Full Moon Final Report V4 0

Full Moon Storage & Delivery of Oxygen & Hydrogen for Lunar Exploration Team Project International Space University Masters Program 2006/2007

© International Space University. All Rights Reserved.

The 2006-2007 Masters Program of the International Space University was hosted at the International Space University Central Campus in Strasbourg, France. The Full Moon ‘lunar gas station cover’ was graciously developed by William Widjaja, and drawn by Alberto Martín Montalbetti. Its purpose is to visualize the Full Moon proposal, and in so doing, inspire the next generation of lunar explorers as to the futures potential. The Full Moon logo was realized by William Widjaja with the input of the team. Please note that while all efforts have been made to ensure accuracy and veracity in this report, ISU does not take any responsibility for the accuracy of its content.

International Space University Strasbourg Central Campus Attention: Publications/Library Parc d’Innovation 1 rue Jean-Dominique Cassini 67400 Illkirch-Graffenstaden France Tel. +33 (0)3 88 65 54 32 Fax. +33 (0)3 88 65 54 47 e-mail. [email protected]

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____________________________Acknowledgements

The members of the Masters’ 2006 Team Project-2 would like to acknowledge Walter Peeters, Nikolai Tolyarenko, John Farrow and Vasilis Zervos of the ISU Faculty for their advice, support and guidance along the way. A special thanks is given to Hugh Hill and Bijal Thakore, the official TP-2 faculty liaisons, for their invaluable input throughout the process. We would also like to take this opportunity to express our gratitude to the following individuals for their valuable input throughout the development of our project: Philippe Achilleas John Blake Arthur Guest Robert Guinness David Gump Ozgur Gurtuna Heinz-Hermann Koelle Bill Larson René Laufer Isabelle Scholl Robert Shishko Gerald Sanders Andrew Tinka

Institut du Droit de l’Espace et des Télécommunications Canadian Operational Research Society International Space University Alumni International Space University Alumni Transform Space Turquoise Solutions (International Space University Alumni) Technische Universität Berlin NASA/Kennedy Space Center University of Stuttgart/Institute of Space Systems International Space University Faculty NASA/Jet Propulsion Laboratory NASA/Johnson Space Center University of California Berkeley

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_______________________________________Authors

Carla Adriana Arregoitia, Mexico/Canada B.Eng.Sci. Chemical/Biochemical Engineering University of Western Ontario

Oluwaseun Bankole, Nigeria B.Tech. Electronic and Electrical Engineering Ladoke Akintola University of Technology

Renée Boileau, Canada B.Ap.Sc. Engineering Physics University of British Columbia

Thomas Bouvet, France Ph.D. Atmospheric Science University of Alberta Rodolphe De Rosée, Belgium/USA M.Eng. Aeronautical Engineering Imperial College London

Israel Ojeda Coronado, Mexico B.Sc.Computer Systems Universidad Autonoma del Estado de Hidalgo Dag Evensberget, Norway M.Sc. Industrial Mathematics Norwegian University of Science and Technology

Alexandre Fréchette, Canada B.Eng. Mechanical Engineering Ecole Polytechnique de Montreal

Hubert Foy Kum, Cameroon Physics and Computer Science University of Buea

Pierre Ghelardi, France Master of Economics Sciences Po University of Strasbourg David T. Haslam, United Kingdom M.Sc. Laser Photonics and Modern Optics The University of Manchester

Kieran Griffith, Republic of Ireland B.Sc. Aerospace Studies Embry-Riddle Aeronautical University Jean-Sébastien Hutt, France Commercial Engineering Ecole Supérieure des Technologies et des Affaires

Kenneth Izomoh, Nigeria M.Eng. Electronic/Telecommunications Engineering University of Port Harcourt

Henrik Karlsson, Sweden Space Engineering Umea University

Alma Krivdić, Bosnia and Herzegovina B.Sc. Applied Mathematics and Computer Science Eastern Mediterranean University

Violetta Kuvaeva, Russian Federation External Economic Activity, Moscow Aviation Institute State University of Aerospace Technologies

Dimitrios Lamprou, Greece Dipl.Ing. Mechanical and Aerosnautical Engineering University of Patras

Scott MacPhee, Canada B.Sc. Major in Physics Dalhousie University

Laure-Hélène Milhome, France Master of Economics Sciences Po University of Strasbourg

Olawale Onifade, Nigeria M.Sc. Computer Science University of Lagos

Thiago Palmieri, Brazil B.Eng. Aerospace Engineering Ryerson University

Shawna Pandya, Canada B.Sc. Honors Neuroscience University of Alberta

Andreea Lavinia Radulescu, Romania B.Sc. Mathematics Trent University

Brian Schoening, USA B.Sc. Aerospace Engineering Purdue University/Northrop Grumman

Karanjeet Singh, Canada B.Eng. Aerospace Engineering Ryerson University

Carlos Uriarte Vega, Spain B. Law and Diploma in Economic Law Universidad de Navarra

William Widjaja, Indonesia B.S. Asia Pacific Management Ritsumeikan Asia Pacific University

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Jun Zhang, People’s Republic of China Space Engineering China Aerospace Science and Technology Corporation

Masters’ 2007

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_______________________________________Abstract

Lunar exploration today is very different from what it once was: previously driven by national glory and political prestige, it has now become a quest for innovation, discovery and scientific return. This quest is defined by a spirit of international cooperation and partnership. The key question driving lunar exploration has changed, too: we no longer ask “How do we get there?” but “How do we stay there?”. Oxygen and hydrogen directly address this question, since these resources are essential as propellants for transportation and as elements for life support systems. Many recent studies have looked into In situ Resource Utilization (ISRU) for extracting these materials from the lunar environment in order to ‘live off the land.’ However, there are missing links in the ISRU chain: while production and extraction methods are well-researched, it is still not clear as to how oxygen and hydrogen can be made easily accessible to the user. The present study addresses these gaps by examining the concept of a ‘lunar gas station,’ beginning with an analysis of the markets, customers and drivers for such a concept. This is followed by a technical assessment of possible storage and delivery options. System selection is performed based on existing concepts for storage and delivery from the production site to the end user. A storage and delivery architecture, although preliminary, is presented based on this assessment, and is one of the first attempts to address this gap in the oxygen and hydrogen supply chain. A subsequent business, legal and ethical analysis evaluates the feasibility of a lunar gas station, suggesting that the Full Moon lunar concept is best realized in conjunction with the burgeoning private space sector.

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________________________________Faculty Preface

Henri Rousseau (1844 –1910) is regarded as one of the greatest PostImpressionist painters. His naïve paintings with brightly-colored wild animals, birds, and jungle vegetation are considered to be some of the most innovative artistic works of the 19th Century. However, despite his fame today, he was rejected by the Parisian artistic establishment during his life. They regarded this self-taught customs officer as a worthless amateur. In fact, following his death, the Petit Journal declared that “…Rousseau’s fame was born of the praise of few and the scorn of many”. It seems that original and creative individuals have frequently been given little encouragement through the ages. Fortunately, there are many contemporary societies, institutions, and places of learning where imagination and creativity are valued and nurtured. The International Space University (ISU) is one such body. Examples of innovative work at ISU are the annual team projects (TPs) associated with the twelve-month Master of Science (M.Sc.) programs in ‘Space Studies’ and ‘Space Management’. Approximately 25% of students’ time during the programs is devoted to this interdisciplinary endeavor. This year, our students had a choice between two projects: “Space techniques supporting archaeology and heritage preservation” (TP1) and “Helium-3 and other planetary resources: support for humanity or folly” (TP2). In the early stages of the Masters program, both teams undertook an extensive literature review, which they formally presented to ISU’s faculty and invited guests in December 2006. With the knowledge acquired from this research, both teams then focused on a very specific aspect of their chosen topic. In the case of TP2, they decided to abandon advanced studies of Helium-3, in favor other lunar resources. Their study, “Full Moon: Storage & Delivery of Oxygen and Hydrogen for Lunar Exploration,” is summarized in this Report and the related Executive Summary. The study considers how lunar-derived oxygen and hydrogen can be stored and delivered on the surface to support future space exploration and utilization. For those readers unfamiliar with ISU TPs, you will note that this Report is faithful to our university’s celebrated ‘3I’ credo: international, intercultural, interdisciplinary. In fact, 29 men and women from 19 different countries—with diverse academic backgrounds—have contributed to this work. It only remains for me to applaud the team, especially those who have been ‘burning the midnight oil’ in recent weeks! Associate Professor Hugh Hill, on behalf of the Resident Faculty.

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_______________________________Student Preface

The 11th ISU Annual International Symposium, held in February 2007, bore the title “Why the Moon?” It was attended by more than 150 space experts from around the world, and being able to participate as full delegates was a wonderful experience for us, the ISU students. For all attendees of this conference it became blindingly obvious that the Moon is once again the focal point of the efforts of space agencies around the world. The Moon is seen both as a stepping stone towards the Red Planet and as a goal in itself. In both cases, the Moon will teach us valuable lessons about surviving in extreme conditions and controlling our environment. In situ resource utilization (ISRU), living of the land, was one of the central topics of the Symposium. Stimulating conversation with scientists and space industry professionals brought up some topics related to ISRU that are not so well covered by current research. This perceived challenge lead us, to choose “designing a lunar gas station” as our project of the year. The project finally found itself titled Full Moon: Storage & Delivery of Oxygen and Hydrogen for Lunar Exploration. The scope of our project reflect the interdisciplinary nature of ISU, covering space science, space engineering, systems engineering, space policy and law, business and management, and space and society. The Full Moon report, executive summary and presentation are the product of 8 weeks of work by our group of 29 students from 19 different countries. The broad scope and interdisciplinary nature of our project makes it quite unique within the extended ISRU literature. It is our belief that the solutions we propose in Full Moon will live on beyond the report and presentation, and form a basis for further research by individual team members, and a reliable reference for students and researchers wanting to keep exploring this exciting area of lunar utilization.

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______________________________Table of Contents

1 INTRODUCTION.......................................................................................... ............20 2 DRIVERS & CONSTRAINTS.................................................................. ................24 2.1 MARKET ANALYSIS FOR REFUELING.......................................................................... 24 2.2 DEMAND GRAPHS...................................................................................................39 2.3 CONSTRAINT FROM COMPETITORS..............................................................................41 2.4 CONSTRAINTS OF LUNAR TOPOGRAPHY...................................................................... 42 2.5 CONSTRAINTS OF THE LUNAR ENVIRONMENT.............................................................. 43 2.6 CONSTRAINTS FROM RESOURCE AVAILABILITY.............................................................44 2.7 CONSTRAINTS OF PRODUCTION..................................................................................47 2.8 FINDINGS: PRODUCTION SCENARIOS...........................................................................51 3 ARCHITECTURE ASSESSMENT.................................................................... .......53 3.1 CHALLENGES OF LUNAR ENVIRONMENT..................................................................... 53 3.2 SYSTEM DESIGN METHOD........................................................................................55 3.3 ASSESSMENT OF STORAGE OPTIONS ON THE LUNAR SURFACE........................................57 3.4 ALTERNATIVE STORAGE POSSIBILITIES........................................................................66 3.5 STORAGE VESSELS..................................................................................................67 3.6 INTERFACING .........................................................................................................67 3.7 ENSURING HUMAN SAFETY AND TANK HEALTH..........................................................69 3.8 TRANSPORTATION OPTIONS........................................................................................70 3.9 DELIVERY SYSTEM EVALUATION............................................................................... 75 3.10 QUALITATIVE DECISION RANKINGS FOR DELIVERY SYSTEMS....................................... 75 3.11 QUANTITATIVE DECISION RANKINGS FOR DELIVERY SYSTEMS.............................. .......78 3.12 FINDINGS: STORAGE AND DELIVERY CONCEPT RANKINGS...........................................83 4 SYSTEM ARCHITECTURE............................................................ ........................85 4.1 THE PROPOSED ARCHITECTURE.................................................................................88 4.2 INTERFACES............................................................................................................96 4.3 OPERATIONS...........................................................................................................99 4.4 THE SYSTEM BLUEPRINTS......................................................................................104 4.5 TECHNICAL RISKS.................................................................................................108 4.6 ADAPTATIONS FOR THE OPTIMISTIC SCENARIO........................................................... 109 5 THE BUSINESS ANALYSIS........................................................... ........................111 5.1 APPROACH AND OVERVIEW.....................................................................................111 5.2 SUPPLY OVERVIEW................................................................................................114 5.3 BUSINESS SOLUTIONS............................................................................................ 115 5.4 BUSINESS RISK ASSESSMENT..................................................................................118 5.5 COST BREAKDOWN ANALYSIS ................................................................................119 5.6 FINANCIAL MODEL .............................................................................................. 124 5.7 SENSITIVITY ANALYSIS...........................................................................................132 5.8 PROMOTION..........................................................................................................134 5.9 RECOMMENDATIONS...............................................................................................135 6 LEGAL & ETHICAL ISSUES............................................................ ....................137

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6.1 OVERVIEW OF THE EXISTING LEGAL FRAMEWORK: THE OUTER SPACE TREATY & MOON AGREEMENT.....................................................................................................138 6.2 LEGAL BARRIERS: A FURTHER ANALYSIS OF NON-APPROPRIATION PRINCIPLES..............140 6.3 ENABLING PRIVATE ENTERPRISE..............................................................................142 6.4 THE PROGRESSION TOWARDS A NEW INTERNATIONAL REGIME ............................ .......143 6.5 ADDITIONAL LEGAL CONSIDERATIONS: INSURANCE, RESPONSIBILITY & LIABILITY.........145 6.6 FULL MOON & ETHICAL CONCERNS.......................................................................148 6.7 CONCLUSIONS.......................................................................................................149 7 CONCLUSIONS ................................................................................................... ...151 8 REFERENCES .......................................................................................... ..............156 A. TANK MATERIAL SELECTION....................................................................... ..165 8.1 EXTRA VEHICULAR ACTIVITIES...............................................................................165 8.2 LSAM ASCENT STAGE MASS BREAK-DOWN...............................................................166 8.3 EXTRA VEHICULAR ACTIVITIES’ CONSUMPTION............................................................166 8.4 CALCULATIONS FOR 3, 6, 9 AND 12 DAYS-MISSION.................................................... 167 8.5 WATER CONSUMPTION:..........................................................................................168 B. LUNAR HYDROGEN....................................................................................... .....169 8.6 RATIONALE FOR SOUTH POLE BASE..........................................................................170 8.7 HYDROGEN EXTRACTION........................................................................................170 C. TANK MATERIAL SELECTION....................................................................... ..172 D. SYSTEM SELECTION DETAILS...................................................................... ..174 8.8 QUALITATIVE DECISION TOOL CALCULATIONS...........................................................174 8.9 QUANTITATIVE DECISION TOOL CALCULATIONS.........................................................175 8.10 SYSTEM OPTIMIZATION METHODS.........................................................................175 8.11 SUPPLY CHAIN MODELING WITH SPACENET............................................................176 E. DELIVERY SYSTEM CALCULATIONS................................................... .........179 8.12 LSAM MODIFICATIONS FOR SURFACE DELIVERY....................................................179 8.13 M-LSAM CALCULATIONS FOR MID-LATITUDE DELIVERIES.....................................179 8.14 WHEELED ROVER CALCULATIONS......................................................................... 180

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_______________________________Index of Figures

FIGURE 1-1: THE FULL MOON STUDY, ITS TECHNICAL COMPONENTS AND ITS INTERFACES WITH THE PRODUCTION FACILITY AND USERS/CUSTOMERS............................................................................................ .......21 FIGURE 2-2: GLOBAL ROADMAP SPACE EXPLORATION................................29 FIGURE 2-3: NEAR AND FAR SIDE LUNAR TOPOGRAPHY..............................30 FIGURE 2-4: ILLUSTRATION OF SHACKLETON CRATER MOON BASE (ESAS 2005)...................................................................................................... ..............31 FIGURE 2-5: NASA PAST AND FUTURE LANDING SITES (ESAS 2005)............31 FIGURE 2-6: SORTIE MISSIONS AT ARISTARCHUS PLATEAU (STEVENS 2007).............................................................................................................. ..................33 FIGURE 2-7: THE LSAM (ESAS 2005).................................................................. .....36 FIGURE 2-8: ANNUAL LOX DEMAND IN TONNES........................................... ....40 FIGURE 2-9: OPTIMISTIC LOX DEMAND IN TONNES......................... ..............40 FIGURE 2-10: BASELINE LH2 DEMAND IN TONNES.......................... ................41 FIGURE 2-11: OPTIMISTIC LH2 DEMAND IN TONNES.................... ..................41 FIGURE 2-12: NEAR AND FAR SIDE LUNAR TOPOGRAPHY............................42 FIGURE 2-13: EPITHERMAL NEUTRON FLUX AT LUNAR POLES .................46 FIGURE 2-14: COLORADO SCHOOL OF MINES EXCAVATOR PROTOTYPE (NASA 2007)........................................................................................... ........................48 FIGURE 3-15: SYSTEM SELECTION METHODOLOGY......................... .............56 FIGURE 3-16: STORAGE ELEMENT IN THE SYSTEM ARCHITECTURE.......58 FIGURE 3-17: TYPICAL LOX CRYOGENIC STORAGE ......................................65 FIGURE 3-18: TITAN EXPLORER THERMAL MODEL, DEEP SPACE TANKS, SUN SHIELD (PLACHTA 2005)....................................................... ...........................66 FIGURE 3-19: ELEMENTS OF THE REFUELING SERVICE.............................. ..70 FIGURE 3-20: MOBILE AND FIXED PLATFORMS.............................................. ..83 FIGURE 4-21: MOBILE STORAGE TANKS............................................ .................91 FIGURE 4-22: DELIVERY CONCEPT COMPARISON BY RANGE AND CAPACITY........................................................................................................... ..........92 © International Space University. All Rights Reserved.

FIGURE 4-23: TUG ROVER WITH TRAILER (BUFKIN ET AL. 1988)................93 FIGURE 4-24: MARS AIRBAG DESCENT BRAKE (STEIN AND SANDY 2003). 94 FIGURE 4-25: NASA LSAM BALLISTIC LANDER AND ATHLETE ROVER ....94 FIGURE 4-26: SUPPLY SYSTEM OVERVIEW......................................................... 96 FIGURE 4-27: INTERFACES BETWEEN PRODUCTION AND STORAGE ELEMENTS........................................................................................ ...........................96 FIGURE 4-28: ARTICULATED HOSE FOR FUEL TRANSFER........................ .....97 FIGURE 4-29: EMERGENCY DELIVERY STORYBOARD......................... .........104 FIGURE 5-30: BUSINESS ANALYSIS LOGIC FLOW CHART............................. 112 FIGURE 5-31: LOW DEMAND SCENARIO OF LOX AND LH2*........................113 FIGURE 5-32: HIGH DEMAND SCENARIO OF LOX AND LH2*.......................113 FIGURE 5-33: MULTI PUBLIC PRIVATE PARTNERSHIP EXAMPLE FOR A LUNAR GAS STATION.................................................................... ..........................116 FIGURE 5-34: COST ANALYSIS APPROACH FOR PRODUCTION, STORAGE, AND DELIVERY SYSTEMS................................................................ ......................119 FIGURE 5-35: COST BREAKDOWN RESULT OVERVIEW................................124 FIGURE 5-36: FINANCIAL MODEL.................................................................. ......125 FIGURE 5-37: NET PRESENT VALUE AND IRR DISCOUNT RATE..................127 FIGURE 5-38: PRIVATE LOW MARKET DEMAND................................. ............129 FIGURE 5-39: PRIVATE HIGH MARKET DEMAND..................................... .......129 FIGURE 5-40: MPPP LOW MARKET DEMAND................................. ..................130 FIGURE 5-41: MPPP HIGH MARKET DEMAND..................................... .............130 FIGURE 5-42: MPPP HIGH MARKET DEMAND ICE PRICE SENSITIVITY. .132 FIGURE 5-43: MPPP LOW MARKET DEMAND ICE PRICE SENSITIVITY....133 FIGURE 5-44: PRIVATE HIGH MARKET DEMAND ICE PRICE SENSITIVITY ................................................................................................................ .......................133 FIGURE 5-45: MPPP HIGH MARKET DEMAND NO ICE PRICE SENSITIVITY ................................................................................................................ .......................134 FIGURE 6-46: BASIS OF THE NON-APPROPRIATION PRINCIPLE................140 FIGURE 6-47: EXPLANATION OF SAFETY ZONES AND EXCLUSIVE ECONOMIC ZONES................................................................................................. ..143

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FIGURE ERROR! NO TEXT OF SPECIFIED STYLE IN DOCUMENT.-48 EVA AND IVA ACTIVITIES CONSIDERED FOR OUTPOST GOX DEMAND (NASA 2005)........................................................................................................... ...................165 FIGURE ERROR! NO TEXT OF SPECIFIED STYLE IN DOCUMENT.-49 LSAM ASCENT STAGE MASS BREAK-DOWN (NASA 2005)................................... .......166 FIGURE ERROR! NO TEXT OF SPECIFIED STYLE IN DOCUMENT.-50 : DISTRIBUTION OF HYDROGEN WITHIN THE FIRST 2M OF LUNAR REGOLITH AS SEEN BY SOURCE: (MAURICE ET AL. 2003), (2004A)............169 FIGURE B-ERROR! NO TEXT OF SPECIFIED STYLE IN DOCUMENT.-51: QUALITATIVE DECISION TOOL INPUT FLOW...................... ...........................174 FIGURE B-ERROR! NO TEXT OF SPECIFIED STYLE IN DOCUMENT.-52: SUPPLY FLOW THROUGH TIME-EXPANDED LINEAR NETWORK MODEL ................................................................................................................ .......................176 FIGURE B-ERROR! NO TEXT OF SPECIFIED STYLE IN DOCUMENT.-53: SPACENET MODEL (HTTP://SPACELOGISTICS.MIT.EDU).............................177

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________________________________Index of Tables

TABLE 2-1: US MISSION ARCHITECTURE FOR LUNAR EXPLORATION (ESAS 2005)...................................................................................................... ..............25 TABLE 2-2: CROSS-COMPARISON OF THE DIFFERENT SPACE POLICIES. .28 TABLE 2-3: NASA TOP 10 LUNAR SITES (ESAS 2005)..........................................32 TABLE 2-4: LANDINGS ON THE MOON PER PARTY AND DATE......................37 TABLE 2-5: LSAM AND LSAM EQUIVALENT ASCENT STAGE PROPELLANT DEMAND PER VEHICLE..................................................................................... .......38 TABLE 2-6: GASEOUS OXYGEN NEEDED FOR A SEVEN-DAY, FOUR CREW SORTIE MISSION (ESAS 2005)........................................................... .......................38 TABLE 2-7: OUTPOSTS CONSIDERED IN THE ANALYSIS (ESAS 2005)...........38 TABLE 2-8: GASEOUS OXYGEN (GOX) FOR OUTPOST DEMAND CALCULATION........................................................................................... .................39 TABLE 2-9: LUNAR SURFACE AVERAGE REGOLITH COMPOSITION (SANDERS ET AL. 2006)............................................................................ ..................44 TABLE 2-10: CONCENTRATION OF SOLAR-WIND IMPLANTED VOLATILES IN REGOLITH (SANDERS ET AL. 2006)........................................ ..........................45 TABLE 2-11: HYDROGEN VOLATILE CONCENTRATION IN POLAR REGOLITH (SANDERS ET AL. 2006)................................................................ ........45 TABLE 2-12: ISRU PROCESSES FOR OXYGEN EXTRACTION FROM REGOLITH (FERTILE MOON 2006; SANDERS 2007)...................... .....................49 TABLE 2-13: PROCESSES FOR OXYGEN AND HYDROGEN EXTRACTION FROM ICE (BLAIR 2002; FERTILE MOON 2006)................................................. ..50 TABLE 2-14: EXTRACTION PROCESS OF SOLAR-WIND-IMPLANTED H2 FROM REGOLITH (FERTILE MOON 2006)..................................................... .......51 TABLE 2-15: THE FOUR ECONOMIC SCENARIOS TO BE ANALYZED WITHIN THE REPORT...................................................................... .........................52 TABLE 3-16: LUNAR TEMPERATURE RANGE TAKEN FROM (ECKART 1999; HEIKEN 1991)..................................................................................... ..........................54 TABLE 3-17: COLD TECHNOLOGIES RESULTS FROM ICES: MANDATORY = X, HELPFUL = O (ID.)..................................................................................... .............62 TABLE 3-18: SUMMARY OF TANK-SUITABLE MATERIAL PROPERTIES (WWW.MATWEB.COM).......................................................................................... ....64

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TABLE 3-19: MOBILE PLATFORM CHARACTERISTICS (ECKART 1999)......71 TABLE 3-20: LUNAR TRANSPORTATION SYSTEMS (APEL 1989).................... .73 TABLE 3-21: FIXED PLATFORMS CHARACTERISTICS (APEL 1987), (TRANSNEFT 2004), (DOPPELMAYR N.D)............................................. .................74 TABLE 3-22: QUALITATIVE METHOD RESULTS.................................. ...............76 TABLE 3-23: QUANTITATIVE SELECTION CRITERIA....................... ................79 TABLE 3-24: WEIGHTED QUANTITATIVE DECISION MATRIX RESULTS....81 TABLE 4-25: DEMAND SCENARIOS.................................................... ....................85 TABLE 4-26: PRODUCTION SCENARIOS............................................... ................86 TABLE 4-27: QUANTITATIVE DECISION TOOL WEIGHTS WITH SECONDARY WEIGHTS..................................................................................... ........88 TABLE 4-28: COMMUNICATIONS CONCEPT COMPARISONS.........................98 TABLE 4-29: TANK DIMENSIONS.......................................................................... .105 TABLE 4-30 TANK DIMENSIONS BY MATERIAL (WWW.MATWEB.COM)...106 TABLE 4-31: ROVER POWER SYSTEM OPTIONS............................... ...............107 TABLE 4-32: ROVER WHEEL OPTIONS................................... ...........................108 TABLE 5-33: FORECASTED REQUIRED INFRASTRUCTURE FOR TOTAL LIFESPAN OF 20 YEARS........................................................................... ................114 TABLE 5-34: HIGH RISK ISSUES FOR BUSINESS ANALYSIS..................... ......119 TABLE 5-35: COST CATEGORY BREAKDOWN................................................... 120 TABLE 5-36: INPUTS AND OUTPUTS FROM ISRU MODEL FOR SUPPLYING LOX AND LH2 (FERTILE MOON 2006)........................................................... .......121 TABLE 5-37: LOX AND LH2 TANKS COST DATA FROM ISRU CASE STUDY (BLAIR 2002)..................................................................................... ..........................121 TABLE 5-38: COST FOR PROPOSED LOX AND LH2 TANKS............................121 TABLE 5-39: PAST ROBOTIC ROVER DEVELOPMENT COST........................122 TABLE 5-40: ROVER SPECIFICATIONS.............................................................. ..122 TABLE 5-41: COST ESTIMATION RESULTS............................. ...........................123 TABLE 5-42: PRICING OF PRODUCT FOR EACH COSTUMER.......................126 TABLE 5-43: CORPORATE TAX RATES FOR SPACE FAIRING COUNTRIES ................................................................................................................ .......................127

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TABLE 5-44: GENERAL ASSUMPTIONS FOR BUSINESS ANALYSIS..............128 TABLE 5-45: FINANCIAL RESULTS................................................. ......................130 TABLE 6-46: SUMMARY OF LIABILITY UNDER THE LIABILITY CONVENTION (ADELTA LEGAL SPACE LAW 2007)..........................................146 TABLE A-47: QUALITATIVE TENSILE STRENGTH TABLE.............................172 TABLE 48: QUALITATIVE THERMAL CONDUCTIVITY TABLE.....................172 TABLE 49: QUALITATIVE DENSITY TABLE........................................................ 172 TABLE B-50: QUANTITATIVE DECISION TOOL FORMULAE.........................175 TABLE B-51: ELEMENTS FOR MODELING A LUNAR CRYOGEN DELIVERY SYSTEM.................................................................................................... ...................177 TABLE C-52: LSAM MODIFICATIONS..................................................... .............179

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______________________________List of Acronyms

ACDPT

GSO He4

Advanced Cryocooler Development Technology Program Autonomous Lunar Transport Vehicle Australian Space Council Italian Space Agency All Terrain Hex-Legged Extra Terrestrial Explorer Business to Business Business to Consumer Business to Government British Petroleum Crew Exploration Vehicle Chinese Lunar Exploration Program Chinese National Aerospace Administration Chinese National Space Agency Cryogenic Operation for the Long Duration Canadian Space Agency Coefficient of thermal expansion Directorate of Defense Trade Controls Export Administration Regulations Expendable Launch Vehicle European Space Agency Exploration Systems Architecture Study Extra-Vehicular Activity Federal Energy Management Program Flexible Optical Solar Reflector Gross Domestic Product Gaseous Hydrogen General Motors Gaseous Oxygen Global Navigation Satellite System Geo-Stationary Orbit Helium 4

HMD ICE

High Market Demand Independent Cost

ALTV ASC ASI ATHLETE B2B B2C B2G BP CEV CLEP CNAA CNSA COLD CSA CTE DDTC EAR ELV ESA ESAS EVA FEMP FOSR GDP GH2 GM GOX / GO2 GNSS

ICES IDEST IGO IMLI IRR ISA ISRO ISRU ISS ITAR ITU IVA JAXA JPL KSC L1 LCH4 LEAD LEO LeRC LH2 LLO LMD LOX LOTRAN LRO LRV LSAM LV Maglev MIT MLI M-LSAM

Estimate Ice, Cloud, and land Elevation Satellite Institut du Droit de l’Espace et des Telecommunication Inter-Governmental Organization Integrated Multi-Layer Insulation Internal Rate of Return International Seabed Authority Indian Space Research Organization In-Situ Resource Utilization International Space Station International Traffic in Arms Regulation International Telecommunications Union Inter-Vehicular Activity Japanese Astronautic Exploration Agency Jet Propulsion Laboratory Kennedy Space Center Lagrange point 1 Liquid Methane Lunar Exploration And Development agency Low Earth Orbit LeClerc Research Center Liquid Hydrogen Low Lunar Orbit Low Market Demand Liquid Oxygen LOcal TRANsportation rover Lunar Reconnaissance Orbiter Lunar Roving Vehicle Lunar Surface Access Module Launch Vehicle Magnetic Levitation Massachusetts Institute of Technology Multi-Layer Insulation Modified Lunar Surface Access Module

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MMH MOSAP MPL MPPP MSFC NASA Nav/Comm NGO NORCAT NPO Energia NPV NTO OFAC OHBSystem OST PPMD PPMS PPP PV RKK Energia RLEP RLV

Metallized Gelled Hydrogen Mobile Surface Applications Pressurized rover Maximum Probable Loss Multi Public Private Partnership Marshall Space Flight Centre National Aeronautics and Space Administration Navigation and Communication Non-Governmental Organization Northern Center for Advanced Technology Research & Production Association Energia Net Present Value Nitrogen Tetra oxide Office of Foreign Assets Control Orbitale Hochtechnologie Outer Space Treaty Propellant Positional Management Devices Propellant Positional Management System Public Private Partnership Present Value S.P. Korolev Rocket and Space Corporation Energia Robotic Lunar Exploration Program Reusable Launch Vehicle

ROI RP-1 RP-1 RSA RSC Energia RSC SDN SEC SEU SG SM SOx SPA-Basin SS SSTO STS SWOT TBD UK UNCLOS UNCOPUO S USML VDMLI VIP VSE

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Return of Investment Refined Petrolium-1/ Rocket Propellant 1 Russian Space Agency Rocket and Space Corporation Energia Russia Space Corporation Energia Specially Designated National Shakelton Energy Corporation Single Event Upset Specific Gravity Service Module Molten Silicate South Pole Aitkin Basin Sun Shields Single Stage to Orbit Space Transportation System Strengths, Weaknesses, Opportunities, Threats To Be Decided United Kingdom United Nations Convention on the Law Of the Sea United Nations Committee on the Peaceful Uses of Outer Space United States Munitions List Variable Density MultiLayer Insulation Vacuum Insulation Panels Vision for Space Exploration

Table of Units

B BP BUSD ºC CSCF F g GGE h ISP K kg km KUSD kW m m2 m3 M mg MUSD MSCF Nm3 ppm t USD

Billion Barometric Pressure Billion US Dollars Celsius Hundred Standard Cubic Feet Fahrenheit Gravity Gasoline Gallon Equivalent Hour Specific Impulse Kelvin Kilogram Kilometer Thousand US Dollars Kilowatt Meter Square Meter Cubic Meter Million Milligram Million US Dollars Thousand Standard Cubic Feet Normal cubic meter of a gas Parts per million Tonne US Dollar

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Full Moon

_____________________________________Chapter 1

1 Introduction

Space exploration today is beginning to take on a new approach via a focus on global partnerships. Numerous reasons have been put forth for returning to the Moon, ranging from scientific exploration to resource exploitation for commercial gain. Regardless of the intent, one thing is clear: lunar activities have only just begun, and as these expeditions increase in scale and duration, so too will the need for supporting infrastructure in areas of life support & habitation and transport. The elements oxygen and hydrogen are invaluable; they address all these needs. In addition to being essential for synthesizing water and air for life support, oxygen and hydrogen are commonly used propellants. Extraction and production methods in which oxygen and hydrogen are extracted from the lunar regolith1 have been well studied, but little research has been done to address how these resources may be made readily available to the user. This report outlines a possible architecture to enable the storage and delivery of oxygen and hydrogen on the lunar surface. Simply put, Full Moon is a proposal for a ‘lunar gas station.’ A lunar production facility alone will not necessarily enable lunar explorers to go where they wish. Full Moon aims to be one of the first proposals with a complete interdisciplinary approach to the storage and delivery architecture on the lunar surface.

1.1.1The Full Moon project Just as purchasing barrels of petroleum from an oil refinery is not the optimal solution for a mother wishing to drive her children to their daily activities, connecting directly to a regolith processing facility is unlikely to be the optimal solution for a lunar explorer wishing to fill up her rover to conduct a geological survey at a distant site if they are not accessible.

1

Regolith is the term used for the layer of loose, heterogeneous material covering solid rock on the lunar surface and elsewhere. 20

Introduction

Figure 1-1: The Full Moon study, its technical components and its interfaces with the production facility and users/customers. Full Moon seeks to fill a gap in the current Moon research by proposing a complete architecture for storage and distribution of oxygen and hydrogen. In the distribution chain production, storage, transportation, customer the report addresses the storage and transportation architecture. The expected demand for lunar delivery of oxygen and hydrogen is estimated based on current plans by space agencies. It is assumed that sufficiently large production facilities are installed on the lunar surface to cover this demand. The feasibility of the project is evaluated technically, economically and legally. To be an economically feasible alternative to simply bringing oxygen and hydrogen along with other supplies from Earth, the cost of producing oxygen and hydrogen gas from the regolith must be sufficiently low that development, launch costs and depreciation costs of a production facility are offset by the high cost of transporting oxygen and hydrogen from the Earth on launch vehicles.

