreneurs and financial institutions to realize returns from venture investments in space technology development efforts. investments in new space transportation systems and satellites. This Innovative businesses will devise products and services making our lives paper explores the major factors impacting future markets that make safer, more rewarding, and more comfortable. From healthcare and use of our space infrastructure. At the top of the list is the high cost of advances in computing to improvements in everyday products, space access. This has been extremely expensive, and will continue to technologies derived from space exploration and technology development be expensive as long as space access remains low on the nation’s provide significant benefits to humankind. And, integrally associated with priority list. Many have tried to address the situation, but all have failed these technologies and the business commercializing them are the laws to make significant progress. While launch prices have generally been relating to inventions, intellectual capital, and innovation… the worldwide reduced over the past several years, they remain well above the elastic landscape of intellectual property laws. The presentation touches on how range of supply and demand. Our best estimate is that it will take an these laws might translate to laws affecting new inventions derived by order of magnitude reduction to significantly expand the market. humans either enroute or in an extraterrestrial environment. Projections about market segments that will represent future winners in space and launch demand forecasts are presented. Future markets, Focusing on business and technical objectives, this presentation introduces outside of traditional strongholds, are explored, including a long-term the different kinds of intellectual property rights protection and valuation view of new commercial space activities, conventional and ambitious mechanisms, including patents, trade secrets, and carefully negotiated future/futuristic activities, and related business aspects. contracts that can have particular relevance to space technology commercialization. The presentation further focuses on selected successful business and legal practices for extracting value from intellectual properties. The presenter further addresses a potential mechanism for returning the financial benefits of technology commercialization to the research base for further space exploration and development.

Different classes and technology areas of recently issued space technology patents and published patent applications: who is receiving these patents (companies and inventors), what the patents cover, and what they mean for further research and development efforts, are addressed. This will help those scientists, engineers, and technology managers responsible for space technology commercialization and technology transfer efforts see better the opportunities that their efforts make possible.

59 60 Public Space Markets – What We Know and What We Potential Effects of Government Regulations on Public Don’t Know Space Travel

Derek Webber Harvey A. Wichman

Futron Corporation, 7315 Wisconsin Avenue, Suite 900W,Bethesda, Claremont McKenna College, 850 Columbia Avenue, Claremont, CA MD 20814 91711 301 347-3441, [email protected] 909 607-7986, [email protected]

Abstract. This paper provides a discussion of the risk and constraint Abstract. The paper provides some background to a comprehensive implications in governing future space flights of the public. The author survey of public space travel that was conducted amongst wealthy explores differences between the proposed regulations by the American Americans by Futron Corporation and Zogby International. Medical Association, FAA and the Space Transportation Association Discussions will include methodology and the level of confidence in ad hoc Committee on Regulations. the findings. The author will identify areas where there is more uncertainty and which might, therefore, provide a useful focus for further research. .

61 62 Creating the Future Space Tourism Movement

Developing the Space Destinations – Near Term Tsuyoshi Saotome Possibilities Crystal Space Place, Inc., 16936 Burbank Blvd, Suite 209, Encino, CA Paola Favata 91316 [email protected] Anastrophe Design, 569 S. Murphy Ave, Sunnyvale, CA 94086 408 733 4915, [email protected] Abstract. The paper presents the current status of three ongoing projects that will introduce people to space and help to generate a new Abstract. This paper explores human well-being and enjoyment space tourism movement. The author will discuss and explain requirements of a space trip and the fundamental physical and relationships among the Space Tourism Society Chapter Japan, the psychological needs that a space accommodation must provide. The Mars Quest simulation project and a future space resort concept, the author discusses potential use of available spacecraft through their Crystal Space Palace. adaptation to ordinary people with the design of a new generation of interiors. Examples are given for The U.S. TransHab Module, The Russian Svezda Module, and the Italian Multi-Purpose Logistics Module.

63 64 Experiences In Securing Insurance For A Reusable Rocket: Implications For Commercial RLVs New Mexico Commercial Space Program

Jeffrey K. Greason, Louis Gomez

XCOR Aerospace, P.O. Box 1163, Mojave, CA 93502 New Mexico Office of Space Commercialization, Santa Fe, NM 661-824-4714, [email protected] The presentation will cover the history of the New Mexico Abstract. Since the summer of 2001, XCOR Aerospace has been Commercial Space Program, including Legislation, site selection for operating a reusable rocket-powered manned test vehicle, the EZ- the Lockheed Martin VentureStar commercial space vehicle, and Rocket. The EZ-Rocket is a test bed for the technology and regulatory current initiatives. compliance issues associated with passenger-carrying reusable launch vehicles. We present a summary of the EZ-Rocket operational Legislation: The New Mexico Office for Space experience to date. XCOR will discuss its insurance experiences with Commercialization was established by the New Mexico Legislature in the EZ-Rocket and its attempts to develop insurance sources for future 1994. The Legislature also established the New Mexico Space RLVs. We discuss areas where the current structure of the insurance Commission, which has the mission of advising the Cabinet Secretary industry pose challenges to the growth of the commercial space for Economic Development on commercial space matters. trasportation market. Lockheed Martin VentureStar Program: New Mexico submitted two sites for consideration by Lockheed Martin: Upham and Oro Grande. Based on Lockheed Martin preferences, New Mexico took the Oro Grande site “off the table.” As a result of the Lockheed Martin VentureStar program, New Mexico identified several issues that required priority effort: a Memorandum of Agreement with White Sands Missile Range, Water, Land Acquisition, Finance, and an Environmental Impact Statement update. Further, Lockheed Martin asked New Mexico to engage with the FAA on Flight Safety.

Other initiatives: New Mexico has been working closely with several other commercial space companies interested in bringing commercial space business to New Mexico.

65 66 The Road To A Spaceport Functions and Requirements In the Advanced Spaceport Environment Maj. Gen. Jay T. Edwards, (USAF-Ret.) Roelof L. Schuiling Executive Director, Oklahoma Space Industry Development Authority th 3545 NW 58 Street; Suite #325 NASA John F. Kennedy Space Center, Mail Code YA-E6 Oklahoma City, OK 73112 Florida, 32899 [email protected] 321-867-9354, [email protected]

Abstract. As the Oklahoma Spaceport moves forward in the licensing Abstract. This paper describes the initial activity of the Advanced process, the Board of Directors, along with the staff, of the Oklahoma Spaceport technology Working Group, a joint federal government, state Space Industry Development Authority (OSIDA) are using the government, military, academia, and commercial working group in the encouragement of state and national leaders to continue in their development of a spaceport technology planning process. The process progress. will be used by the Advanced Spaceport Technology Working Group to identify, coordinate, develop, and demonstrate technologies that will Currently, the agency is in the beginning stages of the Environmental improve spaceport safety and reliability and lower the cost of access to Impact Study (one that is scheduled to last roughly a year-and-a-half. space. The initial phases of this planning process involve systems This paper addresses the processes a new spaceport must go through to definition and performance gap identification by the Advanced meet the requirements set forth by the Federal Aviation Administration Spaceport Technology’s Vision Team. The output of this initiative will (FAA). It also outlines the agency’s goals for the future. then form the input for technical focus teams in developing concepts of spaceport technology gaps and spaceport technology development for future spaceports.

67 68 Launch System Testbed: An Innovative Approach for Electrochemical Evaluation of Alloys for Spaceport Design and Development of Future Launch Structures Design

Bruce T. Vu1, Max Kandula2, Ravi N. Margasahayam2, and Luz Marina Calle1, Louis G. MacDowell1, and Rubiela D. Vinje2 Danielle M. Ford3 1Corrosion Science and Technology, Mail code: YA-C2-T, NASA 1NASA Kennedy Space Center, KSC, FL 32899 Kennedy Space Center, Kennedy Space Center, FL 32899 2Dynacs Inc., Kennedy Space Center, KSC, FL 32899 2Dynacs, Inc., Mail Stop DNX-15, Kennedy Space Center, FL 32899 3Embry-Riddle Aeronautical University, Daytona Beach, FL 32114 (321) 867-3278, [email protected] (321) 867-2376, [email protected]

Abstract. The launch of space vehicles generates extreme conditions, Abstract. Corrosion studies began at the Kennedy Space Center such as vibrations and acoustics that can affect the launch pad, l space (KSC) in 1966 during the Gemini/Apollo Programs with the evaluation vehicles, and their payloads. These acoustic loads are the results of of long-term protective coatings for the corrosion protection of carbon intense acoustic environment generated by the interaction of the rocket- steel. NASA’s KSC Beach Corrosion Test Site, which was established engine exhaust stream mixing with the ambient atmosphere. The at that time, has been documented by the American Society of primary source of structural vibrations and internal loads during launch Materials (ASM) as one of the most corrosive naturally occurring is due to these acoustic loads. Therefore, being able to manage and environments in the world. With the introduction of the Space Shuttle suppress these undesirable conditions is critical to proper functioning of in 1981, the already highly corrosive conditions at the launch pad were vehicle components, payloads, and launch support structures. The goal rendered even more severe by the acidic exhaust from the solid rocker of the Launch System Testbed at NASA-Kennedy Space Center is to boosters. In the years that followed, numerous studies have identified develop new methods to improve the performance of launch systems by materials, coatings, and maintenance procedures for launch hardware reducing noise, vibration, and stress loads generated during launch. and equipment exposed to the highly corrosive environment at the launch pad. The Corrosion Laboratory was established at KSC in 1985 and was outfitted with state-of-the-art electrochemistry equipment to conduct research and materials characterization in many different corrosive environments. This paper will describe the application of electrochemistry in combination with atmospheric exposure to the selection of alloys in a spaceport environment.

69 70 High Altitude Launch for a Practical SSTO Intelligent Launch and Range Operations Testbed

Geoffrey A. Landis1 and Vincent Denis2 Rodney D. Davis and Kevin R. Brown

1NASA John Glenn Research Center, Mailstop 302-1, Command and Control Technologies, 1425 Chaffee Drive, Titusville 21000 Brook Park Road, Cleveland, OH 44135 FL 32780 2International Space University,Parc d'Innovation, 321-264-1193; [email protected] 67400 Illkirch-Graffenstaden Strasbourg, France 216-433-2238; [email protected] Abstract. Today’s space operations rely on a vast network of manual Abstract. Existing engineering materials allow the constuction of activities and human decisions to safely plan missions, configure towers to heights of many kilometers. Orbital launch from a high systems, conduct flights, and support mission analysis. Literally altitude has significant advantages over sea-level launch due to the hundreds of range and launch operations personnel are required to plan reduced atmospheric pressure, resulting in lower atmospheric drag on and execute each mission. The cost of operations is a significant the vehicle and allowing higher rocket engine performance. High- portion of the total lifecycle cost of space flight, but it is difficult and altitude launch sites are particularly advantageous for single-stage to risky to experiment with new, potentially lower-cost operational orbit (SSTO) vehicles, where the payload is typically 2% of the initial approaches during actual space flight operations. A capability for launch mass. An earlier paper enumerated some of the advantages of investigating new operational approaches that promise to significantly high altitude launch of SSTO vehicles. In this paper, we calculate reduce cost and complexity for future launch vehicles is needed. This launch trajectories for a candidate SSTO vehicle, and calculate the paper will describe the concept of an "Operations Testbed" that would advantage of launch at launch altitudes 5 to 25 kilometer altitudes provide such a capability. above sea level. The performance increase can be directly translated into increased payload capability to orbit, ranging from 5 to 20% The objective of this work is to influence the design of future spaceport increase in the mass to orbit. For a candidate vehicle with an initial information systems to enable space flight operations cost reduction payload fraction of 2% of gross lift-off weight, this corresponds to 31% and safety improvement. With support from NASA, we have increase in payload (for 5-km launch altitude) to 122% additional established a pathfinder technology demonstrator for researching payload (for 25-km launch altitude). human centered computing techniques in relationship to intelligent launch and range operations.

The testbed will allow researchers to investigate new human performance and operations techniques using a variety of simulated missions, vehicles, flight anomalies, human controller scenarios, and range operations. With this approach, researchers will simulate experimental operations concepts and analyze alternate approaches to system design to better understand human factors and performance issues. As a result, future designers will be able to produce more effective information and decision support systems for launch, range, and flight operations. The ultimate outcome of this project could be replacement of today’s space flight procedures and systems with intelligent systems that are highly responsive, safe and can accommodate high flight rates and mixed fleet operations with fewer human controllers at much lower cost.

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Boeing Space Launch Initiative Program Overview Selection of Lockheed Martin’s Preferred TSTO Configurations for the Space Launch Initiative Dallas Bienhoff Joshua B. Hopkins The Boeing Company, 5301 Bolsa Avenue MC H012-B202, Huntington Beach, CA 92647-2099 Lockheed Martin Astronautics, Denver, Colorado, USA (714) 896-6435, [email protected] (303) 971-7928, [email protected]

Abstract. This presentation describes the Space Launch Initiative in Abstract. Lockheed Martin is developing concepts for safe, affordable context with the United States Integrated Space Transportation Plan. Two Stage to Orbit (TSTO) reusable launch vehicles as part of NASA’s Its objectives and goals are defined and equated to space transportation Space Launch Initiative. This presentation discusses the options architecture evaluation criteria, or figures of merit. The Boeing scope, considered for the design of the TSTO, the impact of each of these architecture trade space, system definition, and evaluation process are options on the vehicle configuration, the criteria used for selection of defined. Operational and performance requirements are discussed. preferred configurations, and the results of the selection process. More Key architecture attributes and their impact are described along with than twenty configurations were developed in detail in order to related SLI technology risk reduction activities necessary to achieve compare options such as propellant choice, serial vs. parallel burn program goals. Finally, the SLI technology risk reduction activities sequence, use of propellant crossfeed between stages, “bimese” or impact on Boeing candidate architectures is identified. optimized stage designs, and high or low staging velocities. Each configuration was analyzed not only for performance and sizing, but also for cost and reliability. The study concluded that kerosene was the superior fuel for first stages, and that bimese vehicles were not attractive.

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2nd Generation RLV: Overview of Concept Development Process and Results Overview of Orbital Sciences Corporation’s Orbital Space Plane (OSP) for the Space Launch Mark Benton, Jim Berry, Harry Benner Initiative Northrop Grumman Corporation, 1 Northop Way El Segundo,CA, Adrienne E. Wasko1 90245 90245 (310) 331-5384, [email protected] 1Orbital Sciences Corporation, 21839 Atlantic Blvd, Dulles, VA 20166

Abstract. Northrop Grumman Integrated Systems is under contract to (703) 406-5833, [email protected] study the development a 2nd Generation Reusable Launch Vehicle (RLV) architecture for NASA as part of their Space Launch Initiatives Abstract. Orbital Sciences Corporation’s reference Orbital Space Plane effort. The intent of this program is to develop a reusable launch system (OSP) is a new spacecraft design that was developed under contract to that has a significant improvement in system safety and reduction in NASA for the Space Launch Initiative (SLI) Program. The SLI cost of payload delivery when compared to the existing Space Shuttle requirements for the OSP include providing crew escape capability system. The initial part of this effort required the rapid evaluation of a during all phases of flight, transporting astronauts to the International wide design space to identify the types of concepts that might be able to Space Station (ISS), and returning ISS crew members to Earth as part meet NASA’s demanding cost and safety requirements. Twenty three of normal crew rotation or in the event of an on-orbit emergency. concepts, and multiple variants, were originally considered. Through Orbital’s OSP, referred to as a Blended Lifting Body (BLB), is a the application of systems engineering principles and progressively unique concept that is fully responsive to SLI technical and higher fidelity technical analyses this large trade space was reduced to programmatic requirements. This paper describes the OSP design 6 concepts, then 3 then to 1 preferred concepts. This paper reviews the challenges and solutions and provides Orbital’s configuration concepts that were considered, the process by which candidates were highlights. eliminated, and finally an overview of the Northrop Grumman’s preferred concept.

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Systems Engineering Approach to Technology The Successful Development Of An Automated Integration for NASA’s 2nd Generation Reusable Rendezvous And Capture (AR&C) System For The Launch Vehicle National Aeronautics And Space Administration

Sheryl Kittredge, Dale Thomas, Marc Verhage, Charles Smith, Fred D. Roe and Richard T. Howard Leann Thomas Simulation Group (ED19), NASA Marshall Space Flight Center, NASA/Marshall Space Flight Center, MSFC, AL 35812 Huntsville, AL 35812, USA (256) 544-9032, [email protected] 256-544-3512, [email protected]

Abstract. The overall goal of the 2nd Generation RLV Program is to Abstract. During the 1990’s, the Marshall Space Flight Center substantially reduce technical and business risks associated with (MSFC) conducted pioneering research in the development of an developing a new class of reusable launch vehicles. NASA’s specific automated rendezvous and capture/docking (AR&C) system for U.S. goals are to improve the safety of a 2nd-generation system by more space vehicles. Development and demonstration of a rendezvous than an order of magnitude — equivalent to a crew risk of 1-in-10,000 sensor was identified early in the AR&C Program as the critical missions — and decrease the cost an order of magnitude. enabling technology that allows automated proximity operations and docking. A first generation rendezvous sensor, the Video Guidance Architecture definition is being conducted in parallel with the Sensor (VGS), was developed and successfully flown on the Space maturating of key technologies specifically identified to improve safety Shuttle on flights STS-87 and STS-95, proving the concept of a video- and reliability, while reducing operational costs. An architecture based sensor. A ground demonstration of the entire system and broadly includes an Earth-to-orbit reusable launch vehicle, on-orbit software was successfully tested. transfer vehicles and upper stages, mission planning, ground and flight operations, and support infrastructure, both on the ground and in orbit. Advances in both video and signal processing technologies and the Systems design of an architecture emphasizes complete operations and lessons learned from the two successful flight experiments provided a includes all activities/processes that interface with hardware and baseline for the development, by the MSFC, of a new generation of software, focusing on simplicity of the entire system. The systems video based rendezvous sensor. The Advanced Video Guidance Sensor engineering approach ensures that the technologies developed will (AGS) has greatly increased performance and additional capability for synergistically integrate into the optimum vehicle. To best direct longer-range operation with a new target designed as a direct technology development decisions, analytical models are employed to replacement for existing ISS hemispherical reflectors. Components of accurately predict the benefits of each technology toward potential the developed AR&C system are to fly again in 2004. space transportation architectures as well as the risks associated with each technology. Rigorous systems analysis provides the foundation for assessing progress toward safety and cost goals.

The systems engineering review process factors in comprehensive budget estimates, detailed project schedules, and business and performance plans, against the goals of safety, reliability, and cost, in addition to overall technical feasibility. This approach forms the basis for investment decisions in the 2nd Generation RLV Program’s risk- reduction activities. Through this process, NASA will continually refine its specialized needs and identify where Defense and commercial requirements overlap those of civil missions.

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Advanced Checkout, Control, and Maintenance System Next Generation Launch Technologies Program (ACCMS) Breakthrough Technologies Propulsion Activities Overview

Cary Peaden Shayne Swint

NASA Kennedy Space Center, KSC, Florida Next Generation Launch Technologies Program (321) 867-9296, [email protected] NASA/Marshall Space Flight Center, MSFC, AL 35812 (256) 544-4060, [email protected] Abstract. Integration of business systems and processes, ACCMS places the maintainers/sustainers and the decision-makers in a tightly Abstract. The Next Generation Launch Technology Program managed control loop. Vast real-time information on problems, routine represents an integration of technologies within both the Space Launch maintenance results, installations, integrations, etc., can be synthesized, Initiative’s 2nd Generation and the Advanced Space Transportation consolidated, summarized and delivered in near real-time to the persons Program’s 3rd Generation. The overall goal of the NGLT Program is to making the decisions on flight readiness, business operations, or substantially reduce technical and business risks associated with logistics. This also allows creation of an operating environment where developing a new class of reusable launch vehicles. Propulsion has decision information propagates back to real-time operations. been identified as a major enabling technology for reaching the goals and objectives of NASA’s Space Launch Initiative. This presentation describes the propulsion technology activities within the NGLT Program. These key technologies will be matured in parallel with architecture definition and details systems analysis. Development and demonstration of propulsion technologies at the system level will enable a high confidence assessment of the impact of these technologies on increasing reliability and safety margins while decreasing the cost for transportation system development. The specific propulsion activities and how they are supportive of the United States Integrated Space Transportation Plan will be discussed.

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Water Detection and Removal From Shuttle Tiles Standards and Certification for Reusable Space Transportation Robert C. Youngquist

James E. French Spaceport Engineering and Technology Directorate NASA Kennedy Space Center, Florida, 32899 American Institute of Aeronautics and Astronautics, 1801 Alexander (321) 867-1829, [email protected] Bell Drive, Reston, VA 20191 703 264-7570, [email protected] Abstract. Current methods for detecting and removing water from the Space Shuttle tiles have proved inadequate in cases of excessive water Abstract. The development, timing, and use of technical standards for exposure. This paper describes two new tools that are currently being reusable space transportation is still an open issue in the industry. introduced to Shuttle processing to supplement the existing methods. Some suggestions will be provided. A scenario using such standards A capacitive device has been developed to augment the IR camera for the advancement and promotion of this aerospace sector will be method of detecting water in the tiles and a vacuum pump system is shown. The industry can lead the way to its own development in being tested as a likely replacement to the heat lamps currently used to cooperation with pertinent regulatory agencies. Leadership is needed, dry wet tiles. even though progress seems to be on a plateau.

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Development of NASA Technical Standards Program A Nuclear-Powered Laser-Accelerated Plasma Relative to Enhancing Engineering Capabilities Propulsion System

Paul S. Gill1 and William W. Vaughan2 Terry Kammash

1NASA Technical Standards Program Office, NASA Marshall Space Department of Nuclear Engineering and Radiological Sciences, Flight Center, Huntsville, AL 35812 University of Michigan, Ann Arbor, MI 48109-2104 2University of Alabama in Huntsville, Huntsville, AL 35899 (734) 764-0205, [email protected] 1256-544-2557, [email protected] Abstract. Recent experiments at the University of Michigan and other Abstract. The enhancement of engineering capabilities is an important laboratories throughout the world have demonstrated that ultrafast aspect of any organization; especially those engaged in aerospace (very short pulse length) lasers can accelerate charged particles to development activities. Technical Standards are one of the key relativistic speeds. The terrawatt laser at the University of Michigan elements of this endeavor. The NASA Technical Standards Program has generated a beam of protons containing more than 1010 particles at was formed in 1997 in response to the NASA Administrator’s directive a mean energy of over one Mev while the petawatt laser at the to develop an Agencywide Technical Standards Program. The Lawrence Livermore National Laboratory has produced proton beams Program's principal objective involved the converting Center-unique containing more than 1014 particles with maximum energy of 58 Mev technical standards into Agency wide standards and the and a mean energy of about 6 Mev. Using the latter data as a basis for adoption/endorsement of non-Government technical standards in lieu of a present-day LAPPS (Laser Accelerated Plasma Propulsion System) government standards. In the process of these actions, the potential for propulsion device we show that it can produce a specific impulse of further enhancement of the Agency’s engineering capabilities was several million seconds albeit at a fraction of a Newton of thrust. We noted relative to value of being able to access Agencywide the show that if the thrust can be increased to a modest 25 Newtons a fly- necessary full-text technical standards, standards update notifications, by robotic interstellar mission to 10,000 AU can be achieved in about and integration of lessons learned with technical standards, all available 26 years, while a round trip to Mars will be accomplished in about 6 to the user from one Website. This was accomplished and is now being months. In both instances a one MWe nuclear power system with a enhanced based on feedbacks from the Agency’s engineering staff and mass of about 5 MT will be needed to drive the laser, and the recently supporting contractors. This paper addresses the development announced NASA’s Nuclear Space Initiative should be able to address experiences with the NASA Technical Standards Program and the such reactors in the near future. enhancement of the Agency’s engineering capabilities provided by the Program’s products. Metrics are provided on significant aspects of the Program.

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Fusion Ship II- A Fast Manned Interplanetary Space Ion Dynamic Capture Experiments With The High Vehicle Using Inertial Electrostatic Fusion Performance Antiproton Trap (HiPAT)

R. L. Burton1, H. Momota2, N. Richardson1, Y. Shaban1 and G. H. James Martin1, Raymond Lewis2, Suman Chakrabarti1, Miley2 William H Sims1, J Boise Pearson1 and Wallace E Fant3

1University of Illinois at Urbana-ChampaignUrbana, Illinois 61801 1NASA MSFC, TD40, Huntsville, Alabama, 35812 217 244-6223, [email protected] 2R. Lewis Company (MSFC, TD40), Huntsville, Alabama, 35812 2NPL Associates, Inc., 912 W. Armory Ave., Champaign, Illinois 61821 3Cortez III, Huntsville, Alabama, 35812 217-356-5402, [email protected] (256) 544-6054, [email protected]

Abstract. In STAIF 2002, the authors described an inertial electrostatic Abstract. To take the first step towards using the energy produced confinement (IEC) fusion- driven space vehicle concept for manned from the matter-antimatter annihilation for propulsion applications, the interplanetary missions [1]. This design was based on a bank of 10 D-3He NASA Marshall Space Flight Center (MSFC) Propulsion Research fueled IEC reactors operating in parallel to provide 100 MWs of power in a Center (PRC) has initiated a research activity examining the storage of collimated 14-MeV proton beam. Following direct conversion of the beam low energy antiprotons. The High Performance Antiproton Trap power to high voltage electricity, the final thrust was obtained by a bank of NSTAR type ion thrusters operating on krypton giving a specific impulse of (HiPAT) is an electromagnetic system (Penning-Malmberg design) 16,000 seconds and a total thrust of 1020 N. This design resulted in a 500- consisting of a 4 Tesla superconductor, a high voltage electrode metric-ton class vessel. In a recent paper, F. Thio, et al. [2], laid down a confinement system, and an ultra high vacuum test section. It has been challenge to develop a system with propellant exhaust velocity in excess of 100 designed with an ultimate goal of maintaining 1012 charged particles km/s and specific power in excess of 10 kW/kg. They argue that this is with a half-life of 18 days. Currently, this system is being evaluated essential to achieve reasonable bounds on IMLEO for efficient and affordable experimentally using normal matter ions that are cheap to produce, human exploration of the solar system, especially of the outer planets. In relatively easy to handle, and provide a good indication of overall trap response to this challenge we have undertaken a redesign to the 2002 IEC behavior (with the exception of assessing annihilation losses). The ions driven space vehicle with a goal of achieving a one-year round-trip time for a are produced via a positive hydrogen ion source and transported to manned Jupiter mission. The revised design has 500 MWs of proton beam power and delivers a thrust of ~2,500 N with an ISP of ~100,000 sec. HiPAT in a beam line equipped with electrostatic optics. The optics Redesigning several key subsystems retains the overall vehicle weight of ~500 serve to both focus and gate the incoming ions, providing microsecond- metric tons. In this new design the number of IEC reactors is increased, while timed beam pulses that are dynamically captured by cycling the HiPAT a higher ion thruster voltage is employed. The propellant mass is reduced with forward containment field like a “trap door”. Initial dynamic capture the thruster modification and the crew/electronics compartment shielding experiments have been successfully performed with beam energy and weight is also reduced by a change in the cabin design and location. Although currents set to 1.9 kV and 23 mamps, respectively. At these settings up the IEC reactor weight is higher, since this is a small percentage of the overall to 2x109 ions have been trapped during a single dynamic cycle. vehicle weight, the weight reductions already noted generally compensate for this change. The overall vehicle design will be described along with its projected performance for a Jupiter mission. A strategy for developing propulsion-grade IEC power units, considering the status of present IEC experiments, will also be described.

[1] R. Burton, H. Momota, N. Richardson, M. Coventry, Y. Shaban, and G. H. Miley, “High Performance Manned Interplanetary Space Vehicle Using D-3He Inertial Electrostatic Fusion,” Space Technology and Applications Forum (STAIF-2002), edited by M. El-Genk, AIP Conference Proceedings, New York, pp. 819-827, (2002).

[2] Y. C. Francis Thio, G. R. Schmidt, J. Cole, J. F. Santarius, and P. J. Turchi, “Fusion for Space Propulsion, American Nuclear Society 2002 Annual Meeting, the Revival of the Nuclear Energy Option,” Hollywood, FL, p. 401, June 9-13 (2002).

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On the Performance Prediction and Scale Modelling of Nuclear Propulsion Requirements for Science Missions a Motorised Momentum Exchange Propulsion Tether Leonard A. Dudzinski Matthew P. Cartmell1, Spencer W. Ziegler2, and David S. Neill1 NASA Headquarters, Office of Space Science, Nuclear Systems 1Department of Mechanical Engineering, University of Glasgow, Program, Washington DC, WA 20545-0001 Glasgow, G12 8QQ, Scotland, UK (202) 358-3553, [email protected] 2Department of Mechanical, Aerospace & Manufacturing Engineering, UMIST, Sackville Street, PO Box 88, Manchester, M60 1QD Abstract. This paper presents the results of analyses of four reference England, UK solar system science missions that have been defined by NASA +44-141-330-4337, +44-141-330-4343, [email protected] Headquarters Office of Space Science to establish performance requirements for the development of a multi-mission Nuclear Electric Abstract. This paper discusses a programme of research based on the Propulsion (NEP) system. The results show how high level parameters incremental invention of the so-called Motorised Momentum Exchange such as V and initial acceleration can be used to characterize the Tether (MMET) for space vehicle propulsion, and summarises aspects required performance of an NEP system for a range of missions. Using of the predicted performance of hanging, librating, and spinning these figures as a guide, a range of NEP system power levels, specific symmetrical momentum exchange tethers in a circular orbit around the impulses, and technologies can be scoped for applicability to such Earth. A preliminary case for double-payloaded, symmetrical tethers is missions. The paper will also discuss other high level performance also made. This shows that the MMET concept has certain predicted requirements that drive the design of an NEP system. performance advantages over a passive momentum exchange tether. From that stand-point an ESA funded programme of terrestrial scale- model experimentation is discussed. This programme was intended to prove certain practicalities of the motorised concept within a suitably scaled model. To that end a dynamic scaling methodology based on aspects of the Buckingham Pi-theorem was evolved and appropriate dynamic scaling criteria were obtained for both rigid body spin-up and flexural vibrations within the system. The paper outlines the practical design of the scale model which resulted from this work, the running of a set of two-dimensional experimental tests on a large expanse of ice, and the resulting interpretation of results of the tests. The discussion concludes with an overview of new work on initial proposals for de- spin of the payloads after release, and issues relating to post-release de- spin of the centralised motor drive facility.

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MITEE-B: A Compact Ultra Lightweight Bi-Modal Design and Development of the MITEE-B Bi-Modal Nuclear Propulsion Engine for Robotic Planetary Nuclear Propulsion Engine Science Missions John C. Paniagua, James R. Powell, and George Maise

James Powell1, George Maise1, John Paniagua1, and Stanley Borowski2 Plus Ultra Technologies, Inc.,Stony Brook, New York 11790-3350 1Plus Ultra Technologies, PO Box 547, Shoreham, New York 11786 (631) 744-5707; [email protected] 2NASA Glenn Research Center, 21000 Brookpark Road, Cleveland, Abstract. Previous studies of compact, ultra-lightweight high performance Ohio 44135 nuclear thermal propulsion engines have concentrated on systems that only (631) 744-5707 phone/fax, [email protected] deliver high thrust. However, many potential missions also require substantial amounts of electric power. Studies of a new, very compact and lightweight bi- Abstract. Nuclear thermal propulsion (NTP) enables unique new modal nuclear engine that provides both high propulsive thrust and high robotic planetary science missions that are impossible with chemical or electric power for planetary science missions are described. The design is a nuclear electric propulsion systems. A compact and ultra lightweight modification of the MITEE nuclear thermal engine concept that provided only bi-modal nuclear engine, termed MITEE-B (MInature ReacTor EnginE high propulsive thrust. In the new design, MITEE-B, separate closed cooling - Bi-Modal) can deliver 1000's of kilograms of propulsive thrust when circuits are incorporated into the reactor, which transfers useful amounts of it operates in the NTP mode, and many kilowatts of continuous electric thermal energy to a small power conversion system that generates continuous electric power over the full life of the mission, even when the engine is not power when it operates in the electric generation mode. The high delivering propulsive thrust. Two versions of the MITEE-B design are propulsive thrust NTP mode enables spacecraft to land and takeoff described and analyzed. Version 1 generates 1 kW(e) of continuous power for from the surface of a planet or moon, to hop to multiple widely control of the spacecraft, sensors, data transmission, etc. This power level separated sites on the surface, and virtually unlimited flight in planetary eliminates the need for RTG’s on missions to the outer planets, and allowing atmospheres. The continuous electric generation mode enables a considerably greater operational capability for the spacecraft. This, plus its spacecraft to replenish its propellant by processing in-situ resources, high thrust and high specific impulse propulsive capabilities, makes MITEE-B provide power for controls, instruments, and communications while in very attractive for such missions. In Version 2, of MITEE-B, a total of 20 space and on the surface, and operate electric propulsion units. Six kW(e) is generated, enabling the use of electric propulsion. The combination examples of unique and important missions enabled by the MITEE-B of high open cycle propulsion thrust (20,000 Newtons) with a specific impulse of ~1000 seconds for short impulse burns, and long term (months to years), engine are described, including: (1) Pluto lander and sample return; (2) electric propulsion greatly increases MITEE’s ÄV capability. Version 2 of Europa lander and ocean explorer; (3) Mars Hopper; (4) Jupiter MITEE-B also enables the production and replenishment of H2 propellant atmospheric flyer; (5) SunBurn hypervelocity spacecraft; and (6) He3 using in-situ resources, such as electrolysis of water from the ice sheet on mining from Uranus. Many additional important missions are enabled Europa and other Jovian moons. This capability would greatly increase the ÄV by MITEE-B. A strong technology base for MITEE-B already exists. available for certain planetary science missions. The modifications to the With a vigorous development program, it could be ready for initial MITEE multiple pressure tube/fuel element assembly to achieve bi-modal robotic science and exploration missions by 2010 AD. Potential capability are modest. Small diameter coolant tubes are bonded to the surface mission benefits include much shorter in-space times, reduced IMLEO of the MITEE cold frits that enclose the fuel elements. When the MITEE-B is requirements, and replenishment of supplies from in-situ resources. not operating with H2 propellant to generate high thrust, the reactor continues to operate at low thermal, which is transferred to the closed coolant circuit. Three electric power generations are examined for MITEE-B: closed Brayton, Stirling, and a conventional steam cycle with a mini-turbine. The Stirling and steam cycles have the lowest specific masses in kg/kW(e). Both appear practical for MITEE-B.

