ESA UNCLASSIFIED - For Official Use estec

European Space Research and Technology Centre Keplerlaan 1 2201 AZ Noordwijk The Netherlands T +31 (0)71 565 6565 F +31 (0)71 565 6040 www.esa.int

Request for Information:

European Industrial heritage that could be applied to small Mars spacecraft platforms

Prepared by Mars ExPeRT Team

Reference ESA-E3P-EXPE-SOW-015 Issue/Revision 1.0 Date of Issue 01/05/2020 Status Final

ESA UNCLASSIFIED - For Official Use

1 INTRODUCTION 1.1 Background information Space19+ has approved European contributions to the NASA-led Mars Sample Return Campaign, planned for launch in 2026. A Concurrent Design Facility (CDF) study has recently been completed by the European Space Agency (ESA) Human and Robotic Exploration (HRE) Directorate’s Exploration Preparation, Research and Technology (ExPeRT) team which has investigated potential small mission architectures that would further ESA Exploration objectives at Mars post-MSR (in the 2030's) or potentially in- parallel with MSR (if sufficiently low cost).

ESA is therefore releasing this Request For Information to survey the European industrial landscape regarding heritage spacecraft equipment, platforms and relevant spacecraft development experience that could be applied to potential future small and low-cost orbiter missions to Mars. This information will contribute to future strategic planning for Mars exploration within the framework of the European Exploration Envelope Programme (E3P) and interaction with relevant Stakeholders including international partners and ESA Member States.

Information on potential scientific payloads is not requested as part of this RFI.

Important note: All information requested and provided in the frame of this RFI will be used for information and planning purposes only and is not part of an ESA procurement process. This RFI does not bind ESA to any present or future procurements actions nor does it create any rights for respondents in relation to any present or future ESA procurements.

1.2 Reference Documentation The RFI package includes the following reference document:

[RD1] SMARTieS CDF study report

It should be noted that the document listed above can be subject to further evolution and refinement. Furthermore, the CDF study provides reference design concepts to demonstrate feasibility of selected architectures and should not to be considered as prescriptive or pre- empting the overall mission design.

1.3 Mission Architectures The recently completed ESA CDF study SMARTieS identified a set of Mars mission architectures that could potentially meet the following programmatic constraints:

• €125M industrial cost at completion (economic conditions 2022) • Phase B2 kick-off in either Q2 2023 or Q2 2026

Page 2/5 Request for Information: European Industrial heritage that could be applied to small Mars spacecraft platforms Issue Date 01/05/2020 Ref: ESA-E3P-EXPE-SOW-015

ESA UNCLASSIFIED - For Official Use

• Phase B2/C/D in ~4 years • Launch between end 2027 and 2032

Note that no decision has been taken by ESA to undertake any such missions; these studies have been conducted purely for planning and discussion purposes only.

Such programmatic constraints would require a rapid development to flight readiness. As such it is likely that any selected equipment will use existing hardware or be derived from existing developments.

The mission architectures considered in the CDF are described in detail in [RD1] and should be used as general reference cases on which to respond to this request for information. In summary, the mission architectures comprise:

Case 1 Case 2 Mars Communications Constellation Mission Name Mars Science Orbiter Mission Mission Configuration 3 satellite constellation Single satellite Primary: Data relay from surface Primary: Remote sensing missions science investigations Objective Secondary: Remote sensing science Secondary: Data relay from investigations surface missions Mars Trans-AreoSynchronous Orbit Target Orbit Low Mars Orbit (LMO) (TASO) A62 Dedicated launch to Mars A62 Dual Launch to GTO Launch and transfer scenario insertion, chemical propulsion Full Electric Propulsion Transfer on orbit

Science payload mass 7kg (per satellite) 35kg

2 INFORMATION REQUESTED

Information is requested which may be submitted in response to either or both of the mission cases presented. All European spacecraft developers (large scale integrators as well as smaller entities) interested in the subject of this RFI are invited to reply to this request.

The following template should be used as a guide to preparing responses. Responses shall be provided exclusively in electronic format and proprietary information shall be clearly marked.

Section Title Information requested

Page 3/5 Request for Information: European Industrial heritage that could be applied to small Mars spacecraft platforms Issue Date 01/05/2020 Ref: ESA-E3P-EXPE-SOW-015 ESA UNCLASSIFIED - For Official Use

SMARTieS mission Review of the SMARTieS CDF mission architectures (Case 1 architectures review and/or 2) Alternative mission Suggestions for alternative architectures that meet the overall architectures mission and programmatic requirements are welcomed. [optional] Driving requirements Identification of the main driving requirements. Risks and critical Definition of the main risks and critical issues associated with issues meeting the programmatic and technical constraints of the mission. Company heritage Provide examples of relevant company heritage spacecraft platforms/capabilities considered to be largely applicable to the proposed mission architectures. Technical and Identify the key technical and programmatic deltas between programmatic delta company heritage missions and the proposed mission architectures. Technology Indicate potential technology developments that would be developments required to meet mission requirements.

3 SUBMISSION AND EVALUATION

Responses to this RFI should be sent by email to the following addresses by 15 July 2020, 12:00 CET (noon).

[email protected] [email protected] [email protected]

Please include details of a contact person that ESA can contact for further clarification regarding the RFI response, if necessary.

Respondents are asked to prepare submissions against the template given in Section 2. Additional information can be provided as annex. Proprietary information submitted in response to this RFI should be clearly marked.

Once received, ESA will review the submissions and the information will be used to inform future Mars exploration study planning and strategy development. ESA offers to plan for a dedicated meeting with every interested company that will have timely submitted a response to this RFI.

By submitting their responses, respondents agree that ESA may use material received to inform and support its further actions concerning the development of its programmes, which may include releasing part of it in the context of exchanges between ESA and its Member States and of future requests for information or tender actions, subject to the following:

Page 4/5 Request for Information: European Industrial heritage that could be applied to small Mars spacecraft platforms Issue Date 01/05/2020 Ref: ESA-E3P-EXPE-SOW-015 ESA UNCLASSIFIED - For Official Use

• ESA will not release individual RFI responses. • ESA will not release any information that a respondent has clearly identified as being “Proprietary Information” in its submission. Note: If a respondent requires additional non-disclosure arrangements before submitting its RFI response, ESA is ready to discuss this upon specific request.

The submission of RFI proposals does not commit ESA to placing any contract, nor prevents ESA from pursuing or employing those concepts, directly or indirectly, in any present or future ESA activities.

ESA will use any personal data included in the proposals (e.g. names and contact details of the persons having prepared the proposal) exclusively for contacting those persons in case it became necessary in relation to their proposal submission. ESA will not disclose those personal data to third parties without prior agreement of the party that had submitted the proposal.

Page 5/5 Request for Information: European Industrial heritage that could be applied to small Mars spacecraft platforms Issue Date 01/05/2020 Ref: ESA-E3P-EXPE-SOW-015

CDF STUDY REPORT SMARTieS Summary Report for Small Mars Mission Architecture Study

CDF-205(A) April 2020

SMARTieS CDF Study Report: CDF-205(A) April 2020 page 1 of 105

CDF Study Report SMARTieS Summary Report for Small Mars mission ARchiTecture Study

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 2 of 105

This study is based on the ESA CDF Open Concurrent Design Tool (OCDT), which is a community software tool released under ESA licence. All rights reserved.

Further information and/or additional copies of the report can be requested from: S. Vijendran ESA/ESTEC/HRE-XE Postbus 299 2200 AG Noordwijk The Netherlands Tel: +31-(0)71-5653360 Fax: +31-(0)71-5654437 [email protected]

For further information on the Concurrent Design Facility please contact: I. Roma ESA/ESTEC/TEC-SYE Postbus 299 2200 AG Noordwijk The Netherlands Tel: +31-(0)71-5658453 Fax: +31-(0)71-5656024 [email protected]

FRONT COVER

Study Logo for Small Mars Mission Architecture Study

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 3 of 105

STUDY TEAM

This study was performed in the ESTEC Concurrent Design Facility (CDF) by the following interdisciplinary team:

TEAM LEADER D. Binns, TEC-SYE AOCS A. Martinez Barrio, PROGRAMMATICS/ Y. Le Deuff, TEC-MXC M. Sanchez Gestido, AIV TEC-SAG COMMUNICATIONS S. Carrasco Martos, CHEMICAL D. Perigo, TEC-MPC S. Agut Sanz, PROPULSION M. Sousa, TEC-EST CONFIGURATION P. Werk, TEC-MSS ELECTRIC E. Bosch Borras, PROPULSION N. Wallace, TEC-MPE COST G. Cifani, TEC-SYC RISK D. Wegner, TEC-QQD N. Drenthe A. Lopez GS&OPS A. Piris Nino, OPS- SPACE G. Deprez, TEC-EPS OPD ENVIRONMENT F. Piette MISSION ANALYSIS M. Khan, OPS-GFA STRUCTURES A. Saad, TEC-MSS N. Skuppin POWER F. Bausier, TEC-EPM SYSTEMS C. Parfitt, TEC-SYE M. Valencon THERMAL V. Cleren, TEC-MTT

Under the responsibility of: S. Vijendran, HRE-XE, Study Manager A. Mcsweeney, HRE-XE, Study Management Support L. De Backer, HRE-XE, Study Management Support

With the scientific assistance of: C. Orgel, HRE-RS, Study Scientist J. Carpenter, HRE-S, Study Scientist Consultant

With the technical support of: A. Ball, HRE-PEH, Science Payloads S. Kohl, TEC-SPS, Small Satellites

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 4 of 105

R. Walker, TEC-SPC, CubeSats S. Schuster, TEC-MPA, Entry Descent and Landing

The editing and compilation of this report has been provided by: A. Pickering, TEC-SYE, Technical Author

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 5 of 105

TABLE OF CONTENTS

1 INTRODUCTION ...... 9 1.1 Study Background ...... 9 1.2 Study Overview ...... 11 1.3 Study Objectives ...... 12 1.4 Document Structure ...... 12 2 MISSION ARCHITECTURE ...... 13 2.1 Introduction ...... 13 2.2 Programmatic Constraints ...... 13 2.3 Mission Requirements and Design Drivers ...... 14 2.3.1 Common Mission Requirements ...... 14 2.3.2 Design Drivers ...... 15 2.3.3 Design To Cost Approach...... 15 2.4 Selection of Mission Cases ...... 16 2.5 Architecture Options ...... 17 2.5.1 Initial Orbit Injection ...... 17 2.5.2 Launch Vehicles ...... 17 2.5.3 Propulsion Technologies ...... 18 2.5.4 Overview of Potential Launch and Transfer Options ...... 18 2.5.5 Programmatic and Cost Assumptions ...... 20 2.5.6 Technical Assumptions ...... 20 2.6 Mission Architecture Trade-off Results ...... 23 2.6.1 Launcher Assessment ...... 23 2.6.2 Initial Orbit Injection ...... 24 2.6.3 Propulsion Architectures ...... 25 2.6.4 Launch Mass Comparison ...... 26 2.6.5 Transfer Time to Mars ...... 27 2.6.6 Mission Architecture Conclusions ...... 28 2.7 Margin Philosophy ...... 30 2.8 Small Satellite Design Philosophy ...... 30 3 MARS COMMUNICATION CONSTELLATION ...... 33 3.1 Mission Requirements ...... 33 3.2 Mission Analysis...... 33 3.3 Reference Scenario ...... 35 3.4 Systems Analyses and Trade-Offs ...... 36 3.4.1 Spacecraft System Configuration Trade-Off ...... 36 3.4.2 Power Subsystem ...... 42 3.4.3 Communications Performance Analysis ...... 43 3.4.4 Science Instrument Selection ...... 50 3.5 Baseline Design ...... 52 3.5.1 Mission Architecture ...... 52

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 6 of 105

3.5.2 Spacecraft design overview ...... 52 3.5.3 Spacecraft Subsystem Design ...... 53 3.5.4 Configuration ...... 54 3.5.5 System Modes ...... 54 3.5.6 Spacecraft Budgets ...... 55 3.5.7 Equipment List ...... 57 3.5.8 Resources Available to Payload ...... 58 3.5.9 Technology Developments ...... 59 3.6 Programmatics and Risk ...... 59 3.6.1 Requirements ...... 59 3.6.2 Assumptions ...... 59 3.6.3 Baseline Schedule ...... 60 3.6.4 Summary of Risk Register ...... 61 3.7 Cost ...... 62 3.7.1 Methodology ...... 62 3.7.2 Assumptions ...... 62 4 MARS SCIENCE ORBITER ...... 65 4.1 Mission Requirements ...... 65 4.2 Science Objectives ...... 65 4.3 Mission Analysis ...... 66 4.3.1 Orbit Options ...... 66 4.3.2 Orbit Selection ...... 67 4.3.3 Mars Orbit Insertion Trade-Off ...... 67 4.3.4 Mars Orbit Insertion (MOI) Strategy and Lowering to Science Orbit ...... 68 4.4 Reference Scenario ...... 68 4.5 Systems Analyses and Trade-Offs ...... 70 4.5.1 Chemical Propulsion Architecture Trade-Off ...... 70 4.5.2 Communications Performance Analysis...... 71 4.5.3 Science Instrument Selection ...... 75 4.6 Baseline Design ...... 77 4.6.1 Mission Architecture ...... 77 4.6.2 Spacecraft Design Overview ...... 79 4.6.3 Spacecraft Subsystem Design ...... 79 4.6.4 Configuration ...... 81 4.6.5 System Modes ...... 81 4.6.6 Spacecraft Budgets ...... 82 4.6.7 Resources Available to Payload ...... 85 4.6.8 Technology Developments ...... 85 4.7 Programmatics and Risk ...... 86 4.7.1 Assumptions ...... 86 4.7.2 Baseline Schedule ...... 87 4.7.3 Summary of Risk Register ...... 88 4.8 Cost ...... 88 4.8.1 Assumptions ...... 89 5 MARS HARD LANDER ...... 91

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 7 of 105

5.1 Mission Requirements ...... 91 5.2 Reference Scenario ...... 91 5.2.1 Selection ...... 91 5.3 Baseline Design ...... 92 5.3.1 EDL Trajectory ...... 92 5.3.2 EDL System ...... 93 5.3.3 Carrier Vehicle ...... 94 5.3.4 Summary ...... 94 5.4 Alternative Mission Scenarios ...... 96 5.5 Cost ...... 97 5.6 Further Work / Recommendations ...... 97 6 CONCLUSIONS ...... 99 6.1 Achievement of Study Objectives ...... 99 6.2 Further Study Areas ...... 100 6.3 Main Conclusions ...... 101 7 REFERENCES ...... 103 8 ACRONYMS ...... 105

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 8 of 105

This Page Intentionally Blank

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 9 of 105

1 INTRODUCTION 1.1 Study Background ESA’s current Mars exploration programme consists of the flying orbiters Mars Express and the Exomars TGO, while the Exomars rover is planned for launch in 2022. The Nov 2019 ESA Council of Ministers meeting, Space19+, has approved ESA contributions to a Mars Sample Return Campaign, led by NASA, with a launch of the sample retrieval missions planned to occur as early as 2026. ESA’s planned contributions include the Earth Return Orbiter (ERO), the Sample Fetch Rover (SFR) and the Sample Transfer Arm (STA). With Mars Sample Return finally underway, ESA is looking towards the future of European Mars exploration with the horizon goal of participation in an international Human Mars exploration campaign. It is expected that there will still be a number of robotic missions to Mars by international agencies between MSR and human Mars surface missions in order to continue the search of life on Mars as well as prepare the technologies and infrastructure that will be needed and to close scientific knowledge gaps about Mars that will support human exploration. While the vast majority of ESA’s funding for Mars exploration in 2020’s is planned to be invested in Mars Sample Return, there is interest to assess, at Phase 0-level, the possibility of implementing a small mission to Mars in parallel with or soon after the completion of the MSR programme, in order to further the exploration of Mars in areas not directly addressed by MSR. The benefits of targeting a small Mars mission for launch in the late 2020’s are: • Strategic scientific and/or technological knowledge gaps of high value could be filled at a modest cost, in parallel with MSR and much before a future large ESA mission could be envisaged (most likely launch >2033). • Provides additional flight opportunities for exploration science payloads in the 2020’s, following Exomars, since MSR does not. • Concurrent development with MSR (but started slightly later) could allow use of recurring H/W, S/W and ESA/industrial expertise from the recent Exomars and ongoing MSR developments in the small mission => key to keeping costs/risks low for fast implementation. The Rosetta/Mars Express developments provided a successful example of this approach. • The timeframe is more in-line with the planning horizon of international partners • It provides opportunities for smaller ESA Member States to Prime a Mars mission Missions to Mars at the small scale have not been greatly studied within ESA since Mars Express (with the focus on Exomars and MSR), two decades ago. Since then the landscape of technologies (in particular those relevant for small satellite platforms) and launch capabilities (e.g. rideshares) have matured significantly, offering promising new opportunities for low-cost implementations of interplanetary missions. As such, and in order to better understand the scope and constraints of such opportunities, an

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 10 of 105 architectural-level study at in the CDF of such small missions to Mars was requested by the ExPeRT Team of the Directorate of Human and Robotic Exploration. As background it worth defining what is meant by architecture: • Architecture is the fundamental organization of a system, embodied in its components, their relationships to each other and to the environment, and the principles governing its design and evolution (IEEE Std 1471-2000) • Links Systems Engineering & Management by balancing technical and programmatic considerations • It underlies the design’s ability to meet objectives/constraints and satisfy stakeholders by creating a framework of information and options that shows potential to meet objectives, and avoids jumping to a single design point • A successful mission-level architectural study involves addressing, as widely as possible, the trade space of options that could result in an eventual feasible mission design that fits within the key stakeholder constraints across competing factors including: performance, complexity, technical risk, implementation approach, total cost, and cost risk. The key differences between architecture and design are shown below.

Architecture Design

• Dictates what possibilities are • Models, analyses, schematics, allowed, within key constraints, block diagrams, lists, and other while focusing on concepts representations of a design selected to fulfill system objectives • Requirements driven • The broadest framework of the total system (technical and • Starts as iterative process within programmatic) architecture constraints

• Provides the rationale and • What gets built and tested to functional framework for implement the architecture requirements

The focus of this study is on the mission and spacecraft trade-offs and not the detailed design.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 11 of 105

1.2 Study Overview The study considered the following mission cases: • Mission 1: Mars Communications Constellation - Supporting assets on the surface and providing science • Mission 2: Mars Science Orbiter - Science mission driven, remote sensing • Mission 3: Mars Hard Lander/Penetrator - Carrier with a number of small landers released on a ballistic trajectory The focus within the scope of the study was the system level assessment on the first two cases, with the third lander case qualitatively assessed at architecture level. To do this the overall study consisted of 10 sessions shown in Figure 1-1. The first was the kick-off and the final an internal final presentation. In between, 8 design session covered primarily the first two mission cases, the Mars Communication Constellation (MCC) or “Commstellation” and the Mars Science Orbiter (MSO). In parallel to the main study activities the lander case was assessed. All cases are assessed at mission architecture level considering a wide range of possible launch scenarios. The design process was consistent across mission 1 and mission 2 cases. Each mission case starts with a broad assessment of launch and propulsion options, focusing on key sub-systems such as power and propulsion design, with an initial system level allocation for mass. Then a more detailed assessment is made of spacecraft equipment. The needed inventory was checked against the available dry mass for the initially selected launch and mission constraints. Then iterated accordingly.