1.1.2Definition of Project Scope During the “Moon Symposium” held at ISU in February 2007 it became clear from several presentations and from discussions with various experts that little work has been done on storage and transportation of hydrogen and oxygen on the lunar surface. Storing hydrogen and oxygen on the Moon obviously requires different solutions than conventional storage on Earth; similarly there are considerable differences between in-space microgravity storage of hydrogen and oxygen such as the systems attached to the International Space Station (ISS). The lunar surface shares characteristics with both of these environments, but it also presents unique challenges not present in either deep space or on the Earth. Very extensive research has been and is being conducted in the field of In situ Resource Utilization (ISRU) for space applications. ISRU is a wide field, and numerous studies propose and investigate different methods for extracting oxygen from the lunar regolith, as well as for extracting hydrogen from water ice, should it be determined to exist in the Polar Regions. The requirements of potential lunar visitors are not difficult to estimate, there exists a wide variety of proposals for going to the moon by agencies and private companies.

21

Full Moon

In-orbit refueling of spacecraft holds great promise as a cost reducer for space missions; however it falls outside the scope of this project. In order to limit the scope of the project it was decided not to consider inorbit refueling or delivery. This allows the report to focus on surface issues on the Moon. If a cis-lunar economy becomes a reality, it is likely (based on orbital mechanics calculations) that propellant produced on the Moon and shipped from the Moon will be less costly than propellant launched from the Earth, even for satellites in Low Earth Orbit (LEO). In Full Moon, the delivery chain of a potential in-orbit service is considered to end at the lunar launch pad. This decision was also made to avoid interfering with the planned team project of the 2007 Summer Session Program, which will focus on in-orbit servicing of satellites. Lunar ISRU promises to deliver other substances hydrogen. In the near future no manufacturing is place on the Moon, and also any byproduct of production would belong to the production plant, not

than oxygen and expected to take oxygen/hydrogen to the distributor.

The report focuses on the timeframe 2018 – 2047. All future dates are estimated from current agency plans. It was deliberately chosen to focus on the “near-term” since it is believed that this would offer more value to the space community. It is the hope of the authors that the analysis and conclusions of this report will be useful to agency planners and commercial companies considering working along with space agencies to provide a refueling infrastructure in space. It is also hoped that this report will lay the groundwork for future research conducted by individual members of our team.

1.1.3Report Structure (Readers’ Guide) The Full Moon report follows a logical flow so as to equip the reader with an understanding of the lunar operation and business environment. A chapter summary is given here for the convenience of the reader interested in only specific aspects of our proposed architecture and its interdisciplinary evaluation. However, the reader is encouraged to read chapters in the following order as later sections may derive heavily on earlier ones. Chapter 1: Introduction. Chapter 2: Drivers & Constraints. This Chapter aims to clarify the demand for oxygen and hydrogen based on the policy of different space agencies and estimates of the fuel consumption spacecraft capable of landing on the lunar surface. The areas of the Moon most likely to be visited for scientific reasons are identified, similarly the areas most likely to be selected as possible Moon base locations. Chapter 3: Technical Assessment. This Chapter outlines the unique features of the lunar environment and the requirements they place upon any lunar installations. The methods of assessment used in the report are laid out. Various storage and transportation systems are introduced.

22

Introduction

Chapter 4: Recommendations for a System Architecture. Following the analysis of the preceding chapter, this chapter outlines the architecture proposed by the team. The location of customers, production facilities and storage facilities are presented. The transportation systems enabling movement of oxygen and hydrogen between production facilities, storage facilities and customers are described, and the interfaces enabling transfer of propellant are briefly introduced. Chapter 5: Economic Analysis. This chapter examines the economic feasibility of building the system outlined in the previous chapter. The cost of producing, transporting and storing oxygen and hydrogen in situ are compared to the cost of producing oxygen and hydrogen on the Earth and delivering it to the Moon. Chapter 6: Politico-Legal and Ethical Analysis. This chapter addresses the general legal framework of space resource utilization, as well as the specific issues concerned with the implementation of the proposed storage and transportation architecture. Chapter 7: Recommendations and Conclusions. The recommendations of our team, based upon the analysis in this report, are presented. Appendix A: Details tank material selection. Appendix B: Explains the system selection process. Appendix C: Details for modifications to and sizing for delivery vehicles.

23

Full Moon

_____________________________________Chapter 2

2 Drivers & Constraints

In order to develop an adequate storage and delivery infrastructure, it is necessary first to understand the parameters that will drive and limit it. In any environment the primary driver is the market demand, which this chapter assesses by answering the following questions: who are the customers and where are they located? (Section 2.1) and what is their demand for oxygen and hydrogen? (Section 2.2). The ability to service them is subject to constraints, which are categorized as physical, economic or legal. Economic constraints are linked to the presence of competitors (Section 2.3), while physical ones encompass the challenges of the the lunar terrain (Section 2.4), the limitation of resources (Section 2.5) and the production facility’s capabilities (Section 2.7). Legal restrictions originate from national and international legislation and are discussed after the technical solution, in Chapter 6.

2.1Market Analysis for Refueling Conducting a market analysis is not a simple task, especially for a lunar refueling station: the market does not currently exist as humans have not yet returned to the Moon. Considering the substantial budgets required to do so, it is expected that the first customers will be government agencies. It thus becomes necessary to make lunar surface refueling attractive enough to foster both government and private demand, much like the construction of United States government railroads in the 1800s helped generate new industries in their vicinities.

2.1.1Potential Customers Customers were identified by examining the roadmaps of space agencies and future development plans of private companies interested in lunar exploration. Therefore, this market analysis is based on current plans, which are subject to change and which generally under-estimate timelines.

National Aeronautics and Space Administration (NASA) The United States (US), one of the two historic space leaders, remains one of the references in the space sector, especially in terms of policy leadership. After the last Apollo mission in 1976, the US lunar interests declined with only two dedicated lunar robotic missions since then: Clementine (1994) and Lunar Prospector (1998), both of which remained in Low Lunar Orbit (LLO) for remote sensing of the lunar surface. However, this trend was reversed with the announcement of a

24

new “Vision for Space Exploration” (VSE) in 2004 by President Bush, which aims to “return to the Moon by 2020” and “with the experience gained on the Moon”, “take the next steps of space exploration: human missions to Mars and to worlds beyond”. His call upon NASA to “gain a new foothold on the Moon” quickly became the cornerstone of a new global exploration effort of the lunar surface, with many countries following suit by declaring their own lunar intentions. Such leadership on the international space stage was already exercised by the United States in the past two decades by being the main architect of the International Space Station (ISS). After the completion of the ISS in 2010 and the retirement of the Space Shuttle, the US plans to focus on developing the necessary technology (including a new Crew Exploration Vehicle (CEV)) to reach its goal of building a permanent outpost on the Moon. NASA recognizes that these plans are ambitious, as was highlighted by their current administrator: “The United States working alone cannot fulfill the sweeping goals of the VSE; we must maintain the strong international partnerships that have been built during the space station era, and we must extend those partnerships even more broadly to enable a robust human space exploration” (Griffin 2005). This clearly opens the door for international cooperation similar to that for the ISS, so long as it does not lie within NASA’s critical path (Sanders 2007). NASA’s current architecture is summarized in Table 2-1 below: Table 2-1: US Mission Architecture for lunar exploration (ESAS 2005) Key Decisions

4-7 day missions

Permanent base deployment missions & Steady-state base operations Missions to Mars

Architecture 2 missions per year 4 crew per mission Extensive EVA Local mobility (un-pressurized rovers) Mix of exploration technology and science experiments 4 lunar landings per year 6 months crew rotations Logistics mission between crew flights Extended mobility (pressurized rovers) Emphasis on in situ resource utilization To be decided

Date

20182020

20202030+

2035+

Russian Space Agency (Roskosmos) As the other historical and leading space-faring nation, Russia is also considering future lunar missions for the next decades. Even if this vision is, for the moment, not as clear as the US VSE, the space capacities and possibilities of this country have to be taken into account. With its industrial, scientific and technological capabilities, Russia will no doubt be an important player and/or partner for the Moon exploration in the future, just as it was in the past. Russians have acquired great expertise in orbital space stations with MIR, allowing

25

Full Moon

them to participate actively in the construction of the ISS by providing almost half the modules. The national agency, Roskosmos, does not yet have a strategic approach to further space exploration. Visions for the future have mainly come from private companies such as NPO Energia. Nevertheless, it appears that Roskosmos is working on a long-term vision (2020 to 2040) that will consider robotic and manned missions, including a Moon base in 2030 (Casini 2006). It appears that the key challenge to the implementation of this project would be the limited budget Roskosmos, which is linked to Russia’s current economic difficulties. One mitigation strategy used so far has been to increase international cooperation: the European Space Agency (ESA) has launched many of its astronauts on Russian rockets, while much of the Chinese manned program is based on Russian technology. Similarly, the Russians were the first to offer tourist flights to the ISS.

European Space Agency (ESA) The European Space Agency unveiled its space exploration program “Aurora” in 2001 prior to the VSE, with the goal of sending a series of robotic missions to the Moon, Mars, other planets and near-Earth asteroids. The main difference with the VSE were that the program was robotic, aiming to assess the economic potential of space resources, and focused on Mars, with the Moon only as an aside. Nevertheless, ESA revised its program after announcement of the VSE and now includes sending a manned spaceship to the Moon between 2020 and 2025, followed by a human mission to Mars by 2025 to 2030, without much further detail (ESA 2006). This demonstrates ESA’s strong belief in international cooperation, along with its active participation in the ISS (ESA’s Columbus module is due to launch in 2007).

Japanese Aerospace Exploration Agency (JAXA) JAXA issued its “JAXA 2025” vision in 2005 which describes Japan’s interest in constructing and developing lunar surface infrastructure in the long-term, with emphasis on robotic technologies (Casini 2006). Manned missions would occur between 2025 and 2035, even though this would probably occur via international cooperation, given the limited budgets and the specificity of the Japanese space program. This is reflected by the fact JAXA links for a long period ahead its programs with the activities of other space-fairing nations, despite its growing desire for greater autonomy in space. Although the plan’s budget has not yet been approved by the Japanese government, one can expect Japan to be a reliable international partner in the future, as it has been in the past (JAXA’s ISS Kibo module is due to be launched in 2008).

Canadian Space Agency (CSA) Canada’s space efforts are modest but highly specialized, particularly in the field of robotics. Canadians have specialized in developing the Canadarm for the Space Shuttle along with Canadarm 2 for the ISS. These, along with its special status as a “cooperating state” within ESA, underline the importance of Canada’s contribution to NASA and ESA’s space exploration plans. As of today Canada does not have a lunar exploration program of its own; instead it is developing technologies that would complement that of NASA and ESA: from ISRU robotics to scientific instruments for satellite remote sensing of the 26

Moon (Ghafoor 2007). There is no doubt that Canada will develop key technologies for the implementation of the VSE and the Aurora program.

China National Space Administration (CNSA) As the third nation in history to put a man in orbit, China’s space efforts are not to be underestimated given the country’s already large and growing economic resources. According to the declarations of the first Chinese “taïkonaut”, Yang Liwei, at the 2006 International Aeronautical Congress, there is no doubt that China will follow the steps of the Americans on the Moon (Coué 2007). Indeed, ambitious space exploration plans have been announced by the Chinese authorities. Although the details are rather secret, it is known that new additions to the Long March launcher family are being developed, along with more sophisticated Shenzou manned capsules. With respect to the Moon, there are plans for a lunar reconnaissance orbiter, followed by a soft lunar lander in 2010 and a sample return mission by 2030 (Id.). Manned flights and construction of a lunar base would begin after 2030, as confirmed by C. Tangming (2007) at the ISU “Why the Moon?” symposium. Named “Chang’e”, the program aims to promote space science and technology along with international cooperation. However, countries such as the US have expressed concern over the level of involvement of the Chinese military in the program. This reached a climax on the 11th January 2007 when China successfully performed an anti-satellite test. The United States then announced that it was suspending plans for developing cooperative space ventures, including a joint mission to the Moon, which is reflected by NASA’s press statement: “We believe China’s development and testing of such weapons is inconsistent with the constructive relationship that our presidents have outlined” (Gertz 2007).

Indian Space Research Organization (ISRO) As commented by Vajpayee (2005), the Indian space program is also geared towards the Moon: “the country has made significant progress in science and technology and India's scientific development should be strong enough to realize its dream of sending a man on the Moon”. The first step to achieving that goal has already been taken with the upcoming launch of Chandrayaan-1 in 2008, prior to NASA’s Lunar Reconnaissance Orbiter (LRO), which will undertake the full 3-D mapping of the Moon, and aim to confirm the presence of water at the poles. According to Jayaraman (2006), the first manned landings on the Moon would occur in 2020. Taking this into account and considering India’s impressive economic growth, one can easily imagine India as being one of the future major space-faring nations involved in lunar activities. As affirmed by the president of India, Dr. A.P.J. Abdul Kalam, during a visit to the International Space University on 26th April 2007: “I foresee that an important contribution for future of exploration by India would be, space missions to the Moon and Mars founded on space industrialization and international cooperation.”

27

Full Moon

Emerging Countries – the Unidentified Third Party It was decided to consider in the demand analysis an unidentified country whose current contribution is small but who might emerge as an important space faring nation in the near future. Such countries could be Brazil, Indonesia, Pakistan, South Korea, etc. Alternatively, it could represent one of the assumed international partners that may decide not to participate in the international venture, or only partly, such as the Italian Space Agency (ASI).

Agencies Summary The roadmaps described above of the said agencies are summarized in Table 2-2. Table 2-2: Cross-Comparison of the different space policies Agency

Program name

NASA

Vision for Space Exploration

ROSKOSMO S ESA JAXA CNSA ISRO CSA

Policy priority

Permanent base on the Moon as a first step to further manned exploration to Mars Permanent base on the Federal space Moon program Exploration of human resources Aurora

Manned mission

Vision of 2005 White paper Indian space program

Construction of a base on the Moon Manned mission

Space Vision System

Target year 2024

2019 20202025 2030 2017

Manned mission

2020

Shuttle Remote Manipulator System (Canadarm) International Cooperation

2007

Private Enterprise At the present time, few commercial activities have plans to utilize space resources from the Moon. It is hoped nevertheless that the presence of agencies will foster the appearance of commercial activities, becoming in turn lunar refueling customers. As described by the Lunar Exploration and Development Authority (Whittington 2005), this is what NASA aims to encourage: by building the initial facilities and handing them over to private companies for mining and production of resources, a commercial Earth-Moon system could be initiated.

Space Tourism Space and lunar tourism also have the potential of representing significant commercial activity. For example, “Space Adventures” has announced plans in the past for space tourist flights around the Moon aboard a Russian-built Soyuz spacecraft, while Excalibur Almaz and TransOrbital Inc aim to provide not only flights but lunar infrastructure

28

for tourists. However these claims must be taken with precaution given that the timelines provided by these companies are usually not realistic (Peeters 2000). In Figure 2-2 below, one can find a “global roadmap” that summarizes the current plans of agencies and private companies.

Figure 2-2: Global roadmap space exploration

2.1.2Customer Location Now that the customers have been identified, it is fundamental to understand their location, such that the architecture can be adequately designed to cater for them. Although future plans are dynamic in time, identifying where future customers went in the past and will go in the future provides the best possible mapping of the market demand.

Past Landings As part of the “space race” the Soviet Union and the United States successfully landed 13 robotic probes on the surface of the Moon (Luna and Surveyor programs), culminated by NASA’s six manned Apollo landings between 1969 and 1972. Initial missions carried out surveying of the surface, while later ones performed geological experiments and/or returned samples back to Earth, forming the basis of what is known today about regolith and its oxygen and hydrogen components.

29

Full Moon

Figure 2-3: Near and far side lunar topography (http://sse.jpl.nasa.gov/multimedia/display.cfm?IM_ID=804) As shown in Figure 2-3 above, all missions were on the near side of the Moon, targeting mainly the equatorial regions. This is due to the telecommunication relay systems needed to stay in constant communication with Earth when one is on the far side and at the poles. Although these sites have already been visited, they could be revisited in the future, and thus represent potential landing sites.

Future Landings: The United States As explained in Section 2.1.1, the United States main purpose of returning to the Moon is to establish a lunar outpost that will be used for scientific experimentation and as a test-bed to future Mars exploration missions. As detailed in Allen (2007), NASA has decided to build its outpost at Shackleton Crater at the South Pole as shown in Figure 2-4 below.

30

Figure 2-4: Illustration of Shackleton Crater Moon Base (ESAS 2005) As detailed by Allen (2007), NASA believes that such a choice is safe, cost effective, resource plentiful and flexible (see Section 3.1 for discussion on the lunar environment). It follows that NASA will concentrate its efforts on building the outpost there, and, given that NASA plans to lease its facilities in order to foster private enterprise, it makes business sense to produce and store the propellant at the same place. Hence, the storage and delivery business considered in this report should be based at the South Pole. In addition to the outpost, NASA plans to perform “sortie” missions to other locations such as those shown in Figure 2-5 below.

Figure 2-5: NASA past and future landing sites (ESAS 2005)

31

Full Moon

These “Top 10” exploration sites were selected on the basis of ISRU potential, scientific opportunities, “Mars Analog” conditions and geologic diversity, under a “Global Access” or “Anytime Return” policy. The characteristic of these sites are summarized in Table 2-3 below. Table 2-3: NASA Top 10 lunar sites (ESAS 2005) Site South Pole SPA Basin Floor Orientale Basin Floor Oceanus Procellarum Mare Smythii Mare Tranquilitis Rima Bode Central Far Side Highlands Aristarchus Plateau North Pole

Distance from South Pole [km] 0 1 035 2 187

Interest Base, Geology, Mars Analog, Water? Geology, Astronomy Geology, ISRU (Mare & Highlands mix)

2 677

Geology, ISRU (Ti)

2 789 3 018 3 145

Geology, ISRU (Fe) ISRU (Ti) Geology, ISRU (Ti) Geology, ISRU Astronomy

3 492

(Al,

Ca),

3 552

Geology, ISRU (solar wind H)

5 414

Base Alternative, Water?

Regarding missions to these sites, it is important to note the following: • Robotic missions to the surface have been excluded, implying

that they will not be refueled. This decision was taken for two reasons: firstly, robotic missions tend to be small-sized and are only occasionally designed to return to Earth, thus their demand for refueling is deemed negligible; secondly, probes are designed on a one-off basis, for a specific mission, and thus all have potentially different refueling tanks and needs. • On the other hand, manned mission modules will be “mass

produced” and will require the return of the astronauts via their Lunar Surface Access Module (LSAM) ascent vehicles, thus potentially representing a significant market. However, although the said NASA policy does not exclude refueling in principle, it does exclude landing quasi-empty-tanked which would pose a problem to a refueling business. New procedures are needed here, such as establishing the refueling medium on the surface, before the LSAM launches from Earth (typical Earth-Moon trajectories take three to four days). The mobile refuel depot can be equipped with a beacon to guide the LSAM during the descent phase (see Section 4.3.4). This would mean that aborting the mission during descent is not ‘to orbit’ but ‘to surface’. Such an approach may seem as a non-acceptable risk with the current way human space exploration is foreseen. As highlighted by Tolyarenko (2007), “new endeavors require new approaches to ensure success”.

• In the early years, these manned sortie missions will be “Apollo-

type”: LSAMs will be launched from Earth, land at a specific site, 32

perform their mission, then return to Earth, without passing at all by the polar outpost (except those of course which are launched specifically to go to the outpost). This was assumed because the current LSAM design cannot perform multiple launches and landings, and no other vehicle that could do so is currently in design. • As explained in Stevens (2007) these missions of four astronauts

will last seven days and consist of up to six days of science and exploration, within a 75 km radius, using a pressurized rover and driving at maximum speed of 20 km/h (72 m/s), as shown in Figure 2-6 below. Unpressurized rovers (similar to those used during Apollo missions) could also be used, but this would significantly reduce the science return given that the astronauts have a maximum Extra-Vehicular Activity (EVA) time of six to eight hours and would have to return to the LSAM. However, unpressurized rovers, smaller in size, could be launched with the manned LSAM, while pressurized ones would require a dedicated launch (ESAS 2005). Either way, these EVAs that consume water and oxygen could represent an additional demand to the LSAM refueling.

Figure 2-6: Sortie missions at Aristarchus Plateau (Stevens 2007)

Future Landings: International Cooperation As described in Section 2.1.1, the roadmaps of other countries are not as detailed as NASA’s. At the exception of ESA which describes that its robotic program will focus on the equatorial regions, there are no details regarding location. Therefore, it is assumed that as part of their international cooperation with NASA, these countries will land at the same sites as those selected by NASA. Similarly, these countries are assumed to use NASA’s outpost, or build their complementary modules, much like for the ISS.

Future Landings: China and the Unidentified Third Party Given that as of today, the plans of all agencies to return to the Moon, including that of China, are science-driven; China and the unidentified 33

Full Moon

third party are assumed to have the same sites of interest as NASA. However, they are assumed to have build at one point a base of their own at the South Pole, for similar reasons to NASA’s, much like bases from different countries are close to each other in Antarctica.

Future Landings: Private Enterprise and Space Tourism It is difficult to predict which private industries would spawn on the Moon as agencies invest in the infrastructure. In any case, their activity would most probably center on that of a government-owned base, which are all assumed to be built at the South Pole. Similarly, for space tourism: in 2006, space tourism focused on sub-orbital flights and trips to the ISS, but in the future these will undoubtedly expand to include Moon-bound destinations. Only trips landing on the lunar surface are to be considered in this report, and it is assumed that these would all be destined to polar outposts for two reasons: firstly, one expects the rich people to prefer the option of “living like an astronaut” than perform experiments in the field; secondly, it would be cheaper for operators given that the traffic will be greatest at the poles.

Market: Demand Analysis In this part the demand analysis for the lunar fuel is made. Along with the official roadmaps, a set of assumptions is laid down and based on those the demand for LOX-GOX and LH2 on the Moon is estimated for the years 2018-2047. Two main scenarios are developed. The ‘baseline’ scenario, which takes into account landings on the Moon and Lunar outposts from NASA and other agencies. The ‘optimistic’ scenario is the ‘baseline’ with the addition of demand of other parties like China and for Mars exploration after 2035.

2.1.3Demand Scenario General Framework Two demand scenarios are considered: baseline and optimistic.

Baseline Scenario As mentioned in Section 2.1.1 the most likely scheme of lunar exploration is a multinational cooperation led by NASA, in a similar way as the ISS. Willing to return to the Moon in 2018 and develop an outpost on the Moon after 2022, NASA will develop all the technologies lying in the ‘critical path’: transportation and crew vehicles, ISRU, standardization and interfaces, outpost infrastructure, communications (NASA 2005). The rest of the agencies (ESA, CSA, JAXA, and ISRO) are expected to provide modules on the lunar outpost, provide logistics support and conduct their own research on the lunar surface via barter agreements. Russia could have a more active role, by developing a transport vehicle also, but in our analysis RSA is included in the ‘international partners’ of NASA, as the roadmaps of RSA are unclear and mainly coming from private companies (NPO Energia) and not RSA itself (see also 2.1.1). China is considered to follow an individual path in exploration and outpost development after 2030. The Chinese effort is considered to be half that of NASA. This is an assumption, as there is no official governmental announcement. China is included only in the business ‘high demand’ scenario for two reasons:

34

• May want to develop own ISRU capabilities or be independent

from western supplies. Based on the current status of China-US relations, it seems highly unlikely that China will buy supplies from a NASA led consortium. • As mentioned, it is not in the governmental plans yet to go to

land humans on the Moon. China could be a potential competitor of supplies on the Moon.

A third party that is not identified today is also included in the scenarios after 2030. Their effort could be based in already developed technology of other countries by 2030, in a similar way that China is currently using Russian technology for LEO flights. The third party is, as China, included only in the ‘high demand’ scenario, as it is uncertain and could be supplied by China. Space tourism and private flights, are considered to take place after 2030, with destination the outpost. As the there are no definite plans, a conservative assumption has been made that there will be a single tourist flight to the Moon every year. Based on the tourists flying on the ISS, and the order of magnitude higher cost of landing on the Moon, it seems unlikely that there will be high traffic of tourists to the Moon.

The Optimistic Scenario It is clearly stated both in the ESAS report (2005) and the VSE (2004) that the Moon is the first step to go to ‘Mars and Beyond’. Mars human exploration is planned for after 2035 in the NASA roadmap with a vague architecture. In this context, lunar fuel for Mars exploration is weighted as more realistic to raise the demand by an order of magnitude, to examine the business case, than considering a very high traffic between Earth and Moon. High traffic to the Moon would suggest some kind of commercial exploitation of the Moon, which has the same level of uncertainty with exploring Mars. Mars exploration is assumed to begin from the Moon. This includes a refuel station in Moon-Earth L1, and the transportation of fuel from the lunar surface. Mars exploration is too far in the future, and there is no definite information in NASA and ESA roadmaps to make a solid case. To have an estimation of the market size for lunar fuel to go to Mars, indicative data from ESA report are used (Casini 2006), for 20 t payload delivery to Mars using cryogenic stage fuelled from lunar fuel. The mating of the cryogenic stage for Mars delivery and the payload is made in L1 or LEO. In the case nuclear propulsion is used, with LH2 as propellant (Hoffman 1997) the order of magnitude of propellant is the same (~100 tonnes per mission) from a refuel L1 depot to Mars. In the high demand business case, the numbers after 2035 are to indicate an order of magnitude increase of demand, to examine the business case, and cannot be taken as absolute values.

2.1.4Timeline Selection A basic assumption to base the analysis is that agency roadmaps for the Moon exploration in the 21st century go as currently planned.

35

Full Moon

Timeline for the analysis is 2018-2047. In 2018 the science NASA landings on the lunar surface start. In 2023 the building of lunar outpost is starting. From 2023 the demand of LOX/LH2/GOX is constant for NASA given that NASA switches from building to servicing the outpost. 2027 is the year into business, as there is time needed for testing of ISRU (Sanders 2007). From this date 20 years are added to form a business case thus the timeline of the analysis finishes in 2047. Timeline of the analysis spans 40 years from now (2007-2047) which is a very long time for which to predict accurate numbers.

2.1.5Demand from Landings Lunar landings are assumed to be made with the currently proposed LSAM vehicle (ESAS 2005). Future lunar operations, from all parties, and tourism flights are assumed to use LSAM or LSAM-equivalent vehicles (Figure 2-7). The amount of fuel needed, shouldn’t vary considerably if there are different designs from different parties. In any case LSAM design is not frozen and the figures provided in literature may change considerably in the development phase. For this reason differentiating between party landings doesn’t seem to make any sense at this time, and the use of LSAM equivalent is considered to be the simples and safest solution for the type of analysis attempted. The assumed landings are summarized in Table 2-4. LOX and LH2 needed per LSAM ascent vehicle is summarized in Table 2-5 below.

Figure 2-7: The LSAM (ESAS 2005)

36

Table 2-4: Landings on the Moon per party and date Party

Landings/y ear

Period

Type Science on Moon, extensive EVAs Crew shifts on outpost, 4 permanent crew

Source

NASA (early)

2

20182022

NASA (outpost)

4

20232047

International Partners to NASA (early)

1

20232027

Science on Moon

Assumption based on relevant budgets

International Partners to NASA (outpost)

2

20272047

Crew shifts on outpost

Based on ISS historical data

China (early)

1

20302035

Science on Moon

China (outpost)

2

20352047

Crew shifts on outpost

20302035 20352047

Science on Moon Crew shifts on outpost

Third Party (early) Third Party (outpost) Space Tourism/Privat e

1 2

1

20302047

Tourism

(ESAS 2005)

(ESAS 2005)

Assumption – half the NASA effort Assumption – half the NASA effort Assumption Assumption Assumption, comparable to ISS tourism flights

Technical Aspects of Landing Propulsion for the CEV service module (SM) and LSAM ascent module is common in the ESAS report, based on LOX/LCH4 (liquid oxygen/liquid methane), to aid in the aim of Martian exploration as CH 4 can be produced using ISRU on Mars (ESAS 2005). Technology readiness of LOX/LCH4 propulsion is low and in order to achieve the maturity and safety needed for lunar exploration, it has to be tested extensively on the ground and in the early flight of CEV to the ISS. Recently though and after the publication of the ESAS report, there are indications of shifting the propulsion to LOX/LH2 or other than LCH4 storable propellants. For the analysis attempted hereafter, LOX/LH2 propulsion is considered for the LSAM. The amount of propellants needed for a LOX/LH2 reference engine is calculated using the Tsiolkovsky equation (Table 2-5). Mass break-down for the ascent stage of the LSAM can be found in the appendix, along with small discussion on oxidizer to fuel ratio.

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Full Moon

Table 2-5: LSAM and LSAM equivalent ascent stage propellant demand per vehicle Type of propulsion

Oxidizer: fuel ratio (Isp)

Propella nt [kg]

Oxidize r [kg]

Fuel [kg]

LOX/LCH4 LSAM (NASA 2005)

3.5:1 (363)

4 715

3 667

1 048

Using Tsiolkovsky Equation:

mo = m1 ⋅ e

∆u

Isp⋅ g o , mo=initial

total mass, m1=final total mass, ISP=specific impulse, go=9.81m/s2, LOX/LH2 LSAM 8:1 (410) 4 134 3 644 456 (estimate)

2.1.6Demand from Life Support GOX for LSAM and LSAM-equivalent sortie landings is also calculated based on the standard consumption (Wieland 1994). EVA activities of a seven day sortie consider the four hours EVA the first and last day, eight hours EVA for the rest five days by all crew members (ESAS 2005). Table 2-6: Gaseous oxygen needed for a seven-day, four crew sortie mission (ESAS 2005) Activity First and last day EVA Sortie days EVA Total oxygen

Hours crew 4 hours - 4 crew 8 hours - 4 crew

gaseous

Days 2 5

24.5 kg

Cargo flights for the outpost build-up are not considered to use lunar fuel. For the development of the outpost expendable vehicles and landers are expected to be used.

2.1.7Demand from Outposts The transportation vehicle is provided by the US (STS) and all modules were made to fit the STS orbiter. Crew rotation is made by both US (STS) and Russia (Soyuz TMA) vehicles. A similar scheme is assumed to be applied in the lunar outpost development and utilization. The United States already plans to develop the transportation infrastructure and the landing vehicles for the Moon (ESAS 2005). Table 2-7: Outposts considered in the analysis (ESAS 2005)

38

Party

Astrona uts

NASA

4

China

2

Third Party

2

Period 20222047 20352047 20352047

Source (NASA 2005) Assumption, half the NASA effort Assumption

It is not clear what happens to the lunar outpost after 2030 on the official plans (ESAS 2005). Other sources mention it should pass to private funds (Sanders 2007). In any case, for this analysis, landings are considered to happen at the same rate throughout the timeline. At this point we make the assumption that traffic to the Moon and size of lunar outpost, are small and not developing over time, as indicated by historical data of LEO space stations versus expectations.

Outpost Technical Aspects Lunar outposts are considered to support two to four astronauts in a constant basis (four in NASA case, two for China - Third party), and the GOX demand is based on a standard consumption for IVA and EVA (Wieland 1994). No oxygen loss due to airlocks, leaks or other reasons are considered. EVAs are made by team of two (NASA 2005). The weekplan on which the calculations are based can be found in the Appendix A. In Table 2-8 oxygen consumption rates and GOX need per crew and year are summarized. These values are used an as average to estimate the GOX demand of lunar outpost in the analysis scenarios. Table 2-8: Gaseous oxygen (GOX) for outpost demand calculation EVA Rate IVA Rate

Consumption 0.96 kg/crew/day Consumption 0.84 kg/crew/day

Outpost need/crew/year

310 kg/0.31 t

High metabolic rate (Wieland 1994) Normal metabolic rate (Wieland 1994) No leaks, airlocks and other loss, only breathable oxygen calculated

It should be noted that the amount of GOX needed is calculated from other sources to be higher (Sanders 2007), almost double. In this analysis though, the amounts calculated with our method are used.

2.2Demand Graphs Some graphs of the demand along the timeline with the assumptions discussed, are presented in the following pages. Business-cases specific graphs are included in Chapter 5.

39

Full Moon

LOX baseline annual demand

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International Partners landings LOX

CHINA Landings LOX

Tourism/private sector landings LOX

3rd party landings LOX

Figure 2-8: Annual LOX demand in tonnes LOX optimistic annual demand

150.0 140.0 130.0 120.0 110.0 90.0 80.0 70.0 60.0 50.0 40.0 30.0 20.0 10.0

18 20 19 20 20 20 21 20 22 20 23 20 24 20 25 20 26 20 27 20 28 20 29 20 30 20 31 20 32 20 33 20 34 20 35 20 36 20 37 20 38 20 39 20 40 20 41 20 42 20 43 20 44 20 45 20 46 20 47

0.0 20

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100.0

Calendar year NASA LSAM LOX

International Partners landings LOX

CHINA Landings LOX

Tourism/private sector landings LOX

3rd party landings LOX

Mars exploration LOX

Figure 2-9: Optimistic LOX demand in tonnes

40

LH2 baseline annual demand

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International Partners landings LH2

CHINA Landings LH2

Tourism/private sector landings LH2

3rd party landings LH2

Figure 2-10: Baseline LH2 demand in tonnes LH2 optimistic annual demand

LH2 [tonnes]

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International Partners landings LH2

CHINA Landings LH2

Tourism/private sector landings LH2

3rd party landings LH2

Mars exploration LH2

Figure 2-11: Optimistic LH2 demand in tonnes

2.3Constraint from Competitors Now that the demand has been established, it is necessary to assess the threat from competitors. In the case of lunar refueling, competition will come in three forms: direct competition from competitors selling similar products; competition from other resources on the Moon; and competition from Earth-based products. Given that servicing would only start in 20 years time it is difficult to predict what the direct competition will be like, but one expects few competitors given the high investment costs necessary. As of today one can already identify a potential competitor: Stone Aerospace, a Texas-based company, recently announced plans to produce oxygen on the Moon for a Low Earth Orbit (LEO) refueling depot by 2015 (Malik 41

Full Moon

2007). Although the timeline seems unrealistic, the realization of such a project would be a serious threat given that they would use the same production facilities and would only need to diversify to steal market shares. However, this could also be an opportunity for a joint venture to reduce investment and operations costs. Similarly, it is difficult to predict what other resource could drive lunar and space exploration. Potential competitors could be 3He, ISRU metals or solar power, but the technical superiority and financial feasibility of some of these still need to be demonstrated. Concerning competition from Earth, this will depend on the evolution of launching costs from Earth. Importing oxygen and hydrogen could be an alternative to ISRU production, and this is considered in Chapter 5.

2.4Constraints of Lunar Topography Given that the topography is critical to the delivery architecture, it is important to give it special attention. Below is a map of the lunar topography Figure 2-12.

Figure 2-12: Near and far side lunar topography (http://www.lpi.usra.edu/resources/clemen/clemen.html) From Figure 2-12 several geological observations can be made: • The Moon can be crudely identified as being composed of Mare

and Highland regions. The Mare are basaltic plains formed by ancient volcanic eruptions, which given their iron-rich content appear dark to the naked-eye. They cover about 16% of the total lunar surface, mostly on the near-side’s upper-left quadrant (blue/dark-grey regions in map above). Given that they are lowlying, relatively flat regions, covered by four to five m deep regolith (Heiken et al 1991) they are reasonably easily travelable by rovers. It is important to note that not all low-lying regions are 42

Mare, such as the far-side’s (blue/purple/dark-grey zones).

