Space Exploration and Development Systems (SEEDS) 2014

SpacE Exploration and Development Systems (SEEDS) 2014

“Missions and Expeditions to Mars” Orpheus

Executive Summary

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List of authors: Crescenzio Ruben Xavier AMENDOLA Portia BOWMAN Samuel BROCKSOPP Andrea D’OTTAVIO Alex GEE Samuel R. F. KENNEDY Antonio MAGARIELLO Adrian MORA BOLUDA Ignacio REY Alex ROSENBAUM Joachim STRENGE Aurthur Vimalachandran THOMAS JAYACHANDRAN

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ABSTRACT This six crew mission called Orpheus has been designed in order to fulfil some main objectives of space exploration: scientific advancements, technological progress, public outreach and international cooperation. This paper investigates the possibility of exploring the Martian system by using a manned spacecraft, the Crew Interplanetary Vehicle, (CIV) and a cargo vehicle, the Mars Automated Transfer Vehicle (MATV). The main payloads, carried by the high efficiency solar electric MATV, are a two-passenger spacecraft landing on Phobos; an orbital laboratory and a rover network for deployment to the surface of Mars. In order to cope with the constraints imposed for a human mission to deep space, the feasibility study has been performed using a human-centred design approach. The main output of the mission is the preliminary design of the CIV. Furthermore, the main parameters of the MATV as well as the orbital laboratory and the Phobos Lander were estimated. The manned spacecraft is designed to depart from LEO in 2036 using chemical propulsion. Once in Mars proximity, the main manoeuvres will be performed using nuclear thermal propulsion and a bi-propellant chemical system will perform the minor manoeuvres. Orpheus is designed to be a short stay mission in proximity of Mars, including two landings on Phobos; the aim being the return of scientifically valuable samples for analysis on Earth. Finally, the CIV will return to Earth resulting in a total mission duration of 602 days, while the MATV and the rover network will continue to perform scientific investigations on Mars. Significant conclusions of the project are approaches to deal with the some of the most critical challenges encountered at present such as radiation shielding, the boil-off problem, nuclear propulsion systems and long term human spaceflight. The challenges of this project were addressed by working in rigorous collaboration over six months. The result is an innovative concept of a manned Mars mission, leading to outputs which can help to pave the way to the future of manned space exploration.

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ACKNOWLEDGEMENTS We would like to thank all the people who helped and supported us all along these months to finally present our work in ESTEC. Firstly, we would like to thank all the people who organised the SEEDS project in Torino and particularly Mr. Enrico Beruto, Pr. Gianfranco Chiocchia, Pr. Nicole Viola, Pr. Ernesto Vallerani, and all the experts in Thales Alenia Space with special thanks to Mr. Claudio Ferro and Mr. Eugenio Gargioli. Secondly, all the professors and experts in Toulouse have been extremely supportive of all the time and attention given over to putting this project together. We thank particularly Pr. Benedicte Escudier, Ms. Lizy- Destrez, M. Emmanuel Zenou, and M. Alain Lacombe. The help of all the experts from CNES, ALTEC Thales Alenia Space, and Airbus Defence and Space have also been highly appreciated. Last but not least, many thanks to the great team of professors in Leicester for their dedication and involvement in this great adventure. We want to give special recognition to Dr. Richard Ambrosi, and Dr. Nigel Bannister who have been of huge support during our time in Leicester, and also Dr. Ian Hutchinson, Dr. Hugo Williams and Dr. John Bridges.

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FOREWORD & PRESENTATION OF THE AUTHORS Three parallel SEEDS (SpacE Exploration Development Systems) classes are recruited every year at Politecnico di Torino, ISAE Toulouse and University of Leicester, each with a maximum of 15 students. The first 5-6 months are spent at the recruiting Institution and mainly consist of class lectures and exercises providing the general foundations of the various disciplines related to Space Exploration. In the following 6-7 months the students (12 in the 2014 edition) extensively develop the SEEDS Project Work (PW) under the guide of experienced senior tutors. Thanks to its extension the Project Work of SEEDS is one of its main qualifying features. It is divided into three phases, each one performed in a different SEEDS site and dedicated to a special aspect of space exploration. A general topic is identified every year (examples of past topics are the preliminary design of a lunar outpost or a mission to an asteroid). Each PW phase is dedicated to a specific aspect of the selected topic, also taking into account the locally available expertise. The three phases are hosted in a temporal sequence by universities (in Toulouse and Leicester), industries (Thales Alenia Space and ALTEC in Turin) and centres of the European towns associated with SEEDS. During the entirety of the PW the students from Torino, Toulouse and Leicester are grouped together and work in a cross-national team. The PW itself is an advanced and ambitious activity, comparable to a “Phase A Study” and intended to produce scientific reports to be diffused worldwide in the space community. On more than one occasion partial results have been presented by the students to international conferences. For example, after the work described in this project the students intend to present their work at the IAC 2015 in Jerusalem. This Executive Summary followed by a discussion in front of international experts at the ESTEC facilities concludes the course.

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Crescenzio Ruben Xavier Amendola was born in Napoli (Italy) in 1990. He obtained a Bachelor degree in Aerospace Engineering in Italy (Universita’ di Napoli - Federico II) and took part in the Erasmus project in Sevilla (Spain). After the Bachelor Degree, he did an internship at ‘Costruzioni Aeronautiche Tecnam’ (Napoli - Italy), where he was in charge of drawing up the airworthiness manual for a light aircraft. In 2012 Crescenzio moved to France to enrol in the M.Sc. in ‘Aerospace Avionics and echanics’ at ‘Institut Sup rieur de l’A ronauti ue et de l’Espace’ of Toulouse There, he specialized in Aeronautical and Space Structures and he undertook several space related studies. He is both a basketball player and coach, and his main hobbies are traveling and photography.

In order to deepen his knowledge in space systems engineering, he joined the SpacE Exploration and Development Systems (SEEDS) International Master in March 2014: the 6-months project is a ‘’Preliminary Phase A study’’ focused on the ‘’Exploration of ars from its Proximity” All along the mission design, Crescenzio has covered several roles: mission analyst, head of ‘‘Design, echanics and Structure’’ working group and member of Science team, mainly involved in the habitability of Mars and Phobos. He was also in charge of the presentation in all three phases.

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Portia Bowman was born in Kent (UK) in 1992. She completed her undergraduate studies in 2013 with a BSc in Physics with Astrophysics, with a main project working on the analysis of data from an instrument on ENVISAT. She is currently the UKSEDS (UK Students for the Exploration and Development of Space) Careers Officer, and enjoys star-gazing, swimming and outdoor activities.

Portia went on to study on the SEEDS MSc program at the University of Leicester, to graduate in 2014. During the international project phase she filled the roles of Project Manager and Systems Engineer. As Project Manager she organised the work breakdown and resource allocation, ensured the project ran to schedule and that the outputs were delivered on time. She also worked on the detailed Environmental Control and Life Support Systems (ECLSS) design and worked as part of the Science team with a focus on science at Phobos. ◊◊◊

Samuel Ryan Brocksopp was born in Derby, England in 1991. He has a background in physics, having graduated from the University of Leicester in 2013, and is currently the sponsorship officer for UKSEDS (UK Students for the Exploration and Development of Space). His technical experiences involve 4 months within the Rolls Royce Civil Nuclear, analysing construction logistics for heavy pressure vessels, and Submarine divisions. More broadly he has experience teaching STEM subjects to students (aged 11-18) through UKSEDS outreach and the Extreme Physics program.

In October 2013 he joined the Erasmus variant of the SEEDS, SpacE Exploration and Development Systems, MSc program with the intention to enter the growing space sector. During the 6 month international feasibility study, into a manned Mars proximity mission, he has been a lead member of the communication and nuclear electric propulsion subsystems. ◊◊◊

Andrea D’Ottavio was born in Turin (Italy) in 1987 In December 2012, he obtained a M.S.c in Aerospace and Astronautical Engineering in Italy (Politecnico di Torino). From February 2013, Andrea became Assistant Researcher at the Department of Aerospace Engineering of Politecnico di Torino, starting to collaborate with Alenia Aermacchi on the SMAT-F2 Project, as a UAV System and Safety Design Engineer. In September 2013, Andrea started a new collaboration as System Design Engineer with Thales-Alenia Space in the AMALIA Project. In the same period he also obtained the European qualification of Superior Technical for Aircraft Maintenance.

In November 2013 he joined in the Second Level Specialised Master SEEDS, SpacE Exploration and Development Systems an International Master through three different countries. In the first phase Andrea was enrolled in the Mission Analysis Team as a head of Propulsion Systems while for the entire Italian phase, he assumed the role of SEEDS Project Manager, continuing to work in the Mission Analysis Team and in the development of the Nuclear Propulsion Systems. Finally he covered also the role of head of Electrical Power System. He is an athlete of CUS Torino and a volunteer of the Civil and Military Rescue Team of the Italian Red Cross. His main hobbies are music, mountain, traveling and photography.

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Alex Gee was born in Staffordshire (UK) in 1991. He began his undergraduate degree in Physics at the University of Leicester in 2010 where he took part in a final year project focussing on an automated Mars Sample Return mission and upon graduation in 2013 he began the SEEDS Masters course. During the three phases of the SEEDS international study he held positions as Chief Systems Engineer, Chief Thermal Engineer and Head of Science Investigations. During the project he was particularly interested in maintaining budgets and systems interfaces, thermal control of large Methane tanks, and scientific instrumentation specifically Raman Spectroscopy and the use of MOLE like drilling mechanisms to measure the attenuation of radiation by Martian regolith. During university he held positions as finance officer and promotional manager as part of the Ultimate Frisbee club of which he was a keen player. He is a dedicated skier and snowboarder who enjoys travelling, practicing his French and German language skills and taking in local culture. ◊◊◊

Antonio Magariello was born in Stigliano (Italy) in 1985. He completed with honours and on time the MSc studies in Mechanical Engineering at Politecnico di Torino with a thesis developed in collaboration with Universitat Politècnica de Catalunya. In the years after graduation, he was able to work in different fields, enriching his knowledge, dealing with fluid dynamics, construction and mechanical design. During this period, Antonio was engaged in different countries: Spain, Italy and UAE. Due to his passion for Space activities, in 2013 he was enrolled in the Master SEEDS (Space Exploration and Development System). During the three international phases of the master (Toulouse, Turin and Leicester) he contributed for the Mission Analysis in defining the mission concepts, trajectories and landing sites; then in thermal control he was working in thermal loops, radiators and propellant tank boil off; and finally he investigated the nuclear electric propulsion including low thrust trajectories. Antonio had the possibility to interact with professors and experts from industries for achieving these results. ◊◊◊

Samuel Robert Frederick Kennedy was born in Peterborough (UK) in 1992. In 2010 he undertook a Bachelors of Science degree in Physics with Astrophysics at the University of Leicester. His main project was a Mars Sample Return concept study for Airbus Defence and Space. Subsequently, he began his Masters of Science in Space Exploration and Development Systems (SEEDS) at the University of Leicester. During the international SEEDS project he was part of the mission analysis team, specifically the science requirements and electric propulsion systems, head of Environmental Control and Life Support Systems and a member of the science team, developing objectives and payloads. He holds the position of Membership Officer for the UK Students for the Exploration and Development of Space (UKSEDS) and during university he was a course representative. He was also a member of the Astronomy and Rocketry Society, Physics Society, and a part of the interdepartmental Physics football team.

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Adrian Mora Boluda was born in Alicante (Spain) in 1989. He received the diploma of Industrial Engineer from “Escuela Politecnica Superior de Elche” (Spain) in 2013 His studies were completed successfully including two years as student representant, one full academic year as an Erasmus student in Fachhochschule Stralsund (Germany), six months of internship as Android Developer at IT&IS Siglo XXI (Alicante) and a final award as a 5-star student given by his home university in Elche.

Since November 2013 he is enrolled in the sixth edition of the Master SEEDS (SpacE Exploration and Development Systems) at Politecnico di Torino (Italy), which ends in September 2014 with a final presentation at ESTEC (Netherlands). During this specialization master, he has successfully attended lessons and passed exams before starting a 6-month internship in three different countries (Italy, France and United Kingdom). In this internship he has been working for a manned mission to Mars, being part of a systems engineering team. He has worked for this project as a Communications Subsystem Engineer, as a Mission Analyst and as a part of the science team for the last phase in Leicester. ◊◊◊

Ignacio Rey was born in Mendoza (Argentina) in 1981. He studied Computer Science in Argentina and participated in an international program at Politecnico di Torino obtaining a MSc in Software Engineering. He has worked for 6 years as a software developer in Italy, USA and Argentina. His hobbies are bicycling and astronomy, for the latter he took courses at Buenos Aires Planetarium during 2012 and 2013.

He enrolled in the Master SEEDS (SpacE Exploration and Development Systems) in 2013 due to his enormous interest and fascination for space. During the 3 phases of the project he worked in systems engineering, design of the Phobos lander and astronaut locomotion, attitude control systems, radiation, and electric power systems. ◊◊◊

Alex Rosenbaum was born in Paris XII (France) in 1989. He received a M.Sc. in Applied Mechanics and a M.Rc. in Materials Engineering from Ecole Centrale de Nantes in November 2013 after the validation of his final year internship at Rolls Royce Canada. To fullfil his desire for Space Exploration he decided to join in September 2013 an Advanced Master in Space Systems Engineering at the Institut Sup rieur de l’A ronauti ue et de l’Espace in Toulouse where he was able to gain a lot of knowledge concerning all the subsystems of launchers & payloads.

In March 2014 he joined the SpacE Exploration and Development Systems (SEEDS) program for the definition and evaluation of the Step 1 of the program: “Exploration of ars from its Proximity” where he worked successively as a project manager while being part of the system engineer team, a member of the design team, and in the last phase a member of the science team where he was

SEEDS Executive Summary 09/2014 Page 12 mainly involved in In Situ Resources Utilization (ISRU) considerations. He also dealt with RAMS engineering issues all along the project. ◊◊◊

Joachim Heiko Strenge was born in Leer (Germany) in 1987. He obtained his Bachelor of Engineering at the University of Applied Science in Bremen in Aeronautical Engineering after having gathered experience at the German Aerospace Centre as well as in the space industry abroad. In his final year of Bachelor studies, he followed a summer school on human spaceflight in Salon-de-Provence (France). Always drawn to explore other cultures, he pursued his Master of Science at the Institut Supérieur de l'Aéronautique et de l'Espace in Toulouse and graduated in 2014. Joachim is highly enthusiastic about space exploration and an active member of several space related associations, as for example the Community of Ariane Cities and EUROAVIA Toulouse, of which he was the Vice-President during the business year 2013/2014. In his spare time he enjoys rock climbing and debating.

During the SEEDS course, Joachim participated in various areas of research, for example as project head for the development of the mission concept, as system engineer by maintaining the interfaces and budgets of different working groups according to ECSS standards and by organizing a common concurrent design process. Finally he assisted the definition and implementation of scientific requirements to increase the scientific output of the mission. ◊◊◊

Aurthur Vimalachandran Thomas Jayachandran in short Aurthur was born in Chennai,India (1987) from where he graduated in Electronics and Instrumentation Engineering (Anna University). Then he moved to Italy for his aster’s degree in Industrial Engineering and Management (Politecnico Di Torino) during this period he became an Entrepreneur and ventured into many fields such as Golden Infotech Inc-Web contents (India) and TURSO Inc-Bio Fuels (Indo-Poland). Later he worked as Junior Area Manager for Ex-soviet and South-east Asia for Florim ceramics (Italy). Apart from this he was also an active member of many student organisations such as AIESEC and Balon Mundial.

He was enticed by the world of space and decided to do this Master SEEDS which helped him to work in multicultural atmosphere with students having different educational competencies. After having gained experience from various field from technical to management, Aurthur worked in wide field of this project such as Communication, ECLSS, Telemetry, Fault tolerant Failure Analysis, Orbital Assembly, Time line to LEO and Documentations too. In the final phase he was the System Engineer refining the final budgets with help of CDF specialist and produced the cost estimation of the mission.

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TABLE OF CONTENTS

Abstract ...... 5

Acknowledgements ...... 7

Foreword & Presentation of the Authors ...... 9

Table of contents ...... 14

Acronyms ...... 17

1 Introduction ...... 21

2 Mission Overview ...... 22

2.1 Brief ...... 22

2.2 Mission Statement ...... 22

2.3 Mission Objectives ...... 22

2.4 Project Management ...... 23

3 Mission Analysis ...... 25

3.1 Introduction ...... 25

3.2 Development of Interplanetary Trajectories ...... 26

3.3 Definition of the Mission Concept in the Vicinity of Mars ...... 36

3.4 Propulsion ...... 48

3.5 Overall Mission Parameters ...... 63

3.6 Planetary Protection ...... 64

3.7 Critical issues ...... 65

3.8 Conclusion ...... 65

3.9 References ...... 66

4 Science ...... 68

4.1 Objectives ...... 68

4.2 Requirements ...... 70

4.3 Instrument Platforms, Techniques and Selection ...... 70

4.4 Payload Configurations ...... 71

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4.5 Rover Mass Calculations ...... 74

4.6 Additional Scientific Studies ...... 75

4.7 Critical Issues ...... 80

4.8 References ...... 81

5 Systems engineering ...... 83

5.1 Introduction ...... 83

5.2 Methodological Approach of the CIV Study ...... 83

5.3 Results of the CIV Study ...... 86

5.4 Mars Automated Transfer Vehicle Study ...... 89

5.5 Phobos Lander Study ...... 91

5.6 References: ...... 94

6 Spacecraft Design ...... 95

6.1 Design, Mechanics and Structures ...... 95

6.2 Environment Control and Life Support Systems (ECLSS) ...... 101

6.3 Radiation ...... 110

6.4 On orbit Assembly Strategy ...... 112

6.5 Electrical Power Systems ...... 116

6.6 Thermal Control ...... 120

6.7 Communications ...... 126

6.8 telemetry, tracking and command (TTC)...... 131

6.9 Attitude and orbit control system ...... 135

6.10 Critical Issues ...... 138

6.11 References ...... 139

7 Reliability, Availability, Maintanability and safety (RAMS) ...... 143

7.1 Objectives ...... 143

7.2 Reliability ...... 143

7.3 Availability ...... 143

7.4 Maintainability ...... 144

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7.5 Safety ...... 144

7.6 References ...... 145

8 Cost Consideration ...... 146

8.1 Introduction ...... 146

8.2 Cost Analysis ...... 146

8.3 Cost Benefit Analysis: ...... 148

8.4 Conclusion ...... 150

8.5 References ...... 151

9 conclusion...... 152

9.1 Nuclear Electrical Propulsion ...... 152

9.2 References ...... 159

10 Appendix...... 160

10.1 Appendix – Mission Analysis - CIV Mission Phases ...... 160

10.2 Appendix – Mission analysis - Employed Mars Orbits of the Orpheus Mission ...... 168

10.3 Appendix – Science Requirement Matrix ...... 170

10.4 Appendix - System Engineering: SIQ ...... 176

10.5 Appendix – Design Mechanics and Structures – Tanks Trade-Off ...... 177

10.6 Appendix - ECLS ...... 178

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ACRONYMS

AOCS Attitude and Orbit Control System

AU Astronomical Unit

BEE Best Engineering Estimate

BER Bit Error Rate

BOL Beginning Of Life

CH4 Methane

CIV (CM+SM+PM) Crewed Interplanetary Vehicle

CM Crew Module

CMG Control Momentum Gyroscope

CP Chemical Propulsion

CPS Chemical Propulsion System

CRV Crew Re-entry Vehicle

DSN Deep Space Network

ECLSS Environmental Control and Life Support System

ECSS European Cooperation for Space Standardization

EMC Electromagnetic Compatibility

EOL End Of Life

EPS Electrical Power Subsystem

ESTRACK European Space Tracking

GCR Galactic Cosmic rays

HEDM High Energy Density Material

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IMLEO Initial Mass to Low Earth Orbit

IDSN Indian Deep Space Network

ISP Specific Impulse

ISR In-Situ Resources

ISRU In-Situ Resource Utilization

ISS International Space Station

KE Kinetic Energy

LAB Laboratory Module

LEO Low Earth Orbit

LOX Liquid Oxygen

LRPS Liquid Rocket Propulsion System

MATV Mars Automated Transfer Vehicle

ML Mars Lander

MRO Mars Reconnaissance Orbiter

NEO Near Earth Orbit

NEP Nuclear Electric Propulsion

NERVA Nuclear Engine for Rocket Vehicle Applications

NPS Nuclear Propulsion System

NPT Nuclear Propulsion Technology

NRPS Nuclear Rocket Propulsion System

NTPS Nuclear Thermal Propulsion System

NTPT Nuclear Thermal Propulsion Technology

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NTR Nuclear Thermal Rocket

NTRPS Nuclear Thermal Rocket Propulsion System

O2 Oxygen

OBDH On Board Data Handling

OTV Orbital Transfer Vehicle

PBR Particle Bed Reactor

PE Potential Energy

PhL Phobos Lander

PM Propulsive Module

PS Pressurisation System

RDV Rendezvous

S/C Spacecraft

SEEDS Master Course in SpacE Exploration and Development Systems

SEP Space Exploration Program

SLS Space Launch Systems

SM Service Module

SPE Solar particle Events

SRC Sample Return Capsule

SRM Science Requirement Matrix

TC Telecommand

TCS Thermal Control System

TEI Trans-Earth Injection

TM Telemetry

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TOF Time of Flight

TTC Telemetry Tracking & Command

UNOOSA United Nation Office for Outer Space Affairs

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1 INTRODUCTION SEEDS puts major emphasis on a large Project to be sequentially performed through three successive internships in companies and centres of three countries. The SEEDS project in 2014 consisted of a 2 week Preliminary Phase (Torino), and 3, 7 week main phases (Toulouse, Torino and Leicester, in that order). These cities are the home of the 3 participating academic institutions, Institut Supérieur de l'Aéronautique et de l'Espace(ISAE, Toulouse), the Politecnico di Torino and the University of Leicester.

The SEEDS Sc project 2014 covers the first step in a five year study “A Journey from the Earth to Planet ars”

 Step One- Exploration of Mars from its Proximity.

 Step Two- Exploration of the Mars Surface.

 Step Three- Exploitation of Mars Resources ISRU.

 Step Four- Development of Mars Permanent Outpost.

 Step Five- Development of Mars Independent Base.

SEEDS aims to safeguard European knowledge and skills in Space Exploration in the perspective of a renewed emphasis on it in US and EU. Indeed, it is believed that Space Exploration is of prime importance when it comes to technology innovation. Within the next few years, space programs are expected to improve our knowledge of solar energy power; cryogenics; and robotics that are expected to offer great improvements in health care; energy and the environment; everyday technology; and many other areas.

To deliver this Executive Summary it has been decided to use a system engineering approach, not discipline focused as in the academic tradition. That is to say, individual disciplines are not deepened beyond the minimum necessary level.

This report has been written in the most coherent way possible following this approach. That is to say it begins by describing the customer constraints then the top-level requirements, in order to finally introduce the proper mission analysis; the science that will be performed in Mars vicinity; and finally the spacecraft design. It is worth noticing that the ECSS standards have been used where applicable in each section.

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2 MISSION OVERVIEW

2.1 BRIEF The brief given for the 2014 SEEDS course was to design a manned mission to the Mars Proximity. The following points were treated as customer requirements by the students:  To send humans to the Mars-Phobos System.  To lay the framework for future SEEDS missions in the following years.  To land an unmanned payload on the surface of Mars that can be controlled by astronauts in Mars orbit, and must have use after the humans leave.  To land a payload on the surface of Phobos. It was noted early on that the students were to consider the mission as having full international cooperation and that cost should not be a limiting factor. In the preliminary phase, a detailed trade-off was performed where it was decided that the mission would include landing humans on the surface of Phobos. The main reasons for this decision was for useful science to be performed by astronauts rather than having to be automated, for technology demonstration purposes and to increase public interest in space exploration. The following points have been identified as the reasons for why it would be interesting to bring humans to Mars and Phobos. Mars was chosen as it is the most accessible and habitable planet for humans. These are as follows: 1. Science Advancements: A better understanding of the composition of Phobos could suggest its origin. This also could lead to more information about the origin of Mars and Earth. Gathering information related to the composition of Mars lays the foundation for future in-situ resource utilization if possible, as well as continuing the search for biomarkers which draws public interest. A manned mission will allow for research to be performed into the psychological and physiological effects of long term spaceflight on humans. 2. Technological Advancements: Sample preparation, return and validation technology can be advanced. Human spaceflight technology can be improved to set a foundation for future expeditions to Mars which could lead to the potential of setting up human colonies on other planets. In-situ resource utilization techniques can be developed and demonstrated. 3. Public Outreach and International Cooperation: A manned mission will reactivate the wider public interest which could lead to future funding opportunities. International Cooperation is important for any large future endeavours. It is from these imperatives that the mission statement was derived.

2.2 MISSION STATEMENT To perform human exploration of Mars from its proximity; to execute a manned landing on Phobos, including a sample return; to develop and validate techniques to lay the framework for future manned exploration and exploitation of Mars. The mission statement sums up the goals of the mission and has been referred to at all stages of the mission to prioritise work and solve problems by relating the solutions to how well they fit the mission statement.

2.3 MISSION OBJECTIVES The main mission objectives are derived from the mission statement and identify the most important aspects of the mission to achieve the overall goal. The mission objectives are as follows: 1. Sustain human crew in orbit around Mars and return safely. 2. Study the effects on humans of long term spaceflight.

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3. Land humans on Phobos surface. 4. Perform experiments on Phobos surface samples. 5. Search for resources available on Mars and Phobos. 6. Return samples from Phobos surface. 7. Search for biomarkers on Phobos and Mars. 8. Technology development for future missions. 9. Identify possible future landing sites on Mars. 10. Gather material for public outreach. 11. Maintain stable communications with Earth. 12. Land scientific payloads on Mars

2.4 PROJECT MANAGEMENT The project was split into 3 main phases and a short introductory phase. For each phase a different project manager was elected. The overall structuring of the project was planned in Phase 1, with detailed analysis at the start of the following phases. It is important to note that each location had specialist advice available on different topics, so these topics (communication, habitable space design and science) were planned to be worked on in the relevant location so as to utilize the expertise. Phase 1 also included an analysis of the strengths and weaknesses of the group, with the aid of a Belbin test. From this test the group was identified as strongest in the team working ability, but weakest in innovation and the ability to complete tasks. Using this analysis a SWOT (Strength, Weaknesses, Opportunities and Threats) analysis was completed to identify which areas were particularly important to manage. The project manager had the task of creating work breakdown structure, an organisation breakdown structure and a detailed project timeline. As well as this, it was the duty of the project manager to organise weekly group update meetings where work was discussed and shared with the whole group to ensure that all working groups would be aware of how their work would affect other groups. The project manager was to make sure that the work was divided evenly, to deal with interpersonal issues and ensure work satisfaction to maintain a good working environment. Table 2-1 below details each phase.

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Table 2-1: Project Management Phase Preliminary Phase Phase 1 Phase 2 Phase 3 (Location) (Torino) (Toulouse) (Torino) (Leicester) Duration 2 Weeks 7 Weeks 7 Weeks 7 Weeks Working Top-Level Mission Mission Analysis (5) Mission Analysis Science Payload Groups (no. of Definition (12) (Trajectory (2) (Continued Design (7) people in each definition, Propulsion Study) Electric Power group) Propulsion Study) Habitation area Systems (2) Communications design (Structures Nuclear Electric (3) (3) and Propulsion Study(2) Environmental Systems Control and Life Developing project Engineering(4) Support Systems deliverables (12) (ECLSS)(3)) (Presentation and Summary) Thermal Control (2) Systems Engineering(1) Attitude and Orbit Control Systems (AOCS) (1) Systems Engineering (1) Key outputs of Reasons to plan the Detailed Trajectory Detailed Mass and Detailed Mass and the Phase Mission, Mission for the mission, Power budgets of Power budgets for Statement, Mission including departure each system each system Objectives and return dates worked on worked on resulting and delta-v resulting from a from a detailed required. Choice of detailed design. design. An propulsion system. alternative feasibly

Initial mass and study for propulsion power budgets for system. Executive all subsystems. Summary. ESTEC Presentation

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3 MISSION ANALYSIS

3.1 INTRODUCTION The feasibility for a human mission to Mars with the present technology is on the edge of the human possibility. Hence, the mission concept was designed to decrease the gap between current unfeasibility and future feasibility of a manned mission to Mars. A mission to Mars leads to major mass budgets due to the high ΔV re uirements; furthermore the level of safety implies additional mass for redundancy Hence, the mission concept of Orpheus was designed to reduce the initial spacecraft masses by sharing the payload between two. The smallest of both is unmanned and its payload mass has been maximized in order to send payloads not linked with human activity or human safety to Mars: its name is the Mars Automatic Transfer Vehicle (MATV). The manned spacecraft has all the subsystems necessary for sustaining humans for the entire mission duration, the required habitable volume and all the necessary spacecraft`s propellant. The manned spacecraft is called the Crew Interplanetary Vehicle (CIV).

3.1.1 Spacecraft of the Orpheus Mission – the CIV and the MATV The MATV has the purpose of transporting all payloads which are not directly needed for the survival of the crew in vicinity of Mars. Its proposed payload consists mainly of the following elements:  The scientific landing packages to be deployed on the surface of Mars  The Phobos Lander (PhL)  The Laboratory (LAB) The latter two systems are used by the crew once in Mars proximity in order to land on Phobos and to analyse regolith samples. Hence a docking between the two spacecraft once both are in Mars` orbit is re uired This docking manoeuvre will be very critical and a first outside of Earth’s sphere of influence The advantage of using an unmanned transport spacecraft is the use of a high-efficient, low-thrust propulsion system; in particular solar-electric propulsion. On the CIV however, due to human presence, strong transfer time constraints are imposed. Hence, the CIV will have short trajectories with a less efficient propulsion system (lower specific impulse) leading to a penalisation in terms of propellant mass according to the Tsiolkovsky equation. The docking manoeuvre in Mars orbit will be necessary for the exchange of modules between the two spacecraft. In the case of failure of docking the scientific mission fails, as the crew cannot access the PhL; this is an acceptable risk due to fault tolerance. However, the risk of loss of crew in this case is not acceptable. This risk imposes the requirement to equip the CIV with all the necessary propellant for the entire manned mission, rather than the ATV carrying fuel for the CIV’s return trip The prior choice this guarantees failure tolerance. This decision is less mass efficient as the CIV propellant is carried by the CIV itself which has a low efficiency propulsion system. Figure 3-1 shows an artist’s impression of the CIV after the Trans-Mars-Injection (TMI). It can be seen that in this stage of the mission, the methane tanks make up the largest proportion of the spacecraft mass.

Figure 3-1: An artist’s impression of the CIV

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3.1.2 MATV Configuration As previously stated, the MATV main packages are the LAB, PhL, and Mars science packages; the Mars package includes three rovers. After arrival at Mars, the MATV will deploy the landing packages and perform remote sensing of Mars until the arrival of the CIV in Mars proximity. At this point the two spacecraft will dock in order to transfer the payloads for the human mission; the PhL and the LAB to the CIV. Afterwards, the MATV will return to the science orbit to continue its science mission. At the end of the nominal mission phases, the PhL and Lab will not be carried back to Earth, but they will be tugged to and left in a stable graveyard orbit by the CIV. This is necessary since Phobos and Mars have been defined as category 4 with regards to planetary protection, as stated by the Committee on Space Research (2008).

Figure 3-2: An artist’s impression of the MATV

3.1.3 Short or Long Stay Mission Due to the arrangement of the planets, travel to and from Mars is favourable in terms of ΔV(or delta-v) every two years. For a typical Mars mission which includes a return, there are two concepts which can be identified. These are either performing a long stay mission or performing a short stay mission; short stays have a higher ΔV requirement. This choice impacts many aspects of the mission. A long stay mission generally leads to a lower ΔV re uirement since the departure and the return take place in favourable moments of the interplanetary ephemeris. The time in Mars proximity is typically more than one (Earth) year, and the overall mission lasts between two and three years. The short stay generally requires a higher ΔV, the time spent in ars’ vicinity is typically no more than 60 days and the overall mission will last less than two years. The short stay mission implies larger propellant consumption; however, a long stay mission has the major drawback of the crew spending an extended time in space. Effects of microgravity, GCR (Galactic Cosmic Radiation), and psychological isolation will be worsened during this longer duration. For this reason, using a formal trade-off, Orpheus is short stay mission.

3.2 DEVELOPMENT OF INTERPLANETARY TRAJECTORIES This chapter will introduce the development of the trajectory in a linear way. First, imperative trade-offs for the selection of the interplanetary trajectory are introduced; after that, using a set of self-developed tools the actual trajectory will be determined. It is important to understand that the true nature of the interplanetary trajectory design is an iterative one; the choice of the interplanetary trajectory changes the input values of the trade-offs and may change the results, which may then have an impact on the constraints of the interplanetary trajectory. In this report only the final results will be focused on.

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Figure 3-3: Overview of the Mission Concept

3.2.1 Decision Making Process Before being able to numerically analyse trajectories, it is necessary to make basic assumptions concerning the system architecture. The following trade-offs were performed for this purpose:  The choice of utilising high-thrust or low-thrust trajectories for both CIV and MATV  The utilisation of aerocapturing for respectively CIV and MATV when arriving in the vicinity of Mars Trade-offs were performed, according to the recommendation of the European Cooperation for Space Standardization ECSS-E-ST-10C, by carefully defining relevant and non-redundant items and using scientific approaches to weight and score the different items. Sensitivity of weighting and scores was taken into account during all trade-off processes.

3.2.2 Low Thrust or High Thrust Trajectory For both spacecraft, the very basic type of interplanetary transfer orbit - impulse lambert transfer orbit or low- thrust trajectory had to be determined. While this initial trade off only considers only these two options, a more detailed analysis will be performed in Section 3.4. The following items were considered in the trade-off table:  Initial mass to LEO  Psychological constraints  Physiological constraints

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 Risk of mission failure  Past heritage  Future heritage  Design complexity The results of the thrust level trade-off can be found in Table 3-1. Table 3-1: Thrust Level Trade-off Results Solution Low Thrust Lambert Transfer

CIV 4.0 6.0 MATV 5.7 4.3

It can be seen that the CIV will utilise a classical impulse transfer, while the MATV uses a higher efficiency low-thrust propulsion system.

3.2.3 Aerocapturing Trade-off Aerocapturing is the use of atmospheric drag instead of propulsive thrust for the capture manoeuvre of spacecraft. This method has been strongly discussed during the last decades as a potential technique to reduce the ΔV requirements of a given mission considerably. Presently, this technique has never been used in interplanetary flight; because of two main reasons: The Mars capture burn requires a strong reduction in velocity of 1.4 km/s – 2.5km/s. Taking into account the very short time spent in the atmosphere of a planet on hyperbolic trajectories, this leads to high accelerations and heat loads on the spacecraft; this requires a very specific design including a major heat shield. Furthermore, the physiological and psychological loads of such a manoeuvre on the crew would be enormous as indicated by Larson (2005). One of the current limits of interplanetary trajectories is the available orbit determination accuracy. Successful aerocapture requires atmospheric entry of a planet at a very specific angle and altitude due to the exponential density function of the gas layers. Entering the atmosphere slightly too high or too low will lead to a loss of the mission and possibly spacecraft, either by leaving the Martian system on a hyperbolic orbit or by disintegrating in the high atmosphere NASA (2012). As explained earlier, all masses were calculated as the IMLEO and ranked anti-proportional to this value. To be able to estimate the mass of the spacecraft in LEO it was necessary to estimate the mass of a heat shield capable of protecting the spacecraft during aerocapture. Table 3-2: Heat Shield Mass Estimation S/C Payload mass [t] Mass estimate [t]

CIV 400 123 MATV 60 18

The mass of the systems was approximated as a linear scaling of the mass of the heat shield of the currently biggest system to have entered the atmosphere of Mars - the EDL capsule of the NASA Mars Science Laboratory mission as published by the NASA (2012). Scaling up the wet mass of the different spacecraft of the Orpheus mission leads to estimate masses as seen in Table 3-2. The relationship between heat shield mass and payload mass is commonly not linear and the mission profile of MSL and the Orpheus spacecraft is very different. Due to this the estimations have to be treated with care and were solely obtained for the purpose of the given trade-off.

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The following items were considered in the trade-off:  Initial mass to LEO  Risk of mission failure  Design complexity  Unfamiliarity of Mars environment  Past heritage  Future heritage  Human factors The results of the aerocapture trade-off can be found in Table 3-3. Table 3-3: Aerocapture Trade-off Results Solution Aerocapture Powered capture

CIV 3.4 6.6 MATV 4.3 5.7

3.2.4 The Interplanetary Trajectory of the CIV The selection of the launch window is a complex process and was performed using the following steps:  Using a set of assumptions, the different feasible launch windows for the Mars approach and return trip were determined using modified pork chop plots.  An Excel spreadsheet was developed to combine the different pork chop plots into a single matrix summarizing the overall mission propulsion system requirements.  The different launch windows were compared and the final one was selected using a formal trade- off.

3.2.4.1 Development of a Pork Chop Plot The pork chop plot is a common tool to compare the advantages and drawbacks of injecting the spacecraft into interplanetary trajectories at a certain time. For this, the positions of both planets are taken into account. According to the Lambert problem as given in Jordon (1964), for each set of positions as well as the proposed transfer time, a unique interplanetary transfer ellipse can be derived. Comparison of the velocity vectors of this transfer trajectory with the known velocity vectors of the planets at this point allows estimations of the C3 to be made. The values estimated are the required C3 value for injection into this orbit from Earth and the C3 value encountered when arriving at the destination planets sphere of influence. Two kinds of trajectories have been considered:  Common direct transfer orbits  Transfer trajectories utilizing a swing-by manoeuvre at Venus The pork chop plot was created using the software SCILAB using the function set Celestlab, which was created and published by the French Space Agency CNES. It provides both the ephemeris of the planets as well as a function for solving the Lambert problem. A script in SCILAB was developed to calculate both arrival and encounter C3 as a function of the departure date and the time of flight.

‖ ‖ The assumptions given in Table 3-4 were inserted as parameters in the script.

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Table 3-4: Parameters of the Pork Chop Plot script Parameter Value Unit Comment

Gravitational constant 132712440018 km³/s² Brown (1998) of the Sun 01.01.2033 - Departure Window ESA (2004) 31.12.2049 Time Resolution 1 day

Maximum C3 Low C3 requirement as manoeuvre is 50 km²/s² Earth and Mars performed using propulsion system Departure Maximum C3 Mars Low C3 requirement as manoeuvre is 50 km²/s² Encounter performed using propulsion system Maximum C3 Earth High C3 requirement as manoeuvre is 500 km²/s² Encounter performed using passive re-entry

An example of such a pork chop plot for a single launch window is given Figure 3-4.

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Figure 3-4: Example of the Pork Chop Plot for a Single Launch Window in 2035 For verification of the script, the results were compared with a second script, which was provided by CNES. The results were found to be equal up to the range of the numerical accuracy of SCIBLAB. Furthermore, a second script was developed to analyse the possibility of utilising a Venus swing-by manoeuvre. For this, the loops concerning departure date and time of flight were implemented as described above. However, within the inner loop, a third loop was developed, which divides the overall transfer time which consists of the transfer time from the departure planet to Venus and the transfer time from Venus to the destination planet. For each combination of departure date and time of flight, a simple algorithm selects the Venus swing-by trajectory with the minimum requirement of departure C3 as long as the constraints given in Table 3-5 are met.

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Table 3-5: Parameter of the Pork Chop Plot Script for Venus fly-by Parameter Value Unit Comment

Gravitational constant of 132,712,440,018 km³/s² Brown (1998) the Sun Departure Window 01.01.2033 ESA (2004) 31.12.2049 Time Resolution 1 day Maximum C3 50 km²/s² Low C3 requirement as manoeuvre is performed using propulsion system Earth and Mars Departure Maximum C3 Mars 50 km²/s² Low C3 requirement as manoeuvre is Encounter performed using propulsion system Maximum C3 Earth 500 km²/s² High C3 requirement as manoeuvre is Encounter performed using passive re-entry Maximum change of C3 of 127.7 km²/s² Niehoff (1966) the S/C at Venus

The results are a set of additional pork chop plots, which suggest launch windows in periods, in which direct transfer is not available or economic.

3.2.4.2 Entire Mission Trajectory Optimization Process The pork chop(s) obtained from Celestlab gave the information of the departure and encounter C3 for the interplanetary trajectories for a period from 2032 up to 2049. Simulations have been carried out for the following four scenarios, each resulting in a departure and an encounter matrix:  Earth-Mars direct Lambert transfer  Earth-Mars considering Venus swing-by  Mars-Earth direct Lambert transfer C3  Mars-Earth considering Venus swing-by These simulations lead to a set of 8 pork chop matrices, which would lead to an extensive selection process of possible launch windows. Summing up the C3 values directly does not represent the requirements of the propulsion system well; it is necessary to estimate the ΔV for each manoeuvre In order to estimate the ΔV it is necessary to define the starting orbit for each manoeuvre and the speed in the local reference frame. For the departure from Earth has been considered an initial parking orbit in LEO 400 km from the surface. By approximating the parking orbit as circular it was possible to derive, for each Earth departure C3, the corresponding ΔV by subtracting the e uations of the circular orbital velocity and the hyperbolic velocity at pericentre according to Brown (1998).

√ √

A similar calculation has been used for the estimation of the re uired ΔV approaching ars, in this case a capture Orbit need to be defined. As the first approximation this has been considered an orbit which has a pericentre altitude of 298 km and an apogee which equals the semi-major axis of Phobos. The calculation of the re uired ars departure ΔV considers the departure from ars with the graveyard orbit as departure orbit, as will be introduced in Chapter 3.3.2.

