ISSION EQUIREMENTS OCUMENT

a

ESTEC European Space Research and Technology Centre - Keplerlaan 1 - 2201 AZ Noordwijk - The Netherlands Tel. (31) 71 5656565 - Fax (31) 71 5656040 www.esa.int

Marco Polo MRD

E S A D O C U M E N T

document title/ titre du document

ISSION EQUIREMENTS OCUMENT

prepared by/préparé par D. Agnolon reference/réference SCI-PA/2008.001/Marco-Polo issue/édition 4 revision/révision 0 date of issue/date d’édition 02/02/2009 status/état Final Document type/type de document RQ Distribution/distribution SRE-PA, D. Koschny, J. Romstedt, Marco Polo European Science Study

Team

Marco Polo MRD Issue 4 Revision 0 - 02/02/2009

SCI-PA/2008.001/Marco-Polo page ii of xi

A P P R O V A L

Title Titre

Marco Polo MRD issue issue

4 revision revision

0

author auteur

D. Agnolon date date

02/02/2009

approved by approuvé par

D. Agnolon date date

02/02/2009

C H A N G E L O G

reason for change /raison du changement issue/issue revision/revision date/date

Added requirements at the start of the assessment phase

3 0 12/09/2008

Update of the requirements following Phase 1 of the study and baseline selection at the MAR

4 0 02/02/2009

C H A N G E R E C O R D

Issue: 3 Revision: 0

reason for change/raison du changement paragraph(s)/paragraph(s) requirement(s)/specification(s)

Additional explanation about requirements 1.5.1

Modification of the scope (name) of [AD5] 2.1

Addition of acronym definitions (OTS, DOR, FEM, FMECA, MCC, STR, APE, RPE, AME)

3.1

Update of a few terminologies (S/C composite, sampling approaches, absolute/relative pointing errors)

3.2


SCI-PA/2008.001/Marco-Polo page iii of xi


Additional document format requirements (ESATAN, NASTRAN)

4.3 R-SD.FOR-2/3

Update of the mission block diagram 5.2.2

Deletion of the Parent-ID column All All

Added requirements on overall payload attitude, mass, data and power

7.2 R-PL.ACS/MAS/ DAT/POW

Added requirement on the use of OTS equipment 8.1

Incorrect requirement title 9.2 R-TA-4

Added requirement on collision avoidance with NEO 9.3 R-OP.POX-7

Added requirement on mitigation of electric propulsion outage

9.3 R-OP.MIS-5

Added requirement on distance to Earth/Sun 9.3 R-OP.MIS-6

Added requirement on S/C autonomy through solar conjunction

9.3 R-OP.MIS-7

Added requirement on night re-entry 9.3 G-OP.MAS-1

Added generic requirement on S/C configuration 9.4.1 R-SY.GEN-8

Added generic requirement on minimization of launch masses

9.4.1 R-SY.GEN-9

Added requirement on aerothermodynamics fluxes margins

9.4.1 R-SY.MAR-2

Tailoring of power margin from [AD8] 9.4.1 R-SY.MAR-3

Added requirement on protection from micro-meteorites

9.4.1 R-SY.DEB-2

Added requirement on MCC of propulsion module 9.4.1 R-SY.PRM-1

Added requirement on instrument commissioning 9.4.1 R-SY.ORB-5

Added requirement on instrument calibration 9.4.1 G-SY.ORB-1

Rewording of landing accuracy requirement 9.4.1 R-SY.LAN-7

Added requirement on descent/landing telemetry data 9.4.1 R-SY.LAN-10

Added requirement on surface P/L accommodation 9.4.1 R-SY.LAN-11

Added requirement on ERC max. deceleration 9.4.1 R-SY.ERC-8


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Added requirement on the type of ERC re-entry 9.4.1 R-SY.ERC-9

Added requirement on ERC flight instrumentation 9.4.1 G-SY.ERC-2

Added general requirement on minimization of S/C autonomy

9.4.1 R-SY.AUT-1

Added requirement on FMECA analysis 9.4.1 R-SY.AUT-2

Added requirement on safe mode 9.4.2 R-SS.ACS-6

Added requirement on mitigation of STR blinding 9.4.2 R-SS.ACS-7

Added requirement on IMU initialisation and redundancy before/during descent/landing

9.4.2 R-SS.ACS-8/9

Added pointing requirement 9.4.2 R-SS.ACS-10

Added general DHS requirement on TM/TC management

9.4.2 R-SS.DHS-2

Added general DHS requirement on instrument/equipment control

9.4.2 R-SS.DHS-3

Added general DHS requirement on reconfiguration capability

9.4.2 R-SS.DHS-4

Added general DHS requirement on redundancy management

9.4.2 R-SS.DHS-5

Added DHS requirement on storage of surface experiment data

9.4.2 R-SS.DHS-6

Added requirement on compatibility of propulsion system with any attitude/environment

9.4.2 R-SS.PRO.4

Added requirement on mini. Propellant temperature 9.4.2 R-SS.PRO.5

Added general requirement on the performances of propulsion system

9.4.2 R-SS.PRO.6

Update of requirement taking into account new R-SS.COM-8

9.4.2 R-SS.COM.1

Update of requirement to include D-DOR 9.4.2 R-SS.COM.5

Update of weather availability requirement 9.4.2 R-SS.COM.6

Added requirement on maxi. Bit error rate 9.4.2 R-SS.COM.8

Added general requirement on redundancy 9.4.2 R-SS.COM.9

Added general requirement on management of communication functions

9.4.2 R-SS.COM.10

Modification of mechanical load requirement to cover all mission phases

9.4.2 R-SS.STR.2


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Added requirement on structural stiffness 9.4.2 R-SS.STR.4

Added general requirement on structural design 9.4.2 R-SS.STR.5

Added general requirement on minimization of structural mass

9.4.2 R-SS.STR.6

Added requirement on protection from debris at landing

9.4.2 R-SS.MEC.7

Added requirement on monitoring of SATS mechanisms

9.4.2 R-SS.MEC.8

Added requirement on mitigation of sampling unit being blocked

9.4.2 R-SS.MEC.9

Added requirement on determination of soil properties based on mechanism data

9.4.2 G-SS.MEC.3

Update of requirement including more detailed term definition

9.5 R-EN.LAN-1

Added chapter on ground segment, added requirement on compatibility with ESA/NASA deep space network

9.7 R-GS.COM-1

Updated requirement on reliability 9.8 R-PA.REL-1

Added requirement on failure and SPF mitigation 9.8 R-PA.REL-2

Added requirement on overall mission reliability 9.8 R-PA.REL-3

Updated explanation for backup SATS mechanism 9.8 R-PA.RED-1

Update of TBC/TBD list 10

Issue: 4 Revision: 0


Update of AD/RD list 1.3 & 2.1

Additional acronym definitions (CAM, DHS, MAR, RSSD, DoI)

3.1

Update of a few terminologies (sampling strategies, horizontal/vertical relative velocities, sampling site, landing accuracy/ellipse)

3.2

Clarification of mass margin to be used from [AD8] 4.4 R-SD.MAR-2


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Update of mission overview and phases following baseline selection at the MAR

Entire chapter 5 6.1.1, 8.1, 9.1

R-MR-1, 2 R-SO.BAS-1 R-PM.SCH-1, 2 R-LS.ORB-1 R-LS.DEP-1

Update of mission block diagram and S/C naming conventions

5.2.2 in particular but all chapters

Added science operations timeline 5.2.3

Updated planetary protection requirements (Mars fly-by, added information on contamination tracking)

6.2 R-PP.FWD-1, 2

Update of P/L suite following MAR 7.1

Update of P/L resources to include inputs from DoI instrument studies, included in PDD update

7.2 R-PL.GEN-3 R-PL.ACS-1, 2 R-PL.MAS-1, 2 R-PL.POW-1, 2, 3, 4 R-PL.DAT-1, 2

Added requirement on solar elevation angle during global characterization phase

7.2 R-PL.GCR-1

Added requirement on launch margin (uncertain performances of Soyuz on direct escape trajectories)

9.1 R-LS.LAV-1

Clarification on NEO ground observations & alternative targets

9.2 R-TA-4, 5

Clarification on the proximity & science operations, requirements and timelines

9.3 R-OP.POX-1, 2, 3, 4, 5, 8, 9, 10

Added definition of the duration of the global characterization campaign

9.3 R-OP.POX-11

Deleted requirement (unnecessary) and update of related requirement

9.3 R-OP.LAN-1 (Deleted) R-OP.LAN-5 (Updated)

Clarification of sampling rehearsal definition 9.3 R-OP.LAN-3

Definition of surface operations conditions 9.3 R-OP.LAN-5, 6

Added goal to transmit descent pictures to Earth real-time

9.3 G-OP.LAN-1

Clarification 9.3 R-OP.MIS-1

Deleted requirement (no longer applicable for baseline scenario)

9.3 R-OP.MIS-5

Update of transfer geometry (Earth-Sun-NEO ranges, angles, etc.)

9.3 R-OP.MIS-6


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Added requirement on solar conjunction 9.3 R-OP.MIS-8

Update of re-entry conditions 9.3 R-OP.MAS-1

Added CAM mode 9.3 R-OP.SAT-3

Updated following MAR 9.4.1 R-SY.GEN-1, 6

Change of design philosophy (Design-to-cost) 9.4.1 R-SY.GEN-9

Added requirements (moved from R-SY.ERV-X) due to new baseline scenario and deleted related requirements (no longer applicable due to new architecture)

9.4.1 R-SY.GEN-10, 11 (Added) R-SY.ERV-1, 2, 3 (Deleted)

Added specific delta-V margins (20%) for NEO proximity operations

9.4.1 R-SY.MAR-4

Definition of survival probability to micro-meteorites 9.4.1 R-SY.DEB-2

Deleted requirements due to new architecture 9.4.1 R-SY.PRM-1, 2 (Deleted)

Update of gravity field determination accuracy (previous requirement too arbitrary)

9.4.1 R-SY.ORB-1

Deleted requirements (no longer applicable due to new architecture)

9.4.1 R-SY.ORB-3, 4

Update of landing accuracy requirement and clarification of terminology

9.4.1 R-SY.LAN-7

Added landing/touchdown stability requirement 9.4.1 R-SY.LAN-12

Added hover & go clearance requirement 9.4.1 R-SY.LAN-13

Deleted requirement (moved as a goal) 9.4.1 R-SY.LAN-11 (Deleted) G-SY.LAN-1 (Added)

Updated requirement (previously oversizing) 9.4.1 R-SY.ERC-1

Updated requirement (maximum shock load also defined for RF beacons, TBD)

9.4.1 R-SY.ERC-2

Deleted requirement, added and moved to R-GS.REC-1 (requirement was applicable to ground recovery facilities and landing site location rather than on ERC system)

9.4.1, 9.7 R-SY.ERC-3 (Deleted) R-GS.REC-1 (Added)

Deleted requirement, added and moved to R-SY.SEA-1, 2, 3 (requirement was applicable to the sealing system rather than ERC closure system (which might still take care of this function, but this is not required))

9.4.1 R-SY.ERC-4 (Deleted) R-SY.SEA-1, 2, 3 (Added)


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Definition of ERC maximum deceleration through re-entry

9.4.1 R-SY.ERC-8

Definition of measurement frequency (flight instrumentation)

9.4.1 G-SY.ERC-2

Updated sterilization procedures 9.4.1 R-SY.STZ-1

Reformulation of requirement 9.4.2 R-SS.POW-3

Clarification of GNC robustness requirement 9.4.2 R-SS.ACS-3

Extension of possible navigation sensors to mapping LIDAR

9.4.2 R-SS.ACS-2

Updated requirement of far approach navigation distance and time

9.4.2 R-SS.ACS-4

Update of pointing requirements (previously too constraining)

9.4.2 R-SS.ACS-10

Updated requirement to account for all mission phases

9.4.2 R-SS.PRO-2

Definition of bit error rate 9.4.2 R-SS.COM-8

Deleted requirement in line with updated design approach

9.4.2 R-SY.STR-6

Clarification of contamination tracking approach 9.4.2 R-SS.MEC-5, 6

Deleted and moved requirement + Clarification of backup sampling method requirement

