Number: MILESTONE-SEEDS-00064 Title: Mission MILESTONE: Executive Summary Date: 16 September 2015

Contact: [email protected]

Marco BALDELLI, Samuel BROWN, Alberto FERRERO, Oliver HARDY, Rachel HENSON, Lorenzo MENGHINI, Gerard MORENO- Prepared by: TORRES BERTRAN, Jacopo Date: 10/09/2015 PISACRETA, Silvio SILVESTRELLI, Marco VOLPONI, Ameer ZAILANI

Mission MILESTONE: Executive Summary Page i TABLE OF CONTENTS Nomenclature ...... 1 Abstract ...... 2 Acknowledgements ...... 2 1 Introduction ...... 3 1.1 SEEDS ...... 3 1.2 SEEDS VII: MILESTONE ...... 3 1.3 Team Members ...... 4 2 Mission Summary ...... 7 2.1 Mission Statement ...... 7 2.2 Mission Objectives ...... 7 2.3 Mission Architecture ...... 7 2.4 Mission Assumptions...... 7 2.5 Traffic Model ...... 8 2.6 Crew Selection ...... 9 2.6.1 Introduction ...... 9 2.6.2 Crew Training and Mission Roles ...... 9 2.6.3 Selection Criteria Unique to Mars Missions ...... 10 2.7 Landing Site ...... 10 2.7.1 Amazonis Planitia (15N, 155W) ...... 11 2.7.2 Isidis Planitia (12N, 87E) ...... 11 2.7.3 Chryse Planitia (27N, 40W) ...... 11 2.7.4 Elysium Planitia (3N, 155E) ...... 11 2.7.5 Candor Chasma (7S, 71W) ...... 12 3 Mission Analysis ...... 13 3.1 Mission Phases and Scenarios ...... 13 3.1.1 Cargo Phases and Timeline ...... 13 3.1.2 Crew Phases and Timeline ...... 14 3.2 Risk Analysis ...... 15 3.3 Contingency Strategy ...... 16 4 Scientific Objectives and Operations ...... 18 4.1.1 Science Objectives ...... 18 4.2 Science Matrix Breakdown ...... 18 4.3 Scientific Architecture ...... 19 4.3.1 Laboratory Module ...... 19 4.3.2 Static Landers ...... 20 4.3.3 Droppable atmospheric probe...... 21 4.3.4 Rover Scientific Payload ...... 21 4.3.5 Human Deployable Assets ...... 22 4.4 Scientific Timeline ...... 23 5 Power ...... 24 5.1 Outpost Power Consumption ...... 24 5.2 Architecture ...... 24 5.2.1 Nuclear ...... 24 5.2.2 Solar Array and Regenerative Fuel Cells ...... 25 5.2.3 Distribution ...... 25

Mission MILESTONE: Executive Summary Page ii 5.2.4 Total power system launch mass...... 25 6 Module Defining Parameters ...... 26 6.1 Outpost Architecture...... 26 6.1.1 Outpost Layout ...... 26 6.2 Module Physical Design ...... 26 6.2.1 Primary Structure ...... 27 6.2.2 Secondary Structure ...... 28 6.2.3 Structure Mass definition ...... 29 6.3 Thermal Control System...... 29 6.4 Autonomous Electric Power System ...... 31 6.5 Radiation Protection ...... 31 6.6 Planetary Protection ...... 32 6.7 Environment Control and Life Support Systems (ECLSS) ...... 32 6.7.1 Air Revitalisation Subsystem (ARS) ...... 32 6.7.2 Atmosphere management, Thermal Humidity Control and Atmosphere Control and Supply .... 33 6.7.3 Fire Detection and Suppression (FDS) ...... 34 6.7.4 Water Recovery System (WRS) ...... 35 6.7.5 Waste Management System (WMS) ...... 35 7 Habitable Module ...... 36 7.1 Module Sizing ...... 36 7.1.1 Internal Layout ...... 36 7.1.2 Structure Mounting ...... 36 7.2 Environmental Control and Life Support Systems (ECLSS) ...... 36 7.3 Crew Utilities ...... 36 8 Extravehicular Activities (EVA) Module ...... 38 8.1 Crew Transfer Scenarios ...... 38 8.2 Physical Design ...... 38 8.3 EVA Suits ...... 38 8.3.1 Dust Removal System ...... 39 9 Greenhouse module ...... 40 9.1 Greenhouse Layout and Structure ...... 40 9.2 Growth Considerations ...... 40 9.2.1 Growth Mechanism ...... 40 9.2.2 Nutrient Management ...... 40 9.2.3 Illumination ...... 41 9.2.4 Irrigation System ...... 41 9.2.5 Atmosphere Management ...... 41 9.2.6 Robotic Assistance ...... 41 9.3 Waste Management ...... 42 9.4 Crew Diet ...... 43 9.5 Storage ...... 43 9.6 Contingency Considerations ...... 44 10 Node ...... 45 10.1 Node Configuration ...... 45 10.2 Node Structure ...... 45 10.3 Node Mass Budget ...... 46

Mission MILESTONE: Executive Summary Page iii 11 Entry, Descent and Landing (EDL) ...... 47 11.1 Landing Approach ...... 47 11.2 Landing Ellipse ...... 47 11.3 Touchdown Subsystem ...... 48 11.4 Hypersonic Inflatable Aerodynamic Decelerator (HIAD) Subsystem ...... 48 11.5 Retrorocket Subsystem ...... 49 11.6 Sensing Subsystem...... 49 11.7 Heat Shield Ejection Subsystem ...... 49 11.8 Optimisation ...... 49 11.9 Simulation Results ...... 50 11.10 Module Tailoring ...... 52 12 Mars Descent Vehicle (MDV) ...... 54 12.1 Human Descent Phase ...... 54 12.1.1 MDV Sizing ...... 54 13 Ascent ...... 56 13.1 Mission Constraints ...... 56 13.2 Role of the Crew Interplanetary Vehicle (CIV) ...... 56 13.3 Manoeuvres for Ascent ...... 56 13.3.1 Launch and Pitch Over ...... 56 13.3.2 Rendezvous ...... 56 13.3.3 Docking ...... 56 13.3.4 Disposal of Ascent Vehicle ...... 57 13.4 Delta-V Budget ...... 57 13.5 Ascent Duration ...... 57 13.6 Cryogenic Fluid Management ...... 57 13.7 Crew Entry Procedure ...... 58 14 Rover - Exploration and Mobility System ...... 59 14.1 Collection of the cargo...... 59 14.2 Collection of the MDV...... 60 14.3 Manned Configuration ...... 60 14.4 Modes of Operation ...... 61 15 Communication and Data Handling ...... 62 15.1 Data Rates and Rationale ...... 62 15.2 Communications Architecture ...... 62 15.2.1 Cargo Mission ...... 62 15.2.2 Crew Mission...... 63 15.2.3 Communication Satellite Design ...... 64 15.3 Data Handling ...... 65 15.3.1 Storage ...... 65 15.3.2 Computing Power ...... 65 16 In-Situ Resource Utilisation (ISRU) ...... 66 16.1 ISRU Resource Requirement ...... 66 16.2 Chemical Processes...... 66 16.3 ISRU Budget ...... 67 17 Technology Output ...... 69 17.1 Mission Critical Technologies ...... 70

Mission MILESTONE: Executive Summary Page iv 18 Conclusions ...... 72 19 Bibliography ...... 73 Appendix A: Mass Budgets ...... 78 Appendix B: Power Budgets ...... 80 Appendix C: Science Requirements Matrix ...... 82 Appendix D: Technology Readiness Level Evaluation ...... 91 19.1 Mission Enabling Technologies ...... 91 19.2 Mission Enhancing Technologies ...... 94

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NOMENCLATURE ACS - Atmosphere Control and Supply AEPS - Autonomous Electric Power System ARS - Air Revitalisation System ATCS - Active Thermal Control System C&DH - Communications and Data Handling CIV - Crew Interplanetary Vehicle COMM - Communications System COSPAR - Committee On Space Research DHMS - Data Handling Managing System DSN - Deep Space Network ECLS -Environmental Control and Life Support EDL - Entry, Descent and Landing EVA - Extravehicular Activity FDS - Fire Detection and Suppression GCR - Galactic cosmic ray GH - Greenhouse HAB - Habitable module HPGS - High Pressure Gas System ISRU - In Situ Resources Utilization LAB - Laboratory LEO - Low Earth Orbit LMO – Low Mars Orbit MAV - Mars Ascent Vehicle MDV - Mars Descent Vehicle MILESTONE - Mars Initial ExpeditionS Toward a New Era NFT - Nutrient Film Technique POW - Power source RFC - Regenerative Fuel Cell SEEDS - SpacE Exploration Development Systems TCS - Thermal Control System TDRS - Tracking and Data Relay Satellite THC - Thermal Humidity Control TT&C - Telemetry, Tracking and Command TTRD - Time To Recover Damages WMS - Waste Management System WRS - Water Recovery System

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ABSTRACT The Mission MILESTONE (Mars Initial Landing ExpeditionS Towards a New Era) feasibility study focuses on the definition of the first landing mission on the Martian surface. A 500 day transit is assumed for the crew, with the intention of residing on the surface for 60 days. As the mission is intended as a pre-cursor to exploration missions, it is focused on the establishment of a habitable environment to ensure the supply of sufficient resources required to sustain human life. The mission analysis starts and ends in Low Mars Orbit with a detailed design and investigation focused on the surface modules and vehicles, while it does not carry out a detailed analysis on the Earth-Mars transit. As the first human expedition on the Martian surface, the primary objective of Mission MILESTONE is the assembly of a complete outpost, a first step toward a longer presence of mankind on Mars. The outpost is formed of a number of main components: a habitable module, a laboratory, a greenhouse, an ISRU system and a power plant. A pressurised rover is also intended to allow for exploration in future missions, and three communications satellites and an Earth return vehicle are placed in Martian orbit. The outpost is sized for future 500 day scientific and exploration missions that will consist of a crew of six people. While the outpost is intended to be largely self-sufficient in order to reduce the pressure on the launch system, there will be some reliance on Earth resources, with each crew bringing certain consumables. During this first manned mission to the Martian surface, there will also be technology demonstrations, increasing technology readiness levels for the benefit of future missions.

ACKNOWLEDGEMENTS We would like to thank all the people who have helped and supported us during the SEEDS project. Firstly, we would like to thank the mentors at the Politecnico di Torino, in particular Professor Ernesto Vallerani, Professor Gianfranco Chiocchia, Professor Nicole Viola, Mr Enrico Beruto, Mr Eugenio Gargioli, Mr Flavio Bandini and Ms. Annamaria Piras. We would also like to thank all the experts at Thales Alenia Space and ALTEC for their help. Secondly, we would like to thank the mentors at ISAE, Toulouse, in particular Ms. Stephanie Lizy-Destrez and M. Emmanuel Zenou. In addition, we would like to thank the experts at CNES, Thales Alenia Space and Airbus Defence and Space. Finally, we would like to thank the mentors at the University of Leicester, in particular Dr. Richard Ambrosi, Dr Nigel Bannister, Dr Ian Hutchinson, Dr Hugo Williams and Dr John Bridges. We could not have completed this project without the help and support of our mentors, and are extremely grateful for the time they dedicated to us.

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1 INTRODUCTION

1.1 SEEDS SpacE Exploration Development Systems (SEEDS) is an International Masters Program which originated at the Politecnico di Torino in association with Thales Alenia Space. The SEEDS project 2015 covers the second step in a five year study "Conquest by Humans of Mars in Five Steps", with the five steps consisting of: 1. Step One - Exploration of Mars from its Proximity 2. Step Two - Exploration of the Mars Surface 3. Step Three - Exploitation of Mars Resources ISRU 4. Step Four - Development of Mars Permanent Outpost 5. Step Five - Development of Mars Independent Base

1.2 SEEDS VII: MILESTONE SEEDS VII is an international, multidisciplinary group comprised of students from both the Politecnico di Torino and the University of Leicester. The SEEDS VII project covers the second step of “Conquest by Humans of Mars in Five Steps”. The proposed mission to accomplish this task is MILESTONE: Mars Initial Landing ExpeditionS TOward a New Era. The overall aim for the mission was given as "The initial descent, permanence and departure of humans on/from Planet Mars". Mission MILESTONE will be an early stage mission, which will create an outpost on Mars for further human exploration by subsequent missions. This was determined to be the most cost- effective method, as a substantial outpost would be required to sustain human life on Mars, thus a multi- mission outpost is designed to allow sufficient time for experimenting to build a future permanent outpost. Mission MILESTONE’s insignia can be seen in Figure 1, and depicts the Monolith from Arthur C. Clarke’s novel “2001: A Space Odyssey” approaching the Martian surface. The Monolith is tied to the evolution of humankind and our relationship with the solar system. The use of the Monolith pays tribute not only to the media that inspires space exploration but also the work of previous Mars missions beneath the insignia is the Latin inscription of the mission’s primary aim “The first human expedition to the Martian surface”.

Figure 1- Mission MILESTONE Insignia

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1.3 TEAM MEMBERS MARCO BALDELLI Marco Baldelli was born in an Italian town known as Gualdo Tadino in 1985. He graduated in 2014 with a Masters Degree in Mechanical Engineering from the Politecnico Di Torino, Turin. For his master thesis, Marco was working on camera calibration methods which utilizes dynamic method to define the accuracy of the camera calibration. Apart from that, he is interested in many fields including robotics, music as well as fitness. During the SEEDS project, Marco was part of the Habitable module team, defining the structure of the outpost modules, and also of the EVA module, node and science teams. Marco was Presentation Manager, and contributed heavily to the graphical output of the team.

SAMUEL BROWN Samuel Brown was born in Hemel Hempstead, UK in 1993. Sam completed his undergraduate studies in 2014 with a BSc in Physics from the University of Leicester. Afterwards completing his BSc he continued to study at the University of Leicester on the Space Exploration Systems MSc course. Outside of work hours his main interests include skiing, ski touring as well as mountaineering. During the SEEDS project, Sam was involved with the Habitable module team, working on the water recovery system, radiation protection and simulations, as well as being part of the mission analysis, communications and science teams. During the final phase, he was also a systems engineer for the project, and developed a technology roadmap for the mission. ALBERTO FERRERO Alberto Ferrero was born in Torino in 1990. As a proud Torinese, he studied at the Politecnico di Torino and graduated with a Master Degree in Aerospace Engineering in the year 2014. During his first degree studies, he spent a year away in 2013 as Double Degree student at the Instituto Superior Tecnico in Lisbon, Portugal, obtaining the Portuguese title of Engenheiro Aeroespacial. Alberto was Project Manager for the Turin phase, while working in the Habitable module focused on structure and thermal definition, and in the Mission Analysis group. He was also part of the Entry, Descent and Landing (EDL) team, in particular working on the Mars Descent Vehicle. Finally he took part in the definition of the rovers, and worked in the Power and ISRU team.

OLIVER HARDY Oliver Hardy was born in 1991 and grew up in Warrington (UK). He completed a BSc in Physics at the University of Leeds in 2012, with a final project focusing on the analysis of polymer electrolyte batteries. Following graduation he worked for a start- up company, researching and developing microfluidic technologies for industry. Looking to transition into the space industry he enrolled in the SEEDS MSc program at the University of Leicester. During the SEEDS project, he developed the utilities for the crew and their planetary protection requirements, performed the mission analysis for the Mars Ascent Vehicle and worked on the nuclear reactor and was system engineer for the power team. He would like to continue learning as much as he can about space engineering and become a better dancer.

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RACHEL HENSON Rachel Henson was born in Swindon, UK in 1991. She obtained her Bachelors Degree in Physics from the University of Oxford in 2012. In September 2014, she began studying the MSc in Space Exploration Systems. Before starting her Masters, she worked as a trainee patent attorney for Dehns in Oxford for two years, experiencing work with both industrial and academic clients. During the SEEDS project, she was part of the greenhouse, communications and science teams, including working on the landing site selection. She was also part of the team that carried out radiation protection modelling for the outpost. She was also the project manager during the Leicester phase of the project, and developed a technology roadmap for the mission.

LORENZO MEGHINI Lorenzo Meghini was born in Terni, Italy in 1988. He graduated with a MSc in Energetic and Nuclear Engineering from the Politecnico Di Torino in 2014, with a final project focusing on the fuel cell technology developed at the University of California, Irvine. He decided to join the SEEDS program because of its attractive and ambitious international plan in the world of the space exploration. During the project, Lorenzo worked in several systems related to the support of the human life such the air revitalization, the fire detection and suppression and the atmosphere management systems as well as participating to the mission analysis identifying the launch windows. He was involved in the design of thermal and power systems. He was also part of the ascent team and participated in the system engineering consolidation effort. GERARD MORENO-TORRES BERTRAN Gerard Moreno-Torres Bertran was born in the year 1990 in Barcelona, Spain. In year 2012, he graduated with a bachelor degree in Physics from the University of Barcelona. Gerard joined the Physics and Astronomy Department of the University of Leicester, UK in 2014 to read the MSc in Space Exploration and Development Systems (SEEDS). Prior to enrolling to his Masters Degree, he has gained his work experience as a software engineer at Taitus (2012- 2013) as well as Proeng (2013-2014). During the SEEDS project, Gerard was part of the Habitable Module, Entry, Descent and Landing (EDL) and Science teams, including taking the role of System Engineer for the Habitable module.

JACOPO PISACRETA Jacopo Pisacreta was born in Rome, Italy in the year 1985. After having completed his M.Sc. in Astronautic Engineering from the La Sapienza University of Rome in 2013, with a thesis dealing with the engineering and economic issues related to the satellite servicing, he joined the VII SEEDS Program to deepen his knowledge of the Space Sector and gain more experience in the Space Exploration field. Apart from being a space enthusiast, Jacopo is also interested in travelling and he also an eager reader. During the SEEDS Project Work, Jacopo was part of the Greenhouse, Mission and Risk Analysis and Science teams. He took the role of Project Manager during the Toulouse phase, following the development and status of the work packages.

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SILVIO SILVESTRELLI Sivio Silvestrelli was born in Moscow, Russia in 1990. He obtained his Master Degree in Aerospace engineering from the Politecnico di Torino in 2014, performing his master thesis during the Erasmus project at Universidad Politécnica de Madrid concerning about a structural analysis using finite elements of a fuselage section in composite material. During the SEEDS project, he was part of the Habitable module team, analysing the inflatable technology and defining and sizing the structure of the overall modules. He also took the role of System Engineer for the Communications team, defining the link budget. He was also part of the ISRU team, defining and sizing the ISRU and the rovers and on the Power team, working on the power distribution system, defining its architecture and its mass. MARCO VOLPONI Marco Volponi was born in 1989 in Genova, Italy, and raised in Chiavari, a beautiful town on the Mediterranean Sea. He obtained his Master degree from the Università Degli Studi Di Genova in Theoretical Physics in the year 2013. For his master thesis, he worked on SCET (an EFT of QCD) at the National Institute for Subatomic Physics in Amsterdam, The Netherlands. Marco has wide interests, particularly in active radiation protection systems as well as bioregenerative life support systems, such as greenhouse. During the Turin phase, he was the System Engineer of the Greenhouse, analysing the diets and the lighting to minimise the surface, defining the global layout, mass and power budget. In Toulouse, Marco was the System Engineer of the Entry, Descent & Landing, focusing also on the code simulation. In Leicester he worked as Science Manager, organizing the science experiments matrix and coordinating the science team. AMEER ZAILANI Ameer Zailani was born in 1992 in a small kingdom in the South East Asian region, Brunei Darussalam. In 2011, he went to the UK to pursue his Bachelor studies in Aerospace Engineering at the University of Leicester under the scholarship of Brunei Darussalam Government. For his undergraduate thesis, he was working on the control systems of a Quadcopter. Upon his graduation in 2014, he joined the Physics and Astronomy Department at the University of Leicester to study the MSc in Space Exploration and Development Systems. Ameer has previously work as a trainee student for the Royal Brunei Technical Services in 2013. Apart from his engineering background, Ameer has always been interested in travelling. For SEEDS 2015, Ameer has been working under the structural team at which he was involved in sizing and designing the habitable structure. In addition to that, he was also involved in the design of the Extra Vehicular Activities module. In Leicester phase, he was a part of the Power System team where he was studying on the distribution of the power for the MILESTONE outpost.

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2 MISSION SUMMARY

2.1 MISSION STATEMENT The mission statement for MILESTONE is as follows: “To descend humans to the Martian surface; to ensure their survival through the establishment of a long permanence outpost; to conduct in situ scientific operations and exploration; and to allow for their ascent from the Martian surface”. The mission statement summarises the key objectives, intended outcomes and the underlying motivations for the mission. There is special attention to the human aspect of the expedition; ensuring crew survival and safe return to Earth. This mission will also serve as a starting point for the further exploration or possible colonisation of Mars, and so the long permanence outpost is of major importance.

2.2 MISSION OBJECTIVES The mission objectives show the information contained in the mission statement in a schematic way, making it easier to recognise the problems and solutions that the system may present. The mission objectives are as follows:  To safely descend humans on the Martian surface.

 To safely descend cargo on the Martian surface.

 To conduct in-situ scientific operations.

 To conduct in-situ exploration.

 To support human life and provide a habitable environment.

 To safely ascend humans from the Martian surface.

 To establish a long permanence outpost for future human exploration.

These objectives are seen to be of utmost importance for the mission, presenting distinct targets as to what must be achieved.

2.3 MISSION ARCHITECTURE In the preliminary phase, a mission architecture was determined. Mission MILESTONE will be a split mission, with the cargo and crew being sent to Mars separately. In the first stage, separate cargo elements will be descended to Mars and assembled autonomously to form the outpost over a period of two years. With the cargo phase complete, a crew of six will arrive in a Martian Descent Vehicle (MDV) deployed from Low Martian Orbit (LMO) to enter and commission the outpost. The stay on the surface will last 60 days which will initially be spent preparing the outpost for a subsequent 500 day mission, after which there will be the capability to perform in situ exploration and scientific tasks. In addition, four static landers will be deployed from the CIV in order to perform scientific operations. All resources required for the sustenance of the initial 60 day crew will be brought from Earth, with follow up missions being supported by In-Situ Resource Utilisation (ISRU) to produce water and oxygen and a greenhouse for additional food production. The crew will put the outpost into an unmanned state and return to LMO using a Martian Ascent Vehicle (MAV).

