Fast Transit: mars & beyond

final Report

Space Studies Program 2019 Team Project Final Report

Fast Transit: mars & beyond

final Report

Internationali l Space Universityi i Space Studies Program 2019

© International Space University. All Rights Reserved.

i International Space University Fast Transit: Mars & Beyond

Cover images of Mars, Earth, and Moon courtesy of NASA.

Spacecraft render designed and produced using CAD.

While all care has been taken in the preparation of this report, ISU does not take any responsibility for the accuracy of its content.

The 2019 Space Studies Program of the International Space University was hosted by the International Space University, Strasbourg, France.

Electronic copies of the Final Report and the Executive Summary can be downloaded from the

ISU Library website at http://isulibrary.isunet.edu/

International Space University Strasbourg Central Campus Parc d’Innovation 1 rue Jean-Dominique Cassini 67400 Illkirch-Graffenstaden

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ii Space Studies Program 2019

ACKNOWLEDGEMENTS

Our Team Project (TP) has been an international, interdisciplinary and intercultural journey which would not have been possible without the following people: Geoff Steeves, our chair, and Jaroslaw “JJ” Jaworski, our associate chair, provided guidance and motivation throughout our TP and helped us maintain our sanity. Øystein Borgersen and Pablo Melendres Claros, our teaching associates, worked hard with us through many long days and late nights. Our staff editors: on-site editor Ryan Clement, remote editor Merryl Azriel, and graphics editor Andrée-Anne Parent, helped us better communicate our ideas. Juan de Dalmau, Omar Hatamleh, Goktug “G2” Karacalioglu, and Alex Ryan, our ISU academic staff, supported our boundary-breaking concepts. Our knowledgeable experts, listed below, taught us more than we could have imagined about topics ranging from rocket science to space law. We would also like to separately mention our guest lecturer Buzz Aldrin, the second man to walk on the Moon, who awed us with his presence.  Alastair Reynolds (Science Fiction Author)  Angie Bukley (The Aerospace Corporation)  Buzz Aldrin (Astronaut)  Daniel Glover (Former NASA System Engineer)  Dimitra Stefoudi (Space Lawyer)  François Spiero (CNES Strategic Roadmaps)  Gilles Clément (Space Medicine Specialist)  Jacob Cohen (Chief Scientist, NASA AMES)  Jancy McPhee (Human Exploration and Spaceflight)  Jean-François Clervoy (Astronaut)  Jim Green (Chief Scientist, NASA)  John Connolly (NASA Johnson Space Center)  Kathryn Denning (Anthropologist)  Max Fagin (Made in Space)  Michel van Pelt (Cost Engineer)  Pete Worden (Former Director of NASA's AMES Research Center)  Pieter Blue (Maxwell Institute, the University of Edinburgh)  Ryan Weed (Positron Dynamics)  Sebastian Frederiksen and Karl-Johan Sørensen (Space Architects)  Stephanie Thomas (Princeton Satellite Systems)  Steve Brody (Former NASA Program Executive)  William Kramer (Extraterrestrial Environmental Analyst) ISU and Team Project Fast Transit To Mars wish to express their sincere appreciation to NASA, Lockheed Martin Corporation, and the UAE Space Agency for their sponsorship of this project. We shared laughs, tears, and hugs with our fellow SSP19 participants and will always remember this spectacular ISU SSP experience.

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LIST OF PARTICIPANTS

 Aditi Nilvarna  Anastasia Medvedeva  Bowen Han  Cheng Cheng  Daniel Saslavsky  Erin Kennedy  Feng Ji  Gustavo Jamanca Lino  Hamda Alshehhi  Itai Norber  Jacek Wrobel  Jason Dowling  Jennifer Zhu  Jin Young Choi  Julien Villa-Massone  Kathiravan Thangavel  Lisa Kucher  Nathalie Kerstens  Nitya Pandey  Pablo Bedialauneta  Padmdeo Mishra  Praveen Kumar Thakur  Qiang Gao  Raphael Roettgen  Rijin Kv  Sumaya Al Hajeri  Tess Morris-Paterson  Tina Staebler  Tommaso Tonina  Xuemei Zou  Yang Zhu  Yushou Zhao  Zihan Jiao

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ABSTRACT

This document summarizes the research performed as part of the Fast Transit to Mars Team Project within the 2019 Space Studies Program (SSP) at the International Space University (ISU). Our goal was to determine a method for crewed fast transit using continuous acceleration to reduce the mission length from eighteen months to weeks. This will minimize exposure to radiation and microgravity for human travelers. Important project subteams were identified to handle mission feasibility and hazards, mission profile and orbits, aspects of business and law, spacecraft design, human performance in space, and humanities. The choice of these subteams reflected ISU’s spirit of interdisciplinary research. After defining the mission top level requirements, the Team Project (TP) reviewed and evaluated a wide range of propulsion technologies. This main outcome of the first part of our project was to identify two promising propulsion systems, antimatter catalyzed fusion propulsion and magnetic inertial confinement fusion. Despite our current level of technical readiness, we should be capable of achieving our mission of reaching up to 1g acceleration. In the second part of our project, we defined a mission scenario to demonstrate how this technology could be applied to a fast transit to Mars. In the final part of this report, we presented a roadmap to achieve the goal by 2050, identifying new opportunities arising from this technology and discussing potential impacts of this travel on our society. The report concludes with a summary and pledge for decision makers.

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FACULTY PREFACE

“Propulsion physics is certainly my #1 challenge. With the current chemical propulsion physics we have, humans will not move out of our solar system, and will probably get only as far as Mars.” - John Connolly - Mars Study Capability Team Lead at NASA

Developing transportation technologies from the steam engine to the jet, has allowed us to easily traverse the vast distances of our planet. The result is reduced barriers to the movement of people, resources, and ideas. Beyond Earth, our ephemeral connections to the moon and solar system have been shaped by the technology of chemical propulsion. The modest energy densities of chemical bonds have enabled human flights into low Earth orbit and to the Moon, but new methods are required to migrate to Mars and beyond. Power derived through fission, fusion, and antimatter annihilation, will enable rapid transportation and exploration of the solar system. For nine weeks, a team of 33 people from 17 countries, investigated new propulsion technologies and potential effects on a burgeoning space society. The group has sourced hundreds of published journal articles and consulted with scientists, engineers, business professionals, lawyers, entrepreneurs, and policy makers. After conducting thorough research and analysis of existing knowledge gaps, the team identified ten possible propulsion technologies to consider. After an interdisciplinary review, they selected the most viable propulsion system, antimatter catalyzed nuclear fusion, as a technology capable of journeying to Mars in mere days. The team created a developmental roadmap to outline how this new propulsion system could be realized by 2050 and considered its societal impact. This work will empower decision-makers to boost research and development funding now and help bring about a new era within our lifetimes. It has been a great pleasure to work alongside the participants of TP Fast Transit, an ambitious and determined group, eager to make a difference. You are stars shining with a brilliance to surpass great distances.

Geoff Steeves Jaroslaw Jaworski

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PARTICIPANT PREFACE

The SSP experience is something out of this world. For two months, you are placed within a semi-confined environment with hundreds of people only to meet the most incredible individuals, speak to the crème de la crème of the space world, make friends and learn about yourself. This year, being the 50th anniversary of the Apollo 11 landing, gave us big shoes to fill. The moment when the first people walked on the Moon, represented a single moment in history encapsulating what humanity can achieve when challenged with an audacious target. Since then, images and stories about the astronauts have spread across the world and inspired generations to come. The Apollo missions were the culmination of a new era of ambitious thinking. We wanted to exceed our limits and build a case for a Fast Transit to Mars. Realizing that the scope, the propulsion, and the whole mission concept was in our hands, we decided to imagine something inspiring, daring, and possible. Please enjoy this journey where the imagination, technological ingenuity, and perseverance of our team has led us. The sky is not the limit. Thank you to everyone who challenged us and made us become better interplanetary citizens.

“Get your ass to Mars!” Buzz Aldrin ISU, July 2019

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TABLE OF CONTENTS

Acknowledgements ...... iii List of Participants ...... iv Abstract ...... v Faculty Preface ...... vi Participant Preface ...... vii Table of Contents ...... viii List of Figures ...... xi List of Tables ...... xiii List of Abbreviations ...... xiv Part 1: Project Defintion ...... 1 1.1 Introduction ...... 1 1.2 Mission statement and objectives ...... 2 1.3 Review of propulsion technologies ...... 3 1.3.1 Introduction ...... 3 1.3.2 Propulsion system types ...... 4 1.4 Propulsion assessment criteria ...... 15 1.5 Results of propulsion assessment ...... 19 1.6 Conclusion ...... 19 Part 2: Mission Planning ...... 20 2.1 Introduction ...... 20 2.2 Scenario description...... 20 2.3 Mission feasibility analysis ...... 22 2.3.1 Earth area ...... 23 2.3.2 Mars area ...... 23 2.3.3 Orbital infrastructure ...... 23 2.3.4 Surface to orbit ...... 24 2.3.5 Fuel ...... 24 2.3.6 Habitat...... 25 2.4 Orbital mechanics and trajectories ...... 25 2.3.1 Classical orbital mechanics ...... 25 2.3.2 Continuous acceleration spaceflight mechanics ...... 29 2.5 Safety and risk management ...... 39 2.5.1 General Safety Requirements ...... 39

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2.6 Propulsion system design ...... 44 2.6.1 Antimatter Catalyzed Fusion Propulsion System ...... 44 2.6.2 Magnetic Inertial Confinement Fusion (MICF) propulsion system ...... 48 2.7 Spacecraft design considerations ...... 51 2.7.1 Electrical Power System ...... 51 2.7.2 Thermal Control System...... 53 2.7.3 Radiation Shielding ...... 55 2.7.4 Attitude and orbit control ...... 57 2.7.5 Communications ...... 58 2.7.6 Spacecraft design ...... 59 2.7.7 Mars surface transfer module ...... 62 2.8 Human performance considerations ...... 64 2.8.1 Psychological Impact of Spaceflight ...... 65 2.8.2 Exposure to ionizing radiation ...... 66 2.8.3 Microgravity and human health ...... 67 2.8.4 Physiological Challenges on the human body due to microgravity on the long duration space flight ...... 67 2.8.5 Environmental control and life support ...... 68 2.8.6 Conclusion ...... 70 2.9 Business applications ...... 71 2.9.1 Background - Mars settlement ...... 71 2.9.2 Unique Selling Propositions (USPs) and positioning of the FTS ...... 71 2.9.3 Use cases ...... 71 2.9.4 Sizing up the opportunity ...... 72 2.9.5 Fast Transit Spacecraft and propulsion system cost estimation ...... 74 2.9.6 Costs of 10-crew Fast Transit Spacecraft and propulsion system ...... 74 2.9.6 Costs of 10-crew Fast Transit Spacecraft and propulsion system ...... 75 2.9.7 Costs of 100-crew Fast Transit Spacecraft and propulsion system ...... 76 2.9.7 Financing options ...... 76 2.9.8 Execution timeline ...... 77 2.9.9 Illustrative financials ...... 78 2.10 Legal implications ...... 78 2.10.1 Introduction ...... 78 2.10.2 Legality of the Antimatter Catalyzed Fusion propulsion system ...... 79 2.10.3 Environmental Considerations ...... 79 2.10.4 Using space resource ...... 80 2.10.5 Cooperation ...... 80 2.11 Ethical implications ...... 82

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Part 3: The future ...... 84 3.1 Introduction - Beyond Mars, finally going to the stars ...... 84 3.2 Roadmap to 2050 - Future Foresight ...... 85 3.3 A vision of the future ...... 87 3.3.1 Scientific opportunities ...... 88 3.3.2 Science Missions ...... 89 3.3.3 Additional Studies ...... 91 3.4 Approaching the speed of light ...... 91 3.4.1 Travel time calculation example ...... 93 3.5 Summary – A pledge for decision makers...... 94 References ...... 95

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LIST OF FIGURES

Figure 1. An example of a combustion rocket engine, where the thrust (left) results from the exhaust velocity and the mass flow rate...... 3 Figure 2. An example of subsystems in a typical propulsion system...... 3 Figure 3. Schematic representation of chemical propulsion...... 5 Figure 4. VASIMR Engine (Cassady et al., 2014)...... 6 Figure 5. Schematic of solar thermal propulsion (Image courtesy Laboratory of Space Systems, Hoikado University, Japan)...... 6 Figure 6. NERVA nuclear thermal engine of NASA (Credit: NASA)...... 7 Figure 7. Conceptual propulsion system using theta-pinch plasma focus (Czysz, Bruno and Chudoba, 2018)...... 8 Figure 8. The Orion spacecraft (Dyson, 2002)...... 9 Figure 9. Taxonomy of fission-fusion hybrid propulsion concept (Adams, Cassibry and Schillo, 2014)...... 10 Figure 10. A magnetic field in a toroidal configuration confines and directs the hot plasma within the center of the chamber. (Cohen, et al. 2019)...... 10 Figure 11. Direct Fusion drive schematic based on field reversed configuration (Cohen, et al. 2019)...... 11 Figure 12. The basic concept of PCNF propulsion (Weed et al., 2017)...... 12 Figure 13. The 79Kr breeding cycle for the PCNF propulsion (Weed et al., 2017)...... 12 Figure 14. The fuel cycle for the PCNF propulsion (Weed et al., 2017)...... 13 Figure 15. The basic design of the direct matter-antimatter rocket (Morgan, 1982)...... 14 Figure 16. Transit time between Earth and Mars for a constant acceleration propulsion system...... 16 Figure 17. Maximum velocity of a spacecraft in constant acceleration ...... 16 Figure 18. Mass ratio plotted against specific impulse...... 17 Figure 19. Payload ratio versus specific impulse...... 17 Figure 20. Propulsion technologies comparison chart (Czysz, 2018)...... 18 Figure 21. Mission scenario overview with transfer from Earth to Mars...... 21 Figure 22. When two planets’ orbits lie in the same plane, we can use the Hohmann transfer maneuver to transition from one to the other...... 26 Figure 23. Interplanetary transfer trajectories adapted from (Howard, 2005). 1 and 2 represent the orbits of planet 1 and planet 2. The figure on the left shows travel from planet 1 to planet 2. Figure on the right shows travel from planet 2 to planet 1...... 27 Figure 24. Conceptual flight profile for acceleration and velocity over travel time...... 29 Figure 25. Fast-transit spacecraft design flow, indicating external data (grey), design inputs (blue), calculations (light orange) and key outputs (dark orange)...... 31 Figure 26. Closest approach configuration...... 33 Figure 27. 90° configuration example...... 33 Figure 28. The opposition configuration of the planet...... 34 Figure 29. Thrust and fuel flow over travel time (in hours)...... 35 Figure 30. Net acceleration and velocity over travel time (in hours)...... 35 Figure 31. Distance from starting point and velocity over travel time (in hours)...... 36 Figure 32. Thrust and Fuel flow over travel time (in days) ...... 36 Figure 33. Acceleration and velocity over travel time (in days)...... 37 Figure 34. Distance from starting point and velocity over travel time...... 37 Figure 35. Transit trajectory from Earth to Mars and Jupiter (illustration, not to scale)...... 38 Figure 36. Transit trajectory from Earth to Mars and Jupiter (conceptual flight profile)...... 39 Figure 37. Radioisotope breeder fuel cycle part 1 (Weed, 2019)...... 45 Figure 38. The conceptual design of the PCNF rocket engine with payload (credit: Weed, et al. 2017)...... 45 Figure 39. Silicon Carbide moderator arrays (Weed, 2012)...... 46 Figure 40. Cross Section of the main engine design (adapted and modified from Weed, 2019)...... 47 Figure 41. Design of the Engine (adapted and modified from Weed, 2019)...... 47 Figure 42. Main design of the engine (adapted and modified from Weed, 2019)...... 48 Figure 43. Conceptual drawing of MICF thruster (Czysz P A, 2018)...... 49 Figure 44. Cross-section of fuel used in MICF (Adams et al., 2003)...... 49 Figure 45. Model of the magnetic nozzle (Polsgrove et al., 2010)...... 50

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Figure 46. Expulsion of expanding plasma by a magnetic nozzle (Czysz P A, 2018)...... 50 Figure 47. Electrical power system and subsystems...... 52 Figure 48. Most efficient power generation methods for different power requirements and mission time (Angrist, 1982)...... 52 Figure 49. Thermal control system in the spacecraft environment...... 53 Figure 50. Two types of thermal control systems...... 54 Figure 51. Schematic representation of passive two-layer structure for thermal control...... 55 Figure 52. Radiation shielding scheme...... 56 Figure 53. An example of a closed loop control system (Adapted from Angie Buckley (2019), Spacecraft Guidance and Control, International Space University, unpublished...... 58 Figure 54. Engine array, engine plate and propellant tank...... 60 Figure 55. Connector beams...... 60 Figure 56. Proposed spacecraft side view...... 61 Figure 57. Crew module...... 61 Figure 58. Proposed spacecraft in Mars orbit...... 62 Figure 59. Mars Module Subsystems...... 62 Figure 60. Inside the propulsion system...... 63 Figure 61. Structure of Mars module and its configurations. (a) With heat shield (b) deceleration (c) landing...... 63 Figure 62. EDL of Mars Module...... 64 Figure 63. The physiological impact of microgravity. The arrows signify whether these systems are increased or reduced. Image Source: T. Morris-Paterson, 2019...... 64 Figure 64. List of a single astronaut's intakes and products for a single day (Perry and LeVan, 2002)...... 68 Figure 65. Factors influencing spacecraft cabin air quality...... 69 Figure 66. International fund mechanism...... 81 Figure 67. Future Foresight mapping...... 87 Figure 68. Graphic representing the solar system and its objects. The trajectories to selected destinations are highlighted with white arrows, with travel time in days (adapted from: https://www.universetoday.com/142531/an-orbit-map-of-the-solar-system/)...... 88 Figure 69. Diagram showing how gravitational microlensing can be used to evaluate the atmosphere of an exoplanet. Fast transit will enable easier positioning of the telescope at a different location in space (not on Earth as shown in the image). (source: https://teara.govt.nz/en/diagram/8008/gravitational-microlensing)...... 90 Figure 70. Non-relativistic rocket equation (y1) vs Relativistic rocket equation (y2)...... 92 Figure 71. Relativistic rocket equation varying Isp and the mass ratio...... 92 Figure 72. Relativistic rocket equation for specific Isp...... 93

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LIST OF TABLES

Table 1: Technology Readiness Level ...... 4 Table 2. Comparison of advanced radiator concepts. Adapted from (Albert, et al., 1994) ...... 14 Table 3. Propulsion comparison ...... 15 Table 4. Propulsion Technology: Qualitative Ranking, scores from zero for the worst condition to five for the best condition...... 18 Table 5. Summary of the mission plan and its main concepts ...... 21 Table 6. Selection criteria for general mission outline...... 22 Table 7. Launch opportunity windows ...... 28 Table 8. Calculations relating to MICF and ACFP ...... 32 Table 9. Required coasting phase in percent of travel time for different planets...... 38 Table 10. Likelihood of risk occurrences...... 40 Table 11. Risk probability (P) and severity (S) table with impact and counter measurements, with 1 being low and 5 being high probability or severity. Color-code: 1-8 green (low), 9-16 yellow (medium), 16-25 red (high)...... 41 Table 12. Risk assessment matrix of all risk items in Table 11...... 43 Table 13. Review and comparison of space radiator concepts. Adapted from (Albert, et al., 1994)...... 46 Table 14. TRL of Propulsion Subsystem. Derived from Tara Polsgrove, 2010...... 51 Table 15. MICF Engine Parameters Derived from Polsgrove, 2011 ...... 51 Table 16. Mass distribution of the spacecraft...... 59 Table 17. Examples of major stressors experienced during human space flight. Adapted from Kanas and Manzey (2008)...... 65 Table 18. Examples of major stresses experienced during human space flight. Adapted from Kanas and Manzey (2008)...... 65 Table 19. Oxygen, food and water needs. Derived from Jones, (2002) ...... 68 Table 20. Space station water usage. Derived from Jones, 2002...... 70 Table 21. Space Shuttle water usage needs. Derived from Jones, 2002 ...... 70 Table 22. Market segmentation ...... 72 Table 23. Cost calculations for a 10-crew demonstration spacecraft system for Fast Transit Spacecraft and propulsion system...... 75 Table 24. Cost calculations for a 100-crew spacecraft system for Fast Transit Spacecraft and propulsion system...... 75 Table 25. Execution timeline...... 77 Table 26. Illustrative headline financials...... 78 Table 27: Mission profiles for different destinations ...... 89

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LIST OF ABBREVIATIONS

ACFP Antimatter Catalyzed Fusion Propulsion APP Applications ARRA Agreement on the Rescue of Astronauts, the Return of Astronauts and the Return of Objects Launched into Outer Space BUS Business CERN European Organization for Nuclear Research CP Chemical Propulsion CTBT Comprehensive Nuclear-Test-Ban Treaty D-He3 Deuterium-Helium(3) DNA Deoxyribonucleic acid D-T Deuterium-Tritium ECLS Environmental control and life support EDL Entry, Descent and Landing ESA European Space Agency FSO Free-Space Optical Communication FTPS Fast Transit Propulsion System FTS Fast Transit Spacecraft GCR Galactic cosmic rays GCR Galactic Cosmic Radiation HPS Human Performance in Space HUM Humanities ICF Inertial Confinement Fusion IGA Space Station Intergovernmental Agreement IMU Inertial measurement unit Isp Specific Impulse ISS International Space Station ITER International Thermonuclear Experimental Reactor LAN Local area network LEO Low Earth Orbit LIAB Convention on International Liability for Damage Caused by Space Objects LTD Linear Transformer Driver MP Mission Planning MICF Magnetic Inertial Confinement Fusion NASA National Aeronautics and Space Administration NTP Nuclear Thermal Propulsion OST Treaty on Principles Governing the Activities of States in the Exploration and Use of Outer Space, including the Moon and Other Celestial Bodies OTS Orbital Transportation Service PCNF Positron Catalyzed Nuclear Fusion PEL Policy and Law PROP Propulsion PTBT Treaty Banning Nuclear Weapon Tests in the Atmosphere, in Outer Space and Under Water PuFF Pulsed Fission Fusion RAD Radiation Assessment Detector RCS Reaction Control System

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REG Convention on Registration of Objects Launched into Outer Space S/C Spacecraft SPE Solar Particle Events TBR Trapped Belt Radiation TDRS Tracking and Data Relay Satellite TP Team Project TPNW Treaty on the Prohibition of Nuclear Weapons TRL Technology Readiness Level UN COPUOS United Nations Committee on the Peaceful Uses of Outer Space UNOOSA United Nations Office of Outer Space Affairs USA United States of America VASIMR Variable Specific Impulse Magnetoplasma Rocket

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PART 1: PROJECT DEFINTION

Project definition

1.1 Introduction

Interplanetary travel has historically been limited by chemical or electrical propulsion technology and their very long transit times, ranging from six months to several years for celestial bodies in our solar system. Only a small number of scientific missions have been realized, and no human travel endeavors have been undertaken beyond the Moon. With current technology, future visions of exploration, science, and human settlement are not realistic. In this team project (TP), we aim to enable interplanetary solar system travel for humans, as well as the transportation of robotic missions, cargo, and scientific payloads. We want to make outer space travel an everyday possibility. The idea of space travel has inspired humankind for centuries. Exploration of celestial bodies within the solar system and beyond has a new impetus, and Mars is the most attainable due to its proximity, suitability for human presence, similarity with Earth, and potential for housing various unknown life forms. Since the advent of space flight, major efforts have been made to advance scientific knowledge about the Red Planet, sending orbiting probes and robotic missions to its surface. Based on the currently available propulsion methods, traveling to Mars would take at least six months. Long exposures to microgravity and radiation are detrimental for human health and performance, but shorter transit times require new propulsion technologies to be developed which are currently at experimental or conceptual stages with a correspondingly low technology readiness level (TRL). Constant or continuous acceleration propulsion is key. The main benefit would be a drastic reduction in transit times, less exposure to radiation, and more benign conditions for human space travel in comparison to the microgravity environment. Such technologies would also require new developments in areas such as spacecraft design and life support systems. This revamped interest in Mars, and space exploration in general, coincides with a new era marked by the entry of commercial actors willing to engage in space activities. Fast access to Mars and the larger solar system would potentially unlock new commercial opportunities, increasing the attractiveness of these endeavors to private investors. The presence of humanity on other celestial bodies raises a myriad of issues including cultural, ethical, and legal challenges that need to be addressed.