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By implementing the three calculations in Excel, the required delta-v for all the scenarios described above were estimated. The next step was the combination of relevant pork chop plots in order to derive a common mission profile. For this, the Mars encounter and Earth departure were easily summed up, as this operation is always possible, and gave a result of the re uired combined ΔV for Earth departure and Mars encounter. At this point there are four available matrices; the Venus swing-by for both Earth-Mars and Mars-Earth trajectories is a single additional possibility to the direct trajectory. Hence, those tables have been combined in order to obtain only two tables: Earth-Mars (always including Mars Encounter delta-v and the possibility of Venus swing-by) and Mars-Earth (including the swing-by possibility). In case direct approach and Venus swing-by are possible for the same window, the lower value of the overall ΔV has been taken into account It is noteworthy that this operation may enclose the information about nature of the trajectory (it might be direct or a Venus swing-by), however the launch window parameters make it easy to derive the nature after finding a suitable launch window. Having reduced the number of matrices to be considered to two, the next step is the combination between those two tables optimising the minimum ΔV for the return trajectory This operation considers the Earth- Mars trajectory duration and a period of Mars operation of 40 days. Furthermore, an estimated 2 km/s were added for each scenario for the manoeuvres at Mars and a threshold value of 14km/s was introduced to eliminate scenarios with enormous ΔV needs Finally, under recommendation by experts in human spaceflight, it was decided to eliminate all the windows which are in solar minimum due to the higher presence of GCR in the Solar system; it is hard to shield a spacecraft from GCR with limited mass and power. The allowable periods considered are departure dates from one year before the solar maximum and three years after. The resulting table introduces the possible for a short stay mission from 2032 to 2049; they are summarized in Table 3-6. Many scenarios were determined, leading to the final step of the launch window definition - the choice of a reference window with reasonable ΔV and mission re uirements Table 3-6: Initially Identified Launch Windows for a 35 Day Stay at Mars No. Departure ΔV Approach duration Return duration Venus Overall Max. solar date [km/s] [days] [days] flyby duration [days] coeff.

1 16/04/2035 13,6 230 260 no 530 340% 2 05/06/2035 13,1 180 260 no 480 335% 3 10/05/2036 12,2 360 200 yes 600 260% 4 04/07/2037 11,6 280 280 yes 600 765% 5 04/07/2037 9,9 280 330 yes 650 710% 6 02/09/2037 10,8 210 350 yes 600 725% 7 29/11/2045 11,7 280 280 yes 600 260% 8 28/01/2046 12,0 210 290 yes 540 260% 9 02/06/2047 13,2 360 200 yes 600 190% 10 02/06/2047 12,2 360 240 yes 640 190% 11 29/12/2047 13,5 260 240 no 540 280% 12 18/03/2048 13,3 180 240 no 460 290%

From the given windows, opportunities which imply trajectories very close to the sun were eliminated, as they were deemed unfeasible from a thermal control point of view. A maximum threshold value of 280% of the Earth’s solar constant was considered realistic

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3.2.4.3 Launch Window Trade-off Once all the possible launch windows were identified, a trade-off was performed in order to select the best launch window. The four launch opportunities are listed below:  Launch window 3: 05/2036, 602 days with a total ΔV of 12 2 km/s and a Venus fly-by on the approach trip,  Launch window 7: 11/2045, 540 days with a total ΔV of 12 0 km/s and a Venus fly-by on the return trip,  Launch window 10: 06/2047, 640 days with a total ΔV of 12 2 km/s and a Venus fly-by on the approach trip and  Launch window 12: 03/2048, 460 days with a total ΔV of 13 3 km/s and no Venus fly-by. The parameters that have been taken into account in this trade-off are:  IMLEO: Assuming a fixed payload mass of 500t and using the rocker e uation, the ΔV was used to estimate the IMLEO of the CIV  Overall Duration involves several categories, as for example psychology and physiology impact on the crew as well as required amount of resources  Repetitiveness of the mission design. Comparison of the combined pork chop plot indicated that some mission concepts have more often occurring launch windows with similar ΔV re uirements and similar duration of the various mission phases. Those are rated higher, as they allow for higher flexibility of the mission in case of major delays  Maximum solar coefficient has a strong impact on the complexity of the thermal control of the spacecraft  Politics, science and public outreach is one of the main goals of the project. It is rated higher for mission that will happen earlier possible, as the scientific and political impact will be higher, as the likelihood of attempts by other agencies or countries to attempt a similar mission is reduced  Technological Readiness is a condition, which is required to be met before assembling the spacecraft in orbit. The scoring was done, assuming a linear progress of technological advancement since the beginning of spaceflight in 1957  Future heritage intends to forecast the future techniques that this mission can bring to future space application. Indeed, all the mission will develop similar techniques, but earlier mission are ranked higher, as the techniques can be exploited sooner in time  Venus fly-by missions are given a penalty due to the involved risks and the increased design complexity,  Radiation constraints were estimated by a preliminary analysis which concluded, that scenarios with an overall mission duration of more than 650 days are likely to excess the acceptable radiation dose. Hence, only scenarios of less than 650 days are being considered. This is shown in detail in Section 6.3. Table 3-7: Launch Window Trade-off Item Weight 1 2 3 4

Date 05/2036 12/2045 06/2047 03/2048 Delta V 10 5.63 6.87 5.63 1.87 Duration 9.5 4.59 5.11 4.31 5.99

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Radiation constraints 0 5 5 5 5 Venus fly-by 1.4 0 0 0 20 TRL 4.3 4.54 5.06 5.17 5.23 Repetitively 7 6 4 6 4 Solar coefficient 6 4.69 4.69 6.42 4.20 Politics, science and publics 5.25 8.00 4.10 4.00 3.90 Future Heritage 4 5.49 4.93 4.82 4.77 Weighted Scores 5.32 4.99 5.09 4.60

Table 3-7 summarises the results of the launch window trade-off, which led to the decision of using a short stay mission departing in May 2036 with an overall mission duration of 602 days and a Venus fly-by on approach. A visualisation created in SCILAB is given in Figure 3-5, in which the Orbits of Mars (light red), Earth (blue), Venus (yellow) and the CIV (dark red) are shown.

Figure 3-5: Visualization of the CIV Interplanetary Trajectory

3.2.5 The Interplanetary Trajectory of the MATV The calculation of the ΔV re uired for the electric propulsion system was completed using an approximation method in which the equivalent length of travel is calculated for a constant thrust trajectory from an impulsive trajectory. This approximation was completed using NASA technical note “A ethod of Approximating Propellant Requirements of Low-Thrust Trajectories” (NASA, 1966) The study was completed by taking fixed reference values for specific impulse, power-to-mass ratio, efficiency and mass flow rate. These values were taken as reference from Vasimr papers (B. Longmier, et al. 2012) (A. Ilin, et al. 2012). A final value of

SEEDS Executive Summary 09/2014 Page 35 burn duration was also fixed at 2 years of constant thrust, in order to meet with the given time constraint of travel time from Earth to Mars. Using the above values with given numerical approximations for low-thrust trajectories the delta-v budget re uired to calculate the “length of travel” using the length e uation for constant thrust, (NASA, 1966)

√

Where is the length of travel, is the exhaust velocity, is the acceleration (assumed constant), and T is the time of thrust. Using this given length it is possible to calculate an approximate delta V budget for a low- thrust trajectory, using the following delta V equation as given by Zola (1966).

This gives an approximate delta V of 4000m/s, which, when manoeuvres and margins have been included, comes to a final approximation of 8km/s. When compared to the threshold value of 14km/s of the entire CIV mission (including return) this value is appropriate to use as a first order approximation.

3.3 DEFINITION OF THE MISSION CONCEPT IN THE VICINITY OF MARS Being the scientifically and technologically most interesting phase of the missions, the mission concept of the entire system in the vicinity of Mars was studied in detail, especially taking into account the temporal and orbital mechanical requirements involved in a manned mission including two spacecraft. The period of the mission concept in which the spacecraft are in the vicinity of Mars is the most scientifically and technologically interesting phase of the mission. Due to this, the phase was studied in detail; especially taking into account the temporal and orbital mechanical requirements involved in a manned mission which includes two spacecraft.

3.3.1 Definition of Operation Orbits To be able to define the system and ΔV re uirements to analyse the ars manoeuvres, the operational orbits of the spacecraft have to be defined. During the work in this field, the orbits given in the table below were identified. They are explained in detail in Table 10-1 in the Appendix. 1. The capture orbit is the initial orbit after capture. For simplicity, the apocentre was defined to equal the semi-major axis of Phobos' orbit. The perigee was chosen to be 298km altitude over the surface in reference to the capture and operational orbit altitudes of Mars Express. Further detailed analysis may also explore the option of performing capture into an orbit with a higher apoapsis and the use of aerobraking for decreasing the apoapsis for rendezvous with Phobos. The capture orbits are assumed to be non-inclined, as this allows the spacecraft to change into any other inclined orbit without having to change the arguments of nodes. 2. The science orbit has the same semi-major axis and eccentricity as the capture orbit; however it is inclined to allow research of possible future human landing sites. For this, an inclination of 30˚ was chosen according to the requirements defined in Section 4. 3. The orbit of Phobos is used as reference orbit for both the Phobos phasing and rendezvous operations. 4. The CIV Earth departure orbit is the hyperbolic orbit necessary for entering the correct interplanetary trajectory after leaving the Earth's sphere of influence. 5. The CIV hyperbolic arrival orbit is the Mars arrival orbit from interplanetary space. The pericentre is defined to be the pericentre of the capture orbit.

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6. The MATV Earth departure orbit is the hyperbolic orbit necessary for entering the correct interplanetary trajectory after leaving the Earth's sphere of influence. 7. The MATV hyperbolic arrival orbit is the Mars arrival orbit from interplanetary space. The pericentre is defined to be the pericentre of the capture orbit. 8. For full scientific coverage of Mars, also a polar orbit is suggested. It equals orbit 2, but with an inclination of 90˚ Finally, it is not considered in the current mission profile due to the high ΔV requirements involved. 9. The CIV hyperbolic exit orbit at Mars is the necessary orbit to leave Mars' sphere of influence with the correct to enter an interplanetary trajectory intersecting with Earth. 10. The CIV graveyard orbit is the circular orbit, which will be used as graveyard orbit for the PhL and the LAB, see Section 3.3.2. 11. The elliptical transfer orbit connecting the orbit of Phobos with the graveyard orbit.

3.3.2 Definition of the Graveyard Orbit of the PhL and the LAB In order to reduce the overall fuel consumption, it was decided during the preparatory phase that after having performed all the operations on Phobos, the CIV will bring the PhL and the LAB to a graveyard orbit around Mars. The second option would have been for the mentioned payloads to remain attached to the CIV during the burn towards Earth. In order to avoid both biocontamination of Mars' atmosphere and any collision with Phobos, a stable orbit has to be chosen. The resources to conduct a numerical stability analysis of possible graveyard orbits were not available; therefore, an approach for calculating the effort of obtaining a stable orbit had to be found. The approach used in this step does not prove stability by itself, but is a solution based on the definition of graveyard orbits of Earth's retired geostationary satellites. The Inter-Agency Space Debris Coordination Committee (IADC, 2012) proposes a minimum perigee altitude above the geostationary orbit that depends on the solar pressure:

where K is a coefficient which links acceleration and an additional safety margin, is the aspect area and is the solar radiation pressure coefficient. In the proximity of Phobos the gravitational field of the moon has to be taken into account, so this term should be added:

where

In order to compute the value of an iterative method has been used, and the final result of the minimum perigee altitude above the Phobos surface was found to be 488.2 km. While this mathematical approach cannot estimate, how long the proposed orbit would actually be stable, for the framework of this project, it is assumed to be sufficiently stable to be in compliance with current biocontamination rules. Once the CIV has brought the PhL and the Lab in the graveyard orbit, the manned spacecraft can start its interplanetary trip back to Earth.

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3.3.3 Definition of the CIV Parking Orbit at Mars Before discussing the detailed sequence of events, a major trade-off was performed concerning the orbit of the CIV during its stay in the vicinity of Mars. The two following options were considered: 1. After a successful capture in the Martian system, the CIV performs an additional burn to circularise to high orbit with similar orbit parameters as Phobos. The actual distance to Phobos will be discussed in detail later in this report. 2. The CIV stays in the capture orbit, hence decreasing the ΔV re uirements on the massive ship The Phobos Lander deploys from this orbit and performs a major orbit change before performing rendezvous and landing on Phobos. Both options were compared using a formal trade-off, taking into account the following items:  IMLEO according to the mission design estimate of 08/2014 The main driver is the additional ΔV necessary for injecting the massive CIV into the Phobos orbit, and the increased ΔV re uirement when departing from a higher orbit. Scenario 1 has an indicated IMLEO of 6500t compared to 4700t in scenario 2.  Design complexity: taking into account the increased requirements concerning communication, orbit control, and rendezvous timing in the case of scenario 2, in which CIV and PhL will operate in different orbits.  Psychology effects imposed by scenario 2, in which the CIV and PhL operate in different orbits. A main effect is the limited possibilities of returning from Phobos to the CIV.  Future heritage provided for future missions.  Risk of losing the life of Astronauts. As before, this is mainly driven by the complexity and the limited opportunities of the transfer from Phobos to the CIV.  Experimental capabilities of both scenarios. For this, the main driver was the possibility of using remote sensing equipment of the CIV for scientific and landing site selection on Phobos. Furthermore, multiple missions are easier to implement if the CIV and Phobos are in a similar orbit. Table 3-8: CIV Parking Orbit Trade-off Results Solution Scenario 1 Scenario 2.

Total 5.4 4.6

The result of the trade-off is given in Table 3-8. As a result, the scenario 1, transferring the CIV to a Phobos similar orbit - was found to be the preferred solution. The main driver for the choice are the risk of loss of life by sending the crew in the PhL with limited resources to Phobos with CIV being in capture orbit, which is an implied requirement of scenario 2.

3.3.4 Mars Mission Concept and Timeline Following this choice, four different mission sequences were defined in order to study the manoeuvres to be flown in Mars' vicinity in detail: 1. In the first scenario, the MATV transfers the LAB and the PhL after a powered capture into the Phobos orbit, where it waits for arrival and docking with the CIV, which performs a powered capture and transits to the Phobos orbit. After docking, the MATV undocks and transfers to the science orbit. The CIV returns to Earth. Additional system requirements: None; advantages: Use of MATV equipment on Phobos and Mars for long duration; disadvantages: Deployment of Mars lander late in MATV mission.

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2. The MATV performs a powered capture, releases the LAB and the PhL in the capture orbit and then transfers to the science orbit. The CIV docks after a powered capture with the LAB and the PhL, transfers both to Phobos orbit. After the missions, both are released and the CIV returns to Earth. Additional system requirements: PhL and LAB as one unit are self-sustainable; advantages: Long and early science duration of MATV at Mars; disadvantages: No use of MATV equipment of Phobos, CIV must be able to perform orbit changes with the PhL and LAB. 3. The MATV performs a powered capture with PhL and LAB into the science orbit and then returns to the capture orbit after scientific mission and docks with CIV. The PhL is pushed by the CIV to the Phobos orbit after the MATV is released, which returns to the science orbit. The PhL is deployed by the CIV from Phobos orbit. Additional system requirements: None; advantages: Long scientific mission of MATV at Mars; disadvantages: No use of MATV equipment of Phobos. 4. The MATV performs a powered capture into science orbit with the PhL and LAB, the CIV performs capture and transfers to Phobos orbit. The LAB and PhL transfer by themselves to the Phobos orbit. Additional system requirements: PhL able to perform orbit changes with the LAB, PhL and LAB are self-sustainable; Advantages: Long scientific mission at Mars; Disadvantages: No use of MATV equipment of Phobos. These four scenarios were compared with a formal trade off, using the following items:  IMILEO was estimated using the different ΔV re uirements Scoring was performed anti-proportional to the given value.  Risk of loss of live is decreased for scenarios in which access to the additional resources of the MATV are given earlier in the mission.  Science capabilities of the different mission vary, as some profiles may or may not allow the use of the MATV remote sensing equipment also on the surface of Phobos. Additional points were given for scenarios, in which the MATV spends the majority of its lifetime in the science orbits, in comparison with scenarios, in which it spends major timespans ''waiting'' for the CIV in the non-inclined capture orbit.  Design complexity penalties were given to scenarios, which require the PhL and the LAB to operate without the support of one of the main spaceships, especially if they would be required to perform major orbit changes.  Future heritage was evaluated; however, no major differences could be identified.  Psychology ranks scenarios higher, in which access to the additional resources of the MATV is provided earlier, hence giving additional living space to the Astronauts.

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After carefully performing the above trade off, scenario 3 was chosen for the mission. The main driver for this choice was the increased timespan available for the MATV for observing the surface of Mars, especially the fact that this scenario allows to deploy the Mars Lander early in the mission and the limited system requirements of the PhL, which is not required to be able to sustain the LAB without the support of one of the main spaceships. The final mission sequence is shown in Figure 3-6.

Figure 3-6: Operations of both spacecraft in Mars' proximity

3.3.5 Mission Profile in the Vicinity of Phobos After identifying the landing sites of the PhL on Phobos, this section will discuss the operational sequence of the PhL while being in the vicinity of Phobos. Furthermore, it will impose constraints on the position of the CIV during the manned landing phases. The starting point of these operations is arrival of the CIV, including the PhL and the LAB, in the vicinity of Phobos. This will happen one week after the arrival of the CIV in a capture orbit around Mars, to provide sufficient time to dock with the MATV; to transfer the PhL and the LAB; and to phase and circularise its orbit in order to perform a rendezvous with Phobos. Before attempting to land the PhL on Phobos, it is necessary to gather as much remote sensing information about Phobos as possible. For this, the CIV will perform minor orbital changes to reach a quasi-orbit around Phobos, in reference to a mission profile given by Glasmeier (2006). After selecting a suitable landing site, preparations for the first landing attempt will begin. For this, it is foreseen that the CIV will move into a position 200km over the landing site on Phobos and eliminate all relative velocity relative to Phobos. This will lead, due to the coupled rotation of Phobos, to the spacecraft maintaining a certain position over the surface of Phobos. Naturally, the gravitational attraction of Phobos on the CIV will degrade the altitude of the CIV with time, this effect will be counter-measured by the AOCS thrusters The ΔV re uirement for this station-

SEEDS Executive Summary 09/2014 Page 40 keeping was estimated for the worst case scenario - where the CIV spends the entire mission duration in the vicinity of Mars station-keeping over Phobos - as follows:

∫ ∫

This position will allow continuous communication between communication between the CIV and the Astronauts on the surface and also direct observation of the Astronauts with remote sensing equipment. The duration of each landing mission was set in reference to the Apollo missions, a three day landed phase including one day for approach, landing, lift-off and return. Accordingly, the PhL is to be designed for two missions of independent operation, both of the stated duration, including a margin. After the first landing mission and successful docking of the PhL to the CIV, the CIV will continue to quasi-orbit the moon until the beginning of the second landing attempt. After both landing attempts have been completed, the CIV will remain in the quasi-orbit for five days more. This is for both for margin and for exploiting the LAB capability; the selection of Phobos' samples which are the most valuable for return to laboratories on Earth. The Phobos operations are concluded by two engine manoeuvres of the CIV, which lift the spacecraft into the graveyard orbit, as given in Section 3.3.2, where the LAB and PhL are detached. In this orbit the crew will spend four days, implemented as a margin on the return window, before the CIV leaves the Martian system after the Trans-Earth Injection (TEI) burn. An overview of the timeline in the vicinity of Phobos is also given in Figure 3-7.

3.3.6 Astronauts roles and timeline In addition to psychological considerations, the given crew size of six astronauts is also supported by a preliminary analysis of the areas of expertise and the task assignment during the relatively short stay in the vicinity of Mars.

3.3.6.1 Astronaut Roles To define the areas of expertise and roles of the astronauts in the crew, the following assumptions were made:  While all astronauts should be briefed in all relevant emergency procedures, it is assumed that the training of each astronaut will focus on two areas of expertise, a main role as well as a secondary role.  In all fields of expertise which are critical to mission success, at least two astronauts have to be trained. In addition to this fact, all astronauts are briefed in emergency and medical procedures and spacecraft operations, this leads to a Fail-Operational/Fail-Safe strategy. This lead to the definition of the following main roles:  Mission Commander  CIV Pilot  PhL Pilot  Systems Engineer  Physician  Geologist/Scientist And the following secondary roles:  Second in command  Secondary CIV/PhL System Engineer  Secondary Physician

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 CIV/PhL Co-Pilot  Psychologist  EVA specialist Using these assumptions, an example for a crew is given in Table 3-9. The actual field of expertise of the crew however would naturally be different and strongly be driven by the past experiences of the available astronauts as well as political and psychological considerations. Table 3-9: Crew Roles Distribution Example

Astronaut 1 2 3 4 5 6

Mission Systems Primary role CIV Pilot PhL Pilot Physician Geologist Commander Engineer

Secondary Secondary Second in EVA Secondary Co-Pilot System Psychologist role command specialist Physician Engineer

3.3.6.2 Astronaut Timeline The transit phases, which occupy the majority of the mission profile, are the least busy periods for the crew, which strongly shifts the focus of time management of the crew during these phases towards maintaining a good psychological and physiological health. Furthermore, the task of the crew is the maintenance of the spacecraft, which includes observing and analysing the performances of the operating systems as well as regular checkouts of systems in non-operating mode. As an example of a transit phases, the interplanetary journeys will be the longer ones, and it has been used as an example for a timetable during transit periods, as can be seen in Table 3-10. Table 3-10: Transit Phases Time Distribution Task Allocated time (hours per day) Maintenance and checkout of systems 3-6 Physical training and hygiene 2-3 Operational Performance 3 Free time 1-3 Sleep 8

In addition, when close to the arrival at Mars, the crew will prepare itself as well as the CIV for the upcoming weeks of a very dense operational and scientific schedule. During the 35 days, which the crew will spend in the direct vicinity of Mars, the schedule becomes highly dense and mission critical, as the limited amount of men-hour have to be divided into various tasks, as for example:  Performing of major spacecraft manoeuvres of the CIV, e.g. capture burn, rendezvous and docking with the MATV, rendezvous with Phobos, graveyard orbit injection and trans-Earth injection burn.  Performing the manned landing on Phobos, including scientific and mission ground operations.

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 Performing the scientific mission operations related to Mars, especially in the scope of remotely controlling the Mars landing packages in real-time.  Performing remote sensing on Phobos, including the definition of a final landing site of the PhL.  Performing of regolith sample analysis and selection on the LAB. For this, it is assumed, that the sleeping schedule of the Astronauts will be synchronised during this period and that the Astronauts will be divided into three teams of two. Figure 3-7 indicates the occupation of the Astronauts in the period spent in the vicinity of Mars and shows, that approximately 2 weeks of scientific operation are available for both celestial bodies.

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Week 1 2 3

Day 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 Injection of Docking Injection in Phasing of Return MATV Landing site First Landing CIV in Mars CIV and Phobos CIV to to the Detaching selection on Phobos Orbit MATV Orbit Phobos CIV LAB utilization

Team 1 CIV manoeuvres and maintenance Team 2 Preparation of the PhL and landing site selection (14 days) Landing campaigns Team 3 Capability of science on Mars (14 days) Phobos science

Week 3 4 5 Day 18 19 20 21 22 23 24 25 26 27 28 29 30 31 32 33 34 35

Second landing on Phobos Return to the CIV Phobos Sample soil selection in the LAB Detaching of PhL and LAB

LAB utilization Utilization of the LAB (14 days) Margin (4 days) Team 1 CIV manoeuvres and maintenance Team 2 Landing campaigns (17 days) Team 3 Phobos science (17 days) Figure 3-7: Astronauts Occupation During Phobos Operations

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3.3.7 Astronaut locomotion In such low gravity astronauts will not be able to walk as they won’t get enough friction with their feet on the ground. The solution of carbon nanotubes planned by the AENEA mission [SEEDS V – 2011] is not applicable to the dusty surface of Phobos. Moreover, the force that astronauts apply with their feet may cause them to jump high enough to be far from the ground for a meaningful time.

3.3.7.1 Risk of jumping On the surface of Phobos escape speed is 11.4 m/s, or 41 km/h [NASA – Phobos]. The strength of human legs can be estimated by the following assumption: a man who lifts another person of equal weight on his shoulders is strong enough to lift him, but not strong enough to jump to a meaningful height on Earth in such conditions. Is this assumption holds, the strength of human legs can be estimated as the double of the weight. The length of legs (leg + thigh) can be estimated as 49% the height of a person [Exercise Prescription on the Internet]. For the following calculations the base assumption is a human of 70 kg of mass and a height of 1.80 m. The force of legs is then 2 x 70 kg x 9.8 m/s² = 1372 N. The length of legs is 0.88 m. So the maximum work that can be applied by legs on the human body is 1207 J. On Earth this will be partly potential energy and partly kinetic energy. On Phobos, if gravity is neglected, all this energy will be applied as kinetic energy, reaching a speed of 5.9 m/s. This is clearly an overestimation of the speed considering that the astronaut will be carrying extra mass for life support systems of the EVA suit. Jump to orbit is also not possible. The speed of a circular orbit is (GM/r)1/2, while escape speed is (2GM/r)1/2. So escape speed is 21/2 = 1.41 times faster than a circular orbit. If escape speed on Phobos is 11.4 m/s, then 8.06 m/s are needed to get into orbit. However, it is still possible to perform a suborbital flight, so the astronaut could land in a very remote point. For this reason, astronauts shall be explicitly instructed not to jump.

3.3.7.2 Locomotion method Thrusters are required. An important advantage is that the astronaut will be able to quickly come back down to the ground in case of unintentional jump. However, lifting dust from the ground is a risk that must be mitigated. If thrusters fire only upward at an angle of 45º then no dust will be lifted by exhaust gas. It must be carefully designed so that exhaust plumes do not perturb the astronauts work. 3.3.7.2.1 Location of thrusters Thrusters should be located away from the astronaut’s feet to minimize the risk of lifting dust They should also be located away from the astronaut’s head to prevent exhaust gas from disturbing vision or work This leaves only one alternative: locating them around the hips A “ring” around the hips is recommended to keep the plumes away from the body and the astronaut’s hands 3.3.7.2.2 Speed, propellant and mass

Cold gas propulsion can reach a specific impulse of up to 250 s using H2 [Humble, R. W. – 1995]. However, calculations on the showed that using cold gas is impractical due to the extremely low autonomy and very massive tank needed to resist the internal pressure. However cold gas will still be needed when astronauts are close to each other and when they are in the PhL’s proximity A double propulsion system is justified.

With 20 kg of cold gas N2 in a tank at 138 bar a specific impulse of 68 s is achievable [Encyclopedia Astronautica – Nitrogen cold gas]. The mass of the tank has been calculated as 49 kg with a safety margin of 20% 13 m/s of ΔV are usable in these critical conditions, while chemical propulsion is used for the rest. Speeds up to 1 m/s are assumed to be safe to operate on the ground. This is approximate to walking speed, or probably slightly slower. If the astronaut accelerates and stops 100 times, it will need fuel for a Δv of 200 m/s.

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An EVA suit’s mass, including the astronaut’s mass, can be estimated as 223 kg [NASA – 1994]. Assuming a total mass of 500 kg out of which 223 kg are reserved for non-propulsive means, the above propellant+tank mass of 69 kg leaves only 208 kg for chemical propulsion. If HPGP is used as propellant, with an exhaust velocity of 2480 m/s [SSC Group – ECAPS], assuming an engine mass of 1kg and a tank mass of 5%, 197 kg of propellant can be carried. This means that a Δv of 1240 m/s is achievable. Enough mobility for astronauts will then be provided with a non-toxic propellant and a non-toxic exhaust gas. However this doesn’t directly translate into 1240 stops and go’s Thrusters are inclined at 45º, so some part of the energy will be dissipated when the astronaut puts the feet on the ground. A strategy is needed to let the astronaut “walk” maintaining the speed generated by the thrusters while minimizing the energy lost 3.3.7.2.3 Foot and step dynamics The maximum shoe size in the market is 32 cm [Shoe size charts]. Assuming the axis of rotation is 5 cm from the back of the feet, this leaves 27 cm of length of the rotating part for the longest expected foot. The calf muscle is as strong as the leg, capable of lifting twice the human weight. This is verified by a common exercise in gyms, in which the trainee lifts his own weight plus an added weight on his shoulders. However the displacement it can produce is not long. Assuming the foot rotates at most 45º, the path of the foot’s tip is 27 cm long as well. A 70 kg human can then apply a kinetic energy of 370 J. If the dry mass is 275 kg, which will be reached when the tank is empty, the astronaut will be able to push the whole mass upward at a maximum speed of 1.64 m/s (neglecting weight on Phobos’ gravity) This is still too fast, though it is a clear worst case scenario. If the astronaut is trained to exert the minimum needed force with the tip of the foot, using only one foot to half the speed, the speed will be appropriate to perform a small jump of a few m of height. This will be optimized if the astronaut pushes backwards at an angle of 45º.

Figure 3 - The Astronaut exerts a minimum force with the tip of his/her feet [Elaboration on an image credited by NASA] Note that the force is applied while the astronaut is moving forward at 1 m/s This minimizes the astronaut’s capability to apply a meaningful force, simply because he will be moving too fast for his feet to follow the relative speed to the ground, so it will also minimize the achievable vertical speed. Appropriate training will be required. It can be achieved on Earth by using ropes, pulleys and counterweights to counteract most of the astronaut’s weight However, simulating the negligible gravity of 5 7 mm/s² will not be straightforward. A smaller cold gas thruster may be used on Earth, located directly overhead, providing a downward acceleration equal to the Phobos gravity. This will be a realistic simulation in which the astronaut will walk like on Phobos with the aid of the thrusters. It’s important to estimate how much propulsion will be used during steady-state locomotion. This would be better determined empirically during the astronaut’s training If the astronaut jumps at every step with the maximum speed that the calf muscle can provide, then a propellant consumption e uivalent to a Δv=1 64 m/s would be applied at every step! It’s important to train the astronaut to apply a force as small as possible If the vertical speed is estimated at 0.2 m/s at each step, the astronaut would reach a height of 3.5 m at each step (calculated using mgh=½mv2). It is a realistic estimate as the astronaut would be using just 1/8 of the energy that the calf muscles can provide, or 1/4 if using only one foot at a time. Such height may cause vertigo, though this may be prevented if the astronaut is appropriately trained. Each jump would last approximately 70 s with a jump horizontal length of 70 m. Of course the astronaut can use the thrusters to control speed and come to a complete stop at any time, though a meaningful propellant consumption is not expected. Dust may be lifted at each jump, though it is expected to be thrown backwards.

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3.3.7.2.4 Attitude control For the previous method to work it is important to keep the astronaut in the 45º position. Reaction wheels are recommended. Sensors may not be needed if the ACS can be controlled manually by the astronaut, though it is preferable to control the vertical position automatically using EM radars to detect the ground. Reactions wheels are only needed for roll and pitch movements. Yaw movements can still be performed by thrusters. In case vertical speed is too high or distance to the floor is too large, as detectable by an EM radar, thrusters should fire automatically to bring the astronaut back to the ground. In case the astronaut is falling too fast it will be necessary to provide upwards propulsion. This is dangerous as it would lift dust from the ground, so these thrusters should only fire in case of emergency. Once on the ground, downwards thrust is actually provided by thrusters firing upwards at 45º angles. If all thrusters are fired at once then the net force on the astronaut will be 0 However the exhaust gas will “clean” the astronaut’s surroundings by pushing away any dust particles that may be floating around the helmet.

3.3.7.3 Alternate locomotion: dustboarding (snowboard-like) There are uncertainties about the properties of the soil, particularly about the possibility of sinking the feet in the regolith when the astronaut touches the ground with a significant vertical speed. This would be an issue because of the dust lifted when un-burying the astronaut’s feet An alternate concept is sweeping on the ground, which derives from the concept of increasing the base area of shoes. Thrusters will still be needed As the astronaut’s feet do not push the ground, jumps to a meaningful height are not expected by foot dynamics. However terrain irregularities can push upwards during the movement. The main issue in this method would be lifting dust, though this can be minimized if the front shape has the appropriate curvature. Propellant consumption may also be a consequence. The main advantage is precision to reach a specific point on the ground. The hopping and the dustboarding method do not exclude each other. The astronaut may carry the dustboard in a backpack and use it when needed. In fact it may be very useful to combine both methods: hopping to travel long distances, then dustboarding to reach a specific point. Astronauts may in fact be able to improvise a locomotion method, as happened in the Apollo 11 mission, so providing means for two different methods means an extra degree of freedom to improvisation without a meaningful impact on mass.

3.3.7.4 Tether design For safety, a tether with a cross section of 1 cm² will be used. The cable should actually be composed of several threads thin enough for the cable to be flexible. However, the thermo-optical properties of aluminium make it unsuitable for direct sunlight exposure. A layer of 1 mm will be reserved to add a reflective material that protects aluminium from overheating. This leads to a cable of 0.67 cm in radius, out of which 0.10 cm will be the thermo-optical resistant layer. In order to control the astronaut’s acceleration when the maximum length of the tether is reached, it is recommendable to bind it to the PhL with springs. This will also reduce perturbations to the PhL at a point in which its thrusters may control it. During a review of the present work by a former astronaut, an issue emerged: tethers behave unpredictably in very low gravity and they may cause more problems than what they intend to solve. Analysis of the situation has shown that even without tethers the system is still two-failure tolerant, since astronauts can help each other and the Lander spacecraft can move to pick an astronaut up (either it shall be designed to do this automatically, or telecontrolled by astronauts in the CIV). For these reasons, astronauts will be allowed to detach from the tether after following a safety procedure in which the correct functioning of all equipment is validated. This justifies the short tether length of 10 m, which had been previously designed as 1 km.

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3.4 PROPULSION

3.4.1 Introduction This chapter summarises the preliminary design analysis performed for the investigation of the main propulsion sources suitable for the CIV and the MATV mission profiles. On the basis of the mission requirements imposed, the selection of the appropriate propulsion technologies is driven by the trade-off analysis performed for both spacecraft. Through several figure-of-merits - or key parameters - the main critical aspects of each propulsion technology have been considered. The trade-off analysis is focused on the comparison of nuclear and electrical propulsion technologies with Chemical Power Sources (CPSs), the most reliable and utilised system. Appropriate weight-factor has been assigned for each of the figures-of- merit presented below.

Spacecraft Overall Mass This parameter represents one of the most important factors that drive the final result of the trade-off analysis. From the Tsiolkovsky equation, for a given value of the change of velocity ( ) and specific impulse, is possible to understand how the spacecraft’s I LEO strongly affects the amount of propellant requested. Higher values of specific impulse will lead to lower propellant mass, reducing the spacecraft overall mass.

Design Complexity This figure of merit is mainly related to the spacecraft overall mass. It expresses the launch cost and the launch strategy complexity to bring and assembly this mass in LEO. The complexity level of the launch strategy is based on the launcher capacity and availability, aspects which also directly influence the level design complexity of the spacecraft itself. From this point of view, considering Nuclear Propulsion Systems (NPSs), one of the main relevant constraints is the distance at which the NPS should be installed from propellant tanks and habitable modules, in order to reduce the radiation fluence. Also, the storage strategy of the total amount of propellant results to be a critical issue for the spacecraft design complexity and assembly strategy due to the boil-off problem associated with cryogenic fuels.

Technologies Readiness Levels (TRLs) and Past/Future Heritage The TRLs strongly influence the safety and the reliability levels of the propulsion systems considered. Chemical and Electric Propulsion applications are the most mature technologies for space applications, with sufficient TRLs already available today. On the other side, for the Nuclear Thermal Propulsion System (NTPS), actual TRLs result to be very low, due to the fact that no space applications have ever been tested. However, considering the next 20 years, important improvements of the TRLs for NPSs it is assumed. Safety Generally, the safety level is one of the others main critical parameters, which strongly influences the overall mission and spacecraft design, especially if manned payloads are considered. From the propulsion point of view this figure-of-merit is related to the reliability levels of the propulsion systems, which must be able to ensure sufficient safety and survivability levels of the crew. For the MATV Trade-Off Analysis, due to the absence of manned payload, the safety concept has been simply translated in the figure-of-merit of risk of mission failure, related to the systems reliability levels of the MATV.

Psychological Effects Through this figure of merit, is also possible to evaluate the influence that the psychological effects have on the selection of the appropriate propulsion technology. The mission requirements impose to perform manoeuvres with the NPS The idea to have several nuclear reactors “on-board” can produce levels of stress on the crew.

Political Constraints This item regards the possibility to utilise Nuclear Propulsion Technologies (NPTs) directly from LEO. Actually, a clear international normative, which allows or forbids the utilisation of NPTs in LEO doesn’t exist The present normative of the United Nation Office for Outer Space Affairs (UNOOSA) doesn’t identify the worldwide situation about the utilisation of nuclear propulsion technologies in LEO. On the other hand, from

SEEDS Executive Summary 09/2014 Page 48 political point of view, the utilisation of NPTs results to be completely unattractive. Hence, due to the high level of uncertainty, this point strongly influences the overall architecture of the mission and then associated spacecraft architectures, forcing the choice of only CPSs or Electric Propulsion Technologies (EPTs) for the departure manoeuvres. In the following tables show the Trade-Off analysis conducted and each score (from 1 to 10) assigned at each figures-of-merit previously described.

Table 3-11: CIV Propulsion Trade-Off Analysis Figure of merit Weight Factor NPSs Chemical PSs Spacecraft Overall Mass 10 7.65 4,52 Political Constraints 9 4 6 TRL (Past Heritage) 1 2 8

TRL (Future Heritage) 4 8 2 Spacecraft Design Complexity 3 4.5 5.5 Safety 7 4 6 Psychology Effects 3 2 8 Trade-Off result 37 5.24 4.75

Table 3-12: MATV Propulsion Trade-Off Analysis Figure of merit Weight Factor EPTs Chemical PSs Spacecraft Overall Mass 10 7.65 4,52 Manoeuvre Complexity 2 4 6 TRL (Past Heritage) 3 2 8

TRL (Future Heritage) 7 8 2 Spacecraft Design Complexity 7 4.5 5.5 Risk of mission failure 10 4 6 Trade-Off result 39 5.19 4.80

3.4.2 The CIV Propulsion System

3.4.2.1 Introduction As a result of the trade-off analysis, the main propulsion technology selected to accomplish the CIV mission requirements is the Nuclear Thermal Rocket Propulsion System (NTRPS). This is, for some aspects, similar to the concept of the Liquid Rocket Propulsion System (LRPS) except, obviously, for the mechanism that adds heat to the propellant. Indeed, similar feed systems developed for LRPS are typically used to feed the propellant into nuclear reactors.

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If compared with chemical or electrical options, the NTRs result to be the leading option of the propulsion system for a manned mission to Mars. From the physical point of view, according to Newton’s second law, the propulsion systems accelerate matter to provide the requested force (thrust, T) to move and/or rotate the spacecraft around its centre of mass. The momentum variation is directly related to the release and conversion of the Potential Energy (PE) of the reactor mass, into the Kinetic Energy (KE). Unfortunately, not all the PE stored in any type of reactant (chemical, nuclear, ...) can be completely converted in KE, but only a certain fraction of it. How much of this PE is converted depends on the types of PEs tapped. For the purpose of this work, two of the PEs are of interest: the first one is associated to the electroweak force (or Coulomb Force), the second one is associated to the nuclear force. In chemical propulsion, the change in momentum is made by the conversion of the propellant chemical potential into KE. The chemical potential is the PE level associated to the binding energy levels of the different propellant’s chemical species. Hence, the amount of PE that can be converted is directly related to the Coulomb force, which acts on the atomic and molecular scale, with the order of 10’s of electron volts Through the chemical reactions inside the combustion chamber, the binding energy is released and then translated in the associated heat (KE) The main limit of Chemical Propulsion (CP) is the “lowest values” of specific impulse (Ip), linearly related to the exhaust gas velocity – or exit velocity - (Ve) of an ideal expansion. The Ve values are limited because they cannot be much different from the random velocity values (average value, Ve) reached by the molecules inside the thrust chamber, directly related to the amount of PE translated into KE. In the nuclear propulsion, the nuclear strong force considered is the strongest natural force known. This force acts not on the atomic and molecular scale, but among nucleons (protons and neutrons), preventing the spontaneous disintegration of the nucleus due to the Coulomb repulsion. As a result, the PE level that can be translated into KE is much higher than the amount of PE that it is possible to translate from the electroweak force. The Nuclear force has an order of magnitude of MeV: 106 to 107 times larger than the chemical one. This means that the PE translated into the kinetic one is 106 to 107 times larger than chemical ( ).

The values of the SI are linearly related to the associated values of Ve, in this way:

(1)

Whilst, as previously mentioned, Ve is related to the amount of PE converted in KE:

(2)

For which results that:

√ (3)

where: , in other words:

(4) Expression (3) confirms how the Nuclear Power sources are critical for space missions. Through the utilisation of the modern chemical High-Energy-Density-Materials (HEDM) is possible to produce only a modest increase of Indeed, considering the square-root dependence, there are only minor effects on the SI values. To increase significantly the SI is necessary to increase the energy level by order of magnitude that only a nuclear source can produce. For these reasons, through Expression (4), it is also possible to see how the SI obtained with nuclear systems are almost twice the value of the SI obtained with chemical systems. This concept is not only confined at an energetic point of view, but strongly depends also by the molecular mass of the propellant utilised. Indeed, from Equation (2), is also possible to see how the exit velocity of the reactants is inversely proportional to their molecular mass. In NTRs it is possible to utilise propellants which have an atomic mass of 2 (Hydrogen example), instead 9-10 which is typical of the combustion products of the best chemical rockets. This means that:

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√ √

in other words, as confirmed by expression (4):

(5) Thus, to enable manned interplanetary missions, the only physical alternative to the chemical propulsion - which represents the workhorse of the space propulsion technology- it is represented by the adoption of the Nuclear Energy as the main Power Source.

3.4.2.2 The CIV Nuclear Propulsion Systems (NPSs) In a Nuclear Rocket Propulsion System (NRPS) the heat source comes from the nuclear fission reaction. Through different heater exchangers, the heat produced is directly supplied to the propellant, creating hot and high-pressure gas, subsequently expanded through a converging/diverging nozzle. There are two common ways to heat the propellant for a NPRS: in a direct way, through some heat exchangers (Nuclear Thermal technology), or creating electricity (Nuclear Electric technology). For the purpose of this work, according with the results obtained in the Trade-Off analysis, the NPS selected for the CIV is the Nuclear Thermal Propulsion Technology (NTPT). Three main reference NTPSs have been considered. The first one is represented by the Nuclear Engine for Rocket Vehicle Applications (NERVA) developed by NASA from 1958, is considered the main future propulsion framework for the USA Space Exploration Programs (SEP). The Particle-bed Reactor (PBR) is second high innovative NTPT considered. Compared to the others nuclear propulsion technologies, it is characterised by high safety and efficiency levels. Finally, CERMET-Core nuclear rocket is the last nuclear technology evaluated. The latter is more attractive for future applications in Near Earth Orbit (NEO), especially on reusable orbital-transfer vehicles (OTVs). Table 3-13 summarises the main performance parameters of the three reference NTPTs considered for the feasibility study of the CIV nuclear propulsion system.