2.4 MISSION ASSUMPTIONS Mission MILESTONE by design builds on the success of previous missions and facilitates future missions. For Mission MILESTONE to achieve its goals, a number of assumptions must be made as to the work done prior to launch.

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It is assumed that the mission will take place following Mission ORPHEUS, in which humans were landed on the surface of Phobos and a number of exploratory rovers were deployed to the surface of Mars. It is therefore assumed that these rovers are still active on the surface, and can be used during Mission MILESTONE. Mission MILESTONE will re-use the Crew Interplanetary Vehicle (CIV) deployed in ORPHEUS, and so the crew transfer from Earth to LMO will not be considered. It is assumed that all modules will be launched with either a super heavy class or heavy class launcher. This will ensure launch feasibility without detailed analysis of the transfer. It is assumed that MILESTONE will be successful in its set up of the greenhouse and ISRU plant and future missions will be able to utilise these resources. It is assumed that MILESTONE will be landed at a site hazardous to microbes brought from Earth for planetary protection reasons. It is assumed that due to technology developments, the technologies stated in Section 17 will have technology readiness level 8 or 9 by the start of Mission MILESTONE in 2035.

2.5 TRAFFIC MODEL Mission Mission Element Landed Mass EDL Mass (per Launch (per module) (t) module) (t) Window/Class

Cargo Habitable Module (x2) 29.4 12.3 W1-W1/SHC (x2)

Laboratory 29.7 12.3 W1/SHC

ISRU 7.5 6.6 W2/HC

Power 25.2 11.3 W1/SHC

Greenhouse (x2) 28.9 12.2 W1-W2/SHC (x2)

Node 7.6 6.6 W2/HC

EVA Module (x2) 7.8 6.6 W1-W2/HC (x2)

Rovers (x2) 7.8 6.6 W1-W1/HC (x2)

Communication Satellite 0.081 - W1/HC (x3)

Mars Ascent Vehicle 26.7 11.8 W1/SHC (Capsule)

Ascent Module (Airlock) 5.5 6.1 W1/HC

Total 250 130

Crew Consumables (journey) 46.2 -

Mars Descent Vehicle 11.2 7.6

Total 57.4 7.6

Overall Mission 307 138

Table 1 - Mass estimates for mission modules. (The CIV will neither be launched nor landed, as it will be refuelled and reused from Mission Orpheus (1). SHC: Super Heavy Class; HC: Heavy Class)

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Mission MILESTONE is divided in two main phases: Cargo and Crew. The cargo phase is subsequently divided in two launch windows, in September 2039 and in October 2041. The crew mission is planned to launch during November 2042. For the cargo mission each module will be sent separately towards Mars, delivering a total mass of 250 tons to the Martian surface. A breakdown of this mass can be found in Appendix A: Mass Budgets. A total number of 15 launches to Low Mars Orbit (LMO) are required, with one further launch for the Communication Satellites to be deployed into Mars Aerostationary Orbit. The launches will be carried out using two classes of launchers:  Super Heavy lift class: e.g. SLS 2B – ca. 45 tons in LMO (2)  Heavy lift class: e.g. Falcon Heavy – ca. 14 tons in LMO (3) Therefore, due to the large number of launches needed, a sustainable launch campaign strategy has been designed. The strategy for the Cargo phase requires simultaneous launches to be performed during the two planned launch windows. In order to achieve a sustainable campaign, it is assumed that an international collaboration would be fully available, and at least three of the main Space Agencies (e.g. NASA, ROSCOSMOS, and CNSA) will have developed a Super Heavy lift class launcher. Planning for one launch every six months per launch pad, two launch pads would need to be built at each launch site in order to provide the number of launches needed. This allows for six super heavy lift class launches per launch window. For the crew mission, the CIV developed by Mission Orpheus will be reused. It will need a full refuelling with propellant (methane and oxygen) (1) that will be carried out after the return of ORPHEUS in 2038, and following the same launch schedule used previously. The Mars Descent Vehicle is intended to be launched separately with the crew and any required consumables required for the Mars round-trip (around 46.2 t) inside. The maximum number of launches is carried out in the first launch window. The total number of launches for the cargo mission is 7 Super Heavy Class, with 6 in the first window and 1 in the second window, allowing five launch chances for contingency. There are also 7 Heavy Class launches, with five in the first window and two in the second window in a nominal condition. The total number of launches for the Cargo Mission is 15.

2.6 CREW SELECTION

2.6.1 Introduction The selection of the crew is a critical issue that has to be made years in advance of the launch date of the mission. Humans are the most valuable asset for Mission MILESTONE and as a result, analysis of the crew’s physical and psychological health must be performed. For Mission MILESTONE 6 crew members are planned to land on the Martian surface to complete the mission objectives. The roles of the crew members can be selected based on the objectives that the mission has to accomplish, which are:  Complete the assembly of the outpost  Demonstrate the feasibility of in-situ food production and resource utilization  Perform scientific activities and retrieve samples  Explore the Martian surface  Safely descend to and ascend from the Martian surface

2.6.2 Crew Training and Mission Roles While all astronauts should be briefed in all relevant emergency procedures, and will have sufficient operating knowledge to carry out scientific experiments. It is assumed that the training of each astronaut will focus on two areas of expertise, a primary role and a secondary role. By briefing all crew members in assembly procedures and spacecraft operations, a Fail-Operational/Fail-Safe strategy is achieved (1). The area of expertise is required in the three following areas:  Command, control, and vehicle and facility operations and functions (4)

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 Scientific exploration and analysis (4)  Habitability tasks (4) One crew member will be a pilot due to the tasks involved in transit, ascent and descent (5). The role of commander will be assigned to the crew member with the most experience and will not necessarily be the pilot, as has been the case in many space missions (5). The most relevant fields of expertise that are required due to scientific and habitability tasks include mechanical engineering, electrical engineering, geology/planetary science and medicine/psychology (4). Nevertheless, the whole crew has to be trained on the procedures that are required by the assembly phases of the outpost, comprising extravehicular activities (EVA) and intra-vehicular activities (IVA). From analysing the objectives and tasks involved in Mission MILESTONE the following necessary primary and secondary roles are listed in Table 2 (4): Primary Roles Secondary Roles Mission Commander Co-Commander Pilot Co-Pilot Physician Co-Physician Mechanical Engineer Psychologist Geological Scientist Planetary Protection Specialist Electrical Engineer EVA Specialist Table 2 - Crew member roles

2.6.3 Selection Criteria Unique to Mars Missions Due to the unique characteristics of such an extreme mission, space agencies will need to employ some less conventional selection criteria such as genetic screening and precautionary surgery (5). Once a crew member is pronounced genetically free of predisposition to future disease or disorder (for example an increased risk of developing cancer) and is provisionally selected for the crew of the Mars mission, precautionary surgery may be required (5). This is to avoid serious medical emergencies that may result in the failure of Mission MILESTONE; for example the appendix will be removed due to the risk and consequences of appendicitis (5).

2.7 LANDING SITE An important part of the design of Mission MILESTONE is the selection of the landing site for the Outpost and the scientific landers. The scientific and engineering constraints are more rigorous for the Outpost than for the landers due to the high mass and area needed. The key constraints considered when selecting a landing site were:  Elevation: in order to land the modules, it is necessary to have an elevation of at most -3000m (see Section 11).  Latitude: due to delta-v considerations for the CIV (as specified by Mission ORPHEUS (1)), a maximum latitude of ±30° was determined.  Landing area: the combined landing ellipse for all the modules of Mission MILESTONE is approximately 36 km by 56 km, with a desired radius of exploration of 75 km, therefore there must be approximately this area of ground to easily traverse.  Indications of water: due to the need ice for in-situ production of water (see Section 16), it is necessary to land in a region containing sufficient sub-surface ice. Nine landing sites were considered, including three selected by Mission ORPHEUS for their rovers (1). The majority of the sites identified are considered plains, in order to increase the ease of landing and exploration for the crew members. From these, a final five were selected to provide landing sites for the Outpost itself and the four scientific landers. These can be seen in Figure 2.

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Figure 2 - Map of MILESTONE landing sites. From left to right: Amazonis Planitia, Candor Chasma, Chryse Planitia, Isidis Planitia and Elysium Planitia. Image adapted from MOLA topographic map (6)

2.7.1 Amazonis Planitia (15N, 155W) Amazonis Planitia is a plain to the west of Olympus Mons, with a landing region centred on 15N and 155W being selected for the outpost. The elevation is approximately -3500m, and it is a relatively smooth and crater-less surface (from HiRISE images (7)), satisfying the landing criteria. There is evidence of water in Amazonis Planitia (8), with measurements from the Gamma Ray Spectrometer (GRS) on Mars Odyssey placing the minimum water content at 6% in the vicinity of the landing site (9). It is thought to be a Noachian impact basin with Hesperian lava flow, with the most recent surface being Amazonian (8), and therefore would provide an area of geological interest. The history of volcanic activity may have caused fossils or extant life to have been pushed to nearer the surface and to a potentially accessible depth (8).

2.7.2 Isidis Planitia (12N, 87E) Isidis Planitia was the first landing site selected for the scientific landers. This is a fairly cratered (7), Noachian plain, with 2-17% rocks (10). It has a low water percentage of approximately 3% from GRS data (9), which made it unsuitable as a primary landing site. However, its geology makes it an area with good exobiological potential (10), and therefore an interesting landing site for a lander.

2.7.3 Chryse Planitia (27N, 40W) Chryse Planitia was also selected for the scientific landers. The location of the Viking lander, it is thought to have lakes from the late Hesperian/early Amazonian period. It is thought to have shown water erosion in the past, with degraded wrinkle ridges and craters indicating the presence of liquid water in the upper crust (11). Together with the presence of fine grained sediments (11), this was determined to be an area of interest.

2.7.4 Elysium Planitia (3N, 155E) Elysium Planitia has a minimum elevation of -3000m, and therefore was determined to be slightly shallow for the landing of the Outpost. However, the detection of methane in the atmosphere over Elysium Planitia by the Planetary Fourier Spectrometer on Mars Express (12) may indicate the presence of biological life, and there is thought to be ash-covered water ice or even a frozen sea, making it an area of interest to Mission MILESTONE and therefore a landing site for a scientific lander.

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2.7.5 Candor Chasma (7S, 71W) Candor Chasma was the final landing site selected for the scientific landers. It was the only landing site chosen to be a chasm, rather than a plain, and also the only in the southern hemisphere, although it has a near-equatorial latitude. It was chosen due to the presence of signs of water, including injectite megapipes and water-lain sediments seen from HiRISE (13). Recurring slope lineae (RSL) are also known in deep parts of the chasm (13), which may be indicative of water presence. These features made it a region worth scientific exploration despite being unsuitable for landing the main Outpost of Mission MILESTONE.

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3 MISSION ANALYSIS

3.1 MISSION PHASES AND SCENARIOS Mission MILESTONE phases can be divided following the different launch campaigns dedicated to land the cargo and the crew on the Martian surface. Considering the main objective of the mission, the phases start whilst one of the module types, either the cargo or the CIV, is orbiting in LMO.

3.1.1 Cargo Phases and Timeline Each module of the cargo phase is launched separately into Low Mars Orbit at an altitude of 500km. As the Earth-Mars transit is not considered, the feasibility of transporting each module was determined by the launcher limitations discussed in Section 2.5. The orbital stage will be followed by the descent of the cargo module (e.g. the habitable module and its entry descent and landing system). Aerobraking will be used to reduce the mass of propellant required for this phase (see Section 11).

Figure 3 - Cargo mission phases

The cargo mission timeline can be seen in Figure 4. Each part of the cargo phase is critical, and therefore is indicated in red. This is because each step cannot begin until the previous has been completed, for example the transport rovers will not be activated to collect cargo until the integrity check has been completed and the condition of the modules is known. As can be seen in Figure 4, the majority of the time is taken up with the collection of cargo. This timeline is a worst case scenario, and assumes that all of the cargo is only being collected after the second cargo mission, which allows 125 days before the crew mission.

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Cargo mission timeline

Collection of cargo

Activation of transport rovers

Modules’ integrity check

Descent on Mars surface Phases

Orbiting in Low Mars Orbit

0 20 40 60 80 100 120

Time [days] Figure 4 - Cargo mission timeline

3.1.2 Crew Phases and Timeline The major components of the crew phase are the CIV and the Mars Descent Vehicle (MDV), which are assumed to orbit Mars at an altitude of 500 km (i.e. Low Mars Orbit (LMO)). Following a short orbiting phase, the crew (in the MDV) will then undock from the CIV and descend to the Martian surface. The CIV will remain in LMO, where station keeping will maintain its orbit for the return of the CIV. The crew land adjacent to the cargo collection location, with a landing area precision of 19 km by 6 km.

Figure 5 - Crew mission phases

Figure 6 shows the crew mission timeline. In following with the cargo mission timeline this is a worst case scenario, with any variable events (such as the ascent of crew) taking their longest possible duration. This is because the critical (red) events were used to determine the length of time available for the scientific activities and exploration. These activities have been indicated as non-critical in this timeline, as while it is critical that these activities take place, there is a fixed end time for completion and any delays in start time will instead cause a reduced length of time for these operations. Non-critical in this timeline does therefore not mean that they are not critical to the success of the mission, but rather that they have a fixed end point. Due to the time required both for outpost set up (the first six activities) and for Mars departure (the final five

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activities), a total of 47 days are available for scientific activities and exploration. The division of this time into exploration and experimentation can be seen further in Section 4.4. Crew mission timeline Separation of ascent module Dock with orbiter Orbiting in LMO Ascent of crew Ascent preparation

Phases Scientific Activities Exploration of relevant sites Activation of ISRU and Greenhouse Connection of Modules Connection of HAB and POW Travel to base location Descent of the crew Undock from orbiter Orbiting in LMO 0.0 10.0 20.0 30.0 40.0 50.0 60.0

Time [days] Figure 6 - Crew mission timeline

3.2 RISK ANALYSIS The risk analysis is a study of the different scenarios, in order to evaluate their criticality and the possible contingencies that can occur in case of failure detection. With the creation of a parameter, Time To Recover the Damage (TTRD), each scenario has been evaluated identifying the influence of the related phase being accomplished. TTRD can be defined as the product of the criticality level of the damage, the possible time required to recover the damage and the total time allowable for the recovery, which corresponds to the duration of the scenario itself. The time to recover takes four values, with one representing very fast (minutes) to four representing very slow (days). The scenarios were also given a value for the duration of the incident, from one to three representing long (scenario of a duration of a month or more) to short (scenario of approximately a day in length). Finally, the criticality was given a level of one to four, representing minor to crucial levels, as can be seen in Table 3. Criticality Value Description Level Any condition/event that may cause damage to hardware other than Minor 1 conditions previously stated Any condition/event that may cause major to an damage to an Major 2 emergency system, an element not in a critical path, minor personnel injury or minor occupational illness Any condition/event that may cause personnel injury or sever Severe 3 occupational illness Any condition/event that may cause the loss of a module, life sustaining Crucial 4 function or emergency system Table 3 - Criticality levels definition (14)

The TTRD is evaluated as: 푇푇푅퐷 = 퐶푟푖푡푖푐푎푙푖푡푦 ∙ 푇푇푅 ∙ 푑푢푟푎푡푖표푛 (1)

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Then for simplicity, every value of TTRD can be normalised to the maximum value of TTRD equal to 80, which corresponds to the worst case of having a crucial contingency with a very long time to repair in a short duration scenario.

Figure 7 - TTRD levels evaluation

The situations have been divided in three categories according to their TTRD and defined by the impact that they would have on the success of the phase. An "unchanged" situation has a value of 0-0.1125, with detection of damage not affecting functionality. A "degraded" situation has a value of 0.15-3, with detection of damage affecting functionality, but degraded functionality still possible. Finally, a "prejudiced" situation is one with a value of 0.375-1, in which detection of the damage prejudices the functionality of the system.

3.3 CONTINGENCY STRATEGY All critical scenarios that need a crucial recovery intervention have been analysed, including the creation of a contingency strategy. Figure 8 shows this strategy, with Table 4 and Table 5 showing the detailed analysis. Approach Procedure

Automatic software work around: the recovery a) Automatic re-run: the software re-runs, approach tries to solve the damage with a software refreshing the system. intervention (continue nominally) b) Earth override and reprogramming: new commands are sent from Earth, reprogramming the software in order to detect and solve the failure. Redundancy: if the damage persists, the system Isolate failure and activate redundant branch: has to switch to the redundant channel, so a general failure detection, isolation and recovery command is sent from Earth (continue in a approach is applied in order to identify, isolate and degraded manner) recover the failure using the redundant channel of the system. Degraded: if the damage persists, the system has Avoid ripple effect: to continue to work, accepting the loss of the the acceptance of a loss of function must not affect function with the damage (continue in a degraded the functionality of the overall system, especially manner) related to the human support systems. Table 4 - Cargo Phase Recovery Approach

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Approach Procedure

Automatic software work around a) Automatic re-run b) Earth override and reprogramming Redundancy Isolate damage and activate redundant branch Degraded Avoid ripple effect and consequences on human health Recovery: if the damage persists, the crew has to a) Active crewed repair process: the crew follows act directly on the recovery procedure. With the direct the repair procedure in order to solve the intervention, the crew can operate until the damage. completely solution of the damage (continue b) Substitution of failed component: in case of nominally) irreparable components, the crew can substitute it using spares. Table 5 - Crew Phase Recovery Approach

In the cargo mission, if the damage cannot be fixed it is possible to assume the loss of the module, requiring a new launch to replace it. In the crew mission, if the damage persists, the crew is required to start the escape procedure, entering into the emergency scenario. This consists of all the activities the crew shall perform in order to escape safely from the surface in the shortest time possible.

Figure 8 - Contingency Procedures

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4 SCIENTIFIC OBJECTIVES AND OPERATIONS Humans are the crucial to MILESTONE’s scientific investigations. The biological effects of the Martian outpost environment on the human body is of particular interest, in order to help develop further longer-term missions. In the long-term, this science enables the possibility of a future colonisation of Mars, and for comparing the differences between the same analysis performed on Earth, in LEO and in deep space. Another advantage of having humans on the planet is the ability of human beings to discern interesting, unusual and/or unexpected features among a great variety of similar samples. This ability is very difficult to achieve remotely on robotic systems. This input from humans will both improve the quality and quantity of scientific data returned. Last but not the least, sending humans allows a great variety of scientific instruments to be used, instead of relying only on robotics, and therefore improves the possibility of performing analysis by adding increased mission flexibility.

4.1.1 Science Objectives 1. To study the physiological and psychological impact of a Mars surface mission - All of these experiments are meant to improve the understanding of the biological, physiological and physical variation and adaptation to the different gravitational and radiation environments of various forms of life. Spanning from mono-cellular bacteria to plants, which are very important to improve the capability of growing food directly on Mars, to humans. Also, from a psychological perspective, it will be of interest to assess how a crew of six people will react upon living together for an extended time, in an extreme and remote environment like the surface of Mars. 2. To perform in-situ investigations into the Martian environment to support future exploration - Measurements are planned in order to better understand the radiation exposure on the surface of Mars, the residual magnetic field strength and its origin, and the Martian weather and atmosphere, so as to facilitate subsequent missions 3. To collect data to support the identification of landing sites for future human missions - The determination of the composition of Martian surface and subsurface, with a special focus on the presence of water, will help expand the possibilities and capabilities of the human outpost in the future 4. To study the origins and evolution of Mars - The combined study of the soil composition, present and past meteor history, the presence (or not) of life, can help us understand the origin and evolution of the planet, and ultimately of the entire solar system. In order to achieve these goals, a number of experiments have been proposed. One of the main goals of Mission MILESTONE is to set up a long-term outpost that will be used for future long duration missions. As a result the experiments have been divided in three different categories: experiments which will be performed in the 60 day mission and will be continued in the subsequent long-stay missions; the autonomous experiments, which will be set up during the 60 days crew permanence but will run independently; and experiments that are intended for future long-stay missions. The equipment for future missions is included in the Laboratory Module (Section 4.3.1).

4.2 SCIENCE MATRIX BREAKDOWN In order to understand which experiments should be carried out to meet our objectives, and with which instruments, a science requirement matrix has been developed (see Appendix C: Science Requirements Matrix). From the previous main scientific objectives, a list of experiments has been derived, passing through a breakdown process which consists of:  Top Level Scientific Objective → Detailed Science Objective → Science Investigations → Measurement Objective → Measurement Methods

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The full science requirements matrix can be found in Appendix C: Science Requirements Matrix, and an example row can be seen in Table 6. Top Level Detailed Science Science Measurement Measurement Science Objective Investigations Objective Methods Objective To collect data to To study the To study quantity Detect the 5m Mole support the composition of the and quality of presence of sub- identification of Martian sub- water within the surface ice within resources for surface to identify Martian sub- the area of 5m Drill future human resources surface. interest. missions. necessary to support human Mole-head gamma Mars missions. ray spectrometer

Table 6 - Example of the science requirements matrix breakdown

4.3 SCIENTIFIC ARCHITECTURE To maximise the scientific return of Mission MILESTONE a multi-element architecture has been chosen. This consists of a laboratory module, four static landers, droppable atmospheric probes, scientific rover payloads and human deployable assets.

4.3.1 Laboratory Module In accordance with the standardised outpost module (which will be explained in Section 6.2), the laboratory module (LAB) is divided into a rigid part as well as an inflatable part (Figure 9). The rigid part has two bulkheads forming the decontamination area; which is located between the area dedicated to the non- hazardous experiments and the area with the experiments that have been exposed to the Martian environment. This will improve the safety of performing experiments on Martian samples and it will decrease the risk of both forwards and backwards contamination. A dedicated sample delivery and return is designed in order to support the experiments directly coming from the Martian environment.