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SUSEE – An Ultra Lightweight Nuclear Electric Space 20% Efficient InGaAs/InPAs Thermophotovoltaic Cells Power System Based on Conventional Steam Cycle S.L. Murray1, C.S. Murray1, F.D. Newman1, R.R. Siergiej2, 2 2 George Maise1, James Powell1, and John Paniagua1 B. Wernsman , and S.A. Derry

1 1Plus Ultra Technologies, Inc., 25 East Loop Road, Stony Brook, NY Emcore Photovoltaics, Albuquerque, NM 87123, USA 2 (631) 751-2285, [email protected] Bechtel Bettis Inc., West Mifflin, PA 15122, USA (505) 332-5085, [email protected] Abstract. A new approach, termed SUSEE (Space nUclear Steam Electric Energy) for space nuclear electric power is described. SUSEE Abstract. Recent improvements in large area, high efficiency, uses conventional existing reactor and steam cycle technology to monolithic interconnected modules (MIMs) represent a significant step generate electrical power, together with a new, ultra lightweight in the development of thermophotovoltaic (TPV) technology for radiator concept to reject waste heat. It consists of a set of thin flat various power producing applications. The MIM architecture with metal (Al or Be) strips with internal grooved channels to carry the transmissive integrated spectral control offers a desirable high-voltage, condensing steam/water mixture. Design and performance trade low-current output, front-side contacts for simplified packaging, high studies for the SUSEE system (reactor, steam power conversion, and spectral utilization due to a metallic, highly reflective and specular back radiator) are presented as a function of system operating parameters, surface reflector, and a practical method for scale-up to full wafer including steam inlet pressure and temperature, condenser pressure, etc. devices. The n/p/n MIM TPV devices described in this work utilize a Excellent performance in terms of thermal cycle efficiency, radiator tunnel junction and a double heterostructure for improved performance. area, and system specific mass [kg/kW(e)] is achieved. As an example, Lattice-mismatched 0.6 eV, epitaxially grown InGaAs diodes form the a SUSEE system with a steam inlet temperature of 810 K (1000°F) and power-producing element. A power conversion efficiency of 20.6% 68 atm (1000 psi), turbine/generator efficiency of 80%, and condenser and a power density of 0.90 W/cm2 with a silicon carbide radiator pressure of 2 atm (395 K) achieves 23% thermal efficiency, 1.5 m2 operating at 1058oC is achieved for a 4 cm2 (die area) TPV cell radiator area per kW(e), and 2 kg/kW(e) radiator mass (beryllium operating at 26.7oC. construction). Total system specific mass depends on power level and operating lifetime, since criticality and burnup set the minimum reactor size and weight. Assuming the product of power and lifetime > 6 MW(e) years, we calculated the system specific mass (including reactor, radiator, power conversion and plumbing) to be 3kg/kW(e). The SUSEE reactor uses conventional water coolant/steam turbine technology with cermet nuclear fuel (UO2 particles in a metal matrix). 3D Monte Carlo neutronic analyses of SUSEE are described, together with detailed thermal-hydraulic analyses. The SUSEE radiator is described, together with its packaging for launch, deployment in space, and operational startup. The SUSEE radiator construction is very flexible, allowing it to be rolled up into a tight, small-diameter cylinder for launch. Once in space, the radiator is unrolled to form a flat, two- sided radiating surface

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The Status of Thermophotovoltaic Energy Conversion Improved Thermophotovoltaic (TPV) Performance Technology at Lockheed Martin Corp. Using Dielectric Photon Concentrations (DPC)

E.J. Brown, P.F. Baldasaro, S.R. Burger, L.R. Danielson, D.M. DePoy, P.F. Baldasaro & P.M. Fourspring G.J. Nichols, W.F. Topper Lockheed Martin Corp., 2401 River Rd., Niskayuna, New York 12309 Lockheed Martin Corp., 2401 River Rd., Niskayuna, New York 12309 (518) 395-6563, [email protected] (518) 395-7045, [email protected] Abstract. This report presents theoretical and experimental results that Abstract. In a thermophotovoltaic (TPV) energy conversion system, a demonstrate the feasibility of a new class of thermophotovoltaic (TPV) heated surface radiates in the mid-infrared region onto photodiodes energy converters with greatly improved power density and efficiency. which are sensitive in this range. Part of the energy is converted into Performance improvements are based on the utilization of the enhanced electric output. Conversion efficiency is maximized by reducing the photon concentrations within high refractive index materials. Analysis parasitic energy absorption with some form of spectral control strategy. demonstrates that the maximum achievable photon flux for TPV applications is limited by the lowest index in the photonic cavity, and In a TPV system, many technology options exist. Our development 2 efforts have concentrated on flat-plate geometries with greybody scales as the minimum refraction index squared, n . Utilization of the radiators, low bandgap quaternary diodes, front surface tandem filters increased photon levels within high index materials greatly expands the and a multi-chip module (MCM) approach that allows selective design space limits of TPV systems, including: a 10x increase in power fabrication processes to match diode performance. density, a 25% relative increase in conversion efficiency, or alternatively reduced radiator temperature by 40%. Recently, the authors achieved conversion efficiencies of about 17% (radiator 950°C, diodes 50°C) for a module in a prototypic cavity test environment. These tests employed InGaAsSb diodes with 0.52 eV bandgap. Because the spectral control system (front surface filter) was designed to perform at room temperature, the 50°C test condition shifted the diode bandgap wavelength relative to the filter turn-on position. A re-designed filter would result in slightly higher conversion efficiency at 50°C.

This paper provides details of the individual system components and describes the measurement technique used to record these efficiencies.

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Radioisotopic Powered Thermophotovoltaic Energy Lincoln Experimental Satellite (LES) 8/9 Systems Don Maclellan D.M. DePoy, E.J. Brown, P.F. Baldasaro, L.P. Rice MIT Lincoln Laboratory, 244 Wood St., Lexington, MA 02420

781 545-5725, [email protected] Lockheed Martin Corp., 2401 River Rd., Niskayuna, New York 12309

(518) 395-4425, [email protected] Abstract. The primary objective of the LES 8/9 (Lincoln Experimental Satellite 8/9) program was to demonstrate the feasibility of providing Abstract. Thermophotovoltaics (TPV) is a direct energy conversion substantially survivable communications between the National technology in which a heated surface radiating in the mid-IR region Command Authority and the nation's internationally dispersed nuclear illuminates photodiodes that are sensitive in this spectral range. Recent forces. It was imperative that the communications links survive under advances in semiconductor materials, diode fabrication and spectral determined enemy attack, both physical and electronic. Employing control technology have produced measured conversion efficiencies of Multi-Hundred Watt Radio-Isotope Thermoelectric Generators (MHW- about 20% and power densities of about 0.5 W /cm2. el RTGs) was a perfect solution to the problem of spacecraft power, given the alternative of physically fragile solar panels. For this and for other, An alternative to a fossil-fired or solar powered TPV system is one political, reasons we instantly seized the opportunity, to fly, as guinea powered by a radioisotope source. Such a system could prove pigs, the RTG's being developed by the AEC (later ERDA, then DOE). especially useful for remote, unmanned applications. Radioisotope TPV LES 8 and 9 were an outstanding success and led to the establishment systems differ from conventional TPV systems in two ways: the power of major national military communications satellite programs that are of the energy source is fixed and can not vary to meet changing power still being spawned. requirements. Furthermore, the net surface power density would be 2 low; about 0.2 Wel/cm . The second difference is the impact of heat rejection conditions on the overall system design. If heat must be rejected to a rarefied medium (gas or vacuum) or to one at high temperature, the system designer must optimize the balance between large heat rejection structures (increased mass) or higher diode temperatures (more radioisotope fuel required).

Based on technology developed at Lockheed Martin, the authors carried out scoping studies for two cases to bound the performance of radioisotope/TPV systems; each system had a thermal input of 720 Wth with the hot side radiator at 1100°C. The first case was a TPV/radioisotope system rejecting heat to a 20°C water environment. The optimized system produced 123 Wel at a specific power of 19.0 Wel/kg and an overall conversion efficiency of 16.4%. The second case was a system rejecting heat to a 4K vacuum. This system produced 100 Wel at specific power of 10.5 Wel/kg and an overall conversion efficiency of 13.9%.

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Solid Rocket Propellant Fire Characterization Test A Method for the Analysis of Impact Events Involving Results: Interpretation and Modeling Nested Safety Systems

L.W. Hunter1,2, H.N. Oguz1, S.C. Walts1, Y. Chang1 James R. Coleman

1Johns Hopkins University Applied Physics Laboratory, 11100 Johns James R. Coleman, Consulting Hopkins Road, Laurel, MD 20723-6099 P. O. Box 385, Cross, SC, 29436-0385 2443-778-7406, [email protected] Voice & Fax:(843) 753 2612; [email protected]

Abstract. A test was conducted to characterize the environment under Abstract. This paper presents an analysis model devised specifically and near a 200 lbm burning fragment of aluminized solid rocket motor for evaluating sequential impact damage to a space nuclear system propellant burning in open air. The propellant fragment was a cylinder (barrier failure and fuel release for isotopic systems and geometry of radius 10 inch and height 10 inch. The cylinder was oriented with changes for reactor systems) with a nested safety system design. The one flat circular face down, resting on high-clay sand. The undersurface analysis model is a relatively simple deterministic approach using and the curved side surfaces were ignited. The top surface was disruptive work as the measure of damage. This approach offers a inhibited. Radiance was measured with longwave and midwave logical basis for assessing the impact damage to nested safety systems infrared cameras and a spectral radiometer. The longwave radiance from sequential impacts, for designing safety tests, for analyzing implied lower bounds on the spatially resolved temperatures. The subsequent test data, and finally for use in a safety analysis. Because of midwave radiance implied a spectrally resolved spatially averaged the wide variety of conditions that must be considered in a complete emissivity in the midwave region. Rod calorimeters and solid safety analysis and the stochastic nature of the damage states, the cylindrical calorimeters were used to assess the heat flux. Various deterministic approach is only the starting point for a complete safety refractory oxide and metallic witness samples were placed in the analysis. environment to infer local plume temperature benchmarks. All deposits produced by the combustion were analyzed chemically.

This presentation will emphasize progress toward developing a model of aluminized propellant fires burning in open air. The underlying concept of the model is that the plume combustion reactions are non- adiabatic, due to heat conduction and radiation losses. The non- adiabatic combustion concept is implemented in well-stirred zones that shrink as the propellant surface recedes. The degree of reaction of Al(L) droplets in the plume and the resulting heat release also decrease. The spatial extent of each zone as a function of time, and the conditions in each zone, define the fire environment as a function of position and time. Model predictions are compared to the longwave measurements.

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A Brief Discussion of Uncertainty as it Relates to Space Nuclear Systems Initiative: Nuclear Safety Analyses Implementing a Managed Approach to Risk Communication James R. Coleman Victoria P. Friedensen1, Sandra M. Dawson2 James R. Coleman, Consulting P. O. Box 385, Cross, SC, 29436-0385 1Nuclear Systems Initiative, Office of Space Science, Voice & Fax:(843) 753 2612; [email protected] NASA Headquarters, Washington, DC 20546 2 Launch Approval Engineering Group, NASA Jet Propulsion Abstract. This paper defines uncertainty as the probabilistic Laboratory, 4800 Oak Grove Drive, Pasadena, CA 91109-8099 relationship between a set of actual observed (or observable) outcomes 1202/358-1916, [email protected] and the outcomes projected by a model. Some general facts about the mathematics of random variables are presented. It is observed that the major uncertainties in the Cassini safety analysis arise from subjective Abstract. NASA is developing a new model for risk communication judgments regarding environmental transport and human interaction that provides a managed approach to communication and outreach rather than the system or mission. It is concluded that a fair activities related to the development, design, use, and future flight of representation of mission and system safety can be provided only if space nuclear power systems. At this time, this program is being these two types of uncertainties are uncoupled and presented developed at Headquarters and the NASA Jet Propulsion Lab and will separately. later be extended to all the NASA Centers.

This model will feature two types of activities: a procedure for managing development, review and dissemination of risk communication products, and a program that will develop and conduct outreach and engagement activities that will be based on a broad exchange of ideas and information between NASA and the public on the use of these technologies. This approach will cover risk communication, education and outreach, and media relations; disseminate consistent, accurate, quality information on the benefits of the scientific knowledge enabled by space nuclear power; provide for proactive, cooperative engagement with a broad range of potential stakeholders including environmental and social justice organizations; and includes technology education-oriented outreach programs and materials.

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High Temperature Cermet Fuels – A Promising Review of the Historical Capabilities and Testing of Candidate for Space Reactors Composite and Cermet Fuels in Los Alamos

S.K. Bhattacharyya Robert J. Hanrahan Jr., Robert L. Smith III, Jason Morgan

Argonne National Laboratory, 9700 S. Cass Avenue Los Alamos National Laboratory, Materials Science and Technology Bldg. 360, Argonne, IL 60439 Division, MST-6, Metllurg, TA3 MSG 770 Los Alamos NM, 87545 (630) 252-3293, [email protected] 505-667-9560, [email protected]

Abstract. Ceramic (UO2, UN) fuel microspheres (100-200 µ size) Abstract. Fuel materials that are viable for the high temperatures and embedded in metal (Mo, W or their alloys) matrices, are known as aggressive environments of nuclear thermal or bimodal propulsion and cermet fuels. Despite the lack of recent work with these fuels (they power production fall into three basic groups, ceramics (monolithic were developed in the 1960’s as part of the ROVER program), there is oxides, carbides, and nitrides), graphite composites, (carbides or other considerable interest in applications of these fuels to Nuclear Thermal ceramics in a graphite matrix) and cermets (ceramic fuel materials in a Propulsion, Bimodal and Multimegawatt Power applications. This is refractory metal matrix). Of these materials the monolithic ceramics, because the limited data developed during the ROVER and Aircraft particularly carbides, are often cited as possessing the greatest high Nuclear Propulsion programs in the US and information obtained from temperature capability due the rationale that they possess the highest the former Soviet Union suggest that the fuel possesses an impressive melting points. The unarguable difficulty with these materials as set of attributes. monolithic fuels however is that they exhibit practically no ductility and consequently are not appropriate structural materials. The graphite The principal features of cermet fuels are their robustness and strength, matrix composites were extensively investigated during the ability to withstand high temperatures and repeated thermal cycles, Rover/Nerva program because they were both much easier to fabricate ability to retain fission products at high temperatures and potentially and exhibit greater toughness than monolithic ceramics. These high burnups, neutron capture cross-section features that provide materials were extensively tested under the Rover program however the inherent spectral softening and safety in the event of water immersion difficulties experienced with corrosion in flowing hydrogen were only accidents, and projected graceful failure modes at very high partially ameliorated by coatings consequently these fuels are of only temperatures. Its perceived drawback – large weight relative to limited interest today. graphite based cartride fuel - is mitigated by the inherent safety features, which makes an overall system mass comparable. While Another significant issue affects both ceramics and graphite composite binary or ternary carbide fuel have higher temperature capabilities in fuels; these are in most cases thermal reactor fuels, consequently they principle, the practical design advantages of cermet fuels make them require engineered safety systems to maintain subcritical configuration strong contenders for the applications mentioned. in case of a launch accident resulting in submersion. The cermet fuels using tungsten as the matrix (usually isotopically enriched W-184) are Cermet fuels have been successfully fabricated by several methods. necessarily fast reactor fuels which will not be affected by submersion Non-nuclear thermal tests in a hydrogen environment at temperatures in water. Furthermore cermet fuels are fabricable in nearly any up to 2900K, steady state irradiations at 1900K and burnup up to 1.6 conceivable shape, exhibit reasonable ductility, minimal fission product a/o, and transient nuclear tests at very large temperature gradients release, and excellent compatibility with hot hydrogen. (16000 K/s) showed excellent performance. The promise indicated by this older database needs to be realized through a detailed program of We will present a review of the work done at LANL (LASL) during the technology recapture and subsequent development and testing. Rover program on both graphite composite and cermet fuel forms and discuss the areas of research which we have concluded appear most attractive for development of bimondal nuclear thermal propulsion on the basis of this review.

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An Overview of Development and Testing of Uranium Reestablishing Fabrication Capabilities for Space Tri-carbide Fuels Nuclear Power Systems

Samim Anghaie and Travis Knight Jeffrey A. Halfinger, Barry G. Miller, DeWayne L. Husser

Innovative Nuclear Space Power and Propulsion Institute BWX Technologies, 1570 Mt. Athos Road, Lynchburg, VA, 24504 University of Florida, Gainesville, FL 32611-8300 (434) 522-5941,[email protected] (352) 392-1427, [email protected] Abstract. The National Aeronautics and Space Administration’s Abstract. The Innovative Nuclear Space Power and Propulsion recently announced Nuclear Systems Initiative has focused the Institute (INSPI) at the University of Florida has focused on spotlight on the need to reestablish a number of key technologies and improvements in the processing and fabrication of uranium tri-carbide infrastructure within the United States to support development and fuels with the goal of producing net-shape fuel elements. High solid- fielding of nuclear power systems for deep space science missions. phase solubility of uranium carbide in zirconium and niobium carbides Prominent among these are fabrication of highly enriched nuclear fuels provides for high flexibility in using very low to very high uranium including uranium nitride, uranium oxide, as well as some of the more fractions in the fuel. For requirements of compactness, high exotic fuels such as the binary and ternary carbides. Containment performance, and long life, space power reactors require low uranium structures for these fuels are also a top priority given the desire for fractions in the mixed carbide fuel and higher enrichments of uranium. systems that can operate over lifetimes of ten to twenty years. While very little work has been performed over the past decade on space The presence of non-uranium carbides in the fuel allows for gradient nuclear power systems, the nuclear industry has not remained stagnant. coating of fuel pellets and particles with refractory metal carbides, Manufacturing capabilities developed to support space nuclear which act as a robust barrier for containing fission products. No programs in the mid eighties and early nineties have continued to additional coating is necessary as with earlier graphite matrix and operate and advanced the state of the art in support of other nuclear composite fuels, which led to cracking of the coating and mass losses customers. This is the case at BWX Technologies where fabrication of due to corrosion by the hot hydrogen propellant. specialized nuclear fuels in various forms to very high enrichments has been ongoing since 1987. Recently, this expertise was applied to assess Processing and fabrication efforts at INSPI have developed methods of the feasibility of reestablishing manufacturing of highly enriched, cold pressing and UC liquid-phase sintering of near-stoichiometric and uranium nitride (UN) that could meet the specifications of the SP-100 hypo-stoichiometric tri-carbides to produce high-quality, single-phase, program. Over the course of a two-month period, fabrication of UN solid solution samples with less than 5% porosity. In light of the was successfully demonstrated through a process of internal gelation difficulties in extruding solid solution binary carbide fuel elements in and fluidized bed synthesis in a nitrogen-hydrogen atmosphere. This the Rover/NERVA experiments, an innovative space reactor fuel paper presents the results of that effort along with an overview of other geometry, square-lattice honeycomb (SLHC), was developed that key capabilities applicable to the NSI effort. would better lend itself to net-shape fabrication methods available to tri-carbide fuels. Efforts at testing and characterizing their performance under extreme NEP and NTP conditions are currently underway. The development and characterization of these fuels could lead to advanced NEP systems making future long-term, power rich space missions to Mars or other destinations possible.

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Development, Evaluation, and Design Applications of Effect of Long Term, High Temperature Annealing on the an AMTEC Converter Model – I. Strength of Beta”-Alumina Ceramics Development of the AMTEC Converter Model James R. Rasmussen1, Roger M. Williams2, and Adam K. Kisor2 Cliff A. Spence1a Michael Schuller1a Tom R. Lalk1b 1Advanced Modular Power Systems, Inc., 4370 Varsity Drive, Ann 1aCenter for Space Power, and 1bDepartment of Mechanical Arbor, MI 48104 2 Engineering, Texas A&M University, College Station, TX 77843 Jet Propulsion Laboratory, California Institute of Technology, 4800 Oak Grove Drive, Pasadena, CA 91109 1 Abstract. Issues associated with the development of an alkali metal (734) 677-4260 x158 [email protected] thermal-to-electric conversion (AMTEC) converter model that serves as an effective design tool were investigated. The requirements and Abstract. It has been recently reported that subjecting beta”-alumina performance prediction equations for the model were evaluated, and a ceramics to a long term, high temperature anneal for the purpose of modeling methodology was established. It was determined by defining reducing the residual sodium aluminate content within the ceramic the requirements and equations for the model and establishing a results in an apparent increase in the strength of the ceramic as well. In methodology that Thermal Desktop, a recently improved finite- order to examine this hypothesis, a carefully controlled experiment was difference software package, could be used to develop a model that conducted. Ten tubes were cut into 100 rings 1.5 mm long. A third of serves as an effective design tool. Implementing the methodology the rings (randomly selected) were broken in diametral ring fracture within Thermal Desktop provides stability, high resolution, modular tests, while the remaining rings were packaged and shipped to Jet construction, easy-to-use interfaces, and modeling flexibility. Propulsion Laboratory (JPL) where half of them were annealed, and the other half were unpacked and stored under appropriate dry conditions to act a shipping and handling control group. Once the annealing was completed, both groups of rings were repackaged and returned to Advanced Modular Power Systems (AMPS) and broken in diametral ring tests. The annealed group had the lowest strength as indicated by the Weibull characteristic strengths. Weibull characteristic strengths for the unannealed, control, and annealed groups were 376 MPa, 326 MPa, and 294 MPa, respectively. The Weibull moduli of the unannealed and annealed groups were nominally the same at 9.0 and 8.6, respectively. That for the handling control group was lower at 6.8. The lower strength of the annealed ceramics is consistent with earlier work showing a decrease in ceramic strength with increasing grain size.

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AMTEC Response to Changes in Resistive Loading Comparison of Measurement Techniques for Determining the Thermal Emittance of Robert W. Fletcher, Thomas K. Hunt Coupons at Elevated Temperatures Advanced Modular Power Systems, 4370 Varsity Drive Daniel P. Kramer1, Roger G. Miller1, Edwin I. Howell1, Ann Arbor, MI 48108 Donald A. Jaworske2, and Kenneth E.Wilkes3 734-677-4260, ext.202; fax: 734-677-0704; [email protected] 1Mound Power System Technologies, BWXT of Ohio, Inc., Abstract. An important aspect of electric power supply systems is P.O. Box 3030, Miamisburg, Ohio 45343 their inherent response time to rapid changes in loading demands. This 2NASA Glenn Research Center, 21000 Brookpark Road, presentation reviews the experimental response of an Alkali Metal Cleveland, OH 44135 Thermal Electric Converter (AMTEC) system when switched from 3Oak Ridge National Laboratory, P.O. Box 2008, open circuit to stable, resistive loads. Our data show a nominal 35- Oak Ridge, TN 37831 Watt AMTEC converter responded rapidly throughout the power curve. 937-865-3558; [email protected] Response times from open circuit to delivering 90% of peak DC current were within 0.25 milliseconds to 0.85 milliseconds for a range of Abstract. The development of a highly efficient nuclear space power electrically resistive loads at several typical AMTEC operational system requires that all of the available thermal energy emitted from temperatures. Such response times to load changes suggest that the General Purpose Heat Source (GPHS) modules (~250 thermal watts AMTEC may be suitable as a primary power supply, or backup power per module) be utilized in the most efficient manner. This includes supply for critical space applications. defining the heat transfer/thermal gradient profile between the surface of a GPHS module and the surface of the selected converter’s hot end. Control of the radiant heat transfer between the two surfaces can be achieved by regulating how efficiently the converter’s hot end surface transfers heat compared to a perfect blackbody (i.e. its infrared emittance). By oxidizing and/or grit blasting the surface of candidate converter materials it is possible to increase their emittance. L-605 test specimens were WC grit blasted and heat treated at 1023K for 72hours in air and their emittance values at elevated temperatures up to ~1000K were determined using three different measurement techniques (Infrared Camera, Infrared Reflectometer, and a Calorimetric Test Method).

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Ablation Radiation Shields for Nuclear Fusion Rockets First Results of the Gasdynamic Mirror Fusion Propulsion Experiment Luis Coreano1 and Brice N. Cassenti2 William J. Emrich, Jr. Rensselaer at Hartford and Pratt & Whitney, 400 Main Street M.S. 117-16, East Hartford, CT 06108 NASA - Marshall Space Flight Center, Huntsville, Alabama 35812 1 (860) 557-4336, [email protected] (256) 544-7504, [email protected] 2(860) 565-2522, [email protected] Abstract. An experimental Gasdynamic Mirror or GDM device has Abstract. Pulse nuclear propulsion has been the subject of extensive been constructed at the NASA Marshall Space Flight Center to provide studies since the 1960's. Early concepts examined external pulse an initial assessment of the applicability of this technology for propulsion where small critical mass nuclear devices are ejected from propulsion systems. This paper presents the first experimental results the rear of the rocket. A pusher plate absorbs some of the energy form obtained from the machine and an analysis of the types of plasma the detonation, which ablates the plate and provides thrust for the instabilities likely to be encountered. It is intended that this device rocket. It is also possible to have the device detonate in an enclosed operate at higher plasma densities and with much larger L/D ratios than chamber (i.e., internal pulse propulsion). Again, in this case, ablation is previous mirror machines. The high L/D ratio minimizes to a large the primary method for applying the thrust. Ablation can not only extent certain magnetic curvature effects which lead to plasma provide thrust but it can also aid in the dissipation of the heat in a instabilities causing a loss of plasma confinement. The high plasma neutron radiation shield. Since high-energy neutrons will be abundant density results in the plasma behaving much more like a conventional in deuterium-tritium fusion reactions, fusion rockets that use this fluid with a mean free path shorter than the length of the device. This reaction usually are designed with a radiator to dissipate the heat from characteristic helps reduce problems associated with "loss cone" the shield. These radiators usually require a considerable mass. microinstabilities. The device has been constructed to allow a Carbon and tungsten ablative shields may be considerably more considerable degree of flexibility in its configuration thus permitting effective. Ablation and radiation are compared as mechanisms to the experiment to grow over time without necessitating a great deal of dissipate the heat. Although ablation is shown to provide a additional fabrication. considerable mass saving heat loses at the surfaces will create thermal gradients that will adversely effect the ablation rate, and may significantly increase the mass loss.

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A Design Study of a p-11B Gasdynamic Mirror Antimatter Driven P-B11 Fusion Propulsion System Fusion Propulsion System Terry Kammash1, James Martin2, Thomas Godfory2

Chad Ohlandt1, Terry Kammash2, Kenneth G. Powell1 1Nuclear Engineering and Radiological Sciences, University of Michigan, Ann Arbor, MI 48109 2 1 NASA Marshall Space Flight Center, Huntsville, AL 35812 Department of Aerospace Engineering, (734) 764-0205, [email protected] University of Michigan, Ann Arbor, MI 48109 2 Department of Nuclear Engineering and Radiological Sciences, Abstract. One of the major advantages of using P-B11 fusion fuel is University of Michigan, Ann Arbor, MI 48109 that the reaction produces only charged particles in the form of three (734) 764-7573, [email protected] alpha particles and no neutrons. A fusion concept that lends itself to this fuel cycle is the Magnetically Insulated Inertial Confinement Abstract. Fusion gasdynamic mirror (GDM) space propulsion concepts Fusion (MICF) reactor whose distinct advantage lies in the very strong have been previously explored using deuterium, tritium, and helium-3 magnetic field that is created when an incident particle (or laser) beam fuels. This work is a similar design study using the advanced fusion 11 strikes the inner wall of the target pellet. This field serves to thermally fuel combination, hydrogen and boron-11. A GDM using p- B is insulate the hot plasma from the metal wall thereby allowing the optimized for the parameters of temperature, density, fuel ratio, and 11 plasma to burn for a long time and produce a large energy mirror radius. Even after optimization, a traditional GDM using p- B magnification. If used as a propulsion device, we propose using and achieving breakeven appears to be impractical due to antiprotons to drive the system, which we show to be capable of bremsstrahlung and synchrotron radiation losses. A nuclear electric producing very large specific impulse and thrust. By way of validating assisted version of the system is examined and found to decrease the the confinement properties of MICF we will address a proposed size and mass of the system. The optimal plasma temperature is also experiment in which pellets coated with P-B11 fuel at the appropriate reduced by the assistance which decreases the technical requirements ratio will be zapped by a beam of antiprotons that enters the target for magnetic confinement. through a hole. Calculations showing the density and temperature of the generated plasma along with the strength of the magnetic field and other properties of the system will be presented and discussed.

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Implementing a New Line of Medium-Class, Unmanned Considerations and Lessons Learned in Implementing Space Exploration Missions Effective, Low-Cost, Unmanned Space Exploration Missions Thomas H. Morgan and Susan M. Niebur R. Brad Perry1 and Dennis G. Pelaccio2 Solar System Exploration Division, Office of Space Science, NASA Headquarters, Washington, DC 20546 1Earth and Space Science Support Office, NASA Langley Research (202) 358-0828, Fax:(202) 358-3097, [email protected] Center, Hampton, VA 23681 2Science Applications International Corporation, Suite 100, Littleton, Abstract. The NASA Office of Space Science will soon be releasing CO 80127 the first Announcement of Opportunity to propose new, medium-class, 1(757) 864-8257, FAX (757) 864-8894, [email protected] missions for Solar System Exploration. This new mission line, called 2(720) 981-2491, FAX (720) 981-7488, [email protected] New Frontiers, has been created in order to allow greater access to a larger variety of targets in our Solar System. Missions may be of larger Abstract. Since the early 1990s, NASA has implemented a number of scope than the current Discovery program missions; proposed costs for low-cost, unmanned space exploration missions including the recent building, launching, operating, and analyzing data from the New missions associated with the Mars Exploration Program, and also Frontiers missions may be as large as $650M (Fiscal Year 2003 missions in the Discovery and Explorer Programs. This paper dollars) and proposed launch date may be up to 47 months after addresses a number of the system design and implementation development is initiated. New Frontiers missions may use radioisotope considerations and lessons learned that must be addressed for these power sources (e.g. RTGs) and larger launch vehicles than allowed for missions to be successful, and stresses that effective systems Discovery missions, due to the greater allowed cost. Each proposal engineering practices must be adopted and applied consistently opportunity will be focused on proposals for investigations that address throughout all mission project phases. Emphasis is placed on most if not all of the science objectives of one or more of the “Medium highlighting effective approaches that have been successfully used in Class” investigations identified in the recent National Academy past programs. These considerations and lessons learned should be Planetary Science Decadal Survey (National Research Council, 2002). considered in formulating effective future technology demonstration The management of this program is to be similar to Discovery, with a and science exploration mission efforts for NASA’s Nuclear Systems few significant differences. This paper describes the New Frontiers Initiative (NSI) and New Frontiers Program. program, compares it to the successful Discovery program, and discusses the challenges involved in the selection and implementation of New Frontiers missions

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Lessons Learned in Sending an Ion-Drive System to Design Concept for a Nuclear Reactor-Powered Mars Deep Space Rover

David H. Lehman John O. Elliott1, Ronald J. Lipinski2, David I. Poston3

Jet Propulsion Laboratory, 4800 Oak Grove Drive, 1Mission and Systems Architecture Section, Jet Propulsion Laboratory, Pasadena, CA 91109 Pasadena, CA 91109 818-354-2023; FAX: 818-393-4277 [email protected] 2 Advanced Nuclear Concepts, Sandia National Laboratories, Albuquerque, NM 87185 Abstract. Deep Space 1 (DS1), launched on October 24, 1998, was the 3 Nuclear Systems Design, Los Alamos National Laboratory, Los first mission of NASA’s New Millennium program. DS1 was one of Alamos, NM 87545 NASA’s “faster, better, cheaper” missions chartered to flight test (818) 393-5992; [email protected] twelve high-risk, enabling technologies important for future space and Earth science programs on both a fast schedule and a low budget. Abstract. A study was recently carried out by a team from JPL and the Among its firsts, DS1 was the first deep space mission to use ion DoE to investigate the utility of a DoE-developed 3 kWe surface fission propulsion to actually go somewhere (asteroid Braille in July of 1999 power system for Mars missions. The team was originally tasked to and Comet Borrelly in September of 2001) and the first mission to use perform a study to evaluate the usefulness and feasibility of a totally autonomous on-board navigation system. In addition, another incorporation of such a power system into a landed mission. In the of its autonomous systems, called the Remote Agent Experiment, was course of the study it became clear that the application of such a power awarded NASA’s 1999 Software of the Year award. system was enabling to a wide variety of potential missions. Of these, two missions were developed, one for a stationary lander and one for a Concept studies for the project were initiated in July of1995. DS1’s reactor-powered rover. This paper discusses the design of the rover prime mission was successfully completed in September of 1999. mission, which was developed around the concept of incorporating the Advanced technologies flight-tested during the mission included ion fission power system directly into a large rover chassis to provide high propulsion, high-power solar concentrator arrays, three on-board power, long range traverse capability. The rover design is based on a autonomy technologies, two low-mass science instrument packages, minimum extrapolation of technology, and adapts existing concepts and several telecommunications and microelectronics devices. developed at JPL for the 2009 Mars Science Laboratory (MSL) rover, lander and EDL systems. The small size of the reactor allowed its The author was project technical manager of DS1 from the beginning incorporation directly into an existing large MSL rover chassis design, of spacecraft development through the end of the prime mission allowing direct use of MSL aeroshell and pallet lander elements, beefed operations phase. This presentation will describe the lessons learned up to support the significantly greater mass involved in the nuclear on sending an ion-drive system to deep space. power system and its associated shielding. This paper describes the unique design challenges encountered in the development of this mission architecture and incorporation of the fission power system in the rover, and presents a detailed description of the final design of this innovative concept for providing long range, long duration mobility on Mars.