Table 3-13: State of the art of Nuclear Propulsion Systems Performance Parameter NERVA PBR CERMET Power (MW) 1570 1945 2000 Thrust (N) 334,06 333,62 445,27

Propellant H2 H2 H2 Max Chamber Temperature 2361 3200 2507 (TC) Specific Impulse (s) 825 971 930 Chamber pressure (MPa) 3.102 6.983 4.136 Nozzle expansion ratio ( ) 100 125 120 Engine Mass (kg) 10,138 1705 9091 Shield Mass (kg) 1590 1590 1590

The feasibility study of the CIV nuclear propulsion system is mainly based on the performance characteristics of PBR and NERVA nuclear technology. The PBR represents the most attractive and suitable NTP able to ensure higher values of Change of Velocity ( ). On the other side, NERVA results in being more attractive for propulsion systems operating in Near Earth Orbit (NEO), characterised by “low” values of ’s.

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However, considering the staging strategy of the CIV mission profile, both NTPSs have been investigated in parallel. The final choice between them will be performed on the basis of the mission requirements imposed. It is important to highlight that PBR and NERVA design procedures are here outlined presenting the results obtained for only one single engine for each type. Indeed, a low-thrust, clustered engines configuration for the final CIV “engine layouts” has been selected. This is to increase the crew safety and reduce the mission risk. In addition, it is supposed that the time and cost associated to develop and ground test smaller engines, will be less than that required for bigger ones. As a result, the final output of the feasibility study will present two different cluster-solutions: one related to the PBR technology, the other to NERVA propulsion technology. Both NTPTs have been developed considering the parallel between hydrogen and methane as propellants but only this last configuration is included.

3.4.2.3 The Nuclear Propulsion Design Process The feasibility studies of the selected NTPTs start with the evaluation of the propellant thermochemistry characteristics, where the choice of the maximum allowable temperature (Tmax) is the main concern. From Table 3-13, the reference value of chamber temperature (TC) considered for the PBR feasibility study is 3200 K, while the value of 2361 K is considered for NERVA. The following analytical relations allow the values of the specific heat to be accurately estimated, by simply introducing the chamber temperatures previously selected, considering methane (CH4) as the propellant.

( ) [ ] (6)

The temperature parameter is defined as:

while the parameter M represents the molecular mass of the methane, equal to 16.043 g/mol. With the values of the specific heats computed, is now possible to obtain the associated values of the isentropic parameter ( ) and gas constant (R), through the following relations.

[ ]

After to evaluate the values of the specific impulse, is necessary to calculate the respective value of the characteristic exhaust velocity as follow:

√ [ ]

( )

( )

The parameter represents the combustion efficiency which typically ranges from 0.9-0.98. The values of the specific Impulses can be now calculated towards the following relation:

( )

√ ( ) ( ) [ ( ) ] ( )

{ }

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The parameter represents the nozzle efficiency, typically ranges between 0.9 and 0.98. Based on the values of Specific Impulse, Change of Velocity and the Spacecraft Mass Estimations (SMEs) due to the System Engineering Analysis (SEA), the Tsiolkovsky equation allows the estimation of the amounts of propellant mass required to accomplish the provided mission manoeuvres.

( )

(11)

The results should be in accordance with the SEA, to converge to optimized values. From the initial ( ) and the final value ( ) of the CIV overall mass at the beginning and at the end of each single manoeuvre ( ) the propellant mass ) can be simply calculated as the difference between those values. Knowing the initial overall mass of the CIV, utilising the equation reported below, it is possible to estimate the initial thrust level ( ), which directly affects the size and the weight of the PBR Nuclear Reactors. The spacecraft thrust to weight ratio (T/W) considered is equal to 0.1.

(12)

Subsequently, from the definition of SI, the propellant flow rates ( ̇ ) can be estimated as follow:

̇ [ ]

3.4.2.3.1 Sizing the PBR Nuclear Reactor: the Power Core Level Requested The power level requested shapes the size, complexity and cost of any kind of nuclear reactor. Conceptually, nuclear reactors are power unlimited. It is the operative value of the core temperature, which limits the power level, due to material constraints. If the heat produced is removed sufficiently fast, maintaining the temperature values within determined limits, the nominal power can be increased as desired. However, greater power level means greater fuel consumption.

The core power reactor (Pcore) directly depends on the mass flow rate (Equation-13) and by the specific reactor power (P) as showed in the equation below:

̇ (14)

The Specific Reactor Power (SRP) is a function of the chamber temperature selected (Tc). Considering this temperature value in the equations below, is possible to determine the requested SRP which is the power requested to heat 1 kg/s of the propellant flow considered. To get the total amount Pcore required, the SRP is multiplied for the propellant flow rate selected, see equation 14.

[ ]

Once the core power levels have been estimated, the following analytical equations (given by the results of the parametric analysis of the Los Alamos Reactor at the Los Alamos National Laboratory (Humble, 1995)), it is possible to compute the associated values of the PBR Core Radius, considering a 37 fuel element configuration of each PBR core:

(17) – (18) 3.4.2.3.2 Sizing the NERVA Nuclear Reactor: the Power Core Level Requested The NERVA reactors have been designed applying basic physical concepts. All the reactors are assumed homogeneous, with no variation in their performances through the cores. Again, as for the PBR reactors, the sizing parameter is the power level requested (PCORE). Each nuclear technology is also characterised by a certain power density (PD) range. For the NERVA-type reactors the characteristic power density is 1,570 MW/m3. Through the equation below, on the basis of the value of the power core, power density and burnout time or burn duration (tb), is possible to estimate the requested value of the NERVA core volume (VCORE).

(19)

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The NERVA core power level is estimated by reproducing the same steps performed for the PBR technology, considering now a different value of the temperature chamber, equals to 2361 K. Introducing this temperature value, inside equations 14 and 15, means that the respective values of the reactor power required for a given propellant temperature (MWs/kg) can be calculated. Finally, multiplying these values for the associated mass flow rates (equation 14) it is possible to obtain the requested values of the power core for which the NERVA reactors should be designed. The design of NERVA reactors, considers the utilisation of the 90% enriched 235U as the fuel, methane and hydrogen as the propellants, and reflectors, which allow us to reduce the radius of the cores, and therefore the overall mass. For a cylindrical shape, the core dimension is obtained solving the equation below:

( ) ( ) considering that, due to the cylindrical shape, the core volume can be simply also expressed as:

(21) Solving this system of two equations is possible to obtain the values of the NERVA core radius and length (HCORE). From equation-(17), is possible to see how the radius of the core reactors depend by the material buckling parameter ( ) and by the multiplication factor ( ). The first parameter takes into account different nuclear parameters which characterise the neutrons lifetime inside the NERVA reactor, such as:

 Non-leakage Probability (Pnl), defined as the fraction of neutrons that remain in the reactor core region (including reflection). This parameter defines the reactor geometry needed for criticality. The value of this parameter should be chosen to get a critical reactor and then utilised to define the reactor geometry.

 Thermal-Diffusion Length (LCORE) defined as the square root of the ratio between the macroscopic absorption cross section (∑ ), over the transport length (DCORE) which represents the mean distance travelled by neutrons through the core.  Neutron Age ( ), which represents the average lifetime of neutrons inside the core.

√ [ ]

3.4.2.3.3 Nozzle and Reactor Containment Vessel Design On the basis of the values of mass flow rates, chamber pressures and characteristic velocities, is possible to determine the critical dimension of the nozzles. The propellant mass flow rate can be expressed also as a function of the nozzle throat area ( ) and pressure chamber (pc):

̇ { ( ) } [ ] √

Resolving equation 19 for the throat area, remembering at the same time equation 9 and 13, the following is finally obtained,

̇

Before the nozzle exit area ( ) is computed, it is necessary to estimate the nozzle expansion ratio ( ). The first step regards the estimation of the exit Mach number (Me) on the basis of the values of the chamber pressure and of the exit pressure (pe) which, considering the deep space operations coincides with the reference value of the vacuum pressure, assumed equals to Pa.

[ ] (24)

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Substituting the values of Me obtained, it is then possible to estimate the values of the nozzle expansion ratio through the following relation:

[( ) ( )]

Through the definition of the nozzle expansion ratio, considering equation 25, it is possible to calculate the associated value of the nozzle exit area. Considering conical nozzles, the length depends on the divergence half angle ( ): the greater this angle, the shorter the nozzle will be for a given expansion ratio ( ). However, increasing the conical divergence angle means an increase thrust losses due to the increase of the radial components of the exit flow velocity. This effect is taken into account in the nozzle efficiency ( ), which, for the purpose of this work, is considered to be equal to 0.98. Typical values of the conical divergence angle range from 12° to 18°. The length of the nozzle diverging section can be estimated as follow:

To conclude the nozzle design, the total mass has been estimated. Before computing this, it is necessary to evaluate the value of the nozzle wall thickness driven by: the level of the pressure chamber, the Space Safety Factor (SF) and the ultimate tensile strength of the material selected. For this preliminary analysis, a high-temperature material nickel alloy is considered, characterised by an ultimate tensile strength of 310,000,000 Pa and a density of 8500 kg/m3. To be conservative, it is assumed that the pressure level in the Throat section is equal to chamber pressure while, typically, the pressure level in the throat section is around 50% of the pressure chamber. For space applications, a SF of 3 is here considered, to better counteract stress concentrations and thermal effects on the nozzle section. Hence, the requested value of the Throat thickness can be estimated as follow.

To reduce the mass, the nozzle-tapered wall configuration has been selected. The following analytical equation allows the total mass of the nozzle configuration described to be better estimated.

[ ( ) ( ) (( ) ( ) ) ]

The reactor containment vessel, as the name suggests, is the external case containing the nuclear reactor, needed to maintain, inside the core, the pressure level requested (typically from 3 MPa to 8 MPa). Its shape results to be more like a propellant tank. It is assumed that the total vessel volume ( ) is equal to the volume of the reactor (VCORE) plus the two hemispheres at either end of the cylinder.

The associated mass value ( ) has been estimated through the empirical pV/W approach. This approach considers the so-called vessel mass factor ( ) related to the vessel type of material selected. For a total metallic vessel (aluminium) its value is equals to 2500 meters, while if fibre-reinforced composite materials are considered, its values is assumed equals to 10000 meters. Also, the burst Pressure (pb) – the product between the SF and the chamber pressure – and the vessel total volume are considered, as shown in the equation below:

3.4.2.3.4 PBR and NERVA System Pressure Levels Considering the huge amounts of propellant requested to accomplish the mission investigated, the installations of pumped systems are considered. These systems allow consideration of a tank pressure level which typically ranges from 0.2 MPa to 0.5 MPa. For instance, the tank pressure of ARIANE-5 is equal to 0.4 MPa. Without pumps, the propellant tanks shall be pressurised at 8.3 0.05 Mpa, with a high penalty for the

SEEDS Executive Summary 09/2014 Page 55 tank mass budgets. According to the thermal and structural requirements, for the propellant selected (methane), a pressure tank value of 0.9 MPa (9 bar) is considered. The pumped systems are designed to ensure that the requested propellant mass flow rate (kg/s) at the pressure level requested (pC), overcomes all the pressure losses due to leakages and feed systems. The first contribution to the pressure loss is given by the fluid dynamics characteristics of the propellant itself. As the propellant leaves the tanks its velocity goes from zero to the required flow velocity. Considering the propellants selected as incompressible flows the pressure drop due to the increase of the dynamic pressure is simply estimated through the Bernoulli’s equation. The second main pressure loss contribution comes from the lines of the feed system. For Liquid Propulsion Systems (LPSs) the pressure losses due to the feed lines range from 0.35 bar to 0.5 bar. In order to be conservative, a value of 0.5 bar has been considered. The third term is related to the cooling system and it represents the most important pressure drop term for the NPS. In the cooling jacket, the pressure loss can range from 10% to 20% of the chamber pressure (pc). Finally, the last term which contributes to the total pressure loss is the reactor pressure drop; this is equal to 5% of the pressure chamber. This factor derives from the interaction between the heat produced by the fission reaction and the liquid propellant. Considering all the pressure loss contributions it is possible to estimate the pressure level requested at which each propellant tank should be pressurised.

3.4.2.3.5 PBR and NERVA Cooling and Feed Systems Statistically, in LPSs the, nozzles and the combustion chambers make up about the 40% of the total propulsion system mass. Hence, from a statistical point of view, the cooling system takes up about 35%, while the feed system makes up the last 25%. Obviously, NPSs are different but, as a first and rough estimation, and due to the fact that for NPS are not available empirical reference values about these systems, the percentages previously presented are considered. 3.4.2.3.6 PBR and NERVA Pressurisation System The design of the Pressurisation System (PS) depends on the choice made between pressured tanks or pumped-feed systems. For the purpose of this work a pressure-regulated system has been selected. The reason of this choice in mainly focused on the need to maximise the tanks propellant capacities which, considering the huge amount of propellant requested, allows the reduction of the number of launches – and then the number of tanks - necessary to bring the overall propellant mass requested to LEO. As discussed in detail in the Structural and Mechanical chapter, this goal is reached also on the basis of the fairing geometries characteristics of the present and future launchers such as Falcon 9 and SLS. For space applications, the most popular pressurant gas utilised is helium (He), due to its good inert and lightweight characteristics. However, this results in a high expense. Possible valid and economical substitutes can be hydrogen or nitrogen. In space applications, the typical pressure level realised in a pressurant tank is 21 MPa, but higher pressure levels can also be used. In order to better estimate the total mass of the pressurant gas ( ) requested to pressurize the propellant tanks, it is firstly necessary to define the initial ( ) and the final temperature ( ) at which the pressurant gas will be stored. As initial temperature value, the spacecraft (CIV) temperature is considered. On the other side, the stored temperature of He inside the pressurant tanks can be evaluated considering that, during thrust phases, there no sufficient time for the heat to transfer from the tank structures into the propellant. Hence, the final value of the pressurant gas temperature - inside each propellant tank - can be calculated thought the following isentropic relation:

( )

where is the He pressure level at which it is stored inside the pressurisation tanks (21 MPa) and is the final pressure level at which it will be stored inside the propellant tanks (0.9 MPa). Utilising the perfect gas law, the following is obtained:

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where is the final volume of the pressurant required, given by the sum of the propellant tank volume with the volume of the pressurant storage tank, as reported below.

3.4.2.3.7 PBR and NERVA Radiation Shielding For space, NPS radiation shield is required in order to protect the crew, the habitable volumes and other spacecraft equipment sensitive to radiation. Several factors influence the geometry, the composition and the mass of the radiation shield: the size and the nature of the power source, the types of the radiations, the configuration of the CIV platform (the radiation flux level decreases by a factor of 1/d2 from the radiation source), mission duration and the maximum level of radiation dose allowed for each astronaut. Hence, many can be the possible valid configurations for the radiation shield. For the purpose of this work, considering the NPS and CIV configuration, the design of the radiation shield is directly based on the size of the core reactors dimensioned. It is supposed that the radiation shield surrounds the core reactor, protecting the tanks and the habitable modules of the CIV against the radiation flux. To dimension the radiation shield, the following standard cross section has been considered, which a reference mass of 3500 kg/m2, composing of three different layers of material. The first one is 18 cm of Beryllium - neutron reflector material - the second layer is composed of a heavy material in order to shield the gamma rays - for this purpose Tungsten (W) is typically a good candidate - and the last layer is a lighter-weight material, necessary for complete attenuation of the neutron flux. Lithium-hydride (LiH2) is indicated for this last layer due to the combination of the good neutron slowing proprieties of hydrogen, and the absorption cross section performances of lithium.

3-8: Radiation Shield cross section layout

3.4.3 PBR and NERVA Design results The values presented in Table 3-14 are refereed for a single PBR and NERVA Nuclear Thermal Propulsion Engine. Table 3-14: Nuclear Propulsion Results Parameters Source PBR Engine NERVA Engine

Molecular Mass (M) – [g/mol] Ref. Data 2.016 2.016 Liquid Density ( ) – [kg/m3] Ref. Data 421 421 Exhaust Gas Constant (R) – [J/kg*K] Equation (8) 518.2 518.23

Temperature Chamber (TC) - [K] Table-(3.3) 3200 2361

Pressure Chamber (pC) - [MPa] Table-(3.3) 6.9 3.1

Specific Heat (CP) – [J/kg*K] Equation (6) 5958.5 6062.4 Isentropic Parameter ( ) Equation (7) 1.1 1.1

Nuclear Combustion Efficiency ( ) Ref. data 1.15 1.15 Characteristic Exhaust Gas Velocity (c*) – [m/s] Equation (9) 2011.6 2012.4

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Specific Impulse (ISP) – [sec] Equation (10) 588.42 588.42

Thrust/Burn Time (tb) – [hrs] Estimated < 2 <2

Maximum Thrust Level requested (Tinitial) – [kN] Equation (12) 442.5 442.5

Maximum Flow Propellant Mass requested ( ̇ ) – [kg/s] Equation (13) 76.6 76.65 Specific Power Reactor (P) – [MW s/kg] Equation (15) 22 15

Total Power Reactor (PCore) – [MW] Equation (14) 1693 1161

Radius of the Core Reactor (rR) – [cm] Equation (16) 34.8 41.4

Length of the Core Reactor (lR) – [cm] Equation (17) 84.5 87.4 3 Volume Core Reactor (VCORE) – [m ] Equation (19) 0.32 0.74 3 PBR Core Reactor Density ( ) – [kg/ m ] Ref. Data 1600 2300

Weight Core Reactor (MCORE) – [kg] Calculated 514.8 1701

Radiation Shield External Ratio ( ) – [cm] Calculated 0.63 1.14 3 Radiation Shield Volume ( ) – [m ] Calculated 0.73 1.3

Mass of the Radiation Shield ( ) – [kg] Calculated 2032 2046 Reactor Vessel Volume Equation (29) 0.5 3.4 Reactor Vessel Mass Calculated 95.3 957.7

-8 -8 Nozzle Exit Pressure (pe, [Pa]) External Ref. 1.38*10 1.38*10

Nozzle Exit Mach Number (Me) Equation (24) 4.58 4.61 Nozzle Expansion Ratio ( ) Equation (25) 265 287

2 Nozzle Throat Area (At , [m ]) Equation (23) 0.022 0.05 2 Nozzle Exit Area (Ae, [m ]) Calculated 5.92 13.14 Nozzle Efficiency Ref. Data 0.98 0.98 Nozzle conical half Angle ( ) – [deg] Calculated 16.3 16.3

Nozzle Length ( ) – [m] Equation (26) 4.42 6.58

Nozzle Throat Thickness (tw) – [m] Equation (27) 0.009 0.006

Nozzle Exit Thickness (te) – [m] External Ref. 0.01 0.01

Nozzle Mass (mNozzle) – [kg] Equation (28) 1701.6 2678

Total Mass of the Thrust Chamber (mth) – [kg] Calculated 1796.9 3530

Total Mass of the Feed System (mFeed) – [kg] Calculated 1118.6 2200

Total Mass of the Cooling System (mCool) – [kg] Calculated 1576.77 3100 Dry Mass Nuclear Thermal Engine [tonnes] 13.2 21.4

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3.4.4 PBR and NERVA Design Conclusion From the dry mass point of view, considering the same thrust performances requested, the Particle-Bed nuclear thermal engine results to be more efficient then the NERVA type. However, if the amount of power produced - for the same thrust phase - is taken as the reference parameter, the situation is in favour of the NERVA engine. This means that, for the thermodynamic point of view, during the no-thrust phases where both nuclear reactors are switched-off through the insertion or the rotation of the control bars, the NERVA engines request to reject an amount of power less than in the PBR case. Consequently, the size, and the mass, of the radiators for the NERVA NPT will be decreased. The order of magnitude of the power core necessary to be rejected is around hundreds of MW for PBR, and tens of MW for NERVA. For this reason, the amount of power that it is necessary to be rejected becomes the driving parameter for the selection of the appropriate nuclear thermal propulsion technology. In addition, also the value of the chamber temperature (TC) constitutes an important drive parameter, which is directly related to the TRLs of the materials available, again in favour of NERVA. To get more detail on these last concepts refer to the chapter of the Electric Power System.

3.4.5 The MATV Electric Propulsion System (EPS)

3.4.5.1 Introduction Compared with chemical and nuclear thermal propulsion, the electric propulsion technology utilises electromagnetic forces to directly accelerate the reaction mass, creating high exhaust speeds. Three main EPS technologies have been considered for the purpose of this work: the electrothermal technology which uses the electrical energy generated to heat the propellant flow; the electromagnetic technology, where the Lorentz force is utilised to accelerate the propellant to high speeds - such technology sometimes utilises charge-neutral, ionized gas called plasma, so it can also be known as plasma propulsion - and the electrostatic technology. Table 3-15: Space Propulsion Technologies comparisonTable 3-15 briefly highlights the main performance characteristics of the EPS technologies, in comparison with the other propulsion systems evaluated. Table 3-15: Space Propulsion Technologies comparison Type Specific Impulse Thrust Level Thrust Duration [sec] [N] Chemical 200 - 465 0.1 – 12*106 Minutes Nuclear Thermal 550 - 1500 > 12*106 Hours Months (constant) Electrothermal 300 - 1500 Year (intermittent) Months (constant) Electromagnetic 1000 – 10,000 0.0001 - 20 Year (intermittent) Months -Year Electrostatic 2000 – 100,000+ (constant)

The first step of the EPS design process is focused on the evaluation of the most important performance parameters of the thrusters, such as: specific impulse, efficiency, thrust level and propellant type. Through the evaluation of these parameters, it is possible to identify the most appropriate electric propulsion technology, selected between the three shown in Table 3-15. After the selection of the types of thrusters, the second step is focused on the selection of the power source able to satisfy the MATV mass budget. The last two steps of the design process are related to the design of the power condition and thermal management system, respectively.

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3.4.5.2 The Electric Propulsion Design Process The design goal is to maximise the payload mass fraction, defined as the ratio between the payload mass (MPayload) and the MATV overall mass (MInitial). To select the appropriate technology, firstly, it is necessary to select the value of the normalised exhaust gas velocity ( ), which allows the maximisation of the payload mass fraction. Typically, ranges from 0.7 to 1.4. For the purpose of this work, a value of 1.3 has been considered.

The second step consists in the definition of the dimensionless speed. This performance parameter is defined as the ratio between the values of change of velocity requested – derived from the MATV mission requirements - and the reference exhaust gas velocity (V0):

Rearranging the Tsiolkovsky equation, the value of the dimensionless speed can be computed solving the equation below for :

( )

Remembering the previous value selected of , it is possible to estimate the reference exhaust gas velocity (V0), identifying the value of the specific impulse which maximises the payload mass fraction, allow us to select the appropriate electric propulsion technology.

[ ] [ ] –

Once the value of the specific impulse has been evaluated, by readapting equation (11) it is possible to compute the associated propellant mass requested (mProp), considering the MATV initial overall mass and the change of velocity requested to accomplish the mission purposes. Dividing this value for the burn time duration ( ) – due to the MATV mission requirements - is possible to know the associated value of the propellant mass flow rate ( ̇ ). At this point, the value of the system power requested (PS) can be computed as: ̇ ̇ [ ]

3.4.5.3 The selection of the Power Source

Through the definition of the reference exhaust gas velocity (V0) (38) it is possible to compute the power level requested to the EPS in order to ensure the selected SI, then the exhaust gas velocity, which maximise the MATV Payload Mass.

√ [ ]

Solving the equation above for the variable , it is possible to estimate the allowable – maximum – value of the power specific mass (kg/W), necessary to ensure the change of velocity requested, in the indicated burn duration time ( ).

[ ]

Knowing the value, the most appropriate off-the-shelf Electrical Power Source – for instance the Solar Array technology, characterised by a specific power (1/ ) of 41 W/kg - can be evaluated and then selected.

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3.4.5.4 Propellant Tank and Power System Masses From NASA Lewis Research Centre “Electric Propulsion System Design Notes” (Mantenieks, 1990), it is possible to extract some analytical relations based on experimental data, from which it is possible to compute thruster efficiency ( ) and thruster system specific mass (not including propellant system, power source and radiators).

[ ]

The first term of equation above is directly related to the thruster system performances. The thruster specific mass ( ) is multiplied by the efficiency value by converting off-the-shelf electric power to thrust power, while the second term is related to the Power Processor (PP), considering the efficiency value of power thruster conditioning. Hence, the sum expressed by the left-hand-side of equation above, represents the Specific Mass for each single complete Thruster System. To compute equation 40 and equation 41, is necessary to introduce the appropriate constants related to different Thruster types and propellants. For the purpose of this work, three main Thruster technologies have been considered and then compared: Hydrogen Arcjet (C = 5, D = 0), Argon Ion (A = -2.024, B = 0.307, C = 4490, D = -0.781) and Xenon Ion (A = -1.776, B = 0.307, C = 123100, D = -1.198).

For each thruster technology, the overall mass of the thruster (mth) and the power processor (mpp) can be simply estimated by multiplying the value obtained from equation (41), by the power output of the power source selected computed through equation (37). The tank mass values – or inert mass - of the propellant systems (mpr) for hydrogen and xenon propellant types, can be estimated through the empirical relations shown below:

3.4.5.5 Thermal Management System Masses The purpose of this system is to radiate the total amount of heat due to the several inefficiencies of the Electric Propulsion System. This amount of Power can be simply calculated by subtracting the power consumed by the thrusters, utilised to accelerate the reaction matter, from the system’s total power The necessary amount of power consumed by the thrusters depends by the propellant type selected, and by the exhaust gas velocity reached by the thruster system itself. Hence, the first step is to compute the kinetic energy of the propellant matter as follows:

The total thermal power that must be radiated is then calculated:

( )

By knowing the power level which is required to be rejected is directly possible to compute the overall mass (mRad) and the size of the Radiators, on the basis of the material and thermal technology selected.

3.4.5.6 Electric Propulsion System Inert Mass The EPS inert mass is simply defined as the sum of the following contributes:

3.4.6 The MATV Electric Propulsion Results The first table below reports the input parameters considered for the feasibility study of the MATV electric propulsion system. The input data directly come from the Mission Analysis and MATV system engineering

SEEDS Executive Summary 09/2014 Page 61 mass budget prediction. Finally, the table shows the results obtained for two main electric propulsion technologies: hydrogen Arcjet engine and xenon Ion engine.

Table 3-16: MATV Input Data Parameters Source Value

Payload Mass ( ) – [tonnes] System Engineering 61

Maximum Initial Mass allowable ( ) – [tonnes] Guess Value 170 Payload Mass Fraction Calculated 0.35 Burn Duration ( ) – [years] Mission Analysis 2 Change of Velocity ( ) – [km/s]* Mission Analysis 8.4

*To allow some margin for the MATV, an increment of 5% is considered in the Change of Velocity

Table 3-17: MATV Electric Propulsion System Results Hydrogen Xenon Ion Parameters Source Engine Engine

Normalised Exhaust Gas Velocity ( ) – [m/s] Equation- (31) 1.3 1.3 Normalised Characteristic Speed ( ) Equation-(32) 0.456 0.456

Reference Exhaust Gas Velocity (V0) – [m/s] Equation (34) 18,415 18,415

Optimised Exhaust Gas Velocity (Veo) – [m/s] Calculated 23,939 23,939

Computed Exhaust Gas Velocity (Ve) – [m/s] Equation-(35) 9,810 23,939

Specific Impulse (ISP) – [sec] Equation-(35) 1,000 2,444 Thruster Efficiency ( ) Equation-(40) 0.4 0.62

Propellant Mass requested ( ) – [tonnes] Equation-(33) 97.8 50.3

Maximum Flow Propellant Mass requested ( ̇ ) – [g/s] Equation-(36) 1.5 0.8

Propellant Tank Mass estimation ( ) – [tonnes] Calculated 48.8 40.3

System Power Requested (PS) – [kW] Equation (37) 190 370 Maximum allowable Specific Mass ( ) – [kg/W] Equation (41) 0.44 0.11 Thruster Specific Power Requested ( ) – [W/kg] Equation (41) 2.3 8.7

Power Distribution System Mass (mD) – [kg] Calculated 933 3,979

Propellant Kinetic Energy ( ) – [eV] Equation (44) 0.5 392

Thermal Power to be rejected (PREJECT) – [kW] Calculated 112 141

Radiator Mass estimation (MRAD) – [kg] Calculated 45 56.4 Degradation at EOL Assumed 5% 5%

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Solar Array degradation at EOL- [kg/W] Calculated 8.4*10-4 8.4*10-4

Solar Array Mass estimation ( ) – [kg] Calculated 156 309 2 Minimum Solar Panel Area ( ) – [m ] Calculated 1,166 2,309 2 Additional Solar Panel Mass due to degradation ( ) – [m ] Calculated 156 308 Dry Mass Electric Propulsion System [tonnes] 147.6 58.4

Total Mass of Electric Propulsion System [tonnes] 212.6 128.2

3.4.7 The MATV Electric Propulsion conclusion Considering the results obtained the xenon-ion electric propulsion technology results to be the most suitable propulsion system for the purpose of the mission investigated. Starting from an estimated value of the MATV overall mass of 170 tonnes, for this technology, the feasibility study indicates a final value of 128.2 tonnes. Considering the initially specified parameters of the mission, the main critical design element remains the value and the technology selected for the solar arrays, which having an estimated area compared with the ISS value.

3.5 OVERALL MISSION PARAMETERS

3.5.1 CIV Mission Parameters Finally, using the mission concept as well as the propulsion system parameters, it is possible to estimate the overall parameters of the CIV. The propulsion system requirements for the entire mission was developed using a consistent calculation sheet, which summarises all manoeuvres, including margins and ΔV re uired for correction and mid-course manoeuvres. It was developed using an anti-chronological approach, using the following approach. For each mission phase, the final mass of the spacecraft is calculated as the sum of the CIV (e.g. the dry mass of the CIV for the final mission phase), plus the summed masses of the propulsion system tanks and systems, which are separated at the end of this phase:

Furthermore, the total ΔV re uirements is the sum of the ΔV necessary for the main manoeuvres as given by the earlier analysis and additional ΔV for mid-course manoeuvres and station keeping:

Finally, certain phases of the mission require the CIV to perform manoeuvres while being attached to additional modules, noticeably the PhL and the LAB:

Using an estimation of the ISP for the given manoeuvre, it is possible calculate the overall CIV mass, according to the rocket equation:

( )

The value now serves as payload for the prior mission phase. Using these results, the wet masses for the CIV were found, as given in Table 3-18.

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Table 3-18: Mass Estimate of the CIV Mission Phase Mass [t] Included uncertainty

Earth departure 4000 20% Mars capture orbit 640 20% Transfer to Earth 270 20%

3.5.2 MATV Mission Parameter The ΔV re uirements of the ATV were derived in Chapter 3.2.5, while the dry mass and propulsion system budgets were given in Section 5 and Section 3.4.5 respectively. The following table derives the main parameters of the MATV. Mission Phase Mass [t] Included uncertainty

Earth departure 120 20% Mars orbit operation 60 20%

3.6 PLANETARY PROTECTION

3.6.1 Planetary Protection of Mars In addition to the graveyard orbit there are several procedures to be followed in order to avoid biocontamination.

The LAB will have decontamination showers, a UV room and two airlocks to implement biosafety level 4. This will protect the CIV from possible life forms in the Phobos samples. The design of a similar biosafety mobile laboratory has already been implemented on Earth with reasonable power consumption (Vital Probes, Inc. - MMTL Introduction). All samples brought back to Earth will be carried in strictly sealed containers designed to withstand an impact in case of failure during landing, even though a crash is not expected thanks to the use of reliable, widely used crew entry vehicles such as the Soyuz spacecraft.

The MATV service module requires sterilisation prior to launch. At the end of the nominal mission phase it will be disposed of in the science orbit, which has a low periapsis of 298 km. It may perform destructive atmospheric entry or eventually crash on the surface of Mars long after the end of its operational lifetime.

3.6.2 CIV Decommissioning The CIV will return to Earth in a hyperbolic trajectory. While the crew re-enters, the CIV will perform a slingshot with a perigee altitude of 245 km and a hyperbolic excess speed of 4 km/s. This will result in a later heliocentric orbit that allows for a low delta-v manoeuvre of 100 m/s, resulting in a 14:15 orbital resonance with Earth. Even though it intersects the Earth's orbit, it is assumed that by using n-body analysis it is possible to design an orbit which is stable over long time periods. The orbital parameters are , , . Alternatively, a circularization manoeuvre can prevent intersection with Earth's orbit, but this has not been chosen due to the larger delta-v requirements.

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3.7 CRITICAL ISSUES Critical issues have been discovered during mission analysis and these are classed as problems which cannot be solved with current technology. These issues will require major technological research and advancement in order to be used in this mission. The critical issues which have been identified during this mission analysis are as follows:  High speed re-entry of the crew at Earth using the Crew Return Vehicles requires a higher level of detail.  Development of a space qualified, mega-watt scale, nuclear engine including required materials  Design and efficiency of solar arrays required for electric propulsion  Development of qualified Solar Electric Propulsion technology  The low TRL of aerobraking, which could save large amounts of propulsive fuel  Biological protection – risk of forward contamination of Phobos by Astronauts or equipment  Studies of psychological and physiological impacts of long spaceflight mission, current studies to not last as long as the Orpheus mission  The boil-off problem of storing cryogenic propellant in space, with reference to performing a longer stay mission

3.8 CONCLUSION As a result of an extensive development process, a short stay mission concept using two main spacecraft was chosen for Orpheus. The following table summarises its main properties: Table 3-19: Parameter of the overall mission profile Mission Orpheus Summary Specifications CIV Launch date May 2036 MATV Launch date 2034 Overall duration of the CIV mission 602 days Mars vicinity duration of the CIV 35 days Total crew size 6 Landing crew size 2 Aerocapturing No CIV trajectory profile Venus fly by for approach of Mars Lambert transfer for return MATV trajectory profile Continuous low thrust trajectory CIV departure mass 4000t MATV departure mass 120t Phobos landing mission duration 2 landings of 3 days

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3.9 REFERENCES ANGELO J. A. , BUDEN D. (2005) Space Nuclear Power. Florida: Orbit Book Company, Inc.

R. Armellin, P. Di Lizia, G. Di Mauro, M. Rasotto, M. Landgraf (2014) DISPOSAL STRATEGIES FOR SPACECRAFT IN LAGRANGIAN POINT ORBITS http://www.researchgate.net/publication/261437751_DISPOSAL_STRATEGIES_FOR_SPACECRAFT_IN_L AGRANGIAN_POINT_ORBITS BONNET R. M. (2004) The Aurora Programme, Noordwijk: ESA Exploration Advisory Committee BOROWSKI, S. K. (2003) Bimodal Nuclear Thermal Rocket (BNTR) Propulsion for Future Human Mars Exploration Missions. Cleavland: NASA Seal/Secondary Air System Workshop.

BROWN, C. D. (1998) Spacecraft Mission Design. 2 Ed. Reston, Virginia: American Institute of Aeronautics & Astronautics, Inc. BRUNO, C. (2005) Nuclear Space Power and Propulsion Systems. Volume 225. Progress in Astronautics and Aeronautics. Virginia: American Institute of Aeronautics and Astronautics (AIAA).

COMMITTE ON SPACE RESEARCH (2008) COSPAR Planetary Protection Policy [Online] https://cosparhq.cnes.fr/Scistr/PPPolicy(20-July-08).pdf [Accessed June 2014] CUCINOTTA F. A. , KIM Y. MYUNG-HEE (2012) Space radiation cancer risk projections and uncertainties. NASA/TP-2013-217375. Houston, Texas: NASA. Encyclopedia Astronautica - Nitrogen Cold Gas Thruster Module - [Online] Available from: http://www.astronautix.com/engines/colodule.htm [Accessed June 2nd, 2014] Exercise Prescription on the Internet - Body Segment Data [Online] Available from: http://www.exrx.net/Kinesiology/Segments.html [Accessed May 16th, 2014], on turn based on: Plagenhoef, S., Evans, F.G. and Abdelnour, T. (1983) - Anatomical data for analyzing human motion. Research Quarterly for Exercise and Sport 54, 169-178. HUMBLE, HENRY, LARSON J. (1995) Space Propulsion Analysis and Design. 1st Ed. McGraw-Hill Companies.

Humble, Ronald W. et al (1995) - Space propulsion analysis and design - ISBN 0-07-031320 - McGraw-Hill companies Inc. New York IADC (2002) Report on the IADC Activities [Online], Vienna: Inter-Agency Space Debris Coordination Committee ILIN A. V. , CHANG DIAZ F. R. , GLOVER T. W. , CARTER M. D. , CASSADY L. D. , WHITE H. (2012) Nuclear Electric Propulsion Mission Scenarios using VASIMR Technology. [Online] http://www.lpi.usra.edu/meetings/nets2012/pdf/3091.pdf [Accessed June 2014] International Shoe Size Converter Charts [Online] Available from http://www.shoesizingcharts.com/ [Accessed May 16th, 2014] JOHNSON S. A. , BADHWAR G. D. , GOLIGHTLY M. J. , HARDY A. C. , KORANDI A. , YANG T. C. (1993) Spaceflight Radiation Health Program at the Lyndon B. Johnson Space Centre. NASA/TM104782. Houston, Texas: NASA. JORDON, J. F. (1964) The Application of Lambert’s Theorem to the Solution of Interplanetary Transfer Problems. California: Jet Propulsion Laboratory. LARSON W.J., WERTZ J.R. (2005) Space Mission Analysis and Design. Dordrecht, Boston, London: Kluwer Academic Publishers.

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LONGMIER B. W. , SQUIRE J. P. , OLSEN C. S. , CASSADY L. D. , BELLENGER M. G. , CARTER M. D. , ILIN A. V. , GLOVER T. W. , McCASKILL G. E. , CHANG DIAZ F. R. , BERING E. A. (2012) Vasimir VX-200 [Online] http://pepl.engin.umich.edu/pdf/AIAA-2012-3930.pdf [Accessed June 2014] NIEHOFF, J. C. (1966) Gravity Assisted Trajectories to Solar System targets. Vol.3. Issue 9. pp. 1351-1356. Chicago: Journal of Spacecraft and Rockets. NASA (1994) - Space Suit Evolution: From Custom Tailored To Off-The-Rack - [Online] Available from: http://history.nasa.gov/spacesuits.pdf [Accessed May 16th, 2014] NASA (2005) Aerocapture Technology, NASA facts, Huntsville: National Aeronautics and Space Administration NASA (2010), A New Space Enterprise of Exploration, Flagship Technology Demonstration-4, Aerocapture, Entry, Descent & Landing. [Online] http://www.nasa.gov/pdf/458812main_FTD_AerocaptureEntryDescentAndLanding.pdf [Accessed May 2014]. NASA (2012) The MSL Spacecraft, NASA, [Online] http://mars.jpl.nasa.gov/msl/mission/spacecraft/, [Accessed June 2014] NASA Solar System Exploration – Phobos, facts & figures [Online] Available from: http://solarsystem.nasa.gov/planets/profile.cfm?Object=Mar_Phobos&Display=Facts [Accessed May 15th, 2014] SEEDS edition V students (2011) - AENEA: humAn Exploration of a Near Earth Asteroid, Ch. 5: Robotic system - [Online] Available from: http://www.seeds-master.eu/Executive_summary_PW10_Chapter%205-6.pdf Other chapters: http://www.seeds-master.eu/Executive_Summary_PW10_Chapter%201-4.pdf http://www.seeds-master.eu/Executive_summary_PW10_Chapter%207-8.pdf http://www.seeds-master.eu/Executive_summary_PW10_Chapter%209.pdf [Accessed May 14th, 2014] SSC Group - ECAPS CAPABILITIES - [Online] Available from: http://sscspace.com/capabilities-3 [Accessed May 19th, 2014] WILLIAMS, H. (2014) Applied Nuclear Power. Lecture Number 3. Leicester: Leicester University.

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4 SCIENCE The approach used to determine the scientific payload and its platform was to take the scientific objectives as defined in the Introduction and to derive a Science Requirements Matrix (SRM). From the SRM the potential scientific platforms and the possible techniques that could fulfil the requirements were collated, from this document a suite of instruments and platforms was selected that met the main objectives whilst remaining within budget and completing secondary goals. In the following chapter this process, the final payload configuration and the sample selection process are discussed amongst other topics related to the scientific investigation.

4.1 OBJECTIVES Below is a summary of the main scientific objectives and secondary goals.

Objectives: Science Objectives: 1. Collect data in order to support the selection of appropriate future landing sites for humans on Mars 1.1. To characterise the topography of potential Mars landing sites for humans 1.1.1. Analyse and characterise rocks size and abundances in areas of interest 1.1.2. Analyse and characterise slopes in the areas of interest 1.1.3. Investigate the properties of the regolith 1.1.4. Investigate the existence, location and properties of subsurface caves 1.2. To study the level of radiation on the surface of Mars for future human habitability 1.2.1. Study the level of radiation on the surface of Mars 1.2.2. Study the level of radiation below the surface of Mars 1.3. To characterise the atmospheric conditions of Mars 1.3.1. Investigate the dynamics and variability of Mars' weather 1.3.2. Investigate the atmospheric chemistry of Mars 2. To search for ISR (In-Situ Resource) sites on Mars 2.1. To search for a site where water ice is accessible 2.2. To search for CO2 ice

3. To investigate the effects of long term spaceflight on humans. 3.1. Study the changes to muscle mass over mission duration. 3.2. Study the changes to bone density over mission duration 3.3. Study the changes to cardiovascular performance over mission duration 3.4. Study the changes to the central nervous system over mission duration. 3.5. Study the changes to the psychology of humans during mission duration. 3.6. Study the impact of radiation dose on humans during mission duration

4. To study the origin of Phobos

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4.1. To characterize the surface of Phobos 4.1.1. To investigate the composition of the surface 4.1.2. To elaborate topographic mapping of the surface 4.2. To study the gravitational environment of Phobos 4.2.1. To elaborate upon the gravitational map of Phobos 4.3. To characterize the internal structure of Phobos 4.3.1. To investigate the surface structure of Phobos 4.3.2. To investigate the internal structure of Phobos

5. To search for ISR on Phobos. 5.1. To characterise the surface properties of Phobos 5.1.1. Investigate the composition of the regolith / chemical properties of the regolith 5.1.2. Characterise the topography of Phobos 5.1.3. Investigate the mechanical properties of the regolith 5.1.4. Investigate the distribution of different elements and compounds across the surface. 5.2. To characterise the interior composition of Phobos 5.2.1. Investigate the density of the interior

Technological Objective: 6. Develop ISRU (ISR Utilization) technology Additional to the primary objectives a number of secondary goals were outlined, these are scientific investigations that are not paramount to the success of the mission or development of scientific models and data that are of specific interest. Rather they are scientific investigations which would provide useful information particularly to the habitability of Phobos and Mars in general not specifically aiming at human habitability. Below are the goals that were achievable using the instruments selected to complete the mission objectives. Goals: 1. Determine the habitability of Mars 1.1. Search for evidence of surface liquid water and water ice 1.2. Determine the nature and inventory of organic carbon compounds 1.3. Identify features that may represent the effects of biological processes 1.4. Identify suitable geological habitats 1.5. Characterize the local meteorology to provide a comparison with data from orbital measurements 2. Search for evidence of extant life on Mars. 2.1. Determine the nature and inventory of organic carbon compounds 2.2. Identify features that may represent the effects of biological processes 3. Search for evidence of extant life on Phobos. 3.1. Search for organic carbon compounds on the surface.