9.4.2, 9.8 R-PA.RED-1 (Deleted) R-SS.MEC-10 (Added)

Clarification of sample delta-temperature requirement

9.4.2 G-SS.MEC-1

Definition of worst (smallest) case sampling site driving landing accuracy

9.5 R-EN.LAN-1

Update of baseline NEO properties 9.5 R-EN.LAN-2 R-EN.NEO-1, 5

Clarification of the binary case (no longer taken into account for sizing, but analyses required to understand impact on mission performances)

9.5 R-EN.NEO-3

Added JAXA DSN compatibility requirement 9.7 R-GS.COM-1

Update of TBC/TBD list 10


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T A B L E O F C O N T E N T S

1 INTRODUCTION .................................................................................................................1 1.1 Scope of Document ..............................................................................................................................1 1.2 Background ..........................................................................................................................................1 1.3 Documentation architecture .................................................................................................................1 1.4 Issue schedule ......................................................................................................................................2 1.5 Requirements nomenclature.................................................................................................................2

1.5.1 Requirements and goals .............................................................................................................2 1.5.2 Requirement ID and sub-id ........................................................................................................2

2 APPLICABLE AND REFERENCE DOCUMENTS..............................................................3 2.1 Applicable documents..........................................................................................................................3 2.2 Reference document.............................................................................................................................4

3 ACRONYMS AND TERMINOLOGY ...................................................................................6 3.1 Acronyms .............................................................................................................................................6 3.2 Terminology.........................................................................................................................................7

4 STUDY REQUIREMENTS.................................................................................................10 4.1 Study approach and objectives...........................................................................................................10 4.2 Units ...................................................................................................................................................10 4.3 Document format ...............................................................................................................................10 4.4 Margin requirements ..........................................................................................................................10 4.5 Coordinate systems ............................................................................................................................11

4.5.1 Spacecraft Reference Frames...................................................................................................11 4.5.2 Orbital Reference Frames ........................................................................................................11

5 MARCO-POLO OVERVIEW .............................................................................................13 5.1 Top-level Mission Requirements .......................................................................................................13 5.2 Mission Profile ...................................................................................................................................13

5.2.1 Mission overview.....................................................................................................................13 5.2.2 System definition .....................................................................................................................14 5.2.3 Mission phases .........................................................................................................................14

6 SUMMARY OF SCIENCE OBJECTIVES AND PLANETARY PROTECTION REQUIREMENTS..............................................................................................................18

6.1 Science objectives and operations......................................................................................................18 6.1.1 Science objectives ....................................................................................................................18 6.1.2 Science operations....................................................................................................................18

6.2 Planetary protection requirements .....................................................................................................19


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7 PAYLOAD DESCRIPTION AND REQUIREMENTS .........................................................20 7.1 Payload instruments ...........................................................................................................................20 7.2 Payload requirements .........................................................................................................................20

8 PROGRAMMATICS REQUIREMENTS ............................................................................23 8.1 Programmatic requirements ...............................................................................................................23 8.2 Technology requirements...................................................................................................................23

9 SYSTEMS REQUIREMENTS............................................................................................25 9.1 Launch segment requirements ...........................................................................................................25 9.2 NEO target requirements ...................................................................................................................25 9.3 Mission-level operational requirements.............................................................................................26 9.4 Space segment requirements ..............................................................................................................29

9.4.1 Spacecraft system requirements...............................................................................................29 9.4.2 Platform sub-system requirements...........................................................................................35

9.5 Environmental requirements ..............................................................................................................39 9.6 AIV and testing requirements ............................................................................................................40 9.7 Ground segment .................................................................................................................................41 9.8 Product assurance and RAMS requirements......................................................................................41

10 LIST OF TBC AND TBD ...................................................................................................42 10.1 List of TBC ......................................................................................................................................42 10.2 List of TBD ......................................................................................................................................42


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L I S T O F F I G U R E S A N D T A B L E S

Table 5-1: Mission requirements .....................................................................................................................13 Table 5-2: Overview of mission phases ...........................................................................................................16 Table 6-1: Science objectives ..........................................................................................................................18 Table 6-2: Resolution requirements for global characterisation, local characterisation, and context

measurements...........................................................................................................................................19 Table 6-3: Planetary protection requirements ..................................................................................................19 Table 7-1: List of instruments ..........................................................................................................................20 Table 7-2: General Payload requirements........................................................................................................22 Table 8-1: Programmatic Requirements. .........................................................................................................23 Table 8-2: Definition of ESA Technology Readiness Levels..........................................................................24 Table 9-1: Launch Segment Requirements ......................................................................................................25 Table 9-2: NEO target requirements ................................................................................................................26 Table 9-3: Operations Requirements ...............................................................................................................29 Table 9-4. Spacecraft systems Requirements...................................................................................................34 Table 9-5. Platform sub-system Requirements. ...............................................................................................39 Table 9-6: Environmental requirements ..........................................................................................................40 Table 9-7: AIV Requirements..........................................................................................................................40 Table 9-8: Product Assurance Requirements...................................................................................................41 Table 10-1: List of TBC...................................................................................................................................42 Table 10-2: List of TBD...................................................................................................................................42 Figure 3-1: NEO sample return building blocks................................................................................................8 Figure 4-1: Main spacecraft reference frame ...................................................................................................11 Figure 4-2: Local orbital reference frame ........................................................................................................11 Figure 5-1: Block diagram of the Marco Polo mission....................................................................................14 Figure 5-2: Timeline of proximity operations..................................................................................................17 Figure 8-1: TRL scale ......................................................................................................................................24


SCI-PA/2008.001/Marco-Polo page 1 of 53

1 INTRODUCTION

1.1 Scope of Document This is a Mission Requirements Document (MRD) to be used as an Applicable Document in the Marco Polo industrial assessment study. The purpose of the MRD is to provide level 1 (mission) requirements for the Marco Polo industrial system design studies to start in the second quarter of 2008. It includes all functional and performance requirements down to the sub-system level but it also recalls the main science and payload requirements as well as the planetary protection requirements. Recording and tracking of changes as well as giving a brief rationale is very important. The traceability of the requirements is paramount in order to make this document and its associated requirements easy to read and to understand at any stage of the mission assessment and possibly later definition phase, should this mission be selected.

1.2 Background On 18th October 2007 after a meeting of the Space Science Advisory Committee, the candidate missions for the Cosmic-Vision 2015-2025 programme were selected for further assessment and consideration for launch in 2017/2018. The Marco Polo NEO sample return mission was selected as one of the M-class candidate missions [RD6]. During the spring 2008 the mission concept has first been studied in the Concurrent Design Facility at ESTEC (internal pre-assessment) in order to prepare a thorough industrial assessment phase which will be implemented in the summer 2008 and will continue through 2009. After the assessment phase two M-class missions will be selected and will go into definition phase [RD6]. Eventually one M-class mission will be selected in ~ 2011 to go into implementation phase for launch in 2017.

1.3 Documentation architecture The mission requirements document is one of the documents that constitute the foundations of the mission profile for Marco Polo. This MRD is supported by:

• Marco Polo Science Requirements Document [AD1] • Marco Polo Planetary Protection Document [AD3] • Marco Polo Mission Environment Document [AD4] • Marco Polo Consolidated Reference Mission Analysis Document, provided shortly after the

MAR [AD5] • Marco Polo Payload Definition Document [AD6] • Marco Polo Mission Operations Assumptions Document, provided shortly after the MAR

[AD16] • Marco Polo Radio Science Experiment Guidelines, provided shortly after the MAR [AD17] • Marco Polo Science Operations Assumptions Document, provided shortly after the MAR

[AD18]



The following documents contain general requirements for Assessment studies:

• CDF Model Input Specifications [AD7] • Margin philosophy for Science Assessment studies [AD8] • Launcher user’s guide [AD9]

1.4 Issue schedule This issue of the Mission Requirements Document has been established in support of the Industrial Studies of the system design of the Marco Polo mission as proposed by the scientific community in the Cosmic Vision 2015-2025 process. The document is (currently) an open document and regular updates are expected. Particularly, iterative steps with industrial or internal study partners and the ESA study manager are foreseen. Revisions will be published, as required, at the start of, as well as during the system design. The book captain of this document is the Marco Polo study manager and any comment on this document should be directly reported to him. Any conflict between requirements in this document shall be reported to ESA and any modification requires his/her prior approval.

1.5 Requirements nomenclature Items included in this MRD are classified according to the following categories.

1.5.1 REQUIREMENTS AND GOALS R- “Shall” Requirements noted R- are mandatory, shall be verified with an accepted

verification method and shall be complied with. If not complied with, the Agency shall immediately be notified.

G- “Should” Requirements (or Goals) noted G- are desirable requirements with the objective to increase the scientific return or performance of the mission, while minimising the technical complexity and the programmatic aspects of the mission. Goals may be fulfilled under limited favourable conditions. They are highlighted in italic in order to be easily identified.

A requirement may need an explanation. Such an explanation can be found in the “Rationale and/or Comments” column of the requirement tables of this document. This is intended only to assist the understanding of the requirement or its origin (the Parent requirement ID is included wherever applicable)

1.5.2 REQUIREMENT ID AND SUB-ID The requirement ID belongs to the following category:

• AV: AIV Requirement • EN: Environmental Requirement • LS: Launch Segment • MR: Mission Requirement



• OP: Operational Requirement • PA: Product Assurance Requirement • PL: Payload Requirement • PM: Programmatic Requirement • PP: Planetary Protection Requirement • SD: Study Requirement • SO: Science Objective • SS: Sub-system Requirement • SY: Space Segment System Requirement

Sub-ID: The sub-ID is optional and can be created by the author as needed. It should be composed of 3 letters and representative of the item considered. As defined on a case-by-case basis this sub-ID is not defined in the list of acronyms but in the requirement table itself. Sub-ID examples:

• PM.SCH: Programmatic Schedule requirement, • LS.LAV: Launch Vehicle Requirement, • PL.DAT: Payload data rate, • SS.POW: Platform power, • Etc.

2 APPLICABLE AND REFERENCE DOCUMENTS The Applicable Documents (AD’s) listed below shall be complied with, unless conflicting with the mission MRD itself or where specifically stated. Their contents shall be considered as parts of the requirements of this document. A conflict between those documents and the MRD shall be brought to the attention of ESA for clarification. The published ECSS (European Cooperation for Space Standardisation) space standards documents quoted in the MRD are applicable throughout the Marco Polo assessment phase and can be made available by ESA-ESTEC (ECSS Secretariat) or on the Internet at: www.ecss.nl. The published ESA Procedures, Standards and Specifications (PSS) documents can be made available by ESA-ESTEC (ECSS Secretariat).

2.1 Applicable documents

[AD1] Marco Polo Science Requirements Document, [MP-RSSD-RS-001], issue 1e, Document providing Science objectives and science requirements, including measurement specifications

[AD2] Not Applicable (present document)

[AD3] Marco Polo – Assessment of Planetary Protection issues, [SCI-PA/2008.013/Marco-Polo], issue 1.0., Document providing recommendations on Planetary Protection issues relating to the potential NEO targets



[AD4] Marco Polo Environmental Document, [SCI-PA/2008.014/Marco-Polo], issue 1.0

[AD5] Marco Polo Consolidated Reference Mission Analysis, [Ref TBD], issue 1.0, ESOC, document describing the reference transfers and orbit to be used in Phase 2.

