Microsoft Word - [AD1]_MarcoPolo-R_Mission_Requirements_Document

Prepared by David Agnolon

Reference SRE-PA/2011.076/MarcoPolo-R

Issue 3

Revision 2

Date of Issue 17/09/2012

Status Authorised

Document Type RQ

Distribution SRE-F, MP-R Science Study Team

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MarcoPolo-R Mission Requirements Document

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MarcoPolo-R MRD

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Title MarcoPolo-R Mission Requirements Document

Issue 3 Revision 2

Author David Agnolon Date 17/09/2012

Approved by Date

Reason for change Issue Revision Date

To be written

3 2

Issue 3 Revision 2

Reason for change Date Pages Paragraph(s)

To be written

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Table of Contents

INTRODUCTIO N ................................................................. ............................................................... 5

1.1

Background ........................................................................................................................................................................ 5

1.2

Scope of Document ............................................................................................................................................................ 6

1.3

Documentation architecture .............................................................................................................................................. 6

The mission document tree is given here below for information and lists the envisaged documentation (to be confirmed)

available by the time of the Preliminary Requirements Review. ............................................................................................... 7

1.4

Issue schedule .................................................................................................................................................................... 7

1.5

Requirements nomenclature ............................................................................................................................................. 7

1.5.1

Requirements and goals .................................................................................................................................................. 7

1.5.2

Requirement ID ............................................................................................................................................................... 8

MISSION OVERVIEW ........................................................................................................................ 9

2.1

Mission Profile ................................................................................................................................................................... 9

2.1.1

System definition ............................................................................................................................................................. 9

2.1.2

Mission overview ............................................................................................................................................................. 9

2.1.3

Mission phases and operations ..................................................................................................................................... 10

2.1.4

Asteroid proximity operations ...................................................................................................................................... 12

MISSION REQUIREMENTS ............................................................................................................. 13

PROGRAMMATICS REQUIREMENTS ............................................................................................. 13

PLANETARY PROTECTION AND SAMPLE CONTAMINATION REQUIREMENTS ........................... 14

5.1

Planetary Protection requirements ................................................................................................................................. 14

5.1.1

Science contamination requirements ........................................................................................................................... 14

PAYLOAD DESCRIPTION AND REQUIREMENTS ........................................................................... 15

6.1

Model payload instruments ............................................................................................................................................. 15

6.2

Payload requirements ...................................................................................................................................................... 15

TECHNOLOGY REQUIREMENTS .................................................................................................... 16

LAUNCH SEGMENT REQUIREMENTS ............................................................................................ 17

OPERATIONS AND MISSION ANALYSIS REQUIREMENTS ............................................................ 18

SPACE SEGMENT REQUIREMENTS ................................................................................................ 21

10.1

Spacecraft system requirements ...................................................................................................................................... 21

10.2

Platform sub-system requirements ................................................................................................................................. 24

10.2.1

Thermal requirements ................................................................................................................................................... 24

10.2.2

Power requirements ...................................................................................................................................................... 25

10.2.3

GNC/AOCS requirements ............................................................................................................................................. 25

10.2.4

DHS requirements ......................................................................................................................................................... 26

10.2.5

Propulsion requirements ............................................................................................................................................... 27

10.2.6

Communications requirements .................................................................................................................................... 27

10.2.7

Structure requirements ................................................................................................................................................ 28

10.2.8

Mechanism requirements .............................................................................................................................................. 29

10.3

Autonomy and FDIR requirements ................................................................................................................................ 30

10.4

Environmental requirements .......................................................................................................................................... 31

10.5

AIV and testing requirements ......................................................................................................................................... 32

10.6

Ground segment ............................................................................................................................................................... 32

10.7

Product assurance and RAMS requirements .................................................................................................................. 34

COORDINATE SYSTEM ................................................................................................................... 35

APPLICABLE AND REFERENCE DOCUMENTS ............................................................................... 36

12.1

Applicable documents ...................................................................................................................................................... 36

12.2

Reference document ........................................................................................................................................................ 37

ACRONYMS AND TERMINOLOGY .................................................................................................. 38

13.1

Acronyms ......................................................................................................................................................................... 38

13.2

Terminology .................................................................................................................................................................... 40

LIST OF TBC AND TBD .................................................................................................................... 41

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14.1

List of TBC ........................................................................................................................................................................ 41

14.2

List of TBD ........................................................................................................................................................................ 41

Index of Figures and Tables

Figure 1-1: Mission development planning (indicative only and if JUICE launch is 2024) ... 5

Figure 1-2: MarcoPolo-R Phase A/B1 schedule, major milestones for spacecraft and

payload development (indicative only) .................................................................................. 6