International Space University 1 Team Project Final Report

1.2 Mission statement and objectives

Our vision is to enable fast and safe interplanetary travel, allowing humans to attain a sustainable presence in the solar system and beyond. Our goal is to develop a crewed transportation system to Mars by the end of 2050, using continuous acceleration propulsion technologies up to 1g. This mission serves the purpose of advancing humankind from a cultural, societal, economic, and scientific perspective.

Primary Objective: To propose the concept of a fast transportation system which minimizes the exposure to radiation to expand human presence beyond Earth and make humankind a multiplanetary species.

Secondary Objectives:

Propulsion (PROP):  Propose concepts of safe and cost-effective propulsion technology to meet the mission requirements.  Check possibilities and advantages of using in-situ generation of propellants on extra- terrestrial bodies. Mission Planning (MP):  Develop new propulsion systems enabling a new energy era.  Develop a technology roadmap including potential spinoffs for terrestrial applications. Spacecraft (S/C):  Propose a spacecraft concept for the recommended propulsion system and mission scenario.  Develop a new generation of radiation shielding.  Develop advanced composite materials. Human Performance in Space (HPS):  Ensure human survival and adequate life support during the journey.  Map physiological and psychological effects of space travel of various durations on the human body and mind.  Provide recommendations for human sustainability through the journey, on Mars, and other celestial bodies in the solar system. Humanities (HUM):  Suggest ethical implications of allowing an interplanetary human future.  Consider the cultural, social, and political impacts on humanity. Policy and Law (PEL):  Suggest an international cooperation framework.  Discuss necessary updates of existing space law.

International Space University 2 Team Project Final Report

1.3 Review of propulsion technologies 1.3.1 Introduction The basic principle of rocket propulsion comes from Newton’s Third Law, that for every action, there is an equal and opposite reaction. The speed of the propellant, and rate at which it is expelled, results in thrust, accelerating the spacecraft in the direction opposite to that of the expelled propellant. There are different means to generate thrust, which will be discussed in the following sections.

Figure 1. An example of a combustion rocket engine, where the thrust (left) results from the exhaust velocity and the mass flow rate.

Propulsion systems usually consist of several subsystems including the engine, propellant tanks, pumps, throttle, and cooling mechanism.

Figure 2. An example of subsystems in a typical propulsion system.

For our review we will be characterizing propulsion systems using three performance characteristics including specific impulse (Isp), thrust (FThrust), and technology readiness level (TRL). Specific Impulse describes how much thrust is delivered by an engine per unit propellant mass flow rate. Isp is the conventional metric of comparing propellant efficiencies. For example, for a chemical propulsion system Isp is given by:

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1 2 ⋅ 훾 ⋅ 푅 ⋅ 푇 퐼푠푝 = ⋅ √ 𝑔0 (훾 − 1) ⋅ 푀푤

Where γ and Mw are properties of the exhaust gas, R is the ideal gas constant, relating changes in gas temperature to pressure and volume. The gravitational acceleration on the Earth’s surface is g0 and T is the exhaust gas temperature. The thrust force is related to the escape velocity and subsequently Isp by the equation:

퐹푇ℎ푟푢푠푡 = 푣푒 ⋅ 푚̇ = 퐼푠푝 ⋅ 𝑔0 ⋅ 푚̇ The propellant mass flow rate 푚̇ is usually expressed in kg/s. Using these equations, we can state Tsiolkovsky rocket equation (К. Ціолковскій, 1903):

푚0 푚0 Δ푣 = 푣푒 ln = 퐼푠푝 ⋅ 𝑔0 ln 푚1 푚1

Δv is the change of velocity of the vehicle assuming no external forces. The masses m0 , m1 are the initial and final mass of the spacecraft. We use the NASA concept of TRL to describe the relative maturity of potential propulsion technologies. The individual TRLs are summarized as follows:

Table 1: Technology Readiness Level TRL Level Implications

1 Basic principles have been observed and reported.

2 Technology concepts and/or applications have been formulated.

3 Analytical/experimental proof-of-concept research has been performed.

4 Component and/or breadboard laboratory validation has been performed.

5 Component and/or breadboard validation tests in relevant environment have been performed.

6 System/subsystem prototype/model demonstration in a relevant environment has been performed.

7 System prototype function has been demonstrated in a space environment.

8 Completed system flight qualified through ground/space demonstration.

9 Completed system flight proven through successful space mission operations.

1.3.2 Propulsion system types Mission duration, transit distance, and transport mass are the greatest challenges for human spaceflight beyond the moon. Different propulsion technologies offer unique capabilities to overcome these challenges. The propulsion technologies we are considering in this report can be split into four different broad categories: chemical, electric, nuclear, and antimatter. In this project, we are considering a range of technologies from classic chemical propulsion to ones based on nuclear and antimatter derived energy. We will review the following categories, sorted based on

International Space University 4 Team Project Final Report

TRL, for fast transit propulsion and for each we will specify the Isp and thrust as these parameters will help us in our selection. 1. Chemical Propulsion 2. Electric Propulsion 3. Solar Thermal Propulsion 4. Nuclear Fission Thermal Propulsion 5. Magnetic Inertial Confined Fusion (MICF) Propulsion 6. Nuclear Pulse Propulsion 7. Pulsed Fission Fusion Propulsion (PuFF) 8. Direct Fusion Driven Propulsion 9. Antimatter Catalyzed Fusion Propulsion (ACFP) 10. Antimatter-Matter Direct Particle Annihilation In the following pages, we will provide a brief introduction to the systems listed above. Key parameters for each technology are summarized in Table 2. The propulsion systems listed in bold above are the two selected as potential candidates for fast transit to Mars and are deserving of a greater level of analysis. 1.3.2.1 Chemical propulsion Chemical propulsion depends on the release of chemical bonding energy through a reaction between fuel and oxidizer or by decomposition. The high-temperature products of combustion are accelerated through a convergent-divergent nozzle and ejected at high velocity to generate thrust.

Figure 3. Schematic representation of chemical propulsion.

Chemical propulsion is broadly classified into different types based on solid, liquid, or hybrid propellant or based on the number of reactants involved. The energy delivered by a chemical reaction is limited by the reactivity of the propellants. Specific impulse increases with the temperature of combustion and decreases with the molecular weight of the combustion products. Both are limited by thermochemistry, and the structural and thermal limitations of the combustion chamber. The highest specific impulse for a chemical propellant ever tested was 542 seconds with a tripropellant combination of lithium, fluorine, and hydrogen. The highly corrosive nature of fluorine makes it very difficult to handle and reduces the life of propulsion system components. A commonly used propellant combination is liquid hydrogen and liquid oxygen which is capable of a specific impulse of 465 s (Sutton and Seifert, 1950) and has a thrust capability of 2 MN.

International Space University 5 Team Project Final Report

1.3.2.2 Electric propulsion A new kind of electrical thruster with many distinctive benefits is the Variable Specific Impulse Magnetoplasma Rocket Engine (VASIMR). This thruster uses a magnet and two radiowave couplers surrounded by gases like hydrogen, argon, or xenon. The gas is converted to superheated plasma by the radiowave couplers. The plasma is converted into a directed jet by a magnetic nozzle.

Figure 4. VASIMR Engine (Cassady et al., 2014).

In principle, the VASIMR engine can generate specific impulse up to 30,000 s and a thrust level around 80 N with 200 MW of power input (Squire J P et al, 2008). 1.3.2.3 Solar thermal propulsion A solar thermal propulsion system uses electromagnetic radiation from the sun to heat propellant that expands through a nozzle to produce thrust. The engine thrust depends on solar intensity and the surface area of the collector.

Figure 5. Schematic of solar thermal propulsion (Image courtesy Laboratory of Space Systems, Hoikado University, Japan).

The limiting factor for solar thermal engines is how hot they can heat the propellant. Solar thermal propulsion can generate a thrust level up to 4 kN with a specific impulse limited to about 1000 s (Takashi Nakamura et al, 2005)

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1.3.2.4 Nuclear fission thermal propulsion Nuclear thermal propulsion is conceptually like solar thermal propulsion except for the source of heat. In nuclear thermal propulsion, the energy released from nuclear fission is used to heat the propellant. The heated propellant flows through the core of a nuclear reactor and expands through a rocket nozzle to create thrust (Czysz et al 2018). Because of its low mass, hydrogen is mainly used as the propellant. The main drawback to this propulsion system is how to initiate the nuclear fission process and effectively shield the spacecraft and its passengers from nuclear radiation. Political and budget issues in the 1970s led space agencies to shelve development plans for nuclear thermal propulsion (NTP) or rockets (NTR). The traditional solid-core NTR propulsion system, like the nuclear engine for rocket vehicle application (NERVA), had a compact nuclear reactor with a flowing propellant. This propellant cools the reactor and is heated in turn as shown in Figure 6. The fusion reaction provides the heat and in NERVA-type reactors, a pressure of ~70 atm, and temperature material constraints limit the Isp to 880–900 s (Czysz et al 2018).

Figure 6. NERVA nuclear thermal engine of NASA (Credit: NASA).

A potential future NTP development is the concept of using an Americium (Am) engine. In 1999, Rubbia proposed to have Am-242 implemented into a space rocket to heat the rocket propellant (Czysz et al 2018). Such a rocket would have the capacity for a full-duration mission to Mars. This engine may be conceivable within the next 20 or 30 years. The actual NTP test by the USA and the former Soviet Union in the 1970s gave an Isp of 800-900 s and a thrust of 105 N. The nuclear thermal propulsion engine is capable of an Isp around 1000 s and a thrust of 100 kN. 1.3.2.5 MICF propulsion Nuclear fusion occurs when nuclei are forced to combine. If the mass of the combined nuclei is less than the sum of the initial masses, then energy is released. This energy can be used to power nuclear fusion propulsion systems. Nuclear fusion reactions generate the highest amount of energy per initial mass than all other types of nuclear reactions. In terms of energy release per unit mass of fuel, the fusion of deuterium and tritium generates about 360,000,000 MJ/kg compared to 80,000,000 MJ/kg from U-235 fission or just 13 MJ/kg from H2-O2 combustion (Czysz, Bruno and Chudoba, 2018). Efficiency is a major advantage of nuclear fusion propulsion as propellants are heated and accelerated directly in the fusion process. The key challenge in a nuclear fusion reaction is giving

International Space University 7 Team Project Final Report nuclei enough kinetic energy to overcome the coulombic repulsion and initiate fusion. In stars, this occurs when nuclei are heated to millions of degrees Kelvin. Considering different factors, the most efficient fusion fuels are deuterium-tritium, abbreviated as D- T, and deuterium-helium3, abbreviated as D-He3. Nuclear fusion can be considered to be initiated and sustained if the rate of energy production is more than that of energy loss and enough of the energy is captured by the system for further reaction. MICF uses a pulsed magnetic field, which implodes and compresses D-T fuel pellets contained in a shell which reflects neutrons back into the fuel (Betti R, et al, 2010). It produces higher confinement density and burn times compared to magnetic confinement and inertial confinement. Because of reduced cost and operational complexities, this concept is gaining attention among other nuclear fusion strategies.

Figure 7. Conceptual propulsion system using theta-pinch plasma focus (Czysz, Bruno and Chudoba, 2018).

Fuel-injected into the reactor is ionized by the high voltage between the electrodes. The magnetic field squeezes or pinches the plasma current thereby adiabatically increasing the temperature leading to fusion conditions (Betti R, et al, 2010). Mixing with inert hydrogen gas increases the thrust at the expense of specific impulse. Both thrust and specific impulse can be varied as per requirements. MICF engines are expected to have a theoretical specific impulse estimated around 19,000 s and thrust levels of up to 38 kN (Czysz, Bruno and Chudoba, 2018). 1.3.2.6 Nuclear pulse propulsion Nuclear pulse propulsion drive, or external plasma pulsed propulsion, studied as part of Project Orion, was originally considered by scientists (Dyson, 1968). The concept was to avoid the conversion of nuclear energy to electric or thermal energy and instead use it directly by a series of atomic bomb explosions outside the spacecraft. Each small nuclear explosive system is surrounded by propellants. They are sequentially ejected and detonated externally to produce high energy plasma at a certain distance from a pusher plate attached to the vehicle. The expanding propellant, in the form of high-velocity high-density particles, strikes the pusher plate to generate thrust. A series of shock absorbers reduce the load on the vehicle and helps moderate the acceleration. This momentum transfer pushes the rocket forward.

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Figure 8. The Orion spacecraft (Dyson, 2002).

Since no attempt is made to confine the nuclear explosions, this concept tries to circumvent the challenges of material structural and temperature limits associated with conventional confined propulsion systems by not confining the nuclear explosion. Even though the temperature in the explosion cloud can be as high as a million kelvin, the pulsed nature of the concept ensures that the interaction time with the vehicle structure is as low as 0.1 ms and only a small quantity of material is ablated with each blast (Czysz, Bruno and Chudoba, 2018). The specific impulse attainable depends strongly on propellant impingement velocity and the pulse unit fraction decided by the explosion power and the standoff distance. Theoretical estimates for specific impulse are on the order 4000 s (Schmidt, Bonometti and Morton, 2013). The incremental velocity of the vehicle per nuclear detonation also depends on force on the pusher plate which is determined by the standoff distance, the diameter of the pusher plate and energy released per detonation pulse. Properly tuned shock absorbers absorb a large portion of impulsive shock later released to propel the vehicle. The nuclear explosions are properly timed to match with the frequency response characteristics of the pusher plate, shock absorbers, and the spacecraft structure. Freeman Dyson in his paper “Interstellar Transport” (1968) gave a detailed analysis and mission studies using pulsed nuclear fusion to reach Alpha Centauri. Dyson estimated that a pulse rate of 0.33 Hz would yield approximately 1g acceleration over the course of the flight, considering the material property limits and assuming a suitable ablation plate is provided on the pusher plate to reject the excessive heat between the explosions. If operated continuously for 10 days, such a system can propel a spacecraft to a maximum velocity of 10,000 km/s. The concept offers high thrust, in of the order of MN. It also offers high specific impulse compared to traditional chemical propellant rocket engines, with high thrust but low Isp, or modern electric propulsion with low thrust but high Isp. Real engineering demonstration tests are considered impossible in the Earth's environment because of the need for unconfined nuclear explosions. The expected specific impulse that the pulse propulsion is capable of is around 4000 s with a thrust level of 260 kN (Schmidt et al., 2000). 1.3.2.7 PuFF propulsion concept A team led by Adams, et al. (2014) at NASA makes use of a revolutionary idea of combining fission with fusion with the latter assisted by the former. Various techniques exist for combining fission and fusion as illustrated by the schematic in Figure 9. The PuFF makes use of the principle of pulsed two-stage nuclear reactions. The fission/fusion reaction is triggered by the compression of nuclear fuel pellets using intense electrical pulses (Adams, et al., 2017). The target cell consists of nuclear fission fuel, uranium, and nuclear fusion fuel, deuterium-tritium. A lithium cylindrical shell surrounds the uranium which in turn surrounds a central cylinder of D-T mixture. The fission-fusion reaction is self-sustaining until fuels are depleted. Though D-T nuclear fusion is easiest to ignite (Adams, et al., 2017), most of the energy released by the reaction is in the form of high-speed neutrons which cannot be controlled for generating thrust, and also cause damage to the propulsive structure. In the proposed concept, the lithium shell surrounding the fission fuel effectively acts as a moderator and reflector of neutrons. By

International Space University 9 Team Project Final Report varying the thickness of the lithium shell, the thrust can be adjusted at the expense of specific impulse.

Figure 9. Taxonomy of fission-fusion hybrid propulsion concept (Adams, Cassibry and Schillo, 2014).

In concept, PuFF can produce specific impulse up to 20,000s and thrust of 30 kN, which would result in Earth-to-Mars travel time of about 37 days (Adams, et al. 2014). 1.3.2.8 Direct fusion drive using field reversed configuration reactor In the white paper, “Princeton Field Reversed Configuration reactor for spacecraft propulsion,” Cohen, et al. (2019) proposed a novel plasma confinement concept for using nuclear fusion for interplanetary transport in direct drive mode.

Figure 10. A magnetic field in a toroidal configuration confines and directs the hot plasma within the center of the chamber. (Cohen, et al. 2019).

The major challenge in any scheme based on magnetic-field confinement of plasma is the development of a suitable arrangement of magnetic and electric fields that contains the plasma without introducing turbulence to the plasma flow. Cohen, et al. (2019) claimed a field reversed configuration for their direct fusion drive concept, a promising technique to achieve the confinement. This configuration will produce a thrust of ~30 N with a specific impulse of 20,000 s.

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Figure 11. Direct Fusion drive schematic based on field reversed configuration (Cohen, et al. 2019).

The flow of cooler plasma around the core fusion region absorbs energy from the fusion reaction, which is then directed to a magnetically shielded nozzle for accelerating the plasma and ejecting it at high velocity, yielding high thrust and specific impulse. Cohen et al. propose using a D-He3 nuclear fusion reaction, which would produce fewer neutrons compared to the simpler D-T reaction. Neutrons can make nearby materials radioactive, limiting their lifetimes and increasing the shielding mass requirement. Also, being electrically neutral, neutrons cannot be controlled and contribute little to propulsion thrust. Even though the D-He3 reaction requires plasma temperature of at least five times that of the D-T reactions, the production of less energetic neutrons in the former reaction allows for reduced mass in neutron shielding elements. Cohen, et al. (2019) claim that boron-carbide shielding, which also serves as a structural support, may be sufficient. The Cohen team claims that the system can be designed, built, and tested in approximately 12 years at a cost of US$250 million. The high specific power system with high thrust and specific impulse, both of which can be varied, makes the scheme one of the most promising technologies for fast transit of crewed modules between Earth and Mars. Major challenges in developing the technology are the availability of rare and expensive He-3, which is required for the nuclear fusion reaction, and the demonstration of nuclear fusion reaction in a controlled manner. Direct fusion drive based on FRC can generate thrust up to 10 N with a specific impulse in the range of 20,000 s (Cohen et al., 2019). 1.3.2.9 Antimatter catalyzed fusion propulsion system The existence of antimatter was first predicted by Paul Dirac in 1928, with first Positron discovery in 1932 and antiproton discovery in 1955 (Gordon, 2002). The annihilation reaction of matter- antimatter yields more energy (9 × 1016 J/kg) than any other known chemical or nuclear or electric propulsion system (Smith and Webb, 2001). The main issues with this technology are the creation and storage of antiparticles, along with the conversion of the total momentum and kinetic energy, into propulsive energy and shielding from harmful gamma rays. A very small quantity of less than 10 nanograms per year of anti-protons are produced globally, mainly in CERN and Fermilab. This quantity is insufficient for a pure antimatter/matter reactor drive (Gaidos et al., 1997). This makes crewed interstellar travel using antimatter drives unlikely in the near future (Howe and Metzger, 1987; Gaidos et al., 1997; Long, Obousy and Hein, 2011). One other major problem is the conversion of the annihilation products into propulsive force. A possible solution for this could be a hybrid system that uses antimatter and nuclear drives (Gaidos et al., 1998). One hybrid design is the Antimatter Catalyzed Micro Fission/Fusion, but the anti-proton

International Space University 11 Team Project Final Report based fusion or fission based technologies are limited by availability and storage of anti-protons, as well as complex spacecraft designs. Another anti-matter particle is the positron, an anti-electron with positive charge. Positrons are easier to produce and store (Smith and Webb, 2001) compared to antiprotons. Positron Catalyzed Nuclear Fusion (PCNF) is a promising technology using these advantages. Recent technological advances in cold positron production (Weed et al., 2017), the creation of dense deuterium clusters on metallic substrates (Shahriar et al., 2009), and measurement of positron-catalyzed fusion (Weed et al., 2017), have shown that a radioisotope positron catalyzed fusion propulsion system is possible, with a specific impulse, Isp of 106. The PCNF uses a positron source generation concept based on dense deuterium(D) fusion neutron capture reaction (Mills, 1992). This technique which uses a radioisotope breeder fuel cycle can allow for much higher positron source intensities and thrust levels greater than 1 kN, which would be required for an interstellar mission (Weed, 2019). The main chain of reactions required for this propulsion system is given in Figure 12.

Figure 12. The basic concept of PCNF propulsion (Weed et al., 2017).

Figure 13. The 79Kr breeding cycle for the PCNF propulsion (Weed et al., 2017).

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Figure 14. The fuel cycle for the PCNF propulsion (Weed et al., 2017).

The PCNF system works on the basis of positron triggered deuterium fusion (Morioka, 1994), taking place within a dense deuterium layer. The fusion reaction generates highly energetic alpha-particles which are subsequently accelerated and expelled at high velocity generating thrust. As a byproduct, the deuterium fusion produces neutrons which in turn generate more unstable isotopes which decay through fission leading to additional positron production. Once initiated, this will be a self- sustaining process. The specific impulse and thrust per engine for a 5-day mission to Mars comes out to 500,000 s and 26 kN respectively, with an engine dry mass of 1152 kg (Weed, R., 2019). 1.3.2.10 Antimatter-matter direct particle annihilation The direct particle antimatter-matter annihilation and its different types of combinations have been discussed in this paper. Matter-antimatter annihilation provides the highest yield for the conversion of mass into the kinetic energy. The specific energy for anti-matter matter annihilation is 9 * 1016 J/Kg and this type of propulsion offers the highest specific impulse > 106 s (Long, Obousy and Hein, 2011). Major obstacles exist in the development of this technology including antimatter containment, and production rates. Penning traps can store up to 109 antiprotons on the order of 10-18 kg but new means must be identified to store useful quantities of antimatter (Czysz, Bruno and Chudoba, 2018). With respect to antimatter production, Fermilab is capable of producing up to ten nanograms of antiprotons per year (Czysz, Bruno and Chudoba, 2018). The cost to produce antiprotons was estimated to be ~ 64 billion USD per gram, as per NASA 1999 report (Schmidt et al., 1999). Antimatter-matter direct annihilation engine (Figure 15) is theoretically capable of specific impulse up to 30 million s which is the highest of all the propulsion technologies considered and also a high thrust of up to 1.2 million N (Smith, et al., 2001).

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Figure 15. The basic design of the direct matter-antimatter rocket (Morgan, 1982).

1.3.2.11 Propulsion comparison

Table 2. Comparison of advanced radiator concepts. Adapted from (Albert, et al., 1994) Criterion SCR ROF RSR MBR LBR RFR LDR/LSR CPR RBMR

Weight Mod. Mod. High Mod. Mod. Mod. Low Low Low

Reliability Mod. Avg. Mod. Mod. Mod. Good Mod. Excel. Good

Maintenance required Low Mod. Low Mod. Mod. Low Mod. Mod. High

Technology readiness Excel. Mod. Good Poor Poor Mod. Mod. Poor Poor

Life expectancy Good Good Mod. Mod. Poor Mod. Excel. Unkn. Mod.

System complexity Low High Mod. High Mod. Mod. High High High

Area required High High Mod. Mod. Mod. Mod. Mod. Mod. Low

Performance Excel. Unkn. Good Unkn. Unkn. Mod. Good Unkn. Mod.

Life cycle cost Low Unkn. Unkn. Unkn. Unkn. Unkn. Unkn. Unkn. Low

Micrometeoroid Mod. High Low Mod. Mod. Mod. Low Low High vulnerability

Space-constructible radiator (SCR), roll-out fin radiator (ROF), rotating solid radiator (RSR), moving belt radiator (MBR), liquid belt radiator (LBR), rotating film radiator (RFR), liquid droplet radiator (LDR), curie point radiator (CPR), liquid sheet radiator (LSR), rotating bubble membrane radiator (RBMR) Abbreviations: moderate (Mod.), unknown (Unkn.), excellent (Excel.), average (Avg.)