South

Pole-Aitken

Basin

• In contrast, the Highlands are the more elevated regions (green,

yellow, red/light-grey colors) and are generally heavily cratered given their greater geological age (1-4.2 billion years). Apart from steep-sided craters which expose bedrock, the regions are also recovered by regolith, albeit rich in Anorthosites which makes the Highland regions appear brighter to the naked eye. Given that the Moon only exposes its near-side to the Earth, the far-side has served over the years as a meteorite shield, and explains why the far-side is significantly more cratered. Transportation-wise this has two implications: the terrain is impractical to travel through, and the probability of a meteorite impact is greater. • Other geological features exist such as Rilles (lava channels),

Domes, Wrinkle-ridges and Graben, but given that they can only be observed at a local scale, they are deemed to be a challenge only at a local level, which is beyond the “global picture” of this report. Lastly, as can be seen in any Apollo picture, it is essential to remember that in all regions, the lunar surface is jotted with boulders, which makes traveling more sinuous.

• Lastly, it is important to realize from Figure 2-12 the rapid

changes in altitude that are present on the Moon. This is particularly true on the far-side as both the lowest and highest points exist there (respectively -8 km altitude for the South PoleAitken Basin and +8 km for the Leibnitz range (Wlasuk 2000)). Yet again these are parameters to consider when examining possible transportation systems.

2.5Constraints of the Lunar Environment In addition to the topography, several key factors of the environment will affect the delivery and storage architecture, namely: • Illumination levels: • With day/night cycles every 14 days the equatorial regions

prevent the use of solar power on a continuous basis.

• On the other hand, the equator receives up to 70% constant

illumination (Figure 2-4, page 31), which allows extensive use of solar power and incremental build-up of the infrastructure. The only downfall is the low sun inclination which causes problems for judging distances (ESAS 2005). • Strong thermal variations occur as a consequence of the varying

sunlight conditions:

• The equator’s temperature varies between 93 K and 373 K, while

at the equator temperatures are near-constant at 223 K ± 10 K. However, in permanently shadowed craters temperatures are almost cryogenic (50 K), which could cause problems when operating machinery. • The lack of atmosphere:

43

Full Moon • Radiation levels are high without an atmosphere to shield,

requiring all machinery increasing their price. • The

near-vacuum pressurized.

to

condition

be

radiation-resistant,

requires

that

thereby

everything

be

• Lunar dust: • The dust will penetrate everywhere and interfere with machinery,

while its abrasiveness damages fabric.

• Telecommunications issues: • Operations at the South Pole and on the Far-Side require satellite

relays since they are out of sight of the Earth. As part of its outpost infrastructure, NASA plans to deploy such systems and it is assumed these can be used for the delivery and storage infrastructure proposed.

2.6Constraints from Resource Availability Now that the conditions on the Moon have been examined it is necessary to examine the availability of resources. Based on equatorial samples from the Apollo, Luna and Surveyor missions, remote sensing data from Lunar Prospector, Clementine and SMART-1, along with earth-based astronomy scientists have been able to map, on a global level, the resources available on the Moon (Sanders et al. 2006).

2.6.1Resource Availability: Oxygen Table 2-9 below identifies the minerals present on the Moon, and indicates the abundance of oxygen on the Moon. Table 2-9: Lunar surface average regolith composition (Sanders et al. 2006) Regolith Mineral Ilmenite FeO•TiO2 Pyroxene CaO•SiO2 MgO•SiO2 FeO•SiO2 Al2O3•SiO2 TiO2•SiO2 Olivine 2MgO•SiO2 2FeO•SiO2 Anorthite CaO•Al2O3•Si O2

Lunar Surface Average Weight % [Sub-wt %] 15% (98.5%) 50% (36.7%) (29.2%) (17.6%) (9.6%) (6.9%) 15% (56.6%) (42.7%) 20% (97.7%)

As indicated in Table 2-9, oxygen is the most abundant element on the Moon, which bodes well for lunar-oxygen refueling business. However, 44

it is present as a chemical compound and thus must be chemically extracted, which potentially requires large amounts of energy. The only differences between the Mare and Highland regions are the types of minerals present, which implies that different production methods should be used according to the region (see Section 2.7). In addition, different by-products could be obtained, but given that this report focuses solely on storage and delivery and not on production, this aspect becomes a secondary driver in terms of choosing the location for the infrastructure. Lastly, the depth of the regolith, respectively at 4-5 m in the Mare and 10-15 m in the highlands, implies that in terms of extractable volume, the highlands are more interesting, although the more rugged terrain mentioned earlier could pose a greater problem.

Resource Availability: Hydrogen In contrast to oxygen, hydrogen is absent from the chemical composition of the regolith (Table 2-9). Instead, as confirmed by Apollo samples, hydrogen is found as a solar-wind implanted volatile, as shown in Table 2-10. Table 2-10: Concentration of solar-wind implanted volatiles in regolith (Sanders et al. 2006) Solar-Wind Implanted Volatile Hydrogen (H2) Carbon (C) Nitrogen (N2) Helium (He) 3 He

Concentration [ppm] 50 – 100 100 – 150 50 – 100 3 – 50 0.004 – 0.02

As it can be observed the proportions are small, making extraction difficult.

Resource Availability: Water Ice Remote sensing measurements from Lunar Prospector, Clementine, SMART-1 and from Earth have shown much higher hydrogen concentrations than average at the poles and in permanently shadowed craters, as shown in Table 2-11 and Figure 2-13 below. Table 2-11: Hydrogen volatile concentration in polar regolith (Sanders et al. 2006) Volatile Hydrogen (poles & permanently shadowed craters)

Concentration [ppm] 1 500 ± 800

45

Full Moon

Figure 2-13: Epithermal neutron flux at lunar poles (http://lunar.lanl.gov/pages/water.html) Note: Hydrogen concentration is inferred from this data The fact that this hydrogen is present at the poles and permanently shadowed craters has sparked the controversy that this hydrogen could be in the form of water ice (H2O). The reasons are the following: • It is known that the Moon has no stable surface water due to the

lack of an atmosphere that would prevent water molecules from dissociating upon exposure to the Sun’s ultraviolet radiation. However, permanently shadowed craters, particularly at the poles, receive light only from space and the lunar interior, which would allow for “cold traps” to exist where ice could remain stable for indefinite time spans.

• This water could have been deposited within the craters by the

very same meteorites that created them upon impact, and which are known to be rich in water ice. • The

controversy is two-fold: firstly, whether this detected hydrogen compound is water at all; secondly, upon the detection methods used for estimating the quantity (see Appendix A for discussion on the controversy). Lunar Prospector and SMART-1 purposely crashed in such regions in order to confirm the presence of water ice but results were unsatisfactory. Future missions such as NASA Lunar Reconnaissance Orbiter (LRO) and ISRO Chandrayaan-1, both in 2008, aim to confirm this matter, but as Dr. Bill Larson from NASA commented during a teleconference on the 23rd April 2007 “Until we get a robot to go there and scoop the ice out and bring it back to Earth, the controversy will remain”. No such missions are currently planned.

If this polar hydrogen was indeed ice, this implies that it would represent 1.5±0.8 wt% of the regolith, representing a total quantity of 2∙109 tonnes of water if all the hydrogen were water, or 200∙106 tonnes if considering only permanently shaded areas (Javier and Michael 2006), both of which are sufficient for commercial exploitation. From

46

Figure 2-13, North Pole could hold the most water given that it has the highest hydrogen concentration, but South Pole has the largest craters, thus the largest permanently shadowed area. The latter theory is deemed the most accurate by the scientific community and explains why the South Pole is preferred over the northern one for construction of an outpost. Finally, given that water is composed of both oxygen and hydrogen, and that electrolysis is a well-known relatively low energy process, implies that perhaps only exploitation of water ice is necessary for production of oxygen and hydrogen, at the detriment of extracting oxygen or hydrogen volatiles from the regolith (this is discussed in Section 2.7).

2.7Constraints of Production Recalling Chapter 1 the topic of this report focuses only the storage and delivery aspects of lunar surface refueling. However, given that storage and delivery are intrinsically linked to production, it is necessary to perform a brief overview of production, such that accurate production prices and production rates are used as inputs. For oxygen and hydrogen production, NASA is currently developing the following techniques (Sanders 2007): • Oxygen

extraction via chemical reduction of the regolith. Different methods apply to different minerals, which appear in varying quantities across the Moon. The current plan is to deploy a pilot-plant in 2023 and test it until 2027, date at which the technology will be mature enough to produce 10 tonnes of oxygen per year, per production unit (Sanders 2007).

• Hydrogen

and ice-water volatile extraction from regolith. Hydrogen-volatile extraction can be performed anywhere on the Moon, while water extraction can only occur in permanently shadowed craters at the poles. Deployment plans depend very much on the confirmation of the presence of water on the Moon by the Robotic Lunar Exploration Program (RLEP) in 2008 and lunar sortie missions in 2018 (Sanders 2007).

These methods will be explained in further detail in the upcoming sections, but it is important to remember that several hurdles remain in the development of lunar ISRU due to incomplete knowledge of lunar regolith properties related to excavation and transportation; and the high autonomy and reliability required for the machinery.

Production: Oxygen from Regolith According to Sanders (2007) mining oxygen from anywhere on the Moon will require: • Excavating regolith from the mining site and transporting it to

processing chambers. As explained by Sanders (2007) the latest design by the Colorado School of Mines and the Northern Center for Advanced Technology (NORCAT) have demonstrated high excavation rates up to 150 kg of regolith per hour for a total mass of less than 50 kg (see Figure 2-14 below), to which one

47

Full Moon

must 150 kg for support hardware allocations and on-board power system. • Excavation needs to be followed by pre-processing duties such as

regolith sorting, crushing and/or beneficiation2 • Chemically, electrically and/or thermally extracting oxygen from

the mineral compounds present in the regolith

• Removing and recycling the reagents used • Collecting and purifying the oxygen produced for further storage • Removing waste products out of the processing plant • Potentially continue the processes for extraction of by-products

such as silicon, iron, titanium, etc. However, recalling that this report focuses solely on storage and delivery and assumes that production is performed by NASA (or one of its sub-contractors), production of by-products are not relevant to this report.

Figure 2-14: Colorado School of Mines excavator prototype (NASA 2007) Three extraction processes are currently under laboratory development and evaluation at Sanders (2007), all of which have been extensively studied in the “FERTILE Moon” ISU Team Project report 2006: 1) Hydrogen Reduction of Ilmenite, which is only worthwhile performing in the Mare regions where Ilmenite proportions are high. This is not compatible with setting up operations at the South Pole as desired in Section 2.1.2. Furthermore, the chemical process requires an extra input of hydrogen, which if not produced locally must be imported from Earth. 2) Carbothermal Reduction of silicates using Methane or Carbon Monoxide. As silicates are abundant especially in the Highlands and at the poles, this is a suitable method for oxygen production at the South Pole. However, if one observes the equations involved one will realize that if hydrogen is to be produced in addition to oxygen a constant supply of methane or carbon monoxide is necessary; or if all the methane/carbon monoxide is

2

In the mining industry beneficiation refers to a variety of processes where extracted “ore” from mining is reduced to particles that can be separated into mineral (for further processing) and waste. 48

to be recovered in the process, then a constant input of hydrogen is required. 3) Electrowinning of the regolith using Molten Silicate Electrolysis. As opposed to the previous two methods which are purely chemical, this is an electro-chemical process and requires no additional reagents. This is a method of choice as all types of regolith minerals can be processed, implying that this method is suitable for production at the South Pole. Table 2-12 gives a summary of their properties based on FERTILE Moon (2006) and the latest technology developments at NASA (Sanders 2007): Table 2-12: ISRU processes for oxygen extraction from regolith (FERTILE Moon 2006; Sanders 2007) ISRU Process (Oxygen from regolith) Hydrogen reduction of Ilmenite (FERTILE Moon 2006) Carbothermal reduction of Silicates (Sanders 2007) Molten Silicates Electrolysis (FERTILE Moon 2006)

Regolit h excavat ion rate [kg/h]

Reagen t Output

Specifi c mass3

Specifi c power4

Efficien cy5

0.15

1.93

1.41%

~0.1

1.35

~14%

0.065

1.5

21.4%

H2(g) 150 O2(l)

15

CH4(g) or CO(g) O2(l) None

10

O2(l)

As it can be observed from Table 2-12, molten silicate (SOx) electrolysis is the ideal method as it requires the lightest production plant and has the best extraction efficiency. At an excavation rate of 10 kg of regolith per hour this represents, for a production of 10 tonnes of oxygen per year, excavating a football/soccer field to a depth of 0.4 cm (0.04 m) each year (based on example given in Sanders 2007). Considering that the maximum baseline demand is of the order of 40 tonnes of oxygen per year (Section 2.1.2), this means that four football fields would have to be harvested per year to a depth of 0.4 cm (0.04 m), which is an insignificant surface area in industrial terms, and implies that production can remain close to the base.

3

4 5

MassFacility (t ) ProductionRate (t / year ) Plant Power ( kW ) Production Rate (t / year ) Mass Product (kg ) MassFeedstock (kg )

49

Full Moon

In terms of power supply, building of the infrastructure at the poles implies that there is potential for 70-80% yearly constant illumination, making solar panels a solution of choice, which is in accordance with NASA’s outpost-building plans (NASA 2005, 2007). From FERTILE Moon (2006), a solar power system on the Moon will produce power at a minimum rate of 1.82 kW/m2 and will weigh 3.02 kg/kW. Thus, a molten silicate electrolysis plant producing 10 tonnes of oxygen per year will weight 650 kg and require 15 kW of power, to be provided by 8.24 m2 of solar panels, weighing 45.3 kg. It follows that as the demand grows new production plants or expansions will have to be built, along with new power supplies.

2.7.1Production: Oxygen & Hydrogen from Water Ice As described in FERTILE Moon (2006) the actual process of water ice extraction is relatively simple, requiring heating of the regolith at low temperatures (< 873 K) for the water ice volatile to escape. It is then captured and made to condense as liquid water. This water is then electrolyzed into hydrogen and oxygen, which compared to the previous three methods, is an energy-cheap process. However, the major difficulty remains in operating machinery in the high vacuum, 40-100 K temperature environment offered by the permanently shadowed craters at the poles. Based on Blair (2002) and FERTILE Moon (2006) the properties of such a production facility are summarized in Table 2-13 below: Table 2-13: Processes for oxygen and hydrogen extraction from ice (Blair 2002; FERTILE Moon 2006) ISRU Process (Oxygen and Hydrogen from ice) Water extraction from regolith (FERTILE Moon 2006) Electrolysi s, liquefactio n& liquefier radiator (Blair 2002) Total extraction process

Required regolith excavation rate [kg/h]

210

Outp ut

Specific Mass [per tonne H2O)

Specific Power [per tonne H2O]

Efficien cy

H2O(l)

0.039

16.4

1%

H2(l) N/A

210

11.1% 0.01061

0.8

O2(l)

88.9%

H2(l)

0.111%

O2(l)

0.0496

17.2

0.889%

50

From Table 2-13, the production plant as a whole is slightly lighter than the oxygen production plants, at the expense though of a 12 times higher energy consumption. However, the main advantage remains that both oxygen and hydrogen are produced, which may outweigh the energy issue given that solar power can easily be harnessed at the poles. In terms of harvesting area, a 10 tonnes production of oxygen per year would require excavation of 21 football fields to a depth of 0.4 cm (0.04 m) (or one football field to a depth of 8.4 cm - 0.084 m). This would also produce 1.25 tonnes of hydrogen in the process.

2.7.2Production: Hydrogen Implanted Volatiles

from

Solar-Wind

The extraction of solar-wind implanted hydrogen follows the same process as extracting water ice volatiles; except that the regolith needs to be heated up to 1 173 K (see FERTILE Moon 2006 for further details). For that reason the method is more energy expensive as can be seen in Table 2-14 below. It is important to note that the energy consumption per kilogram of regolith is highly dependent on the assumed concentration of hydrogen entrapped. Numbers given here are for the equatorial concentration of 1 500±800 ppm. Table 2-14: Extraction process of solar-wind-implanted H2 from regolith (FERTILE Moon 2006) ISRU Process Hydrogen Extraction

Outpu t

Specific Mass

Specific Power

Efficien cy

H2

3.4

124

0.004%

Table 2-14 clearly indicates that extracting hydrogen from solarimplanted volatiles is an energy consuming process. Producing 1.25 tonnes of hydrogen per year would require 155 kW which is more than ten times more the energy needed to produce ten tonnes per year of oxygen using molten silicate electrolysis mentioned earlier; not to mention that the production plant would weight 6.5 times more! This suggests that perhaps extraction of hydrogen in this manner is uneconomical, as claimed by Blair (2002).

2.8Findings: Production Scenarios Based on this research, it can be assumed that NASA will build a base at the South Pole and that the other countries will follow. Since the majority of the demand is at the pole, it was decided to produce at the poles too, thereby making use of NASA facilities and satisfying the US philosophy of fostering private enterprise. With that decision in mind, two production scenarios were identified depending on the presence of ice: •

In the case that ice is confirmed, oxygen and hydrogen will be produced from melting of the ice and electrolysis of the water

•

If there is no ice, only oxygen is produced using Molten Silicate Electrolysis as it is efficient and requires no reagents. It is thus assumed that hydrogen is imported from Earth

51

Full Moon

In either case it was demonstrated that only a small area needs to be harvested on a yearly basis to satisfy demand. In addition, based on the agency roadmaps two demand scenarios were identified: a baseline scenario and an optimistic one. This implies that in total four scenarios are to be considered as shown in Table 2-15 below. Table 2-15: The four economic scenarios to be analyzed within the report Demand

Baseline

Optimisti c

Resource Availability Ice No-Ice Oxygen & Oxygen hydrogen production production <40 tonnes/year <45 tonnes/year molten silicate total electrolysis water electrolysis Oxygen & Oxygen Hydrogen production production <140 <165 tonnes/year tonnes/year total molten silicate water electrolysis electrolysis

It is then the role of the system architecture to satisfy these scenarios within the constraints laid by the lunar environment, and determine their financial viability. Lastly, it is important to note that hydrogen extraction from solar-wind volatiles has been omitted from the “No-Ice” scenario, since it would require an extremely expensive infrastructure to satisfy a small demand of the order of five tonnes per year for the baseline scenario and 20 tonnes for the optimistic scenario. In the “No-Ice” case, it is assumed that hydrogen is supplied from Earth, but distributed using the same oxygen delivery infrastructure.

52

_____________________________________Chapter 3

3 Architecture Assessment

LOX and LH2 storage and delivery solutions for the Moon are not mere transfers of Earth or microgravity space technology, however one should learn from the meticulous lessons previously learned. Materials, energy requirements and the challenges specific to the lunar environment have to be assessed. In this chapter, the challenges of storing oxygen and hydrogen on the lunar surface are discussed. One must determine which materials can be used to survive this extreme environment, whilst providing a cost effective storage capability, keeping the product contaminant free and ready for use. Delivery must make the storage solution useful, mobile and adaptable to market demands it must serve. In order to do this, analytical tools will be used. Those will be based on criteria one can predict, such as complexity of devices and technology readiness levels. There are specific ideas and solutions regarding storing and mobility on the Moon and these obviously should be considered when identifying potential solutions for the currently planned and future direction of human utilization of the Moon. Building on this, the trade space of how to mobilize the storage solutions and then find a way to trade off merits and disadvantages of future term solutions have to be defined. This is a problem of predicting future developments and technology readiness, compared to the standard approach of trading off existing technologies on their historical performance and profitability. One must analyze the harsh realities of storage and transportation on the lunar surface of hydrogen and oxygen, specifically focusing on the challenges of the lunar environment. The following section discusses the challenges of storage and delivery in the various forms, be that gaseous, cryogenic liquid, or perhaps more simply as liquid or ice water, with special regard for the specific problems associated with the lunar environment and marries this storage form with a workable delivery solution.

3.1Challenges of Lunar Environment The delivery and storage systems on the Moon hold specific challenges. The Moon’s reduced gravity, the impinging radiation, near vacuum environment and the rapidly changing thermal regimes must all be considered. Lunar dust has been shown to be one of the biggest

53

challenges facing lunar exploration. NASA considers it to be one of the harshest challenges humanity must tackle in returning to the Moon and using it successfully (Allen 2007). Lunar seismic activities and magnetic field effects on lunar surface systems are regarded as negligible and are not discussed further.

Reduced but Apparent Gravity Gravity on the Moon is 1/6th of that of Earth. This reduces the structural requirements of all equipment and makes transport issues easier and more energy efficient. For the storage and delivery of the propellants, convection and evaporation although existing may need augmentation by mechanical systems to overcome the surface tension effects and laminar flows in fluids (pumps, vacuum feed, etc).

Radiation Environment The Moon is not protected by a dense atmosphere and a magnetosphere as Earth. This allows electromagnetic and ionizing radiation to reach its surface. Electromagnetic radiation emitted by the Sun is responsible for thermal variations. Ionizing radiation consisting of solar wind, solar cosmic rays and galactic cosmic rays should be considered when selecting materials and when designing automated systems (Eckart 1999).

Lunar Atmosphere The lunar atmosphere, or exosphere, is about 14 orders of magnitude less dense than Earth’s atmosphere, thus vacuum conditions for materials and thermal design (Eckart 1999). Vacuum, as with the space environment of LEO, imposes stronger design criteria on tanks solutions. Non-metallic materials outgassing and leaking are the most pronounced.

Temperature The Moon creates, in general, the very same thermal conditions as a spacecraft has to deal in Earth orbit. The day-night cycle is 14 Earth days. The Moon also has specific cases such as deep craters which shield all sunlight around the year. Shadowed polar craters like the Schackleton crater have an average temperature of 40 K. These prolonged cold periods can be used to aid the system to be more efficient, but, impose specific design requirements different than tanks designed for orbit, as they are not optimized for long periods of constant eclipse. Table 3-16 below lists the estimated average temperature of different locations and their monthly range. Table 3-16: Lunar temperature range taken from (Eckart 1999; Heiken 1991) Polar Shadow Othe ed r Craters [K] [K]

54

Equatorial Fron t

Back

L imb

[K]

[K]

[K]

MidLatitudes [K]

Average Temperatur e Monthly Range

40

220

254

256

255

220
0

10

140

140

140

110

Meteorites Lunar surface is under constantly impacted by meteorites and micrometeorites. The velocity of these objects hitting the Moon is in between 13 km/s and 18 km/s. Although more data are required in order to estimate actual risks of collision, it is obvious that such collision would be disastrous to a storage or delivery system (Eckart 1999). Debris shields, energy absorbing materials and damage tolerant structures, like the ones used in manned ISS modules today, could be used in the critical components of the system under study.

Lunar Dust Lunar dust has many damaging impacts on vehicles designed for operation in space. On the other hand, regolith can be used as radiation and meteoroid shielding as well as thermal insulation. One problem is the sharpness of the individual grains because of the lack of erosion effects, as there would be on Earth. Sharpness makes them damaging to moving and sealing parts. This scratching of intimate equipment compromises overall lifetime. Concerning continuous operation in the lunar environment, lunar dust presents a highly abrasive challenge and can be highly detrimental to certain materials. Sealing materials, solar arrays, optical properties of surfaces and thermal coatings can all be impaired by lunar dust.

3.2System Design Method The delivery and storage systems must be designed for vacuum environment. The choice between systems should not be limited to the characteristics of individual technologies alone. These systems interface with each other and with the production facilities and consumers, which define their function. To be meaningful for agency plans that encompass decades, the delivery and storage systems must also be responsive to long term changes in agency goals and in the market or the scale of operations. To make the decision-making process less haphazard for our delivery and storage systems, we considered using qualitative and quantitative decision matrices and linear network modeling. Our system selection process went through the following stages, as illustrated in Figure 3-15.

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Figure 3-15: System selection methodology

3.2.1Qualitative Decision-Making To choose which systems to research in more detail, a decision method was developed to compare all the options using qualitative criteria with a simple better/worse ranking. This method is useful when little quantitative information is known about the choices, but there is some general knowledge of each system. Using a matrix format reduces prejudices towards “favorites” and confusion between long lists. The criteria were selected based on qualities that matched perceived consumer needs and business drivers. The spreadsheet tallied the results in a table illustrating the relative ratings of each option and a net “ranking”. This method was useful for quickly eliminating the extremely unsuitable options from a large number of choices without detailed evaluations of individual criteria. Based on this method, there is no “best” solution. Each of these options may be optimal for one or more consumers in one market phase or another. To choose rationally between these disparate elements for the most reliable and economical solution, more quantified information is needed. To that end, a quantitative method was also developed.

3.2.2Quantitative Decision-Making As more details of the systems became known, comparisons of options were made quantitatively. While there is not enough data available in literature for comparing, the complete life cycle costs of systems that have never been built, some available quantities can be selected to represent the key criteria identified by business and technical needs. To evaluate the quantities according to the relative importance of the parameters, they are first normalized and then weighted. A sample calculation used in the Quantitative Decision Tool is shown in Equation 1.

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Equation 1: Sample parameter calculation Parameter OperationalEfficiency ≡

Quantity

Normalize d

Weighted

power ( kW ) 1 × ×3 mass (kg ) max ( Powersp )

The weights were determined by polling two groups of people – the team project members and the ISU faculty. The standard deviation of the weights was higher for the team members than for faculty. Finally, this method is only suited to comparing the options for one system in isolation. As stated in the beginning, the interfaces between system elements (especially storage and delivery) must be considered as part of a rational design.

3.2.3Futures: Supply Chain Modeling A complex system with many governing criteria is best chosen based on supply chain modeling. Supply chain modeling is a computer method for representing real world scenarios, predicting outcomes and optimizing a complete system. A supply chain model describes the system in some detail, including a consistent “utility function” (usually cost or performance metrics) for each system component. So, the final step for selecting a lunar cryogenic storage and delivery system in this project was to combine the team knowledge in a concurrent engineering session. While the system selection process did not include optimization, by using a rough qualitative method to eliminate the worst design choices early and a quantitative method to compare the remaining options it was possible to narrow the scope of work before beginning concurrent engineering and to develop useful tools step-wise along with competence in designing these tools before the most complex step of visualizing the system with a concurrent engineering approach.

3.3Assessment of Storage Options on the Lunar Surface Storage systems must meet the challenges of a lunar environment.

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Figure 3-16: Storage element in the system architecture

3.3.1Challenges of Stored Materials There are challenges inherent in coming up with a useful oxygen and hydrogen storage solution on the Moon. An effective solution must be delivered. It has to be cost effective based on minimizing energy input, launch mass, and development cost, whilst maximizing the ability to store the produced or supplied products on the Moon to maximum efficiency. First the three main types of storage have to be discussed: gaseous, liquid (cryogenic), or water (combined, liquid or ice). Something consistent with all lunar challenges is that there are no truly ideal temperature ranges for any phase of storage, but on the other hand, energy is needed to keep the chosen storage method within the desired temperature range.

Gaseous Gasses are far less dense than the liquid phase. Storing as a gas means that less can be stored in a given volume. Since cryogenic liquids vaporize too in the tank, dealing with cryogenic storage means dealing with gaseous phase as well.

Cryogenic Liquids Many space missions require large quantities of cryogenic propellants.High-energy propulsion systems will be required for space based transfer vehicles and manned lunar and Mars missions (Gaby n.d.). Due to the high probability of their use, there continues to be a strong interest in developing the technologies necessary for the management of cryogenic fluids in space environment, although there are significant challenges involved in their long-term storage.

A Definition of Cryogenic Storage Cryogenic liquids are the ones stored between 7 K and 207 K with boiling points below 183 K. For some elements, lower temperatures can make them storable in the solid phase. The reason for doing this is the increased density of the stored substance, allowing smaller tanking solutions rather than larger vessels for liquids and gases. Hazards associated with cryogenic liquids include human exposure (cold burns / frostbite), material compatibility, high pressure,

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explosions and implosions, toxicity and asphyxiation hazard (Desert Research Institute 2007). At temperatures above the critical temperature, it is impossible to keep cryogenic liquids under pressure (see Appendix B for vapor pressure and critical points of oxygen/hydrogen). The standard method used to deal with such cryogenic fluids, is to permit a fraction of the liquid to boil-off. Through this method the latent heat of vaporization taken from the liquid keeps it at the appropriate temperature in the liquid state (Turner 2005). Cryogenic storage introduces technological problems regardless of the tank location (Earth, orbit, Moon). Energy input is required in the system to keep it within the temperature margins that keep fuel liquid and allow for transfer through piping. Storing cryogenic liquids adds extra hazards to an already complex problem and the hazards to the crews that will come into close proximity with these tanks need to be addressed. However this decision can be based on the likely location where the tanks are placed. If the tanks are to be positioned in the dark of craters of constant shadow, the 40 K (Luna Gaia 2006) temperatures can be used to keep LOX and LH2 almost passively, (Blair 2002). There will be a need to actively cool the LH2 a further 10 K down to 30 K; however, the temperature needs of the LOX are provided by the environmental conditions. To put this in context, to store water in the same environment requires raising its temperature by 237 K. Before choosing the best solution for storing of gasses on the Moon, the tried and tested ways the commercial and space sectors tackle the problem on the Earth have to be identified.

3.3.2LOX Properties and Considerations Liquid oxygen a strong oxidizer. It will react with nearly all organic materials and metals, usually forming an oxide. One of the primary concerns in the handling of cryogenic oxygen is the possibility of a combustion reaction if the oxygen comes in contact with incompatible materials; however, on the Moon this is less likely due to the lack of ambient atmosphere. It can also cause many structural materials to become brittle. LOX density is similar to that of water at 1.41 kg/m3.

3.3.3LH2 Properties and Considerations Liquid hydrogen is an energetic fuel with the lowest atomic weight of any substance. Its low density leads to large tanks, which a lunar storage infrastructure must accommodate. Hydrogen is also known to leak through materials due to its small molecular size. Organic materials used in seals are particularly susceptible to these leaks, therefore materials must be used that will ensure containment of hydrogen in all storage tanks and transfer mechanisms, to alleviate any and all safety concerns, and to prevent significant transfer losses (Turner 2005).

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3.3.4Thermal Control Thermal control and regulation are necessary to assure proper operation of the storage system. In considering thermal control technologies, there are two main approaches to the problem: active and passive. In short, active requires energy for operation, but offers a greater range and tighter temperature control. Passive, on the other hand, essentially uses intelligent designed technologies to transfer heat and/or exclude heat naturally based on radiation, reflection or conduction of the structures. The ability to regulate temperatures is more limited due to the environmental regime the system operates in, but overall costs less to run and can be more durable due to no powered parts. In order to store LOX and LH2 in a cryogenic state, tanks must be able to maintain a temperature of 20 K for LH2 and 90 K for LOX. As a result, the temperature of the storage tank must be regulated and controlled. Cryocoolers are devices needed to reach cryogenic status. However, the cooling of the tanks can be favored by the lunar environment since locations like the lunar South Pole have ambient temperatures of 40 K6. In such regions, less power would be needed to maintain the tanks at proper temperature. Placing the tanks in other regions of the lunar surface would require greater amount of power to cool down the tanks, as the difference between day and night temperatures could vary as much as 250 K. Cooling failure of the tanks can lead to pressure increase over the tank rupture limit. Storage location depends on following criteria: • Physical parameters like: temperature fluctuations on storage

sites, pressure and radiation • Distance or nearness to production and fueling sites • Operational parameters like: accessibility and dust.

To determine storage locations, it is needed to analyze the above mentioned criteria based on two possible scenarios regarding water existence on the lunar South Pole. If water ice exists on the lunar South Pole then all the operations, from production to delivery, would be executed on the lunar South Pole, including storage. If no ice/water exists on the South Pole, storage location will change from the lunar South Pole, since production and customer exploration sites in the long term would be near the lunar equator. In case of using only South Pole as storage location, LOX and LH2 will be stored, but not water. The best possible storage location for LOX and LH2 at South Pole is the Shackleton crater since it has both permanently illuminated and shadowed areas. A detailed analysis of eclipse durations showed that two areas close to the rim of Shackleton crater, only 10 km apart, are collectively illuminated for more than 98% of the time (Fristad 2004). These regions are close to several areas of probable permanent shadow that may harbor ice deposits. Storage tanks are designed to be transportable and usable anywhere on the surface of the Moon. To satisfy design requirements, tanks have to be able to sustain all possible temperatures in area between the South Pole and equator and possibly beyond equator towards the North 6

By ‘ambient’ is meant, the temperature that a body reaches under the conduction-radiation dominated heat transfer in vacuum conditions. 60

Pole. As shown in Table 3-16, temperature on the South Pole fluctuates over a range of 183 K while in the equatorial area fluctuation is even more, up to 250 K. Furthermore, to prevent boil off greater than needed (0.1%), temperatures of fluids should not to vary more than 36 K for LOX and 6 K for LH2. These high demanding conditions require the use of both passive and active cooling. As LOX and LH2 have different physical properties, thermal control for LOX tank will be somewhat different than for LH2. LH2 tank will need stronger active cooling system than LOX tank what will result in higher energy consumption. Following, is a description on the design of the LOX tanks to satisfy the previously mentioned requirements.

Passive Thermal Control As mentioned previously, passive thermal control system do not require power to be functional. In that perspective, the general recommendations for cryogenics storage are as following: • Lightweight, low thermal conductivity cryogenic tank struts and

support concepts • Low

thermal conductivity cryogenic tank penetrations, i.e., instrumentation feed through, feed-lines, vent lines

• Lightweight, insulating thermal protection schemes for use on

the Moon • LH2 tank needs to be protected from LOX tank in order to achieve

passive cooling

• Single shade can offer passive cooling protection for LOX tanks

(NASA 2003) The proposed solution will depend on type of tank materials however, it is suggested to have doubled and in some cases, triple tank walls. That is due to the brittleness of materials at cryogenic temperatures, as mentioned above, and the level of vaporization which should be between 0.1% and 0.5% (Tolyarenko Personal Conversation 2007). Because tanks will be very complex it will be required to use advanced passive control management. Example of such a system is the Cryogenic Operation for the Long-Duration (COLD) system. It comprises of insulation plus some other features explained below: • Vacuum Insulation Panels (VIP) • Variable Density Multi-layer Insulation (VDMLI) • Propellant Positional Management Devices or System (PPMD or

PPMS)

• Sun Shields (SS) • Settled Pressure Control (see later)

As Kutter (2005) states, “COLD technologies enable passive 0.01% per day boil off. COLD technology is affordable for early phases of long term storage and can be combined easily with active thermal control to support long term conditions exceeding one year. Combining both systems could enable thermal protection for LOX and LH2 tanks which in theory would last for an infinite lifetime. Table 3-17 below gives

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lifetime of each of systems’ components. The COLD technologies inserted into ICES will parallel the mission-duration needs (Kutter 2005). Table 3-17: COLD technologies results from ICES: Mandatory = X, Helpful = O (Id.). COLD Technologies

D ays X X O O

Mission duration Week Month Year s s X X X X X X X X X O X X

Vacuum Panel Insulation Settled Pressure Control Vapor-cooled Points Variable Density MLI Propellant Position O O X Management System Sun Shield O O X Enhanced technologies Pre-launch Sub Cooling O 0g Pressure Control O Cryocooler O

>Year X X X X

X

X

X

X

X O O

X X X

The specific characteristic of some of COLD components are described in the following sections.