Figure 1-1: Study Plan and Flow

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 12 of 105

1.3 Study Objectives Objectives of the CDF study are: 1. Provide a portfolio of potential transfers to Mars for launches between 2026 and 2032. 2. Assess the feasibility and technical scope of 1-2 small satellite orbiter or a mother/daughter(s) spacecraft combination missions to Mars, within the cost and schedule constraints. 3. Assess the feasibility and technical scope of delivering small lander(s) to the surface of Mars within the cost and schedule constraints. 4. Design the missions to a maximum cost at completion of 250 MEuros. 5. Investigate existing technologies and equipment for use in Mars missions within the development timeframe. 6. Define the “small satellite approach” to be adopted, considering reliability, FDIR, ECSS tailoring, COTS approach, operations/ground segment, margins philosophy etc. 7. Define schedule, risks and programmatic aspects of the mission compatible with a CM2022 decision and launch between 2026 and 2032. 8. Evaluate the landscape for industrialisation of such small missions, considering alternatives to the European LSIs as Prime Contractor. 9. Highlight the main operational constraints. 1.4 Document Structure Compared to a normal CDF report the layout is different. This is an executive summary report that serves to describe at architecture level the multiple mission and spacecraft concepts. This report is supported by detailed technical presentations, available on request, for spacecraft sub-systems. This report has four main parts: • Mission Architecture – this includes an assessment of the programmatic constraints, identification and definition of common mission requirements, definition of the mission cases and their relevance, discussion of the possible launch approaches and selection, radiation assumptions, and the margin philosophy • Mars Communications Orbiter – in addition to the applicable common mission requirements, the specific communications system level requirements are established. Several system level trades are performed, the constellation configuration, electric propulsion options, the performance of the communications network, and science priorities and instrument complement. • Mars Science Orbiter – The focus in this case is identifying the possible payload for what would be considered the most optimal mission from a programmatic standpoint. The common mission requirements are extended to case specific requirements. Therefore propulsion options trades are included. Instruments are identified with prioritisation. • Mars Hard Lander – This section covers the reference scenario and selection. A high level sizing case is defined including, trajectory, EDL system and assumptions on the carrier spacecraft.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 13 of 105

2 MISSION ARCHITECTURE 2.1 Introduction The initial driving constraints stipulated by the customer for the SMARTieS study are largely programmatic rather than scientific or technical and do not impose any particular kind of mission architecture. With a launch timed to complement the upcoming Mars Sample Return missions at the end of the 2020s, applying these constraints would mean having an opportunity for a European mission to Mars that provides some science return to the Mars science community at a time when currently no new science data is expected. With no technical or scientific drivers for this mission the research question to be answered by the CDF study is very broad: “With a ESA CaC of only 250 MEuros and the wish to launch a mission to Mars by 2032, what kinds of ESA-led mission concepts could be possible, if any?” Given the large trade space of potential options that could be considered for this programmatic and cost driven mission, a strategy with which to approach the mission architecture definition is devised. Initially, an assessment of the programmatic and cost constraints is developed into a set of high level mission requirements and the design drivers are formulated. A consultation with ESA Mars experts reveals some mission themes which are turned into three distinct mission cases. These mission cases are representative of three different types of mission and are general enough to cover a wide variety of scenarios that scientists might like to see in a small Mars mission. These cases are studied with the intention that the focus remains on understanding the types of mission architectures that are compatible with the programmatic constraints, while demonstrating at high level that feasible technical solutions exists for mission concepts within those constraints. The architectural trade space is then analysed for key components of the mission architecture, such as launch scenario and the means of transfer to Mars. Various trade- offs are conducted at mission level in order to condense the options into a set of reference mission scenarios that are the basis for the study. In the following sections this strategy is broken down and detailed further with rationale provided for design decisions that should be revisited in the future if new information relevant to the mission architecture emerges. 2.2 Programmatic Constraints The following constraints were provided by the customer at the start of the study: • ESA cost at completion of 250 MEuro (economic conditions 2022) • Phase B KO as early as 2023 • B/C/D in ~4 years • Launch between end 2026 and 2032

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 14 of 105

• Total transfer time to the final target orbit should not exceed 3 years Given these constraints, the focus of the study is therefore on a low cost, short mission development schedule with a cost-driven spacecraft design and mission architecture. 2.3 Mission Requirements and Design Drivers In the following sections the programmatic constraints are developed into top level mission requirements and drivers in order to guide the development of the overall mission architecture approach. 2.3.1 Common Mission Requirements Top level mission requirements that are derived from the customer constraints are captured in Table 2-1.

Common Mission Requirements ID Statement Justification Constraint provided by customer. Intention is for the mission to be flown in the same time frame as The mission shall be compatible with a MIS-010 Mars Sample Return mission to launch date between 2026 and 2032 provide additional opportunities for European in-situ science that is not currently being considered. This is not considered a hard requirement, but it is thought that a The mission shall be designed such that “low-cost, short development time” the selected final target orbit can be mission should have a reasonably MIS-020 reached after a maximum of 3 years short transfer duration so as to after launch remain appealing to the science community as well as reduce operations costs. Constraint provided by customer. The total mission cost at completion MIS-030 For a mission to Mars this is shall be <250 MEuro considered “low-cost”. Constraint provided by customer. A There shall be a 4 year allocation for low cost mission should have a phases B/C/D condensed development schedule MIS-040 Note: Refer to Mars Express and is based on previous successful implementation phases mission programmatic approaches such as Mars Express. A fast development schedule implies that spacecraft hardware and Spacecraft units shall have achieved MIS-050 payload should be based on heritage TRL 7/8 by PDR or high TRL hardware and should have a sufficiently high TRL by PDR. Table 2-1: Common requirements independent of mission architecture

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 15 of 105

2.3.2 Design Drivers The mission requirements imply a certain set of architectural drivers, particularly given the main research question that is the focus of the study (see Section 2.1). The main drivers for this mission are therefore cost and schedule. Using programmatics and cost as the main drivers isn’t necessarily the “normal” CDF approach, where the science case is usually the basis for a Mars mission design. Therefore it is emphasised that this study would not be driven by mass, performance or science return.

The study therefore aims to address what can be achieved for a certain cost when there are no initial performance requirements placed on the subsystems. Because of these drivers, the technical team is instructed to consider equipment that has high heritage already, a has a robust development plan through existing programme such as the Mars Sample Return mission. Important note on science return: whilst these design drivers remain the priority for ensuring a low cost, short development time mission, it is still important that there is useful and attractive science return available from the mission. With this in mind, the science team helped to guide the evolution of the spacecraft design, ensuring that valuable science could be produced within the mass, power and data envelopes under consideration and suggesting representative target orbits at Mars that would enable such missions as well as reasonable targets for minimum payload allocations. 2.3.3 Design To Cost Approach The design to cost approach taken in the study is very iterative. With a target of 250 MEuros the cost team provided an initial assessment of the cost breakdown into 5 key areas: • Industrial costs • Launcher costs • Operations costs • ESA costs • Margin Although the mission architecture definition is not driven by mass, the allocated industrial cost for the mission gives a rough order of magnitude of the size of spacecraft that would be feasible within this constraint. Therefore, in order to start the technical iterations a preliminary allocation for mass was made. In order to push costs down in a realistic way the following assumptions were therefore taken:

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 16 of 105

1. ESA costs can be brought down to a realistic minimum, using heritage from previous programmatically constrained missions such as Mars Express. 2. Launch costs will be minimised, with a focus on rideshare opportunities and dual launch options rather than a dedicated launch of an expensive launcher. 3. The dry mass target of the spacecraft in the final Mars orbit will be ~300kg. This mass estimation gave us a rough indication that a rideshare or dual launch would be feasible. However it should be emphasised that this mass estimate is not a hard requirement and the cost and schedule constraints remain the mission drivers. This means that, in theory, a high heritage subsystem with a large mass should be selected over a lower mass, low TRL option. This assumption works when the resultant spacecraft mass is not increased to the point where the allocation for launch is exceeded. 2.4 Selection of Mission Cases Because the mission is designed to cost, the study retains a very broad view of the types of mission architectures that would be feasible within the cost constraints so as not to discard feasible solutions too early on in the definition process. An early brainstorming session was held with ESA Mars experts to get an idea of the types of small missions that are of interest to scientists and to the Mars Exploration Programme. Following this consultation the following cases were put forward for consideration to the CDF.

Case 1 Case 2 Case 3 Mission Mars Communications Mars Science Orbiter Mars Hard Lander / Name Constellation Mission Mission Penetrator Mission Single satellite Carrier module + Configuration 3 satellite constellation science orbiter EDL of n landers Science instrument suite Data relay and secondary Science hard lander Objective and secondary data science instrument demonstration relay Considered Areostationary/ Low Mars Orbit Ballistic entry Orbits Trans-Areosynchronous

Outline schematic [Not to scale]

Table 2-2: A summary of the mission cases under study These mission cases provided the starting point from which the reference mission architectures were defined.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 17 of 105

2.5 Architecture Options Given the wide scope of architectures that would be available for a cost driven mission to Mars and the various mission cases under study, there are many potential options for the launch and transfer scenario. In order to condense such a wide range of options an initial qualitative assessment was made of launch vehicles, rideshare options and propulsions technologies that are likely to be available in the given timeframe. 2.5.1 Initial Orbit Injection The initial orbit into which the spacecraft is injected dictates the delta-V requirements needed to transfer to Mars. Options put forward for trade-off are shown in Figure 2-1. There are several other orbit options that may be considered, most of which would be opportunistic rather than regular. At this early stage of study, opportunistic launches to orbits outside of those described here are not further considered.

Rideshare options

Direct LEO GTO L2 injection

Figure 2-1: Initial orbit options overview 2.5.2 Launch Vehicles 2.5.2.1 Programmatic launcher assumptions Some assumptions are made when down selecting a candidate launch vehicle for this kind of cost driven mission architecture. It was therefore assumed that: 1. On an ESA led mission there would be a preference for a European launch vehicle. Other launch vehicles can be considered if a European vehicle of the same class is not available. 2. The cost of a shared launch (dual or rideshare) is within budget and scales proportionally with the wet mass of the spacecraft. 3. As a secondary passenger, the spacecraft would have no control over the orbit it is inserted into by the launch vehicle. 2.5.2.2 Launcher options The launch vehicles and corresponding injection orbits depicted in Figure 2-2 were considered. Whilst other similar options may become available opportunistically, the launch scenarios given here may be considered as representative for alternative launch vehicles of similar cost and performance. Soyuz launches would not be available from Kourou within the timeframe of this mission and therefore cannot be further considered.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 18 of 105

Launcher

H3-X00 Vega-C Ariane 62

Launch to Mars insertion Launch to LEO Launch to GTO 2 2 (C3 = 10 km /s )

Launch to Earth escape Launch to L2 2 2 (C3 = 0 km /s )

Figure 2-2: Launch vehicle options overview 2.5.3 Propulsion Technologies In the case that direct injection to Mars orbit is not possible the spacecraft requires the capability to achieve Trans-Mars Injection of its own accord. There are fundamentally 3 options available. The propulsion architecture options put forward for trade-off are shown in Figure 2-3.

Propulsion Architecture

Chemical Electric Hybrid Propulsion Propulsion

CP kick stage CP kick stage EP spacecraft CP spacecraft EP spacecraft

Figure 2-3: Propulsion architecture options overview 2.5.4 Overview of Potential Launch and Transfer Options The assessments made in this chapter show that there are a multitude of potential options available for the launch and transfer options for getting to Mars. With the aid of some preliminary mission analysis to assess the required transfer stages, these options are summarised in the schematic shown in Figure 2-4. The combination of options amounts to 16 different scenarios for the 3 mission cases described in Section 2.4.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 19 of 105

Mars LMO sol 4 Areostationary Aerobraking Aerobraking Aerobraking + + + EP spiral down to Mars to LMO Mars down spiral EP to LMO Mars down spiral EP to LMO Mars down spiral EP to LMO Mars down spiral EP to LMO Mars down spiral EP CP CP CP s / CP ) EP spiral down to to ASO down spiral EP to ASO down spiral EP to ASO down spiral EP to ASO down spiral EP to ASO down spiral EP km 0 = capture ( Vinf s / km 3 = Vinf ) 10 Ballistic coast Ballistic coast Ballistic coast Ballistic coast ~ = 3 C heliocentric transfer ( EPSpiral toMars Capture EPSpiral toMars Capture EPSpiral toMars Capture EPSpiral toMars Capture EPSpiral toMars Capture EPSpiral toMars Capture ) 0 = 3 escape C ( CP Kickstage CP Kickstage ) 00 X - 3 H ( CP Kickstage ) 00 X - GTO 3 H ( CP Kickstage Dedicated Launch Launch Dedicated Dedicated Launch Launch Dedicated LEO Launch Ballistic Hyperbolic Entry Hyperbolic Electric Propulsion Electric Chemical Propulsion Chemical

Figure 2-4: A summary of launch and transfer options for the mission casesMission Architecture Trade-off Assumptions Summary

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 20 of 105

The assumptions made in the selection of the mission architectures for study are summarised here. 2.5.5 Programmatic and Cost Assumptions

Design to Cost Assumptions Ref Assumption 1 ESA costs can be brought down to a realistic minimum, using heritage from previous programmatically constrained missions such as Mars Express. 2 Launch costs will be minimised, with a focus on rideshare opportunities and dual launch options rather than a dedicated launch of an expensive launcher. 3 The target dry mass of the spacecraft in the final Mars orbit is ~300kg Note: this mass target has been derived from the cost requirement (MIS-030) and is subject to change provided the cost requirement is met.

Programmatic Launcher Assumptions Ref Assumption 1 On an ESA led mission there would be a preference for a European launch vehicle. Other launch vehicles can be considered if a European vehicle of the same class is not available. 2 The cost of a shared launch (dual or rideshare) is within budget and scales proportionally with the wet mass of the spacecraft. 3 As a secondary passenger, the spacecraft would have no control over the orbit it is inserted into by the launch vehicle.

2.5.6 Technical Assumptions 2.5.6.1 Launch Performance Assumptions The following assumptions were made for the launch vehicles under assessment.

Launch Performance Assumptions Launcher Orbit Insertion Mass performance Comment [kg] C3 = 10 km2/s2 1286 Estimated based on a H3-X00 2400kg performance to 2 2 C3 = 0 km /s 1828 GTO Provided by mission Vega-C LEO 2375 analysis team. GTO 5000 See RD[2]. Ariane 62 Provided by mission L2 2700 analysis team. Table 2-3: Launch Performance Assumptions

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 21 of 105

2.5.6.2 Propulsion assumptions The mission architecture trade-offs did not presume any particular thruster technology solution. In order to make a rough comparison between electric and chemical propulsion options the assumptions given in Table 2-4 were used. The electric parameters are loosely based on a nominal operational point of the PPS- 1350E engine, which was initially assessed as potentially being a suitable candidate. The chemical parameters are based on the heritage Lisa Pathfinder kick stage. Trade-off and selection of the specific propulsion architecture for each mission type is given in the relevant chapter later in this report.

Propulsion Assumptions Type Isp Thrust Comment 4 x 22 N thrusters Loosely based on Lisa Pathfinder Chemical propulsion 308 s used for orbit kick stage (~220kg) insertion Based on PPS-1350E performance at 1900W discharge power. In reality these numbers would not Electric propulsion 1787 s 107 mN be static over the transfer but were adequate to assess the initial architecture. Table 2-4: Propulsion Assumptions An additional assumption for an electric propulsion architecture is that the transfer trajectories to Mars are calculated as time optimal transfers (as opposed to delta-V optimal). This is due to the programmatic constraint of a 3 year transfer duration. 2.5.6.3 Delta-V assumptions The delta-V assumptions for each of the mission case options is broken down by transfer phase and depicted in Figure 2-5. The values given here do not include margins (see margin philosophy Section 2.7). For the case where the spacecraft uses chemical propulsion to achieve Mars orbit insertion, gravity losses of ~22% are assumed. This is due to the low thrust levels assumed for the spacecraft chemical propulsion system use for Mars orbit insertion. A higher thrust level could bring gravity losses down further.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 22 of 105

s s s / / / Mars LMO s s s s s m m m / / / / / 120 120 120 m m m m m Aerobraking Aerobraking Aerobraking sol 3000 3000 3000 3000 3000

4 – – – – – s s s s s / / / / / m m m m m Areostationary s s s / / / 1000 1000 1000 1000 1000

m m m – – – – – s / 1247 1247 1247

m – – – s EP spiral down to Mars to LMO Mars down spiral EP to LMO Mars down spiral EP to LMO Mars down spiral EP to LMO Mars down spiral EP to LMO Mars down spiral EP / CP CP CP 1747 )

km – Spiral to ASO Spiral to ASO Spiral to ASO Spiral to ASO Spiral to ASO 0 = CP capture ( Vinf s / s s s s s / / / / / km m m m m m 3 = s 5700 5700 5700 5700 5700 /

- - - - - Vinf m ) 4000

- 10 Ballisticcoast Ballisticcoast Ballisticcoast Ballisticcoast ~ = 3 C heliocentric transfer ( s s EPSpiral toMars Capture EPSpiral toMars Capture EPSpiral toMars Capture EPSpiral toMars Capture EPSpiral toMars Capture / / m m 1975 EPSpiral toMars Capture 1975

)

- - 0 = 3 escape C s ( / m ) 3968

CP Kickstage CP Kickstage - 00 X - s 3 / H ( m ) 3481

CP Kickstage - 00 X - GTO 3 H ( Dedicated Launch Launch Dedicated CP Kickstage Dedicated Launch Launch Dedicated LEO Launch Ballistic Hyperbolic Entry Hyperbolic Electric Propulsion Electric Chemical Propulsion Chemical

Figure 2-5: A summary of launch and transfer option delta-V assumptions for the mission cases

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 23 of 105

2.6 Mission Architecture Trade-off Results The architectural trade-off was made for each of the combinations provided in Figure 2-5 and from this trade-off a reference architecture was selected for study for each mission case. To evaluate these architectures, the assumptions presented in Section 0 were applied. In the following sections the launch vehicle, initial orbit injection and propulsion architecture are qualitatively assessed and then compared to a quantitative assessment of the launch and transfer scenario performance. A reference mission architecture is then selected to apply to the mission cases. 2.6.1 Launcher Assessment An assessment of the launch vehicles selected for study is given in Table 2-5.

Launch Assessment Conclusion Vehicle Small JAXA launch vehicle performance H3-X00 oversized as a dedicated launcher to Mars and cost is considered high unless a rideshare • Potential for opportunity becomes available. opportunistic launch as a

piggyback. Could be considered as a piggyback option • Dedicated launch is should programmatics allow and opportunities too expensive become available within the mission timeframe. At the lower limit for launcher size, the Vega-C estimated mass at Mars is very low and transfer to Mars considered difficult from LEO with • Technical and current propulsion technologies. programmatic feasibility deemed Note: A dedicated launch to GTO is considered to be low. to be outside of the nominal operational range of the Vega-C so not considered in this study. Too expensive for a dedicated launch but likely to be in budget for a rideshare (assuming a proportional mass/cost ratio). • Rideshare likely to Ariane 62 be within budget In the case of a rideshare to GTO or L2, this • Case covers covers larger vehicle options (e.g. Ariane 64, rideshare Falcon 9 etc.). opportunities on larger launch vehicles. Performance to GTO = 4500 kg Performance to L2 = 2450 kg Table 2-5: Launch vehicle assessment

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 24 of 105

2.6.2 Initial Orbit Injection An assessment of the orbits that the spacecraft could be injected into from the launch vehicle is given in Table 2-6.

Initial Orbit Assessment Conclusion Options No dedicated small sat launch options for ~300kg class satellites. • Easiest transfer for a Piggyback would be needed on another Mars small sat Direct mission (this would be the easiest transfer for a • No clear opportunities injection to small sat, but opportunities are few). for piggyback Mars • Dedicated launch too transfer expensive Oversized dedicated launcher could be cost and mass inefficient and likely over budget, however could be considered as a joint mission with a partner. Getting from LEO to Mars requires high delta-V and additional propulsion, which may be too great for a small satellite to accomplish on its • Delta-V requirements own. very high for small sat LEO capabilities Long duration spent in Van Allen belts • Dedicated launch expensive Wet mass of the small satellite would require a dedicated launch. History of spare launch capacity (>1000kg) on Ariane 5 was assessed, regular opportunities for • Regular opportunities rideshare to GTO (at least one per year). for rideshare GTO • Design for GTO covers opportunistic launches Design for GTO covers opportunistic launches for higher orbits. for higher orbits, like L2. L2 is an ideal orbit to hitch a ride to, as it is closest to Earth escape.

Much fewer opportunities to rideshare to than • L2 closest to Earth GTO, especially with mass required. It could not escape be demonstrated that there would be any L2 • realistic rideshare opportunities to L2 in the No regular given time frame for this mission. opportunities for rideshare

Could be programmatically simpler to ride with an ESA science mission to L2 than a commercial mission to GTO. Table 2-6: Qualitative assessment of rideshare options

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 25 of 105

2.6.3 Propulsion Architectures A preliminary assessment of the propulsion architectures is given in Table 2-7.

Propulsion Preliminary Assessment Conclusion Technology • High-TRL thruster options available • Lower launch mass than CP for the same mass at Mars

• Long transfer durations • Launch mass lower than Solar • Thruster lifetime a concern for CP and high TRL Electric options available. Propulsion • Flexible launch window (many opportunities for rideshare) • Flexible launch windows. • Flexible arrival dates • Operational flexibility (e.g. changing of orbits at destination)

• Traditional method for interplanetary transfer • Satellite could require a kick stage (e.g. Lisa Pathfinder kick stage) for anything other than a dedicated launch. • Launch mass higher than for EP, less flexible Chemical • Higher launch mass than solar electric launch windows. Propulsion propulsion for the same mass at Mars • Could be considered if EP • Specific launch windows needed (unless not feasible. waiting in a parking orbit) • Specific arrival dates (coincide with global dust storm season for given mission timeframe.)