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Table 3. Propulsion comparison Propulsion Technology Specific impulse (sec) Thrust capability (N) TRL

Chemical Propulsion 465 2,000,000 9

Electric Propulsion (VASIMR) 30,000 80 5-6

Solar Thermal Propulsion 1000 4000 4-5

Nuclear Fission Thermal Propulsion 1000 100,000 5-6

Nuclear Pulse Propulsion 4000 260,000 2-3

Magnetic Inertial Confinement Fusion 19,000 38,000 2-3 (MICF) Propulsion

Direct Fusion Driven (DFD) 10,000 2,000 2-3

Pulsed Fission Fusion Propulsion 20,000 30,000 2-3

Antimatter catalyzed Fusion 500,000 26,000 2

Antimatter-matter Direct Particle 30,000,000 1,200,000 1 Annihilation

1.4 Propulsion assessment criteria

The previous section included the outcome of the literature research we conducted as a technological assessment of the ten different propulsion systems. We considered concepts ranging from classic chemical propulsion to advanced antimatter propulsion. Each propulsion system has been analyzed to provide its:  Fundamental concept  Performance characteristics  Strengths and weaknesses  Costs To achieve the TP objective, which is interplanetary fast transit, the following functional requirements of the propulsion system were identified:  A sufficient thrust-to-weight ratio, which determines the acceleration and the transit duration,  Sufficient specific impulse, which determines the maximum achievable velocity and mass ratio.

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Figure 16. Transit time between Earth and Mars for a constant acceleration propulsion system.

According to the Figure 16, a constant acceleration of 0.01g results in a flight time of 35 days, which drops to approximately 7 days with 0.25g. Further increases in acceleration do not yield substantial reduction in transit times, and are not worth the significant increase in mass, cost, and complexity of the propulsion system. The maximum velocity of a spacecraft in constant acceleration is plotted below in Figure 17. The velocity ranging from 148 to 743 km/s can be achieved with constant acceleration varying between 0.01g to 0.25 g, resulting in a transit time varying between 35 and 7 ḍays respectively.

Figure 17. Maximum velocity of a spacecraft in constant acceleration

In Figure 18, mass ratio, which is a ratio of the vehicle’s initial mass compared to its final mass after engine shutdown, is plotted against specific impulse. Specific impulse varies from 450s for the most efficient chemical propulsion to 100,000 s for antimatter-based propulsion. Note that for the fast transit application, the minimum specific impulse should be above 10,000 to limit mass ratio.

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Figure 18. Mass ratio plotted against specific impulse.

We can also understand the specific impulse’s significance from the plot of the payload ratio—the mass of payload divided by the initial mass of vehicle—versus specific impulse as shown in Figure 19 below. We see that, to have a payload ratio from 0.05 to 0.3 in a transit mission to Mars in less than 30 days, the specific impulse would have to be greater than 10,000.

Figure 19. Payload ratio versus specific impulse.

Based on the above requirements, it can be seen that a propulsion technology with specific impulse greater than 10,000 and an acceleration of 0.01 g to 0.25 g is required to meet the fast transit requirement. Figure 20Figure 19 shows the jet power versus specific impulse of different technologies. The jet power can be defined as the product of specific impulse and thrust. This demonstrates why nuclear fusion technology is a promising candidate.

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Figure 20. Propulsion technologies comparison chart (Czysz, 2018).

We also identified other requirements in the trade off analysis including those pertaining to: human performance in space, humanities, technology readiness, law, and business. To successfully achieve the mission, we needed an integrated system analyzing different disciplines and perspectives. Taking into consideration all the above requirements, we evaluated each propulsion technology through teams dedicated to each discipline shown in Table 4. Table 4. Propulsion Technology: Qualitative Ranking, scores from zero for the worst condition to five for the best condition.

Evaluation by

thematic teams:

matter matter

-

Total scores genic)

Chemical Propulsion (Cryo Propulsion Electric (MPD) Solar Thermal Propulsion NuclearFission ThermalPropulsion (NTP) NuclearFission Pulse Propulsion(Orion) Magnetoinertial confinement Fusion DrivenRocket (MICF) FusionDirect Driver (DFD) Fission Pulsed Fusion Propulsion Antimatter Catalyzed Fusion Antimatter ParticleDirect Annihilation Propulsion 0.0 3.5 3.7 3.8 4.1 4.2 4.1 4.1 5.0 0.0

Spacecraft - 2.3 2.8 2.8 0.8 4.8 2.1 2.1 2.9 4.6

Mission Planning - - 0.0 0.0 3.0 3.0 0.0 1.0 4.0 5.0

Business 1.0 0.0 0.0 3.5 3.0 5.0 2.0 0.0 4.0 2.0

Legal 5.0 5.0 5.0 5.0 3.0 4.0 3.0 2.0 5.0 5.0

HPS and Humanities 2.4 2.9 3.2 2.6 0.5 4.3 2.8 0.2 3.0 2.4

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1.5 Results of propulsion assessment

The initial screening was carried out based on the specific impulse of each technology, since it was the primary criterion for meeting the mission requirements. The second level of screening was according to the TRL, in which a minimum of TRL of 2 was considered in order to realistically achieve the mission by 2050. Also, the propulsion system must have a sufficient thrust-to-weight ratio, which is determined iteratively by integrated mission analysis by taking into account the spacecraft mass and mission duration. The final screening process relies upon the business applications, legal implication, and effects on human performance. Salient points based on the trade analysis are the following:  Chemical propulsion (CP) has a specific impulse of 465 s, which is incredibly low compared to the minimum required specific impulse of 10,000 s, and so has been screened out.  Antimatter-matter direct particle annihilation was ruled out due to its very low TRL, high shielding requirements against gamma-rays, and extremely high propellant cost with the existing schemes of production and storage of antimatter.  Even though nuclear fission thermal propulsion has a higher specific impulse compared to chemical propulsion, it is still not sufficient to meet the fast transit requirement. Also, it requires handling a large quantity of radioactive materials, and leads to a high risk in the launch, in-orbit assembly, and transit phase of the mission.  The Electric Propulsion and Solar Thermal Propulsion offers low mass and low-cost design due to their inherently high specific impulse. They did not fulfil the fast transit requirement of the mission due to their low thrust capability and were excluded.  The direct fusion drive was ruled out due to its low thrust-to-weight ratio despite its high specific impulse.  We disregarded nuclear fission pulse propulsion due to its low specific impulse, high dry mass, high radiation effect, and varying acceleration which negatively impacts human performance onboard the spacecraft.  Finally, the PuFF has also been filtered out due to its low specific power and relatively high development cost.  Although antimatter-matter direct particle annihilation had the highest specific impulse, we removed it from our list due to its very low TRL, high shielding requirements against gamma- rays, and extremely high propellant cost.

1.6 Conclusion

We selected two technologies that could potentially fulfill the mission requirements, the ACFP and MICF as a backup propulsion technology. The backup option will be developed in parallel with the primary option, considering the technology risk involved due to the low TRL of the primary option.

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PART 2: MISSION PLANNING

MISSION PLANNING

2.1 Introduction

In mission planning, considerations of orbital mechanics and planet configurations are taken into account to determine the best option. All subsystems, as well as business, legal, cultural, psychological, and physical considerations are brought together to set up the best mission from a common point of view. We discuss these considerations in the second part of the report, starting in the following chapter 2.2, Mission Scenario. We will describe our planned 10-person demonstration mission to Mars in 2050. In chapter 2.3, we conduct a feasibility analysis of different mission options and explain why this mission scenario was chosen based on input, constraints, and considerations from different teams based on the seven ISU disciplines. In chapter 2.4, we describe the orbital mechanics needed for this mission and other mission scenarios in detail. We consider risks and safety considerations in chapter 2.5, taking into account the whole mission from ground tests up until successful flight and providing a risk assessment. In chapter 2.6, we discuss the two chosen propulsion systems, followed by the spacecraft design in chapter 2.7. We consider the implications of fast transit propulsion on human health and performance in chapter 2.8. Chapter 2.9 reviews business considerations, while chapter 2.10 studies legal ones. This section of the report closes with the implications that fast transit to Mars might have on humanity in chapter 2.11.

2.2 Scenario description

Our mission is to demonstrate fast and safe travel between the Earth’s orbit and Mars’ orbit. Passengers arrive at the Fast Transit Spacecraft (FTS) by existing means. This may be any transportation system that is feasible to transport humans to Earth’s orbit (e.g. Soyuz). The FTS with the antimatter propulsion system can either be docked to an orbital space station where Orbital Transportation Services (OTS) may dock or the OTS may dock directly to the FTS. The merit of a space station is, that passengers can arrive from different areas of the Earth without having to meet for OTS all in one place. This will save travel time and reduce environmental impact. During the waiting period, the FTS will be in processes of preventive maintenance, undergo checks and will be refueled by a special refueling system (without humans onboard). The OTS will undergo a maintenance process and will return to Earth as soon as it is fully crewed again. When all pre-flight checks have been passed successfully the FTS will undock using a Reaction Control System (RCS) with cold gas thrusters and do a maneuver into the starting position for the fast transit. The Antimatter Catalyzed Fusion Propulsion System (ACFPS) will be started and the FTS will depart with continuously increasing acceleration towards Mars’ orbit. The maximum

International Space University 20 Team Project Final Report acceleration reached is 0.24 g. Around the middle of the trip, the spacecraft will be turned around with a slow angular rate. After fulfilling the turn the ACFPS accelerates in the opposing direction than the velocity of the FTS, therefore decelerating the FTS. At the best configuration of Mars and Earth, Mars is reached after 4.3 days. In Mars’ orbit either a space station is available or the OTS dock directly to the FTS. The transfer process of passengers is the same as performed in Earth’s orbit with existing infrastructure. After refueling new passengers may board for the travel from Mars to Earth. A summary of the mission scenario can be found in Figure 21. At the best configuration, a flight to Mars and back is possible without refueling on Mars’ orbit.

Figure 21. Mission scenario overview with transfer from Earth to Mars.

Table 5. Summary of the mission plan and its main concepts

Mission Concept: Fast and safe travel between Earth’s orbit and Mars’ orbit.

Payload: 10 metric tonnes

Design: Existing infrastructure for surface to orbit travel Interplanetary transit ship powered by antimatter catalyzed fusion propulsion system

Passengers: 10

Orbit: Earth's orbit: circular low Earth orbit (LEO) Mars’ orbit: elliptical orbit capture, circular orbit docking

Demo Mission: 2050

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2.3 Mission feasibility analysis

To choose the most appropriate mission plan, we performed a feasibility analysis using a set of two- dimensional matrices representing on one axis all possible scenario options, and on the other axis all thematic department constraints. These scenarios are related to travel from Earth to Mars and back to Earth. The general plan is to transfer a spacecraft containing 10 astronauts between the orbits of Earth and Mars by 2050, expanding the number of passengers to 100 by 2055. A general overview of the selection criteria can be found in Table 6, where the selected options are highlighted in grey. Table 6. Selection criteria for general mission outline.

From To HPS BUS PROP MP / HUM S/C

Mars’ orbit - - - - - (spaceline)

Mars’ surface Launch needed: Launch from a surface Earth’s orbit - - Not feasible and - involves high forces on (relaunch) higher risk the structure

Launch needed: Launch from a surface Mars’ moons - - Not feasible and - involves high forces on higher risk the structure

Not cost More Mars’ orbit effective due to Longer exposure to - - the number of travel time (spaceline) radiation passengers (10)

Not cost More Launch needed: Launch from a surface Mars’ surface effective due to Longer Moon’s orbit exposure to Not feasible and involves higher forces the number of travel time (relaunch) radiation higher risk on the structure passengers (10)

Not cost More Launch needed: Launch from a surface effective due to Longer Mars’ moons exposure to Not feasible and involves higher forces the number of travel time radiation higher risk on the structure passengers (10)

Not cost More Launch needed: Launch from a surface Mars’ orbit effective due to Longer exposure to Not feasible and involves higher forces the number of travel time (spaceline) radiation higher risk on the structure passengers (10)

Not cost More Launch needed: Launch from a surface Moon’s Mars’ surface effective due to Longer exposure to Not feasible and involves higher forces surface the number of travel time (relaunch) radiation higher risk on the structure passengers (10)

Not cost More Launch needed: Launch from a surface effective due to Longer Mars’ moons exposure to Not feasible and involves higher forces the number of travel time radiation higher risk on the structure passengers (10)

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2.3.1 Earth area We considered departing from or arriving to Earth’s surface, Earth’s orbit, the lunar surface, or lunar orbit. We discarded departing from or arriving to Earth’s surface, because an acceleration greater than 1g is required to overcome the Earth’s gravity at the surface, and our propulsion options are providing an acceleration lower than 1g. Therefore, take-off or landing directly on the Earth surface is not possible. We also rejected using the lunar surface or orbit for departures or arrivals, as we assumed most passengers departing to or returning from Mars would prefer to arrive to or depart from Earth. Any Moon option would require a chemical rocket to launch passengers from Earth’s surface to lunar orbit, adding 2 to 3 days to the total travel time, which would increase radiation exposure and weaken the value of business propositions. Taking off from or landing on the lunar surface would also increase loads on the spacecraft and exposure to lunar surface conditions, raising the risk of engine or spacecraft failure. Finally, a spacecraft capable of landing requires added mass for landing gear, a landing guidance system, and a redundancy system for propulsion in critical phase of flight, precise engine control to perform Entry, Descent and Landing (EDL) with an engine not designed for that purpose). As a consequence of this process, we selected a departure and arrival at Earth orbit. The target Earth orbit must be sufficiently high to not be significantly affected by atmospheric drag. It should also allow for a fast departure to Mars. This leads to the concept of a highly elliptical orbit. However, the final orbit, especially its apogee, will be limited by the chemical propulsion vehicle that will transfer passengers from Earth surface to the interplanetary spacecraft and back. 2.3.2 Mars area We considered departure and arrival from/to Mars surface, Mars orbit, and Mars moons. For reasons similar to Earth’s moon discussion, we discarded the direct landing options, to avoid additional loads on the spacecraft and the detrimental surface environment. Also, it is significantly more complex to design a spacecraft capable of landing on Mars, with additional considerations including thermal control and engine control for EDL, added mass for landing gear, guidance to specific point, and redundancy system for propulsion in critical phase of flight. Mars moons were also discarded to avoid additional travel time and complexity for the travelers. 2.3.3 Orbital infrastructure For passenger transfer, we considered the use of existing or future orbital infrastructure, including ISS, lunar Gateway, a future Earth orbital station and a lunar surface station.  Station in Mars orbit  Station in Lunar orbit (lunar Gateway)  Station in Earth orbit  Station at Moon surface  Transfer ships with docking possibilities In the previous analysis, we determined that moons or moon orbits will not be used to park the interplanetary spaceship for passenger transfer. Therefore, we will discard the lunar Gateway and any station at a Moon surface.

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Transferring passengers first to an Earth orbiting station and then to a lander has an important advantage over simply transferring passengers directly to a lander. Passengers can save up to 24 hours of atmospheric travel by landing close to their final destination. This might not be problematic for a crew of only 10 people. However, our demonstration mission is about testing all systems for future long-term commercial operations for 100 people, where the additional time may be significant. In this scenario, the interplanetary spacecraft would dock with an orbiting station (“orbiting spaceport”) in Earth orbit and offload all passengers to this orbiting spaceport. Several landers will be already docked onto the orbiting spaceport. These landers can belong to and be operated by other companies or organizations. Passengers will then transfer to the relevant landers (according to their purchased ticket). Following a pre-established schedule, landers will undock from the orbiting spaceport and each re-enter using specific trajectories to land at separate points on Earth. This will help passengers save time to get to their final destinations. In the same way, the landers will bring outgoing passengers from different areas of the world to the orbiting spaceport before Mars departure. Should several competing companies operate interplanetary spacecraft (with several destinations— not limited to Mars), it is conceivable to have one or several orbital spaceport operated independently by one or several third parties, and providing spaceport services (transfer passengers from interplanetary spaceships to local landers). While an orbiting spaceport around Earth can enable fast passenger dispatch to/from final locations at Earth’s surface, Mars will not likely have a dense human presence. By 2050, we imagine Mars with one city expanding locally over time rather than many cities all over the globe. Therefore, we do not choose to dispatch travelers to/from several locations on Mars in favor of a direct spacecraft-to- surface lander. It is therefore decided to discard any Mars orbiting station system to transfer passengers. 2.3.4 Surface to orbit We consider this part out of scope of our operations. We anticipate that in 2050, there will be sufficient infrastructure to transfer passengers from a spaceship parked in orbit to the surface and back, both at Earth and Mars. This is a condition for the fast transit business concept to make economic sense. Without an existing human presence on Mars supported by surface-to-orbit capabilities, there will be no market demand for the mass fast transit solution. As a result, it is reasonable to predict an existing infrastructure. For example, we could imagine a fleet of SpaceX Starship-class vehicles operating locally at Mars, using CH4 / O2 fuel, able to carry 100 outgoing passengers from planet surface to the interplanetary spaceship, and in turn, carry to the surface 100 incoming passengers from the interplanetary spaceship. For Earth, smaller-sized vehicles could carry travelers back and forth from the orbital spaceport to their landing pads on Earth. We would not take part in the operations of these surface-to-orbit vehicles. 2.3.5 Fuel Despite the probability of existing infrastructure for passenger loading and unloading, it is not clear that local fuel extraction and supply chain will be developed and reliable during our operation, especially if exotic fuel is required (such as He3 or Deuterium). As a result, we argue that in-house fuel extraction and supply chain must be developed and operated.

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Our main propulsion option relies on a supply of Deuterium. On Earth, Deuterium can be extracted from the ocean. At Mars, Deuterium can be extracted from Mars surface or from Mars moons. In both cases, the Deuterium can be processed by a surface plant and then transferred to the interplanetary spaceship by a local fueler that will take-off from the surface using its own chemical propulsion and deliver fuel to the interplanetary spaceship parked in orbit. The required infrastructure is therefore composed of the following components, on each target planet (Earth and Mars in this scenario):  Extraction plant  Processing plant  Refueling spacecraft assembly/maintenance facility  Launch/landing pad and supporting infrastructure We will refuel by docking at the interplanetary spacecraft. 2.3.6 Habitat Similar to local surface-to-orbit infrastructure, we consider that habitat on the surfaces will be existing in sufficient number and quality or be in the process of being developed. We consider this topic out of the scope of our study.

2.4 Orbital mechanics and trajectories

Orbital mechanics is a field of study used to calculate the flight path of an object in space, when influenced by the gravity of other objects. It is used to plan the flight path of a spacecraft, also known as a trajectory. In traditional space flight, the amount of energy available in a spacecraft is very limited due to the use of the existing propulsion technology, namely chemical propulsion. In traditional missions going from a departure planet to a destination planet, the trajectory is set on departure by a short impulse or engine burn lasting a few minutes. Another impulse slows down the spacecraft upon arrival and inserts it into orbit. Between those two short engine burns, there is a long coasting phase with no engine burn, lasting from several months to several years, depending on the distance to destination. In fast transit spaceflight, the engine burn happens during the entire flight, with the spacecraft performing a 180° rotation about half-way through the trip. The spacecraft accelerates continuously to high speeds for the entire first half of the trajectory, and then following the rotation, decelerates to the relatively low orbital speed of the destination planet. Fast transit mechanics differ from traditional orbital mechanics because gravity is no longer the most significant force acting on the spacecraft during cruise. To guide the reader through these different concepts, we will first discuss traditional orbital mechanics, and then cover fast transit mechanics. 2.3.1 Classical orbital mechanics This section describes traditional orbital mechanics, when the major force acting upon a spacecraft during most of the spaceflight is the gravity of a large body, typically the Sun. The most energy-efficient path to travel from one planet to another is through a Hohmann transfer maneuver (Hohmann, 1925). The orbits of the planets involved must lie in the same plane and must be positioned just right for a Hohmann transfer to be used. We can derive the periodicity of any two planets being in such a configuration to identify the launch window. We divide the travel path into

International Space University 25 Team Project Final Report three parts: the hyperbolic departure trajectory relative to the home planet, the transfer ellipse relative to the sun, and the hyperbolic arrival trajectory relative to the target planet.

Figure 22. When two planets’ orbits lie in the same plane, we can use the Hohmann transfer maneuver to transition from one to the other.

To determine the delta-v requirement at departure, we first match the velocity of the spacecraft at the home planet’s sphere of influence to that required to initiate the outbound elliptical transfer phase, and then specify the periapsis radius of the departure hyperbola. The required ∆v at departure is:

2 ⋅ 휇푠푢푛 푅2 Δ푣푑푒푝푎푟푡푢푟푒 = √ ⋅ (√2 ⋅ − 1) = 푉퐷 − 푉1 (1) 푅1 (푅1 + 푅2)

(Howard, 2005) And the required ∆v at arrival is:

2 ⋅ 휇푠푢푛 2 ⋅ 푅1 Δ푣푎푟푟𝑖푣푎푙 = √ ⋅ √1 − = 푉2 − 푉퐴 (2) 푅2 푅1 + 푅2

(Howard, 2005)

Where R1 and R2 are orbital radii of planet 1 and 2, respectively, around the Sun in near-circular orbit and μSUN is the Sun’s gravitational constant, VD is departure velocity and V1 is planet-1 escape velocity at a specific altitude similarly VA is arrival velocity and V2 is planet-2 escape velocity. The purpose of an interplanetary mission is for the spacecraft not only to intercept a planet’s orbit but also to rendezvous with the planet at the correct time and place. For a rendezvous to occur at the end of a Hohmann transfer, the location of planet 2 in its orbit at the time of the spacecraft departure from planet 1 must be such that planet 2 arrives at the apsidal line of the transfer ellipse - the most distant point in the orbit - at the same time the spacecraft does. Phasing maneuvers are not practical for human missions, because heliocentric orbits– when transferring from one planet to another within the solar system–require large phasing periods of several months or years and such

International Space University 26 Team Project Final Report long waiting time is not realistic for crewed spaceflights. The time required for the phase angle to return to its initial value is called the synodic period, which is denoted as Tsyn:

푇1 ⋅ 푇2 푇푠푦푛 = (3) 푇1 − 푇2 (Howard, 2005) where T1 and T2 are the orbital periods of Planet 1 and 2 with respect to the Sun. Figure 23 depicts a mission from planet 1 to planet 2. Following a heliocentric Hohmann transfer, the spacecraft intercepts and rendezvouses with planet 2. Later, it returns to planet 1 by means of another Hohmann transfer. The major axis of the heliocentric transfer ellipse is the sum of the radii of the two planetary orbits, R1+R2. The time t12 required for the transfer is one-half the period of the ellipse.

휋 ⋅ (푅1 + 푅2) transfer time: 푡12 = 푅 + 푅 (4) √ 1 2 2 ⋅ 휇푠푢푛 (Howard, 2005)

Figure 23. Interplanetary transfer trajectories adapted from (Howard, 2005). 1 and 2 represent the orbits of planet 1 and planet 2. The figure on the left shows travel from planet 1 to planet 2. Figure on the right shows travel from planet 2 to planet 1. For a transfer trajectory between Earth to Mars with low energy, the phase angle between Earth and Mars when the spacecraft reaches Mars is determined by:

휙푃ℎ푎푠푒 푎푛푔푙푒 푠푝푎푐푒푐푟푎푓푡 푎푡 푀푎푟푠 = 휋 − 푛푚푒푎푛 푚표푡𝑖표푛 표푓 푒푎푟푡ℎ ⋅ 푡12 (5) The total required wait period at Mars to reach the same phase angle can be obtained by the following equation; −2 ⋅ 훷 − 2 ⋅ 휋푁 푡푤푎𝑖푡 = (6) 푛푀푎푟푠 − 푛퐸푎푟푡ℎ

Φ is the phase angle between Earth and Mars and nMars and nEarth are the mean motions of Earth and Mars respectively in rad/s. 6 6 9 3 2 Using R1 = 149.6 x 10 km, R2 = 227.9 x 10 km, 휇푠푢푛= 132.71 x 10 km /s , Tearth = 365.26 days, Tmars = 24 30 687.99 days, mearth = 5.974 x 10 kg, msun = 1.989 x 10 kg, we can obtain the Earth-Mars launch opportunity parameters as seen in Table 7.