Figure 9 - Laboratory Module

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4.3.2 Static Landers Prior to the descent of the crew in the Mars Descent Vehicle, the CIV will deploy four static landers, on the Martian surface in locations of scientific interest. The landers will carry out complimentary science experiments for the characterisation of possible future landing or exploration sites. The selected locations are shown in Table 7 and were shortlisted as candidates for the outpost landing site, as discussed in Section 2.7. Candor Chasma Elysium Planitia Isidis Planitia Chryse Planitia Latitude 6.6 S 3.0 N 12 N 26.7 Longitude 71 W 154.7 E 87 E 40 W Table 7 - Landing site of the static landers

The static landers are all of equal size and have the same scientific payloads. The scientific payload on the static landers is composed of the instruments stated in Table 8. Instrument Measurement Objective Reference Instrument Gamma Ray Neutron Elemental composition Mercury Gamma Ray and Spectrometer Neutron Spectrometer (MGNS) (15) Mole Head Gamma Ray Sub-surface ice detection, elemental Instrument proposed by Spectrometer composition depth profiling Ambrosi et al. (16) (17) Sub-Surface Seismometer Seismic characterisation Instrument being developed by Microdevices Laboratory, Jet Propulsion Laboratory (18) Environmental Monitoring Suite Atmospheric monitoring REMS (Mars Science (Wind/Temperature/Pressure/U Laboratory) (19) V/Humidity) Mole (5m) Excavation and sample retrieval Moon/Mars Underground Mole (MMUM) (20) Laser Induced Breakdown Ranged elemental composition ChemCam (Mars Science Spectrometer (LIBS) Laboratory) (21) Rock Abrasion Tool Sample homogenisation Rock Abrasion Tool (RAT) Mars Exploration Rover (MER) (22) Alpha Particle X-Ray Elemental composition Rosetta Lander, Philae, Alpha Fluorescence Spectrometer Particle X-Ray Spectrometer (APXS) (23) Raman Spectrometer Elemental composition and detection ExoMars RLS (24) of organic bio-markers Magnetometer Magnetic environment Smart Digital Magnetometer characterisation HMR2300, Honeywell (25) Robotic Arm Sample retrieval Phoenix Lander (26) Panoramic camera Local environment physical ExoMars PanCam (27) characterisation Radiation assessment detector Radiation environment RAD (Mars Science (RAD) characterisation Laboratory) (28) Table 8 - Breakdown of the scientific payload on the static landers

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The overall parameters of the static landers are described in Table 9. Science Payload Mass (kg) Landed Mass (kg) Maximum Power (W) 34.2 162.7 142.8 Table 9 - Mass and power breakdown of the static landers

4.3.3 Droppable atmospheric probe The entry, descent and landing system designed for this mission is largely dependent on the Martian atmospheric density, as most of the deceleration is provided by the Hypersonic Inflatable Aerodynamic Decelerator (HIAD). It is necessary to have an accurate value of the density in order to reduce the landing ellipse. To achieve this a droppable atmospheric probe will be used. Every module will be fitted with one of these probes, which will descend before the module, in order to sound the atmosphere and give back the density profile to the cargo module before its descent. This probe has been designed to follow the same trajectory as the cargo modules. This has been done by using a HIAD scaled in proportion to the probe mass. This was tested in the trajectory calculator used for the descent, as discussed in Section 11. The probe is equipped with:  An accelerometer that will be used to calculate the acceleration over the decent, which will allow the calculation of the density  Data storage  A Ka-band antenna to send the recorded data to the communication satellite  A regenerative fuel cell to provide power  A HIAD The possibility of having a blackout during the decent was considered. However it was concluded that the chances of this happening are low. Due to the design of the decent module, the speed of the probe will never go above 3.5 km/s, which is below the threshold at which the electronic density of the surroundings will disrupt communications on the Ka band. Once the density profile is known it will be possible to choose the correct starting point for the module to begin its descent.

4.3.4 Rover Scientific Payload In order to meet the science goals of Mission MILESTONE, and reduce the science mass and power requirement, three different payloads have been designed for the rover: one permanent, and two specific payloads for different science objectives. The equipment for the specific payloads can be removed by the crew members during the crew mission to meet the needs of specific scientific objectives.

4.3.4.1 Permanent Payload The permanent payload for the rover will provide scientific data for the lifetime of the rover, and includes a number of sensors primarily aimed at radiation and magnetic monitoring, which will permanently be active during both the cargo and crew missions. This includes a Cherenkov detector and a Geiger counter for radiation monitoring, and a magnetometer for mapping the magnetic field. While the Geiger counter will be able to provide a count rate for alpha, beta, gamma and x-ray radiation hits on the rover, the Cherenkov detector will be able to provide more detailed information about particle energy, to help the abundance of solar particles and galactic cosmic rays to be better mapped (29). The data collected by these sensors can be transmitted from the rover to the Communication Satellite during the cargo mission, and to the Outpost Control Centre during the crew mission, allowing it to be analysed to provide a detailed map of the radiation and magnetic field in the landing area. The permanent payload has a mass of 4 kg and requires a power of 13.5 W, making it suitable to be permanently affixed on the rover.

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4.3.4.2 Planetary Science Payload The first specific payload for the rover is designed to enable the astronauts to achieve planetary science and geological objectives. The details of this can be seen in Table 10. Instrument Measurement Objective Reference Instrument Laser Induced Breakdown Ranged elemental composition ChemCam (Mars Science Spectrometer (LIBS) Laboratory) (21)

Alpha Particle X-Ray Elemental composition Rosetta Lander, Philae, Alpha Fluorescence Spectrometer Particle X-Ray Spectrometer (APXS) (23) Next Generation Sample Analysis of trace gases, chemical Sample Analysis at Mars Analysis at Mars and isotopic composition of Instrument Suite (Mars Science (GC/MS/Tunable Laser atmosphere and samples Laboratory) (30) Spectrometer) X-ray diffractometer/fluorescence Elemental composition CheMin (Mars Science spectrometer Laboratory) (31) Drill head gamma ray Sub-surface ice detection, Instrument proposed by Ambrosi spectrometer elemental composition depth et al. (16) (17) profiling Drill (5m) Excavation and sample retrieval ExoMars Drill (32) Rock Abrasion Tool Sample homogenisation Rock Abrasion Tool (RAT), Mars Exploration Rover (MER) (22) Table 10 - The planetary science payload for the rover

The planetary science payload for the rover has a mass of 120kg and a power of 1.3kW.

4.3.4.3 Environmental Measurements Payload The second specific payload is the environmental measurement equipment. This includes pressure, temperature and humidity sensors, together with a sonic anemometer for measuring wind speed. The environmental measurement payload can be attached to the rover by the crew members, and may be used constantly during the crew mission or alternatively may be used only when an EVA is being carried out. The combination of equipment will allow for the weather patterns around the Outpost to be better tracked and understood, in order to gain a greater understanding of Martian weather. In addition, it can help to predict weather fronts, which may be necessary when planning crew EVAs. The payload has a total mass of 1.6 kg and a power requirement of 5.5 W, and therefore again is suitable for near-constant use on the rover.

4.3.5 Human Deployable Assets

4.3.5.1 Mini-Greenhouse The structure of the mini-greenhouse structure consists of two parts: an aluminium cylindrical base of 0.5 mm thickness, designed to house the subsystems related to the internal atmosphere control, and a polycarbonate dome of 3 mm thickness for atmosphere containment and UV protection (Figure 10). The mini-greenhouse contains a water regenerative system which uses a water tank of 1.1l and guarantees a buffer for 15 days of operation. The power supply relies on solar arrays that provide 40 W. In the base of the greenhouse there is a CO2 tank which ensures 15 days of supply. The mini-greenhouse also includes various sensors to monitor the sprouting and growth of the plants. It contains a pressure sensor, a humidity sensor, a thermometer, a Geiger counter, a chlorophyll measurer and a camera to film all the stages of the plant development. The mini-greenhouse has a total mass of 58 kg.

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Figure 10 - Mini-Greenhouse

4.3.5.2 Atmosphere and radiation monitoring The other Human Deployable Assets will mainly monitor the atmosphere, studying its composition, temperature, pressure, humidity, together with its variation during the days and the season. The sensors will also collect data about the radiation environment, both looking at the sky and the surface. The atmospheric sensors are based on a LIDAR instrument, similar to that on the Phoenix lander (33) . The radiation monitoring assets are based on multiple instruments, spanning from Geiger counters to Cherenkov telescopes.

4.4 SCIENTIFIC TIMELINE In order to assess which experiments will be performed during Mission MILESTONE and to what extent, a preliminary timeline was produced. The main experiments to be performed were selected and their equipment gathered, then the time needed to complete each analysis was estimated and divided in to overhead time, time per single acquisition and number of acquisitions needed. The total analysis time can be seen in Table 11. The experiments were then divided into two categories, based on the total time needed; the short ones (i.e. less than two hours), which comprise the use of IR Spectrometry, X-Ray Fluorescence and Multi Laser Raman Spectrometry, which should be used on all the samples collected as a preliminary investigation. The long duration experiments (over two hours in duration), which should be performed only on the interesting samples, have been assumed to be carried out on only 20% of the samples. In addition, to more accurately estimate the time required, a concurrency factor (c.f.) has been applied to the total acquisition time. This considers that a crew member, during the time that the equipment takes to analyse a sample, can begin a second analysis. The average time to analyse a sample was calculated to be 472 minutes. Considering two crew members will be fully assigned to the laboratory for ~40 days of the MILESTONE mission, the total number of samples that can be analysed is 81. Four to six EVAs with the rover shall be performed to collect these samples, each one collecting up to 10 samples (coring included). The remaining analysis capability shall focus on samples collected on EVAs around the outpost. Instrument Total analysis time (minutes) IR Spectrometer 130 X-ray Fluorescence 130 X-ray diffraction 180 Multi-laser Raman 130 Gas-Chromatography 240 Mass Spectroscopy 360 Antibody Array 180 Table 11 - Data for evaluating the timeline

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5 POWER

5.1 OUTPOST POWER CONSUMPTION The outpost power supply will provide power to the node, EVA modules, greenhouses, laboratory, habitable modules and ISRU plant. The rovers will have the ability to be recharged by the outpost but it will not be their primary power source, both the Mars Ascent Vehicle and Descent Vehicle will have their own power supply. A breakdown for the power requirements for each module can be found in Appendix B: Power Budgets. The outpost will operate on a duty cycle with three modes of operation when manned; nominal, minimal and survival. Nominal is the preferred mission mode, minimal is a reduced power mode that will still allow for the completion of Mission MILESTONEs objectives and survival which is a drastically reduced power setting that is only intended for keeping the crew alive so they can address the issue or escape. The power levels for each mode and the duty cycle are shown in Figure 11. Outpost Power Consumption 200 180 160 140 120 100 80

Power(kWe) 60 40 20 0 1 2 3 4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21 22 23 24 Hour of the day (Martian Hour) Nominal Minimal Unmanned Afer Crew Survival

Figure 11- Outpost Power Consumption

Mission MILESTONE will have a peak power in nominal manned mode of 175kWe, in minimal mode of 50kWe and in survival mode of 11kWe (all with margin). The duty cycle has also been prepared for the outposts operations during the period between two crews allowing for ISRU operations and food production. The unmanned period has a peak power of 135kWe.

5.2 ARCHITECTURE Power for the outpost will be provided by an 85kWe nuclear reactor and 19 units of solar arrays of 300m2 each, coupled with regenerative fuel cells. Each system is designed to be double failure tolerant in providing nominal power and in particular the solar power plant is sized to provide survival power in the event of dust storm at any point on the Martian year.

5.2.1 Nuclear The Nuclear reactor is a modular heat pipe based system (34) with a closed Brayton cycle power conversion unit. The reactor core will use uranium fuel rods surrounded by control drums to control the reactors criticality. The reactor will launch in a subcritical state and not be active until the crew arrive. The reactor is a fast spectrum reactor and will not require a moderator.

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Both heat transfer and heat rejection are performed with heat pipes. The heat pipes are a sealed system with low maintenance, modularity and inherent redundancy. Two back-up power conversion units are included to provide cold redundancy. To reduce the harmful radiation dosage that the crew may receive as a result of the reactors fission products the reactor will be positioned 100m from the outpost and employ a shadow shield. The shadow shield is a layered tungsten, to shield from gamma rays, and lithium hydride to shield from neutron leakage (35). In the area covered by the shield, radiation can be reduced to near Martian background levels. The area covered by the shield should be directed toward the outpost. In the unshielded region the radiation levels will be higher and crew time should be tightly controlled within 1km of the unshielded region (36). Future missions may think to cover the unshielded side of the reactor with Martian soil to provide adequate shielding (36). The reactor will produce 85kWe, weigh approximately 4.8 tons and its volume will likely be comparable to the design specifications of the SP-100, i.e. to fit in one third of the payload bay of the now discontinued spaces transport system shuttle (37).

5.2.2 Solar Array and Regenerative Fuel Cells In order to follow the duty cycle needs, solar arrays will be coupled to a storage system made up of regenerative fuel cells. A simple and reliable configuration of static solar arrays has been selected over sun tracking arrays. The solar panels orientation has been optimized for the outpost latitude so that the energy production is maximised at the least convenient period of the year. A mathematical model has been produced to calculate this. The panels point southward and the tilt angle is 34.5°. The solar panels are to be deployed before crew arrival either with the assistance of the rovers or with a self- deploying system. A dust removal system is also implemented in the structure (38) (39), thus a conservative and constant dust degradation factor has been applied. It is to be noted that the angular tilt of the panels, which was derived to favour the lowest period of available solar energy, will aid the dust removal operation. Regenerative fuel cells are used as an energy storage system to ensure a high density of energy and power. The regenerative fuel cells will use a supply of water which over time will experience losses and can be compensated for by water produced in situ. The issue of stack degradation with time, the cells reliability and their performance will need to be improved before a Martian mission. The mass of the regenerative fuel cells will be 810kg without margin, this was calculated using a NASA model.

5.2.3 Distribution The power system is comprised of the nuclear reactor, solar array and regenerative fuel cells. To manage the power production from each, the outpost has a power distribution system. The distribution system is modelled after the distribution system aboard the ISS (40). The distribution system has an efficiency of 76% for nuclear distribution and 74% for power distributed from the solar arrays. The overall mass of the distribution system is just over 8 tons (with margin) which is split between nearly 5 tons to be launched with the power system and 3 tons to be distributed amongst the modules.

5.2.4 Total power system launch mass The total mass of the power producing system to be launched (solar panels, regenerative fuel cells and nuclear reactor) is 20 tons with a 40% system level margin. The elements of the power distribution system which will be launched with the power system have a mass of 5 tons with 40% mass system margin. This distribution mass needs to be shipped as part of the power system, while the other components of the distribution system will be integrated in the respective modules. The overall volume to be fitted into the super heavy launcher fairing is about 130 m3 (without margins) (2), so this value is comfortably within the fairing constraints.

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6 MODULE DEFINING PARAMETERS

6.1 OUTPOST ARCHITECTURE

6.1.1 Outpost Layout As the outpost of Mission MILESTONE involves a number of modules, it is necessary to define a simple, safe and extendable layout for the outpost. A number of possible configurations for extension were considered from Bell, 2007 (41). Due to the shape of the modules (cylindrical with end caps for connection), the extension capability has been reduced to in-plane, i.e. without a possibility for vertical extension. This is also the simplest mode of extension. It has been determined that an approximately circular node which can be attached to four modules is the best for the outpost of Mission MILESTONE. This will allow the outpost to initially take the form of a cross with airlocks at the end caps of two of the modules, which can then be extended to form a closed loop with the addition of further modules in the future.

Figure 12 - Milestone outpost layout

The arrangement chosen, and shown in Figure 12 was done so with a view to a minimal configuration, i.e. the fewest launches possible from Earth. It was therefore decided that only a single node would be sent to form the outpost around. The laboratory has been chosen as one of the end modules (which are not connected to an EVA module) as this allows the dedicated contamination zone discussed in Section 4.3.1 to be at one end. The habitable modules have each been connected to an EVA module to provide two survival configurations, i.e. configurations in which the crew can be kept alive and able to reach the ascent vehicle. The greenhouse modules have been spatially separated for redundancy. However, this arrangement also ensures frequent passage of the crew members through the greenhouse, providing psychological benefits due to the interaction with growing plants.

6.2 MODULE PHYSICAL DESIGN Each module has a specific role and requirements. This section of the report looks at the mathematical models of the design that were applied to the main part of the modules. The generic structure of the module, the thermal control system, radiation protection, power considerations and planetary protection were analysed.

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6.2.1 Primary Structure The approach for designing the module’s structure has been driven by the limits on mass and volume given by the launcher fairing and the minimum volume required for a crew of six. As such, the structure of the modules incorporates both rigid and inflatable sections. The main characteristic of the inflatable structures is the possibility to have a greater volume than the one offered by traditional rigid metallic space modules with the same mass. The inflatable structure can be packed up to 60% of its inflated size in the longitudinal direction. Whilst the module is in the launcher fairing, the inflatable part is totally packed and all the subsystems of the module are mounted in the rigid part. Once the modules are landed on Mars and fully inflated, the subsystems are redistributed along the total module length and in the inflatable part the secondary structures (e.g. floor and ceiling) are mounted by the crew to create a living volume. In order to have a unified design and therefore increase the ease of production, the same structure definition is applied to all the main modules of the Mission MILESTONE. This approach was intended due to the industrial benefits that would be introduced by series production, and the subsequent cost reductions.

Table 12 shows the size of the module, considering both the rigid and inflatable part. Considering the fairing of an SLS 2B, which has a length of 19 m and a diameter of 8.5 m, the total length and the diameter allocated to the module are respectively 15.7 m and 4.5 m in order to have free volume for the Entry, Descent and Landing systems.

Length at launch 13 m Module diameter 4.5 m Rigid part length 11.2 m Inflatable length when fully inflated 4.5 m Inflatable length when packed 1.8 m Total module length 15.7 m Table 12 - Structure module sizing

The total volume allocated in the rigid part to all the subsystems, devices and equipment is 102 m3. The living volume (i.e. the remainder of the volume) when the module is fully inflated is 147 m3.

Figure 13 - Main modules structure definition

The primary structure sustains the primary loads, such as the launch loads, the landing loads and the pressure containment, and provides the external geometry of the module. The module’s primary structure

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consists of shell, rings and longerons and the end caps, which form the closing parts of the cylinder and the docking mechanism. The design of the primary structure is based on some general considerations:  The shell has to maintain the pressure of the module. In the rigid part, it is made of a layer of aluminium and a layer of insulation material. The inflatable part can be packed during launch and is made of several layers of Kevlar, Polyamide and Nomex for structural restraint, and layers of Beta Cloth, Kapton and Nextel for insulation. It will be inflated when the crew lands on Mars.  8 longerons are sized to resist to the launch and entry load of 10 Earth-g. For the rigid part, they are made in aluminium 6061 with a C-shaped cross-section of 32 cm2. Extendible longerons are present in the inflatable part to be deployed when the inflatable part is fully extended, and are sized in order to sustain the load of the structure on Mars.  The rings are sized to sustain the landing load of 3.5 Earth-g, and are made from aluminium 6061 with a double-T shaped cross-section of 30 cm2. The distance between one ring and the following is 1.5 m, so in total there are 12 rings.  The end-caps close the modules and provide the standard connection capability. They sustain the internal pressure and they are made of aluminium 6061.

Figure 14 - Primary structure elements

6.2.2 Secondary Structure The primary structure addresses the main operative loads, such as the internal pressure load or the mechanical action that the structure has to face during its mission. The secondary structure sustains the secondary loads, distributes loads to the primary structure and provides mounting provisions for the payloads and units. The secondary structure includes the panels for the floor and the ceiling of the module, and the ribs that connect the panels and sustain the loads of the equipment. The design of the secondary structure follows two main considerations:  The sizing of the ribs assumes empty squared section bars of Aluminium 6061 that define the internal layout of each module. The ribs connect the longerons with the floor and the ceiling of the module via the rings, so each rib has a length of 1.5 m.  Plastic composite material plates form the floor and the ceiling. Each panel has a length of 1.5 m and they are connected to the ribs.

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Figure 15 - Secondary structure elements

6.2.3 Structure Mass definition Considering the dimensions of the main modules of Mission MILESTONE, the habitable module, the greenhouse and the laboratory, it is possible to size the primary and the secondary structure. Table 13 summarises the structure mass and the margin that have been applied. Item Mass [kg] Contingency Mass w/ Margin [kg] Margins Rigid shell 8101 5% 9316 Inflatable shell 1685 5% 1938 Longerons 1012 5% 1164 Rings 1143 5% 1318 Primary (Shell, Longerons and Rings) 11941 13732 Ribs 391 5% 462 Floor 560 5% 661 Ceiling 280 5% 331 Secondary (Ribs, floor and ceiling) 1232 1454 Table 13 - Mass of structural elements

After applying the system margins, the primary structure has a mass of 13.7 tons with a 15% margin, and the secondary structure has a mass of 1.45 tons with an 18% margin.

6.3 THERMAL CONTROL SYSTEM The goal of the Thermal Control System (TCS) is to keep the temperature of the components and of the cabin around a fixed value, transporting heat to the rejection system via a transport system. The design has been performed in order to maintain the thermal balance over both that the power that has to be transported and rejected is the one that is absorbed by the system, 푞푠푦푠푡, and the heat load due to the crew operating inside the pressurised volume, 푞푏표푑푦. Considering the formulation of Dubois, 푞푏표푑푦 considers having 6 US male crew members performing hard activities, as worst case scenario. Generally, the air in the cabin is kept at 25°C through the air conditioning system, and the TCS subsystems limit the temperature of the electric components to rise above 50°C. The components are connected to the collection loop via cold plates. For safety reasons a water loop is considered for the inner collection loop, while for the outer one uses ammonia can be used, which has the benefit of a lower freezing point.