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3.2. Search for organic carbon compounds near the surface (0.1-2 meters). 3.3. Search for compounds that indicate biological processes have occurred.

4.2 REQUIREMENTS From the top level Science Objectives from the list in 4.1 an SRM was formed with a number of second tier science objectives that drove the choice of necessary science investigations. These investigations had their own measurement objectives and some options of techniques to meet these objectives. The SRM, which can be found in the appendix, does not include the technological objective or the human psychology and physiology requirements, the human science investigations are discussed in 4.6.6 Human Studies.

4.3 INSTRUMENT PLATFORMS, TECHNIQUES AND SELECTION The next step in defining the scientific investigations was to further define the instrument payloads and the various platforms they would be mounted to. Due to the definition of the mission scope and the available craft there are several platforms that are an intrinsic part of the mission. These are: remote sensing capabilities on the MATV, the ability to transport and manually deploy instrumentation on the surface of Phobos, and a remote sensing payload mounted to the exterior of the lab for remote sensing on the surface of Phobos. The unknown platform is the landed payload on the Martian surface, from the early stages of the project a large amount of mass was reserved in budgets to be utilised as landed payload. The following chapter follows the decisions in choosing the optimal suit of instruments for the pre-defined platforms and the iterative process of making initial investigations into what instruments may be useful for meeting the set science objectives and what effect on the trade-off of types of platform it has. Although the following sections follow instrument selection first and assess the choice of Mars landed platform afterwards, it should be understood that both processes were taking place side by side. Upon completion of the SRM a new document was formed to highlight the instrument techniques that could fulfil the objectives, once it was completed a selection process was completed to help choose the instrument techniques required to achieve the objectives. This approach took into account several key factors listed below from most to least important:  Objective importance  Quality of results  Landed or orbital payload  Other factors Once the table was constructed it allowed for an objective numerical approach to instrument selection to occur, this aimed to select the best instruments in terms of cumulative mass, technique, result quality and objective importance. The output of the selected instruments can be found as part of each payload configuration, for each respective payload see:  Table 4-1:ExoMars Class Rover Instrument Payload  Table 4-2: Network Rover Instrument Payload  Table 4-3: MATV Scientific Instrument Payload  Table 4-4: CIV/Orbital Lab Remote Sensing Instrument Payload

Instrument Platforms With a set of objectives to obtain, budgets to adhere to and an idea of the likely techniques to fulfil the objectives, a trade off to determine what payloads to land on the surface of Mars was performed. The final choice comes in the form of what type of payload was to be landed on Mars.

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The trade-off was conducted by identifying the main Mars objectives that depend on landed payload assigning an importance which is reflected in the objective weighting that is applied. Another weighting was applied, this was a technique such that it assesses the effectiveness and ease that the objective can be completed. A large number of payload types and combinations were derived from one large rover to a large network of small static landers. If a payload category could meet the objective it scored a 1 if not it scored a 0, its score was then multiplied by both the objective and technique weighting factors. The trade-off considered many platform types, these included;  Minor Rover Network  Minor Lander Network  Major Lander Network  Major Lander  MER Class Rover + Minor Lander Network  MER Class Rover Network  Major Rover The output of the trade-off drove towards networked measurements and larger rovers but the clear winner was the Minor Rover Network which consisted of three rovers of approximately MER class size. However, as the design matured it became clear that the option of one larger rover and several smaller rovers had been neglected, which ended up fulfilling more science requirements. For expanded information see section 4.5 Rover Mass Calculations.

4.4 PAYLOAD CONFIGURATIONS The final payload configurations were now possible to define, below are several tables that describe the instruments present and the objectives they meet for each instrument platform. Due to the limited scope of this project it was only possible to use previously launched or proposed instruments as a reference that could potentially meet the science requirements. The instruments chosen were selected based on suitability to the objective required to be met, choosing more modern instruments where possible and if at all possible instruments that could meet multiple objectives. Using the reference instruments, a mass and power budget could be generated. Margins were applied to values; the extent of the margin applied is based on heritage, suitability and simplicity. See parametric analysis of rover mass in 4.5 Rover Mass Calculations for rover margins.

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Table 4-1:ExoMars Class Rover Instrument Payload

Reference Instrument Type Mass (kg) Power (W) Reference Instrument

Weather WUSTL (2014), Space Sci Rev 1.3 1 REMS Measurements (2012)

Radiation 1.5 4.2 RAD WRMISS (2008), NASA (n.d) Measurements

Subsurface 43rd Lunar and Planetary Science radiation and 2 3.5 MOLE/HP3 Conference (2012) thermal probe

Stereo Imaging 0.9 6 MASTCAM DiBaise, D. (n.d)

Laser Induced Lunar and Planetary Science XXXVI Breakdown 5.62 3 CHEMCAM (2005) Spectroscopy

Sample Collection ExoMars 78 70 Baglioni, P. (2013) and Analysis Sample Analysis

Total 89.3 87.7

Table 4-2: Network Rover Instrument Payload Mass Power Reference Instrument Type Reference (kg) (W) Instrument

WUSTL (2014), Space Sci Rev Weather Measurements 1.3 1 REMS (2012)

Radiation Measurements 1.5 4.2 RAD WRMISS (2008), NASA (n.d)

Subsurface radiation and 43rd Lunar and Planetary Science 2 3.5 MOLE/HP3 thermal probe Conference (2012)

Stereo Imaging 0.9 6 MASTCAM DiBaise, D. (n.d)

Laser Induced Breakdown Lunar and Planetary Science 5.6 3 CHEMCAM Spectroscopy XXXVI (2005)

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ISRU Demonstration 8.5 22.5 MIP Kaplan, D. et al. (2001)

Total 19.8 40.2

Table 4-3: MATV Scientific Instrument Payload Reference Instrument Type Mass (kg) Power (W) Reference Instrument

Gamma Ray 30.5 32 GRS + HEND Boynton, W. et al. (2004) Spectroscopy

Raman Spectroscopy 33 47.3 N/A N/A

UV Spectroscopy 27 - IUVS Blau, P. (2011)

IR Spectroscopy 33 47.3 CRISM JHU/APL (2014)

HD Imaging 65 60 HiRISE McEwen, A et al. (2007)

Context Camera 10 20 - Estimate

Ground Penetrating 4 25 EMIS Samson, C. et al. (n.d) Radar

IR Imaging 11.2 14 THEMIS Christensen, P. et al. (2001)

Total 213.7 245.6

Margin 1.2 1.2

Total with margin 256.4 294.7

Table 4-4: CIV/Orbital Lab Remote Sensing Instrument Payload Mass Power Reference Instrument Type Reference (kg) (kg) Instrument

Boynton, W. et al. Gamma Ray Spectroscopy 30.5 32 GRS + HEND (2004)

Raman Spectroscopy 33 47.3 N/A N/A

Smith, D. Zuber, M. Laser Altimetry 9.6 30 LOLA (n.d)

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Part of K-Band Ranging 5 - communications N/A +5kg.

McEwen, A et al. HD Imaging 65 60 HiRISE (2007)

Samson, C. et al. Ground Penetrating Radar 4 25 EMIS (n.d)

Christensen, P. et IR Imaging 11.2 14 THEMIS al. (2001)

Total 158.3 208.3

Margin 1.2 1.2

Total with margin 190.0 250.0

4.5 ROVER MASS CALCULATIONS Once the mass budget had been further defined it was possible to size the rovers in order to maximise the scientific capability of the rovers. The initial design of the rovers was such that they were all of the approximately the same size but through an iterative process of assessing rover size and the desired instruments it was found that it would be preferable to have two smaller Mars Exploration Rover class rovers (Network Rovers) and one larger ExoMars class rover, as the instruments comprising the sampling suite were complementary to each other and therefore needed to be done at the same site on the same samples. Using a parametric approach the overall rover payload mass including descent stages was calculated. This was done for the two smaller rovers by finding data from previous Mars rovers and using the parametric equation gained from this data to size the rover based on the instrument mass. The rover mass then allowed the entry decent and landing (EDL) system to be sized, again based on a parametric analysis. The issue with this method is that due to the few numbers of past systems, there is not enough data to make the parametric analysis a reliable estimate. It is for this reason that a margin of 1.5 has been included in both parametric analyses. For the Exomars class rover, the instrument mass is almost the same as that of Exomars, so it is reasonable to assume that the mass of the rover will be similar to that of Exomars, hence the margin applied is lower. The results can be found in the table below. Table 4-5: Mars Deployable masses Rover sizing Rover Mass + EDL mass : from EDL sizing from Including rover Instrument Instruments inc Rover Mass inc uncertainty and EDL Rover Type Mass (kg) Margin uncertainty (kg) Margin Uncertainty (kg) uncertainty (kg) ExoMars Class 89.3 1.2 372 1.2 1434 1806 Network Rover 19.8 1.5 325 1.5 1246 1571

As described by the mission analysis section the rovers are to be deployed by the MATV. Before the mission is launched a number of provisional landing sites will be selected (discussed in 4.6.1) and upon arriving in the science orbit the MATV will perform remote sensing on the sites to collect data to be used in conjunction

SEEDS Executive Summary 09/2014 Page 74 with the landed payload. The rovers will then land and conduct the desired scientific investigations, for an approximate duration of 3 years.

4.6 ADDITIONAL SCIENTIFIC STUDIES Alongside the main studies to drive the instrument selection and the science to be completed there were a number of other studies performed in order to fully define the mission, which follow.

4.6.1 Mars landing sites As three rovers will land on Mars surface, a study of the possible landing sites has to be taken into account in order to maximize the achievement of the scientific objectives already set. Mars has an extensive surface area, but only some places may be suitable for landing. In order to choose the proper sites, some criteria have been set to consider a location to be a landing site:  Safe from surface hazards such as rocks and boulders.  Good place to achieve the objectives of the mission.  Has been already mapped in detail.  No slopes.  Protected from dust storms.  Within +30/-30 degrees of the equator due to delta-v limitations. Taking into consideration these characteristics and reviewing previous studies on Mars, some spots have passed the selection process and will serve as possible landing sites for the Orpheus mission. These landing sites are purely prospective and it is expected that a larger number of sites should be taken into account and undergo a proper scientific landing site selection process in a later phase if this mission was to go ahead.

4.6.1.1 Hellas Planitia It is the deepest basin registered on Mars. Its borders start from 29 degrees south from the Martian equator, which is inside the range of the landers capability (+30/-30 degrees) which was set as the limit on the fuel expenditure of the craft. At this basin, the pressure is the highest registered on the red planet, this allows for easier production of water from water ice (NASA 2000). This characteristic would allow the future human landed crew to have liquid water with fewer complications than in other places of the planet.

4.6.1.2 Melas Chasma In the Valles Marineris, it is the largest canyon in the planet. Its main features are the thick layered outcrops, which will help to study the composition of mars; the abundance of hydrated minerals and the existence of fluvial features. At this place it is known to be a massive ~1100 km2 deposit of kieserite (mineral used as fertiliser on Earth) and poly-hydrated sulphates (Weitz et al., Chojnacki and Hynek 2006).

4.6.1.3 Athabasca Valles Athabasca valley was one of the proposed landing sites for the MER mission. It is theorised to have ground water (past or present) that erupted onto the surface from Cerberus Fossae. It is a place where great science can be achieved in small extension, which makes things easier for a rover. It also meets the landing safety requirements needed for future human missions (McEwen 2001).

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Figure 4-1: Landing sites map (Encyclopaedia Brittanica) Moving from the left of Figure 4-1: Landing sites map (Encyclopaedia Brittanica)the landing sites are shown respectively as Athabasca Valles, Melas Chasma and Hellas Planitia.

4.6.2 Orbital Laboratory The orbital lab was designed to select the optimal samples for Earth return as the system is mass limited in the mass that can be returned to Earth. The lab is also used for preliminary science investigations and biological protection. The lab is modelled as being similar in mass, volume and power to the Columbus module on the ISS, having a structural mass of 10.3 tonnes before margin allowing for 4.7 tonnes of interior scientific equipment, biosafety measures and remote sensing to be attached this brings the total mass including margin to 18 tonnes. A configuration was devised where the lab met biological safety requirements for both the astronauts and any samples or equipment brought aboard. The lab is designed to have several instruments which would conduct preliminary scientific investigations into the samples, the data from which will be transmitted back to Earth. The findings from this investigation into the extracted samples would allow for the selection of the most scientifically valuable samples for the mission’s limited sample return mass of 60kg. The devices used were optimised for limited space rather than mass, as use of the MATV means that volume becomes the limiting factor. These instruments overlap to some degree with the handheld analyser tool used by the astronauts on the surface; however the measurements made in the lab are much more accurate due to the increased complexity of the instruments on board. Below is a table of the components of the lab. Table 4-6: Laboratory Configuration Breakdown

Mass with Instrument Type Mass (kg) Uncertainty uncertainty (kg)

Raman 32.92 1.2 39.504 X-Ray 40 2 80 Mass spectrometer 5 1.2 6 SEM 400 1.2 480 Optical Microscope 80 1.2 96

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Decontamination zone 32 2 64

Samples container 60 2 120

Sample analysis 30 2 60

Total mass excluding Total mass with 679.92 945.504 uncertainty (kg) uncertainty (kg)

It is clear from the results of Table 4-6: Laboratory Configuration Breakdown and Table 4-4: CIV/Orbital Lab Remote Sensing Instrument Payload that the expected payload mass of the lab is around 1.5 tonnes, significantly under the mass budget of 4.7 tonnes.

4.6.3 Earth Curation Facility Upon safe return of the samples to the Earth’s surface they will be taken to a curation facility, this should be a site designed solely for receiving of the samples and maintaining the highest levels of biosafety. The facility shall have all the necessary procedures and equipment to receive samples, separate into different smaller samples ready for distribution to institutes where the samples can be analysed by state of the art equipment in order to gain the largest possible scientific return.

4.6.4 Phobos Landing Sites In order to best maximise scientific return it was important to visit at least two scientifically interesting sites on Phobos, therefore to extract the most valuable samples it was important to select two sites that had high probabilities of containing interesting samples, different types of samples and were safe places for humans to land. The surface of Phobos is diverse, which makes the selection of a scientific landing site non-trivial. The factors considered for the choice of the landing site are both operational and scientific and are shown below:  Presence of craters,  Grooves,  Mars ejecta,  Abundances of dust,  Light condition on the surface,  Communication with CIV and ground stations on Earth. To maximize both the safety and the scientific output of the mission, the selection of the precise landing areas will be performed by the ground station after the arrival of the CIV in the vicinity of Phobos, when high resolution pictures and various other remote sensing measurements are available. The crew will also have the ability to select a specific point suitable to land during the landing process, much like on the Apollo missions. Naturally this process would, in the years before the mission, be supported by extensive analysis of potential landing sites, the development of landing site models and careful selection of proposals from the scientific community. For the purpose of this study, a general survey of interesting landing sites is performed, leading to the selection of two landing sites, which are utilized as reference for the further study of the Orpheus mission. In order to select landing sites with the highest feasible scientific output it is necessary to find sites, which fulfil the science requirements as well as the landing capability of the lander. The figure below shows the places of high interest on the Phobos surfaces and the lightning condition of them. Using Hopkins, J. (2011)

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& Murray, J. B. (2006) it is possible to identify several landing sites that can be explored during the 36 days of stay in Phobos proximity. Hopkins provides the surface map of Phobos, given in Figure 4-2 including scientifically interesting sites and areas of continuous sunlight (in yellow) and visibility of Mars (in red for partial visibility and in green for full visibility).

Figure 4-2 Phobos Visibility and Sunlight Scientifically interesting sites can be found on both the east and west sites of Phobos. On the western hemisphere three interesting sites were identified:  The Stickney crater,  The Limtoc crater inside the Stickney crater,  The grooves created by Mars' ejecta. Studies suggest that the Stickney crater including the Limtoc crater give the best access to deeper layer of Phobos material, as well as an insight in the crustal structure, making it a good site for studies concerning the origin of Phobos. Hopkins suggests the structure of Phobos' upper layer as well as its composition with other known celestial bodies is currently assumed to be the most efficient was to study the origin of Phobos. However, in the eastern dorsa of Phobos (e.g. the Kepler Dorsum) and mainly in the grooves on those dorsa, it is believed that Mars ejecta can be found, making the grooves a good location for a better study of the Red Planet, possible biosignatures and the interaction between Mars and its moon. On the eastern side the following areas of interest were analysed:  The Zunil crater  The Tooting crater

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Due to difficulties in calculating the landing ellipse of the Phobos lander it is not possible to precisely decide the exact point of landing. However, the given mobility of Astronauts as well as the extensive time of each landing mission is assumed to allow taking sample at a pre-defined site. As pointed out earlier, this report will only recommend two landing sites, while the actual landing site selection would be performed using detailed pre-mission analysis as well as during the mission by updating the selection models with available remote sensing data. The following two landing sites for the two manned landings are chosen as reference:  Stickney crater, with possibility of exploring the Limtoc crater  Zunil or Tooting crater, very likely the zone with the most important concentration of Mars ejecta

These sites have been selected on their ability to likely provide samples to analyse the origin of Phobos and analyse the potential mass ejecta. Below is an adapted version of Figure 4-2 Phobos Visibility and Sunlight however this figure only focuses on the landing sites selected which are highlighted by green circles.

Figure 4-3 Selected Landing Sites

4.6.5 Sample Selection Process The selection of the most scientifically valuable samples for detailed analysis on Earth, this is done using a four step process of Identification, Collection, Evaluation and Transportation.

1. Identification This is the process by using existing data on Phobos, remote sensing and the astronauts using the handheld scanning tools. This is to increase the percentage of useful samples selected by the astronauts for further sampling. 2. Collection The selected sample is then removed using a variety of simple tools and the samples are placed into suitable biosafe containers. The sample containers and astronauts then pass through a decontamination area to stop cross contamination of either Phobos or the craft. Upon docking with the lab (attached to the CIV) the samples are passed to the lab for analysis. 3. Evaluation The samples removed from Phobos are then screened to select the most scientifically valuable samples, the data collected is transmitted to Earth for further analysis. This is done using the instrument suite in the lab,

SEEDS Executive Summary 09/2014 Page 79 as defined above. Once the samples have been selected (up to the limit of 2 x 30kg, the space for samples in the two return capsules) they are sealed into suitable containers for return to Earth. 4. Transportation The sample containers are then stored for the return trip to Earth and return to the Earth’s surface inside the crew re-entry vehicle. When the samples have been retrieved from the landing site they will be taken to the sample curation facility where they will be split into smaller sample sizes if necessary and then distributed to the necessary research facilities that will conduct the in depth analysis.

4.6.6 Human Studies One of the objectives of the mission is to study the effects of long term spaceflight on human physiology and psychology. This is expected to be mainly done through observations and simple tests during the mission and rigorous assessments pre and post mission. The system shall include a Crew Healthcare System as defined in the ECLSS sub-system to provide healthcare, treatment and environmental monitoring in order to maintain astronaut health. Some of these pieces of equipment and on-board health checks will provide useful insight into the physiological effects of long-term spaceflight. The astronauts will be exercising as part of their regular routine in order to minimise muscle loss however a mission of this duration has never been undertaken in microgravity and the final condition of the astronauts musculoskeletal system is one of the most important investigations of the mission as well as changes to the cardiovascular system and organs. Therefore the majority of the equipment for human studies will be part of the ECLSS system and take the form of regular health checks. Other aspects such as psychology will be monitored using computer programs and communications with Earth, as well as pre and post flight analysis by psychologists. The final aspect of the human studies is radiation dose, each astronaut will be fitted with a radiation dose monitoring device which will remain with them for the entire mission duration.

4.7 CRITICAL ISSUES The science investigations utilise a number of existing technologies, however a few new technologies would need to be developed and these are discussed below. Remote Raman Spectrometry This involves the use of Raman Spectroscopy at significant range; in this case this would be at altitudes of the order of hundreds of kilometres. Theoretically the device would behave as most Raman spectrometers however the optical component is a telescopic lens to focus on the distant surface. The laser required would need large amounts of power which is available due to the large amount of power available in Mars orbit from the MATV resulting from the propulsion system choice. Another issue with orbital Raman spectroscopy is the pointing. As the laser beam would cover a very small surface area of the target object and it is travelling at large speeds it is important that an advance pointing system be present. Although there are a number of experiments into using ranged Raman spectroscopy these have demonstrated at distances up to 120m (Sharma, S. 2007). Therefore one of the required developments for the mission is the completion of a stand-off remote sensing Raman Spectrometer. Handheld Analysis Tool This is the device to be used by the astronauts in order to get in situ quick analysis on a potential sample to be able to differentiate between samples so as to select the most appropriate. The device is expected to be similar to existing handheld Raman spectrometers and XRF spectrometers for the general market. Therefore with the relevant funding a flight qualified version would be possible to develop. Sub-surface Radiation Probe One of the desired instruments is that of a sub-surface radiation probe, initially this device was a variant of previous thermal probes such as: Beagle 2’s PLUTO; the descoped HP3 from Exo ars now proposed to be launched as part of NASA’s 2016 Insight mission; and the larger proposed oon and ars drill MMUN. These probes, alongside others in development, provide a simple drilling technique that requires long

SEEDS Executive Summary 09/2014 Page 80 operation times but low power requirements which have been speculated to drill down to depths of 5m (EPSC 2001). With suitable increases in the size of the probe and the development of the necessary detectors it was deemed possible to attach a simple radiation probe that could measure the total ionising dose with depth. This would give values for the attenuation of radiation with depth below the Martian surface whilst the thermal probes gave data on thermal conductivity and temperature. It was thought that with time a more comprehensive radiation probe could be developed giving more accurate results and providing data on the species of incoming particles. Therefore the development of a probe to attach to a HP3 mole like probe would be useful in attaining some of the mission’s objectives

4.8 REFERENCES WUSTL, 2014. Rover Environmental Monitoring Station [Online] http://an.rsl.wustl.edu/mer/help/Content/About%20the%20mission/MSL/Instruments/REMS.htm

Space Sci Rev, 2012. REMS: The Environmental Sensor Suite for the Mars Science Laboratory Rover [Online] http://cab.inta-csic.es/rems/wp-content/uploads/2013/04/REMS.-Space-Science-Reviews-2012.pdf

WRMISS, 2008. Calibration of the Radiation Assessment Detector (RAD) for MSL [Online] http://wrmiss.org/workshops/thirteenth/Hassler.pdf

NASA, (n.d). Radiation Assessment Detector (RAD) [Online] http://mars.jpl.nasa.gov/msl/mission/instruments/radiationdetectors/rad/

43rd Lunar and Planetary Science Conference, 2012. Insight: Measuring The Martian Heat Flow Using The Heat Flow and Physical Properties Package (HP3) [Online] http://www.lpi.usra.edu/meetings/lpsc2012/pdf/1445.pdf

DiBaise, D. no date. A Zoom Lens for the MSL Mast Cameras: Mechanical Design and Development [Online] http://www.infernolab.com/txt/2012AMS_Zoom%20MC%20Lens%20paper_2012-03- 14_as_submitted.pdf

Lunar and Planetary Science XXXVI. (2005) CHEMCAM Science Objectives For The

Mars Science Laboratory (MSL) Rover [Online] http://www.lpi.usra.edu/meetings/lpsc2005/pdf/1580.pdf

Baglioni, P. (2013) Rover development status [Online] http://robotics.estec.esa.int/ASTRA/Astra2013/Presentations/Baglioni_0000003.pdf

D. I. Kaplan et al (2001) The Mars In-Situ-Propellant-Production Precursor (MIP) Flight Demonstration [Online] http://www.lpi.usra.edu/meetings/marsmiss99/pdf/2503.pdf

W.V. Boynton et al. (2004) The Mars Odyssey Gamma-Ray Spectrometer Instrument Suite

Blau, P. (2011) MAVEN Instrument Overview http://www.spaceflight101.com/maven-instrument- overview.html

JHU/APL (2014) CRISM Specifications http://crism.jhuapl.edu/instrument/design/specifications.php

McEwen, A et al. (2007) Mars Reconnaissance Orbiter’s High Resolution Imaging Science

Experiment (HiRISE) https://wiki.umn.edu/pub/ShackeltonCraterProject/CameraInfo/HiRISE_Manual.pdf

Samson, C. (n.d) Electromagnetic Induction Sounder for Subsurface Mapping of Resistivity and Magnetic Susceptibility on Mars [Online] http://www.gve.on.ca/17_Samson.pdf

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Christensen, P. et al. (2001) The Thermal Emission Imaging System (THEMIS) for the

Mars 2001 Odyssey Mission [Online] http://www.mars.asu.edu/christensen/docs/christensen_themis_ssr.pdf

NASA (2000) – Making a splash on Mars [Online] Available from: http://science.nasa.gov/science- news/science-at-nasa/2000/ast29jun_1m/ [Accessed September 3rd, 2014]. Cathy Weitz et al. – Paleo-lake in Melas Chasma as a potential landing site for MSL [Online] Available from: http://marsoweb.nas.nasa.gov/landingsites/msl/workshops/2nd_workshop/talks/Weitz_Melas_MSL_2worksh op.pdf [Accessed August 27th, 2014]. M. Chojnacki and B. Hynek – East Melas Chasma landing site rationale [Online] Available from: http://webgis.wr.usgs.gov/msl/docs/1st_LSWorkshop/Chojnacki_1st_MSL_workshop.pdf [Accessed August 28th, 2014]. Alfred McEwen – Athabasca Valles: Evaluation as MER Landing Site [Online] Available from: http://webgis.wr.usgs.gov/mer/March_2002_presentations/McEwen/McEwen_ASM-MER3-27-02.pdf [Accessed August 27th, 2014]. Encyclopaedia Britannica – Global topographic map of Mars produced from high-resolution laser [Online] Available from: http://www.britannica.com/EBchecked/media/70946/Global-topographic-map-of-Mars- produced-from-high-resolution-laser [Accessed September 3rd, 2014]. J. B. HOPKINS et al. Comparison of Deimos and Phobos as destinations for human exploration and identification of preferred landing sites, AIAA SPACE 2011 Conference \& Exposition, California, 2011 J. B. Murray et al New evidence of the origin of Phobos' parallel grooves from HRSC Mars Express, Lunar and Planetary Science XXXVII, Milton Keynes, 2006 Smith, D. Zuber, M. (n.d) Lunar Orbiter Laser Altimeter Fact Sheet [Online] http://lunar.gsfc.nasa.gov/lola/images/LOLA_Fact_Sheet.pdf

SHARMA, S. (2007) New trends in telescopic remote Raman spectroscopic instrumentation Publisher - Elsevier EPSC. (2011) Measuring Heat Flow on Mars: The Heat Flow and Physical Properties Package on GEMS [Online] http://meetingorganizer.copernicus.org/EPSC-DPS2011/EPSC-DPS2011-379-1.pdf

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5 SYSTEMS ENGINEERING

5.1 INTRODUCTION The purpose of the strict systems engineering approach, which was employed during the development of the Orpheus mission, was to create a common interface for the different subsystem working groups in order to ensure compliance with interface requirements. Furthermore it compared the actual system budgets with budget constraints. The key element of the system engineering approach employed was the head system engineer, who organised weekly meetings with the different entities using a fixed questionnaire to create a common and consistent interface. The systems engineering approach was strictly followed for the development of the Crew Interplanetary Vehicle (CIV) for the purpose of providing a pre-phase-A study of the crewed spacecraft. The other elements of the Orpheus mission, as for example the Phobos Lander (PhL) and the laboratory (LAB) were studied as overall concepts in conceptual feasibility studies only, as given for example in Section 5.5, Section 5.4, and in Section 4.6.

5.2 METHODOLOGICAL APPROACH OF THE CIV STUDY

5.2.1 Standards Formally correct system engineering approach can strongly reduce the overall engineering effort, especially for a field of engineering, which is strongly iterative. For this reason, it was decided to implement standardized approaches throughout the entire study of the Orpheus mission. As reference for this, the guidelines published by the “European Cooperation for Space Standardization” (ECSS) were followed as far as considered reasonable for a feasibility study. The reference documents, which were considered for the different subsystems are mentioned in the respective chapters. For the global system engineering approach, the following documents were found relevant:  ECSS-E-ST-10C System engineering general requirements - March 2009  ECSS-E-ST-10-04C Space environment - November 2008  ECSS-E-ST-10-06C Technical requirements specification - March 2009 Relevant approaches and guidelines were employed throughout the entire system engineering process of the “Crew Interplanetary Vehicle” (CIV) during the Turin phase. Utilization on the development of the “Phobos Lander” (PhL) and the “ ars Automated Transfer Vehicle” ( ATV) was disregarded due to the very limited timeframe of the overall project.

5.2.2 Implementation The guidelines of ECSS were followed as presented here: 1. At the beginning of the work on the Orpheus mission a consistent margin and uncertainty policy as given in the following chapter was defined. 2. For the mass of the CIV including the different subsystems, and their respective uncertainties, a trend analysis was prepared to be able to analyse the evolution of those properties during the design process. This allows maturity estimates of the overall system design by analysing the convergence behaviour. Furthermore, these graphs were provided to the different subsystem design groups to serve as a valuable asset to their design process. 3. ECSS system engineering guidelines, which are relevant and to be respected also by the different subsystem working groups, were identified and communicated to the respective working groups. 4. Being the common interface between the efforts of the different working groups, the system engineering group identified the most important interfaces between the different working groups in

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initial meetings during the second week of the Turin phase. These interfaces, which are to be provided and documented according to ECSS, were added into a formal System Interface Questionnaire (SIQ). For example the most critical design elements and the employed tools for the design process are included. The SIQ was developed as a generic tool; however it was adapted to the specific needs and requirements of the different working groups. An example of a SIQ is given in appendix 10.4. 5. Using the SIQ in weekly scheduled meetings with all working groups, the properties of all interfaces including their uncertainties were updated. 6. Then, the new values were added to the respective budgets and trend analysis tools. In case of major changes, the new values were communicated to the interfacing working groups. 7. Finally, if certain input values, like for example system or interface properties, are urgently needed by a certain working group, the system engineering team communicates this needs to the group providing the value. This was done to ensure an overall smooth and linear design process of the CIV.

5.2.3 Margin and Uncertainty Policy According to the ECSS recommendation, both margins and uncertainties were applied to the subsystem interface properties. Furthermore, for global system budgets, an additional overall system margin was added. The uncertainties of subsystems were chosen according to the respective working group estimations in the SIQ, margins were chosen as given in the table below: Table 5-1: Overview of the applied margin and uncertainty policy Property Value Source

Margin of subsystems estimated by rough 50% Estimation estimation (e.g. by physical principles) Margins for preliminary designed 20% Larson (1999) subsystems (e.g. by similarity analysis) Subsystem margin for preliminary designed 5% Larson (1999) subsystems (sub-subsystem level) Overall system margin (e.g. for CIV) 5% Larson (1999) Dependant on the maturity of the Uncertainties of a system 20%...100% design, by estimation of the subsystem working group

The following figure visualizes the structure of the applied margins and uncertainties.

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Figure 5-1: Margin and Uncertainty approach on system level The following values were considered to represent the mission design, according to Daimer-Benz Aerospace (1996):  Best Engineering Estimate (BEE): Best estimate of the interface value without applying any margins or contingency, except “hidden margins” within subsystems  Contingency Value: Best estimate value with added contingency, assuming the maximum value for the given uncertainty:

 SPEC value: The contingency value including additional subsystem and system margins to obtain the maximum possible interface value taking into account the future detailed design of the system. This value is communicated to the different working groups as interface value to ensure a feasible design if in case of changes of the BEE value.

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5.3 RESULTS OF THE CIV STUDY

5.3.1 CIV Mass Budget The following table displays the mass budget at the end of the feasibility study: Table 5-2: Mass Budget of the CIV BEE incl. SPEC value Item BEE [tons] Margin Uncertainty Margin [tons]

TCS 8.1 20% 23% 9.8 12.0

Structures 49.5 20% 20% 59.4 71.5

EPS 5.0 20% 15% 6.0 6.9

Astronauts 0.5 20% 20% 0.6 0.8

Resources 34.4 20% 20% 41.3 49.5

SPE shelter 0.0 20% 20% 0.0 0.0

COMS + Avionics 0.3 20% 100% 0.3 0.7

ECLSS 18.9 20% 20% 22.7 27.2

AOCS 3.7 20% 20% 4.4 5.3

Capsule 8.4 20% 20% 10.1 12.1

Propulsion 35.4 20% 20% 42.5 51.0

5%

Total 164.2 26% 18.6% 206.9 236.9

Furthermore, to understand the evolution of the CIV design as well as its maturity, the trend of the value was displayed and analysed. This can be seen in Figure 5-2.

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Mass [t] CIV total mass including margins

450

400

350

300

250

200 Estimation

150

100

50

0 1 2 3 4 5 6 7 8 9 week

Figure 5-2: Trend of the CIV dry mass Naturally, this analysis was not only done for the overall mass budget, but also developed for the different subsystems to serve as additional input to evaluate subsystem evolution and maturity. An example of this is shown for the structural subsystem in Figure 5-3.

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Mass [t] Structures and Habitational volume 90.00

80.00

70.00

60.00

50.00

40.00 Estimation 30.00

20.00

10.00

0.00 1 2 3 4 5 6

Week

Figure 5-3: Mass trend of the structural subsystem

5.3.2 CIV Power budget The Table 5-3 displays the power budget of the CIV. Table 5-3: Power budget of the CIV Peak Power Average Power Duty Contingency Item with margin Power with (max) [kW] Cycle margin [kW] margin [kW]

AOCS 1.4 50% 1.2 1.7 0.8 OBDH+Comms 1.4 25% 1.2 1.6 0.4

Science payloads 2.0 50% 1.2 2.4 1.2

ECLSS 25.0 70% 1.2 30.0 21.0 TCS 15.6 100% 1.2 18.8 18.8 Docking Systems 0.1 100% 1.2 0.1 0.1 Propulsion subsystem 10.0 1% 1.2 12.0 0.1 Total before System level margin 66.6 42.5 Total including margin 69.9 44.6

SPEC value 97.2 57.5

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Additionally, for the power budget, a trend analysis was performed, the status in July 2014 can be seen below:

Average Power [kW] CIV Power estimation 120.00

100.00

80.00

60.00 Estimation 40.00

20.00

0.00 1 2 3 4 5 6 7 8 Week

Figure 5-4: Trend of the Power Budget of the CIV

5.3.3 Conclusion The development of the system budgets shows clearly the growing maturity of the current phase A study. Constantly decreasing overall uncertainties and spacecraft mass estimates within the uncertainties predicated in the prior design steps show that proper system engineering approaches were used by all working groups. For very few occasions, the uncertainty level of one week exceeds the one of the prior week, usually due to major changes in the subsystem concept or in case of major changes of interface values.

5.4 MARS AUTOMATED TRANSFER VEHICLE STUDY The Mars Automated Transfer Vehicle, conceptually shown in Figure 5-5 is a novel space vehicle, which serves the main purpose of transporting heavy payload into the vicinity of Mars. For this purpose it utilizes a combination of a chemical stage for injecting the system into a highly elliptical Earth orbit and a solar electrical propulsion system for interplanetary and Martian operations.

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Figure 5-5: Artists impression of the MATV While the propulsion study is given in Section 3.4.5, this report gives an overview of the dry mass estimate of the system, which was derived assuming similar system performances as given for the CIV in the following chapter. This leads to the power and mass budget as given in Table 5-4 and in Table 5-5. Table 5-4: Power budget of the MATV Item Power (nominal) Contingency Power with [kW] margin margin [kW] AOCS 0.5 1.2 0.6 COM & OBDH 0.5 1.2 0.6 ECLSS, TCS, Docking systems 6.5 1.2 7.8 Lab science systems 13.5 1.2 16.2 Biosafety systems 10 1.2 12 Electric propulsion 400 1.2 480 5% Total 431 26% 568.92

Table 5-5: Mass budget estimate of the MATV Item BEE [t] Margin BEE incl. Margin EPS 5.3 20% 6.4 TCS 1.4 20% 1.6 LAB 15.0 20% 18.0 Mars Lander 5.0 20% 6.0 Docking 3.0 20% 3.6 Service Module 3.7 20% 4.4 Phobos Lander 15.0 20% 18.0 Remote sensing equipment 0.5 20% 0.6 5%

Total 48.9 26% 61.6

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5.5 PHOBOS LANDER STUDY The Phobos Lander (PhL) is based on the heritage of Apollo’s Lunar Excursion odule (LE ) and the European Automated Transfer Vehicle (ATV). The purpose of this study is to ensure the feasibility of a human landing on Phobos. Since it could compromise the feasibility of the whole mission, it was performed at the beginning of the activities in this project.

5.5.1 Stability on the Ground The estimated gravitational acceleration on the surface of Phobos is 5.7 mm/s2 (NASA – Phobos). A spacecraft of 15 t in this environment is expected to weigh 86 N, which is the equivalent of 9 kg on Earth. This force is obviously permanent, while solar wind and solar radiation pressure are expected to be negligible over a small, human scale surface area. Due to the lack of atmosphere there are no expected external perturbations. There are only two forces that the PhL will have to counteract. The first is the movement of astronauts moving inside. The second is thrust from decompression valves in preparation for EVAs, though this is trivially solved by locating the valves on top of the roof, thrusting downwards. For these reasons the PhL is not expected to detach from the ground. However rotations are a concern, as they could cause the spacecraft to tumble. The proposed solution is:  Locate the centre of mass as low as possible.  Powerful momentum wheels or control-moment gyroscopes to counteract astronaut torques.  Astronaut handles must be as close as possible to the centre of mass.  The span of landing legs shall be at least 4 m to each side. The first item will be further discussed in this document. More details on the remaining 3 items are provided in section 6.9 – AOCS. The stability and safety of astronauts on the ground is discussed in subsection 3.3.1 – Astronaut Locomotion.

5.5.2 Crew Resources The short duration of the stay in the Phobos surface does not justify any closed loop. Each landing is expected to last at most 4 days.

5.5.2.1 Oxygen, Food and Water Based on oxygen consumption measurements in (Larson, 1999), the mission duration and crew size, a total of 16 kg of oxygen including a 50% safety margin will be required per landing. An Aluminium 7075-T6 tank can be designed with an unloaded mass of 27 kg, which leads to a total mass of 43 kg. Water needs, as specified in (Larson, 1999), can be estimated as 7.5 L per person per day if dish washing and clothes washing are neglected. It is a reasonable assumption for a small spacecraft. This leads to a total of 78 L including a margin of 30%. Food needs of 0.62 kg/person/day (Larson, 1999) lead to a total of 6 kg, including a margin of 20%. All these numbers are per landing. However when the airlock or hatch is open air will be lost. By the dimensions of the PhL, the number of excursions and the initial pressurization, the total mass is 621 kg for tanks, 65 kg of O2 and 261 kg of N2. Total mass is 947 kg for air and containment, including a margin of 50%. Total mass for crew resources is then 1031 kg, out of which 621 kg are reusable and 400 kg are per landing. The alternatives of performing water electrolysis or burning lithium perchlorate have been considered, but their benefit is negligible compared to the unavoidable need for N2 tanks. The mass of an hypothetical airlock is not justified.

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5.5.2.2 Propellant As described in the section subsection 3.3.7 – Astronaut Locomotion, 197 kg of liquid propellant and 20 kg of gaseous propellant per astronaut and per spacewalk are needed. For landing two astronauts that will perform 4 spacewalks each, a total of 870 kg of propellant will be needed. For the PhL’s propellant budget see section 5.5.4.

5.5.3 Electric Power The Apollo Lunar Excursion Module (LEM) had an electric system supplying 50 kWh (Encyclopaedia Astronautica – Apollo LM). The PhL follows the design of Apollo’s LE in most parameters, though in this case the PhL is designed to land at least 2 times. Either rechargeable batteries must be used, or batteries must be replaced upon docking with the Lab. If Sony 18650-like rechargeable Li-ion batteries are used, with a capacity of 6.66 Wh and a mass of 44 g per cell, an energy density of 151 Wh/kg is achieved. For 50 kWh this translates into 331 kg of batteries. For mission flexibility purposes and due to the relatively small difference in mass, rechargeable batteries are preferred, even if non-rechargeable could be less massive. Applying a margin of 30% to the energy requirements, 430 kg of batteries must be carried on board the PhL.

5.5.3.1 Thermal Control Electric heaters may be needed, though electric power is expected to be enough. Radiators are not needed since a large amount of N2 and O2 compressed at 200 bar is available - decompressing a small amount of gas and releasing it into the internal environment will suffice for cooling.

5.5.4 Propulsion A total wet mass of 15 tons has been assumed as total for the PhL. Astronauts and all resources to be used on the Phobos surface are assumed to be included in this number. Since the propellant will be carried in pressurized space, all toxic propellants or oxidizers are ruled out (MMH, UDMH, Aerozine 50, HNO3, N2O4).

RP-1 can be burned using H2O2, and in this case a specific impulse of up to 290 s is expected to be reached in vacuum (Moon, Y. et al – 2014). An additional 1% of propellant mass should be foreseen to be lost due to its inevitable decomposition of 1% per year (Sutton, 2001). Exhaust gases are not excessively toxic.

However, in case of incomplete combustion some H2O2 can be released into the environment, and it can be corrosive for EVA suits. Alternatively, a commercial monopropellant is advertised as High Performance Green Propellant (HPGP) (SSG Group – ECAPS Capabilities). Its specific impulse, based on elementary calculations from the exhaust velocity advertised in the webpage, is 255 s It’s considered safe enough so that pressurized suits are not required to handle it (SSG Group – ECAPS Capabilities). For the rest of the calculations in this mass, a specific impulse of 255 s will be taken as a reference assuming HPGP will be used. In case RP-1 is chosen at a later stage of development, mass budgets will still apply as it has a higher specific impulse. Propulsion re uirements, based on mission analysis results, must account for a ΔV = 269 m/s per round trip. This number includes approach, landing, lift-off, return to the CIV and relevant margins. The Tsiolkovsky rocket equation outputs a dry mass of 13500 kg (rounded up), so there is room for 1500 kg of propellant. An additional propellant mass may be needed to use for attitude control. Even though the propellant use of such manoeuvres can be calculated with simple math, the needs to control attitude and the frequency of such manoeuvres cannot be easily estimated, especially in this low gravity environment. The driving assumption will be an equal propellant mass to that of the rest of the mission, which is clearly an overestimation. The conclusion is that 3 tons of propellants are needed for each landing, leading to a dry mass of 12 tons.