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Table 7. Launch opportunity windows Launch opportunity between Earth to Mars with Low Energy Transfer

Parameters Values

Synodic Period between Earth to Mars (푇푠푦푛) 797.9 days

One-way Travel Time from Earth to Mars 258.82 days

Total wait period at Mars 453.8 days

The phase angle between Earth to Mars for Launch opportunity 44.57 degree

Sphere of Influence for Earth 925 km x 106 km

Sphere of Influence for Mars 577 km x 106 km

Total trip for Human Mission with Low Energy Transfer 2.66 years

If a spacecraft launches towards Mars from Earth’s orbit at an altitude of 300 km, the hyperbolic excess velocity V∞ would be 2.943km/s and the speed would be vc = 7.726 km/s. Finally, ∆v required for departure hyperbola is following

2 푣∞ km 훥푣 = 푣푐 (√2 + ( ) − 1) = 3.590 . (7) 푣푐 s

Using the rocket equation 훥푣 훥푚 − ⋅퐼 = 1 − 푒 푔 푠푝 = 0.705 (8) 푚 it indicates that 70% of the spacecraft’s mass should be propellant mass for a low energy transfer. From these calculations, we conclude that a mission to Mars using low energy transfer trajectory and chemical propulsion imposes the following constraints:  A limited launch opportunity approximately once every two years  A long travel and wait time  No phasing maneuver possible for human mission due to long period heliocentric orbit  A high radiation exposure  Nano-gravity for a long duration  A high mass for ECLS  A long communication delay from 12 to 25 min one-way Low-energy alternatives to the Hohman transfer approach exist, but only increase flight durations. NASA has flown several missions to Mars using an opposition trajectory; roughly that of a heliocentric Hohmann transfer to Mars. These low energy transits take approximately 250 days and 180 degrees transit around the Sun. In practice, there is a very small energy/mass penalty, for Mars direct entry missions, to reduce the transit time closer to 200 days (Takuto, 2010). The transfers for Mars Odyssey, Spirit and Opportunity trajectories were 200, 208, and 202 days respectively. The energy/mass penalty to arrive at Mars much faster than 200 days with chemical propulsion grows exponentially, as can be derived from the rocket equation (8), quickly becoming impractical. These derivations show the limitations that traditional spaceflight technologies combined with biological

International Space University 28 Team Project Final Report constraints place on a human mission to Mars. To achieve fast transit, we will need superior technology. 2.3.2 Continuous acceleration spaceflight mechanics This section covers continuous acceleration spaceflight, which refers to a spacecraft traveling in interplanetary space with its own thrust producing an acceleration significantly greater than that produced by the Sun’s gravity. We will first introduce the concept of continuous acceleration spaceflight, and then how this concept guided our technology selection process from a mission planning perspective. Finally, we will introduce a more accurate way to model continuous acceleration spaceflight. 2.3.2.1 Continuous acceleration spaceflight mechanics concept The general concept of continuous acceleration is for a spacecraft to accelerate continuously during the trip and perform a 180° rotation at about half the distance to decelerate towards the target. The following Figure 24 represents the acceleration in light blue and the velocity in dark blue. If the spacecraft travels 1g for half the trip, it would travel at -1g following rotation for the second half of the trip. The velocity would increase from 0 to top velocity until the midpoint of the journey and decrease to zero during the second half.

Figure 24. Conceptual flight profile for acceleration and velocity over travel time.

Using this concept, we accelerate from Earth to a mid-point between Earth and Mars, and then decelerate until the destination resulting in the trip time

푑푡표푡푎푙 푑𝑖푠푡푎푛푐푒 [m] 푡푡푟𝑖푝 푡𝑖푚푒 [s] = 2 ⋅ √ m . (9) 푎 [ ] 푎푐푐푒푙푒푟푎푡𝑖표푛 s2 Equation (9) integrates a continuous acceleration and deceleration between origin and destination. 2.3.2.2 Continuous acceleration propulsion To select viable engines among the potential systems, we set a pass/fail criteria, where both available burn time and trip time are computed and compared. If the available burn time is larger than the trip time, then the engine is considered viable, as it will be capable of accelerating during the total trip time to arrive at the destination at the right velocity. This selection criterion ensures the key criterion of our team project of providing continuous acceleration during space travel is met. We require the spacecraft to be able to reach Mars from Earth and back at any point in time, meaning for any planet configuration between closest or furthest approach. As a result, our Trip

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Time is based on the furthest approach distance between Earth and Mars. This method assumes a complete refueling at the destination. Computations are primarily based on a high-level requirement such as the payload mass, which is derived from the number of passengers on board. We defined our mission as accommodating a crew of 10. The rest of the spacecraft mass can be derived from the payload requirement. The payload mass is determined by a certain mass per passenger to consider the mass of the person themselves along with their luggage, food, and life support systems. This is discussed further in the spacecraft systems section of this report. In the following discussion, we will be using the normal rocket equation concepts to calculate propulsion physics and flight dynamics. We consider this assumption acceptable for two reasons. First, the propulsion physics for nuclear and antimatter follow the usual rocket equations, such as conservation of momentum, despite relativistic speeds of the exhaust, as discussed in section 3.4 Approaching the speed of light. Second, the spacecraft flight dynamics follow non-relativistic physics because the relative velocity to references Earth and Mars is too low to have relativistic effects. For example, a transit from Earth to Mars at the furthest approach gives a maximum velocity of 180 km/s, which is only 0.06% of the speed of light so the relativistic effects would be negligible. The primary consideration that drives spacecraft design is desired acceleration, and it is achieved by selecting the best propulsion configuration. To calculate spacecraft acceleration, engine thrust is multiplied by the number of engines, and divided by the spacecraft mass.

m 퐹푡표푡푎푙 푒푛푔𝑖푛푒 푡ℎ푟푢푠푡 [N] 푎푠푝푎푐푒푐푟푎푓푡 푎푐푐푒푙푒푟푎푡𝑖표푛 [ 2] = (10) s 푚푠푝푎푐푒푐푟푎푓푡 푚푎푠푠 [kg] (Roger, 1971) The two key design inputs are number of engines and thrust derived from engine type selection. Another key consideration is the energy available in the propulsion system, to sustain acceleration for a long period of time. We use the term “propellant” to describe the energy source that powers the propulsion system, which is being ejected at a certain velocity to produce thrust. The design input for this is the propellant mass. This value is adjusted to reach the desired trip time. The total propellant mass is divided by the total engine propellant flow rate to obtain the available burn time.

푚푝푟표푝푒푙푙푎푛푡 푚푎푠푠 [kg] 푡 [s] = 푎푣푎𝑖푙푎푏푙푒 푏푢푟푛 푡𝑖푚푒 kg (11) 푚̇ [ ] 푝푟표푝푒푙푙푎푛푡 푓푙표푤 푟푎푡푒 s (Roger, 1971) All these parameters affect other spacecraft masses such as the engine mass, propellant mass, propellant tank mass, and spacecraft structure mass, which recursively impact spacecraft acceleration. As a result, a loop process akin to trial and error is necessary to achieve acceptable parameters.

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2.3.2.3 Continuous acceleration propulsion selection and spacecraft design

Figure 25. Fast-transit spacecraft design flow, indicating external data (grey), design inputs (blue), calculations (light orange) and key outputs (dark orange).

Figure 25 shows how design parameters interact with each other. For our selection process, we chose to exclude any propulsion systems not able to cover the Earth-Mars distance at their closest approach, 59 million kilometers, with continuous acceleration. The pass/fail criteria is Available Burn Time > Trip Time. To achieve this selection process, we adjusted the two following parameters: engine count and propellant mass. This is a repetitive process until the pass/fail criteria is met. In some cases, it is not possible to pass the criteria and the results diverge with increased engine count or propellant mass. When that is the case, we considered the engine unviable and discarded it from the mission analysis. A secondary condition is to have acceleration stronger than the sun’s gravity. If that were not the case, the spacecraft would not be able to accelerate in the opposite direction of the sun. For example, at closest approach, there is a Sun - Earth - Mars alignment. If the sun’s gravity is stronger than the spacecraft acceleration, the spacecraft will drift towards the Sun despite its own thrust. The selection process yielded several viable engines, so we expanded the mission requirements. We increased the required trip distance to an Earth-Mars oppositional configuration, as seen in Figure 1, to yield engines that would allow a larger launch window than just the closest approach. Results Based on analysis of multiple propulsion technologies, including chemical, solar, nuclear, and antimatter, we identified two that may be capable of delivering 10 people to Mars by 2050 with continuous acceleration. MICF and ACFP are the most promising fast transit options, but both involve high mass and high cost, as shown in Table 2.

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Table 8. Calculations relating to MICF and ACFP Parameters MICF ACFP

퐼푠푝 in seconds 19,000 500,000

Thrust in N 38,000 26,150

Dry mass in kg 115,000 1,150

TRL 2.5 2

Economic criteria ³He: $400/g - $1000/g D: $1/g - $10/g

Safety criteria Radiation protection Confinement and radiation

Mass flow rate kg/s 0.408 0.080

Engine restart Yes Yes

Payload ratio 0.008 0.051

Acceleration at start 0.006 g 0.2 g (maximum mass)

Fuel consumed one-way in kg 863,755 27,589

Based on Table 8, we conclude that the travel time and cost for ACFP is significantly less than that of MICF, even though the TRL is slightly better for MICF. Both of our proposed propulsion technologies are not limited to a certain launch window, as they are both capable of accelerating continuously during the entire flight from Earth to Mars and back in any planetary configuration. A high energy transfer trajectory is made possible by these technologies. Common flight trajectories In the Figure 26 and Figure 27 a 0.2g continuous acceleration spacecraft—reflecting the ACFP technology - is used to simulate trajectories from Earth to Mars, using actual planetary orbital data. The following simulations are based on the conceptual method, a simple acceleration for half of the trip and deceleration for the other half of the trip, in an Excel spreadsheet. This model, in addition to VR and future simulations can be downloaded at http://fast-transit.space . Closest approach configuration occurs when the three bodies are in the following alignment: Sun - Earth - Mars. In our example, the distance is 57 million kilometers. However, this distance will vary due to the elliptical nature of planetary orbits and depending on the position of the planets on their orbits. At 0.2g, the trip time when using conceptual modeling is 3.9 days. The 90° configuration occurs when the Earth, the Sun, and Mars form a 90° angle. It represents the half-way mark between closest and longest approach configurations. The distance between Earth and Mars in our example is 303 million kilometers. At 0.2g, the trip time when using conceptual modeling is 9.1 days. In the opposition configuration in Figure 29 the planetary alignment is Mars - Sun Earth. A straight trajectory as represented below is not realistic due to the flight path crossing the Sun. A safe trajectory involving a minimum distance from the sun is required. This will result in a curved trajectory. However, our simplified model only provides straight lines to approximate travel time. At 0.2g with this example configuration, the trip time when using conceptual modeling is 10.4 days.

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Figure 26. Closest approach configuration.

Figure 27. 90° configuration example.

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Figure 28. The opposition configuration of the planet.

2.3.2.3 Mission planning fuel policy 90,000 kg of propellant is required to reach Mars from Earth - or vice-versa - in the worst-case situation, in opposition configuration, when the total distance is 628 million km - 400 million km as a direct line plus a large additional distance to curve around the Sun. Our spacecraft would have a tank capacity of 90,000 kg of propellant. Under an optimal Earth-Mars configuration, it is theoretically not required to have full tanks. For example, during the closest approach, with a distance of 59 million km, only 27,589 kg of propellant would be required. To ensure the best safety margins possible and a contingency plan with a possibility to return to Earth in the middle of the trip, we foresee all departures with full tanks, until the safety and reliability aspects are controlled over years of experience. 2.3.2.4 Continuous acceleration spaceflight mechanics modeling There are two major reasons why the conceptual method in Figure 28 is not practical. The first is the existence of a thrust and an acceleration profile varying with time. The second applies when traveling to distances further to the spacecraft capabilities in constant acceleration mode. To simulate the thrust profile and a full flight from Earth to Mars, we developed an Excel-based model. This model represents the spaceship and respective engine physics including thrust, fuel flow, mass, acceleration, net acceleration, velocity, and distance covered. The time step for this was set to 60 seconds, and approximately 6000-time steps are necessary to cover the entire flight. This model can be downloaded at http://fast-transit.space ACFP ramp-up Looking at Antimatter Catalyzed Fusion propulsion, our analysis indicated that the engine requires a period of 20 hours to increase thrust from 0% to 100%. Given uncertainties about such futuristic engines, we assumed that ramp up will be linear as shown in Figure 29.

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Figure 29. Thrust and fuel flow over travel time (in hours).

When leaving Earth orbit towards Mars, we modeled the spacecraft with 15 engines and calculated thrust and fuel flow as presented in the graph above. For simplicity reasons, this was done by considering only the sphere of influence of the Sun and neglecting the Earth one. There is a linear ramp up for 20 hours (Weed, 2019), and then the thrust is constant.

Figure 30. Net acceleration and velocity over travel time (in hours).

Net acceleration as shown in Figure 30 is calculated by subtracting the Sun gravity (which we define as acceleration due to the Sun's gravity) from the spacecraft engine acceleration. Acceleration increases linearly for 20 hours, and then increases at a much slower rate. This slow increase is solely due to propellant mass decrease over time. Velocity increases according to acceleration.

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Figure 31. Distance from starting point and velocity over travel time (in hours).

The distance from the starting point increases according to velocity as shown in Figure 31. The more advanced model of the spacecraft acceleration, which is discussed throughout this section of the report, is based on a more precise engine model and the respective velocity and position. This model shows that the trip time is increased by about 10 hours due to the engine ramp up phase. ACFP full flight modeling This section will model with more precision the full flight of a continuously accelerating spacecraft, as depicted in the following graphs:

Figure 32. Thrust and Fuel flow over travel time (in days)

Thrust and fuel flow increase from 0 to maximum capacity in about 20 hours, and then provide continuous thrust until switched off, as shown in Figure 32.

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Figure 33. Acceleration and velocity over travel time (in days).

Acceleration increases quickly during the first 20 hours, and then increases slowly from 0.20g to 0.24g over four days. Velocity increases to 350 km/s and then reduces to almost zero relative to Mars to be captured in Mars orbit. The spacecraft travels 59 million kilometers in 4.3 days, from Earth orbit to Mars orbit, as shown in Figure 34.

Figure 34. Distance from starting point and velocity over travel time.

Coasting phase for larger distances In case of distant destinations such as Jupiter’s moons, Pluto, or the heliopause, it might happen that the travel time is longer than the available burn time. In such cases, it is necessary to accelerate for half of the available burn time - where spacecraft acceleration is significantly stronger than Sun gravity - then have a coasting phase where Sun gravity is the only force providing acceleration and traditional orbital mechanics are applicable, and finally a deceleration phase where spacecraft acceleration is again significantly stronger than Sun gravity. This approach could also be used to lower spacecraft cost in Mars missions but has not been considered due to our team project requirement to use continuous acceleration for the entire duration of the trip in an Earth Mars trajectory.

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For example, on a fast transit to Jupiter at its closest approach to Earth, fifteen ACFP engines could provide total burn duration of 10.9 days with 0.332 g acceleration, leaving approximately 24% of travel time in the coasting phase without engine burn. This will increase travel time to Jupiter from 10% to 20% depending on the engine thrust profile.

Table 9. Required coasting phase in percent of travel time for different planets. Coasting of travel Coasting of travel Travel Time closest Travel Time furthest Planet time closest time furthest approach (days) approach (days) approach (%) approach (%)

Jupiter 14 17 24 38

Saturn 20 22 46 51

Uranus 29 31 63 64

Neptune 38 38 71 71

Pluto 39 49 72 78

The following figure shows the trajectory of transit to Mars as an almost straight line—slightly curved by sun gravity, but not discernable at this scale. The transit to Jupiter is shown on a more curved trajectory - curve exaggerated for illustration purposes - due to the longer time required and the additional effect of Sun’s gravity.

Figure 35. Transit trajectory from Earth to Mars and Jupiter (illustration, not to scale).

The following figure represents the acceleration in light blue. For example, it depicts +1g for the first part of the trip, 0g during coasting, then -1g following the spacecraft rotation for the final part of the trip. The velocity in dark blue increases from 0 to top velocity at the end of acceleration phase and during coasting and reduces to zero during the deceleration part.

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Figure 36. Transit trajectory from Earth to Mars and Jupiter (conceptual flight profile).

2.5 Safety and risk management

In our TP, safety is the most important concern. NASA defines safety as “freedom from those conditions that can cause death, injury, occupational illness, damage to or loss of equipment or property, or damage to the environment” (2000). Assuring safety on our proposed mission will require on-the-ground testing, such as of the engines, as well as in-orbit testing of the fully assembled spacecraft. The spacecraft will be assembled in Earth’s orbit as it will be too heavy and too big to be launched from Earth's surface in constructed form. We will need to consider safety during the construction period and minimize risk to humans by using robotic assembly where feasible. There are risks beyond just safety. Our primary risk is that the propulsion technology may not be ready or available in time - or is even not feasible at all. The success of this mission is also dependent on the availability of surface to orbit transportation, which we do not address in this mission design. 2.5.1 General Safety Requirements 2.5.1.1 Ground testing and development Adequate safety precautions must be taken into account for testing of radioactive components of the propulsion system to avoid personnel harm during the event of inadvertent failure. This is also addressed in the legal section 2.10 of this report if there is no possibility of conducting full-scale testing on the ground, alternative options could be investigated for in-orbit testing or testing on the Moon. 2.5.1.2 Launch Launching from any planetary surface carries risk, but there are additional constraints associated with launching nuclear or radioactive materials. Firstly, it is important that there is no residual radioactivity from ground testing in any part of the launched spacecraft parts. Secondly, if any radioactive material needs to be launched it needs to comply with the launching states regulatory requirements. Depending on the amount of radioactive material different precautions need to be considered and safety procedures need to be followed to get a launch approval (NASA, 2017a). In our opinion it is advisable to take this kind of precaution and responsibility into account for any potential environmental hazardous material launched.

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2.5.1.3 Assembly Integration and test All tests which can be conducted on Earth should be carried out on Earth from a business point of view. Also, assembly of the S/C should be tested on Earth. The final assembly will then take place in orbit. 2.5.1.4 Demo missions We propose a fully robotic demonstration mission to precede any crewed mission to assess radiation levels. Uncrewed missions are important for human safety and can be used to conduct experiments and as preliminary cargo transports. Earlier programs, such as Apollo, used the same approach (NASA, 2005). 2.5.1.5 Flight safety For safety reasons there needs to be at least two main engines. The current engine design is using 15 engines. The engines need to have a fully redundant infrastructure. This could be achieved in having two redundant circuits for fuel and infrastructure for each half of the engines. Redundancy is needed for all support systems and all mission-critical systems. There needs to be an escape vehicle for the crew in case of an emergency which can sustain the crew until a secondary spacecraft can reach them. This means that the escape vehicle shall also be maneuverable even when undocked from the rest of the spacecraft. This can be achieved by minor electrical guidance systems. Apart from these precautions redundant systems need to be put in different places. This ensures that in the case of an emergency it is less likely that both systems are affected. Crew members need to be trained on how to behave in the case of an emergency. 2.5.1.6 Radiation Besides the built-in radiation shielding required for human survival, shielding redundancy needs to be considered in critical areas such as reactors and around living areas, to protect against internal (such as a reactor problem) or external (such as meteorite impacts) damage. For reactor malfunctions, a rapid shutdown sequence is critical to prevent shielding damage and excessive radiation damage which can lead to food and water contamination and health damage to the crew. 2.5.1.7 Operational phase As mentioned in our mission statement, we seek a safe trip to Mars. This means that a loss of a spacecraft is not acceptable and for a mission beyond 2050, where passengers are on board. NASA has implemented risk management which reflects probability, frequency, and severity (NASA, 2017b). ESA also uses a very similar risk matrix (Preyssl, 1999). Therefore, we would like to recommendation to adapt the risk management matrix to our space travel.

Table 10. Likelihood of risk occurrences. Score Likelihood of occurrence (p)

5 Near certainty p > 80%

4 Highly likely 60% < p 80%

3 Likely 40% < p 60%

2 Low likelihood 20% < p 40%

1 Not likely p 20%

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Table 11. Risk probability (P) and severity (S) table with impact and counter measurements, with 1 being low and 5 being high probability or severity. Color- code: 1-8 green (low), 9-16 yellow (medium), 16-25 red (high).

P S Risk # Risk (1-5) (1-5) (P*S) Impact Mitigation

1 Propulsion technology not 4 5 20 Might delay project or cause monetary Increase researchers, collaborations for funding, technology feasible, developed, or tested loss. verification, and technology validation. by 2050

2 Fatal failure of propulsion 2 5 10 May result in casualties, progress Use adequate precautions during testing. system such as explosion, slowdown, or progress cancellation. Develop strong versions of structural components for concept structural failure, or radiation demonstration. leakage during tests Properly train personnel.

3 Failure of spacecraft structure 2 3 6 Might stop or slow down progress, Conduct sufficient modeling and analysis. during tests delay project, and lead to monetary Proper inspection and quality assurance activities. loss. Train and certify personnel

4 Fatal failure of transport rocket 1 5 5 Might stop or slow down progress, Conduct sufficient tests. during launch operations of might delay or terminate project. Run proper inspection and quality assurance activities. spacecraft parts Transfer risk to insurance.

5 Failure during in-orbit 3 5 15 Might lead to mission failure, death of Train personnel extensively. Validate assembly and test assembly and testing crew, and monetary loss. procedures. Use robots for in-Orbit assembly.

6 Any critical failure during a 1 5 5 Might lead to mission failure, death of Train personnel extensively. Provide escape vehicle and transit flight crew, and monetary loss. redundancy of critical systems.

7 Any non-critical failure during a 2 3 6 Might lead to mission abortion, injuries Provide redundancy of critical systems. transit flight of crew, or longer travel time.

8 High radioactive exposure 1 5 5 Might lead to death of crew. Radiation shielding and monitoring.

9 Cost overruns and/or funding 2 5 10 Might delay or terminate the project. Plan extensive scenario analyses and stakeholder support. running out

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# Risk P S Risk Impact Mitigation (1-5) (1-5) (P*S)

10 Orbital access restrictions 2 4 8 Might lead to less use than expected or Support stakeholder, debris mitigation programs. in the worst case to inviability.

11 Loss of control of fusion 1 5 5 Might lead to death of crew. Plan rapid shutdown and injection system for radioactive failure, reaction Have escape vehicle.

12 Failure of life support system 1 5 5 Might lead to death of crew. Provide redundancy of critical systems and safeguards.

13 Crew (astronauts) erratic 1 3 3 Might lead to loss of control of mission Do extensive psychological screening. and/or dangerous behavior and lead to deaths or injuries.

14 Crew (passengers, after 2050) 2 1 2 Might lead to loss of control of mission Do psychological screenings of passengers, observe space laws, erratic and/or dangerous and lead to deaths or injuries. and use attendants to guide passengers. behavior

15 Exposure to Microgravity 3 1 3 Might lead to bone or muscle loss. Use artificial gravity for long duration flights such as missions to Pluto and exercise.

16 Exposure to high acceleration 1 5 5 Might lead to possible cardiovascular Plan rapid shutdown possibility for the propulsion system. (>>1g) in case of system failure failure and/or death of crew.

17 Crew and/or passengers (after 1 5 5 Might lead to injuries or death of crew. Do psychological screenings, observe space laws, and use 2050) societal breakdown attendants to guide passengers.

18 Travel, propulsion system, or 2 3 6 Might lead to monetary loss, delay or Use environmental radiation protection. spacecraft noncompliant with stop of the project. Art. 9 (OST)

19 Hull impact by debris and/or 2 5 10 Might lead to injuries or death of crew. Build a robust spacecraft hull. object

20 Unsuccessful docking 1 1 1 Might lead to damage of docking Develop automated docking, collision detectors, and a robust adapter or lead to hull damage. docking adapter.

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Table 12. Risk assessment matrix of all risk items in Table 11.

4, 5, 8, 11, 5 2, 9, 19 5 1 12, 16 17

4 10

3 13 3, 6, 18 Severity

2

1 20 14 15

1 2 3 4 5

Probability

High > 16 Medium 9 - 15 Low < 8

Communication guidelines need to be established to make sure that any incident is known by all key personnel. This is especially important as one team might consider something non-critical whereas it might have a high impact on some other technical team. Everything which is out of the norm needs to be recorded and communicated. To ensure communication lines a communication officer is recommended. In general, our goal is to automate all spacecraft systems to a level where human intervention is only needed when a failure occurs. Human piloting capability will be built in as an override function.