[AD6] Marco Polo Payload Definition Document, [SCI-PA/2008.002/Marco-Polo], issue 3.0, Document providing description and resources of the reference science payload suite

[AD7] CDF Model Input Specification, [CDF-IFS-001], issue 3.1, and associated Excel workbooks 'Mission Input Issue 3 rev 1.xls' and 'data exchange.xls'

[AD8] Margin Philosophy for SCI-PA Assessment Studies, [SCI-PA/2007.022], issue 1.0

[AD9] Soyuz-Fregat 2-1b from the Guiana Space Centre User’s Manual, issue 1.0, June 06. Note: The launch vehicle performances to be used in this study are defined in [RD3]

[AD10] ECSS-E-10 series, available from http://www.ecss.nl





[AD15] ECSS-E-60A, available from http://www.ecss.nl

[AD16] Marco Polo Mission Operations Assumptions Document, [Ref TBD], issue 1.0, ESOC, document describing the reference mission operations, communication strategy and ground segment

[AD17] Marco Polo Radio Science Experiment Guidelines, [Ref TBD], issue 1.0, ESOC, document describing the reference Radio Science Experiment architecture and performances

[AD18] Marco Polo Science Operations Assumptions Document, [Ref TBD], issue 1.0, RSSD, document describing the reference science and instrument operations

The above documents are subject to updates; [AD1] through [AD6] are expected to be matured by the Agency during the contract throughout Phase 1 (the study phases are defined in the statement of work of this contract). ESA may provide a final version of each document to the Contractor at the start of Phase 2.

2.2 Reference document The Reference Documents (RD’s) listed below are given as complementary information and background data related to the Marco Polo Mission or the ESA standards.



[RD1] Marco Polo CDF IFP material, Final presentation material of the internal ESA assessment study (Summary of the CDF study report [RD2])

[RD2] Marco Polo CDF study report, CDF-72(A), Final report of the internal ESA assessment study. Note: will be provided by ESA at the negotiation meeting

[RD3] Mission analysis guidelines for Marco Polo, [SCI-PA/2008.027/Marco-Polo], ESOC, document providing reference transfer trajectories for Phase 1 of this study

[RD4] Study overview of the Near-Earth Asteroid Sample Return, [SCI-PA/2007.004/DA], ESA Technology Reference Study Summary

[RD5] Deimos Sample Return Technology Reference Study, [SCI-A/2006.010/DSR], ESA Technology Reference Study Summary

[RD6] Comic Vision Presentation to Industry

[RD7] Marco-Polo: NEO sample return mission proposal, M.A. Barucci and M. Yoshikawa on behalf of the European and Japanese NEO science community, June 2007

[RD8] http://www.esa.int/techresources/ESTEC-Article-fullArticle_par-50_1151301926473.html, TRL definition

[RD9] Phase II of the Small Main-Belt Asteroid Spectroscopic Survey, S.J. Bus and R.P. Binzel, Icarus 158, 106–145, 2002



3 ACRONYMS AND TERMINOLOGY

3.1 Acronyms AD Applicable Document AIV (AV) Assembly, Integration and Verification AME Absolute Measurement Error AOCS Attitude and Orbit Control System APE Absolute Pointing Error CAM Collision Avoidance Mode CDF Concurrent Design Facility CoG Centre of Gravity CoM Centre of Mass CoP Centre of Pressure CSG Centre Spatial Guyanais (Kourou, French Guyana) DHS Data Handling System DM Descent Module DoI Declaration of Interest DOR Differential One-way Ranging DSM Deep Space Manoeuvre ECSS European Cooperation for Space Standardisation EN Environment requirement ERC Earth Re-entry Capsule ERV Earth Return Vehicle ESA European Space Agency ESTEC European Space research and TEchnology Center FEM Finite Element Model FMECA Failure Modes, Effects and Criticality Analysis GA Gravity Assist GNC Guidance, Navigation and Control GTO Geostationary Transfer Orbit HEO Highly Elliptical Orbit HGA High Gain Antenna IR Infrared ITT Invitation To Tender JAXA Japan Space Exploration Agency LEOP Low Earth Orbit Phase LS Launch Segment requirement MAR Mission Architecture Review MCC Mid-Course Correction M-class Medium Class MOS Margin Of Safety MR Mission Requirement MRD Mission Requirements Document



NEO Near Earth Object OP Operational requirement ORB Orbiter OTS Off-The-Shelf PA Product Assurance requirement P/L Payload PP Planetary Protection (requirement) ppb particle per billion ppm particle per million PRM Propulsion Module RAMS Reliability, Availability, Maintainability and Safety RD Reference Document RF RadioFrequency RPE Relative Pointing Error RQ Requirement RSE Radio Science Experiment RSSD Research and Scientific Support Department (ESA) SATS Sample Acquisition and Transfer System SRE-PA Advanced Studies and Technology Preparation Division S/C Spacecraft SD Study requirement SM Sampling Module SO Science Objective SPF Single Point Failure SRP Solar Radiation Pressure SS Sub-System SSAC Space Science Advisory Committee STR Star Tracker SY System TBC To Be Confirmed TBD To Be Determined TPS Thermal Protection System TRL Technology Readiness Level TRS Technology Reference Study VIS Visible

3.2 Terminology The space segment of a NEO sample return can be elementarily broken down into the following elements:



Outbound propulsion

ModuleOrbiter Descent

ModuleEarth Return

VehicleEarth Re-

entry Capsule

PRM ORB DM ERV ERC

Sampling Module

SM

Figure 3-1: NEO sample return building blocks

Using the above elementary elements a large number of combinations are possible and shall be investigated (including single composite) in Phase 1. The role of each element is briefly described below. Spacecraft composite: The combination of several System elements (S/C) during the launch,

NEO surface and proximity operations and cruises phases. Outbound Propulsion Module (PRM): The propulsion module used for the outward transfer Orbiter (ORB): Element used for orbital operations (no landing or descent capability) Descent Module (DM): The actual space element which descends towards the surface of the

NEO Sampling Module (SM): Module used for sampling operations (it has to be delivered via the

descent module). It includes the sampling mechanism itself but might not include the whole sample acquisition and transfer system as the sample may have to be transferred through the ERV as well

Earth Return Vehicle (ERV): The space element which returns to Earth after collection of a sample. It carries the ERC back to Earth and is in charge of placing it on a convenient re-entry trajectory. It has a propulsive capability. As the gravity of NEOs is in general very low, the Earth Return Vehicle (or the sampling module) will always encompass the re-ascent capability as well (no rendezvous and capture/docking capability is foreseen)

Earth Re-Entry Capsule (ERC): The capsule which performs the re-entry through the Earth atmosphere. The sample container is located inside the ERC. It may be a purely passive element (i.e. no active propulsive or braking capability)

In addition, the following definitions/clarifications may also be useful in the frame of this study: Sample acquisition and transfer system (SATS): All equipments involved in the sampling,

sample transfer, sealing and containment functions of the spacecraft. This is considered to be a spacecraft sub-system (i.e. not a payload element) and thus should be under ESA responsibility (incl. development and procurement cost) in the frame of this study

Hover and Go: Sampling/landing strategy defined by the following characteristics: - Sampling operations during GNC guided hovering - Surface contact only by sampling mechanism - No nominal impact (no impact energy for sampling/ascent)

Touch And Go: Sampling/landing strategy defined by the following characteristics: - Sampling operations during GNC guided impact - Surface contact by sampling mechanism and landing structure - Nominal impact (impact energy used for sampling/ascent)

Full landing (short stay): Sampling/landing strategy defined by the following characteristics:



- Sampling operations after GNC guided soft landing - Surface contact by sampling mechanism and landing structure - No significant change in illumination/thermal conditions of landing

Full landing (extended stay): Sampling/landing strategy defined by the following characteristics:

- Sampling operations after GNC guided soft landing - Surface contact by sampling mechanism and landing structure - Significant change in illumination/thermal conditions of landing

Payload: The payload is defined as the ensemble of physical element from the input signal collection up to the digital electronic signal production. The sample acquisition and transfer system and its equipment will not be considered as part of the payload. It will be treated as a mechanism sub-system and is under the direct responsibility of ESA. This is also the case of any other mechanism which is required by the different in-situ payload elements to ensure their proper positioning (e.g. robotic arm, carousel, etc.)

Absolute pointing error: Angular separation between the actual instantaneous generalised pointing vectors of the S/C and the commanded or desired generalised pointing vectors

Relative pointing error: Angular separation between the actual instantaneous generalised pointing vectors of the S/C and the median generalised pointing vectors defined over a time interval containing the reference time instant. Also known as pointing stability.

Absolute measurement error: Angular separation between the actual and measured generalised pointing vectors of the S/C

Horizontal or Lateral landing/touchdown velocity: The component of the touchdown/velocity parallel to the average local slope at the asteroid landing site

Vertical landing/touchdown velocity: The component of the touchdown/velocity perpendicular to the average local slope at the asteroid landing site

Sampling (or landing or touchdown) site (candidate): Safe area scientifically suitable for sample collection, defined on the NEO surface by the ground teams, which feature the following: the maximum characteristic size of all surface hazards (rocks, boulders, crater rims, etc.) is 50 cm

Landing accuracy: Downrange semi-major axis of the landing ellipse Landing ellipse: Ellipse centred on the sampling site which envelops all possible landing

locations reachable by the spacecraft at 3-σ confidence level taking into account the variation/error of input parameters: navigation (sensors, filter, etc.), gravity model and physical property (rotation axis, period, etc.), solar radiation pressure, control (thruster), S/C modelling (thruster alignment, etc.), etc.

Sampling spacecraft: In Phase 2 of the study, the sampling spacecraft is the main spacecraft (composite) and support all functions listed above: PRM, ORB, DM, SM, ERV

In this entire document, the terms “landing” and “touchdown” (i.e. with immediate re-ascent) are both indifferently used to indicate the phase when the spacecraft (or the sampling tool only) touches the surface. At this stage it is not determined yet which approach is selected. This will be defined in the final issue of this document.



4 STUDY REQUIREMENTS

4.1 Study approach and objectives R-SD.OBJ-1 The objective of this assessment study is to establish a feasible and cost-efficient

mission design fulfilling the requirements of this document and all other applicable documents (AD) within the constraints of the Cosmic-Vision 2015-2025 programme. As the time and cost constraints of the Cosmic Vision programme are very challenging a scientifically sound, technically simple and robust and cost-effective mission scenario, spacecraft design and development plan shall be produced. The results of this study

will be used in the decision process of the Cosmic Vision programme in 2009. R-SD.OBJ-2 The main output of the study will be a set of Technical Notes, as defined in the

statement of work, describing in details the space segment design and various trade-offs performed and which will serve as an input to the Cosmic-Vision selection process.

R-SD.OBJ-3 Specific attention will also be given to the Technology Development Plan which will be included in the programmatics data package of this study. Its objective is to identify technologies which are on the critical path of the spacecraft development and for which a timely and urgent technology development programme is required. Particular attention shall be given to R-PM.TEC-1 and R-PM.TEC-2

R-SD.OBJ-4 Detailed cost breakdown shall be included with the cost per item at the subunit level, including the envisaged TRL

4.2 Units R-SD.UNI-1 All documentation shall use the SI International System of Units, as specified in

[AD12] Part 1A, sections E-2 to E-6

4.3 Document format R-SD.FOR-1 The output of the system design and mathematical models shall be compatible

with the specifications provided in [AD7] R-SD.FOR-2 All structural Finite Element Model (FEM), if applicable, shall be delivered to

ESA in NASTRAN format R-SD.FOR-3 Any thermal mathematical model shall be delivered to ESA in ESATAN format

4.4 Margin requirements R-SD.MAR-1 The margin philosophy standard to all D/SCI assessment studies and defined in

[AD8] shall be applied. R-SD.MAR-2 Section 2.1.1 of [AD8] shall be applicable for the mass margin philosophy.



4.5 Coordinate systems R-SD.COR-1 All Marco Polo Reference Coordinate Systems (Reference frames) shall be

orthogonal and right-handed.

4.5.1 SPACECRAFT REFERENCE FRAMES R-SD.COR-2 For the description of the Marco Polo space element(s) and of its (their) major

components, the following S/C Reference Coordinate Frames shall be used:

+ ZS

+ YS

+ XS

Optical instruments

HGA

+ ZS

+ YS

+ XS

Optical instruments

HGA

Figure 4-1: Main spacecraft reference frame

4.5.2 ORBITAL REFERENCE FRAMES R-SD.COR-3 For the definition of the Marco Polo spacecraft attitude along its (their) orbit, the

following Local Orbital Reference Frame shall be used:

+ZL

+XL +YL

+ZL

+XL +YL

Figure 4-2: Local orbital reference frame

+ZL is oriented towards nadir and +XL is oriented along the velocity vector of the spacecraft.