Figure 1-3: Preliminary MarcoPolo-R mission document tree ................................................... 7

Figure 2-1: Block diagram of the MarcoPolo-R mission .................................................... .......... 9

Figure 2-2: Asteroid proximity operations for the 2022 mission scenario (indicative only) 12

Figure 11-1: Main spacecraft reference frame ............................................................................. 35

Figure 11-2: Local orbital reference frame .......................... ......................................................... 35

Figure 13-1: Baseline spacecraft architecture (note that the sealing function of the SATCS

might also be performed within the ERC) ........................................................................... 40

Table 2-1: Overview of baseline mission phases ......................................................................... 12

Table 3-1: Resolution requirements for global characterisation, local characterisation, and

context measurements ........................................................................... ................................. 13

Table 6-1: List of instruments, all located on the sampling spacecraft ................................... 15

Table 7-1: Definition of ESA Technology Readiness Levels and TRL scale ............................ 17

Table 14-1: List of TBC .................................................................................................................... 41

Table 14-2: List of TBD ................................................................................................................... 41

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1 INTRODUCTION

1.1 Background

In 2011, after a meeting of the Space Science Advisory Committee, the M3 candidate missions for

the Cosmic-Vision 2015-2025 programme were selected for further assessment and consideration

for launch in the early 2020’s. The MarcoPolo-R Near-Earth Asteroid sample return mission was

selected as one of the M-class candidate missions. The nominal launch date for M-class missions

is now 2024, but the missions shall also be compatible with opportunities in 2022.

During the fall 2011 the mission concept was studied in the Concurrent Design Facility at ESTEC

(internal pre-assessment) in order to prepare a thorough industrial assessment phase. The

purpose of the ongoing assessment phase is to assess the feasibility of this mission. After the

assessment phase ending in 2013, one M-class mission will be selected. If adopted, the mission

would proceed into definition and implementation (Phases B1/B2/C/D). In parallel with the

system level activities, an Announcement of Opportunity will be issued calling for the provision of

scientific instruments by the science community through ESA. It is anticipated that the selection

of the science instruments will be confirmed by February 2013.

Currently, references to the payload point to the Payload Resources and Requirements Document

(PRRD, [RD16]). After the payload for MarcoPolo-R is selected, this will be replaced by individual

Experiment Interface Documents – Part B (EID-B) for each instrument.

The overall mission development planning is shown in the figure below. The dates and various

activities are indicative only. The planning will be updated following the assessment phase and the

Contractor shall provide its own bottom-up development schedule.

Figure 1-1: Mission development planning (indicative only and if JUICE launch is 2024)

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Figure 1-2: MarcoPolo-R Phase A/B1 schedule, major milestones for spacecraft and payload

development (indicative only)

1.2 Scope of Document

This is a Mission Requirements Document (MRD) to be used as an Applicable Document in the

MarcoPolo-R industrial assessment study. The purpose of the MRD is to provide all high-level

mission-level requirements (including S/C and payload, launcher, ground segment and

operations) necessary to achieve the science goals detailed in [AD2] for the MarcoPolo-R

industrial system design studies running through 2012/2013. It includes functional and

performance requirements down to the sub-system level which can be defined at this stage.

Later on in the course of the definition phase, it will result into two self-standing documents, i.e.

the Mission Requirements Document and the System Requirements Document.

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.

This is issue 3.2 of the MRD. It has been updated after the baseline selection review in the course

of the assessment phase. It will be reviewed as part of the assessment phase and will be updated

following the Preliminary Requirement Review at the end of 2013.

1.3 Documentation architecture

The mission requirements document is one of the documents that constitute the foundations of

the mission profile for MarcoPolo-R.

This MRD is specifically supported by:

 MarcoPolo-R Science Requirements Document [AD2]

 MarcoPolo-R Mission Environment Document [AD4]

 MarcoPolo-R Planetary Protection Document [AD5]

 MarcoPolo-R Consolidated Reference Mission Analysis Document [AD6] (to be provided

in the course of task 2, [RD1] to be used prior to that)

 MarcoPolo-R Payload Resources and Requirements Document [RD16] (note that this

document supersedes the Payload Definition Document, previously referred to as [AD3],

no longer applicable)

 MarcoPolo-R Margin philosophy Document [AD13]

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The mission document tree is given here below for information and lists the envisaged

documentation (to be confirmed) available by the time of the Preliminary Requirements Review.