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High Efficiency Thermoelectrics in NEP Conceptual Design of a 100-kWe Space Nuclear Reactor Reactor Power Systems Power System with High-Power AMTEC

Daniel T. Allen, Saeid Ghamaty & Norbert B. Elsner Mohamed S. El-Genk and Jean-Michel Tournier

Hi-Z Technology, Inc., 7606 Miramar Road, San Diego, California, Institute for Space and Nuclear Power Studies and Department of USA 92126 4210 Chemical and Nuclear Engineering +1 858 695 6660, @hi-z.com The University of New Mexico, Albuquerque, NM, 87131 (505) 277 – 5442, Fax: (505) 277 – 2814, [email protected] Abstract. Thermoelectric space reactor power systems that utilize Multi-Layer Quantum Well (MLQW) technology are presented and Abstract. Alkali Metal Thermal-to-Electric Conversion (AMTEC), discussed in the context of Nuclear Electric Propulsion (NEP). although currently at a Technology Readiness Level-3 (TRL-3), has an Quantum wells are one of the recent developments in low-dimensional excellent potential for use in Space Nuclear Reactor Power (SNRP) thermoelectric materials that show a factor of 2.5 increase in the systems for NASA’s deep-space exploration missions. In addition to thermoelectric figure of merit. This breakthrough in converter operating at a conversion efficiency > 20%, representing the highest performance promises higher efficiency power generating devices. The fraction (> 60%) of Carnot efficiency of all other static and dynamic MLQW under development at Hi-Z Technology, Inc., applied to space conversion technology options, the relatively high heat rejection radiator systems provides the design flexibility traditionally available with temperature (650–700 K) reduces the size and mass of the radiator and of thermoelectric conversion in reactor power systems with higher the SNRP system. A high-power AMTEC unit design has been developed performance. The reactor concept evaluated is the Heatpipe Power and optimized for operating at reactor exit temperatures < 1180 K and radiator temperature < 680 K. Depending on the reactor exit temperature, System (HPS) reactor. the nominal electrical power of the AMTEC unit, measuring 594 mm x 410 mm x 115 mm and weighting 44.3 kg, could be as high as 5.6 kWe, with a margin of > 5% for an additional load-following increase. A conceptual design of a 100 kWe SNRP system with these high-power AMTEC units is developed and presented in this paper. The total mass of major subsystems, including the converters, nuclear reactor, shadow radiation shield, and radiator, is calculated and compared with that for the SP-100. Despite the large specific mass of the AMTEC units compared to the SiGe thermoelectrics in the SP-100 system, the lower masses of the reactor, radiation shield, and radiator make the present AMTEC-SNRP system > 26% lighter, for the same electrical power. An optimized AMTEC-SNRP system could potentially operate at a specific power > 30 We/kg (or specific mass < 33 kg/kWe), use non-refractory structures of super-steel alloys with well-know properties, relatively low density, low Ductile-To- Brittle (DTB) transition temperatures, and good compatibility with space and planetary environments containing CO2 and oxygen. The radiator area for the baseline 100 kWe AMTEC SNRP system is < 27 m2, which, together with operating the potassium heat pipes in the radiator sonic limited shortly after a reactor shutdown, would extensively prolong the cool-down time of the reactor from several days to many months, before freezing the sodium in the reactor’s heat pipes (371 K).

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Power Needs For Planetary Surface Exploration Review of Power System Options for Human Planetary Exploration S. Hoffman1 and J. George2 Robert L. Cataldo 1Science Applications International Corporation, Houston, TX, USA; 2 NASA Johnson Space Center, Houston, TX NASA Glenn Research Center 281-483-9364, [email protected] Cleveland, OH 44135

Abstract. Over the last several years, the Exploration Office at the Abstract. Strategies for human planetary surface exploration of the NASA Johnson Space Center has investigated a number of scenarios Moon, Mars and most recently Callisto, a moon of Jupiter, have been for the exploration of the Moon and Mars by humans and robots. proposed. A design goal of any human mission is to reduce risks, but it These scenarios have included specific analyses of the power is particularly critical in human exploration missions beyond cislunar requirements for various systems that are required for or significantly space. One method of reducing risk is “know before you go”, where a enhance these exploration activities. Nuclear and non-nuclear power 800 day cache of life support and ascent vehicle propellants are solutions have been identified for these activities. The presentation will manufactured and stored prior to crew departure from Earth. Life discuss functional needs, and where possible, numerical estimates, for support materials; oxygen, water and buffer gasses and spacecraft the power usage of three of these systems: life support for a fixed return fuel is not imported from Earth, but rather manufactured from habitat, surface mobility (i.e., rovers), and in situ resource utilization. Mars’ indigenous resources. This scenario has many impacts on the In addition, the Exploration Office has developed a draft set of nuclear surface systems that are sent and deployed years prior to crew power requirements for human exploration missions. Highlights of departure. Specific impacts are levied on the power system such as self these requirements, as they apply to planetary surface exploration, and deployment or tele-operated with time delay impacts, untended associated rationale will be presented. operation and need for autonomous system controllers are some such issues.

Characteristics and features of nuclear versus solar power systems, including requirements for deployment and maintenance will be compared. Nuclear power systems have an advantage of lower mass, volume, and deployed area as contrasted to solar power systems. Although the nighttime energy storage requirements are less demanding for a Mars base versus a lunar base resulting in lower battery mass, solar arrays are large due to Mars’ distance from the Sun and the solar insolation attenuation caused by atmospheric conditions. Alternative power sources for surface mobility systems, ranging from long range pressurized rovers to piloted and robotic utility rovers, will also be compared.

An assessment of past mission studies and proposed power systems in will be complied comparing and contrasting different concepts and technologies. Technology roadmaps or blueprints will be evaluated suggesting the most promising technology to encompass a wide and varied set of mission and the difficulty of achieving each technology path.

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SAFE Testing Nuclear Rockets Economically Nuclear Thermal Rocket Exhaust Conditioning In Open Cycle and Closed Cycle Systems Steven D. Howe, Bryan Travis and David K. Zerkle Stanley V. Gunn Los Alamos National Laboratory, P.O. Box 1663, MS-H845 Los Alamos, NM 87545 Rocketdyne-Retired, 20300 Tau Place, Chatsworth, CA, 91311 505-665-9367, [email protected] Abstract. Environmentally compliant approaches for the ground testing of nuclear thermal rocket engines are critical to the successful development of Abstract. Several studies over the past few decades have recognized nuclear propulsion systems in the 2000's. To meet these requirements, a NTR the need for advanced propulsion to explore the solar system. As early propulsion test facility will be required to prevent the release of TBD amounts as the 1960s, Werner Von Braun and others recognized the need for a of radioactivity under normal operating conditions, as well as possible nuclear rocket for sending humans to Mars. The great distances, the accidental events in the propulsion system or the core of the test article. Before facility design can commence, failure modes and effects analysis should be intense radiation levels, and the physiological response to zero-gravity performed for the expected test articles. In addition, the test facility should be all supported the concept of using a nuclear rocket to decrease mission useful in the conduct of development testing of both non-nuclear components time. These same needs have been recognized in later studies, and subsystem tests, as well as appropriate nuclear subsystem tests. Finally, especially in the Space Exploration Initiative in 1989. One of the key the facility should be adaptable to NTR thrusts ranging from 15,000 lbf to questions that has arisen in later studies, however, is the ability to test a 125,000 lbf. nuclear rocket engine in the current societal environment. Unlike the Rover/NERVA programs in the 1960s, the rocket exhaust can no longer The nozzle exhaust gas temperatures are expected to range from 2300 to 3000 be vented to the open atmosphere. As a consequence, previous studies K, with sustained entrained fission products less than two percent of the rate of fission fragments being produced in the reactor core. However, if abnormal have examined the feasibility of building a large-scale version of the deterioration of the fuel elements develop, the test operator has the Nuclear Furnace Scrubber that was demonstrated in 1971. We have responsibility to terminate the test if the ejected radioactivity exceeds preset investigated an alternative that would deposit the rocket exhaust along limits. Further, if the reactor coolent supply system should fail, the test with any entrained fission products directly into the ground. The conductor must have access to an emergency coolent flow system to permit a Subsurface Active Filtering of Exhaust, or SAFE, concept would allow controlled shutdown of the operating test system. variable sized engines to be tested for long times at a modest expense. A system overview, results of preliminary calculations, and cost The exhaust gases processing approach presented herein employs indirect estimates of proof of concept demonstrations are presented. The results cooling of the gases using either, for an open cycle system. two (first water, indicate that a nuclear rocket could be tested at the Nevada Test Site for then liquid nitrogen) heat exchangers, or a closed cycle system, three (first water, next liquid nitrogen, then liquid hydrogen) heat exchangers. The under $20 M. scrubber system employs beds of activated charcoal to remove the fission gases, primarily xenon and krypton gases, and any entrained fuel element particles. With the closed cycle system, it is possible to recondense the exhaust gases, return them to the run tank and continue operation for as long as the supply of heat exchanger coolents will permit.

The NTR sytstem technology advances achieved under the Rover and NERVA programs had reached a level that would support the design of a NTR flight propulsion system that would not need any engine radiation shield. Thus the design of the foreseen test facility should accomodate realistic test radiation exposure upon the propellant inlet ducting and the flight-type tank bottom. Finally, a medical-realistic limit must be established for test crews working in the test stand area.

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Technologies to Improve Ion Propulsion System Life Fusion Propulsion Through a Magnetic Nozzle and Efficiency and Open Divertor

Ira Katz, John R. Brophy, John R. Anderson, James E. Polk Craig H. Williams1a, Pavlos G. Mikellides2, Ioannis G. Mikellides3, 4 1b Jet Propulsion Laboratory/ California Institute of Technology Richard A. Gerwin , Ian J. Dux 4800 Oak Grove Drive, Pasadena, CA 91109 818-393-6948, [email protected] 1a Space Transportation Project Office and 1bSystems Engineering Division, NASA Glenn Research Center, Cleveland, OH 44135 Abstract. The performance of nuclear electric spacecraft critically 2 Department of Mechanical and Aerospace Engineering, depends on the ability of the electric propulsion system to reliably Arizona State University, Tempe, AZ 85287 provide high ISP thrust with great efficiency. Missions to the outer 3 Science Applications International Corporation, planets will require thrusters to operate for the order of ten years, San Diego, CA, 92121 several times the life of the state of the art NSTAR thruster. Propulsion 4 Department of Plasma Physics, DOE Los Alamos National Lab, system efficiency is a multiplying factor in the overall system “alpha”, Los Alamos, NM 87545 the determining parameter in how well a nuclear electric spacecraft (216) 977 – 7063, [email protected] performs on deep space missions. In order to help make NEP a reality, we are developing models of the physical processes that control ion Abstract. A revised magnetic nozzle and open poloidal bundle divertor propulsion system life and efficiency. concept is proposed for a small aspect ratio, spherical torus reactor to In this paper we present results from a series of new computer codes be used in direct nuclear fusion space propulsion. A preliminary developed at JPL that model the performance limiting and erosion analysis of convergent/divergent flow through a magnetic nozzle mechanisms of ion thrusters. These codes model extraction grid ion illustrated the importance of high efficiency, where a power conversion optics, the discharge chamber, and both discharge and neutralizer efficiency of only 75% could lead to a doubling of exit area ratio hollow cathodes. Basic plasma physical processes including ionization, requirements over ideal conditions. Preliminary MHD simulations have electron transport, and charge exchange, are solved in the governing shown that as much as 50% of the mass flow could penetrate into the equations. The grid ion optics and erosion model have been validated plasma-magnetic field boundary layer, leading to a thermal-to-directed with NSTAR life test results. As an example of the power of the jet power loss of 30% and a spacecraft payload mass ratio loss of ~ computer models, we present a design for a set of 7000 second Isp ion 60% of ideal. The importance of understanding the plasma-magnetic thruster grids for operation at 12kW on the 40 cm NEXT discharge field interface and the inability of space/time dependent MHD chamber. These grids have a projected throughput life in excess of simulations to resolve fine-scaled, gradient-driven micro-instabilities is 1000kg when fabricated from carbon-carbon materials. currently under study. Preliminary plasma interface broadening mechanism studies have concluded that significant attachment of the plasma onto the magnetic field can occur at the nozzle inlet, that anomalous resistivity is likely to be a significant issue in the plasma regime of interest, that Rayleigh-Taylor instabilities may not be of immediate concern (but cannot be discounted), and that full Hall current effects must be incorporated into the ongoing simulations.

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Test Facilities in Support of High Power Electric Thermal Stress Calculations for Heatpipe-Cooled Propulsion Systems Reactor Power Systems

Melissa Van Dyke, Mike Houts, Thomas Godfroy, Ricky Dickens, Richard J. Kapernick1a, Ray M. Guffee1b James J. Martin, Patrick Salvail, and Robert Carter 1a Nuclear Systems Design, Decision Applications Division Marshall Space Flight Center, National Aeronautics and Space 1b Design Engineering, Engineering Sciences and Applications Division Administration, Huntsville, Alabama, 35812 Los Alamos National Laboratory, Los Alamos, NM 87545 (256) 544-5720, [email protected] 505-665-0526; [email protected]

Abstract. Successful development of space fission systems requires an Abstract. A heatpipe-cooled fast reactor concept has been under extensive program of affordable and realistic testing. In addition to tests development at Los Alamos National Laboratory for the past several related to design/development of the fission system, realistic testing of years, to be used as a power source for nuclear electric propulsion the actual flight unit must also be performed. If the system is designed (NEP) or as a planetary surface power system. The reactor core to operate within established radiation damage and fuel burn up limits consists of an array of modules that are held together by a core lateral while simultaneously being designed to allow close simulation of heat restraint system. Each module includes a single heatpipe surrounded from fission using resistance heaters, high confidence in fission system by 3-6 clad fuel pins. As part of this development effort, a partial array performance and lifetime can be attained through non-nuclear testing. of a candidate heatpipe-cooled reactor is to be tested in the SAFE-100 Through demonstration of systems concepts (designed by DOE experimental program at the Marshall Space Flight Center. The partial National Laboratories) in relevant environments, this philosophy has array comprises 19 3-pin modules, which are powered by resistance been demonstrated through hardware testing in the High Power heaters. This paper describes the analyses that were performed in Propulsion Thermal Simulator (HPPTS). The HPPTS is designed to support of this test program, to assess thermal and structural enable very realistic non-nuclear testing of space fission systems. performance and to specify the test conditions needed to simulate Ongoing research at the HPPTS is geared towards facilitating research, reactor operating conditions. development, system integration, and system utilization via cooperative efforts with DOE labs, industry, universities, and other NASA centers. Through hardware based design and testing, the HPPTS investigates High Power Electric Propulsion (HPEP) component, subsystem, and integrated system design and performance.

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Direct-Drive Gas-Cooled Reactor Power System: Mechanical Design and Fabrication of a SAFE-100 Heat Concept and Preliminary Testing Exchanger for Use in NASA's Advanced Propulsion Thermal-hydraulic Simulator S. A. Wright1, R. J. Lipinski1, T. J. Godfroy2, S. M. Bragg-Sitton2, 2 M. K. Van Dyke Robert W. Carter1, Ray M. Guffee1, Russell L. Rosmait2, Pat Salvail3

1 Sandia National Laboratories, P.O. Box 5800, Albuquerque, NM 1Los Alamos National Laboratory, Los Alamos, NM 87545 87185 2 Engineering Technology, Pittsburgh State University, Pittsburgh, KS 2 Marshall Space Flight Center, National Aeronautics and Space 66762 Administration, Huntsville, Alabama, 35812 3NASA Marshall Space Flight Center, Huntsville, AL 35812 (256) 544-1104, [email protected] (505) 667-1374, Fax: (505) 667-3559, [email protected] Abstract. The SAFE-100 reactor is a heatpipe-cooled fission power Abstract. This paper describes the concept and preliminary source proposed for both nuclear electric propulsion (NEP) and component testing of a gas-cooled, UN-fueled, pin-type reactor which planetary surface missions. One possible power conversion uses He/Xe gas that goes directly into a recuperated Brayton system to arrangement is the coupling of a Brayton cycle to the heatpipe reactor. produce electricity for nuclear electric propulsion. This Direct-Drive A key component in this reactor/Brayton system is the heat exchanger. Gas-Cooled Reactor (DDG) is designed to be subcritical under water or This heat exchanger is directly attached to the heatpipe core and wet-sand immersion in case of a launch accident. Because the gas- transfers heat to the Brayton cycle working fluid. A series of cooled reactor can directly drive the Brayton turbomachinery, it is performance tests are planned at NASA on a 19 heatpipe module, possible to configure the system such that there are no external surfaces SAFE-100 core with a representative heat exchanger functioning as the or pressure boundaries that are refractory metal, even though the gas primary heat removal system. These tests provide an opportunity to delivered to the turbine is 1144 K. The He/Xe gas mixture is a good develop a generic heat exchanger design, which could be used on other heat transport medium when flowing, and a good insulator when SAFE reactors, as well as developing manufacturing processes and stagnant. Judicious use of stagnant cavities as insulating regions allows techniques used in the heat exchanger fabrication. This paper details transport of the 1144-K gas while keeping all external surfaces below the mechanical design of the SAFE-100 heat exchanger and describes 900 K. At this temperature super-alloys (Hastelloy or Inconel) can be the prototyping studies and manufacturing operations that were used instead of refractory metals. Super-alloys reduce the technology performed during fabrication. risk because they are easier to fabricate than refractory metals, we have a much more extensive knowledge base on their characteristics, and, because they have a greater resistance to oxidation, system testing is eased. The system is also relatively simple in its design: no additional coolant pumps, heat exchanger, or freeze-thaw systems are required. Key to success of this concept is a good knowledge of the heat transfer between the fuel pins and the gas, as well as the pressure drop through the system. This paper describes preliminary testing to obtain this key information, as well as experience in demonstrating electrically heated testing of simulated reactor components.

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Progress in the Development of High Efficiency Thermal Stability Characterization of Skutterudite Segmented Thermoelectric Unicouples Antimonide, Arsenideand Phosphide

Thierry Caillat, Jeff Sakamoto, Jean-Pierre Fleurial and Jeff Snyder Jeff Sakamoto, Thierry Caillat, Jean-Pierre Fleurial, and Virgil Shields

Jet Propulsion Laboratory, California Institute of Technology, 4800 Jet Propulsion Laboratory, California Institute of Technology, 4800 Oak Grove Drive, MS 277/207, Pasadena, California Oak Grove Drive, MS 277/207, Pasadena, California (818) 393-6693, [email protected] 818 354-0407, [email protected] Abstract. High-efficiency, segmented thermoelectric unicouples

incorporating advanced thermoelectric materials with superior Abstract. This paper presents and discusses the latest results obtained thermoelectric figures of merit are currently being developed at the Jet on the development of highly efficient, segmented thermoelectric Propulsion Laboratory (JPL). These segmented unicouples include a unicouples at the incorporate include a combination of state-of-the-art combination of state-of-the-art thermoelectric materials based on thermoelectric materials based on Bi Te and novel, high ZT p-type p- 2 3 Bi Te and novel p-type Zn Sb , p-type CeFe Sb -based alloys and n- type CeFe Sb -based alloys and n-type CoSb -based alloys developed 2 3 4 3 4 12 4 12 3 type CoSb -based alloys developed at JPL. The paper reports at JPL. The maximum predicted thermal to electrical efficiency is about 3 investigations concerning the thermal stability of these novel materials. 15% for a hot-side temperature of 975K and a cold-side temperature of The primary emphasis is on characterizing sublimation behavior of Sb, about 300K. Various segmentations have been explored and several As, and P from the skutterudite compounds. Additionally, methods of unicouples have been fabricated and tested. A new thermal and suppressing sublimation are also presented and discussed. electrical testing set-up and procedure is described in this paper and some tests results reported. In addition, a number of initial life test results on unicouple components are presented and discussed. Finally an overview of the development progress made to date and future work is presented.

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Potential Improvements in Skutterudite Thermoelectric Engineering Nanostructures for Efficient Properties due to Solid Solution Formation Thermoelectric Power Conversion

Virgil Shields, Thierry Caillat, Jean-Pierre Fleurial, Andrew Zoltan, Gang Chen Leslie Zoltan, Matthew Tuchscherer. Mechanical Engineering Department, Massachusetts Institute of Jet Propulsion Laboratory Technology, Cambridge, MA 02139 California Institute of Technology 617-253-0006, [email protected] 4800 Oak Grove Drive, Pasadena, California 91109 Abstract. Efficient solid-state energy conversion based on the Peltier 818-354-9506, [email protected] effect for cooling and the Seebeck effect for power generation calls for materials with high electrical conductivity, high Seebeck coefficient, Abstract. High temperature skutterudite thermoelectric semiconductors and low thermal conductivity. Identifying materials with good possess attractive transport properties. These materials have the thermoelectric figure-of-merit has proven to be an extremely potential for achieving high figures of merit in a temperature range of challenging task. In nanostructures, quantum and classical size effects 973 to 1273 K. The binary compounds however tend to possess and interface effects provide opportunities to tailor the electron and thermal conductivities that are too high. The formation of solid phonon transport through structural engineering. Quantum wells, solutions, or alloying, however, is well known to substantially reduce superlattices, quantum wires, and quantum dots can be employed to the thermal conductivity of semiconductors. Due to the relatively large change the band structure, energy levels, and density of states of numbers of skutterudite compounds, solid solutions and related phases electrons, leading to potentially improved energy conversion capability exist that can potentially offer many possibilities for optimizing the of charged carriers compared to those of their bulk materials. Interfaces thermal and electrical properties for a range of thermoelectric can be used to filter electrons and to reflect phonons. This presentation applications. In this work several of these skutterudite solid solutions will discuss both theoretical and experimental investigations on various systems have been investigated consisting of Co-Ni-P, Co-As-P, Co- size and interface effects on the electron and phonon transport for solid- Sb-P and Co-Sb-As along with the results due to various dopants. The state energy conversion applications. effects on the thermoelectric properties of Co1-X NiXP3, CoAs3-YPY, CoSb3-YPY and CoSb3-YAsY are presented which indicate a reduction in thermal conductivity. Seebeck and electrical resistivity measurements are also discussed.

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Looking Back on Rover Nuclear Thermal Rocket – An Established Space Propulsion Technology C. Paul Robinson Milton Klein Sandia National Laboratories, P. O. Box 5800, Mailstop 0101

Albuquerque, NM 87185-0101 NASA/AEC Space Nuclear Systems Office (505) 844-7261, [email protected] Group Vice President (retired), Electric Power Research Institute, Abstract. The Rover Program was a major research, development, and 48 Politzer Drive, Menlo Park, CA 94025 test program to build a nuclear reactor powered rocket engine that (650) 329-9261, [email protected] could be used for interplanetary travel and deep space maneuvers. The program was established in 1958, continued until 1973, and ended due Abstract. From the late 1950s to the early 1970s a major program to an NASA decision to abandon its plan for human exploration of successfully developed the capability to conduct space exploration Mars (which had been seen as the follow-on to the Apollo lunar using the advanced technology of nuclear rocket propulsion. The explorations.) It was also clear then that the increasing restraints on program had two primary elements: pioneering and advanced testing of the rocket engines—by exhausting the hydrogen coolant into technology work---Rover---at Los Alamos National Laboratory and its the air—would no longer be environmentally acceptable. Rover was a contractors provided the basic reactor design, fuel materials unique technology test bed and produced advances in many fields, development, and reactor testing capability; and engine development--- including materials science, control systems, cryogenic pumps, and NERVA---by the industrial team of Aerojet and Westinghouse building remote assembly and disassembly. The unique circumstances of the on and extending the Los Alamos efforts to flight system development. test configurations also allowed many “firsts” to be achieved in “add- on experiments” to the reactor/rocket tests. This presentation describes the NERVA program, the engine system testing that demonstrated the space-practical operation capabilities of nuclear thermal rockets, and the mission studies that point the way to most effectively use the NTR capabilities. Together, the two programs established a technology base that includes proven NTR capabilities of (1) over twice the specific impulse of chemical propulsion systems, (2) thrust capabilities ranging from 10,000 to 250,000 lbf, and (3) practical thrust-to-weight ratios for future NASA space exploration missions, both manned payloads to Mars and unmanned payloads to the outer planets.

The overall nuclear rocket program had a unique management structure that integrated the efforts of the two government agencies involved— NASA and the then-existing Atomic Energy Commission. The lessons learned through the use of that structure provide an important template for the future.

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Looking Backward, Looking Forward Nuclear Thermal Rocket Ground Test Potential at NASA Glenn’s Plum Brook Station Gary L. Bennett Brian P. Willis, Mark R. Woike NASA Glenn Research Center/Plum Brook Station, 6100 Columbus 5000 Butte Road, Emmett, Idaho 83617-9500 Ave, Sandusky, OH44870 Telephone/Fax: 1+208.365.1210, [email protected] (419) 621-3202, [email protected] Abstract. With the advent of NASA's Nuclear Systems Initiative Abstract. NASA Glenn Research Center’s Plum Brook Station, a (NSI) it is appropriate to review recent flight programs in space nuclear 6300 acre federal reserve located in rural Erie County (50 miles west of power to ascertain what lessons can be drawn that may be of general Cleveland), plays host to the Hypersonic Tunnel Facility, HTF. The applicability to NSI to aid in its success. genesis of the HTF was the Hydrogen Heat Transfer Facility, HHTF, conceived and constructed in support of the NERVA program. Central to the design philosophy of the experimental infrastructure of the test site is a large graphite storage heater. This 3.5MW induction heater is the only large-scale, non-reacting heater core that is operational with NASA. At a bulk temperature of 2760C the 27,216kg of graphite is storing approximately 137,150MJ of energy. An oral presentation on both facilities will be offered. Topics to be covered will include the historical background of the original HHTF, a current reality description of the HTF experimental infrastructure, as well as some potential future uses focused on NTR development activities.

The Hydrogen Heat Transfer Facility was designed to process 7kg/sec of gaseous hydrogen through the graphite storage heater resulting in a very high enthalpy (~2000C) gas stream. A high pressure motive force (~68 atmospheres) was used to accelerate the heated stream through a rocket nozzle. Anecdotal evidence indicates that the facility was proof tested using nitrogen gas. The use of hydrogen was never attempted partly because the NERVA program was terminated.

The Hypersonic Tunnel Facility was created from the ashes of the HHTF. In this configuration the graphite heater is used to create a very high enthalpy gas stream consisting of nitrogen to which is added oxygen. This creates a synthetic air gas stream that is accelerated through a convergent-divergent nozzle to hypersonic speeds. In this fashion the facility is capable of producing Mach 5, Mach 6 and Mach 7 flight-enthalpy flows.

Combined with the corporate knowledge of the HTF Team, the existing experimental infrastructure is well suited to support nuclear thermal rocket testing. An existing (smaller) non-reacting heater core could be modified to support core configuration studies as well.

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Recent Developments in Mixed Ionic and Electronic High Power AMTEC Converters for Deep-Space Conducting Electrodes for the Alkali Metal Thermal Nuclear Reactor Power Systems Electric Converter (AMTEC) Mohamed S. El-Genk and Jean-Michel Tournier Robert W. Fletcher1, 2, Johannes W. Schwank2 Institute for Space and Nuclear Power Studies and Department of 1Advanced Modular Power Systems (AMPS), 4370 Varsity Drive Chemical and Nuclear Engineering Ann Arbor, Michigan 48108 The University of New Mexico, Albuquerque, NM, 87131 2Department of Chemical Engineering, University of Michigan, 2300 (505) 277 – 5442, Fax: (505) 277 – 2814, [email protected] Hayward Street, 3030 H.H. Dow Building Ann Arbor, Michigan 48109-2136 Abstract. A high electrical power, Alkali Metal Thermal-To-Electric 734-677-4260, ext. 202; fax: 734-677-0704; [email protected] Conversion (AMTEC) unit design is developed, and performance estimates as functions of the beta”-alumina solid electrolyte (BASE) Abstract. Mixed ionic and electronic conducting electrodes (MIECEs) temperature (or anode vapor pressure), condenser temperature, and the have recently gained more attention in the development of the Alkali type of the working fluid (sodium or potassium) are calculated and Metal Thermal Electric Converter (AMTEC). The advantage of discussed. The Na and K-AMTEC units, measuring 410 mm x 594 mm MIECEs, as their name implies, is that they allow both electronic x 115 mm in outside dimensions, are identical except for the type of the transport and ionic transport within the matrix of their materials. This BASE and working fluid. The peak efficiency of the Na-AMTEC at a lowers charge transport resistances, which lowers overall electrical BASE temperature of 1123 K is 29.2%, decreasing to 26.8% at a BASE resistances of the electrochemical system, and, thus, can improve temperature of 1073 K. The corresponding specific powers of the Na- reaction rates for a given apparent or superficial surface area of the AMTEC unit are 76 and 54 We/kg, respectively. For nominal electrode. The AMTEC system is a self-contained, self-regenerating operation at 85% of peak electrical power at BASE temperatures of compact technology that electrochemically converts heat directly to 1123 K and 1073 K, the conversion efficiency of the Na-AMTEC electricity. Our latest developments for AMTEC MIECEs are decreases to 26.7% and 24.5%, respectively, but the corresponding presented. These include electrodes formulated from various blends of specific powers increase significantly to 125 and 91 We/kg, titanium nitride or molybdenum with titanium dioxide. The general respectively. The Na-AMTEC nominally generates 4.0 and 5.6 kWe formulation, application, processing and testing methods for these when operating at BASE temperatures of 1073 K and 1123 K, electrodes are presented. Measured power densities of selected respectively. When operating at the same anode vapor pressure of 76.8 MIECEs are given. The physical morphology, and composition of our kPa, the BASE temperature in the K-AMTEC is only 1002 K, versus MIECEs are also described. Finally, a possible operational mechanism 1123 K for the Na-AMTEC, generating only 2.0 kWe at a specific for these electrodes is proposed. electrical power of 45 We/kg. In addition to the lower specific power, the specific radiator area for the K-AMTEC is significantly larger than for the Na-AMTEC because of the lower conversion efficiency (~ 22.5%). The Na-AMTEC operates at higher conversion efficiency and higher electrical power than the K-AMTEC for condenser temperatures > 620 K.

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Design Optimization of High-Power, Liquid Anode Recent Developments in Mini-Electrode Test Cell AMTEC Testing

Jean-Michel Tournier and Mohamed S. El-Genk J.R. Rasmussen and T.K. Hunt

Institute for Space and Nuclear Power Studies and Department of Advanced Modular Power Systems, Inc., 4370 Varsity Drive Chemical and Nuclear Engineering Ann Arbor, MI 48108 The University of New Mexico, Albuquerque, NM, 87131 (734 677-4260), [email protected] (505) 277 – 5442, Fax: (505) 277 – 2814, [email protected] Abstract. The mini-electrode test cell (METC) was originally Abstract. A high power, liquid-anode, Alkali Metal Thermal-To- developed as a vehicle for quickly and easily evaluating alkali metal Electric Conversion (AMTEC) unit design is developed and optimized. thermal-to-electric converter AMTEC cell components (electrodes, Optimized parameters include: number of BASE elements; spacing current collectors, electrolytes, etc.) in a prototypical AMTEC between BASE elements and their arrangement in rows and columns; environment. Since its introduction in 1996, the METC has undergone electrode’s surface area per BASE element; and number and diameter two major design changes to increase its reliability and the size of cells of the orifices in the internal radiation heat shield. The effects of these it can accommodate. In addition, the methodologies for testing and parameters on the size, mass and performance of the AMTEC unit analyzing data have evolved, resulting in the ability to extract cell and when operating at BASE and condenser temperatures of 1123 K and electrode parameters B (termed the temperature independent current 670 K, respectively, are discussed. The specific power of the Na density), R0 (the series cell resistance), G (termed the electrode mass- AMTEC unit is maximum at a WRh1.5 electrode’s surface area of transport parameter), and a (the electrochemical transfer coefficient). 140.8 cm2 per BASE element. A total number of 128 BASE elements, Finally, tests have been developed to measure: the electronic electrically connected in series, deliver the nominal power of 5.6 kWe conductivity of beta”-alumina; the isolation resistance between the cell which is 84% of the peak power (6.7 kWe). The BASE elements are electrodes and the cell case; and permeation of sodium through arranged in 4 rows of 32 elements each, resulting in a nearly square mechanical breaches in the high/low sodium pressure boundary. footprint, minimum mass and parasitic heat losses, and maximum specific power and efficiency. At the nominal operating current of ~100 A, the radiation and conduction parasitic heat losses amount to only 7% of the total heat input. The optimized AMTEC unit design measures 594 mm x 410 mm x 115 mm, weighs 44.3 kg, and has an efficiency of 27.6% at the nominal power of 5.6 kWe (a specific power of 125 We/kg). For a 100-kWe power system with a nuclear reactor core exit temperature of ~1180 K, 18 of these units would provide excellent redundancy and reliability in energy conversion.