R-SD.COR-4 When the radius of the NEO is used it shall reflect the radius of the equivalent sphere for the estimated NEO mass and an average density of ~ [1100-1500] kg.m-

2 (see R-EN.NEO-4) unless explicitly mentioned (i.e. for dynamics analyses a 2:1:1, 4:2:1 or more irregular shape models shall be used).



5 MARCO-POLO OVERVIEW

5.1 Top-level Mission Requirements The Marco Polo mission has two main top-level objectives which relate to its science objectives: Requirement

ID Title Statement

(Defines a requirement unless specified otherwise)

Rationale and/or Comments

MR: Mission Requirements R-MR-1 Primary mission

objective Marco Polo shall safely return to Earth a sample from the primitive NEO 1999 JU3 (Cg-type)

Primary target defined at the MAR. Optional targets considered in Phase 1: 2001 SG286 (D-type) 2001 SK162 (T-type) 1989 UQ (C-type)

R-MR-2 Science requirements The mission shall enable characterization of the NEO at different scales (e.g. dm, cm, mm)

Ref: [AD1], [AD6], [AD17], [AD18]

Table 5-1: Mission requirements

5.2 Mission Profile This paragraph describes the overall mission profile which has been selected at the MAR, based on the outputs of Phase 1, consultation with the European Science Study Team and an ESA internal review process. More details can be found on the mission profile and operations in [AD5], [AD16], [AD17] and [AD18].

5.2.1 MISSION OVERVIEW A single spacecraft on which is mounted the re-entry capsule is launched by Soyuz-Fregat 2-1b from Kourou on a direct escape trajectory to 1999 JU3, a primitive Cg-type NEO. The baseline mission is launched in December 2017/January 2018 with a backup in November 2018 or November 2019 (TBC). The spacecraft approaches the NEO in 2021/2022 (TBC). The spacecraft is then inserted in the vicinity of the asteroid. All manoeuvres are carried out via a chemical propulsion system. The detailed transfers to be considered (including back-up) are described in [AD5]. The first 3 to 4 months are dedicated to orbital science global observations and hazard mapping. A number of potential sampling sites are then characterized at high resolution. After a number of sampling rehearsals (which can be done jointly with the local characterization campaign) the spacecraft attempts to sample ~ 30-100 g of surface material on the most suitable one (i.e. yielding the best compromise between science return and risk-mitigation). If it is confirmed, via a high-fidelity verification technique, that a scientifically meaningful sample has been collected, the sample is transferred to the re-entry capsule and sealed (this can also be done before transfer to the ERC) and the spacecraft is ready for departure back to Earth or can carry on orbital science if more



time is available. If not, the spacecraft can undertake up to 5 sampling attempts. The spacecraft is also capable of collecting three different samples and keep them separated. The total stay duration of the spacecraft is in the order of 8 months. The spacecraft returns to Earth in December 2024, both in the baseline and backup scenario, and releases the re-entry capsule which undertakes a high-speed Earth re-entry at Ventry ~ 12.1 km.s-1.

5.2.2 SYSTEM DEFINITION The reference system breakdown, selected at the MAR is as defined in the diagram below. It essentially consists of two elements: the sampling spacecraft and the re-entry capsule. The sampling spacecraft is defined as a single spacecraft, the S/C composite, and embeds all primary functions: outbound and inbound propulsion, orbiting, descent, sampling and re-ascent.

Figure 5-1: Block diagram of the Marco Polo mission

5.2.3 MISSION PHASES Typical mission phases are listed hereafter. Detailed operations and their associated timeline and sequence are described in [AD16] and [AD18].

Phase name Orbit/location Main events Pre-launch activities Launch pad Battery charging

L-8 hours to launch

Marco Polo System

Launcher Segment Ground SegmentSpace Segment

Earth Re-entry Capsule

Main composite: Sampling spacecraft

ESA Ground stations and communications

Subnet

Mission and Science Operations

Mission Operations Centre (ESOC)

Science Data Archiving and Dissemination (ESAC)



Phase name Orbit/location Main events TBD

Launch (LEOP) Ascent trajectory, cislunar and direct escape

Duration: Launch + 3 days Ignition Soyuz 1st stage Jettisoning of Soyuz 1st stage Ignition of Soyuz 2nd stage Jettisoning of Soyuz 2nd stage Ignition of Soyuz 3rd stage Jettisoning of Soyuz 3rd stage Ignition of Fregat upper stage Launch into escape hyperbola by Fregat

S/C Acquisition Interplanetary Separation from Fregat upper stage S/C and P/L commissioning Deployment of solar panels and HGA if applicable 1st determination of S/C trajectory TBD

Interplanetary cruise Interplanetary Hibernation TBD

Solar conjunctions TBD Autonomous mode TBD

Fly-bys Earth, Mars (TBC) Fly-by altitude TBD

Fly-by date: TBD Instrument calibration TBD

S/C checkout Intermediate interplanetary cruise

Partial S/C wake-up System and P/L checkouts TBD

DSM Intermediate interplanetary cruise

S/C wake-up DSM burn TBD

Rendezvous with NEO

Approach trajectory

S/C wake-up Rendezvous burn, date TBD Far approach Initial fly-by, characterization of target Instrument commissioning Duration: ~ 2 weeks (TBC) TBD

Initial characterization

Far “formation flying”-like position (~ 5 km altitude) between the Sun and the NEO

Insertion within Hill sphere of NEO Start of proximity operations Initial characterization First mapping and determination of physical properties (rough mass, rotation period, axis orientation, shape, etc.) Orbit maintenance Duration: ~ 2 weeks (TBC) TBD

Gravity field campaign

Few km (TBD) altitude self-stabilizing orbit

Gravity field campaign on a terminator orbit (~ 1 month) Preliminary mapping when favourable conditions + laser measurements Delta-DOR Duration: ~ 1 month (TBC) TBD

Global characterization

Few km (TBD) altitude controlled

Global mapping of the NEO at medium spatial resolution on a controlled 09:00 – 21:00 orbit with a



Phase name Orbit/location Main events orbit local solar elevation between 30 and 60 degrees (~ 2

months) Global mapping (all orbital instruments on) Communications to Earth via HGA (~ 8 hours a day) Orbit maintenance Duration: ~ 2 months (TBC) TBD

Local characterization

Few 100 m (TBD) altitude

Local mapping of 5 selected sampling sites at high spatial resolution Duration: ~ 1 month (TBC) TBD

Sampling rehearsals TBD Might be combined with local characterization TBD

Sampling operations Descent and surface

Descent Touchdown/landing Sampling Identifying local contextual information Duration: ~ 10 weeks (2 weeks/attempt), incl. re-ascent

Re-ascent Re-ascent to safe position (e.g. ~ 5 km formation-flying-like position)

Ascent burn Re-acquisition of the S/C TM by the Earth TBD

Extended orbital operations

TBD Same as global characterization TBD

Earth return transfer injection

NA Departure manoeuvre TBD

Interplanetary return cruise

Interplanetary Hibernation TBD

ERC release Cis-lunar space S/C wake-up Release (and spinning, TBC) of the ERC TBD

Direct Earth re-entry Entry point: TBD Landing site: TBD

Entry at 120 km altitude Landing on Earth TBD

Retrieval operations NA Observation of re-entry Locating and retrieval of capsule via RF beacons TBD

Table 5-2: Overview of mission phases



Proximity operations ~ 8 monthsRendez-vous

~ 2 weeks

Departure

Far initial

characterization

Radio science

experiment


Local characteriza

tion + rehearsals

Sampling attempts

Spare month

21 3 4 5 6 70 8 months

Proximity operations ~ 8 monthsRendez-vous

~ 2 weeks

Departure

Far initial

characterization

Radio science

experiment


Local characteriza

tion + rehearsals

Sampling attempts

Spare month

21 3 4 5 6 70 8 months

Figure 5-2: Timeline of proximity operations



6 SUMMARY OF SCIENCE OBJECTIVES AND PLANETARY PROTECTION REQUIREMENTS

6.1 Science objectives and operations Below are recalled the main science objectives. The detailed scientific requirements and operations can be found in [AD1] and [AD18] and are to be considered fully applicable.

6.1.1 SCIENCE OBJECTIVES Requirement

ID Title Statement



SO: Science Objectives SO.BAS: Baseline Science Objectives R-SO.BAS-1 Sample return The Marco Polo mission shall return a

sample from the NEO 1999 JU3 Main objective of the mission. In-situ investigation shall enable the safe operation and manoeuvring of the spacecraft in close proximity to the NEO and safe collection of the sample(s)

R-SO.BAS-2 Global and local characterization

The Marco Polo mission shall place the samples in their global and local context

Determine global and local properties of the NEO, select the scientifically best landing site

R-SO.BAS-3 Primary science requirements

All science requirements defined in [AD1] as “shall” shall be fulfilled

SO.ADD: Additional Science Objectives G-SO.ADD-1 Complementary

science The Marco Polo mission should provide complementary science results not achievable from the samples themselves

Additional NEO science (shall not jeopardize R.SO.BAS-1/2/3)

G-SO.ADD-2 Additional science requirements

The compliance with all science requirements defined as “should” in [AD1] should be sought for

Additional NEO science (shall not jeopardize R.SO.BAS-1/2/3)

Table 6-1: Science objectives

6.1.2 SCIENCE OPERATIONS The science operations shall serve the previous mission objectives. In particular three phases have been identified in [AD1] as follows: • ‘Global characterisation’ means to measure the properties of the complete NEO, on a global

scale;



• ‘Local characterisation’ is the characterisation of up to 5 dedicated areas which are identified as potential sampling sites;

• ‘Sample context’ are measurements being performed at the actual sampling site.

A dedicated RSE campaign of ~ 1 month might be necessary in addition to the above (TBC). Spatial resolution for

imaging in the visual Spatial resolution for VIS/IR spectrometer

Spatial resolution for mid-IR instrument

Global characterisation Order of dm Order of m Order of 10 m Local characterisation Order of mm Order of dm Order of dm Context measurements Tens of μm - -

Table 6-2: Resolution requirements for global characterisation, local characterisation, and context measurements

6.2 Planetary protection requirements These are the scientific requirements related to planetary protection (e.g. PP category, contamination issues, etc.). The engineering requirements related to it will be specified in the proper element requirements (e.g. AIV for clean room environment, spacecraft for level of sterilization, etc.). Requirement

ID Title Statement


Rationale and/or

Comments PP: Planetary Protection Requirements PP.GEN: PP General Requirements R-PP.GEN-1 General PP The Marco Polo mission shall comply with all planetary

protection measures listed in [AD3] Ref. [AD3]

PP.FWD: Forward PP Requirements R-PP.FWD-1 Forward PP

Category The Marco Polo mission shall fulfil the requirements of planetary protection cat. III

Mars fly-by in baseline trajectory probability of impact with Mars < 10-4

R-PP.FWD-2 Forward PP contamination

Even though the forward planetary protection is cat. III, contamination of the sampling area (landing site) or the sample itself by the spacecraft shall be avoided or at least the potential contaminants shall be identified, controlled, tracked, documented and readily identifiable from the NEO sample material. The tracking of these contaminants shall start as of the S/C manufacturing process.

Ref. [AD1] & [AD3]. Relates to the science contamination. See R-SY.STZ-1

PP.BCW: Backward PP Requirements R-PP.BCW-1 Backward PP

Category The return leg of the Marco Polo mission shall be classified as planetary protection cat. V unrestricted return

Ref. [AD3]

Table 6-3: Planetary protection requirements



7 PAYLOAD DESCRIPTION AND REQUIREMENTS

7.1 Payload instruments Here is a brief description of the instruments. The reader will refer to the Payload Definition Document [AD6] for more detailed data (interfaces, performances, requirements, etc.). This payload has been defined such as to fulfil the science requirements defined in [AD1]. Any non-compliance with the science requirements (to be traded with mission profile and operations) shall be reported to the Agency. Baseline Payload Wide Angle Camera Conventional low-resource camera with a wide angle in order to image the

whole NEO from the science orbit. Located onboard the main spacecraft. Narrow Angle Camera Low-resource narrow angle camera used for medium/high-resolution science

measurement and hazard mapping from orbit. Located onboard the main spacecraft.