• Could be considered if transfer duration is too long for EP technologies • Considered operationally Hybrid CP • Likely to not be efficient for mass or cost complex, costly and mass + EP inefficient • More complex (and therefore more costly) operationally and programmatically

Table 2-7: Assessment of propulsion technology options

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 26 of 105

2.6.4 Launch Mass Comparison In order to compare the mission architectures quantitatively, a summary of the launch mass versus the dry mass in Mars orbit has been made in Table 2-8. For the dedicated launches the launch mass is assumed to be the maximum capability of the launch vehicle. For the rideshare options, the launch mass needed to achieve 300kg in the target orbit is selected. In those cases, the launch mass needed is less than half of the capability of the Ariane 62, meaning that a rideshare or dual launch as a secondary passenger is achievable. Note that for the Mars Communication Constellation mission, 300kg was allocated as the total mass for all 3 satellites in the constellation.

Launcher Launch conditions target Launch Orbit Target Mars Propulsion Architecture Launch kg / Mass at Mars Dry Mass orbit/kg target Mars Science Orbiter

H3-X00 Dedicated C3 =10 LMO CP 1286 824 H3-X00 Dedicated C3 =0 LMO EP 1828 1066 Vega-C Dedicated LEO LMO CP 2375 242 Vega-C Dedicated LEO LMO Hybrid 2375 285 Ariane 62 Rideshare GTO LMO CP 1337 300 Ariane 62 Rideshare GTO LMO EP 650 300 Ariane 62 Rideshare GTO LMO Hybrid 1015 300 Ariane 62 Rideshare L2 LMO EP 514 300 Mars Communication Constellation H3-X00 Dedicated C3 =0 TASO EP 1828 1201 Vega-C Dedicated LEO TASO Hybrid 2375 348 Ariane 62 Rideshare GTO TASO CP 1497 300 Ariane 62 Rideshare GTO TASO EP 577 300 Ariane 62 Rideshare GTO TASO Hybrid 935 300 Ariane 62 Rideshare L2 TASO EP 457 300

Table 2-8: Launch mass vs dry mass in Mars orbit

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 27 of 105

The key takeaways from the table are: • All but 2 of the options meet or exceed the target of putting 300kg dry mass in Mars orbit within the capability of the launch vehicle. • There is considerable variation in the launch mass needed to achieve 300kg dry mass in Mars orbit. • For the rideshare options, an EP transfer to Mars provides the lowest launch mass when compared to a CP or hybrid architecture. 2.6.5 Transfer Time to Mars For chemical propulsion architectures the transfer time to Mars is <1 year and is thus well within the requirement of reaching Mars within 3 years (aerobraking manoeuvres may add some months to this, but this will not affect the 3 year transfer limit). For the low thrust electric propulsion architecture, transfer times can be much longer. Using the assumptions given in Section 2.5.6.2 in order to make an initial assessment the results for the total transfer time to Mars are given in Table 2-9. For the H3-X00 launch vehicle, while the transfer time is acceptable, the cost of the launch would be prohibitive unless a rideshare opportunity was found. For a rideshare Ariane 62 launch and a 300kg dry mass at Mars, the calculated transfer times are also within the 3 year limit. These numbers indicate that an EP architecture is possible. Further refinement of the transfer times are included in the specific mission cases.

Launcher Propulsion Propulsion Architecture Launch target Launch target orbit/kg target Dry Mass at Mars at Mars Dry Mass Mars Target Orbit Target Mars Launch conditions Transfer yr. Time / Mars Science Orbiter H3-X00 Dedicated C3 =0 LMO EP 300 1.83 Ariane 62 Rideshare GTO LMO EP 300 2.72 Ariane 62 Rideshare L2 LMO EP 300 1.83 Mars Communication Constellation H3-X00 Dedicated C3 =0 TASO EP 300 1.42 Ariane 62 Rideshare GTO TASO EP 300 2.2 Ariane 62 Rideshare L2 TASO EP 300 1.42

Table 2-9: Transfer times to Mars using a low thrust electric propulsion architecture

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 28 of 105

2.6.6 Mission Architecture Conclusions This chapter has analysed potential launch vehicles, initial orbit injections and the propulsion architectures required to define a mission architecture for a small satellite Mars mission. The qualitative and quantitative assessments have been considered in the frame of the whole architecture and the following conclusions have been made: 2.6.6.1 Launch vehicle conclusion: • H3-Xoo launch vehicle: o The H3-Xoo dedicated launch option far exceeds the dry mass needed in Mars orbit. However, the implication of building an 800 - 1200kg satellite in order to make the launch scenario mass-efficient is beyond the cost and programmatic constraints of this study. o The H3-X00 launch vehicle could be considered if a rideshare opportunity arises with JAXA and was programmatically feasible, however none such opportunities have yet been identified. o The H3-X00 launch vehicle is therefore discounted for further analysis in this study. • Vega C launch vehicle: o The Vega-C is only compatible with a launch to LEO. The transfer to Mars is considered not feasible from LEO with an EP architecture. o The Vega-C dedicated launch option does meet the required target mass in Mars orbit for the Mars Communication Constellation, but not for the Mars Science Orbiter. o Whilst the performance is met for the Mars Communication Constellation, the dedicated launch costs and the cost of a hybris propulsion system are thought to be too high. Therefore the Vega-C is discounted for further analysis. • The Ariane 62 launch vehicle: o Rideshares to both GTO and L2 have the capability to put 300kg in Mars orbit and is well within the capability of the launch vehicle. o The cost of such a rideshare is assessed to be within budget. o The Ariane 62 launch vehicle is selected as the reference launcher for the SMARTieS study

2.6.6.2 Injection orbit conclusion: Given the selection of the Ariane 62 launch vehicle the down selected injection orbits are either GTO or L2. The assessment provided in this chapter has the following conclusion: • L2 injection orbit: o An Ariane 62 rideshare to L2 has the capability to put 300kg in Mars orbit and is well within the capability of the launch vehicle.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 29 of 105

o There are deemed to be few opportunities with spare capacity in the considered timeframe to the L2 point which put into doubt the launch window stipulated in the programmatic constraints. o A spacecraft design that is able to reach Mars from L2 will not be able to reach Mars from GTO without modification to the propulsion system, giving the least flexibility for the mission. o L2 could be considered if an opportunity arose but will not be further considered in this study. • GTO injection orbit: o An Ariane 62 rideshare to GTO has the capability to put 300kg in Mars orbit and is well within the capability of the launch vehicle. o It is assumed that there is the potential for many rideshare opportunities on commercial missions to GTO on the Ariane 62 (and potentially other larger launch vehicles) o A spacecraft design that is able to reach Mars from GTO will also be able to reach Mars from L2, giving the greatest flexibility for the mission. o GTO is selected as the reference injection orbit for the SMARTieS Study 2.6.6.3 Propulsion architecture conclusion: Given the selection of an Ariane 62 launch to GTO there is still the possibility of either an EP, CP or Hybrid architecture. The assessment provided in this chapter has the following conclusion: • Hybrid Chemical and Electric Propulsion: o A hybrid solution gives a launch mass that is between a fully EP and fully CP solution, but 2 different propulsion architectures is an inefficient use of dry mass. o Transfer times are within the 3 year limit o Considered operationally complex, costly and mass inefficient and therefore not within the spirit of a low cost mission and not considered further. • Electric Propulsion o An EP architecture provides the lowest launch mass when compared to a CP or hybrid architecture. o An EP architecture allows a transfer time to Mars that is within the 3 year limit o High TRL technologies are identified that could meet mission requirements. o Provides flexibility in launch and arrival dates o Whilst programmatically and technically interesting the technical feasibility and cost considerations of an Electric Propulsion architecture need to be considered further. o It is concluded to select EP as the reference propulsion architecture for the Mars Communication Constellation case. • Chemical Propulsion:

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 30 of 105

o A CP architecture allows a transfer time to Mars that is within the 3 year limit o High TRL technologies with good heritage are identified that could meet mission requirements. o Provides flexibility in launch and arrival dates o Whilst programmatically and technically feasible, the development cost and complexity of a chemical kick stage that may not be off the shelf needs to be considered further. o It is concluded to select CP as the reference propulsion architecture for the Mars Science Orbiter case. 2.7 Margin Philosophy The margin philosophy applied for this study is “Margin Philosophy for Mars Exploration Studies” and can be found on the vCDF server: W:\SMARTieS_Study\SMARTieSMiscellaneous\Margin Philosophy, in summary: • Equipment/unit level o 5% for fully developed items o 10% for items to be modified o 20% for items to be developed • System level o A system margin of 20% will be added to the total system mass for EP architectures only o A system margin of 30% will be added to the total system mass CP architectures only • Margin Tailoring: o MAR-PWR-020: In case of electric propulsion, the design maturity power margins from MAR-PWR-010 shall be 10% applied to the overall Solar Electric Propulsion system. - This margin has been waived and is now 0%. This is because the power cannot exceed the maximum input of the PPU and so a margin would be worthless to add here. o No power system margin as per MAR-PWR-040/50 is considered during eclipse. The margin is instead assumed in thermal model to keep s/c at acceptable minimum temperature at end of eclipse, considering pre-heating before eclipse. o For a Martian aerobraking phase, a factor of 2 shall be applied as margin between target and limit loads. 2.8 Small Satellite Design Philosophy The approach taken to the design of a small, low cost satellite should be different to that of a conventional satellite. Table 2-10 shows identified areas that can be challenged in the design of a small satellite.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 31 of 105

Fixed Variable ∆v required to reach operational orbit Payload  required accuracy, mass, data generated, data quality Propellant mass (depending on Prop system Pointing accuracy  AOCS equipment  choice and overall S/C mass) pointing accuracy Tank size (depending on propellant mass) Communication S/S  data rate, data volume Size solar panels (depending on P/L power Equipment choice (COTS vs. high rel, and prop system power needs) redundancy) Operations  number of contacts

Table 2-10 : Design areas that can be challenged for small satellite design For a constellation of small satellites, redundancy might be taken at constellation level rather than at satellite level. For single satellite missions a “smart redundancy” approach could be applied, where redundancy is only taken for critical systems. This also includes incorporating redundancy from an architectural level, such as the utilisation of strategic cross strapping, and only incorporating fully redundant systems when the mass impact is low (e.g. including 4 reaction wheels instead of 3).

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 32 of 105

This Page Intentionally Blank

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 33 of 105

3 MARS COMMUNICATION CONSTELLATION 3.1 Mission Requirements The design of the spacecraft constellation is strongly linked to the programmatic constraints and mission design drivers identified in Section 2.2 and Section 2.3. In addition to the general architectural mission requirements given in Section 2.3.1 and following the mission architecture trades performed in Section 2.6, the Mars Communications Constellation case study also has some specific mission requirements given in Table 3-1.

Mission Requirements ID Statement MIS-MCC-010 The mission architecture shall comprise 3 communication relay satellites. The Mars Communications Constellation shall provide continuous coverage for MIS-MCC-020 surface (equatorial and moderate latitudes) and in-orbit assets on Mars. Following arrival at the target orbit, the mission shall provide communication MIS-MCC-030 services for a minimum of 2 years (goal: 6 years). Note: the navigation service capability should be assessed The total mass delivered to the target orbit shall be 300 kg (TBC), including all satellites and all margins. MIS-MCC-040 Note: this mass target has been derived from the cost requirement (MIS-030) and is subject to change provided the cost requirement is met. MIS-MCC-050 The mission shall be compatible with existing Mars Relay Network protocols.

The mission shall be compatible with a shared launch on an Ariane 6.2 launch MIS-MCC-060 vehicle. The mission shall be compatible with an interplanetary transfer to a Mars MIS-MCC-070 Areosynchronous Orbit (TBC) from a Geostationary Transfer Orbit (GTO). The mission shall achieve the interplanetary transfer stipulated in MIS-MCC- MIS-MCC-080 070 using an Electric Propulsion Architecture. Table 3-1: Mars Communications Constellation mission specific requirements 3.2 Mission Analysis It was identified early on that a 3 satellite constellation would need to be in an Areostationary or an Areosynchronous type orbit around Mars in order to meet the requirement for continuous ground coverage (MIS-MCC-020) across the entire planet. A trade-off was made between the following orbits: • Areostationary (radius 20,428 km) • Trans-Areostationary Orbit (TASO) (preliminary assumption: radius 21,000 km) In analogy to the well-known geostationary orbit, there exists an areostationary orbit, which would allow communications satellites to hover 17,032 km above the Mars

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 34 of 105 equator and remain stationary with respect to the rotating planet. On the areostationary orbit, there are four equilibrium longitudes spaced by 90 degrees at which spacecraft could be positioned at (near-)zero manoeuvre cost. At all other longitudes, costly stationkeeping is needed to prevent a spacecraft from drifting far from its assigned location. For global coverage (excluding the polar regions), all four equilibrium longitudes must be utilised; omitting one would already lead to a significant and permanent coverage gap. Global coverage of all low to moderate latitudes with only three satellites is however possible. The orbit should then be located somewhat (e.g., 600 km) beyond the areostationary altitude. This trans-areostationary orbit (TASO) is not subject to the strong perturbations present in the areostationary orbit. On the TASO, three satellites spaced by 120 degrees could provide uninterrupted coverage up to around 60 degrees of latitude, requiring only minimal stationkeeping to maintain their spacing. As their altitude is less than 4% higher, the link budget is unaffected; but because all satellites on a TASO will be slowly drifting westwards rather than remaining stationary, surface locations using the network would have to re-point their antennas regularly.

Areostationary TASO

EP Delta-V cost to • 1000 m/s if T/m at capture > • 1000 m/s if Thrust to mass ratio reach Orbit from a 4 0.15 mN/kg at capture > 0.15 mN/kg sol Mars Insertion • • Orbit 1100 m/s otherwise 1100 m/s otherwise

4 equilibrium stationary points 3 drifting points

Stable points in orbit

Station-keeping Up to 37 m/s/year for satellites not Occasional small adjustments to cost at stable equilibrium points maintain 120 deg separation With 3 satellites in the constellation and 4 equilibrium stationary points around Mars in an Areostationary Orbit, there is a significant delta-V cost in maintaining a satellite out of those points. Using 3 of the 4 points would produce and unequal spacing and therefore impact on the Result continuous coverage requirement. The delta-V impact is reduced significantly with the TASO orbit and because the points are drifting, full gorund coverage is achieved, and therefore the TASO orbit is selected as the reference orbit for the Mars Communication Constellation. Table 3-2: Trade-off and selection of reference orbit for the Mars Communication Constellation

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 35 of 105

3.3 Reference Scenario The reference scenario chosen for study for the Mars Communication Constellation is summarised in Table 3-3.

Reference scenario for Mars Communication Constellation Number of satellites 3 in constellation Launch scenario Ariane 62, rideshare Initial injection orbit Geostationary Transfer Orbit Means of transfer to Electric Propulsion Mars Orbit GTO to C3 = 0 (spiral up) 3700 m/s C3 = 0 to Mars Capture 5700 m/s (interplanetary transfer) Delta-V 1000 m/s (assuming if T/m at capture > 0.15 Mars Capture to TASO mN/kg) Total delta-V 10.4 km/s (without margin) Mars orbit type Trans-Areostationary Orbit (TASO) Period 92394 s = 25.665 Earth hours Radius 21,000 km (optimal orbit still TBD) Longitude drift westwards, -360 deg of longitude in 25.3 days

Mars Orbit

Uninterrupted coverage: • Coverage Provided up to latitude +/- 48 deg for minimum elevation 10 deg • Provided up to latitude +/- 15 deg for minimum elevation 20 deg • NOT PROVIDED for minimum elevation 30 deg C3=0 Vinf = 3km/s LEO GTO C3=~10 Vinf = 0km/s TASO 4 sol Mars SSO (escape) (heliocentric transfer) (capture)

EP Spiral to Mars Capture - 5700m/s

Table 3-3: Summary of mission architecture reference scenario for Mars Communication Constellation

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 36 of 105

3.4 Systems Analyses and Trade-Offs 3.4.1 Spacecraft System Configuration Trade-Off The stipulation that there should be 3 spacecraft in the constellation and that they should arrive at Mars using electric propulsion does not presume any specific spacecraft configuration. A summary of the options considered are presented in Table 3-4.

Option 1 Option 2 Configuration Overview Description EP Mothercraft + 2 identical CP EP carrier + 2 identical CP small sats small sats Key points • Mother contains propulsion to • Propulsion carrier contains get all 3 satellite to Mars propulsion to get all 3 satellite to • Mother contains all major Mars systems for communications to • Each satellite contains all major Earth systems for communications to • Small sats communicate with Earth Earth through intersatellite link • Small sats have capability to to Mother manoeuvre into operational orbit • Small sats have capability to • 2 different satellite designs manoeuvre into operational orbit • 2 very different satellite designs Option 3 Option 4 Configuration Overview 3 identical CP small sats with EP Description 3 identical independent EP small sats booster add-on • Each small sat makes its own Key points • All 3 satellites essentially have way to Mars the same comms functionality • Small thruster must meet • All satellites are “tethered” and transfer duration requirements travel together to Mars • 3 x EP and power systems • One satellite has an extra EP • propulsion system “add-on” Less efficient in terms of dry mass? • Small sats have capability to • manoeuvre into operational More flexibility for MOI orbit • All satellites have same design • Number of satellites can be rescoped more easily than the other configurations. Table 3-4: Spacecraft system configuration trade-off options The results of the spacecraft configuration trade-off are shown in Table 3-5. The results show that option 4 (3 identical independent EP small sats) is the clear winner to take as the reference case for study.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 37 of 105

Criteria Option 1 Option 2 Option 3 Option 4 All 3 satellites are Includes release essentially the same mechanism for small but include Includes release No additional sats, intersatellite separation mechanism for small complexity associated links are required, 2 mechanism for each sats, 2 different with joining very different small sat. Includes a Configuration satellites to design spacecraft, all satellites to design potentially complex Complexity (main difference is systems identical and (main difference is tethered power propulsion) independent propulsion and system and comms payload) additional propulsion stage 2 3 1 4 Highest cost for Highest cost for EP Development power subsystem Medium hardware Medium hardware system hardware, and Hardware hardware. Highest and development cost and development cost lowest development Cost (input development cost for cost overall from cost EP. team) 4 3 1 4 Could be more Likely to need a dual Likely to need a dual Likely to need a dual flexible, e.g. multiple Launch launch launch launch strategy rideshares. 3 3 3 4 Operation of 3 Operation of 1 Operation of 1 Operation of 1 satellite during satellite during satellite during satellite during transfer transfer transfer transfer 1 satellite type to 2 different satellite 2 different satellite 1 satellite type to operate during Operational types to operate types to operate operate during operational phase. considerations during operational during operational operational phase. EP Development - including cost phase. Development phase. transfer, CP on orbit? Operations +++ ++ Development + Development + depending transfer Operations +++ Operations ++ Operations ++ duration 1 4 4 2 Redundancy Redundancy provided by having provided by having multiple identical Mothercraft requires Mothercraft requires multiple identical spacecraft, Redundancy full redundancy full redundancy spacecraft, EP stage potentially no concept requires full redundancy in the redundancy propulsion system 1 1 3 4 Mothercraft power Mothercraft power Dedicated power for Tethered power for for comms and EP, for EP, small sat EP EP, complex solution small sat power sized power sized Plenty of power (control , stability, EP/Power according to power according to power available at Mars connectors) needs at Mars needs at Mars orbit for each S/C. 3 3 1 4 Will not provide an advantage in the Higher data capacity Higher data capacity Higher data capacity Comms overall data capacity in Earth link as each in Earth link as each in Earth link as each Performance as all data transferred s/c has DTE s/c has DTE s/c has DTE through mother 1 4 4 4 Score 15 21 17 26 Table 3-5: Spacecraft configuration trade-off resultsElectric propulsion thruster trade-off

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 38 of 105

The electric propulsion technology selection was a key driver in the spacecraft design. The trade-off and assessment of candidate technologies is provided in this section. 3.4.1.1 Assumptions • Direct transfer from GTO to Mars TASO delta-V 10.4 km/s (without margin) o GTO to Earth Escape Velocity transfer with constant thrust at Optimum Electric settings for each thruster; the solar panels are sized for this case o Interplanetary Transfer with degraded power available on board, following 1/r2 law between Earth and Mars distance. EP systems performance is degraded with available power. o Spiral-down into Mars TASO with constant thrust at Optimum Electric settings for each thruster with the available power on-board. • Mars-Sun distance = 1.666 AU (worst case) • Earth-Sun distance = 1 AU • Thruster Qualification safety factor = 1.1 • EoL disposal ΔV is negligible • De-tumbling by injection module • Maximum of 1 thruster firing at any point in the mission 3.4.1.2 Design drivers Design drivers for the EP system are as follows: • Cost is main driver of the system -> EP system shall prioritize the minimization of the global cost of the spacecraft • Thruster minimum lifetime -> high-delta v mission in many cases requires multiple thrusters to include lifetime spares • Limit power due to large solar panels and difficulty to allocate in fairing • Limit system complexity • Total transfer time within 3 years • High TRL solutions as of 2020 desired and TRL7/8 by PDR. 3.4.1.3 Technology options The architectural trade space that was explored was based solely on European technology and included Kaufman type gridded ion engines (GIE), Hall effect thrusters (HET) and radio frequency ion thrusters (RIT).