Figure 1-3: Preliminary MarcoPolo-R mission document tree

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 MarcoPolo-R mission. 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 MarcoPolo-R 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. Their impact on the mission

technical complexity and programmatic aspects (e.g. cost, risk) shall be limited so as to

stay within the M-class boundaries. Goals may be fulfilled under limited favourable

conditions. They are highlighted in italic in order to be easily identified.

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A requirement may need an explanation. Such an explanation can be found in italic below the

requirement itself. This is intended only to assist the understanding of the requirement or its

origin (the Parent requirement ID might be sometimes included wherever applicable).

1.5.2 Requirement ID

The requirement ID, preceded by R- or G-, belongs to the following category:

 AIV: AIV

 AUT: Autonomy and FDIR

 COM: Communication sub-system

 CON: Contamination

 DHS: DHS sub-system

 ENV: Environmental

 GNC: GNC/AOCS sub-system

 GRS: Ground Segment

 LAS: Launch Segment

 MEC: Mechanism sub-system

 MIS: Mission

 OPS: Operational

 PLD: Payload

 POW: Power sub-system

 PPR: Planetary Protection

 PRM: Programmatic

 PRO: Propulsion sub-system

 PQA: Product/Quality Assurance, RAMS

 SCI: Science

 STD: Study

 STR: Structure/Mechanical sub-system

 SYS: Space Segment System

 TEC: Technology

 THE

: Thermal sub-system

Should a requirement be deleted in a later update of this document, its number will not be reused

and should a requirement be introduced in between other requirements, the last digit will be used

(e.g. R-YYY-115 between R-YYY-110 and R-YYY-120).

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2 MISSION OVERVIEW

2.1 Mission Profile

This paragraph describes the baseline MarcoPolo-R system and its mission profile selected at the

Baseline Selection Review of the assessment phase in June 2012, in consultation with the

European Science Study Team, the industrial teams and through an ESA internal review process.

2.1.1 System definition

The system breakdown is defined in the diagram below.

Figure 2-1: Block diagram of the MarcoPolo-R mission

2.1.2 Mission overview

MarcoPolo-R is a sample return mission to the primitive Near-Earth Asteroid (NEA) 1996 FG3

which is a binary system. A spacecraft composite made of a main spacecraft, also called the

sampling spacecraft and carrying the Earth re-entry capsule (ERC), will be launched by a Soyuz-

Fregat MT 2-1b launch vehicle from Kourou on a direct escape trajectory to 1996 JU3.

The mission design and its development needs to be compatible with launch opportunities from

2022 onwards up to 2025 for programmatic reasons. Therefore, appropriate transfers for the

launch slots between 2022 and 2025 have been identified with relevant backup opportunities and

are all described in [RD1].

The 2022 mission scenario is described hereafter. Mission transfers, swing-bys and associated

dates and duration slightly vary across the various opportunities. The reader will refer to [RD1] for

the full analysis and the requirements in this document refer to the most demanding case for each

parameter.

The launch takes place in March 2022. After one Venus Swing-by in October 2024, the spacecraft

rendezvous with the NEA in February 2026.

MarcoPolo-R

Launcher

Segment

Ground

Segment

Space

Segment

Earth Re-entry

Capsule

Sampling

spacecraft

Mission

Operations Centre

(ESOC), Ground

Stations and

Communications

SubNet

Science Operations

Centre (SOC),

Science planning,

Data Archiving and

Dissemination

(ESAC)

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In the 2022 scenario the proximity operations last for ~ 190 days (but for other opportunities it

ranges from 180 through 390 days). The proximity/science operations are described in 2.1.4. In a

nutshell, this phase will include instrument calibration, far and close observations of the binary

system with the various science instruments, gravity field determination and hazard mapping

above roughly 5 km altitude. Five sampling site candidates are then characterized at high

resolution at about 250 m (TBC) distance to the surface during a few minutes in order to

determine the most suitable sampling site (i.e. yielding the best compromise between science

return and risk-mitigation).

After one successful sampling rehearsal, the spacecraft, which will be designed to cope with

surface hazards (e.g. large clearance to the surface), navigates towards the finally selected

touchdown/sampling site and collects hundreds of g of surface material. The sampling strategy is

based on a touch and go approach to lower cost and risk. Therefore the spacecraft performs a soft

touchdown of the surface for a few seconds (in the order of 2-5 seconds) and then takes-off

immediately after that to move into a safe position away from the surface.

If it is confirmed, via a reliable verification technique, that a scientifically meaningful sample has

been collected, the sample is transferred to the re-entry capsule and sealed. The spacecraft is then

ready for departure back to Earth or can carry on orbital science if desired. If the sampling

operation was not successful, the spacecraft can undertake 2 more attempts (3 attempts in total).