Variable Density Multi-layer Insulation (VDMLI) VDMLI is one of the best thermal insulations in a vacuum, having no problems relating to density control and performance, application labor, and difficulty covering small and large scales. It is based on micro-machine or micro-molded structures as stated by NASA (2006). VDMLI is an improved insulation which “should provide lower thermal conductivity, lower specific thermal conductivity, vacuum compatibility, layers inherently attached to each other that support themselves, efficient assembly and provide structural reliability” as stated by (NASA 2007). Moldings techniques used for VDMLI guarantee low thermal conductivity for materials and low outgassing. VDMLI is improved MLI so it will still have interior and outer layers. As it is very important to minimize the thermal impact from outside on fluids inside, a very low absorptivity/emissivity (α/ε) ratio will be needed. That can be achieved by using silvered Teflon as a surface finish on the outer blanket layer as described by Gilmore (2002). As the author explains further “when Teflon is used, however, it should be bonded to a durable support material such as Kapton because Teflon will lose all mechanical strength over time as a result of the effects of charged particles and thermal cycling”. Interior layers should be made of aluminized Mylar (aluminized from both sides) to enable low emittance and generation of minimal amount of particulate contaminants as stated by (Gilmore 2002). As suggested by (Tolyarenko Personal Conversation 2007), Kapton and Dacron could be options too. Furthermore as an insulation system one should not neglect the micro sphere insulation system consisting of microscopic glass spheres within the inner vessel, vessel whose outer skin is made out of stainless steel

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or aluminum. A heat transfer analysis will show that there has to be a balance between the cryocooler capacity and the insulation thickness mainly due to the shifting day night temperatures on the Moon.

Sun Shields Mobile tank requires sun shields which can protect it both at the South Pole and equator. As a solution two types of Sun Shields will need to be used: one reflecting light from Sun and the other reflecting reflected sunlight from the Earth. One solution proposed for this issue is to use conical radiators which reflect sunlight directly into the space (Chui et al. 2005). Each of these conical radiators would have inner radiator, intermediate shield and conical sunshield. It is important to mention that both types of radiators would be placed on the tank. One big problem for Sun Shields is lunar dust. During Apollo missions astronauts had to clean it every couple of hours. Sophisticated and long lasting solutions would have to be provided before Sun Shields could be used.

Settled Pressure Control As mentioned above, it is necessary to use pressure control to keep cryogenic fluids in liquid state. LOX needs to be maintained at a pressure of 5.03 MPa7, and LH2 at 1.30 MPa (Harvard University 2003). Thus exceeding the critical pressure would cause rapture of the tanks and explosion may occur. Maintaining high pressure, especially for LH2, will decrease the cooling energy requirement, but needs thicker walls to withstand the pressure, thus higher launch mass is required (see appendix B for the vapor pressure of LOX and LH2). The trade off between higher pressure and lower cooling requirement would have to be examined more thoroughly. Necessary equipment includes sensors which will be described later in this report.

Active Thermal Control Active thermal control includes mainly cryocoolers. Refrigeration Systems are active thermal control systems with the purpose of cooling liquids to achieve liquefaction point, zero boil-off and densification of cryogens. Some the proposed requirements include: Cryocooler systems with cooling capacity greater than 10 W in the 10 K-40 K range Small scale tank pressure control and/or integrated tank boil off control and liquefaction technologies for liquid oxygen or liquid hydrogen As an example for such system, analysis will be made on the LockheedMartin four-stage cryocoolers developed for JPL’s Advanced Cryocooler Technology Development Program (ACDTP). • As previously mentioned, cryogenic tanks need less energy

consumption to regulate and maintain cryogenic temperature when placed in the permanent shadow of the lunar South Pole crater. Power consumption includes the power to transfer and transport liquid oxygen and liquid hydrogen to and from the storage tanks. Thus a robust power infrastructure is necessary. 7

Pa = N/m2 63

Some options for the power could be regenerative fuel cells combined with solar arrays, fission reactors and Radio Thermal Generators (RTGs).

3.3.5Materials With regards to materials, aluminum-lithium (Al-Li) alloys seem to be the preferred solution. They are lighter8 than conventional aluminum alloys and there is heritage from use in the tanks of the Space Shuttle and aviation industry. Research has also shown that there is also preference for Teflon, stainless steel for hydrogen tanks since it is resistant against hydrogen brittleness. Also composite materials especially for oxygen tanks can also be considered as solutions. The usual shape of the tank is cylindrical with double wall. This inter-vessel space is often used for insulation. There is an inter-tank shielding which reduces the heat transfer from the oxygen tank to the hydrogen tank. Shades shield are highly recommended and useful for protecting the propellant tank from solar or planetary albedo and infrared radiation. These shade shields work best when used together with Multi Layer Insulation (MLI) blankets. When deciding to use Teflon as a shell or outer casing material a desired option is to use FOSR (Flexible Optical Solar Reflector) as an MLI cover sheet. FOSR provides a low solar absorbance and it protects from tearing and gives strength to Teflon due to its Nomex scrim (a material used in the MLI layers for tear limitation). Table 3-18: Summary of tank-suitable material properties (www.matweb.com) Material Al-Li alloys (AA8XXX) Composite Materials (carbonepoxy, aviation fibers – space rated, low outgassing resin) Teflon Stainless steel (for cryogenic applications)

Density [kg/m3]

Thermal Conductivity [W/m/K]

2 550

93.5 (T = 298 K)

1 750

2-21

~500 (fiber dependent)

2 200

0.23

25 – 28

8 030

16.3 - 21.5

up to 1 300

3.3.6Cryogenic Storage Technologies

Tensile strength [MPa] ~450 (temper depending)

Methods:

Existing

Between 1970 and the mid-1980s a high percentage of cryogenic research was conducted at NASA Lewis Research Center (LeRC) focusing on cryogenic storage, supply and transfer in support of deepspace exploration programs. Research and testing involved LH2 tank 8

Low density Lithium (Li), decrease the density of the aluminum-lithium alloy (by ~10%). 64

thermodynamic studies, tank pressurization testing, no-vent cryogenic fill, tank thermal control with MLI blankets and in-space propellant technology management work (Thomsik, 2000).

3.3.7Liquid Oxygen Oxygen requires special equipment for handling and storage. A typical storage system consists of a cryogenic storage tank, one or more vaporizers, a pressure control system and all piping necessary for the fill, vaporization and supply functions. The cryogenic tank is constructed, in principle, like a Dewar Cylinder, such as a Thermos™ bottle. There is an inner vessel surrounded by an outer vessel. Between the vessels is an annular space that contains an insulating vacuum medium. This annular gap keeps heat away from the liquid oxygen held in the inner vessel.

Figure 3-17: Typical LOX cryogenic storage Cryogenic fluids are usually stored in properly insulated containers designed to minimize the loss of product due to boil-off. A Dewar flask is the most common container for cryogenic fluids Figure 3-17 above. It is a double-walled, evacuated container made of metal or glass, with a vacuum between the walls. In the space environment, one has a natural vacuum and one can replace the double walled Dewar with a single walled vessel, better still one can wrap the single walled vessel in MLI. • Larger quantities of cryogenic fluid require double-walled metal

containers of evacuated construction. • Exposed glass should be taped to minimize the flying glass

hazard if the container should break or implode.

• Liquids should be transferred from the metal Dewar vessels with

special transfer tubes or pumps designed for that particular application.

3.3.8Cryogenic Storage Technologies

Methods:

Emerging

Programs undertaken at NASA Glenn Research Center in the last decade studied new and improved denser forms of LH2 and LOX.

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Propellants of particular interest are subcooled cryogenic propellants. This is due to the fact that they have significantly higher density, a lower vapor pressure, and improved cooling capacity over the normal boiling point cryogens. Higher density propellants enable additional propellant to be encased in a given volume, which results in improved performance for a launch vehicle by decreasing its overall weight and size (Thomsik 2000). “Density improvements of 10% for LOX and 8% for LH2 are expected to reduce the gross lift-off weight of a launch vehicle system by up to 20%” (Id.). In Figure 3-18 below is an example of passive storage of LOX and LH2 applied to space mission concepts. The surface properties used were based on JPL experience.

Figure 3-18: Titan explorer thermal model, deep space tanks, sun shield (Plachta 2005)

3.4Alternative Storage Possibilities The following storage options were considered.

3.4.1Slush Hydrogen Considerations

Properties

and

Hydrogen in a solid-liquid mixture, or slush, offers the advantages of higher density and heat capacity, when compared to the normal boiling point of LH2. This increase in density has the potential to reduce storage volume and consequently overall tank mass. “At a solid fraction of 50%, slush hydrogen provides a 15% advantage of increased density over that of normal boiling point hydrogen” (Friedlander et al. 1991). There are still several issues which must be resolved before slush hydrogen can be used for space missions. These include the transfer capability in low/micro gravity, verification and testing of insulation systems, the definition and testing of any additional components required for these space vehicles, as well as the ability for long term storage (Friedlander et al. 1991). Other forms of higher density hydrogen other than slush include liquidgelled hydrogen, atomic hydrogen, metalized-gelled, Earth storable

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(NTO, MMH, RP-1) and high-energy density propellants (Thomsik 2000). The research required to detail each high density hydrogen storage option is beyond the scope of this report; therefore only slush hydrogen will be considered as a storage option within the confines of our proposed architecture.

3.4.2Storage as Water Storing as water has the advantage of least pressure stressing the containing vessel, yet it still requires thermal control to keep it within the optimum range, i.e., liquid state with minimum vapor pressure, for the lunar environment. A closed container is still necessary as boil off will occur just like with any other method, but specialist containment tanks need be no more complicated than an oil drum, because pressure loading is minimal, compared to cryogenic liquids (Blair 2002) The argument against storing as water is that one must electrolyze it and then liquefy to use it as propellant. Having LOX and LH2 means that combining to produce water actually gives you energy useful for batteries etc, whilst still providing an essential life support component.

3.5Storage Vessels Storage vessels have to be well designed to fulfill their task and one of the requirements with regards to obtaining a durable design involves the choice of appropriate materials as well as choosing a certain shape and capacity. The larger the tank, the bigger the thermal capacity and the less difficult it is to thermally control, i.e., more efficient use of power, and on top of all that, a lower leak/evaporation rate (Domashenko 2002). A spherical tank offers the smallest area of contact for the same volume, thus the less thermal capacity of walls easier thermal control. Cylindrical tanks are less efficient in this aspect, but are easier to manufacture and handle. It is also suggests that evaporation can be used to chill tanks to the required temperature, and that larger tanks have a larger thermal capacity, going as large as demand allows makes sense (Domashenko 2002). Oversized tanks therefore offers benefits in thermal control as well as expansion mitigation.

3.6Interfacing When transferring the LOX and LH2, the interface must be durable and properly fix so that it can withstand the pressure and do not leak when transferring LOX and LH2. The connectors must also follow a universal standard to ease the junction between the tanks and the transportation system. In order to realize the benefits of cryogenic propellant transfer, one must ensure the robustness and reliability of the transfer process (Chato 2006). Several technical challenges arise when attempting to fill cryogenic tanks in low gravity. Thermal energy stored in the tank walls causes high vapor generation rates, the distributions of liquid and vapor within the tank are uncertain, and the operating pressure must be kept low, in

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order to minimize required tank mass. These considerations have to be taken into account in the overall storage design.

3.6.1Transfer Guidelines A fill in Earths gravitational environment involves a tank with a top vent kept open, in order to vent the vapor which is generated during the filling process. This venting is responsible for maintaining low tank pressure. In low gravity conditions, the position of the ullage (or unfilled space in the tank) is unknown relative to the position of the vent. This unknown ullage position can lead to venting liquid instead of vapor, and consequently large amounts of liquid being dumped overboard.

Filling Strategy The no-vent fill approach is one of the most promising methods. Initially the cryogenic tank undergoes “chill-down” (the tank wall temperature is decreased to the boiling point temperature of the fluid being transferred), followed by spray injection and fluid mixing to achieve the desired thermodynamic state in the receiver tank, which allows for filling without the need for venting. The no-vent fill also has the potential for high rate transfers (Chato 1991). When discussing the creation of an orbital propellant depot, it is concluded that research in no-vent fill transfers has matured the technology to the point where it is the recommended approach (Chato 1991). As it is the objective of sustainable fuel accessibility architecture to retain as much fuel as possible, venting of boil-off is not considered a viable option for storage, except in the case that emergency release is required. Based on the high value of hydrogen on the lunar surface and the available research regarding filling, the no-vent fill will be the assumed method applied to the chosen architecture.

Propellant Management The chosen tank design including pump and piping arrangement will be identical, only different in scale. As stated above, LH2 is deemed to be the more valuable of the two resources, as it is either brought from Earth or mined from the lunar surface from a smaller resource base. Due to this higher value, the LH2 tanks will contain more significant amount of insulation and protection, as its loss is the case of boil-off and it would pose a more serious problem for the lunar infrastructure (where a loss of LOX could be much more readily re-supplied). The major components involved in the storage and transfer of cryogenic materials include the tank, suction and discharge lines (piping), relief valves, back pressure valves, centrifugal pumps, strainers, shut-off valves, drainage valves and check valves. Due to the complexity and interconnectivity of these systems, they will be discussed here only on a conceptual level, although aspects of these systems will be considered in the storage system design. In terms of overall system architecture, it is recommended that the lunar storage facility should be located adjacent to the control and servicing facility. This setup would allow for efficient refueling and defuelling of a lunar lander during service procedures.

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3.7Ensuring Human Safety and Tank Health In order to control the storage facilities a series of monitoring infrastructure strategies must be installed. The two parts to be installed is the measuring infrastructure and the communication infrastructure to monitor the health of the storage facilities In order to be able to measure health of the storage tanks in space the following five methods can be used: pressure sensors, strain sensors, optical sensors, compression mass gauge and leak sensors. Some of the challenges that must be taken into account when handling fluid storage and transfer involve the design of a well thought integration of pumping and storage system equipped with the appropriate maintenance and health monitoring solutions. Key requirements propose the use of new technology valves for cryogenic applications (LOX/LH2) with the main purpose of minimizing thermal losses, and leakage especially pressure drops. Some of the solutions include shut off control valves, flow control valves, leak proof couplings using robust sealing technology and compatible with LOX and LH2, low power and lightweight pump for a reasonable flow rate (up to 2 l/min), pressure control sensors and integrated tank boil off control sensors, automated umbilical systems designed for high reliability and safety and appropriate for ground to flight interfaces (NASA 2003). In terms of monitoring methods high consideration must be given to location of joints as well the number of sensors for a successful detection and this is highly critical in detecting and monitoring leakage. One of the best ways of leak detection is by pressure variations applied with a systematic leakage detection method as well as a so called point source method (application specific controllers). An application specific controller is preferred allowing for an easier and more flexible control system without any functionality interference of other system components in case one fails. Sensor selection must take into account sensors for the following function monitoring: temperature monitoring, pressure controlling and monitoring, fluid velocity, liquid level monitoring, leakage detection and last but not least structural integrity.

3.7.1Assessment of delivery systems Delivery system is the element linking the production, storage and users elements of the architecture as Figure 3-19 shows. It follows that it has to accommodate the “boundary conditions” of three interfaces, in term of location, production and demand rates, form of gas to transport, mobility capabilities. Interactions with the environment add another “virtual interface”, which further constrains the delivery system. Within those constraints, the delivery system should be selected to maximize the benefit to users of refueling services. After presenting the criteria, possible transportation options will be discussed. They include both mobile and fixed transportation platforms. The section will conclude on an evaluation matrix that on the advantages and disadvantages of each transportation system.

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Figure 3-19: Elements of the refueling service

3.8Transportation options The following sections will briefly transportation options explored.

describe

of

the

possible

3.8.1Mobile platforms Mobile vehicles allow very flexible in the sense they can accommodate personalized user demands regarding location and time for delivery. However, the transportation capacity is limited and energy efficiency (amount of energy used per volume of fuel delivered) is low compared to fixed platforms.

Level of autonomy Manned vehicles allow real time human control, which is crucial for complex mission requiring fast interactivity with the environment. However, the life support systems installed onboard human pressurized platforms to accommodate astronauts significantly add complexity, which translates into additional costs and a lesser level of reliability. Using un-pressurized human vehicles, designed to carry astronauts in space suits, allow remove the vehicle built-in life support system while maintaining a high level of interactivity. However, it will go shorter distances, as astronauts need to return to their base for space suit life support system servicing. It follows that it is not fit to explore remote areas. Robots can be appropriate to accomplish tasks of relatively low complexity, requiring a lesser level of interactivity. Autonomous rovers are however complex, expensive, and not prone to repair (Eckart 1999). Remotely controlled is an intermediate solution between the capabilities of fully autonomous robots and human vehicles. The higher level of interactivity remotely controlled vehicles entail makes them capable of more complex missions than an autonomous robot. The interactivity is however limited by the time delay needed to transmit communications between humans and the vehicle. The following Table 3-19 gives a description of the mobile systems considered in this project.

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Table 3-19: Mobile platform characteristics (Eckart 1999) Delivery System

Rang e [km]

Speed [km/h]

Terrain Capabilit y

Capacity [kg]

Wheeled rover

100s (custo mizabl e)

5-20

Medium

100 to 20 000

Tracked vehicles

100s (custo mizabl e)

<wheele d rovers

High

500

Ballistic

aroun d the Moon

>6 000

Terrain independ ent

1 000 – 3 000

30

150

Limited

7 000

Mechanical hopper

Characteristics Wheels are fast and efficient on smooth, hard surfaces but lose traction on loose soil such as the sand-like regolith Safe and reliable Can be made fully autonomous Moving parts must be designed with tolerances to withstand thermal expansion due to the extreme range of temperatures on Moon surface Low power requirements Good terrain capabilities Excellent floatation characteristics in the lunar soil Immune to thermal fluctuations Remotely controlled if not automated Unreliable due to its high complexity Large power requirements (10 kW/t) Well-suited for transportation across rough terrain Fairly easy to control, and can be fairly autonomous Fast and accurate landing (within 100 m) Low energy efficiency The concept of the mechanical hopper lies between walkers and ballistic vehicles Can be fully autonomous Energy efficient Technical feasibility of such vehicles has not been demonstrated as yet

71

Ballisticwheeledwalker

72

2500

>6 000

Medium

4 000

System under consideration is a modified version of the NASA LSAM. This is a ballistic delivery system that operates between lunar orbit and the lunar surface with the All Terrain Hex Legged Extra-Terrestrial Explorer (ATHLETE) base for surface maneuverability (currently under development by NASA). High complexity.

3.8.2Fixed platforms Fixed infrastructures are built on Earth when large quantities of material need to transit between two fixed locations such as a production site to a distribution center. A similar approach is likely to be used on the Moon if a refueling station is requiring more than several tonnes per day over a long period of time. The reason behind this is that this quantity has to justify the large investment required for building and maintaining the infrastructure. That is why fixed platforms are not preferred during the preliminary phases of a lunar exploration and utilization. At more advanced phases, the benefit of pooling large volumes of LOX and LH2 on a fixed transportation facility may outweigh or complement the flexibility brought by mobile transportation platforms. Many alternatives exist for a fixed transportation system. Five means of conveyance have been identified: cable system, monorail/maglev train, pipelines and mass drivers. Each of them has different range capability and is listed in Table 3-20 below. Table 3-20: Lunar Transportation Systems (Apel 1989) Mean of conveyance Cable System Monorail/Maglev Train Pipelines Mass Drivers

Short Range Cargo Transportation Yes Yes

Long Range Cargo Transportation No Yes

Yes No

Yes Yes

Even though many solutions exist for the short range or long range situations, is it important to list their advantages and disadvantages as well as their energy requirements in order to evaluate which one are really compatible with lunar requirements. Table 3-21 below gives a more detailed view of the fixed systems considered in this project.

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Table 3-21: Fixed Platforms Characteristics (Apel 1987), (Transneft 2004), (Doppelmayr n.d) Delivery System

Range [km]

Cable system (Gondola)

10 (pole to pole) there can be multiple poles

Pipeline

Mass driver

Mag-lev trains

Unlimited

Speed [km/h] 20 (on Earth)

∝ diameter

Terrain Capability

Capacity

Terrain independent

1 500 000 kg/h (on earth)

Level ground

High (large amounts of fluids at a flow rate up to 10 million tonnes/year

around the Moon

5 300

High

300 tonnes

5400

600

Medium

-

Characteristics Can overcome difficult terrains with inclination up to 45º Thermal expansion and contraction of the cable could lead to fatigue in the material and premature failure Fully autonomous and reliable The pipe also acts like a tank and needs to be filled before used High infrastructure cost Device that accelerates cargo with extremely high g-load (up to 10000 g) to very high velocities, in the order of magnitude of 5,300 km/h, needs means of deceleration/landing Dual use system that can transport both cargo and personnel High infrastructure cost High energy efficiency [only 0.17-0.25 kW∙h/(Mg∙kg)] Low energy efficiency [specific energy consumption between 0.165 and 0.392 kW∙h/(Mg∙kg)] High system costs

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Off-surface vs. Underground Off-surface systems represent an interesting option when a large terrain inclination needs to be overcome. This situation is present when the propellant has to be carried from a crater to higher ground. Offsurface systems also offer advantages against dust problems that represent main concern of NASA for ground equipment. On the other hand, these systems are more complex to install and also more difficult to maintain. The cable system is by definition a system that is above ground. But the monorail/Maglev train, steel belt conveyor, pipelines and mass drivers can all be either on the ground or above. Only the pipelines can be installed underground without requiring tremendous effort.

3.9Delivery System Evaluation Since many of the systems are novel concepts on the Moon, it is often not possible to assess them based on historical data on performance and cost. Therefore, they should be assessed according to criteria relevant to users, in an objective and systematic manner that integrates the multiplicity of criteria in a single metrics.

3.9.1Evaluation Criteria Relevant criteria for evaluation and selection of the transportation system include mission requirements (the drivers), and constraints. Drivers pertain to quantitative and quantitative performance. An ideal transportation system would match perfectly all users’ demands regarding performance, and suit their evolution through the anticipated lunar development scenario. On the other hand, the effort it takes to implement the delivery system should be minimized. Following this logic, we identified eleven criteria which are directly applicable to transportation vehicles (Table 3-22). Note that two parameters that were considered but not retained are “functionality” and “development effort” because of overlap with other parameters.

3.10Qualitative Decision Delivery Systems

Rankings

for

In the early research stage, the qualitative decision tool was used to select out the options least suitable for lunar cryogen delivery, based on a simplified set of criteria, derived from those presented earlier.

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ImplementationEase of

Simplicity

Robustness

Final Ranking

Pipeline, surface Wheeled Ballistic Ballistic wheeledwalker Tracked Gondola Rail, surface Hopper Conveyor belt Pipeline, subsurface

Scalability

System

Operational Efficiency

Criteria

Technical Readiness

Table 3-22: Qualitative method results

5 9 8

8 4 0

4 9 6

4 9 8

9 3 4

7 2 8

#1 #2 #3

7

2

7

7

2

0

#4

6 2 3 4 0

3 7 5 1 6

8 3 2 5 1

6 2 3 5 0

1 7 6 0 5

1 4 5 3 6

0

0

0

0

0

0

#4 #4 #7 #8 #8 #1 0

Table 3-22 above highlights the relative merits of each system. The initial ranking is based on intuitive opinions about the systems before they were rated. Comparing the initial and final rankings, it is apparent that the system does not agree in some respects with intuition – based on these criteria, fixed systems seemed globally preferable to mobile ones. This points out that the fixed and mobile systems must be considered separately since they represent solutions for different phases in the overall cryogen system development. With this in mind, the simple ranking was used to reduce the research effort, eliminating the least suitable fixed and mobile systems.

Analysis Ratings The simple qualitative method favored the pipeline system over other systems mainly because of its efficiency, simplicity and its technical readiness. However, this method was not taking in account for the overcapacity of such system neither the energy nor the cost required for building the infrastructure. Thus selecting a system base solely on this tool would not allow taking into consideration for the different weights of each criteria and would be likely to indicate wrong conclusion. But used as a cut off tool, it can clearly identify the systems that are inappropriate and allow focusing all efforts on possible scenario. Keeping that in mind, conveyor belt system will not be investigated further because of its technical immaturity and the development effort that would be required in order to make it sustain the lunar environment. This system is also unsuited to carry either gas or liquid in limited amount.

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The wheel rover scored the highest in the initial iteration. It was selected as one of the transportation options for transporting propellant. Among the highlights of the wheeled rover concept is its advanced technical maturity relative to other mobile vehicles, as its technology has already be proven. Technical maturity also translates into a good reduced development effort. Wheeled rover also scored highest for operational efficiency and is the most scalable system. On the robustness side, it is weaker than the tracked rover and the ballistic vehicle. Indeed, the moving parts are unstable in a wheeled rover. It is however considered more robust than hoppers. Simplicity is an ingredient of robustness. It is therefore not surprising to see wheeled rovers scoring in the median range in term of design complexity too. The tracked rover scores well on the robustness contest, as it has been designed for the military with demanding requirements in terms of reliability and sturdiness. However it performs badly in term of technical maturity. Indeed, it has never been space qualified. In addition, the large number of parts it involves makes it the hardest system to develop (Eckart 1999). Other major downsides of the tracked vehicle include a low energy efficiency and little scalability. Therefore from the above discussion it was decided to drop this concept for the final design. Ballistic vehicles obtained the second highest rating. They have several disadvantages. First, they need a lot of propellant and are considered energy inefficient (Eckard 1999). Inaptitude for scalability is another downfall. Indeed, modifying the capacity of the payload bay is difficult. However, good performance on the other criteria compensate for those shortcomings. Robustness is also a key driver of the overall rating, due to very few moving parts, smooth and stable overall structure, and built in obstacle avoidance capability. The ballistic-wheeled-walker is in the middle of the overall ranking. The main advantage of this type of vehicle is that it is multi-tasked to both fly for 1000’s km and then also do on surface maneuvering once landed on the surface. Since it has walking capabilities it can also encounter terrains well. It is not technically as mature as a ballistic or a wheeled vehicle, but once it does it will be a very efficient vehicle. Hoppers on the contrary score poorly. Regarding technological maturity, hoppers have remained to conceptual stage as yet. They are also one of the least robust of all options, due to numerous moving parts and susceptibility to obstacle hitting. Based on these results, subsurface pipeline, gondola, tracked and hopper were eliminated from the pool of possible transport designs and the choices for the delivery system was reduced to 6 options. Comparing the initial and final rankings, it is apparent that the system does not agree in some respects with intuition – based on these criteria, fixed systems seemed globally preferable to mobile ones. This points out that the fixed and mobile systems must be considered separately since they represent solutions for different phases (and accordingly different user needs) in the overall cryogen system development.

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3.11Quantitative Decision Delivery Systems

Rankings

for

As more details of the systems became known (see Table 3-23), they were used as inputs in the quantitative tool with criteria related to cost, performance and robustness to determine the best of the remaining delivery system choices. A representative metric for each of the criteria was selected based on consistently available data. The criteria were manipulated so that higher numbers correspond to “better” ratings and normalized. Finally, the ratings were weighted based on team members’ opinions and on “expert” (faculty and external contacts) opinions separately.

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Table 3-23: Quantitative selection criteria

How much it performs

effortImplementation

Criteria

Parameters Technical readiness (mission cost) Operational efficiency (wise utilization of resources Scalability (expandability of a system) Adaptability (Multitasking capabilities)

Definition of “best” Ready the earliest

Related quantity Years to develop

Weights (team : faculty) 6.4 : 8.4

Least energy-intensive to Specific power [W] run

7.0 : 6.2

Easiest to scale up or down to change in Marginal cost (%) to double output demand

6.3 : 7.7

Adaptable to other uses

5.9 : 6.3

Other applications

Ease of Implementation

Lowest capital cost to Mass install

Freedom of location (accessibility)

Least restrictive

Range [km], terrain compatibility, max. grade, regions covered

8.2 : 7.2

Delivery Time

Quickest

Time to deliver fixed amount to a fixed distance

6.3 : 7.5

6.1 : 8.6

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performsHow well it

Criteria

Parameters Reliability

Safety Level of Autonomy Resilience to Lunar Environment

Definition of “best”

Related quantity

Fewest moving parts, lowest maintenance Maintenance cost requirements Is it manned? Is it space-proven Least likely to injure tech? Is it free of hazardous someone materials and properties? Is it stable? Does it require a human? Can it be Most automated tele-operated? Is it fully automated? Hardiest in lunar Low tolerances (thermal expansion); environment (radiation, imperviousness to dust; terrain vacuum, extreme capability temperatures)

The results for the qualitative and quantitative matrix methods are compared in calculated values for each system along with criteria and their relative weights.

80

Weights (team : faculty) 6.0 : 7.3

6.2 : 4.3 7.8 : 7.3 7.9 : 7.8

below. The table shows the

Table 3-24: Weighted Quantitative Decision Matrix Results

Contrary to the qualitative method the quantitative decision rankings lists the ballistic solution first. This is the result of adding weighting factors to the criteria. The ballistic solution is indeed a proven technology based on the Lunar Lander that has the ability to cover the whole Moon surface while being cost efficient. However, this solution requires a lot of energy to deliver to the customers. Although the pipeline system would be more efficient operationally it still falls second when taking into consideration for the infrastructure deployment effort and cost. This system is then followed by the ballistic-wheeled-walker, the Maglev train and wheeled rover. It should be noted that the quantitative method was not optimized for a specific scenario. This means that inputs for a system are based on average values and that the weighting is not associated to mission requirements. Moreover, system characteristics have a range that depends on their configuration which reduces the accuracy of the method in some cases up to ±25%. Also, if two outposts are 1 km away from each other, the weight for freedom of location would be reduced greatly compared to a scenario where delivery at the equator at different locations is required. This

81

method has also other limitations. One of these limitations is shown by the delivery time criterion. Because it is calculated using the speed and the capacity of a system without considering the number of travels that could be done in one day. That is detrimental to a mass driver or any other fixed infrastructure that are almost continuous application. It is important to remember that the decision matrix tool is just an aid to evaluate the relative merits of individual systems. However, it is not suitable to evaluate the supply system architecture as a whole, and cannot be easily adapted to include combinations of systems, which are of interest in a phased market or to compare between different future scenarios. To make recommendations regarding the supply system architecture and its evolution in time to accommodate the shifting lunar base plans from various agencies, a more complex decision method should be considered, such as the “supply chain model” presented in the beginning of this chapter. However, such a sophisticated tool has not been implemented in this work. In addition, no decision making tool can ever totally substitute sound thinking and critical analysis of skilled engineers. The multi-criteria decision methods show the relative merits of individual systems, but are not easily adapted to include combinations of systems, which are of interest in a phased market or to compare between different future scenarios. To make recommendations that are more relevant to the shifting lunar base plans from various agencies, a more complex decision method was considered.

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Figure 3-20: Mobile and fixed platforms

3.12Findings: Storage and Delivery Concept Rankings Storage and delivery on the lunar surface present interesting challenges to current and emerging technologies. However, it seems prudent to go with those technologies both optimized for the space sector already, but taking advantage of the lunar environment to make ones task easier. Al-Li alloys therefore seem to suggest a workable tank material, at least in the near term, from which to design a solution in parallel with a delivery system capable of meeting the early and later phases of demand.

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Delivery systems should make use also of the lunar environment specifics, where possible, but also take heed of basic constraints imposed by the harsh dusty environment. One may consider systems more favorable, the fewer number of moving parts they may have. Energy efficiency and time constraints will play a big role in the trade space. Having assessed these systems, one question remains to be addressed: which system architecture will best address the needs of the station?

84

Introduction

__________________________________Chapter 4

4 System Architecture

With knowledge of the drivers and constraints for a lunar liquid oxygen and liquid hydrogen supply system, and armed with a large number of options for storage and delivery elements from the previous chapters, a complete system architecture can now be selected. The selection is based on the perceived demand for the baseline scenario described in Section 2.1.1 with a single production facility at a south polar base in the event of ice at the South Pole (“Ice Scenario”) and without (“No-Ice Scenario”). The system architecture must satisfy key objectives and the identified gaps in current literature for this baseline demand and consider potential future adaptations for the optimistic higher demand future (also in Section 2.1.3). In this chapter, an “element” is defined as a major system component, including vehicles (e.g. a cryogen delivery rover) and infrastructure (e.g. a base habitat module). “Customer” refers to the end user of the cryogen (e.g. NASA), “ascender” is a vehicle operating between the lunar surface and lunar orbit, “consumer” is the element accepting the cryogen (e.g. a NASA ascender), and “astronaut” refers to any human in space. The system architecture is based on certain assumptions including demand and production infrastructure. It is also assumed that all manufactured items must be shipped from Earth (i.e. no in situ manufacturing for system elements and spare parts).

The Demand • As previously discussed in Section 2.1.1, it is assumed that a

manned lunar base, mid-latitude manned exploration missions and Mars exploration are all potential customers for LOX and LH2 refueling. Table 4-25 recapitulates the demand and location of those customers. Table 4-25: Demand scenarios

Scenari o

Locatio n

Baseline

Local

Major (Minor) Customers Life support, ascent vehicles (human exploration missions)

Rang e [km]

2

Demand [tonnes] Per Per delivery year

4

<40

85

Full Moon

Baseline

Equator

Optimisti c

Local

Ascent vehicles (human exploration missions) Baseline and propellant for Mars missions, life support for additional bases

3 000

4

<8

2

???

100

• The objective of the proposed architecture is to meet this

demand to the closest possible.

The Production Infrastructure The production facility is regarded as a starting point from which the storage and delivery systems are conceptualized. Table 4-26 summarizes the boundary conditions imposed by production regarding the type of fuel produced, the production rate and the location production. Table 4-26: Production scenarios Scenario

Production Location South Pole (in Ice at South permanently shadowed Pole crater) Anywhere in South Pole No ice area

Output LOX LH2

and

LOX

Note that the storage and delivery architecture does not vary in these alternate (“ice” and “no ice”) scenarios. They are assessed in the economic analysis, as the existence of on site hydrogen determines whether or not hydrogen will need to be transported from Earth. It is assumed that regardless of whether hydrogen is brought from Earth or is produced on site, the storage and delivery concepts remain the same in each timeline. The production rate is looked at as a variable which can be increased or decreased, dependent only on the amount of power being supplied to it. With such a view of power requirements, and to have the flexibility to upscale production to meet higher demand, power supply could be achieved with a nuclear reactor, or with solar panels located in a permanently illuminated area close by the production facility. It should be noted however that NASA currently considers the lunar nuclear power infeasible. Also, the production facility is viewed to be made moveable by the gas station transportation systems, in the case that ice or regolith supplies become too low in the vicinity to warrant the associated transportation costs, or in case the base must be moved. It follows that the storage/delivery system should be flexible enough to accommodate relocations of production.