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2.6 Propulsion system design

In this section we compare the suitability and design of the two preferred propulsion systems (antimatter catalyzed fusion and magnetic inertial confinement fusion) from Part 1 for our selected mission. 2.6.1 Antimatter Catalyzed Fusion Propulsion System Current state of the art propulsion systems fail to meet the requirements of 21st-century space missions. The antimatter propulsion system is a very promising candidate for safe human transport with reduced transit times, delivering faster scientific outputs, increasing the payload mass to permit more cargo/piloted missions, and reducing overall transit costs. Each design is a trade-off between mass and complexity, but they all share an increased specific impulse (Isp) well above the most ambitious electric ion-propulsion. We present a positron-based propulsion mechanism that uses positron radioisotope sources in conjunction with an annihilation-catalyzed fusion which includes a key design for a propulsion demonstration which utilizes various fast transit to Mars mission criteria. The main reason for using positrons from antimatter to catalyze the fusion is that they are 2000 times easier to produce than anti-protons. The basic design concept has been given in section 1.2. In this section, we propose a design that can carry 10 humans as payload by 2050. The D-D fusion will create neutrons that can be captured with a one meter thick or ten atmospheric pressure Krypton 78 gas. In this propulsion system, the initial reaction of positron annihilation with matter will produce kaons and pions; kaons decay into muon in 20 ns. These are captured by Deuterium to make mD, which further interacts with D to make ³He + ¹n + muon (non-consumed). Although the fusion rate for D-D muon-catalyzed fusion is only about 1% of the fusion rate for D-T muon-catalyzed fusion, this still gives about one D-D nuclear fusion every 10 to 100 picoseconds. This system produces less energy but generates sufficient fusion by-products to generate the propulsive thrust (Weed, et al. 2017). Muons are similar to electrons but 207 times larger and are highly unstable subatomic particles. Muon catalyzed fusion from antimatter would multiply the production of energy from antimatter. positron emission: 78Kr → 78Kr + 1 e⁺ + 0.94 MeV of kinetic energy positron annihilation: e⁺ + matter → pion (5%) or kaon (95%) kaon decay: kaon → muon (80%) in 20 nsec muon capture: muon + D → mD Fusion: mD + D → ³He + ¹n + muon (non-consumed) (0.07 – 1.5 nsec) muon decay: muon + time → electron + neutrinos (2,200 nsec) The above basic reactions summarize the fuel cycle for antimatter catalyzed fusion. Such a radioisotope breeder fuel cycle paves the way for enormous positron source intensities and the desired thrust for interstellar travel. The isotope 79Kr decays into isotope 79Br and is further synthesized to produce carbon tetrabromide. The fusion cycle is initiated with the help of the Bromine isotope which emits positron and neutrino to catalyze the D-D fusion reaction cycle, continuously.

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Figure 37. Radioisotope breeder fuel cycle part 1 (Weed, 2019).

The main design parameters corresponding to human mission rated spacecraft are: Dry engine mass: 1150 kg Thrust per engine: 26150 N (Weed, 2019) Specific impulse: 500,000 s Flow rate per engine: 0.07997 kg/s These parameters can be used to calculate the final number of engines required to carry 10 humans to Mars in a 5 to 15-day one-way trip. The number of days will vary depending on the relative position of Sun, Earth, and Mars, as well as the final orbit design.

Figure 38. The conceptual design of the PCNF rocket engine with payload (credit: Weed, et al. 2017).

The thermal management system is a critical subsystem of the propulsion engine. Since this type of propulsion system gives high thrust and operates at high temperature, it is necessary to implement heat space radiators for this propulsion type. The tube-fin structure which carries the circulated coolant liquid is the most common design. The tube walls must be thick enough to avoid micrometeoroid penetration. The mass of the radiator could be as much as half of the total propulsion system mass. Below is a summary of space-based radiator technology.

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Table 13. Review and comparison of space radiator concepts. Adapted from (Albert, et al., 1994). Criterion SCR ROF RSR MBR LBR RFR LDR/LSR CPR RBMR Weight Mod. Mod. High Mod. Mod. Mod. Low Low Low Reliability Mod. Avg. Mod. Mod. Mod. Good Mod. Excel. Good Maintenance required Low Mod. Low Mod. Mod. Low Mod. Mod. High Technology readiness Excel. Mod. Good Poor Poor Mod. Mod. Poor Poor Life expectancy Good Good Mod. Mod. Poor Mod. Excel. Unkn. Mod. System complexity Low High Mod. High Mod. Mod. High High High Area required High High Mod. Mod. Mod. Mod. Mod. Mod. Low Performance Excel. Unkn. Good Unkn. Unkn. Mod. Good Unkn. Mod. Life cycle cost Low Unkn. Unkn. Unkn. Unkn. Unkn. Unkn. Unkn. Low Micrometeoroid Mod. High Low Mod. Mod. Mod. Low Low High vulnerability Space-constructible radiator (SCR), roll-out fin radiator (ROF), rotating solid radiator (RSR), moving belt radiator (MBR), liquid belt radiator (LBR), rotating film radiator (RFR), liquid droplet radiator (LDR), curie point radiator (CPR), liquid sheet radiator (LSR), rotating bubble membrane radiator (RBMR) Abbreviations: moderate (Mod.), unknown (Unkn.), excellent (Excel.), average (Avg.)

It is important to note that in the positron and Deuteron number densities, the fusion rate relies only on the overlap. The positrons must be cooled or moderated to guarantee optimal overlap with the fuel design. They are incredibly energetic or heated (mean power~250keV) when generated and it is therefore hard to constrain their motion. One major challenge to date is the ability to use realistic electrical and magnetic fields to regulate these very hot positrons. In a process called moderation, they must be cooled down before all the positrons could be used as such. To date, the moderation process had an efficiency of less than 1%. There are new approaches, such as silicon- carbide arrays, that enhance the efficiency of cooling moderators by multiple levels. Combining the method of charged particle extraction through electromagnetic field-assisted moderation, as seen in Figure 39, the use of efficient moderation will facilitate intense and concentrated positron pulses that can deposit a significant amount of energy into fuel targets.

Figure 39. Silicon Carbide moderator arrays (Weed, 2012).

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Figure 40. Cross Section of the main engine design (adapted and modified from Weed, 2019).

The cross-section of the basic design for antimatter catalyzed engine is illustrated in Figure 40. The further studies on advanced engine model enhancement could consist of performing simulations of positron deuterium annihilation, plasma combustion, electromagnetic fields, and fuel fusion. Exterior shell simulations are also essential. This would help to deliver better values for the system's specific impulse. (Long, Obousy and Hein, 2011)

Figure 41. Design of the Engine (adapted and modified from Weed, 2019).

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Figure 42. Main design of the engine (adapted and modified from Weed, 2019).

The above CAD engine design is the preliminary design and an illustration of how the ACFP engine may look like in the future. This proposed design could be used as a basis to help develop manufacturing by 2050 for the expected final launch of the piloted ten-person crew mission to Mars. The performance numbers of the main design parameters corresponding to human mission rated spacecraft are based on current technological estimations, however, advances in miniaturization and performance of the different components in this design are advancing rapidly. By 2050, the current models will be much more powerful and efficient enough to transport humans to Mars and beyond with near 1g acceleration. 2.6.2 Magnetic Inertial Confinement Fusion (MICF) propulsion system MICF (Polsgrove et al., 2011) is an attempt to incorporate the best of two confinement concepts: Magnetic Confinement and Inertial Confinement which attempts to minimize the cost, mass, and complexity of the fusion propulsion system. The technology is based on the transient magnetic field that squeezes or pinches plasma leading to fusion. This replaces the high-power bulky laser system with less expensive and lighter capacitors. In Z pinch fusion, a transient current of very high intensity (up to 10 MA) is directed along the longitudinal axis of the reactor for about 10 ns which induces a radial magnetic field (Cuneo, 2001). The resulting Lorentz force squeezes plasma by increasing the radial magnetic pressure. Z pinch is thus the radial implosion of cylindrical plasma under the influence of a strong magnetic field created by axial current flow through the plasma. The intense current required to create the radial magnetic field is provided by rapidly discharging the bank of capacitors. Fuel (usually D-T) is injected into the reactor and is surrounded by an annular stream of Lithium. The lithium liner serves as a current return path as well as a radiation shield. The reaction between neutrons and lithium produces tritium thus adding further fuel to the fusion reaction.

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Figure 43. Conceptual drawing of MICF thruster (Czysz P A, 2018).

Fuel for MICF The main fuel for MCF is deuterium although D-H3 also can be considered if H3 can be produced from the moon or outer planets (Cuneo, 2001) The arrangement of the plasma target is shown in the figure below (Adams et al, 2003). The primary fuel composed of the D-T mixture is surrounded by the main energy source (Deuterium) and hydrogen or lithium in the outer most layer.

Figure 44. Cross-section of fuel used in MICF (Adams et al., 2003).

Using D-T makes the ignition easier because of the low ignition temperature whereas the major portion of energy release comes from the D-D reaction in the surrounding layer. The minimizes the required for a quantity of tritium which is a scarce element. Also, the D-D reaction generates neutrons of much lower energy compared to that in D-T reaction (2.45 MeV vs 14.1 MeV). Using hydrogen or lithium act as neutron moderator thus shielding the structure from harmful neutrons as well as converting otherwise wasted neutron kinetic energy into useful charged particle kinetic energy. The following reactions are expected to take place in the plasma upon reaching fusion conditions (Miernik, 2013). D + T → He4 (3.5 MeV) + n (14.1 MeV)

D + D → T (1.01 MeV) + p (3.02 MeV) 50% D + D → He³ (0.82 MeV) + n(2.45 MeV) 50% D + He³ → He4 (3.6 MeV) + p (14.7 MeV)

T + T → He4 + 2n + 11.3 MeV

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Magnetic nozzle Axial thrust for the spacecraft is generated by ejecting the hot plasma at high velocity by making use of a magnetic nozzle (Polsgrove et al, 2011) as shown in the figure below. It is made of a series of carefully contoured conductor rings that create a three-dimensional paraboloid type magnetic field.

Figure 45. Model of the magnetic nozzle (Polsgrove et al., 2010).

Each ring is composed of two separate conducting rings and various structural, cooling and neutron moderator/shielding mechanisms. A high-temperature superconducting ring generates the initial seed magnetic field whereas a conventional conductor carries the electrical current by back emf induced during plasma expansion. An yttrium based superconductor alloy having a transition temperature of approximately 90 K and kept cooled by liquid nitrogen is used as the seeding coil. The second coil made of metal matrix composite consisting of molybdenum in the titanium diboride matrix provides good electrical and strength properties at high temperature. Thermal coolant (Fluorine-lithium-beryllium) flowing through the channels inside the ring assemblies and C-C structure supporting the coils executes the dual role of heat removal as well as the capture of gamma rays and neutrons. The plasma created by the pinch being electrically conductive interacts with the magnetic field. The magnetic field is pushed back by the expanding plasma thereby compressing it and exerting thrust on the vehicle structure. As the plasma expands and cools down the compressed magnetic field springs back and accelerates the plasma out of the nozzle. By doing this the magnetic nozzle converts most of the radial momentum of isotropically expanding plasma into axially directed momentum and thus minimize loss and maximize thrust. (Czysz P A, 2018)

Figure 46. Expulsion of expanding plasma by a magnetic nozzle (Czysz P A, 2018).

Electrical energy generation system The electrical energy required to power the next pulse and initial magnetic field by tapping off a portion of the current that is induced in the coils during the expansion of the plasma cloud within

International Space University 50 Team Project Final Report the magnetic nozzle (Polsgrove, 2011). The electrical energy is stored in a capacitor bank and the magnetic field is stored in a superconducting magnetic energy storage device. The discharge time of 100 ns is met by capacitor banks with low capacitance. The parameters of the propulsion system working on based on MICF is given in Table 15.

Table 14. TRL of Propulsion Subsystem. Derived from Tara Polsgrove, 2010. Sub-System TRL

High temperature Z-pinch 4

Intense electrical pulse power 4

Magneto-hydrodynamic electricity 5

Thermo-nuclear equations of state 3

Dynamic plasma radiation shielding 3

Advanced structures 2

Reaction containment 2

Table 15. MICF Engine Parameters Derived from Polsgrove, 2011 Thrust per engine / N 38,000

Specific impulse / s 19,500

Engine dry mass / kg 115,000

Propellant mass flow rate / kg/s 0,204

2.7 Spacecraft design considerations 2.7.1 Electrical Power System The electric power system of a spacecraft is a network of electric and electronic components deployed to generate, store, and distribute power to the different systems and components of the spacecraft. The power system is a crucial component of the spacecraft, and failure of the power system will directly lead to failure of the mission. With the advances in human space exploration and the increasing complexity of spacecraft operations, power system requirements for space missions have become stricter. The electric power system can be divided into three main parts: power generation, storage, and management. The power generation is the subsystem that creates the necessary energy for operation of the spacecraft. There are three common power generation techniques. The chemical power generation systems make use of fuel cells to generate electricity (Schmidt-Rohr, 2018). Solar power generation systems take advantage of the photoelectric effect to generate electricity with

International Space University 51 Team Project Final Report sunlight (Zhang, 1996). Traditional nuclear generators make use of nuclear decay to heat metal and use the Seebeck effect to generate electricity (Rowe, 2006). Power storage is the system that stores the energy generated by the power generation system. It is usually composed of batteries or big capacitors. Finally, the power management system controls the power distribution and the status of all systems, (including the power system).

Figure 47. Electrical power system and subsystems.

Different missions have different power requirements. In the Apollo era, the lunar module’s system used three fuel cells to supply and manage electric power, in which power requirements were 65 kilowatt-hours at a rate of 4kw to satisfy a 35-hour lunar stay time (Campos, 1972) The International Space Station (ISS) power system uses solar panels. There are eight solar array wings used in the ISS now, each one made of about 33,000 solar cells; when fully extended, the wings are 35 meters long and 12 meters wide and can generate 84-120 kw. Lithium-ion batteries are used to store the electric power. (Wright, 2016) According to Angrist’s (1982) estimation, a one-week mission to Mars would take require 2000-4000kW. Apart from power requirements, efficiency is one of the main drivers of the power system design. Figure 48 shows the most efficient power generation methods for different power requirements and mission time.

Figure 48. Most efficient power generation methods for different power requirements and mission time (Angrist, 1982).

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For the power requirement of a Mars mission, only nuclear dynamic systems meet the requirement. Since we have adopted a nuclear propulsion scheme, the same nuclear source will provide the necessary energy for the nuclear thermal power system and will also greatly improve the efficiency of the spacecraft propulsion system. The estimated power generation for the main propulsion system of the spacecraft is 1 MW (Weed, 2019). This covers the power generation requirements for the spacecraft operation during the transit period. For spacecraft operation while cruising (i.e., with no active propulsion), we will use cryogenic H2/O2 fuel cells as a second power source. Cryogenic hydrogen/oxygen are stored in two cryogenic propellant tanks and burned in an expansion engine driving the turbine to generate electric power. The fuel cell generates water, that can be consumed by the crew, the same way as in Apollo’s lunar module (Campos, 1972). 2.7.2 Thermal Control System The main task of the thermal control system is to adjust the temperature for each system of the spacecraft. Different instruments and systems onboard, including humans, have different temperature requirements and it is not safe to operate them outside the required temperature ranges.

Figure 49. Thermal control system in the spacecraft environment.

The thermal control system maintains the internal and external temperature of the spacecraft within a given range and reduces the impact of space temperature fluctuations by managing the heat dissipation of the spacecraft. There are three modes of heat transfer in nature: radiation, conduction, and convection. In space, there is no heat convection between the spacecraft and the outside environment because the outside environment is a vacuum, and conduction requires a medium. The same principle applies to convection. Therefore, the main factor affecting the temperature change of the spacecraft is heat radiation. To control the temperature of a spacecraft, we must balance heat received from external and internal sources, and heat radiated to space. There are two commonly used spacecraft thermal control methods, passive and active. Passive thermal control systems use elements that control the temperature of the spacecraft with no

International Space University 53 Team Project Final Report external control, just by taking advantage of physical processes and properties of materials. Active thermal control systems make use of sensors and actuators to continuously adjust the temperature to the desired one. Large spacecraft with restrictive thermal requirements usually incorporate both passive and active technologies. One example of such an implementation is the ISS.

Figure 50. Two types of thermal control systems.

The ISS thermal control system consists of a passive thermal control system and an external active thermal control system. The passive thermal control system includes insulation, surface coatings, heaters, and heat pipes. When the thermal load exceeds the capacity of the passive thermal control system, the active thermal control system starts. The active thermal control system consists of two independent loops, which both use mechanically pumped liquid ammonia in closed-loop circuits. The active thermal control system is capable of rejecting heat up to 70 kW, and provides 14 kW of heat dissipation (Boeing, [8]). For a Mars mission, there is a need to deal with extreme thermal conditions, as the spacecraft will be moving either to or away from the Sun. The challenge for the thermal control system is to provide enough heat-dissipation capability during the hot operating phase and enough heating during the cold phase to maintain a constant temperature of the equipment in the spacecraft. The spacecraft moves from Earth orbit to Mars orbit. There is a period during the transit when the spacecraft is close to the Sun, and this results in a dramatic temperature gradient on the spacecraft between the surfaces facing the Sun and the ones in the shadow. This is the worst-case scenario, and only happens when the transit to Mars is done when Earth and Mars are in opposition or close to opposition. To achieve control during this phase, we will use a passive thermal control and an active thermal control combination system. For the passive thermal control system, we will use a two-layer structure. The first layer consists of multilayer insulation blankets, which are designed to minimize radiative exchanges in the vacuum of space. They are composed of several layers of aluminized plastic film (e.g., Mylar or Kapton) and can reach an effective emittance value of about 0.001 (Fortescue, et al,. 2011). The second layer uses heat pipes and two-phase systems (capillary pumped loops [CPLs]), in which a volatile working fluid that is circulated by capillary action in a porous core structure transfers heat in the form of latent heat of vaporization (Faghri, 1995).

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Figure 51. Schematic representation of passive two-layer structure for thermal control.

In case the spacecraft thermal loads exceed the capacity of the passive thermal control system, the active thermal control system will begin to operate to maintain thermal equilibrium. The active thermal control system uses mechanically pumped two-phase loops, which are based on CPLs, adding a mechanical pump to the liquid circuit, which will increase the heat transfer capacity and make the loop less sensitive to instability caused by bubbles in the liquid line or nucleation within the evaporator. To dissipate the heat generated by the equipment, a wing-like structure is placed on the outside of the spacecraft, and the heat generated by the electrical components is conducted through the radiator and finally discharged into the space in the form of heat radiation. 2.7.3 Radiation Shielding The pace of human space exploration is gradually moving from near-Earth space to the Moon, Mars, and deeper space. In the planning and implementation of crewed deep-space exploration missions, the first thing to consider is the safety and health of astronauts. Radiation is one of the main risk factors that threaten the health of astronauts in crewed deep space flight. This has become one of the main bottlenecks restricting humankind's move into deep space, and the "invisible blade" that harms the health of astronauts. Fast transit technology, by dramatically reducing travel times, would significantly reduce the risk of radiation to astronauts. In space flight, ionizing radiation is mainly derived from trapped belt radiation (TBR), galactic cosmic rays (GCR), and solar particle events (SPE) (Spillantini, 2014). GCR bombard the solar system in an even distribution from elsewhere in the galaxy, possible starting with supernovae and similar explosive events. The heavy particles account for about 98% of GCR, with the remainder being electrons. SPE is derived from the high-energy charged particle flow generated by solar flares, mainly composed of protons, so it is also called the solar proton event. The outbreak of solar particle events is random and unpredictable at present, but it has a certain correlation with the solar activity cycle. In low Earth orbit (LEO), spacecraft are largely protected by the Earth's magnetic field and the protection of the atmosphere, so the radiation dose to astronauts is mainly derived from geomagnetic capture radiation, which is radiation trapped in Earth’s magnetic field that hits the spacecraft, and galaxy cosmic radiation. In LEO, the contribution of GCR to total dose is highly correlated with orbital height and inclination (Spillantini, 2014). Reaching Mars means leaving the protection of the Earth's magnetosphere. On the way to and from the Earth, the main sources of radiation are the GCR and the SPE.

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There are two approaches for radiation shielding, passive and active. Passive shielding refers to the use of spacecraft bulkheads as shielding materials. The principle is that charged particles gradually lose energy as they pass through the shielding material, and when the thickness of the shielding material is greater than the range of the particles therein, the incident particles will be deposited in the shielding material, protecting the spacecraft and the astronauts. The radiation dose in the cabin is severely reduced (Spillantini, 2014). Passive shielding is not sufficient to ensure the radiation safety of crew during deep space exploration. Increasing the thickness of the cabin body is costly and interaction of the radiation with the shielding material can generate secondary radiation, which increases the difficulty of protection and brings new problems. Active shielding creates an artificial strong magnetic field around the living space of the spacecraft that deflects radiation particles. The main advantage of active protection technology is that it is lower mass than passive protection technology, lower construction cost, and can effectively reduce the damage of space radiation to astronauts and on-board equipment. The superconducting magnetic ring proposed by Piero Spillantini (2011), sponsored by ESA, pointed out that for the protection system of large-scale spacecraft, using passive protection, has a mass of 3359 kg, while the mass of using active protection in the form of a three-coiled magnetic ring is 700 kg. The active protection mass is smaller than the passive protection by a factor of three. Our spacecraft for fast transit to Mars adopts a combination of passive protection and active protection. We designed a passive system using an aluminum layer and water layer. The aluminum layer provides the necessary structural support and preliminary radiation protection for the passenger compartment, and the water layer further strengthens the protection system. Our approach to active protection involves a superconducting coil around the passenger compartment to generate an artificial strong magnetic field (B=10T). The power required to generate such a magnetic field is produced by the power generated by the propulsion system (the antimatter catalyzed fusion drive generates power by itself, in the order of 1 MW per engine).

Figure 52. Radiation shielding scheme.

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2.7.4 Attitude and orbit control The ability of an interplanetary spacecraft to accurately orient itself and follow a pre-designated path is critical to mission success. The systems developed to execute these tasks are called the guidance, navigation and control system (GNC) and they effectively have seen a huge technological shift in the last three decades; from operations highly dependent on ground support to largely autonomous control systems with adaptability built-in (Truszkowski et al, 2009). The attitude of a spacecraft is its orientation in space (Markley and Crassidis, 2014). Orbital control of a spacecraft occurs when the operational orbit is maintained within the parameters of acceptable accuracy (Mazzini, 2015). If performed correctly, orbital control maneuvers enable an efficient use of fuel necessary for longer duration space missions. Prior to configuring the attitude and orbit control systems for a spacecraft, we need to define the requirements to achieve the overarching mission objectives. For typical interplanetary exploration, there are three parts of the mission that require attention from a navigation and guidance perspective: i) escape from the departure planet, ii) heliocentric flight, and iii) arrival at the destination planet (Tang and Conway, 1995). However, this project is proposing two rendezvous points to connect with established Earth and Mars in-orbit stations, and therefore the requirements are split into five distinct phases: i) escape from Earth’s surface and in-orbit rendezvous with Earth’s in-orbit station, ii) escape from Earth’s in-orbit station, iii) orbit to orbit transfer, iv) in-orbit rendezvous with Mars’s in-orbit station, and v) entry to Mars atmosphere, descent, and landing. The following technical section will discuss each of these areas from both an attitude and guidance perspective. Technical Approach i) Escape from Earth’s surface and in-orbit rendezvous with Earth’s in-orbit station. This section of the guidance system will work similarly to the current techniques used for crewed vehicle docking with the International Space Station (ISS). This area can further be split into the lift- off, orbital maneuvering, and docking phases. The lift-off, also known as boost/ascent, phase is pre-computed and the role of the guidance system at this stage is to take the vehicle through the atmosphere with optimal efficiency (minimizing fuel use) and maintain structural loads within the design parameters. The guidance algorithm determines the throttle based on the altitude, to control the g-forces felt by the crew. Once through Earth’s atmosphere, steering and throttle values adjust to remain within predefined constraints such as vehicle altitude, velocity, flight path angle, and orbital plane (Kulkarni and Krishnakumar et al, 2005). The first attempt to link a capsule to a target vehicle was performed by Neil Armstrong and Dave Scott, where the Gemini VIII capsule and the Agena spacecraft tried to dock, but the Gemini VIII capsule lost control resulting in an aborted mission (Hacker and Grimwood, 1977). Now successful docking is a regular feature of both crewed and uncrewed capsules rendezvousing with the ISS. The Soyuz docks with the ISS using the same method developed for the MIR space station (IGLA), which was replaced by the KURS system. The system enables control from either the ground or from the active on-board spacecraft computer (local control by a cosmonaut/astronaut, if required). The system is radar based and determines relative attitude and position from directional antennas. The relative position and strength between the radar pulse, antenna, and transponder enable rendezvous and docking (Hinman and Bushman, 1991). ii) Escape from Earth’s in-orbit station. A spring-loaded mechanism has typically been used to release spacecraft from the ISS into a trajectory that is reliable and repeatable. The primary role of the attitude control system at this stage is to gauge the spacecraft position with reference specifications, and to continually adjust until the correct attitude is achieved. For the most part, closed loop control systems are in use with

International Space University 57 Team Project Final Report spacecraft. These follow a continual process of sensor feedback, calculation of error, control signal for adjustment and output as seen in Figure 53.