Laser altimeter Laser altimeter used to determine the global topography of the NEO. Located onboard the main spacecraft.

Visible/Near Infrared Spectrometer

Spectrometer used to determine the mineralogy on the NEO surface. Located onboard the main spacecraft.

Mid-Infrared Spectrometer Spectrometer used to determine the mineralogy on the NEO surface and the surface temperatures. Located onboard the main spacecraft.

Radio Science Experiment Radio Science Experiment making use of the communication system of the spacecraft to accurately determine the gravitational field of the NEO. Located onboard the main spacecraft.

Neutral Particle Analyzer Used to identify the processes related to the exposure to space environment and collisions. Located onboard the main spacecraft.

Close-up camera Provides the local microscopic context of the sampling area. Located onboard the main spacecraft on a deployable device (e.g. sampling mechanism).

Additional Payload Thermal sensors Provides surface temperature measurements. Located onboard the main

spacecraft in the landing feet (if applicable). Electric Field Sensors Provides electric field measurements. Located onboard the main spacecraft in

the landing legs (if applicable). APXS Used to determine the chemical composition on the NEO surface. Located

onboard the main spacecraft on one of the landing feet (if applicable). Volatile sensor/microscope Used to investigating the presence of volatiles, specifically water-ice, and also

silicate mineralogy within the NEO shallow surface. Located onboard the main spacecraft on one of the landing feet (if applicable), i.e. no mole or deployment device.

Table 7-1: List of instruments

7.2 Payload requirements Requirement

ID Title Statement



PL: Payload Module Requirements



PL.GEN: General Payload System Requirements R-PL.GEN-1 Model payload All primary scientific objectives shall be

addressed by using a set of model payload instruments defined in [AD6]

Instruments characteristics defined in the PDD shall be used as a reference

R-PL.GEN-2 Synergy science/GNC

The potential use of the science instruments as GNC sensors shall be addressed and, if identified, subsequent payload modifications shall be proposed

Lower resources by using science equipment as GNC sensors

R-PL.GEN-3 Optional payload

The accommodation of additional instruments defined in [AD6] shall be addressed to the greatest extent and implications at system level shall be identified and assessed. These instruments shall not bring additional risk to the mission

R-SO.BAS-3, G-SO.BAS-1/2

R-PL.GEN-4 Planetary Protection

All instruments defined in [AD6] shall be compliant with the planetary protection standards defined in [AD3]

To be addressed in [AD6] as well. See R-PP.GEN-1

PL.ACS: Payload attitude requirements R-PL.ACS-1 Absolute

Pointing Error The absolute pointing error of the orbital payload, during data acquisition, shall be less than 1.25 mrad 3-σ, for the yaw, roll and pitch angles (TBC)

From [AD6], driven by narrow-angle camera

R-PL.ACS-2 Relative Pointing Error

The relative pointing error of the orbital payload shall be less than 0.025 mrad over 1 s and 0.1 mrad over 7 s 3-σ, deg for the yaw, roll and pitch angles (TBC)

From [AD6], driven by narrow angle camera and mid-IR spectrometer

PL.MAS: Payload mass requirements R-PL.MAS-1 Orbital payload

mass The spacecraft shall provide an orbital payload mass allocation of 22.5 kg (TBC)

From [AD6] including maturity margins

R-PL.MAS-2 Surface payload mass

The spacecraft shall provide a surface payload mass allocation of 3.7 kg (TBC)


PL.POW: Payload power requirements R-PL.POW-1 Orbital payload

operating power

In operating mode the spacecraft shall provide an orbital payload power allocation of 108 W (TBC)


R-PL.POW-2 Orbital payload non-operating power

In non-operating mode the spacecraft shall provide an orbital payload power allocation of 35.5 W (TBC)


R-PL.POW-3 Surface payload operating power

In operating mode the spacecraft shall provide a surface payload power allocation of 30.8 W (TBC)


R-PL.POW-4 Surface payload non-operating power

In non-operating mode the spacecraft shall provide a surface payload power allocation of 5.2 W (TBC)


PL.DAT: Payload data requirements R-PL.DAT-1 Total orbital

payload data volume

The spacecraft design and operations shall cope with a total orbital payload data volume of 383.6 Gbit to be transmitted to Earth within the period defined in R-OP.POX-11 and no later than at the end of pre-sampling operations defined in R-OP.POX-2 (TBC)

From [AD6] without any margins ([AD8] guidelines to be implemented), compressed data volume

R-PL.DAT-2 Total surface The spacecraft design and operations shall From [AD6] without any



payload data volume

cope with a total surface payload data volume of 21.7 Gbit over five sampling attempts (TBC)

margins ([AD8] guidelines to be implemented) , compressed data volume

PL.GCR: Global characterization requirements R-PL.GCR-1 Local solar

elevation During global characterization in the visible range, the measurements shall be performed with a local solar elevation angle between 30 and 60 degrees

From [AD1]

Table 7-2: General Payload requirements



8 PROGRAMMATICS REQUIREMENTS

8.1 Programmatic requirements The following requirements are related to the schedule, planning and development of the mission as a whole, the space segment only and/or the specific technologies. Requirement

ID Title Statement



PM: Programmatic Requirements PM.SCH: Schedule Requirements R-PM.SCH-1 Schedule The launch of Marco Polo shall nominally occur in

2017 or 2018 to 1999 JU3 [RD6]

R-PM.SCH-2 Schedule Backup launch dates in 2018 or 2019 to 1999 JU3 shall be considered (minimum 1 backup launch date)

R-PM.SCH-3 Mission duration The total mission duration shall not exceed 8 years Operational costs PM.TEC: Technology and equipment Requirements R-PM.TEC-1 Technology

Readiness Only technologies at a minimum of TRL 5 in 2009 shall be used

[RD6]. In principle, TRL 5 in 2011. However for M-class missions, no technology development can be afforded until 2011 unless see R-PM.TEC-2

R-PM.TEC-2 Technology Readiness

Technologies not fulfilling R-PM.TEC-1 may be used in specific cases if and only if: No alternative fulfilling R-PM.TEC-1 has been

identified They also have a strong generic interest for future

ESA science and exploration missions Activities are ongoing/planned to bring the

considered technology up to TRL 5 by 2011

Potential future applications shall be identified

R-PM.TEC-3 Procurement European equipment shall be preferred as far as practicable

R-PM.TEC-4 OTS equipment OTS equipment shall be used wherever possible For critical technologies, see R-PM.TEC-1/2

Table 8-1: Programmatic Requirements.

8.2 Technology requirements R-PM.TEC-1 and R-PM.TEC-2 above describe the technology requirements themselves. The following table defines the ESA TRL levels which have to be used in the frame of this study.



Figure 8-1: TRL scale

ESA Technology Readiness Levels 9 Actual system "flight proven" through successful mission operations 8 Actual system completed and "flight qualified" through test and demonstration (ground or

flight) 7 System prototype demonstration in a space environment 6 System/subsystem model or prototype demonstration in a relevant environment (ground or

space) 5 Component and/or breadboard validation in relevant environment 4 Component and/or breadboard validation in laboratory environment 3 Analytical and experimental critical function and/or characteristic proof-of-concept 2 Technology concept and/or application formulated 1 Basic principles observed and reported

Table 8-2: Definition of ESA Technology Readiness Levels



9 SYSTEMS REQUIREMENTS

9.1 Launch segment requirements For performances of the launch vehicle it is referred to [AD9] and [RD3]. Requirement

ID Title Statement



LS: Launch Segment System Requirement LS.LAV: Launch Vehicle Requirement R-LS.LAV-1 Launch Vehicle The launch vehicle shall be Soyuz-Fregat 2-1 b [AD9], [RD3] R-LS.LAV-2 Launch margin A launch margin of 5% shall be considered Launch margin defined as

ratio in percent between: - launcher performance for

the target orbit ([AD5]) and

- total spacecraft wet mass (incl. all other margins: maturity, system, propellant, etc.) + launch vehicle adapter

LS.LAS: Launch Site Requirement R-LS.LAS-1 Launch site Launch site shall be CSG (Kourou, French

Guyana)

LS.ORB: Launch and Orbit Parameters R-LS.ORB-1 Orbit insertion Launch vehicle shall directly insert the

spacecraft composite on an escape hyperbola to 1999 JU3

[AD5]

LS.DEP: Departure R-LS.DEP-1 Departure

strategy The departure conditions are as defined in [AD5].

[AD5]

Table 9-1: Launch Segment Requirements

9.2 NEO target requirements 1999 JU3 shall be considered as the baseline target for Phase 2 of the study. The following requirements shall apply to new targets only (possibly identified in a later stage of the mission definition phase). Requirement

ID Title Statement



TA: Target Mission Requirement R-TA-1 NEO

detection The NEO target shall be bright enough so that it can be detected by the spacecraft onboard camera at a sufficient distance to ensure optimal rendezvous

See R-SS.ACS-4



manoeuvres R-TA-2 NEO

properties The NEO physical and orbital properties shall be such that self-stabilizing (i.e. no control of the S/C) orbits are possible

Taking into account all perturbations (SRP, tumbling movement, etc.)

R-TA-3 NEO rotation period

The rotation period of the selected NEO shall be greater than 2.5 hours and lower than 5 days (120 hours)

R-TA-4 NEO characteristics

It shall be possible to identify the rough size, the rotation period, the spin state, the binarity and the rough shape of the NEO before entering into implementation phase (~ 2011-2012)

Ground observations with relative visual magnitude lower than m=22 shall be possible

R-TA-5 NEO target selection

Should the targets listed in R-MR-1 not be suitable, other targets shall be considered. These shall belong to the taxonomy types: C and related (according to SMASS II classification), B, D or T

SMASS II classification defined in [RD9]

Table 9-2: NEO target requirements

9.3 Mission-level operational requirements This chapter defines the mission operations requirements. The following phases are covered in this chapter: post-launch or LEOP, cruise (outbound/inbound), DSM, GA, rendezvous, proximity operations, cislunar and re-entry. Requirement

ID Title Statement



OP: Operational System Requirement OP.POX: Proximity Operations Requirement R-OP.POX-1 Duration of

proximity operations

The duration of the proximity operations about the NEO shall be at least 8 months and not exceed 24 months

Mitigation of risk, feasibility of science operations, maximizing rehearsals, etc. Definition: the proximity operations start when the spacecraft enters within the Hill radius of the NEO

R-OP.POX-2 NEO global characterization

The minimum duration for pre-sampling operations shall be at least 4 months and a half (TBC)

Covers far initial characterization, RSE and global/local characterization campaign (from [AD18])

R-OP.POX-3 NEO local characterization

The local characterization campaign to image the 5 candidate sampling sites at high resolution shall be at least 4 weeks (TBC)

Duration of local characterization campaign (from [AD18])

R-OP.POX-4 Hazard mapping The global characterization orbit shall be such that all hazardous surface features (rocks, boulders, steep slopes, crater rims, etc.) can be determined and precisely located with an accuracy better than 50 cm on a global scale

Preliminary mapping of hazards

R-OP.POX-5 Hazard mapping The local characterization campaign shall be such that all surface hazardous features larger than 10 cm will have been mapped in the vicinity of the 5 sampling site candidates

Mapping of hazards on the landing site.



Requirement ID

Title Statement (Defines a requirement unless

specified otherwise)


(~size corresponding to the expected landing/touchdown accuracy)

R-OP.POX-6 Re-attempt Proximity operations shall allow a minimum of 5 sampling attempts

Mitigation of failure

R-OP.POX-7 Collision avoidance with NEO

Proximity operations shall be planned so as to avoid collision with the NEO in case of failures on the S/C or in ground operations. The S/C shall support these activities

R-OP.POX-8 Science resolution requirements

The proximity operations shall be such as to fulfil the resolution requirements listed in [AD1] and in Table 6-2

R-OP.POX-9 Proximity operations operational margin

An extra month shall be included at the end of the proximity operations for margin purposes and to prepare for return manoeuvres

R-OP.POX-10 Global characterization conditions

Proximity operations shall be such that 100 % of the illuminated surface areas can be imaged in the visible range with a local solar elevation angle between 30 and 60 degrees

See R-PL.GCR-1.