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Technology Development Program for an Advanced Technology Concept for a Near-Term Closed Brayton Potassium Rankine Power Conversion System Cycle Power Conversion Unit Compatible with Several Space Reactor Designs John Foti, Dave Halsey, Tim Bauch and Glen Smith Bruce B. Bevard1, Graydon L. Yoder2 Hamilton Sundstrand, A United Technologies Company, 1Oak Ridge National Laboratory, P.O. Box 2008, Oak Ridge TN. 4747 Harrison Avenue, Rockford, IL 61125-7002 37830 815-394-2153; [email protected] 1865-574-0279, [email protected], 2865-574-5282, [email protected] Abstract. There is a need in the space science community for nuclear- powered electric propulsion systems to enable high-value, deep space Abstract. The major goal of space reactor power system designers is and planetary exploration. Certain missions are driven by once-in-a- to increase the specific power (kWe/kg) of the overall reactor power lifetime or highly infrequent occurrences that require the near-term system. During the early days of the U.S. space power program, development of a flight-capable nuclear space power and electric Rankine cycle power conversion technology was vigorously pursued as propulsion system in order to take advantage of the scientific an approach for achieving extremely favorable specific powers – opportunity. The broader applicability of Brayton power systems to the particularly for system power levels on the order of 100 kWe and commercial and military aircraft markets has provided fertile ground larger. As a result, liquid-metal Rankine cycle power conversion for the continued development and implementation of new technologies technology is relatively mature (compared to other dynamic power applicable to a closed Brayton cycle space Power Conversion Unit conversion technologies), with in-space flight demonstration being the (PCU). One concept for effectively achieving a near-term Brayton principal remaining impediment to its near-term application. The space power capability is based on the development work associated unique technology issue associated with a space-based Rankine system with the Integrated Power Unit (IPU). This unit embodies the state of is the control and management of two-phase liquids. While many the art in turbomachinery, generators, bearing systems and electric aspects of two-phase management have been demonstrated for short power management and distribution capability that can readily be periods of time in zero-gravity with water and organic working fluids, evolved into a closed Brayton cycle PCU. This paper provides an the demonstration of liquid-metal two-phase systems in space remains overview of aircraft-based Brayton power system technologies, their the fundamental obstacle to its use. ORNL was recently awarded a implementation into the IPU and one approach for leveraging this contract with NASA to design a small-scale flight demonstration capability into a near-term closed Brayton cycle space power program to demonstrate Rankine cycle two-phase fluid management in conversion unit. space. An initial goal of this work is to demonstrate two-phase interface control by means of an early program space flight demonstration with a simplified system employing surrogate materials (such as Plexiglas and Freon).

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A Closed Brayton Power Conversion Unit Concept for Energy Conversion Options for Advanced Radioisotope Nuclear Electric Propulsion for Deep Space Missions Power Systems

Claude Russell Joyner II Mohamed S. El-Genk

Discipline Chief, Mission Analysis & Vehicle/Propulsion Integration Institute for Space and Nuclear Power Studies and Chemical and Pratt Whitney, Ms 712-67, PO 109600, West Palm Beach, Florida Nuclear Engineering Department 33410-9600 The University of New Mexico, Albuquerque, NM, 8713 561-796-3159, [email protected] (505) 277 – 5442, [email protected]

Abstract. Nuclear electric propulsion has the potential to provide Abstract. Static and dynamic energy conversion technologies for increased payload fractions and reduced mission times to the outer Advanced Radioisotope Power Systems (ARPSs) are reviewed and planets. One of the critical engineering and design aspects of nuclear their impact on the system’s total mass and specific electrical power electric propulsion at the power levels (> 25 kWe) is the mechanism and the amount of 238PuO2 fuel needed for the heat source are that is used to convert the thermal energy of the reactor to electrical assessed and compared. Conversion technologies considered are power. The use of closed Brayton cycles has been studied over the Segmented and cascaded Thermoelectric, Alkali-Metal Thermal-to- years and shown to be the optimum approach for power requirements Electric Conversion, and Free Piston Stirling Engines (FPSEs) and, for over 20 kWe and is scalable to higher power levels from 100 kWe to comparison, SiGe thermoelectric. Estimates for a 100 We ARPS multi-megawatt levels. The closed Brayton cycle engine power indicate that when using SiGe thermoelectric, operating between 1273 conversion system is the most flexible for a wide range of power K and 573 K, 8 General Purpose Heat Source (GPHS) modules would conversion needs. The long operating life requirements for nuclear be required and the system’s specific power is ~ 4.6 We/kg. Using electric propulsion demands high reliability from the dynamic aspects STE converters, operating between 973 K and 373 K, 5 GPHS modules of the CBC Engine. Pratt & Whitney’s history of designing long-life are required and the ARPS’s specific power is ~ 7.28 We/kg. The next turbo-machines and the use of integrated control/health management generation STE converters that could operate between 1273 K and 573 systems will serve to over-come many of these related design issues. K, for a projected system efficiency of 13.8%, decrease the number of GPHS modules needed to 4 and increase the system’s specific power to This paper will discuss the use of a closed Brayton cycle engine as the ~ 9.9 We/kg. With cascaded SiGe-STE converters, operating between power conversion unit on a gas-cooled nuclear reactor and the design 1273 K and 373 K, the system’s efficiency could be as much as 16%, trends relative to its use for powering electric thrusters in the 25 kWe to requiring only 3 GPHS modules, for an estimated specific power of 100kWe power level. The general design attributes, scalability to a 10.7 We/kg. This specific power is more than twice that for SOA range of power requirements, and integration issues for using the CBC RTG. With the current version 1.0 of FPSEs, the 100 We ARPS needs Engine and other PCU elements on a Deep Space Mission spacecraft only two GPHS modules, but its specific power (4.1 we/kg) is slightly will be discussed. Some mission trade study trends will also be lower than that of SOA RTG (4.6 We/kg). Future introduction of discussed. versions 1.1 and 2.0 engines, with slightly higher conversion efficiency and significantly lower mass, could increase the system’s specific power to ~ 7.5 We/kg, using the same number of GPHS modules as version 1.0 engines. With Na-AMTEC and K-AMTEC, the 100 We ARPS needs 3 and 4 GPHS modules, respectively, for an estimated specific power of 5.3 and 5.8 We/kg, respectively.

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The Art of Dynamic System Testing - As Applied to a Continuing Development for Free-Piston Stirling, Stirling Convertor Radioisotope Power Systems

Songgang Qiu Allen A. Peterson, Songgang Qiu, Darin L. Redinger, John E. Augenblick, and Stephen L. Petersen Stirling Technology Company, 4208 W. Clearwater Ave. Kennewick, WA 99336-2626 Stirling Technology Company, 4208 W. Clearwater Ave. (509) 735-4700 x 101, [email protected] Kennewick, WA 99336-2626 (509) 735-4700 x 113, [email protected] Abstract. The Department of Energy (DOE) has selected Free-Piston Stirling Convertors for future, advanced Radioisotope Space Power Abstract. The Department of Energy (DOE) has selected Free-Piston Systems. Stirling Radioisotope Generators (SRGs) will be employed for Stirling Convertors for future, advanced Radioisotope Space Power deep space and planetary-lander missions, where high efficiency and Systems. Stirling Radioisotope Generators (SRGs) will be employed for decreased isotope-usage make them more attractive than current deep space and planetary-lander missions, where high efficiency and Radioisotope Thermoelectric Generators (RTGs). Stirling Technology decreased isotope-usage make them more attractive than current Company (STC) has been working with the DOE and NASA Glenn Radioisotope Thermoelectric Generators (RTGs). Stirling Technology Research Center (GRC) to evolve the DOE-funded, STC Technology Company (STC) has been working with the DOE and NASA Glenn Demonstration Convertor (TDC) into Flight-Prototype (FP) status. Research Center (GRC) to evolve the DOE-funded, STC Technology DOE recently announced the system integration contractor for Phases II Demonstration Convertor (TDC) into Flight-Prototype (FP) status. and III of the SRG contract and STC has been working with STC has also been working with Westinghouse to develop and Westinghouse to develop and implement a Quality Assurance (QA) implement a Quality Assurance System to control and document system. As part of the QA system, the latest STC Flight Prototype fabrication and testing of FP units. This paper describes recent machines undergo performance mapping and verification testing. progress and status of the Flight Prototype, presents an overview of NASA GRC is performing independent verification testing on STC recent performance test data, and discusses system integration efforts Technology Demonstration Convertors. For a dynamic system, the guided by the newly announced System Integration Contractor for quality of the test data is affected by many factors, such as the setup of Phases II and III of the SRG program. the facility, the calibration of the instrumentation and the data acquisition system, data sampling rate and the length of sampling time, and data processing technique. This paper presents a detailed description of the test setup and data acquisition system used for the Flight Prototype Stirling convertor testing. The advantage and the disadvantage of the test system are discussed. The planned improvement of the test system is also discussed. The data processing technique for the dynamic system of the Flight Prototype Stirling convertor is addressed. The uncertainty analysis of all measured parameters is presented.

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Stirling Convertor Performance Mapping Test Results for Future Radioisotope Power Systems Nuclear Electric Propulsion Design Factors For Deep Space Robotic Missions Songgang Qiu, Allen A. Peterson, Franklyn Faultersack, Darin L. Redinger, and John E. Augenblick Joe Bonometti1, Eric Stewart1, Jeff Dilg1, Larry Kos1, Lee Mason2 and Stirling Technology Company, 4208 W. Clearwater Ave. Gary Langford1

Kennewick, WA 99336-2626 1 (509) 735-4700 x 101, [email protected] Marshall Space Flight Center, National Aeronautics and Space Administration, Huntsville, AL 35812 2 Abstract. The Department of Energy (DOE) has selected Free-Piston Glenn Research Center, National Aeronautics and Space Stirling Convertors for future, advanced radioisotope space power Administration, Cleveland, OH systems. Stirling Radioisotope Generators (SRGs) will be employed 256-544-4019, [email protected] for deep space and planetary-lander missions, where high efficiency and decreased isotope-usage make them more attractive than current Abstract. In the conceptual design of Nuclear Electric Propulsion Radioisotope Thermoelectric Generators (RTGs). DOE recently (NEP) systems, the integration of the entire power and propulsion announced the system integration contractor for Phases II and III of the system with the rest of the spacecraft is essential to achieving mission SRG contract and STC has been working with Westinghouse to success. The focus of conceptual studies is to realistically determine develop and implement a Quality Assurance (QA) system. As part of the spacecraft requirements, characteristics, and trade space. The paper the QA system, the latest STC Flight Prototype machines undergo describes the initial concept definition required by the mission systems performance mapping and verification testing. This paper describes the analysis, and then discusses more detailed design and integration issues most current performance mapping results for units fabricated to date, found during further examination of the initial concept. with comparisons to Stirling thermodynamic simulation models. Mapping tests performed and presented to aid in system integration efforts include effects that result from variations in: internal charge pressure, cold end temperature, hot end temperature, alternator temperature, input power, and variation of control voltage.

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The Impact of Mission Performance Requirements on Hardware Based Technology Assessment in Support of the Development of an Early-Flight Space Fission Near-Term Space Fission Missions Reactor Mike Houts, Melissa Van Dyke, Tom Godfroy, James Martin, David I. Poston Shannon Bragg-Sitton, Ricky Dickens, Pat Salvail, Eric Williams, Roger Harper, Ivana Hrbud, Robert Carter Nuclear Systems Design Group, Los Alamos National Laboratory Los Alamos, New Mexico, 87545 NASA MSFC, TD40, Marshall Space Flight Center, Alabama, 35812 505-667-4336; [email protected] (256)544-7143, Fax: (256)544-5926; [email protected]

Abstract. One of the lessons learned from past space fission reactor programs Abstract. Fission technology can enable rapid, affordable access to is that a small first step should be taken to successfully deploy a space fission any point in the solar system. If fission propulsion systems are to be system. The “size” of the first step is primarily determined by the mission developed to their full potential; however, near-term customers must be performance requirements, most importantly power, lifetime, and mass. This identified and initial fission systems successfully developed, launched, paper discusses the benefits of keeping these requirements as simple as and utilized. Successful utilization will most likely occur if frequent, possible. Several issues are presented that demonstrate how lower-power, significant hardware-based milestones can be achieved throughout the shorter-lifetime, and higher-mass requirements can significantly decrease program. Achieving these milestones will depend on the capability to development risk for a fission system. perform highly realistic non-nuclear testing of nuclear systems. This paper discusses ongoing and potential research that could help achieve these milestones.

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A 100-kWt NaK-Cooled Space Reactor Concept for an MPD Thruster Performance Analytic Models Early-Flight Mission James Gilland1, Geoffrey Johnston2

David I. Poston 1OAI/NASA Glenn Research Center 2Univ. of Notre Dame/NASA Glenn Research Center Nuclear Systems Design Group, Los Alamos National Laboratory NASA Glenn Research Center, Cleveland, OH 44142, USA Los Alamos, New Mexico, 87545 (440) 962-3142, [email protected] 505-667-4336; [email protected]

Abstract. A stainless-steel (SS) sodium-potassium (NaK) cooled reactor could Abstract. Magnetoplasmadynamic (MPD) thrusters are capable of potentially be the first step in utilizing fission technology in space. The sum of accelerating quasi-neutral plasmas to high exhaust velocities using all system-level experience for liquid-metal-cooled space reactors has been Megawatts (MW) of electric power. These characteristics make such with NaK, including the SNAP-10a, the only reactor ever launched by the US. devices worthy of consideration for demanding, far-term missions such This paper describes a 100-kWt NaK reactor, the NaK-100, which is designed as the human exploration of Mars or beyond. Assessment of MPD to be developed with minimal technical risk. In additional to NaK technology thrusters at the system and mission level is often difficult due to their heritage, the NaK-100 uses a proven fuel-form (SS/UO2) and is designed for status as ongoing experimental research topics rather than developed simplified system integration and testing. The pins are placed within a solid SS prism, and the NaK flows in an annulus between the pins and the prism. The thrusters. However, in order to assess MPD thrusters’ utility in later nuclear and thermal-hydraulic performance of the NaK-100 is presented, as missions, some adequate characterization of performance, or more well as the major differences between the NaK-100 and SNAP-10a. exactly, projected performance, and system level definition are required for use in analyses. The most recent physical models of self-field MPD thrusters have been examined, assessed, and reconfigured for use by systems and mission analysts. The physical models allow for rational projections of thruster performance based on physical parameters that can be measured in the laboratory. The models and their implications for the design of future MPD thrusters are presented.

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Nuclear Electric Systems Analyses and Technology Recent VASIMR Accomplishments Development Options for Outer Planet Missions Franklin R. Chang-Díaz,1 Jared P. Squire,1 Timothy Glover,1 Andrew 1 1 1 2 Benjamin Donahue1, Michael Cupples2, Shaun Green2 Petro, Verlin Jacobson, Andrew Ilin, Roger Bengtson, Boris Breizman,2 Wallace Baity,3 Richard Goulding,3 Mark Carter,3 Oleg 1Boeing Phantom Works, MC JW-63, 499 Boeing Blvd., Huntsville, AL Batischev,4 Roderick Boswell.5 35824 1 2 SAIC, 6725 Odyssey Dr., Huntsville, AL 35806 Advanced Space Propulsion Laboratory, NASA, Lyndon B. Johnson [email protected] Space Center, Code CB-ASPL, Houston, TX, 77058 2 Dept. of Physics, The University of Texas at Austin, Austin, TX. 78713 3 Abstract. Nuclear Electric Propulsion (NEP) uranium-fuel nuclear The Oak Ridge National Laboratory, Oak Ridge, TN. 37831 4 fission reactors with advanced power generation, power processing and Massachusetts Institute of Technology, Cambridge, MA. 02139 electric propulsion systems would enable scientifically rich robotic 5 Australian National University, Canberra, Australia exploration missions. Benefits would include: 281-792-5536, [email protected] · Visits to multiple destinations Abstract. The Variable Specific Impulse Magnetoplasma Rocket · Provision for high power for science at destinations (VASIMR) is being developed to address requirements for fast, high- · Maintain position and operations over long durations power interplanetary space transportation. Its electrodeless architecture · Reductions in launch window constraints relies on radio frequency (RF) waves to create and accelerate a plasma · Provisions for high power, high-rate broadband communications in a magnetic nozzle. The physics and engineering of this concept · Reductions is spacecraft weight continue to be developed, along parallel paths, by a NASA-led research · Reach more distant objects with orbiters team, involving industry, academia and government facilities. Major · Faster trips physics accomplishments in 2002 include the characterization of the · Two way sample return missions first stage helicon source and the experimental demonstration of complete propellant burnup in both deuterium and helium. Major The objective of the paper will be to identify and present technology accomplishments include the demonstration of a high sound technical options for the formulation of exciting solar system temperature superconducting magnet prototype and development of a exploration missions. Missions investigated include Titan and Neptune point design for a 1MW engine module. This paper reports on these destinations for a variety of trip times and science payload allotments. results and discusses other scientific issues such as plasma detachment The work to be presented has been done under the NASA and the use of multiple propellants to efficiently cover a wide specific MSFC/Boeing/SAIC In-Space Technology Assessment (ISTA) impulse range. Contract. Mr. Donahue is a senior member of the Boeing Phantom Works Advanced Programs Division and has been active in evaluating advanced propulsion systems for space transfer for the last 15 years with Boeing. The ISTA program analyses presented in this paper include established strategies for creating system wide optimizations and trade studies which are intended to facilitate good technology investments for the Nuclear System Initiative (NSI). These assessments are intended to provide insight into return on investment analysis, and aid in setting forth technology roadmaps and maturation activities to enable exploration mission with 10 fold science return as compared to present chemical propulsion exploration vehicles.

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MW-Class Thruster Experiments at NASA GRC Advanced 35 W Free-Piston Stirling Engine for Space Power Applications Michael R. LaPointe J. Gary Wood and Neill Lane OAI, 22800 Cedar Point Road, Cleveland, OH 44142 (216) 433-6192, [email protected] Sunpower, Inc. 182Mill Street Athens Ohio 45701 740-594-2221 ext. 509 [email protected] Abstract. As the lead NASA center for electric propulsion, the Glenn Research Center is developing MW-class electric thrusters to meet a variety of future mission applications. This paper Abstract. This paper presents the projected performance and overall discusses the pulsed, high power MPD thruster test facility that design characteristics of a high efficiency, low mass 35 W free-piston has been established at GRC, and presents preliminary voltage- Stirling engine design. Overall (engine plus linear alternator) current measurements obtained with a MW-class baseline MPD thermodynamic performance greater than 50% of Carnot, with a thruster. Fabrication and testing of a flexure based thrust stand are specific power close to 100 W/kg appears to be a reasonable goal at this being completed, and the stand is expected to be in service in the small power level. Supporting test data and analysis results from near future. The combined measurements of thrust, voltage, and exiting engines is presented. Design implications of high specific current will provide sufficient information to determine thruster power in relatively low power engines is presented and discussed. efficiency, and the near-term goal of the GRC test program is to design and demonstrate gas-fed MPD thrusters with efficiencies in excess of 50%. The close collaboration between modeling and experiment is anticipated to lead to the development of more efficient MW-class MPD thrusters capable of meeting the diverse and demanding in-space propulsion requirements envisioned by NASA mission planners.

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Developments in Turbo-Brayton Power Converters Thermoacoustic Space Power Converter

Mark V. Zagarola, Christopher J. Crowley and Walter L. Swift Emanuel Tward 1, Michael Petach1, and Scott Backhaus 2

Creare Incorporated, P.O. Box 71, Etna Road, Hanover, NH 03755 1Space and Technology Division, TRW, One Space Park, Redondo 603-643-3800, [email protected] Beach, CA 90278 2Condensed Matter and Thermal Physics Group, Los Alamos National Abstract. Design studies show that a Brayton cycle power unit is an Laboratory, Los Alamos, NM 87545 extremely attractive option for thermal-to-electric power conversion on 1310-812-0389; [email protected] long-duration, space missions. At low power levels (50 to 100 We), a Brayton system should achieve a conversion efficiency between 20% Abstract. A thermoacoustic power converter for use in space in the and 40% depending on the radiative heat sink temperature. The conversion of radioisotope-generated heat to electricity is under expected mass of the converter for these power levels is about 3 kg. development. The converter incorporates a thermoacoustic driver that The mass of the complete system consisting of the converter, the converts heat to acoustic power without any moving parts. The acoustic electronics, a radiator, and a single general purpose heat source should power is used to drive a pair of flexure bearing supported pistons be about 6 kg. The system is modular and the technology is readily connected to voice coils in a vibrationally balanced pair of moving coil scalable to higher power levels (to greater than 10 kWe) where alternators. Initial tests of the small ~100W thermoacoustic driver have conversion efficiencies of between 28% and 45% are expected, the demonstrated good efficiency. An alternator matched to the driver is exact value depending on sink temperature and power level. During a now under construction. A description of the system and the results of recently completed project, key physical features of the converter were development tests are presented. determined, and key operating characteristics were demonstrated for a system of this size. The key technologies in these converters are derived from those which have been developed and successfully implemented in miniature turbo-Brayton cryogenic refrigerators for space applications. These refrigerators and their components have been demonstrated to meet rigorous requirements for vibration emittance and susceptibility, acoustic susceptibility, electromagnetic interference and susceptibility, environmental cycling, and endurance. Our progress in extending the underlying turbo-Brayton cryocooler technologies to thermal-to-electric power converters is the subject of this paper.

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Reliability Assessment Approach for Stirling Mission Concept for a Nuclear Reactor-Powered Mars Convertors and Generators Cryobot Lander

Ashwin R. Shah1, Jeffrey G. Schreiber2, Edward Zampino2, and John O. Elliott1, Ronald J. Lipinski2, David I. Poston3 Timothy Best2 1Mission and Systems Architecture Section, Jet Propulsion Laboratory, 1Sest, Inc., 18000 Jefferson Park, Suite 104, Middleburg Heights, OH Pasadena, CA 91109 44130 2 Advanced Nuclear Concepts, Sandia National Laboratories, 2NASA Glenn Research Center at Lewis Field, 21000 Brookpark Road, Albuquerque, NM 87185 Cleveland, OH 44135 3 Nuclear Systems Design, Los Alamos National Laboratory, Los (440) 234-9173; [email protected] Alamos, NM 87545 (818) 393-5992; [email protected] Abstract. Stirling power conversion is being considered for use in a Radioisotope Power System (RPS) for deep space science missions Abstract. Recently, a team from JPL and the DOE carried out a study because it offers a multifold increase in the conversion efficiency of to investigate the utility of a 3 kWe surface fission power system for heat to electric power. Quantifying the reliability of an RPS that Mars landed missions. In the course of the study it became clear that utilizes Stirling power conversion technology is important to develop the application of such a power system was enabling to a wide variety and demonstrate the capability for long-term success. A description of of potential missions. Of these, two concepts were developed, one for a the Stirling power convertor is provided, along with a discussion about stationary lander and one for a reactor-powered rover. This paper some of the key components. On-going efforts to understand discusses the design of the lander mission, which was developed component life, design variables at the component and system levels, around the concept of landing a cryobot on the Mars north polar ice and related sources and nature of uncertainties is discussed. The cap. The cryobot is designed to bore through the entire 2-3 km requirement for reliability is discussed, and some of the critical areas of thickness of the ice cap, providing a picture of the Martian climate concern are identified. A section on the objectives of the overall spanning more than a million years of Martian history. The high development of the performance model and computation of reliability sustained power available from the reactor system proves to be an ideal is included to highlight the goals of this effort. Also, a viable physics match for this mission design, enabling a level of science return based reliability plan to model the design level variable uncertainties at unavailable from any alternative power sources. The lander design is the component level and the system level is outlined, and potential based on a minimum extrapolation of technology, drawing heavily on benefits elucidated. The plan encompasses interaction of different the existing concepts in development at JPL for the 2009 Mars Science disciplines, maintaining the physical and probabilistic correlations at all Laboratory (MSL) lander and EDL systems. This paper describes the the levels, and a verification process based on rational short-term tests. unique design challenges encountered in the development of this Additionally, both top-down and bottom-up coherency has been mission architecture and incorporation of the fission power system in maintained in order to follow the physics based design process and the lander, and presents a detailed description of the final design of this mission requirements. The outlined reliability assessment approach trailblazing science mission. provides guidelines to improve the design and identifies governing variables to achieve high reliability in the SRG (Stirling Radioisotope Generator) design.

]

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A Fission Powered Mars Telecommunications Orbiter Reaching the Outer Planets with Nuclear Electric Mission Concept Propulsion: Trades, Sensitivities, and the case for a Neptune System Explorer Erik N. Nilsen Muriel A. Noca, Robert C. Moeller

Jet Propulsion Laboratory, California Institute of Technology, 4800 Jet Propulsion Laboratory, California Institute of Technology, Oak Grove Drive, Pasadena, CA 91109 Pasadena, CA 91109 (818)354-4441, [email protected] 818-393-0950, [email protected]

Abstract. The Mars Program is performing ongoing studies to Abstract. Over the last year, a large effort that involved several NASA investigate concepts for high data rate telecommunications orbiters to agencies and DOE was initiated to evaluate the mission benefits and support future Mars surface and orbital science missions. NASA understand the sensitivities of Nuclear Electric Propulsion (NEP). This recently completed a mission study to examine the use of a fission paper first describes the sensitivities of the mission design parameter reactor powered telecommunications orbiter to enable high priority space (i.e. the trades between propulsion system characteristics, power science at Mars and to validate key technologies in power, propulsion level, system efficiencies, and flight times). It then illustrates the and telecommunications. This paper details the objectives of the findings for a conceptual Neptune System Explorer mission. A point mission, the key constraints that influenced the mission design, and the design for this mission is presented, using a 100-kWe Power and flight and ground system implementation. Issues that influenced the Propulsion Module (designed in parallel by a NASA MSFC lead effort) mission architecture are discussed, as well as the technology along with a representative science payload. This mission features a assumptions. Key mission and technology trades are listed and the Fly-by Nereid, a capture and 10-month stay around Triton, a transfer to decision criteria are developed. This mission concept was developed to an elliptical polar orbit around Neptune and science in this orbit for 12 examine the science benefit of a reactor powered telecommunications months, and finally a transfer into a low equatorial circular orbit to orbiter. No decision on power sources would be made until after study Neptune’s rings. The system features a very high downlink data completion of an Environmental Impact Statement. rate from Neptune (several 10s of Mbps), and a full complement of science instruments. Variations in power levels around the design point are investigated. This analysis shows where the technologies should be headed to fully take advantage of the NEP capabilities.

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NEPTranS; A Shuttle-Tended NEP Interplanetary Technology for Space Reactor Applications Transportation System F.W. Wiffen and S.J. Zinkle John O. Elliott1, Roy Y. Nakagawa1, Thomas R. Spilker1, Ronald J. Lipinski2, David I. Poston3, Dean W. Moreland4 Metals and Ceramics Division, Oak Ridge National Laboratory Oak Ridge, TN 37831-6138 1Mission and Systems Architecture Section, Jet Propulsion Laboratory, 865 481-0822, [email protected] Pasadena, CA 91109 2 Advanced Nuclear Concepts, Sandia National Laboratories, Abstract. The resurgence of interest in space nuclear power requires Albuquerque, NM 87185 the examination of technologies needed for the design, construction, 3 Nuclear Systems Design, Los Alamos National Laboratory, Los and safe operation of compact fission reactors. Nb-1Zr alloy is an Alamos, NM 87545 attractive candidate for this application, especially for reactors cooled 4Payload Safety Group, Johnson Space Center, Houston, TX 77058 by liquid alkali metals, and was the choice material in the SP-100 (818) 393-5992; [email protected] program. This paper will examine the readiness of Nb-1Zr for application, with emphasis on mechanical properties and the effects of Abstract. Recently, a study was performed by a team from JPL and irradiation on the alloy. Similarities and differences in the high the DoE to develop a mission architecture for a reusable NEP temperature tensile and creep behavior of Nb-1Zr and PWC-11 will be Interplanetary Transfer Vehicle, a “Space Truck”. This vehicle is examined. Data generated in the 1960s and in the 1980s will be designed to be used for delivery of payloads from Earth to a variety of compared, to search for clues to the disappointing performance of the destinations, including Mars and Venus, dependent on mission needs. PWC-11 alloy. Radiation effects technology was given limited In addition to delivering payloads to the target bodies, the vehicle is attention in the SP-100 program because of the high-temperature designed to perform autonomous rendezvous and capture of sample operating regime. To fill this gap, a thorough review of the available return capsules at the destination for return to Earth. In order to literature on irradiation effects in niobium alloys will be presented and maximize the utility of the vehicle, its design is optimized for servicing used to suggest the appropriate application range and conditions for between missions with the Space Shuttle. Fuel tanks, ion thrusters, and niobium alloys and to identify limitations that radiation effects may Power Management and Distribution electronics are all on-orbit place on the use of Nb-1Zr. This will identify open questions that must replaceable units, located at the payload interface end of the spacecraft be addressed in future experimentation. to ensure a minimal radiation dose to the Shuttle and crew during maintenance and resupply operations. Operational flexibility is maximized through the use of replaceable fuel tanks and thrusters, allowing tailoring of fuel load to any given destination and payload mass. This paper discusses the preliminary design developed for the NEP Interplanetary Transfer Vehicle, including its configuration and design features, and outlines the concept for mission design, including discussion of unique requirements for launch, deployment and operations with the Space Shuttle, and rendezvous and servicing by the Shuttle in Earth orbit following a return from each target destination.

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Advanced Thermionic Converter Technology Program Effects of Collector Temperature on the Performance of Grooved Electrode Thermionic Converters James R. Luke Yoichi Momozaki and Mohamed S. El-Genk Institute for Engineering Research and Applications New Mexico Institute for Space and Nuclear Power Studies/Department of Chemical Institute of Mining and Technology 901 University Blvd. SE and Nuclear Engineering, The University of New Mexico, Albuquerque, New Mexico 87106 USA Albuquerque, NM 87131, USA 505-272-7275, Fax: 505-272-7297, [email protected] (505) 277-5442, FAX: (505) 277-2814, [email protected]

Abstract. A thermionic energy converter (TEC) is a direct energy Abstract. Experiments are conducted to compare the performance of conversion device, which converts heat to electricity with no the smooth electrodes, grooved emitter, and grooved collector moving parts. Thermionic converters are well suited to space converters at the same emitter and collector temperatures. The focus of nuclear power systems because of their high power density, high the experiments is to investigate the effects of collector temperature on heat rejection temperature, and immunity to radiation. Several the performance of grooved electrode converters. The converters with recent advances in thermionic energy conversion technology have various electrode configurations and a 0.5 mm gap are tested at emitter greatly improved the efficiency of these devices. A research temperatures, T = 1473 - 1673 K, collector temperatures, T = 773 - program was undertaken to independently confirm these advances, E C 1023 K, and cesium pressures, PCs = 10 – 500 Pa. All electrodes are and to extend them to converters with practical geometry. The planar and fabricated with polycrystalline Molybdenum. The grooved recent development of a stable cesium/oxygen vapor source has led electrodes have concentric macro-grooves 0.5 mm wide, 0.5 mm deep, to a significant improvement in performance. The addition of a and 1.0 mm apart. The output electrical power density and the small amount of oxygen to the cesium vapor can increase the conversion efficiency are calculated from the measured I-V curves and emission current by a factor of three or more. The beneficial the maximum power density and conversion efficiency are determined effects of oxygen are stable and reproducible. A TEC with a cold from the calculated P -V and h-V envelopes. At T = 1473 K, the seal has been invented, which greatly simplifies construction, D E electrical power density for the smooth electrodes converter at the operation, and maintenance of the TEC. Electron reflection from optimum collector temperature (873 K) is 2.25 W/cm2 at h = 13.8 %, the collector has been shown to reduce the performance of TEC's. 2 Reflection suppressing materials were produced and tested. One decreasing to 1.46 W/cm at h = 12.4 % as TC increases by only 75 K sample showed evidence of reflection suppression, increasing the to 948 K. However, the corresponding decrease in the electrical power density for the grooved collector converter is only 0.04 W/cm2 (from average output voltage by 0.16 V. Another sample did not. 2 2 Research in this area is ongoing. 1.24 W/cm at h = 11.4 % and TC = 873 K to 1.20 W/cm at h =10.7 % and TC = 948 K) and that for the grooved emitter converter is 0.12 2 2 2 W/cm (from 0.82 W/cm at h = 8.1 % and TC = 873 K to 0.70 W/cm at h = 6.8 % and TC = 948 K). At the higher emitter temperatures of 1573 K and 1673 K, similar results are obtained. In all experimental conditions, however, the grooved emitter converter shows the lowest power density and conversion efficiency, followed by the grooved collector converter, and the smooth electrodes converter, except for TE = 1673 K and TC = 1023 K at which the converter with a grooved collector operates at the highest power density (2.38 W/cm2) and conversion efficiency (14.0 %).

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Application of Electrometer Technology to Materials Experimental Results from a 2 kW Brayton Power Evaluation for Future Planetary Spaceports Conversion Unit

C.I. Calle1, C.R. Buhler2, J.G. Mantovani3, E.E. Groop1, and A.W. David Hervol1, Lee Mason2, Arthur Birchenough3 Nowicki4 1Mechanical Systems Branch, Glenn Engineering and Scientific 1Electromagnetic Physics Laboratory, NASA, Kennedy Space Center, Support Organization , Analex Corporation at NASA Glenn Research FL 32899 Center, MS 301-5, 21000 Brookpark Road,Cleveland, OH 44135 2Swales Aerospace, Merritt Island, FL, USA 2Thermo-Mechanical Systems Branch, Power and On-Board 3Florida Institute of Technology, Melbourne, FL, USA Propulsion Technology Division,NASA Glenn Research Center, 4Dynacs, Inc., Mail Stop DNX-15, Kennedy Space Center, FL 32899 21000 Brookpark Rd, Cleveland, OH 44135 (321) 867-3274; [email protected] 3Electrical Systems Development Branch, Power and On-Board Propulsion Technology Division, NASA Glenn Research Center, Abstract. Future spaceports on dusty and dry planetary environments, 21000 Brookpark Rd, Cleveland, OH 44135 such as the Martian or lunar environments, may be hindered by the 1216-433-9624, 216-433-8311, [email protected] build-up of electrostatic charge that may generate unwanted electrostatic potentials. In an effort to evaluate suitable materials for Abstract. This paper presents experimental test results from operation these environments, an electrometer sensor technology was developed. of a 2 kWe Brayton power conversion unit. The Brayton converter was This technology and its associated environmental simulators have made developed for a solar dynamic power system flight experiment planned possible the characterization of the electrostatic response of possible for the Mir Space Station in 1997. The flight experiment was spaceport materials exposed to planetary regolith simulants. A cancelled, but the converter was tested at Glenn Research Center as “triboelectric series” ranking of these materials according to their part of the Solar Dynamic Ground Test Demonstration system which electrostatic response has been developed. Some material candidates included a solar concentrator, heat receiver, and space radiator. In with low electrostatic response when in contact with the regolith preparation for the current testing, the heat receiver was removed and simulants have been identified. Candidate materials for future planetary replaced with an electrical resistance heater, simulating the thermal spaceports can be characterized with this technology. input of a steady-state nuclear source. The converter was operated over a full range of thermal input power levels and rotor speeds to generate an overall performance map. The converter unit will serve as the centerpiece of a Nuclear Electric Propulsion Testbed at Glenn. Future potential uses for the Testbed include high voltage electrical controller development, integrated electric thruster testing and advanced radiator demonstration testing to help guide high power Brayton technology development for NEP.