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5.5.5 Geometry and Dimensions For stability reasons the PhL will have a cylindrical shape in a horizontal position, breaking the heritage from the Apollo LEM and more analogous to the ATV. It shall be designed to be launched in an Ariane 5 vector. The dimensions of the PhL are 6.4 m length and 4.5 m diameter. The centre of mass must be located at 1/3 of its height. Vertical landing legs will absorb the landing shock, and diagonal landing legs shall be spanned at 4 m to each side of the vehicle.

5.5.6 Landing Dynamics An important aspect is avoid lifting abrasive dust, based on the Apollo experience, but also considering that in this negligible gravity environment it would take a very long time for it to settle back. For this reason rocket engines must not fire when the vehicle is at a close distance to the ground. At a height of 200 m it is possible to perform a good terrain reconnaissance and avoid the presence of boulders or obstacles. The PhL may reach a zero relative speed to the ground and be left in free fall, touching down at a speed of 1.5 m/s. This shock is easy to absorb using spring-damper systems, which shall have a damping ratio of at least 0.7 to avoid bouncing back. No anchoring is foreseen due to the uncertainties about the mechanical properties of the regolith. The stability on the ground will be discussed in section 6.9 – AOCS.

5.5.7 Mass 3 tons have been reserved for propellant. The mass of tanks, as the propellant is not expected to be cryogenic, can be safely estimated as 5% of this mass, leading to a total of 3.15 tons. Two shelters for solar storms are needed. As they have been estimated as 250 kg each, as discussed in the preparation phase, 0.5 tons are reserved for radiation protection. The total of crew resources is 1031 kg for life support resources and 870 kg for propulsion. Additionally 430 kg will be needed for batteries. The above numbers leave a total of 9 tons for structure and systems. The Apollo Lunar Module had a wet mass of 15 t and a dry mass of 4.2 t (Encyclopaedia Astronautica – Apollo LM). This proves the PhL is more than feasible with the current mass budget. It may be required to reduce the habitable volume, but based on the heritage of the Apollo LEM it is known that it can be reduced down to 7 m3 while the current value is around 100 m3. Currently feasibility is not compromised by volume based on heritage from the ATV. Since the feasibility of the PhL is proven and the study of the Orpheus mission focuses on the CIV, no further study is needed on the PhL subsystems.

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5.6 REFERENCES: LARSON, W.J. (1999) Mission analysis and design. 3rd Edition. El Segundo: Kluwer Academic Publishers DAIMLER-BENZ AEROSPACE (1996) Columbus requirements document, Revision: 3-B, Daimler-Benz ESA (2009) ECSS - Space engineering general requirements, Third issue, ESA Requirements and Standards Division European Cooperation for Space Standardization (2009) ECSS-E-ST-10C, System engineering general requirement, Noordwijk: ESA NASA – Phobos: Facts & Figures [Online] Available from: http://solarsystem.nasa.gov/planets/profile.cfm?Object=Mar_Phobos&Display=Facts [Accessed May 15th, 2014] Larson, Wiley J. et al (1999) Human Spaceflight: Mission Analysis and Design. New York: McGraw-Hill Companies, Inc. ISBN 0-07-236811-X. Encyclopedia Astronautica – Apollo Lunar Module [Online] Available from: http://www.astronautix.com/craft/apollolm.htm [Accessed May 19th, 2014]

Yongjun Moon, Chul Park, Sungkwon Jo, Sejin Kwon (2014) Design specifications of H2O2/kerosene bipropellant rocket system for space missions [Online] Available from: http://www.sciencedirect.com/science/article/pii/S1270963814000170 [Accessed May 21st, 2014] Sutton, George P.; Biblarz, Oscar (2001) Rocket Propulsion Elements 7th edition. New York: John Wiley and Sons, Inc. ISBN 0-471-32642-9. SSC Group – ECAPS CAPABILITIES [Online] Available from: http://sscspace.com/capabilities-3 [Accessed May 19th, 2014] SEEDS edition V students (2011) AENEA: humAn Exploration of a Near Earth Asteroid, Ch. 5: Robotic system [Online] Available from: http://www.seeds-master.eu/Executive_summary_PW10_Chapter%205-6.pdf Other chapters: http://www.seeds-master.eu/Executive_Summary_PW10_Chapter%201-4.pdf http://www.seeds-master.eu/Executive_summary_PW10_Chapter%207-8.pdf http://www.seeds-master.eu/Executive_summary_PW10_Chapter%209.pdf [Accessed May 14th, 2014] Genta, Giancarlo (2012) Introduction to the Mechanics of Space Robots. New York: Springer. ISBN 978-94- 007-1795-4. European Space Agency - ATV Configuration [Online] Available from: http://www.esa.int/Our_Activities/Human_Spaceflight/ATV/ATV_configuration [Accessed May 20th, 2014] S.A. Miedema, G.H.G. Lagers, J. Kerkvliet (2007) - AN OVERVIEW OF DRAG EMBEDDED ANCHOR HOLDING CAPACITY FOR DREDGING AND OFFSHORE APPLICATIONS [Online] Available from: http://www.dredgingengineering.com/dredging/media/lecturenotes/miedema/2007_wodcon/mooringoverview. pdf [Accessed June 25th, 2014]

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6 SPACECRAFT DESIGN

6.1 DESIGN, MECHANICS AND STRUCTURES

6.1.1 Introduction This section’s objective is to describe the external design of the spacecraft. The preliminary design of the spacecraft was performed using CATIA V5 for the 3D and AutoCAD 2012 for the 2D. All the elements are referenced and described in Figure 6-1. The first study of this part deals with the propellant tanks as it could have a huge impact on the overall design of the CIV. For sizing the nuclear propulsion tanks, the resulting choice of methane became a system driver. Two additional tanks have been provided in order to perform minor manoeuvres - with chemical propulsion - in Mars proximity: in this case, the bi-propellants for the systems are methane and oxygen. The use of chemical propellant for departure stage from LEO will require further analysis. After that, the attention has been focused on the Habitation Modules (HAB), both externally and internally. The main drivers for this part were the radiation protection, and the optimization of space. Finally, the different other elements of the spacecraft have been studied, mainly in terms of attachment (for solar panels, radiators) and accessibility (Nodes 1 and 2).

6.1.2 Elements of the spacecraft: External Configuration

6.1.2.1 2D model of the CIV Figure 6-1 gives a representation of the CIV. The thermal radiators are not visible in this figure, as they are positioned in the lateral axis. The table below gives the signification of the acronyms of the different modules from left to right:

Acronyms Signification

CRV 1 & 2 Crew Rescue Vehicle 1 & 2 N1 & N2 Node 1 & 2 I1 & I2 Inflatable Part 1 & 2 R1 & R2 Rigid Part 1 & 2 HAB 1 & 2 Habitation Module 1 & 2 LAB Laboratory Module

Figure 6-1: 2D model of the CIV

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6.1.2.2 Propellant Tanks Existing studies (M. Sippel et al, 2013) are under development to determine the feasibility of composite fuel tanks for use in the space industry. The challenge in developing a cryogenic CFRP (Carbon Fiber Reinforced Plastic) tank is finding a solution for the problems caused by differences in thermal expansion coefficients (CTE) on a microscopic scale. If a liner is required, another challenge is to overcome the differences in CTE of the liner with respect to the structural shell. It has been considered using two cryogenic propellants for the nuclear system; methane, which needs to be stored at about 100K and at less than 9 bars of pressure, and hydrogen - at 20K. Taking into account these issues, an acceptable compromise for the tanks needed for this mission has been found. The mechanical characteristics that have led the choice all along the study refer to cryogenic temperatures. After an accurate analysis of the references, the best option is to use a Liquid Crystal Polymer (LCP) as a liner for the tank, and in particular the Vectran LCP fiber. Indeed, thanks to its promising mechanical properties at low temperatures and its good adhesion on a laminate base, this material proved to be the number one choice for this application. Concerning the composite that will used to sustain the mechanical pressure and/or forces, the choice is Toray T700S carbon fibre fabric and Araldite epoxy 5056. Mechanical properties of this chosen composite:  Density: 1.6 g/cm3  Tensile Modulus: 203 GPa  Tensile Strength: 3500 MPa  Specific Tensile Strength: 2188 MPa  Specific Modulus: 1.27E12 cm2 s-2 Therefore, in Mars proximity, it will require 20 tanks of Methane for nuclear propulsion. To pressurize those 20 tanks of propellant, it has been decided to use 5 small tanks of pressurized gas (in this case Helium) on the rear of each tank. The tanks have been sized in order to fit into the Falcon Heavy launcher. They will have a length of 11.4m and a diameter of 4m, with a dry mass of 0.75t. The mass at launch will be 53t. Concerning the two additional tanks for chemical propulsion, they have the same characteristics of the methane tanks. Because of the ratio between oxygen and methane within the tank – 3.4:1, as in the combustion chamber – the mass at launch is 115t. So, in order to be able to lift those tanks, the SLS 130 must be used. As previously mentioned, those tanks will feed the chemical engine in order to perform low energy manoeuvres in Mars proximity.

6.1.2.3 Habitable Modules (HAB) For now, there are several aerospace companies currently working on the concept of inflatable structures for space related purposes. There are many advantages to this kind of structures such as a reduction in terms of mass and initial volume. The Orpheus concept of Habitation Module (HAB) will have a toroidal shape (see Figure 6-2) in order to optimize the available volume in terms of rigid/inflated ratio. A volume of 90m3 per astronaut has been provided, plus all the necessary volume for contingency. The assumption applied for the needed volume (see section 6.2) is approximately 650m3. This volume will be divided into two separate HABs, each of them connected by a node (N2). This node will provide an airlock system, which could be used in case of emergency (fire in one HAB for example).

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Figure 6-2: Habitation Modules

The proportion between rigid/inflatable is approximately ⅓ The material used for the rigid part will be an aluminium alloy. Concerning the inflatable part, there will be multiple layers of different materials (total thickness of 0,45m, average areal density 8 ⁄ ) that provide blanket insulation, protection from orbital and meteoroid debris, optimized restraint layer and a redundant bladder with a protective layer. The inflatable material composition is provided in the list below [NASA, 1999].

 The outer layers protect multiple inner bladders, made of a material that holds in the module’s air The shell also provides insulation from temperatures in space that can range from +121°C in the Sun to -128°C in the shade.  The second part is made of Nextel, a material that insures debris protection.  Into the shell there is a layer of woven Kevlar that holds the module’s shape  Three bladders of Combitherm hold the air inside.  The innermost layer, forming the inside wall of the module is Nomex cloth, a fireproof material that also protects the bladder from scuffs and scratches.  Moreover, some layers of Vectran insure mechanical properties that are useful against meteoroids, debris and radiation. Four large windows coated with a UV protection film will support both celestial and Martian viewing, dimensioned 30cm x 30cm and located respectively on the front and back of the two different HAB of the CIV. The material used for the window will be Udelpolysulfone (SOLVAY, 2014), the same material that is currently used for the face shield visor for the astronaut space suits. In the Orpheus design, the propellant tanks of methane and methane/oxygen have been fixed around the HAB circumference, see Figure 6-3. An additional methane/oxygen tank has been put between the nuclear reactor and N1, for better thermal control – reducing the amount of Sun incidence - and as additional shielding from the nuclear thermal engine core. The length of each tank is 11.4m; under pressure at 9 bars and with a thickness of 1mm; whereas the length of the two HAB is 10m (5m x 2) and the length of N2 is 2m. Therefore, it is not possible to cover the total length of two HABs with one tank. It has then been decided to position 21 tanks in two separate crowns all around the HAB with minor longitudinal displacement respective to each other. The spatial position of the tanks has been optimized in order to both provide the astronauts with enough area to perform maintenance for the HABs/N1 and in order to avoid any problems during the jettisoning of stages.

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Figure 6-3: Fixation of the tanks around the HAB

6.1.2.4 Miscellaneous equipment The node in front of the HAB 1 (N1) has been considered as support structure for several components. The characteristics of the node are described below. The airlock system is a section of the N1 cylinder, as can be seen in Figure 6-4, with a diameter of 3m. The length of the airlock section is determined by the airlock hatch diameter, including margin.  Airlock hatch: 1.3m. (See NASA Quest Joint Airlock)  0.1m of margin each side of the hatch.  Diameter: 3m  Length: 1.5m The N1 will be critical for several reasons. First of all, the Crew Re-entry Vehicles will be attached to it in order to re-enter safely. These vehicles are based on the configuration of the Soyuz 7K-LOK, which has been designed for hyperbolic re-entry. This will be adapted for 3 astronauts. Two EVA ports will be provided for maintenance purposes, and in cases of emergency.

Figure 6-4: Node 1 representation All the trusses necessary for the fixation of the tanks, solar arrays, and radiators will be made of carbon/epoxy composites and deployed in orbit, using an existing CNES technology firstly used for the deployment of solar arrays itself, but applicable to the trusses, as shown in Figure 6-5. Inside each truss many systems could be installed post deployment.

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Figure 6-5: Solar array trusses The final rendering of the CIV, with all the several components assembled, can be seen in Figure 6-6.

Figure 6-6: Final rendering of the CIV at arrival in Mars vicinity

6.1.3 Elements of the spacecraft: Internal Configuration The internal configuration has been done in collaboration with the ECLSS team. Depending on the resources to fit inside the HAB, and the personal space needed per astronaut several different configurations for the HAB have been thought of.

The major outputs are as followed:  Inside the rigid part of the HAB, a corridor of 1.8m will be dedicated for the displacement of the astronauts. The rest of the volume of the rigid part (32m3) will be used as storage for subsystems, see Figure 6-7.  The HAB is divided into three floors and attachment points will be provided, in order to fix the required equipment. The rigid part will allow an easy access to the intermediate floor as it will be open on the flanks. The separation between two floors will most probably be done using a net.  Each floor will be divided in different areas: the equipment and the subsystems will be attached on the inflatable part of the structure trough nets and Velcro, while on the rigid parts the subsystems will be fixed trough mechanical interfaces.

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Figure 6-7: Internal configuration of the HAB

6.1.4 Inspection and regenerative tools During the 600 days of travel, several kinds of structural problems could potentially occur on the CIV. In order to prevent and repair damage, the crew should be capable to perform EVA and maintenance.

One method that allows to save mass and to validate new technologies in the deep space is the use of the regenerative tools, such as a 3D printer. It is able to produce structures, utilizing on board plastic base resources and CAD models sent post launch by ground control. These parts can be rapidly manufactured.

6.1.5 Mass The components, described in the previous sections, are summarised with their respective Best Engineering Estimate (BEE) masses and total masses after applied margin and uncertainties in Table 6-1. Table 6-1: CIV, post-Earth departure, mass breakdown by component. Component Best Estimate (t) Margin Intermediate (t) Uncertainty Total (t)

HAB1 15.00 1.1 16.5 1.2 19.80

HAB2 15.00 1.1 16.5 1.2 19.80

Node 1 6.27 1.3 8.1458 1.5 12.22

Node 2 1.00 1.3 1.3 1.2 1.56

Truss - Solar 1.48 1.3 1.924 1.2 2.31

Truss - Radiator 1.48 1.3 1.924 1.2 2.31

Truss - Tanks 0.45 1.3 0.585 1.5 0.88

Robotic Arm 0.63 1 0.63 1.05 0.66

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BE Total 42.81 11.00 49.46 Total 59.54

6.1.6 Conclusion The CIV, in the configuration of post-Earth departure, presents a significant engineering and logistical challenge. All the structural components require several launches utilising different types of launcher. The present geometry and distribution are the results of an iterative process and should remain very similar moving forward. Refinements such as the internal configuration of the habitation modules and radiation shielding may be required in the future. The logistics of orbital construction, especially considerations for the Earth departure stage, is a key focus for future review. The expected frequency of launches will be a factor towards the mission feasibility. Cost analysis is currently at a low maturity stage and values have to be considered very critically. To reiterate a complete cost analysis is not a primary output and beyond the scope of the study. However future efforts are necessary ‘size’ the financial re uirements more realistically The output of the ‘Design, echanics and Structures’ is the preliminary proposal for the CIV in post-Earth departure configuration. Mass estimations, geometric design and assembly logistics have been evaluated in order to provide a complete analysis of the CIV.

6.2 ENVIRONMENT CONTROL AND LIFE SUPPORT SYSTEMS (ECLSS)

6.2.1 Introduction In the design of the ECLSS, first requirements of the systems were identified in order to keep the astronauts alive and healthy; using the ECSS document ECSS-E-ST-34C as a guide. The key areas of focus for the ECLSS design are:  Provision of Resources  Environmental Control  Waste Management  Special Operations  Health Related Services  Crew Safety (e.g. in case of fire)  Habitation Layout In this mission design, there is a crew of 6 that need to be supported by the ECLS system for 610 days (Mission duration plus 1 week extra to account for a week in Earth orbit before leaving or on return). In this section the focus is on the design of the ECLSS system for the CIV only, the Phobos Lander and Lab will have separate systems that have not been considered. This is an area for further study.

6.2.2 Provision of Resources To begin the design of the ECLSS one of the first items to be defined are the inputs and outputs of a human, and the conditions required for the human to stay alive. Considering these items allowed us to establish the resources needed for a long-duration flight which will then allow us to determine the ECLS systems to use - whether regenerative or not. The ECLS systems must provide the inputs and deal with the outputs of a human, but also must provide additional things to allow the astronauts to live comfortably, such as sleeping areas. The Figure 6-8 & Figure 6-9 below show the mass and volume of each category of resources and small systems, not including water and atmospheric gas generation as these will be likely be regenerative due to

SEEDS Executive Summary 09/2014 Page 101 the large masses involved. Food is included as a resource as it is not expected to be technically feasible to grow food using waste for this mission so all food for the entire mission must be launched.

Figure 6-8: Mass of Resources

Figure 6-9 Volume of Resources

6.2.2.1 Open or Closed Loop An open loop system is one in which resources must be resupplied from Earth. A closed loop system attempts to use waste products to generate useful resources to get rid of the need for resupply. In this section of the report, it is determined whether to have an open or closed loop for each system. 6.2.2.1.1 Water For an open loop system the amount of water needed, based on 28.8kg of water per astronaut per day (Wieland, 1994), for a crew of 6, is 106 tonnes of water for the entire mission (610 days). Clearly this is unfeasible to launch from earth, in fact after about 20 days it becomes mass efficient to take a recycling system with a water recycling efficiency of 90%.The water that needs to be launched in this case is 11 tonnes. This efficiency value is a linear estimate of efficiency based on the efficiency of the technology improving. Theoretically, there are systems that have been built on Earth that provide recycling efficiencies up to 95%. It is reasonable to assume that if this technology can be proved for space use, it will be available

SEEDS Executive Summary 09/2014 Page 102 in 20 years (2036) at 93%, which means that a recycling system is need after 17 days and 6 tonnes of water needs to be launched. It is worth considering water for use as radiation shielding in the event that propellant does not fulfil the shielding requirements. In this case, more water may need to be launched. Having calculated the amount of water needed and that water recycling systems are required, is must be decided which technologies are required to recycle the water. There are two basic approaches to water regeneration, (1) phase changes or (2) physically separating the components through filtration or reverse osmosis. As the mission requires so much water both these approaches will need to be used, phase changes to recycle urine and filtration to handle the less contaminated water such as shower water. Potable water will need to be processed more than shower water so as to be suitable for drinking. For this mission, values provided by NASA have been used to estimate the amount of water needed per astronaut per day. This was split into the amount of water input into the concentrated feed (VCD recycling) and water input straight into the dilute feed to be recycled by multi-filtration. The mass of the water including tanks has been calculated using a packing factor of 1.02. The mass of the water recycling systems (2571kg) was taken from table 6.9 of Hanford, A. J, 2005, but could potentially be decreased to 1000kg if advanced technology can be used, however this was not used for the initial estimates in order to have a worst-case estimate.

6.2.2.1.2 CO2

Regenerative CO2 filtration is the act of using regenerative filter technology to remove CO2 from the air and provide fresh air to the environment. After the CO2 is scrubbed from the air, the filters release it. This process is therefore regenerative. Once the CO2 is separated from the air the O2 can be removed from the CO2 and reused. Using a 4-Bed Molecular Sieve (4BMS), the need for replacements disappears and so it is suitable for longer duration missions. The system mass is 30kg/p, which for a mission of 6 means a total mass of 180 kg (Larson and Pranke, 1999). The mass of this regenerative system quickly surpasses use of LiOH cartridges in terms of mass efficiency, by day 18 in space. As such for this mission 4BMS will be used. 6.2.2.1.3 Oxygen Based on 6 astronauts consuming 0.84kg of oxygen per day (Wieland, 1994) the masses required for the storage of pressurised oxygen, and the mass required for electrolysis of oxygen from water were calculated over the duration of the mission. For the oxygen storage, a packing factor of 1.9 was used (Larson and Pranke, 1999,Table 17-9) and resulted in a final mass of 5800kg, whilst electrolysis resulted in a total final mass (including a Sabatier process, and assuming an efficiency of 0.9) of 4300 kg. The system from which the estimated mass of the electrolysis system was derived is the Oxygen-Generation Assembly (OGA) on board the ISS. The OGA uses electrolysis of water from the water processing system of the ISS to produce enough oxygen to sustain a crew of six indefinitely. The mass of this system is taken as 35kg per person (Larson and Pranke, 1999). Along with this, the Sabatier process will use a system estimated at 38kg per person (Larson and Pranke, 1999) The mass of stored oxygen was compared with the mass of water with an electrolysis system and with the mass of water with electrolysis and water recovery from Sabatier. The stored oxygen has a packing factor of 1.9 and the oxygen regeneration system has a reference mass from the OGA. It is seen that past durations of approximately 140 days it is more mass efficient to use an oxygen generation system. Using the Sabatier process to reclaim water is clearly more mass efficient, however these estimates do not include the mass of any power or thermal cooling systems needed and more systems also increase the complexity of the overall ECLSS.

6.2.3 Environment Table 6-2 shows the environmental requirements for the habitable volume and suggested technologies. The environmental control system should be able to meet these requirements. The mass, volume and power estimates used for the environmental control system is shown in Table 10-6 found in the appendix.

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Table 6-2: Environmental Control Requirements Atmospheric Required Pressure 101325 Pa Pressure control1,2 Pressure Upper Limit 114657 Pa (nominal) Pressure Lower Limit 99900.0 Pa (nominal) Pressure Lower Limit 87992.0 Pa (degraded) Pressure Lower Limit 53328.9 Pa (emergency) Atmospheric Earth Composition Nitrogen 78.084% Composition control3 Oxygen 20.946% Trace Gases 0.97% Water vapour ~1% Oxygen Partial Pressure 21224 Nitrogen Partial Pressure 79119 Atmospheric Comfortable Humidity level 60 % Humidity control Upper Humidity Limit 70 % Lower Humidity Limit 30 % Trace gas/odour Max Gas Concentration Below Spacecraft Maximum control Allowable Concentrations (SMAC) Levels4 Particulate control Airborne particulate removal Maintain Average Concentration and disposal: CCAA Filters of 0.05mg/m3 HEPA Filters Atmospheric Cabin airspeed 0.08-0.1 m/s Circulation System Airborne Airborne microbe removal Microorganism and disposal: CCAA Filters

1 Barratt and Pool, 2008

2 Stack Exchange, 2013

3 Barratt and Pool, 2008

4 NASA, 1999

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control HEPA Filters Noise Level5 60 dB

6.2.4 Waste Management In the previous sections the possibility and importance of recycling air and water were discussed, this chapter is concerned with the management of contagious and non-contagious solid waste. Contagious waste can consist of biological waste and food waste whilst non-contagious waste may consist of clothes, packaging, etc. The wastes that are more critical are prioritized by their infection hazard. The collection of solid waste will be either manual or automatic depending on the type of waste. For example the crew will dispose of cartons or food waste manually whereas biological waste will be automatically removed and managed. The main purpose and critical functions of the waste system is not the regeneration of useful resources but rather to provide crew health and meet planetary protection requirements. Waste is described in one of the following states during the duration of the mission:  Unprocessed (no changes)  Volume reduced (tins and cans crushed to smaller sizes)  Dried (water removed to prevent microbe activity)  Sterilised (destruction of existing microbes)  Stabilised (non-reactive with other materials)  Mineralised (converted to forms that could be extracted as resources)  Bio signature free (prevents interference with search for life)

6.2.4.1 Present Technologies 6.2.4.1.1 Volume reduction A mechanical compactor can be used to reduce the volume of waste by crushing the materials and increasing the density in this case helping to manage space. Plastic waste can be heated and melted, and has the potential to be used in a 3D printer. This heating and melting process currently has many space qualified systems which are able to perform these operations. 6.2.4.1.2 Drying The drying process has many ways of removing heat in order to prevent microbial activity. These methods include thermal drying, lyophilisation, and microwave freeze drying, all of which has been space qualified. 6.2.4.1.3 Food Production from Waste Food production from waste is crucial for future missions to Mars. With the present technology readiness level, food production in space is not reliable enough to be included as part of a closed loop to provide food for the entire mission. However, it is a great possibility in future and thus performing a validation for this technology during the mission is incredibly valuable. The current process of food production from waste is the cultivation of single cell protein or fungi and subsequent conversion to an edible food. The next process which is possible is recovering nutrients from waste and then converting these to foods via a hydroponic system.

5 ESA, 2008

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6.2.5 Special Operations The spacesuits of the crew members shall be able to provide protection of the crew in all the phases of the mission. These phases go from launch phase, transfer phase, operation phase, re-entry phase, and also consider factors such as contingency extravehicular activity (EVA) and Launch Escape Abort (LEA) phases. The most significant mission critical situation is the Phobos surface operation capability for EVA. The productivity and lives of the crew depends on the performance of these suits. The expected characteristics of the suits should provide high mobility and durability with minimum leakage and will need a significant amount of dust protection. These characteristics should be provided at very low maintenance.

6.2.5.1 Procedures for EVA suit 1. Pre-donning suit preparation 2. Portable Life Support System (PLSS) preparation 3. Suit Donning and Final Check 4. Egress/Ingress 5. Mid-EVA rest period 6. Post-EVA Securing 7. Non-Routine Maintenance

6.2.5.2 Pre- Breathing The astronauts will need to pre-breathe pure oxygen in a standard pressure environment; in order to avoid decompression sickness when moving from the pressure of the spacecraft to the lowered pressure of the space suit. Design of a space suit which does not require this procedure is ideal for future space suits. The current procedure for the Phobos Landing is a pre-breathing phase in the Phobos Lander which would take more than 3 hours before an EVA on the surface could take place, this time is sufficient for pre-breathing. For contingency EVA from the CIV for repairs an airlock will be used to pre-breathe.

6.2.6 Health Related Services Larson, W. J and Pranke, L.K, 1999 states that a Crew Healthcare System (CHS) must provide three services. The CHS must diagnose and treat sick crew members, maintain optimal health in the crew, and monitor the environment in order to provide warnings about exposure to hazards. The crew must be able to operate this suite in at least a limited mode. An expert (i.e. a medical doctor or surgeon) should be included in the crew to perform operations and maintenance of the system. These systems will be contained in the medical suite in one of the habitation modules. As the mission is a long term expedition to Mars, with little to no abort options the system must provide long- term and emergency care, which includes a suite to provide capability of major surgery (Larson, W. J and Pranke, L.K, 1999). The system will provide a quarantine system in order to separate a crew member if needed, in case of emergency such as a detected contagious disease or a mentally disturbed crew member. The crew member which is primarily responsible for the use and maintenance of the medical suite will be trained to perform all procedures which are foreseeable to be needed during the mission. In addition to providing care for injuries and illness the system must preserve the wellbeing of the crew. This includes maintaining circadian rhythms by using the correct on board lighting and by providing countermeasures against the micro-gravitational environment. Maintaining circadian rhythm is important as this is the process which regulates the sleep cycle and therefore allows the crew to mimic the sleep cycles of Earth i.e., falling asleep and waking with the same period of sleeping and waking hours each day. This cycle is essential as it ensures that the crew get enough sleep each night and also that they sleep the same hours as each other, which allows for the crew to interact during the day, which counteracts feelings of isolation which can occur during this type of deep space human mission.

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6.2.7 Crew Safety There are many factors that must be considered that come under the “Crew Safety” category From an ECLS point of view the systems must be able to respond to uncontrolled depressurisation and pressurisation events, fires, hazardous atmospheres and radiation exposure.

6.2.7.1 Fire Suppression Fire burns differently in microgravity, molecular diffusions means that the flame draws oxygen toward it and pushes the combustion products away. Flames can burn at a lower temperature and with less oxygen than on Earth, which means fire suppression systems must be adjusted for this. Fire detection systems would need to be built into the ventilation system (as on the ISS) as smoke does not rise in space. The current methods of fire suppression on the ISS rely heavily on the crew to extinguish the fire. First, the ventilation system is turned off to stop the spread of the fire and the power cut off to the unit, then the astronauts can extinguish the fire either using CO2 or water-based foam fire extinguishers. A water-mist system is being developed for use on future long duration missions, which has the added benefit that water will be available to refill the system and cleanup is easy to do using the de-humidifier system.

6.2.7.2 Hazardous Atmosphere There is a risk that the atmosphere of the CIV could become contaminated. The ECLSS must be able to detect this and provide a means of protecting the astronauts from it, such as pure air masks. It must also be able to restore the atmosphere to appropriate levels.

6.2.7.3 Pressurisation and Depressurisation Events The habitable volume could become over or under pressurised. The system should be able to detect these events, recover from these events and restore to normal operation. There also needs to be provisions included to treat the crew for over-pressurisation sickness, hypoxia and decompression sickness. The system also needs to alert the crew to these events.

6.2.8 Habitation The configuration of the habitation space (HAB) is detailed design and it is not possible to achieve a complete layout of the two habitation volumes at this time. However, it is possible to give an overview of the areas and systems which will be required on board, and a preliminary draft of how the layout may appear. The HAB is split into two volumes of 324m3 which each consist of a central rigid tube of 5m length and 3.4m diameter, which each support an inflatable toroid-like structure. The rigid structure houses a volume of 90.8m3 and is the surface which systems must be attached to if they require a solid mount. The total habitable volume is 648m3, this was calculated by taking a value of 90m3 for each crew members, this totalled and given a 20% uncertainty gives this value. The value of 90m3 was calculated by looking at human spaceflight and Mars mission sources, mainly the ISS, Skylab, MIR and Mars 500. Required Systems were assigned regions of space in the spacecraft based on the requirement for location in the rigid or inflatable section (i.e., the need for a solid mount) and on the forecasted volume required. The ECLS systems require a solid mount as they consist of high mass complex components which will be kept in racks similar to those in the ISS. The ECLS system was split into two separate systems each of which support a single HAB. This was due to redundant safety, so that in the case one of the systems breaks or a habitable volume is compromised, there is a second which can be used until the broken system is repaired. The location of the sleeping cabins and storage areas were designed primarily with regards to radiation protection. It was decided that the sleeping cabins would be best located in the centre of the craft, with solid angles protected by volumes of food, water, waste, and supplies provides the best possible protection of the crew from Galactic Cosmic Radiation (GCR). The volume of the ECLS systems takes up the majority of the volume of the rigid structure when space for an access corridor is included. 1 metre of material is required to protect the sleeping cabins; this means that a thin layer of protection must be fixed onto the outer side of the rigid tube. The rest of the systems (non ECLSS) will be kept in the inflatable volume.

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As a preliminary draft it was decided that the HAB volumes would be split into 3 “floors” which each would be split into two volumes. The various areas required would be split amongst these volumes. These areas are approximate and will possibly change when more precise volumes for the required areas, access points and storage are known. An example layout for the habitation volumes can be seen below. HAB 1 is listed first, and this is the habitable module situated furthest from the engine. Longitudinal sections can be seen in Figure 6-10, with port side shown on the left and starboard side on the right. A cross section is shown in Figure 6-11, which also shows a possible configuration for the position and arrangement of the sleeping cabin, in which 3 crew will sleep at a time in a 2 metre long compartment. Figure 6-12 and Figure 6-13 shows the same as above, for HAB 2. The cross section in Figure 6-11 shows a possible configuration for the position and arrangement of racks in which the ECLS systems will be kept within the rigid structure. The layouts below were chosen in order to create separate work-like and home-like habitation volumes in order to provide distinct environments for the crew. A toilet has been placed in each HAB for redundancy and separation and the sleeping cabins will need to be placed and designed in such a way to minimise noise. This is essential as sleep is a key factor in the performance of the crew.

Figure 6-10 Longitudinal Section of HAB 1 Figure 6-11: Cross-section of HAB 1

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Figure 6-12: Longitudinal section of HAB 2 Figure 6-13: Cross-section of HAB 2

6.2.9 Conclusion Table 6-3 shows the overall mass, volume and power budgets. The breakdown of budgets of the ECLS systems can be seen in the appendix. The design of the ECLS systems has a low maturity as a conceptual study is being performed and therefore it is not possible to perform a detailed design. Redundancy is very important in an ECLS system. Having an ECLS system failure would be catastrophic as the mission would be a failure. Each system must be able to survive two single point failures. There should be two separate systems to perform each critical task that work in different ways, so as to reduce the risk of two identical parts failing. For this reason the mass, power and volume estimates have been doubled for each system to account for the redundant systems. This is a good enough approximation as although the systems will function differently, the mass volume and power estimates should be roughly the same. There is also 20% uncertainty included in this estimate. Unfortunately detailed redundancy calculations are beyond the scope of this project. The most critical areas of research going forward would be to perform a more detailed design of individual ECLS systems as well as research into different technologies for each of the most crucial areas, water regeneration, oxygen production and waste recycling, due to the reasons stated above.

Table 6-3: Mass, volume, and power budges of ECLSS and resources Items Mass (kg) Volume (m^3) Power (w)

(Including calculations for spares, integration etc) 7500 36 8800 System Total Above with redundancy 15000 70 17600

Resources Including food, water, oxygen, etc 38000 190 n/a Total

Overall Total 53000 260 17600

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6.3 RADIATION

6.3.1 Introduction Space radiation is a major issue for long-term interplanetary missions. This study aims to improve the astronaut health and reduce the risk of exposure-induced death or cancer (REID) as much as reasonably achievable.

6.3.2 Effect of trajectory and mission epoch on radiation The mission epoch has been chosen in the year 2036, part of the period of maximum solar activity that is expected to start in 2034. Despite the fact that solar particle events (SPEs) are far more dangerous in the short term and far more frequent during this period, the relatively low energy of particles makes them easy to shield. The advantage of the selected mission epoch is the solar modulation of galactic cosmic rays (GCR), effectively reducing the dose by 1/2. Short-stay trajectories to Mars perform a Hohmann transfer either on approach or on return, while the other trait is a longer trajectory that usually crosses Venus' orbit or even uses a gravitational assist from the mentioned inner planet. In the case of this mission, the Venus fly-by occurs during the approach trip and the Hohmann transfer is performed for returning to Earth. This gives a wide availability of propellant during the longest part of the trajectory, and leaves the spacecraft unprotected only for a shorter time.

6.3.3 Using propellant for radiation protection After the first burn for Earth departure, the spacecraft will still have 1300 tonnes of methane available. Most of this propellant will not be used until the arrival to Mars takes place. Use of propellant for radiation protection has been studied in Spillantini, P ( 2014). Given the geometry of the spacecraft, approximately 1/2 of the solid angle will be protected by methane tanks. Their diameter is 4 m and they are arranged in 2 layers, so that the outer layer covers all gaps. Combining this geometry with the density of 422 kg/m3, this provides 169 g/cm2 of methane shielding.

6.3.4 Habitable space design and sleeping compartments The remaining half of the solid angle, the front and back of the CIV will not be protected by methane. In this case the only protection available will be the HAB. Based on the typical density of 5 g/cm2 of inflatable modules, and based on the knowledge that the deep-space version of the HAB provides a layer of 3 cm of water, a density of 8 g/cm2 has been assumed. Sleeping compartments are a strategic space for radiation protection, given that a small space is easier to protect and the significant fraction of the astronauts' time spent in this space. 10 tonnes of water is already available for ECLSS, plus 8 tonnes of food. Through the duration of the mission these numbers are not expected to change significantly as water and food are replaced with waste.

6.3.5 Effectiveness of shielding against GCR [Cucinotta, 2013.a] shows that, in deep space, aluminium does not provide meaningful shielding beyond what can be achieved with 20 g/cm2. At average solar minimum, considering a constraint of 3% risk of radiation induced cancer, the same reference concludes that with aluminium shielding of 20g/cm2, 256 days in space are safe for 55-year-old never-smoked males. Since this mission is planned at solar maximum the duration can be doubled, arriving to 512 days, though this is still below the 602 days that this mission is expected to last. ESA's Human Mission to Mars (HMM) [L. Bessone, D. Vennemann et al., 2004] estimates the dose equivalent based on the thickness of water shielding. For a trajectory during solar maximum with 8 g/cm2 the expected dose is 220 mSv/year. However these numbers may be outdated and superseded by the measurements of MSL [Cucinotta, 2013.b].

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6.3.6 GCR Attenuation analysis A Monte Carlo analysis has been performed using standard tools, including Geant4, SPENVIS and a local copy of GRAS [G. Santin et al. 2005] [S. Agostinelli et al. 2003]. Planar slabs have been used to represent the geometry of the spacecraft, which can be seen in Figure 6-14:

Figure 6-14: Planar representation of the spacecraft Simulations show a dose-equivalent attenuation of 40% in the path that traverses methane tanks and the polymer shell of the inflatable habitation modules. There is a further attenuation of 80% if the sleeping compartments are shielded with 1m (100 g/cm2) of water. The percentages on Figure 6-14 show the equivalent dose that a hypothetical detector in that point would perceive, compared to a detector in the outside. Unfortunately in the current configuration only 1/2 of the solid angle is covered with methane tanks, and they will not be available during the return trip to Earth. In the second configuration, when only the polymer shell and the water wall will be available, simulations show that there is no meaningful attenuation behind the polymer shell within uncertainties. The wall of water reduces the dose-equivalent by 10%. From these analyses it's evident that shielding the sleeping compartments is a worthwhile configuration. It is desirable to maximize the covered solid angle and time spent inside. Given the small attenuation in the direction in which methane is not available, it's straightforward to conclude that the mass of the caps of the cylinder is not justified, so they have been removed. If 6 sleeping compartments are shielded at the same time, the cylinder can be designed with an inner diameter of 3 m and an outer diameter of 5 m. Given its length of 2 m, it would require 25 tonnes of water. Since the ECLSS system has 10 tonnes of water and 8 tonnes of food, such a configuration would require 7 extra tonnes for radiation protection. However, if only 3 compartments are shielded and astronauts have a sleeping schedule, these resources are enough to provide the required shielding. A trade-off analysis by the entire team concluded that the sleeping schedule is preferable, though it will be broken during the short stay in ars’ proximity to sustain the science-related activities (for this purpose 3 extra, unshielded sleeping compartments will be provided). The impact on radiation dose is not expected to be meaningful because of the short duration. The resulting cylinder, with an inner radius of 75 cm for 3 crew members at a time, provides a size just slightly smaller than standard 1 place beds on Earth, which have a width of 90 cm.

6.3.7 Solar radiation Solar particle events are very dangerous but relatively easy to shield. The inflatable structures, with 8 g/cm2 do not provide enough protection for this, though the inner, rigid part has been assumed to reach an equivalent protection of 20 g/cm2. Additional protection can be reached by using inflatable suits to be filled with the water that is already available for ECLSS [Spillantini, P. – 2014]. Covering the whole astronaut’s body with an additional 20 g/cm2 would require only 400 kg of water per crew member. The discomfort to the crew is considered tolerable since it doesn’t last more than a few days in the worst case, and only 1 day in the average case.

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6.4 ON ORBIT ASSEMBLY STRATEGY

6.4.1 Introduction On orbit assembly is a huge challenge for this mission. Indeed, as Orpheus is a multi-modular mission, it will be needed to rely on different launch and launchers. In the analysis it has been considered several type of still non-existing launchers, or with a low TRL (SLS 70t, SLS 130t, Falcon Heavy). It is essential to put in place a timeline for the on-orbit assembly strategy in order to comply with the launch date of the mission (May 2036). In this section is proposed a strategy that could be adopted to fulfil the mission requirements.

6.4.2 On-orbit assembly strategies For the moment, there are not many studies about what an “on-orbit assembler” would look like Four options have been considered to be a possible space assembly strategy (see Figure 6-15).  Self-Assembly: Each module performs its own rendezvous and docking operations.  Single Tug: A dedicated, reusable space tug performs all assembly operations.  Multiple Tugs: Each tug performs only a portion of the assembly transfers; therefore, multiple tugs are required to complete the assembly task.  In-Space Refuelling: A single tug performs all assembly operations, but is refuelled after a certain number of transfers (new propellant tanks are launched or the tug is refuelled from an orbiting depot).

Figure 6-15: Assembly strategies (GRALLA and DE WECK, 2007)

6.4.3 Assembly model overview Figure 6-15 has presented the assembly strategies that were mentioned above, in a schematic way. It gives an idea of how the assembly would have to be performed, the frequency of module launches against tug/tank launches that should be planned, as a concept.

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Figure 6-16: Models of the space tug (left) and self-assembled module (right). (GRALLA and DE WECK, 2007)

Two different vehicles are considered for this comparison: module and Tug. Figure 6-16 shows notional models for both vehicles: the space tug (left) and a self-assembled module (right). A module assembled by a tug would consist simply of the ‘module’ element, with no extra structure, tank, propellant, or engines In the self-assembly strategy, it can be noticed that the payload mass is limited not only by the capacity of the launcher, but also for the additional systems that have to be present to allow a successful assembly. With the tug strategy the payload capacity in each launch are maximized, leaving the assembly task to the tug. The results of the comparison shown in Figure 6-17 clearly indicate that in-space refuelling of tugs, as modelled here, is the best assembly strategy for nearly all assembly tasks. In tasks with very few modules to be assembled, self-assembly often has a lower overhead mass. The single-tug and multiple-tug strategies rarely have lower overhead mass values than either self-assembly or in-space refuelling. For the Orpheus mission, it is assumed that for the date of the Earth departure towards Mars a space tug with on-orbit refuelling capability will be available.

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Figure 6-17: Strategy comparison depending on the numbers of modules launched.