R-OP.POX-11 Global characterization duration

The nominal duration of the global characterization campaign shall be 8 weeks (TBC)

See [AD18].

G-OP.POX-1 Multiple sampling It should be possible to collect one sample at three different locations and keep them separated

See [AD1]

G-OP.POX-2 Duration between collection of two separate samples

Between collection of two samples or two sampling attempts, two weeks should be dedicated to review of the first sampling/attempt operations

in order to evaluate the science interest of the first acquired sample and analyse telemetry data from the previous sampling phase

G-OP.POX-3 Separate lander It should be possible to deliver a separate in-situ lander

OP.LAN: Descent and landing/touchdown operational Requirement R-OP.LAN-1 (Deleted)

Descent and landing/touchdown conditions

All NEO descent, landing/touchdown and sampling operations shall occur during solar illumination

Use of cameras as descent imager

R-OP.LAN-2 Descent During descent a minimum of 5 pictures of the surface shall be taken

To acquire the attitude during descent in case a problem occurs

R-OP.LAN-3 Sampling rehearsals

A full sampling attempt shall be preceded by at least one successful descent rehearsal

A descent rehearsal does not need to involve surface touchdown. Its aim is to validate the convergence of the navigation filter. Details TBC

R-OP.LAN-4 Touchdown/landing The mission design shall enable NEO landing/touchdown and sampling operations at any latitude and longitude as long as R-OP.LAN-5 is fulfilled

R-OP.LAN-5 Surface operations conditions

The surface operations on the asteroid shall occur on a fully illuminated location

In order to get context information in the visible range (close-up imager)



Requirement ID




R-OP.LAN-6 Surface operations duration

The surface operations shall be as short as possibly allowed by the sampling method and in no case longer than half the rotation of the NEO

No extended stay (over night). Minimization of sampling operations

G-OP.LAN-1 Descent The 5 descent images should be transmitted real-time to Earth

If not possible, the images shall be stored onboard the S/C

OP.MOP: Mission Operations Requirement R-OP.MIS-1 Jettisoning The jettisoning strategy of any element shall

ensure collision avoidance with other spacecraft elements or the asteroid at any stage of the mission

Mostly relate to the Soyuz-Fregat stage (due to direct escape), ERC and possibly sampling tool if necessary

R-OP.MIS-2 Communications, operations

No critical S/C operations shall be performed if the Sun-Earth-S/C angle is lower than 5o

e.g. descent, landing, local characterization, etc.

R-OP.MIS-3 Communications, operations

No standard S/C operations shall be performed if the Sun-Earth-S/C angle is lower than 2o

R-OP.MIS-4 Re-entry strategy The re-entry capsule shall be released by the ERV from the return hyperbolic trajectory and directly enter the Earth atmosphere

No Earth orbit insertion

R-OP.MIS-5 (Deleted)

Thrust arcs For electric propulsion the ratio “thrust-to-coast” arcs shall be less than 85 %

To allow corrections in case of thruster outages

R-OP.MIS-6 Distances The spacecraft shall support the success of all operations with a minimum distance to the Sun of 0.85 AU, a maximum of 1.7 AU and a maximum distance to Earth of 1.7 AU during proximity operations and 2.5 AU in cruise (TBC)

Envelope of maximum/minimum distances. Detailed geometry to be used for sizing defined in [AD5]

R-OP.MIS-7 Solar conjunction The spacecraft shall function autonomously throughout any solar conjunctions on cruise to and back from the NEO

R-OP.MIS-8 Solar conjunction duration

The duration of a Solar conjunction or when the Sun-Earth-S/C angle is lower than 2o during the transfer to and from the asteroid shall be limited to 50 days

OP.MAS: Mission analysis Requirement R-OP.MAS-1 Re-entry conditions The re-entry velocity and flight path angle of

the ERC, defined by mission analysis, shall be such that heat fluxes and heat loads do not exceed respectively 15 MW/m2 and 250 MJ/m2

Re-entry conditions to be defined at 120 km altitude

G-OP.MAS-1 Night re-entry Mission analysis should ensure a night re-entry of the ERC

To ensure good observations of the chemical species involved in re-entry

OP.SAT: Satellite operational modes Requirement R-OP.SAT-1 Satellite operational

modes An operations profile shall be defined covering all mission phases from launch until the end of the mission

R-OP.SAT-2 Autonomy The S/C shall autonomously perform nominal operations when ground intervention is not possible

R-OP.SAT-3 Redundancy, safe A redundant functional path shall be triggered Safe mode if in a safe



Requirement ID




and collision avoidance mode

and used upon any failure, independently from ground control. It shall allow continuity of the nominal mission timeline and performances. If no redundant path exists or if the failure is too critical, the spacecraft shall enter Safe Mode or Collision Avoidance Mode

position or Collision Avoidance mode if FDIR triggered during critical phase, i.e. descent, touchdown/landing, sampling, re-ascent.

Table 9-3: Operations Requirements

9.4 Space segment requirements

9.4.1 SPACECRAFT SYSTEM REQUIREMENTS This chapter defines the general system requirements applicable for the sampling spacecraft and the ERC or the space segment as a whole (general, margin, etc.). Requirement

ID Title Statement

(Defines a requirement unless Specified otherwise)


SY: Space Segment System Requirement SY.GEN: General System Requirement R-SY.GEN-1 Lifetime The spacecraft lifetime shall be 7 years in space

and 1 year for pre-launch ground activities (AIT/AIV/storage) in a controlled environment and all space-based elements shall be designed and sized for this duration

R-SY.GEN-2 LV compatibility The space segment shall be compatible with a Soyuz-Fregat 2-1 b launch from Kourou

including fitting within the LV fairing and interface with LV adapter. See R-LS.LAV-1

R-SY.GEN-3 Compatibility of the S/C with backup launch and NEO

The space segment shall be sized for the mission scenario which is most demanding (i.e. worst case ΔV and launch mass for either primary or backup)

Besides ΔV, worst cases of other parameters shall also be considered (e.g. NEO rotation rate, NEO size, etc.)

R-SY.GEN-4 Launch vehicle integration

The design of the space segment shall cope with an integration on a Soyuz-Fregat 2-1 b from CSG

e.g. vertical integration atop the Soyuz third stage on the launch pad, etc.

R-SY.GEN-5 In-situ surface package

An additional in-situ surface package or any additional science requirements [AD1] shall not place any critical constraints on the sample return spacecraft design and operations

See R-SO.BAS-1/2/3, G-SO.ADD-1/2

R-SY.GEN-6 Stack The space segment shall be composed of two spacecraft: ERC + main composite (or sampling spacecraft)

R-SY.GEN-7 Design standards All spacecraft shall be design according to [AD10], [AD11], [AD12], [AD13], [AD14] and [AD15] unless otherwise specified in this

Any conflict between the requirements in this document and the ECSS



Requirement ID


Specified otherwise)


document. standards shall be reported to ESA

R-SY.GEN-8 Configuration The spacecraft system shall accommodate all equipments and sub-systems

R-SY.GEN-9 S/C design approach

A design-to-cost and risk minimization mission design approach shall be followed

Reduction of cost and risk

R-SY.GEN-10 ERC The sampling spacecraft shall accommodate, transport and release the ERC back to Earth

R-SY.GEN-11 Observation of ERC release

The sampling spacecraft shall take images of the separated ERC after release

SY.MAR: Spacecraft Margin Philosophy R-SY.MAR-1 Margin The mission design shall follow the margin

philosophy defined in [AD8] Section 2.1.1 of [AD8] for mass margins

R-SY.MAR-2 Aerothermodynamics margins

For aerothermodynamics fluxes computation, 100 % and 20% margin shall be added respectively to the radiative and convective parts of the raw data (from numerical analysis w/o margins)

R-SY.MAR-3 Power margins The spacecraft composite power sub-system shall be sized assuming a power and energy system margin of 20% assuming a battery cell failure and a loss of solar cell string

Tayloring of R-P-3 in [AD8]

R-SY.MAR-4 Propellant margin for proximity operations

A margin of 20% will be added to all delta-Vs used for NEO proximity operations (incl. orbit maintenance, descent and ascent, hold-on thrust, etc.)

SY.DEB: Debris mitigation R-SY.DEB-1 Protection from

debris The S/C shall be protected from damage due to debris and dust encountered during proximity operations

e.g. spacecraft jettisoning, release, landing on NEO, etc.

R-SY.DEB-2 Micro-meteorites The S/C shall be designed to withstand the predicted micro-meteorite environment throughout the mission with a probability larger than 0.998

The micro-meteorite environment is as defined in [AD4]

SY.PRM: Propulsion system R-SY.PRM-1 (deleted)

MCC manoeuvres The propulsion system shall be able to perform mid-course corrections at any time before NEO and Earth arrival

R-SY.PRM-2 (deleted)

Proximity operations manoeuvres

The propulsion system shall cope with the proximity operations manoeuvres requirements derived from gravity field campaign, controlled orbits or descent/touchdown and re-ascent phases

Delta-V and thrust

SY.ORB: Orbiting System Requirement R-SY.ORB-1 Determination of

gravity field The knowledge of the gravity field of the asteroid shall be such as to fulfil R-SY.LAN-7

To enable accurate landing/touchdown

R-SY.ORB-2 Science orbital payload accommodation

The main spacecraft shall accommodate the science orbital payload and enable fulfilling all orbital primary science measurements

Payload as defined in [AD6]

R-SY.ORB-3 (deleted)

Observation of separation (if

The orbiter shall take images of the separated modules (e.g. propulsion, descent/landing, etc.)

Applicable only if orbiter separate from other



Requirement ID




applicable) modules R-SY.ORB-4 (Deleted)

Observation of touchdown/landing (if applicable)

The orbiter shall observe the touchdown/landing manoeuvre

Applicable only if orbiter separate from descent/landing module

R-SY.ORB-5 Instrument commissioning

The main spacecraft shall allow for commissioning and calibration of the orbital science instruments during the transfer to the NEO

G-SY.ORB-1 Instrument calibration

On the way to the NEO it should be possible to calibrate the relevant instruments via fly-bys

SY.LAN: Decent/Landing System Requirement R-SY.LAN-1 Landing/touchdow

n conditions If a full landing is considered the sampling spacecraft shall approach the surface with a maximum horizontal velocity of 5 cm.s-1 (TBC)

To limit the magnitude of forces on the touchdown/landing system and to mitigate toppling

R-SY.LAN-2 Landing conditions (if applicable)

If a full landing is considered the sampling spacecraft shall approach the surface with a maximum vertical velocity of 30 cm.s-1 (TBC)

Not necessarily applicable for “touch and go” or “hover and go” approaches, see 3.2. To be traded against landing system and contamination (last braking manoeuvre)

R-SY.LAN-3 Landing conditions (if applicable)

The sampling spacecraft shall have a maximum attitude of +/-10o with respect to the local surface at landing/touchdown (TBC)

Not necessarily applicable for “touch and go” or “hover and go” approaches, see 3.2. To mitigate toppling and to limit constraints on spacecraft geometry (e.g. solar panels)

R-SY.LAN-4 Re-attempt capability

The sampling spacecraft shall have a multi-attempt capability (up to 5)

To mitigate sampling failure (also covers multiple sampling requirement). See R-OP.POX-6 (+ G-SS.MEC-2)

R-SY.LAN-5 Landing system (if applicable)

If a full landing is considered the sampling spacecraft the landing/touchdown system shall be able to cope with hazardous features with diameters up to 50 cm and worst case landing/touchdown conditions (VV=30 cm.s-1, VH=5 cm.s-1,θ =10o)

Not necessarily applicable for “touch and go” or “hover and go” approaches, see 3.2. Landing system sized for worst case landing and re-attempt capability. VV depends on selected approach. See R-EN.LAN-1, R-SY.LAN-1/2/3

R-SY.LAN-6 Dust and debris All equipments of the sampling spacecraft shall be protected from damage caused by dust and debris resulting from the landing phase

See R-SY.DEB-1

R-SY.LAN-7 Landing accuracy (if applicable)

If a full landing (short or long term) is considered the landing accuracy of the sampling

1.5 m corresponds to ~ S/C radius



Requirement ID




spacecraft shall be better than the radius of the sampling site – 1.5 m, 3-σ value (TBC)

See R-SY.LAN-5 and all related definitions (3.2) For worst case (smallest) available sampling site (R-EN.LAN-1), landing accuracy shall be better than 1 m.