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 39 of 105

Table 3-6 Electric propulsion technology options for trade-off Due to the large number of options a down-selection was made to take forward for trade-off, mainly based on the existing high TRL of candidate technologies and the recommendation of the Electric propulsion team. The options taken forward for trade- off were the PPS-1350/E, PPS-X00, T5 and T6 engines. 3.4.1.4 Thruster performance

Electric Propulsion Thruster Parameters Thruster Type Power Power Isp/s Thrust/mN range / W efficiency [at max power] [at max power] PPS-1350/E Hall effect [700, 2500] 0.9 1752 141 PPS-X00 Hall effect [200, 1000] 0.87 1614 54 T5 Gridded ion [150, 750] 0.9 3068 25.5 T6 Gridded ion [2700, 5000] 0.9 3884 145 Table 3-7: Electric Propulsion Thruster Parameters In order to make an initial assessment of the thrusters under consideration, a parametric calculation tool was used to provide estimates of the electric propulsion and power subsystem mass for a given satellite dry mass. Initially an allocation of 100 kg per satellite was made and the subsystem teams made preliminary estimations for their respective subsystems. In order to meet the TRL requirements high heritage equipment that would respect the cost and schedule constraints was selected. Using this methodology, it immediately became clear that the mass of the power subsystem, in particular heritage solar array technology, and existing electric thrusters would not be able to support a small satellite on such a large transfer. For a 100kg satellite, more than 91% of the mass would need to be allocated to the EP and power subsystem. Taking into account the minimum mass needed for the remaining subsystems an iterative approach

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 40 of 105 was taken in order to converge on an optimised mass solution for the power and EP systems. The results of this intial assessment are given in Table 3-8.

Number Time to Satellite Satellite Total Power of Total EP orbit [y] dry wet impulse required Thruster thrusters subsystem 1 mass mass required @ Earth per mass thruster [kg] [kg] [MNs] [W] satellite ON PPS-1350/E 5.59 2 85.20 2778 2.05 347 688.6 PPS-X00 6.23 7 141.69 889 5.99 376 784.9 T5 4.18 2 68.32 833 8.3 302 457.7 T6 4.98 1 74.09 6452 1.84 388 521.1

Table 3-8: Electric Propulsion Thruster Trade-off Options It is important to note here that the parametric model was used for assessment purposes and does not represent the final design. The numbers are based on an assumed dry mass of all the subsystems, including system margin and excluding the EP system and solar arrays of 221.55kg. An assessment of the model output provides the following conclusions: • Use of a single T5 engine would mean a transfer time of 8.3 years, which far exceeds the <3 year transfer time requirement and is therefore not considered further.

• Using PPS-X00 technology would require 7 thrusters, which would be hard to accommodate, and would mean a transfer time of 5.99 years, which far exceeds the <3 year transfer time requirement and is therefore not considered further.

• Using PPS-1350E technology would require 2 thrusters to accomplish the required total impulse, which is costly, and the satellite wet mass is close to 690kg. With 3 satellites and a launch adaptor to consider, the constellation is likely to exceed the mass of a rideshare or dual launch and would require a dedicated launch. If a dedicated launch could be considerd then the constellation would be the primary passenger and would have control over the launch, rather than need to be the passenger on a shared launch to GTO. A dedicated Ariane 62 launch would be able to launch the constellation to C3 = 0 km2/s2 and the delta- V needed for the transfer would be drastically reduced. This would mean that the number of PPS-1350E thrusters could be reduced to 1 per satellite, instead of the 2 needed to transfer from GTO. This option is discussed in the sensitivity analysis given in Section 3.4.1.5. • Despite the high cost, the T6 gridded ion engine is technically the best solution of the four options in the trade-off. This is due to the fast transfer times when compared to the T5 and PPS-X00 and the low wet mass when compared to the PPS-1350E option.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 41 of 105

While the T5 and the PPS-X00 are discarded, two viable solutions result from the analysis given above. The major deciding factor in the selection of two techncically feasible solutions is the cost and programmatics. A cost sensitivity analysis was therefore performed between the T6 option (1x T6 per satellite transfer from GTO to TASO, dual launch) and the alternative PPS-1350E option (1 x PPS-1350E per satellite transfer from C3 = 0 km2/s2 to TASO). 3.4.1.5 Electric propulsion architecture sensitivity analysis As described in the preceding section, two technically viable electric propulsion architectures resulted from the initial propulsion analyses. Since this mission is cost driven, in order to form a baseline design the cost and programmatics of each option had to be considered. The results are presented in Table 3-9. The cost is not highlighted with a specific figure. A relative indication is given between both options.

Electric Propulsion Architecture Cost Trade Off Parameter PPS-1350/E T6 Number of thrusters 1 1 per satellite Launch Scenario Dedicated Ariane 62 Dual Launch Ariane 62 Mars transfer C3 = 0 km2/s2 to TASO GTO to TASO The thruster technology is more Whilst the thruster technology is expensive than the PPS-1350E, cheaper than the T6, the and even though there is a dedicated launch causes the Cost shared launch the cost still mission cost to exceed MIS-030. exceeds MIS-030. Cost is Cost is slightly more than the T6 slightly less than PPS-1350/E option. thruster option. Launch flexibility Dependant on opportunities Programmatics SEP System TRL 7 SEP System TRL 7 Table 3-9: Electric Propulsion Architecture Cost Trade Off An assessment of the cost and programmatic aspects provides the following conclusions: • Use of a single PPS-1350/E engine would require a dedicated launch, which is programmatically simpler, and potentially offers more launch mass margin, launch date flexibility and no co-passenger constraints are foreseen for a slightly higher than the T6 option. • Use of a dual launch with a single T6 engine offers a small cost saving, where the higher cost of the T6 engine is off-set by the cost savings of a dual launch. Whilst both options are similar, at this stage of study the cheaper option is selected and therefore the T6 option (1x T6 per satellite transfer from GTO to TASO, dual launch) is selected as the study baseline.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 42 of 105

3.4.2 Power Subsystem 3.4.2.1 Assumptions and design drivers For an electric propulsion small spacecraft the power system design is a driver for mass and configuration. The maximum operating power required for the T6 engine at 1AU was used to drive the sizing of the solar array1, whilst the nominal on orbit operations is the driver for the sizing of the PCDU and battery. The mission drivers of high TRL and low cost are a factor for all aspects of the spacecraft design, the following assumptions were also made: • Worst case distance to sun: 1.666 AU (sun irradiance ratio 0.36 compared to 1AU) • Solar array Sun aspect angle: assumed 0° with SADM 1 degree of freedom (system, AOCS) • Coverglass thickness: 125µm 3.4.2.2 Trade-offs The trade-offs given in Table 3-10 were performed for the design of the power subsystem design architecture.

Power Architecture Trade Off Equipment Trade-off Conclusion Rigid Rigid panels selected with a power density of 70 W/kg vs Solar array as they are more adapted and mature for high power Semi-rigid technology wing. European flexible SA not yet available or mature vs enough, especially in this (low) power range. Flexible 1 wing Solar array 2 wings of 4 rigid panels each selected for symmetry of vs configuration configuration and ease of AOCS control 2 wings Low cost Si, 3J Solar cells vs 4G32 cell chosen to optimize mass, cost and feasibility High efficiency 4J Bus voltage and Dual regulated bus 100V/28V for compatibility with PCDU MPPT EP PPU and MPPT for large variations or SA architecture vs illumination conditions (distance to sun and S3R temperature variations) Table 3-10: Power architecture trade-off summary

1 Note that depending on the arrival date at Mars the minimum power needed to operate the T6 engine (in this case 2200W) at Mars can become the major driver for the sizing of the solar array, not the maximum operating power at 1AU. This means that the solar array mass and power values given in this report could be underestimated by approximately 20%. This should be corrected in future iterations of this mission case. The conclusions of the study are not changed.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 43 of 105

3.4.3 Communications Performance Analysis Following the design-to-cost approach, the communication subsystem, which is the primary ‘payload’ for the communication constellation, is designed to maximise communications performances within the low cost constraints. The communication subsystem has to adhere to requirement MIS-MCC-050 (The mission shall be compatible with existing Mars Relay Network protocols) and should follow a design to cost approach. 3.4.3.1 Assumptions The assumptions (and their origin) of the communication subsystem design are as follows: • General: o No performance requirements are given. The design approach is to see what performance can be accommodated within the cost constraints. However, similar performances to previous Mars missions are initially targeted. • Earth link: o X-Band, 50W TWTA – [system] o 1m HGA – [system] o Occultation of 80 min every orbit for 1/3 of the Martian year – [mission analysis] o Superior conjunction interrupting comms for SEM angles < 2 deg, degradation up to 5 deg - [link budget] o Maximum data rate with 3 dB link margin – [ECSS] o 6h of G/S contact time (EL 10 deg) – [operations] o 0.5 Gb / day generated data - [system] o No redundancy - [system] o New Norcia 35m ground station at 10 degrees of elevation • Mars link: o UHF Proximity-1 – [heritage of current Mars rovers] o LGA with the boresight pointed to Mars o Surface asset with LGA and 9W RF power - [Curiosity Rover as reference] o No subsystem redundancy – [System – redundancy at spacecraft level]. Note: The values given above refer to a single spacecraft, not the whole constellation. 3.4.3.2 Communications subsystem baseline design The baseline design and communications links for the communications subsystem is presented in Figure 3-1 and Figure 3-2.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 44 of 105

Figure 3-1: Comms subsystem equipment layout

Figure 3-2: Earth – Mars Communications links

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 45 of 105

3.4.3.3 Communications Performance Analysis 3.4.3.3.1 Earth, Sun and Mars range

Figure 3-3: Range plots for Earth, Sun and Mars in the mission period

3.4.3.3.2 Baseline Earth link performance

Figure 3-4: Baseline Earth link performance – stored data volume

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 46 of 105

Figure 3-5: Baseline Earth link performance – bit rate

Figure 3-6: Baseline Earth link performance – latency

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 47 of 105

Considering 0.5 Gb generated on-board (or returned from another Mars asset) per day, these results show that for about 50% of the time the data can be relayed back to Earth within one day (during next ground station contact). For the remaining time, the latency can be as high as 52 days for data collected by April 2028. This is caused by the range between Mars and Earth, which can vary significantly during the year. On top of higher propagations losses for the communications link with Earth during this period, for up to 3 weeks the superior solar conjunction interrupts the communication completely. However, surface assets plan for this conjunction and should reduce their data output accordingly. Outside of these peak periods, it can be seen that the curves reach zero which means that a higher data downlink is possible, which will then be dedicated to the Science data downlink (approx. 1 Gbit/day for Science for 60% of the year). The use of “selective data downlink” allows the system to prioritise the downlink of science instrument data when the data downlink budget is low. This is an important concept of operation when the science data is required quickly e.g. for monitoring daily weather, clouds, dust activities. During the less favorable season the stored on-board data reaches its maximum at around 27 Gb in the second half of 2030. 27 Gbits of data storage doesn’t represent a constraint considering the baselined onboard computer (OBC). It should be noted that such a poor relay performance is not considered ideal for a dedicated Mars telcommunications orbiter, but may be sufficient to demonstrate the concept if the cost constraints are maintained. Future iterations of this concept should look to improve significantly on this performance if possible. Further, with these contraints, the poorer data-rate performance at high range could be offset by the increased operational flexibility provided by having multiple orbiters available.

PARAMETER Value Notes

RANGE [AU] 2.7 RANGE [km] 399426315.6 FREQUENCY [MHz] 8425 TX POWER [W] 50 TX ANTENNA GAIN [dBi] 36.61 0.5 deg Point. Err. TX LOSSES [dB] 2 TX EIRP [dBW] 51.60 Calculated PATH LOSSES [dB] 282.98 Calculated ATMOSPHERE LOSS [dB] 0.50 RX POL. LOSS [dB] 0.20 RX G/T [dBK] 51.70 DEMOD. LOSS [dB] 1.00 Estimation MOD. LOSS [dB] 1.00 S/N0 [dB] 46.22 REQUIRED Eb/No [dB] 0.14 Turbo 1/4 MINIMUM MARGIN [dB] 3.00 Standard ESA MAX BIT RATE [dBHz] 43.08 MAX BIT RATE [Mbps] 0.020 MAX BIT RATE [Kbps] 20.313 Figure 3-7: Preliminary Earth link budget @ 2.67 AU

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 48 of 105

3.4.3.3.3 Baseline Mars link performance

Figure 3-8: Baseline Mars link data rate performance@ 100% duty cycle

Figure 3-9 Baseline Mars link data volume performance @ 100% duty cycle

PARAMETER Value Notes ALTITUDE [km] 17032.0 17032.0 ELEVATION [deg] 5.0 RANGE [km] 19850.7 FREQUENCY [MHz] 400 TX POWER [W] 9 TX ANTENNA GAIN [dBi] -3.15 TX LOSSES [dB] 1 TX EIRP [dBW] 5.39 Calculated PATH LOSSES [dB] 170.44 Calculated ATMOSPHERE LOSS [dB] 5.00 Atmosphere + MP RX POL. LOSS [dB] 0.20 RX G/T [dBK] -20.62 5 dBi, 11.23 deg off-bor. DEMOD. LOSS [dB] 1.00 Estimation MOD. LOSS [dB] 0.00 S/N0 [dB] 36.73 REQUIRED Eb/No [dB] 6.20 RS in Elektra MINIMUM MARGIN [dB] 3.00 Standard ESA MAX BIT RATE [dBHz] 27.53 MAX BIT RATE [Mbps] 0.001 MAX BIT RATE [Kbps] 0.567 Figure 3-10: Preliminary Mars link budget @ 5 deg elevation

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 49 of 105

3.4.3.3.4 Baseline Mars link navigation capabilities By taking advantage of the existing UHF Proximity link: • Navigational information can be derived using the 2-way Doppler measurement provided by the baselined transponder. The signal transmitted by one node is frequency locked by another node and transmitted back to the first one. The first node can then calculate the difference between the frequency of transmitted and received signals. Given that the orbiter’s orbit is precisely known, the Doppler shift profile of the transmitted signal from the Mars surface will be specific to a set of locations (the real one plus ambiguity). After a few passes over the same asset, the ambiguity can be resolved and the accuracy increased. This has been used for other assets on the surface or during EDL. The magnitude of the Doppler (frequency) shift it directly proportional to the relative speed between the two elements. In a TASO / Areostationary orbit the relative speed with respect to a rover or a static asset is low or virtually none which makes it difficult to observe a significant Doppler shift. In this scenario this technique might provide lower accuracy than current measurements taken by orbiters in low Mars orbit. • A more complex system might use direction finding techniques by arraying UHF antennas on the orbiter in order to detect the angle of arrival of a signal. However, given the distance between the orbiter and the surface (~17000 km), if the system is able to discriminate the angle of arrival with 1 deg of precision, it represents best case scenario (around Martian equator) a region with a radius of around 150km. A more in-depth study of the available techniques is needed to assess what precision can be achieved using the current architecture baseline. 3.4.3.3.5 Improving the communications subsystem performance To achieve better performances in terms of data volume / latency in the Earth downlink the following aspects of the design can be considered: • Increase transmitted power (data volume directly proportional) • Increase antenna size (data volume proportional to the square) • Increase G/S contact time (data volume directly proportional) • Use Ka-Band with same antenna size (data volume increases by a factor of ~3) • Use MSPA to access more than one spacecraft at the same time • Variable data generation rate taking advantage of the “good season” for communications. To achieve better performances in terms of data volume / latency in the Mars return link: • Increase antenna size (data volume proportional to the square) • Use higher bands with Medium Gain Antennas (e.g. X or K-band, however these would have to be supported by the assets too). To achieve lower mass: • X-Band transponder integrated with UHF transceiver (similar to IRIS and additionally with UHF TX)

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 50 of 105

3.4.4 Science Instrument Selection The Mars Communication Constellation satellites has a primary objective of providing a communications network around Mars. However, an allocation was made for a secondary science payload on each satellite. As each satellite in the constellation is identical, the same science payload will fly on each of the three satellites. 3.4.4.1 Assumptions The following assumptions and preliminary allocations were made in order to inform the selection of the science instruments: • The scientific payload should not occupy more than 5.83 kg, (plus 20% margin = 7 kg.) • The scientific payload and data relay from surface assets should not produce more than 0.5 Gbits/day in average data volume • The science data will be stored on a dedicated instrument data handling unit. • Science payload selection should build on ESA’s heritage of past/present planetary missions with high TRL. • The science payload should perform simultaneous day and night-time full-disc monitoring of Mars. • The payload assumes that in normal operations the spacecraft always presents the same face towards Mars. • The science data requires “selective downlink” – prioritized instruments, alternatively prioritized observations – and data compression with storage capacity. 3.4.4.2 Selection of the scientific objectives The scientific objectives enabled by the TASO orbit and the possibility of doing multi- point observations using the constellation are in Table 3-11.

Objective Benefit Weakness Priority Science 1 : Global Coordinated multi-point monitoring of space EP greatly influences the Medium measurements on plasma weather at high instrument(s) performance Priority dynamics and atmospheric loss temporal resolution Long-term observation of the Martian Moons eXploration Science 2 : Martian moons and their (MMX) in the mid-2020 (JAXA): High Observation of the dynamics, dust environment, would cover all proposed science Priority Martian moons composition, origin objectives, if successful Coordinated multi-point Long-lived orbiters with global measurements on global diurnal coverage will provide the Science 3 : Global, weather conditions: largest volume of atmospheric High km-scale weather temperature, composition, data to support model Priority monitoring winds, dust, clouds, water development and validation (Low content Mars Orbit) Table 3-11 : Scientific Objectives trade-off for the Mars Communication Constellation

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 51 of 105

As the global, km-scale weather monitoring is the objective combining high priority from the Mars Exploration Program Analysis Group (MEPAG), in support of preparation for future human exploration, and benefits the most from the multi-point observation capability of the constellation, Science Objective 3 will be the baseline for this mission. The main scientific objective is to monitor the daily weather conditions, evolution of dust storms, clouds as well as the energy balance of the atmosphere. 3.4.4.3 Selection of the instrument suite It was decided to use an integrated instrument suite, based on high heritage planetary instruments (VMC (Venus Express), NOMAD UVIS (TGO), MARA (MASCOT)), which follows the weather monitoring concepts used on typical Earth Geostationary Satellite (Meteosat, SEVIRI & GERB). An integrated instrument suite with multiple apertures and channels, but in a single unit with shared electronics, provides an efficient combination of sub-assemblies with high TRL, albeit lower integration readiness levels. Table 3-12 shows the final selection of the instrument suite for the Mars Communication Constellation Satellite.

Baseline Dimensions Investigation Heritage Power Mass TRL Science (cm) (W) (kg) x y z

Daily global weather Wide-Angle Imaging VMC (Venus 1.0 10.0 10.0 10.0 7 monitoring, dust storm, Camera Suite (VIS) Express) clouds 10.0 Atmospheric Wide-Angle Imaging VMC (Venus 1.0 10.0 10.0 10.0 7 composition O2, H2O, Camera Suite (NIR) Express) CO, CO2, N-species

Wide-Angle Imaging NOMAD UVIS Ozone (250-270 nm), 6 < 2 20.0 10.0 10.0 7 Camera Suite (UV) (TGO) Aurora effects

Thermal IR Radiometer MARA Temperature of the Radiometer 5 1.0 12.5 12.5 15.0 7 +Thermal (MASCOT) atmosphere radiator Table 3-12 : Final selection of the science instruments of the Mars Communication Constellation, per satellite

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 52 of 105

3.5 Baseline Design 3.5.1 Mission Architecture The mission architecture baselined for the Mars Communications Constellation remains the same as the reference case described in Table 3-3 except for one minor difference. The dry mass of the spacecraft increased past the initial allocation of 300 kg which meant that a rideshare option was not possible and a dual launch scenario would be needed.