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Initial Tests of a Thermoacoustic Space Power Engine NASA GRC Stirling Technology Development Overview Scott Backhaus1 Lanny G. Thieme and Jeffrey G. Schreiber 1Condensed Matter and Thermal Physics Group, Los Alamos National Laboratory, Los Alamos, NM 87545 NASA Glenn Research Center at Lewis Field, MS 301-2 505-667-7545, [email protected] 21000 Brookpark Road, Cleveland, OH 44135 (216) 433-6119 [email protected] Abstract. Future NASA deep-space missions will require radioisotope- powered electric generators that are just as reliable as current RTGs, Abstract. The Department of Energy, Lockheed Martin (LM), Stirling but more efficient and of higher specific power (W/kg). Technology Company, and NASA Glenn Research Center (GRC) are Thermoacoustic engines at the ~1-kW scale have converted high- developing a high-efficiency Stirling Radioisotope Generator (SRG) for temperature heat into acoustic, or PV, power without moving parts at potential NASA Space Science missions. The SRG is being developed 30% efficiency. Consisting of only tubes and a few heat exchangers, for multimission use, including providing spacecraft onboard electric thermoacoustic engines are low mass and promise to be highly reliable. power for NASA deep space missions and power for unmanned Mars Coupling a thermoacoustic engine to a low mass, highly reliable and rovers. efficient linear alternator will create a heat-driven electric generator suitable for deep-space applications. Conversion efficiency data will NASA GRC is conducting an in-house supporting technology project be presented on a demonstration thermoacoustic engine designed for to assist in developing the Stirling convertor for space qualification and the 100-Watt power range. mission implementation. Preparations are underway for a key thermal/vacuum system demonstration and unattended operation during endurance testing of the 55-We Technology Demonstration Convertors (TDC’s). Heater head life assessment efforts continue, including verification of the heater head brazing and heat treatment schedules and evaluation of any potential regenerator oxidation. Long-term magnet aging tests are continuing to characterize any possible aging in the strength or demagnetization resistance of the permanent magnets used in the linear alternator. Testing of the magnet/lamination epoxy bond for performance and lifetime characteristics is now underway. These efforts are expected to provide key inputs as the system integrator, LM, begins system development of the SRG.

GRC is also developing advanced technology for Stirling convertors. Cleveland State University (CSU) is progressing toward a multi- dimensional Stirling computational fluid dynamics code, capable of modeling complete convertors. Validation efforts at both CSU and the University of Minnesota are complementing the code development. New efforts have been started this year on a lightweight convertor, advanced controllers, high-temperature materials, and an end-to-end system dynamics model. Performance and mass improvement goals have been established for second- and third-generation Stirling radioisotope power systems.

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Component-Level Dynamic Modeling and Test Results SP-100 Overview And Lessons Learned – Prime for a Stirling Convertor Contractor Perspective Dr. Sterling Bailey Songgang Qiu and Allen A. Peterson Lockheed Martin Consultant, 16510 Bonnie Lane, Los Gatos, CA Stirling Technology Company, 4208 W. Clearwater Ave. 95032 Kennewick, WA 99336-2626 408-356-5520, [email protected] (509) 735-4700 x 101, [email protected] Abstract. The SP-100 Program was the last major space nuclear Abstract. The Department of Energy (DOE) has selected Free-Piston power technology development program conducted in the United States Stirling Convertors for future, advanced radioisotope space power or elsewhere. The program began with 3 years of technology systems. Stirling Radioisotope Generators (SRGs) will be employed for assessment and competitive design studies involving many contractors, deep space and planetary-lander missions, where high efficiency and DOE, and NASA laboratories, and universities. In 1986 the SP-100 decreased isotope-usage make them more attractive than current prime contract was awarded to General Electric Aerospace, which later Radioisotope Thermoelectric Generators (RTGs). Changing from static became Martin Marietta and then Lockheed Martin. The program was power conversion to a dynamic system is a big step for space-power terminated in 1994 because of budget constraints and lack of a specific applications. The DOE recently announced the system integration mission. While GE was the system contractor and performed the contractor for Phases II and III of the SRG contract. To aid in the majority of the design and development scope, substantial technical system integration effort, Stirling Technology Company (STC) has work was also performed by Westinghouse and several national developed component-level dynamic models for the internal moving laboratories (LANL, ANL, ORNL, WHC, Sandia, JPL, and LRC). In sub-assemblies of its Flight Prototype (FP) Stirling Convertors. These addition, over 400 vendors and suppliers across the US worked with dynamic models have been validated against test data from operating GE to provide special materials and services. JPL and LANL served as prototypes of the FP machines. This paper describes the modeling the government’s program managers in addition to performing approach, discusses the validation-test setup, and compares predictions technical work. to measured results. Conditions under which the model was exercised The technology selected was a fast spectrum, UN pin fuel reactor are also discussed, such as during launch-loading or potential impact cooled by liquid lithium with conductively coupled thermoelectric loading. Though this paper only addresses sub-assembly dynamics, power conversion. Brayton power conversion was also used in several system-level dynamics are addressed in another paper by the same design studies. In addition, Stirling and Rankine power conversion authors. were evaluated for high power missions. Design efforts were concentrated on a 100 kWe design and the ground engineering test. System designs were also developed for 5 kWe up to 30 MWe power levels. The majority of the program focused on technology development through design, analysis, fabrication and testing of hardware to meet the system and subsystem requirements. Based on the successes and shortcomings of the work performed by GE, its subcontractors, and the DOE and NASA labs observations are made from the prime contractor’s perspective that hopefully will enhance the efficiency and success of future space nuclear power programs. Lessons learned from SP-100 will be discussed relating to 1) effective use of a broad spectrum of national resources, 2) recapture of prior technology, 3) impacts of programmatic direction and changes, 4) design and development philosophy and implementation, and 5) mission commitment. 171 172

Fuel Development for SP-100 Component Fabrication Lessons Learned From The SP100 Nuclear Space Power Program James D. Stephen Edwin D. Sayre GE, Retired 17110 Wild Way, Los Gatos, California 95030 Engineering Consultant to Lockheed Martin Company, 218 Brooke 408-395-2313, [email protected] Acres Drive, Los Gatos, CA 95032 408 356 2769, [email protected] Abstract. The fuel for SP-100 was a sealed pin design with uranium nitride pellets in niobium alloy cladding. This design provided the Abstract. The SP100 Nuclear Space Power System generates 100 irradiation stability and compatibility with the liquid lithium coolant kilowatts of electrical power by a high temperature 1400K (2061F) fast required to meet the system specifications for long-term performance at reactor cooled with liquid lithium that flows through a heat exchanger 1350K reactor core outlet temperature. Development of the SP-100 upon which bonded thermoelectric generators produce electricity. The fuel pin design proceeded from the extensive database and analysis optimum refractory metal alloy for the components of this system is methodologies available from the DOE fast reactor fuel program, and niobium/1%zirconium alloy with highly controlled oxygen and carbon was conducted by GE and the national laboratories working together. content and grain size.

UN fuel reduces the reactor size, and therefore the reactor mass, for The most complex and difficult to fabricate components for the original space applications. Taking advantage of the high thermal conductivity SP100 system or any subsequent similar system, e.g., transfer heat from of UN fuel pellets, the SP-100 fuel was designed to operate at low liquid lithium to a gas turbo-generator system are: 1.Rhenium/niobium temperature, where release of fission gases from the pellet is low. The alloy bonded fuel cladding, 2. Thermoelectric, (TE) cells, 3. small void volume required to contain the released fission gas Thermoelectromagnetic, (TEM) Pump, 4. Thin walled, niobium alloy, permitted the design of a short, sealed fuel pin, and hence a small heat exchanger. reactor core. The fuel and cladding manufacturing processes were qualified for The cladding was a Nb-1Zr tube with a Re liner bonded to the inner production. Since there are still no suppliers for rhenium tubing it must surface. The Nb-1Zr provided the compatibility with lithium and be made in house as it was. Tubing made by rolling , welding and sufficient creep strength to contain the internal pressure produced by drawing rhenium sheet meets the high quality demand for bonding to the released fissions gases. The liner was added to provide a barrier to the niobium alloy cladding by Hot Isostatic Pressure (HIP) diffusion minimize nitrogen migration from the fuel and interaction with the bonding. Flat plate thermoelectric cells with compliant pads are cladding. extremely difficult and costly to fabricate. They are also extremely difficult to bond to the walls of the TEM pump and the thin walled heat Over seventy fuel pins were irradiated under the SP-100 program. This exchanger. It is recommended that a tubular thermoelectric system HIP paper presents the lessons learned from the design, analysis, and bonded be evaluated for a future system. The square. sharp cornered, irradiation testing and examination of fuel pins conducted through the niobium alloy, ducts were the most difficult components to be close of SP-100 in 1994, with emphasis on data on UN pellet swelling, successfully manufactured for the TEM pump. A process for bonding fission gas retention and release, nitrogen migration, chemical the thermoelectric cells to the ducts was not developed before the interactions, cladding irradiation-induced swelling, and other results program was cancelled. A process for producing the thin walled which may be applicable to a future space nuclear power program. niobium alloy heat exchanger by HIP bonding was developed.

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Review of SP-100 Program Thermoelectric Converter Hydrodynamic Electrode Model for MPD Thruster Development Subrata Roy J.M. Reyes and L.E. DeFillipo Computational Plasma Dynamics Laboratory, Kettering University, Lockheed Martin Astronautics, 230 Mall Blvd, King of Prussia, PA Flint, MI 48504-4898 19406 810-762-9949;[email protected] 610-354-1794, [email protected] Abstract. Accurate sheath modeling is of considerable interest to the Abstract. Near-term NEP missions envisioned for the exploration of effective design of ionized flow in several magnetohydrodynamic the outer planets have modest power demands of less than 100kWe, (MHD) applications including space propulsion thrusters and high- with longer term sample return missions requiring 200 to 500kWe, and speed air vehicles. In particular, an electrode sheath (fall) voltage future manned missions requiring multi-megawatts. Taking advantage model is necessary to predict the total thruster voltage, which in turn is of the inherent modularity of thermoelectric conversion, design required to predict the total efficiency for on-board propulsion configurations from 10kWe to 300kWe were developed during the thrusters. Understanding plasma wall interaction is crucially important SP100 program. Conductively coupled thermoelectric devices have to improve the high power thruster efficiency. In high-speed air been successfully fabricated, completing up to 23,000 hours of testing vehicles, the interaction of the near-field flow around a supersonic and under typical service conditions. Made with same SiGe alloy as the hypersonic vehicle and an applied magnetic field acting on the ions space-proven RTG unicouples, the results showed that these devices produced at the bow shock wave can actually produce beneficial effects are expected to have the same reliable performance exhibited by the on drag and heat transfer. However, these effects can be further RTG systems with over 27 years of continuous service. This paper controlled by the existence of plasma sheath near the leading surface of reviews the accomplishments of the SP100 program in thermoelectric the air vehicle. Present status of the space propulsion and hypersonic power conversion, the materials and processes of device fabrication, flow research reflects a dearth of consistent numerical models to lessons learned during the program, and options for both SiGe device understand the effect of near wall plasma interaction with a magnetic improvement and advanced thermoelectric converters. field. The anomalies are due to the choice of Bohm’s criterion as the boundary condition for both plasma and sheath using a non-consistent model. There are several theoretical sheath models available in the literature. However, none of these models include the ionization and recombination process in the presence of neutrals that is of practical significance. In this paper, a finite element discretized one-dimensional formulation of plasma–sheath dynamics, using multi-fluid equations for a partially ionized plasma, was documented. Based on the experimental data for multiple ionization of xenon gas, a third order polynomial has been used as a fit to describe ionization processes. Such a polynomial has been used to self-consistently calculate the rate of ionization in the plasma dynamic equations. Simulation results show a dominant role of recombination near the wall in the presence of neutrals. The formulation is applied to a simplistic electrode model. The number densities of electrons, ions and neutrals along with ion and neutral velocities, sheath potential and electron temperature profiles are presented.

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Large Carbon-Carbon Grids for High Power, End-to-End Demonstrator of the Safe Affordable High Specific Impulse Ion Thrusters Fission Engine (SAFE) 30: Power Conversion and Ion Engine Operation Jay Polk1, John Brophy1, Vince Rawlin2, George Williams2 Ivana Hrbud1, Melissa Van Dyke2, Mike Houts2, Keith Goodfellow3 1Jet Propulsion Laboratory, California Institute of Technology, 4800 Oak Grove Drive, Pasadena, CA 91109 1ERC, Inc., NASA MSFC Group, Huntsville, AL 35812 2 NASA Glenn Research Center, 21000 Brookpark Road, 2NASA Marshall Space Flight Center, Propulsion Research Center, Cleveland, OH 44135 Huntsville, AL 35812 (818) 354-9275, [email protected] 3NASA Jet Propulsion Laboratory, Pasadena, CA 91109 (256) 544-2465, [email protected] Abstract. NASA is investigating high power, high specific impulse propulsion technologies that could enable ambitious flights such as Abstract. The Safe Affordable Fission Engine (SAFE) test series multi-body rendezvous missions, outer planet orbiters and interstellar addresses Phase 1 Space Fission Systems issues in particular non- precursor missions. Ion engines can efficiently operate at very high nuclear testing and system integration issues leading to the testing and specific impulse, but accelerating voltages of many kV result in greater non-nuclear demonstration of a 400-kW fully integrated flight unit. ion optics wear rates due to high energy ion bombardment. These The first part of the SAFE 30 test series demonstrated operation of the lifetime issues can be addressed by the use of ion optics grid materials simulated nuclear core and heat pipe system. Experimental data such as carbon with very low sputter yields and optics designs which acquired in a number of different test scenarios will validate existing minimize the flux and energy of ions to the grid surfaces. This paper computational models, demonstrated system flexibility (fast start-ups, reports performance and lifetime predictions based on numerical multiple start-ups/shut downs), simulate predictable failure modes and modeling for carbon-carbon composite grids with an active beam operating environments. The objective of the second part is to diameter of 60 cm operating at up to 30 kWe and 14,000 s with krypton demonstrate an integrated propulsion system consisting of a core, propellant. The numerical results are compared with the measured conversion system and a thruster where the system converts thermal performance of a set of carbon-carbon grids tested on a 75 cm diameter heat into jet power. This end-to-end system demonstration sets a krypton ion engine. precedent for ground testing of nuclear electric propulsion systems. The paper describes the SAFE 30 end-to-end system demonstration and its subsystems.

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Numerical Simulations of the Pulsed Inductive Thruster Overview of NASA Multi-Dimensional Stirling Convertor Code Development and Validation Effort Pavlos G. Mikellides Roy C. Tew1, James E. Cairelli1, Mounir B. Ibrahim2,Terrence W. Arizona State University, College of Engineering and Applied Sciences Simon3, and David Gedeon4 Department of Mechanical and Aerospace Engineering, P.O. Box 876106, Tempe, AZ 85287-6106 1Thermo-Mechanical Branch, NASA Glenn Research Center, (480) 727-6215, [email protected] MS 301-2, 21000 Brookpark Road, Cleveland, OH 44135, USA 2Department of Mechanical Engineering, Cleveland State University, Abstract. Numerical modeling of the Pulsed Inductive Thruster Cleveland, OH 44115, USA exercising the magnetohydrodynamic code, MACH2 aims to provide 3Department of Mechanical Engineering, University of Minnesota, bilateral validation of the thruster’s measured performance and the Minneapolis, MN 55455, USA code’s capability of capturing the pertinent physical processes. 4Gedeon Associates, Athens, OH 45701, USA Computed impulse values demonstrate excellent correlation to the 1216-433-8471, [email protected] experimental data for a range of energy levels and helium propellant- mass values. The effects of the vacuum tank wall and mass-injection Abstract. A NASA grant has been awarded to Cleveland State scheme were investigated to show trivial changes in the overall University (CSU) to develop a multi-dimensional (multi-D) Stirling performance. computer code with the goals of improving loss predictions and identifying component areas for improvements. The University of Minnesota (UMN) and Gedeon Associates are teamed with CSU. Development of test rigs at UMN and CSU and validation of the code against test data are part of the effort. The one-dimensional (1-D) Stirling codes used for design and performance prediction do not rigorously model regions of the working space where abrupt changes in flow area occur (such as manifolds and other transitions between components). Certain hardware experiences have demonstrated large performance gains by varying manifolds and heat exchanger designs to improve flow distributions in the heat exchangers. 1-D codes were not able to predict these performance gains. An accurate multi-D code should improve understanding of the effects of area changes along the main flow axis, sensitivity of performance to slight changes in internal geometry, and, in general, the understanding of various internal thermodynamic losses. The commercial CFD-ACE code has been chosen for development of the multi-D code. This 2-D/3-D code has highly developed pre- and post-processors, and moving boundary capability. Preliminary attempts at validation of CFD-ACE models of MIT gas spring and “two space” test rigs were encouraging. Also, CSU’s simulations of the UMN oscillating-flow rig compare well with flow visualization results from UMN. A complementary Department of Energy (DOE) Regenerator Research effort is aiding in development of regenerator matrix models that will be used in the multi-D Stirling code. This paper reports on the progress and challenges of this multi-D code development effort.

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Development of a Dynamic, End-to-End Free Piston Radiator Concepts for Nuclear Powered Brayton Stirling Convertor Model Conversion Systems

Timothy F. Regan1, Scott S. Gerber2, Mary Ellen Roth3 Devarakonda Angirasa1, Lee S. Mason2, and Richard K. Shaltens3

1Sest, Inc. 18000 Jefferson Park, Suite 104 NASA Glenn Research Center, MS 301-2, 21000 Brookpark Road, Middleburg Heights, Ohio, 44130 Cleveland, OH 44135 2Zin Technologies, Inc. Brookpark Ohio, 44142 1SEST, Inc., (216) 433-3914, [email protected] 3NASA Glenn Research Center, Cleveland Ohio, 4413 2NASA GRC, (216) 977-7106, [email protected] 3NASA GRC, (216) 433-6138, [email protected] Abstract. A dynamic model for a free-piston Stirling convertor is being developed at the NASA Glenn Research Center. The model is an Abstract. This paper reviews the technical background for the end-to-end system model that includes the cycle thermodynamics, the development of Brayton power conversion system for Nuclear Electric dynamics, and electrical aspects of the system. The subsystems of Propulsion (NEP). The Brayton system is outlined in terms of its interest are the heat source, the springs, the moving masses, the linear subsystems and their components. The heat rejection subsystem is alternator, the controller and the end-user load. The envisioned use of examined. Some potential concepts are developed for the heat rejection the model will be in evaluating how changes in a subsystem could subsystems. The requisite analyses for the design of radiators were affect the operation of the convertor. The model under development developed. Theoretical and numerical aspects are discussed. will speed the evaluation of improvements to a subsystem and aid in Technological issues with reference to the radiators are highlighted. determining areas in which most significant improvements may be found. One of the first uses of the end-to-end model will be in the development of controller architectures. Another related area is in evaluating changes to details in the linear alternator.

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Concepts for a Capillary-Pumped Heat Engine Radioisotope Power Systems for NASA Space Science Missions R. B. Williams Therese M. Griebel1a, Raynor Taylor1b, Bernard Edwards2, Los Alamos National Laboratory, P.O. Box 1663, MS K575, Los Steve Oleson3 Alamos, NM 87545 505-665-1593, [email protected] 1a,b Office of Space Science, NASA Headquarters Washington DC, 20546, USA Abstract. A scoping thermal analysis was done to generally consider 2 Microwave and Communication Systems Branch, NASA Goddard the thermal-hydraulic feasibility of capillary pumped heat engines. The Space Flight Center, Greenbelt, Maryland 20771, USA analysis was motivated by recent advances in nano-scale materials 3Power and Propulsion Office, NASA Glenn Research Center, science that have made it increasingly practical to manufacture high Cleveland, Ohio 44135, USA porosity wicks with a median pore diameter on the order of a few nanometers. Capillary-pumped heat engines (CPHE) are proposed to operate constructively like a Rankine cycle and to be generally feasible Abstract. NASA is investing in the development of Radioisotope for wick evaporation rates of about 1 watt per square centimeter, Power Systems (RPS) as a part of the Nuclear Systems Initiative to assuming wick material thermal conductivity coefficient of a few W/m- reestablish the capability to produce radioisotope power systems for K. An architecture for a radioisotope driven CPHE is discussed. future solar system exploration missions and to increase the efficiency and range of application for these systems. RPS can provide continuous power for more than twenty years, and can be used in regions of space where solar power is not feasible. Coupled with other new and exciting technological advances, RPS will enable new classes of space science missions that have never been considered before. Some examples described herein include Radioisotope Electric Propulsion, sensor webs and enhanced communication capability for increased scientific data return.

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Coated Particles Fuel Compact-General Purpose Heat Pin-Type Gas Cooled Reactor for Nuclear Electric Source for Advanced Radioisotope Power Systems Propulsion

Mohamed S. El-Genk and Jean-Michel Tournier Steven A. Wright and Ronald J. Lipinski

Institute for Space and Nuclear Power Studies and Department of Sandia National Laboratories, Lockheed Martin Corp. US Department Chemical and Nuclear Engineering of Energy, PO Box 5800, MS 1146 Albuquerque, New Mexico 87185 The University of New Mexico, Albuquerque, NM, 87131 505-845-3014, [email protected]; 505-845-7311, (505) 277-5442, Fax: (505) 277-2814, [email protected] [email protected]

Abstract. Coated Particles Fuel Compacts (CPFC) have recently been Abstract. This paper describes a point design for a pin-type Gas- shown to offer performance advantage for use in Radioisotope Heater Cooled Reactor concept that uses the fuel pin developed for SP-100. Units (RHUs) and design flexibility for integrating at high thermal The Gas-Cooled Reactor is designed to operate at 100 kWe for 7 years efficiency with Stirling Engine converters, currently being considered plus have a reduced power mode of 20% power for a duration of 5 for 100 We, Advanced Radioisotope Power Systems (ARPS). The years. The power system uses a gas-cooled, UN-fueled, pin-type particles in the compact consist of 238PuO2 fuel kernels with 5-mm reactor to heat He/Xe gas that flows directly into a recuperated Brayton thick PyC inner coating and a strong ZrC outer coating, whose system to produce electricity. Heat is rejected to space via a thermal thickness depends on the maximum fuel temperature during reentry, the radiator that unfolds in space. The reactor contains approximately 154 fuel kernel diameter, and the fraction of helium gas released from the kg of 93.15 % enriched UN in 313 fuel pins. The fuel is clad with kernels and fully contained by the ZrC coating. In addition to rhenium-lined Nb-1Zr. The pressures vessel and ducting are cooled by containing the helium generated by radioactive decay of 238Pu for up the 900 K He/Xe gas inlet flow or by thermal radiation. This permits to 10 years before launch and 10-15 years mission lifetime, the kernels all pressure boundaries to be made of superalloy metals rather than are intentionally sized (> 300 mm in diameter) to prevent any adverse refractory metals, which greatly reduces the cost and development radiological effects on reentry. This paper investigates the advantage schedule required by the project. The reactor contains sufficient of replacing the four iridium-clad 238PuO2 fuel pellets, the two rhenium (a neutron poison) to make the reactor subcritical under water floating graphite membranes, and the two graphite impact shells in immersion accidents without the use of internal shutdown rods. The current State-Of-The-Art (SOA) General Purpose Heat Source (GPHS) mass of the reactor and reflectors is about 750 kg. with CPFC. The total mass, thermal power, and specific power of the CPFC-GPHS are calculated as functions of the helium release fraction from the fuel kernels and maximum fuel temperature during reentry from 1500 K to 2400 K. For the same total mass and volume as SOA GPHS, the generated thermal power by single-size particles CPFC- GPHS is 260 W at Beginning-Of-Mission (BOM), versus 231 W for the GPHS. For an additional 10% increase in total mass, the CPFC- GPHS could generate 340 W BOM; 48% higher than SOA GPHS. The corresponding specific thermal power is 214 W/kg, versus 160 W/kg for SOA GPHS; a 34% increase. Therefore, for the same thermal power, the CPFC-GPHS is lighter than SOA GPHS, while it uses the same amount of 238PuO2 fuel and same aeroshell. For the same helium release fraction and fuel temperature, binary-size particles CPFC-GPHS could provide ~ 5 % higher specific thermal power, but at ~ 12% higher mass than single-size particles CPFC-GPHS.

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Liquid Metal Cooled Reactor for Space Power Design Development Analyses in Support of a Heatpipe- Brayton Cycle Heat Exchanger Abraham Weitzberg Brian Steeve1, Melissa Van Dyke1, Alok Majumdar1, Dalton Nguyen1, 9116 Cranford Drive, Potomac, MD 20854 Melissa Corley3, Ray M. Guffee2, Richard J. Kapernick2, 301-299-3434, [email protected] 1NASA Marshall Space Flight Center, Huntsville, AL 35812 Abstract. The conceptual design is for a liquid metal (LM) cooled 2Los Alamos National Laboratory, Los Alamos, NM 87545 nuclear reactor that would provide heat to a closed Brayton cycle 3Mechanical Engineering, Stanford University, Stanford, CA 94309 (CBC) power conversion subsystem to provide electricity for electric propulsion thrusters and spacecraft power. The baseline power level is 100 kWe to the user. For long term power generation, UN pin fuel with Abstract. One of the power systems under consideration for nuclear Nb1Zr alloy cladding was selected. As part of the SP-100 Program this electric propulsion or as a planetary surface power source is a heatpipe- fuel demonstrated lifetime with greater than six atom percent burnup, at cooled reactor coupled to a Brayton cycle. In this system, power is temperatures in the range of 1400-1500 K. The CBC subsystem was transferred from the heatpipes to the Brayton gas via a heat exchanger selected because of the performance and lifetime database from attached to the heatpipes. This paper discusses the fluid, thermal and commercial and aircraft applications and from prior NASA and DOE structural analyses that were performed in support of the design of the space programs. The high efficiency of the CBC also allows the reactor heat exchanger to be tested in the SAFE-100 experimental program at to operate at relatively low power levels over its 15-year life, Marshall Space Flight Center. A companion paper, “Mechanical minimizing the long-term power density and temperature of the fuel. Design and Fabrication of a SAFE-100 Heat Exchanger for use in The scope of this paper is limited to only the nuclear components that NASA’s Advanced Propulsion Thermal-hydraulic Simulator”, presents provide heated helium-xenon gas to the CBC subsystem. The principal the fabrication issues and prototyping studies that, together with these challenge for the LM reactor concept was to design the reactor core, analyses, led to the development of this heat exchanger. An important shield and primary heat transport subsystems to meet mission consideration throughout the design development of the heat exchanger requirements in a low mass configuration. The LM concept design was its capability to be utilized for higher power and temperature approach was to assemble components from prior programs and, with applications. This paper also discusses this aspect of the design and minimum change, determine if the system met the objective of the presents designs for specific applications that are under consideration. study. All of the components are based on technologies having substantial data bases. Nuclear, thermalhydraulic, stress, and shielding analyses were performed using available computer codes. Neutronics issues included maintaining adequate operating and shutdown reactivities, even under accident conditions. Thermalhydraulic and stress analyses calculated fuel and material temperatures, coolant flows and temperatures, and thermal stresses in the fuel pins, components and structures. Using conservative design assumptions and practices, consistent with the detailed design work performed during the SP-100 Program, the mass of the reactor, shield, primary heat transport, reactor instrument and control, and additional structure totaled approximately 1100 kg.

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Revolutionary Concepts for Human Outer Planet “Bimodal” Nuclear Thermal Rocket (BNTR) Exploration (HOPE) Propulsion for an Artificial Gravity HOPE Mission to Callisto

Patrick A. Troutman1, Kristen Bethke2, Fred Stillwagen1, Darrell L. Stanley K. Borowski1, Melissa L. McGuire1, Lee M. Mason1, Caldwell, Jr3., Ram Manvi4, Chris Strickland5, Shawn A. Krizan3 James H. Gilland2 and Thomas W. Packard3

1NASA Langley Research Center, Hampton, VA 23681 1NASA Glenn Research Center, 21000 Brookpark Road, 2Princeton University, Dept. of Mechanical and Aerospace Cleveland, OH 44135 Engineering, Princeton, NJ 08544 2Ohio Aerospace Institute, 22800 Cedar Point Road, 3Analytical Mechanics Associates, Inc., 17 Research Drive, Hampton, Cleveland, OH 44142 VA 23666 3Analex Corporation, 1100 Apollo Drive, Brookpark, OH 44142 4Jet Propulsion Laboratory, Pasadena, CA 91109 Phone: (216) 977-7091, E-mail: [email protected] 5Swales Aerospace, 1224T1 N. Wright St. Hampton, VA 23681 (757) 864-1954, [email protected] Abstract. This paper summarizes the results of a year long, multi- center NASA study which examined the viability of nuclear fission Abstract. This paper summarizes the content of a NASA-led study propulsion for Human Outer Planet Exploration (HOPE). The HOPE performed to identify revolutionary concepts and supporting mission assumes a crew of six is sent to Callisto, Jupiter’s outermost technologies for Human Outer Planet Exploration (HOPE). Callisto, the large moon, to establish a surface base and propellant production fourth of Jupiter’s Galilean moons, was chosen as the destination for facility. The Asgard asteroid formation, a region potentially rich in the HOPE study. Assumptions for the Callisto mission include a launch water-ice, is selected as the landing site. High thrust BNTR propulsion year of 2045 or later, a spacecraft capable of transporting humans to is used to transport the crew from the Earth-Moon L1 staging node to and from Callisto in less than five years, and a requirement to support Callisto and back to Earth in under 5 years. Cargo and LH2 “return” three humans on the surface for a minimum of 30 days. Analyses propellant for the Crew Transfer Vehicle (CTV) is pre-deployed at performed in support of HOPE include identification of precursor Callisto prior to crew departure using low thrust, high power, nuclear science and technology demonstration missions and development of electric propulsion (NEP) cargo and tanker vehicles powered by vehicle concepts for transporting crew and supplies. A complete hydrogen magnetoplasmadynamic (MPD) thrusters. The CTV is surface architecture was developed to provide the human crew with a powered by three 25 klbf BNTR engines which also produce 50 kWe of power system, a propellant production plant, a surface habitat, and power for crew life support and spacecraft operations. To counter the supporting robotic systems. An operational concept was defined that debilitating physiological effects of long duration space flight (~855 provides a surface layout for these architecture components, a list of days out and ~836 days back) under “0-gE” conditions, the CTV surface tasks, a 30-day timeline, a daily schedule, and a plan for generates a “1-gE” artificial gravity environment via rotation of the communication from the surface. vehicle about its center-of-mass at a rate of ~4 rpm. After ~123 days at Callisto, the “refueled” CTV leaves orbit for the trip home. Direct capsule re-entry of the crew at mission end is assumed. Dynamic Brayton power conversion and high temperature uranium dioxide (UO2) in tungsten metal “cermet” fuel is used in both the BNTR and NEP vehicles to maximize hardware commonality. Technology performance levels, vehicle characteristics and requirements for CTV reusability are presented and discussed.

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Bimodal Nuclear Electric Propulsion for Human High Power MPD Nuclear Electric Propulsion (NEP) Missions to the Solar System for Artificial Gravity HOPE Missions to Callisto

Leonard A. Dudzinski Melissa L. McGuire1, Stanley K. Borowski1, Lee M. Mason1, and James Gilland2 NASA Glenn Research Center Systems Engineering Division 1NASA Glenn Research Center, 21000 Brookpark Rd., Cleveland, OH (202) 358-3553, [email protected] 44135 2NASA Glenn Research Center/Ohio Aerospace Institute, 22800 Cedar Abstract. This paper presents the results of recent analyses of Bimodal Point Road, Cleveland, OH 44142 Nuclear Electric Propulsion (BNEP) technology to support human (216) 977-7128, [email protected] missions to the solar system. Two missions are considered: a human mission to Mars and a human mission to the Jovian moon Callisto. The Abstract. The following paper documents the results of a one-year results show that such missions are possible with smaller electric multi-center NASA study on the prospect of sending humans to propulsion systems than single-mode Nuclear Electric Propulsion, thus Jupiter’s moon, Callisto, using an all Nuclear Electric Propulsion providing an alternate evolutionary path from science class systems to (NEP) space transportation system architecture with human class systems. These analyses have been enabled by recent magnetoplasmadynamic (MPD) thrusters. The fission reactor system analytical tool development, and identify technology benefits that were utilizes high temperature uranium dioxide (UO2) in tungsten (W) metal not previously quantifiable. matrix “cermet” fuel and electricity is generated using advanced dynamic Brayton power conversion technology. The mission timeframe assumes on-going human Moon and Mars missions and existing space infrastructure to support launch of cargo and crewed spacecraft to Jupiter in 2041 and 2045, respectively.

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High Power Nuclear Electric Propulsion (NEP) for The Use of Today’s Space Propulsion Technologies To Cargo and Propellant Transfer Missions in Cislunar Facilitate the Deployment of a Hybrid Nuclear Space Propulsion and Power System

Robert D. Falck and Stanley K. Borowski Claude Russell Joyner II

NASA Glenn Research Center, 21000 Brookpark Road, Cleveland, OH Discipline Chief, Mission Analysis & Vehicle/Propulsion Integration 44135 Pratt Whitney, Ms 712-67, PO 109600, West Palm Beach, Florida (216) 433-2295, [email protected] 33410-9600 561-796-3159, [email protected] Abstract. The performance of Nuclear Electric Propulsion (NEP) in transporting cargo and propellant from Low Earth Orbit (LEO) to the Abstract. Pratt & Whitney recently designed a hybrid nuclear first Earth-Moon Lagrange point (EML1) is examined. The baseline propulsion and power concept called ESCORT to meet NASA mission NEP vehicle utilizes a fission reactor system with Brayton power needs in regards to science exploration missions, future manned conversion for electric power generation to power multiple liquid exploration missions and for other possible in-space architectures. This hydrogen magnetoplasmadynamic (MPD) thrusters. Vehicle hybrid nuclear design with combined power and propulsion was characteristics and performance levels are based on technology defined in collaboration with NASA to meet a wide range of mission availability in a fifteen to twenty year timeframe. Results of numerical requirements. The hybrid ESCORT design concept was originally trajectory analyses are also provided. configured to deliver 25 kWe power for spacecraft systems, but was most recently examined for use as a planetary surface power plant capable of delivering up to 160 kWe electrical power using a common reactor design.