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The system that collects and transports the heat is the Active Thermal Control System (ATCS). For sizing the ATCS, a rough estimate based on the heat collected by the loop has been performed. The reference heat is considered to be produced by the electrical components and connected to the cold plates, 푞푠푦푠푡, and the one collected by the air conditioning system, produced by the CM, 푞푏표푑푦.The mass and the power consumption of all the components of the ATCS has been based on Human Space Flight: Mission Analysis and Design (42). The ATCS is includes heat exchangers, cold plates, pumps and accumulator, heat pumps, plumbing valves, instruments and control and the fluids. All the components have been sized for the power known and they are all sized knowing the power that must to be dissipated. The heat rejection is performed through radiators that receive the heat from the external fluid loop. To size the radiators the following assumptions have been made: The total power to be dissipated considers two other factors. The first is the heat leakages from the module to the Martian environment, considering a polyamide layer thickness of 10 cm per module, with a thermal conductivity of 0.04 Wm-1K-1. The second is heat rejected through convection of the radiators surface, depending on the surface itself and on an average wind speed of 7 ms-1 and an air density related to the landing site altitude. The heat to be rejected is therefore:

푞푟푎푑 = 푞푠푦푠푡 + 푞푏표푑푦 − 푞푙푒푎푘 − 푞푐표푛푣푒푐푡푖표푛 (2) The temperature that the radiators face is half that of the sky, i.e. -122°C on the hottest day, and the ground, i.e. 31°C on the hottest day, with a view factor of 0.5 (half sky, half ground). The radiators working temperature is considered to be 30°C. The specific mass of the radiators is 8.5 kg m-2, for two sided deployable radiators. With these considerations the area was sized and consequently the mass of the radiators and all the sub- components of the ATCS for all the modules. The inputs for the sizing model used are only the heat (the power absorbed) that the electrical components produce in each module, and the metabolic heat due to the crew. A global margin of 15% has been considered.

Figure 16 - Thermal loop schematic

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6.4 AUTONOMOUS ELECTRIC POWER SYSTEM The Autonomous Electric Power System supplies electrical energy to the modules during the period between the cargo mission and the crew arrival. The primary electrical need during this phase is for maintaining the temperature of module components, which is generally needed to be much higher than the Martian surroundings. This will be achieved with heaters at strategic points along the length of the module. Additional energy will be required for functions such as powering the communication links and fluid handling. Hybrid thermal-electric solar panels with concentrators will provide this power (43), which have been included in Section 17 as a technology that needs developing. The solar panels will be located on the top of the modules and the water already present inside the modules will act as a thermal sink. Regenerative fuel cells will store the solar excess produced in the daytime to supply a minimal and flat electric load during night time. The required performance of the panels was compared with the available water for use as a heat sink and the internal temperature of the modules. The performance required is for over 44% efficiency which is in agreement with NASA estimations (43). Results are based on high performance photovoltaic panels, as data for hybrid thermal-electric with solar concentrators is not available and the effects due to a dust storm have not been analysed. Furthermore, the temperature to be maintained inside the modules and the temperature gain achievable with the solar panels have been assumed, and these heavily influence the results, thus the performance of this technology will play a key role for Mission MILESTONE. Once the crew arrives the modules will be powered by the outpost centralised power system (see Section 5).

6.5 RADIATION PROTECTION A major factor into the safety of the crew members for Mission MILESTONE will be the long and short term effects of radiation, and as such it is necessary to establish a protection system. The majority of the danger from radiation will be exposure during the cruise phase to Mars because there is no natural shielding from the Martian atmosphere. However Mission MILESTONE is only dealing with the mission from Low Mars Orbit, with Mission ORPHEUS already having provided radiation protection in the CIV. The sources of radiation that the crew members are exposed to are solar particles and galactic cosmic rays (GCRs). Solar particle fluxes will vary during the life-cycle of the sun and will also increase due to solar particle events such as solar flares and coronal mass ejections; however solar particles are easier to shield against than GCRs. Galactic cosmic rays are much more difficult to provide protection against because of the secondary particles produced by interactions with heavy elements. A forward Monte Carlo simulation of the radiation environment on the Martian surface was modelled using SPENVIS MEREM models on the Geant4 platform (44) (45) (46). The sensitive elements of the simulation were modelled as ‘humans’ made of water located in a geometric model of the habitable module as it was designed at the time. The radiation protection employed was using a layer of water within the shell of the module and this thickness was varied during to simulate the effect that this water layer had. The relative effect of this water layer was analysed to accommodate for changes in the module design and is shown in Figure 17. The water thickness layer was then determined by used by any excess water from buffers for the habitable modules and the greenhouses. The thickness of the water layer that would be used for radiation protection is 10.41cm. As can be seen in Figure 17 this thickness of water used for radiation protection would reduce the relative dose to 70% of the dose without any protection besides the module structural components. The ~30% reduction in the dose received by crew members on Mission MILESTONE is considered acceptable. However, the dose received during the surface stay phase of a Mars mission is minimal compared to the transit phase, thus research should be focused on reducing the radiation dose received during transit (47).

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Radiation Dose as a function of water shielding thickness 130 120 y = -21.38ln(x) + 168.63 R² = 0.9051 110 100 90 80 70 60 As % of 0 Value 50 Log. (As % of 0 Value) 40 30

Dose Dose as of0mm % shielding dose 20 10 0 0 100 200 300 400 500 600 700 800 900 1000 1100 Water shielding thickness (mm) Figure 17 - The effect of the water shielding layer to reduce radiation dose

6.6 PLANETARY PROTECTION MILESTONE will be a category V restricted mission with respect to the Committee on Space Research (COSPAR) guidelines (48). Category V restricted requires the highest level of protection as it is an Earth return mission to a planet deemed to potentially host life. Each level of protection includes the requirements of the lower levels, and category V is unique in the respect that it requires Earth based containment and analysis facilities. To address forward contamination all exteriors and landers will need to be cleaned to a to-be-determined level of allowable biological burden (49). For humans performing EVAs, a total quantified biological burden cannot be determined and protective measures will need to be put in place (50). One crew member will need to be responsible for Planetary Protection measures being followed. In particular, care will need to be taken to avoid human error related contaminations caused by tiredness or ill health. Items taken from the outpost will need to be sterilised (49). For backward contamination, the policy should be to break the chain of contamination. This means no surface that comes directly into contact with Mars should be returned to Earth. This requires samples to be cached and for the disposal of the Mars Ascent Vehicle. Precursor missions with robotic probes to the landing site can do an initial assessment of signs of extant life to better understand the risk of contamination. Backward contamination also needs to be factored into medical procedures, with capabilities for quarantine should a crew member become affected (50). For both forward and backward contamination, the presence of humans, and the subsequent requirement for return to Earth increase the complexity with regard to planetary protection and wherever possible, contact with the Martian surface should be minimised. Planetary protection needs to be considered at all stages. All planetary protection is dependent on the situation, and the landing sites will need to be as well understood as possible prior to the crew’s arrival.

6.7 ENVIRONMENT CONTROL AND LIFE SUPPORT SYSTEMS (ECLSS)

6.7.1 Air Revitalisation Subsystem (ARS) The air revitalisation system is responsible for keeping an appropriate amount of O2 in the modules atmosphere, as well as removing the CO2 generated by the crew. The system is based on three main components:

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 The oxygen generation assembly, wherein a polymer electrolyte membrane electrolyser produces oxygen through water electrolysis  The carbon capture assembly, which is based on pressure swing adsorption technology to separate CO2 from air  The carbon reduction assembly, which consists of a Sabatier reactor and enables a reduction in the water needs. For future missions utilising the outpost, the ISRU plant will be used to extract water to fulfil the crew requirements and produce oxygen for the ascent module. In this scenario water will be supplied to the ARS for producing O2 for breathing, but this oxygen may be directly supplied by the ISRU with the ARS simply being used as a buffer. In order to size the ARS, a contingency margin of 5% has been applied to the requirements for the 60 day mission. Such value of contingency margin has been selected because the calculation is based on data available for the technologies now used on ISS, which have TRL 9 in LEO; however further technology improvements will follow to increase performance and reliability and make it suitable for use on Mars.

6.7.2 Atmosphere management, Thermal Humidity Control and Atmosphere Control and Supply Due to the scientific and exploration objectives of Mission MILESTONE, a high number of EVAs are envisaged for the crew, including at least four with the use of the exploration rover (see Section 4.4). The procedure to perform such an activity is quite complex and physically stressing. The difference between the atmosphere in the outpost and the atmosphere that is set inside the spacesuit strongly influence the physical stress caused by an EVA. This would normally be remedied by the crew undergoing a number of hours of pre-breathing, to prevent a shock transition. However, in order to increase the number of possible EVAs, it would be appropriate to reduce the pre-breathing time. This parameter depends on the ratio between the partial pressure of the inert gas in the outpost atmosphere and the spacesuit total pressure. Pre-breathing time can be eliminated when this ratio is 1.2 or less (51)). Mission MILESTONE EVA capability relies on pressurised spacesuits which work at 57.9 kPa. It means that an outpost pressure of 90.5 kPa gives zero pre-breath EVAs. Therefore, the outpost will work at around one atmosphere, with a composition of 79% of nitrogen and 21% of oxygen, of which all consumables will be brought from Earth. As Mission MILESTONE will establish a long permanence outpost, leakage has to be considered not only for the duration of the mission and for the time between cargo landing and crew arrival, but also for the possible future missions. The data for leakage was from the ISS, and this amounted to 0.227 kg/day (52). Due to the lack of data for inflatable modules, the same estimates have been used for the inflatable section. However, it is believed that this data can be considered a conservative value, as it considers the pressurised volume of the entire ISS in 2011, about 900 m³ which is much higher than the pressurised volume of a single module. The mission for the module can be split in three different periods:  Period One: time from module landing to crew arrival  Period Two: time from the crew arrival to crew departure for Mission MILESTONE  Period Three: time from crew departure to next mission The leakage is influenced by the pressure and composition of the atmosphere, and all three periods need to be considered. According to the strategy for Fire Detection and Suppression (see Section 6.7.3), Table 14 summarises the atmosphere in the outpost during the three periods.

Time (years) Pressure (atm) Nitrogen (%) Oxygen (%)

Period One 3 0.5 100 0 Period Two 60 days 1 79 21 Period Three 2 0.5 95 5 Table 14 - Atmosphere composition and pressure for each mission period

The time between Mission MILESTONE and the following mission has been assumed to be two years. All the gasses needed will be brought in tanks pressurised at 204 bar and 26.7 °C (53) and they will form the HPGS (High Pressure Gas System). In order to take into account the lower leakage at a lower pressure, the leakages have been considered proportional to the module pressure. Considering the volume of an empty

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pressurised module and the mass of the tanks, the total mass of the HPGS was computed for each module. The mass of the tanks have been considered proportional to the mass of gas according to the NASA BVAD, 2015 (54). For nitrogen, an average value of 1.124 kg of tank per kg of gas has been used; for oxygen a value of 0.364 kg of tank per kg of gas has been used. A contingency margin of 5% has been applied directly to the quantity of gas to be stored, which is in line with the conservative leakage values used. Thermal Humidity Control (THC) and Atmosphere Control and Supply (ACS) systems have been based on those on the ISS, and sized on the basis of data from (55).

6.7.3 Fire Detection and Suppression (FDS) Fire has to be seriously taken into account in the design of a Martian outpost. Experimental studies have been carried out exploiting the short period of time given by flights on the KC-135 (parabolic flight) (56) and the results are shown in Figure 18. Flammability range and flame spread increase to a maximum between Earth gravity and microgravity, showing a peak around the Martian gravity environment. The FDS system is based on portable water-foam fire extinguishers for all the modules in manned configuration, and the generation of an inert atmosphere to prevent a fire event when the outpost is unmanned. Water-foam extinguishers are the most effective in fire suppression, as the recovery phase is localised and easily accomplished. The greenhouse modules will follow a different strategy because the generation of an inert atmosphere has more issues and could damage the plants. The best solution for these modules is therefore to isolate them from the rest of the outpost and implement an automated system. The automated fire suppression system for the greenhouse adds complexity and mass to the module, but it is necessary in this case because the module needs to start operating before the crew arrive in future missions. The system will be based on CO2 pressurised tanks, a distribution network and sprinklers located in strategic points.

Figure 18 - Flame spread rate and flammability limits depending on gravitational acceleration (56)

In order to generate an inert atmosphere, oxygen needs to be removed. This can be done by oxidising hydrogen in a H2 combustor. Hydrogen can be produced by the water electrolyser already present in the ARS. Water resulting from the reaction can be recovered by the thermal and humidity control system, while other possible sub-products derived from the hydrogen combustion will be filtered. The oxygen produced by water electrolysis will be stored and used to fill the atmosphere when the next crew reaches the outpost. Fire detection will be made by sensors. Smoke sensors will be located inside each module, while the greenhouse will be equipped also with flame detectors. The probability of having false alarms must be reduced due to the presence of the automated system and the time delay in communication with the Earth ground segment.

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6.7.4 Water Recovery System (WRS) For Mission MILESTONE a water recovery system has been designed to recover potable water from all waste water streams, to provide some of the ~30kg/day needed by the crew members (57). The water recovery system chosen is based on the current water recovery system used aboard the ISS, which consists of both the urine processor assembly (UPA) and the water processor assembly (WPA) (58). The urine processor assembly is responsible for recovering water from the urine produced by the crewmembers of Mission MILESTONE and is expected to recover 85% of the water content from pre-treated urine (59). The effluent water from the urine processor assembly is combined with waste water from other sources that is fed to the water processor assembly. The waste brine stream from the urine processor assembly is then stored and processed by the waste management system described in Section 6.7.5. The urine processor components of the water recovery system would only be present in the habitable modules as they are not required elsewhere. The water processor assembly would be responsible for treating any other waste water (excluding urine), such as waste water from scientific experiments, washing and condensate from the thermal humidity control (THC) system (Section 6.7.2). The ISS water recovery system, on which the Mission MILESTONE WRS is modelled on is designed to process ~142kg/day; however due to the water provision requirement of the mission, the ISS WRS has been parametrically scaled for 180kg/day (59). The Mission MILESTONE water recovery system has been designed to recover a total of 87% of the water content of the waste water streams, and as a result some water needs to be taken from Earth to ensure the provision of enough water right until the end of the mission (60) (59). The extra water needed to be provided each day to supplement the water recovery system is 3.9 kg per crew member per day. For Mission MILESTONE the mass of water that will be launched from Earth will be ~1.4 tons. Spare water for Mission MILESTONE will be used for radiation protection as stated in Section 6.5. In future missions the supplemental water will need to be provided by the ISRU system. Advanced water recovery systems are being developed that incorporate biological processes to treat water and it is estimated that they will be much more efficient. The Advanced Water Processor (AWP) is being developed by NASA’s Next Generation Life Support Project and is expected to be able to recover ~95% of the initial wastewater volume to produce potable water (61). A similar system is being developed by ESA and is specified in Section 17. If the advanced water recovery systems are flight ready by Mission MILESTONE and perform as expected, the amount of supplemental water will just be 1.5 kg per crew member per day, or just 540 kg for the whole mission.

6.7.5 Waste Management System (WMS) For a 60 day mission compaction, sterilisation and storage alone is an appropriate method of waste management (62). In order to further reduce the amount of waste, oxidation can be introduced. Oxidation processes require additional CO2, trace gas, ash and water vapour. However, oxidation is oxygen intensive, with 1kg of oxygen needed to process 1kg of waste, which is not feasible for Mission MILESTONE (62). However, for longer duration missions with ISRU and greenhouse related oxygen production (such as the long term mission the outpost is sized for) this may become viable. General waste, from left over food and packaging will be collected in bin bags and then treated using a Heat Melt Compactor (HMC) (63). Standard astronaut waste contains over 20% water and 20% plastic (64). The HMC heats the waste in order to kill any microbes that may be present, evaporates useful water from the waste and then compacts it into a 1 inch thick tile of 9 inches square. Removal of the water content from the waste also prevents odour and microbial growth. The plastic content of the waste aids the bonding of the waste into a compacted tile. These tiles can be used as additional radiation shielding or incinerated. A trash compactor will be included in each habitable module. The Greenhouse modules have their own waste management system detailed in Section 9.3.

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7 HABITABLE MODULE The Habitable Module (HAB) is the core of the life support systems for the outpost, also providing the central control area to the entire outpost as it houses the communication core. Its dimensions are the same as the other primary living modules, with a diameter of 4.5 m and a total length when inflated of 15.7 m. The HAB. It is designed to sustain the life of the all 6 crew members for a period of 60 days. Two HAB modules have been included for redundancy, and in nominal operations both modules can be considered to house 3 crew members.

7.1 MODULE SIZING

7.1.1 Internal Layout Having the same external structure as the other modules, the habitable module has a unique internal layout due to its functions of support the life of the crew. The total volume has been divided in liveable volume and volume that can be used for allocate all the subsystems to the crew support. The section of the living volume has a hexagonal shape, with a maximum ceiling height of 2.3 m and a floor width of 4 m (shown in Figure 21). The total walkable floor surface is 63 m². From the analysis of the crew support systems, it is possible to allocate an internal volume. The biggest section of the volume is dedicated as common area that can be used for leisure activity, as a dining room and in which the communication devices and the command room are also present. Three bedrooms are present, equipped with bunk beds. In this way the maximum capability is of 6 crew members per time, although during normal operation only 3 people are living in each habitable module, to ensure private space for the crew members. Figure 19 summarises the definition of the living volume.

Figure 19 - Habitable module internal layout

7.1.2 Structure Mounting Once the crew has landed, the first task to operate the HAB is to inflate the inflatable part and mount the panels of the floor and the ceiling. In this way the HAB is fully extended and pressurised. All the furniture and the electrical equipment are stored in the rigid part during the travel while they are equally distributed all long the structure during the mounting operations. The life related subsystems are allocated in the volume above the ceiling and beneath the floor. This way the internal operations of the crew are not obstructed by the subsystems, while maintenance is ensured by the removable panels of the structure.

7.2 ENVIRONMENTAL CONTROL AND LIFE SUPPORT SYSTEMS (ECLSS) The ECLSS system is based on the subsystems set out in Section 6.7. This has a total mass of 6.2 tons, 56% of which is the wet mass, with the remaining forming the dry mass (i.e. the physical systems).

7.3 CREW UTILITIES The crew utilities are the support structures and equipment that the crew will interact with inside the habitable module. The utilities have been divided into subsystems to cover the different functions needed by the crew: food and galley, waste collection services, hygiene, clothing, personal, housekeeping, operational supplies, sleep accommodation and crew health care. For each subsystem the masses and volumes were calculated

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for both the consumable utilities, such as food and hygiene products, and the permanent elements, such as the dedicated medical suite and toilet facilities. The crew utilities were adapted and resized from a 500 day Martian surface mission compiled by Stillwell et al. (64). The food and galley subsystem encompasses the stored food the crew will consume over the 60 day mission and the facilities to prepare and store the food i.e. microwave ovens and freezers. The freezers are large and consume power but will reduce food spoilage when the greenhouse is fully producing food. The clothing system includes both a washer and dryer to allow for clothing reuse. On the ISS clothing is not reused and is worn for a period and then disposed of. With the higher gravity than on the ISS increasing clothing contact with skin and the extra exertion of the body as a result of being on Mars compared to in orbit, clothing will not last as long. During a 60 day mission it is feasible to bring non reusable clothing. However, for crews occupying the outpost over 500 days a washing and drying machine becomes a viable option to an increasingly large clothing mass. Comparing to a comparable Mars design mission (65), a washing machine and dryer was shown to be lower in mass than non-reusable clothing, and saved the crew’s time over manual washing. Crew health care involves both exercise equipment and a dedicated medical suite. The crew will need to perform both cardiac and resistive training to compensate for the reduced gravity they will experience on Mars (66). The inclusion of a medical suite is a NASA requirement for a Mars mission and should provide: basic life support, first- aid capability, clinical diagnosis, ambulatory care, private audio/video telemedicine, autonomous advanced life support, basic surgical care and palliative care, imaging and sustainable advanced cardiac life support and advanced trauma life support (67). A medical suite is provided in each habitable module to provide redundancy. Mission MILESTONE will require a physician on the crew and dedicated mission flight surgeons to provide telemedicine support. The inclusion of telemedicine support does not remove the need for autonomous medical care. In the event of a medical emergency, the crew of MILESTONE will need to act autonomously and decisively to make decisions due to the potential for communications delays, and part of this is ensuring all crew and stakeholders are aware of the procedures for a medical event and the potentially fatal consequences. The total mass of the utilities is 5.34 tons (with a 5% margin) with 1.795 tons being consumed by 60 day mission and 3. 55 tons remaining for future usage. A 5% margin was used as the utilities are comprised mostly of single unit items, and those which have high technology readiness levels for example stored food.

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8 EXTRAVEHICULAR ACTIVITIES (EVA) MODULE The Extravehicular Activities Module (EVAM) allows crewmembers to access the Martian environment.

8.1 CREW TRANSFER SCENARIOS The crew transfer and mobility scenarios were evaluated prior to designing the EVAM. This step was considered in order to primarily determine the functions of EVAM. The crew transfers identified are listed below:  Transfer of crew from Mars Descent Vehicle (MDV) to the outpost after landing on Mars  Transfer of crew from rover to Outpost  Transfer of crew from rover to Martian surface  Transfer of crew from Martian surface to rover  Transfer of crew from Outpost to Martian surface  Transfer of crew from Martian surface to Outpost  Transfer of crew from Martian surface to the Mars Ascent Vehicle (MAV)  Transfer of crew from rover to the Mars Ascent Vehicle (MAV)  Transfer of crew between modules in the outpost

8.2 PHYSICAL DESIGN Figure 20 shows EVA module layout for MILESTONE.

Figure 20 - EVAM Layout

Similar to the habitable module, the EVA module is divided into a rigid part of length 3.6m and an inflatable part, with a total length of 5.9m and a cross-sectional area of 9.6m². As seen in Figure 20, the solid part was designed to provide access to the rover and is also responsible for connecting the EVA module to the outpost. The inflatable part of the EVA module is divided into two compartments; the inner compartment, which has a length of 1.2m, and the outer compartment, which has a length of 1.1m. The inner compartment was designed to provide access to the EVA suits while the outer compartment was dedicated to house the EVA suits, forming a suit lock. The outer compartment also contains a door to provide access to the Martian surface. The two inflatable compartments are separated by a bulkhead that contains two suit ports, from which astronauts will be allowed to don and doff the suits directly into the pressurized volume (rear-entry method). This architecture was considered as a contamination mitigation method for the EVA modules of the EVA module as well as a measure to reduce pressure loss. A door was also included in the bulkhead to allow access to the EVA suits for maintenance and repair purposes. However, it is also worth mentioning that the outer compartment will be equipped with dust removal systems in order to reduce risk of backward contamination.