3.5.2 Spacecraft design overview An overview of the context of the spacecraft design in shown in Figure 3-11.

Solar Array (1 DOF)

Science HGA Thruster Instruments (2 DOF) To Earth

UHF (Fixed) Axis - Y

X-Axis

Solar Array (1 DOF)

Figure 3-11: Spacecraft design overview for Mars Communication Constellation

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 53 of 105

3.5.3 Spacecraft Subsystem Design Table 3-13 shows the main characteristics of the baseline design of one of the Mars Communication Constellation Satellite.

Mars Communication Constellation Satellite (per satellite) – System baseline summary System Overview Mass Dry Mass (incl. margin) 449 kg (per satellite) Wet Mass 610 kg (per satellite) Dimensions Stowed 1118 mm x 1200 mm x 2795 mm Deployed 1118 mm x 10144 mm x 2795

mm Instruments Wide-Angle Camera (VIS, NIR, UV) Thermal IR Radiometer Common Electronics

Subsystem Components

4x Reaction Wheel (1 Nms) Pressure Transducers, Filters, Valves 2x IMU Monopropellant Pipes Chemical AOCS 4x Star Tracker 1x Monopropellant Tank Propulsion 2x Star Tracker Electronic Unit 8x Monopropellant Hydrazine CHT-1 6x Sun Sensor Thruster

1x UHF Helix Antenna (fixed) 1x EP Filter Unit 1x UHF Receiver/Transmitter 1x EP Thrust Pointing Mechanism UHF RFDN 1x Mechanical Pressure Regulator 1x X-Band HGA (1 m diameter) - 1x T6 Gridded Ion Thruster (2 Degrees of Freedom) Electrical Comms 1x T6 PPU 1x X-Band HGA APM Propulsion 1x PRE Card 3x X-Band LGA (fixed) 1x Xenon Tank MT S-XTA90 1x X-Band Transponder 1x Xenon Flow Controller 1x X-Band TWTA (50W) Harness, Piping and Misc. X-Band RFDN

2x 12.5 m² Solar Arrays, 4G32 Ammonia Heat Pipes cells MLI 1x Secondary Battery (1137 Whrs) PNC Black Paint 1x Power Conversion and SSM Coating Distribution Unit (PCDU)-MPPT Power Thermal Thermistors Architecture. Thermal Fillers 2x Solar Array Drive Mechanism (SADM) Thermal Washers 1x Solar Array Drive Electronics Heaters for Spacecraft, Thrusters and (SADE) Tanks Data Composed of 6 main CFRP panels and 1x On-Board Computer (OBC) Structures Handling shear walls Table 3-13: Mars Communication Constellation Satellite (x3) – System baseline summary

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 54 of 105

3.5.4 Configuration

Figure 3-12 Stowed Configuration

Figure 3-13 Deployed Configuration 3.5.5 System Modes Four System modes are used to define the major states in which the spacecraft might be throughout its lifetime. It is used to circle the critical driving factor of each phase of the mission. They are also used to calculate the power budget as those system states do not have the same power consumption, and thus, dissipation/heating requirements. - Low Earth Orbit Phase (LEOP): From launcher umbilical separation to sun- pointing attitude. All equipment are OFF except for essential equipment. S/C powered by battery only (battery fully charged). AOCS actuators/sensors are ON. TT&C up- and down-link. Instruments are OFF. - Safe Mode (SAFE): Maintaining all equipment in a safe state. Solar arrays sun pointing. TT&C up- and down-link. - Transfer Mode (Transfer): EP Module ON. AOCS actuators/sensors to ensure sufficient attitude knowledge and pointing. - Nominal Operations Mode (Nominal): Providing the nominal global communication coverage on Mars. Providing nominal science operations. Providing the nominal communication to Earth.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 55 of 105

3.5.6 Spacecraft Budgets This section will summarise the budgets made during the study. It includes Mass, Power, and Delta-V budgets. The margin philosophy used to calculate the system budgets is applied as stipulated in RD[1]. 3.5.6.1 Mass Budget The following assumptions are made in the calculation of the mass budget: • It is noted that the harness mass was assumed to be 5% of the dry mass. • 2% was added on top of the propellant masses given by the propulsion teams to account for the leftover and unusable propellant in the tanks. • As the main propulsion system is electric, a system margin of 20% was used. It follows MAR-MAS-040 in the Margin Philosophy of RD[1]. Table 3-14 shows the mass budget for one of the three identical Mars Communication Constellation Satellites, per subsystem, including dry and wet masses, and the relevant maturity margins and system margin.

Mass (kg) Attitude, Orbit, Guidance, Navigation Control 7.96 Chemical Propulsion 6.95 Communications 30.08 Electric Propulsion 70.39

Instruments 7.00 Power 158.90 Structures 56.64 Thermal Control 13.31 Data Handling 5.25 Harness 5% 17.82 Dry Mass 374.27 System Margin 20% 74.85 Dry Mass incl. System Margin 449.13 CPROP Propellant Mass 4.70 CPROP Propellant Margin 2% 0.09 EPROP Fuel Mass 153.80 EPROP Fuel Margin 2% 3.08 Wet Mass 610.80 Table 3-14: Mass budget of one of the Mars Communication Constellation Satellite Table 3-15 shows the mass of the whole constellation, including the launch adapters. This table is to be used as a comparison to the selected launcher performances.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 56 of 105

Mass (kg) Constellation Total Dry Mass incl. System Margin 1347.39

Constellation Total Wet Mass 1832.40

Launch Adapter 360

Launch Mass (Wet Mass + Adapter) 2192.40 Ariane 62 to GTO Performances 5000 Ariane 62 to GTO Performances (excl. margins) 4500 Launch Mass left for the main passenger 2307.6 Table 3-15 : Constellation Launch Mass Budget

3.5.6.2 Power Budget The power budget of one of the Mars Communication Constellation Satellite is presented in Table 3-16. The Power budget is based on the System modes defined in 3.5.4. The following table does not include the 30% system margin advised in MAR-PWR-040 as the system margin was applied by the Power expert on top of what is presented in Table 3-16. However, it does include the equipment maturity margin. (W) (W) (W) (W) System Mode Transfer LEOP SAFE Nominal Total 2837.4 563.0 561.6 548.6 Attitude, Orbit, Guidance, Navigation 20.9 14.5 7.9 20.9 Control Communications 24.2 69.0 24.2 91.0 Data Handling 15.0 15.0 15.0 15.0 Instruments 5.0 5.0 5.0 25.0 Power 28.4 33.4 46.0 43.0 Thermal Control 26.9 426.1 463.5 353.7 Electric Propulsion 2717.0 0.0 0.0 0.0 Table 3-16 : Power Budget of one of the Mars Communication Constellation Satellite Note: The “Transfer” power consumption value presented here above is applicable for a Transfer Phase near Mars. 3.5.6.3 Delta-V Budget The final budget presented in Table 3-17 is the delta-V budget. The margin requirements used for each type of manoeuvre was identified. Please note that a colour scheme blue/green is used to denote the execution of a manoeuvre using respectively Chemical Propulsion or Electrical Propulsion.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 57 of 105

Mission Manoeuvre Delta-v Budget Value Unit Phase type From GTO to C3 = 0 km2/s3 deterministic 3700 m/s From C3 = 0 km2/s2 to Vinf = 0 km/s deterministic 5700 m/s From Vinf = 0 km/s to Areostationary/TASO deterministic 1000 m/s Margin for stochastic Interplanetary Approach Transfer 0 m/s Navigation Manoeuver - MAR-DV-100 or 150 Margin on stochastic delta-v - MAR-DV-040 0 % Margin on deterministic delta-v - MAR-DV-010 5 % Total det. and stoch. Manoeuvres 10920 m/s Specific Additional Margin for Electric Propulsion during 10 % Margin from Cruise - MAR-DV-120 Mars Total det. and stoch. Manoeuvres incl. EP Exploration 11490 m/s margin Studies Nominal lifetime 2 yrs Extended lifetime 4 yrs Orbit maintenance delta-v per year 2 m/s/yr Orbit Maintenance Orbit maintenance delta-v 12 m/s Margin on orbit maintenance delta-v - MAR-DV- 0 % 020 Total orbit maintenance delta-v 12 m/s Disposal manoeuvre 0 m/s Disposal Margin on disposal manoeuvre 0 % Total disposal manoeuvre 0 m/s Total delta-v without margin 10412 m/s Total delta-v including margin 11502 m/s Table 3-17 : Delta-V Budget for the Mars Communication Constellation Satellite 3.5.7 Equipment List The equipment list of the Mars Communication Constellation Satellite is presented in Table 3-18.

No. of TRL Manufacturer Countr units /Supplier y AOGNC RW Honeywell HR04 1Nms 4 7 Honeywell USA IMU Sensonor STIM300 2 7 Sensonor Europe STR Sodern Auriga Electronic Unit 2 9 Sodern Europe STR Sodern Auriga Optical Head 4 9 Sodern Europe SUN LENS Bison 64 6 8 LENS Europe COMMS UHF Helix Antenna 1 9 Tryo Group Europe UHF Receiver 1 9 JPL (Electra) USA UHF Transmitter 1 9 JPL (Electra) USA X Band HGA 1 9 SENER Europe X Band LGA 3 9 RYMSA (TRYO Group) Europe X Receiver 1 4 IMT Europe

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 58 of 105

X Band TWTA 1 8 TAS Europe X Transmitter 1 4 IMT Europe UHF RFDN N/A 9 Radiall Europe X RFDN N/A 9 Radiall Europe X APM – HGA ADPM 1 7 SENER Europe CPROP Monopropellant Tank 1 9 Monopropellant Pipes N/A 9 Monopropellant Hydrazine Thruster 8 9 DH Onboard Computer (OSCAR) 1 9 Airbus Europe INS Imaging Suite 1 7 Europe PWR Battery (18650 NL) 1 6 ABSL Europe PCDU 1 5 ADS Crisa Europe Solar Array 2 5 STI Europe SADE (Elektra-4 or -5) 1 9 Kongsberg Europe SADM (Karma-4TG or -5TG) 2 9 Kongsberg (or Ruag) Europe STR Main Structure N/A 6 Europe TC Ammonia Heat Pipes N/A 9 Europe MLI N/A 9 Europe Black Paint N/A 9 Europe SSM Coating N/A 9 USA Thermistors N/A 9 Europe Thermal Fillers N/A 9 Europe Thermal Washers N/A 9 Europe Heaters N/A 9 Europe EPROP EP Filter Unit 1 8 ADS-CRISA Europe Mechanical Pressure Regulator 1 5 Moog – Bradford Europe Engineering PRE Card 1 5 RUAG Europe Xenon Tank MT S-XTA90 1 6 MT Aerospace Europe Xenon Flow Controller 1 8 Moog – Bradford Europe Engineering T6 PPU 1 8 ADS-CRISA Europe T6 Gridded Ion Thruster 1 8 QINETIQ UK Europe T6 PPM Thrust Pointing Mechanism 1 8 RUAG Europe Table 3-18 Equipment List of the Mars Communication Satellite 3.5.8 Resources Available to Payload After completing the iterative design process, the final design of the satellite allowed for the mass, power, volume and data that would be available to a strawman payload to be allocated. This resource allocation is described in Table 3-19.

Minimum Data Downlinked to Power (W) Mass (kg) Volume (m3) Earth (Gbits/day) 25.0 5.8 6 x 10-3 0.5 Table 3-19 : Resources available to a strawman payload for the Mars Communication Constellation Satellite

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 59 of 105

Note: The column “Minimum Data Downlinked to Earth” isn’t applicable during occultation. In these periods, the data downlinked will be 0 Gbits/day. Refer to 3.4.3 for more details. 3.5.9 Technology Developments Throughout the study, two types of technology developments were tracked: • Baseline technology development, necessary for the pursuit of the baseline design described in this report. • Alternative technology developments, necessary to reach higher system performances than the ones described in this report. In addition to standard adaptations that are required for all missions as part of normal work, Table 3-20 lists the technology developments necessary for the baseline design.

Equipment Supplier Current Comments Name (Country) TRL Delta-qualification for interplanetary Sensonor IMU Stim 300 7 environment. Already envisaged for M- (Europe) ARGO May require lifetime and radiation RW HR04 1 Nms Honeywell (USA) 7 harness upgrade X Band IMT (Europe) 4 EM delivered in 2021 (TRL 6) Transponder Modifications to accommodate the 1m X HGA & APM SENER (Europe) 7 dish Table 3-20 : Technology developments required for baseline equipment It should also be noted that, in general, development of flexible or hybrid solar array technology and longer lifetime electric propulsion thrusters would be a major enabler of low mass interplanetary satellites. 3.6 Programmatics and Risk 3.6.1 Requirements The mission aims to be a low-cost short timeframe project leveraging extensive technology heritage. The target date of the mission also is to be in-line with MSR. It was also recommended to consider the lessons learned and approach taken by Mars Express (when re-flying Rosetta technology) including scheduling and team sizes. On the basis of these assumptions the following sections assess the impact at programmatic level. 3.6.2 Assumptions To develop a schedule the assumptions are based on CHEOPS and Mars Express. The Mars Communication Constellation is made up of three identical spacecraft. The first will require some development, but the 2nd and 3rd spacecraft are recurrent units. In this case, it is assumed that reviews have a nominal duration of 30 days. To compress the schedule, it is proposed that the ESA contingency before launch campaign

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 60 of 105 start/Acceptance Review is reduced from 6 months to 4 months after the 3rd satellite AIT activities. The launch campaign duration is increased from 3 months to 4 months for a campaign of 3 satellites. Taking into account that the constellation shall comprise three satellites and considering that no existing platform has been identified, a hybrid model philosophy is the less risky approach (or the most conservative) with: • An STM: Structural and Thermal Model with a platform structure representative of the flight design, and dummy for equipment or STM for the most critical instruments. For Thermal test, the STM (including dummies) will be representative for a thermal point of view. • An EFM: Electrical and Functional Model, with EM, or EQM of the equipment of the satellite for early verification of electrical and functional verification • A PFM (sat 1) ProtoFlight Model for the 1st satellite for protoqualification • Two FM Flight Models for the 2nd and the 3rd third satellite (acceptance tests). Another factor taken into account is items which have a TRL <7. These have been identified to be: • X-Band Receiver and Transmitter: TRL 4. This is based on the M-ARGO RD[4] technology for purposes of being more mass optimal with TRL 6 in 2021. • Xe Propellant tank: TRL 6 • Battery, PCDU and solar arrays: TRL 6. An adaptation of each is required when compared to other missions. • Main Structure: TRL 6. This is custom for this mission. 3.6.3 Baseline Schedule

MARS CHEOPS EXPRESS SMARTIES Remarks 2013-2019 1999-2003 CDF To Phase A 6 mths 9 mths Phase A 4 mths 7 mths Phase B1 4 mths 3 mths 7 mths Phase B2 7 mths 7 mths 9 months (18 months studies + 1 month CDR) With STM (6 months Phase C 21 mths 18 mths 19 mths manufacturing + 8 months AIT) With EFM (9 months) AIT activities starts 4 months after Phase D 29 mths 16 mths 30 mths beginning of phase D, for a duration of 16 months.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 61 of 105

MARS CHEOPS EXPRESS SMARTIES Remarks 2013-2019 1999-2003 AIT FM sat2 starts just before beginning of Evt tests PFM sat1 AIT FM sat3 starts just before beginning of Evt tests FM sat2 Implementation 57 mths 41 months 58 mths Total (Phases B2/C/D) (4 yrs 9 mths) (3 yrs 5mths) (4 yrs 10 mths) Table 3-21: MCC Preliminary Schedule B2 takes into account the Mars Express schedule as well as CHEOPS for comparison. This approach should take into account the following considerations: • Phase B2 KO as early as 2023 should be possible. Taking into account the CHEOPS’ schedule, perform Phase A and B1 before CM2022 (November 2022) is credible. No major risks identified. However, there is no margin to the finalise Phase 0 and to prepare the ITT before the end of the year 2020. • B/C/D in ~4 years. This is based on the Mars Express’ schedule, however it must be noted that in that case: o The Prime was a Major European Manufacturer with strong experience in Scientific and Telecom satellites o 80% of items were similar to Rosetta, and 20% from other projects. o Mars Express platform and CPS were based on an Eurostar PF o The Prime managed the technical interfaces between the spacecraft and science payload and between the spacecraft and launcher. o The time from concept to awarding the design and development contract was cut from about five years to little more than one year. o Early selection of scientific instruments. o Modification one of these conditions may impact significantly the schedule. • For the Mars Communication Constellation with 3 satellites, the requirement of B/C/D in ~4 years is not realistic. Any schedule compression will increase the AIT costs and the risks. In this case, this requirement may need to be reconsidered. 3.6.4 Summary of Risk Register The following were identified as critical: 1. Long mission life time – lack of redundancy could lead to degradation of communications service. 2. Launch date – the earlier part of the target launch period 2028, could well be at risk, when the baseline currently is not considering an existing platform.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 62 of 105

3. Service reliability – due to the mass restrictions of the spacecrafts, there may not be capacity to allow for full redundancy at sub-system level. This in turn highlights the risk of service degradation if one spacecraft is lost due to sub- system failures. Risk assessment shows that case 1 of the MCC is feasible from risk viewpoint if cost- related risks are not considered. The main driver from the risk viewpoint is the guarantee of service availability over the lifetime with an acceptable likelihood (due to: GTO waiting/ radiation, galactic radiation, small S/C, no on-board redundancy per S/C, …). Note that navigation services would need all 3 spacecraft. 3.7 Cost 3.7.1 Methodology The cost estimation approach undertaken was: • Parametric cost estimation method o SPICE method and tool, supplemented by ESA standard CERs o Estimate detailed down to equipment level o CEDRE database used for calibration of SPICE o Further important references used - TGO & EnVision aerobraking assumptions - M-ARGO for Transponder • Design and technical parameters from the OCDT model • Accuracy of estimate: between +/-30% 3.7.2 Assumptions Three satellites providing inter-system redundancy, no intra-system redundancy for mission duration of 24 months. The mission architecture options investigated for cost estimation were the T6 Gridded Ion Engine with a shared commercial launch to GTO (the baseline design) and PPS1350 Hall Effect Thrusters with a dedicated launch (alternative option). Further assumptions are: • Model Philosophy: o 1 STM, 1 EM and 1 PFM (note: for non-recurrent cost estimation) • All costs estimated in K-Euro on price basis 2019 • All equipment is assumed have a TRL of 5 or higher by the start of Phase B2 (the technology development costs required to reach this level are not part of the SMARTIES project budget). • Launch prices have been based on an Ariane 6.2 launch. Two scenarios were estimated; one with a dedicated launch at ESA agreed rates, and a dual launch opportunity to GTO

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 63 of 105

• At the request of the study management team, a reduced ESA internal cost share was considered based on the expectation of utilizing a slimmed-down ESA project team (similar to Mars Express). • Typical percentage project-level risks were identified • As per a study management team suggestion, the Science Operations are limited in scope due to the simplicity of the payload. The assumed cost estimation scope is: • Industrial Costs of Platform MAIT and Development during implementation Phase (B2, C/D, E1), including Payload to S/C AIT • Launcher price • Industrial cost risk margins related to above • Mission Operations costs (to be refined with ESOC) • Science Operations cost • ESA Internal costs • ESA Internal cost risk margins. Not included in the estimates are: • Payload Development and MAIT • Phase A and B1 costs, which are assumed to be covered under the General Study Program Budget • Specific technology developments funded through advanced technology development programmes such as TDE or E3P/ExPeRT. Cost has been a clear driver for a number of trade-offs throughout this study. The most important point to note is the relative difference between the SEP PPS-1350/E and T6 options. The latter offers a of saving of 3% comparably. Should schedule and programmatics become prioritised, PPS-1350/E is a possible alternative. Nevertheless the selected T6 dual launch option for the MCC exceeds the cost constraints.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 64 of 105

This Page Intentionally Blank

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 65 of 105

4 MARS SCIENCE ORBITER 4.1 Mission Requirements The design of the Mars Science Orbiter spacecraft is strongly linked to the programmatic constraints and mission design drivers identified in Section 2.2 and Section 2.3. In addition to the general architectural mission requirements given in Section 2.3.1 and following the mission architecture trades performed in Section 2.6, the Mars Science Orbiter case study also has some specific mission requirements given in Table 4-1.