Figure 53. An example of a closed loop control system (Adapted from Angie Buckley (2019), Spacecraft Guidance and Control, International Space University, unpublished. iii) Orbit to orbit transfer. The guidance systems suitable for a mission of this nature have already been developed for deep space satellites. However, the most useful development has been from asteroid interception systems, which have been designed with high velocity tracking (Hawkins, et al., 2012). These systems use two tracking methods: ranging and two-way Doppler. Ranging is the precise measurement of radio signals travelling to and from the spacecraft proving the distance to the ground station. Two- way Doppler measures the signal’s change (Doppler shift) to give the velocity along that line-of-site (range-rate). The random error on range is approximately 1 meter; impressive accuracy explaining why the European Space Agency has been using this system (Delta-DOR) since 1986 and has proven its capability as far as Venus (Maddè, et al., 2006). iv) in-orbit rendezvous with the Mars in-orbit station. The guidance systems in use today have evolved from proportional navigation, which were developed for the homing phase of missile flight. The principle of this system is based on rapidly rotating parabolic mirror and electronics with infrared radiation (IR) to identify errors, then compensate by applying a moment to a gimballed mirror to ensure the target is maintained. This system can also be used as a rendezvous device (Yuan and Hsu, 1994). Proportional navigation may be a useful tool for shorter range targeting for the in-orbit rendezvous, alongside the docking procedures mentioned in section i). v) entry to Mars atmosphere, descent, and landing. We assume that entry, descent, and landing (EDL) capabilities from Mars orbit will be in place by 2050, based on the current developments of EDL techniques for propulsive landing on Earth. 2.7.5 Communications All communications will be bi-directional. Onboard the spacecraft will be a local area network (LAN) which will use the Wi-Fi-equivalent standard in 2050. On this LAN, all sub-system computers will be able to exchange information. The docking modules will communicate with the LAN through a Controller Area Network (CAN) bus connection to the docking computer. The docking computer will then communicate with the necessary sub-system computers to send the data.

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Multiple spacecraft can communicate with each other using a star network mesh topology. The number of simultaneous connections from spacecraft to spacecraft will be limited to two. The reason for this limitation is to balance between redundancy and latency. Too many connections in this mesh network would result in lower latency of the communications. When in range, the spacecraft can connect to the ascent/descent module. This would be used as a data relay between spacecrafts for when the connections are greater than two. The spacecraft connects to Earth directly or through data relays. The direct communications would be received by a designated ground station depending on the spacecraft location and Earth rotation. The data relays include the Earth Ascent & Descent Module, the Lunar Orbiting Platform Gateway, and the Tracking and Data Relay Satellite (TDRS) (NASA, 2001), or anticipated follow on missions. The spacecraft connects to Mars through data relays, and eventually directly when habitats and ground stations are established on the surface of Mars. The data relays include the Mars Ascent & Descent Module, and new Mars Orbiters. Given that the spacecraft is traveling in space and not entering or exiting any atmospheres, the array of various communication transceivers and antennas will be mounted on the exterior of the spacecraft, with pan and tilt mechanisms to adjust the positioning. 2.7.6 Spacecraft design The spacecraft has been designed to meet all the criteria of the different subsystems and mission requirements. The number of engines is 15, in order to reach Mars orbit in a maximum time of 13 days, with a crew of 10 people. The spacecraft designed is a minimal mass vehicle that can carry 10 people safely to Mars. The structure mass ratio is 37%. This is a slightly higher number than the usual ratio used, because of the requirement for passive radiation protection. This brings the maximum structural mass of the spacecraft to 70 tons of kg, as the payload is 10 tons, the mass of all engines is 17.25 tons, the propellant mass is 90 tons, and the tank mass is 6.75 tons. The payload includes the mass of the astronauts and the mass of the life support system necessary to support them for one trip to Mars. Future spacecraft would have a higher fraction of the total mass allocated to the life support system, as the number of astronauts increases, and the system has to be reliable for multiple missions.

Table 16. Mass distribution of the spacecraft. Payload (kg) 10,000

Engine mass (kg) 17,250

Propellant (kg) 90,000

Tank mass (kg) 6,750

Spacecraft structure (kg) 70,000 (max.)

Total mass (kg) 194,000

The array of 15 engines is attached to the engine plate as shown in Figure 54. The plate serves as an attachment point as well as a passive radiation protection element.

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Figure 54. Engine array, engine plate and propellant tank.

The propellant tank contains cryogenic deuterium in liquid form. The tank serves as a passive radiation shield as well. The tank is contained in a half-sphere-shaped structure. The half-sphere structure allows for easy attachment of the engine plate to the spacecraft and acts as further passive radiation shielding. The engine plate, tank, and half-sphere structure are made of composite materials, to bring the mass down while ensuring structural integrity. The sphere-shaped structure is connected to the payload and systems bay by eight steel beams. These beams carry part of the loads on the spacecraft from the thrust of the propulsion system. The central beam is hollow and built with composite materials. This central bar also carries loads, but its main function is to serve as an interface between the engine and tank part and the rest of the spacecraft. It can transfer power and fluids between both parts of the vehicle.

Figure 55. Connector beams.

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The cylindrical part attached to the connector beams is the payload and systems module. It is built in aluminum and can hold the basic payload needed to carry 10 people to Mars. It is also a systems hub, holding most of the spacecraft’s subsystems. It also has docking capabilities for crew and cargo transfer.

Figure 56. Proposed spacecraft side view.

The radiators are part of the thermal control loop and are used in case heat must be removed from the spacecraft. The crew module is at the top of the spacecraft. To further separate the engine array and the crew, we will use five composite beams. This gives structural integrity, radiation protection (by increasing the distance between crew and radiation source) and is mass-efficient. The central beam is like the one connecting the payload bay and the propellant tank. In addition to transporting fluids, power, and data, crew can move from the crew module to the payload bay through this beam. The crew module consists of four cylindrical hollow enclosures, where the living quarters, science equipment, and common habitat are hosted. The crew module can hold 10 people and is built with aluminum for radiation protection.

Figure 57. Crew module.

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The top part of the crew module has a cupola, similar to the one on the International Space Station. This cupola is only accessible when solar activity is low and no single particle events are predicted. The cylindrical part of the crew module is actively protected from radiation by magnetic fields outside the spacecraft, generated by superconductors and powered by the power system that uses the propulsion system to generate electrical energy. The total structural mass of the spacecraft is 67.17 tons.

Figure 58. Proposed spacecraft in Mars orbit. 2.7.7 Mars surface transfer module The Mars module is an independent spacecraft pre-positioned on Mars orbit. After transferring the crew on the carrier to the surface of Mars, the Mars module returns to Mars orbit with the humans that will transit from Mars to Earth. As an independent spacecraft, the Mars module has several subsystems, as shown in Figure 59.

Figure 59. Mars Module Subsystems.

The design criteria of some of these subsystems is similar to the spacecraft subsystem design, such as GNC, thermal control system, and life support system. This section will provide a broad description of the propulsion system, electric power system, and the structural design. Because of the presence of water on Mars, solar energy can be used to electrolyze water to produce hydrogen and oxygen. Cryogenic hydrogen/oxygen expansion engines can be used for the propulsion of the Mars Module (Victor, 2010). The propellant tanks are placed on top of the engine and

International Space University 62 Team Project Final Report connected to it for propellant supply. The propellant tanks are filled on the surface of Mars. The Mars module is equipped with four main engines and four rail attitude control engines for propulsive landing and ascent and attitude adjustment. To optimize the layout of the propellant tank structure, the propellant tank adopts a parallel supply mode, as shown in Figure 60.

Figure 60. Inside the propulsion system.

As the Mars module is an independent system from the transit spacecraft, electric power is required to operate the module’s subsystems and perform functions such as control, information transmission, and communications. The primary power source is from the hydrogen-oxygen engine, which pushes the turbine to rotate and work to generate electricity. The designed Mars module exterior structure is shown in Figure 61. The entire structure can be divided into three capsules, namely the propulsion capsule, transition capsule, and crew capsule. The propulsion capsule has the engines used for powered landing and ascent. While performing the EDL maneuver, the engines are covered by a heat shield. After the maximum heat load phase has been passed, the heat shield is discarded, and a supersonic parachute is released to slow the module down. After the capsule enters the subsonic regime, the engines are ignited, and propulsive landing takes place. If the propulsion system is decelerated throughout the re-entry process, the propellant that the propulsion system needs to carry will be greatly increased. Therefore, a lighter parachute is used to complete the initial deceleration of the re-entry process, and the propulsion system is started after the parachute is thrown.

Figure 61. Structure of Mars module and its configurations. (a) With heat shield (b) deceleration (c) landing.

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The EDL process is shown in the figure below.

Figure 62. EDL of Mars Module.

2.8 Human performance considerations

Human performance in space can be divided into two over-arching approaches; the first is medical oversight of astronauts, where the primary concern is the health and well-being of the astronaut, and secondly the biomedical research for us to better understand how the human body adapts to microgravity environments. This section will view these two areas as interchangeable, and each learns from and complements the other. In addition, these areas can further be divided (from a medical oversight and biomedical research perspective) into psychological and physiological adaptations, either as a result of reduced gravity, or impact of the spaceflight arena itself. Figure 63 outlines the physiological impact.

Figure 63. The physiological impact of microgravity. The arrows signify whether these systems are increased or reduced. Image Source: T. Morris-Paterson, 2019.

The psychological impact of spaceflight has only been highlighted as an area that requires significant attention over the last 20 years. Prior to this, the bodies and individual who governed space

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activities were not receptive to astronaut or cosmonaut reports of feeling sub-optimal. In most instances today, it is accepted that space is considered an extreme environment and therefore a variety of psychological changes will occur. “The most critical problems facing humans in long duration spaceflight, after the biomedical problems, are the psychosocial and psychological problems” Oleg Atkov, Russian Cosmonaut (237 days aboard Salut 7, 1984). 2.8.1 Psychological Impact of Spaceflight A successful human mission is characterized by robust morale, healthy psychological and psychosocial consequences for the crewmembers their mission control support staff and family and friends back home on Earth. In practice this is difficult to achieve as life in an extreme environment will always require some adaptation, and the very definition of an extreme environment is that it does not provide physiological and psychological comfort (examples on Earth include the polar regions, high mountains (altitude), and underwater habitats (Kanas and Manzey, 2008). To sustain human life beyond the Earth it is useful to control stressors and the potential damaging consequences they may bring. Stressors are elements of the environment that affect humans, usually in a negative and disturbing manner. In space, people usually suffer from four kinds of stressors: physical, habitability, psychological, and interpersonal (Kanas and Manzey, 2008). Examples of such stressors are in the Table 14

Table 17. Examples of major stressors experienced during human space flight. Adapted from Kanas and Manzey (2008). Habitability Psychological Interpersonal

Vibration Isolation/confinement Personality conflict

Ambient light Heterogeneous crew Cultural differences and language Temperature/ pressure Danger and risk barriers

Air quality Monotony Personal preference for leadership Tension of workload or followership vs. their crew role Stresses are the reactions of a person to one or more stressors. In space, there are four kinds of stresses that affect human beings: physiological, performance, interpersonal, and psychiatric (Kanas and Manzey, 2008). Examples of these stresses are listed in Table 18.

Table 18. Examples of major stresses experienced during human space flight. Adapted from Kanas and Manzey (2008). Physiological Performance Psychiatric Interpersonal

Space sickness Disorientation Mental disorder Tension Vestibular problems Visual illusions /depression Territorial behavior Sleep disturbances Attention deficits Muscle weakness Isolation Bodily fluid shifts Error proneness Hallucinations Bone loss and hypocalcaemia Psychomotor problems

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Behavioral issues are common among crew members, regardless of how well the crew are selected and trained. As the duration of a mission increases, astronauts need to spend a longer amount of time away from their families, friends, and their natural environment. The lack of typical daily routines also reinforces feelings of isolation, as travelers will spend much more time in a confined spacecraft with a reduced amount of control over their environment (NASA, 2009). During the Russian Soyuz and Salyut space station missions in the 1970s to 80s, there were several occasions when cosmonauts reported uneasiness. Most of these issues were found to be psychological issues and in several cases the missions were terminated early to prevent escalation in behavioral and psychiatric problems (Kanas, 2008). A vital concern which needs to be considered is that every space traveler can be vulnerable to psychological issues, independently from their level of training. In terms of our proposed fast transit to Mars, in the best-case scenario, our proposed mission propulsion system can reduce the Earth to Mars journey duration from six to seven months to less than five days. We expect that this reduction in transit time will significantly mitigate exposure to psychological stressors, although they will still need to be considered. 2.8.2 Exposure to ionizing radiation Space is a dangerous environment for humans, particularly for long duration travel involving exposure to high levels of ionizing radiation, with its associated increased risks of cancer, cataracts, and cardiac damage. Astronauts outside the protection of the Van Allen Belts are exposed to the high energy Galactic Cosmic Radiation (GCR), and there is an intermittent risk of Solar Particle Events (SPE). Ionizing radiation is the most significant barrier to long-term space missions, including setting up a permanent human presence on Mars (Bacal and Romano, 2016). For this project we need to consider both potential sources of radiation from space weather and the propulsion system. Radiation exposure is usually described as Grays or rads (100 rads = 1 Gray). The anticipated impact of this radiation on human health is measured as Sieverts (Sv) or rems (100 rems = 1 Sv). Organ specific limits for astronauts over a 30 days flight have been defined by NASA as 0.250 Sv for blood forming organs, 1 Sv for the eyes, and 1.5 Sv rem for the skin. Entire career limits range up to 0.5 Sv for blood forming organs and the eyes; and 6 Sv for the skin. For comparison, three x-rays deliver approximately 0.001 Sv. A six-month journey (182 days) to Mars would expose an astronaut to approximately 0.3 Sv of radiation from the space environment (Nicogossian, et al. 2016), or 0.002 Sv per day. A recent retrospective analysis of the effects of ionizing radiation on astronaut and cosmonaut health (Reynolds, 2019) failed to find evidence that historical doses of space radiation pose any excess mortality risk for astronauts and cosmonauts. However, it is important to note that future missions of deep space exploration will likely offer much greater doses of space radiation than historical doses, which will lead to a different risk profile for future astronauts and cosmonauts. An important motivation for faster travel to Mars is that career limits for current astronauts (for blood forming organs) are between 1.5 to 4.0 Sv for males and 1.0 to 3.0 Sv for females (the range is lower for younger astronauts, and is also lower for females) (Nicogossian, et al. 2016) which, with current technology, may be reached in a single round trip to Mars. There is a need for a greater understanding of these risks, new in-transit monitoring (such as blood-based biomarkers) and new repair DNA technologies (CRISPR-type technologies may enable preflight DNA comparison and repair). Countermeasures for ionizing radiation under investigation include: ● medications such as vitamins with antioxidant properties that can retard cell division, giving the body time to fix damage before harmful mutations can be duplicated (Okunieff, 2008);

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● radiation protection vests that include micro dosimeters, which are being evaluated on NASA’s Exploration Mission-1; and ● space weather monitoring to avoid high levels of sunspot activity. The proposed propulsion system will result in faster transit times to Mars, which will significantly decrease the risks of ionizing radiation exposure from travel (there is still a need for shielding on the surface of Mars, however this we consider this out of scope of this project). Adequate shielding will need to be incorporated in the spacecraft design to reduce crew radiation exposure. 2.8.3 Microgravity and human health One of the major benefits of International Space Station research is an improved understanding of how the human body behaves in the space environment over 3-12 months missions and what kind of countermeasures are required for successful human space flights. The term microgravity refers to an environment where the apparent weight is negligible. During a space flight, astronauts experience a multitude of gravitational forces with the highest potential forces coming from re-entry. In the Mercury spacecraft era 7.6 to 11.1 +Gx were experienced, 4.3 to 7.7 +Gx in the Gemini projects, and 3.3 to 6.8 +Gx in the Apollo (Kumar and Norfleet, 1992). The most pervasive physiological stressor in all human spaceflight is microgravity. There are mainly four kinds of gravity fields astronauts will experience during a journey from Earth to Mars: 1. Gravitational force of approximately between 1 g to 3 g during launch (based on the space shuttle) (Crippen, 2010). 2. Microgravity experience of approximately 0.2 g between the transition from Earth orbit to Mars orbit. 3. On the Mars surface, astronauts will be exposed to approximately one third of Earth gravity. 4. After returning to Earth they will need to readapt to 1g. 2.8.4 Physiological Challenges on the human body due to microgravity on the long duration space flight The transition between different gravity fields is complex and affects the spatial orientation, head- eye and hand-eye coordination, balance, locomotion, and increased risk of experiencing motion sickness. In addition, there are many negative effects of microgravity on human health. A fast transit to Mars would decrease the duration of physiological changes from reduced gravity. If this transit were possible at 1 g acceleration, then we assume the negative effects of microgravity would be avoided. This acceleration is unlikely to be obtainable for the entire transit duration and currently the effects of long duration travel at constant gravity levels higher than microgravity but less than 1g need further investigation. Research is still required to determine the relationship between different gravity levels and effects on human health.

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Figure 64. List of a single astronaut's intakes and products for a single day (Perry and LeVan, 2002). 2.8.5 Environmental control and life support To support life in harsh environments such as space, we need to equip ourselves with systems that can produce our basic biological needs such as air, water, and food. For this to succeed we need the use of environmental control and life support (ECLS) in our vehicles. For life support systems for various duration missions, the only concrete example we can build on for our spacecraft's journey to Mars is the International Space System (ISS). Here are the numbers of a human’s needs:

Table 19. Oxygen, food and water needs. Derived from Jones, (2002). Standard crewmember needs kg

Oxygen 0.84

Food solids 0.62

Water in food (rehydration) 1.15

Food preparation water 0.76

Drinking water 1.62

Total oxygen, food and water mass 4.99

In terms of air requirements for ECLS, the following subsystems are required: - Filtration - Adsorption (single pass/VSA/TSA) - Catalytic oxidation - LiOH chemisorption - Condensing heat exchange

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Since breathing is a very fast process and creates carbon monoxide as a byproduct, an open loop system (meaning oxygen supply must be renewed) cannot be considered for long durations because of the huge quantity required (0.84 kg/crewmember; Table 19). Therefore, in the future a system to infinitely recycle air should be developed (for example in a closed loop system, where oxygen is recycled from the carbon monoxide by-product). A potential way of this becoming a reality is using carbon filters and electricity from the power system to reverse the process. Using such systems ensures the autonomy of the spacecraft and enables greater transit distances to further explore our solar system.

Figure 65. Factors influencing spacecraft cabin air quality.

Air quality is important beyond respiration. Proper temperature and humidity levels and circulation are needed to control bacteria spreading and to mitigate the lack of sweat vaporization, particulate settling, and odor dissipation in microgravity. The air purification systems deployed on ISS are the best example we have for ECLS that can maintain human life for long durations. The current technology of absorption of catalytic oxidation is a good basis to start form but we need to complement it to overcome its inefficiencies and further reduce its reliance on consumable resources. The water requirements from an ECLS unit, (for example, for the ISS) include 2.4 kg/day for drinking water and hydrating food (required as dehydrating food can reduce food resupply mass by two thirds). From Table 20 and Table 21 we can see that the water used for hygiene is around 25 kg/day, approximately 10 times higher than the water needed for drinking and hydrating food. Estimated required energy ranges from 2,300 to 3,200 kcal per day with an additional 500 kcal/day are needed on EVA days. Caloric requirement considerations for ECLS are based on basal energy expenditure (BEE) by the following formula (Clément, 2011): 푊표푚푒푛 퐵퐸퐸 = 655 + (9.6 ∗ 푊푒𝑖𝑔ℎ푡 [푘𝑔]) + (1.7 ∗ 퐻푒𝑖𝑔푡ℎ [푐푚]) − (4.7 ∗ 퐴𝑔푒 [푦푒푎푟푠]) 푀푒푛 퐵퐸퐸 = 66 + (13.7 ∗ 푊푒𝑖𝑔ℎ푡 [푘𝑔]) + (5 ∗ 퐻푒𝑖𝑔푡ℎ [푐푚]) − (6.8 ∗ 퐴𝑔푒 [푦푒푎푟푠])

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Table 20. Space station water usage. Derived from Jones, 2002. Water usage (kg) Minimum Nominal Maximum

Food Preparation 0.40 0.76 0.91

Drinking 0.21 1.62 1.77

Consumed Total 0.61 2.38 2.68

Shower 1.82 2.73 2.73

Dishwasher 3.63 5.45 5.45

Hand washing 3.63 4.09 4.54

Toilet flush 0.00 0.50 0.73

Laundry 0.00 12.50 12.50

Hygiene total 9.08 25.27 25.95

Total 9.69 27.65 28.63

Table 21. Space Shuttle water usage needs. Derived from Jones, 2002. Water usage (kg) Minimum Nominal Maximum

Food Preparation 0.73 0.89 1.22

Drinking water 0.27 1.70 3.57

Consumed Total 1.00 2.59 4.79

Station Hygiene Total 9.08 25.27 25.95

Total Shuttle and Station 10.08 27.86 30.74

2.8.6 Conclusion In summary there are a number of important considerations and countermeasures required to support and protect human health during fast transit travel. These include protection from ionizing radiation, physiological changes due to microgravity and psychological effects from space travel. In addition, nutritional and life support requirements need to be considered. One advantage of fast transit travel is a reduction in mass required for food and life support.

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2.9 Business applications 2.9.1 Background - Mars settlement At the core of our business opportunity analysis, we assume that there will be a permanent human settlement on Mars by 2050. Both NASA (Foust, 2019) and SpaceX have announced plans for human missions to Mars in a 2030s and 2020s timeframe, respectively. SpaceX’s President Gwynne Shotwell has noted that cargo flights, using its next generation Starship vehicle, could commence as early as the 2025 Mars-Earth close approach (Dunn, 2019). Elon Musk has stated that SpaceX would use every close-approach launch window going forward (Musk, 2019), and has also recently given an expected run rate of Raptor—the engine used by Starship—production (Musk, 2019), as of the end of 2019, that implies a production rate of 12 fully-reusable Starships p.a. Assuming previously quoted capacities of 100 metric tons (pure cargo) or 100 passengers (crewed version), it becomes clear that, by 2050, there is scope to lift enough cargo from Earth to establish critical infrastructure (e.g. habitats, ECLS, food production and storage systems, power generation systems, propellant production plants, local transport, local communications, etc.) as well as to bring thousands of potential settlers. In summary, even with a very conservative approach to SpaceX announcements, we expect a Mars settlement could have of over 10,000 people supported by critical infrastructure (see above) by 2050 and having started to engage in commercial activities. If this scenario is achieved, we would expect the size of the settlement to continue growing at a double-digit percentage rate. Under this scenario, we see commercial demand for the FTS. In addition to NASA and SpaceX, new space-faring nations, such as China, India, and the UAE have also stated their intentions to conduct Mars exploration missions and devoted resources to it, increasing the chances of multiple settlements appearing on Mars by 2050 (ISRO, 2017; Jones, 2019). 2.9.2 Unique Selling Propositions (USPs) and positioning of the FTS The key strength of the FTS is its speed advantage of four days’ transit time at close approach versus approximately 180 days for CP. This speed results in further advantages including less radiation exposure and reliance on a convenient orbital window as even at the maximum distance between Earth and Mars, transit time would be a mere 13 days. FTS is more expensive than CP. Natural market segmentation will exist between the FTS and CP. Where time-sensitive and relatively cost-insensitive customers would use the FTS, others would use the cheaper CP. Use cases for CP would likely include the transportation of: large-scale cargo, passengers and crew seeking lower costs, less time-sensitive passengers and crew, and possibly even ‘pleasure cruises’ in space. A current terrestrial comparison would be ocean travel versus air travel as most cargo uses the former, while passengers and crew use the latter. 2.9.3 Use cases In this section, we will lay out the key use cases for the FTS using government, corporate, and private customers. Visits As the Mars settlement grows, we would expect a subsegment of settlers to have demand for an occasional quick return visit to Earth Relatives of settlers might also want to visit their family on

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Mars. There could even be time-sensitive family travel that could not wait for a convenient orbital window, such as the need to attend a wedding, birth, or funeral, or visit a sick relative.