R-SY.LAN-8 Landing system configuration (if applicable)

The landing system footprint shall fit within the fairing of the launch vehicle without requiring deployment mechanisms

Only if landing legs are envisaged

R-SY.LAN-9 ERC accommodation

The sampling spacecraft shall accommodate the ERC in a safe location such that the risk of collision of the ERC TPS with the NEO surface is minimized

R-SY.LAN-10 Preservation of telemetry data

After touchdown/landing the full set of telemetry data acquired during descent shall be protected from overwriting until receipt on Earth is confirmed

R-SY.LAN-11 (deleted)

Surface Payload accommodation

The touchdown/ landing element shall accommodate the science payload and enable fulfilling all surface science measurements


R-SY.LAN-12 Stability The landing/touchdown system shall guarantee the static and dynamical stability of the spacecraft at landing/touchdown and during sampling operations

R-SY.LAN-13 Hover and go/touch and go

If the spacecraft performs sampling without fully landing on the surface, the clearance with respect to the local surface shall be at least 2 m (TBC)

G-SY.LAN-1 Surface Payload accommodation

If a full landing is considered the sampling spacecraft should accommodate the in-situ science payload and enable fulfilling all surface science measurements


SY.ERV: Earth Requirement R-SY.ERV-1 (deleted)

Rendezvous and capture

Rendezvous and capture/docking capability shall not be considered

Due to the technology development constraints

R-SY.ERV-2 (deleted)

ERC transport The ERV shall carry the ERC back to Earth

R-SY.ERV-3 (deleted)

Observation of ERC release

The ERV shall take images of the separated ERC

SY.ERC: ERC System Requirement R-SY.ERC-1 Heat fluxes The ERC heat shield shall be able to withstand

heat fluxes up to 15 MW/m2 Compatible with TPS technology in Europe

R-SY.ERC-2 Earth landing shocks

The ERC and its sub-systems shall be designed such that the shock loads on:

- the sample container are lower than 800 g half-sine for ~ 5 ms (TBC) (100 Hz) or 200 g for 20 ms (25Hz)

- RF beacons and related batteries are lower than TBD g half-sine for ~ TBD ms

Limit adequate for the sample. See [AD1]

R-SY.ERC-3 ERC recovery The ground segment recovery facilities and



Requirement ID




(Deleted) selected landing site shall allow ERC recovery within 24 hours post-landing

R-SY.ERC-4 (Deleted)

ERC sealing system

The ERC sealing system shall prevent: Less than 0.1 ppm of terrestrial water Dust or liquid particles larger than 0.3 μm

to penetrate into the sample container

Design of the ERC sealing system. See [AD1]

R-SY.ERC-5 ERC function The ERC shall safely deliver the sample to the Earth surface according to R-SY.ERC-2 and shall allow retrieval of the sample within the capsule

See R-SY.ERC-2

R-SY.ERC-6 ERC stability The ERC shall be statically and dynamically stable at all aerodynamic regimes

ERC dimensions, mass and CoM, CoP location

R-SY.ERC-7 ERC Beacons The ERC shall provide power to RF beacons for at least 5 days after landing

R-SY.ERC-8 Maximum deceleration

The maximum deceleration of the ERC through re-entry shall be 80 g (TBC)

To ensure integrity of the ERC up to (excluded) landing

R-SY.ERC-9 Ballistic entry The ERC re-entry shall be ballistic, non-propulsive and non-controlled

G-SY.ERC-1 ERC design The ERC design should ensure that the sample is never exposed to temperatures higher than +40oC. For less than one minute +80oC is acceptable

See [AD1]

G-SY.ERC-2 Flight instrumentation package

The ERC should be equipped with a flight instrumentation package to perform the following measurements during entry and descent: heat flux, temperature and pressure on both front and back shields at a frequency of 0.5 Hz (TBC)

This information will be stored (no real-time communications). If implemented, the maximum allowable shocks on the memory shall be defined and included in R-SY.ERC-2

SY.SEA: Sample Sealing Requirement R-SY.SEA-1 Sample container

sealing The sample sealing system shall prevent: Less than 0.1 ppm of terrestrial water Dust or liquid particles larger than 0.3 μm

to contaminate the sample up to 24 h after landing

See [AD1]

R-SY. SEA -2 Sample cross-contamination

No particles larger than 100 µm from any of the collected samples shall contaminate another sample

See [AD1]

R-SY. SEA -3 Sample container integrity

The sample sealing system (see R-SY.SEA-1) shall withstand a shock load up to 800 g half-sine for ~ 5 ms (TBC) (100 Hz) or 200 g for 20 ms (25Hz)

See [AD1]

SY.STZ: Sterilization Requirement R-SY.STZ-1 Sterilization

requirement The spacecraft elements shall cope with the following sterilization and cleaning procedures: -Alcohol cleaning -TBD -Etc.

Mitigation of forward contamination, see R-SS.PRO-1. See [AD1]

SY.AUT: Autonomy Requirement



Requirement ID




R-SY.AUT-1 Autonomy requirement

S/C autonomy shall be limited to the minimum level needed to ensure achievement of all mission goals

R-SY.AUT-2 FMECA A FMECA of the system level design shall be carried out to determine failure modes and recovery options

Table 9-4. Spacecraft systems Requirements.



9.4.2 PLATFORM SUB-SYSTEM REQUIREMENTS This table defines all key S/C sub-system requirements (pointing, thermal, power, data handling, telecommunications, etc.). Requirement

ID Title Statement



SS: Platform Sub-System Requirement SS.THE: Thermal System Requirement R-SS.THE-1 Spacecraft thermal

control The spacecraft thermal control system shall be designed to cope with the spacecraft sub-system thermal requirements, including payload

R-SS.THE-2 Space environment The spacecraft thermal control system shall cope with the space environment throughout the mission (incl. surface operations if applicable)

See [AD4]

SS.POW: Power System Requirement R-SS.POW-1 Spacecraft power The spacecraft power system shall cope with the

power requirements of the various spacecraft sub-systems, including payload, at any stage of the mission

i.e. power generation, storage, conditioning and distribution

R-SS.POW-2 Battery requirements

The spacecraft battery shall be sized for worst case eclipses and the descent/landing/re-ascent phase

Eclipses will be avoided for nominal operations but may be encountered if control of the spacecraft is lost

R-SS.POW-3 Solar array sizing The sizing of the solar arrays shall allow the S/C to stay on a self-stabilized orbit

Solar radiation pressure. See R-OP.POX-4

SS.ACS: AOCS/GNC Requirement R-SS.ACS-1 Payload line of

sight The spacecraft AOCS system shall comply with the payload pointing requirements and all applicable R-SY.LAN requirements

Payload stability requirements defined in [AD6] and R-PL.ACS-1/2

R-SS.ACS-2 Navigation sensors The GNC descent approach shall be based on a vision-based or mapping LIDAR navigation system

R-SS.ACS-3 Robustness of GNC system

During descent the GNC navigation system shall be robust to all NEO surface features larger than 25 cm (i.e. guaranteeing R-SY.LAN-1, 2, 3 and 7)

25 cm = 50 cm feature minus 50 % margin

R-SS.ACS-4 NEO acquisition The spacecraft navigation system shall be able to acquire the target from 300 000 km distance and 2 weeks before the start of proximity operations

To enable accurate localization and insertion manoeuvres

R-SS.ACS-5 Stability at landing/touchdown

The AOCS/GNC system shall guarantee the stability of the spacecraft at landing/touchdown and during sampling operations

Also valid for “hover and go”

R-SS.ACS-6 Safe mode attitude The AOCS system shall be able to maintain, during safe mode, the composite in a Sun-pointing attitude using a minimum of the onboard resources while ensuring power generation and ground communication for vital equipment

R-SS.ACS-7 STR blinding The AOCS/GNC system shall be able to autonomously detect blinding of the STR by Sun



Requirement ID




or planets or small bodies (NEO, etc.). In such event the AOCS/GNC system shall make attitude estimation based on IMU integration, if applicable

R-SS.ACS-8 IMU initialisation The spacecraft composite design shall enable the initialisation and accurate calibration of IMU, if applicable, before descent/landing

R-SS.ACS-9 IMU redundancy If applicable, the sampling spacecraft shall provide triple redundancy and majority voting during descent, landing and sampling

R-SS.ACS-10 Pointing during observations

During observations, the AOCS system shall be able to orient the payload line of sight towards nadir while maintaining a communication link with the ground via the HGA and providing power from the solar arrays

SS.DHS: Data Handling System Requirement R-SS.DHS-1 Data handling The spacecraft data handling system shall cope

with the data transfer and storage requirements of the various spacecraft sub-systems, including payload, at any stage of the mission

R-SS.DHS-2 TC/TM management

The DHS shall be able to receive TC and send TM from/to ground

R-SS.DHS-3 Instrument/equipment command

The DHS shall be able to command the instruments and equipment onboard

R-SS.DHS-4 Reconfiguration capability

The DHS shall provide reconfiguration capabilities in case of failure detection

R-SS.DHS-5 Redundancy management

The DHS shall manage the redundancy for the relevant sub-systems

R-SS.DHS-6 Experiment cycle During surface operations, the DHS shall provide the capability to store the scientific and housekeeping data generated by a complete experiment cycle

Experiment cycle corresponds to measurements of all surface instruments ([AD6])

SS.PRO: Propulsion System Requirement R-SS.PRO-1 Sample

contamination by thrusters

Contaminants or undesired particles coming out of the thrusters shall have a density lower than 1014 (1013 as a goal) molecules/cm2 on the NEO surface

Typical unwanted contaminants: organics, MMH, Fe-rich metal grains, most volatiles, etc. See [AD1]

R-SS.PRO-2 Sizing of propulsion system

The propulsion system shall cope with all operations and associated Delta-V/thrust requirements incl.: transfer (MCC, etc.), orbital operations, initial fly-bys, gravity field campaign, controlled orbits, characterization of 5 different sampling sites, sampling rehearsals, descent/touchdown and re-ascent phases, etc.

R-SS.PRO-3 By-products The by-products of the propulsion system shall be characterized and reported at ppb level

R-SS.PRO-4 S/C attitude vs propulsion

The propulsion sub-system shall be designed to be compatible with any operational S/C attitude. The main engine and any reaction control

Particularly for NEO surface environment



Requirement ID




thruster shall be thermally qualified for such environmental conditions

R-SS.PRO-5 Propellant freezing point

The propulsion sub-system thermal design shall assure that the minimum predicted temperatures of any wetted component or surface contacting the propellant remain at least 10oC above the maximum freezing point of the onboard propellant

R-SS.PRO-6 Performances of propulsion system

The performances of the propulsion system in terms of total impulse and margin shall satisfy the requirements imposed by the mission, the trajectory analysis and the overall system requirements

SS.COM: Communication System Requirement R-SS.COM-1 2-way

communications 2-way (uplink and downlink) communication with the ground shall be provided at all time ensuring a margin as defined in R-SS.COM-8. Whenever this is not possible due to mission constraints, R-SS.COM-3 shall apply

Except during Earth occultation or when the Sun-S/C-Earth angle is too small. This does not fully apply to the ERC.