8.3 EVA SUITS The Z-series space suits were chosen as the Mission MILESTONE EVA suits, as it is currently being developed and would be beneficial to the mission (68). Z-series suits are pressurised garments designed by

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NASA for the purpose of planetary exploration. It utilises suit port technology which acts as an access point to/from the suits. In a report produced by Amy Ross in 2012, the first generation of the Z-series (Z-1) weighs about 73 kg (68). The main features of these suits are a soft upper torso, three-bearing shoulder design, a modular dual-axis waist, 2-bearing hip and modified lower leg patterning (68). All of the main features mentioned were included in the Z-1 architecture in order to tackle the weaknesses in previous space suits designs. The Z-1 PLSS contains the Variable Oxygen Regulator (VOR) that enables the user to control the pressure with up to 84 settings (69). This innovation will improve the EVA suits operations such as reduced pre-EVA breathing as well as in suit pressure decompression treatment. Such technology is currently at TRL 3 and NASA is aiming to increase the TRL to 6 by the year 2027 (70). As mentioned previously, the suits will be used in suit locks, which improve the pre-EVA breathing, airlock consumables and contamination mitigation (71). These should allow crew members to enter their suits and leave the bulkhead within 10 minutes in a low gravity environment (70).

8.3.1 Dust Removal System The inflatable outer compartment includes a dust removal system, with two modes of operation. In nominal mode, the inlet flow pump compresses filtered CO2 taken from the Martian atmosphere and sends this to the blower pipes, which will be used to blow the dust away from the spacesuit. The preliminary dust cleaning phase allows the collection of the dust in to the dust collector tank. In order to remove the remaining particles, a secondary dust cleaning phase is defined. During this phase, the EVA crew members use the vacuum cleaner pipes to suck the trapped dust particles, which are then directly ejected into the Martian atmosphere.

In intensive cleaning mode, the inlet flow pump compresses N2 available from the tank and sends this to the blower pipes which will be used to blow the dust away from the spacesuit. This cleaning phase allows the collection of the dust in to the dust collector tank. In this mode, during the secondary cleaning phase, the vacuum cleaner pipes are connected to the cyclone dust separator which divides the gas from the particles in order to recover as much N2 as possible. The EVA crew use the vacuum cleaner pipes to suck the trapped dust particles. The two modes will be activated in relation to the mode of the inflatable outer compartment (IOC). If the IOC is in nominal, the mode of the dust removal system will also be in nominal. If the IOC is in maintenance mode, the dust removal system mode is switched to deep cleaning mode. The dust removal system is depicted in Figure 21.

Figure 21 - EVAM Dust Removal System

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9 GREENHOUSE MODULE

Figure 1- Rendering of the greenhouse module

The greenhouse modules (GH) have been introduced mainly for the sake of future missions: in fact, the greenhouse is convenient against carrying food from Earth only after an extend permanence on the Martian surface. So, its benefits will become evident after some long permanence missions; nevertheless, it will guarantee a solid base for investigating the growth of plants on another planet, a key point for a sustainable planet colonisation. In addition, access to plants has been shown to have psychological benefits for the crew members (72).

9.1 GREENHOUSE LAYOUT AND STRUCTURE The greenhouse modules share the same primary and secondary structure with the Habitable Module: so the section of the internal volume has a hexagonal shape, with a maximum height of the ceiling of 2.3m and a floor dimension of 4m. In addition it has also a tertiary structure which consists in a series of four racks, extended for all the length of the rigid part, separated with two corridors, in order to guarantee accessibility for both the crew and the robot. The driving factor for the size is the cultivable surface: the configuration assumes the average height available to each plant is 58cm, compared to the 45cm that was used to compute the minimum necessary volume. This additional height will guarantee a better visibility and accessibility of each plant, but also a slightly higher distance from the LED during growth.

9.2 GROWTH CONSIDERATIONS

9.2.1 Growth Mechanism The greenhouse will use a hydroponic growth system, specifically the nutrient film technique (NFT). In NFT, plant roots are kept in contact with a small amount of water (a few cm of depth), which is continually cycled around the system. It has been found that it is possible to grow most crops with NFT, including potatoes which usually require a solid growth substrate to grow successfully (73). The benefit of NFT is that the steady stream of water and nutrients past the plant roots allows them to absorb as much water as they require. A gravity-assisted system will be employed, in which the water/nutrients flow through the plant system assisted by gravity before returning to a liquid tank, where additional nutrients can be introduced before the liquid is recycled through the plants.

9.2.2 Nutrient Management In order to ensure correct growth of the crops in the greenhouse, an enriched nutrient solution will be introduced into the water flowing around the hydroponic system. These nutrients can be added in the form of salts by means of a diluted solution, namely an initial ionic solution (more diluted) and a replenishment ionic solution, such as those listed in the NASA BVAD (74).

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9.2.3 Illumination From a trade-off, it was determined that LED lighting would be best suited for the greenhouse. This allows the frequency of lighting to be tailored to ensure optimum crop growth. The lighting system has been sized based on current Earth-based technology, the O-S Zelion®HL (75).

9.2.4 Irrigation System

9.2.4.1 Water Circulation NFT systems require a large amount of water circulation. In order to do this, there is assumed to be 0.6 channels/m² of greenhouse, each with a length of 9 m. From Earth-based studies (76), 1l/min of water flux has been found to be appropriate.

9.2.4.2 Water Recovery The Water Recovery System for the greenhouse will be as discussed in Section 6.7.4, and has a total mass of 922 kg and a power requirement of 1.5 kW. Part of the water input is absorbed by the crops while growing, and forms both edible and inedible water content (74). The rest of the water is then transpired by the crops, and a 95% efficiency of water recovery from transpiration is assumed. As the waste management system (see Section 9.3) also recovers a large amount of water, a total of 14.5 kg/day are recovered, requiring only 31.2 kg/day to be introduced from an external source.

9.2.4.3 Water Storage Due to the use of a hydroponic growth medium, water is a critical component in the greenhouse and it is necessary to provide a number of buffers. The total water used by the greenhouses is 400kg, as each channel has a depth of 5 mm of water at all times (77). It is necessary to provide a tank in which the nutrient solution can be mixed with the water before being circulated, which must hold the volume of the entire system (in order to replenish it when a cleaning flush is carried out), and additionally the daily water usage - the "in-use" tank. It is assumed that cleaning flushes will be carried out at regular intervals, fully removing and replacing the water in the NFT system. This tank has a volume of 1.3 m³. Each greenhouse will have a tank of half this size for redundancy and efficiency. A second tank will be considered as a water buffer, if the in-use water were to become contaminated in any way, and will store water for 60 days of operations and two flushes. This tank has a volume of 8.2 m³. A third tank will also be provided in which the water can be purified before being returned to either tank one or tank two. This tank is sized to be able to contain all the water for the "in-use" tank, plus the water in the system, such that if there were a contamination just before a flush all the necessary water could be purified. This tank has a volume of 2.3 m³. This water will not in practice be stored in tanks, but will be stored as radiation protection around the habitable modules, as discussed in Section 6.5.

9.2.5 Atmosphere Management The atmospheric composition of the greenhouse will be the same as for the rest of the outpost to increase the ease of movement for the crew and to allow for a consistent air system throughout the outpost. It is therefore provided by the systems stated in Sections 6.7.1, 6.7.2 and 6.7.3. However, if it became necessary to increase crop production, it has been considered to increase the carbon dioxide levels in the greenhouse, as discussed in the NASA BVAD (74).

9.2.6 Robotic Assistance A dexterous multifunction robot is planned in each module to partially automate all the operation needed to be performed in the greenhouse, from the planting of the seeds to moving the plants from germinator trays to growing trays, to help in checking the growth and the maturation of the fruits. It is also envisaged to help humans harvesting the crops and disposing of non-edible parts of the plants. The best solution would be to have a dexterous main arm, which could switch the end effector between different ones, like a tongs, a camera system, a spray etc. As an example, the KUKA LBR IIWA 14 R820 has been taken into account (see Figure 22). It has a mass 29kg and can carry up to 14kg, with a precision of

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one millimetre. The robot could be fixed to a telescopic tube which could move along rails on the ceiling, so it can operate together with humans, without mutual hindrance (an example is given in the picture below).

Figure 22 - Illustration on the possible use of the KUKA robot

The mass budget of the robotic system has been kept equal to the reference in the BVAD 2015 (74), which is surely a conservative estimate, including the rails. In order to estimate the power, a value of 5kW per module has been allocated to take into account any additional robotic subsystems; this value is more than the double the power required by the robot itself.

9.3 WASTE MANAGEMENT The configuration of the waste management system of the greenhouse is depicted in Figure 23 and its budgets are reported in Table 15. The system will be capable of using both the greenhouse waste stream and the O2 produced by the crops to restore some of the CO2 necessary to the growth chamber. It will also treat a part of the waste stream incoming from the habitable modules, mainly the dried faeces, urine and toilet paper produced by a crew of six. The waste management system is formed from:  A steriliser, where the biomass wet base and the habitable module dry base are treated  A drier, where the majority of the water is recovered  A condenser, that condenses the water removed in the drier  A size reduction and compactor, for reducing the waste size  A batch incinerator, which burns the resulting dry mass The by-products of the burning are ashes, which are collected inside storage, and a mixture of hot gases. The noxious gases are extracted (mainly NOx and particulates), hot air is returned to the drier and carbon dioxide is fed back to the growth chamber of the greenhouse.

Mass [kg] Volume [m³] Power [kWe] Cooling [kWth] ESM [kg] 982 9.2 5 4.6 1800 Table 15 - Properties of the Waste Management System

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Figure 23 - Final configuration of the waste management system

Each greenhouse produces 1.87 kg/day of O2 and the incineration process will need 1.52 kg of it to burn all the fuel coming from the waste streams. The system will be able to recover 7.27 kg/day of water and will produce 2.09 kg/day of CO2 for the growth chamber. The ash produced amounts to 0.114 kg/day, and at the end of the 500 days mission it is expected to have to store around 58 kg. In the future, it would be beneficial to introduce nutrition recovery from the ashes and to design a system to recover N2 from the harmful NOx.

9.4 CREW DIET The crew diet was decided from a number of considerations, one of which was the required dietary inputs stated by NASA (74), which require approximately 3000 kCal, 400g carbohydrate, 100g fat and 120g protein per crew member per day, together with vitamin and mineral limits. A number of diets were compared, with their nutritional benefit being assessed using data from the SELFNutritionData website (78). The diet chosen requires the smallest area of greenhouse, and introduces a number of high calorie and fat foods which would be brought from Earth, such as bresaola, tuna (preserved in olive oil), salame milano and chocolate. The total diet provides 2.2kg of food per astronaut per day, with 560g of this being provided from Earth. This includes both the high calorie foods and dried wheat, rice and beans, as these were deemed too difficult and resource consuming to grown in the greenhouse. The total requirement of food for transport from Earth is 1875 kg, which provides 76% of the calorie content per crew member per day. The crop growth and masses can be seen in Table 16.

Crop Mass (g) Crop Mass (g) Cabbage 220 Snap bean 150 Carrot 100 Onion 50 Celery 100 Strawberry 220 Lettuce 50 Pea 50 Green onion 50 Tomato 220 Pepper 220 Radish 120 Red beet 110 White potato 200 Table 16 - Crop variety with daily consumption masses

9.5 STORAGE In order to minimise crop variation due to the Martian environment, it was decided that all seeds required would be brought from Earth. All crops would therefore be first-generation Martian crops, with no use of seeds produced (other than for scientific experiments). The total storage required is 42kg, which provides

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enough crops for 6.72 crew members for 500 days (which accounts for crew cross-over time for the long duration mission), including the mass of potato seeds. Any crops required for storage will be stored in a freezer, with a total of 2.2 m³ of storage being required per greenhouse. This provides sufficient storage for 98 days worth of crop growth, which corresponds to the 68 day average growth period for crops, and an additional 30 day storage of 'fresh' food. Freezing was chosen as it was a power and mass-efficient method of storage, and also causes minimal physical changes to the food produced, which may provide psychological benefits for the crew.

9.6 CONTINGENCY CONSIDERATIONS Before sizing the greenhouse, a 15% margin was applied to the surface required for production. This is intended to take into account the uncertainties of growing plants on another planet, an operation never tried before. Under optimal lighting and nutrient condition, the actual yield of plants can be boosted by up to 10% (74). This, combined with the margin, shall be enough to compensate for possible complications. 76% of the energy amount is provided by the food brought from Earth, the following plan for contingencies has been chosen:  double the food to be brought from Earth;  use the nominal surface to be cultivated, but split in two greenhouse modules;  bring from Earth the same type and amount of food the greenhouse is designed to produce in a month, lyophilized, as an emergency buffer for 1 month. With this configuration, the volume and mass is kept at the minimum, and in the worst case scenario (complete stop of the production of both greenhouse modules), the crew has 30 days of eating a normal diet to repair the failure. The system is therefore two-failure tolerant. It is also thinkable, even if is like a life-or- death compromise, that in case of unrecoverable loss of both the greenhouse modules, the crew could wait for the subsequent launch window (which is up to two years) eating only the food brought from Earth together with vitamins and minerals. The values of this system configuration, per greenhouse module, are shown in Table 17:

Cultivated Area (m²) 93 GH Cultivated Volume (m³) 36 Water per day (kg/d) 200 O2 per day (kg/d) 1.9 CO2 per day (kg/d) 2.6 Inedible Dry Mass (kg/d) 0.8 Inedible Water Content (kg/d) 7.3 Inedible Total Mass (kg/d) 8.1 Edible Water Cont. (kg/d) 6.3 PPF per day (mol/d) 4165 Table 17 - Characteristics of each greenhouse module

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10 NODE

10.1 NODE CONFIGURATION The node of MILESTONE consists of a cylindrical rigid structure and four inflatable structures. It is used to connect separate modules together.

Figure 24 - MILESTONE Node

As seen in Figure 24, the rigid structure acts as a central point that connects the four inflatable structures. These inflatable structures will be responsible for connecting the modules to the central rigid part. This design was considered in order to optimize the volume needed to connect the modules together without exceeding the diameter constraints of the fairing. The node has a rigid diameter of 2.6m and a rigid height of 3m. The inflatable sections have an inflated length of 2.3m, and a diameter of 2.4m at the end cap. The node has a total volume of 60m³ and a mass of 7.64 tons.

10.2 NODE STRUCTURE Due to its functional requirement, the node design is different from any other modules included in Mission MILESTONE. As seen in Figure 24, the node structure essentially consists of a vertical cylinder that is connected by a group of inflatable structures. All of the primary and secondary structures are similar to any other modules, however, they are arranged in different manner. The driving parameter in sizing the structure was the available volume in Falcon Heavy Launcher (3). In addition, the space needed to avoid contact between the collected modules was another important design constraint.

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10.3 NODE MASS BUDGET The node mass budget has been estimated, and can be seen in Table 18 and Figure 25.

Subsystem Type Mass [t]

Primary Structure 5.21 Secondary Structure 0.53 ECLSS 1.07 ATCS (Active Thermal Control System) 0.61 AEPS (Autonomous Electric Power System) 0.16 Communications 0.07

Total 7.64 Table 18 - Node mass budget

Node Module Mass Budget

2% 1% 8% PRIMARY 14% STRUCTURE SECONDARY STRUCTURE ECLSS 7% ATCS

68% AEPS

COMMUNICATION

Figure 25 – Node mass breakdown

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11 ENTRY, DESCENT AND LANDING (EDL)

11.1 LANDING APPROACH Several configurations were taken into consideration, with the general approach always containing the following steps:  The cargo modules orbit in a 200 km in a circular orbit. They make a Hohmann transfer to a 100 km orbit, where atmospheric drag begins the aerobraking process.  A Hypersonic Inflatable Aerodynamic Decelerator (HIAD) is then inflated to increase the drag and therefore the efficiency of the aerobraking. For this phase, it is assumed that the angle between the direction of nose and the velocity vector is kept equal to zero by the trajectory control system.  At a certain altitude above the Martian surface, the HIAD and the TPS are ejected using springs, which provide a simple and reliable solution for separating the stages. This is the end of the aerobraking phase.  A constant flight path angle descent is then performed, using retrorockets to slow down the module for the last part of the entry manoeuvre. This is divided in two phases. First, a more powerful retrorocket is used to completely halt the module. Subsequently, smaller retrorockets are used to land vertically, and to control the orientation of the module.  For the last stage of the landing, it is necessary to take into consideration the unavoidable damage that the ground will suffer due to the retrorockets. In order to ensure a safe touchdown, the module will be granted an impulse in the horizontal direction at a certain altitude that will be optimized with a python™ script. Then, as the module falls, wheels will be deployed. This way, the module will fly away from the damaged ground. The wheels will have to provide a sufficiently large surface that the ground does not crumble under the force of the impact. This force will be softened by oleo- pneumatic shock absorbers. The EDL phase, starting from the LMO at 100 km has been simulated using a code written in python™. The code has been used to trade-off between several possible configurations and then to optimise the winning one. The final configuration is shown in Figure 26.

Figure 26 - Schematic definition of the EDL phases

11.2 LANDING ELLIPSE The landing ellipse was evaluated using the same simulation, taking into consideration the uncertainties for the position and velocity of the rocket when the descent starts. The uncertainty values used are 5 m/s for the velocity and 400 m for the position. Using the simulation as a propagator and considering these errors, it was

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possible to roughly estimate the size of the landing ellipse, which was found to be an ellipse with a semi-major axis of 10 km, and a semi-minor axis of 3 km.

11.3 TOUCHDOWN SUBSYSTEM In order to avoid debris damage on the bottom of the module, due to the impingement of the plume on the ground which will throw rocks up in the air, and also in order to avoid the possibility of landing inside the hole formed after the turning off the engines (as pointed out in section 5.9 of the NASA Mars Design Reference Architecture 5.0 – Addendum (4)), the following touchdown approach has been chosen:  Using the main retrorocket, the module achieves a zero velocity and hovers over the Martian surface. At this stage, the main retrorocket is jettisoned and the vertical descent is performed using the secondary retrorocket.  At an altitude of 15m above the landing site, a horizontal impulse of the secondary rocket is performed to achieve a speed of 6 ms-1 in the longitudinal direction. This is done in order to avoid the crater created by the vertical retrorockets. The retrorockets are then turned off, after which a free fall starts.  Immediately after that, the shock absorbers are released from their stowed configuration, in which they are maintained until the shutdown of the vertical retrorocket in order to avoid heat damage to the wheels.  At the touchdown, the impact is softened by the shock absorbers. This will ensure that the structure of the module does not suffer accelerations greater than its structural limit, set at 3.5 Earth-g.  As the module has a horizontal velocity, wheels are used to grant mobility in the longitudinal direction during the touchdown. This also helps to reduce the impact pressure so that the ground does not crumble. Eventually, brakes are used to stop the module. Figure 27 shows the touchdown phases graphically.

Figure 27 - Touchdown phases

11.4 HYPERSONIC INFLATABLE AERODYNAMIC DECELERATOR (HIAD) SUBSYSTEM The HIAD is an inflatable structure used to increase the drag during the aerobraking phase of the descent. The inflation is carried out using inert pressurised gas, and Figure 28 shows the dimensions of the HIAD both in its deployed and packed configurations. The inflatable part of the HIAD packages like an umbrella. The material used is based on SIRCA-15 (79). There is also a possibility of using the HIAD as acoustic shielding during launch. The size of the HIAD has been chosen considering the amount of drag required for the aerobraking phase, optimised in the simulation. The structural mass is related to maximum aerodynamic pressure that the module faces during the aerobraking, also given by the simulation. Moreover a rigid frontal thermal protection system (TPS) is required to absorb the heat flux of the entry. It has been sized knowing the heat load given by the simulation. Considering that the HIAD and the TPS can be described as a sphere-cone nose, the drag coefficient of the aerodynamic decelerator has been evaluated integrating the flux around the surface, in the hypothesis of Newtonian flow (80). This assumption fits relatively well with the hypersonic flow regime.

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Figure 28 - HIAD dimensions and packed configuration

11.5 RETROROCKET SUBSYSTEM The size of the retrorocket has been performed considering:  Main thruster has to stop the module above a certain altitude after a constant flight path angle descent.  Secondary thrusters have to sustain the module during the vertical descent and guarantee the longitudinal speed necessary to avoid the hole on the ground due to the impingement of the thrusters themselves.  Reaction control thrusters are sized in order to maintain the angle of attack always equal to zero during the all phase and to orient the module in the proper position for the last part of the descent. To size the propulsion system accurately, it was taken into account not only the mass of the engines and of the propellant but also the mass of the tank, the cryocooler system (section 13.6), the micrometeoroid protection and the insulation system. The main and secondary retrorockets work with LCH4 and LO2 propellant, considered with an Isp = 360s.

11.6 SENSING SUBSYSTEM It will be necessary to provide the system with several different sensors, as well has having a better understanding of the Martian surface and atmosphere. Some of the sensors that will be used are:  Inertial navigation system: using gyroscopes and accelerometers it is possible to estimate the position of a spacecraft via dead reckoning. This method is susceptible to cumulative errors, which is why it is combined to other sensors to provide further precision.  Barometric sensor: this sensor is used as an altimeter in Earth based aircraft. However it will be necessary to study the Martian atmosphere to much greater depth in order to modify this sensor and achieve a sufficiently high TRL  Radar: using a radar it is possible to estimate the altitude of the module.  Lidar: using the same principles as a radar, it is possible to estimate the altitude of the module using a light beam and measuring the scattered light from this beam.

11.7 HEAT SHIELD EJECTION SUBSYSTEM Leaf springs will be used to force the heat shield apart upon release. They offer a simple and reliable solution for a low weight. The speed of the ejected bodies will be of 1 ms-1, which will offer enough time to separate the heat shield from the module before landing takes place.

11.8 OPTIMISATION Once the simulation had been implemented, it was necessary to find an optimum configuration of parameters that would allow the module to reach the touchdown phase with a minimum speed while carrying the maximum amount of weight. Different approaches were used in order to validate the results: a Monte Carlo simulation and a brute force approach. Since the physics of the two models is the same, the optimisation performed by both gave the same results.