Mission Requirements ID Statement The mission architecture shall comprise a single orbiter providing science MIS-MSO-010 observations of the Martian surface from ~300km altitude. Science observations shall target potential landing sites for future human MIS-MSO-020 missions between Martian latitudes of 30°N and 30°S. Following arrival at the target orbit, the mission shall provide science MIS-MSO-030 operations for a minimum of 4 years The total mass delivered to the target orbit shall be 300 kg (TBC), including all satellites and all margins. MIS-MSO-040 Note: this mass target has been derived from the cost requirement (MIS-030) and is subject to change provided the cost requirement is met. The mission shall provide capability to support relay communications from MIS-MSO-050 surface assets The mission shall be compatible with a shared launch on an Ariane 6.2 MIS-MCC-060 launch vehicle. The mission shall be compatible with an interplanetary transfer to a Low MIS-MCC-070 Mars Orbit (TBC) from a Geostationary Transfer Orbit (GTO). The mission shall achieve the interplanetary transfer stipulated in MIS- MIS-MCC-080 MCC-070 using a Chemical Propulsion Architecture. Table 4-1: Mars Science Orbiter mission specific requirements 4.2 Science Objectives The main objective of the Mars Science Orbiter is to participate to the characterization of Human landing sites. It primarily consists in mapping the thermophysical properties, and composition of the surface focusing on hydrous minerals and characterizing hazards for future landed missions. Secondarily, it aims at analysing the change detection on the surface (e.g., new impacts, activities, atmospheric composition).

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 66 of 105

4.3 Mission Analysis 4.3.1 Orbit Options The science team requested a Martian orbit that could be as low as possible without introducing issues for station-keeping, lifetime or planetary protection. Mission analysis of orbits around 300 km showed the following characteristics: • Lower than 300 km: Lifetime issues • Significantly higher than 300 km: Resolution of science data diminishes • Close to 300 km: near 13:1 synchronicity of orbital period with sidereal Mars day, which means very large gaps in surface coverage. Depicted in Figure 4-1 over the course of 7 days.

Figure 4-1: Mean semi major axis 3696 km (300 km mean altitude) • Mean altitude wrt. equator 320 km: No strong synchronicity between orbital period and Mars rotation, groundtrack tightly covers entire Mars surface already after 7 days.

Figure 4-2: Mean semi major axis 3716 km (320 km mean altitude)

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 67 of 105

4.3.2 Orbit Selection Given the characteristics of the options for orbits around 300 km described above, an orbit with the following elements was selected as a reference for the Mars Science Orbiter: • Mean semi-major axis: 3716 km • Mean eccentricity: 0.009 • Mean inclination: 92.76 deg • Node drifts +360 deg / 1 Mars year (687 Earth days) • Mean argument of periares: 270 deg • Periares altitude over areoid: ~308 km • Apoares altitude over areoid: ~378 km • Node crossings ~15:00/03:00 mean local solar time • True local solar time of node crossings varies mainly due to Mars orbit eccentricity • Stationkeeping requirements < 2 m/s/year, may be combined with wheel offloadings.

Figure 4-3 Orbital characteristics of selected baseline orbit 4.3.3 Mars Orbit Insertion Trade-Off For a chemical propulsion satellite there are various options for performing the Mars orbit insertion. Considerations given to aerobraking and aerocapture and a baseline MOI insertion strategy are presented in Table 4-2.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 68 of 105

Aerobraking Aerocapture ExoMars TGO, several NASA Heritage Never flown before missions Duration of science orbit ~ months up to a year ~ hours acquisition Delta V required for: Delta V required for: - MOI to 4 sol orbit: 1165 - Perigee lift manoeuvre and Delta V m/s apogee correction - Aerobraking: Up to 120 manoeuvre after m/s aerocapture pass: ~200 m/s Heat loads during Low, no aeroshell needed Higher, aeroshell is needed atmospheric pass Control authority during No control authority needed Lift or drag modulation is needed atmospheric pass Mass in science orbit Driven by fuel needed Driven by TPS mass Table 4-2: Mars orbit insertion trade-off Aerocapture is an attractive option due to the quick orbit acquisition and low delta-V required. However, it is considered not feasible within the cost and time constraints as it has not flown before. Therefore aerobraking was chosen for the Mars Science Orbiter. 4.3.4 Mars Orbit Insertion (MOI) Strategy and Lowering to Science Orbit For MOI using chemical propulsion, the post-MOI orbit is assumed as a 250 km x 96,000 km HEO with a period of 4 sols. The initial inclination can be selected as required for the final target orbit, however, the node must also be set such that the desired final node crossing times are obtained. The strategy to obtain node and inclination is TBD. It may require some delta-v for inclination adjustments. Circularising this HEO down to the final low Mars orbit would cost 1320 m/s of impulsive delta-v without the use of aerobraking. For a small satellite solution, the savings in delta-V given by using aerobraking are necessary, despite the operational challenges and additional time of doing so. The delta-V needed during the aerobraking phase is drastically reduced to 120 m/s. • Recommendations for aerobraking design: o Leave factor of 2 as margin between target and limit loads o Define ballistic parameter B=m/(CD*A) for stable aerobraking attitude - “stable” with respect to direction of incident flow o Scale values in Excel workbook to obtain duration o Add several weeks to allow for walk-in and walk-out o Add some time for contingencies and safe modes - Every safe mode requires another walk-in. 4.4 Reference Scenario The reference scenario chosen for study for the Mars Science Orbiter is summarised in Table 4-3.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 69 of 105

Reference scenario for Mars Science Orbiter Number of satellites 1 Launch scenario Ariane 62, rideshare Initial injection orbit Geostationary Transfer Orbit Means of transfer to Chemical Propulsion Mars Orbit GTO to C3 = 10 km2/s2 1975 m/s C3 = 10 km2/s2 to Mars Capture Ballistic coast (interplanetary transfer) Delta-V 1247 m/s (assuming 22% gravity Mars Capture to 4 sol losses) 4 sol to SSO Up to 120 m/s (aerobraking) Total delta-V 3.342 km/s (without margin) Mars orbit type Mars Sun Synchronous Mean semi-major axis 3716 km Mean inclination: 92.76 deg ~15:00/03:00 mean local solar Node crossings time Further details see Section 4.3.2

Mars Orbit

C3=0 C3=~10 Vinf = 3km/s Vinf = 0km/s GTO (heliocentric TASO Mars SSO LEO (capture) 4 sol (escape) transfer)

Aerobraking CP Kickstage - 1975m/s Ballistic coast CP – 1247m/s 120 m/s

Table 4-3: Summary of mission architecture reference scenario for Mars Science Orbiter

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 70 of 105

4.5 Systems Analyses and Trade-Offs 4.5.1 Chemical Propulsion Architecture Trade-Off 4.5.1.1 Assumptions For the reference case scenario an assessment was made of the performance of the Lisa Pathfinder kick stage with a chemical propulsion spacecraft to Mars SSO. The following assumptions were made: • Bi-propellant HZ system • 4 x 22 N thrusters used for Orbit insertion (combined Isp 308s) • On orbit AOCS vis 8 x 1N monopropellant thrusters. • 3 equal sized propellant tanks (2 for Hydrazine and 1 for MON) • Tanks isolated after MOI for remaining mission • Single use at MOI allows baseline with a single pressure regulator • Redundancy at check valve level and thruster valve level only. • LPF kick stage can lift 500kg to C3 = 10 km2/s2 4.5.1.2 Initial iteration Using the assumptions provided above, the mass performance of the system would allow a spacecraft of ~320kg in the Mars target orbit. An initial iteration of the spacecraft design and mass budget showed that the mass would be close to, but exceed this value and at this stage of study it would be prudent to have an architecture more robust to mass increases. 4.5.1.3 Trade-off options Trade-off options for the chemical propulsion transfer are presented in Table 4-4.

Kick Launch and Option Modification Impact stage Transfer Requalify LPF for entire transfer to This would require tanks to be GTO to Mars Mars. completely filled, structural, thermal Option 1 LPF SSO dual and environmental requalification and launch Remove spacecraft a requalification for the increased CP architecture lifetime. needed for MOI This would require tanks to be GTO to Mars Requalify LPF for a completely filled. Thought to be a less Option LPF SSO dual higher launch complicated requalification than the 2 launch mass. using LPF for the full transfer, but development costs still high. C3 = 10 km2/s2 Remove KS A dedicated launch removes the need Option to Mars SSO on completely and for a kick stage entirely but the launch None 3 dedicated have a dedicated itself is more costly. Programmatically launch launch more straightforward. Table 4-4 Chemical propulsion architecture trade-off options

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 71 of 105

The trade-off of chemical propulsion architectures mainly comes down to cost and programmatic complexity. Following a detailed cost assessment, the cost delta between all three options was shown to be not significant, but all options exceeded the cost target. In the case where the costs are similar then the programmatically more straightforward option should be chosen. Although a dedicated launch option was ruled out earlier in the study, the costs associated with the chemical kick stage are such that the dedicated launch now becomes more appealing. Therefore the dedicated launch was chosen as the baseline launch and transfer architecture for the Mars Science Orbiter. 4.5.2 Communications Performance Analysis The primary payload on the Mars Science Orbiter is the scientific instrumentation, however a secondary objective is for the orbiter to provide a communiations relay to assets on the Mars surface. The communication subsystem therefore has to adhere to requirement MIS-MCC-050 (The mission shall provide capability to support relay communications from surface assets) and should follow a design to cost approach. 4.5.2.1 Assumptions The assumptions (and their origin) of the communication subsystem design are as follows: • General: o No performance requirements are assumed. The design approach is to see what performance can be accommodated within the cost constraints. • Earth link: o X-Band, 50W TWTA – [system] o 1m HGA – [system] o Occultation of 40 min every orbit (115 min) – [mission analysis] o Superior conjunction interrupting comms for SEM angles < 2 deg, degradation up to 5 deg - [link budget] o Maximum data rate with 3 dB link margin – [ECSS] o 6h of G/S contact time (EL 10 deg) – [operations] o 0.5 Gb / day generated data - [system] o Redundant transponder and TWTA - [system] o New Norcia 35m ground station at 10 degrees of elevation • Mars link: o UHF – [heritage of current Mars rovers] o LGA with the boresight pointed to Mars o Surface asset with LGA and 9W RF power - [Curiosity Rover as reference] o No subsystem redundancy – [System].

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 72 of 105

4.5.2.2 Communications subsystem baseline design The baseline design and communications links for the communications subsystem is presented in Figure 4-4.

Figure 4-4: Comms subsystem equipment layout 4.5.2.3 Communications Performance Analysis The applicable range plots for Earth, Mars and the Sun are given in Figure 3-3. 4.5.2.3.1 Baseline Earth link performance

Figure 4-5: Baseline Earth link performance – stored data volume

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 73 of 105

Figure 4-6: Baseline Earth link performance – latency Considering 0.5 Gb generated on-board (or returned from another Mars asset) per day, these results show that for about 50% of the time the data can be relayed back to Earth within one day (during next ground station contact). For the remaining time, the latency can be as high as 140 days for data collected by April 2028. This is caused the range between Mars and Earth which is at its highest value during the Martian year. On top of higher propagations losses for the communications link with Earth during this period, for up to 3 weeks the superior solar conjunction interrupts completely the communication. However, surface assets plan for this conjunction and should reduce their data output accordingly. During the less favorable season the stored on-board data reaches its maximum at around 70 Gb in the second half of 2028. 70 Gbits of data storage doesn’t represent a constraint considering the baselined onboard computer (OBC).

PARAMETER Value Notes

RANGE [AU] 2.7 RANGE [km] 399426315.6 FREQUENCY [MHz] 8425 TX POWER [W] 50 TX ANTENNA GAIN [dBi] 36.61 0.5 deg Point. Err. TX LOSSES [dB] 2 TX EIRP [dBW] 51.60 Calculated PATH LOSSES [dB] 282.98 Calculated ATMOSPHERE LOSS [dB] 0.50 RX POL. LOSS [dB] 0.20 RX G/T [dBK] 51.70 DEMOD. LOSS [dB] 1.00 Estimation MOD. LOSS [dB] 1.00 S/N0 [dB] 46.22 REQUIRED Eb/No [dB] 0.14 Turbo 1/4 MINIMUM MARGIN [dB] 3.00 Standard ESA MAX BIT RATE [dBHz] 43.08 MAX BIT RATE [Mbps] 0.020 MAX BIT RATE [Kbps] 20.313 Figure 4-7: Preliminary Earth link budget @ 2.67 AU

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 74 of 105

4.5.2.3.2 Baseline Mars link performance During a pass (in a 330km orbit) over a surface asset, the Mars link performance is influenced by: • Limit on the maximum data rate by the Electra UHF transceiver (2 Mbps) • Range from the surface asset to the orbiter, from ~1265km at 5 deg of elevation to 330km at 90 deg of elevation. • Radiation pattern of the LGA both on the surface asset and orbiter

Figure 4-8: Baseline Mars link data rate performance, 0.5Gb/pass average

PARAMETER Value Notes ALTITUDE [km] 330.0 ELEVATION [deg] 5.0 RANGE [km] 1265.5 FREQUENCY [MHz] 400 TX POWER [W] 9 TX ANTENNA GAIN [dBi] -3.15 TX LOSSES [dB] 1 TX EIRP [dBW] 5.39 Calculated PATH LOSSES [dB] 146.53 Calculated ATMOSPHERE LOSS [dB] 5.00 Atmosphere + MP RX POL. LOSS [dB] 0.80 RX G/T [dBK] -27.20 DEMOD. LOSS [dB] 1.00 Estimation MOD. LOSS [dB] 0.00 S/N0 [dB] 53.46 REQUIRED Eb/No [dB] 6.20 RS in Elektra MINIMUM MARGIN [dB] 3.00 Standard ESA MAX BIT RATE [dBHz] 44.26 MAX BIT RATE [Mbps] 0.027 MAX BIT RATE [Kbps] 26.690 Figure 4-9: Preliminary Mars link budget @ 5 deg elevation

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 75 of 105

4.5.2.3.3 Baseline Mars link navigation capabilities By taking advantage of the existing UHF Proximity link: • Navigational information can be derived using the 2-way doppler measurement provided by the Electra transponder. The signal transmitted by one node is frequency locked by another node and transmitted back to the first one. The first node can then calculate the difference between the frequency of transmitted and received signals. Given that the orbiter’s orbit is precisely known, the Doppler shift profile of the transmitted signal from the Mars surface will be specific to a set of locations (the real one plus ambiguity). After a few passes over the same asset, the ambiguity can be resolved and the accuracy increased. This has been used for other assets on the surface or during EDL. • This is also the technique used by Argos system in LEO to track wildlife transmitters with an accuracy better than 250m radius. 4.5.2.3.4 Improving the communications subsystem performance To achieve better performances in terms of data volume / latency in the Earth downlink the following aspects of the design can be considered: • Increase transmitted power (data volume directly proportional) • Increase antenna size (data volume proportional to the square) • Increase G/S contact time (data volume directly proportional) • Use Ka-Band with same antenna size (data volume increases by a factor of ~3) • Variable data generation rate taking advantage of the “good season” for communications To achieve better performances in terms of data volume / latency in the Mars return link: • Increase antenna size (data volume proportional to the square) • Implement an Antenna Pointing Mechanism (APM) • Use higher bands with Medium Gain Antennas (e.g. X or K-band, however these would have to be supported by the assets too) To achieve lower mass: • X-Band transponder integrated with UHF transceiver (similar to IRIS and additionally with UHF TX) 4.5.3 Science Instrument Selection The Mars Science Orbiter has a primary objective of returning useful science data, however the mission remains cost and schedule driven rather than science driven. In order to return the most useful science for the resources available, the science team were able to define the most useful Mars orbit for their science needs, but were given a preliminary allocation to meet for the science payload. The target orbit at Mars is SSO with an LTAN of 15:00, 3:00 true local time and is defined in detail in Section 4.3.2.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 76 of 105

4.5.3.1 Science objectives The main objective of the Mars Science Orbiter is to participate to the characterization of Human landing sites. It primarily consists of mapping the thermophysical properties, and composition of the surface focusing on hydrous minerals and characterizing hazards for future landed missions. Secondarily, it aims at analysing the change detection on the surface (e.g., new impacts, activities, atmospheric composition). 4.5.3.2 Assumptions The following assumptions and preliminary allocations were made in order to inform the selection of the science instruments: • The scientific payload should not occupy more than 25 kg, (plus 20% margins = 30 kg) • The scientific payload should not consume more than 75W of power at any instant. • The scientific payload and data relay from surface assets should not downlink more than 0.5 Gbit/day in average. • Science payload selection should build on ESA’s heritage of past/present planetary missions with high TRL. • SSO with node crossing the equator at 15:00 and 03:00 true local times allows high quality science data acquisition. • The payload assumes that in normal operations the spacecraft always presents the same face towards Mars within 0.1°. • The science operation assumes daily mean observation of heritage instruments from previous missions. • The instrument unit houses its own processing electronics for image compression, and a buffer allowing accumulation of payload data for 140 sols before it is transferred to the platform for downlink. • “Selective downlink” is given for high priority science data. 4.5.3.3 Selection of the instrument suite Following the definition of the primary and secondary objectives of the science payload, High TRL heritage instruments from ESA’s heritage planetary missions and European contributions to non-ESA missions were selected in order to build the instrument suite given in Table 4-5.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 77 of 105

Baseline

Dimensio Data Power ns (cm) Volume Investigation Justification Heritage (W) Mas TRL Science s (Gbits/ day) (kg) x y z Op Stan s d-by Thermal IR Surface Radiometer/ MERTIS composition Imaging , Spectrometer MERTIS temperature Data coverage 18. 18. 13. of high (BepiColom 13 5 3.4 5 , thermal Imaging 0 0 0 TIS bo) 0.04 inertia, rock Spectrometer resolution spectral and abundance, TIR Radiometer optical data is atmospheric limited (few science

%). The High instruments resolution address high colour priority imaging of the surface, Visible Imaging science CaSSIS 67. 43. 42. CaSSIS 7 0 15.2 0.084 9 change System questions for (TGO) 5 8 0 human detection, exploration. stereo imaging, ESA has geological extensive context experience in cameras and MicrOmega spectrometers (Hayabusa- MicrOme NIR . 2/MASCOT) 15. 10. 12. 5 2 2.0 0.89 5 Mineralogy ga Spectrometer , 0 0 0 MacrOmega (MMX)

Table 4-5: Final selection of the Instruments of the Mars Science Orbiter 4.6 Baseline Design 4.6.1 Mission Architecture The mission architecture baselined for the Mars Science Orbiter differs from the reference scenario studied due to the outcome of the trade-offs performed in Section 4.5.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 78 of 105

Baseline mission architecture for Mars Science Orbiter Number of satellites 1 Launch scenario Ariane 62, dedicated Initial injection C3 = 10 km2/s2 Means of transfer to Chemical Propulsion Mars Orbit C3 = 10 km2/s2 to Mars Capture Ballistic coast (interplanetary transfer) 1247 m/s (assuming 22% gravity Mars Capture to 4 sol Delta-V losses) 4 sol to SSO Up to 120 m/s (aerobraking) Total delta-V 1367 m/s (without margin) Mars orbit type Mars Sun Synchronous Mean semi-major axis 3716 km Mean inclination: 92.76 deg Node crossings ~15:00/03:00 mean local solar time Further details see Section 4.3.2

Mars Orbit

Coverage Groundtrack tightly overs entire Mars surface within 7 days

C3=0 C3=~10 Vinf = 3km/s Vinf = 0km/s GTO (heliocentric TASO Mars SSO LEO (capture) 4 sol (escape) transfer)

Aerobraking Ballistic coast CP – 1247m/s 120 m/s

Figure 4-10: Summary of baseline mission architecture for Mars Science Orbiter

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 79 of 105

4.6.2 Spacecraft Design Overview An overview of the context of the spacecraft design is shown in Figure 4-11.

Solar Array (1 DOF)

Science HGA (2 DOF) To Earth Launcher interface Instruments

UHF (Fixed) Axis - Y

X-Axis

Solar Array (1 DOF)

Figure 4-11: Spacecraft design overview for Mars Science Orbiter 4.6.3 Spacecraft Subsystem Design Table 4-6 shows the main characteristics of the baseline design of the Mars Science Orbiter Satellite.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 80 of 105

Mars Science Orbiter – System baseline summary System Overview Mass Dry Mass (w/ 369.6 kg margin) Wet Mass 597.5 kg Dimensions Stowed 1334 mm x 1500 mm x 2668 mm Deployed 1334 mm x 6836 mm x 2668 mm Instruments MERTIS - Thermal IR Radiometer/Imaging Spectrometer CaSSIS - Visible Imaging System MicrOmega - NIR Spectrometer Camera Electronics & Payload and Relay Data