86

Introduction

The System Objectives Because the storage and delivery system must support human exploration, human safety is paramount. The system should provide service to the predicted consumers with this as the key design criteria. A requirement for human safety is that critical supplies must either accompany a human mission or be present and confirmed before the mission is launched from Earth. The latter method is called “pre-delivery.” Any system for use with humans in space must not be susceptible to a “single point failure” (i.e. there must be a built-in redundancy to prevent a critical system from shutting down for any potential failure). Like fuel systems on Earth, a lunar fueling system can be fullservice (no customer action required) or self-service (the customer is required to perform part or all of the connection to and operation of the refueling system). Since self-service implies customer (i.e. astronaut) training and responsibility for the correct operation of the system, creating a human safety issue, a fullservice system is recommended. Further, since a fully-automated is complex and potentially could pose a risk to human safety if it fails, tele-operation is recommended. Because of the cost of implementation and maintenance a delivery system is expected to be flexible enough to perform more than one task (Larson teleconference April 2007). To fill identified gaps in NASA reports (Sanders 2007) and other research, the system architecture must meet the following objectives: • Local delivery at the lunar South Pole for fuel and life

support • Delivery to human scientific missions at the equator • Ability of the delivery system to multi-task (modularization) • High level of human safety, especially:

a. Elimination of single point failures (redundancy) b. Unmanned operation to eliminate EVA and astronaut training c. Pre-delivery d. Tele-operation To weight existing concepts according to these criteria, the selection methods detailed in Chapter 3 were used.

The System Selection Method The delivery components for the supply system architecture were selected based on the results of the literature review combined with the quantitative decision tool described in 3.11. The results were factored with secondary weights (shown in Table 4-27) to better represent the importance of key parameters for each

87

Full Moon

scenario. The criteria with secondary weights (original team weights factored by two) are highlighted in the table.

Ease of Implementation

Safety

Freedom of Location

Delivery Time

Reliability

EnvironmentResilience to Lunar

Baseline: local Baseline: midlatitudes Optimistic :

Scalability

Scenario

Operational Efficiency

Criteria 

Technical Readiness

Table 4-27: Quantitative decision tool weights with secondary weights

1 2.8

7.0

6.3

5.9

6.1

8.2

6.0

7.8

7.9

1 2.8

7.0

6.3

5.9

6.1

8.2

1 2.0

7.8

7.9

6.4

7.0

1 2.6

5.9

6.1

8.2

6.0

7.8

7.9

This chapter proposes a supply system architecture to suit the perceived needs of a South Pole lunar base and equatorial human exploration missions, justifies this selection and describes the interfaces within the system and between the system and external elements. Operations are then described in detail, followed by the implementation plan (“The System Blueprints”). Technical risk is evaluated. Finally adaptations are proposed to meet the optimistic high demand scenario.

4.1The Proposed Architecture The main elements of our proposed supply system architecture for LOX and LH2 on the lunar surface are storage tanks and delivery systems.

4.1.1The Storage Solution The assessment in this section aims to provide a set of recommendations regarding the storage of oxygen and hydrogen on the lunar surface. Research in this area emphasizes the need for long duration storage of oxygen and hydrogen on the lunar surface. This long duration storage requirement stems from the needs of future manned lunar outposts and their associated transportation vehicles. The following section will provide recommendations for the near-term scenario regarding which form of oxygen and hydrogen should be stored, the tank material

88

Introduction

used for construction, tank shape, tank size (based on customer demand), tank weight, tank thermal control, and storage location. A brief long-term storage assessment will be provided later, in the context of an infrastructure which may develop based on our near-term recommendations. The reasons for selecting particular storage options (such as tank shape and material) are not as apparent as the reasons for selecting a particular transportation system (such as ballistic vs. rover vs. pipeline). Usually the desired thermodynamic requirements for a given tank guide these selections, and as detailed thermodynamic analysis is beyond the scope of this report, a qualitative comparison of the storage options was conducted in order to come to a final down-selection of the tank specifications. The following set of storage objectives were chosen as comparison criteria in order to come to a final set of recommendations regarding the storage design: 1) Reduce tank weight 2) Maintain structural integrity for extended periods in the lunar environment 3) Maximize capacity 4) Minimize energy consumption 5) Maintain human safety and tank health

Location Storage locations are inherently connected to the thermal conditions in a given lunar environment. An ideal location is one which makes storage of LOX and LH2 “easiest”, i.e. where is less costly energetically to store. Since the production facility and associated infrastructure is primarily located at the South Pole, the desired storage location for any LOX or LH2 is on the surface, within a permanently shadowed crater. To be stored in stable conditions LOX requires a temperature of 90 K, and LH2 requires 20 K. As shadowed locations on the South Pole can reach temperatures as low as 40-50 K, if the location is properly chosen, the LOX could be initially cooled with passive thermal control (zero energy expenditure) whereas LH2 would undergo continuous, however minimal, active cooling. If cryogenic storage is to take place for long periods of time under thermal conditions such as those at the equator when illuminated, the ideal storage conditions would be subterranean. At a depth of one meter beneath the surface, the tank surroundings could be maintained in near-constant conditions at approximately 238 K (Heiken et al. 1991). If cryogenic storage is required to be mobile at the equator under thermal conditions during the lunar day, a storage facility (subterranean or surface) is recommended to effectively block incoming solar radiation, and therefore minimize tank heating. In essence the goal would be to construct an ideally insulated structure, effectively creating an area of permanent shadow.

89

Full Moon

Tank Shape The tank shape was chosen to be spherical, as a sphere is a theoretically ideal pressure vessel. Spherical tank offers the least thermal capacity and inertia, as it contains the maximum volume with the least surface area, thus least wall material.

Tank Material The tank material chosen for construction is aluminum-lithium alloy (Al-Li). This material was chosen due to its technical readiness for application, as well as its contribution to reducing landing mass. Although carbon composites were also considered, it was thought that the probability of micro-cracking was too high (due to differences in the coefficient of thermal expansion of the carbon fibers) which means that its use would have too high of a risk for the early phases of the project (Scatteia et al. 2005). The thermal cycling endured by the tanks may be extensive, therefore a material that has proven to be reliable under such conditions must be chosen.

Thermal Control Variable Density Multi-Layer Insulation is selected since it is one of the best thermal insulations in a vacuum, and because there are no problems related to density control and performance, nor with covering on small or large scales. VDMLI should have lower thermal conductivity, vacuum compatibility efficient assembly and provide structural reliability. An outer layer blanket should also be installed, consisting of VDMLI, finished with silvered Teflon, to minimize surface heating. After comparing the options for the storage of oxygen and hydrogen, LOX and LH2 were chosen from the forms considered for the near-term scenario. Although more energy intensive to store, these forms are almost certain to be used to fuel near-term spacecraft and will therefore require delivery in these forms. In the No-Ice Scenario, it was thought that hydrogen would be transported from Earth in the form of LH2, in order to reduce the launch mass of the transportation vehicle. Lunar oxygen would be produced and stored only when a specific demand is identified, in order to minimize required LOX storage time, and hence minimize the energy required for refrigeration. Less energy intensive forms of storage (such as water and ice) were also considered in this scenario, but a more complicated infrastructure would be required for their use. Energy storage devices (such as regenerative fuel cells) would be required to retain the energy created during water formation, and energy for electrolysis would be required in order to return the oxygen and hydrogen to their LOX and LH2 forms for delivery. This would require an infrastructure that must maintain storage facilities for water, ice, LOX and LH2, as well as have the facilities to convert the substances into their alternate forms respectively. It is much easier to produce the substances one need and maintain them for a limited amount of time, than to overcomplicate the infrastructure unnecessarily in the initial phases.

90

Introduction

In the Ice Scenario, LOX and LH2 production would only take place when a demand has been identified, and the LOX and LH2 would only require storage pre-delivery. Ice could be separated from the regolith, and stored as is, or could be melted into discrete sized blocks in precisely measured quantities to simplify production of the appropriate amounts of LOX and LH2. Storage of ice would be in a permanently shadowed crater, or in a suitably cold area (protected from solar heating), until the time a demand has been identified. The storage tanks, sized for each of the cryogens, are shown in Figure 4-21.

Figure 4-21: Mobile storage tanks To delivery the storage tanks, a delivery system is required.

4.1.2The Delivery Solutions The delivery system will supply cryogens to consumers at the lunar South Pole base (i.e. “local delivery”) and to human missions in LSAM9-type ascenders at the equator. Emergency deliveries have also been considered. The delivery system options can be quickly narrowed down for each of these three applications by comparing the element capabilities with the requirements (see Figure 4-22).

9

“Lunar Surface Access Module” – a NASA design concept for human lunar exploration that carries humans and cargo between lunar orbit and the lunar surface. 91

Full Moon

Figure 4-22: Delivery concept comparison by range and capacity

4.1.3Local Delivery Based on the assumptions for the baseline scenario, since production and the lunar base will be co-located, all demand will be met from the South Pole. The objectives specific to local delivery supporting ascent modules from the base (and incidental amounts for human exploration missions) include: a capacity of four tonnes per delivery and a range of two kilometers. The delivery system should be capable of moving other supplies or performing other tasks. Fixed infrastructure is not expected to be in place in the nearterm. Mobile systems were considered more suitable since flexibility is an important criterion in the likely case that the base architecture is changed or expanded. From the mobile concepts in 3.8.1, the predicted demand is best met using wheeled rovers. While it is possible to fix cryogen tanks directly to the rover, the system will be more flexible with removable tanks. To simplify the operations (and avoid having separate cranes or other loading systems for the tanks), the storage tank(s) will be carried on a trailer, allowing the rover freedom to perform other tasks (see Figure 4-23). By separating the mobile element from the payload element, the rover has the power to deliver other cargo, tow habitat modules or perform ad hoc inspections without a trailer. It can also be designed with a seat and a manual on-board operation system as a back-up un-pressurized human rover. However, it would be oversized and therefore inefficient for exploration.

92

Introduction

Figure 4-23: Tug rover with trailer (Bufkin et al. 1988) According to the criterion, the wheeled rover is excellent in technical readiness and scalability, although it is among the weakest in freedom of location and delivery time, neither of which is critical for local deliveries (see Table 4-27). The wheeled rover does not have the range or terrain capability to service the equator, however.

4.1.4Equatorial Delivery Delivery to the lunar equator implies a range of 2 500 km. The required capacity is four tonnes per delivery as for the local case. Again, construction of a fixed infrastructure such as roads or pipelines is not feasible in this time frame, limiting the selection to mobile systems with a high degree of freedom of location. Rovers are too limited by terrain and distance. Ballistic rockets, mass drivers and a ballistic-wheel-walker system were considered for service to the equator. However, this introduces the problem of descent speed. The velocity required for a trajectory to reach the equator from the pole is 1 660 m/s. The kinetic energy from an uncontrolled descent at this speed would crush the vehicle and the payload. If the journey was to be completed in “hopper” stages, even at moderate distances (500 km) the impact velocity is still 900 m/s. Current energy absorption methods for uncontrolled descents include air bags, a hard shell and retro-rockets. Airbags have only been designed for final velocities after aerobraking, such as the Mars Exploration Rover airbags (shown in Figure 4-24) designed for 25 m/s (Stein and Sandy 2003). Likewise, hard shells for descent capsules are designed for impacts at speeds reduced by aero-braking and retro-rockets. A shell designed to absorb the total impact energy of a sub-orbital lunar trajectory would be prohibitively heavy. A novel but undeveloped solution is a mass catcher, a funnel-like device that directs and slows the descent vehicle through mechanical friction, but that implies infrastructure placed strategically throughout the region – again unlikely in the near term.

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Figure 4-24: Mars airbag descent brake (Stein and Sandy 2003) On the other hand, ensuring a controlled descent imposes an enormous propellant cost. Using the M-LSAM (described as the NASA LSAM with the ATHLETE rover in 3.8.1) as a basis for calculation, 30 tonnes of propellant are needed to deliver a four tonne payload to the equator and return the delivery vehicle to the pole (see calculation in Appendix ). This is within the capacity of the M-LSAM and is eight times less costly in propellant than delivery from Earth.

Figure 4-25: NASA LSAM ballistic lander and ATHLETE rover (www.astronautix.com) (www.jpl.nasa.gov) For supporting human missions, the cryogens must be predelivered to the site. To prevent landing accidents, the mobility of the M-LSAM will permit the human mission to land kilometers away from the supplies and the ATHLETE to complete the last stage of the delivery. None of the other concepts are considered technically viable for delivery to the equator without a quantum leap in design. If a much larger market appears than is anticipated (such as an 94

Introduction

equatorial base), infrastructure such as a network of storage sites and road or rail networks or mass catchers, could be justified. It is also possible to use the M-LSAM configuration for emergency applications.

Emergency Delivery System Since all human missions landing at the equator will need to be fully-fueled, the only requirement for them will be emergency support. All human missions will carry sufficient oxygen and hydrogen for fuel and life support with a safety margin. In an emergency, the M-LSAM is best suited to deliver quickly and “to the door.” If it is assumed that NASA adopts the M-LSAM, it will be particularly suited to this phase because it will be a proven technology and will be reusing equipment that is already on site (a small marginal cost to make operational). This system combines the fastest, most certain way to deliver emergency supplies (less than two hours to the equator) with the ability to deliver to a specific location, so that human safety is not further jeopardized. The M-LSAM also does not require a launch pad. The drawback in this phase is the extreme inefficiency that can only be justified for human emergencies. Based on the development assumption, the M-LSAM is excellent in the technical readiness and freedom of location criterion, which are critical for emergency use. The ballistic component, once refueled gives rapid (less than two hours) point-to-point delivery as far as the equator. The ATHLETE is a combination wheeled walker rover that can roll along level terrain at low power and climb over obstacles in a second mode. The ATHLETE can be tele-operated either by a stranded human mission or, if the mission is unable to do this, by the lunar base using communications relay stations in orbit. (The combination LSAM and ATHLETE is hereafter referred to as the M-LSAM.) This choice was made using the quantitative decision tool with an increased weight for technical readiness and freedom of location. (The adjusted weights were shown in Table 4-27.) For this system of storage and delivery elements to be compatible with each other and with the consumer elements, the interfaces must be standardized.

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Figure 4-26: Supply system overview

4.2Interfaces It is critical that all the systems have compatible (and therefore standardized) interfaces. These interfaces will be used between the storage fixed tanks and the delivery tanks, and between the delivery tanks and the consumers. The general layout of this is shown in Figure 4-27.

Figure 4-27: Interfaces between production and storage elements

4.2.1Interfaces with the M-LSAM The M-LSAM has two refueling hoses (one for LH2 and one for LOX) articulated and actuated by a robotic arm, remotely operated from the lunar base. This operated hoses dock either on the full tanks at the production-storage site to fill up the M-LSAM tanks or on the ascender tanks at delivery.

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The armed hose is shown in Figure 4-28.

Figure 4-28: Articulated hose for fuel transfer

4.2.2Interfaces with Local Servicing Rovers Upload of full tanks on rover and download of empty tanks from rover at the production-storage facility: juxtaposition of rover to fixed stand, and transfer with conveyor belts. The lunar base has the same armed hoses as on the M-LSAM, used to transfer the life support fuel from the tank brought by a rover. Two robotic arms (with internal piping) will be installed in each unit to facilitate cryogen transfer. These robotic arms will extend from the fuel port of the tank, and will be autonomously or remotely controlled, with the capacity to be operated manually. The need for two arms is due to the highly explosive nature of the two propellants, and since pipe cleaning procedures will be minimal, each fuel must have its own delivery system.

The filling process is described as follows: 1) The “hands” of the arms are docked with the appropriate ports on the receiving tank, and are locked in place. One connection is made to remove vapor created during tank chill-down, and the other for removing tank residuals and tank filling. 2) Any residual fuel remaining in the tank is drained and stored for later use. 3) A sprayer head is injected into the receiving tank from one of the “hands”, and a small amount of fuel is sprayed into the tank interior to complete the tank chill-down. The newly created vapor is evacuated to the tanks compressor, where it is condensed and stored for later use. 4) Once tank chill-down is completed, tank filling begins. Once the tank is filled, all valves are closed, both arms are unlocked, and servicing is complete. It should be noted that each tank has individual valves that must be opened for each interaction. All interfaces should be

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standardized in order to facilitate docking and transfer procedures. In the event a customer wishes to be serviced via a vented fill, this can be accomplished without the need for the sprayer and the second arm, as thermal energy will be removed via the vapor evacuating to space.

4.2.3Communications A communications infrastructure is necessary to support this supply system architecture. The communications system is assumed to be deployed and operating when lunar cryogen supply operations begin. The communications system will link the production facility at the pole to the mid-latitude customers via relay stations that are either ground stations on Earth or orbiting the Moon. Communications is needed to perform telemetry, command and voice communication between the delivery systems and the base. Continuous service is required, since voice and data transfer are needed for monitoring and commanding the mobile systems during travel and refueling. Three earth ground stations would be needed to ensure 24 hour communications. It is expected the communication system will receive and transmit in the S band and Ka band (Bufkin et al. 1988). A pros and cons analysis is performed of the following communication concepts in Table 4-28. To extend the local delivery system range, “a tower antenna would be required at the base, on the vehicle or both” (Bufkin et al. 1988). This is called direct line communications and has a range of 50 km and requires that there be no obstacles between the two points. The propellant production facility will be in a crater, so this method of communication might not be ideal. To eliminate a large tower antenna, Bufkin et al. (1988) recommend relay stations. In early lunar base development stages, relay stations from Earth are recommended. With expanded lunar activity, a network of relay satellites and GNSS-type systems in Low Lunar Orbit (LLO) may be justified. Or satellites can be placed in the Earth-Moon L1 point that will allow continuous communications for the near side (Id.).

Table 4-28: Communications concept comparisons

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Concept

Advantages

Disadvantages -Short distances only Direct line -Cannot relay through +Quick and easy setCommunicati obstacles up ons -Large tower antennas required +Large tower antennas not required -Cannot be used for Relay +Good for local relaying to equator and Stations coverage (e.g. South farther Pole main base) Relay Satellite / +Large coverage Earth Relay

-Need a more complex infrastructure, like putting satellites in orbit

With a complete and compatible system architecture, now the operations aspect can be reviewed.

4.3Operations The system will provide local delivery to the South Pole base, predelivery at the equator and emergency services. The servicing method must be autonomous in all cases. The following sections describe how the system architecture will be used to satisfy customer service and safety.

4.3.1Storage Tank Rotation The storage system can be described as five functional units (assigned letters A to E) in order to simplify the architectural description. Unit-A is a fixed large scale storage unit partnered with the production facility. This large unit will allow for production when all other tanks are in use. Unit-B will be used to support the needs of the lunar base. It is logical to assume that the base may request or require that one unit be filled and operated constantly for “back-up” lunar base support, for the psychological well-being of the inhabitants, and to be utilized if an instantaneous need arises in an emergency situation. Therefore one unit will be assumed to be filled and in constant use, located in the permanently shadowed crater in order to reduce the energy demand. It should be noted that according to Nortunado (2007), continuous operation of cryogenic systems offer greater reliability compared to numerous thermal transient cycles, therefore units traveling between the South Pole and the equator will be more heavily taxed than those in the South Pole environment. Unit-C and Unit-D are mounted on mobile trailers and would be used for customer servicing, moving between production facility and launch area or lunar base. The function of Unit-E would be that of a back-up unit in case Units B to D shows signs of fatigue and is about to fail. The lunar 99

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environment and its effect on a set of tanks with identical designs cannot be predicted. The thermal conditions experienced by this tank will be more or less constant, given that it will be located in a permanently shadowed crater (the exception being when the production facility is being moved). Unit-E would not be moved or used unless it is deemed necessary. Unit-E can be moved by the same transportation system as the production facility. The tanks B, C and D will be rotated in order to equalize the type and amount of stresses across the tanks, as to prevent one tank from being stressed to the point of failure from constant use in a more taxing role. Thermal considerations when choosing the tank materials will therefore be of primary importance.

4.3.2Servicing Methods The options for servicing methods are numerous and it is likely that a combination of all of these aspects may be the only solution that it viable regarding procedures as complicated as those involved in fuel storage, transportation and transfer. Below you will find a brief description of the advantages and disadvantages of each, regarding the possible options.

Hands-on Servicing In order to provide hands-on servicing, the M-LSAM would either require a human pilot that could facilitate transfer upon landing or an-unpiloted M-LSAM would be set-up with a delivery system that the customer could easily understand and use. A human pilot has the advantages of quick decision making capacity, unique problem solving ability, as well as efficient motor control. Also a human has the ability to facilitate repairs in the case of mechanical malfunction. These aspects are not something that can easily be reproduced in remote or autonomous systems. However, the use of a human pilot increases the risk involved with every delivery, increases the launch mass and fuel expenditure. This would also require the expenditure of additional fuel to balance the weight of the human, the additional support structure, EVA equipment and fuel required for life support. A manned M-LSAM is therefore not recommended, as the reactivity and the skills of humans can be brought with an efficient teleoperation system.

Remote Servicing Due to communications issues with distant customers, either the on-duty ground station or the South Pole production facility would be responsible for remote control. In order to minimize delay times, it is thought that ground station control is the better choice. In terms of piloting, due to the delay times associated with remote control from the ground station, this is not really an option. Should something unpredicted occur, there is simply too much risk involved due to the time required for appropriate reaction. One counter-argument to this is the concept of passing remote M-LSAM control to the customer on the ground. This would result in minimal delay time, but would place the responsibility for M-LSAM landing, payload delivery and M-LSAM return on the 100

Introduction

customer. It would likely require extensive training in advance for the “pilot”. Once on the ground, remote control is much simpler. Delay time is not the primary issue and fuel transfer could easily be directed remotely (via the use of robotic transportation means and transfer lines) from the ground station, South Pole station or the customer. One advantage of this is that there is no need for human ExtraVehicular Activity (EVA) to facilitate this type of servicing. Should an EVA not be possible by the customer, refueling can still be accomplished.

Autonomy Autonomous servicing is highly dependent on the abilities of the intelligent systems in use. With autonomous piloting, there is potential for extremely accurate and safe landings, and depending on the level of program complexity, there is the possibility of quick reactions in the case of emergency. Autonomous service, as in remote service, would require minimal additions to launch masses, and no additional infrastructure for life support. Autonomous fuel transfer should be possible with the use of standardized systems as long as sufficient and accurate sensor data is being supplied to the program in use. This type of servicing has no issues associated to communication delay time. One obvious disadvantage to remote servicing as well as autonomous servicing is that they are limited in their ability to problem solve as well as to react quickly in case of emergency. Although autonomous programs can be made adaptable, unpredictable situations may arise, which a given program is not able to compensate. Autonomous systems may also behave in an undesirable way, due to things such as Single Event Upsets (SEU), as sensor data could be misinterpreted. Sensor malfunction is another worry in this case. If autonomy is the desired method, all transfer processes should be closely monitored remotely to ensure delivery, as well as provide the option for manual override if necessary. In conclusion, it is recommended to use a remotely controlled M-LSAM, so as to associate advantages related to unmanned vehicles, while keeping a good level of human interactive capability to cope with occurrence of contingencies.

4.3.3Local Service with Wheeled Rovers A wheeled rover is used as a “tow truck” and would be used to move around mobile tank unit trailers, LOX and LH2 tanks and supporting equipment, on a wheeled platform, as presented on Figure 4-27. The tank units would be picked up at the production and storage facility in the shadowed crater, and brought at the lunar base (life support) or to an ascender coming for refuel in the lunar base close vicinity (a few kilometers). It would come back empty at the production facility and would leave the mobile tank unit there for refueling. Between delivery events, the wheeled rover would be available for other tasks, as explained in the introduction.

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4.3.4Equatorial Service with the M-LSAM The M-LSAM will serve as the fuel transportation unit, commuting between the launch area on the South Pole (near the productionstorage site), and the consumers, located in the equatorial region on the near-side of the Moon. The M-LSAM will service one customer at a time. This is a realistic estimate, as only a few lunar ascenders (customers) will come to the Moon every year in the near-term (low demand) scenario. The launch area will be located at a safe distance (a few hundred meters) from the production facility. The M-LSAM will use its walking capabilities (ATHLETE) to travel between the launch area and the production facility for refuel from the larger fixed storage unit. In addition it can traverse between the landing location and the customer upon landing. The main risk associated with delivering fuel on the lunar surface is the necessity of its supply to the customer. The customer’s survival is dependent on the delivery of the fuel, if their ascender has been launched with only enough fuel to arrive on the lunar surface. With this in mind, the majority of the risk should be on the shoulders of the fuel supplier, and not on the customer. The M- LSAM will land in the area of the target (either the ascender on the way there or the production-storage facility on the way back) and move to within approximately 100 m using the ATHLETE to prevent creating a dust cloud at the target, provided the targets are properly beaconed. The additional close range mobility to get right to the target will be provided by the ATHLETE rover module of the M- LSAM. We need communication between lunar base (manned) and the MLSAM and the customer ascender via Earth.

Pre-arrival Delivery In the pre-delivery scenario, communications between the customer and the supplier will be established prior to M-LSAM launch, in order to determine their targeted landing location as well as to meet their needs post landing. Also, visual observations of the targeted landing area will be provided by the customer from orbit, in order to facilitate the safe landing of the M-LSAM. The launch trajectory will be calculated from data provided from in orbit customer observations and observations made by the ground stations on Earth. The M-LSAM will be launched, and will land at the predicted customer landing site. Upon arrival, a homing beacon will be activated to aid the customer in locating the fuel supply on the surface, as well as assist a landing within an acceptable range for fueling. By delivering the fuel before the customer touches down, we minimize the risk for the customer, and maximize the risk for the provider. Should the M-LSAM not be able to launch for whatever reason, the customer will still have sufficient ability to abort. Should the customer land outside of the acceptable range, servicing will be impossible, and a back-up LSAM will have to be sent for servicing in its place.

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Post-arrival Delivery In the post-arrival delivery scenario, communications between the customer and the supplier will also be established prior to M-LSAM launch. This will avoid unnecessary launches, oversized launch payloads and inaccurate trajectories. The launch trajectory will be calculated using observational data received from the ground stations on Earth, which will provide the exact location of the customer on the lunar surface. Data will be provided by the customer regarding lunar surface properties surrounding their landing site. From this data a preferred landing target for the M-LSAM will be determined. Upon determination of the preferred landing site, it is the responsibility of the customer to activate a homing beacon at the desired location. This homing beacon will be provided by the fuel supplier, and will be brought from Earth. By delivering the fuel post-arrival, the risk is minimized for the supplier. In the event that the customer decides to abort, no fuel has been expended by the supplier. Should the M-LSAM not be able to launch, a back-up LSAM would be launched in its place. Should no LSAM be able to launch, the customer is left to fend for his/her self. This scenario is acceptable provided an emergency service is available to mitigate contingencies on the part of customers.

4.3.5Emergency Services with the Modified LSAM Standard safety procedures for launch will be followed prior to each M-LSAM delivery, including vehicle inspection and operational testing on-site. Upon passing safety inspection, the M-LSAM will be cleared for launch and will depart on its predetermined launch trajectory. It will be requested that the customer performs a simple step-bystep inspection of the M-LSAM prior to and after filling procedures have taken place. This inspection will be made as simple as possible, but will increase the probability of discovering a problem should one arise. It will also ensure customer safety in the event of system damage/failure. Identical launch procedures regarding determination of ideal trajectory will take place, and operational tests will be completed via ground station control. After customer and operational tests have been completed, the M-LSAM will return to the launch area. The emergency service would be a modify service from the equatorial general service use. Emergency delivery of propellant will occur when the tank of a user fails. Therefore, on top of delivery fuel, new tanks should be delivered. The payload consists of one standard 3 600 kg LOX tank, one standard 450 kg LH2 tank and the associated tank subsystems. This differs from the general delivery service, where fuel is carried in M-LSAM own tanks. It is also recommended to specifically dedicate an M-LSAM for emergencies, which would be available anytime very shortly. This emergency M-LSAM would be stationed in stand-by with full tanks in the shadowed crater. As Figure 4-29 indicates, assistance in equatorial regions could be brought less than two hours. 103

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Figure 4-29: Emergency delivery storyboard Admittedly, the emergency service would be very expensive to operate. Minimizing human safety issues would however encourage lunar exploration missions and activities on the Moon in general.

4.3.6Maintenance Maintenance routines must be designed for minimum human interaction and training. This implies modular construction so that repairs and replacements are simple sub-system swaps. It also requires a self-diagnostic system for the vehicles to eliminate time-intensive trouble-shooting. All vehicles will be maintained at the South Pole production-storage site. All spare parts must be on site, delivered from Earth. The rovers and M-LSAM will have monitoring sensors communicating with the base constantly providing information on the health of the vehicle and fuel storage (e.g. pressure and temperature). Cleaning will be performed on a regular basis to prevent contamination and degradation of moving parts from the abrasive lunar dust. This can be performed initially by humans and later by robotic washers (similar to the automatic car wash at gas stations). A back-up rover can be used to tow a broken rover to the base for repairs. Minor repairs can be performed at the delivery site using the tele-operation system (for getting back to the base).

4.4The System Blueprints The delivery and storage systems were designed in collaboration. This section gives details of what is needed to implement the 104

Introduction

proposed system architecture development plans.

given

the

projected

lunar

4.4.1Storage Implementation Storage Units will be brought from Earth as parts and will be assembled on-site. The spherical tanks will be brought in sections (imagine an orange cut in half, and then each half sliced into four identical pieces). The method of assembly is not described here, although it is assumed that a method for surface assembly can be devised. The tank casing including subsystems will be modular, and will also be integrated and assembled once on the lunar surface. This will maximize volumetric efficiency of the tank related cargo, and therefore allow for more cargo to be transported.

Number of Tanks The number of tanks required depends on the number of customers needing servicing, and the desired flexibility of the system. The overall depot capacity is only related to the number of tanks when determining tank sizing. The objective of the storage architecture is to avoid hindering the ability to produce fuel at any given time, given the chance of multiple customers or emergency needs. Four identical 3 600 kg tanks for LOX and four identical 500 kg tanks for LH2 will be required. The tank mass dimensions are based on projected customer needs. The specifications for the tanks are found in the Table 4-29. The tank masses include a 3% loss margin for LOX and a 9% loss margin for LH2. The specifications were calculated using the equations in Appendix B. Table 4-29: Tank dimensions Contents 3 600 kg LOX 500 kg LH2

Volume [m3] 3.2

Radiu s [m] 0.9

7.0

1.2

Eight tanks (or four “units”, consisting of one LOX tank and one LH2 tank) are necessary in order to provide a system with the flexibility to support a lunar base, provide fuel servicing to customers near the base and at the equator, as well as provide minimal redundancy. One must design a system that is capable of handling a period of high turnover in the event of its occurrence, to provide sufficient instantaneous capacity and redundancy, as well allow for growth.

The Tank Design Tank masses were calculated based on tank volume, material density and calculated thickness for the materials considered. The materials chosen for assessment were also compared based on their coefficient of thermal expansion (CTE), as well as thermal conductivity. High thermal conductivity and high coefficients of

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thermal expansion were regarded as negative attributes, as higher thermal conductivity means higher energy consumption, and a higher coefficient of thermal expansion mean more thermal stress effects due to temperature cycling. These values are listed in the table below. Thicknesses of the inner and outer vessels were assumed to be the same. For safety purposes, tank thickness was not allowed to be less than 1.5 mm for LOX and 2.0 mm for LH2. The material volume required for the construction of the inner vessel was doubled to approximate the amount of material needed for both the inner and outer vessel construction. Table 4-30 Tank dimensions by material (www.matweb.com) Tank Thickness Material

Teflon Carbon composite Stainless Steel Aluminumlithium alloy

Tank Mass

Therma l Conduc tivity [W/m/K]

CTE [μm/m/K]

LOX [ mm]

LH2 [mm]

LOX [ kg]

LH2 [kg]

1.5

2.0

167

457

0.2

14x10-5

2.1

3.4

77

209

50- 100

6

3.6

5.9

250

564

17

15.9

1.5

2.0

85

190

88

23.6

As one can see from the above, although aluminum lithium alloy does not have the ideal thermal characteristics, its advantages in mass are apparent. Stainless steel was found to require the highest mass for tanks overall. In terms of thermal properties, Teflon has the lowest thermal conductivity, as well as the lowest CTE. Aluminum-lithium alloy has the highest CTE of the materials assessed, and has arguably the highest thermal conductivity. With effective passive thermal insulation, thermal stress effects and heat flux through the tank walls should be minimized. Therefore, based on the available data, aluminum lithium alloy is the recommended choice for tank material. All tanks are designed with identical sensors, compressors, valves, ports, supports for ease of maintenance and use. All sensors are modular and are removable to facilitate maintenance and repair.

4.4.2Local Delivery System Implementation Two rovers are needed: one for servicing (assuming low demand, i.e. one costumer at a time) and one for back-up for continuous support in case of a break-down or during maintenance cycles. The wheeled rover is fully tele-operated. It will connect to a tanker at the production-storage site, tow the tanker to the consumer site and connect for transferring fuel. After replenishing the consumer, the rover returns the tanker to the production site. The tanker will be in line for refilling. The delivery of the lunar propellant shall be done through the use of wheeled rovers. In that perspective, many constraints arise

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with regards to the rover design. First, the rover shall be able to cover a range of approximately two km. It must also be able to carry a payload of 4 tonnes. On the way to the destination site, the rover will travel over rough terrain and through regolith. The maximum inclination due to terrain shall be 30º. The rover must withstand a maximum temperature range between 200K and 365 K. These constraints assure the success of the mission and also reduce the risk of failure for the rover.

Key Subsystems – Power The rover needs power for mobility. This can be provided by photovoltaic cells, batteries or fuel cells. Fuel cells were chosen since they will be also used for thermal control of the propellant tanks (see Table 4-31). Photovoltaic cells will be added to this subsystem, providing redundancy and increasing reliability. The total power consumed by the rover to cover a distance of five kilometers (two kilometers to and from delivery site and one kilometers safety margin) is 10.5 kWh (see calculation in Appendix B). Table 4-31: Rover power system options Concept

Advantages +LOX and LH2 already in production on the Fuel Cells Moon, so abundant supply is available +Self-contained, no Batteries production facility required

Solar Cells

Disadvantages -Fuel cell technology still needs development; therefore current reliability status is unknown. -Heavy (Earth-supply consideration)

-Can only be used during daylight, cannot be the primary power system. -If more power is needed, +Abundant unlimited larger bulkier solar panels supply of solar radiation. are required, not feasible on +Technology is reliable a moving rover. -Power deteriorates with exposure to lunar environment (radiation, dust), high maintenance

Wheel Configurations The purpose of the rover wheels is to allow the rover to overcome obstacles and drive through the lunar sand without much resistance. They are also responsible for assuring a smooth journey to avoid damage to the rover and the propellant tanks. Increasing the number of wheels improves the redundancy of the system. However, fewer larger wheels make a simpler, lighter design. The high, thin wheels have less “bulldozing” resistance and better clearance, but have packaging problems (Bufkin et al. 1988). After analyzing all those options, it is recommended that the rover be designed with small wheels to increase stability and redundancy (see Table 4-32). Also the materials of the wheels

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should be metallic, but at the same time flexible. The NASA Lunar Roving Vehicle (LRV) is designed with metallic wire mesh wheels (Young 2007) (Rubber and plastic cannot be used due to extreme temperatures). Table 4-32: Rover wheel options Concept Smaller wheels Large / lighter wheels

Advantages + Increase the consistency of the rover + Small, so can be made redundant

Disadvantages - Will cover lesser surface area in a given time - Need many small wheels (more parts, so more maintenance)

+ Are simpler and are light in weight - Have packaging problems + Less “bulldozing” resistance

4.5Technical Risks Technical risks have been identified for the storage and delivery elements.