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The parameters that were optimised are:  Thrust in Newtons of the main retrorocket  Thrust of in Newtons of the secondary retrorockets  Altitude at which the HIAD is ejected, so aerobraking stops  Altitude at which the main retrorocket is ejected  Altitude at which retrorockets are initiated, which is equal to the HIAD jettisoning altitude  The diameter of the HIAD, and therefore its drag coefficient The value used for the optimization is 푣푚−2, where 푣 stands for the arrival speed, and 푚 is the mass that is landed when the touchdown takes place.

11.9 SIMULATION RESULTS This section presents the result of the simulation and the following optimisation made for the heaviest module of the mission MILESTONE. The graphs show the trajectory of a payload, of 39 tons in LMO, entering in the Martian atmosphere, following the physics described in the section above. The landing altitude is supposed to be at MOLA level of -3000 m. In the first phase, the HIAD is deployed in order to brake against the Martian atmosphere, reducing the velocity up to an altitude of -2000 m, as given by the optimisation. The HIAD diameter has been optimised to a value of 23 m. At this point, the module ejects the HIAD and continues the descent with the retrorockets with a constant FPA up to an altitude of 200 m; then, the main engine is ejected and a vertical descent is performed. The total entry phase lasts around 1000 s. Considering the topographic values given by MOLA satellite from NASA (81), it is possible to identify the areas on Mars where land with the selected system can take place. Further considerations on the bearing capacity of the soil and the possible presence of resources are necessary.

Figure 29 - Definition of possible landing area related to MOLA altitude (81)

Figure 30 and Figure 31 show the velocity and heat flux variation for the descent configuration.

Figure 30 - Velocity variation during aerobraking deceleration

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Figure 31 - Heat flux variation during aerobraking deceleration

The maximum dynamic pressure that the module experiences is 447 Pa, reached around at 35 km altitude after ~780 seconds. The maximum deceleration instead is about 1.8 gEarth, reached at similar values of altitude and time. The heat peak occurs at an altitude of ~57 km. Table 19 summarises the results obtained with the simulation. Parameter Value Altitude MOLA [km] Time [sec] Total Heat flux 2.48W/m2 57 690 Maximum Aerodynamic pressure 447 Pa 35 780 Maximum g load 1.8 35 780 Table 19 - Maximum thermal, aerodynamic and structural loads during aerobraking phase

Table 20 shows the mass budget for the EDL system related to a module mass, including the EDL, of 39 tons in LMO. The propulsion system includes the propellant for the first Hohmann transfer. EDL avionics includes the mass of the trajectory control system, so the propellant mass of the controllers. The landing system mass includes the shock absorbers and the landing wheels. The mass of the retrorocket systems considers the mass of the retrorockets themselves, the propellant and the tanks.

Subsystem Mass[kg] Propulsion system 1045 EDL avionics 1066 Communication/Avionics 29 HIAD 6464 Retrorocket system 4561 Landing System 2455 Total 15620 Table 20 - EDL System Definition

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EDL MASS BUDGET

Propulsion system EDL avionics Communication/Avionics HIAD Retrorocket system Landing System

7% 16% 0% 7% 0%

29% 41%

Figure 32 - EDL System Mass Budget

Figure 33 shows EDL configuration of the module as previously described.

Figure 33 - EDL System Components

11.10 MODULE TAILORING Having demonstrated the capability of landing large payload on the Martian surface with the selected architecture, simulations have been run in order to properly define the EDL sub-system mass for each module of Mission MILESTONE. Firstly, several simulations have been performed, varying the mass in LMO from 10.5 tons up to 44 tons in order to evaluate the total mass on the Martian surface. For each configuration, the code optimises the different sensitive parameters (i.e. the amount of thrust of the primary and secondary thrusters and the HIAD diameter) to achieve the largest possible landed mass. As shown in Figure 34, the simulations show the EDL system mass can be considered as proportional to the entry mass.

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Figure 34 - EDL and landed Mass related to the mass in LMO

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12 MARS DESCENT VEHICLE (MDV)

12.1 HUMAN DESCENT PHASE The MDV will arrive in Mars orbit attached to the CIV. The interplanetary trip ends when a 500 km LMO is achieved and stabilised. After, the crew enters in the MDV, the CIV is switched into unmanned mode and the EDL phase starts. The MDV consists of a scaled habitable module that would allow a crew of six members to survive for 21 days. This accounts for the possibility that the MDV lands far from the designated point, given that the landing ellipse has semi-axis of 10km and 3km. With this approach, the MDV has the same entry, descent and landing system as the other modules. The life support supply also allows the MDV to be used as a pressurised exploration vehicle for future missions to allow for multiple day exploration, although it will not be used for this function during Mission MILESTONE. The EDL system achieves a maximum deceleration compatible with the manned entry requirements. It is possible to assume its safety level is acceptable, since the crew arrival is supposed to be around two years after the first module landings, so around 10 modules have already safely landed on Mars using this system. Due to the fact that the MDV has wheels (used for the landing), it can be carried by the rover to the outpost location. During the initial assembly and set up phase of the outpost, the crew would perform EVA activities through the MDV’s suit-lock while living in the MDV itself. Two suit locks are present on the MDV since a two crew member EVA is foreseen, as discussed previously. A standardised docking port is also present, and using this the MDV can connect to an EVA module, allowing the crew to enter in the outpost once it is operative. This happens when at least one HAB is connected to one EVA module and the power plant is operative, as shown in Figure 35.

Figure 35 - Minimal survival configuration

In case of failure of the two suit locks, the MDV is equipped with six EVA suits that allow the crew to exit directly from the MDV to the Martian surface. This scenario foresees the MDV depressurisation and it is intended only in the case of critical failures.

12.1.1 MDV Sizing In order to size the MDV, the habitable module has been scaled considering the shorter duration of the mission of only 21 days. These 21 days ensure a contingency margin period to complete the set-up of the outpost, while this is not operational. The dimensions of the module have therefore been reduced, to have a diameter if 4m and module length of 6m. For simplicity, the module only has a rigid structure.

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All the subsystems related to the human support were sized from the equivalent habitable module systems, considering the reduced stay. Figure 36 show the masses of the different subsystems. Considering the reduced duration of the mission, the MDV has different subsystems. For example, only oxygen tanks have been considered for the Air Revitalisation System. The details can are shown in Table 34 in Appendix A: Mass Budgets. Moreover, the MDV is designed to have 4 deployable solar arrays in order to produce the amount of power required for the subsystems, estimated around 2 kW. The power is guaranteed by 4 solar arrays with a radius of around 1.3 m, considering a solar constant on Mars of 98 Wm -2, and a specific mass of the array of 36 Wkg-1. As seen in Section 14.2, for mission MILESTONE the MDV can be brought by the rover far from the outpost as an exploration habitat for the crew, after the completion of the outpost assembly. In the mass budget, the systems related to the EVA activity are also considered.

0% MDV Mass Budget 3% 5% 2% PRIMARY STRUCTURE SECONDARY STRUCTURE 29% 56% ECLSS ATCS 5% AEPS COMMUNICATION SPACE SUIT

Figure 36 – MDV habitable subsystems mass budget

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13 ASCENT

13.1 MISSION CONSTRAINTS The Mars Ascent Vehicle (MAV) will be landed fully fuelled with no in-situ propellant production, using the landing system described in Section 11. Both the fully fuelled MAV and its landing system will fit inside a single SLS-2B launch. The MAV is a separate module from the Mars Descent Vehicle, as described in Section 12, which allows for the MAV to be sent as part of the cargo phase with its ensured integrity being a prerequisite for the human mission to take place. The MAV has been sized with the focus on a being able to be landed, launched and able to rendezvous with its target, and as such a detailed docking procedure has not been developed. This study aims to show that a non-ISRU reliant vehicle is feasible, and further work will have to prioritise docking capabilities. The MAV has been calculated as a single stage to orbit vehicle a more mature development may select a two stage configuration to provide a mass saving at the cost of design complexity.

13.2 ROLE OF THE CREW INTERPLANETARY VEHICLE (CIV) The CIV is the unmanned crew interplanetary vehicle designed by Mission ORPHEUS (1), which will return the crew to Earth (1). The CIV serves no active role during the Surface Activities stage of the mission other than to be accessible by the MAV. To reduce the complexity of the MAV manoeuvres, the CIV will stay in a 500km circular Low Mars Orbit (LMO). The orbit will also need to remain in a higher inclination than the launch site inclination to enable two launch windows per sol and be able to maintain the orbit; effects of precession will be compensated for to avoid costly manoeuvres by the tightly mass constrained MAV. For contingency it can be assumed that the CIV may be able to perform a rescue of the MAV should it fail to rendezvous under its own power.

13.3 MANOEUVRES FOR ASCENT

13.3.1 Launch and Pitch Over To convert the radial velocity away from Mars to a tangential velocity to put the MAV in orbit a pitch over manoeuvre is used. This is achieved with two initial thrusts, one to cause the vehicle to take off of 350 kN, and one to cause the vehicle to turn of 150 kN.

13.3.2 Rendezvous Having pitched over and achieved a tangential motion, the MAV will need to rendezvous with the orbiting CIV in LMO. The MAV will perform a Hohmann transfer to reach the CIV at the point of the circularising second burn, as this is the most fuel efficient manoeuvre. For the MAV to reach the CIV directly with the transfer, the CIV will need to be 297o behind the MAV at the first point of transfer. The MAV has been sized with additional propellant to correct the rendezvous position by ±10o. Due to the rotation of Mars and the need to use the MAV at any point during the 60 day mission, the CIV cannot be relied upon to be in position at time of launch. To prevent the need for costly inclination changes and to minimise the time between potential launches a parking orbit has been proposed. With a 200km parking orbit the MAV can launch twice a day and then wait a maximum of 16 hours in the parking orbit to begin transfer directly to the CIV assuming some degree of station keeping from the CIV.

13.3.3 Docking Having rendezvoused with the CIV the MAV must then dock so the crew can transfer themselves and the samples. The docking procedure has not been developed and will need to be in time for the mission. For the purposes of this stage of development, the rendezvous process has been examined in some detail and the docking given some basic assumptions. The docking process will be automated with the option for crew intervention. The Communication Satellites in Areostationary orbit will allow for relative position tracking for the CIV and MAV. The final alignment will most likely be conducted using camera alignment and/or some form of on board radar or lidar, though this is beyond the scope of the current report.

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13.3.4 Disposal of Ascent Vehicle For the purposes of planetary protection, the MAV will be disconnected from the CIV and left in orbit, breaking the chain of contact between the contaminated MAV and the Earth’s atmosphere. A 500km orbit is sufficiently high to provide a graveyard orbit so no other manoeuvres will be required.

13.4 DELTA-V BUDGET The Mars Ascent Vehicle will have a total delta-V of 4.1 km s-1 which covers the launch, pitch over, transfer to 500km, circularising burn, corrective manoeuvre and an additional compensation for performing the transfer whilst in the Martian atmosphere. The delta-V is comparable to values given for a Mars ascent given by Whitehead (82) and Condon et al. (83).

13.5 ASCENT DURATION The duration of the Mars Ascent Vehicle’s rendezvous procedure is dependent on the path the MAV takes. The potential paths are shown in Figure 37.

Figure 37 - Ascent and Rendezvous alternative paths

A launch can take place twice a sol resulting in a launch window every 13 hours. The mandatory pitch over manoeuvre is of the order of minutes. A transfer directly to the CIV will take approximately one hour. Alternatively a transfer to 200km parking orbit will take an hour to reach the parking orbit, up to 16 hours in the parking orbit waiting for a transfer point and one additional hour to get to 500km. The time required to correct the position on orbit by ±10o is 3 hours. The docking process has not been examined in detail but is of the order of minutes (84). The total time for the longest path from launch site to CIV is 34 hours. The ascent vehicle has life support systems for three days which covers the longest possible path and allows for an additional missed launch window and parking orbit transfer window.

13.6 CRYOGENIC FLUID MANAGEMENT The MAV will use a liquid methane propellant with liquid oxygen oxidiser. The propellant will be stored cryogenically so that the propellant will remain liquid until the point of ignition. The MAV will be launched to Mars as part of the cargo mission and so cryogenic fluid management is a necessity. To prevent boil off during the cargo mission phase till the point of launch active cryogenic management tanks have been sized for mass and power based on the mathematical model presented in Kittel et al. (85). The tanks will be cooled with solar powered cryocooler unit.

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13.7 CREW ENTRY PROCEDURE The MAV will require a separate EVA module, detailed in section 8, but one which is raised up to allow for crew ingress to the vehicle. This is a necessity due to the tight mass restrictions on the MAV not allowing for an on board airlock and the bulkiness of the space suits and planetary protection policy not allowing direct access from the Martian surface. Using an EVA module will enable the crew to enter two at a time using the suit lock mechanism and having to remove the suits for each additional pair entering. In the event that a crew member is incapacitated or that everyone needs to enter quickly, the EVA module doors can be depressurised and the crew can enter in full suits. Entering in full suits will expose the crew to Martian dust and will require cleaning and health monitoring once on the CIV and possible extensions to quarantine on return to Earth.

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14 ROVER - EXPLORATION AND MOBILITY SYSTEM The rovers discussed in Section 4.3.4 are the mobility system for Mission MILESTONE, and are initially used in the collection of the cargo, and later for exploration and scientific purposes. The sizing of the mobility system has been performed considering the collection phase of the cargo modules once they have landed. In fact, considering the landing ellipses analysed in the EDL section (see Section 11), the modules land on the Martian surface in a wide area and two collection rovers are planned to bring them on the selected outpost location. The rovers are assumed to be able to travel autonomously during the cargo mission, but also have the ability to be driven by the crew.

14.1 COLLECTION OF THE CARGO The collection rover has a primary objective to bring all the modules in the selected outpost location. The sizing of the power source required has been performed considering the two launch windows for the cargo mission. All the modules have to be in position and working before the crew departure, so this leaves around 640 days for the first window, and only 90 days for the second one. The design case has been considered as having just one rover working in the second launch window in which for contingency reasons eight modules have been landed. Figure 38 shows an example of a contingency launch campaign in the second launch window.

Figure 38 - Contingency launch campaign for the second window

The total distance that one rover has to cover in 90 days, to collect all the modules in the worst case landing ellipses, is around 280 km. This implies a required average velocity of collection of 0.13 kmh-1. The power required for the rover has therefore been evaluated considering that the rover has to move constantly at the average speed while carrying the heaviest module, for example one greenhouse of 30 tons. The total mechanical power can be evaluated as (86):

푃푚푒푐푐 = 푉푎푣푒푟푎푔푒(푚푟표푣 + 푚퐺퐻)푔푚푎푟푠(푓0 + tan 훼) (3) where 푓0 + tan 훼 = 휇 is the friction coefficient evaluated considering a slope of α=30° and a coefficient f0=0.05 (86). The mechanical power has to be transmitted through electrical motors, considering an efficiency of 90%. The total electrical power is intended to be collected by deployable solar arrays that charge Regenerative Fuel Cells (RFCs). The charging of the RFCs is assumed to have a 70% charge efficiency from the solar arrays, which are sized assuming four deployable solar arrays. These solar arrays were sized based on a solar constant of 152 Wm-² and a specific mass of 1.8 kgm-². In this way, the power requested by the collection phase can be given to the rover continuously (24 hours per day) for all the collection period.

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On the basis of these assumptions, 3.3 kW of power is provided to the rover. This comes from 3.7 kW of RFC charge power, which are charged from four 1.9m radius solar arrays in a flower configuration. The total mass of solar arrays is 93.5 kg, assuming an 18% margin.

14.2 COLLECTION OF THE MDV When the MDV lands near the outpost location, it has to be collected by the transport rover and brought to the outpost. As shown in Figure 38, the distance to be covered is around 20 km. If the MDV were to move at the average speed for the cargo phase, it would take around one week. Considering the time line of the mission, the time to bring the MDV to the outpost has been allocated as 6 hours, so the speed of the rover must be increased to around 3.6 kmh-1. This allows a feasible power increase for the rover, while maintaining a sufficiently fast transit.

Figure 39 - MDV Landing Scenario

The rover is assumed to have six wheels to sustain the landing, four of which are motorised and can be activated independently. These four wheels therefore have their own dedicated electrical motor and gear box. As the collection of the MDV is the most power demanding in terms of the electrical power and torque applied to the wheels, with only the RFCs being used to supply power. The solar arrays for the rover have been sized such that they can recharge the RFCs during the cargo collection phase. The total energy required for the rover is 389 kWh, assuming a 35% margin due to uncertainty of the landing position and based on the criticality of this phase. The electrical power is 34 kW, which is required to transport the MDV in 6 hours, with 8.5 kW/wheel. This therefore requires 1136 kg of RFCs, with 18% margin, based on an assumption of 1150 Wh/kg with 10000 hours of life. The wheels have a radius of 0.5m, and have a total mass (for tires and shock absorbers) of 450 kg. The four powered wheels also require 258 kg of electrical motor mass in total (87). The wheels have a speed of 1 rads-1, which was constrained by the collection, and a torque of 7 kNm. The total mobility and landing system therefore has a mass of 3.4 tons, with a 20% margin.

14.3 MANNED CONFIGURATION The rover is capable to autonomously perform the collection phase, but it is also equipped by an unpressurised cockpit that can host two crew members. This allows a manned configuration in which a crew member, wearing the space suit (see Section 8.3), to drive the rover in order to perform exploration or mount a scientific instrument away from the outpost. In order to ensure the operation of the rover both during the manned and autonomous modes of operation, the following components are required:  Inertial Measurement Unit (IMU (88)): provides information on three axes about its position, velocity and acceleration, which enable the rover to make precise vertical, horizontal and yaw movements  Navigation Cameras (NavCams (89)): are black and white cameras which use visible light to gather panoramic and 3D imagery, and work in cooperation with the hazard avoidance cameras to provide a complementary view of the soil  Hazard Avoidance Cameras (HazCams (89)): are black and white cameras which use visible light to gather 3D imagery, which helps to prevent the rover crashing into unexpected obstacles by working with software that allows the rover to “think on its own” making autonomous safety decisions  Guidance Navigation & Control (GN&C (90)): controls the movements of the vehicle

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 Rover Computer Element (RCE (91)): is the brain of the rover, processing all the data coming from the sensors and actuators and using this to check any factors required to keep the rover “alive” and also scheduling and preparing data for communication sessions  Data recorder: stores all the thermal, power, navigation and communication data. This device is already included in communication and data handling system (see Section 15). The total mass of the avionics systems has been evaluated as 35 kg. Busses and cabling are also considered as the 25% of the avionics mass. Moreover, the avionics mass also includes the permanent scientific instruments (see Section4.3.4). Their mass has been evaluated as around 4 kg, which is a sufficiently small mass to not affect the global performance of the rovers. Finally, the chassis has been sized host the RFCs, which are considered to have a specific volume of 200 kWhm-3, and the avionics. Aluminium alloy plates, reinforced by 12 booms that are capable to resist to the landing loads, form the case for the rovers. The overall mass of the rover has been estimated as 7.8 tons, including all subsystems.

PRIMARY STRUCTURE ATCS AEPS MOBILITY SYSTEM COMMUNICATION AVIONICS

Figure 40 - Collection rover mass budget pie chart

14.4 MODES OF OPERATION Table 21 summarises the different modes of operation in which the rover can run. The performance is detailed in terms of maximum daily range and average speed, related to the power source. Mode of Operation Power Source Working Period Average Daily Speed Range Autonomous Collection RFC + Solar 24 hours, continuously 0.13 kmh-1 3.12 km (with module attached) Arrays 3.42 kW Autonomous Exploration RFC + Solar 24 hours, continuously 0.68 kmh-1 16 km (without module attached) Arrays 3.42 kW Autonomous Collection of RFC 7 hours, recharge 3.6 kmh-1 25 km MDV 41.3 kW needed

Autonomous Exploration RFC 24 hours, recharge 3.2 kmh-1 77 km Maximum Range (only 16.2 kW needed Rover) Manned Exploration RFC 4.2 hours, recharge 17.5 kmh-1 74 km Maximum Speed 92 kW needed (Rover + 2 Astronauts in suit) Table 21 - Rover modes of operation

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15 COMMUNICATION AND DATA HANDLING A communication and data handling system has been determined for the both the elements of Mission MILESTONE, and for the long-permanence outpost established on the Martian surface. A large number of communication links for communication over different distances were required.

15.1 DATA RATES AND RATIONALE In order to determine the most appropriate medium for each of the communication links, the required data rate was first determined. The data rates in Table 22 were used to establish a peak data rate for transmission for each link. The required data rates were determined for both during a nominal situation and for during a contingency, where it is foreseen that lower-rate telemetry, tracking and command (TT&C) and audio would be the only data transmitted. Data rates were determined from the International Telecommunication Union (92). Data Type Data Rate for sending (bps) Audio 45,000 Video (high resolution) 2,500,000 Video (low resolution) 200,000 Photo (sending 1 photo 1,667 every 10 minutes) Photo (sending 1 photo a 12 day) TT&C (high resolution) 200,000 TT&C (low resolution for 1,000 contingencies) Scientific Data 250,000 Table 22 - Required data rates for each data type - TT&C is tracking, telemetry and command.

15.2 COMMUNICATIONS ARCHITECTURE Two architectures for the communication system have been determined, one for the cargo mission (Figure 41) and another for the crew mission (Figure 42). The architecture is based on an adaptation of Earth cellular networks for the Martian surface, primarily using 4G LTE links between on-surface elements, and using a number of orbiting elements to relay communication back to Earth. Both surface and orbital elements of the Earth Ground Station are used, in the form of the Tracking and Data Relay Satellites (TDRS) and the Deep Space Network (DSN). Optical communication from Mars to Earth will need to communicate with the TDRS due to the high attenuation that it would suffer from rain in the Earth’s atmosphere, and can then be relayed using a different link type to the Earth's surface.