Handling Unit Subsystem Components 4x Reaction Wheel Honeywell HR04 (1 Nms) 1x Bipropellant Pressure Regulator 2x IMU Sensonor STIM300 1x Bipropellant Pressurant Tank 4x STR Sodern Auriga (Optical Chemical 3x Bipropellant Propellant Tank AOCS Head) Propulsion 4x Bipropellant Thruster 2x STR Sodern Auriga (Electronic 8x Monopropellant Hydrazine Thruster Unit) Transducers, Valves, Filters, Pipes 6x Sun Sensor LENS Bison 64 1x UHF Helix Antenna (Fixed) 1x UHF Receiver/Transmitter UHF RFDN 1x X-Band HGA (1 m diameter, 2 Degrees of Freedom) Comms Ammonia Heat Pipes 1x X-Band HGA APM MLI 3x X-Band LGA PNC Black Paint 2x X-Band Transponder SSM Coating 2x X-Band TWTA (50W) Thermistors X-Band RFDN Thermal Thermal Straps 2x 3.56 m² Solar Arrays Thermal Fillers 1x Secondary Battery (1279 W.h) Thermal Washers 1x Power Conversion and Heaters for Spacecraft, Thrusters and Distribution Unit (PCDU). MPPT Tanks Power Architecture. 2x Solar Array Driving Mechanism (SADM) 1x Solar Array Drive Electronics (SADE) Data 1x On-Board Computer (OBC) Composed of 6 main CFRP panels and Structures Handling (internally redundant) shear walls Table 4-6: Mars Science Orbiter Satellite– System baseline summary

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 81 of 105

4.6.4 Configuration

Figure 4-12: Stowed Configuration

Figure 4-13: Deployed Configuration 4.6.5 System Modes Four System modes are used to define the major states in which the spacecraft might be throughout its lifetime. It is used to circle the critical driving factor of each phase of the mission. They are also used to calculate the power budget as those system states do not have the same power consumption, and thus, dissipation/heating requirements. - Low Earth Orbit Phase (LEOP): From launcher umbilical separation to sun- pointing attitude. All equipment are OFF except for essential equipment. S/C powered by battery only (battery fully charged). AOCS actuators/sensors are ON. TT&C up- and down-link. Instruments are OFF. - Safe Mode (SAFE): Maintaining all equipment in a safe state. Solar arrays sun pointing. TT&C up- and down-link. - Transfer Mode (Transfer): CP Module ON. AOCS actuators/sensors to ensure sufficient attitude knowledge and pointing. - Nominal Operations Mode (Nominal): Providing nominal science operations. Providing the nominal communication to Earth.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 82 of 105

4.6.6 Spacecraft Budgets This section will summarise the budgets made during the study. It includes Mass, Power, and Delta-V budgets. The margin philosophy used to calculate the system budgets is applied as stipulated in RD[1]. 4.6.6.1 Mass Budget The following assumptions are made in the calculation of the mass budget: • It is noted that the harness mass was assumed to be 5% of the dry mass. • 2% was added on top of the propellant masses given by the propulsion teams to account for the leftover and unusable propellant in the tanks. • As the main propulsion system is chemical, a system margin of 30% was used. It follows MAR-MAS-040 in the Margin Philosophy of RD[1]. Table 4-7 shows the mass budget for the Mars Science Orbiter, per subsystem, including dry and wet masses, and the relevant maturity margins and system margin.

Mass (kg) Attitude, Orbit, Guidance, Navigation Control 7.96 Chemical Propulsion 41.37 Communications 34.30 Instruments 34.24 Power 74.48 Structures 63.39 Thermal Control 9.78 Data Handling 5.25 Harness 5% 13.54 Dry Mass SC 284.30 System Margin 30% 85.29 Dry Mass SC incl. System Margin 369.59 CPROP Fuel Mass 124.80 CPROP Fuel Margin 2% 2.50 CPROP Oxidizer Mass 97.67 CPROP Oxidizer Margin 2% 1.95 CPROP Pressurant Mass 1.00 CPROP Pressurant Margin 2% 0.02 Total Wet Mass SC 597.53 Launcher Interface (remain. on launcher or jetizoned) 64 Launched Mass 661.53 L2 Launch Performance 2700 L2 Launch Performance (incl. margins) 2430 Launch Mass remaining for the Secondary 1768.47 Passenger

Table 4-7 : Mass Budget of the Mars Science Orbiter

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 83 of 105

4.6.6.2 Power Budget The power budget of the Mars Science Orbiter is presented in Table 4-8. The Power budget is based on the System modes defined in Section 4.6.5. The following table does not include the 30% system margin advised in MAR-PWR-040 as the system margin was applied by the Power expert on top of what is presented in Table 4-8. It eased the application of the relevant margins following MAR-PWR-050 however, it does include the equipment maturity margin.

(W) (W) (W) (W) System Mode Transfer LEOP SAFE Nominal 290.7 310.4 309.1 295.4 Total Attitude, Orbit, Guidance, Navigation 20.9 14.5 7.9 20.9 Control Communications 37.2 82.0 37.2 104.0 Chemical Propulsion 1.8 0.0 0.0 0.0 Data Handling 15.0 15.0 15.0 15.0 Instruments 19.0 19.0 19.0 61.0 Power 28.4 33.4 46.0 43.0 168.5 146.5 184.0 51.5 Thermal Control Table 4-8 : Power Budget of the Mars Science Orbiter 4.6.6.3 Delta-V budget The final budget presented in Table 4-9 is the delta-V budget. The margin requirements used for each type of manoeuvre was identified.

Mars Mission Manoeuvre Delta-v Budget Science Unit Phase type Orbiter From Vinf = 3km/s to 4 sol (includes 22% deterministic 1247 m/s gravity loss) Aerobraking from 4 sol to SSO 330 km, stochastic 120 m/s 15:00H-03:00H

Margin for stochastic Interplanetary 10 m/s Transfer Approach Navigation Manoeuver - MAR- DV-100 or 150 Margin on stochastic delta-v - MAR-DV- 0 % 040 Margin on deterministic delta-v - MAR- 5 % DV-010 Total det. and stoch. Manoeuvres 1439 m/s Nominal lifetime 4 yrs Orbit Extended lifetime 0 yrs Maintenance Orbit maintenance delta-v per year 2 m/s/yr

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 84 of 105

Mars Mission Manoeuvre Delta-v Budget Science Unit Phase type Orbiter Orbit maintenance delta-v 8 m/s Margin on orbit maintenance delta-v - 0 % MAR-DV-020 Total orbit maintenance delta-v 8 m/s Disposal manoeuvre 0 m/s Disposal Margin on disposal manoeuvre 0 % Total disposal manoeuvre 0 m/s Total delta-v without margin 1375 m/s Total delta-v including margin 1447 m/s Table 4-9 : Delta-V Budget for the Mars Science OrbiterEquipment List The equipment list of the Mars Communication Constellation Satellite is presented in Table 4-10.

Number TRL Manufacturer Country of units /Supplier AOGNC RW Honeywell HR04 1Nms 4 7 Honeywell USA IMU Sensonor STIM300 2 7 Sensonor Europe STR Sodern Auriga Electronic Unit 2 9 Sodern Europe STR Sodern Auriga Optical Head 4 9 Sodern Europe SUN LENS Bison 64 6 8 LENS Europe COMMS UHF Helix Antenna 1 9 Tryo Group Europe UHF Receiver 1 9 JPL (Electra) USA UHF Transmitter 1 9 JPL (Electra) USA X Band HGA 1 9 SENER Europe X Band LGA 3 9 RYMSA (TRYO Europe Group) X Receiver 2 4 IMT Europe X Band TWTA 2 8 TAS Europe X Transmitter 2 4 IMT Europe UHF RFDN N/A 9 Radiall Europe X RFDN N/A 9 Radiall Europe X APM – HGA ADPM 1 7 SENER Europe CPROP Monopropellant Thruster 8 9 Bipropellant Pipes N/A 9 Bipropellant Pressure Regulator 1 9 Ariane Group Europe Bipropellant Pressurant Tank 1 9 Bipropellant Propellant Tank 3 9 Bipropellant Thruster 4 9 MOOG DH Onboard Computer (OSCAR) 1 9 Airbus Europe INS Imaging Suite 1 7 University of Bern Europe Spectrometer Suite 1 5 CNES Europe Thermal IR Suite 1 5 University of Europe Muenster Camera Electronics 1 9 Europe Payload and Relay Data Handling Unit 1 9 Europe PWR Battery (18650 NL) 1 6 ABSL Europe

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 85 of 105

Number TRL Manufacturer Country of units /Supplier PCDU 1 5 ADS Crisa Europe Solar Array 2 5 STI Europe SADE (Elektra-4 or -5) 1 9 Kongsberg Europe SADM(Karma-4TG or -5TG) 2 9 Kongsberg (or Ruag) Europe STR Main Structure N/A Europe TC Ammonia Heat Pipes N/A 9 Europe MLI N/A 9 Europe Black Paint N/A 9 Europe SSM Coating N/A 9 USA Thermistors N/A 9 Europe Thermal Fillers N/A 9 Europe Thermal Washers N/A 9 Europe Heaters N/A 9 Europe Thermal Straps N/A 9 Europe Table 4-10 : Equipment list of the Mars Science Orbiter 4.6.7 Resources Available to Payload After completing the iterative design process, the final design of the satellite allowed for the mass, power, volume and data allocations for a strawman payload described in Table 4-11:

Minimum Data Power (W) – Mass Volume (m3) Downlinked to Operation/Stand-By (kg) Earth (Gbits/day) 61 19 31.4 0.15 0.5 Gbits Table 4-11: Resources available to payload for the Mars Science Orbiter Note: The column “Minimum Data Downlinked to Earth” isn’t applicable during occultation. In these periods, the data downlinked will be 0 Gbits/day. Refer to Section 4.5.2 for more details. 4.6.8 Technology Developments Throughout the study, two types of technology developments were tracked: • Baseline technology development, necessary for the pursuit of the baseline design described in this report. • Alternative technology developments, necessary to reach higher system performances than the ones described in this report. In addition to standard adaptations that are required for all missions as part of normal work, Table 4-12 lists the technology developments necessary for the baseline design.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 86 of 105

Supplier Current Equipment Name Comments (Country) TRL Delta-qualification for interplanetary IMU Stim 300 Sensonor (Europe) 7 environment. Already envisaged for M-ARGO May require lifetime and RW HR04 1 Nms Honeywell (USA) 7 radiation harness upgrade EM delivered in 2021 X Band Transponder IMT (Europe) 4 (TRL 6) Modifications to X HGA & APM SENER (Europe) 7 accommodate the 1m dish

Table 4-12 : Technology developments required for baseline equipment 4.7 Programmatics and Risk The assessment of the Mars Science Orbiter (MSO) is complementary to that of the Mars Communications Constellation. Again, the aim is to be a low-cost short timeframe project leveraging extensive technology heritage. The target date of the MSO mission is to be in-line with Mars Sample Return (MSR) as this is one further option for a small Mars mission. 4.7.1 Assumptions The MSO is a single spacecraft. To develop a schedule the assumptions are based on CHEOPS and Mars Express. To compress the schedule, it is proposed that the ESA contingencies before launch campaign start/Acceptance Review is reduced is 6 months. The launch campaign duration is 3 months. Considering that no existing platform has been identified, a hybrid model philosophy is the less risky approach (or the most conservative) with: • A STM: Structural and Thermal Model with a platform structure representative of the flight design, and dummy for equipment or STM for the most critical instruments. For Thermal test, the STM (including dummies) will be representative for a thermal point of view. • An EFM: Electrical and Functional Model, with EM, or EQM of the equipment of the satellite for early verification of electrical and functional verification. • A PFM: ProtoFlight Model with protoqualification tests. Another factor taken into account, is items which have a TRL <7. These have been identified (and are similar to the MCC aside from increased equipment level redundancy) to be: • Spectrometer Suite: TRL 5. Some design optimisation needed, but heritage is drawn from the MMX mission. • Thermal IR Suite: TRL 5. Re-design of BepiColombo MERTIS instrument. However with less stringent thermal requirements.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 87 of 105

• 2 x X-Band Receiver and Transmitter: TRL 4. This is based on the M-ARGO RD[4] technology for purposes of being more mass optimal with TRL 6 in 2021. • Xe Propellant tank: TRL 6 • Battery, PCDU and solar arrays: TRL 6. An adaptation of each is required when compared to other missions. 4.7.2 Baseline Schedule

MARS CHEOPS EXPRESS SMARTIES Remarks 2013-2019 1999-2003 CDF To Phase A 6 mths 9 mths Phase A 4 mths 7 mths Phase B1 4 mths 3 mths 7 mths Phase B2 7 mths 7 mths 9 months (20 months studies + 1 month CDR) With STM (6 months 19 mths manufacturing + 9 Phase C 21 mths 18 mths (21 with LSP KS) months AIT: Impact LSP Kick stage +1 month ) With EFM (9 months) AIT activities starts 4 months after 20 mths beginning of phase D, Phase D 29 mths 16 mths for a duration of 18 (22 with LSP KS) months (impact LSP Kick Stage + 2months). Implementation 57 mths 41 months 48 mths (Phases (4 yrs 9 mths) (3 yrs 5mths) (4 yrs 4 mths) B2/C/D) Table 4-13: Mars Science Orbiter Preliminary Schedule B2 takes into account the Mars Express schedule as well as CHEOPS for comparison. This approach should take into account the following considerations: • Phase B2 KO as early as 2023 should be possible. Taking into account the CHEOPS’ schedule, perform Phase A et B1 before CM2022 (November 2022) is credible. No major risks identified. However, there is no margin to the finalise Phase 0 and to prepare the ITT before the end of the year 2020. • B/C/D in ~4 years. This is based on the Mars Express’ schedule, however: o The Prime was a Major European Manufacturer with strong experience in Scientific and Telecom satellites o 80% of items were similar to Rosetta, and 20% from other projects.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 88 of 105

o Mars Express platform and CPS were based on an Eurostar PF o The Prime managed the technical interfaces between the spacecraft and science payload and between the spacecraft and launcher. o The time from concept to awarding the design and development contract was cut from about five years to little more than one year. o Early selection of scientific instruments. o Modification one of these conditions may impact significantly the schedule. • For the Science Orbiter Mission the requirement for Phases B/C/D in ~4 years seems feasible but it is very dependent of the design of the satellite, the organization of the Project (ESA) and the experience of the industrial Prime Contractor. 4.7.3 Summary of Risk Register The following were identified as critical: 1. Mass budget – the constraints on mass budget were set initially at 300 kg, this was considered too low and was increased to 500 kg. However the final result is around the latter and could impact the launch scenario. The system mass margin of 30% provides a good mitigation to mass increases in later phases. 2. Planetary Protection - low operational orbit and a failure resulting in re-entry. 3. Redundancy – propulsion system and HGA pointing mechanism not fully redundant could lead to trajectory errors and degradation of comms link performance respectively. Potentially reducing mission lifetime. 4. Aero-braking – required for operational orbit. There is the potential for damage to the spacecraft and control issues. 5. Cost overrun – the cost constraint could be exceeded. A consequence of needed spacecraft design. 6. Launch date – the earlier part of the target launch period 2028, could well be at risk, when the baseline currently is not considering an existing platform. The risk assessment shows that the MSO is generally feasible beside cost-related risk concerns. The main drivers of risk are related to the relatively low operational orbit: e.g. aero-braking, orbit instabilities, planetary protection. But also due to mass constraints, requiring some reduction of on-board redundancy. Furthermore, the study showed that the cost budget was slightly exceeded. 4.8 Cost The cost estimation approach undertaken was: • Parametric cost estimation method o SPICE method and tool, supplemented by ESA standard CERs o Estimate detailed down to equipment level o CEDRE database used for calibration of SPICE o Further important references used - TGO & EnVision aerobraking assumptions - M-ARGO for Transponder

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 89 of 105

• Design and technical parameters from the OCDT model • Accuracy of estimate: between +/-30% 4.8.1 Assumptions A single spacecraft is considered with redundancy with a 48 month mission duration including aero-braking. The mission architecture options investigated for cost estimation were chemical propulsion system using a dedicated launch and a chemical propulsion system with a Lisa Pathfinder Kickstage on a shared commercial launch to GTO. Further assumptions are: • Model Philosophy: 1 STM, 1 EM and 1 PFM • All costs estimated in K-Euro on price basis 2019 • All equipment is assumed have a TRL of 5 or higher by the start of Phase B2 (the technology development costs required to reach this level are not part of the SMARTIES project budget). • Launch prices have been based on an Ariane 6.2 launch. Two scenario’s were estimated; one with a dedicated launch at ESA agreed rates, and a rideshare opportunity to GTO • At the request of the study management team, a reduced ESA internal cost share was considered based on the expectation of utilizing a slimmed-down ESA project team (similar to Mars Express). • Typical percentage project-level risks were identified • CP + Dedicated Launch – baseline: o For a similar cost with respect to the alternative scenario a dedicated launch is available, mitigating risk of additional LPF delta qualification. For launch costs up to €100M, this is a more cost effective option. o In the future update on Arianespace & ESA Frame contract prices for Ariane 6.2 launch is expected • CP + LPF + Dual launch – alternative: o Shared launch at commercial launch prices in cost/kg o It is assumed the LPF kickstage requires minor engineering and a PFM, which is an optimistic case. The assumed cost estimation scope is: • Industrial Costs of Platform MAIT and Development during implementation Phase (B2, C/D, E1), including Payload to S/C AIT • Launcher price • Industrial cost risk margins related to above • Mission Operations costs (to be refined with ESOC) • Science Operations cost • ESA Internal costs • ESA Internal cost risk margins.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 90 of 105

Not included in the estimates are: • Payload Development and MAIT excluded • Phase A and B1 costs, which are assumed to be covered under the General Study Program Budget • Specific technology developments funded through advanced technology development programmes such as TDE or E3P/ExPeRT.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 91 of 105

5 MARS HARD LANDER Due to time limitations the study of the hard lander was performed with a limited amount of experts. It contained experts on systems, mission analysis, EDL and cost. The baseline mission was chosen to consist out of three hard landers, each with an entry mass of 70 kg and aimed to each land 50 kg on the surface with an impact velocity of 20 m/s. During the CDF study the EDL trajectory and EDL equipment of the hard landers was analysed. Additionally, a high-level design of the carrier spacecraft was performed and a high level cost estimation was made. 5.1 Mission Requirements

Mission Requirements ID Statement

MIS-MHL-010 The mission architecture shall comprise 3 hard landers The total mass delivered to the surface shall be 50 kg (TBC) per hard MIS-MHL-020 lander, including all margins The speed of each hard lander upon impact on the Martian surface shall be MIS-MHL-030 less than 20 m/sec The hard landers shall be able to lander at 0 km altitude MOLA elevation MIS-MHL-040 or lower Table 5-1: Mars Hard Lander mission specific requirements 5.2 Reference Scenario 5.2.1 Selection Reference scenario for study: Dedicated Ariane 62 launch to C3=10, ballistic coast towards Mars hyperbolic entry, 3 hard lander probes of each 70 kg at entry.

5.2.1.1 Justification A chemical propulsion transfer was chosen as reference scenario due to commonality with previous Mars carrier vehicles and because an electrical propulsion carrier vehicle would result in a relatively large carrier vehicle that would not be used after delivery of the landers. A dedicated launch with Ariane 62 to C3=10 was chosen following the same reasoning as for the Mars Science Orbiter. The cost of a rideshare to GTO + LPF kick- stage is similar compared to a dedicated launch to C3=10. Since the dedicated launch allows a less complicated spacecraft design and it would provide increased flexibility in launch date and launch mass compared to using the LPF kick-stage, this was chosen as the reference scenario.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 92 of 105

Three hard landers were studied for redundancy reasons and to enable science that benefits from simultaneous measurements in different locations (e.g. weather monitoring, seismicity). Future work could look into whether it is possible to include a fourth lander while staying within acceptable cost. The entry mass of 70 kg draws large similarity with ESA’s Beagle 2 mission (2003) which performed a semi-hard landing. The study focussed on analysing if all the requirements could be met using the same aeroshell, parachutes and EDL GNC as Beagle 2. Using the same – high heritage – equipment as Beagle 2 is in line with the main drivers for the mission: cost and schedule. Beagle 2 had an entry mass of 68.84 kg, of which 33.18 kg is landed mass and 35.66 kg is the EDLS mass. The EDL system of Beagle 2 consisted of: heatshield + back cover (17.81 kg), disk-gap-band (DGB) pilot parachute (0.5 kg), ringsail main parachute (2.76 kg) and airbags including a gas generator (14.59 kg). Beagle 2 landed at ~-3.4 km MOLA elevation with a touchdown velocity of 16 m/sec. The airbags are not needed for the Mars Hard Lander (MHL) which would be designed to sustain the higher g-loads on landing.