When used as a hybrid propulsion and power system for in-space transportation, the ESCORT design incorporates many proven technologies from the current RL10 liquid rocket engine, aircraft power generation and others technologies currently in development for the next generation of space propulsion systems. This includes the turbopumps that will supply liquid hydrogen during propulsive mode, the closed Brayton engine for generating power, tanks designed for propellant stages in use today, a carbon-carbon large area ratio nozzle, and mission enabling integrated health management and controller system technologies.

This paper/presentation will review the applicable propulsion and power generation technologies and current nuclear fuels design approaches for low cost testing and manufacture that will be integral to successfully deploying a nuclear in-space propulsion or power system design. The technology readiness level of the ESCORT hybrid system relative to these technologies will be presented. Additionally, the scalability of the ESCORT design for delivering power from 25 kWe to 500 kWe for in-space missions will be discussed.

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Deep Space Propulsion Requirements Development “24 Hours to the Moon” Using LOX-Augmented Nuclear Thermal Rocket (LANTR) Propulsion Melvin J Bulman1 and Stanley K. Borowski2 Stanley K. Borowski 1GenCorp Aerojet, D5276 P.O. Box 13222, Sacramento, California 2 NASA Glenn Research Center, NASA Glenn Research Center, NASA Glenn Research Center, 21000 Brookpark Road, Cleveland, OH Cleveland, OH 44135 (916) 355-3451, [email protected] (216) 977-7091, [email protected] Abstract. Deep space propulsion is very demanding. The requirements begin with the launch vehicle interface and ascent into earth orbit. The Abstract. This presentation examines the feasibility and requirements propulsion system must operate in the harsh environment of space for for routine “commuter” flights to the Moon. The availability of nuclear many years with little or no maintenance. It’s primary function is to propulsion and power systems, together with “lunar-derived” oxygen transport, and in some cases, return its payload from distant locations. (LUNOX) produced from iron-rich volcanic glass beads (“orange soil”) A secondary function maybe to provide electrical power for the discovered during the Apollo 17 mission, are key to reducing the size, payload during the trip or at it’s destination. cost and complexity of lunar space transportation systems. The LANTR concept is an enhanced nuclear rocket that can leverage the benefits of With the new interest in nuclear propulsion for space, this paper uses LUNOX in a unique way. LANTR utilizes the divergent nozzle section first principles to derive the propulsion system requirements for as an “afterburner” into which oxygen is injected and supersonically different classes of missions from small science to large crewed combusted with nuclear preheated hydrogen emerging from the Planetary/Lunar missions. Metrics are defined to evaluate different engine’s sonic throat -- essentially “scramjet propulsion in reverse.” active propulsion options. A model of the ideal propulsion system is By varying the oxygen-to-hydrogen mixture ratio, the LANTR engine defined for each mission class. These ideal propulsion systems are the can operate over a wide range of thrust and specific impulse values basis of requirements for an in space transportation system architecture while the reactor core power level remains relatively constant. An that minimizes development costs and accelerates system availability. “evolutionary” mission scenario is outlined that utilizes Shuttle-derived heavy lift vehicles to launch NTR-powered lunar transfer vehicles The development road map demonstrates a cost effective capability (LTVs), surface nuclear power systems and LUNOX production plants. with growth by leveraging on early systems through a Preplanned By increasing the supply of LUNOX to lunar landing vehicles (LLVs) Product Improvement (P3I) process. Recommendations are made for initially, then to the orbiting LTVs, a “reusable” space transportation the development of a sustainable deep space transportation system to architecture with increased payload delivery capability can be realized. support Human exploration and expansion beyond low Earth Orbit. As LUNOX production capacity increases and initial outposts grow to commercially viable settlements, the LANTR concept can enable a rapid “commuter” shuttle capable of 24 hour “one-way” trips to and from the Moon. A mobile nuclear electric propulsion (NEP) tanker / propellant depot is a key infrastructure element that can provide a convenient staging point for lunar orbit operations. A vast deposit of “iron-rich” volcanic glass located at the southeastern edge of the Sea of Serenity could supply sufficient LUNOX to support daily commuter flights to the Moon for the next 9000 years!

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The International Space Station Habitat Lessons Learned in Maintenance of the International Space Station Patricia Mendoza Watson, Mike Engle William W. Robbins, Jr. NASA Johnson Space Center, Houston Texas 77058, Mail Code OM 281-483-7770; [email protected] Space Station Program Office Johnson Space Center, Houston, Texas 77058 Abstract. The International Space Station (ISS) is an engineering (281) 244-7722, [email protected] project unlike any other. The vehicle is inhabited and operational as construction goes on. The habitability resources available to the crew Abstract. The International Space Station (ISS) began development in are the crew sleep quarters, the galley, the waste and hygiene 1984. On-orbit assembly and operations began in 1998, and ISS is now compartment, and exercise equipment. These items are mainly in the a 330,000-pound operational orbiting laboratory. The ISS Program still Russian Service Module and their placement is awkward for the crew has several years of assembly ahead, with fifteen years of operations to to deal with. ISS assembly will continue with the truss build and the follow. However, the experience to date has proven valuable in addition of International Partner Laboratories. Also, Node 2 and 3 will identifying lessons in developing a logistics support infrastructure, and be added. The Node 2 module will provide additional stowage volume maintaining a permanently orbiting facility. Understanding what has and room for more crew sleep quarters. The Node 3 module will been successful in ISS, as well as not so successful, will help new space provide additional Environmental Control and Life Support Capability. exploration programs. ISS lessons will help new programs effectively The purpose of the ISS is to perform research and a major area of embed supportability in design and management, and control life cycle emphasis is the effects of long duration space flight on humans, a result cost through effective programmatic requirements and prudent early of this research they will determine what are the habitability design investments. These lessons can be grouped into three major requirements for long duration space flight. areas. The first is the programmatic lessons in establishing and managing an acquisition logistics office. The second area is design strategies. The third area is lessons in operational maintenance. Human space exploration and colonization of space is dependent on the ability to sustain a long duration space-based vehicle that is funded, designed, built and operated by a consortium of international partners. The lessons that are emerging from the ISS program are of value to the next generation of space vehicle development managers.

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Machining in Microgravity Robots and Humans: Synergy in Planetary Exploration

Graylan Vincent Geoffrey A. Landis

Department of Aeronautics & Astronautics, University of Washington, NASA John H. Glenn Research Center, mailstop 302-1, 21000 Brook Box 352400 Seattle, WA 98195, USA Park Road, Cleveland, OH 44135 206-985-0086, [email protected] (216) 433-2238; [email protected]

Abstract. A CNC mill was flown aboard NASA’s KC-135 Abstract. How will humans and robots cooperate in future planetary “Weightless Wonder” microgravity research aircraft to investigate the exploration? Are humans and robots fundamentally separate modes of effect of gravity on the machining process and to demonstrate the exploration, or can humans and robots work together to synergistically feasibility and functionality of a CNC mill in a weightless environment, explore the solar system? It is proposed that humans and robots can such as aboard the International Space Station. The experiment work together in exploring the planets by use of telerobotic operation to hypothesis was that the surface roughness of milling cuts made in expand the function and usefulness of human explorers, and to extend microgravity would be of higher quality than cuts made in a te range of human exploration to hostile environments. gravitational environment due to increased chip removal. The technical problems associated with microgravity machining (such as the chip removal and collection process), and the engineering solutions to these problems were also evaluated in this experiment.

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Novel Amine-Functional Membrane for Metabolic CO2 Because it is Hard: The Crucible of Space as a Source of Removal from Spacesuit Breathing Loop Innovation

Michael D. White Chung-Yi A. Tsai1, Ipek Guray2, Xia Tang1, Tim Nalette3, Catherine 3 4,5 5 Thibaud-Erkey , C. Jeffrey Brinker , and George Xomerita Blank Rome Comisky & McCauley LLP, 900 17th St., NW, Ste. 1000 Washington, DC 20006 1United Technologies Research Center, East Hartford, CT 06108 (202) 530-7421, [email protected] 2Worcester Polytechnic Institute, Worcester, MA 01609 Abstract. We choose to go to the moon. We choose to go to the moon 3Hamilton Sundstrand Space System International, Windsor Locks, CT in this decade and do the other things, not because they are easy, but 4Sandia National Laboratories and the 5University of New Mexico, because they are hard, because that goal will serve to organize and Albuquerque, NM 87106 measure the best of our energies and skills, because that challenge is one that we are willing to accept, one we are unwilling to postpone, Abstract. Life support systems for space suits or habitats require and one which we intend to win… . (Excerpt from President Kennedy’s effective removal of metabolic CO2 and moisture from the breathing September 12, 1962 Speech at Rice University). loop with no loss of O2. Conventional techniques, using either metal hydroxides or metal oxides, require after-mission regeneration or The payoff from the exploration of space is what is learned in making replacement, thus limiting mission duration. A membrane device offers the journey, not just in reaching the destination. The creation, a novel approach to the problem, and would allow continuous protection and utilization of intellectual property assets from space separation of CO2 from the breathing loop, while simultaneously exploration are addressed. venting the CO2 directly to the vacuum of space. Such a membrane device does not require regeneration and therefore, extended extravehicular activity (EVA) for space exploration can be realized without limitation.

We developed a unique porous membrane material with amine- functional pore surfaces using surfactant templating techniques for CO2 removal from the breathing loop of spacesuits. We use evaporation during coating to induce the formation of surfactant micelles that self- organize into desirable pore structures. The membrane, with secondary amine functional groups inside the membrane pores, showed a dual-gas separation factor at 90oC that was ten times higher than the ideal Knudsen separation factor. A CO2 flux of 4 GPU (1GPU (gas permeation unit) =10-6 cm3 (STP)/cm2-s-cmHg) was observed. When the temperature was reduced to 20oC, the mixed-gas separation factor increased two fold, indicating enhanced CO2 surface diffusion via active amine adsorption sites at lower temperatures. An amine- functional membrane could offer a unique alternative to NASA’s current capability for regeneratively removing CO2 from spacesuit breathing loop.

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A Critique of Theoretical Explanations of Gravity Update on an Electromagnetic Basis for Inertia, Shielding Phenomena Gravitation, the Principle of Equivalence, Spin and Particle Mass Ratios R. Clive Woods Bernard Haisch1, Alfonso Rueda2, L. J. Nickisch3 and Jules Mollere4 Department of Electrical and Computer Engineering, 2128 Coover Hall, Iowa State University, Ames, Iowa 50011–3060 1California Inst. for Physics & Astrophysics, 901 Mariners Island (515) 294 3310, Fax (515) 294 8432, [email protected] Blvd., Ste. 325, San Mateo, CA 94404 2Dept. of Electrical Eng., California State Univ., Long Beach, CA Abstract. Podkletnov & Nieminen (1992) have published experimental 90840 results which they interpreted as evidence of gravity shielding. Their 3Mission Research Corp., Monterey, CA 93940-5776 experiment requires the gravitational field to be measured above a 4Henderson State Univ., Arkadelphia, AR 71999-0001 sample of “high-temperature” superconductor YBCO in the form of a [email protected], 650-593-8581, fax: disk cooled below 70K, magnetically rotated at 5000 rpm, and 650-595-4466, www.calphysics.org simultaneously levitated magnetically using two separate high frequency excitations. Weight changes (in a test mass) of the order of Abstract. A possible connection between the electromagnetic quantum 1% were reported. vacuum and inertia was first published by Haisch, Rueda and Puthoff (1994). If correct, this would imply that mass may be an Their experimental results appear to contradict conventional electromagnetic phenomenon and thus in principle subject to gravitational theory, because they suggest evidence of modifications to modification, with possible technological implications for propulsion. the gravitational field in a laboratory-based system that does not require A multiyear NASA-funded study at the Lockheed Martin Advanced general relativity for accurate analysis. This experiment is therefore Technology Center further developed this concept, resulting in an potentially highly important scientifically because of the enormous independent theoretical validation of the fundamental approach (Rueda technological implications for the design of current transportation and Haisch, 1998ab). Distortion of the quantum vacuum in accelerated vehicles and handling methods for bulk materials if gravity reference frames results in a force that appears to account for inertia. modification (and, in particular, gravitation reduction) were We have now shown that the same effect occurs in a region of curved demonstrated to be feasible. spacetime, thus elucidating the origin of the principle of equivalence (Rueda, Haisch and Tung, 2001). A further connection with general More recent reports by De Aquino (2001, 2002) suggest a basis for relativity has been drawn by Nickisch and Mollere (2002): zero-point understanding the Podkletnov & Nieminen (1992) observations. These fluctuations give rise to spacetime micro-curvature effects yielding a reports will be discussed in the light of fundamental physical laws, and complementary perspective on the origin of inertia. Numerical the theoretical basis required for substantial gravitational shielding will simulations of this effect demonstrate the manner in which a massless be re-examined. fundamental particle, e.g. an electron, acquires inertial properties; this also shows the apparent origin of particle spin along lines originally De Aquino, F., arXiv physics/0112081 (2001). proposed by Schrödinger. Finally, we suggest that the heavier leptons (muon and tau) may be explainable as spatial-harmonic resonances of De Aquino, F., arXiv physics/0201058 (2002). the (fundamental) electron. They would carry the same overall charge, but with the charge now having spatially lobed structure, each lobe of Podkletnov, E., & Nieminen, R., Physica C 203, 441 (1992). which would respond to higher frequency components of the electromagnetic quantum vacuum, thereby increasing the inertia and thus manifesting a heavier mass.

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It’s All Gravity…. Transient Inertial Effects and Stationary Forces

P. A. Murad James F. Woodward

1441 Montague Drive, Vienna, Virginia, USA (703) 759-2028, [email protected] Departments of History and Physics California State University Abstract. Newtonian gravitation adequately predicts planet and Fullerton, California 92834 satellite motion. Gravitational anomalies and the wish to travel at e-mail: [email protected] relativistic speeds, however, imply that gravity should be integrated within a unification framework that may include electricity and Abstract. The theoretical basis for prediction of transient inertial mass magnetism. Thus, new theories are needed that predict currently shifts that follow from the strong form of Mach’s principle (that both accepted phenomenon as well as anomalies to prepare the necessary inertial reaction forces and the origin of inertial mass arise from the groundwork for experimental validation needed for advanced gravitational action of chiefly distant matter in the universe) is technology propulsion schemes and far-term missions. A primary recapitulated. The generation of stationary forces is then addressed. deficiency is that we are obviously limited within the confines of our Experimental work in progress is described. own solar system and a different gravity model may be applicable elsewhere in the cosmos. The model proposed here follows previous ideas proposed by Murad, Dyatlov, and Jefimenko for a universal gravitation model with an intrinsic radial force term coupled with angular momentum. Including angular momentum may explain several spin symmetries seen in some anomalous gyroscopic experiments and throughout the universe regarding planets that orbit around the sun; moons that orbit larger planetary bodies; and the rotation about each planetary axis.

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New Experiments With Spinning Metallic Discs Quantum Vehicle Propulsion

Konstantin Mazuruk and Richard N. Grugel Jerry E. Bayles

Marshall Space Flight Center, MS- SD46, Huntsville, AL 35812 (256) 544-9165 [email protected] Gravitational Research, 2825 Pioneer Road, Medford, OR 97501-9642 TEL. 541-535-9263; [email protected] Abstract. A number of recent advanced theories related to torsion properties of the space-time matrix predict the existence of an Abstract. This paper presents a solution of what gravity is and a interaction between classically spinning objects. Indeed, some method of building a vehicle that uses quantum principles to cause experimental data suggest that spinning magnetic bodies discernibly that vehicle to jump through space much as an electron jumps through interact with Earth's natural fields. If there are interactions between space. While this is not a new idea, the methodology to achieve this rotating bodies then nuclear spins could be used for detection. Thus, has not been forthcoming until now. The formulae presented in assuming a spinning body induces a hypothetical torsion field, a sensor this paper are based on a new science created by this author wherein the based on the giant magnetoresistance effect would detect local changes. gravitational action is defined both in quantum mechanical and Experimentally, spinning a brass wheel shielded from Earth's magnetic electrodynamic terms. field showed no measurable change in signals; with no shielding a Faraday disc phenomenon was observed. Unexpected experimental measurements from the non-axial Faraday disc configuration were recorded and a theoretical model was derived to explain them.

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Very Large Propulsive Effects Predictedfor A 512kV Field-Force Acceleration Using Type II Superconductor Rotator Glen A. Robertson David Maker1 and Glen A. Robertson2 265 Ita Ann Ln., Madison, AL 35757 17807 Hilton Dr., Huntsville, AL, 35802 [email protected] 2Gravi Atomic Research, 265 Ita Ann Ln., Madison, AL 35757 2256-430-6875, [email protected] Abstract. It has been shown (Robertson, 2002) that an electromagnetic force density model can predict the experimental gravity like force data Abstract. An equation was developed from an Ungauged GR (Maker reported by Podkletnov (1992, 1997, & 2001). In this paper, it is 2001) that predicts a negative gravity propulsive force with the pulse suggested that the gravitational field-force generated by the speed coming out of the integral of w times V times dq/dt times sin2q superconductor on a test sample would affect the superconductor in a divided by 1-V/512kV. V is the electric potential, w is the azimuthal similar manner. Given this assumption, the expected maximum angular velocity of the electron cloud, dq/dt the frequency of polar velocity obtainable would be proportional to the applied or induced angle oscillation of the electron cloud. Note that if V=512kV this magnetic field and inversely proportional to the density of the shielded equation is singular implying that large effects are possible near 512kV mass, which includes the superconductor. Whereby very large especially if w is also large and dq/dt is in phase with V. This equation magnetic fields would be required to achieve reasonable velocities for appears to have been verified in several experiments for both above and large space vehicles. below 512kV so there is a high likelihood that these large propulsive effects can be created near 512kV. Podkletnov, E. and R. Niemen, ”A Possibility of Gravitational Force Shielding by Bulk YBa2Cu3O7-X Superconductor,” Physica C, Vol. 203, pp. 441 – 444, (1992).

Podkletnov, E., “Weak Gravitation Shielding Properties Of Composite Bulk Yba2cu3o7-X Superconductor Below 70K Under E.M. Field,” cond-mat/9701074 v3, (1997).

Podkletnov, Evgeny and Giovanni Modanese, “Impulse Gravity Generation Based on Charged YBa2Cu3O7-X Superconductor with Composite Crystal Structure,” LANL physics/0108005, (2001).

Robertson, Glen A., “Force Density Model of Gravity Forces from Type II Superconductors,” Submitted to Physica C, August, (2002).

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High-Speed Computational Applications For Space Reliability Methods for Shield Design Process Radiation Shielding Analysis R.K. Tripathi and J.W. Wilson John E. Nealy1, Brooke M. Anderson2, John W. Wilson3, and Garry D. Qualls3 NASA Langley Research Center, Hampton, VA 23681 757 864 1467; [email protected] 1Old Dominion University, Norfolk, VA 23508, USA 2Swales Aerospace, Hampton, VA 23666, USA Abstract. Providing protection against the hazards of space radiation 3NASA Langley Research Center, Hampton, VA 23681, USA is a major challenge to the exploration and development of space. The 757-864-1422, [email protected] great cost of added radiation shielding is a potential limiting factor in deep space missions. In this enabling technology, we have developed Abstract. Expanding knowledge of the complexities of the space methods for optimized shield design over multi-segmented missions radiation environment and its interactions with matter, coupled with involving multiple work and living areas in the transport and duty greater burdens associated with budgetary and time constraints, have phase of space missions. The total shield mass over all pieces of given impetus to the need for application of more sophisticated equipment and habitats is optimized subject to career dose and dose analyses in more abbreviated time spans. Recent work at NASA-LaRC rate constraints. An important component of this technology is the in this area has resulted in development of high efficiency algorithms estimation of two most commonly identified uncertainties in radiation coupled with high-speed computers and visualization hardware and shield design, the shielding properties of materials used and the software to analyze space radiation effects and shielding methodologies understanding of the biological response of the astronaut to the for advanced missions. Special interfacing with CAD solid models and radiation leaking through the materials into the living space. The largest 3-D immersive visualization equipment plays a major role in this uncertainty, of course, is in the biological response to especially high endeavor. Recent applications have included analyses for EVA in a charge and energy (HZE) ions of the galactic cosmic rays. These CAD-modeled STS space suit, for vector flux exposure in an ISS uncertainties are blended with the optimization design procedure to habitation module, and for preliminary exposure predictions within a formulate reliability methods for shield design process. The details of conceptual habitation module at an Earth-Moon libration point. the methods will be discussed. Execution times for these heretofore rather lengthy analyses have been reduced from matters of hours to matters of minutes.

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Development of Collaborative Engineering Immersive Shield Design of a Gateway Space Station Environments for Spacecraft Design Concept

Robert C. Singleterry Jr.1, Bradley D. Johns2, Kwok Y. Fan2, F. McNeil Chris A. Sandridge1, Brooke M. Anderson1, Aric R. Aumann2 Cheatwood3, Garry D. Qualls4, Todd A. Wareing5, John McGhee5, Shawn Pautz5, Anil Prinja6, Frederick Gleicher6, Greg Failla7, Jaroslaw 1NASA Langley Research Center, Hampton, VA Sobieszczanski-Sobieski1, John W. Wilson1 2Analytical Services and Materials, Inc., Hampton, VA

1Structures and Materials Competency, NASA Langley Research Center, Hampton, VA USA 23681 Abstract. The Immersive Design and Simulation Laboratory at NASA 2Swales Aerospace, NASA Langley Research Center, Hampton, VA Langley Research Center has recently developed an interactive, virtual USA 23681 reality simulation that accelerates the calculation, evaluation, and 3Space Access & Exploration Program Office, NASA Langley Research design of shield concepts for aerospace vehicles. This tool brings Center, Hampton, VA USA 23681 together 30 years of radiation shielding algorithms into a tools that can 4Systems Engineering Competency, NASA Langley Research Center, be easily used by the engineer and the physicist. The tool is designed Hampton, VA USA 23681 to run in a Cave Automatic Virtual Environment (CAVE), which is a 5Transport Methods Group, CCS-4, Los Alamos National Laboratory, projection based virtual environment in which the users are completely Los Alamos, NM USA 87545 surrounded by three-dimensional images. In the CAVE, the user can 6University of New Mexico, Chemical and Nuclear Engineering, navigate or walk anywhere in the virtual spacecraft and interactively Albuquerque, NM USA 87131 compute directional doses and dose distributions for various radiation 7ICEM CFD Engineering, Berkeley, CA USA 94705 environments. The user can then augment the environment by (757) 864 1437, [email protected] rearranging vehicle contents or by adding additional shielding. The radiation calculations can then be repeated to understand the effects of Abstract. The hazards of ionizing radiation in space continue to be a the augmentation. The radiation calculations are run interactively using limiting factor in the design of missions, spacecraft, and habitats. raytracing techniques and the HZETRN direct transport solutions. Shielding against such hazards is an enabling technology in the human and robotic exploration and development of space. If the design of the This presentation will describe the interactive, immersive simulation radiation shielding is not optimal for the mission, then excess mass will environment and describe how it is used to evaluate radiation shielding be launched, mission costs will be higher than necessary, and useful on a Gateway Space Station concept that is proposed for future Moon payload will be reduced. This reduced mission capability scenario can missions. Videos will be shown of the CAVE environment in action. be repeated for other technical design disciplines. These other disciplines can also effect each other’s requirements. A collaborative engineering environment can optimize ALL design parameters for cost (or another parameter) producing a spacecraft that will meet all mission requirements at the minimum cost. This paper will describe two environments that include space radiation analyses and the inclusion within a larger optimization methodology.

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Radiation Shielding Analysis for Deep Space Missions Gateway/L1 Modeling And Radiation Analysis

Giovanni De Angelis1, Martha S. Clowdsley2, John E. Nealy1, Brooke M. Anderson1, John E. Nealy2, James R. Geffre3, Garry D. R.C. Singleterry3, Ram K. Tripathi3, John. W. Wilson3 Qualls4, Shawn A. Krizan5, Pat Troutman4

1Old Dominion University, Norfolk, VA 23508 1Swales Aerospace, 1224T1 N. Wright St. Hampton, VA 23681 2College of William and Mary, Williamsburg, VA 23185 2Old Dominion University, NASA Langley Research Center, Hampton, 3NASA Langley Research Center, Hampton, VA 23681 VA 23681 757 864-1423, [email protected] 3NASA Johnson Space Center, Houston, TX 77058 4NASA Langley Research Center, Hampton, VA 23681 Abstract. An environment for radiation shielding analysis for manned 5Analytical Mechanics Associates, Inc., 17 Research Drive, Hampton, deep space mission scenarios has been developed. The analysis is VA 23666 performed by dividing a mission scenario into three possible different (757) 864-8459, [email protected] phases, namely the interplanetary cruise phase, the final planetary approach and orbit insertion, and the surface phase. In the first phase Abstract. The radiation environments at Earth-Lunar L1 are vastly only Galactic Cosmic Rays and Solar Events particles are used, in the different than what has been encountered before by humans, thus second phase the effects of trapped radiation belts are also taken into analysis of the predicted dose rates will need to be examined if the account, and in the third phase also the effect of the planetary Gateway scenario is to be properly evaluated. The environments to be environment is considered. Planetary surfaces and atmospheres are considered will be those of deep space. Recently a CAD model of the modeled based on results from the most recent targeted spacecraft. The three vehicles of the Gateway concept was developed to define the dose results are coupled with mission design visualization techniques. shielding properties inherent to the vehicles. The models consist of approximately 50-100 components that describe the directional shielding provided for the astronaut. The goals were to provide simple models so that near real-time analysis can be performed while the solid angles subtended by the modeled elements were compatible with those of the true vehicle models. The analysis will consist of determining the variable dose though out each of the Gateway vehicles. .

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Deep Space Environment and Shielding ISS as a Platform for Environmental Model Evaluation

J.W. Wilson1, J.E. Nealy2, G. de Angelis2, M.S. Clowdsley3, and F.F. Craig P. Hugger1, Garry D. Qualls2a, John W. Wilson2b, Frank A. Badavi4 Cuccinotta3, Mark R. Shavers3, and Neil Zapp4

1NASA Langley Research Center, Hampton, VA 23681 USA 1 Swales Aerospace, Mail Stop 186A NASA Langley, Hampton, VA 2Old Dominion University, Norfolk, VA 23508 USA 23681 3College of William and Mary, Williamsburg, VA 23185 USA 2aData Analysis and Information Branch and 2bAnalytical and 4Christopher Newport University, Newport News, VA 23602 USA Computational Methods Branch, 757 864-1414, [email protected] NASA Langley, Hampton, VA 23681 3 Johnson Space Center, Houston, TX 77058 Abstract. Mission scenarios outside the Earth's protective magnetic 4Lockheed-Martin Space Operations – Space Radiation Analysis shield are being studied. Included are high usage assets in the near- Group, Houston, TX 77058 Earth environment for casual trips, for research, for commercial and for 757 864 1654; [email protected] operational platforms, in which career exposures will be multimission determined over the astronaut's lifetime. In addition, the exploration Abstract. The International Space Station provides researchers with the beyond these near Earth operational platforms will include single unique opportunity to develop and test radiation environment models missions of long duration to planets, asteroids, and planetary satellites. and computational procedures. Since having manned missions into The interplanetary environment is evaluated using convective diffusion space is an ongoing part of NASA exploration of space, it has become theory. Local environments for each celestial body are modeled by necessary to conceive new methods to determine the safety of the using results from recent targeted spacecraft studies, and integrated into astronauts in their environment. One of the major environmental the design environments. Design scenarios are then evaluated for these hazards in space is the amount of radiation an astronaut will be exposed missions. The underlying assumptions in deriving the model to during their mission. NASA Langley Research Center, working in environments and their impact on mission exposures with various conjunction with Johnson Space Center, is developing a process that shield materials will be discussed. accurately simulates the radiation environment in the International Space Station. The environment includes the shielding provided by the ISS and the radiation environment of the ISS. The radiation shielding is modeled through the use of CAD software. Then, ray-tracing software is used to calculate thickness of the shielding for the model. A radiation environment is generated from data acquired from the NASA-GSFC National Space Science Data Center and transport codes for heavy ions and electrons. The radiation dose can then be calculated from the thickness files and the radiation environment. Various radiation dose experiments were and still are being performed on the International Space Station. Analysis of the radiation shielding environmental model can be compared to data obtained from the ISS. This enables researchers a way to perfect their methods for predicting radiation environments and allows them to create a tool that can be used for future endeavors into space.

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Space Radiation Shielding Calculation Models The Martian Radiation Environment Experiment For LEO Satellite (MARIE) on the 2001 Mars Odyssey Spacecraft

Myung-Won Shin, Myung-Hyun Kim William Atwell

Department of Nuclear Engineering, Kyung Hee University The Boeing Company, 13100 Space Center Blvd, MS: HS2-10, YongIn-Shi, Gyeongki-do, 449-701, Korea Houston, TX 77059 Tel:+31-201-2562, e-mail:[email protected] 281/226-5751, [email protected]

Abstract. Two approximate calculation models for a cosmic radiation shielding in LEO satellites were proposed. They are a sectoring method Abstract. The 2001 Mars Odyssey spacecraft was launched to and a chord-length distribution method. In order to simulate cosmic Mars on April 7, 2001, and reached Mars in late October 2001. radiation environments, IGRF model and AP(E)-8 model were used. Following arrival at Mars, a series of aero-braking maneuvers When an approximate method was applied in this study, complex were performed to place the spacecraft in a ~400 km altitude structure of satellite was described into multiple 1-dimensional slabs, orbit at Mars. One of the experiments on board the spacecraft is and the pre-calculated dose-depth conversion function was introduced the Martian Radiation Environment Experiment (MARIE), which to simplify the calculation process. Verification calculation was was designed to measure the galactic cosmic radiation (GCR) performed for orbit location and structure geometry of KITSAT-1 and environment in the energy range from 20– 500 MeV/u during compared with detailed 3-dimensional calculation results and the cruise phase to Mars and while in Mars orbit. The experimental values. The calculation results from approximate method were estimated conservatively with acceptable error. However, results experiment is also capable of measuring solar particle events for satellite mission simulation were underestimated in total dose rate (SPE) in the same energy range. During the cruise phase a compared with experimental values. communication problem with the MARIE was encountered from early August until early March 2002 when communication was re-established. In this paper the MARIE experiment is discussed in detail. The GCR measured data are compared with the GCR model environment. Several solar particle events that have been observed by MARIE are compared with the near-Earth GOES satellite SPE data. These data represent a “first ever” GCR and SPE measurements in the vicinity of Mars and will be used to establish some guidelines for future human missions to Mars. The SPE data can also be utilized to compare with existing SPE interplanetary propagation models.

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Surface Environments for Exploration History of Space Tourism

M.S.Clowdsley1, G.DeAngelis 2, F.F.Badavi 3, J.W.Wilson 4, Lorenzo Ortega III R.C.Singleterry 4, and S.A.Thibeault 4 1Spaceflight Unlimited, 312 Helene Drive, Bellevue, NE 68005 1College of William and Mary, Williamsburg, VA 23185 2Old Dominion University, Norfolk, VA 23508 3Christopher Newport University, Newport News, VA 23601 Abstract. This paper reviews the history of Space Tourism. Beginning 4NASA Langley Research Center, Hampton, VA 23681 with Akiyama Toyohiro, a Japanese news reporter through Lance Bass’ (757)864-1099; [email protected] efforts to secure the next tourist flight into space. Each step of the training and the business processes required to successfully arrange for Abstract. The ability to accurately model the radiation environment at the flight is covered. Particular attention is devoted to the other any time for planetary surfaces is a necessity in the evaluation of health burgeoning aspects of space tourism including tourists visiting launch risk to astronauts on deep space missions. Several examples, including sites, museums and participating in space training in the United States models for Lunar, Martian, and Jovian environments, are discussed. In and in Russia. these models, the differential spectra for neutrons as well as charged ions are evaluated. Since surface radiation environments are made up of a combination of free space particles scattered through the planetary atmosphere and backscattered particles from the surface, these models include a dependence on both altitude and surface material as well as time in the solar cycle.

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Benefit Estimation Model for Tourist Spaceflights Business Context of Space Tourism

Robert A. Goehlich Harrison H. Schmitt

Technical University Berlin, Institute of Aero- and Astronautics, University of Wisconsin-Madison, P.O. Box 90730, Albuquerque, NM Spacecraft Technology, 87199 Secr. F6, Marchstrasse 12, 10587 Berlin, Germany (505) 823-2616, [email protected] Tel: +49-30-314 79 464, Fax: +49-30-314 21 306, eMail: [email protected], www.Robert-Goehlich.de Abstract. Broadly speaking, two types of potential commercial activity in space can be defined. First, there are those activities that represent Abstract. It is believed that the only potential means for further an expansion and improvement on services with broad existing significant reduction of the recurrent launch cost, which results in a commercial foundations such as telecommunications. The second type stimulation of human space colonization, is to make the launcher of potential commercial activity in space is one that may offer a type of reusable, to increase its reliability, and to make it suitable for new service with few or any existing commercial foundations such as space- markets such as mass space tourism. But such space projects, that have based remote sensing. Space tourism clearly belongs in the first long range aspects are very difficult to finance, because even politicians category of potential commercial activity in space. Roles in would like to see a reasonable benefit during their term in office, cooperation with the private sector that might be considered for NASA because they want to be able to explain this investment to the taxpayer. include 1) acceleration of the "Professional-in Space" initiative, 2) research and technology developments related to a) a "Tourist This forces planners to use benefit models instead of intuitive Destination Module" for the Space Station, b) an "Extra Passengers judgement to convince sceptical decision-makers to support new Module" for the payload bay of the Space Shuttle, and c) a "Passenger- investments in space. Benefit models provide insights into complex rated Expendable Launch Vehicle," 3) definition of criteria for relationships and force a better definition of goals. In the paper a new qualifying candidate space tourists, and 4) initiatives to protect space approach is introduced that allows to estimate the benefits to be tourism from unreasonable tort litigation. As baseline information for expected from a new space venture. The main objective why humans establishing fees, the cost of a possible tourist flight should be fully and should explore space is determined in this study to “improve the quality objectively delineated. If it is correct that the marginal cost of each of life”. For clearness, this main objective is broken down in sub Space Shuttle flight to Earth-orbit is about $100 million and the objectives, which can be analysed in respect to different interest effective Shuttle payload is about 50,000 pounds, then the marginal groups. Such interest groups are the operator of a space transportation cost would be roughly $2,000 per pound. system, the passenger, and the government. For example, the operator is strongly interested in profit, while the passenger is mainly interested in amusement, while the government is primarily interested in self- esteem and prestige. This leads to different individual satisfactory levels, which are usable for the optimisation process of reusable launch vehicles.