6.4.4 Number of launches Concerning the Earth departure two options were considered: 1. Use of Cryogenic fuel. 2. Use of non-Cryogenic fuel. The number of launches changes considerably depending on these options because the initial mass to LEO changes with the fuel selection. This is shown in the Table 6-4. Cryogenic fuel has been chosen as primary choice for the reference mission. This leads to an IMLEO of about 4000 tons; this will require 56 launches the breakdown of which can be seen in Table 6-5. Table 6-4: Initial mass to LEO for Earth departure stage.

6.4.5 Time Line to LEO The mass values of the mission suggest that the spacecraft is not going to be put into LEO with a single launch. Considering that the biggest projected launchers are expected to be able to carry between 50 and 130 tons to LEO (NASA, 2012; SPACEX, 2014), the spacecraft will have to be divided into different modules that could dock together and perform the assembly directly into orbit, as it can be seen in Table 6-5.

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Table 6-5: Timeline to LEO

It is worth notifying that the peak number of launches for each year is indicated, and it can be seen that the last year before departure will require a lot of infrastructures in terms of ground segment.

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6.4.6 Conclusion The on orbit assembly procedure has been thought even if it was not entirely entering into the scope of the project, as it appeared to be extremely critical for mission feasibility. A timeline has been created in order to realize how huge the internal cooperation should be in order to realize this project, not only in terms of cost, but also in terms of infrastructures (Ground segment).

6.5 ELECTRICAL POWER SYSTEMS

6.5.1 Power budget The power requirements of the CIV, based on the results of the systems engineering activities, have been estimated at 44.6 kW. This number is the best engineering estimate (BEE) plus margins, though not including uncertainties.

6.5.1.1 Solar Array Sizing Since this is a mission planned for 20 years in the future, it has been assumed that today’s solar cell technology under research will be available at the time of manufacturing. III-V multi-junction solar cells can reach an efficiency of 44.7% (Philipps, S. P. and Bett, A. W. - 2013), which for the solar constant of Mars at apoapsis translates into a generated power of 220 W/m2. Solar panels must be sized not only for the power requirements of the spacecraft, but also for charging the batteries. In the planned orbits the spacecraft will spend 12% of its orbit time in eclipse in Mars orbit and 50% of its time in Earth’s orbit These two cases have been carefully compared Since Earth’s orbit is shorter due to its low altitude and the solar constant is higher, calculations show that ars’ orbit is the worst case and the solar panels have been sized for it. Due to the space radiation environment, a degradation of 15% is expected over an 8 year period. This is based on the SOHO solar observatory, which is located in Earth-Sun L1 (De Donder, E – 2012.), a representative location for the deep-space radiation environment. It is also representative of the solar cycle for the mission profile, which will be assembled in orbit over a period of 5 years (maximum) and will stay in deep space for almost 2 years during a solar maximum. SOHO also experienced the 2001-2003 solar maximum at the end of the sampled period. Also since SOHO spent 8 years in deep space, it is a worst-case scenario for GCR radiation, which will be partially reduced in LEO thanks to the Earth’s magnetosphere This 15% over 8 years is being taken as a whole (as opposed to a per-year calculation) because it is not uniform. During solar minimum the flux of GCR is doubled, and during solar maximum SPEs are more frequent. (De Donder, E – 2012) implies that SPEs have a meaningful impact on the efficiency of solar arrays. The missions for both the CIV and MATV were considered to be of similar length and exposure with regards to the degradation of the solar panels. This is due to the CIV having a Venus fly by and the construction of the MATV in LEO which would cause degradation to the efficiency of the panels. The CIV would likely be up in orbit before the MATV and being exposed for a longer period, however the MATV would be travelling during the solar maximum period, when SPE radiation is higher. For these reasons the degradations in both cases are being considered as equivalent, and in any case this simplifying assumption is a worst case scenario for the MATV. An error of 1% has been assumed for pointing accuracy. Spacing of solar cells when laid out on an array decreases the covered area by 15% (Larson, Wertz 2005). This is a nominal value. The efficiency of individual cells also degrades with temperature by 0.5% per ºC above 28ºC (Larson, Wertz 2005). The thermal control team has calculated during the previous phase that the extreme temperatures will be 80 5 ºC in the closest point to the Sun (85% of Venus’ semi-major axis), - 47ºC in Mars’ orbit before eclipse, and -186 ºC in ars’ orbit when returning into sunlight from eclipses (Patel – 2005) states that lower temperatures increase power output, so the lower temperatures in ars’ orbit will not be an issue. On the other hand, in the perihelion of the trajectory of the CIV the efficiency due to

SEEDS Executive Summary 09/2014 Page 116 temperature will be 99.5% x (80-28) = 54% of the nominal efficiency. As the solar constant at that distance will more than double that at Earth’s proximity, this effect can be neglected In Earth’s proximity temperature can also be very high. Even if the solar constant is not as high as in the perihelion of the trajectory, planetary albedo and infrared can raise the temperature substantially. (Larson, Wertz 2005) suggest a 15% nominal value. Since the solar arrays have been sized for ars’ proximity and the solar constant on Earth doubles this value, in this case this effect can also be neglected. Finally, the I-V curve may also cause efficiency losses. However, the above mentioned efficiency of 44.7% (Philipps, S. P. and Bett, A. W. - 2013) is assumed to be the result of an already optimized curve.

6.5.1.2 Battery Sizing Mars orbit is the worst case for the duration of eclipses. However, the spacecraft will be assembled in LEO prior to departure. Since the assembly time has been estimated for 5 years, the batteries of the CIV will undergo 29500 cycles of charge and discharge. This has to be accounted for. The MATV will only stay for 2 years in LEO (worst case) so the total number of cycles is 11500. Boschetti, D. suggested in his lecture titled Electric Power Systems and EMC that at a depth of discharge (DoD) of 30%, the FADE of batteries has been estimated to 0.3 after 30000 cycles. This is consistent with the numbers in (Patel, 2005). For the MATV it is 0.8. So at EOL batteries of the CIV need to store 62 kWh of energy and 40.5 kWh in the MATV. This translates to BOL powers of 206 kWh and 51 kWh, respectively. Note that batteries in the MATV will not be used for propulsion purposes. Different battery types have been considered. Mass of batteries grows with lower costs, but the launch costs, estimated at $5000 per kg (Wall, M. – 2012), are not exceeded by the higher cost of lighter batteries. Sony 18650 batteries can store 6.66 Wh per cell with a mass of 42g [8, p43]. This translates into 1.3 tonnes of batteries for the CIV and 401 kg for the MATV. For each charge/discharge cycle an average of 20% of the energy is dissipated as heat (Patel 2005, p219). This requires oversizing the batteries by about 20%, as most of this energy is dissipated during discharge. Oversizing solar arrays needs probably a smaller factor, but since not enough information is available for this case, the worst case is a safe assumption so they will be oversized by 20% as well. These numbers lead to a calculated area of 516 m2 of solar panels for the CIV, with a total mass of 584 kg. Batteries will be 1970 kg. For the MATV, calculations of solar panels are analogous, but based on the higher power requirement the area is 4400 m2 and the mass is 4900 kg. A margin of 20% is included in both numbers.

6.5.1.3 Voltage ECSS‐E‐ST‐20C establishes a voltage of 110V-120V for high power systems. The ISS uses 124V according to the manufacturer [Boeing, International Space Station Electric Power System]. 120V will allow reuse of several designs and comply with the standard, so it has been taken as a reference.

6.5.2 Electric Power Systems Architecture Together with the Power Budget, the interface requirements constitute one of the most important definition steps for the definition of the EPS architecture. All the data collected from the main steps have been translated in the Electric Power System Requirements, which shall comply with the ECSS standards. In more details, for the purpose of this work, reference standards include: ECSS-E-ST-20C, ECSS-E-ST-50-14C and ECSS-E-ST-20-08. Generally speaking, any type of EPS architecture can be conceptually organised in three different major elements: the Primary Power System which basically comprises the Primary Power Sources, the Power Management and Distribution System which mainly provides functions of Charge Control, Power Conditioning, Regulation and Distribution and, finally, the Storage System mainly related to batteries.

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6.5.3 The Fully Regulated Direct Energy Transfer Bus

Figure 6-18: Conceptual EPS Architecture [Image from Larson, W. J. - 1999] In accordance with the ECSS-E-ST-20C, for higher power load budget the Nominal Bus Voltage selected is 120 Vdc. Several electrical power sources exist for space applications. For the purpose of this work the Photovoltaic-Battery System results to be the most appropriate one to accomplish the mission requirements. On the basis of this choice, the most appropriate EPS architecture selected results to be the Fully Regulated DET Transfer bus, also simply known as regulated bus. Its general baseline architecture is shown in the figure below. On the basis of the investigation of the ISS EPS architecture, for the purpose of the mission investigated, the EPS Primary Power Subsystems are: Solar Array Wings (SAWs) with a total mass of 584 kg associated with Shunt Sequential Units (SSUs) in order to control the output voltage value of the SAs, and the SA drive mechanisms in order to satisfy the pointing requirements. Finally, eight DC Switching Units (DCSUs) for

Figure 6-19: Fully Regulated Bus reference architecture [Image from Patel, M. R. – 2005]

Power distribution completes the Primary Power section. The Secondary Power Subsystems selected are: four Main Bus Switching Units (MBSUs) necessary for manage and distribute the Primary Power. Twelve DC/DC Converter Units (DDCUs) which represent the real workhorse of the Secondary Power Subsystems, and 200 Remote Power Controller Modules (RPCMs) which represent the direct interfaces with the electrical equipment, or loads. The final overall mass estimated for the EPS results to be 5,81 tonnes, out of which 406kg are actually part of the TCS budget, leading to a total of 5.2 tonnes.

6.5.4 EPS Harness One of the concerns of the current study is that in case of failure and overcurrent, cables might overheat and catch fire. The Federal Aviation Administration has performed numerous tests on this subject. Approved test methods are defined by the requirements in (FAA Doc. No. 5066, 1964). Several standard wires typically used in aerospace applications have been approved under such tests (Ochs, 2006), and NASA has also researched on insulation materials (Langley RC, 2008). Kapton insulation has been widely used for aerospace applications, including the Apollo Lunar Module (NASA History Office, 2009) and the Space Shuttle (NASA, 2005). Being non-flammable makes it a very good insulator for space, but it is not very resistant to mechanical bending and can break. Both sources cited above express concerns about this. According to FAA tests, Polytetrafluoroethylene is a good non-flammable alternative (Ochs, 2006). The currents of main power lines vary from 0.01A to 39A, for that reason several different wire thicknesses have been chosen based on the recommendations of Patel (2005) and classified according to American Wire

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Gage. Such recommendations are based on operating temperatures of 70ºC in vacuum, which is a very harsh environment when compared to the CIV’s pressurized living space The thermal power dissipated by main power lines has been calculated as 0.4 mW.

6.5.5 Electromagnetic Compatibility The applicable EMC issues on this phase of a study are at the system requirements level. This is based on ECSS-E-ST-20-07C. Margins are 20 dB for safety-critical equipment and 6 dB for mission-critical. Most of the defined test points are ECLSS equipment. Each module has been matched to the launch vehicle it will be loaded on The user’s manual of Ariane 5 defines all EMC requirements, for that reason it is expected to be the same with Ariane 6, used to launch crew resources. Other launchers, such as Falcon 9, have an EMC assessment performed by the company prior to launch. Radio frequency compatibility (RFC) requires no susceptibility to bands in 7.1 GHz, 8.4 GHz, 32 GHz and 34 GHz as they will be used for communications, transmitting at high power. DC magnetic field compatibility is also important during assembly in LEO, so the spacecraft’s magnetic dipole shall not exceed 10 Am2.

6.5.6 Nuclear Thermal Propulsion as Electrical Power Source From the trade-off analysis performed for the selection of the Propulsion System in section 3.4, it was defined that the main propulsion source for the CIV will be the Nuclear Thermal Rocket Engine (NTRE). From the beginning of the Electric Power System analysis, the selection of the most appropriate architecture was driven by the possibility to extract a certain amount of power. This corresponds to the CIV Housekeeping Power Level requested of 44 kW, directly from the Nuclear Propulsion System. Indeed, besides providing high thrust and high Specific Impulse, the NTR also represents a rich energy source because it contains substantially more enriched U-235 fuel in its core reactor than is consumed during the primary propulsion manoeuvres. Hence, a preliminary feasibility study has been performed in order to properly answer this issue, which can strongly influence not only the EPS architecture of the CIV, but also its overall mass and design. This is due to the change of the primary power source, represented by the SAWs. To extract some kW from the NTRE, it is necessary to select the Bimodal Nuclear Thermal Rocket (BNTR) propulsion technology, which is actually indicated by NASA only for NERVA Nuclear Thermal technology, constraining the choice to this technology. The term “Bimodal” indicates that the NTRE can be configured in two different operational modes: Thrust mode and Power Production mode, requiring the insertion of the control bars to be able to manage the level of the Power Core released by the reactor. When switched-off by inserting the control bars, 10’s of kilowatts of electrical power (kWe) can still be generated for housekeeping needs, like crew life support and high data- rate communications. With a more advanced generation of BNTR, the electrical power levels that can be extracted can be very high: from ~100’s kWe to 1 MWe, helping to solve the boil-off problem of the propellant. During the high thrust “propulsion phase”, 100’s of Wt are produced and removed using CH4 propellant pumped through the engine’s reactor core During the power generation mode the BNTR’s reactor continues to operate but at power levels equal to the 7% of the nominal power level reached during the thrust phases. The thermal power produced by each core reactor during the no-thrust phases, is directly related to the heat, or thermal power, produced by the decay process of fission products generated during the thrust phase. Through the following empirical relation, knowing the Burn-Time duration ( ) of each single manoeuvre performed by the NTRE and the associate Elapsed Time ( ), which represents the time that passes until the next manoeuvre, it is possible to compute the thermal power produced by the decay process ( ̇ ). ̇

[( ) ( ) ] ̇ Considering the NERVA engine designed for the purpose of this mission, the insertion of the control bars inside the core reactor greatly reduced the thermal power from the thrust level of 1161 MWt to 81.3 MWt.

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Energy generated in the reactor fuel assemblies (equation above) is supposed to be removed utilising a “closed” gas loop (Figure 6-20) that carries the methane, due to the following energy balance.

( ) ̇ ̇ ( )

Resolving for , it is possible to calculate the temperature reached by the methane inside the core in the steady state condition, expressed by the equation directly above. Obviously the temperature reached by the methane cannot reach higher values due to the core and turbine material constraints, but should be optimised for the Brayton cycle inside the turbine. Also the dissociation temperature of the methane itself constitutes an important operative temperature constraint. As a result, the methane temperature inside the nuclear core reactor should be managed in order to match these temperature constraints, reaching the appropriate value of the mass flow rate. From the analysis conducted, the system should be designed in order to reject a maximum power level of 83.11 MWt, for which the temperature of the methane goes from the initial value ( ) of 119 K, to the final optimised one ( ) of 1,362 K, with a propellant mass flow rate (inside the core) which passes from the nominal thrust level of 76.6 kg/sec, to the 11 kg/sec for the power generation mode. Through this regulation, the core temperature goes from the 2361 K of the nominal thrust phase, to the 1,118 K of the power generation phase. Considering an efficiency of 30% to translate the thermal power into electrical, the optimised methane mass flow rate is able to extract 187.87 kWt which, as a first approximation, corresponds to 56.4 kWe, more than the power budget loads estimated for the CIV. This means that, in order to verify the CIV EPS requirements, ~13 kWe should be deviated to the radiators in order to be rejected.

Figure 6-20: NERVA Bimodal Operation concept [Elaboration on an image credited by NASA]

6.5.7 Conclusion The analysis performed reveal good results for the Electric Propulsion System. However, compared with photovoltaic-system architecture, due to the high power level that must be rejected by the radiators, the adoption of the architecture proposed in Figure 6-20 is not justified. This is due to the increase of the mass and especially in the complexity level of the CIV. Also the risk for the safety of the crew is increased. For these reasons, the use of photovoltaic arrays has been chosen for power generation.

6.6 THERMAL CONTROL

6.6.1 Overview and Topics Due to the temperature range systems can operate at and to provide a suitable habitat for humans it is necessary to develop a comprehensive Thermal Control System (TCS) accommodating the needs of all aspects of the system. This was obtained through a combination of passive techniques (i.e. MLI) and active

SEEDS Executive Summary 09/2014 Page 120 techniques (i.e. thermal loops) through available technology where possible and techniques that appeared feasible in the near future.

6.6.2 Study Approach By adhering to constraints provided by other systems such as EPS, ECLSS and Mechanical Design it was possible to develop a system to manage waste heat and provide heating to the areas that require it whilst remaining within power and mass budgets. The approach was to utilise passive thermal control where at all possible and active thermal control in all other cases. Through calculations of geometry, thermal characteristics, thermal environment and phase it was possible to develop a suitable system for nominal operations and to identify its capability at handling unexpected or peak loads. The mathematical approach is based on the following very simple equations:  The Stefan Boltzmann Radiation Equation:

Where P = Radiative Power (kW), = Stefan Boltzmann's Constant (5.670 ×10-11 kW m−2 K−4), Emissivity ε, Area (m2), Temperature (°K)  The Solar Absorption Equation:

Where P = Absorbed thermal load (kW), = Stefan Boltzmann's Constant (5.670 ×10-11 kW m−2 K−4), Absorptivity α, Area (m2); and Sc= Solar Constant (kW/m2)  The Specific Heat Capacity Equation:

E = Total energy required to raise the tank by the temperature specified (kJ), M = Mass (kg), Cp = Specific Heat Capacity (kJ/[kg °K]), ΔT = Temperature Change (°K).  Clausius-Clapeyron equation:

( ) ( )

Where P1 = Pressure 1 (Bar), P2 = Pressure 2 (Bar), ΔH = Enthalpy of Vaporisation, R = Universal Gas Constant (8.314 J/[mol K]), T1 = Temperature 1 (°K); and T2 = Temperature 2 (°K).

6.6.3 Pointing In order to maintain the low temperatures required for the cryogenic fuel tanks it would be best to design a craft configuration that accommodated all the systems minimising the area facing the Sun. Therefore it was decided to point the longitudinal axis of the CIV in the direction of the Sun in order to minimise heat loads and the complexity of the TCS.

6.6.4 Habitable Volume Heat Loads The Habitable Volume (HAB) generates the most heat waste through the use of electrical equipment. In order to size the loops and radiators it is important to calculate the thermal loads to be rejected, this is done by making the following assumptions:  All electrical power generated by the Electric Power System is converted into heat within the HAB (57.33kW);  The exterior of the HAB is covered in currently available MLI with insulative properties of Absorptivity (α) and Emissivity (ε) which are 0.09 and 0.27 respectively;

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 Heat loads from the Sun are absorbed by the surface area on the front of the HAB due to pointing;  The sides and rear of the CIV are rejecting heat into deep space.

6.6.4.1 Heat Loads Knowing the MLI properties, the surface area perpendicular to solar flux and the maximum value of the solar constant along the trajectory, it is possible to calculate the heat load by the Sun on the front of the craft by using the Solar Absorption Equation. This value is approximately 31.2kW. In a similar way it is possible to calculate the heat rejected by the CIV, this is done by using the Stefan Boltzmann Radiation Equation and inputting the radiating area, the MLI's emissivity and the average temperature of 21.75°C (the ECLS System is providing induced convection through a fan system in such a way that the air temperature through the craft is approximately constant). The value for heat rejection is approximately 31.6kW Summarizing the heat load due to electrical equipment, human and solar heat loads, removing the radiated heat by the HAB it is found that the heat needed to be removed and rejected by the TCS is 49.91kW. The aim of the TCS is to collect this heat, transport and finally reject into deep space through radiators.

6.6.4.2 Cooling Loops The subsystems which achieve the heat transportation are included in the Thermal Control System (TCS). The heat waste is finally rejected into deep space trough Radiators. The heat transportation happens in two different ways:  Passive transportation  Active transportation The passive transportation of heat is achieved by heat pipes which are able to bring the heat from hot equipment to cold plates or heat exchangers. Due to the absence of gravity it is enough a temperature difference between two points that the heat pipes works autonomously without the utilization of any moving components and therefore they have high reliability and long life. The Active Thermal Loop (ATL) needs to achieve large heat transportation; in fact it forces the heat transfer by pumping a working fluid into the loops. It is used when very stringent thermal requirements are applicable or where high dissipating equipment is accommodated. Finding inspiration in the ISS TCS, BOEING (2006), Orpheus’ CIV includes three types of active loops: external loops, internal low temperature loops and internal moderate loops.

6.6.4.3 External Loops One external active thermal control loop consists of an ammonia loop (due to the attractive thermal properties and working temperature range) that collects heat from the interface heat exchangers and from external electronic equipment mounted on cold plates (each ammonia loop contains five cold plates, where two are for redundancy). The heated ammonia circulates through large panels located on the exterior of the spacecraft, releasing the heat by radiation to space that cools the ammonia as it flows through them. A different fluid that may be utilized but is still in development is HFE7200. The external loop is a pressurized loop (20.7 bars) in order to avoid the freezing of the ammonia when it passes through the radiators. The nominal ammonia flux is 0.736 kg/s. The nominal outlet from the radiator is 2.8°C in order to avoid the freezing of the internal fluid in the heat exchanger and to avoid the internal loop running below the dew point (water is used). The outlet temperature is monitored in order to regulate the angle of the radiators maintaining desired temperatures. In this way radiators may have a variable Sun facing angle, in the nominal condition, this angle is 0° (parallel to incoming solar flux) to maximize the heat rejection capability. Due to the CIV configuration, made of two separated toroidal volumes, two loops are chosen to reject the half of the entire nominal power. In addition, having two loops there is also a hot redundancy; in fact the two loops are connected by valves which may connect them in case of specific needs, each loop is thought to be always single failure tolerant.

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The loop is composed by Heat exchanger, Pump module (including pump, gas trap, filter and accumulators), Cold plates (large and small), Stepper valves, Solenoid valves, Flux and pressure sensors, temperature sensors one way valve, motors and beams. The level of uncertainties chosen is 1.05 for the equipment and 1.25 for the pipes and the NH3 mass due to the uncertainty in the flow path of the pipes and their respective NH3.The overall mass estimated (including uncertainties) is 1.52 tonnes for the loop A (the loop servicing the HAB nearest to the radiators) and 1.58 tonnes for the further loop (loop B, which has slightly longer loop). There are three different levels of temperature for the external fluid as the different colours in Figure 6-21 represent, where red, green and blue symbolise hot, moderate and low respectively.

Figure 6-21: Active Thermal Control Architecture

Table 6-6: Key for Figure 6-21 Pressure and flux Accumulator Filter sensors (nr. 3)

Gas trap Radiator Cold Plate DDCU Solenoid valves Cold Plate MBSU Motor and Beam (ON/OFF) Stepper valves Exchanger with Internal Low One way valve temperature loop (mixer) Temperature Exchanger with Internal Pump Moderate Low temperature loop sensors (nr. 3)

The power budget for one external loop is mainly due to the consumption of the pump. However there is also an small electrical power needed for the valves, in fact, in order to guarantee the one failure tolerance, the ON/OFF valves are solenoid valves, it guarantees in case of failure that the valve closes autonomously; only one redundant valve is needed for this kind of valve. Regarding the regulating valves (mixing valves) it is thought to have in total 4 valves for one loop (3 redundant) in order to guarantee one failure tolerance in case the valve fails in both modes: always open or always closed. For these valves a Stepper-like valve has been chosen. The estimated power budget for each loop is 0.65 kW for the loop A. With regards to the second loop, the power will be a little higher since the pipe length is greater so there are more losses which require more power; the power budget for the second loop is estimated to be 0.76 kW.

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For each required measurement it has been considered that three detectors are needed with two for redundancy in order to ensure that the information obtained is correct even in the case one sensor fails and gives a wrong output. As it has been shown above, there are two different kinds of exchangers (low temperature exchangers and moderate temperature exchangers) for different thermal control purposes which exchange heat flux with the internal loop.

6.6.4.4 Internal Loops The internal heat waste is removed in two ways, through cold plates and heat exchangers, both of which are cooled by the external loop with circulating ammonia loops on the outside of the craft. The internal spacecraft has to manage heat for ECLS and scientific equipment. It drives two different needs in term of heat management. Using ISS heritage the internal loops were derived, the internal heat is managed by two separated loops interfacing with one external loop: the Low Internal Loop (LIL) and the Moderate Internal Loop (MLI). LIL is used mainly for ECLS for the habitable module. The water is exchanging heat with ammonia with the lowest temperature (2.8°C); this is necessary also to achieve the possibility to keep the spacecraft above the dew point to avoid water condensation. The external loop, with an intermediate temperature is then entering into the second heat exchanger to collect heat load from the MIL: the nominal minimum temperature for the water in the MIL is 9°C. The internal loops mass and power evaluations have been calculated starting from the Orion Multi-Purpose Crew Vehicle (MPCV) scaling the power consumption estimating the losses in the pipe. Heat pipes are evaluated only for the MIL; cold plates and fans are instead evaluated in the ECLS System. The final results are 824 kg and 1.06 kW for the LIL, 1580 kg and 2.03 kW for the MIL.

6.6.4.5 Radiators The radiators chosen are proposed future lightweight panels suggested by Hanford and Ewert (1996). The radiators have a mass of 7.5 kg/m2 and have a stated emissivity of 0.86.Using the Stefan Boltzmann Radiation Equation it was possible to calculate the area of radiators required and the mass of these radiators, which are respectively; 262 m2 and 1.96 tonnes (including the structural component). Each of the two external cooling loops has its own radiator wing, on each wing there are two radiator units, (for redundancy). Each radiator unit is made in a configuration of 6 panels wide and 24 panels long giving dimensions of approximately 13.4m long and 4.9m wide (each panel has 0.56 m length and 0.81 m width).

6.6.5 Cryogenic Pressurised Tank System The Cryogenic pressurised tank system is a low temperature unit; there are 22 units which need to be maintained at low temperatures to keep the fuel in fluid phase. The fuel (methane and oxygen) are pressurised to 9 bars, which drives the temperature range constraint change according with the Clausius- Clapeyron equation, Chemteam (n.d). In addition a 15 degree margin was applied. The operational mode range is 133.6 K - 105.7 K.

6.6.5.1 Insulation Approach The approach taken is to heavily insulate the front of the tank which would receive the largest thermal load from the Sun (Advanced MLI Thermal Engineer (n.d) absorptivity α 0 024), the sides of the tanks utilised currently available types of MLI (α 0 27) and the rear of the tanks are white painted (emissivity ε 0 93) to act as heat rejection surfaces. To keep a constant temperature inside the tanks due to the absence of gravity it is necessary to mix up the fluid. Different solutions are proposed, such as fan, heat pipes and Stirling motors. Utilising this approach to insulation, the fuel tanks maintain an equilibrium temperature at the maximum temperature that were calculated during the closest approach to the Sun; this was to allow for margin in the event of pointing errors and manoeuvres.

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6.6.5.2 Tanks Pointing Accuracy Pointing accuracy became a key driver as it was important in maintaining the desired tank temperatures. For this reason a constraint was derived such that the maximum pointing error was calculated. It has been defined as to increase the equilibrium temperature of the tanks by one degree above the set upper normal operating temperature. This was done by calculating the thermal loads on the area perpendicular to the incoming solar heat loads and the heat radiated by all other radiative surfaces. It was possible to calculate the constraint on pointing accuracy, this value is 11°. To ensure that the temperature of the tanks does not raise more than one degree above the maximum temperature limit an additional margin has been added, this drove the AOCS pointing error constraint to be 10°.

6.6.6 Phases During the 30 hour swing-by of Venus which is the period when at least 1 Watt thermal is absorbed by a tank and assuming worst orientation, the system can operate nominally, in fact the tank temperature raises by only 0.29K. The system may remain in a worst orientation when at the closest point to the Sun for 2.64 hours before raising the tank temperature by one degree; this factor can be used to estimate maximum burn time. Finally in Mars eclipse in which the tanks are left with no thermal inputs, it takes 0.72 years to drop the temperature of the tanks by one degree. Therefore it is considered that the tanks have enough thermal inertia to avoid freezing during all the following mission phases.

6.6.7 Construction Scenario Keeping the tank in fluid phase became critical during the LEO assembly scenario due to the combination of infrared, albedo effects and solar incidence. The LEO orbit was split into quarters as an approximation, during these orbits the tanks will remain with the Advanced MLI coated front face pointing in the direction of the Sun (even in eclipse). There are therefore three scenarios described below. Through calculations and research it was decided that by launching the tanks towards the end of the construction phase, decreasing the time in orbit, it may be possible to rely on thermal inertia and avoid the boil off. By freezing the propellant fluid the volume decrees, so it is possible to add excess fluid (maximizing the launcher capability in tem of mass and volume). The extra fluid provided will cool the useful mass due to latent heat. In addition the overall mass has to change phase from solid to liquid. With this solution it was theorised that its maximum it would be possible to remain in LEO for 208 days which may be sufficient but due to the complexities of phase changes this idea may not be feasible, therefore other techniques such as cooling to just above freezing point thus increasing thermal inertia without the problems of phase changes. Furthermore it could be possible to increase the insulation of the tanks by covering more of the tanks in Advanced MLI or Sun Shields increasing the time of permanence in LEO or launching additional tanks.

6.6.8 Photovoltaic thermal control For the thermal control of the solar panels the cases in hottest and coldest cases were assessed. The closest approach to the Sun (hot scenario) it is considering the following equation: . The equilibrium temperature achieved is 353.5 K (80.5°C). For the cold scenario it has been calculated the solar array temperature at the end of the eclipse which is 86.8 K (-186.2°C). Both these temperatures are acceptable according to Larson and Wertz (2005) and according to industry experts the speed at which the temperature changes in these cases is not an issue.

6.6.9 Summary In summary the tanks have been designed to self-regulate their temperature based on thermal inertia and correct craft pointing. The photovoltaic cells have been calculated to remain within an acceptable temperature range. The CIV has water based internal fluid loops that interface through a heat exchanger

SEEDS Executive Summary 09/2014 Page 125 with external Ammonia loops which transport the heat to the radiators. At the radiators the heat is rejected into space. The mass and power needed for TCS components are respectively 9.88 tonnes and 15.6 kW.

6.6.10 Future Areas for Research Assuming all approaches are reasonable including their margins, within budgets and constraints it is considered that with the discussed configuration and the development of the necessary equipment it is a suitable system. The areas of research are mainly the development of the insulation needed for the front ends of the fuel tanks. Other areas for research are mainly into the use of new coolants such as HFE7200 which may prove a more useful coolant fluid in the future. It can be considered that if a reliable thermodynamic loop system were to be developed in the future which may reduce the radiator mass.

6.7 COMMUNICATIONS

6.7.1 Introduction The communication systems are responsible for the transfer of command, telemetry and bulk data between all assets in the Orpheus programme and the Earth segment. Described in this section are the critical links necessary to facilitate the mission profile for all assets that are physically separated. Additional considerations for storage of sufficiently high fidelity data, for archive and retrospective analysis purposes, have been studied but are not shown here. The purpose of this study has been to size hardware (mass, power and performance) and understand key driving features necessary to meet mission requirements. Considerations of physical, technical and legislative constraints have been applied to make the result as applicable and useful as possible for future work. Based upon a first order analysis two systems were determined to be viable for the mission, Ka band transceivers and optical packages. X band had been disregarded due to large bandwidth consumption and restricted spectrum (Kwok A. 2009). Through a trade study it was seen that optical communications present the best overall performance from a system level perspective as shown in later subsections. The study utilises references regarding previous flight heritage and proposed missions. Specifically the NASA proposal for Mars Laser Communication Demonstration (MLCD) that would have been a demonstration of high throughput optical links, up to 100Mbps (Townes et al 2004), at distances of 1 AU. In terms of flight proven hardware there are few examples of high performance communication packages between Mars and Earth today. NASA’s ars Reconnaissance Orbiter (MRO) launched in 2005 was primarily designed with science capabilities in mind, but carries on board an impressive 100W X band package and an additional 34W Ka band demonstrator. At its minimum operational distance MRO can nominally reach 5.2Mbps while dropping down to 0.5Mbps at its greatest (Taylor et al 2006). Optical inter-orbit communications have been successfully performed between JAXA’s OICETS LUCE (Laser Utilizing Communication E uipment) and ESA’s ARTE IS OPALE (Optical Payload Laser Experiment) packages. Transmitting at wavelengths from 800nm to 850nm a maximum throughput of 49.37Mbps was achieved (Takashi 2012) Furthermore direct oon to Earth transmissions have been performed by NASA’s LADEE LLCD (Lunar Laser Communications Demonstration) to two Earth based telescopes with a downlink of 620Mbps and an uplink of 20Mbps (Boroson et al).

6.7.2 Environment Definition To size the system a thorough understanding of the operational environment is required, knowledge of the separation between transceiver terminals, propagation losses and alignment error are crucial. As it should be mission analysis is the dominant factor in achieving mission feasibility. However, what is best from a mission perspective may not be ideal from a communications point of view. This can be seen in Figure 6-22 where the separation between Earth and Mars varies significantly with time. For the launch window in mid-2036, with an interplanetary transfer of the order 300 days, the Mars operations phase is spatially placed at a time where the separation between Earth and Mars is approximately 1.8 AU.

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Such distances place huge minimum constraints of the size of hardware necessary to perform high-speed communications. For example a package similar to the 100W X band transmitter in MRO would achieve a transfer rate below 1Mbps. This would be considered undesirable for a human class Mars mission.

Figure 6-22: Earth-Mars separation in AU as a function of time, beginning June 2036. Vertical highlights (orange) show Mars proximity phases for different launch dates. Lower line (red) is the separation at which MRO peaked to 5.2Mbps. Upper line (green) is the separation of lowest MRO performance 0.5Mbps.

6.7.2.1 Atmospheric Attenuation As electromagnetic signals propagate through a medium their intensity is reduced, typically as a function of distance. This reduction is described as path losses of a transmission, which include propagation through free space and interactions with the dense atmosphere of Earth. Where possible these losses should be minimised to allow for higher channel capacity.

∑

The Ka band, with frequencies around 32GHz, can pass readily through Earth’s atmosphere under ‘clear sky’ conditions. They are however heavily attenuated by increased water content, such as clouds. To calculate the effects the atmosphere has on transmitted signals CNES’ PROPAGATION (Lacoste 2010) software package was used. This implements the semi-empirical models outlined by the International Telecommunication Union (ITU).

6.7.2.2 Communications Availability At present there is no distributed network of assets to provide continuous communications with spacecraft at Mars. A main requirement for workable communications is an unobstructed path between terminals. Eclipse phases during the mission result in a duty cycle of less than 100%. Thus a terminal has to be sized to achieve the desired average throughput by using periods of higher peak performance. There are few software packages, even fewer that are readily accessible, capable of performing the analysis needed for the mission. Therefore scripts were developed in MATLAB using celestial ephemerides data, provided by JPL Horizons, to determine important information regarding availability windows during the

SEEDS Executive Summary 09/2014 Page 127 mission phases. Combining trajectory outputs of the CIV from NASA’s GMAT (General Mission Analysis Tool) in the script availability as a function of time can be determined. The output of the script can be seen in Figure 6-23. During the critical period of Phobos operations the CIV has at maximum 86% availability every Phobos orbit.

Figure 6-23: Plot of CIV communications availability, either available or not, as seen by Earth. The periods of zero availability (red line) represent eclipses due to Mars. The continuous (blue) line is eclipse due to Phobos, of which there are none.

6.7.3 System Sizing

6.7.3.1 System Requirements  Uplink: BER ≤ , Downlink: BER ≤ (Bit Error Rate)  Doppler: 0.1 mm/s over 60 sec.  Ranging: 1- 10 m over 60 to 600 sec.

6.7.3.2 Data Rate Analysis In order to understand the level performance required of the system, an estimation of the quantity of data that is to be transferred and received per day by each phase is necessary. Upon review of the goals and achievements of notable missions, such as Apollo and the ISS, comparisons can be drawn against project Orpheus. A key feature is the need to relay live HD video of the Phobos lander as it makes first contact, back to Earth. With a system budget, time frame and availability duty cycle, from a previous section, data transfer rates can be calculated that drive the design of the communication systems. These are shown in Table 6-7 and Table 6-8 for the CIV and Phobos Lander respectively. It can be seen that the mission is feasible within near future technologies, with a maximum sustained downlink rate of 29.5Mbps.

Table 6-7: CIV data volumes per day per mission phase. Phobos Proximity is in bold. CIV Downlink Uplink Downlink Uplink

Orbital Availability Data per Data per Req-Rate Req-Rate Phase Period (%) day (Mbits) day (Mbits) (Mb/s) (Mb/s) (Hr)

Assembly phase (C-E-1) 1.5 0.95 4800000 2020100 58.48 15.59

Transfer Phase to mars (C-I-1) 24 0.95 4000 2000 0.05 0.02

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Capture Orbit Phase (C-M-1) 1 0.50 5000 2000 2.78 1.11

Phobos Transit Phase (C-M-2) 10 0.80 5000 20000 0.17 0.69

Phobos proximity Phase (C-M- 7.5 0.86 2005000 2020100 29.47 89.07 3)

Transfer Phase to Earth (C-I-2) 24 0.95 6000 3000 0.07 0.04

Earth proximity (C-E-2) 24 0.95 4000 2000 0.05 0.02

Table 6-8: Phobos lander data volumes per day per mission phase. Phobos proximity is in bold. Phobos lander Downlink Uplink Downlink Uplink

Orbital Availability Data per Data per Req-Rate Req-Rate Phase Period (%) day (Mbits) day (Mbits) (Mb/s) (Mb/s) (Hr)

Assembly phase 1.5 0.95 2000000 20 24.37 1.54E-04

Transfer Phase to mars 24 0.95 20 4 8.12E-05 4.87E-05

Capture Orbit Phase 1 0.50 20 4 3.70E-03 2.22E-03

Phobos Transit Phase 10 0.80 20 4 2.31E-04 1.39E-04

Phobos proximity Phase 7.5 0.86 2000000 100 29.39 4.41E-03

Graveyard Orbit phase 10 0.9 20 0 2.06E-04 0.00E+00

6.7.3.3 Link Analysis. Key parameters of the communications hardware were ascertained after a sufficiently detailed link analysis was performed. Major factors required were the sustained data rates, as described in the previous section, and signal losses along the whole path link. The following describes the problem, using reference data from case studies where available. The link equation, in its general form

( ) ( )

where; is the received carrier power (Watts), and is the transmission power (Watts). Further details can be found in Hemmati and Chen (2005), pg 89. Improvements in the received signal power can be accomplished through adjustment of the design variables, such as transmit power, wavelength and system efficiencies. For digitally coded systems a useful number to work with is the Energy-Per-Bit to Noise Density ( ), which described the performance of the link. Where is the bit rate and the noise power density .

( )

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For a desired bit error rate (BER), such as , a of 10.5 dB is necessary for BPSK and QPSK modulation schemes. An additional link margin of 3 dB is applied size the system to mitigate for non-ideal weather conditions.

6.7.3.4 System Level Comparisons To better understand which system is most suited a side-by-side comparison between Ka band and optical laser must be undertaken. Using the method outlined by Fielhauer and Boone (2012) a Figure of Merit (FoM) can be established for a direct comparison between technologies at the system level, independent of distance. The system-level metric includes the space-base terminal performance (be it RF or optical) and the ground receiver aperture and performance.

( ⁄ )

Where D represents the distance between the terminal apertures, bpsMax is the maximum transmitted information rate, mass (kg) and power (Watts). This is applicable to ‘clear sky’ conditions Figure 6-24 shows the results of the system level comparison. It can be seen that direct to Earth transmissions, at distances of 2.35 AU, the optical laser package clearly has the best system level performance. It is noteworthy that an X band (8GHz) system outperforms a Ka band (32GHz) system for a given peak data rate, due to the greater atmospheric attenuation Ka band experiences. However, an X band system such as this would vastly exceed the legislated 10GHz bandwidth allocation for space science return. An optical communication system is selected going forward.

Figure 6-24: System level comparison of communication packages for different terminal separations, close (1.34 AU) and far (2.35 AU). MRO’s X band and Ka and systems are shown for reference.

6.7.4 System Architecture Top-level system architecture, as displayed in Figure 6-25, was made outlining the primary links necessary to meet the mission requirements. Some assumptions have been made about the ground segment, but it is beyond this study to go into detailed analysis. At present there is an array of existing infrastructure for RF downlinks, such as DSN (NASA), ESTRACK (ESA), SDSN (RKA), CDSN (CNSA) and ISDN (ISRO).

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However it is the assumption that a sufficiently capable receive terminal, whether Earth direct or Space relay to surface, is available.

Figure 6-25: Communication system architecture showing primary communication links between assets.

6.7.5 Conclusion The scope of the study was to size hardware capable of sufficient data transfer to complete the mission requirements. Considerations of physical, technical and legislative constraints have been applied to make the result as applicable and useful as possible for future work. The output of the study is that a 2W 1500 nm laser based optical system with a system mass 199 kg and power 317 W meets the necessary communication requirements through all phases of the mission.

6.8 TELEMETRY, TRACKING AND COMMAND (TTC).

6.8.1 Introduction TTC is one of the mission critical sub-systems and the amount of valuable scientific information returned is based on its performance. The system should be able to provide two ways information flows throughout all the mission phases, this data should be error checked using a high performance error handling system. The TTC sub-system was designed using new techniques and the principles as described by L, Benjamin (1997), P, Fortescue (2011) and NASA (n.d).

6.8.2 Requirements By referring to the data requirements of the communication system, it is necessary to assess the data and calculate the actual amount to be processed or analysed. This is difficult to predict for future missions, as in 20 years the speed and lifetime efficiency of the following components would change significantly. There are no missions of this scale, as such it is not easy to predict the required data handling and the effects on avionics designed for a deep space manned mission. The required technological advancements could be gained throughout the 20 years preceding the mission, through precursor missions or ground based TRL validations. The requirements could be classified in overall functions as shown in Figure 6-26: Transition to Components and from these requirements it is possible to derive the major functions of avionics. Various missions in the past have helped to identify the necessary components of the system. In Figure 6-26 it is possible to see the transition from functions to components.

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Figure 6-26: Transition to Components

6.8.3 TTC Architecture The Architecture of TTC is not simple; there are some components that are critical such as processors, data storage and components designed to deal with single event upsets. Due to the added complexity of the wirings and interfaces that have to be connected to the micro controllers with efficient paths and parallel connections, the TTC architecture shown in Figure 6-27 is non-trivial.