4.5.1Storage The highest overall risk is refrigeration failure, which would lead to cryogen loss. The best mitigation is a dual system. Explosion, while it has a high potential for damage, is improbable since there are no reactants in the lunar environment (i.e. no atmosphere). This is resolved with separation of volatiles and mitigation strategies for individual components. Structural failure and transfer arm malfunctions are less severe, but slightly more probable because of the complexity of the systems. Modularization and a local supply of pre-tested spare parts are good mitigation strategies. Sensor failure is the most likely, but has little impact with minor mitigation by redundancy. The least likely or severe is a valve malfunction – this is well-proven technology and a back-up regulated valve can mitigate this.

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4.5.2Delivery The technical risks associated with the wheeled rover and LSAMATHLETE system are low to medium. The wheeled rover is a low risk since it is a renovation of a relatively simple proven technology. The LSAM is low risk because it is necessary for the higher level purpose of supplying the lunar base from Earth, so the development is assured by NASA. Both the LSAM and the ATHLETE are based on proven technology. The simplicity of the LSAM ballistic component and its close relation to decades of history in rocket propulsion also greatly reduce the risk. However, adding the complexity of the new ATHLETE rover is a moderate risk – during use, the articulated joints and wheeled drive train imply multiple failure modes, but these are mitigated by the redundancy in components (multiple arms and independent wheel shafts) and the increased terrain capability and the utility of the whole system to delivery to an area without the rover component.

4.6Adaptations Scenario

for

the

Optimistic

Given an optimistic future, with potential customers supplying cryogens to orbit for Mars missions, the proposed system architecture will scale well.

4.6.1Scaling Up the Architecture In the optimistic scenario, production will remain at the South Pole. The storage and delivery system will be scaled up to meet an estimated 150 tonnes demand per year. In later phases of the mission, there might be a high demand for propellant. In order to meet that demand, the delivery system will need to be scaled up. This would mean increasing payload capacity, turn-around time or the number of functioning vehicles at one time. Deliveries in the local South Pole would increase and the individual delivery demands would be higher. To meet this, the mobile storage tanks should be scaled without exceeding the load of the local wheeled rover and trailer. The number of rovers may also increase if the lunar base expands and demands become concurrent, although it is unlikely to exceed the carrying capacity of the original system. A third rover is recommended to provide more redundancy for the higher demand.

4.6.2New Services The future of lunar based propellants.

4.6.3Recommendation: Architecture

A

Proposed

The recommended system architecture consists of modular storage tanks for LOX and LH2 with standardized interfaces to a

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fixed production-storage site and mobile delivery systems. For the local baseline scenario, wheeled rovers with trailer storage tanks are considered the most technically mature delivery system that can easily be scaled to suit future needs at the South Pole. For the equatorial baseline scenario, the M-LSAM (NASA’s LSAM lander combined with the ATHLETE rover) is recommended for either predelivery for human exploration missions or for emergency “doorto-door” service anywhere in the southern hemisphere. In the event that a Mars mission demand scenario develops, the local delivery system can be scaled up to meet the needs of orbital service vehicles. This begs the question - can a business case be built to support this architecture?

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_____________________________________Chapter 5

5 The Business Analysis

In Chapter 3, the primary shape and structure of storage tanks and delivery infrastructure that make up the lunar gas station were introduced. Once the preliminary system architecture has been selected (Chapter 4), it is essential to evaluate the economics of establishing them and turning it into ‘regular business’. The idea of a lunar gas station brings a financially lucrative image to mind, but the burden of heavy initial investment can impose a barrier to entry into market. Through a preliminary economic evaluation carried out in the subsequent sections, the cost of delivered LOX and LH2 on the lunar surface is determined. Costing of products derived in space still remains an unmastered art; hence a variation of market demand (projections, See Chapter 2) and business models is evaluated. It is assumed for all through the analysis that the production facility for LOX and LH2 is in operation and that the lunar gas station storage components are established in its vicinity. The costing, financial strategy determination and risk analysis that follow derive heavily from assumptions (see Section 5.1) and the reader is requested to consider these. Different business partnerships options are evaluated to identify the candidate organizations interested in embarking on turning the lunar gas station to reality in the next 20 years. Finally, pricing (based on two Market projections) is followed by considerations for new taxing strategies.

5.1Approach and Overview To assist the reader in understanding the factors that influenced the final business solution the process flow of the analysis carried out is explained in Figure 5-30. Current roadmaps detailed in Chapter 2 were taken as market demand for storage and delivery services. To match this demand each component of the storage and delivery process was costed and the cost of the final product determined. Given that in the absence of the said services, one would have to rely on launching from Earth (product baseline), it was used to compare if the business would be profitable for various combinations of market players. Risk, income expectations and partnerships (business models) were used to complement the previous pricing to result in a financial strategy to market.

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Figure 5-30: Business analysis logic flow chart As seen recently, with the completion of the ISS, heavy investment ventures in space business are highly susceptible to the level of International cooperation. Chapter 2 briefly looked at different scenarios pertaining to international cooperation. These three scenarios: No International Cooperation, Moderated level of International Cooperation and Strong International Cooperation, have reflected in the financial analysis to help assess the combination of business partnership best suited to own and operate the lunar gas station. Before one engages in determining if the lunar gas station is profitable, some important considerations regarding the market demand described in Chapter 2 (Sub Section 1.3) must be revisited. Some modifications and assumptions used for the economic analysis are highlighted and limitations explained in the section that follows.

5.1.1Market Demand Overview Current Projections According to the market research performed the baseline demand will consist primarily by NASA (NASA 2004) and its international partners as shown in *Data obtained from the Market demand Analysis Chapter 2. which represents a 22 tonnes demand of LOX and 3 tonnes of LH2 per year. The analysis carried out assumed a 20 year economic life of business, starting in 2028. This timescale albeit too short for such a business, is only analyzed until the year 2047 due to the difficulty in predicting how cis-lunar market would behave in that time. it As the lunar resources market develops, having several international customers and a more diverse space industry, would lead to higher market demand projection, as shown in *Data obtained from the Market demand Analysis Chapter 2. The amounts needed on the lunar surface are expected to be 112 tonnes of LOX and 21 tonnes of LH2 per year.

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Low LOX, LH2 Demand Market Scenario 30.0 20.0

LH2

10.0

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tonnes

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*Data obtained from the Market demand Analysis Chapter 2.

Figure 5-31: Low Demand Scenario of LOX and LH2*

High LOX, LH2 Market Demand Scenario 200.0 tonnes150.0 100.0 50.0 0.0

LH2

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*Data obtained from the Market demand Analysis Chapter 2

Figure 5-32: High Demand Scenario of LOX and LH2*

For the timeline considered in Figure 5-31 and Figure 5-32 the amounts of LOX and LH2 required are different to those deduced in Chapter 2, as only USA and International partners were considered to establish the costumer baseline. On the other hand, an international consortium of USA, Russia, ESA, China and third parties resulting in Figure 5-32.

Evolution of Market and its Players The initial market projections are based on primary lunar bases operations that would be established and run by governmental agencies. As the lunar bases operations develop, a business operating the storage and delivery operations may also be able to increase the customer base to a multitude of other businesses carrying out commercial activities. In the economic analysis, this is reflected as Business to Government, and Business to Business approaches as follows: The Business to Government (B2G) approach is the concept of an entity whose primary customer is the government (FEMP 2003). It addresses the fact that most high technological endeavors are started with government support and that the market is often developed by government. This results for the low market projection are given in Figure 5-31. The Business to Business (B2B) approach is the concept of an entity whose primary customer is another business, non-governmental related. A lunar gas station facility under such an approach would be operated by business partners who are interested in servicing

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commercial as well as scientific mission traffic. This results in the high market projection are given in Figure 5-32.

5.2Supply Overview The cost estimation of the lunar gas station addresses the storage, the delivery and the production part of the architecture. For the purposes of the business analysis of the proposed infrastructure, the cost of the production facility and the process had to be determined. To reflect the infrastructure required to supply the market demand projected in *Data obtained from the Market demand Analysis Chapter 2., the number of production plants, rovers and storage tanks to be developed over a 20 year lifetime is provided in Table 5-33. Table 5-33: Forecasted required infrastructure for total lifespan of 20 years Demand Supply infrastructure to be setup forecasted required* Scenario Average per Production Storage Rovers year (t) plants(t) tanks (t) (t) High Demand 111 36 16 LOX, LH2 Low Demand 25 12 8 LOX, LH2 * For detailed explanation please refer to Chapter 4

20 8

The business case selected that the service of the company should be only through the delivery in a short term radius of operation up to two kilometers, which correspond to a delivery time in approximately two hours.

5.2.1Lunar Oxygen and Hydrogen Supply Service Two options were considered for the approach of a lunar gas station. The first is a self service concept requiring customers to come to the gas station, which is not acceptable due to the variance in customer needs. The second approach is a full service where the gas station goes to the customer using rovers to deliver the product. The full service option better addresses the need of the customer to explore various regions of the Moon and was chosen as the path forward.

5.2.2Assumptions The economic analysis derives heavily from assumptions made for the lunar gas station system architecture (Chapter 4). The reader is encouraged to refer be aware of these along with ones specific to the economic analysis, elaborated as under: •

All values are expressed in USD FY 2006 unless stated otherwise and were translated using NASA New Start Inflation Index (NASA 2005).

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•

The market scenario for the lunar gas station where the operator company performs is a natural monopoly.

•

The interfaces between the production, storage and delivery were assumed to be negligible compared to capital costs.

•

The simplest system of storage and delivery were used for the project cost analyses, and long distance services are not included in this business case.

•

Operation costs are assumed to be 12% of the total capital cost per year.

•

It is assumed that the production facility is operational in year 2028 and a communication network is already set up, resulting in lower operational costs

•

A low dependency on astronaut EVA is assumed, and costs associated are considered negligible

5.3Business Solutions National space policy typically emphasizes the use of commercially available goods and services as a means of encouraging and increasing private investments in space activity. There is already a pronounced interest by the commercial space sector in cis-lunar activities including re-fuelling in orbit as well as on the moon surface (refer to details in Chapter 2). However, the major factor inhibiting entry is the high capital required for investment and the associated risks. Here, the main types of business solutions are examined to analyze their effectiveness for establishing a lunar gas station.

5.3.1Public Cooperation Development Phase

Scenarios

for

The possibility of the idea to come to fruition relies heavily on the socio-political scenarios i.e. the extent to which agencies and private industry are partnering and cooperating, often dictated by policy changes in their countries. The three main business solutions that are assessed for different socio-political scenarios are detailed as in Figure 5-33. It is assumed that production facilities are already set up, primarily by NASA or a US space industry integrator. This assumption for ‘first to market’ is purely based on evidence from published documents, which suggests that plans of NASA and its industry partners are further developed than those of several agencies that have merely registered an interest in lunar activities.

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Figure 5-33: Multi Public Private Partnership example for a Lunar Gas station

No International Cooperation The lunar gas station established in a socio-political environment allowing no international cooperation would be owned by a single entity private investment or government. Three possible investors are shown in Figure 5-33 and are a private company such as a space integrator, a national space agency, and a company based on different joint venture concepts. Although this might be the more attractive ‘First to market’ approach, it would require a sustained flow of heavy funds. Even if the high initial investments are financed, the space agency needs to maintain a strong commitment to this program and build an important support based on opinion of the taxing paying public. Furthermore, the market may be very small and it will be difficult to find others customers of the lunar gas station. Consequently, in this scenario, it is not sure if the project can be profitable.

Strong International Cooperation This cooperation would be based on a strong multilateral cooperation between space agencies (NASA, Roskosmos, ESA, CSA, and JAXA, maybe ISRO or CNSA) and potentially some private companies. This scenario corresponds to stronger interrelationships than is currently seen for the International Space Station. The application of this international cooperation will depend on the evolution of the lunar exploration plans of the space agencies and their will to cooperate for a common goal.

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Moderate International Cooperation A strong international cooperation can be difficult to implement due to a large number of reasons: internal political issues, problems of specific bilateral of multilateral relationships, non confidence between partners, political rivalry and cultural differences. In the last case, the study assumes a moderate situation resulting in some level of international cooperation as an efficient compromise. This model is often referred to a Multi-Public Private Partnership (MPPP) (Zervos 2005). Referring to Figure 5-33, the venture first begins with the management of the project being taken by different space agencies sharing the development of the facilities. NASA can be in charge of production facility and transportation, ESA would take charge in the storage development and Roskosmos of the delivery, and accordingly negotiate their use of the lunar gas station. This relationship would lead to a change in pricing policy as well to different partners. For example, for the main investor (NASA) the price would be set as price of the product minus a percentage discount. This discount could be determined by the amount of investment over a timeframe that NASA expects to get the invested capital back for the production infrastructure set up. The percentage of their investment in the overall lunar gas station would also come into play in calculating the discount. This would be done likewise for the other partners.

5.3.2Private Joint Venture Operational Phase Considering the investment it makes economic sense to form a joint venture to lessen the financial burden and risk on the individual company. Offsetting the initial investment of the government reduces the “payback” required by the private member of the PPP. It also pulls together the expertise base of the merging companies where each company’s specialization can be exploited within the venture. (Hennesy 1992) The role for the government again can take various forms. The two depicted in the Figure 5-33 are national agency to a joint venture of international space integrators and the other is a joint venture of a US space integrator and a big oil company.

International Joint Venture The national space agency contracts its main space integrator to build up the infrastructure on the Moon surface for lunar gas station. When the operation starts, the main national space integrators build up the international joint venture to operate the lunar gas station. According to the investment contributions, it can share the high risk and high cost operation. The disadvantage of this case is inherent in the difficulty to cooperate between different national space communities. There are lot of limitations such as Exports Control Regulations, technology transfer issues and security issues.

USA Joint Venture When discussing the business of a lunar gas station the first thought to come to mind is would big earth oil companies be interested? The big oil companies would bring forward many benefits for the joint venture.

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Not only there is a solid link in business sector, but also these companies maintain extremely high revenue streams and spend it heavily on corporate image. Investing in the lunar gas station can therefore be profitable to the oil companies since the project reinforces their advertising strategy. Practically, their brand names could be displayed on the lunar gas station infrastructure while their support to the project could be highlighted in their advertisements or in media. The financial resources provided by these companies are difficult to estimate precisely and certainly could not cover all the costs of the lunar gas station project. In exchange for the financial assistance from the government within the PPP approach, the private company will be expected to pay back some portion of that investment to the government. One method that accommodates the financial risk undertaken by the private company is for the private company to pay a royalty to the government. This method will enable the government to receive a percentage of the company’s revenues, once the project becomes commercially successful, regardless of the company’s profits or expenses. The length of time required is dependent upon the amount of government investment, how much they expect in return, and the level of profitability of the company. For the Galileo program, a good model of a PPP scenario, the royalties are evaluated initially around 2% to 5% (PriceWaterhouseCoopers 2001).

5.3.3Business Solutions Summary This lunar gas station undertaking will require a heavy investment during the project life-cycle especially at the initial phase. From the analysis, it is deduced that the case of no cooperation results in high investment and risk, thus making it highly improbable. It is difficult to implement a strong international cooperation historically, which suggests a low probability for this scenario. Based on the above preliminary assessment, one can realistically expect the lunar gas station to happen with minimal international cooperation, possibly at a smaller extent than international partnerships within the ISS project.

5.4Business Risk Assessment A risk assessment carefully examined and can be Appendix D. It consists of identifying the hazards present and then evaluating the extent of the risk they pose. By carrying out a risk assessment, it should be easier to assess whether or not enough precautions are considered to reduce the potential risk and to assess whether or not more controls are needed to prevent harm. The highest foreseeable risks are given in Table 5-34.

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Table 5-34: High risk Type of Risk Feasibility of the financial cost models Market demand too small to justify investment (overestimation) Partnership failure

Unacceptable responsibility service

of

level of refueling

Current NASA policy does not allow landing without fuel to return

issues for business analysis Mitigation Finance R&D to come with more realistic options Not be too ambitious in infrastructure, keep it simple and do not project the market too far forward Establish and sign an International Governmental Agreement to bind the parties that belong to the contract Maintain appropriate insurance coverage and create proper standards for lunar surface refueling Work far in advance to developing the business model to ensure that NASA can find a common ground regarding this policy.

5.5Cost Breakdown Analysis For space missions for lunar exploration and commercialization cost estimation is difficult and still remains an uncharted territory (FERTILE Moon 2006). Figure 5-34 suggests the approach that was adopted for the estimation of cost of different components within our supply service. Different components have been costed by analogy based on a review of past lunar missions (both manned and unmanned) and current space exploration proposals. This data is used to create cost equations and assumptions that can later be used to calculate elements of production, storage and delivery cost. The accumulation of all these cost categories resulted in a final product price.

“ISU FERTILE Moon” cost Model

Space Mission cost data

Specific cost Equation & Assumptions

Estimate Production System cost

Estimate Delivery System cost Estimate Storage System cost

Total product cost

Figure 5-34: Cost analysis approach for production, storage, and delivery systems

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5.5.1Cost Categories To understand the business in supplying lunar derived products, the costs of production, storage and supply need to be less than that of launching products that are ‘ready-to-use’ from the Earth. In order to do this seemingly simple calculation, one has to determine the cost of the LOX and LH2 to the lunar gas station operator. Although, the operator may not be producing the LOX and LH2, it is expected that they operate in a manner similar to petroleum gas stations on Earth and price the final product accordingly. The cost components considered include capital costs required to develop, construct and launch required facility components, whilst operating costs considered are labor, costs for buying energy to run the facility, consumables and maintenance. Although the latter category provides a wholesome view of the costs involved, the capital costs form the bulk of the overall costs. The different variables that are significant for calculating the cost largely depend on the level of the technology or process, its technological readiness and complexity and described in Table 5-35. Table 5-35: Cost category breakdown Cost Element Factors Involved in Parametric Costing Term Capital Cost Mining Equipment Mass and complexity of lunar Development and production facility * Production Storage Equipment Mass and complexity of lunar storage Development and facility Production Delivery Equipment Mass and complexity of lunar delivery Development and facility Production Launch Cost Specific Earth-Moon transportation cost (USD/kg) Operations Cost Maintenance and Consumable consumption rate (kg/yr), Consumables Cost specific Earth-Moon transportation cost (USD/kg) Energy Cost Specific power of lunar facility, specific cost of power (USD/kW) Labor Cost Tele-operation costs, EVA costs * Assumed that the facilities has already being constructed

5.5.2Launch Cost The mass of the system is the driving component due to the high cost of launch from earth. Based on a survey of current launch masses and launch costs, and not considering the cost of launching to the moon using systems that are under development, such as the CEV, the average specific launch cost is 42 400 USD/kg (FERTILE Moon 2006). Using this value with inflation, the launch cost for the present analysis has been estimated to be USD 43 200/kg.

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5.5.3Capital Cost Production Cost A tool was developed at ISU in 2006, in order to assess the cost effectiveness of producing lunar oxygen and hydrogen in-situ by comparing it to the price baseline of launching all materials from Earth. This tool, named FERTILE Moon (2006), is capable of providing an adequate cost for the production of lunar oxygen and hydrogen and was used as an input to the cost analysis described here. The production facility demand and processes were matched with production processes described in Chapter 2 so as to give a close comparable estimate. Table 5-36: Inputs and outputs from ISRU model for supplying LOX and LH2 (FERTILE Moon 2006) Inputs Outputs Production Time 30 days Process Cost: USD 26 Million Hydrogen 333 kg Process Final USD 49 Demand Cost Million Oxygen Demand 2 676 kg Electrolysis yes Option

Storage System Cost As lunar propellant storage represents a missing link within lunar roadmaps, it is hard to come by credible estimates of such equipment and limited studies have looked into the issue of development costs. The tank specific costs were sourced from NASA Exploration Team Case study for commercial lunar ice mining as listed in Table 5-37 (Blair 2002). Table 5-37: LOX and LH2 tanks cost data from ISRU case study (Blair 2002) Development First Unit Mass Tank Cost Cost [USD [kg] [USD Million] Million] LH2 Tank 450 8.4 0.78 LH2 Tank Specific 18 KUSD/kg 1.67 Cost KUSD/kg LOX Tank 1 999 17.5 2.0 LOX Tank Specific 8.8 KUSD/kg 1 KUSD/kg Cost Using the tanks specific cost figures from the Table 5-37, the development cost of the storage solution chosen within this study (Chapter 3 and 4) is calculated by multiplying the specific cost of the mass and thereby obtain the cost numbers in the Table 5-38. The two types of tanks will be used for storage service. Table 5-38: Cost for proposed LOX and LH2 tanks

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Tank LH2 LOX

M ass [kg] 445 218

Life Time [years] 10 10

Development Cost [MUSD] 8 1.9

Unit Cost [MUSD] 0.7 0.2

Delivery Capital Cost – Rovers Rovers have been extensively used for space missions and are developed individually, depending on their main tasks. With the systems selected for the given study, a review of past rover development costs was undertaken as shown in Table 5-39. Table 5-39: Past Robotic Rover Development Cost M Life Developing Distance/day Mission ass Time cost [m/day] [kg] [days] [MUSD] Spirit 150 90 100 400§ Opportuni 150 90 100 400+ ty Apollo Moon 92km/day 208 720 190* rover (12.6km/h) (LRV)^ + (Marsha 2006) (FY 2004) * (Williams 2006) (FY 1971) ^Lunar Rover Vehicle These rover costs are particularly high-end, given that the Mars Rovers Spirit and Opportunity are both designed to be autonomous. The rovers selected within the proposed architecture are not expected to perform the same level of scientific tasks and are closer to the cost of the LRV used during the Apollo missions (here, converted to FY2006 for consistency) as shown in Table 5-40. The unit cost ratio compared to the development cost is 12% and 30% according to the (NASA Cost Estimating Handbook 2005) and (Koelle 1996), the rover unit cost is assumed 16% of the development cost. The rover development cost is 190 million, then the unit cost is 31 million according to specifications described in detail in Chapter 4 , with the launch cost included (USD 110 million), the unit cost is assumed to be USD 141 million per rover. Mission

Rover

Table 5-40: Rover specifications Mass Lifetime Launch Developme [t] [years] cost nt cost [MUSD] [MUSD] 2.6

5

110

190

Unit cost [MUSD ] 141

5.5.4Operation cost The operation cost is assumed to be 12% of capital cost of system. This includes the maintenance and consumables cost as well as the labor cost according to space program historical experiences.

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5.5.5Summary of Cost Estimation The cost of producing the LOX and LH2 would affect the final price of the product delivered to the consumer. It is only possible to fix this production cost estimate if it can be determined whether the LOX is derived from regolith or ice core mining. This, as suggested previously in Chapter 3, awaits confirmation of the presence of ice on the lunar poles. This gives two options for the costing of different market demands, as used in Chapter 3 and 4, the ‘no-ice’ and ‘ice’ options. As shown in the Table 5-41 the total capital and operation costs that were identified are presented including both scenarios by having ice or no ice on the Moon and also the scenarios having a high or low market demand.

Items Total Capital Cost

Total Operatio n Cost

Table 5-41: Cost estimation Product System Scenario Cost [MUSD] Low market + Ice 3780 Low Market + No 4300 ice High Market + 24035 Ice High Market + 158774 No Ice Low market + Ice 4500 Low Market + No 5800 ice High Market + 13512 Ice High Market + 9505 No Ice

results Storage System Cost [MUSD] 256.6 256.6

Delivery System Cost [MUSD] 1888 1888

533

4720

533

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30 30

226 226

64

568

64

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In conjunction with Table 5-41, Figure 5-35 gives the cost breakdown for the storage and delivery solutions in the architecture proposed (Chapter 4).

Ice/ no-ice: The main difference in cost for these two scenarios is

that the two mining methods costs are different, the no-ice option being higher. For the operator of the lunar gas station, delivery and storage costs do not change but would reflect in the pricing for the final product.

Mining: The price of the product has been calculated by the tool

FERTILE Moon (2006) and the reader is requested to refer to this model to understand how its underlying assumptions have affected the final costs presented in Table 5-41.

Storage system: The storage system is relatively a low proportion

of the final price. This is an effect of the assumptions taken during the design of the storage system (tanks are extremely light weight), as well as the fact that the active thermal control system has not been included in costing. To compensate, electricity charges to be paid to the lunar base (to run cryocoolers) were estimated and factored in.

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Delivery: The delivery system represents a significant portion of the costs, because as demand rate increases by the years, more numbers of rovers would be required to provide more fuel. Given that the rover life time is only five years, replacement may be needed. Cost per Year MUSD

200000 180000 160000 140000 120000 100000 80000 60000 40000 20000 0 High demand no ice

High demand no ice

High demand Ice

High demand Ice

Low demand no ice

Low demand no ice

Low demand ice

Low demand ice

Scenario Operational cost

Capital cost

Production cost

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Figure 5-35: Cost breakdown result overview

5.6Financial Model The financial model assesses where the required funds come from (revenues and financing) and what they are used for (recurring and non-recurring expenses). This information is used to calculate the performance of a private sector business using valuation metrics such as the income statement documents (to show profits and losses) and the cash flow statement. The statements incorporate assumptions on the project’s capital strategy, which is the choice of the debt and equity proportion used for funding. These pro-forma statements require four types of financial inputs that in turn rely on outputs from the demand and engineering analyses. These inputs are: •

Revenue inputs (functions of the market share)

•

The costs of investments (capital expenditures), operations, storage and delivery

•

Products sales and general administrative inputs

•

Taxes and royalties

These required outputs lead the development of this integrated engineering and economic modeling. The business economic viability is measured using the financial selection criteria such as the net present value (NPV) and the internal rate of return (IRR).

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The way these assumptions are used and how they get to our financial analysis are shown in the following graphic (Figure 5-36).

Figure 5-36: Financial Model

5.6.1Calculating Annual Costs and Revenues Private Case The financial model used here is a generic way to utilize the two principal financial accounting documents: a balance sheet, which shows the profits and losses of the revenue and information and a cash flow statement that characterizes the venture’s annual cash flows. These documents are used to calculate the performance of a private company. It should be noted that the analysis incorporates the assumption that the capital invested in the project is 100% equity funded. The proforma statements require different types of financial inputs that in turn rely on outputs from the demand and engineering analyses (see Chapter 2 and 3). These inputs are: •

Annual Revenue = (price of product) * (number of total sales)

•

Annual Cost = (capital cost) + (operation cost)

•

Gross Profits= (Annual Revenue) – (Annual Cost)

•

Insurance = x * (Annual total Cost)

•

Taxes = Income tax rate/(1 – Income tax rate) * (Gross Profits)

•

Contingency = y * (Gross Profits)

where x is a percentage

where y is a percentage

These required outputs lead the development of an integrated engineering and economic modeling where the final profits would be obtained.

MPPP For the MPPP Case, in addition to the above mentioned inputs, Royalties to the government were added: •

Royalties= z * (Gross Profits)

where z is a percentage

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5.6.2Pricing Approach It has been assumed that NASA is the largest investor followed by ESA and Roskosmos. With this assumption, it is necessary to establish different prices of products according to the size of the contribution coming from each of the main investors. For example NASA may invest mainly in plant facilities, ESA for storage systems and Roskomos for delivery parts. The prices, based on the potential percentage of involvement, have been identified and listed in the Table 5-42 for each costumer. Table 5-42: Pricing of product for each costumer Customers Discount of price of products (Market Share) LMD+IC LMD+No HMD+IC HMD+No E ICE E ICE NASA (60%) 18% 20% 25% 100% ESA (15%) 1% 1% 1% 1% Roskosmos 9% 9% 5% 5% (15%) Other Parties 0% 0% 0% 0% (10%)

5.6.3Selection Criteria In order to select which scenario would be better for the project, the NPV and IRR selection criteria have been chosen to validate the business case.

Net present value (NPV) The NPV indicator is useful for an investor to measure the current value of a project and gives an input to choose a profitable investment. Assuming the project will last T years, the NPV will be: NPV = CO + Σ [Ct / (1+r)t ] With CO = Initial investment at time 0, which is usually negative (PV cost) if it is an investment. If the investment is viable, the NPV has a positive value. To know the value of the NPV, we can use the internal rate of return (Brealey 2003).

Internal rate of return (IRR) The IRR is defined as the discount rate which makes NPV=0. Applying the IRR discount rate into the NPV equation, we can deduce the following equation for the internal rate of return: Σ Ct / (1+IRR)t = CO As we can see graphically in Figure 5-37, which shows the financial result in a typical investment scenario with an IRR of 39%, and the IRR is greater than the opportunity cost of capital of 25% for the project and which makes the business scenario economically viable. If the opportunity cost of capital is greater than IRR, then the NPV is negative. Therefore, when comparing the opportunity cost of capital with the IRR of this project, is basically asking if this project has positive NPV.

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The IRR for an investment is the maximum allowable discount rate that would yield the value considering the cost of capital and risk. This is sometimes referred to the breakeven rate of return.

IRR 39%

at

Figure 5-37: Net present value and IRR discount rate. The preliminary analysis results in an opportunity cost of 25% as the rate of return. This first order financial analysis asserts that a Lunar Gas Station would represent an economic opportunity in a Private or in a MPPP case.

5.6.4Taxes, Insurance and Discount Factor Corporate Taxes A key component of lunar commerce will be the manner of which companies are taxed on the income generated by using space resources. Current corporate tax laws pose a major threat for any future private space endeavor. Table 5-43 reviews the current corporate tax rates in few countries. The current value of 35% was used within the financial analysis and estimating the pricing. Table 5-43: Corporate Tax Rates for Space fairing countries Income Country Tax (%) Limitation [USD] United States 35 > 18 333 333 Germany 25 United 33 > 2 000 000 Kingdom France 33 -

Insurance Insurance protection is needed to offset the high risks and launch failures associated with space business. The insurance rate is generally in the range of 17% to 22% according a study carried out by Futron on

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major brokers and underwriters. Depending on the capacity of the placement, this rate is even higher between the range 30% to 35% (Futron 2003). The rate proposed for this project is 20%.

Discount Rate The discount rate for this project is assumed to be 25%. This is high because of the advanced technology and risky nature of space systems.

5.6.5Cash Flow Analysis and Profitability Analysis Based on the cost of capital committed to the development phase of the lunar project, two scenarios were considered for the financial analysis. It is assumed that there will be ice available at the Lunar South Pole and the lunar base is located near the pole. In contrary, the second case assumes that there is no ice available on the Moon. The availability of ice would have a significant effect on the profitability of the lunar gas station. Moreover, comparing the different levels of tax rate, a strategic choice can be made on the location of the production and development. These analyses are based on the general assumptions for the project’s business plan as shown in Table 5-44 below. Table 5-44: General Assumptions for Business Analysis Items Assumptions of Input Data Economic Life Propellant (Products)

Mass

Capacity Plant/tank/Rover life time Launch Cost (EarthMoon) Income Tax Rate Insurance Rate Contingency Rate Discount Rate Royalties Rate Financing methods

20 Years (2028-2047) Low Market Demand (35 t/year) High Market Demand (Average 111 t/year) 16% margin for production, 1-2 backup for the redundancy of the Storage and Delivery system 7 years/10 years/5 years 43 200 USD/kg 35% of Gross Profits 10% of the Total Cost 20% of Gross Profits 25% 5% Only equity

Two business scenarios has been considered, low market demand and high market demand, while the cash flow graphs for both the private (Figure 5-38 and Figure 5-39) and the MPPP (Figure 5-40 and Figure 541) business ventures are as shown below. This graphics were generated using the general assumptions in Table 5-44, the financial spreadsheets results shown in the Appendix D and the financial results stated in Table 5-45 expressed in values of the NPV and IRR.

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Figure 5-38: Private Low Market Demand

Figure 5-39: Private High Market Demand The cash flow graphs in Figure 5-40 and Figure 5-41 show the MPPP business scenarios, which identify the most profitable business opportunity.

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Figure 5-40: MPPP Low Market Demand

Figure 5-41: MPPP High Market Demand From the forecasted cash flow from the low and high market demands with ice and no ice at the lunar South Pole shows the NPV and IRR results as follows: Table 5-45: Financial results

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Private business

Presence of Ice No Ice Presence of Ice

MPPP No Ice

Low Market Demand [LMD] NPV= -701M IRR= 16% NPV= -4.3B IRR= *

High Market Demand [HMD] NPV= 1.96B IRR= 39% NPV= -28B IRR= *

NPV= 1.9B IRR= * NPV= 1.4B IRR= *

NPV= 6B IRR= * NPV= 776M IRR= *

* Not applicable. For a 20 years life time, the private business for the low market demand (ice and no ice) shows negative NPVs, while IRR is less important than the opportunity cost of capital. This is also the case for private business in high market demand but with no ice. In contrary, NPVs are positive for all MPPP scenarios and for a private business with high market demand and ice. These assumptions make the MPPP business the most profitable scenario. For the business case of high market demand with ice for the private, the business opportunity favors the business case with the ice scenario, with an NPV of 1.9B and an IRR of 39%, which is greater than the opportunity cost of capital for this project. Though the profitability is on a return at a longer projected period, this is the only recommended business case for the private business venture. The others private case does not meet the business proposal in the long run of the 20 year period. Though the MPPP business scenarios are all profitable, the profitability trend of the low market demand is lower compared to the high market demand trend on the MPPP chart. The ice scenario will be more profitable if it can be more realistic and proven, In this case, LH2 would not need to be transported from Earth to the lunar base, which makes it more economic cost effective than without ice.

Financial analysis conclusion An international cooperation like in the MPPP business case is recommended for the commercial companies that are investing in the Lunar Gas Station. This will enable a reduction of inherent risks associated with the business. More, MPPP has the potential to mitigate the risks of the huge cost of investment that private business can’t do and encourage the commercial industries in space business. The cost estimations and revenues are optimistic for the 20 years life time of this proposed project due to the following reasons, which may act as limitations: •

The maximum price of the products is identified for this financial model

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•

The interfaces between the production, storage and delivery were not taken into consideration

•

The simplest system of storage and delivery were used for the project cost analyses, such as the long distance services, are not included in this business case

•

Cost estimation approach is made very broadly

5.7Sensitivity Analysis Now that we have identified the good business scenario and assuming an in-elastic demand in the business cases, a sensitivity analysis is used to analyze the impact of uncertain parameters on the investment financial forecasts and the conditions for financial viability. Key parameters to test the sensitivity include the price of products and the discount rate.