15.2.1 Cargo Mission The communication system architecture for the cargo mission can be seen in Figure 41. This demonstrates not only the nominal situation, but also a number of contingency communication links. These have been introduced to maximise the communication link to Earth, ensuring 100% availability of the link (other than during solar conjunction). The contingency links operate at the lower data rate discussed in Section 15.1. Only three modules and one transport rover are shown for simplicity; there would actually be 11 modules and two transport rovers. The nominal arrangement is shown with solid lines, with the different contingency links shown as dotted lines. On each element, the boxes are used to identify the types of data that will be required, in the most general terms. The four elements shown represent respectively a speaker, a microphone, a screen and a camera as example items. This is used to show where photos/videos would be made and where they would be shown,

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as this is not necessarily a two-way process. The use of audio is shown here for consistency with the crew mission, but audio transmission would not be used in the cargo mission.

Figure 41 - Cargo mission architecture

In the cargo mission, each module is equipped with an omnidirectional 4G LTE band antenna, which has a mass of 4.3 kg and requires a power of 4.1 W. The transport rovers are equipped with both an omnidirectional antenna and a parabolic antenna (for communication with the communication satellite), which have a total mass of 8.8 kg and require a power of 9.6 W.

15.2.2 Crew Mission The crew mission architecture can be seen in Figure 42, with four levels of contingency shown as well as the nominal situation. As previously, each contingency link will operate at the lower data rates shown in Section 15.1, with the exception of the link between the CIV and the Earth Ground Station which will transmit data at the same rate as from the Communication Satellite to the CIV (6.63 Mbps). The three Communication Satellites are also planned to transmit data between themselves at this rate via Ka-band communications. During the crew mission, two contingencies are envisaged to ensure that the crew do not become stranded and unable to communicate with Earth. These are shown as contingencies two and four (see Figure 42), and prevent single point failures in the on-Mars communication network.

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Figure 42 - Crew mission architecture

Mission Element Antenna Type Antenna Diameter (m) EVA Crew Omnidirectional - Omnidirectional - Exploration Vehicle Parabolic (4G LTE) 0.5 Module Omnidirectional - Omnidirectional - Outpost Control Centre Parabolic (Ka-band) 2 Descent Module Omnidirectional - Ascent Module Omnidirectional - Omnidirectional - CIV Omnidirectional (Ka-band) - Parabolic (optical) 0.35 Table 23 - Communication antennae for each element of the crew mission. Where not stated, omnidirectional antennae are for 4G LTE.

15.2.3 Communication Satellite Design The Communication Satellites provide a key part of the Mission MILESTONE communication network. The Communication Satellites will have four antennae: a 0.35m diameter optical antenna for communication with Earth; a 2m diameter parabolic Ka-band one for communication with the Outpost Control Centre; a 2m diameter parabolic Ka-band antenna for communication with the other communication satellites and the CIV; and a 3m diameter parabolic 4G LTE antenna for use in the cargo mission. As the frequency of 4G LTE is only approximately 800 MHz, from an Areostationary orbit the parabolic antenna has a footprint of over 10000 km (calculated based on (93)), which will be more than sufficient to cover the range of exploration without reorientation being required.

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The Communication Satellites will be solar powered, with regenerative fuel cells being used to provide additional battery power (assuming the same solar panels and fuel cells in use as on each module during the cargo mission). The overall mass of each Communication Satellite is therefore 81.4 kg, requiring a power of 105.9 W from the solar panels/fuel cells.

15.3 DATA HANDLING After data rates for each link were established in the overall communication architecture, the data storage for each mission element was established using the periods of use of each link and the amount of each type of data (i.e. audio, video) being sent per day. To size the non-volatile storage, the ingoing and outgoing links for each mission element were analysed and it was assumed that the outgoing link is ineffective so all data produced must be stored. The non-volatile memory size for each mission element has to be capable of storing two weeks of data input without any outgoing data link. It was determined that two weeks of data storage was needed on the assumption that a solar conjunction lasts 8.13 days for Ka-band (94), and therefore two weeks was seen to be an appropriate storage length to allow the sending of both new and backlogged data when the link is restored. The non-volatile memory for the EVA crew is an exception and it is required they would only capable of storing data for one day at a time, because the data could be transferred elsewhere before carrying out more EVAs. In addition there is no non-volatile storage provided on the communications satellites.

15.3.1 Storage The total storage required by each module had a 50% margin memory added to the calculated value, in compliance with the ECSS standards (95). In addition, two storage units will be provided to each mission element, with the same data being written onto each, in order to provide redundancy of data in the case that one of the units were to become damaged. The storage required for each mission element can be seen in Table 24. The storage is foreseen to be provided by off-the-shelf equipment, such as that provided by Airbus Defence and Space (96) and Surrey Satellite Technology (97). Mission Element Storage Needed (GB) EVA 20 Exploration Vehicle 337 Module 348 Outpost Control Centre 1477 Descent Module 90 Ascent Module 134 CIV 155 Communication Satellite 0 Transport Rover 312 Total 5994 Table 24 - Final storage requirements for the mission elements

15.3.2 Computing Power For the on-board computing, the mass and power has been approximated using a current space qualified OSCAR computer by Airbus Defence and Space (98). The mass and power of the OSCAR computer is 5kg and 15W (98). It is intended to have two computers for redundancy. However, these computers will only be used for the processing of communications and additional computers may be needed for general operation of the outpost.

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16 IN-SITU RESOURCE UTILISATION (ISRU)

16.1 ISRU RESOURCE REQUIREMENT Mission MILESTONE will bring all resources necessary for the survivability of the crew for the 60 days of permanence on Mars from Earth. However, due to the intention to create a long permanence outpost, it is intended to provide in-situ resource utilisation (ISRU) for a number of key resources. When the crew of Mission MILESTONE leaves the planet, the ISRU is intended to collect all the resources that are necessary for the next mission, before the departure from Earth of the next crew. From all the previous evaluations it is possible to analyse all the resources that have to be collected, which are shown in Table 25

Activity Water [kg] CH4 [kg] O2 [kg] N2 [kg] Water quantity to activate two GHs 200 Water quantity for the WRS of the two HABs 200 Water quantity for the two Sabatier cycles of the ARS 366 (one per HAB) MAV Propellant 5361 MAV Oxidant 14950 Radiation protection for water shield (2 HABs) 28000 N2 leakage from modules 93 Radiation protection for water shield (LAB) 14000 TOTAL 42766 5361 14950 93 Total per day (to be carried out over 540 days) 79.2 9.9 27.7 0.17 Total per hour 3.3 0.4 1.15 0.007 Table 25 - ISRU resource requirements to prepare for the following mission

It is assumed that there are 540 days before the next crew departure from when MILESTONE crew leaves the outpost, so it has been possible to evaluate the quantity of resources that the ISRU has to collect in terms of kg/day or kg/hour. The ISRU process was chosen on the basis of two main considerations: a large mass of water is required; and the MAV requires large masses of CH4 and O2. Water, CH4 and O2 can all be linked using the Sabatier cycle, so this was chosen to size the ISRU system. Rapp, 2008 (99) was used as a reference system for the ISRU.

16.2 CHEMICAL PROCESSES

To accomplish the production listed above, the ISRU must extract CO2 from the Martian atmosphere and water from the Martian soil using the Sabatier and electrolysis processes. Figure 43 shows a schematic of the water and methane production of the ISRU. The water is extracted from the Martian soil using an excavator system. Through the electrolysis process, the water is decomposed into oxygen and molecular hydrogen as shown in the following formula:

퐻2푂 → 퐻2 + 푂 (4) The oxygen is stored in the MAV tanks using a liquidiser system in order to reduce the storage space and to avoid gas leaks. The molecular hydrogen is used in the Sabatier process. Considering the amount of molecular hydrogen required for the production of methane, the process produces 1.65 kgh-1 of oxygen, greater than the 1.15 kgh-1 required for the fuel combustion of the MAV. This extra oxygen can be stored and exploited by the crew. The Sabatier reaction is described as follows:

퐶푂2 + 4퐻2 → 퐶퐻4 + 2퐻2푂 (5)

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Figure 43 - The ISRU schematic process for the production of liquid oxygen and methane

The carbon dioxide is obtained from the Martian atmosphere and the molecular hydrogen from the electrolysis -1 -1 process. The products of the reaction are methane and water. Using 1.14 kgh of CO2, 0.41 kgh of methane and 0.93 kgh-1 of water can be produced. The 0.93 kgh-1 of water produced is not enough to sustain the water demand of 5.45 kgh-1 required by the outpost. In order to reach the demand, the ISRU extracts more water from the Martian Soil which is directly passed into the water cycle, after passing through a filtering and contamination system.

16.3 ISRU BUDGET In order to size the ISRU system, the quantities stated in Section 16.1 are combined with two main assumptions: the ISRU is working 24 hours per day; and it extracts a percentage of kg of water per kg of Martian soil equal to 6% (99). Table 26 summarises the values used to estimate the mass and the power of the ISRU plant, based on the mass of each resource to be produced per hour. From Table 26 it is clear that water excavation and extraction is the most power intensive process, however it is crucial to enable a future mission. A fully fuelled, capable of being landed, MAV (section 13), has been demonstrated. However a method for delivering full water for a 500 day mission has not been developed. Process kg per kg material/h kW per kg material/hour Water excavation and extraction 1500 12

CO2 acquisition 50 1,2

CO2 for Sabatier 12.5 0.16 Water electrolysis 11.2 2.4

Liquefying O2 35 1.1

Liquefying CH4 85 3 Table 26 - ISRU plant devices mass and power estimation (99)

After applying a 30% margin on the mass and 20% margin on the power, the ISRU has a mass of 7.5 tons and consumes 60kW of power during operation. This includes a storing system sized in the same manner as for the MAV (Section 13.6). As can be noticed from Figure 44, the largest contribution is given by the water extractor and processor, which has to produce the water for the outpost.

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Liquefying ISRU Mass Budget (LO2, LCH4) Storage 1% 7% Electrolysis Sabatier 0% 0% CO2 acquisition 0%

Water Extraction for Water Sabatier Extraction 27% for the outpost 65%

Figure 44 - ISRU mass budget pie chart

Finally, the long term mission resources demand have been considered, based on the daily resource losses based on the system designs (i.e. gas leakage). The results is shown in Table 27. Since the resources for a manned phase are lower, the ISRU plant has been sized for the worst case scenario, i.e. the values seen in Table 25.

Activity Water [kg] CH4 [kg] CO2 [kg] N2 [kg] Human water cycle 44 Growing plants 31.2 Sabatier water for O2 2.5 Requested by plants 0.49 Daily leakage of Modules 0.172 Production/day 77.7 0 0.49 0.172 Total/h 3.23 0 0.2 0.007 Table 27 - ISRU resource requirements for a long stay manned mission

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17 TECHNOLOGY OUTPUT For Mission MILESTONE, a number of technologies which are not yet fully developed have been identified that are required for the mission. The maturity of a technology is stated by the technology readiness level; with TRL 1 being the least mature and TRL 9 technology being at the highest level of maturity. Seen below is a detailed breakdown of the technology readiness levels, as defined in the ESA Technology Readiness Level Handbook for Space Applications. (100)  TRL 1 – Basic principles observed and reported  TRL 2 – Technology concept and/or application formulated  TRL 3 – Analytical and experimental critical function and/or characteristic proof-of concept  TRL 4 – Component and/or breadboard validation in laboratory environment  TRL 5 – Component and/or breadboard validation in relevant environment  TRL 6 – System/subsystem model or prototype demonstration in a relevant environment (ground or space)  TRL 7 – System prototype demonstration in a space environment  TRL 8 – Actual system completed and “flight qualified” through test and demonstration (ground or space)  TRL 9 – Actual system “flight proven” through successful mission operations For Mission MILESTONE, the technology readiness level for systems and sub-systems needs to be at least TRL 8 because it involves human spaceflight. As a result a number of technologies that need to be matured before the launch for Mission MILESTONE have been categorised into mission enhancing, mission enabling and mission critical technologies (Section 17.1). The current timeline for maturing technologies were identified using the ESA Exploration Roadmaps 2015 and the 2015 NASA Technology Roadmaps (101) (43). Mission enhancing technologies are classed as such if it would enhance the efficiency of the mission but do not determine the feasibility of Mission MILESTONE. Two examples enhancing technologies are shown in Table 28. A complete list of the mission enhancing technologies can be found in Appendix D: Technology Readiness Level Evaluation. Minimum Time Description/ Capability Current Final to Mature Reference Technology Performance Goal TRL TRL Technology Agency (Years) Recycling of grey water and urine, Water with the possibility of also combining recycling the water from the waste 3 5 5 ESA system management. High efficiency and low maintenance desired. Mars Pressure Maintain mobility with increased Garment capability over current extravehicular 2 6 7 NASA System mobility unit. (PGS) Layup Table 28 - Examples of enhancing technologies for Mission MILESTONE

Mission enabling technologies are classed as such if a future 500 day missions would not be feasible without. Examples of enabling technologies are shown in Table 29. A complete list of mission enabling technologies can be found in Appendix D: Technology Readiness Level Evaluation.

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Minimum Time Description/ Capability Current Final to Mature Reference Technology Performance Goal TRL TRL Technology Agency (Years) Waste Recycling and disposal of organic management waste from the greenhouse and 3 8 6 ESA /batch future integration with human waste. incinerator Advanced Pressurized structure designs with shielding integrated micro-meteoroid orbital debris, radiation, and permeability structures 3 8 7 ESA (radiation, protection, electrical harnessing, meteoroids, thermal control, and sensor dust) subsystems. Table 29 –Examples of enabling technologies for Mission MILESTONE

17.1 MISSION CRITICAL TECHNOLOGIES The mission critical technologies were also identified, and can be seen in Figure 45. Technologies are classed as mission critical if the 60 day surface stay of Mission MILESTONE would not be feasible without. The mission critical technologies must have a TRL 8 by 2035 to be ready in time for Mission MILESTONE. As can be seen in Figure 45, there are several technologies that could not be identified in the ESA Exploration Roadmaps 2015 but could be found in the 2015 NASA Technology Roadmap. However, particular attention should be paid to the need for an ascent capsule as shown in Figure 45. The complete unit of an ascent capsule was not found in either the ESA Exploration Roadmaps 2015 or the 2015 NASA Technology Roadmap (101) (43).

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Figure 45 - Mission critical technologies TRL roadmap

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18 CONCLUSIONS Mission MILESTONE will be a short duration mission to the surface of Mars, landing a crew of six humans on the Martian surface in order to establish a long-term outpost for future exploration. The mission will take the form of a split mission, with the cargo being launched in 2039 and 2041. The cargo mission will deliver 14 modules to the Martian surface and three communication satellites to Areostationary Orbit. The Martian surface elements are composed of two habitable modules, two greenhouses, a laboratory, two EVA Modules, a node, two rovers, a Mars Descent Vehicle, a Mars Ascent Vehicle and associated EVA module, an ISRU plant, and a power plant. The crew mission will follow in 2042, once the modules have been successfully assembled at the chosen outpost location, allowing a 60-day period of human surface operations and exploration. The outpost established by Mission MILESTONE is intended to allow future long-duration exploration, with an aim to reducing the dependence on Earth for resources. The presence of a greenhouse and an ISRU plant are key to this reduction in reliance, as they will provide food, water and oxygen which are crucial to human survival. The scientific operations taking place in Mission MILESTONE are led by four key science objectives:  To study the physiological and psychological impact of a Mars surface mission;  To perform in-situ investigations into the Martian environment to support future exploration;  To collect data to support the identification of resources for future human missions; and  To study the origins and evolution of Mars. The science objectives defined the scientific architecture of the mission, in particular introducing a number of static landers to allow for scientific operations over a larger portion of the Martian surface, and a number of entry probes to better analyse the atmosphere in order to verify the entry, descent and landing process. In order for Mission MILESTONE to be feasible, a number of technologies identified must be matured, with a particular focus on the mission critical technologies. However, given the proper time and resources, Mission MILESTONE could be a robust and important step in human space exploration.

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50. Law, Jennifer. Planetary Protection For Human Exploration Missions: A Flight Surgeon's Perspective. s.l. : NASA, 2015. 51. National Aeronautics and Space Administration. Advanced Life Support Program - Requirements Definition and Design Considerations. [Online] Crew and Thermal Systems Division, 1996. [Cited: 16 July 2015.] https://taskbook.nasaprs.com/peer_review/prog/old/ALSREQ96.html#Heading110. 52. R.N. Schaezler, A.J. Cook, D.J. Leaonard, A. Ghariani. Trending of overboard leakage of ISS cabin atmosphere. 53. National Aeronautics and Space Administration. HSF - The Shuttle. [Online] Human Spaceflight, 2002. [Cited: 16 July 2015.] http://spaceflight.nasa.gov/shuttle/reference/shutref/orbiter/eclss/cabinpress.html. 54. NASA. Life Support Baseline Values and Assumptions Document. March 2015. 55. Lobascio, Cesare and Lamantea, Matteo. Environmental Control and Life Support, Slides of the course in Human Space System II. 56. Space Based Fires & Suppression Systems. Thomas, John W. April 2013. 57. Wieland, P. O. Designing For Human Presence in Space: An Introduction to Environmental Control and Life Support Systems (ECLSS). s.l. : NASA Marshall Space Flight Center, 2005. NASA/TM—2005–214007. 58. Lobascio, Cesare. Environmental Control & Life Support Lectures. s.l. : Politecnico di Torino, 2015. ECLS – 6 (WRM). 59. Carter, D. Layne. Status of the Regenerative ECLSS Water Recovery System. s.l. : NASA, Marshall Space Flight Center, 2009. 2009-01-2352. 60. Regeneration of water at space stations. Grigoriev, A.I., et al. s.l. : Acta Astronautica, 2011, Vol. 68. 61. Alternative Water Processor Test Development. Pickering, Karen D., et al. s.l. : 43rd International Conference on Environmental Systems, 2013. 62. Maxwell, Sabrina and Drysdale, Alan. Assessment of Waste Processing Technologies for 3 Missions. Society of Automotive Engineers. 2001. 63. Turner, Mark and Fisher, John. Generation 2 Heat Melt Compactor. Tucson, Arizona : 44 international Conference on Environmental Systems, 2014. 64. Stilwell, Don, Boutros, Ramzy and Connolly, Janis. H. Crew Accommodations. [book auth.] Wiley Larson and Linda Pranke. Human Spaceflight: Mission Analysis and Design. 1999. 65. Mars or Bust, Inc. Martian Habitat Design. 2003. 66. Williams, David. R. HRET Human Adaptation and Habitability Counter Measurements. 67. Hamilton, D. Autonomous Medical Care for Exploration Class Space Missions. 68. Ross, Amy. Z-1 Prototype Space Suit Testing Summary. Houston, Texas : s.n., 2012. 69. Barta, D. J. Next Generation Life Support Project: Development of Advanced Technologies for Human Exploration Missions. Houston, Texas : s.n., 2012. 70. NASA. 2015_nasa_technology_roadmaps_ta_6_human_health_life_support_habitation. [Online] 2015. [Cited: 9 06 2015.] http://www.nasa.gov/sites/default/files/atoms/files/2015_nasa_technology_roadmaps_ta_6_human_h ealth_life_support_habitation.pdf. 71. Boyle, R., M., et al. Suitport Feasibility - Human Pressurized Space Suit Donning Tests with the Marman Clamp and Pneumatic Flipper Suitport Concepts. Houston, Texas : s.n., 2012. 72. Hublitz, Inka. Engineering concepts for inflatable Mars surface greenhouses. s.l. : Division of Astronautics, Technische Universität München, Germany. 73. Crop productivities and radiation use efficiencies for bioregenerative life support. Wheeler, R M, et al. s.l. : Advances in Space Research, 2008, Vol. 41. 74. Anderson, Molly S, et al. Life Support Baseline Values and Assumptions Document. s.l. : NASA, 2015. NASA/TP-2015-218570.

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75. Osram Sylvania. ZELION® HL Horticulture LED Grow Light Fixtures. [Online] Osram Sylvania, 2013. [Cited: 09 September 2015.] http://www.osram-americas.com/en-us/products/display-optic- specialty-lighting/Pages/ZELION-HL-Horticulture-LED-Fixtures.aspx. 76. Issue 01: NFT Culture. Hydroponics. [Online] [Cited: 24 04 2015.] http://www.hydroponics.com.au/issue-04-nft-culture/. 77. Engineering Verification of the Biomass Production Chamber. Prince, R P, et al. Houston, Texas : NASA Conference Publication 3166, 1988, Vol. 1. 78. Nutrition Facts and Analysis. SELFNutritionData. [Online] Condé Nast. [Cited: 23 April 2015.] http://nutritiondata.self.com/. 79. Alicia M. Dwyer Cianciolo, Jody L. Davis, David R. Komar, Michelle M. Munk, Jamshid A. Samareh, Julie A. Williams-Byrd, and Thomas A. Zang. Entry, Descent and Landing Systems Analysis Study: Phase 1 Report. 2010. NASA/TM-2010-216720. 80. Ferrero, Alberto. CALCULATION AND OPTIMIZATION OF AERODYNAMIC COEFFICIENTS FOR LAUNCHERS AND RE-ENTRY VEHICLES. Lisbon : Instituto Superior Tecnico, 2014. 81. mola.gsfc.nasa.gov. [Online] 82. Mars Ascent Propulsion Trades with Trajectory Analysis. Whitehead, John. C. Fort Lauderdale, Florida : Joint Propulsion Conference and Exhibit, 2003. AIAA-2004-4069. 83. A.Balanis, Constantine. Antenna Theory, Analysis and Design, III edition. s.l. : wiley, 2005. 84. Autonomous Rendezvous and Docking Technologies - Status and Prospects. Wertz, James. R and Bell, Robert. Orlando,Florida : SPIE AeroSense Symposium, Paper No 5088-3, 2003. 85. Kittel, P, Salerno, L.J and Plachta, D.W. Cryocoolers for Human and Robotic Missions to Mars. s.l. : Kluwer Academic/Plenum Publishers, 1999. 86. Genta, Giancarlo. Introduction to the Machenism of Space Robots. s.l. : Springer, 2012. ISBN 978- 94-007-1795-4. 87. Greiffenberger, ABM. Motors for Electric Vehicle. Austria : s.n., 2015. 88. Northrop Grumman. LN-200S Inertial Measurement Unit (IMU). s.l. : Northrop Grumman Systems Corporation, 2013. 25441_022013. 89. Mars Exploration Rover Engineering Cameras. Maki, J. N., et al. s.l. : Journal of Geophysical Research, 2003, Vol. 108. 90. Berkelman, Peter, et al. Design of a Day/Night Lunar Rover. Pittsburgh, Pennsylvania : Carnegie Mellon University, 1995. CMU-RI-TR-95-24. 91. BAE Systems. RAD750 radiation-hardened PowerPC microprocessor. Manassas, Virginia : BAE Systems, 2008. PUBS-08-B32-01. 92. International Telecommunication Union. Spectrum for IMT. [Online] [Cited: 01 July 2015.] https://www.itu.int/ITU-D/tech/MobileCommunications/Spectrum-IMT.pdf. 93. University of Hawaii Department of Physics and Astronomy. Antenna Introduction/Basics. [Online] [Cited: 08 July 2015.] http://www.phys.hawaii.edu/~anita/new/papers/militaryHandbook/antennas.pdf. 94. Williams, W. Dan, et al. High-Capacity Communications From Martian Distances. s.l. : National Aeronautics and Space Administration, 2007. NASA/TM-2007-214415. 95. European Cooperation for Space Standardization. Space Engineering: Software Engineering Handbook. s.l. : ECSS Secretariat, ESA-ESTEC, 2013. ECSS-E-HB-40A. 96. Airbus Defence and Space. Solid State Recorders for Space Applications. s.l. : Airbus Defence and Space, 2014. 97. Surrey Satellite Technology Ltd. Mass Memory Unit Data Sheet. s.l. : Surrey Satellite Technology Ltd, 2014. 0207694 v001. 98. Airbus Defence and Space. OSCAR: compact, powerful and versatile On Board Computer based on LEON3 core. s.l. : Airbus Defence and Space, 2014. 99. Rapp, Donal. Human Missions to Mars: Enabling Technologies for Exploring the Red Planet. s.l. : Springer, 2008. ISBN 978-3-540-72939-6.