Table 22. Market segmentation. Chemical propulsion Fast transit propulsion system High volume and non-time sensitive Timeline sensitive cargo (such as cargo Goods transport key hardware replacement) Scientific and some commercial High-value cargo payloads Tourists High net-worth tourists Crew Essential specialists Human transport Settlers VIPs Scientists Business travel

Tourism The main tourism use case is that some Earth inhabitants, even without any family links to Mars, may want to visit Mars as an exotic tourism destination. Likely incremental spacecraft applications include specialized chartered, unscheduled tourist excursions (for ultra-high net worth individuals). Longer-term, other tourism use cases will likely emerge. Business travel Like Earth, despite the availability of ever more sophisticated videoconferencing, some business will continue to be conducted in person. Time-sensitive cargo We expect that most cargo will continue to travel to and from Mars on cheaper but much slower chemical propulsion-driven spacecraft. There will still be demand for the transport of time-sensitive goods on our fast transit system, ranging from urgent medical supplies to perishable luxury food items requested by ultra-high net worth individuals. Other opportunities Other potential business use cases or revenue opportunities, which we do not have time to discuss in further detail, could include: military uses such as rapid troop movement, scientific uses such as sending probes more quickly to distant celestial bodies much quicker; and settlement beyond Mars, such as on Saturn’s moon Titan, which would take seven years to reach from Earth using current propulsion systems (Wohlforth and Hendrix, 2016). 2.9.4 Sizing up the opportunity In the following section, we estimate the revenue for the first full year of commercial operations in 2055, by estimating both potential demand volume and pricing. Please note that the numbers given are highly illustrative and should be seen as midpoints of what are likely large ranges of potential outcomes. Demand volumes We project a global Martian population of 20,000 by 2055. We then assume a portion of this population would have demand and means for a return visit to Earth p.a. who would use fast transit as opposed to chemical propulsion. While we looked at, for example, Concorde market shares of the transatlantic long-haul market that was approximately 2%, this comparison was not ideal as Concorde’s improvement over regular flight times was less than three times that of the fast transit system’s twenty-fold improvement over chemical propulsion. A more suitable comparison would be the historical numbers of transatlantic air passengers versus transatlantic sea passengers when airlines started taking market share from ocean liners. While we could not find good data within the

International Space University 72 Team Project Final Report timeframe of this project, we would recommend this area as a subject for future research. If we assume that 10% of the Martian settlers would like to return to Earth p.a., and half of those would choose fast transit, then we would expect 1000 roundtrip seats—or 2000 single flight leg seats p.a.— in demand from this source, considering only Mars outbound demand. Assuming symmetric outbound demand from Earth, we would add another 2000 single flight seats, for a total of 4000 single flight seats demand from ‘visits.’ Our 50% market share assumption for fast transit is roughly based on anecdotal evidence that jet aircraft's market share of transatlantic travel was around 50% as soon as the first jet service was introduced, in 1958 (Taneja, 1971). We made a rough estimate of tourism demand by considering the known reservations and inquiries for suborbital flights by Virgin Galactic – which are, respectively, 603 and 2500 (US Securities and Exchange Commission, 2019), respectively for a total of 3103. We are only comparing a suborbital flight that provides only a few minutes of space experience with an on-the-ground visit to another planet. We found it plausible that a higher fraction of the approximately 20 million millionaires (MGMResearch, 2019) in the world may find value or interest in the latter experience and estimate an annual demand of 1000 roundtrip seats and 2000 flight one-way seats from this source. Since historically millionaires have sought exotic travel experiences, this tourism application might be one of the most important for the FTS. For business demand, we estimate another 2000 one-way seats. We considered evidence (Taneja, 1971) regarding the percentage of airline first and business class seats bought by companies rather than the passengers themselves. We determined this figure to be around 65%. We also assume that a good part of the demand from 'visits' is also paid for by companies (or other organizations, incl. governments and agencies), specifically we assume 3000 out of 4000 seats. We will assume another 2000 seats for other business travel demand, giving us a total of 5000 seats purchased by someone other than the actual passenger, or a ratio of close to the 65% mentioned above. This then results in an overall demand of approximately 8000 one-way seats p.a. as of 2055 (and growing from there) between Earth and Mars for passenger travel. For time-sensitive cargo, we assume that an incremental 20% of weight is available, similar to the ratio of cargo versus passenger space on terrestrial aircraft (Taneja, 1971) and used at a price that is essentially equivalent to the average price per weight charged to passengers. We simply gross up our passenger revenue estimate by 20% in order to account for cargo revenue). We did not estimate for revenues from scientific or military uses as this project may be funded by civil or military sources who then would benefit from the technology without having to pay for its continued use as seen in the following “funding” section. Pricing To estimate the price of crewed single tickets or cargo, we took two approaches. First, we estimated our cost base and, assuming a reasonable margin, arrived at a price. Specifically, we assumed US$70 thousand in propellant costs, US$145 thousand in licensing fees for the FTPS., and US$30 thousand for other operational costs such as flight operations, maintenance, or sales To arrive at a 30% operating margin, we required a ticket price of US$350 thousand, although this price would come down over time as the business scaled up and operating costs decreased. Second, we considered the relative pricing of FTPS versus CP. While estimating CP pricing in 2055 is very speculative, we can use Musk’s comments regarding potential one-way ticket prices to Mars as a starting point (2019), specifically that a ticket may cost less than US$500 thousand and perhaps even less than US$100 thousand. Given SpaceX plans to fly humans to Mars as early as the late 2020s (Dunn, 2019), we can assume that prices will have come down significantly by 2055 simply due to the increasing scale of operations. If we assume a US$70 thousand per one-way CP ticket to Mars, for example, then the FTS ticket price of US$350 thousand would be five times more

International Space University 73 Team Project Final Report expensive – similar to the differential between air freight and ocean freight on Earth under an average scenario (Freightos, 2019). Revenue Using our above assumptions about (8000 single leg seats / 80 single leg flights (given 100 seats per spacecraft), assuming a 100% load factor), we arrived at 2055 as the approximate first year of commercial operations with a revenue of US$2 billion as seen in the illustrative financials below. After 2055, revenue should grow roughly in line with the size of the Mars settlement, which we expect to grow at a rate in the teens through the end of the 2050s. 2.9.5 Fast Transit Spacecraft and propulsion system cost estimation Estimating the cost of research and development, production, and operation of spacecraft and propulsion systems is complex. Widely used methodologies to budget the costs of space programs, include: bottom-up costing techniques, analogy cost estimation, and parametric cost estimation methods (Kanipe, 2014). Bottom-up costing relies on detailed engineering analysis of each process involved in the project and produce the most accurate results, yet they are the most time consuming to implement. On the other hand, analogy estimation uses simple comparisons with analogous systems produced or programs implemented in the past, to derive cost estimates. Finally, parametric models aim to establish non-linear statistical relationships between cost drivers and indirect measures of cost or output, based on historical information on space project costs. We chose parametric techniques to estimate the costs of developing, producing, and operating a fast transit spacecraft and propulsion system for a 10-crew demonstrator spacecraft to enter in operations by the year 2050, and a 100-passenger spacecraft to operate from 2055 onwards providing commercial spaceflight operations. Parametric techniques, as compared to bottom-up and analogy costing methods, represent a good compromise between results’ accuracy and model simplicity. Our model (see Table 23and Table 24) derives the required number of man-year units (WYr) to develop, produce, and operate different space systems, based on their total mass (M/kg), system sensitivity parameters A and B estimated from historical data, and expansion coefficients to account for factors like the maturity of the technology (TCS-Transcostsystem, 2013). For the purposes of budgeting the FTS and ACFP, three types of costs were considered: (1) development costs (research and development); (2) production, and; (3) operation costs such as (testing and ground operations, etc). 2.9.6 Costs of 10-crew Fast Transit Spacecraft and propulsion system Table 20 presents the results of our cost estimation exercise. Developing, producing, and operating a 10-crew demonstrator FTS, along with a first-of-its-kind ACFP would cost approximately US$84 billion. About 90 percent of the total cost would emerge from the research and development stage, and US$45 billion alone to develop the ACFP, which is consistent with its current low TRL. The remaining US$32 billion would cover development of the spacecraft systems for the 10-crew demonstrator craft. According to the results of the model, production costs would add up to approximately US$5 billion, most due to expenses in spacecraft manufacturing.

International Space University 74 Team Project Final Report

Table 23. Cost calculations for a 10-crew demonstration spacecraft system for Fast Transit Spacecraft and propulsion system. Cost WYr A B M / kg US$B Spacecraft system 32.5 84,818 1,113 0.383 82,000 Propulsion* 45.9 29,932 277 0.48 17,250 Sub-total development costs 78.3 114,750 99,250 Spacecraft system 4.0 10,463 0.16 0.98 82,000 Propulsion system** 0.2 582 3.2 0.535 17,250 Sub-total production costs 4.2 11,045 Sub-Total operation costs 1.5 … Total costs 10-crew demonstrator and propulsion system 84.1 Source: own estimations of spacecraft and propulsion system parameters, based on parameters extracted from TCS-Transcostsystems (2013) Notes: * Uses a factor of 4 based on liquid propulsion development cost parameters. ** Based on production costs for cryo-engines. Operation costs are estimated with Space Shuttle operation costs, using 2x and 2.5x factor, yet subtracting flight costs.

Table 24. Cost calculations for a 100-crew spacecraft system for Fast Transit Spacecraft and propulsion system. Cost WYr A B M / kg / US$ B Spacecraft system 10.1 26,409 1,113 0.383 552,000 Propulsion* 17.1 11,161 277 0.48 115,000 Sub-total development costs** 27.2 37,570 667,000 Spacecraft system 26.0 67,799 0.16 0.98 552,000 Propulsion system*** 0.6 1,632 3.2 0.535 115,000 Sub-total production costs 26.6 69,430 Sub-total operation costs 2.5 … Total costs 100-crew spacecraft and propulsion system 56.3 Total program costs w/o contingencies 140.4 Contingencies (15%) 21.1 Total program costs 161.4 Source: own estimations of spacecraft and propulsion system parameters, based on parameters extracted from TCS-Transcostsystems (2013) Notes: * Uses a factor of 4 based on liquid propulsion development cost parameters. ** Scale-up factor is 0.15 of total R&D costs. *** Based on production costs for cryo-engines. Operation costs are estimated with Space Shuttle operation costs, using 2x and 2.5x factor, yet subtracting flight costs. 2.9.6 Costs of 10-crew Fast Transit Spacecraft and propulsion system Table 23 presents the results of our cost estimation exercise. Developing, producing, and operating a 10-crew demonstrator FTS, along with a first-of-its-kind ACFP would cost approximately US$84

International Space University 75 Team Project Final Report billion. About 90 percent of the total cost would emerge from the research and development stage, and US$45 billion alone to develop the ACFP, which is consistent with its current low TRL. The remaining US$32 billion would cover development of the spacecraft systems for the 10-crew demonstrator craft. According to the results of the model, production costs would add up to approximately US$5 billion, most due to expenses in spacecraft manufacturing. 2.9.7 Costs of 100-crew Fast Transit Spacecraft and propulsion system In the case of a 100-crew spacecraft development, we estimated its development, production and operation to cost US$56 billion. Development costs have been computed as 15% of the total greenfield development costs, assuming that only a fraction of that would be required to scale-up the demonstrator spacecraft. Following this logic, development costs would account for approximately 50 percent of the total costs. Production costs are significantly higher than the demonstrator 10-crew FTS, as the increment in total spacecraft mass is eightfold. Production costs for the propulsion system are estimated in the vicinity of US$0.6 billion. TOTAL PROGRAM COSTS AND CONCLUSION Total costs for both the demonstrator 10-crew spacecraft and the scaled-up version for commercial space operations would approximate US$160 billion, including contingencies (Table 21). About 45 percent of the total costs before contingencies can be attributed to the development of the propulsion system for both spacecrafts, and 33 percent to developing the spacecraft systems. Estimating the costs of a long and ambitious technology development program and demonstration mission is a very challenging task. Historically, space programs have typically been ridden with cost overruns and delays. For instance, the Space Shuttle Program costs were over their original estimate by a multiple of 10, and the cost of the ISS was 15 times its original budget (Denissov, Laforteza and Wynn, 2014). Despite the contingencies already included the total cost calculation, such risks need to be contemplated when budgeting a program of this nature. 2.9.7 Financing options Through our initial assessment of the cost of this mega-project, we forecast that it will likely require total investments of around US$200 billion and span a multi-year development period ending by 2050. The annual investment for Fast Transit to Mars would contribute an additional 44 percent to annual global government investments in space exploration, which totaled $14.6 billion in 2017 (Euroconsult, 2018). Because of the astronomical cost, the project will likely need public funding sourced from a single country such as the US, China, UAE, India, or by members of European Union. It will be used initially for research and subsequently for development of all new reusable systems and subsystems including entirely novel propulsion method(s), spacecraft, orbital stations, and space refueling stations. Due to the size of the project both in research and funding, there is no precedent on how to finance such a large project and historical multinational co-operations to fund the European Organization for Nuclear Research (CERN) and International Thermonuclear Experimental Reactor (ITER) do not provide a financing model that can be used as an example. The Fast Transit to Mars project would be five times larger than both CERN and ITER combined. As this would likely be the largest singular space project ever undertaken by any country or group of countries, it would require a close cooperation among all participating entities. Based on our calculations and best practices, we believe that it will require initial public funding, consisting of first round financing through series A funding, followed by series B and C funding in subsequent decades. By approximately the early to mid-2040s, the project will likely reach an advanced stage, allowing for additional funding through private public partnerships. Specifically, we assume that of the total

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US$200 billion, US$170 billion would be financed by public funding and US$30 billion by private funding. In comparison, the Space Shuttle program cost approximately US$196 billion (2011 dollars), the Apollo program US$144 billion (2017 dollars), the Manhattan project US$23 billion (2018 dollars), and the F35 fighter jet program US$1.5 trillion. Based on our estimates, by the year 2055 the project will start generating income, through transporting cargo, passengers, and research projects, between Earth, Mars, and potentially other destinations. 2.9.8 Execution timeline We believe that by the end of 2030 we will be able to test this new propulsion system on Earth. By the end of 2040, the project will launch autonomous non-crewed flights to Mars, and by the end of 2050 initial flight demonstration with a crewed spacecraft. The initial crew will have 10 members, but by 2055, we envision commercial operations shuttling 100 persons spacecraft between Earth and Mars.

Table 25. Execution timeline. 2020 2030 2040 2050 2055

Development R&D Propulsion In-space Demonstration Start of commercial timeline phase technology demonstration flight of crewed operations demonstration, phase, including spacecraft to Mars (100-person ground tests non-crewed flight (10-person spacecraft) to Mars spacecraft) Financing Seed Series A Series B Series C Series D Type financing (demonstration of (in space (minimum value (commercialization) early propulsion technology product and scaling business prototype) demonstration) demonstration) up Financing Public Public Public Public and private Public and private source

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2.9.9 Illustrative financials

Table 26. Illustrative headline financials. Flight leg seats 8000

Ticket price/flight leg seat 350,000

Total in Per seat Investment 2055

US$ 1m US$ 1k US$ 1bn ROI

Revenue passengers 2800 0.350 Public 161 1%

Revenue cargo 560 0.070 Private 34 3%

Revenue total 3360 0.420 Total 195

Propellant cost -560 -0.070

Other opex -240 -0.030

Licensing -1552 -0.194

Operating profit (OP) 1008 0.126 EV/OP multiple 34x

% operating margin 30%

Our illustrative financials show that both public - via the licensing fee - and private stakeholders could achieve acceptable returns. The enterprise value to operating profit multiple as of 2055 is also reasonable, considering the expected ongoing revenue and profit growth.

2.10 Legal implications 2.10.1 Introduction After overviewing the fast transit mission planning in the previous section, where we recommended the use of ACFP, we will now examine key aspects of international law underlying the journey. Today, international law stands in both international and national documents (ICJ, 1945), which are: 1. International treaties, 2. Custom and general principles of international law, 3. Non-binding international instruments, inter alia soft law, such as guidelines, recommendations, reports, declarations and opinions of the most qualified publicists, 4. Domestic laws and regulations.

International Space University 78 Team Project Final Report

There are five main treaties to which all states involved should abide: 1. The Treaty on Principles Governing the Activities of States in the Exploration and Use of Outer Space, including the Moon and Other Celestial Bodies (OST), 2. The Agreement on the Rescue of Astronauts, the Return of Astronauts, and the Return of Objects Launched into Outer Space (the Rescue Agreement or ARRA), 3. The Convention on International Liability for Damage Caused by Space Objects (the Liability Convention), 4. The Convention on Registration of Objects Launched into Outer Space (the Registration Convention), 5. The Agreement Governing the Activities of States on the Moon and Other Celestial Bodies (Moon Treaty or Moon Agreement). In this chapter, we examine how existing law, related to the peaceful exploration and use of outer space, applies to our mission scenario. We also analyze the use of ACFP through the existing international legal framework, currently governing outer space activities. 2.10.2 Legality of the Antimatter Catalyzed Fusion propulsion system After careful assessment of the ACFP, we concluded that the propulsion system does not breach the important principles broad scope of five space treaties. We made sure that the use of ACF ensures peaceful and secure use of outer space, and international cooperation and understanding (Article III of the OST). We also ensure that ACFP will not contaminate outer space and trigger adverse changes in the environment (Article IX of the OST). 2.10.3 Environmental Considerations “Throughout the ages, [hu]mankind has, for economic and other reasons, constantly interfered with nature. In the past, this was often done without consideration of the effects upon the environment. Owing to new scientific insights and to a growing awareness of the risks for [hu]mankind—for present and future generations—of pursuit of such interventions at an unconsidered and unabated pace, new norms and standards have been developed.” Gabčíkovo-Nagymaros Project case (1997) “In most areas—and definitely in the space sector—increased environmental awareness is an imperative without which any activity will eventually become impossible” Lotta Viikari (2010)

Launch and re-entry are the most dangerous stages of every space mission. Preparing a spacecraft for lift-off and overcoming aerodynamic pressure requires not only thorough preparation, but also a series of adequate risk assessments, including an environmental one. Due to possible complications during the mission’s launch phase, none of the potentially dangerous elements within the spacecraft should be activated on Earth during lift-off. They should also not contaminate or pose any danger to the planet’s environment during the re-entry phase. We prohibit the disposal in or contamination of the orbital environment with ionizing radiation waste that presents as free neutrons. To reduce risks of outer space pollution, we shall position our spacecraft in a safe orbit within the LEO region. The fast transit mission shall be in compliance with international space law (Article IX of the OST) and basic principles of international environmental law, like sustainable development, prevention, precaution, no harm, polluter pays, incorporated into space law through Article III of the OST.

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By 2050, we suggest that all state parties to the mission ratify corresponding safety policies to ensure efficient emergency readiness and response. We also call on the state parties involved to create a budget for scientific research on a safe emergency return of the mission. 2.10.4 Using space resource We aim to create fast interplanetary travel for humans, robotic missions, cargo, and scientific payloads. Our mission’s success also depends on the possibility of space resource mining, which could be important for refueling the spacecraft. Our spacecraft needs two elements, deuterium and krypton, which can be found on celestial bodies, the closest being the Moon. Their extraction and use will not only reduce the mission’s cost, but also keep our transportation services fast and efficient, allowing us to have regular flights in the future. We propose an array of rules for mining on the Moon and other celestial bodies:  According to Article VI of the OST, state actors are responsible for all activities of their governmental and non-governmental entities, if registered under a certain government. The use of space resources should be authorized by the state. To ensure the legal character of mining activities, a company, directly or indirectly involved in mining and/or overseeing it, should be based in, registered in, and licensed by a country where legislation allows such activities,  The authorization or licensing of a company is only possible after a submission of a clear feasibility plan, which will identify all foreseeable risks, as well as methods for avoiding, preventing and mitigating them,  Space resources should be used for peaceful purposes only,  A company’s presence on the Moon or any other celestial body can only be temporary and purely functional, as all territorial appropriation or sovereignty claims are prohibited,  Mining activities should be beneficial for all humankind, regardless of the economic and scientific development of any certain country. The benefits of using space resources should be shared among all states,  The mining activities of any country should not interfere with another state's ongoing space exploration or space resource activities,  Mining activities should not pose harmful risks to humans, the space environment, or cultural heritage, such as the footprints of Neil Armstrong left on the Moon, or other celestial bodies. 2.10.5 Cooperation “Yours will be a future where human beings have pushed farther into the universe, not just to visit but also to stay. To me, public diplomacy and (international) cooperation in space go together like peanut butter and jelly... When we go up to cislunar space, it’s going to give our international partners an opportunity to be with us, because no venture into deep space is going to be done by one nation... The future of space policy is built on international cooperation” Charles Bolden, a former Administrator of NASA, retired United States Marine Corps Major General, and former NASA astronaut

“States are free to determine all aspects of their participation in international cooperation in the exploration and use of outer space on an equitable and mutually acceptable basis” (UN, 1996) Space Benefits Declaration, para. 2

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Due to its high cost, the development of a constant acceleration propulsion technology is hardly possible without international cooperation. The most sustainable way to achieve a significant reduction of travel time to Mars by 2050 should be based on a common effort, reinforced by corresponding policies and legal mechanisms. The latter are diverse and flexible in form and substance, but we would like to highlight two of them specifically for our mission: A multilateral agreement or a set of agreements among participating states that are legally binding, legally non-binding, or the combination of both. Example: The Space Station Intergovernmental Agreement (IGA) The Space Station Intergovernmental Agreement originally brought together five space agencies and their respective International Space Station (ISS) programs to explore Earth’s orbital environment and assure humankind’s peaceful and permanent presence in outer space. It has three branches (Von Der Dunk and Brus, 2006): the IGA establishing a long-term international cooperative framework, the Memoranda of Understanding signed by the cooperating agencies of state parties involved, and further implementing arrangements—such as contracts and subcontracts— involving private actors. Multilateral coordination mechanism This mechanism has a more sophisticated and administrative-heavy structure, requiring more human resources to establish a dedicated international organization within the UN umbrella, or one of a regional or interregional character. Another possible method of cooperation includes creating a specialized fund by the United Nations Office of Outer Space Affairs (UNOOSA) to support the development and implementation of fast interplanetary travel projects. Each member state involved will financially contribute to the new UNOOSA fund. The sum of the contribution will be converted to units adjusted for different factors like the level of economic and technological development, population, and GDP. The larger the contribution of a Member State is to the system, the more privileges and chances it would get in terms of setting strategies, participating in the UN decision making process, or accessing high-level meetings. These units will be transferred to an international fund, financing selective space projects or long- term space programs. The fund will also supervise the redistribution of benefits from these space activities to member states on an equal basis.

Figure 66. International fund mechanism.