KEYWORDS: Benefit Model, Space Tourism, Reusable Launch Vehicle

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The Pioneer Rocketplane XP Aircraft Space Stations For Dummies, Part 1: Accounting For Space Mitchell Burnside Clapp Christopher Lee Martens Pioneer Rocketplane, 1607 Mission Drive, #301, Solvang, CA 93463 (805) 693-8222, FAX: (805) 693-8122, [email protected] Mutual Space, Ltd., PO Box 2245, Crestline, CA 92325-2245 (909) 338-3358, [email protected] Abstract. The Pioneer XP is a four-seat fighter-sized vehicle powered by two jet engines and two rocket engines, enabling it to reach altitudes Abstract. The next generation of multi-use space stations will be of 350,000 feet. The XP will operate from ordinary airfields within the completely built using private, commercial funding. In order to well-established rules and practices for experimental aircraft. The XP understand how private industry will take over the role usually reserved does not use any launch assist: No airdrop, no towing, no aerial for governments, it is necessary to take a brief look at the history of propellant transfer, and no developmental engines. It will achieve up to space expenditures up to this point. This paper will examine the 6,000 ft/s and 350,000 feet altitude. differences between the perceived high cost of space and its actual cost. It will also address the major misconceptions which have hindered Safety, existing aircraft rules, and the constraints of cost and schedule private investment and lay the groundwork to demonstrate numerous force the Pioneer Rocketplane team to select oxygen and kerosene as ways in which the cost of space can and will be greatly reduced. An the main propellants, to use a simple and reliable propellant feed application of this cost cutting approach to space, shown in terms of system, to employ safe, already certified aircraft components, and to cost to payback, will be examined using an overview of a Multi-use plan on an incremental flight test approach. Space Station Design. Although we plan to fly the aircraft soon, affordably, and safely, Pioneer is also developing an aircraft that can be upgraded to higher performance in the future. The thermal protection system, the wing propellant tanks, and the other systems are designed with the possibility of eventual upgrade in mind from the beginning, so that the option for the long-range aircraft can be preserved without requiring expensive reinvestment after the XP flies. It is our intention not merely to reach 350,000 feet, but to do it with an aircraft that is traceable to not only a satellite launch system but also a long range transport aircraft in follow- on versions. The XP itself is intended for passenger travel, promotions and sponsorhsip, and research and observation applications.

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Human Mission from Planet Earth: Technology Energy Policy for Habitation and Terraforming Mars Assessment and Social Forecasting the Future of Space Thomas Meyer1, Christopher P. McKay2 Eligar Sadeh 1Boulder Center for Science & Policy, Box 4877, Boulder, CO 80306 2 University of North Dakota, Dept. of Space Studies, Grand Forks, ND NASA Ames Research Center, MS 254-3, Moffett Field, CA 94035 58202-9008 303-494-8144, [email protected] (701) 777-3462, [email protected] Abstract. Of all the planets in the Solar System, Mars is the most Abstract. This paper advances scenarios for an evolutionary approach attractive candidate for future colonization. It has all of the resources to the establishment of a Human Mission from Planet Earth (H-MFPE) necessary to sustain life in some accessible form on the surface. It has involving exploration and utilization of the Moon and Mars. Of critical abundant amounts of water, adequate sunlight, soil suitable for plant importance, are the concepts of robotic/human and Moon/Mars growth, a rotation rate similar to that of Earth, a thin but useful synergies. The technological, scientific, political, and economic atmosphere, a manageable radiation environment, and a climate that impacts and consequences related to H-MFPE are presented and does not preclude human habitation. A significant difference between discussed. Earth and Mars is that Mars does not have stored energy resources such as fossil fuel and atmospheric oxygen. Thus all power and chemical fuels needed on Mars must either be imported or produced from primary energy sources. If Mars is to become home to substantial populations of millions or even billions of people, then energy technologies and a strategy for infrastructure investment that will minimize costs over the long-term must be developed. As a way of estimating the energy requirements for a large population on Mars with an assumed standard of living comparable to that of the United States, we can take the per capita energy consumption of the United States (less wastefulness) and add to that an estimate of the amount of energy necessary to compensate for the harsher conditions and environmental deficiencies as they currently exist on Mars. Mechanical approaches to this compensation will generally involve additional expenditures of energy, including energy to offset the cold, to prepare breathable air, to mine water-ice or desalinate brine sources, to operate life support systems, to operate greenhouses, to maintain pressurized vehicles, buildings and cities, to protect against radiation, to manufacture protective clothing, and many other measures that may be needed in conjunction with endeavors such as mining, manufacturing, agriculture, transportation, business and recreation. In this paper we evaluate the cost of sustaining a substantial Martian colony entirely by mechanical means and compare it with a strategy that incorporates a program of global climate modification and development of a biosphere that would ameliorate the harsh physical and environmental conditions and thereby reduce the overall energy requirement for a Martian colony.

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The International Space Development Authority A System Design for a Compact, Renewable and Energy Efficient, Oxygen, Waste Recycle and Food Supply for Declan J. O’Donnell Manned Travel to Mars Untied Societies in Space, Inc., 499 South Larkspur Drive, Castle Rock, George F. Erickson Colorado 80104 303-688-1193, Fax: 303-663-8595, [email protected] Nanotube Engineering, 4 Hopi Lane, Los Alamos, NM 87544-3813 505-672-9818, [email protected] Abstract. The architecture of space governance may feature a nation in space at some future date. In the interim and in order to prepare for an off world estate, the United Nations and the space faring nations Abstract. Conventional methods of growing food will not by possible for manned travel to Mars. Issues of energy efficiency (waste heat), may join in the formation of a special legal authority to coordinate and volume, low or micro-gravitational environments, oxygen generation help finance space development. The legal basis for this is discussed, as and adaptability to life on Mars or the Moon are just some of the well as the structure of the entity for international imprimatur. challenges that need to be addressed before a manned trip to Mars can be considered. This paper outlines a system concept that attempts to answer these questions and several other systems engineering issues associated with this type of endeavor.

In order to send six people to Mars, ~ 300 square meters of conventional farm would be required to sustain the food and oxygen needs of the crew. Conventional white light hydroponic farming would require nearly one megawatt of electrical energy and nearly all of that would need to be rejected to space as waste heat. At room temperature, this would require a heat rejection panel of nearly 2400 square meters. A strain of Blue-Green algae, Spirulia Platensis can provide all of the 11 critical amino acids and lipids needed for sustaining life. In addition, the algae does not suffer from geo-tropism, does not need a dark cycle (24 hour a day growth without a dark cycle) and can be successfully grown using a combination of red and blue LED (light emitting diodes). In addition, these algae can be grown in a low- pressure atmosphere (1-PSI absolute). Thus, the waste heat requirement for this system is reduced to approximately 30 kilowatts. Using a small energy efficient heat pump, the radiator area for this heat source could be less than 5 square meters. In addition, the 6 - 500 gallon production tanks can serve as radiation shielding for the crew by placing them on the perimeter of the ship.

By using existing technology applied to several key system issues, it appears that a critical hurdle to the goal of manned trips to Mars might be solvable. In addition, this technology is readily adaptable to a continued life on the moon or mars.

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The Next Logical Step – Post ISS Pathways To Colonization

Douglas A. O’Handley David V. Smitherman, Jr.

Department of Physics, Santa Clara University, Santa Clara, CA 95053 NASA, Marshall Space Flight Center, Mail Code FD02, Huntsville, AL (408) 736-5937, [email protected] 35812, 256-961-7585, [email protected] Abstract. As the completion of the International Space Station soon becomes a reality, what is the next step going to be to sustain the human Abstract. The steps required for space colonization are many to grow exploration of space? With the discovery of potential ice at the poles of the from our current 3-person International Space Station, now under Moon by the lunar Prospector Mission managed by Ames Research Center construction, to an infrastructure that can support hundreds and and the recent report of the Space Studies Board of the National Research eventually thousands of people in space. This paper will summarize the Council, “New Frontiers in the Solar System – An Integrated Exploration author’s findings from numerous studies and workshops on related Strategy,” that recommends a return to the moon of a robotic mission to subjects and identify some of the critical next steps toward space investigate the Aiken Basin, the Moon as a destination for exploration colonization. Findings will be drawn from the author’s previous work seems to be returning to the realm of targets for both robotic and human on space colony design, space infrastructure workshops, and various exploration. There is obvious interests in the Moon from Europe with the studies that addressed space policy. In conclusion, this paper will note launch SMART 1 mission of the European Space Agency early in 2003 that significant progress has been made on space facility construction and the Lunar Penetrator mission to be launched in 2003 and the SELENE through the International Space Station program, and that significant missions of Japan, and then the interest by India in an orbiter and supposed efforts are needed in the development of new reusable Earth to Orbit lurking interests of China. transportation systems. The next key steps will include reusable in The idea of a colony of robots working cooperatively could set the course space transportation systems supported by in space propellant depots, for future construction of radio observatories and optical observatories on the continued development of inflatable habitat and space elevator the Moon, exploration of the poles and establishing a solar power technologies, and the resolution of policy issues that will establish a transmission technology from the Moon. These technologies are well future vision for space development. within our current technology base. Their development and use on the Moon would enhance any future exploration of Mars. The concept of in-situ resource utilization has become an important part of current architecture for both the moon and the eventual human exploration of Mars. Confirmation of water ice will make the human mission easier than previously suspected. The push for the human exploration of Mars requires a base on the Moon to test the equipment to be sent to Mars. Investigation of the human physiological effects expected on a human mission on Mars is also possible in an environment much similar to the actual risks than those found in Earth based analogs. Finding a solution for the harsh radiation of the Moon would clearly reduce the radiation challenges to be encounter on the surface of Mars. The solution to power requirements for a self-sustaining colony on the Moon to survive a lunar night will enhance the reliability of any power system taken to another body in our Solar System. This paper shall exam work at ORBITEC which was completed under a NASA Institute for Advanced Concepts grant to establish parameters for a self-sustaining human base on the moon.

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A View of Future Human Colonies on Mars New Concepts for Permanently Manned Lunar Bases, Report of the Lunar Base Design Workshop, held in Robert Gustafson, Eric Rice, Daniel Gramer, Brant White Noordwijk, The Netherlands from 10-21 June 2002 Orbital Technologies Corporation (ORBITECTM) Barbara Imhof1, Susmita Mohanty2, Hans Jurgen Rombaut3, Space Center, 1212 Fourier Drive, Madison, WI 53551 Paul J. van Susante4, Jim Volp5 (608) 827-5000, [email protected] 1 Architect, Institut fuer hochbau II, TU-Vienna, Karlsplatz 13/270 2, Abstract. In a recent feasibility study, ORBITEC conceptualized A-1040 Vienna and Principal, ESCAPE*spHERE, grosse mohrengasse systems and an evolving architecture for producing and utilizing Mars- 38/1, A-1020, Vienna, Austria based in-situ resources utilization (ISRU) propellant combinations. 2 Industrial Designer, Principal, Moonfront LLC, 550 Battery Street, The propellants will be used to support the propulsion and power Suite 804, CA94111, San Francisco, USA systems for ground and flight vehicles that would be part of Mars 3 Architect, Principal, Lunar Architecture, Utrecht, The Netherlands exploration and colonization. The key aspect of the study was to show 4 Civil Engineer, Graduate Research Assistant, Colorado School of the benefits of ISRU, develop an analysis methodology, as well as Mines ,department of Engineering, 1500 Illinois street, Golden, CO provide some guidance to propellant system choices in the future based 80401, Colorado, USA upon what is known today about Mars. The study time frame includes 5 Astrophysicist, Young Graduate Trainee, ESA/ESTEC, Keplerlaan 1, the early unmanned and manned exploration period (now to 2040) and 2200 AG, Noordwijk, The Netherlands a colonization period that occurs from 2040 to 2090. As part of this 4 tel. +1-(303)-216-0632, e-mail : [email protected] feasibility study, ORBITEC developed two different Mars colonization scenarios, namely a low case that ends with a 100-person colony and a Abstract. This paper presents some of the results of the Lunar Base high case that ends with a 10,000-person colony. A population growth Design Workshop, held in Noordwijk, The Netherlands from 10-21 model, mission traffic model, and infrastructure model was developed June 2002. Six groups designed six different lunar bases according to for each scenario to better understand the requirements of future Mars six different scenario’s. The main findings have to do with the colonies. This paper will outline the characteristics of the Mars horizontal and vertical movement in 1/6G in shirt sleeve environment. colonies that ORBITEC envisions under both colonization scenarios. Different concepts for interaction and “contact” with the green inside This would include a discussion of the flow of people and materials the base. Another topic was the use of water ice which can be used in between the Earth and Mars, the infrastructure requirements of the several forms. Organic growth using small elements to create a larger colonies, a potential colony base configuration, and the mission space in the desired shapes was important for several reasons. Man and requirements of the colonies. its activities were the central spine in this design workshop. The results are useful together with the already existing more engineering oriented design studies for the next steps in lunar base development. In the talk, the results will be discussed and compared to previous designs

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Stages in the Terraforming of Mars: the Transition to Terraforming Mars: Can We Feed Ourselves If We Go? Flowering Plants James M. Graham and Kandis Elliot James M. Graham Department of Botany, University of Wisconsin-Madison, Madison, WI Department of Botany, University of Wisconsin-Madison, Madison, WI, 53706, USA 608-262-0657, [email protected] 608-262-0657, [email protected]

Abstract. The process of the biological terraforming of Mars can be Abstract. Scientists and engineers have recognized for a long time that compared to the process of primary ecological succession on terrestrial transporting food supplies from Earth to extraterrestrial colonies is not barren rocks. Each stage in the succession alters the environment in a feasible option over the long run because of weight, distance and such a way that the next stage in the process becomes possible. The ultimately costs. It has therefore been accepted that colonists on the initial stage in terraforming Mars will be dominated by microorganisms Moon and Mars will have to be capable of feeding themselves from and lichens. The initial stage will begin the process of removing carbon greenhouse-grown crops. A number of greenhouse designs have been dioxide from the Martian atmosphere, adding oxygen and nitrogen, and proposed, and the designs fall into two categories: inflatable adding organics to the regolith to produce a true Martian soil. The greenhouses manufactured on Earth and transported to Mars and second stage will be dominated by bryophytes, simple plants such as greenhouses manufactured on Mars using local materials and resources. mosses and liverworts, which will draw down the carbon dioxide level Both types come in small, spartan versions and large-scale architectural of the Martian atmosphere and raise the level of oxygen. The carbon masterpieces, the latter envisioning beautiful mall-like vistas in which dioxide removed will be locked up in peatlands and permafrost. The grain fields and orchards are interspersed with fishponds and goat critical limiting factors for the introduction of flowering plants are the pastures, while strawberry-munching gardeners oversee the scene. We level of oxygen in the atmosphere and the lack of animal pollinators. propose some guidelines for the construction of greenhouses on Mars The majority of flowering plants require a minimum oxygen level of 20 with the emphasis on the biological challenges presented by growing to 50 mbar. Most flowering plants require these minimal oxygen levels plants and other organisms on a continual basis to sustain a human to support aerobic respiration in their roots and germination of their population. We regard greenhouses to be the first, small terraformed seeds. Many flowering plant species also must have animal pollinators surface area on Mars and will suggest how the knowledge gained in to complete reproduction. Certain aquatic plants and arctic plants, their operation may be applied to the later course of biological however, are highly tolerant of anoxic conditions. Some of these same terraforming of the entire planet. arctic plants can successfully reproduce without animal pollinators by employing one or more alternate reproductive mechanisms such as vegetative propagation, apomixis, autogamy and anemophily. Thus by judicious selection of existing terrestrial plants and possibly genetic engineering, it may be possible to circumvent critical limitations and introduce flowering plants to Mars at an earlier stage in terraforming.

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Terraforming Mars: The Fluorine Bottleneck and the Artificial Biogeochemical Cycles for Mars Importance of Sample Return Penelope J. Boston, Benton C. Clark III New Mexico Institute of Mining & Technology, 801 Leroy Place, Lockheed Martin Astronautics, POB 179, Denver, CO 80201 Socorro, NM 87801-4796 USA (505) 835-5657; [email protected]

Abstract. Fluorinated compounds are some of the most powerful Abstract. Earth cycles its materials over geologically significant time members of the greenhouse gases. Fluorine (F) has not been detected periods using the massive engine of plate tectonics. On Mars, plate on Mars, either in the atmosphere or in the regolith fines (soils). tectonic motions have not existed for a very long time, if ever. What Martian meteorites are low in F, from 10-60 parts per million (ppm). mechanisms can be devised to replace at least some of the recycling Although F could be exported to Mars from Earth, the transportation functions performed on Earth? The possibilities include use of costs would be prohibitive for the quantities needed to seriously microorganisms tailored to minimize material losses to sediments, deep enhance the retention of surface-emitted IR by the martian atmosphere. injection of microorganisms or nanomechanisms to remobilize and However, there is a distinct possibility that the martian fine-grained volatilize sediments, a biota designed to keep materials sequestered in regolith is significantly enriched in this element. Other halogens, the biological rather than lithological component of the system, including both chlorine and bromine were detected far above expected creation of tidal forces via artificial satellites, selective impacts to values by the Viking soil analyzers. If fluorine is similarly enriched, its revolatilize crustal materials, re-awakening volcanism, or abandoning concentration may be 100 times that in the martian rocks. Another the notion of recycling and opting to bring in fresh material via towed unknown factor of considerable importance is the chemical state of asteroids or comets. Are any of these methods feasible in terms more fluorine in the soil. Chlorine is likely present as water-soluble salts, as concrete than science fiction? We will present a plausibility “trade-off” is bromine. Fluorinated salts typically have low solubility. If enriched, study of suggested mechanisms from the ridiculous to the sublime. and if easily extracted and separated, the martian resource in F could be considerable. If neither of these cases is true, the long-term cost of terraforming the red planet could be higher by orders of magnitude. Because of its low concentration (not detected so far), and the need to determine its exact chemical state(s), as well as those of potentially interfering compounds, it is essential to conduct a return of martian regolith material so that proper analyses can be accomplished within sophisticated soil geochemistry laboratories here on Earth. Once that is done, this aspect of terraforming feasibility can be evaluated with confidence. The martian regolith fines have been shown to be uniform over large areas, if not all of the planet, and would be a universal source of fluorine, as well as chlorine, for the production of highly efficient greenhouse gas.

ISRU Development Strategy & Recent Activities to Evaluation of Private Sector Roles in Space Resource Support Near & Far Term Missions Development

1 2 2 Russell S. Baird, Gerald B. Sanders, and Thomas M. Simon Elisabeth S. Lamassoure , Brad R. Blair , Javier Diaz , Mark Oderman3, Michael B. Duke2, Marc Vaucher3, Ramachandra Manvi1, 1 Energy Systems Division, NASA Johnson Space Center, Houston, TX and Robert W. Easter 77058, USA 1Mission and Systems Architecture Section, Jet Propulsion Laboratory, Pasadena, CA 91109 (281) 483-9013, [email protected] 2 Center for Commercial Applications of Combustion in Space (CCACS), Colorado School of Mines, CO 81401 3 Abstract. The practical expansion of humans beyond low Earth orbit CSP Associates, Inc., Cambridge, MA 02139 into near-Earth space and out into the solar system for exploration, 818 393 6933, [email protected] commercialization, tourism, and colonization will require the effective utilization of whatever indigenous resources are available to make these endeavors economically feasible and capable of extended operations. This concept of “living off the land” is called In-Situ Abstract. An integrated engineering and financial modeling approach Resource Utilization (ISRU). The resources available for ISRU has been developed and used to evaluate the potential for private sector applications vary widely, depending upon the location. However, investment in space resource development, and to asses possible roles there are resources, technologies, and processes that are common to of the public sector in fostering private interest. This paper presents the multiple destinations and ISRU-related applications. These resources modeling approach and its results for a transportation service using range from carbon dioxide (CO2) and water vapor found in human propellant extracted from lunar regolith. The analysis start with careful habitats (surface & spacecraft) and in the Martian atmosphere, to case study definition, including an analysis of the customer based and water (ice and hydrated minerals) and various oxygen, carbon, and market requirements, which are the basis for design of a modular, metal-bearing resources found on comets and asteroids, and in scalable space architecture. The derived non-recurring, recurring and planetary surface materials at numerous destinations of interest operations costs become inputs for a “standard” financial model, as (Moon, Mars, Titan, and Europa). Many parties are investigating the used in any commercial business plan. This model generates Pro common technologies and processes to effectively extract and use Forma financial statements, calculates the amount of capitalization these resources. This paper will discuss how ISRU is enabling for required, and generates return on equity calculations using two both near and far term human exploration missions, and present a valuation metrics of direct interest to private investors: market summary of recent and on-going ISRU work sponsored by the enterprise value and multiples of key financial measures. Use of this NASA/Johnson Space Center. Technology development activities model on an architecture to sell transportation in Earth orbit based on that will be described in detail include an advanced CO2 freezer lunar propellants shows how to rapidly test various assumptions and acquisition system, a multi-fluid common bulkhead cryogenic storage identify interesting architectural options, key areas for investment in tank, and a variety of microchannel chemical reactor concepts. Recent exploration and technology, or innovative business approaches that advanced Sabatier reactor concept development activities in preparation for later, end-to-end system testing will be described as could produce an economically viable industry. The same approach well. This paper will also discuss an ISRU-based strategy to enable can be used to evaluate any other possible private venture in space, and extensive robotic and human surface exploration operations and a conclude on the respective roles of NASA and the private sector in related on-going demonstration program for a fuel cell based power space resource development and solar system exploration. plant for rover applications. Technology commonalities between ISRU, Environmental Control and Life Support System (ECLSS), and Extra Vehicular Activity (EVA), applications will also be presented.

Investigation into Uses for Lunar Regolith Surface Mine Design And Planning For Lunar Regolith Production Charles Horton, Carlos Gramajo, Lance Williams, Andenet Alemu, Alex Freundlich and Alex Ignatiev Leslie Sour Gertsch and Richard E. Gertsch

Texas Center for Superconductivity and Advanced Materials, Department of Geological and Mining Engineering and Sciences, University of Houston, 724 Science and Research Bldg 1 Michigan Technological University, Houghton, MI 49931 Houston, Texas 77204-5004 906-487-2233; [email protected] 713-743-3621, fax 713-747-7724, [email protected] Abstract. Terrestrial surface mine design and planning techniques are applied to the production of lunar regolith for manufacturing makeup Abstract. Any sustained effort on the Moon will require use of in-situ gases for the life-support system of a lunar base. Two scenarios are resources as much as possible to reduce costs. Power generation is a examined, due to the uncertainty of whether bound hydrogen sensed primary concern and solar cells would be a convenient source of power near the lunar poles is from cometary ice deposited in cold traps (#1), for any lunar activity. Lunar regolith is an easily accessible in-situ or to hydrogen implanted within regolith grains by the solar wind (#2). resource. The lunar regolith consists primarily of silicon dioxide (about Scenario #1, with a total production requirement of 44 tonne/day of half) with aluminum oxide, iron oxide, calcium oxide and magnesium regolith, could be handled with four groups of four 6-m3 capacity oxide comprising the majority of the residual. The regolith can be slushers (drag scrapers), each group extending 100 m from a single refined to extract silicon, aluminum and other potentially useful metals processing module. Scenario #2 (4,382 tonne/day) could be while also producing oxygen. Silicon and aluminum can be used in the accomplished with three powered bowl-type scrapers (capacity 24 m3) fabrication of lunar based solar cells. In addition, lunar regolith can be gathering the regolith into long windrows more amenable to direct melted to form a glassy substrate upon which the silicon solar cells are processing. The present orebody model is extremely thin (1 m), deposited. The glassy regolith melt is electrically insulating providing although broad in extent; this prevents usage of high production-rate good isolation for solar cells. In addition, films of the lunar regolith are systems such as large draglines. Multiple-machine systems provide transparent at thicknesses less than a half micron. Measurements of the greater flexibility in scheduling than larger single units. The main index of refraction indicate a value suitable for optical coatings. The difference between these two systems is the speed of haulage. Being regolith film could be used in antireflection coatings, transparent simpler, the slusher system is likely to be easier to adapt directly to protective coatings and possibly even in electronic devices as insulating lunar mining. films. We will present results of our investigation into lunar regolith material properties and some alternative uses for the lunar regolith.

Operation, Modeling and Analysis of the Reverse Water Extraction of Water from the Martian Atmosphere Gas Shift Process Matthew A. Schneider and Adam P. Bruckner Jonathan E. Whitlow1 and Clyde F. Parrish2 Department of Aeronautics and Astronautics, University of 1Department of Chemical Engineering, Florida Institute of Technology Washington, Box 352400, Seattle, WA 98195-2400 Melbourne, Florida 32951 (206) 543-6143, [email protected] 2Spaceport Engineering and Technology Directorate, YA-C3 NASA John F. Kennedy Space Center, Florida 32899 Abstract. We report on experiments to validate a concept for (321) 674-7354, [email protected] extracting water vapor from the Martian atmosphere by adsorption. The adsorber investigated is zeolite 3A, a synthetic molecular sieve that Abstract. The Reverse Water Gas Shift (RWGS) process is a has a cage-like microstructure with an effective aperture of 3Å. This candidate technology for water and oxygen production on Mars as part micropore size allows water molecules to enter but excludes carbon of the In-Situ Space Resource Utilization (ISRU) initiative. This paper dioxide molecules, which comprise 95% of the Martian atmosphere. focuses on the operation and analysis of the RWGS process, which has Although the literature on zeolites is extensive, there is little been constructed and operated at Kennedy Space Center. While the information available on type 3A zeolite at Mars ambient conditions, investigation of the RWGS process is on-going, a summary of results i.e., temperatures of ~150-250 K and water vapor pressures ranging obtained from the operation to date is presented. In addition, simulation from ~10-3 Pa to ~1 Pa. To obtain the desired adsorption data we models of the RWGS process have been developed and description of developed a continuous flow Mars environmental simulation facility the models is also included. capable of precisely generating and controlling for long durations the desired pressures, temperatures, and humidities. The experiments conducted to date have produced measurements of the magnitudes and trends of the uptake kinetics and equilibrium capacity of a small-scale packed bed of zeolite 3A at conditions approaching those on Mars. Although the data obtained to date are preliminary, they show that zeolite 3A is capable of capturing significant quantities of water vapor at these conditions.

Solar Energy for In-Situ Resource Utilization in Space High Temperature Interaction Between H2, CH4, NH3 and Ilmenite Takashi Nakamura, John A. Case, and Connie. L. Senior Giovanni De Maria1, Bruno Brunetti2, Giuseppe Trionfetti1, Daniela Physical Sciences Inc., 2110 Omega Rd, Suite D Ferro2 San Ramon, CA 94583 925-743-1110; FAX: 925-743-1117; [email protected] 1Diparimento di Chimica, Università La Sapienza, P.le A. Moro 5, 00185 Roma, Italy Abstract. In the past 10 years, Physical Sciences Inc. (PSI) has been 2Istituto per lo Studio dei Materiali Nanostrutturati – CNR c/o developing the Optical Waveguide (OW) system for solar energy Diparimento di Chimica, Università La Sapienza, P.le A. Moro 5, utilization. In this system, solar radiation is collected by the 00185 Roma, Italy concentrator which transfers the concentrated solar radiation to the OW [email protected], 39-064462950 transmission line. The OW transmission line transports the solar radiation to the location of solar energy utilization. Applications of this Abstract. Over twenty different chemical and physical processes have system include: material processing, plant lighting and power been proposed for the oxygen production from lunar materials in the generation in space. In this paper, a review of our work conducted last years. They can be grouped into major categories characterized by during the last 10 years pertinent to in-situ resource utilization in space solid-gas interaction, pyrolisis, silicate-oxide melt, aqueous solution is presented. and coproduct recovery. Among these processes the hydrothermal reduction of ilmenite occupies a prominent place. A kinetic study of this reaction was carried out in our laboratory by Bardi et al. and subsequently a solar energy sound assisted fluidized bed process for the oxygen extraction at high temperature was worked out. In this optique it appeared of interest to study the mechanism and the kinetic of some reactions involving other reducing agents (CH4, NH3) with the aim to compare the relative merits of different oxygen extraction processes. For each compound, the reduction process was investigated at different temperatures values and different values of gas flux and pressure, utilizing a commercial thermobalance, Ugine Eyraud. Setaram B 60 coupled with a quadrupole mass-spectrometer (Balzers QMA 410) The processes can be described according to the following relations: Reduction with hydrogen: FeTiO3(s) + H2(g) Fe(s) + TiO2(s) + H2O(g) Reduction with methane: CH4(g) C(s) + 2H2(g) FeTiO3(s) + H2(g) Fe(s) + TiO2(s) + H2O(g) C(s) + H2O(g) CO(g) + H2(g) Reduction with ammonia: FeTiO3(s) + 2/3NH3(g) Fe(s) + TiO2(s) + H2O(g) + 1/3N2(g)

even if some parallel reactions occur owing to the change of the components ratio with the evolution of H2 reduction process; in particular the catalytic effect induced by the metallic iron on the decomposition of methane and ammonia has been evidenced.

Visions and Possibilities for Future Exploration of The Initial Nine Space Settlements Space Anita E. Gale1, 2a and Richard P. Edwards1, 2b Robert A. Cassanova1, Ron Turner2, and Patricia Russell3 1Co-Founders, Space Settlement Design Competitions 1NASA Institute for Advanced Concepts, 555-A 14th Street, 18506 Upper Bay Rd., Nassau Bay, TX 77058 2a 2b Atlanta, GA 30318 Senior Project Engineer, and Principal Engineer, Space Shuttle 2ANSER Analytic Services, Inc., 2900 South Quincy Street Suite 800 Program, The Boeing Company, Houston, TX Arlington, VA 22206 (281)226-5691, [email protected] 3Universities Space Research Association, 10227 Wincopin Circle, Suite 212, Columbia, MD 21044 Abstract. The co-authors describe a chronology of space infrastructure 1404-347-963;3 [email protected] development illustrating how each element of infrastructure enables development of subsequent more ambitious infrastructure. This is Abstract. The NASA Institute for Advanced Concepts (NIAC) is likened to the “Southern California freeway phenomenon”, wherein a sponsored by NASA HQ to inspire, solicit, select, fund and nurture new freeway built in a remote area promotes establishment of gas revolutionary advanced concepts for aeronautics and space. NIAC has stations, restaurants, hotels, housing, and eventually entire new funded concepts that explore the possibilities of significant communities. The chronology includes new launch vehicles, inter-orbit breakthroughs and paradigm shifts applicable to the exploration of vehicles, multiple LEO space stations, lunar mining, on-orbit space. This paper summarizes NIAC funded advanced concepts that manufacturing, tourist destinations, and supporting technologies directly relate to colonization of planetary surfaces by humans and their required to make it all happen. The space settlements encompassed by agents. the chronology are in Earth orbit (L5 and L4), on the lunar surface, in Mars orbit, on the Martian surface, and in the asteroid belt. Each space settlement is justified with a business rationale for construction. This paper is based on materials developed for Space Settlement Design Competitions that enable high school students to experience the technical and management challenges of working on an industry proposal team.

Colonization of Venus Implications of Outside-the-Box Technologies on Future Space Exploration and Colonization Geoffrey A. Landis Theodore C. Loder III NASA John Glenn Research Center, mailstop 302-1, 21000 Brook Park Road, Cleveland, OH 44135 Institute for the Study of Earth, Oceans, and Space 216-433-2238; [email protected] University of New Hampshire; Durham, NH 03824 603-862-3151, [email protected] Abstract. Although the surface of Venus is an extremely hostile environment, at about 50 kilometers above the surface the atmosphere Abstract. In general, planning for future manned space exploration of Venus is the most earthlike environment (other than Earth itself) in either to the moon, Mars, or an asteroid has depended on a somewhat the solar system. It is proposed here that in the near term, human linear extrapolation of our present technologies. Two major prohibitive exploration of Venus could take place from aerostat vehicles in the cost issues regarding such planning are payload lift and in-flight energy atmosphere, and that in the long term, permanent settlements could be generation. The costs of these in both engineering and actual flight made in the form of cities designed to float at about fifty kilometer costs, coupled with the planning necessary to carry out such exploration altitude in the atmosphere of Venus. have prevented us from actively moving forward. Although, it will be worthwhile to continue to plan for such exploration using "present" technologies, I recommend that planning be concerned mainly with mission strategies and goals utilizing both present technology and total new energy breakthroughs. There are presently in research and development an entire suite of relevant outside-the-box technologies which will include both zero point energy generation and antigravity technologies that will replace our present solar/nuclear/fuel cell energy technologies and liquid/solid fuel rockets. This paper describes some of these technologies, the physics behind them and their potential use for manned space exploration. The companies and countries that first incorporate these technologies into their space programs will lead the way in exploring and colonizing space.