R-SS.COM-2 Orbiter relay If a separate lander is implemented, 2-way communications shall be relayed through the orbiter

Avoid duplication

R-SS.COM-3 Real-time TM Real-time data shall be provided directly to Earth during descent and landing allowing monitoring of the major events

Beagle 2 recommendations: get a minimum health status of the S/C throughout descent

R-SS.COM-4 ESA telecommunication standards

All S/C shall comply with ESA telecommunication standards defined in [AD14]

ECSS

R-SS.COM-5 Ranging and Doppler

The communication system shall support the two-way Ranging and Doppler measurements of the S/C throughout all mission phases and ΔDOR when high-precision navigation is required (i.e. RSE campaign)

Standards defined in [AD14]

R-SS.COM-6 Weather availability

The link budgets of the spacecraft to ground shall be calculated for a weather availability of 95%

R-SS.COM-7 RSE equipment The RSE equipment shall be considered as part of the telecommunication equipment unless it is only used for science

R-SS.COM-8 Bit error rate The maximum bit error rate shall be better than 10-5

R-SS.COM-9 Communication system redundancy

The communication system shall implement full redundancy except for the HGA

R-SS.COM-10 Simultaneous handling of communication functions

The telecommunication system shall be capable of simultaneously handling telemetry, ranging and telecommands

SS.STR: Structure Requirement R-SS.STR-1 Structure loads The spacecraft structure shall support the launch

environment [AD9] with MOS as defined in



Requirement ID




[AD12] which will be agreed with the agency R-SS.STR-2 Mechanical loads The spacecraft structure shall support the

mechanical static and dynamic loads encountered during its entire lifetime, including: manufacturing, handling, transportation, testing, launch and in-orbit operations (incl. touchdown/landing on the NEO)

R-SS.STR-3 Sample container The structure directly containing the sample (sample container) shall be able to withstand the loads as defined in R-SY.ERC-2

R-SS.STR-4 Structural stiffness The structural stiffness shall guarantee fundamental frequencies of the S/C within the requirements of the LV to avoid dynamic coupling

R-SS.STR-5 Structural design The structural design shall be based on simple load paths

R-SS.STR-6 (Deleted)

Structural mass The structural mass shall be as low as possible

SS.MEC: Mechanism Requirement R-SS.MEC-1 Sampling

mechanism The sampling mechanism shall be able to collect a minimum of 30 g of NEO material with a goal of 100 g

i.e. 30 g at once minimum and a goal of ~ 100 g if G-SS.MEC-2 is considered. See [AD1]

R-SS.MEC-2 Sampling mechanism

The sampling mechanism shall be able to collect cm-sized pebbles , plus a large amount (104) of small particles (100s of μm-sized to mms-sized)

Pebbles up to 3 cm. See [AD1]

R-SS.MEC-3 Sampling mechanism

The sampling mechanism shall be compatible with NEO material having the properties defined in R-EN.NEO-2 and [AD4]

See R-EN.NEO-2

R-SS.MEC-4 Sampling verification

The sampling mechanism (or any other part of the spacecraft) shall be equipped with autonomous means of verifying suitable sampling

Determine whether the sampling operation was successful or not (+ rough sample mass and volume estimate)

R-SS.MEC-5 SATS cleanliness All mechanisms involved in the SATS chain shall be cleaned so as to fulfil levels specified in R-SS.PRO-1 and according to the standards defined in R-SY.STZ-1

See R-SS.PRO-1 for quantification + see [AD1], R-PP.FWD-2

R-SS.MEC-6 Tracking of contaminants

Witness plates shall be implemented in order to track the possible in-situ contaminants and shall be returned to Earth

e.g. propellant, S/C outgassing, etc. See [AD1], R-PP.FWD-2

R-SS.MEC-7 Protection against debris

All mechanisms shall be protected against debris generated by landing and sampling on the NEO

Guarantee mechanism function

R-SS.MEC-8 SATS Monitoring The SATS system shall be able to monitor its torque, thrust and penetration depth (if applicable)

R-SS.MEC-9 Blocking of the sampling unit

In case a sampling unit gets blocked during sampling and repeated attempts to execute the emergency dislodge manoeuvre are unsuccessful,



Requirement ID




it shall be possible to release the blocked unit

R-SS.MEC-10 Backup sampling method

A highly reliable backup sampling method shall be implemented to ensure a sample can be collected should the main sampling mechanism fail

For the backup sampling method, a sample mass requirement of ~ g is acceptable (R-SS.MEC-1)

G-SS.MEC-1 Sample temperature

After collection of a sample,the design of the sample collection and distribution unit (incl. container) should ensure that the sample is never exposed to temperatures higher than +40oC. For less than one minute +80oC is acceptable

Sample temperature and collection process. See [AD1]

G-SS.MEC-2 Sampling at different locations

The sampling mechanism should allow to collect one sample from 3 different locations and keep them separated (from a contamination viewpoint)

See [AD1], G-OP.POX-1

G-SS.MEC-3 Determination of soil’s properties

Through the monitoring of its various sensors’ information as a function of penetration progress, the SATS should support the determination/estimation of the soil’s mechanical properties

Soil properties: density, porosity/compaction, hardness/cohesion, cementation, etc.

Table 9-5. Platform sub-system Requirements.

9.5 Environmental requirements The following table defines the environment-related requirements that the spacecraft has to comply with throughout the mission. Requirement

ID Title Statement



EN: Environmental Requirement EN.LAV: Launch Vehicle Environment Requirement R-EN.LAV-1 Launch

environment The launch environment used for the design of the spacecraft shall be as defined in ([AD9])

EN.EMC: Electromagnetic Compatibility Requirement R-EN.EMC-1 EMC The electrical design shall comply with the

requirements of [AD11]. Tailoring of these requirements may be proposed and need to be justified

EN.GEN: General Environment Requirement R-EN.GEN-1 Environment

requirement The mission design shall be compliant with the general space environment (cruise) requirements defined in [AD10] and the mission-specific (NEO) environment requirements defined in [AD4]

Mission environment

EN.LAN: Landing/touchdown Site Environment Requirement R-EN.LAN-1 NEO surface

environment It shall be assumed that the NEO surface features at least one sampling site of ~ 5 m diameter

[AD4]

R-EN.LAN-2 NEO surface The maximum surface temperature of the [AD4], temperature on



temperature sampling site to be considered for the design of the spacecraft shall be 400 K

1999 JU3 at perihelion

G-EN.LAN-1 Landing site selection

The selected sampling site should feature hazardous features smaller than 10 cm

Landing site minimizing risk. Operational requirement. For sizing purposes R-EN.LAN-1 shall be used

EN.NEO: NEO Environment Requirement R-EN.NEO-1 NEO rotation

period The mission design shall be compatible with a NEO rotation period of ~ 7.7 hours +/- 10%

R-EN.NEO-2 Surface properties The surface properties which shall be assumed for the design of the sampling mechanism are as follows: Grain Size: Sub-µm mm up to 3 cm Shape: Very angular Cohesion: 0.1 – 1 kPa Compressive strength: Up to 100 kPa Shear strength: 1 – 2.6 kPa Bulk density: <1.2 – 1.8 (max) Angle of internal friction: 30° - 50°

No solid rock. Only loose material shall be assumed.

R-EN.NEO-3 NEO shape For the analysis of proximity operations (including landing/touchdown) different shape and gravitational models shall be used incl. worst cases (e.g. spherical, 2:1:1, irregular with singularities). It will also be assumed that the NEO may be rotating around its non-principal rotation axis

NEOs have very irregular shapes. Tumbling movement. See [AD4]. Binary shall not be assumed for sizing purposes but analyses are required to verify the impact on mission success

R-EN.NEO-4 NEO density The NEO density shall be assumed to be between 1100 and 1500 kg.m-3

[AD4]

R-EN.NEO-5 NEO size A NEO diameter of 900 m +/- 20 % shall be assumed

[AD4]

Table 9-6: Environmental requirements

9.6 AIV and testing requirements Requirement

ID Title Statement



AV: AIV System Requirement AV.LAV: Sterilization Requirement R-AV.LAV-1 Integration into LV The AIV activities of the S/C shall be compliant

with the integration requirements of a Soyuz-Fregat 2-1 b from CSG

AV.STZ: Sterilization Requirement R-AV.STZ-1 Sterilization

facility and equipment

The AIV facilities shall enable the implementation of sterilization and cleaning procedures as defined in R-SY.STZ-1

Table 9-7: AIV Requirements



9.7 Ground segment Requirement

ID Title Statement



GS: Ground Segment Requirement GS.COM: Ground communication Requirement R-GS.COM-1 Compatibility with

ESA deep space network

The mission shall use and be compatible with the standards of the ESA deep space network as well as the NASA and JAXA deep space network

GS.REC: Ground recovery facilities/site Requirement R-GS.REC-1 ERC recovery The ground segment recovery facilities and

selected landing site shall allow ERC recovery within 24 hours post-landing

9.8 Product assurance and RAMS requirements Requirement

ID Title Statement



PA: Product Assurance Requirement PA.REL: Reliability Requirement R-PA.REL-1 Reliability Reliability considerations shall be used as

additional criteria to trade-off candidate concept options such that R-PA.REL-3 is achieved

Note: if R-PA.REL-3 is TBD, the mission architecture selection shall be such that overall mission reliability is maximized.

R-PA.REL-2 Failure and SPF Failure propagation shall not be allowed and SPF shall be removed (or minimized if full SPF removal cannot be guaranteed)

Redundancy to be considered to ensure SPF tolerance

R-PA.REL-3 Mission reliability The overall mission reliability shall be better than TBD

PA.RED: Redundancy Requirement R-PA.RED-1 (Deleted)

Mechanism redundancy

A redundant sampling mechanism or a backup sampling device less complex than the baseline system shall be implemented

To minimize the risk of sampling failure. R-SS-MEC.1 might not be completely fulfilled with this backup mechanism (Agency must be informed)

R-PA.RED-2 Mechanism redundancy

The design philosophy for all mechanisms shall be as defined in [AD12]

Objective: Single Point Failure Tolerant

Table 9-8: Product Assurance Requirements



10 LIST OF TBC AND TBD

10.1 List of TBC Chapter Requirement ID Description 5.2.1 & 5.2.3

NA Orbit/location and events of the mission overview and mission phases

7.2 R-PL.ACS-1,2 Absolute/Relative Pointing Error of the payload line of sight 7.2 R-PL.MAS-1,2 Mass allocation for the orbital/surface payload 7.2 R-PL.POW-

1,2,3,4 Power allocation for the orbital/surface payload

7.2 R-PL.DAT-1,2 Total generated data volume allocation for the orbital/surface payload

9.3 R-OP.POX-2,3, 11

The duration depends on the NEO physical properties (rotation period and orientation of the spin axis) which are usually not accurately known beforehand and on the operational strategy (see above)

9.3 R-OP.MIS-6 Earth-Sun-NEO ranges and relative angles 9.3 R-OP.LAN-3 Definition of “sampling rehearsal” 9.4.1 R-SY.LAN-

1,2,3,7 All landing/touchdown-related requirements will be consolidated once the landing/touchdown system (structure, mechanism and GNC) requirements are better known

9.4.1 R-SY.LAN-1,2 Clearance of the S/C with respect to the surface for a hover & go approach (depends on size of the S/C and solar panels in particular as well as attitude requirement)

9.4.1 R-SY.ERC-2 The shocks that the sample container can withstand upon landing on Earth are to be confirmed by further analysis

9.4.1 R-SY.ERC-8 Maximum ERC deceleration through re-entry 9.4.1 G-SY.ERC-2 Measurement frequency of the ERC flight instrumentation

package 9.4.1 R-SY.SEA-3 The shocks that the sample sealing can withstand upon landing

on Earth are to be confirmed by further analysis Table 10-1: List of TBC

10.2 List of TBD Chapter Requirement ID Description 5.2.3 NA Main events of the various mission phases 9.4.1 R-SY.ERC-2 Maximum allowable shocks on the RF beacons 9.4.1 R-SY.STZ-1 Level of sterilization to be defined in [AD3] based on

contamination requirements 9.8 R-PA.REL-3 Overall mission reliability

Table 10-2: List of TBD