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2.11 Ethical implications

“The future belongs to those who believe in the beauty of their dreams.” - Eleanor Roosevelt A reason for humankind to aspire to extend our presence into space is because “one cannot remain in the cradle forever” (Tsiolkovski, 1911). Former NASA Director Michael Griffin has stated that we explore space not only because of economic, policy, military, and scientific rationales. (Griffin, 2007). However, in our view the main reason for us to go to space must be curiosity, “the fundamental human trait of wishing to expand our knowledge” (Cambridge, 2019). Our team’s curiosity has motivated our challenge of going to Mars and beyond faster than before. Exploration and expanding civilizations in other worlds have deep roots in human history. The exploratory nature of space is multi-faceted and the drivers for this exploration can include technology, science, adventure, geopolitics and inspiration. Our motivation to explore the unknown is a complex combination of factors, and Pyne (2006) argues that each theme can unfold differently. Historically, the movement of humans across land masses has enabled connection of different human societies, leading to knowledge exchange and to a mindset shift from ego-centrism to global awareness. This has sometimes been accompanied by technology and scientific advancement. However, the impact on both sides have not always been symmetric, with wealth accumulation, introduction of disease, religious conversion, slavery, and other factors leading to the rise and fall of civilizations. Mars, similarly to Antarctica, the deep oceans, and the Earth orbit, has no pre-existing human presence. This is a distinctly different scenario from many of the historical analogies of discovery, colonisation and settlement. Starting from an uninhabited planet gives us freedom and permission to create a society with new values and rules. However, the culture we establish on Mars will inevitably be influenced by a culture of settlement. Questions that need to be considered are: Will political conflicts on Earth be transported on other celestial bodies? Will we carry our prejudices with us? Will the society on Mars eventually diverge from Earth? As Pyne (2006) points out, the most plausible prognosis is that the future will resemble the past. Assuming a sustained habitation scenario on the surface of Mars by 2050, in the following years we expect that a Fast Transit Spacecraft fleet is likely to be shuttling around one hundred passengers per return trip between the Red Planet and Earth. With the journey taking merely days, this distances in our Solar system would be drastically reduced. Several historical analogies to this have been experienced in the past. A remarkable example is the development of aviation, where a fast and economical way to travel large distances revolutionized many aspects of society. Initially, aircraft only provided a travel option for the wealthy, associating the idea of traveling with luxury and prestige. Decades later, thanks to economies of scale, improvements in technology, and falling ticket costs, it was possible for more people to fly around the globe in hours. This led to effects including increased and easier interconnectedness between people, facilitated efficient economic transactions for people and products, and enabled previously inaccessible recreation possibilities. The consequences of aviation were not all beneficial, in addition to environmental degradation, it has changed the nature of warfare and provided society with a means of destruction on a scale never imagined. However, this mode of fast transport has given humanity a sense of mastery and control of their destiny as members of a much larger and united world (Hudson and Pettifer, 1979). Further revolutions in travel speed in space will inevitably bring new opportunities and challenges to humanity. Initially, with high costs of the system just emerging from research and development, only governments and wealthy individuals will be able to access this journey. Like aviation development, as cost and technical barriers are overcome, more and more people will be able to make the trip to

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Mars. Astronauts will eventually be replaced with passengers and crew, with a minimum level of selection and training required. The existing human habitation on Mars will be eventually connected to Earth by frequent and regular commercial spacecraft, just as the continents and regions on Earth have been linked by commercial aircraft. This will result in greater interplanetarization, analogously to globalization, with flows of people, goods and ideas between Earth and Mars, and ultimately beyond. With a fast transit technology, there would be no need to stop at Mars. There are multiple destinations in our solar system. These may include visiting the moons of Jupiter, or a quick mining missions to an asteroid, or sending scientific instruments to the outer reaches of our solar system to help study the mysteries of the universe. There are a range of ethical concerns surrounding human exploration to Mars. These include both positive and negative consequences to our home planet and the universe. Questions which need to be addressed include: How long do we stay? Can everyone go? What if we pollute Mars? Will going to Mars to help Earth? What if we find life on Mars? An additional ethical concern is the potential for cross contamination between Mars and Earth. We have aimed for an inclusive approach for our Fast Transit mission. From our analysis, we have identified key features related to linguistic integration, gender inequality, racism, and extreme exclusivity (Drake, 2018). We have sought to address these issues through inclusion, ensuring that our mission was ethically sound through close consultation with our thematic teams. In addition, to ensure an ethical environment for the mission, we would establish an “Ethical Committee on Fast Transit.” We believe it could serve as a forum where relevant experts could discuss ethical issues with regards to our missions and greatly contribute to the success of fast transit.

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PART 3: THE FUTURE

THE FUTURE

3.1 Introduction - Beyond Mars, finally going to the stars

In this final section of the report, we discuss the future with fast transit technology, how this can be achieved and what the potential impacts of this travel will be on our society. In section 3.2, we will investigate the development roadmap for nuclear fusion technologies for interplanetary travel. By 2020, humanity needs relevant policies at the international level to support topics such as space debris, space tourism, and the use of space resources. By 2030, we will require:  Development of advanced Antimatter Catalyzed Fusion Propulsion system.  A breakthrough for Magnetic Inertial confinement fusion (MICF)-propulsion system research and development to reach TRL4/5 (breadboard model tested).  An operational Lunar Gateway.  The first human mission to Mars orbit.  A resolution for space debris.  More accessible space tourism. By 2040, our propulsion system should achieve:  TRL6/7 (operation in the space environment), with several human performance, engine restart, and shut-down tests.  Technology allowing in-orbit assembly, optimization, and cost reduction.  Enough efficient, clean, and safe energy to power several cities of around 80,000 people. By 2050, our propulsion system should:  Achieve maturity and reach TRL8/9, allowing us to launch a crew of 10 people to Mars. By 2060, we:  Shall launch up to 100 people per trip to Mars once per month. In section 3.3, we will analyze the impacts of a more accessible space travel to Mars on human society, and how the development of a fast-transit technology will benefit our exploration of the solar system in general. We will present a table listing travel time ranges assuming a 0.2 g

International Space University 84 Team Project Final Report acceleration. We will also use our fast transit propulsion system to help scientific research in areas like gravitational microlensing, solar science at the heliopause, planetary science, radiation sensors, human survivability, and gravitational waves. In section 3.4, we will analyze a potential propulsion system in both non-relativistic and relativistic settings. We will then compare the efficiency of propulsion systems of different Isp. Finally, with a propulsion system capable of providing 1 g acceleration, reaching 99% of the speed of light, we will calculate the time needed to reach the nearest star system, Alpha Centauri.

3.2 Roadmap to 2050 - Future Foresight

The first nuclear fission generator was developed in 1950 as an early prototype reactor followed by the 2nd generation in the 1960s mainly used for commercial power reactors. At the same time, fusion technology started to emerge into the R&D sector where the 1970s saw the birth of an early preliminary fusion reactor design. In the 1990s the 3rd generation advanced fission reactors were developed. Around the same timeframe, the preliminary tritium experiment in England saw the development of the world’s first controlled release of fusion power which ultimately led to the first successful production of helium plasma. The reactors were then enhanced in 2010 towards generation 3+ to achieve more economic efficiency. By 2040, we foresee the development of a fusion compact reactor that will not consume space and able to be fit on large vehicle but be capable of powering a city of 80,000 people (Future Outlook 100 Global Trends for 2050, 2017). This propulsion system will be disruptive in different areas, such as transportation, energy, and finally outer space and interplanetary travel. The below timelines provide future foresight and trend analysis which influences the mission plan of this report: 2020: TRL2 – Research of Propulsion Technology (Application) In this phase, further research must be conducted towards a proof of concept for the propulsion technology. Operations will move beyond LEO space stations and towards a lunar gateway (NASA, 2019). This technology has been tested on in-orbit servicing specifically on active debris removal (Surrey University, 2018). Thanks to the development of propulsion technology, private space tourism will most likely grow New terms to collectively call non-astronaut space traveler like “participants” reflects the trend (FAA, 2019). The trend has an impact on law. As an example, UN COPUOS has included space resources utilization as an official agenda item to be discussed under the legal subcommittee. Many national laws are now moving towards creating legal frameworks for utilization of space resources, in addition to national policies, like the UAE Space Policy (UAE, 2016), and regional strategies like the European Space Agency (ESA) Space Resource Strategy (ESA, 2018). 2030: TRL4 – Development of Technology The US National Ignition Facility (NIF) and the International Thermonuclear Experimental Reactor (ITER), estimated that the breakthrough for Inertial Confinement Fusion, which is known as ICF, will start with the miniaturization process, improvements in laser and superconducting technologies. These progresses in fusion technologies will also have implications in antimatter catalyzed fusion and related propulsion technologies. It is expected that the progress of the propulsion system R&D is going to reach TRL4 in both MICF and ACNF systems. Accomplishment of TRL4 means validation of components in the laboratory environment.

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The lunar Gateway, which is NASA’s deep-space outpost, will become operational in 2030’s. Eventually, such a plan will trigger drastic enhancement of in-situ space resource utilization, as well as changes in the international legal framework. The UN COPOUS will most likely issue the legally binding principles on the use of in-situ space resources and voluntary guidelines for space mining. Potential guidelines may discuss bringing of resources from Mars to Earth. Consequently, these changes in international legal frameworks will create a safe environment for national government to actively engage in space resources utilization business. Global Exploration Roadmap issued by the International Space Exploration Coordination Group, which is known as ISECG states that the first human mission to Mars will be in operation in 2030. SpaceX, which is led by Elon Musk, also announced plans for a cargo mission to Mars as well as a crewed mission to the moon. This will develop several experiments and increase the knowledge base related to human performance in space (ISECG, 2019). By 2030’s, the amount of space debris would be at a very high level, which will require the mitigation of space debris mandatory. Member states will have to remove debris as a prerequisite to conducting a launch, increasing demand on in-orbit servicing for active debris removal or refueling services to lengthen the lifetimes of in-orbit satellites. Additionally, In-orbit servicing will become fully operational and backed-up by the lunar Gateway. Space tourism will also become easier and more accessible. Hopefully, people will become more passionate about space travel. It may be possible that suborbital flights will replace aviation services. Humanity will get used to traveling faster. 2040: TRL6/7 - Technology Demonstration and Uncrewed Mission As announced by Lockheed Martin, the 4th generation propulsion system will be highly cost effective and safe. Also, this system will produce only minimal waste (Ministry of Cabinet Affairs and the Future, 2017). We expect that the propulsion system to reach TRL 6/7 when a ground testing and space demonstration mission is completed. This demo mission will aim to test human performance in various situations such as high accelerations and maneuvering. Also, the engine restart and shut- down will also be tested. In 2040, advancement of spacesuits as well as exoskeletons should provide in onboard autonomous crew life support systems and extreme weather conditions on Mars. Space-built spacecraft will also appear. This means that we will potentially in-space assembly, optimization, cost reduction, and testing and operation of the spacecraft. Together with the changes in 2030’s made in international legal framework, national government will actively engage in space mining. It will become more common and easier than in the past to bring materials back to Earth. Accordingly, legal frameworks will change to the direction which will cover topics beyond in-situ utilization. 2050: TRL 8/9 – Launch and Operations of the Mission In 2050, we expect our propulsion system to be mature and reach TRL 8/9. In this era, the first crewed space mission, which is Fast Transit to Mars will launch with 10 crew members. The success of this mission will open the door for humankind to access space beyond Mars. Space tourism on Mars will emerge in the value chain of the space economy. Untapped demand will be integrated to serve the first “spaceline” transportation service to Mars. Increasing demand on space tourism will positively trigger the enhancement of spacesuits. Scientific research of exo-planets will advance with probes from the edges of the solar system and un-crewed exploration missions to exo-planets. 2060 and beyond: TRL9 – Technology in Full Operations The propulsion technology will be fully operational and standardized. Spacecraft will now transfer up to 100 people per trip. The first Mars habitation will have been built with plans for other planets and moons in the solar system. Crewed missions to exo-planets will become possible. Other Trends Impacting the Transitional Environment of Interplanetary Travel

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There are more trends to consider in the future. Below the map describes subjects of interplanetary travel research.

Figure 67. Future Foresight mapping.

3.3 A vision of the future

Exploration of civilization has deep roots in human history. The nature of space has a variety of drivers, including: exploration, technology, science, adventure, geopolitics, and inspiration. Our motivation to explore the unknown is a complex combination of those factors, and Pyne argues that each of these can unfold differently (Pyne, 2006). Human built vehicles have been on Mars for several decades and it will not be long before the first human sets foot on the red planet. Any presence, may it be robots or human, carries a risk of contaminating the planet, which we must keep in mind. Assuming a sustained habitation scenario on the surface of Mars by 2050, the Fast Transit Spacecraft fleet will be shuttling ten passengers per return trip between Earth and Mars. With this journey taking only a few days, the travel time for the distances in our Solar system could be drastically reduced. Several historical analogies to this were experienced in the past. One example is the development of aviation, where a fast and ultimately economical means of traveling large distances revolutionized aspects of global society. Initially, air-travel only provided an option for the wealthy, associating the idea of traveling with luxury and prestige. Decades later, thanks to economies of scale, improvements in technology, and falling ticket costs, flying around the globe was accessible to those in the middle-class. This led to increased and easier interconnectedness between people, facilitating efficient economic transactions for people and products, and enabling previously inaccessible recreation possibilities. The consequences of aviation were not always beneficial. In addition to environmental pollution, it changed the nature of warfare, and gave humanity a means of destruction on an unprecedented scale, for example aerial bombings in WW II. At the same time, this means of transport interlinked humanity on a larger scale (Hudson, 1979). Even though space-travel will likely always be more costly than air-travel, if not only due to the large distances covered, once the cost and technical barriers are overcome, space travel, for example a trip to Mars, may become possible for more people. Passengers would need a minimum level of

International Space University 87 Team Project Final Report selection and training, but less than the training required of current astronauts. The existing human habitation on Mars will be eventually connected to Earth by frequent and regular commercial spacelines, just as the continents and regions on Earth have been linked by commercial airlines. This will result in greater “interplanetarization” with flows of people, goods, and ideas between Earth, Mars, and ultimately beyond. With a fast transit technology, there are a huge number of fascinating destinations in our solar system that are beyond Mars. With only a few weeks to visit a moon of Jupiter, or a quick mining mission to an asteroid, or a scientific mission to the outer reaches of our solar system, researching the mysteries of the universe will be far more accessible. 3.3.1 Scientific opportunities A constantly accelerating spacecraft allows for faster travel than conventional methods which work with short accelerating boosts and orbital maneuvers such as Hohmann transfers or gravity assist. The high velocities which can be achieved over a period of days or weeks enable reaching remote destinations in several weeks instead of several years. With a constant acceleration propulsion technology humans may be able to settle on Mars as a second home. This will lead to changes in research, society, and utilization of space resources. Space agencies and science organizations could achieve faster progress towards new discoveries, our human understanding of life, of our solar system, and of our universe to a new level. Furthermore, space business will have the opportunity to become achievable with real and sustained operations in fields such as transportation sector and space mining.

Figure 68. Graphic representing the solar system and its objects. The trajectories to selected destinations are highlighted with white arrows, with travel time in days (adapted from: https://www.universetoday.com/142531/an-orbit-map-of-the-solar-system/).

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Table 27: Mission profiles for different destinations. Approximate travel Distance from Earth time (days) at Mission profile (million km) continuous 0.2g Common Human Closest Furthest Closest Furthest destinations Settle Mining exploration Science approach approach approach approach Planets

Mercury X 77 222 5 8

Mars X X X 59 401 4 10

Venus X X 40 261 3 8

Jupiter X 629 928 13 16

- Europa X X 629 928 13 16

- Io X X 629 928 13 16

- Ganymede X X 629 928 13 16

- Callisto X X 629 928 13 16

Saturn X 1261 1500 19 20

- Titan X 1261 1500 19 20

Uranus X 2570 2856 26 28

Neptune X 4310 4390 34 35

Pluto X X 4550 7250 35 44

- Charon X X 4550 7250 35 44

Other objects

Ultima Thule X 6500 42

Ceres X X 260 560 8 12

Kuiper Belt X 4500 35

Main asteroid belt X X 120 6

Science missions

Study of atmosphere X 2000 23 0 of exoplanets Solar science at X 12000 15000 57 64 heliopause 3.3.2 Science Missions Since the potential for discovery is significant, we would like to describe the following three science missions, as proposed by Dr Jim Green (NASA) during a presentation to our project team in July 2019 during ISU SSP19 in Strasbourg, France. Gravitational Microlensing In recent years, there has been considerable progress in the study of extrasolar planets, through new tools such as the Kepler Space Telescope (Borucki, W.J., et al. 2010). One of the most common ways to detect exoplanets is the transit method, where an exoplanet transiting in front of its star is detected by measuring the reduction in brightness of the star. Using this or any other existing methods, astronomers have difficulty determining with any accuracy the atmospheric composition of exoplanets and cannot clearly establish the presence of life supporting environment or direct proofs of life. An alternative method consists in placing a probe aligned between the exoplanet and our Sun. A lensing effect occurs when a large portion of light from the exoplanet is deflected and concentrated

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(Turyshev, 2017). Positioning a probe at a given location can help collect enough light to support new discoveries in the field, which could have important consequences for humanity and our understanding of our place in the universe (see Figure 69 below).

Figure 69. Diagram showing how gravitational microlensing can be used to evaluate the atmosphere of an exoplanet. Fast transit will enable easier positioning of the telescope at a different location in space (not on Earth as shown in the image). (source: https://teara.govt.nz/en/diagram/8008/gravitational-microlensing).

Solar Science at the Heliopause Scientific studies on the Sun are important to improving our understanding of the solar system. Solar winds emanating from the sun, creates a surrounding bubble in interstellar space known as the heliosphere. The edge of the heliosphere is known as the heliopause. What little we know about this region is from the Voyager space probes, which after travelling around 40 years have encountered the heliopause in 2012 and 2018 respectively. To measure the evolution of the heliosphere, it is necessary to place several probes in specific points around the heliopause for an extended period. Probes delivered by fast transit could help improve our knowledge of the heliosphere and may be the only way to perform such a study. Planetary Science The solar system is composed of planets and many smaller celestial bodies. Our understanding of the solar system can be significantly improved by sending out probes. Our fast transit propulsion system can reach destinations far from the Earth and perform exploration, sample collection, experimentation, and preparation for future human settlements. To this day, space probes have been sent to all the planets and many other celestial bodies, but they have rarely entered orbit, or landed on a surface. Doing so will allow for more precise and long duration measurements. The New Horizons probe reached Pluto in nine years, representing a relatively short time considering the distance involved. A set of fast transit probes could help explore solar system bodies with significantly more detail by slowing down, being captured in orbit, and possibly landing. All of this would occur in a very limited time-frame.

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3.3.3 Additional Studies Survival of Human Organs in Space A fundamental prerequisite for human long-term exploration of the universe is crew survivability and health. This includes medical procedures commonly performed on Earth, such as surgical interventions and transplants. Further research is required regarding the creation and implementation of artificial organs, the survivability of tissues and organs in deep space and on celestial bodies, the effects of radiation and gravity variation on the human body, the use of medical equipment for space applications, and more. Sensors for Radiation During space missions, high-energy radiation might arrive from the environment and from the propulsion system. It is fundamental to measure the radiation dose absorbed by the crew during the mission, in order to create countermeasures to prevent overexposure. For this reason, further development of radiation sensors, shields, and countermeasures is needed. Cosmology - Gravitational Waves In the future, the Laser Interferometer Space Antenna (LISA) mission will be further extending its capabilities. This represents the first step in developing great scale space-time measurements in orbit. The interferometers might be also mounted on different spacecraft using the fast transit propulsion system, allowing more accurate measurements thanks to the greater relative distance. This would support the research regarding the origin and development of our universe.

3.4 Approaching the speed of light

In the previous section we considered many of the interesting destinations and potential scientific missions we could consider a 0.2 g fast-transit system. Here we will consider where such technologies could ultimately bring us, beyond our solar system into interstellar space. The greatest obstruction to interstellar travel is the massive distance between stars. We usually use the unit light year, representing the distance light travels in a year through a vacuum, to denote the distance to a specific star. Rockets that can reach speeds approaching the speed of light would make crossing interstellar distances far more practical. In this section, we present a mathematical analysis of the rocket equation in both non-relativistic and relativistic settings. In the end, we use our results to calculate the time needed to reach the closest star system to ours, Alpha Centauri. To begin, we revisit the non-relativistic Tsiolkovsky rocket equation (К. Ціолковскій, 1903):

푚0 푚0 Δ푣 = 푣푒 ln = 퐼푠푝 ⋅ 𝑔0 ln 푚1 푚1 where: 훥푣 is the change of velocity of the vehicle (assuming no external forces)

푚0 is the initial total mass before burning

푚1 is the final mass after using the propellant that generated the 훥푣

푣푒 = 퐼푠푝 ∗ 𝑔0 is the velocity of the exhaust material

푣푒 퐼푠푝 = is the specific impulse [s] 푔0 2 𝑔0 = 9.81 m/푠 is the standard gravitational acceleration on Earth’s surface

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As rockets reaching about 30% the speed of light this equation becomes insufficient to predict the speed of a rocket and a modified ‘relativistic rocket equation’ must be utilized. The equation is as follows:

퐼푠푝 ⋅ 𝑔0 푚0 Δ푣 = 푐 ⋅ tanh ( ⋅ ln ) 푐 푚1 7 Comparing the two rocket equations above, after defining 퐼푠푝 = 3 ⋅ 10 s; 1 < 푚푎푠푠 푟푎푡𝑖표 < 5, we see both as a single variable function of the mass ratio:

Figure 70. Non-relativistic rocket equation (y1) vs Relativistic rocket equation (y2).

At the beginning, when the mass ratio is close to one, the change in speed that can be provided is negligible compared to the speed of light. Both equations have similar results for values of mass ratio close to one. As the amount of fuel burned increases, the disparity between the non-relativistic (y1) and the relativistic (y2) functions becomes significant and grows as the mass ratio increases. Consistent with the theory of special relativity, nothing can travel at the speed of light or faster.

Figure 71. Relativistic rocket equation varying Isp and the mass ratio.

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In case of considering the 퐼푠푝 as a variable and the relativistic rocket equation as a two variables Δ푣 function, we can plot the graph against the mass ratio and 퐼 : 푐 푠푝

Figure 72. Relativistic rocket equation for specific Isp.

푚 As 0 increases, as seen in the graph, the velocity of the spacecraft increases rapidly according to 푚1 the 푙표𝑔 function. Independently from the 퐼푠푝, the asymptotic nature of the functions leads major mass ratio increases for minor 훥푣/푐 improvement. In addition to this, the velocity of an object influences its momentum. The faster it travels, the more momentum it can possess. This represents a challenge when approaching the speed of light as the thrust would remain constant while the effect of the thrust force to accelerate will decrease. It is highly unlikely for us to carry human at a speed that is near the speed of light. 3.4.1 Travel time calculation example As an example, let us estimate the time needed to reach the nearest star system, Alpha Centauri based on a 0.2 g: If we accelerated at 0.2g to 99% of the speed of light, then remained at 99% of the speed of light for most of the journey, and finally decelerated at 0.2g to 0 speed: we need to accelerate to 2.97 ⋅ m 108 , which means we need 4.8 ⋅ 2 = 9.6 years to accelerate and decelerate; we would have s travelled 2.25 ⋅ 1016푚. We will travel at 99% of the speed of light for 1.882 ⋅ 1016푚 for about 1.99 years. To reach Alpha Centauri, we would need nearly 11.5 years. Conclusion To travel at a speed near the speed of light, we would need an engine with a practical 퐼푠푝 able to reach near 107s.

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3.5 Summary – A pledge for decision makers

Fifty years after humans set foot on the Earth Moon, space is experiencing a renaissance, as people worldwide are awed by the promise of advancing rocket technology opening up the Earth orbit in a broad way, returning humanity to the Moon in 2024, and even taking us to Mars by 2040. Current development roadmaps imply that there will soon be a permanent human presence on the Moon and Mars, which may become sizable before 2050. At the same time, research is ongoing globally on advanced technologies that could further open up space, including fusion and antimatter. Given this historic set of circumstances, we believe it is once more time to set ourselves a hard challenge: we propose to develop a spacecraft capable of taking humans safely from Earth to Mars within one week, and to do so by the year 2050. In the decade after this we envisage regular connection flights between Earth and Mars. This spacecraft would likely use antimatter-catalyzed fusion propulsion and accelerate at acceleration between 0.2 and 1.0g. It would operate between the planets’ orbits, while more conventional vehicles (such as chemically propelled) would provide shuttle services to and from the planets’ surfaces. There are several technological challenges to achieve our goal of fast transit. The most important one is the lack of affordable production methods of antimatter which might negatively affect the technology readiness level of the propulsion system. In addition, there could be unforeseen challenges associated with scaling up the system from the already demonstrated cubesat level to the size required for interplanetary human transit. These challenges are to be tackled with innovative solutions and collective wisdom of international cooperation. Most probably international cooperation might finance the development of the system, through its first crewed demonstration flight in 2050 with a total expected expense of approximately US$160 billion. Antimatter propellant might be provided by institutions like CERN or its spin-offs, an existing example of international cooperation, and we hope this will be possible at scale by the year 2040. On the political front, we aim to conclude a Mars cooperation agreement by 2040. Once the capability of the new system is proven, around 2045-50, we propose to license the technology to private investors, in the hope that they will enable regular services between Earth and Mars, ensuring a permanent close link between humans despite the distance. In short, just as the race to the Moon captivated and inspired all of humankind and provided tremendous spin-off benefits on all levels to society, we have confidence that this new program will have the same effects. With this potential reward in mind, history once again calls on us to be bold and to unite in the pursuit of a dream. We hope that this report will be useful, inspirational and will help accelerate efforts to make fast travel to Mars and beyond real in our lifespan. The team would be happy to help provide further information. Please contact us if more details are required.

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International Space University 102 Team Project Final Report

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International Space University 103 Fast Transit: mars & beyond

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