Figure 6-27: Architecture of TCC The parallel connections are used in multi-tasking when dealing with large data flows to resolve bottleneck issues with parallel controllers. All the systems and instruments are to be connected in parallel and should follow the Failure Mode Effects Analysis FMEA recommendations. The purpose of using this recommended approach is because it will help to identify faults and failures. Figure 6-27 is the architecture that was designed for this mission. The architecture is an optimisation of both centralised and decentralised

SEEDS Executive Summary 09/2014 Page 132 architectures, where the interfaces of various systems are minimised by the use of a centralised architecture and the decentralised architecture allows us to find faults and isolate or recover depending on the type of failure. The decentralised architecture also allows parallel computation and depicts the growth path for selective technology upgrade. The distributed computing systems (DCS) acts as the top level and the data flows down to each subsystem of the system making it a decentralised approach. The computing devices consist of the TM/TC which is the telemetry and telecommand is interfaced with the data processing and handling systems.

6.8.3.1 Telemetry downlink The information from all the systems is collected, processed and then modulated; this information is later transmitted to the ground station via the communication system. This is all the data regarding the condition of subsystems and allows for the determination of the performance of the subsystem at instruments level.

6.8.3.2 The Command uplink If there are any faults or errors in the data received then the ground station would transmit the necessary instructions and procedures to rectify the problem.

6.8.3.3 The Ranging transponder This allows for simple tracking of position and the current trajectory or orbit. If any errors are found instructions for corrective manoeuvres would be transmitted.

6.8.4 Space Protocols The instructions, information and commands should all follow the current space protocols defined by the international Communication standards. The space communication protocol standards (SCPS) have recommended that the systems should conform to the Consultative Committee for Space Data Systems (CSDS). The Interplanetary Internet (IPN) should conform to the CCSDS file delivery protocol (CFDP) to minimise the error. The system should be able to perform autonomously and this helps us to minimise the latency for the communication. It is critical to enable the use of Delay Tolerant Network (DTN).

6.8.5 On Board Data Handling (OBDH) The ESA standard for OBDH is in the highest form of engineering standards and already the use of military standard BUS 1553 B have been proven with the Ariane rockets. The OBDH would probably need this type of bus to perform its tasks. The future bus system itself would act as a compression and storage devices.

6.8.5.1 Processor The processor would be a fast 32-bit space qualified device for long duration deep space flight. It should be able to provide functions such as decoding the information, operating as a central terminal unit for the bus and should be able to distribute the correct signals to the intended receivers. Note that the receivers here are the sensors and actuators. The processors could further be classified on their functions for specific tasks such as; Payload Processor, Distributor processors, Communication processors and other sub functions.

6.8.5.2 Storage The storage is based on the data that is being generated and to deduce that it is necessary to know what type of data is being collected and processed for information. The telemetry data is classified into three main categories:

 Housekeeping data

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 Attitude control data  Payload data

These different data can be sent to the ground segment for further analysis. It is common to pre-process data to send only useful data to the ground segment due to the limited data transfer available.

6.8.6 Fault Tolerant System Consideration The mission requires reliable performance of all the systems without any need of logistical support. The long duration mission requires particular expertise in these areas to provide safety, reliability and autonomy. This includes a new technology plan to guide the development of the next generation fault tolerant computing technology. In this mission the fault tolerance and failure shall be nullified with the help of redundancy.

The redundancies are done by the following techniques.  Detection and isolation of a failure.  Performing hardware reconfiguration.

The faults are usually detected with the help of self-monitoring systems, which have to provide all the vital data of housekeeping. The faults are broadly classified in the following section and this helps to analyse the type of faults.

6.8.6.1 Classification of Faults: For any type of fault tolerance it is critical to identify and foresee the type of faults which would help us to solve them when they occur using a specified set of procedures that have been already predefined.  Permanent faults  Transient (temporary) faults  Intermittent (not continuous) faults  Timing faults  Latent (hidden) faults  Worst-case fault scenarios with a lower probability of occurrence

It is noted that the manual fault should never occur and the system should be designed to withstand this error. All the faults should be classified at component level, this would help to send the right procedures on board when the fault occurs.

6.8.7 Conclusion The mission would need very low operating costs for the entire life cycle for the components with low power and mass. The mission plan for a long-duration manned flight must concentrate on an extended view of operations and the development of detailed documentation, which leads to a vast amount of data to be stored and analysed. The performance of the primary flight control systems requirements could be such as designed and utilised in existing systems for deep flight missions especially Mars based exploration such as:

 Mission reliability level of 0.97 success probability at 10 years with planned maintenance.  The processor throughput would be 100 million instructions per second (MIPS).

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6.9 ATTITUDE AND ORBIT CONTROL SYSTEM

6.9.1 Introduction For some spacecraft in the Orpheus mission, AOCS is a critical subsystem. Particularly, the PhL needs a strong and fast AOCS to ensure its stability on the ground with very low gravity.

6.9.2 Spacecraft Dynamics

6.9.2.1 Phobos Lander (PhL) The Phobos Lander is the greatest concern for this mission due to stability While estimates indicate it can’t unexpectedly fly away from Phobos, rotations could be a problem causing it to tumble. 6.9.2.1.1 Moment of Inertia The PhL has a cylindrical shape, though the internal mass distribution is intentionally non-uniform to lower the centre of mass. The driving assumption will be that the lowest 1/3 of the volume will contain 1/2 of the mass, i e 7 5 t contained in the lowest 1 5 m of the PhL This is consistent with the PhL’s design

Figure 6-28: Size of the PhL for subsystems and living space The centre of mass can then be estimated at 1.5 m from the bottom and 3 m from the top, as illustrated in Figure 6-28. The total moment of inertia has been estimated as 21000 kg·m2. The vehicle’s X axis has been defined over its length, the Y axis is vertical and the Z axis is horizontal following the right hand rule. 6.9.2.1.2 Torques If astronaut’s handles for locomotion are located at 1 m from the floor, which is a reasonable choice, tor ues exerted by astronaut locomotion can be estimated to be applied at this distance. Estimating the strength of human arms as 700 N, based on the assumption that a person may walk on his hands but not jump under these conditions, this is trivially a torque of 700 N·m. The maximum expected angular acceleration is 0.21 rad/s2 = 12º/s2. The PhL will need to counteract torques around the X and Z axes. Rotations about the Y axis can be ignored in station mode as this axis is perpendicular to the ground. 6.9.2.1.3 Sizing of actuators Sizing is based on existing gyroscopes from bibliography [EADS Astrium – 15-45S] [Takeuchi, H et al. – 2011], from which it’s clear that scaling up linearly from the tor ue is an overestimation of mass and power. A torque of 700 N·m is 16 times larger than the reference, so 16 times more power and mass translates into 400 W and 188 kg, respectively. From figure 1 it may seem intuitive to assume that the worst torques will be those applied around the X axis. This is false as the landing legs will be deployed at a 4m horizontal span. Assuming 22 N thrusters of reference [SSC Group - ECAPS] are used then 11 engines will have to fire at the same time and in the same direction. As a worst case estimate, if multiple 22 N engines are used and their mass is 1 kg each, then 44 engines are needed to counteract torques in both senses around the X and Z axes. A good design would use fewer and stronger engines if this can save mass, though 44 kg is really not meaningful for the PhL’s mass budget. 6.9.2.1.4 Desaturation A meaningful benefit of staying in the ground is that CMG desaturation does not require thrusters. While astronauts are in EVA or sleeping the system may slow down the spinning wheels slightly. This will apply a

SEEDS Executive Summary 09/2014 Page 135 torque on the PhL, but if the torque is controlled and not large enough to lift any landing legs from the ground then it will only create a non-uniform weight distribution The spacecraft’s weight and non-uniform reaction forces from the ground will be able to counteract this torque, effectively desaturating the CMGs without any propellant expenditure. 16 h per day are expected to be available for desaturation Gravity is almost negligible, but with the PhL’s large mass, a weight of 86 N is not. Therefore desaturation is not a critical issue.

6.9.2.2 MATV In early phases of its mission, up to docking with the CIV, the MATV will be designed to be able to use the PhL’s actuators Since the PhL has been designed for perturbations much larger than what the MATV+PhL+Lab+ML complex is expected to withstand, this is retained non-critical. The MATV service module will detach after delivering the Lab and PhL to the CIV, and at this point the ML will have already detached as well. Since the service module itself after detachment is not a critical element for the Orpheus mission, and since it is not expected to require better pointing accuracy or withstand larger perturbations than already existing Mars orbiters, the task of sizing its actuators can be neglected to concentrate on the CIV.

6.9.2.3 CIV This is a critical vehicle for AOCS. It will withstand meaningful perturbations due to the presence of astronauts inside and the docking of the MATV first and the PhL later. Its sun-inertial attitude has to be maintained with a certain pointing accuracy to prevent cryogenic propellant from boiling. 6.9.2.3.1 Disturbance environment in Mars orbit Mars-GRAM has been used successfully during the aerobraking operations of Mars Global Surveyor, Mars Odyssey and Mars Reconnaissance Orbiter [Justh, H.L. – 2010] It’s based on the University of ichigan Mars Thermospheric General Circulation Model (MTGCO) [Justh, H.L. – 2010] [Bougher, S. W.]. Based on calculations from the TGCO the density of ars’ atmosphere at 240 km can be estimated as 3 27 x 10-12 kg/m3. At 300 km is has been overestimated as equal to those at 240 km due to the lack of more precise information. The CIV will stay 7 days in this orbit with a wet mass of 1342 tons before moving into Phobos’ orbit From the mission analysis results, the capture orbit will have a periapsis of 300 km at 4 km/s. Its external dimensions, have an external surface of 3700 m2 (only one side of the cylinder) when moving sideways on the Martian atmosphere. Solar arrays are 500 m2, leading to a total of 4200 m2. By assuming the distance from the centre of pressure to the centre of mass equal to the length of the spacecraft, the resulting torque is manageable, but the accumulated angular momentum over an orbit is not. However the approximately cylindrical shape offers an advantage: by changing slightly the roll attitude a different face of the spacecraft can be made normal to the flow.

Figure 6-29: Variations of the 'yaw' angle change the surface area normal to the flow Using the above technique and making solar arrays parallel to the flow, the centre of pressure may be moved towards the front of the spacecraft. Just a very small angle will suffice, and it wouldn’t violate the pointing constraints required for thermal control, which have specified a maximum of 11 degrees. This way the position of the centre of pressure can be regulated and saturation of CMGs can be prevented.

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6.9.2.3.2 Phobos’ proximity After performing a burn at apoapsis to circularize the orbit, the CIV will have a wet mass of 594 tons. Atmospheric density will decrease exponentially. It is meaningful for Phobos over 50 million years, but during the duration of this mission it will be negligible for the CIV. The magnitude of perturbations in this environment will then be assumed to be equal to the interplanetary trajectory, where radiation pressure is the dominant force. 6.9.2.3.3 Interplanetary trajectory In the worst case, during the return trajectory, the wet mass becomes almost equal to the dry mass of 300 tons. Solar radiation pressure is the dominant force. The interplanetary medium has a very low density whose pressure is 3 orders of magnitude smaller. Calculating radiation pressure as W/cR2, where W is the solar constant on Earth (1360 W/m2), c is the speed of light in vacuum and R is the distance to the Sun in AU (treated as dimensionless), it results to vary between 1.6 μPa at ars’ aphelion and 12 2 μPa at the CIV’s perihelion. On solar arrays the expected forces range from 0.65 mN to 4.9 mN. If their distance to the centre of mass is estimated at 3 m, based on the design of the spacecraft, and a maximum angle of 10º as required for thermal control, the torque varies between 0.34 mN m and 2.6 mN m. If such torques are kept constant for a month, the accumulated angular momentum varies between 88 N·m·s and 6600 N·m·s depending on the distance to the Sun. (This is an overestimation assuming 100% reflection. Actual numbers will be smaller due to the absorption of solar arrays). These numbers have a double meaning. They imply that perturbations are meaningful and capable of saturating CMGs, though at the same time they imply that external forces are usable for desaturation. For the above reasons the attitude control system must be able to point with accuracy of 1º and the spacecraft must be designed to tolerate misalignments of up to 10º on each side of the Sun pointing line. Radiators can be used to perform a similar desaturation for the remaining orthogonal axis. 6.9.2.3.4 Moment of Inertia and sizing of actuators The CIV’s moment of inertia is 516 x 106 kg·m2. CMGs can be sized equal to the International Space Station (240 Nm). Even if the mass and the moment of inertia of the CIV are higher, rotations aren’t as fast in the most commonly used control modes. However during manoeuvres the solar-inertial attitude will be temporarily broken. The thermal control results indicate that temperature of propellant increases by 1 ºK every 30 minutes if the appropriate attitude is not maintained. Rotations will be designed for a maximum temperature rise of 10 ºK. This means the whole manoeuvre can take at most 300 minutes, or 5 h. From the mission analysis team, the longest burn will take 2.5 h. This leaves only 2.5 h for preparation and realignment. If 1 h is left as safety margin, the CIV must be able to rotate 90º in 45 min. This translates into 0.033º/s. To reach the mentioned angular velocity in 10 s, the required angular acceleration would be 0.0033 º/s2 = 0.06 mrad/s2. At a distance of 17.5 m from the centre of mass, thrusters would need a force of 11 kN. Using several engines based on the Space Shuttle Orbiter’s thrusters of 3 9 kN, with a force of 11 kN, a burn time of 10 s and an Isp of 290 s [Braeunig, R.,2006.b], propellant required is 3.9 kg/s or 39 kg per 10 s burn. Each manoeuvre will take 4 burns, or 160 kg. Using data from the mission analysis results, at least 8 manoeuvres will be performed, so the total propellant for rotating would be 1300 kg, or 2600 kg if a safety factor of 2 is applied.

6.9.3 Minor manoeuvres and station keeping of the CIV The nuclear reactor cannot be made critical for manoeuvres of a few m/s, which may be needed for station keeping, phasing or mid-course corrections. The total ΔV of non-nuclear manoeuvres is 1.35 km/s, though this number is meaningless due to staging. More precise calculations show a required propellant mass of 184 tons. As methane is already available in the CIV, such manoeuvres will be performed using CH4/O2 chemical engines. Liquid oxygen is similar to liquid methane in cryogenic properties. The Russian RD-0234- CH engine is the reference design.

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6.9.4 Summary on AOCS Even though AOCS is a critical subsystem of some of the elements in this mission, in all cases there are beneficial conditions that can be used to simplify the design. The critical elements are the Phobos Lander (PhL) due to its stability on very low gravity, and the CIV due to its thermal control requirements for cryogenic propellant. Given the conditions studied here, it’s possible to assure that designing attitude control systems for these spacecraft is feasible with current technology.

6.10 CRITICAL ISSUES During the study of the CIV design, critical design elements have been identified, which can also be referred as sensitive point in terms of TRL, or need further study. 1. The design of the tanks is one sensitive point because for the moment, it does not exist composite tanks able to resist at cryogenic temperatures. Nevertheless it is strongly believe that in a not so far future, such tanks will be available for space use 2. The Radiation issue. Analysis has been performed using GRAS (Geant Radiation Analysis for Space) in order to optimize the shielding for the astronauts, however the calculations made had to be taken with a lot of precautions as it would have been too long to make a model of the whole system. 3. On orbit assembly timeline. The precise timeline have not been entirely described. Indeed for a complete analysis it should be certain that each module sent in orbit should have the needed power/fuel to “survive” These calculations have not been included in the scope of this project as the customer constraints were stating that the CIV could be considered fuelled and assembled in LEO. 4. Boil off issue. This issue has been considered very carefully but as it is a very versatile issue depending on the pressure, temperature, a special attention should be dedicated to this problem if any key figure is modified.

The objective of identifying these issues is to be aware that it will need further attention in case someone wants to re-use the data of this project. At the end, the spacecraft design, on a system engineering point of view, has been finished within the deadline and is functional. A Concurrent Design Facility session has been performed at the end of the project in order to ensure the reliability of the key figures.

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6.11 REFERENCES SIPPEL, M., KOPP, A., MATTSSON, D., FRATERS, A. , KOUSSIOS, S., Advanced Cryo-Tanks Research in CHATT, 5th European Conference for Aeronautics and Space Sciences (EUCASS) 2013. MITAL, S. K. et al.: Review of Current State of the Art and Key Design Issues with Potential Solutions for Liquid Hydrogen Cryogenic Storage Tank Structures for Aircraft Applications, NASA/TM—2006-214346. (2006). FRATERS, A., Current state of the art knowledge in the structural aspects of advanced composite cryogenic tanks, CHATT Deliverable D3.1.1, 2012. MATTSSON, D., Report on specification, materials selection and theoretical design, CHATT Deliverable D3.2.1, 2012. UdelPolysulfone Design Guide, SOLVAY Advanced Polymers, 2014. TORAYCA, n.d., T700S DATA SHEET NASA – TransHab concept – International Space Station History http://web.archive.org/web/20070601021217/http://spaceflight.nasa.gov/history/station/transhab/ [Accessed June 05th, 2014] BARRATT, M.R., and POOL, S.L., 2008. Principles of Clinical Medicine for Space Flight. USA: Springer ESA, 2008. Environmental Control and Life Support Systems – ESA-HSO-COU-030 2.0. [pdf] ESA. Available at: European Cooperation for Space Standardization, 2008. ECSS-E-ST-10-11C Space engineering Human factors engineering. Noordwijk: ESA Requirements and Standards Division. European Cooperation for Space Standardization, 2008. ECSS-E-ST-34C Space engineering Environmental Control and Life Support. Noordwijk: ESA Requirements and Standards Division. HANFORD. A. J, 2005. Advanced Life Support Research and Technology Development Metric – Fiscal Year 2005. Available at: http://ston.jsc.nasa.gov/collections/trs/_techrep/CR-2006-213694.pdf [Accessed 30/06/14] LARSON, W.J. and PRANKE, L. K., (1999) Human Spaceflight Mission Analysis and Design. McGraw-Hill NASA, 1999. Spacecraft Maximum Allowable Concentrations for Airborne Contaminants. [pdf], Houston, Texas: Lyndon B. Johnson Space Center . Available at [Accessed 3 July 2014] Stack Exchange, 2013. Air Temperature and Humidity inside the ISS [Online]. Available at < http://space.stackexchange.com/questions/2539/air-temperature-and-humidity-inside-the-iss> [Accessed 3 July 2014]f Ed. Paul O. Wieland 1994. Designing For Human Presence in Space: An Introduction to Environmental Control and Life Support Systems (RP-1324). NASA MSFC. Marshall Space Flight Center, Alabama CUCINOTTA, F. et al. (2013.a) How Safe Is Safe Enough? Radiation Risk for a Human Mission to Mars. [Online] http://www.ncbi.nlm.nih.gov/pmc/articles/PMC3797711/ [Accessed 16th July 2014].

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G. SANTIN, V. IVANCHENKO, H. EVANS, P. NIEMINEN, E. DALY (2005) GRAS: A general-purpose 3-D modular simulation tool for space environment effects analysis, IEEE Trans. Nucl. Sci. 52, Issue 6, 2005, pp 2294 - 2299, http://dx.doi.org/10.1109/TNS.2005.860749 [Accessed 17th July 2014]

S. AGOSTINELLI et al. (2003) Geant4 - A Simulation Toolkit, Nuclear Instruments and Methods A 506 (2003) 250-303, http://dx.doi.org/10.1016/S0168-9002(03)01368-8 [Accessed 17th July 2014]

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CENTER FOR SCIENCE AND ENGINEERING PARTNERSHIPS [CSEP] Radiation Laws (n.d) [Online] http://csep10.phys.utk.edu/astr162/lect/light/radiation.html (Accessed 11th June) LARSON, W. and WERTZ, J. (2005) Space Mission Analysis and Design. 3rd Ed. El Segundo, CA: Microcosm Press

BOEING (2006) Active Thermal Control System (ATCS) Overview, [Online] www.boeingtravel.com (Accessed 9th June 2014) HANFORD, A. and EWERT, M. (1996) Advanced Active Thermal Control Systems Architecture Study, [Online] http://ston.jsc.nasa.gov/collections/trs/_techrep/TM104822.pdf (Accessed 14th June 2014) NIST (2011) Methane [Online] http://webbook.nist.gov/cgi/cbook.cgi?ID=C74828&Mask=4#Thermo-Phase (Accessed 16th June 2014)

CHEMTEAM (n.d) The Clausius-Clapeyron Equation [Online] http://www.chemteam.info/GasLaw/Clasius- Clapeyron-Equation.html (Accessed 16th June 2014)

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SOLAR MIRROR (n.d) Table of absorptivity and emissivity of common materials and coatings, [Online] http://www.solarmirror.com/fom/fom-serve/cache/43.html

PTABLE (2014) Specific Heat Capacity Table, [Online] www.ptable.com (Accessed 30/06/14) Braeunig, Robert A. (1998 - 2006) Spacecraft & Launcher Specifications: Space Shuttle Orbiter [Online] Available from: http://www.braeunig.us/space/specs/orbiter.htm [Accessed June 13th, 2014] NASA (2002) - Space Shuttle Reaction Control System [Online] Available from: http://spaceflight.nasa.gov/shuttle/reference/shutref/orbiter/rcs/overview.html [Accessed June 13th, 2014] NASA History Office (2004) The SkyLab Program [Online] Available from: http://history.nasa.gov/apollo/skylab.html [Accessed June 13th, 2014] BRAEUNIG, Robert A. (2006.b) - Spacecraft & Launcher Specifications: SkyLab [Online] Available from: http://www.braeunig.us/space/specs.htm [Accessed June 13th, 2014] Boeing company (2006) Space Exploration: ISS Motion Control Subsystem [Online] Available from: http://www.boeing.com/assets/pdf/defense- space/space/spacestation/systems/docs/ISS%20Motion%20Control%20System.pdf [Accessed June 16th, 2014] BURT, R. ; Loffi, R. (2003) Failure Analysis of International Space Station Control Moment Gyro [Online] Available from: http://adsabs.harvard.edu/abs/2003ESASP.524...13B [Accessed June 16th, 2014] EADS Astrium - Control Momentum Gyroscope CMG 15-45S [Online] Available from: http://www.astrium.eads.net/en/equipment/cmg-15-45s.html [Accessed June 18th, 2014]

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HIROSHI TAKEUCHI et al (2011) Development of the Anti-Rolling Gyro 375T (Rolling Stabilizer for Yachts) Using Space Control Technology [Online] Available from: https://www.mhi.co.jp/technology/review/pdf/e484/e484070.pdf [Accessed June 18th, 2014] J. COOK et al (2011) ISS Interface Mechanisms and their Heritage [Online] Available from: http://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20110010964_2011009594.pdf [Accessed June 25th, 2014] SSC Group - ECAPS CAPABILITIES [Online] Available from: http://sscspace.com/capabilities-3 [Accessed June 25th, 2014] Justh, H.L. (NASA - 2010) Mars Global Reference Atmospheric Model 2010 [Online] Available from: http://ntrs.nasa.gov/search.jsp?R=20140003184 [Accessed June 27th, 2014] Bougher, S. W. - Comparative Terrestrial Planet Thermospheres - [Online] Available from: http://data.engin.umich.edu/tgcm_planets_archive/thermo.html [Accessed June 27th, 2014]. Cited by [Justh, H.L. - 2010]. Additional references that this entry is based on: Bougher, S. W., and R. G. Roble, Comparative terrestrial planet thermospheres: 1. Solar cycle Variation of Global Mean Temperatures, J. Geophys. Res., 96, 11045-11056, 1991. BOUGHER, S. W., and R. G. ROBLE, Thermosphere, pp. 819-825, in The Encyclopedia of Planetary Sciences, Eds. J. H. Shirley and R. W. Fairbridge, Chapman and Hall, London, 1997. BOUGHER, S. W., G. M. KEATING, R. W. ZUREK, J. M. MURPHY, R. M. HABERLE, J. HOLLINGSWORTH, and R. T. CLANCY, Mars Global Surveyor Aerobraking : Atmospheric Trends and Model Interpretation, Adv. in Space Research, 23, #11, 1887-1897, 1999. BOUGHER, S. W., S. ENGEL, R. G. ROBLE, and B. FOSTER, Comparative Terrestrial Planet Thermospheres : 2. Solar Cycle Variation of Global Structure and Winds at Equinox, J. Geophys. Res., 104, 16591-16611, 1999. BOUGHER, S. W., S. ENGEL, R. G. ROBLE, and B. FOSTER, Comparative Terrestrial Planet Thermospheres : 3. Solar Cycle Variation of Global Structure and Winds at Solstices, J. Geophys. Res., 105, 17,669-17,692, 2000. BOUGHER, S. W., S. ENGEL, D. P. HINSON, and J. M. FORBES, Mars Global Surveyor Radio Science Electron Density Profiles: Neutral Atmosphere Implications, Geophys. Res. Letters, Vol. 28, 16, 3091-3094, 2001. CLACK J.P.C. et al (NASA - 1971) NASA SPACE VEHICLE DESIGN CRITERIA (GUIDANCE AND CONTROL) - SPACECRAFT AERODYNAMIC TORQUES [Online] Available from: http://www.abbottaerospace.com/download/reference_data/nasa/space_vehicle_design_criteria/NASA%20- %20sp8058%20-%20Space%20Vehicle%20Design%20Criteria%20- %20Spacecraft%20Aerodynamic%20Torques.pdf [Accessed June 17th, 2014] BELLUSCIO, A. (NASASpaceflight - 2014) SpaceX advances drive for Mars rocket via Raptor power [Online] Available from: http://www.nasaspaceflight.com/2014/03/spacex-advances-drive-mars-rocket-raptor-power/ [Accessed June 30th, 2014] TOWNES et al, 2004, The Mars Laser Communication Demonstration, NASA KWOK, A., 2009, Rev. B Frequency and Channel Assignments, DSN Telecommunications Link Design Handbook 810-005 TAKASHI, J., 2012, Optical Inter-orbit Communication Experiment between OICETS and ARTEMIS, NICT BOROSON, M., D et al, No Date, Overview and Status of the Lunar Laser Communication Demonstration LACOSTE, F., 2010, ITU-R PROPAGATION MODELS SOFTWARE LIBRARY, CNES. < http://logiciels.cnes.fr/PROPA/en/logiciel.htm>

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TAYLOR, J., Lee, D.K., Shambayati S., 2006. Mars Reconnaissance Orbiter Telecommunications. NASA, DESCANSO design and performance summary series, article 12 HEMMATI, H., Chen C., 2005. Deep Space Optical Communications. JPL, Deep Space Communications and Navigation Series. IPPOLITO, J., L., 2008, Satellite Communications Systems Engineering, Wiley "HORIZONS Web-Interface." HORIZONS Web-Interface. Web. 04 June 2014. . FIELHAUER, B., K and Boone, G., B, 2012, Concurrent System Engineering and Risk Reduction for Dual- Band (RF/optical) Spacecraft Communications, NASA ITU REC 840-5, 838-4 and 676-9 International Telecommunication Union. "MATLAB." - The Language of Technical Computing. Web. 04 June 2014. . ITU REC 840-5, 838-4 and 676-9 International Telecommunication Union. Patel, Murkund R. (2005) Spacecraft Power Systems, ISBN 0-8493-2786-5 Federal Aviation Administration (FAA) - [Doc. No. 5066, 29 FR 18291, Dec. 24 1964, as amended by Amdt. 25–11, 32 FR 6912, May 5, 1967; Amdt. 25–23, 35 FR 5676, Apr. 8, 1970] [Online] Available from: http://www.gpo.gov/fdsys/pkg/CFR-2002-title14-vol1/pdf/CFR-2002-title14-vol1-sec25-869.pdf NASA Langley Research Center (2008) Technology gateway Vol.1 Issue 6 [Online] Available from: http://technologygateway.nasa.gov/docs/TOA_LARC06_RP46Insulation_08web.pdf Ochs, Robert I.; FAA (2006) Electric Wire Insulation Study: Flammable Properties and Testing Methods [Online] Available from: http://www.fire.tc.faa.gov/pdf/materials/OCHS-Wire_Insulation.pdf Boeing Company - International Space Station Electric Power System [Online] Available from: http://www.boeing.com/assets/pdf/defense- space/space/spacestation/systems/docs/ISS%20Electric%20Power%20System.pdf NASA History Office (2009 - 2011) Apollo 11 Flight Journal [Online] Available from: http://history.nasa.gov/ap11fj/21day6-tei.htm NASA (2005) Audit Report IG-05-023 - Space Shuttle Orbiter wiring inspection [Online] Available from: http://www.hq.nasa.gov/office/oig/hq/audits/reports/FY05/ig-05-023.pdf De Donder, E. / SpaceWeather.eu (2012) Space Radiation Effects p.15 [Online] Available from: http://www.spaceweather.eu/ro/repository/download?id=Tutorial_Rad_Effetcs_ED- 1332501225.&file=Tutorial_Rad_Effetcs_EDD.pps [Accessed April 18th 2014] Simon P. Philipps, Andreas W. Bett (2013) III-V MULTI-JUNCTION SOLAR CELLS: MARKET, TECHNOLOGICAL STATUS AND RESEARCH TRENDS. Fraunhofer Institute for Solar Energy Systems Wall, M. / NBCnews (2012) NASA’s huge new rocket may cost $500 million per launch [Online] Available from: http://www.nbcnews.com/id/49019843/ns/technology_and_science-space/#.U43a1LV4VhE [Accessed June 3rd, 2014] GRALLA, E.L. and DE WECK ,O.L. 2007. Strategies for on-orbit assembly of modular spacecraft. [Online] Department of Aeronautics and Astronautics, Massachusetts Institute of Technology. Available at: http://strategic.mit.edu/docs/2_20_JBIS_60_6_219_Tugs.pdf [Accessed 3 July 2014]. Benjamin, A. L, Lala, J. H. (AIAA 1997) Advanced Fault Tolerant Computing for Future Manned Space Missions. (IEEE). [Volume 2, 26-30 Oct. 1997 Page(s):8.5 - 26-8.5-32 vol.2].

P, Fortescue, G, Swineral and J, Stark. (2011) Space craft : Systems Engineering 4th edition, Ed. - Wiley

NASA (n.d) Computers in Spaceflight: The NASA Experience [http://history.nasa.gov/computers/contents.html].

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7 RELIABILITY, AVAILABILITY, MAINTANABILITY AND SAFETY (RAMS) RAMS engineering is a powerful tool used to support the management of “failure risks”

7.1 OBJECTIVES  To guarantee the compliance of the system functions and performances with the specification over the complete mission duration.  To guarantee the mission success within an expected margin.

7.2 RELIABILITY The reliability assessment is based on probability theory to provide the likelihood of mission success. Differing levels of system redundancy and diversity can then be applied to improve the mission’s chances When reliability analysis of the mission is performed, it is distributed among the various sub-systems of the functional chain. They are treated as independent, segregated and that there is no command failure between them. A reliability specification of has been chosen. The following methodology for computation reliability has been implemented: The credibility of the numerical data is of first importance. Generally reliability tests are limited for costs reason. But as it can be seen in the example, the reliability goals help to define a test plan. In this part the worst-case scenario of the whole mission is analysed. It has been chosen to use a mission time of 600 days, which is approximately 14 400 hours. Depending of the frequency of the maintenance, a reduced value for the calculations can be used Therefore: P1 = exp (-λ*T) The failure rate, is roughly 8000 Failure in Time (FIT) for a GNC computer, and one FIT is 1.10^-9s. Let’s take the example of the GNC As the configuration must be Fail Operational/Fail Safe (FO/FS), it is necessary to have at least 3 computers. For reliability reasons, let’s make the calculations when the 4 computers are working, then: After calculations: Probability of success of the mission (for GNC): 0,99997 With 3 computers (one failure): Probability of success of the mission (for GNC):0,97 With this total probability of the GNC equipment it is achieved a satisfying probability of success. It is retained that a level of redundancy between 3 to 4 systems is necessary. A predictive reliability analysis has been performed because it is based on failure rate λ (they are not measured data but estimated). This is how it has been decided the reliability level of all the equipment. Fault masking will be used to determine the outputs for GNC.

7.3 AVAILABILITY The needed probability of a device of being available at a time is related to the level of redundancy that should be applied for this device. However redundancy should not be limited to simple duplication of systems. The following questions have been answered during the analysis: What implementation? Define the physical layout for the connections (buses). Which failure mode? Structure of the redundancy based on the failure modes.

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Which failure propagation? Protect the elements upstream and downstream. Which common mode? Assure the independency of the failures. What type of redundancy? Define the operational process.

7.4 MAINTAINABILITY A dedicated time per day for each astronaut will be used for maintenance purposes. The CIV design has been thought in order to answer to the maintenance questions. For example the tanks have been implemented in a way that the astronauts could do EVA to evaluate eventual damages.

7.5 SAFETY

7.5.1 Fail Operational / Fail Safe Fault Tolerance is the ability of the system to withstand an unwanted (upset or hazard) event and maintain a safe condition. There are several actions to put in place in order to respect this requirement:  It is necessary to segregate the resources to avoid failure propagation not only between functional chains but also between nominal, redundant, safe mode resources. The avoidance of single point failure is essential.  This requirement implies big consequences for the power supply architecture. Indeed it must be designed in a way that permits all the equipment to be operational in case of failure. It implies a separate line of power for all the equipment linked to the safe mode.

7.5.2 Safe mode The possible causes for safe mode are:  Double failures i.e. unexpected environmental conditions (radiation, solar activity).  Combination of a single failure and an operational failure.  Design error, none intentional. The Safe Mode is a very particular configuration, which once activated, should prevent any other failures by activating only the very basic functions of the CIV (thermal, power). The Management System Unit (MSU) receives the triggering logic in order to know when to activate the safe mode. The consequences of the safe mode should be different depending on the current mission phase the CIV is in. If only the GNC part is considered, the safe mode will be triggered if two computers of the GNC are in failure. The implication in terms of architecture design is that it is needed to insert this safe mode to the system. So to take into account:  What sensors and actuators should be connected to perform safe mode in all the phases.  The criteria that will trigger the safe mode.  Make sure that the safe mode unit is able to compute this criterion. So it should be connected to the communication processor.

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7.6 REFERENCES Ley, W. et al. (2009), Handbook of Space Technology, New Jersey: John Wiley & Sons, Ltd. Wertz, James R., Larson Wiley J. (2010), Space Mission Analysis and Design III, Space Technology Library. Gonnaud JL., Pascal V. (1999), ATV Guidance Navigation and Control, Proceedings 4th ESA International Conference on Spacecraft Guidance, Navigation and Control Systems, ESTEC, Noordwijk, The Netherlands.

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8 COST CONSIDERATION The cost of the mission was of serious consideration in the feasibility of the mission. Using historical data from the entirety of NASA’s missions, it is known that the cost of operation for missions is 20 times more than that of the development cost. It is important to undertake a cost analysis before the mission can be proposed, to determine the feasibility of the mission. What follows is a basic estimation of cost, but more detail needs to be provided before the mission can be proposed.

8.1 INTRODUCTION The Orpheus mission is the first of many missions to pave the way for future Mars exploration missions. This brings serious consideration for evaluating and comparing the total cost for the design, development, testing and validating of the system requirements.

8.2 COST ANALYSIS There are large uncertainties concerning the development cost of new technologies necessary for this kind of mission and this makes it very difficult to provide a cost estimate. The most practiced way of determining the cost of a mission and its feasibility is by learning from the past missions. In the following sections several approaches will be considered. Please note that all cost and prices are dated for August 2014.  Mass Cost Ratio  Development and Operation Cost Ratio  Launch Cost Ratio

8.2.1 Mass Cost (M-C) Ratio The M-C ratio is a reliable source of determining the cost of past missions. The ISS project is considered as a reference because of the duration of the orbital assembly (5 years). The cost of the ISS project is estimated to be $150 billion (NASA,n.d). For example the ISS required international cooperation and the use of 37 Shuttle launches in order to be assembled (NASA,n.d).The Space Shuttle orbiter had a dry mass of 80 tons with a payload capacity of 20 tons (NASA,n.d); which means there was a total mass in LEO between 2960 and 3700 tons. This mass does not consider the Russian contribution which consisted of 97 launches, 4 Japanese HTVs and 5 European ATVs. It could be concluded here the feasibility of our mission in terms of initial mass at LEO (4000 tonnes) and its cost ratio of assembly/maintenance is almost similar to the one of the ISS but with the overhead cost of current inflation.

8.2.2 Development and Operational Cost Ratio The total ISS mass in LEO is approximated 3700 tonnes; this required around $150 billion for; developing technology, launches, assembly, operations and maintenance. From this number, around $20 billion, not including the Russian and European contributions was spent on the launch costs. The remaining funds were distributed in the period of the next 20 years and mainly used for development. The Apollo mission would have cost around $170 billion if it had happened in 2005; this includes the cost of developing all of the technology needed to land humans on the Moon (Congressional Budget Office report, 2004).The shuttle program had a cost of $209 billion for the entire project life cycle which ended in 2010. It was operated for a total of 135 flights, yielding an average cost per launch of more than $1.5 billion. These facts show that it is reasonable to assume the development cost for Orpheus would not be greater than $200 billion; but compared to the other missions it could be acceptable. The MATV could be a cost factor in terms of development and validation for all the systems and science payloads. The setting up of ground stations in order to facilitate operations and provide infrastructure is a huge cost but can be re-used in future SEEDS missions.

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8.2.2.1 Cost of Developing Nuclear Engine The cost of development for the entire mission was previously assumed to be $200 billion. It is worth wondering if it is a reasonable cost. It is assumed that almost all the development and design costs for a reasonable TRL have already been invested such as launch system (SLS), communication system (LADEE), ECLSS (ISS) and inflatable modules (Bigelow and TransHAB). The cost of developing a nuclear engine has some background now, as Russia has already invested $600 million into this kind of program (Seth D. Baum, 2009) which is the basis for the development cost of nuclear engine technology.

8.2.3 Launch Cost Ratio: The launch cost is estimated by adding up the cost of all the necessary launches in order to put the spacecraft of 4000 tonnes into orbit. For this calculation, the SLS launcher of NASA with a payload of 130 tons to LEO is used. This launcher has an estimated cost of $500 million per launch. This allows the comparison of the launches required for the mission from the table below considering that there are no constraints in terms of volume or design for the payload.

Table 8-1: Launches without any constraint using SLS 130 Estimated mass in EARTH Departure Launchers needed LEO (propulsion type) (SLS 130) (tonnes)

Chemical 4000 31

It is considered that the orbital assembly will last 5 years and the cost of inflation was considered to be 5% for Europe and 10 % for the USA. The budget has to be pre allocated by no later than 2029 and from this date inflation overrun is considered. The inflation cost was considered to be around 28 billion Euro for European launchers and 35 billion Euro for American ones. The difference is then 7 billion Euros. It would be extremely interesting if Europe could develop its own heavy launchers for this mission as the inflation for the next year 20 years in Europe is found to be stable and lower than that of U.S.A. The number of launchers needed is not optimised because constraints such as the payload fairing or volume were not considered. The SLS development program cost is projected to be 5.5 billion Euro but NASA have reused the engines of the shuttle. The 7 billion Euros obtained above indicates that there may be enough margin for the European Space Agency to design their own heavy launcher. The launch cost comparison for our mission is shown in Table 8-2. Launchers chosen are optimised for this mission. This allows us to go into a detailed analysis of our cost.

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Table 8-2: Entire CIV and MATV launch cost.

Cost Estimation for Orpheus mission using chemical propulsion for Earth departure

launchers cost per Total cost Total cost per launch Launcher needed launch per launch Time line to LEO cost in used before cost in $ cost in $ € (million) departure (million) (million)

Launch of supports for Assembly (space tugs/ robotic arms/ SLS 70 2 500* 1000 770 deployable structures)

Launch SLS 130 22 500* 11500 8855 (cryo-fuels) Falcon Launch of MATV 3 85 255 196.4 heavy Launch of Node 1 Falcon 9 1 62 62 47.74 with radiators and solar panels.

Launch of Nuclear engines Falcon 9 1 62 62 47.74

Launch of Node 2 Falcon 2 85 170 130.9 and HAB 1 and 2 heavy

Resources Ariane 5 1 135 135 104

Falcon 20 85 1700 1309 heavy Launch of methane SLS 130 2 500 1000 770

Launch of crew Soyuz 2 80 160 123.2

Total launch cost (Nearest 100 Million) 56 $ 16100 € 12400

*Please note that the price of SLS is 500 million dollars per launch. The cost of all the launches is predicted to be 16 billion Euros (estimation in August 2014). Please note that the mission is not cost optimised in terms of launchers choice. Indeed, it has not been considered to use the cheapest launcher for each subsystem, but preferably the most reliable and available on the market for feasibility.

8.3 COST BENEFIT ANALYSIS: The Cost Benefit Analysis (CBA) is a process which can aid in decision making. The paradigm of the CBA works on the simple “if” function as shown below (Seth D. Baum,2009). If the benefits are larger than the costs then the decision is acceptable. If the costs are larger than the benefits then the decision is not acceptable.

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To determine the benefits, they have to be classified in two forms, one is non valuable market return and the other is valuable market return. For this mission both were analysed as follows.

8.3.1 Non Valuable Market Return (NVMR) A non-valuable market return is a gain that cannot be expressed in terms of financial indicators.

8.3.1.1 Long term Astrological threats Terrestrial life will never survive in the long term future due to following facts: Our star the Sun as any other star will change its composition and vary its radiative output with respect to time. The rate at which the Sun is converting hydrogen into helium generating radiative heat is increasing; this increase in heat load is making the Earth warmer. It is known that in around one billion years the Earth will be totally uninhabitable. If the Earth is to remain habitable its inhabitants must adapt and develop new technology to survive in this type of environment (Ward PD, Brownlee D). In 7 billion years the Sun will destroy the Earth as described by Caldeira K, Kasting J (1992) this drives the need to develop human space exploration and the future colonisation of different planets as the only solution to the survival of the human race and other species.

8.3.1.2 Immediate threats The immediate danger to life comes from threats such as rapid climate changes, disruptive technologies, nuclear warfare, pandemics etc. These threats would mostly affect the life only on Earth, leading to the solution of space colonisation where life could still persist. (Yudkowsky E, 2008).

8.3.1.3 The cost for the loss of human life It is a major drawback of losing life and is incomparable to the cost associated with it. The idea is found to be wrong when loss of human life occurs even in the pursuit for saving a larger amount of life. The CBA would reject any of these ideas pertaining to the loss of human lives.

8.3.2 Valuable Market Return (VMR) A valuable market return is a gain that can be expressed in terms of financial indicators.

8.3.2.1 Substantial wealth redistribution from rich to poor This mission would help future missions to land on the surface of Mars. This mission would directly infuse resources to evolve advanced technology which would give many industrial opportunities for their research and development. The products directly or by spin off would help to create new market shares and would directly help in creating high wage jobs and expanding business. To provide some examples: the closed loop water cycle technology of ISS being used in Africa where they have water scarcity thus providing and improving better quality of life. The manned moon landing alone employed 400,000 people and financed more than 20,000 industrial firms and universities (Allen, Bob).