Space Colonization—Benefits For The World Space Colonization - Economic Drivers and Justifications W. H. Siegfried Gordon Woodcock The Boeing Company, Integrated Defense Systems, Huntington Beach, California 92647 Gray Research, Inc. 714-896-2532, [email protected] 675 Discovery Drive, Suite 302, Huntsville, AL 35806 (256) 880-9708, [email protected] Abstract. We have begun to colonize space, even to the extent of early space tourism. Our early Vostok, Mercury, Gemini, Apollo, Skylab, Spacehab, Mir and now ISS are humankind’s first ventures toward Abstract. The idea that people can find permanent homes off Earth colonization. Efforts are underway to provide short space tours, and has been in science fiction almost since science fiction began, and has endeavors such as the X-Prize are encouraging entrepreneurs to provide been a matter of serious technical discussion since the 1960s. This new systems. Many believe that extended space travel (colonization) paper defines settlements as distinct from outposts or bases. A key will do for the 21st century what aviation did for the 20th. Our current distinction is that a settlement must be, or become, largely self- concerns including terrorism, hunger, disease, and problems of air sufficient since no terrestrial government is expected to take on a perm- quality, safe abundant water, poverty, and weather vagaries tend to anent and growing commitment to support it for the indefinite future. overshadow long-term activities such as Space Colonization in the minds of many. Our leading “think tanks” such as the Woodrow Notions, popular in some circles, that grants of property rights to Wilson International Center for Scholars and the Brookings Institute do corporations or other wealth entities would encourage them to develop not rate space travel high on lists of future beneficial undertakings even advanced space transport systems to support settlements, without the though many of the concerns listed above are prominently featured. It need for understanding their economic viability, are dismissed. is the contention of this paper that Space Colonization will lead toward solutions to many of the emerging problems of our Earth, both A space settlement is economically different from historical settlements technological and sociological. The breadth of the enterprise far on Earth. Humans can survive and procreate on Earth with only the exceeds the scope of our normal single-purpose missions and, most primitive technology; humans and prehumans obviously did for therefore, its benefits will be greater. some millions of years. Any other location in the solar system requires enough technology to produce and contain/control air to breathe, water to drink, and food to eat, as well as other necessities. The underlying question is whether such a level of technology can be reproduced, or even maintained, by the number of people it supports, or like the present-day space station, is a far larger number of people required?

Present and future space transportation costs, and projected transportation requirements for settlements on the Moon or Mars, are discussed. Simple examples of input/output analysis are used to estimate effects of representative levels of self-sufficiency. Technical approaches to reaching such levels are discussed. Needs for income to support settlements are described, along with discussion of possible avenues for reaching such levels of income.

Triggering Events for the First Space Settlement Prospects for Revolutionary Technology for Space Colonization Anita E. Gale1, 2a and Richard P. Edwards1, 2b Edward McCullough 1Co-Founders, Space Settlement Design Competitions, 18506 Upper Bay Rd., Nassau Bay, TX 77058 The Boeing Company, 5301 Bolsa Avenue HO13-C321Huntington 2a 2b Senior Project Engineer, and Principal Engineer, Space Shuttle Beach, CA 92647-2099 Program, The Boeing Company, Houston, TX (281)226-5691, [email protected] Abstract. During the last half of the Twentieth Century, a host of technologies and disciplines, which had witnessed millennia of slow or Abstract. We know where humankind is now in its limited ability to no growth, suddenly went exponential. Photography, chemistry and venture into space, and we can envision technologies that include quantum mechanics combined to produce a new industrial revolution. routine space flight and large human populations in space; the This nexus of technologies set loose the fields of electrical and challenge is to figure out how to get from where we are now to what we mechanical engineering on courses of what appears to be unbounded can envision. Although the technical challenges of space infrastructure improvement. Largely through the efforts of a brilliant xray development will be significant, the factors most responsible for crystallographer, the structure of deoxyribose nucleic acid (DNA), the preventing us from surmounting those challenges are politics and store house of genetic information, was discovered. Within decades, economics. Various rationales have been proposed by other authors progress was made in understanding the cellular processes of life and and are summarized, with assessments of the hurdles involved in each. the human genome was sequenced. The macro world was not spared In an effort to make Space Settlement Design Competitions for high these revolutions. The centuries old technology of printing was school students as realistic as possible, the co-authors developed a extended to 3 dimensions with inks of polymers, ceramics, wood and compelling rationale for building the first community in space and the metals. These and other technologies have affected other technologies infrastructure required to support it, which passes the tests of economic so that now at the dawn of the twenty first century, one technology necessity and political appeal. after another is assuming an exponential trajectory. This paper surveys some of these technologies and speculates on what they portend for Space Colonization.

Solar Cells Using Lunar Resources The Development of ISRU and ISSE Technologies Leveraging Canadian Mining Expertise Alex Freundlich, Charles Horton, Andenet Alemu, Carlos Gramajo, Lance Williams, Alex Ignatiev Dale S. Boucher1, Jim Richard2, Erick Dupuis3

Texas Center for Superconductivity and Advanced Materials, 1Northern Centre for Advanced Technology Inc. 1400 Barrydowne University of Houston, Houston, Texas 77204-5002 Road, Sudbury, Ontario, Canada, P3A 3V8 713-743-3621, [email protected] 2Electric Vechile Controllers Ltd. 2200 Valleyview Road, Hammer, Ontario, Canada, P3N ILI Abstract. Silicon solar cells have been thoroughly optimized for 3Canadian Space Agency, 6767 route de l’aéroport, St-Hubert, Québec, terrestrial conditions. Solar cells fabricated on the Moon using lunar Canada, J3Y 8Y9 resources require different optimizing conditions and design. The (705) 521-8324 x202, [email protected] purity of the initial source material will be much less than the photovoltaic grade silicon used terrestrially for solar cells. In addition, Abstract. Future space missions to planetary bodies, both manned and the fabrication methods used will require novel substrates and non robotic, will require the efficient utilization of in-situ resources to standard deposition techniques. We report the results of our ensure longevity and success. In Situ Resources Utilization (ISRU) and calculations and modeling of silicon solar cells within the constraints of In Situ Support Equipment (ISSE), while requiring the development of the Lunar environment. The effects of decreased purity of the silicon new technologies and methods for commodity extraction, will still rely on the solar cell design will be presented. Cell designs to maximize upon some method of mining technology for the harvesting and efficiency and data on the fabrication of thin film Si cells on lunar transformation of the raw materials prior to processing. regolith substrates will be presented. The Northern Centre for Advanced Technologies Inc. (NORCAT), in partnership with Electric Vehicle Controllers Ltd. (EVC), is presently engaged in the development and adaptation of existing mining technologies and methodologies for use extra-terrestrially as precursor and enabling technologies for ISRU and for use as ISSE in support of longer term missions. More specifically, NORCAT and EVC, in partnership with MD Robotics and under contract to the Canadian Space Agency, are developing a drill and sample handler system for sub surface sampling of planetary bodies, specifically Mars.

The partnership brings to the table some formidable world leading expertise in space robotics coupled with world leading expertise in mining technologies. The system is an integrated package consisting of a drill, a sample manipulator, a sample triage capability, and a sample preparation station.

This paper will explore the results of that development and highlight other potential space exploration technologies based upon mining technology for use in ISRU and ISSE.

ISRU Reactant, Fuel Cell Based Power Plant for Optimized ISRU Propellants for Propulsion and Power Robotic and Human Mobile Exploration Applications Needs for Future Mars Exploration and Colonization

Russell S. Baird, Gerald Sanders, Thomas Simon, and Kerri McCurdy Eric Rice, Robert Gustafson, Daniel Gramer, Martin Chiaverini, Ron Teeter, and Brant White Energy Systems Division, NASA Johnson Space Center Houston, TX 77062, USA Orbital Technologies Corporation (ORBITECTM) (281) 483-9013, [email protected] Space Center, 1212 Fourier Drive Madison, WI 53551 Abstract. Three basic power generation system concepts are generally (608) 827-5000 considered for lander, rover, and Extra-Vehicular Activity (EVA) assistant [email protected] applications for robotic and human Moon and Mars exploration missions. The most common power system considered is the solar array and battery Abstract. In a recent feasibility study for the NASA Institute for system. While relatively simple and successful, solar array/battery systems Advanced Concepts (NIAC), ORBITEC conceptualized systems and an have some serious limitations for mobile applications. For typical rover evolving an optimized architecture for producing and utilizing Mars- applications, these limitations include relatively low total energy storage based in-situ resources utilization (ISRU) propellant combinations. capabilities, daylight only operating times (6 to 8 hours on Mars), relatively short operating lives depending on the operating environment, and The propellants will be used to support the propulsion and power rover/lander size and surface use constraints. Radioisotope power systems systems for ground and flight vehicles that would be part of future are being reconsidered for long-range science missions. Unfortunately, the Mars exploration and colonization. The key aspect of the study was to high cost, political controversy, and launch difficulties that are associated show the benefits of ISRU, develop an analysis methodology, as well with nuclear-based power systems suggests that the use of radioisotope as provide guidance to propellant system choices in the future based powered landers, rovers, and EVA assistants will be limited. The third upon what is known today about Mars. The study time frame includes power system concept now being considered are fuel cell based systems. the early unmanned and manned exploration period (through 2040) and Fuel cell power systems overcome many of the performance and surface a colonization period that is postulated to occur from 2040 to 2090. As exploration limitations of solar array/battery power systems and the part of this feasibility study, ORBITEC developed two different Mars prohibitive cost and other difficulties associated with nuclear power colonization scenarios, a low case that ends with a 100-person colony systems for mobile applications. In an effort to better understand the (an Antarctica analogy) and a high case that ends with a 10,000-person capabilities and limitations of fuel cell power systems for Moon and Mars colony (a Mars terraforming scenario). A population growth model, exploration applications, NASA is investigating the use of In-Situ mission traffic model, and infrastructure model was developed for each Resource Utilization (ISRU) produced reactant, fuel cell based power scenario to better understand the requirements of future Mars colonies. plants to power robotic outpost rovers, science equipment, and future This paper will present the overall results of the study. ISRU proved to human spacecraft, surface-excursion rovers, and EVA assistant rovers. be a key enabler for these colonization missions and the ISRU This paper will briefly compare the capabilities and limitations of fuel cell propellant, carbon monoxide and oxygen, proved to be the most cost- power systems relative to solar array/battery and nuclear systems, discuss effective propellant combination. the unique and enhanced missions that fuel cell power systems enable, and discuss the common technology and system attributes possible for robotic and human exploration to maximize scientific return and minimize cost and risk to both. Progress made to date at the Johnson Space Center on an ISRU producible reactant, Proton Exchange Membrane (PEM) fuel cell based power plant project for use in the first demonstration of this concept in conjunction with rover applications will be presented in detail.

Tailored Force Fields for Space-Based Construction Distributed Power Sources for Mars Colonization

George H. Miley and Yasser Shaban Narayanan M. Komerath, Sameh S. Wanis, Joseph Czechowski Department of Nuclear, Plasma, and Radiological Engineering, Fusion School of Aerospace Engineering, Georgia Institute of Technology, Studies Laboratory, University of Illinois at Urbana-Champaign, Atlanta, Georgia 30332-0150, USA. Urbana, Illinois 61801 404-894-3017, [email protected] 217-333-3772, [email protected] Abstract. One of the fundamental needs for Mars colonization is an Abstract. In Space, minor forces exerted over long periods can abundant source of energy. The total energy system will probably use a produce major results. Force fields of various kinds can be used to mixture of sources based on solar energy, fuel cells, and nuclear build large structures, superseding the human-intensive construction energy. Here we concentrate on the possibility of developing a techniques of today. In this paper we consider how several techniques distributed system employing several unique new types of nuclear now used in diverse fields can be generalized and applied to Space- energy sources, specifically small fusion devices using inertial based construction. Radiation pressure exerted by coherent optical or electrostatic confinement [1, 2] for MW level fixed stations combined ultrasonic beams on scattering objects is today used in microscale with portable “battery type” proton reaction cells [3]. positioning. Standing-wave fields offer important advantages, with the radiation force in a standing wave field being as much as 3 orders of magnitude greater than that of the source. The strong analogy between optical/electromagnetic and acoustic radiation is used to extend a [1] G. H. Miley, “The Inertial Electrostatic Confinement Approach to microgravity flight result from acoustic standing wave fields to Fusion Power,” Current Trends in International Fusion Research, edited electromagnetic fields. Walls of complex shape can be formed by Emilio Panarella, Plenum Press, NY, . pp. 135-148 (1997). automatically. The power requirements for use of such fields over different parts of the spectrum are compared and considered. [2] George H. Miley, "The Inertial Electrostatic Confinement (IEC) as Generation of sufficient power to overcome gravitational jitter at the a Unique Space Power Source,” Specialist Workshop on Advanced solar orbit of earth appears to be feasible with various types of fields Space Propulsion, Joint Propulsion Laboratory, Pasadena, CA, 18-20 and particles. The interim architecture to bootstrap an economy which May 1994. will permit large-scale construction projects is briefly considered. [3] G. H. Miley, C. Castano, A. Lipson, S.-O. Kim, and N. Luo, “Progresss in Development of a Low Energy Reaction Cell for Distributed Power Applications,” Proceedings of ICONE 10: 10th International ASME Conference on Nuclear Engineering, Arlington, Virginia, Track 8, pp. 1-7, April (2002).

Architecture Studies for Commercial Production of Space Colonization Using Space-Elevators from Phobos Propellants From the Lunar Pole Leonard M. Weinstein

Michael B. Duke1, Javier Diaz1, Brad R. Blair1, Mark Oderman2, and Advanced Measurement and Diagnostics Branch, NASA Langley Marc Vaucher2 Research Center, Hampton, VA 23681, USA (757)864-5543, FAX: (757)864-8315, [email protected]

1 Center for Commercial Applications of Combustion in Space (CCACS), Colorado School of Mines, CO 81401 Abstract. A novel approach is examined for creating an industrial 2CSP Associates, Inc, Cambridge, MA 02139 civilization beyond Earth. The approach would take advantage of the Phone: 303-670-2763; email: [email protected] unique configuration of Mars and its moon Phobos to make a

transportation system capable of raising mass from the surface of Mars Abstract. Two architectures are developed that could be used to to space at a low cost. Mars would be used as the primary location for convert water held in regolith deposits within permanently shadowed support personnel and infrastructure. Phobos would be used as a source lunar craters into propellant for use in near-Earth space. In particular, of raw materials for space-based activity, and as an anchor for tethered the model has been applied to an analysis of the commercial feasibility carbon-nanotube-based space-elevators. One space-elevator would of using lunar derived propellant to convey payloads from low Earth terminate at the upper edge of Mars' atmosphere. The tip velocity orbit to geosynchronous Earth orbit. Production and transportation relative to the ground would only be 0.52 km/sec. Small craft would be system masses were estimated for each architecture and cost analysis launched from Mars' surface to rendezvous with the moving elevator was made using the NAFCOM cost model. Data from the cost model tip, and their payloads detached and raised to Phobos. The carrier were analyzed using a financial analysis tool reported in a companion vehicle would then detach and land for reuse. Another space-elevator paper (Lamassoure et al., 2002) to determine under what conditions the would be extended a comparable distance outward from Phobos to architectures might be commercially viable. Analysis of the launch craft toward the Earth/Moon system or the asteroid belt. Release architectural assumptions is used to identify the principal areas for from the outward elevator tip at 3.52 km/sec would give a hyperbolic further research, which include technological development of lunar velocity of about 2.6 km/sec, which is the Hohmann elliptical transfer mining and water extraction systems, power systems, reusable space velocity needed to reach the Earth/Moon system and also the transfer transportation systems, and orbital propellant depots. The architectures velocity to reach the inner portion of the asteroid belt. This velocity and commercial viability are sensitive to the assumed concentration of boost would greatly reduce total propellant needs for space ice in the lunar deposits, suggesting that further lunar exploration to transportation. The outward elevator tip could also be used to catch determine whether higher-grade deposits exist would be economically arriving craft. These space-elevators would allow low cost movement justified. of people and supplies from Mars to Phobos and from Phobos to interplanetary space. In addition, large quantities of material obtained from Phobos could be used to construct space habitats and also supply propellant and material for space industry in the Earth/Moon system as well as on and around Mars.

Business Approach To Lunar Base Activation Space Base: Design Problem Identification December 1979 Harrison H. Schmitt Laurie Barlow, AIA University of Wisconsin-Madison, P.O. Box 90730, Albuquerque, NM 87199 L.Barlow & Company, Architecture Planning & Design 505 823 2616, [email protected] 1441 Huntington Drive South Pasadena, CA. 91030 626-286-3255 Abstract. It remains unlikely that any government or group of governments will make the long-term funding commitments necessary Abstract. This study concerns the development of a space base to to return to the Moon in support of scientific goals or resource house 100 people in low-earth orbit. The primary thrust of the study is production. If a lunar base is to be established within the foreseeable the application of architectural design methods to solve the problem of future, it will support of commercial production and use of unique form & variety in a totally self-contained artificial environment. The energy resources Business plan development for commercial uniqueness of this problem should provide new options for design production and use of lunar Helium-3 requires a number of major steps, solutions. including identification of the required investor base and development of fusion power technology through a series of business bridges that The concept development consists of exploring the background issues provide required rates of return of space industrialization. General design criteria are subsequently developed from these issues. The central concept is developed for . solving the gravity issue for large groups of people for 65% of their time without building giant infrastructure. It employs the "cocooning" idea for relief from the "space" environmental experience, which could become overwhelming over time.

It employs a method of providing artificial gravity in living and recreation areas only that do not require rotation of large areas of the physical structure, and this gravity provision can be expanded in modules. This leaves the majority of the structure free for functional zero-gravity production and assembly, as well as research and astronomical observation.

This establishes the segregation between gravity habitat and 0-g habitat in a way that minimizes the necessary rotation mass, and hence the cost, of constructing a space base. It utilizes each gravity environment to its fullest advantage, both from a human factors standpoint and from a utilitarian standpoint. It allows for a means of growing the structure in an incremental way as needed for gravity and non-gravity environments.

AQUAPLEX An Interplanetary Rapid Transit System An environmentally aware model Lunar Settlement between Earth and Mars

Darel W. Preble1 Kerry Nock1, Michael Duke2, Robert King3, Mark Jacobs4, Lee Johnson3, Angus McRonald1, Paul Penzo1, Jerry Rauwolf4, and Chris 1Space Solar Power Institute, Space Solar Power Workshop, 2557 Betty Wyszkowski1 Jean Drive, Jonesboro, GA 30236 770.477.9143; [email protected] 1Global Aerospace Corporation, 711 West Woodbury Road, Suite H, Altadena, CA 91001 Abstract. The construction and operation of a replica Lunar settlement, 2Center for Commercial Applications of Combustion in Space, including CELSS, on earth can provide many lessons in in-situ Colorado School of Mines, Golden, CO 80401 resource utilization, telerobotic operation, and a wealth of other 3Engineering Division, Colorado School of Mines, 1500 Illinois St., essential insights into key Lunar operational principles while building Golden, CO 80401 important bridges to earthly industries businesses and other interests. 4Science Applications International Corporation, 1501 Woodfield One among these is the reduction of the water demanded by traditional Road, Suite 202N, Schaumburg, IL 60173 models of lunar operation. By severely reducing water requirements, 626-345-1200, [email protected] not including drinking water, of course, a larger settlement may be operated with the same amount of water. Hypes and Hall and all other Abstract. A revolutionary interplanetary rapid transit concept for examples found in the open CELSS literature propose quantities of transporting scientists and explorers between Earth and Mars is hygiene water far in excess of what would be needed in actual presented by Global Aerospace Corporation under funding from the operation using simple technologies now readily available within NASA Institute for Advanced Concepts (NIAC) with support from the environmentally aware communities. A careful study will show that Colorado School of Mines (CSM), Science Applications International using zero water toilets, low water showers, CO2 dry cleaning Corporation (SAIC), and others. We describe an innovative architecture machines, low water use, energy efficient washing machines and other that uses highly autonomous, solar-powered, xenon ion-propelled new environmentally aware standard hardware and processes, hygiene spaceships, dubbed Astrotels; small Taxis for trips between Astrotels water use can be slashed and precious process water can be made and planetary Spaceports; Shuttles that transport crews to and from available for more critical needs. Equally important, through involving orbital space stations and planetary surfaces; and low-thrust cargo the environmental community in this fascinating exercise, bridges can freighters that deliver hardware, fuels and consumables to Astrotels and be built to the environmental community which has heretofore not been Spaceports. Astrotels can orbit the Sun in cyclic orbits between Earth strong supporters of the space research and development community. and Mars and Taxis fly hyperbolic planetary trajectories between The Space Solar Power Workshop sees great opportunity in the near Astrotel and Spaceport rendezvous. Together these vehicles transport term to advance the prospects for Lunar settlement through this replacement crews of 10 people on frequent, short trips between Earth initiative. and Mars. Two crews work on Mars with alternating periods of duty, each spending about 4 years there with crew transfers occurring about every two years. We also discuss the production of rocket fuels using materials mined from the surfaces of the Moon, Mars and the Martian satellites; the use of aerocapture to slow Taxis at the planets; and finally the life-cycle cost estimation.

A.A.M. Delil ...... 1 Borowski, S...... 89 Chen, G...... 132 Easter, R. W...... 240 Gertsch, R. E. ... 242 Hopkins, J. B...... 74 Abe, Y...... 52 Borowski, S. K. Chernyshova, M. Edwards, B...... 184 Gerwin, R. A. ... 124 Horton, C...... 241, 255 Alemu, A...... 241, 255 .. 190, 192, 193, 195, 196 A ...... 47 Edwards, J. T...... 67 Ghamaty, S...... 117 Houts, M...... 125, 150, Allen, D. T...... 117 Boston, P. J...... 238 178 Cheung, K...... 49 Edwards, R. P. 248, Gill, P. S...... 83 Anderson, B. M. Boucher, D...... 256 253 Howard, R. T...... 78 .. 258 ...... 152, 192 ...... 211, 214, 216 Chiaverini, M. Gilland, J. Bragg-Sitton, S. El-Genk, M. S.118, Howe, S. D ...... 121 ...... 225 ..... 190 Anderson, J. R. 123 ...... 150 Clapp, M. B. 138, 139, 144, 166, 185 Gilland, J. H. Howell, E. I...... 108 Angelis, G. D. .. 215 Bragg-Sitton, S. Clark III, B. C.. 237 Elliot, K...... 236 Glazer, S...... 37 Hrbud, I...... 150, 178 Angelis, G. De.217 M...... 127 Clowdsley, M. S. Elliott, J. O. 116, 160, Gleicher, F...... 213 ...... 215, 221 Hugger, C P...... 218 163 Godfory, T...... 112 Anghaie, S...... 103 Bray, D...... 40 Clowdsley, M. S. Hulsey III, W. N. Elsner, N. B...... 117 .. 125, 150 Angirasa, D...... 182 Brinker, C.J...... 201 ...... 217 Godfroy, T...... 59 Brophy, J...... 17, 177 Emrich, Jr., W. J. ... 98, Godfroy, T. J. ... 127 ...... 140 Augenblick, J. E. Coleman, J. R...... 110 Hunt, T. K...... 146, 147 Brophy, J. R...... 123 99 Goehlich, R.A.. 223 ...... 107 Brown, E. J...... 95 Engle, M...... 197 Hunt, T. K. 214 ...... 109 Aumann, A. R. Brown, E. J...... 93 Coreano, L. Goodfellow, K.178 Erickson, G. F. 230 Hunter, L. W...... 97 Backhaus, S...... 158, Brown, M...... 48 Corley, M...... 188 Graham, J. M... 235, ...... 213 Iacona, E...... 28 169 Failla, G. 236 Bruckner, A. P.244 Crowley, C. J.... 157 Badavi, F. F...... 221 Falck, R. D...... 193 Ibrahim, M. B.. 180 Cuccinotta, F. A. Gramajo, C. 241, 255 Brunetti, B...... 246 Ignatiev, A... 241, 255 Badavi, F. 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P... 87 Friedensen, V. P...... 123 Bevard, B. B...... 141 Diaz, J...... 240 ...... 100 Halsey, D...... 142 Katz, I. Case, J. A...... 245 Bhattacharyya, S. G. H. Miley ...... 85, 260 Kawamura, H..... 52 Cassanova, R. A. Dickens, R... 125, 150 Hamdan, M...... 39 G.DeAngelis...... 221 ...... 51 K...... 101 ...... 247 Dilg, J...... 148 Hanrahan Jr, R. J. Kim, J. Gale, A. E..... 248, 253 ...... 102 Bica, D...... 53 Cassenti, B. N.. 109 Donahue, B...... 153 Kim, M...... 219 Ganapathi, G.B.36 Harper, R...... 150 Bienhoff, D...... 73 Chakrabarti, S.... 86 Douglas, D...... 42, 45 King, R...... 266 Gedeon, D...... 180 Henderson, H. T. Birchenough, A. 45 Kislov, N...... 44 Champion, J. L. Dudzinski, L. A...... 39 ...... 168 ...... 88, 191 Geffre, J. R...... 216 Kisor, A. K...... 106 Chandrasekhar, ...... 51 35, 36, 42 George H. Miley ...... 260 Henry, C. Birur, G. C. Duke, M...... 266 Kittredge, S...... 77 P...... 42 ...... 119 ...... 28 ...... 43 George, J. Herman,C. Biter, W. Duke, M. 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Kramer, D. P..... 108 Mazuruk, K...... 207 Novak, K. S...... 35 Preble, D. W...... 265 Schreiber, J. G. Strickland, C. .... 189 ...... 159, 170 Kreisel, J...... 58 McCullough, E. Nowicki, A.W. 167 Prinja, A...... 213 Sudante, P.J.V. 234 ...... 254 Schuiling, R. T. 68 Krizan, S...... 189 O’Donnell, D. J. Qiu, S...... 145, 146, 147, Sunada, E. T...... 35 McCurdy, K...... 257 ...... 229 171 Schuller, M...... 105 Krizan, S. A...... 216 Suzuki, K...... 52 McGhee, J...... 213 O’Handley, D. A. Qualls, G. D...... 218 Schwank, J. W.137 ...... 38 ...... 52 Kroliczek, E...... 231 Suzuki, M. McGrath, P. L.... 36 Qualls, G. D...... 211, Sefiane, K...... 41 Ku, J...... 46, 49 Swanson, T...... 42 McGuire, M. L. Oderman, M...... 261 213, 216 Sena, J. T...... 54 ...... 40 .. 45 Kurwitz, C...... 190, 192 Oderman, M...... 240 Rasmussen, J. R. Senior, C. L...... 245 Swanson, T. D...... 106, 140 Lalk, T. R...... 105 McQueeney, T. . 42 Oguz, H. N...... 97 Shaban, Y...... 85 Swift, W. L...... 157 Rauwolf, J ...... 266 Lamassoure, E. ... 266 ...... 43 ...... 260 Swint, S...... 80 McRonald, A. Oh, S. Rawlin, V...... 17, 177 Shaban, Y. S...... 240 Tadrist, L...... 41 Medis, P...... 39 Ohlandt, C...... 111 Redinger, D. L. Shah, A. R...... 159 Landis, G. A..... 200, Mikellides, I. G. Oleson, S...... 184 ...... 146, 147 Shaltens, R. K.. 182 Takahash, S...... 52 249 ...... 124 Osiander, R...... 45 Regan, T. F...... 181 Shavers, M. R.. 218 Tang, X...... 201 Landis, G. A...... 71 Mikellides, P. G. Ottenstein, L.. 46, 49 Reid, R. S...... 57 Shields, V..... 130, 131 Taylor, R...... 184 Lane, N...... 156 ...... 124, 179 Rice, E...... 233, 258 Teeter, R...... 258 . 85, 260 Packard, T. W.. 190 Shin, M...... 219 Langford, G...... 148 Miley, G. H. Rice, L. P...... 95 Paniagua, J...... 89, 91 Shumway, D.A. 30 Tew, R. C...... 180 LaPointe, M. R. Miller, R. G...... 108 Richard, J...... 256 ...... 155 Paniagua, J. C.... 90 Siegfried, W. H. Thibaud-Erkey, Moeller, R. C.... 162 .... 85 Richardson N...... 251 Lehman, D. H.. 115 Parrish, C. F...... 243 C...... 201 Mohanty, S...... 234 Riehl, R. R...... 32 Siergiej, R. R...... 92 Thibeault, S. A. Lewis, R...... 86 Mollere, J...... 204 Parrish, K...... 37 Ring, P. J...... 54 Simon, T...... 257 ...... 221 Lipinski, R. J. .. 116, Pastukhov, V. G. Robbins, Jr., W. Momota, H...... 85 Simon, T. M...... 239 ... 170 127, 160, 163, 186 ...... 47 Thieme, L. G. Momozaki, Y. .. 166 W...... 198 Simon, T. W...... 180 Loder III, T. C. 250 Patzold, J. D...... 36 Thomas, D...... 77 Moreland, D. W. Robertson, G. A. Sims, W. H...... 86 ...... 42 ..... 35 Thomas, L...... 77 Lovas, A...... 163 Pauken, M. T...... 209, 210 Luke, J. R...... 165 Singleterry Jr., R. Thomson, S...... 37 Morgan, J...... 102 Pautz, S...... 213 Robinson, C. P. MacDowell, L...... 133 C...... 213 Topper, W. F...... 93 Morgan, T. H.... 113 Peaden, C...... 79 G...... 70 Roe, F. D...... 78 Singleterry, R. C. Tournier, J. . 138, 139, ...... 86 Mosier, C. L...... 34 Pearson, J. B...... 215, 221 185 Maclellan, D...... 96 Rombaut, H.J. .. 234 Murad, P. A...... 205 Pelaccio, D. G. 114 Smith III, R. L. 102 Tournier, T. M.118 Maise, G...... 89, 90, 91 Rosmait, R. L... 128 Murray, C. S...... 92 Penzo, P...... 266 Smith, C...... 77 Travis, B...... 121 Majumdar, A. ... 188 Ross, D. A...... 42 Murray, S. L...... 92 Perry, R. B...... 114 Smith, G...... 142 Trionfetti, G...... 246 Maker, D...... 209 Roth, M. E...... 181 N. Richardson...... 85 Petach, M...... 158 Smitherman, Jr., Tripathi, R. K.. 212, Mantovani, J. G. Roy, S...... 176 215 Nakagawa, R. Y...... 42 ...... 167 Peters, W. D. V...... 232 ...... 163 ...... 204 ...... 216 Rueda, A. Snyder, J...... 129 Troutman, P. Manvi, R...... 189, 240 Petersen, S. L. .. 146 Nakamura, T. .... 245 Russell, P...... 247 Troutman, P. A. Peterson, A. A.146, Sobieszczanski- Margasahayam, Nalette, T...... 201 ...... 189 147, 171 Sadeh, E...... 227 Sobieski, J. .. 213 R. N...... 69 Nealy, J. E. . 211, 215, Sakamoto, J...... 129, 130 Tsai, C. A...... 201 Phillips, C. J...... 35 Spence, C. A. .... 105 Maria, G. De. .... 246 216 Salvail, P...... 125, 128, Tsuyoshi ...... 177 Polk, J. 150 Spilker, T. R...... 163 Marinicã, O...... 53 Nealy, J. E...... 217 Saotome...... 64 Polk, J. E...... 123 Salvail, P. G...... 55 Stahlhofen, A. A...... 38 Neill, D. S...... 87 .... 36 Marland, B...... 29 Tsuyuki, G. T. Ponappan. R...... 42 Sanders, G...... 257 Martens, C. L. .. 226 Newman, F. D... 92 Tuchscherer, M. Ponnappan, R..... 44 Sanders, G. B...... 239 Stebe, K.J...... 50 Nguyen, D...... 188 ...... 131 Martin, J... 86, 112, 150 Sandridge, C. A. Steeve, B...... 188 Nichols,G. J...... 93 Pop, G...... 53 ...... 247 ...... 214 Turner, R. Martin, J. J...... 125 Stewart, E...... 148 Nickisch, L. J. .. 204 Poston, D. I...... 116, Sayre, E. D...... 54 Tward, E...... 158 Mason, L...... 148, 168 149, 151, 160, 163 Stillwagen, F..... 189 Niebur, S. M...... 113 Schmitt, H. H. . 224, Van Dyke, M... 150, Mason, L. M..... 190, Powell, J...... 89, 91 263 ...... 161 Stoian, F. D...... 53 178 192 Nilsen, E. N. Powell, J. R...... 90 Schneider, M. A...... 162 Stoica, V...... 53 Van Dyke, M. K. Mason, L. S...... 182 Noca, M. A...... 244 Powell, K. G...... 111 ...... 127 Nock, K...... 266 Stouffer, C...... 38 Maydanik, Y. F.47 Schneider, W...... 45 Van Dyke, V..... 188

Vasiliev L. L...... 31 White, M. D...... 202 Woodward, J. F...... 206 Vasiliev, L...... 31 Whitlow, J. E.... 243 Wright, S. A. .... 127, 240, 261 62 Vaucher, M. Wichman, H. A. 186 Vaughan, W. W. Wiffen, F.W...... 164 Wyszkowski, C...... 83 Wilkes, K. E...... 108 ...... 266 Vékás, L...... 53 Williams, C. H. Xomerita, G...... 201 Verhage, M ...... 77 ...... 124 Y. Shaban ...... 85 Yoder, G. L. Verlaat, B...... 33 Williams, E...... 150 ...... 141 Williams, G...... 17, 177 Vincent, G...... 199 Youngquist, R. Williams, L. 241, 255 Vinje, R. D...... 70 C...... 82 Williams, R...... 40 Volp, J...... 234 Zagarola, M. V. Williams, R. B.183 ...... 157 Vu, B. T...... 69 Williams, R. M. Zampino, E. Z. 159 Walts, S. C...... 97 ...... 106 Zapp, N...... 218 Wanis, S. S...... 259 Willis, B. P...... 136 Zay, B. J...... 42 Wareing, T. A.. 213 Wilson, J. W..... 211, Zerkle, D. K...... 121 Watson, P. M.... 197 212, 213, 215, 217, 218, 221 Ziegler, S. W...... 87 Webber, D...... 61 Woering, A. A... 33 Zinkle, S. J...... 164 Weinstein, L. M...... 136 ...... 131 ...... 262 Woike, M. R. Zoltan, A. Weitzberg, A. ... 187 Wood, J. G...... 156 Zoltan, L...... 131 Wernsman, B...... 92 Woodcock, G. .. 252 White, B...... 233, 258 Woods, R. C...... 203