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100. European Space Agency. Technology Readiness Levels Handbook for Space Applications. s.l. : European Space Agency, 2008. TEC-SHS/5551/MG/ap, Iss. 1 Rev. 6. 101. —. ESA Exploration Technology Roadmaps. 2015. 102. al., Jenkins B. M. et. Combustion properties of biomass. s.l. : Fuel Processing Technology, 1998.

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APPENDIX A: MASS BUDGETS Subsystem Mass (kg) Primary Structure 13732 Secondary Structure 1454 ECLSS 6390 ATCS 1215 AEPS 1002 C&DH 34.8 Utilities 5606 Total 29496 Table 30 - Habitable module mass budget

Subsystem Mass (kg) Primary Structure 13146 Secondary Structure 1454 ECLSS 3197 ATCS 2289 AEPS 1621 C&DH 21.6 GH Subsystem 7157 Total 28886 Table 31 - Greenhouse mass budget

Subsystem Mass (kg) Primary Structure 14292 Secondary Structure 1454 ECLSS 4366 ATCS 2930 AEPS 1681 C&DH 34.8 Science Payload 4969 Total 29727 Table 32 - Laboratory module mass budget

Subsystem Mass (kg) Primary Structure 5680 Secondary Structure 429 ECLSS 1064 ATCS 318 AEPS 318 C&DH 21.6 Total 7830 Table 33 - EVA module mass budget

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Subsystem Mass (kg) Primary Structure 6677 Secondary Structure 454 ECLSS 2827 ATCS 420 AEPS 297 C&DH 28.5 Space Suits 526 Total 11229 Table 34 - Mars descent vehicle (MDV) mass budget

Subsystem Mass (kg) Primary Structure 464 ECLSS 2237 ATCS 1632 AEPS 3403 C&DH 37 Avionics 28 Total 7801 Table 35 - Rover mass budget

Subsystem Mass (kg) Capsule 3931 Crew 540 Samples 165 Rocket 1998 Propellant 14238 Total Mass Landed 26651 Total Mass Launched from Mars Surface 20872 Table 36 - Mars Ascent Vehicle mass budget

Subsystem Mass (kg) Primary Structure 4208 Secondary Structure 331 ECLSS 249 ATCS 0 AEPS 0 C&DH 0 Space Suits 752 Total 5540 Table 37 - Mars Ascent Vehicle EVA module mass budget

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APPENDIX B: POWER BUDGETS

Subsystem Power (W) ECLSS 5591 ATCS 411 C&DH 119 Utilities 12540 Total 18662 Table 38 - Habitable module power budget

Subsystem Power (W) ECLSS 2368 ATCS 578 C&DH 39 GH Subsystem 21707 Total 24692 Table 39 - Greenhouse module power budget

Subsystem Power (W) ECLS 1068 ATCS 61 C&DH 39 Total 1168 Table 40 - EVA module power budget

Subsystem Power (W) ECLSS 1068 ATCS 121 C&DH 97 Total 1286 Table 41 - Node power budget

Subsystem Power (W) ECLS 1068 Utilities 1746 ATCS 86 C&DH 64 Total 2964 Table 42 - Mars Descent Vehicle power budget

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Subsystem Power (W) ECLSS 1492 ATCS 555 C&DH 30 Utilities 3110 Science Payload 17326 Total 22513 Table 43 - Laboratory power budget

Subsystem Power (W) Water extraction 50700 MAV propellant Production 9100 Storage 296 Total 60096 Table 44 - ISRU plant power budget

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APPENDIX C: SCIENCE REQUIREMENTS MATRIX

Objective 1: To study the physiological and psychological impact of a Mars surface mission Science Objectives Science Investigations Measurement Objectives To analyse the effect of reduced gravity 1.1.1.1 environment on Human Skeletal Muscle (Biopsy) To analyse the effect of long-term lung 1.1.1.2 function in reduced gravity environment To monitor electrical 1.1.1.3 activity of the cardiovascular system To monitoring the musculoskeletal, biomechanical and neuromuscular human physiology to better understand the effects 1.1.1.4 of reduced gravity environment on the muscular system (To Measure the strength To analyse effects on 1.1.1 of isolated muscle specific organs To study the groups in arm and physiological impact of legs) 1.1 a Mars mission on the To measure the human body changing in 1.1.1.5 Crewmember body mass To monitor the 1.1.1.6 hormonal activity and demineralisation To monitor the risk of chronic illness by 1.1.1.7 performing periodical auscultations of the body To determine the levels of radiation received to the skin, eyes and blood- 1.1.1.8 forming organs of crewmembers and ways to mitigate exposure To monitor the electrical activity of the To analyse effects on 1.1.2 1.1.2.1 brain, the neuro- neurological functions vestibular control posture, balance and

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motion sensory coordination To measure the effect of the exposure of crewmembers to 1.1.2.2 cosmic radiation on brain activity and visual perception. To measure hematologic and basic metabolic panels, 1.1.3.1 blood gases, cardiac and liver markers and To analyse the 1.1.3 urine analytes biochemical effects To analyse tissue growth and factors 1.1.3.2 influencing the tissue growth in reduced gravity environment To evaluate the behaviour of the CM 1.2.1.1 under stressful To analyse the conditions 1.2.1 emotional effects on To evaluate the mood humans of the CM and 1.2.1.2 estimate the effects on the performances To identify and characterize To study psychological interpersonal and 1.2.2.1 1.2 impact of a Mars cultural factors that mission on humans may affect the CM interactions To identify and To analyse the social 1.2.2 characterize dynamics of the crew interpersonal and cultural factors that 1.2.2.2 may affect the interactions between the CM and the ground support personnel To study plants adaptation to the 1.3.1.1 Martian environment (height, surface of leaves, etc.) To study the impact of To study statistical the Mars outpost To study the growth and 1.3 1.3.1 1.3.1.2 plants mutations environment on the life morphology of plants (linked to radiation) of plants To study the Chlorophyll types ratio 1.3.1.3 and quantities, to determine plants spectrum absorption

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To exploit the feasibility of using 1.3.2.1 Martian soil in pots as substrate (nutrients provided) To exploit the feasibility of using 1.3.2.2 Martian soil in pots as full growth medium (no nutrient provided) To exploit the To study and exploit the possibility of growing possibility of growing 1.3.2 directly on Mars plants using the Martian 1.3.2.3 surface (water, soil nutrients, atmosphere provided) To study properties and possible toxicity of 1.3.2.4 the crops cultivated in the three Martial soil experiments To study possible DNA mutation in all 1.3.2.5 the three Martian soil growth experiments To determine the quantity and quality of 1.3.3.1 the seeds produced by the plants in the GH To study the ("Martian") seeds 1.3.3.2 germination and plant To study the growth 1.3.3 development of seeds in To determine the DNA the Martian environment 1.3.3.3 mutation from the parent seeds To repeat the three test above with the seeds born from 1.3.3.4 plants cultivated using Martian soil as a substrate To study the crops yield and composition 1.3.4.1 under variation of light, To study the crop water and nutrients 1.3.4 production amount To study the crops 1.3.4.2 organoleptic properties To determine To determine the To study differences between impact of the Mars pharmaceuticals 1.4 1.4.1 1.4.1.1 the effects of the outpost environment administration effects on administration of some on other living beings animals common medicines on

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(bacteria and little Earth and in the animals) Martian outpost environment To study and test the effects of radiation 1.4.1.2 mitigation and bone decalcification mitigation drugs To study how the Martian outpost 1.4.2.1 environment physically affects animal brains To perform tests on To determine the memory and neurological effects of 1.4.2 1.4.2.2 spatial/time the Martian outpost awareness and environment on animals perception To determine behavioural 1.4.2.3 modification respect to Earth To perform "five- senses" tests on animals, to compare To determine the effects 1.4.3.1 them with results on of the Martian outpost Earth and with the 1.4.3 environment on animal modification in senses and sensorial humans ones capability To determine 1.4.3.2 equilibrium (inner ear) modifications To study animals willingness, capability and success of 1.4.4.1 To study the effects of reproducing in the the Martian outpost Martian outpost 1.4.4 environment on animal environment reproduction To study DNA mutation in embryos 1.4.4.2 generated from animal reproduction To study bacteria (Escherichia Coli, Saccharomyces 1.4.5.1 Cerevisiae) growth, To study embryos reproduction & development and evolution/mutation/ada 1.4.5 organism growth in the ptation Martian outpost To study embryos environment development (salamander, mice) 1.4.5.2 and animal growth from Earth frozen embryos

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To study embryos development (salamander, mice) and animal growth 1.4.5.3 from animal reproduction in the Martian outpost environment To study the effect of 1.4.5.4 aging in the Martian outpost environment To determine changes To determine variation in groups dynamics in 1.4.6 in social dynamics in 1.4.6.1 mice, to possibly animals compare with humans

Objective 2: To perform in-situ investigations into the Martian environment to support future exploration Science Objectives Science Investigations Measurement Objectives Determine the energy deposited by the To analyse the radiation 2.1.1 2.1.1.1 Martian radiation environment on the base environment on the base Determine the energy To analyse the radiation deposited by the Characterise the 2.1.2 environment on the base 2.1.1.2 Martian radiation Martian radiation surroundings environment on the environment 2.1 base surroundings

To analyse the Determine the energy

existence of radiation of deposited by the 2.1.3 2.1.1.3 the Martian soil on the Martian soil radiation base on the base Determine the energy To analyse the deposited by the existence of radiation of 2.1.4 2.1.1.4 Martian soil radiation the Martian soil on the on the base base surroundings surroundings Determine the intensity and time To study the variation of Martian 2.2.1 electromagnetic 2.2.1.1 electromagnetic environment on the base environment on the base Determine the To study the Martian To study the intensity and time 2.2 magnetosphere electromagnetic variation of Martian 2.2.2 2.2.1.2 environment on the base electromagnetic surroundings environment on the base surroundings To study the interactions Determine the energy of the Martian deposited by the 2.2.3 2.2.2.3 magnetosphere with Martian radiation radiation environment on a

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zone where the magnetosphere is present

To record the daily 2.3.1.1 weather of the To study the Martian plains meteorological 2.3.1 characteristics on the Mars surface

To record the variation in wind 2.3.1.2 conditions on the Martian surface To study the To study seasonal meteorological 2.3 2.3.2.1 variations of Martian characteristics of Mars temperature

To study seasonal 2.3.2.2 variations of Martian To study seasonal pressure 2.3.2 variations of Martian To study seasonal meteorology 2.3.2.3 variations of Martian humidity To study seasonal variation of Martian 2.3.2.3 atmosphere composition To the study of the To study the atmosphere of Mars composition and from top to bottom, in 2.3.3 2.3.3.1 particulates in the air on a variety of observing the Martian surface modes, including stellar occultation.

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Objective 3: To collect data to support the identification of resources for future human missions Science Objectives Science Investigations Measurement Objectives

Determine the bulk atomic structure of the 3.1.1.1 area of interest to identify the presence of resources

Detect the elemental composition of samples from the area of interest to 3.1.1.2 identify the presence of resources, including hydrogen, To study the oxygen, carbon and composition of the To study the nitrogen Martian surface to composition of the 3.1 3.1.1 identify resources Martian surface within Detect the crystalline necessary to support the area of interest structure of samples human Mars missions 3.1.1.3 from the area of interest to identify the presence of resources

Detect the molecular composition of samples from the area of interest to 3.1.1.4 identify the presence of resources including water, methane and carbon dioxide

Detect hydrated minerals within 3.2.1.1 samples from the area of interest To study the composition of the To study quantity and Martian sub-surface to 3.2 3.2.1 quality of water within identify resources the Martian sub-surface Detect the molecular necessary to support composition of sub- 3.2.1.2 human Mars missions surface samples from the area of interest

Detect the presence 3.2.1.3 of sub-surface ice

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within the area of interest Detect the transportation of gas 3.2.2.1 To study the quantity through the sub- and quality of gaseous surface resources within the Detect the elemental 3.2.2 Martian sub-surface composition from the area of interest to 3.2.2.2 identify the presence of gases including hydrogen, oxygen and nitrogen

Objective 4: To study the origins and evolution of Mars Science Objectives Science Investigations Measurement Objectives To measure the composition of crater 4.1.1.1 features to determine the meteorite source To measure global 4.1.1.2 crater density on the Martian surface To study the To study the current To measure the composition of 4.1 and historical meteor 4.1.1 seismic activity meteorites on the 4.1.1.3 flux and sources caused by impacts Martian surface from meteorites To track the production of impact craters in order to 4.1.1.4 determine correlation with seismometer readings To detect the elemental composition at To study the 4.2.1.1 composition and depth different depths of layers within the throughout the 4.2.1 Martian sub-surface Martian subsurface To detect the crystalline structure at To collect data on the 4.2.1.2 different depths Martian subsurface to throughout the 4.2 improve understanding Martian subsurface of the evolution of To perform Mars radioactive dating of 4.2.2.1 rock samples from To study the origin of regions of interest Mars and its historic 4.2.2 To perform existence within the radioactive dating on solar system 4.2.2.2 regolith and surface samples from regions of interest

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To perform radioactive dating of 4.2.2.3 samples from the Martian sub-surface from areas of interest To determine the organic presence in 4.3.1.1 selected samples (e.g. SFFs) To determine the 4.3.1.2 organic composition of selected samples To collect localised To detect the data on the Martian To study the presence of metabolic activity surface and past or current life within 4.3.1.3 4.3 4.3.1 inside selected subsurface to improve the landing region of samples knowledge of the interest specific landing region To detect the organic bio-signature on 4.3.1.4 selected samples and discriminate the sources To sequence the DNA 4.3.1.5 like sample eventually been found To characterise the 4.4.1.1 crustal thickness and To study seismic activity structure of Mars 4.4.1 and the sources on the surface of Mars To characterise the 4.4.1.2 crustal composition of Mars To collect data of To characterise the 4.4 seismic activity on the 4.4.2.1 crustal thickness and Martian surface structure of Mars To study seismic activity To characterise the 4.4.2 and the sources within 4.4.2.2 crustal density of the sub-surface of Mars Mars To characterise the 4.4.2.3 internal density of Mars

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APPENDIX D: TECHNOLOGY READINESS LEVEL EVALUATION

19.1 MISSION ENABLING TECHNOLOGIES Minimum Time Description/ Capability Current Final to Mature Reference Technology Performance Goal TRL TRL Technology Agency (Years) Recycling and disposal of Waste organic waste from the management/batch greenhouse and future 3 8 6 ESA incinerator integration with human waste. Means to transform and Material increase the technology management and value of materials brought 3 8 10 ESA recycling from Earth to minimise waste. High efficiency LEDs for High Efficiency providing light at optimal 4 - - - LEDs plant growth. Lightweight, anti-microbial clothing for extended use. Advanced clothing 5 8 2 NASA Reduces clothing related waste and lint production. Remove odour from clothes, Laundry freshening improve hygiene and 3 7 4 NASA system duration of clothing life. Reuseable wet wipes for Reuseable housekeeping. Can be housekeeping washed for waste reduction, 2 8 1 NASA wipes but maintain hygienic wetness. Water containment and privacy structures added to a Full body shower 6 6 4 NASA full body washing, lit and ventilated shower. Freezing unit for storage of Food storage excess crop production and 2 - - - freezers uneaten food. Table 45 - Life support and asset protection technologies

Minimum Time Description/ Capability Current Final to Mature Reference Technology Performance Goal TRL TRL Technology Agency (Years) Thermal and electric energy Hybrid thermal- can be collected for the same electric solar solar panels which can rise 1 8 4 NASA concentrator the temperature of a fluid panels thanks to solar concentrator

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Dust removal Remove the dust periodically system for solar accumulated on the array 2 - - - array surface

System able to deploy solar Solar array array on Mars automatically deployment system 2 - - - or with the help of a Martian for Martian surface rover Table 46 - Novel energy production and storage technologies

Minimum Time Description/ Capability Current Final to Mature Reference Technology Performance Goal TRL TRL Technology Agency (Years) Robotic arm for planting seeds, checking the “vegetable Greenhouse readiness level” and harvesting tending robotic 3 - - - crops. Able to manipulate growing arm trays position and move around the module. Table 47 - Automation and robots technologies

Minimum Time to Description/ Capability Current Final Reference Technology Mature Performance Goal TRL TRL Agency Technology (Years) Pressurized structure designs with integrated micro-meteoroid orbital Advanced shielding debris, radiation, and structures (radiation, 3 8 7 ESA permeability protection, meteoroids, dust) electrical harnessing, thermal control, and sensor subsystems. Landing mechanisms which are capable of Low speed landing surviving multiple mechanisms 2 5 3 ESA attempts/hopping, and (passive/active) landing on poorly known soil stiffnesses. Allow the module to land on tires, allowing further Lightweight elastic tire mobility, which have a 2 - - - material sufficiently low pressure to minimise damage to the Martian surface. Table 48 - Advanced structures and mechanisms

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Minimum Time to Description/ Capability Current Final Reference Technology Mature Performance Goal TRL TRL Agency Technology (Years) A device that reduces the volume of trash (paper and plastic products, residual foods, beverages Trash management and containers, used 4 6 4 NASA system housekeeping and hygiene wipes) and stabilizes it for safe storage and disposal. System for removal of oxygen produced in excess from the Oxygen Removal greenhouse and for 3 - - NASA System purging oxygen from the outpost during unmanned operation. Table 49 - Systems and processes technologies

Minimum Time to Description/ Capability Current Final Reference Technology Mature Performance Goal TRL TRL Agency Technology (Years) A miniature gamma ray spectrometer Sub-surface Gamma (approximately 0.2 kg) for 2 - - - Ray Spectrometer use on a sub-surface mole. A stethoscope to enable both medical and non- Stethoscope medical crew members to 6 8 5 NASA listen to internal ausculation sounds. In-flight analysis platform including at least blood Suite of laboratory and metabolic panels, analysis platforms and 6 8 5 NASA blood gases, cardiac and assays liver markers and urine analysis. Handheld Raman spectrometer for use by Mobile Raman the crew on EVA for 5 - - - Spectrometer qualitative sample analysis. X-ray diffractometer for X-ray Diffractometer use in a Mars-based 5 - - - laboratory.

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Gamma ray neutron spectrometer for use on Gamma Ray Neutron both static landers and 6 - - - Spectrometer rovers on the Martian surface. Miniature gamma ray neutron spectrometer for Drill/Mole Head use on drills/moles to Gamma Ray Neutron 3 - - - allow for subsurface Spectrometer measurements without disturbing samples. Mole for obtaining Mole (5m) subsurface samples from 3 - - - a depth of up to 5m. Drill for use on rovers to Drill (5m) obtain samples from a 4 - - - depth of up to 5m. Table 50 - Science instrumentation technologies

19.2 MISSION ENHANCING TECHNOLOGIES Minimum Time to Description/ Capability Current Final Reference Technology Mature Performance Goal TRL TRL Agency Technology (Years) Recycling of grey water and urine, with the possibility of also Water recycling combining the water from 3 5 5 ESA system the waste management. High efficiency and low maintenance desired. Increase time between crew maintenance task, Closed loop CO2 reduce power, control 2 6 4 NASA recovery system CO2 concentrations to lower levels Increase the percentage Oxygen recovery 3 8 9 ESA of O2 recovered from CO2 Maintain mobility with Mars Pressure increased capability over Garment System 2 6 7 NASA current extravehicular (PGS) Layup mobility unit. Table 51 - Life support and asset protection technologies

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Minimum Time to Description/ Capability Current Final Reference Technology Mature Performance Goal TRL TRL Agency Technology (Years) Increase in cryocooler Cryocooler and performance; zero boil-off 2 8 5 ESA cryofluid management system demonstration and qualification. Increasing cooling capacity to greater than 150 W at 90 K, with High capacity 90 K specific power of less 4 6 2 NASA cryocooler than 10.6 W/W and specific mass of less than 0.35 kg/W. Table 52 - Thermal, TPS and ATD aspects technologies

Minimum Time to Description/ Capability Current Final Reference Technology Mature Performance Goal TRL TRL Agency Technology (Years) Miniaturised antennas will Miniaturised antennas reduce the mass required for on-surface for small planetary assets, 3 8 7 ESA planetary such as landers or communication probes. Table 53 - Communication, remote sensing and imaging technologies

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