8.3.2.2 Discount Rate: The longer the term of the project the more important the discounting factor becomes. Since in space exploration the decision taking terms are so long and crucial the discounting is even more important. The discount rate can be expressed as the rate at which both the costs and its benefit loses value over time (Schelling TC). To find the relation between uncertainties, discount rate and benefits we use the CAPM model.

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Let us assume that,

Size of benefit as B Number of years as N Discount rate as D

Then the Capital Asset Pricing Model (CAPM) would be

B*exp(-D*N)

The lower the discount rate, the more important future costs and benefits are. The discount rate is made to match the market interest rates depending on whether one uses the interest rates from bonds, stocks, or other financial instruments. The discount rate decreases as the uncertainty decreases.

8.4 CONCLUSION While evaluating both the VMR and NVMR it clearly shows and explains that space exploration with no loss of life is the ultimate solution. United States Court of Appeals judge Richard Posner gave human survival a value of $600 trillion which he himself has said to be underestimated value (Posner R 2004). This gives out the CBA importance of space exploration for the main factor of survival of human race. The total cost is estimated to be around 16 billion Euros for launches and is predicted that it would need around 150 billion Euros for developing various systems. The cost analysis shows that although the price of the mission is high it is minimal in comparison to the valuation of the human race and as a starting point for future further manned exploration and colonisation. This mission therefore is beneficial not only financially and scientifically but also in terms of the survival of terrestrial life.

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8.5 REFERENCES

NASA,n.d.(http://www.nasa.gov/centers/kennedy/about/index.html) Space Shuttle Orbiter NASA,n.d. ( http://www.braeunig.us/space/specs/orbiter.htm) A Budgetary Analysis of NASA’s New Vision for Space, September 2004, Congressional Budget Office report. Seth D. Baum, 2009 Cost-Benefit Analysis of Space Exploration: Some Ethical Considerations. Allen, Bob (ed.). "NASA Langley Research Center's Contributions to the Apollo Program". Langley Research Center. NASA. Retrieved August 1, 2013. Ward PD, Brownlee D. The life and death of planet Earth. New York: Henry Holt;2002. Caldeira K, Kasting J. The life span of the biosphere revisited. Nature 1992;360:721–3. Yudkowsky E. Artificial Intelligence as a Positive and Negative Factor in Global Risk In Bostrom N, Ćirković, eds. Global Catastrophic Risks. Oxford: Oxford University Press;2008:308-345. Posner R. Catastrophe: risk and response. New York: Oxford University Press;2004. Schelling TC. Intergenerational and international discounting. Risk Analysis 2000;20(6):833–7. Newell R, Pizer W. Discounting the distant future: how much do uncertain rates increase valuations? Journal of Environmental Economics and Management 2003;46:52–71. Federal Aviation Administration ,Second Half of 2009,Semi- Annual Launch Report Commercial Access to Space from Cecil Field, Florida. SpaceX, Press Conference. (www.spacex.com. Retrieved 2011-04-16). NASA, 2008, NASA Cost Estimating Handbook .

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9 CONCLUSION

A manned mission to the Martian moons would require a large international effort; nevertheless, most of the critical aspects have been proven feasible by the current study. The radiation issue and boil off problem have not been definitely solved but they can be mitigated to a meaningful extent. An in-orbit assembly strategy has also been proposed, and an overall cost comparable to already existing projects.

In the Table 9-1 it is possible to find all the key figures of the mission.

Table 9-1: Key figures of the mission Mission Orpheus Summary Specifications CIV Launch date May 2036 MATV Launch date 2034 Overall duration of the CIV mission 602 days Mars vicinity duration of the CIV 35 days Total crew size 6 Landing crew size 2 CIV trajectory profile Venus fly by for approach of Mars Lambert transfer for return Overall ΔV required for the CIV 12.5km/s2 MATV trajectory profile Continuous low thrust trajectory CIV departure mass 4000t MATV departure mass 120t Phobos landing mission duration 2 landings of 3 days Payload deployed on mars Network of 3 rovers Sample returned from Phobos 2*30kg

In terms of teamwork, it has been an enriching experience for all the members working for the project. There have been lessons in teamwork, team management and project management; which have been decisive for the accomplishment of the main objective of this internship: to create a preliminary study of a manned mission to the Martian moons.

Outputs of the current study include the need to perform further research on psychological stability of smaller crew sizes, and resistance to accelerations in a hyperbolic entry after 600 days in microgravity. A study concerning Nuclear Electrical Propulsion has also been performed and as it is a very promising, but still with a very low TRL, technology an overview can be found in Section 9.1.

9.1 NUCLEAR ELECTRICAL PROPULSION To assess that the study utilises the best technologies available, alternative options have to be considered. The reference missions performed by ESA and NASA typically use hydrogen nuclear thermal propulsion (NTP), storable and cryogenic propellants for manned spaceflight. Another technology suitable is nuclear electrical propulsion (NEP) that could potentially reduce the IMLEO significantly. Objectives within this chapter:  To size the power & propulsion systems

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 To maintain key Project Orpheus features  Minimise IMLEO (< 6500 t)  Understand the effects of intermediate orbits  To provide preliminary departure/arrival data

Constraints:  Total ission time ≤ 610 days, ars stay ≤ 60 days  CIV dry mass = 210 tonnes

9.1.1 Low Thrust Mission Overview To start with the study of a NEP, it is necessary to understand the orbital mechanics of the manoeuvres. In case of impulsive manoeuvre, the gravitational effect is negligible and the Lambert transfer is achievable with one burn only. A low thrust electrical propulsion system requires long burn durations to achieve significant changes in velocity, thus manoeuvres cannot be considered impulsive. In case of low thrust the trajectory is not an ellipse, but a spiral. Previously calculated mission orbital dynamics are no longer valid for such a propulsion system and thus have to be revised, for example the studied ephemeris and the launch dates. In fact, since the trajectories are different the time of flight (ToF) will be different too.

9.1.2 Analytical Methods To simplify the initial problem ΔV estimations have been made using Figure 9-1 in which the gravitational effect to the low thrust has been introduced, essentially accounting for losses. A pessimistic value of 2.3 has been utilized (Turner) and the nominal ΔV for the Mission Orpheus has been implemented in the next table.

Figure 9-1: Velocity increment loss factor as a function of spacecraft thrust-to-weight ratio for electric propulsion. (Turner, Rocket and Spacecraft Propulsion, 2009 Third Edition)

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Table 9-2: ΔV breakdown by phase. Comparisons between nominal impulsive and proposed low thrust trajectories. Stock (Impulsive) Nuclear Electric Propulsion

Phase Initial ΔV (km/s) Duration Corrected ΔV (km/s)

C-I-1 Transfer Phase to Mars 4.23 360 days 9.73

C-I-2 Transfer phase to Earth 3.02 206 days 6.95

C-M-1 Mars Capture Orbit Phase 0.20 5 days 0.46

C-M-2 Mars Phobos Transit Phase 0.72 3 days 1.64

C-M-3 Phobos proximity phase 0.15 20 days 0.35

C-M-4 Graveyard orbit 0.23 7 days 0.54

C-E-2 Earth Arrival 0.15 1days 0.35

Total 8.70 602 days 20.01

9.1.3 Numerical Methods for Ephemeris The low thrust ΔVs are defined using the nominal mission. It is necessary to show the correct ephemeris of the planets in order to achieve the target. A model in MATLAB has been developed starting with a random departure date. The first step has been the development of a propagator. It is based on the restricted two body equation (an adequate level of approximation for the actual level of detail): ̈ ⃗. There is future scope to implement n- body calculations. This has been demonstrated to be acceptable at modelling planetary orbits with elliptical shape and conserving mechanical energy The next step has been to approximate the low thrust propulsion as a sequence of small impulsive manoeuvres. Between each impulsive manoeuvre the propagator moves the spacecraft in the Solar System. At present the low thrust model is simplified to have the ΔV vector always in the same direction of the spacecraft velocity vector. This presents a significant limitation since the Earth and Mars are in two different orbital planes that are inclined by 1.85 degrees. However, given suitable stopping conditions the model works well and approximate launch windows arriving at the proximity of Mars can be determined. Shown in Figure 9-2 are the resulting orbits with multiple trajectories with varying Earth departure dates and constant force.

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Figure 9-2: Multiple trajectories (green) with constant thrust (200N) for various Earth departure dates. Complete ellipse (blue) is Earth's orbit; partial ellipse (red) is Mars’ orbit over 1 Earth year. Cross symbols show final position after 1 Earth year. From the model developed in MATLAB it is possible to obtain the overall ΔV and the required force. However due to the approximation made and early development of the model, the values of ΔV previously defined with the gravity correction factor are used going forward. The necessary force calculated is of order 300 N. At this point it is useful to define the mass at each phase. To calculate this it is necessary to have an estimation of the CIV dry mass, minus the Nuclear Thermal Reactor, equating to 197 tonnes.

9.1.4 System Architecture and Sizing This section concerns the sizing of the nuclear electrical power system in term of systems mass and power (thermal and electrical). The first calculation is about the cluster of electrical propulsion which the CIV needs. The prolusion system chosen is the VASMIR, which has a specific impulse of 3000 s and can throttle the level of thrust. According with the VASMIR actual technology (AD ASTRAROCKET COMPANY) and the required force estimated it derives that it is necessary to implement a cluster of 84 VASMIR. The cluster will need of 16.8MW in term of electrical power. The basics of the Nuclear Electrical Engine is to exploit the heat produced from the nuclear fuel and generate electrical power due the consequential difference in temperature. There are different thermodynamic cycles that have been taken into consideration. In this study Stirling and the Brayton cycles have been considered. According with Mason (2001), the Brayton cycle is chosen due to its higher efficiency. Figure 9-3 shows the architecture of the NEP system. Due to mechanical constraints, the upper temperature of the cycle is 1400 K Mason (2001) and the lower temperature is 550 K to avoid the solidification of the Lithium coolant. Lithium has been chosen due to its very good thermal properties and liquid phase in the temperature working range of the Brayton loop. The resulting efficiency is 58%. Which a known required electrical power (16.8 MW), it is possible to evaluate the Nuclear Thermal power (30.7 MW) and consequentially the heat waste to be rejected which is 13.9 MW.

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Figure 9-3: Nuclear electrical production (NEP) architecture. The heat to be rejected into space drives directly the dimension and the mass of the heat radiators. In order to decrease the mass, it was thought to add a refrigeration thermodynamic cycle requiring additional mechanical power. The benefit of this loop is to reject heat waste through use of hot radiators, increasing their radiation output per unit area. Due to the increase of system complexity and additional hardware required, it results does not give any real befit. The overall rejection area is about 3100 m2. In the next table there is the summary of the estimation mass for the NEP. Table 9-3: NEP dry mass breakdown system BEE (Tonnes) BEE with margin 20% (kg) BEE with Uncertainties 10% (kg)

CIV Mass 197 197 197

Reactor 39 47 52

Heat exch lit-gas 42 50 55

Radiator 19 22 24 cluster_VASMIR 38 45 50

Radiation shielding 10 12 14

Brayton loop 35 42 47

Total 380 411 433

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At this point, implementing the Tsiolkovsky rocket equation, the defined ΔV, the force and the dry mass it is possible to define the mass for each phase and the IMLEO: which is 811 tons.

Table 9-4: NEP mass at beginning of each mission phases Vehicle Mass Force Phase Adj Duration Acceleration (t) (N)

C-I-1 Transfer Phase to Mars 350 Days 3.22E-04 m/s2 811 261

C-I-2 Transfer phase to Earth 170 Days 4.47E-04 m/s2 527 249

C-M-1 Mars Capture Orbit Phase 11 Days 4.84E-04 m/s2 583 292

C-M-2 Mars Phobos Transit Phase 37 Days 5.14E-04 m/s2 574 295

C-M-3 Phobos proximity phase 20 Days 2.00E-04 m/s2 543 108

C-M-4 Graveyard orbit 12 Days 5.19E-04 m/s2 536 278

C-E-2 Earth Arrival 6 Days 6.66E-04 m/s2 416 277

The electrical propulsion cluster has been designed to provide a force of 300 N, so, in order to deal with this constraint, the duration of each phase has been modified. The new overall duration of the mission is 606 days, which is in any case under the constrained limit of 610 days.

9.1.5 Conclusion The proposed high efficiency electrical prolusion presents a viable alternative for the nominal Orpheus mission. The idea is to have a cluster of low thrust propulsion devices in order to achieve an acceptable force and rendezvous with the target within time constraints. The energy required from the cluster is large, but achievable with well-known nuclear technology available on Earth for more than 60 years. The solution described needs an IMELO mass of 811 tonnes against the 3981 tonnes of the nominal Orpheus mission. Another benefit of this kind of technology is that it can be parked in LEO or in L2 after each use, serviced and potentially reused to lower mission costs for the future SEEDS programs. To reduce manned transfer times it is possible to have the spacecraft parked in L2, a small chemical crewed spacecraft can be sent to the main spacecraft that then starts the mission from L2 saving propellant. The journey from LEO to L2 and vice versa can be achieved with the spacecraft when it is still unmanned. If this manoeuvre will be introduced, an additional ΔV of 8 km/s must be taken in consideration (it should be the phase C-E-2 in the previous table). This considers also the gravitational loss effect of Earth. In this new scenario, the spacecraft will need additional propellant and, in case of Orpheus mission, the IMLEO will be 1103 tonnes. However the technology to be developed for this kind of spacecraft is definitively more expensive. The TRL is not a level that makes the solution feasible at present, but it is highly probable that this will become a viable alternative for future manned interplanetary missions.

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Figure 9-4: Artistic view of the CIV in case of Nuclear Electric Propulsion System Figure 9-4 shows an artistic concept of the CIV with the Nuclear Electric Propulsion system. The habitable module is unchanged. The tanks around the spacecraft shown are unchanged; however, they will change since cryogenic fuel is not being used for NEP. The aim of the picture is simply to show the huge presence of radiators in front of the spacecraft; located at the end is the nuclear electric system. The main object of the figure is to give an idea about the change in architecture, and the size of the system when compared with the nominal solution, previously proposed in this work.

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9.2 REFERENCES TURNER, M., Rocket and Spacecraft Propulsion, 2009, Third Edition, Praxis MASON (2001) A Comparison of Brayton and Stirling Space Nuclear Power Systems for Power Levels. OTTING (n.d.) Brayton power conversion system technology development. AD ASTRAROCKET COMPANY (2009) http://www.adastrarocket.com/AIAA-2010-6772-196_small.pdf ANGELO Space Nuclear Power (Orbit, a foundation series). BRUNO Nuclear Space Power and Propulsion Systems (Progress in Astronautics and Aeronautics).

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10 APPENDIX

10.1 APPENDIX – MISSION ANALYSIS - CIV MISSION PHASES

Spacecraft Phase of Flight Operational Mode Duration

CIV Assembly phase Station keeping, Assembly t < 5 years (C-E-1)

Objectives To assemble the CIV and to ensure the operability of the systems

Interfaces with External Systems Launchers

External Environments

● Microgravity ● Residual atmosphere ● Atomic oxygen ● Frequent light-eclipse transitions ● Earth's Albedo ● Infrared thermal radiation from both sides of the planet ● Electrostatic charges due to UV ● Plasma ● Van Allen Belts & South Atlantic Anomaly radiation ● Reduced radiation from GCR and SPE ● Abundant orbital debris

Remarks

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Spacecraft Phase of Flight Operational Mode Duration

CIV Transfer phase to Mars Orbit control 360 days (C-I-1)

Objectives To transfer the CIV from LEO to Mars proximity

Interfaces with External Systems

External Environments

EARTH ORBIT ENVIRONMENT ● Microgravity ● Residual atmosphere ● Atomic oxygen ● Frequent light-eclipse transitions ● Earth's Albedo ● Infrared thermal radiation from both sides of the planet ● Electrostatic charges due to UV ● Plasma ● Van Allen Belts & South Atlantic Anomaly radiation ● Reduced radiation from GCR and SPE ● Abundant orbital debris

MARS ORBIT ENVIRONMENT ● Microgravity ● Possible Atomic oxygen ● Gravitational perturbations from Phobos ● Possible dust lifted from Phobos ● Daylight-eclipse transitions ● Infrared radiation from Mars and Phobos ● Mars’ and Phobos’ albedo ● GCR and SPE ● Meteoroids ● Solar plasma and possible electrostatic charges. ● Potential high Mars atmospheric (ionosphere) interactions

INTERPLANETARY MEDIUM ENVIRONMENT ● Microgravity ● Crossing of the Van Allen radiation belts during first hour of departing from Earth. ● High SPE and GCR radiation. ● Solar radiation pressure ● Constant sunlight ● High vacuum ● Micro-meteoroids ● Solar power decreasing steadily

VENUS PROXIMITY ENVIRONMENT ● High solar constant ● High SPE levels ● Planetary Albedo

Remarks

Venus fly-by

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Spacecraft Phase of Flight Operational Mode Duration

CIV Capture orbit phase Position keeping, Science on 5 days (C-M-1) Mars, Docking

Objectives To insert the CIV in capture orbit and to perform docking with the MATV

Interfaces with External Systems MATV, PhL, Lab

External Environments

Mars, Phobos MARS ORBIT ENVIRONMENT ● Microgravity ● Possible Atomic oxygen ● Gravitational perturbations from Phobos ● Possible dust lifted from Phobos ● Daylight-eclipse transitions ● Infrared radiation from Mars and Phobos ● Mars’ and Phobos’ albedo ● GCR and SPE (no magnetosphere) ● Meteoroids ● Solar plasma and possible electrostatic charges. ● Potential high Mars atmospheric (ionosphere) interactions

MARS ATMOSPHERE AND SURFACE RADIATION ENVIRONMENT ● CGR and SPE reduced due to shielding by the planet ● Gravity environment

Remarks

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Spacecraft Phase of Flight Operational Mode Duration

CIV Phobos transit phase Several manoeuvres, incl. RDV 3 days (C-M-2) with Phobos

Objectives To transfer the CIV from capture orbit to Phobos’ proximity

Interfaces with External Systems PhL, Lab

External Environments

Phobos, Mars MARS ORBIT ENVIRONMENT ● Microgravity ● Possible Atomic oxygen ● Gravitational perturbations from Phobos ● Possible dust lifted from Phobos ● Daylight-eclipse transitions ● Infrared radiation from Mars and Phobos ● Mars’ and Phobos’ albedo ● GCR and SPE (no magnetosphere) ● Meteoroids ● Solar plasma and possible electrostatic charges. ● Potential high Mars atmospheric (ionosphere) interactions

MARS ATMOSPHERE AND SURFACE RADIATION ENVIRONMENT ● CGR and SPE reduced due to shielding by the planet ● Gravity environment

Remarks

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Spacecraft Phase of Flight Operational Mode Duration

CIV Phobos proximity phase Continuous station keeping 20 days (C-M-3)

Objectives To perform remote science on Phobos and to ensure the safety of the multiple docking between the CIV and the PhL

Interfaces with External Systems PhL, Lab

External Environments

Phobos, Mars MARS ORBIT ENVIRONMENT ● Microgravity ● Possible Atomic oxygen ● Gravitational perturbations from Phobos ● Possible dust lifted from Phobos ● Daylight-eclipse transitions ● Infrared radiation from Mars and Phobos ● Mars’ and Phobos’ albedo ● GCR and SPE (no magnetosphere) ● Meteoroids ● Solar plasma and possible electrostatic charges. ● Potential high Mars atmospheric (ionosphere) interactions

MARS ATMOSPHERE AND SURFACE RADIATION ENVIRONMENT ● CGR and SPE reduced due to shielding by the planet ● Gravity environment

Remarks

During the entire phase, regular ignitions of the CIV main engine necessary for station keeping, due to the gravity of Phobos

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Spacecraft Phase of Flight Operational Mode Duration

CIV Graveyard orbit phase Orbit control 7 days (C-M-4)

Objectives To transfer the CIV with the PhL and the Lab to a graveyard orbit and to release them

Interfaces with External Systems PhL, Lab

External Environments

Phobos, Mars MARS ORBIT ENVIRONMENT ● Microgravity ● Possible Atomic oxygen ● Gravitational perturbations from Phobos ● Possible dust lifted from Phobos ● Daylight-eclipse transitions ● Infrared radiation from Mars and Phobos ● Mars’ and Phobos’ albedo ● GCR and SPE (no magnetosphere) ● Meteoroids ● Solar plasma and possible electrostatic charges. ● Potential high Mars atmospheric (ionosphere) interactions

MARS ATMOSPHERE AND SURFACE RADIATION ENVIRONMENT ● CGR and SPE reduced due to shielding by the planet ● Gravity environment

Remarks

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Spacecraft Phase of Flight Operational Mode Duration

CIV Transfer phase to Earth Orbit control 206 days (C-I-2)

Objectives To transfer the CIV from the graveyard orbit to Earth proximity

Interfaces with External Systems

External Environments

Mars, Deep space environment MARS ORBIT ENVIRONMENT ● Microgravity ● Possible Atomic oxygen ● Gravitational perturbations from Phobos ● Possible dust lifted from Phobos ● Daylight-eclipse transitions ● Infrared radiation from Mars and Phobos ● Mars’ and Phobos’ albedo ● GCR and SPE (no magnetosphere) ● Meteoroids ● Solar plasma and possible electrostatic charges. ● Potential high Mars atmospheric (ionosphere) interactions

MARS ATMOSPHERE AND SURFACE RADIATION ENVIRONMENT ● CGR and SPE reduced due to shielding by the planet ● Gravity environment

INTERPLANETARY MEDIUM ENVIRONMENT ● Microgravity ● Crossing of the Van Allen radiation belts during first hour of departing from Earth. ● High SPE and GCR radiation. ● Solar radiation pressure ● Constant sunlight ● High vacuum ● Micro-meteoroids ● Solar power decreasing steadily

Remarks

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Spacecraft Phase of Flight Operational Mode Duration

CIV Earth proximity End of operations 1 days (C-E-2)

Objectives To ensure the end of operations and the safety of the Earth

Interfaces with External Systems Re-entry capsule

External Environments

Earth, Deep space environment

EARTH ORBIT ENVIRONMENT ● Space environment as specified before ● Heat and kinematic loads during re-entry ● Increasing gravity levels ● Increasing density of atmosphere ● Landing surface properties

INTERPLANETARY MEDIUM ENVIRONMENT ● Microgravity ● Crossing of the Van Allen radiation belts during first hour of departing from Earth. ● High SPE and GCR radiation. ● Solar radiation pressure ● Constant sunlight ● High vacuum ● Micro-meteoroids ● Solar power decreasing steadily

Remarks

This phase ends when the Re-entry capsule will be detached from the CIV

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10.2 APPENDIX – MISSION ANALYSIS - EMPLOYED MARS ORBITS OF THE ORPHEUS MISSION Table 10-1: Operational orbits in the vicinity of Mars Orbital data

1 CIV Mars Capture orbit CIV Mars Capture orbit Pericentre altitude 298 Semi-major axis [km] 6,531.50 Pericentre radius 3687 Period [s] 16,026.32 Apocentre altitude 5987 Period [h] 4.45 Apocentre radius 9376 V_Pericentre [km/s] 4.08 Inclination 0 V_Apocentre [km/s] 1.61 2 MATV Mars Science orbit MATV Mars Science orbit Pericentre altitude 298 Semi-major axis [km] 6,531.50 Pericentre radius 3687 Period [s] 16,026.32 Apocentre altitude 5987 Period [h] 4.45 Apocentre radius 9376 V_Pericentre [km/s] 4.08 Inclination 30 V_Apocentre [km/s] 1.61 3 CIV Mars Phobos orbit CIV Mars Phobos orbit Pericentre altitude 5845.42 Semi-major axis [km] 9,376.00 Pericentre radius 9234.42 Period [s] 27,563.91 Apocentre altitude 6128.58 Period [h] 7.66 Apocentre radius 9517.58 V_Pericentre [km/s] 2.17 Inclination 1.093 V_Apocentre [km/s] 2.11 4 CIV Earth Departure orbit CIV Earth Departure orbit Pericentre altitude 400 Semi-major axis [km] 20,847.30 Pericentre radius 6771 Semi-minor axis [km] 18,115.20 C3 [km/s]² 19.12 Ecc 1.32 Inclination 0 ΔV compared to 400km 4.03 circular 5 CIV Hyperbolic arrival orbit CIV Hyperbolic arrival orbit Pericentre altitude 298 Semi-major axis [km] 1,222.62 Pericentre radius 3687 Semi-minor axis [km] 4,754.95 C3 [km/s]² 35.03 Ecc 4.02 Inclination 0 V_Pericentre 7.63 6 MATV Hyperbolic arrival orbit MATV Hyperbolic arrival

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orbit Pericentre altitude 298 Semi-major axis [km] 42,828.30 Pericentre radius 3687 Semi-minor axis [km] 18,149.65 C3 [km/s]² 1 Ecc 1.09 Inclination 0 V_Pericentre 4.92 7 MATV Earth Departure orbit MATV Earth Departure orbit Pericentre altitude 400 Semi-major axis [km] 398,600.40 Pericentre radius 6771 Semi-minor axis [km] 73,781.39 C3 [km/s]² 1 Ecc 1.02 Inclination 0 V_Pericentre 10.90 ΔV compared to 400km 3.22 circular 8 MATV Mars polar orbit Mars Science orbit Pericentre altitude 298 Semi-major axis [km] 6,531.50 Pericentre radius 3687 Period [s] 16,026.32 Apocentre altitude 5987 Period [h] 4.45 Apocentre radius 9376 V_Pericentre [km/s] 4.08 Inclination 90 V_Apocentre [km/s] 1.61 9 CIV Hyperbolic exit orbit from graveyard Hyperbolic exit orbit orbit Pericentre altitude 6628 Semi-major axis [km] 2,786.49 Pericentre radius 10017 Semi-minor axis [km] 12,496.59 C3 [km/s]² 15.37 Ecc 4.59 Inclination 0 V_Pericentre 4.89 10 CIV Graveyard orbit Graveyard orbit Pericentre altitude 6628 Semi-major axis [km] 10,017.00 Pericentre radius 10017 Period [s] 30,438.34 Apocentre altitude 6628 Period [h] 8.46 Apocentre radius 10017 V_Pericentre [km/s] 2.07 Inclination 1.093 V_Apocentre [km/s] 2.07 11 CIV Graveyard transfer orbit Graveyard transfer orbit Pericentre altitude 5,987.00 Semi-major axis [km] 9,625.71 Pericentre radius 9234.42 Period [s] 28,672.37

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Apocentre altitude 6628 Period [h] 7.96 Apocentre radius 10017 V_Pericentre [km/s] 2.20 Inclination 1.093 V_Apocentre [km/s] 2.03

10.3 APPENDIX – SCIENCE REQUIREMENT MATRIX

Table 10-2: SRM-Objective 1

OBJECTIVE 1. To collect data in order to support the selection of appropriate future landing sites for humans on Mars

Science Objectives Science Investigations Measurement objective Instruments

1.1 To characterise the Analyse and 1.1.1a. Measure the size-frequency Visual Imaging topography of potential Mars characterise rocks function of rocks in a relevant area landing sites for humans size and abundances in areas of interest 1.1.1b. Measure the size-frequency Visual Imaging function of rocks globally

1.1.1 1.1.1c. Measure the rock coverage in Visual Imaging a relevant area

1.1.1d. Measure the rock coverage IR Imaging, EM Sounding globally

Analyse and 1.1.2a. Measure the slopes on the Visual Imaging characterise slopes sub-metre scale in the areas of interest

1.1.2 1.1.2b. Measure the slopes on the Visual Imaging, EM 10's of metres scale Sounding, Visual Imaging

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Investigate the 1.1.3a. To measure the composition IR Spectroscopy, Visual properties of the and compound abundances of the Imaging, LIBS, Surface regolith surface Sample Composition Analysis

1.1.3 1.1.3b. Measure the mechanical Visual Imaging, Thermal behaviour of the soil Probe

Investigate the 1.1.4a. Measure the mass EM Sounding, EM existence, location distribution of Mars' upper crust Sounding 1.1.4 and properties of submartian caves

1.2 To study the level of 1.2.1. Study the level 1.2.1a. Measure global surface UV Spectrometry, Gamma radiation on the surface of of radiation on the radiation dose Ray Spectrometry Mars for future human surface of Mars habitability

1.2.1 1.2.1b. Measure localised surface Radiation Detector radiation doses

1.2.2. Study the level 1.2.1c. Measure localised subsurface Radiation Detector of radiation below radiation doses 1.2.2 the surface of Mars

1.3 To characterise the 1.3.1 Investigate the 1.3.1a. Measure the wind speed and Weather Property Measure atmospheric conditions of dynamics and variation of wind speed on Mars Mars variability of Mars' weather

1.3.1b. Measure daily and seasonal Weather Property Measure variation of pressure simultaneously using a global network

1.3.1

1.3.1c. Measure daily and seasonal Weather Property Measure, variation of temperature IR Imaging simultaneously using a global network

1.3.1d. Measure daily and seasonal Weather Property Measure variation of humidity simultaneously using a global network

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1.3.1e. Measure an air density profile Weather Property Measure of Mars during descent in varying locations 1.3.1f. Measure a temperature profile Thermal Probe, Weather of Mars during descent in varying Property Measure locations

1.3.1g. Measure an pressure profile Weather Property Measure of Mars during descent in varying locations

1.3.1h. Analysis of physical weather Visual Imaging, IR Imaging, phenomena Visual Imaging, Weather Property Measure

1.3.2 Investigate the 1.3.2a. Measure the chemical Raman Spectroscopy, IR atmospheric composition of the Martian Spectroscopy, Atmospheric chemistry of Mars atmosphere Composition Analysis

1.3.2 1.3.2b. Measure the variation of the Raman Spectroscopy, IR composition of the Martian Spectroscopy, Atmospheric atmosphere with time Composition Analysis

Table 10-3 SRM-Objective 2

OBJECTIVE 2. To search for ISR sites on Mars

Science Objectives Science Investigations Measurement objective Instruments

2.1 To search for a site on To determine the 2.1.1a. To determine surface EM Sounding, IR Mars where water ice is availability of composition globally Spectroscopy, Gamma Ray accessible accessible water ice Spectroscopy

2.1.1b. To determine surface EM Sounding, Thermal composition in relevant areas Probe

2.1.1

2.1.1c. To determine sub-surface EM Sounding, IR composition globally Spectroscopy, Gamma Ray Spectroscopy, Seismometer 2.1.1d. To determine sub-surface EM Sounding, Thermal composition in relevant areas Probe, Surface Sample

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Composition Analysis 2.2 To search for accessible IA.3.a To investigate 2.2.1a. To test the feasibility of Atmospheric Composition CO2 ice on Mars the production of propellant production on the soil of Analysis, Propellant ISRU 2.2.1 propellant Mars feasibility

To investigate the 2.2.2a. To test the feasibility of life Atmospheric composition creation of buffer consumables (buffer gas) production analysis, Buffer gas ISRU gas (a mixture of on the soil of Mars feasibility demonstration nitrogen and argon) on Mars surface (K. 2.2.2 R. Sridhar, et. al, 2000)

Table 10-4 SRM-Objective 4

OBJECTIVE 4. To study the origin of Phobos

Science Objectives Science Investigations Measurement objective Instruments

4.1 To characterize the To investigate the 4.1.1a. To detect globally the Raman Spectroscopy, surface of Phobos composition of the elemental composition of Phobos GRS surface including rare and heavy elements.

4.1.1b. To measure the elemental Astronaut, Handheld abundance of Phobos including the surface analyser tool, 4.1.1 abundance of rare and heavy Mass Spectrometer, elements. Raman Spectroscopy, X- Ray Spectroscopy, Earth Sample Analysis

4.1.1c. To identify the compounds Raman Spectroscopy, that make up the regolith of Phobos GRS

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4.1.1d. To measure the compounds Astronaut, Handheld abundance that make up the regolith surface analyser tool, of Phobos Mass Spectrometer, Raman Spectroscopy, X- Ray Spectroscopy, Earth Sample Analysis

To elaborate 4.1.2b. To determine the topography HD Cameras, Laser topographic mapping of Phobos (& use it to determine Altimeter, EM sounder of the surface landing sites) 4.1.2

4.2 To study gravity in To elaborate a 4.2.1a. To globally map the gravity of Ranging System - KBR Phobos 4.2.1 gravitational map of Phobos Phobos 4.3 To characterize the To investigate the 4.3.1a. Characterise the crustal HD Cameras, IR Imager, internal structure of Phobos surface structure of structure and thickness Seismometer, Astronaut, Phobos Handheld surface analyser tool, Earth Sample 4.3.1 Analysis

To investigate the 4.3.2a. To characterise the internal Ranging System - KBR, internal structure of density of Phobos Seismometer, Astronaut 4.3.2 Phobos

Table 10-5 SRM-Objective 5

OBJECTIVE 5. To search for ISR on Phobos

Science Objectives Science Investigations Measurement objective Instruments

5.1 To characterise the To investigate the 5.1.1a. To detect globally the Raman Spectroscopy, surface properties of composition of the presence of Water, Hydrogen and GRS Phobos regolith / chemical Oxygen. properties of the regolith

5.1.1b. To measure the abundance of Astronaut, Handheld 5.1. 1 Water, Hydrogen and Oxygen. surface analyser tool, Mass Spectrometer, Raman Spectroscopy, X-Ray Spectroscopy, Earth Sample Analysis

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5.1.1c. To detect significant amounts Raman Spectroscopy, of organic molecules to investigate GRS, Astronaut, Handheld the potential of carbon for fuel surface analyser tool, Mass Spectrometer, Raman Spectroscopy, X- Ray Spectroscopy, Earth Sample Analysis

To characterise the 5.1.2a. Taking images to map the HD Cameras, IR Imager, topography of Phobos characteristics of surface & EM sounder, Seismometer, 5.1.2 subsurface features with respect to Astronaut planning future landing sites. To investigate the 5.1.3a. Determine the sinkage / Astronaut, Penetrometer, mechanical properties pressure function of Phobos regolith. Earth Sample Analysis 5.1. 3 of the regolith

To investigate the 5.1.4a. To detect the presence of GRS, Astronaut, Handheld distribution of Water, Hydrogen and Oxygen.. surface analyser tool different elements and compounds across the surface. (In particular 5.1.4b. To measure the abundance of Astronaut, Handheld difference between Water, Hydrogen and Oxygen. surface analyser tool, Mass stickney crater(blue Spectrometer, Raman unit) and red unit) Spectroscopy, X-Ray Spectroscopy, Earth Sample Analysis

5.1. 4 5.1.4c. To detect significant amounts Raman Spectroscopy, of organic molecules to investigate GRS, Astronaut, Handheld the potential of carbon for fuel surface analyser tool, Mass Spectrometer, Raman Spectroscopy, X-Ray Spectroscopy, Earth Sample Analysis

5.2 To characterise the To investigate the 5.2.1a. To Measure the density to Ranging Method - KBR, interior composition of density of the interior determine if there is water ice in the Seismometer, Astronaut Phobos core and to investigate the potential for mining any resources 5.2.1

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10.4 APPENDIX - SYSTEM ENGINEERING: SIQ

Thermal Control Subsystem Spacecraft Element (if apl.) Date

CIV Thermal Control 16.07.2014

Critical design Reasons/ Interfaces / Properties Current estimate elements (incl. lowest Sources TRL)

Radiator Panel Choice Using standard Current Mass Estimate Calculated 7.66 tonnes absorptivity and (Uncertainty?) BEE emissivity values for (Average uncertainty radiators combined 23%) with a NASA proposal for reduced kg/m^2.

Pumps, loops, Using ISS and ISS-like Current Power Best estimate with is compressors and heat scaled systems. consumption estimate 15.64W nominally, which exchangers Internal loop hard to (Uncertainty?) includes 36.3% define, seeking expert uncertainty. advice.

Surface coatings and Tanks-Advanced MLI Max heat radiated by 92.15 kW insulative materials on end, standard MLI radiators (kW) during closest solar on tank and white paint approach. at the rear end. CIV - Standard MLI sides and end with Adv MLI on front.

Tanks - internal heat To be designed, Threshold duration in 2hrs 38min per Kelvin flow proposing heat pipes worst case attitude - (10% unc.) and pump. closest sun approach

System engineering Validation, reference System engineering Validation, reference tools sys., sources tools sys., sources

Application of Met with Thales head of Confirmed ideas had no appropriate margins TCS and engineer major flaws, provided data Excel spreadsheets and uncertainties. The working on Orion’s TCS on certain items. with iterative design of input data was found craft design and from reputable sources. insulation distribution.

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10.5 APPENDIX – DESIGN MECHANICS AND STRUCTURES – TANKS TRADE-OFF In order to evaluate the best solution for the design of the tank structure able to carry 982t, a trade-off has been performed in order to evaluate the dimensions and the number of the elements. The three solutions are strongly linked to the launch capability (SpaceX Falcon Heavy vs. NASA SLS): in fact, a constraint of the tanks is the dimension of the launcher’s fairing

Item Weight 1 2 3

Launch cost [M$] 1753 3105 1741

Mass eff. [%] 100 100 100

Volume eff. [%] 93.08 45.59 93.85

N° of launches 15 6 15

Radius [m] 1.75 2.60 2.30

Length [m] 11.40 11.88 4.80

Launch cost 10 5.83 3.29 5.87

Mass efficiency 2 5.00 5.00 5.00

Volume efficiency 2 6.00 2.94 6.05

GCR (considering radius) 4 3.95 5.86 5.19

GCR (considering length) 4 6.09 6.35 2.56

TCS (considering length) 5 6.16 4.15 4.69

Staging possibility 4 6.25 2.50 6.25

Assembly 5 3.33 8.33 3.33

Design complexity 5 5.00 5.00 5.00

Weighted Scores 5.36 4.72 4.92

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10.6 APPENDIX - ECLS Table 10-6- Mass and Volume Estimates for Systems Mass (kg) Volume (m^3) Power (we) Items (2 s.f) (2 s.f) (2 s.f)

Atmospheric Pressure Control 120 0.26 71

CO2 Removal (4BMS) 180 0.42 540

O2 Generation (Electrolysis + Sabatier) 440 1 3300

Air Gaseous Trace Contaminant Control 86 0.4 190 Subsystem Atmosphere Composition Monitoring Assembly 54 0.09 104

Sample Delivery System 35 0.04 0

Fire Detection 1.5 0 1.5

Fire Suppression System 6.8 0.04 0

Common Cabin Air Assembly 120 0.5 530

Avionics Air Assembly 12 0.03 180 Thermal Control and Atmosphere Circulation 10 0.02 61 Circulation Atmospheric Microbial Control 100 0.27 0

Internal Thermal Control System 89 0.18 120

Solid Waste Collection 36 0.13 14 Waste Solid Waste Treatment 350 0.86 0 Subsystem Toilet 180 8.7 1100 All Water All Water Systems (MF + VCD) 2600 4.6 1100 Systems

(Including calculations for spares, integration etc) 7500 36 8800 Total The above with redundancy 15000 73 17600

Table 10-7: Total Mass and Volume Estimate Items Comments Mass (kg) (2 s.f) Volume (m^3) (2 s.f)

Dry Resources All supplies, not including water and oxygen 24000 180

Water Including tanks 10000 10

Oxygen Including tanks 430 1

Nitrogen Including Tanks 3400 3.6

SEEDS Executive Summary 09/2014 Page 178

Systems With Redundancy 15000 69

Total 53000 260

Table 10-8: Resources included. (Reproduced from Larson and Pranke, 1999) Food Food Freezers Conventional Oven Microwave oven Kitchen/oven cleaning supplies Sink, spigot for hydration of food & drinking water Dishwasher Cooking/Eating supplies(pans, dishes, plates etc) Waste Collection System Waste collection system WCS supplies (toilet paper, cleaning solutions, filters etc) Contingency Faecal and urine collection mittens/bag Personal Hygiene Shower Hand wash/mouthwash tap Personal Hygiene kit Hygiene supplies Clothing Clothes Washing Machine Clothes Dryer Rest Provisions Personal Stowage Sleep Provisions (restrains only) Sleep Space Housekeeping Vacuum (prime and 2 spares) Disposable wipes for housecleaning Trash compactor/trash lock Trash Bags Operational Supplies and Restraints Operational Supplies (diskette Ziploc, Velcro, tape) Restraints and Mobility aids Maintenance: Repairs in Habitable areas Hand tools and accessories

SEEDS Executive Summary 09/2014 Page 179

Spare parts/equipment & consumables Test Equipment (Oscilloscopes, Gauges etc) Fixtures, Large machine tools. glove boxes, etc Photography Equipment (still &video cameras, lenses, etc) Film (Assuming everything is digital) Crew Healthcare Exercise Equipment Medical/ Surgical/Dental suite Medical/ Surgical/Dental consumables

Table 10-9- Mass and Volume Estimates for Each System Mass (kg) Volume (m^3) Power (we) Items (2 s.f) (2 s.f) (2 s.f)

Atmospheric Pressure Control 120 0.26 71

CO2 Removal (4BMS) 180 0.42 540

O2 Generation (Electrolysis + Sabatier) 440 1 3300

Air Gaseous Trace Contaminant Control 86 0.4 190 Subsystem Atmosphere Composition Monitoring Assembly 54 0.09 104

Sample Delivery System 35 0.04 0

Fire Detection 1.5 0 1.5

Fire Suppression System 6.8 0.04 0

Common Cabin Air Assembly 120 0.5 530

Avionics Air Assembly 12 0.03 180 Thermal Control and Atmosphere Circulation 10 0.02 61 Circulation Atmospheric Microbial Control 100 0.27 0

Internal Thermal Control System 89 0.18 120

Solid Waste Collection 36 0.13 14 Waste Solid Waste Treatment 350 0.86 0 Subsystem Toilet 180 8.7 1100 All Water All Water Systems (MF + VCD) 2600 4.6 1100 Systems

(Including calculations for spares, integration etc) 7500 36 8800 Total The above with redundancy 15000 73 17600

SEEDS Executive Summary 09/2014 Page 180

Table 10-10: Total Mass and Volume Estimate Items Comments Mass (kg) (2 s.f) Volume (m^3) (2 s.f)

Dry Resources All supplies, not including water and oxygen 24000 180

Water Including tanks 10000 10

Oxygen Including tanks 430 1

Nitrogen Including Tanks 3400 3.6

Systems With Redundancy 15000 69

Total 53000 260

SEEDS Executive Summary 09/2014 Page 181