5.3 Baseline Design

5.3.1 EDL Trajectory The entry trajectory for MHL was modelled using the same aeroshell and parachutes as Beagle 2. A chemical propulsion Earth escape was studied, with escape occurring on the 27th of November in 2028 and arrival at Mars on the 2nd of October 2029. The entry velocity for this transfer is ~5.6 km/sec. The trajectory is divided into three entry phases: (1) Hypersonic entry: Mach > 1.4, (2) DGB parachute: 1,4 > Mach > 0,4 and (3) Ringsail parachute: 0.4 < Mach. The trajectory was modelled for a flight path angle (FPA) ranging from -11 degree up to -16 degrees. The drag coefficients were adapted to match the terminal velocity of Beagle 2, it needs to be noted that they could be slightly overestimated. For each FPA, the peak g-load, peak heat flux, total heat load and velocity at certain altitudes was calculated. The results are shown in the table below:

Intertial FPA [deg] -11 -12 -13 -14 -15 -16 Relative FPA [deg] -11.4 -12.4 -13.5 -14.5 -15.6 -16.6 peak g-load 32.7 35.1 38.2 41.5 44.6 88.5 peak heat flux [kw/m²] 395 441 485 523 557 588 total heat load [kJ/m²] 29519 25931 23706 22110 20877 19873 v@0km [m/s] 19.7 19.7 19.7 19.7 25.2 483.5 v@-1km [m/s] 18.7 18.7 18.7 18.7 18.7 18.7 v@-2km [m/s] 17.8 17.7 17.7 17.7 17.7 17.7 v@-3km [m/s] 16.8 16.8 16.8 16.8 16.8 16.8 v@-4km [m/s] 16.0 16.0 16.0 16.0 16.0 16.0 Table 5-2: Entry, Descent and Landing trajectory

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 93 of 105

It can be concluded that for FPA between -11 and -14 degrees the impact velocity requirement is met (<= 20 m/sec at 0 km MOLA). Some general remarks are that the landing site dispersion is dependent on the FPA and increases for shallower entries. Using Delta DOR on the carrier spacecraft with 1 baseline improves the dispersion by a factor of 3 compared to no Delta DOR. Delta DOR with 2 baselines improves the dispersion by a factor of 10 compared to no Delta DOR. In Beagle 2 Lessons Learned it is recommended to use Delta DOR, it will therefore be implemented with at least 1 baseline. Future work could consider the use of Delta DOR with two baselines, once more information is known about the required landing ellipse size. Note: after finalisation of the CDF, a newer version of the Mars Climate Database was implemented, this newer version has higher densities in the lower regions of the atmosphere. This results in slightly lower velocities between 0 and -4 km MOLA altitude than stated in the above table. The update has not been implemented in the calculations shown here.

5.3.2 EDL System Using the peak heat flux and total heat load calculated it could be analysed whether the thermal protection system (TPS) used on Beagle 2 would be sufficient for the MHL. The TPS material used on Beagle 2 is Norcoat-Liege, which has a heat flux material limit of 2 MW/m². This is well within the MHL limit (maximum heat flux of 560 kW/m²). The nominal heat load that Beagle 2 was designed for is 19592 J, which is lower than the MHL heat load calculated for each hard lander. This is due to the use of shallower FPA (Beagle 2 entered at 15.8 degrees FPA). Therefore, additional TPS material needs to be added to the aeroshell. The mass of the additional material ranges from 5.17 to 0.67 kg for a FPA of -11 to -15 respectively. The same DGB pilot parachute and ringsail main parachute as Beagle 2 are used, they have a diameter or 8 m and 10.4 m respectively. The DBG parachute is deployed by a mortar, and the ringsail parachute is deployed by the pilot parachute. The aeroshell is released once the main parachute has been deployed. Due to the high landing speed and high g-loads, a crushable attenuation structure is added to the lander. It was chosen to use the same material as ESA’s Schiaparelli lander, which is an aluminium honeycomb sandwich structure. This material has a crush strength of 170 kPa and density of 16 kg/m³. A maximum vertical load of 120 g’s and a maximum lateral load of 80 g’s was allowed since this shows to be the right balance between allowable g-load and crushable structural mass This is more than double of the allowable loads on Schiaparelli. This results in a crushable structure with a thickness of 0.308 m, which over an area of 0.33 m2 adds to a total crushables mass of 1.63 kg. Increasing the allowed g-loads would likely result in a thinner and lighter crushable structure. It needs to be noted that this is an initial mass estimation and therefore allowable g-loads could be tailored in further design phases by slightly increasing or decreasing the crushable mass and material. Regarding GNC equipment during EDL, the same equipment as Beagle 2 is considered for the MHL. Beagle 2 used two accelerometer for parachute deployment, which will also be used for the MHL. Additionally Beagle 2 had an altimeter to trigger airbag deployment. Since the MHL does not have any airbags it could be argued that there is no need for an altimeter, however in the Beagle 2 Lessons Learned [RD-7] it is

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 94 of 105 recommended to also use the altimeter to trigger the parachute release before touch down. The hard landers should be more robust against variable impact velocities compared to Beagle 2 (semi-hard lander), so in order to save cost the altimeter could be discarded and parachute deployment could rely on accelerometer measurements. On the other side, increasing the robustness against varying impact velocities also adds cost as all equipment will need to be designed for a worst case g-load. Therefore, at this point it was decided to include an altimeter for parachute release triggering, but it is recommended to revisit this trade in a more detailed design phase.

5.3.3 Carrier Vehicle A high level study of the carrier vehicle able to carry at least 3 lander probes was performed, the focus is on the transfer scenario and the spacecraft itself is considered to be a scaled down version the Mars Science Orbiter. The baseline was chosen to be a CP transfer due to similarity to all past carrier vehicles for Mars landers. Some requirements on the carrier vehicle are that it shall be able to release each lander probe separately and that it shall contain Delta DOR navigation with 1 baseline. The following adaptions are made starting from the Mars Science Orbiter. The mass of the CP subsystem is reduced to an estimated 4 kg, since there is no more need for a Mars orbit insertion. The power subsystem mass is reduced to 58 kg due to lower power requirements and the science instruments were removed. This results in a 228 kg spacecraft.. When adding a 30% margin on the spacecraft, the three entry probes and the fuel mass, the launch mass to C3=10 becomes 513.3 kg excluding the adapter. It needs to be noted that this is a rough estimation, and that future studies should design the carrier vehicle in more detail. In order to achieve a spread in landing sites between the different hard landers, they need to be deployed from the carrier vehicle sequentially, while leaving several days in between each lander deployment. Several days would be needed to adjust and confirm accuracy of the carrier trajectory. Therefore, the first lander probe has to be released ~10 days or possibly more before entering Mars orbit, meaning that the sizing of the lander its battery will be influenced. Also, additional fuel has to be brought on the carrier in order to enable the trajectory adjustments.

5.3.4 Summary The equipment and mass budget shown for the entry probe satisfies entry at flight path angles of -14 and -13 degrees. As seen in Table 5-5 a system margin of 5% is applied on the EDL system. This is lower than expected due to the high reuse of equipment from flight proven Beagle 2 EDL. The system margin on the lander is 30%, this margin accounts for the large uncertainties in the lander design.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 95 of 105

Mars Hard Lander Entry Probe – System baseline summary Mass Entry Mass 70.0 kg Landed Mass 42.84 kg Dimensions Aeroshell 0.934 x 0.934 x 0.523 m Hard Lander Not studied 1x Aeroshell 1x Disk-Gap-Band parachute Entry, Descent and 1x Ringsail parachute Landing System 1x Crushable structure 1x Altimeter 1x Accelerometer Table 5-3: Mars Hard Lander – System baseline summary

Mars Hard Lander Carrier Vehicle – System baseline summary

Mass Dry Mass (w/ margin) 520.46 kg Wet Mass 527.19 kg Entry Probes 3x Probes AOCS Same as Mars Science Orbiter Communications Same as Mars Science Orbiter Data Handling Same as Mars Science Orbiter Power 1x Solar Array (no detailed design, total mass estimated at 25 kg) 1x Secondary Battery 1x Power Conversion and Distribution Unit (PCDU) 1x Solar Array Driving Mechanism (SADM) 1x Solar Array Drive Electronics (SADE)

Chemical Propulsion No detailed design, total mass estimated at 4 kg

Thermal Same as Mars Science Orbiter Structures Same as Mars Science Orbiter Table 5-4: Mars Hard Lander – System baseline summary

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 96 of 105

Mass (kg) Aeroshell 21.95 Disk-Gap-Band parachute 0.53

Ringsail parachute 2.90

Crushable structure 1.79

EDLS mass 27.16 System Margin 5% 1.36 EDLS mass incl. System Margin 28.52 Landed mass 31.91 System Margin 30% 9.57 Landed mass incl. System Margin 41.48 Dry Mass Entry Probe (EDLS + LANDER) incl. 70.0 Margin Table 5-5: Mass Budget of one Mars Hard Lander Entry Probe

Mass (kg) Attitude, Orbit, Guidance, Navigation Control 7.96 Chemical Propulsion 4.00 Communications 34.30 Power 58.00 Structures 98.61 Thermal Control 9.92 Data Handling 5.25 Harness 5% 10.87 Dry Mass SC 228.17 System Margin 30% 68.45 Dry Mass SC incl. System Margin 296.62 3x Entry Probe incl. System Margin 210 Dry Mass SC + Entry Probes incl. System Margin 506.62 CPROP Propellant Mass 6.60 CPROP Fuel Margin 2% 0.13 Total Wet Mass SC 513.35

Table 5-6: Mass Budget of the Mars Hard Lander Carrier Vehicle

5.4 Alternative Mission Scenarios An alternative to the baseline CP carrier vehicle would be to decrease the current wet mass needed in C3=10 from 513 kg to 500 kg and use the Lisa Pathfinder kick-stage for a rideshare launch to GTO. The launch mass in this case would be 1397 kg excluding the adapter. There is also an alternative when using electrical propulsion for the carrier vehicle: a scaled down version of a Communication Constellation satellite could be used. This implies a rideshare to GTO and the use of 1 T6 engine on the carrier. A rough estimation

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 97 of 105 results in a 434 kg spacecraft + 210 kg for 3 lander probes. The launch mass to GTO would be slightly higher than 900 kg excluding the adapter. Using an EP spacecraft solely as carrier implies that a relatively large mass is sent to Mars which will then not be used once at Mars. Therefore it was also considered make the EP spacecraft enter into Mars orbit for scientific or communication purposes after releasing the lander probes. However, trajectory wise this would be challenging, since the EP spacecraft needs to be in a trajectory with a near surface perigee while releasing the landers, it needs to be able to quickly lift its perigee and perform an MOI after the landers are released. Detailed trajectory analysis would have to be performed to see if this would be possible with EP or if a hybrid system would be needed to perform the MOI. This would again complicate the design, therefore making the fully CP reference scenario more attractive.

5.5 Cost References cases of previous landers and probes were used, e.g., Huygens and Beagle 2 extrapolating the mass cost relationship. The science orbiter can serve as a reference case for the carrier spacecraft cost estimation. 5.6 Further Work / Recommendations It can be concluded that 50 kg landed mass per hard lander cannot be achieved when using a hard lander design based on Beagle 2 (Requirement MIS-MHL-020). The landed mass is equal to 41.5 kg including system margin, which is still higher than the 33.3 kg landed mass from Beagle 2. Based on this information several recommendations for future work are stated below: Carrier spacecraft: More attention needs to be put into the sizing of the carrier spacecraft and especially in the mechanism for the sequential deployment of three lander probes. A future study needs to look at the spread of the landing sites required and the effect thereof on the system. EDL: It has been shown that the Beagle 2 EDLS could be used if additional TPS material is added to the aeroshell. This is because a shallower entry flight path angle compared to Beagle 2 is needed in order to comply with the impact velocity requirement. It is recommended for future studies to take into account all Beagle 2 EDLS lessons learned RD[3], where relevant. Lander: This study has not looked at the lander itself; therefore more research is needed towards a lander design. Particular attention needs to be focused on equipment and payloads that can survive 120-g impact load.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 98 of 105

This Page Intentionally Blank

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 99 of 105

6 CONCLUSIONS 6.1 Achievement of Study Objectives 1. Provide a portfolio of potential transfers to Mars for launches between 2026 and 2032. – Achieved. Several transfer options have been identified. It presented a particular challenge as many different combinations of launcher, launch configuration and propulsion system were taken into account. A number of examples were taken forward, primarily to investigate ride-share and hence require a flexible propulsion system. 2. Assess the feasibility and technical scope of 1-2 small satellite orbiter or a mother/daughter(s) spacecraft combination mission to Mars, within the cost and schedule constraints. – Achieved. The study provided a number of spacecraft concepts for: Mission 1, a three satellite communications constellation. Mission 2 a science orbiter. A technically feasible baseline for the Mars communications constellation exceeds cost and preferred launch rideshare capabilities, generally as a consequence of EP system requirements (a more expensive engine is required to meet mass budget limitation on the launcher). For the science orbiter with CP, a dedicated launch is the preferred approach, on balance between cost and complexity. An option is to use the LPF propulsion module, which could allow a shared launch. However, the cost of LPF (and possible delta-qual) results in a similar mission price as a dedicated launch. 3. Assess the feasibility and technical scope of delivering small lander(s) to the surface of Mars within the cost and schedule constraints. – Partially Achieved. Complexity and the number of options for mission 1 and 2 meant that the lander assessment was only done at system level due to study resources and time limitations. A sizing exercise was completed, giving an indication of mass and cost. 4. Design the missions to a maximum cost at completion of €250M. – Partially Achieved. Mars Communications Constellation exceeded cost constraint with T6 dual launch GTO baseline. Dedicated launch for science orbiter is slightly over. However, the possibility of exceeding this cost cap must be taken into account. 5. Investigate existing technologies and equipment for use in Mars missions within the development timeframe. – Achieved. Technologies are baselined with high TRL where possible. Options to optimize are identified in critical areas. This extra effort for technology development could allow the reduction in mass, making ride- share a more feasible option.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 100 of 105

6. Define the “small satellite approach” to be adopted, considering reliability, FDIR, ECSS tailoring, COTS approach, operations/ground segment, margins philosophy etc. – Partially Achieved. Ops, margins and small sat approach options were identified. Further work will be needed to show FDIR, ECSS tailoring. During the systematic assessment, some off the shelf equipment was adopted. 7. Define schedule, risks and programmatic aspects of the mission compatible with a CM2022 decision and launch between 2026 and 2032. – Achieved. A schedule has been defined, commensurate with these launch dates, but not without risk for the Mars communications constellation. Consideration of shared dev. models was addressed. Graceful degradation of mission performance is an option to make savings at spacecraft level in terms of redundancy. 8. Evaluate the landscape for industrialisation of such small missions, considering alternatives to the European LSIs as Prime Contractor. – Not Achieved. The study has focused on the technical design. 9. Highlight the main operational constraints. • Achieved. Associated with ensuring data return for science and surface assets. Without extensive GS time, downlink latency is high. In general operations costs could be considered independent of spacecraft size, and more driven by usage, hence downlink data latency. Re-use of MSR ERO GS is possible but requires further asssesment. Further recommendations on transferring the project team to a very similar GS setup could be considered further in a next phase. 6.2 Further Study Areas During the study there are aspects that were identified either as out of scope or relevant for further analysis in the next phases. These are: • Elaborate the EP transfer time, based on the baseline T/m. • Elaborate lander design and concept, in particular the carrier spacecraft. • Identify options to increase payload mass. • Perform iteration of solar array sizing calculations for the the MCC T6 engine option, taking the minimum power for the T6 as 2200W at Mars. • Science planning needs assessment to refine solutions to improve data latency. This would include improvements to comms sub-system and assumptions for science operations. Comms relay performance requires improvement to be competitive with other services. • Definition of electric propulsion gimbal pointing mechanism, current assumptions is extrapolated and is low TRL. This option and alternatives require development. • Possibility of using SEP for the science orbiter could be investigated .

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 101 of 105

• Move towards a non passive thermal control system could be investigated to optimize power demand. This is less relevant for SEP systems, as propulsion is a significant power driver. • A62 was considered as the baseline launcher, but in light of the design results, PSLV could be an option to consider or establish as a back-up launcher. • A more detailed assessment of operations costs is required, to ensure that the assumptions taken are representative of small missions. 6.3 Main Conclusions The study assessed low cost options for a small Mars mission which could be launched in commercial rideshare configurations into Earth orbit. The original intention was to identify a low cost piggy back missions, however the technical analysis has shown that, for feasible spacecraft designs, dual or dedicated launches can be expected. The Mars Communications Constellation (MCC) was baselined as an SEP spacecraft. A dual launch or dedicated launch is considered technically feasible each with differing SEP technologies. A dual launch requires a more expensive SEP thruster whereas a cheaper alternative technology pushes the mission to a more expensive dedicated launch. However, neither approach fits the cost constraints with relatively small differences. In particular: • The ride-share was envisioned to keep the launch cost low. However, the study has shown that the spacecraft becomes more complex to ensure a feasible design, off-setting the majority of these launch cost savings. • The large delta-V needed for the Mars transfer and limitations in European technology for solar electric propulsion and solar arrays contribute to a required spacecraft mass in excess of what is possible for cheaper, lower mass launch options. • The study also analysed a dedicated launch option, for a minimal extra cost compared to the dual launch to GTO scenario. The extra cost could be justified for a simplified spacecraft and provide launch opportunity flexibility, thereby reducing schedule risks. The Mars Science Orbiter (MSO) makes use of a dedicated launch, as on balance, the flexibility of this option is favoured over the risk of using the LPF propulsion stage with potential re-qualification needed to take advantage of a shared launch to GTO option.. In particular: • To launch to GTO, and ensure a technically feasible spacecraft design, a chemical propulsion kick stage is needed to get onto a Mars trajectory. However, the total mission costs including a kick stage, even when using heritage solutions, is as expensive as a dedicated launch to Mars directly. • As with the MCC, a dedicated launch offers more flexibility on launch date and allows a simplified lower cost spacecraft. • Consideration should also be given to a SEP option for the MSO. This could be a feasible spacecraft and provide mission flexibility, but would likely lead to a more

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 102 of 105

expensive spacecraft design. However, a more detailed analysis would be required beyond this study for confirmation. Overall the study has identified a wide range of potential small Mars mission architectures including orbital and lander missions. However, from the resulting analysis, there appears to be only a small range of technically feasible options possible within the programmatic constraints. The use of rideshare constrains the capabilities of the spacecraft significantly. The 250 MEuro cost requirement is certainly exceeded for a payload between 10 – 30 kg for the MCC and is borderline for the MSO. More mass optimisation may improve this by developing key technologies in Europe such as advanced power systems including low mass solar arrays and lower-cost, high lifetime SEP thrusters. If there is further flexibility on mission price, a small mission to Mars over the next 10 years could be possible within the range of 250-350 MEuro.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 103 of 105

7 REFERENCES RD[1] Margin Philosophy for Mars Exploration Studies RD[2] Ariane 6 user’s manual, Issue 1 Revision 0, March 2018; https://www.arianespace.com/wp-content/uploads/2018/04/Mua- 6_Issue-1_Revision-0_March-2018.pdf RD[3] Beagle-2: Lessons Learned and Management and Programmatics, First Edition Publication, 2004 RD[4] M-ARGO study final report, CDF-171(A), March 2017 RD[5] TEC-SYC Cost Risk Procedure, TEC-SYC/5/2010/PRO/HJ, February 2010 RD[6] ESA Cost Engineering Charter of Services, Issue 4, TEC-SYC/12/2009/GRE/HJ RD[7] ECSS-E-AS-11C Adoption Notice of ISO 16290, Space Systems – Definition of the Technology Readiness Levels (TRLs) and their Criteria of Assessment, dated 1 October 2014. To be supersede by EN16603-11.

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 104 of 105

This Page Intentionally Blank

ESA UNCLASSIFIED – For Official Use SMARTieS CDF Study Report: CDF-205(A) April 2020 page 105 of 105

8 ACRONYMS

Acronym Definition COTS Commercial Off The Shelf CP Chemical Propulsion ECSS European Cooperation for Space Standardization EDL Entry, Descent and Landing EDLS Entry, Descent and Landing Stage EP Electric Propulsion FDIR Fault Detection, Isolation and Recovery GTO Geostationary Transfer Orbit MCC Mars Communication Constellation MPPT Maximum Power Point Tracker MSO Mars Science Orbiter MSR Mars Sample Return SA Solar Array SEP Solar Electric Propulsion TASO Trans-Areostationary Orbit

ESA UNCLASSIFIED – For Official Use