Sensitivity of Price of Product This parameter is very important to influence the market and the private company business. The minimum price of products which makes the business profitable can then be compared with the potential competitor price and its return expectations. The following figures show the results of these analyses for two scenarios. The Figure 5-42 represents a case of MPPP with high market demand with presence of ice:

Figure 5-42: MPPP High Market Demand Ice Price Sensitivity The Figure 5-43 represents a situation of MPPP for a low market demand with ice presence:

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Figure 5-43: MPPP Low Market Demand Ice Price Sensitivity According to the two scenarios, the range of the product’s price is about USD 22.9 million/t to USD 43.2 million/t for a low market demand. In case of a high market demand, this range will be about USD 18.6 million/t to USD 43.2 million/t.

Sensitivity of Discount Rate Private companies use NPV and IRR to account for the perceived risk of the venture: the higher the uncertainty, the higher the discount factor required. Two scenarios were chosen to analyze the sensitivity of discount rate on NPV. One is the private business case for the high market demand and ice; another is the MPPP for high marked demand and no ice. For the first scenario, we can see on Figure 5-44 that the tendency of NPV decreases slowly with a discount rate increase, which range is from 25% to 35%, comparing to from 10% to 25%. So the range of discount rate from 25% to 35% is suggested to evaluate the high risks in private business scenario.

Figure 5-44: Private High Market Demand Ice Price Sensitivity

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For the second scenario, the Figure 5-45 shows a similar curve but with a higher decrease of the NPV for discount rates of 10 to 30%. Here the suggested discount rate for the MPPP is around 40-60%.

Figure 5-45: MPPP High Market Demand No Ice Price Sensitivity

5.8Promotion In order to mitigate marketing risks, it is important that a well structured and defined promotion campaign is established. By having one that promotes the investment and benefits to the costumers by getting the product, the risk can be reduced by assuring a sufficient market for the products. Therefore, a public promotion campaign that shows benefits with respectable cost reduction is very important and it should demonstrate cost-benefit relation of LOX and LH2 instead of using Earth based propellants and also making emphasis that are not readily available and are a life-saving necessity on the Moon. A global promotion campaign will be less efficient than a segmented one (Kotler, 1988) because the message will not have the same effect on different types of targets. The promotion campaign should focus on three main targets: •

Investors in the Lunar Gas Station

•

Customers of the Lunar Gas Station

•

Public opinion

Investors The investor target is relatively small when considering the MPPP model developed previously. Actually, in this scenario the investors are limited to the government agency and national space integrator. So we will focus our promotion campaign on customers and public opinion.

Customers The potential customers have already been analyzed and has been stated that is will be mostly space agencies such as who has lunar missions slated for 2020. However private companies may be involved. For example the US firm Space Adventures, who is planning to launch lunar orbit flights. It will be more efficient to segment the promotion campaign in order to be more efficient with the potential customers. 134

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After some time, private customers can emerge. At this point one can develop a specific promotion campaign for this target which will be based on three main slogans: •

the profitability of using the Lunar Gas Station

•

increase for the global corporate image

•

the image of state-of-the-art technology used for commercial purposes and that can attract customers for the customers

•

the potential market is important

•

response to potential technical risks are prepared and day-today risks are highly controlled

•

lots of potential spin-offs can be implemented for both space and non-space activities in a relatively short-term

In order to diffuse the campaign, it will be build strong partnerships with enterprises like this: •

Automobile industry such as GM, Honda and Chrysler since they are leaders in hydrogen automobiles

•

Oil industry such as Chevron and/or BP since they have research on hydrogen production and gas stations

•

Shackleton Crater Expedition which has plans for a low orbit gas station presently

•

Air Liquide since they supply oxygen, hydrogen and many other gases and services to most industries, including the supply to Ariane launchers and other spacecraft.

Public opinion Convincing the public is very important to get a relevant promotion campaign, because the project, by the actions of space agencies, will be partly financed by public taxes. So space agencies will need a continuous and strong support to achieve the Lunar Gas Station project. To realize this goal, different elements are proposed: •

The creation of a website to promote and educate the public. This could help to increase the outreach process with the public.

•

A strong communication with the press, magazines and TV to promote a state-of-the-art realization

•

Develop public relation with journalists to speak about the project

5.9Recommendations Although preliminary, it is the hope of this first-order evaluation will help bridge the gap that exists within plans for development of lunar

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bases and their sustained operation and recommendations drawn would aid future business plans for a lunar gas station. For the private business, the financial analysis resulted that the economic viability can only be justified if the high market demand and the value of the price of products are consistent on a long term period. This illuminate that private company maybe can make profit if the ice is found at the lunar South Pole and if there is a high market demand. Based on the financial analysis results, the business case that showed a better and more feasible investment that meets the proposed business requirements is the MPPP model; which showed to be the most profitable in all the proposed market scenarios. The economic viability is on a shorter term compared to the solely private investment case, and this makes the MPPP the most attractive option for the high technology space business scenario. Therefore a MPPP is the recommended business solution. According to the MPPP model price analysis, even when the product price drops 40% due the accessibility on the Moon surface of lunar LOX and LH2 compared with the price baseline of the product taken from the Earth to the Moon, it is has been found that the scenario is attractive for a business opportunity based on the assumption that the demand does not vary according to the increase of the availability of the product. If the presence of ice is demonstrated at the Lunar South Pole, the cost of developing and operating a lunar gas station would decrease significantly, providing an attractive opportunity for commercial uses.

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_____________________________________Chapter 6

6 Legal & Ethical Issues

Space activities have changed tremendously since the first artificial satellite was launched by the former Soviet Union. These ventures no longer consist solely of scientific research, but have since developed into technologies capable of meeting the needs and demands of humanity. Inevitably, such development also gives rise to the commercialization of space activities, necessitating a legal framework. However, while the existing legislation is well defined for developed exploratory and scientific endeavors, the law is not nearly as clear for developing commercial space activities. In fact, two of the treaties governing space activities, the 1967 United Nations’ Treaty on Principles Governing the Activities of States in the Exploration and Use of Outer Space Including the Moon and Other Celestial Bodies, (commonly referred to as the Outer Space Treaty or OST) and the 1979 United Nations’ Agreement Governing the Activities of States on the Moon and Other Celestial Bodies (commonly referred to as the Moon Agreement), give rise to multiple, and sometimes ambiguous, interpretations regarding the legality of commercial Moon and space activities, as will be shown. Given the importance of commercial activity to the success of the proposed project, it is important to identify and clarify any points of contention within these treaties, and propose alternate suggestions where the legislation is vague regarding the ownership of lunar land and the exploitation of lunar resources. This chapter will therefore briefly review the salient points of the Moon and Outer Space Treaties, identifying sources of contention within these documents and proposing alternatives that will enable commercial Moon activities. This chapter also discusses other legal considerations essential to the realization of the proposed project, including insurance, liability and export control regulations, focusing on US protocols as a case study, given that this country is the most capable of realizing the proposed storage and delivery framework. Finally, given that many legal issues arise because of underlying ethical issues, this chapter will conclude with a review of the pertinent ethical questions that surround the exploitation and alteration of the lunar environment, briefly addressing how these issues may be resolved.

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6.1Overview of the Existing Legal Framework: The Outer Space Treaty & Moon Agreement The 1967 OST was the first international space treaty, forming the cornerstone of international space law and providing a basis for subsequent international legal documents. This treaty is considered to be a platform on which a house of more precise international legal documents can be built. In analyzing the OST, one should understand that it is a treaty of principles and is thus subject to broad interpretation only. The OST primarily addresses national activities rather than private activities, simply because only states had space-faring abilities at the time of negotiations. Moreover, as the OST was drafted during the Cold War, the concern amongst the major space-faring nations, the Soviet Union and the United States, was that the first man on the Moon not be able to claim it in the name of his country. Yet beyond forbidding any one nation to lay claim to celestial territories, the OST also strives to promote both the peaceful use and freedom of access to space, which are reflected in Article IV and Paragraphs 1 and 2 of Article I, respectively (OST 1967). The former states that “the exploration and use of outer space, including the Moon and other celestial bodies, shall be carried out for the benefit and in the interests of all countries, irrespective of their degree of economic or scientific development, and shall be the province of all mankind.” In this sense, it could be said that a lunar gas station would serve the interests of all countries since it would reduce the cost of lunar ventures, making space easier to access for all. Additionally, Paragraph 2 declares that “outer space shall be free for exploration and use by all States without discrimination of any kind, on a basis of equality and in accordance with international law, and there shall be free access to all areas of celestial bodies,” which can lead to the understanding that it is possible to explore and use lunar resources. Moreover, the OST (1967) establishes guidelines on a subject’s sovereignty, jurisdiction and resource appropriation in space. There are three key Articles that address commercial Moon activities (White 2001): Article II declares that outer space, including the Moon and other celestial bodies, is not subject to national appropriation by claim of sovereignty, by means of use of occupation, or by any means. The non-appropriation princicple is discussed later in this chapter in greater detail. Article VIII establishes that a State Party on whose registry an object launched into outer space is carried shall retain jurisdiction and control over such object, and over any personnel thereof, while in outer space or on a celestial body. Article IX, along with other largely applied international laws, prohibits interference with the activities of another State Party, meaning that jurisdictional authority under the OST provides most of the protection traditionally associated with property rights, with a few relatively insignificant limitations, making it a "quasi-territorial" jurisdiction. 138

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From the discussion above, one can derive multiple conclusions on the subject of a state’s rights in space. Firstly, locations can be occupied by space objects on a first-come, first-served basis, given that the OST promotes the equality of access to space and does not forbid human presence on celestial bodies. Next, nations will retain jurisdiction over their space facilities and personnel in, irrespective of nationality, and that these nations have the right to conduct their activities without the harmful interference from other states Furthermore, even if commercial space enterprises seemed like a distant possibility in 1967, the notion was still addressed on a broad level, owing to conflicting ideologies. The Soviet Union, arguing from a socialist perspective, contended that space activities should be carried out solely and exclusively by states, while the capitalist Americans rejected this argument, instead proposing Article VI. This article declared that States Parties to the Treaty would bear responsibility for national activities in outer space, whether or not such activities were carried out by governmental or non-governmental agencies, thus acknowledging that commercial space activities might take place in the foreseeable future. Although these aspects of the OST suggest that commercial activities are permitted where they facilitate access to space and serve to benefit humanity, the Treaty remains ambiguous on the subject of commercial activities, an essential consideration for potential commercial activities such as lunar mining. For this reason, another treaty, the 1979 Moon Agreement was proposed to restrict the exploitation of the Moon's resources by any single nation. This treaty, however, was not signed by any of the space-faring nations, having received only eleven ratifications and five signatures in total. This treaty was thusly received mostly due to Article 11, Paragraph 1, which declares that the Moon and its natural resources are the “common heritage of mankind (Moon Agreement 1979).” This terminology was and remains problematic as there is no further explanation of the principle: it could be interpreted as either a common resource (meaning a resource is freely available for all members to use) or as common property (which is owned by all members, but still requires permission to use). In this sense, it is similar to the Deep Sea Bed Regime (Convention on the Law of the Sea 1982), which also refers to this principle. The Moon Agreement also attempted to address the subject of resource exploitation, expanding upon the principle of the freedom of scientific investigation addressed in Article I of the OST (1967). Article 6 of the 1979 Moon Agreement declares that the States Parties to the Treaty have the right to collect, remove and use lunar resources for scientific purposes, on the condition that these samples are made available to other interested states and the international scientific community for scientific investigation. However, this Article does not address the collection of resources for economic gain. In sum, the Moon Agreement (1979) failed to satisfy the international community’s interests with respect to ownership in private enterprise, hence contributing to its widespread rejection.

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6.2Legal Barriers: A Further Analysis of Non-Appropriation Principles As space activities continue to evolve, barriers and discrepancies on the subject of commercial activities within the OST (1967) and the Moon Agreement (1979) become increasingly apparent. In particular, property rights and use of resources for economic gain have risen to the forefront of pressing issues in international space law, with UN members beyond the space powers calling for discussion of these issues within the United Nations Committee for the Peaceful Uses of Outer Space. At the international level, most disputes regarding these issues concern the principles of non appropriation by any means, stated within Article II of the OST (1967), and the principle of common heritage of humankind principle established by the article 11 of the Moon Agreement (1979). Given the importance of property issues to the success of commercial enterprises, this section explores the Treaties and precedents that could pertain to commercial activities, first within the context of land ownership, and secondly within the context of resource ownership for economic gain.

6.2.1Non-Appropriation of Celestial Bodies Article II of the OST (1967) and Article 11 of the Moon Agreement (1979) declare that outer space, including the Moon and other celestial bodies, is not subject to national appropriation by any means, including claims of sovereignty, use or occupation. This essentially means that outer space is an international zone and that no country can claim property rights to space, the Moon or other celestial bodies (see ). As a point of reference, this framework is similar to the high seas regime (Convention on The Law of the Sea 1982), rather than the airspace regime, which is deemed to belong to the State over which the airspace lies, as per the Chicago Convention of 1944.

Figure 6-46: Basis of the Non-Appropriation Principle Sources of international law are depicted by different sizes to convey their level of importance (not to scale).

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This non-appropriation principle is generally accepted as having customary value, meaning that the principle is so universally upheld that is also applicable to non-signatories of a treaty. This means that even if a state wishes to withdraw from the OST (1967), the weight of international legal opinion indicates that the treaty's provisions will be upheld by international law. The transition of this principle into a generally upheld principle partially stems from the fact that no state has claimed or attempted to claim sovereignty on the Moon or any other celestial body. As such, while both the United States and the Soviet Union have planted flags on the Moon as symbolic gestures, neither nation has ever laid claim to any part of the Moon’s surface. Theoretically, it could be argued that states have not made territorial claims because they have yet to amass the economic and technological resources necessary to perfect a territorial claim under international law. However, an evolving commercial space sector has rendered the OST and the non-appropriation principle at least partially anachronistic: as the commercial satellite industry evolved, it became apparent that there was a need to appropriate property in geosynchronous orbit. Because the bulk of international opinion shifted to favor this opinion, it is now possible to claim property in geosynchronous orbit. In this scheme, both companies and states can claim a volume of space for their satellites, legally exclude others from this space, and of course can make a private profit from use of this space. The International Telecommunications Union (ITU) is the international body designating these property rights based upon a claim and a fee. Despite these arguments, it is argued here that the principle of nonappropriation is upheld as customary law, because both of the arguments introduced above are somewhat flawed. In the first instance, if one hypothesizes a case where a nation is indeed technologically and financially capable of launching an expedition to stake a claim on lunar land, the non-appropriation principle is still generally followed by enough nations that defying it would risk initiating international conflict. The argument is tautological: even if the non-appropriation principle does not have customary value, this belief is so widely held that any nation laying claim would risk international conflict, thus further solidifying this principle as part of international custom. In the second instance, even if precedent demonstrates that is possible to overturn international law by way of popular opinion, the current international sentiment regarding the appropriation of lunar property is not about to follow suit. If it did, the space powers would be essentially given free reign over lunar land claims while the remaining OST signatories received nothing in turn. The appropriation of satellite slots, adversely, benefited the majority of parties involved. From these arguments, it is clear that the non-appropriation principle is widely established and firmly upheld. However, it can be said the lack of acceptance of the Moon Agreement (1979) establishes a precedent that the prohibition of property rights on celestial bodies, along with are unacceptable to most nations. Hence, one can conclude that there is a need to replace the existing solution with an alternative that addresses the issue of lunar enterprise and property rights.

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6.2.2Appropriation of Economic Benefit

Lunar

Material

for

While the OST (1967) and Moon Agreement (1979) clearly do not support appropriation of land, the documents are not as clear on the subject of resource appropriation for economic gain. In the previous discussion on the Moon Agreement (1979), it was noted that lunar materials may be removed and subsequently collected for scientific benefit. However, based on precedent, there is a possibility that an element of economic exchange may be possible: after all, Article 6 does not explicitly forbid the use of materials for economic gain once it has already been claimed for scientific use. For instance, in December 1993, the Russian Federation sold a very small sample of the lunar material from the Luna probe 16 mission for $442 000 USD while a few milligrams of lunar dust were sold for $42 000 USD at an auction in California in the same year. This establishes a precedent for a government owning and selling a resource extracted from a celestial body in outer space. In compliance with the principle of nonappropriation of celestial territory, the seller did not claim the territory from which the sample was taken in either case, but they clearly claimed ownership of the material as per and derived economic benefit from the sale of this material. Given this precedent, it might be possible to remove lunar samples for economic gain if they are also used for scientific purposes.

6.3Enabling Private Enterprise Thus far, the discussion has identified inadequacies within the existing legal framework. This section concentrates on identifying solutions that will enable private enterprise and lunar resource exploitation, and will be followed by a logical progression for applying these solutions.

6.3.1Enabling Lunar Land Use: Economic Zones and Safety Zones

Exclusive

As has been noted, if a facility is established on the Moon, property rights become an issue, since it is not legal to appropriate land. However, because the OST (1967) and general principles of international law legally protect a state’s activities from interference by other parties, there are solutions for protecting commercial activities. In the discussion that follows, exclusive economic and accompanying safety zones are proposed as means for appropriating land. Upon the establishment of a lunar facility, one can introduce the concept of an exclusive economic zone, whereby only one party is able to use a specific area of lunar land for economic gain. In this case, all other states will have free access to the zone as long as they do not initiate commercial activities there. This enables commercial activity without violating the principle of non-appropriation. Of course, one must address how to allocate these economic zones. Based on the earlier analysis of the OST (1967), places on celestial bodies are allocated on a first-come, first-served basis, hence it makes sense to designate these exclusive economic zones in the same way.

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When introducing an exclusive economic zone, one can introduce a “safety zone” around the facility as agreed by international parties, to prevent potential damage and destruction to the facility. This would permit the proprietor of the facility to exert a measure of control over activities taking place within the vicinity of the facility (see ). Based on the precedent used for off-shore drilling platforms based on Earth and around the International Space Station, these safety zones can be extended to 500 meters.

Figure 6-47: Explanation of safety zones and exclusive economic zones. However, this proposal does not settle all the issues concerning land claims and interference with a state’s activities. For example, an outside party could initiate activities in an area outside of a state’s economic exclusive zone, even if is clear that the state planned to extend its activities to that area. Here, it is not clear if the principle of prohibition against harmful interference claimed by the article IX of OST (1967) would apply.

6.4The Progression Towards International Regime

a

New

In response to these impediments, one can either develop a new legal framework or attempt to amend the OST to permit the appropriation of property and commercial exploitation of lunar resources and to also establish an international lunar authority to oversee commercial lunar activities and lunar resource exploitation usage.

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6.4.1Legalizing the Appropriation of Land and Resources To enable commercial activity, the non-appropriation principle needs to be recalled and replaced with new regulations. However, to amend the OST (1967), a conventional amendment is required, meaning that the new text must be accepted by the majority of the signatories (IDEST 2007). While possible in theory, this is nearly impossible in practice because space-faring signatories are greatly out-numbered by the remaining signatories of the OST (Id.). Hence, it makes sense that these non-space-faring members would not be especially interested in enabling commercial activities on the Moon, as this would only widen the economic gap between the OST (Id.) members. As a potential solution, the space-faring nations wishing to undertake commercial lunar enterprises could pay taxes to OST (Id.) members demonstrating a clear plan to access space, but lacking the finances and resources to do so. This proposal would facilitate the acceptance of any conventional amendments, while enabling commercial activities on the Moon. However, once appropriation is allowed, lunar resources should not be left unregulated, given that disputes could lead to armed conflicts and militarization of outer space. Hence, a regulatory body is necessary to regulate exploratory and commercial activities on the Moon. Two possibilities are explored here. The first solution consists of operating within the Moon Agreement (1979) to create a new lunar authority to oversee lunar resource use, while the second consists of defining regulations within a new legal framework altogether. The Moon Agreement (1979) prohibits appropriation of lunar land and resources for economic gain as long as no regulatory body exists to regulate this sector. Thus, in order to exploit lunar resources, members must abide by Article 11, Paragraph 5, which declares that States Parties to this Treaty undertake to establish an international regime, including appropriate procedures and regulations to govern the exploitation of lunar resources of the moon. The first step in implementing this solution would be the signature and ratification of the Moon Agreement by space-faring nations, after which point an international regime to govern exploitation of Moon natural resources could be established (IDEST 2007). As previously noted, once the non-appropriation and common heritage of mankind issues were addressed, the Moon Agreement would likely be more widely accepted. The new body would be responsible for the following activities (Sattler 2004): •

Developing guidelines for lunar exploration

•

Issuing licenses and permits for space activities on the Moon

•

Overseeing construction and mining operations on the Moon and other celestial bodies

•

Coordinating habitation and the placement of structures on the Moon

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•

Managing accident liability and establishing of a specialized space tribunal to resolve accident liability and legal claims related to lunar activities

To aid in the development of the new governing body, one could refer to the establishment of the International Seabed Authority (ISA), which was established under the 1982 United Nations Convention on the Law of the Sea, and implemented in 1994. ISA, together with the ITU, could serve as a guideline for the operations of the new lunar activities’ regulatory body. Both of these organizations successfully address similar regimes: ISA regulates the activities on the seabed and ocean floor and controls the exploitation of maritime resources, while the ITU oversees the distribution of geosynchronous orbit allocations. In some sense, the new lunar authority would take upon the duties of both, regulating lunar activities, overseeing the allocation of lunar land use and controlling the exploitation of lunar resources. Alternatively, this authority could be established outside of the frame of the Moon Agreement, under a new international treaty governing the exploitation of Moon resources to solve legal questions that cannot be answered by the existing legal documents. Regardless of the method, such an authority is crucial if commercial lunar activities are to be realized. However, the development and negotiations process is expected to extend for quite some time, hence justifying the adoption of temporary national legislation in the interim.

6.4.2Temporary National Legislation Any new national legislation should authorize the exploitation of lunar resources until the adoption of the relevant international regime (IDEST 2007). This proposal is based on precedent. During the negotiations of the Convention on the Law of the Sea, the United States temporarily adopted the 1980 Deep Sea Bed Hard Mineral Resources Act to be able to exploit deep seabed resources. France followed suit in 1981, adopting the Act on Deep Seabed Mineral Resources. While these immediate laws take effect, States shall begin the negotiation of a new international regime. National legislation shall then be amended as needed when the new international regime enters into force as national law will then be in compliance with national law. The need to adopt temporary national legislation is proposed because the more concrete solutions take time to develop. Thus, it makes sense to apply temporary solutions while developing longer term solutions, rather than be limited in executable activities until more steadfast solutions are implemented.

6.5Additional Legal Considerations: Insurance, Responsibility & Liability Assuming, for the moment, that land and resource appropriation becomes a non-issue, there are still numerous legal considerations a party must consider in embarking upon the Full Moon proposal. Although most issues are presented in an international light, special attention is given to US policy and regulations given that NASA is the most likely and able to realize the Full Moon proposal. 145

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6.5.1International Guidelines As per Article VI of the OST (1967), no state permits private activities in space without insurance, which serves to place the insured back into the financial position it would have been in had a loss had not occurred. This is because this Article declares that states are responsible for authorizing and supervising the space activities conducted by non-governmental entities. Meanwhile, Article VII of OST declares: Each State Party to the Treaty that launches or procures the launching of an object into outer space, including the Moon and other celestial bodies, and each State Party from whose territory or facility an object is launched, is internationally liable for damages to other State Parties to the Treaty or to its natural or juridical persons by such object or its component parts on the Earth, in air space or in outer space, including the Moon and other celestial bodies. (OST 1967) Furthermore, the 1972 Liability Convention establishes rules for the resolution of personal injury and property damage issues at the international level, the results of which are summarized in Table 6-46. Table 6-46: Summary of liability under the Liability Convention (Adelta Legal Space Law 2007) Article

Type of Damage

Type of Liability

Article II

Damage on Earth or to aircraft in flight Damage to another space object not on the surface of the Earth Damage to another space object in outer space which subsequently causes damage on Earth or to aircraft in flight More than one launching State

Absolute Liability

Article III Article IV

Article V

Fault Liability Joint and Several Absolute Liability Joint and Several Liability

Lastly, with regards to the topic of legal responsibility, Article 1 of the 1975 Registration Convention establishes that a launching state has to maintain a registry of all objects launched space and create national legislation establishing a registration-permit system for private entities conducting space activities. This is to help establish fault in case of damage or destruction in case.

6.5.2US Guidelines Federal Aviation Regulations (FAR) govern liability and responsibility in the United States. With private commercial activities limited to launch and satellites operations, there is a lack of analogue cases for the establishment of insurance clauses for a lunar gas station. However, launch operations must still be addressed and are discussed here. According to FAR, there is a three-tiered system of liability risk sharing between the American government and a launch operator. The government covers a statutory ceiling of USD 1.5 B, beyond which the launch operator is personally liable up to an amount of USD 500 M,

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beyond which exists a Maximum Probable Loss (MPL) statutory ceiling, which is the maximum insurance coverage available in the world’s market. Since the launch operator is responsible for both public and property safety, the operator has to demonstrate the financial ability to compensate Maximum Probable Loss (MPL) from claims by a third party for damages claimable under the Liability Convention. As the Full Moon project would most likely be initiated by the US government, it could be assumed that the government itself would be the launch operator until a private company took control and maintenance of the lunar gas station. However, after a private company took ownership of the gas station, it would be very difficult to insure the maintenance of the station at this point, as no underwriter would want to take on the risk more than 10 years after the start of operations, due to the increased risk of technology failure due to depreciation over time. At this point, the United States government would be responsible for claims exceeding the MPL amount but only up to the statutory ceiling of USD 1.5 B mentioned.

6.5.3US Export Control Regulations Given that the United States is heavily in favor of international cooperation for lunar missions, Export Control Regulations become an extremely relevant topic, as these guidelines address the unauthorized export of certain controlled items, information or software to foreign persons or entities in the U.S. and abroad. These guidelines exist to protect US national security and foreign policy interests. There three different Export Control Regulations are summarized here: Export Administration Regulations (EAR), which address “dual-use” items, information and software designed for commercial purposes but having military applications Office of Foreign Assets Control (OFAC), which establishes economic sanctions and regulations prohibiting trade with and/or the transfer of payments, property or anything of value to certain sanctioned or embargoed countries or “Specially Designated Nationals” (SDNs) of those countries. International Traffic in Arms Regulations (ITAR), which are enforced by the State Department’s Directorate of Defense Trade Controls (DDTC). ITAR addresses the export and import of defense-related articles and services specifically designed, developed or modified for military applications and listed in the United States Munitions List (USML). These regulations state that information and material pertaining to defense and military-related technologies may only be shared with American Persons unless they are excluded and exempt with a DDTC approval. Special attention is given to ITAR here, because of its importance in regulating the freedom of information sharing between American and non-American entities. The strictness of these regulations, and accompanying penalties, ranging from fines up to USD 1 M and ten years’ imprisonment, often result in difficulties in international partnerships.

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It is for this reason that the regulations have been relaxed in the case of international partnerships. In the case of the International Space Station, the export control framework states that NASA is obligated to deliver, disclose or transfer technology, data and commodities essential for meeting its obligations in the program. This model can serve as a reference for potential partnerships between NASA and international entities in enacting the Full Moon proposal.

6.6Full Moon & Ethical Concerns Ethical issues refer to those dilemmas that appeal to one’s sense of ‘right’ or ‘wrong.’ This concept is too subjective for most, so academics and institutions alike have tried to formalize the term into a more objective definition. The United Nations’ Commission on Ethics and Technology (COMEST), for example, defines ethics as ‘a critical examination consisting in moral reasoning (COMEST 1998),” whereas Arnould (2007) defines the term as “an evaluation of what is right or wrong according to philosophical interrogation concerning the consequences of an activity.” Space ethics refers to ethics in the context of human exploration of outer space, and the topic is particularly relevant as the Full Moon project touches on a medley of controversial issues, including the commercialization and exploitation of lunar resources and the need to preserve the lunar environment. Bearing that in mind, how does one decide whether or not to proceed with the proposal? A logical approach is to look at the issues associated with Full Moon and oxygen and hydrogen usage, values reflected in each course of action and then weight them against the standard values that are generally held in high regard by society (or the organization or organism in question) and evaluate which outcome is more reflective of these values. According the non-appropriation principle discussed earlier in this chapter, space and lunar resources are considered as belonging to all of humanity. As such, questions arise as to whether it is justifiable to lay claim to lunar resources. In the context of the Full Moon project, the options entail going ahead with the project and laying claim to lunar resources with the intent of using them for commercial purposes, or abiding by the nonappropriation principle set out in the OST. In rejecting the Full Moon proposal to avoid appropriating and commercializing lunar resources, one is making a statement of belief that all of humanity has equal claim to these resources. Alternatively, by going ahead with the Full Moon proposal, one is making the values statement that appropriating lunar materials to enable lunar exploration and habitation is of greater value than having equal access to these resources. Broadly, the discussion then evolves into a values judgment: what is more important to society: equality of access to resources or the spirit of exploration and scientific discovery? Problematically, these are not issues that are easily compared.

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Moreover, the ethics involved of the Full Moon proposal is not as simple as this. Other values come into play: the Full Moon proposal, for example, encourages international cooperation, fostering new networks, and on a broad level, promoting global harmony. At the same time, the project has the potential to harm, or at least permanently alter, the lunar environment. NASA (2000) itself said that "the footprints left by the astronauts in the Sea of Tranquility are more permanent than most solid structures on Earth. Barring a chance meteorite impact, these impressions in the lunar soil will probably last for millions of years.” It is therefore very important to consider the impact of a settlement (mining, vehicular traffic) on the lunar environment. The increase in space debris and the environmental consequences of the transfer of resources that do not exist on Earth are also to be taken into account before exploiting and mining lunar resources. Industrial operations could, for example, result in a significant atmosphere around the Moon which would degrade its natural state and interfere with others' scientific and industrial operations as well, thereby interfering with the initially stated value of the spirit of exploration and scientific discovery. The international scope of the exploitation and mining of space resources adds complexity to the problem, because of the range of values systems and cultural ethics that come into play. For example, a socialist might choose not to pursue the Full Moon project based on the infringement of equal access to resources. Alternatively, a capitalist might invest in the project because of the economic benefit involved. There is also a need to measure the risks involved with such a program: disaster and failure represent practical considerations that might cause one to reach a different conclusion to an ethical issue. Lastly, ethical issues might reveal undesirable options that themselves offer new considerations. For example, one of the main concerns with the commercialization of space resources is that private companies act for their personal and economical interests and not for the good of all mankind. However, these companies could also establish a settlement as part of their venture, and in doing, promote exploration and scientific return. In short, ethical implications are not easily resolved, especially not in the context of the Full Moon project, but at the very least highlight the values and issues at stake, and offer an added perspective to practical matters, for example the existing legal debate with respect to the nonappropriation principle and its continued enforcement.

6.7Conclusions From the legal discussion discussed in this chapter, several points are immediately clear. Firstly, while there is an existing legal framework governing state and private activities in space, it is not at all conducive to commercial enterprises wishing to lay claim to lunar land and lunar resources. This is relevant because the Full Moon project proposes to claim ownership of oxygen and hydrogen extracted from the lunar environment for use in its storage and delivery architecture. According

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to the non-appropriation and common heritage principles of the Outer Space Treaty and the Moon Agreement, this is not permissible. Despite this, the fact that the Moon Agreement received such little acceptance, combined with the precedent set by the ownership of property in geosynchronous orbit, indicate that the international community is becoming increasingly open to the idea of allowing celestial land and resources to be appropriated. Moreover, there are international principles that permit the Full Moon concept to operate legally without violating the non-appropriation principle, such as the concepts of exclusive economic and safety zones. However, these are but temporary solutions. Ultimately, the entire legal framework must be either amended or replaced with a new framework permitting the appropriation of land and resources, and establishing a regulatory body to oversee the ownership and exploitation of lunar and resources. While these solutions are being implemented, nations should adopt their own temporary national legislation, based on the precedent set by the Convention on the Deep Sea Bed. Meanwhile, nations should be aware of other international guidelines of direct bearing on the Full Moon proposal, including insurance and liability guidelines and regulations dictating the freedom of information sharing in a proposal that will most likely be based on international cooperation. This is especially important in the context of American Export Control Regulations, which are extremely strict, and not especially conducive to international partnerships. For this, a model similar to that established for the International Space Station should be followed to facilitate the flow of information between international partners. Lastly, while ethical issues pertaining to the Full Moon project are not easily resolved, they are valuable in highlighting the issues and values as stake, and in adding a new perspective to practical discussions such as legal matters.

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_____________________________________Chapter 7

7 Conclusions

With regards to space exploration, the Moon is The Next Big Thing. In recent years, the international space community has boldly and firmly declared that lunar colonization is a challenge and a priority, to be embarked upon in the spirit of exploration, enterprise and ambition. Hence, the Apollo-era question of “How can we get there?” has since evolved to “How do we stay there?”. As noted in Chapter 1, first steps toward ‘taming’ any new, uncharted territory lie in establishing an adequate life support infrastructure, followed by a transportation network to ensure mobility across the territory. However, it is all-too-often forgotten that supporting elements themselves require supporting elements: these supporting elements must themselves be meticulously assessed and planned for, so that they can be designed, developed and implemented in a way that meets the territory’s needs. Similarly, the first missions to establish a permanent human presence on the Moon, beginning with NASA in 2018, will require transportation and life support infrastructures of their Moon—which will in turn require supporting elements, namely oxygen and hydrogen for use as propellants in vehicles, and also for air and water in life support systems. While production and extraction methods for these resources have been extensively studied, there is a distinct dearth as to how to make these readily accessible to the user. Full Moon attempted to address this need by proposing a storage and delivery architecture for these elements, and evaluating its potential through a comprehensive technical, business and legal analysis, the results of which are summarized here. As shown, the difficulties posed by the lunar environment are numerous, and are summarized as follows: • • • • •

Extreme temperature gradients Long lunar night lasting 14 Earth days High radiation exposure Dust abrasion Rough terrain

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These challenges, in conjunction with the need to supply oxygen and hydrogen locally at the South Pole and globally to the equatorial region, mean that that any system architecture needs to be flexible based on demand scenarios. For example, if the optimistic scenario reaches fruition and the Moon becomes a Mars resupply station, demand for hydrogen and oxygen will increase, meaning that the storage tanks must be able to hold a greater volume. Taking all this into account, the Full Moon proposal offers a simple yet reliable storage and delivery architecture to meet these needs, based on a brief yet thorough technical assessment of available options. Briefly, after identifying potential options for each of the storage and delivery systems, the assessment consisted of qualitative, quantitative and weighted analyses based on criteria such as performance, reliability, cost and efficiency, a decision-making matrix and extensive consultation with experts in the field. The remaining candidates for each system were then integrated, using a ‘concurrent engineering’-like approach, into a final structure that was chosen for its simplicity of design, flexibility to adapt to different scenarios and compatibility with other system components. The final system architecture, designed for a basic production sitestorage-delivery-storage scenario is as follows: • • •

•

•

•

A large storage tank directly attached to the production facility in the Lunar South Pole region. The tank is located in an area of inside permanent shadow, protected by sunshields if necessary. A set of smaller tank trailers that are can be towed using a lunar rover. Passive thermal control supplemented by active cooling allow the tank to maintain cryogenic temperatures. The active cooling is powered by a fuel cell. Storing in a permanently shadowed crater reduces energy spent for thermal control. Two delivery options: o A ballistic vehicle for delivery to higher latitudes. The ballistic vehicle is based on the principles of rocketry and as such its fuel usage is high; to deliver 4 tonnes of payload, 30 tonnes of fuel is required. o A rover for delivery to local sites near the South Pole. The rover transports tank carts by towing them. An efficient interface between the storage tank and delivery vehicle, utilizing a no-vent propellant transfer scheme. The mechanisms required for no-vent transfer is located inside the tank module. Standard interface design that are made known to all parties wishing to travel to the Moon, such that problems with varying transfer schemes and nozzle design are minimized.

The transportation vehicles were chosen on based on their overall value to the customer. The ballistic vehicle was chosen as global transport option for delivery to equatorial and distant sites, because it is the fastest option, travelling from the South Pole to the equatorial region in the space of hours, and is based upon reliable, proven technology. Other alternatives, such as land-based delivery, would take 152

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weeks, and would involve considerably more risk given the unpredictability of the lunar terrain. Thus, although the system is more costly and consumes more propellant than land-based delivery, it is the most efficient system for the task required of it. The local transportation system consists of a lunar rover. The rover is preferable to fixed installations such as e.g. pipelines. Any fixed transportation infrastructure would be very massive and thus the launch costs make them unattractive. Lunar manufacturing technology enabling the production of e.g. pipes on the Moon could swing this balance in the farther future. The simplicity of this system is an advantage. It reduces design costs and involves few components, thereby improving reliability. To enable storage of LOX and LH2 for indefinite periods in the tanks an active cooling system is necessary. By including this in the storage unit further standardization, some of the technological complexity of an active refrigeration system is alleviated. This system is scalable, and can supplemental tank units, rovers and ballistic transport vehicles may be launched from Earth. Thus, if demand doubles, the system can adapt to cater for this additional demand. Importantly, the Full Moon solution is implementable on a time scale that reflects the roadmaps of the major space agencies: a semipermanent human presence is set to begin as early as 2023, while the first phase of Full Moon is planned for 2027—meaning that the solution will come at a time when the demand for oxygen and hydrogen is increasing, and the market becomes more conducive to the Full Moon proposal as a business concept. However, even if the market demand proves to be definite and growing, commercial success is not guaranteed. In fact, due to the high capital investment, uncertainty of the market and long-term return on investment, private ventures have a low-probability of succeeding due to the high level of risk involved. This is not to say that the Full Moon proposal should be shelved: on the contrary, Multi-public-private-partnerships, however, are deemed to be the best solution for success for the following reasons:

•

The responsibilities for the various aspects of the lunar infrastructure are broken up amongst multiple partners—thereby mitigating financial risk & liability.

•

Private entities can initiate commercial activities more easily under this structure, as they only pay operations costs and payback to government and government agencies, who pay for start-up costs.

•

Joint ventures with related industries such as Earth oil companies could form the basis for corporate partnerships allowing the private enterprise from the knowledge base of the

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established partner, and the latter from the image of space exploration. Yet while the existing technology levels and corporate framework are conducive to the Full Moon proposal, the present legal framework is extremely cumbersome to the project—and all lunar commercial activities—owing to unfavorable legislation regarding property rights and resource exploitation. Despite this, commercial activities are still feasible using existing legal principles such as safety and exclusive economic zones, which respectively protect a party’s facility and their right to exclusively conduct commercial activities within a specified area. This is, however, a temporary solution. A best solution would entail the creation of a new legal framework enabling commercial activities by rescinding the non-appropriation and common heritage principles and by further creating a regulatory body to oversee the use and exploitation of lunar resources. Since this is expected to be a lengthy process, temporary national legislation could be adopted to enable these commercial activities while existing international treaties are amended, or while new treaties are developed and approved together. Importantly, ethical considerations could shed light on these and other important discussions regarding the Full Moon process. Any entities that choose to take on the Full Moon proposal would also do well to consider the various legal documents considering liability, responsibility and freedom of exchange of information, all of which are important for the applicability of the Full Moon proposal. In short, the Full Moon concept offers a formidable challenge, which, if realized, would take humanity another step towards establishing a permanent presence on the Moon by facilitating accessibility to resources essential to lunar exploration and survival, namely oxygen and hydrogen. What is currently proposed as a simple storage and delivery architecture for oxygen and hydrogen could be developed and iterated, leading to solutions for newer challenges. As the market develops, future Full Moon iterations could address the following challenges: • • • • • •

Creating an architecture for resupply at Lagrange points and in Low Lunar Orbit for exploratory missions extending beyond the Moon Incorporating new technologies to offer better, more efficient services Incorporating new data and market analyses to update and/or extend the Full Moon concept beyond the present 2050 scenario Designing a method for route optimization should lunar demand increase to the point where multiple point-to-point deliveries are no longer efficient Adapting the Full Moon architecture for resupply at future targets such as Mars or the asteroids Adapting the Full Moon architecture for storage and delivery of other lunar resources

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The Full Moon proposal, in conclusion, is a next step in establishing a permanent human presence on the Moon – but it is also a stepping stone towards the future of human colonization beyond the Earth.

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_____________________________________Chapter 8

8 References

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1 INTRODUCTION.......................................................................................... ............20 1.1.1 The Full Moon project ................................................. .......................20 1.1.2 Definition of Project Scope................................................. .................21 1.1.3 Report Structure (Readers’ Guide)................................................ .......22 2 DRIVERS & CONSTRAINTS.................................................................. ................24 2.1 MARKET ANALYSIS FOR REFUELING.......................................................................... 24 2.1.1 Potential Customers......................................................................... ....24 2.1.2 Customer Location............................................................................ ...29 2.1.3 Demand Scenario General Framework..................... ..........................34 2.1.4 Timeline Selection.................................................................... ............35 2.1.5 Demand from Landings................................................................... .....36 2.1.6 Demand from Life Support............................................................ .......38 2.1.7 Demand from Outposts................................................. .......................38

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2.2 DEMAND GRAPHS...................................................................................................39 2.3 CONSTRAINT FROM COMPETITORS..............................................................................41 2.4 CONSTRAINTS OF LUNAR TOPOGRAPHY...................................................................... 42 2.5 CONSTRAINTS OF THE LUNAR ENVIRONMENT.............................................................. 43 2.6 CONSTRAINTS FROM RESOURCE AVAILABILITY.............................................................44 2.6.1 Resource Availability: Oxygen............................................. ................44 2.7 CONSTRAINTS OF PRODUCTION..................................................................................47 2.7.1 Production: Oxygen & Hydrogen from Water Ice................................50 2.7.2 Production: Hydrogen from Solar-Wind Implanted Volatiles...............51 2.8 FINDINGS: PRODUCTION SCENARIOS...........................................................................51 3 ARCHITECTURE ASSESSMENT.................................................................... .......53 3.1 CHALLENGES OF LUNAR ENVIRONMENT..................................................................... 53 3.2 SYSTEM DESIGN METHOD........................................................................................55 3.2.1 Qualitative Decision-Making.................................................... ...........56 3.2.2 Quantitative Decision-Making................................................. ............56 3.2.3 Futures: Supply Chain Modeling....................................... ..................57 3.3 ASSESSMENT OF STORAGE OPTIONS ON THE LUNAR SURFACE........................................57 3.3.1 Challenges of Stored Materials.................................................... ........58 3.3.2 LOX Properties and Considerations ................................. ..................59 3.3.3 LH2 Properties and Considerations................................................... ..59 3.3.4 Thermal Control....................................................... ...........................60 3.3.5 Materials............................................................................................ ..64 3.3.6 Cryogenic Storage Methods: Existing Technologies............................64 3.3.7 Liquid Oxygen .................................................................................. ...65 3.3.8 Cryogenic Storage Methods: Emerging Technologies......................... .65 3.4 ALTERNATIVE STORAGE POSSIBILITIES........................................................................66 3.4.1 Slush Hydrogen Properties and Considerations..................................66 3.4.2 Storage as Water...................................................................... ............67 3.5 STORAGE VESSELS..................................................................................................67 3.6 INTERFACING .........................................................................................................67 3.6.1 Transfer Guidelines............................................................................. .68 3.7 ENSURING HUMAN SAFETY AND TANK HEALTH.......................................................... 69 3.7.1 Assessment of delivery systems ................................... ........................69 3.8 TRANSPORTATION OPTIONS........................................................................................70 3.8.1 Mobile platforms............................................................................. .....70 3.8.2 Fixed platforms.......................................................................... ..........73 3.9 DELIVERY SYSTEM EVALUATION............................................................................... 75 3.9.1 Evaluation Criteria................................................................. .............75 3.10 QUALITATIVE DECISION RANKINGS FOR DELIVERY SYSTEMS....................................... 75 3.11 QUANTITATIVE DECISION RANKINGS FOR DELIVERY SYSTEMS.............................. .......78 3.12 FINDINGS: STORAGE AND DELIVERY CONCEPT RANKINGS...........................................83 4 SYSTEM ARCHITECTURE............................................................ ........................85 4.1 THE PROPOSED ARCHITECTURE.................................................................................88 4.1.1 The Storage Solution....................................................................... .....88 4.1.2 The Delivery Solutions................................................................ .........91 4.1.3 Local Delivery............................................................ .........................92 4.1.4 Equatorial Delivery...................................................... .......................93 4.2 INTERFACES............................................................................................................96 4.2.1 Interfaces with the M-LSAM............................................ ....................96 4.2.2 Interfaces with Local Servicing Rovers......................................... .......97 4.2.3 Communications.............................................................. ....................98 4.3 OPERATIONS...........................................................................................................99

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4.3.1 Storage Tank Rotation...................................................................... ....99 4.3.2 Servicing Methods............................................................................. .100 4.3.3 Local Service with Wheeled Rovers.............................. .....................101 4.3.4 Equatorial Service with the M-LSAM....................... .........................102 4.3.5 Emergency Services with the Modified LSAM............................ ........103 4.3.6 Maintenance.............................................................. ........................104 4.4 THE SYSTEM BLUEPRINTS......................................................................................104 4.4.1 Storage Implementation....................................................... ..............105 4.4.2 Local Delivery System Implementation................................... ...........106 4.5 TECHNICAL RISKS.................................................................................................108 4.5.1 Storage........................................................................................... ....108 4.5.2 Delivery..................................................................... ........................109 4.6 ADAPTATIONS FOR THE OPTIMISTIC SCENARIO........................................................... 109 4.6.1 Scaling Up the Architecture......................................... ......................109 4.6.2 New Services............................................................................. .........109 4.6.3 Recommendation: A Proposed Architecture .......................... ............109 5 THE BUSINESS ANALYSIS........................................................... ........................111 5.1 APPROACH AND OVERVIEW.....................................................................................111 5.1.1 Market Demand Overview..................................... ............................112 5.2 SUPPLY OVERVIEW................................................................................................114 5.2.1 Lunar Oxygen and Hydrogen Supply Service......................... ............114 5.2.2 Assumptions............................................................... ........................114 5.3 BUSINESS SOLUTIONS............................................................................................ 115 5.3.1 Public Cooperation Scenarios for Development Phase......................115 5.3.2 Private Joint Venture Operational Phase......................... ..................117 5.3.3 Business Solutions Summary................................................ ..............118 5.4 BUSINESS RISK ASSESSMENT..................................................................................118 5.5 COST BREAKDOWN ANALYSIS ................................................................................119 5.5.1 Cost Categories...................................................... ...........................120 5.5.2 Launch Cost............................................................................ ...........120 5.5.3 Capital Cost............................................................................ ...........121 5.5.4 Operation cost......................................................... ..........................122 5.5.5 Summary of Cost Estimation..................................................... .........123 5.6 FINANCIAL MODEL .............................................................................................. 124 5.6.1 Calculating Annual Costs and Revenues....................................... .....125 5.6.2 Pricing Approach...................................................................... .........126 5.6.3 Selection Criteria...................................................................... .........126 5.6.4 Taxes, Insurance and Discount Factor.......................................... .....127 5.6.5 Cash Flow Analysis and Profitability Analysis..................................128 5.7 SENSITIVITY ANALYSIS...........................................................................................132 5.8 PROMOTION..........................................................................................................134 5.9 RECOMMENDATIONS...............................................................................................135 6 LEGAL & ETHICAL ISSUES............................................................ ....................137 6.1 OVERVIEW OF THE EXISTING LEGAL FRAMEWORK: THE OUTER SPACE TREATY & MOON AGREEMENT.....................................................................................................138 6.2 LEGAL BARRIERS: A FURTHER ANALYSIS OF NON-APPROPRIATION PRINCIPLES..............140 6.2.1 Non-Appropriation of Celestial Bodies............................... ...............140 6.2.2 Appropriation of Lunar Material for Economic Benefit.....................142 6.3 ENABLING PRIVATE ENTERPRISE..............................................................................142 6.3.1 Enabling Lunar Land Use: Exclusive Economic Zones and Safety Zones...................................................................... .................142 6.4 THE PROGRESSION TOWARDS A NEW INTERNATIONAL REGIME ............................ .......143

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6.4.1 Legalizing the Appropriation of Land and Resources........................144 6.4.2 Temporary National Legislation............................... .........................145 6.5 ADDITIONAL LEGAL CONSIDERATIONS: INSURANCE, RESPONSIBILITY & LIABILITY.........145 6.5.1 International Guidelines.................................................. ..................146 6.5.2 US Guidelines.................................................................... ................146 6.5.3 US Export Control Regulations....................................................... ...147 6.6 FULL MOON & ETHICAL CONCERNS....................................................................... 148 6.7 CONCLUSIONS.......................................................................................................149 7 CONCLUSIONS ................................................................................................... ...151 8 REFERENCES .......................................................................................... ..............156 A. TANK MATERIAL SELECTION....................................................................... ..165 8.1 EXTRA VEHICULAR ACTIVITIES...............................................................................165 8.2 LSAM ASCENT STAGE MASS BREAK-DOWN...............................................................166 8.3 EXTRA VEHICULAR ACTIVITIES’ CONSUMPTION............................................................166 8.4 CALCULATIONS FOR 3, 6, 9 AND 12 DAYS-MISSION.................................................... 167 8.5 WATER CONSUMPTION:..........................................................................................168 B. LUNAR HYDROGEN....................................................................................... .....169 8.6 RATIONALE FOR SOUTH POLE BASE..........................................................................170 8.7 HYDROGEN EXTRACTION........................................................................................170 C. TANK MATERIAL SELECTION....................................................................... ..172 D. SYSTEM SELECTION DETAILS...................................................................... ..174 8.8 QUALITATIVE DECISION TOOL CALCULATIONS...........................................................174 8.9 QUANTITATIVE DECISION TOOL CALCULATIONS.........................................................175 8.10 SYSTEM OPTIMIZATION METHODS......................................................................... 175 8.11 SUPPLY CHAIN MODELING WITH SPACENET............................................................176 E. DELIVERY SYSTEM CALCULATIONS................................................... .........179 8.12 LSAM MODIFICATIONS FOR SURFACE DELIVERY....................................................179 8.13 M-LSAM CALCULATIONS FOR MID-LATITUDE DELIVERIES.....................................179 8.14 WHEELED ROVER CALCULATIONS......................................................................... 180

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____________________________________Appendix B

A.Tank Material Selection

8.1Extra Vehicular Activities 1.1 Weekly timetable for astronauts on the outpost

Figure Error! No text of specified style in document.-48 EVA and IVA activities considered for outpost GOX demand (NASA 2005)

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8.2LSAM ascent stage mass break-down

Figure Error! No text of specified style in document.-49 LSAM ascent stage mass break-down (NASA 2005)

Mass break down of the LOX/LCH4 propelled LSAM, and ISP for LOX/LH4 engine are used to calculate the LOX/LH2 needed. Oxidizer/Fuel ratio for LOX/LH2 is considered to be 8:1. Engines with even higher ratios of up to 10:1 are desirable, as the amount of LH2 needed is reduced. The benefit of high ratio is twofold: 1. Smaller tanks for LH2 which has very low density thus high volume and structural mass. 2. In the case there is no ice in the lunar South Pole, lander ascent LH2 will be carried from Earth, and only LOX will be refueled on the lunar surface. The less LH2 needed, the greater the benefit in launch mass on Earth.

8.3Extra vehicular activities’ consumption The following values are based on an average metabolic rate of 136.7 W/person and with a respiration quotient of 0.87 (which correspond to the molar ratio of CO2 generated to 02 consumed) and an average build: •

Oxygen consumed = 0,84kg/(person-day)

During an EVA, the activity level is getting greater so the value is changing to: •

Oxygen consumed = 0,96kg/(person-day)

1) Without rovers o

Distance: 1-3km

o

Without recharge 166

Introduction

  o

Length: 4-hour uninterrupted operation Oxygen consumed = 0,16kg/person

With recharge  Length: 6-8 hour EVA  Oxygen consumed = 0,32kg/person

2) With single rover o Distance: 10-15km (walk-back) 3) With two rovers (the second rover is provided to the single rover, so the walk-back requirement can be avoided) o Distance: 20-30km max o Length: 8 hours max (=crew member-suited physiological guidelines) o Oxygen consumed = 0,32kg/person According to NASA and Lockheed Martin, 4 persons will be there at the same time. Then, in high activity metabolic load, the oxygen consumption for 4 persons will be 4*0.96=3.84kg/day

8.4Calculations for 3, 6, 9 and 12 daysmission About the rovers: •

Maximum rover’s speed is about 20 km/hour

•

Unpressurized rovers need to return to the base at the end of each day

•

Pressurized rovers allow longer mission (3, 6, 9 or 12 days mission)

The following calculations consider that: •

Density of oxygen: 1.429g/L=1.429kg/m3

•

Density of Liquid oxygen: 1.141g/L=1141kg/m3

1) Oxygen consumed for a 3 days mission with 4 persons: 4*0.96*3= 11.52kg Volume of Oxygen=8.06m3 Volume of Liquid Oxygen=0.01m3=10L 2) Oxygen consumed for a 6 days mission with 4 persons: 4*0.96*6= 23.04kg Volume of Oxygen=16.1m3 Volume of Liquid Oxygen=0.02m3=20L 3)

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Oxygen consumed for a 9 days mission with 4 persons: 4*0.96*9= 34.56kg Volume of Oxygen=24.28m3 Volume of Liquid Oxygen=0.03m3=30L 4) Oxygen consumed for a 12 days mission with 4 persons: 4*0.96*12= 46.08kg Volume of Oxygen=32.24m3 Volume of Liquid Oxygen=0.04m3=40L According to NASA’s exploration architecture study, surface missions lasting four days will have EVAs each day. The first (landing) and last (takeoff) days will likely have shorter duration EVAs of four to six hours, while the middle two days will each have a full 6- to 8-hour EVA period. Longer-duration sortie missions of up to seven days would likely require at least one day of rest without planned EVAs. In the last case, the oxygen consumed for the four persons will be: 4*0.96*6 + 4*0.84*1=26.4kg The volume will be: 26.4/1.429=18m3

8.5Water Consumption: 3.9kg/(person-day) (included drink, water in food and metabolized water) 3 days mission for 4 persons: 46.8L 6 days mission for 4 persons: 93.6L 9 days mission for 4 persons: 140.4L 12 days mission for 4 persons: 187.2L

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____________________________________Appendix B

B.Lunar Hydrogen

Despite the hydrogen controversy, this report assumes existing data from Lunar Prospector and Clementine and lays emphasis on how hydrogen can be obtained from these resources. An unusual discovery from Clementine was the non-uniformity of lunar tectonics and mineralogy; hence analysis from Apollo sample cannot infer the entire Moon is dry-bone (2004a). Mean concentration of solar wind implanted hydrogen is as low as 18 ppm (Maurice et al. 2003, Lawrence et al. 2007). Because surface regolith grain exposure time to solar fluence or meteorites/comets impacts far exceeds time needed to saturate a grain hence additional impacts or solar wind does not alter hydrogen concentration irrespective of latitude or longitude. (Maurice et al. 2003)

Figure Error! No text of specified style in document.-50 : Distribution of Hydrogen within the first 2m of lunar regolith as seen by Source: (Maurice et al. 2003), (2004a)

The graph shows a mean solar wind implantation of 50±20 ppm of Hydrogen with concentrations above mean corresponding to small dark craters while below mean concentrations correspond to points that have lost their hydrogen due to heat generated during impact and ultra-violet radiation from sun that dissociate water molecules (Maurice et al. 2003). Permanently cold, dark shadows near the south poles show a concentration of 190 ppm of hydrogen (equivalent to 1.5 ± 0.8% of water) Hydrogen concentration increases towards the lunar poles and significant concentrations are associated with large impacts craters with shaded bottoms. Since the South Pole has the largest, coldest and

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oldest craters, it is believed to have significant deposits that can be economically extracted. (2004b) Supposing all hydrogen in the lunar poles is trapped as water ice, Feldman and colleagues estimate total volume could be two billion tons of water in the upper three feet. A rough estimate of water ice deposits that can be economically extracted exist in the bottoms of very cold, permanently dark shaded areas of craters (-233 °C) near the south pole average to 1.5%. The best estimate for the total shaded area would give 200 tons of water equivalents. (Javier and Michael 2006) Because of the low resolution of Lunar Prospector and the fact that Clementine was not specifically deigned to detect water ice while water may exist as ice or crystal in lunar soil, the possibility of existing small areas with even greater concentrations cannot be ruled out. (2004b, Maurice et al. 2003) The controversy about lunar hydrogen centers on many unanswered questions some of which include the mere existence of water deposits that can be accessed, its form, distributed concentrations in regolith, concentrated deposits, and prospecting technique.

8.6Rationale for South Pole base Detail knowledge of lunar polar resources is essential for lunar exploration programs; therefore a combined study of water ice near the South Pole would have a greater effect for both exploration and scientific applications. (Lawrence et al. 2007) other merits include availability of clean power from near permanent sunlight, habitable zone without 14 day lunar nights saving billions of dollars worth spending on nuclear reactor to support life at the equator and midlatitudes regions, mild and favorable temperatures of about -50° ± 10° C and possibility of having water near the south poles would be great economic and exploration advantage.

8.7Hydrogen Extraction. Hypothetically, hydrogen can be extracted from equatorial regolith by heating the soil up to 700°C with 90% of the implanted hydrogen driven off making collection difficult. On the other hand, polar water ice requires two orders of energy less at 100° C using the base’s solar source to evaporate melted ice which is easy to collect. Because hydrogen concentration is small about 50 ppm, extraction is not economically worthwhile but it turns out that the straight forward method for extracting hydrogen can also be used to extract carbon present in about 110 ppm. The method is still useful to extract oxygen by reducing iron oxide in the regolith using either hydrogen or carbon yielding oxygen and methane as products usable in methane/oxygen engine and more efficient than hydrogen. (Ruiz et al. 2004) An optimal performance of the extraction system is achieved by mining a high soil maturity site with soil grains less than 20 um and in the upper 10cm layer soil with possibility of getting hydrogen concentration of about 100ppm. (Team Project Final Report 2006) A propellant production plant size is determined by ice concentration, mining equipment and performance in that order. Javier and Michael

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(2006) suggest an ISRU production plant with a capacity of 900 metric tons of water per year from which about half of the water could be electrolyzed to yield hydrogen and oxygen. Assuming south pole based-plant, power system becomes a trivial issue due to the high performance of the solar system. Though wireless power transmission from solar energy and nuclear energy plants is suggested as options for providing energy in the permanently dark shadows, however more research into problems associated with astronauts or robots working in extremely dark and cold temperatures is still required. (Javier and Michael 2006) Available space, continuous sunlight and no atmosphere on the moon makes solar power generation more efficient so that higher temperatures can be attained at a shorter time with the capability of scaling up as demand increases. (Houdashelt et al. 1989) Many extraction methods exist but the least complex requiring less equipment though higher energy compared to De Beni Carbon-Iron Process, is electrolysis and it is assumed in this report to be best for the production of hydrogen and hydrogen on the basis of complexity. Chemically, two molecules of water yields tow molecules of gaseous hydrogen and a more of gaseous oxygen. Judging from the 249 KJ per mole as the dissociation free energy, the calculated minimum specific power of the process is 3.84 KWhrs/kgH2O(Team Project Final Report 2006).

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____________________________________Appendix C

C.Tank Material Selection

The following tables were used as an aid to choosing storage tank materials. Table A-47: Qualitative Tensile Strength Table Al–Li Alloy Al-Li Alloy Composi te Material

Composite material

Teflon

Stainles s Steel

Result

Rankin g

L

U

L

1

#1

U

L

1

#1

L

0

#2

0

0

Teflon Stainless Steel

Table 48: Qualitative Thermal Conductivity Table Al–Li Alloy Al-Li Alloy Composi te Material

Composi te Material

Teflon

Stainles s Steel

Result

Rankin g

U

U

U

3

#1

U

L

1

#2

L

0

#3

0

0

Teflon Stainles s Steel

Table 49: Qualitative Density Table Al–Li Alloy

Composit e Material

Teflon

Stainles s Steel

Result

Rankin g

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Al-Li Alloy Composi te Material Teflon Stainles s Steel

U

U

L

2

#1

L

L

0

#2

L

0

#3

0

0

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____________________________________Appendix D

D.System Selection Details

This appendix provides the details on the calculations used for selecting the delivery system. In the last sections, it expands on methods for modeling the complete system using academic software called “SpaceNet.” Effort, performance and robustness are considered key criteria.

8.8Qualitative Decision Tool Calculations The Qualitative Decision Tool was used to evaluate each system pairwise with another by criteria. A set of tables was created (one per criteria) to compare the systems by the defined “best.” The relation between the inputs and results is shown in Figure B-Error! No text of specified style in document.-51. The “ranking column tallies the number of ‘U’s across the row with the system and the number of ‘L’s in the column under the system.

Figure B-Error! No text of specified style in document.-51: Qualitative Decision Tool input flow While this method allowed for contradictory selections, this did not occur often and was made apparent by systems with identical tallies. The results were checked to counter major errors. This method was iterated to ensure that the concepts for the criteria and particularly “best” within that criteria were clearly the same as that intended by the constraints.

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8.9Quantitative Decision Tool Calculations The Quantitative Decision Tool relied on the subjective choice of appropriate quantities and formulae to ensure that “best” was represented consistently by a higher number. The formula for each criteria is given in Table B-50. Table B-50: Quantitative Decision Tool formulae Parameter Technical Readiness Operational Efficiency

Formula = 1 / time [years] to develop = 1 / specific power = mass [kg] / power [kW] Scalability = 1 / (marginal cost to double output, as % original cost) Adaptability = # of alternate applications (listed) Ease of Implementation = 1 / mass [kg] Freedom of Location = % coverage of lunar surface = area (range) / lunar surface area Delivery Time = capacity [kg] / speed [km/h] Reliability = maintenance cost per annum for terrestrial analog [USD] Safety = # of ‘yes’ answers to 5 key questions Resilience to Lunar = sum of ratings [1 to 3] for each of Environment expansion tolerance, dustimpermeability and terrain capability Each of the calculations was normalized and weighted. In the end, the formulae were simple and the calculations were fast. The weights for each faculty member or student were entered in one sheet, averaged among the faculty and students separately and the standard deviations calculated using built-in MS Excel formulae. Secondary weights were used in the concurrent engineering session. The scenario-specific results were the weighted results multiplied by the secondary weights.

8.10System Optimization Methods Some insight into multi-criteria decision making and system optimization was provided by John Blake, President of the Canadian Operational Research Society, Robert Shishko, Principal System Engineer and Economist at NASA Jet Propulsion Laboratory and Ozgur Gurtuna, President of Turquoise Technologies Solutions. This is a synopsis of their advised approach to this problem in more generalized scenarios. Blake suggested that a simple linear node network may be implemented in Excel using the Solver package (e-mail 2007-APR-14). Shishko noted that the MIT SpaceNet modeler (http://spacelogistics.mit.edu/) is designed for similar problems including a lunar surface-specific network (teleconference 2007-APR-23). The hazard with these methods is that they are only at a heuristic stage and have not been verified with real world scenarios. These experts agree that further work is needed in this field for space applications to aid with better program decisions.

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To optimize the cryogen distribution system, the problem can be broken down in stages: comparing individual technological delivery solutions; comparing different locations and number of storage sites; describing the different architectures as a network model; running an algorithm on the variations in the model and; weighting the results based on the importance to the consumers and to the business case.

The Linear Network Model for the Complete Cryogen System Computer modeling can be used to make decisions that consider combinations of systems and time. The model represents the supply and demand as nodes in a network linked by transportation elements. Each transportation element can be characterized by a utility function (such as or including cost). The problem for lunar cryogen storage and delivery system optimization is called a “transportation problem,” more particularly a “trans-shipment hub problem” because supplies can be held in storage nodes (John Blake e-mail 2007-APR-14). A sample iteration of a time-expanded network is shown below.

Figure B-Error! No text of specified style in document.-52: Supply flow through time-expanded linear network model From this, the optimum number and location of storage sites could be deduced. The transport systems with the best overall utility solutions can also be selected.

8.11Supply Chain Modeling with SpaceNet Supply chain modeling could solve the lunar cryogen distribution problem more realistically. This section describes a method to lay out the problem in the modeling package SpaceNet, but the method can be modified to suit any similar programs.

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Figure B-Error! No text of specified style in document.-53: SpaceNet model (http://spacelogistics.mit.edu) The model is a description of the worldview for the lunar cryogen system. The model includes a set of nodes representing predicted production sites and fixed and mobile consumers on a spherical surface and links (transportation elements such as the delivery systems described in the Assessment chapter. Table B-51: Elements for modeling a lunar cryogen delivery system Model Components • terrain difficulty map • supply elements (LH2, LOX) with transfer and transport losses • production node map with rates • fixed and mobile consumer node map with demand rates & priority • consumer revisit frequency

Variable Inputs • number of storage nodes • storage node capacity • transport capacity • transport speed & range • transport terrain capability • transport system autonomy • transport complexity • transport development lead time

Output • optimal location of storage node(s), • optimal delivery system design (including hybrids) across phases or different market scenarios • maximized system life cycle utility (or minimized cost)

The following assumptions can be made to simplify the problem: • There is a great circle route minimum path between any 2 surface

nodes, with the addition that the routes are weighted by a percentage of the distance for obstacle avoidance and terrain profile, depending on vehicle type.

• Consumers are defined as: small volumes delivered to robotic

rovers across regions; moderate high priority volumes delivered to a habitat and; high volumes delivered to launch vehicles.

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wheeled rover). • Demand for the full scale base, human rovers and launch vehicles

could exceed local (lunar) supply in the far term.

The “utility function” for each delivery system element should reflect these assumptions. Each element will have a “utility” defined by costs and/or performance characteristics. The engine will iterate the algorithm over variations of the model until it finds the optimum “utility” for the whole system. This can be minimum life cycle cost or maximum performance, for example. For the lunar cryogen system, both consumer and business case needs must be considered. The “utility” function must combine these needs by either translating all the requirements into costs or another consistent metric. Blake recommended first optimizing site locations, delivery system to meet consumer requirements, optimize the utility function for each system type node, 3 nodes, 10 nodes, say). The systems should at different phases in the predicted market.

then scaling each then iterating to (repeat for single also be compared

For node optimization, for example, LogicNet (http://www.logictools.com), a commercial supply chain modeling software, uses a mass balance equation at each node, i.e. outflow - inflow <= demand. Other algorithms are possible.

Results The benefits of modeling a complete system in this detail can be used for scaling the system to different predicted markets, choosing compatible system elements and visualizing delivery schedules. The hazard is reliance on the results before they have been verified with results from real world cases. This is difficult for space projects, but with care lessons learned from complex projects like the International Space Station could provide a baseline for methodically planning largescale space project.

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____________________________________Appendix E

E.Delivery System Calculations

The following material provides more detail on the delivery system selections.

8.12LSAM Delivery

Modifications

for

Surface

In order to use the NASA LSAM for deliveries from the lunar south pole to the equator, the following modifications are recommended: Table C-52: LSAM modifications

Range

Objective or Requirement 3 000 km

Payload Capacity

4t

Stage Type

Reusable

Fuel Transfer System

Deliver fuel to customer without changing or removing tanks Maneuver within 100 m of target

Surface Mobility

Capabilities of Present LSAM Round trip low lunar orbit to lunar surface 500 kg

reusable combined ascender/lander No external fueling/refueling system No surface local mobility capability

Recommended Modification Remove ascender, use main propulsion LOX/LH2 tanks for customer fuel Add tele-operated hose arm

Attach the ATHLETE wheeledwalker rover

8.13M-LSAM Calculations for Mid-Latitude Deliveries “Delta-V” Calculation The total “Delta-V” required for a ballistic vehicle traveling between two points on the lunar surface is given by

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where μ = 4.90279882 × 1012 m3/s rm = radius of the Moon = 1 735 km β = top angle (in radians) of the triangle connecting departure point, arrival point and centre of the Moon. This equation takes into account the initial burn required to take off and the final burn required to land in a controlled fashion (ISU 2000). For the lunar pole to equator range of 2 500 km, β = 1.41 radians. Based on this, ΔVideal = 3 121 m/s. Without a controlled landing, the impact velocity is half of ΔVideal.

Propellant Calculation  − ∆V    I g sp o  

The final mass ratio, µ final = exp

where Isp = the specific impulse of the RL-10 engine = 462 s go = 9.8 m/s2 This gives a final mass ratio of 0.48 The final mass of the vehicle, m f = mi × µ final , i.e. mi =

mf

µ final

.

The dry mass (i.e. final returned mass) of the stripped-down LSAM (no ascender) with the ATHLETE rover on board is mf2 = 6 210 kg. From this, the initial mass at the delivery location is, so mi2 = 12 940 kg. Adding the 4 t payload, the final mass for the delivery journey is then mf2 = 16 940 kg = mf1, which is the final mass for the delivery leg. Dividing again by the mass ratio, the mi1 = 35 290 kg, the initial wet mass for the vehicle with payload. Subtracting the initial vehicle dry mass, the total propellant needed for this trip is 29 080 kg. Since the maximum propellant that can be carried by the LSAM is 30 319 kg (NASA presentation 2007), this is within the limits; however, the vehicle structure should be optimized to allow a safety margin.

8.14Wheeled Rover Calculations Mass Estimation The estimated mass of the rover, m rover = c ⋅ m payload where c = 0.65 (Bufkin et al. 1988) mpayload = 4 000 kg + 5 000 kg (tank mass) = 9 000 kg(approximately) 180

Introduction

From this, the mass is 5 850 kg, and the total loaded mass is m rover = 5 850 + 9 000 = 14 850 kg.

Power Estimation The estimated power required for the rover, Ptotal = a ⋅ m rover ⋅ d where a = 0.1412 (Bufkin et al. 1988) mrover = total mass of the rover (calculated above) = 14 850 kg d = total distance traveled = 2 x range (2 km) + safety margin (1 km) = 5 km Based on this, total power requirement is 10.5 kWh.

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