The spacecraft departs from the asteroid in August 2026 and returns to Earth in April 2029 after

one Venus swing-by in November 2027. The re-entry capsule is then released and undertakes a

high-speed Earth re-entry at V

entry

of maximum ~ 13 km

-1

(in other scenarios the re-entry velocity

is lower ~ 11-12.6 km.s

-1

). The capsule will be retrieved on the ground at the Woomera test range

which has already been used for Hayabusa landing.

2.1.3 Mission phases and operations

Phase name Su-phase name Orbit/location Main events

Launch

Pre-launch Launch pad

 L-8 hours to launch

Launch and escape

(LEOP)

Ascent trajectory

and direct escape

 Duration: Launch + 3 days

 Launch into direct escape orbit by Fregat MT

S/C Acquisition

(LEOP + S/C

commissioning).

Interplanetary

 Separation from Fregat MT upper stage

 Deployment of solar panels

 S/C and P/L commissioning

 Electric propulsion sub-system

commissioning

 1

determination of S/C trajectory

 1

launcher dispersion correction manoeuvre

Outbound cruise

Interplanetary

cruise

Interplanetary

 Various electric propulsion thrust and coast

phases

 1

year: daily contact with S/C to monitor

health status of electric propulsion system

 After one year, only weekly coverage of the

S/C

 System and P/L check-out

Venus (x0 or x1 or

x2) or Earth fly-by

(x0 or x1) (generic)

Venus/Earth fly-by

 Interruption of electric propulsion

 Possible targeting manoeuvres with chemical

propulsion

 Determination of post fly-by S/C trajectory

 Possible recovery manoeuvre

Asteroid

rendezvous

Rendezvous with

1996 FG3

Approach

trajectory, from ~

2000000 km

(TBC) to 500 km

 SEP and far approach

 Far range imaging (both for science and

navigation)

 Asteroid localization

 Targeting manoeuvres

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Phase name Su-phase name Orbit/location Main events

Arrival at 1996 FG3

500 km distance to

asteroid

 S/C checks

 Date of arrival in the 2022 scenario: 8

February 2026

Proximity

operations

Close approach

trajectory

500 to 100 km

distance to asteroid

 Start of proximity operations

 P/L commissioning and calibration

 7 days

Transition to far

global

characterization

100 to 10 km

distance to asteroid

 Start of asteroid mapping

 10 days

 Binary system mapping

Far global

characterization

Far “formation-

flying” (~10 km)

between the Sun

and the asteroid

 Initial characterization

 Initial mapping and determination of physical

properties

 Orbit maintenance

 10 days

Radio science

~ 3 to 5 km

distance to the

asteroid surface

 Gravity field campaign on a radio science

orbit

 Preliminary mapping when favourable

conditions

 20 days

Global

characterization,

GCP

~ 5km distance to

the asteroid surface

 Global mapping of the asteroid primary and

secondary at medium spatial resolution on a

TBD orbit with a local solar elevation between

30 and 60 degrees

 All instruments on (except CuC)

 Communication to Earth via HGA (~ 8

hours/day)

 No science during communication

 Orbit maintenance on a regular basis

 35 days

Local

characterization,

LCP

250 m distance to

the asteroid surface

 Local mapping of 5 selected sampling sites at

high spatial resolution

 15 days, ~ 3 days/site

Sampling/descent

rehearsals

Descent phase up

to ~ 100 m

 Sampling attempt down to at least 100 m

 Up to validation of critical events/GNC

algorithms, except sampling

Descent and

sampling, SAM

Descent and

surface

 Switch to autonomous descent mode

 Descent

 Descent interruption

 S/C position check by Earth via pictures

analysis

 Descent continuation

 Touchdown

 Sampling

 CuC operations

 21 days, ~ 3 attempts, 7 days/attempt

Ascent

Re-ascent to safe

position

 Ascent burn

 Transition to Safe Position

Post-sampling LCP

250 m distance to

the asteroid surface

 Local mapping of the site which has been

sampled

 3 days

Extended GCP

Same as GCP

 Same as GCP

 Transfer of the collected sample to ERC

Departure from

1996 FG3

Far distance to the

surface

 Start-up of the SEP system

 Injection of return trajectory

 Date of departure in the 2022 scenario: 16

August 2026

Inbound cruise

Interplanetary

cruise

Interplanetary

 Same as outbound cruise

Venus fly-by

Same as outbound

fly-by

 Same as outbound Venus fly-by

ERC release Cis-lunar space

 S/C wake-up

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Phase name Su-phase name Orbit/location Main events

 Release and spinning of ERC 4 hours before

entry

Earth arrival

Direct Earth re-

entry

Entry point TBD

Landing site:

Woomera test

range

 4 hour coast arc by ERC

 S/C Earth avoidance manoeuvre

 Entry at 120 km altitude

 Parachute-free entry

 Landing at Woomera

 Date of Earth arrival in 2022 scenario: 18

April 2029

ERC recovery

operations

 Observations of re-entry

 Localization and recovery of capsule

 Transfer to curation facilities

Table 2-1: Overview of baseline mission phases

2.1.4 Asteroid proximity operations

The proximity operations shall serve the previously defined high-level mission requirements as

well as science requirements defined in [AD2]. In particular three phases have been identified as

follows: Global Characterization Phase (GCP), Local Characterization Phase (LCP) and context

measurements, the latter being performed during sampling operations (SAM) and therefore very

short (in the orders of seconds). A so-called radio science phase is also implemented in order to

get the gravity field of the binary system to the level required in [AD2].

The science operations around the asteroid, to be further defined in [AD15], are part of the

proximity operations. Other operations are: instrument commissioning, descent and sampling,

etc. All proximity operations for the 2022 scenario are shown in the figure below.

Mission scenarios in 2024 and 2025 are less demanding as the duration of the proximity

operations is far longer, i.e. resp. 275 and 290 days. In the 2023 scenario, the proximity

operations last for 180 days. The latter is the sizing case. So, the 18 days margins shown in the

figure below become 9 days in the 2023 scenario.

The detailed description of the proximity and science operations as well as the associated

requirements will be detailed in the Science Assumptions Document and the Mission Operations

Assumptions Document ([AD14] and [AD15]).

Figure 2-2: Asteroid proximity operations for the 2022 mission scenario (indicative only)

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3 MISSION REQUIREMENTS

The MarcoPolo-R mission has two main top-level objectives which relate to its science objectives:

R-MIS-010 MarcoPolo-R shall safely return to Earth a sample from the primitive Near-Earth

Asteroid 1996 FG3.

Main objective of the mission.

R-MIS-020 The mission shall perform global and local characterization of the Near-Earth

Asteroid to retrieve the sample context information at the resolutions defined

below.

Spatial resolution

for imaging in the

isual ran

Spatial resolution

for VIS/IR

ectrometer

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

1 mm - -

Table 3-1: Resolution requirements for global characterisation, local characterisation, and context

measurements

 ‘Global characterisation’ means to measure the properties of the complete NEA, on a

global scale;

 ‘Local characterisation’ is the high-resolution 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.

Besides the science objectives, in-situ investigations defined in R-MIS-020 are meant to

support the safe operation of the S/C in close proximity to the NEA and the safe collection

of the sample(s).

4 PROGRAMMATICS REQUIREMENTS

The following requirements are related to the schedule, planning and development of the mission

as a whole.

R-PRM-010 The launch of the MarcoPolo-R baseline mission to the binary asteroid 1996 FG3

shall nominally occur by 2022.

Note: The M3 mission is intended as a backup candidate for the 2022 launch slot,

depending on JUICE schedule. Although in the science programme the baseline

launch slot for M3 is 2024, all current study planning should remain compatible

with a 2022 launch. The final decision on the nominal launch date will be made

following JUICE mission adoption.

R-PRM-020 The MarcoPolo-R mission shall also be compatible, in all ways, with launch

opportunities in 2024 and 2025.

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“All ways”: Launch mass; Sizing of: propellant tanks, thermal control, power,

data downlink, etc.

G-PRM-030 The MarcoPolo-R mission should also be compatible, in all ways, with launch

opportunities in 2023.

R-PRM-040 The total mission duration shall not exceed 8 years.

Mission duration is defined as the time between launch and landing of the Earth

re-entry capsule at the end of mission.

R-PRM-050 Deleted.

R-PRM-060 Deleted.

R-PRM-070 Deleted.

R-PRM-080 Deleted.

R-PRM-090 Deleted.

5 PLANETARY PROTECTION AND SAMPLE

CONTAMINATION REQUIREMENTS

5.1 Planetary Protection requirements

These are the high-level requirements related to planetary protection and sample contamination

(e.g. PP category, S/C cleanliness, contaminant tracking, etc.). The engineering requirements

related to it will be specified in the relevant category requirements (e.g. AIV for clean room

environment, spacecraft for level of sterilization, etc.).

R-PPR-010 The MarcoPolo-R mission shall comply with all planetary protection measures

listed in [AD5].

R-PPR-020 The MarcoPolo-R mission shall fulfil the requirements of planetary protection

category II.

R-PPR-030 The return leg of the MarcoPolo-R mission shall be classified as planetary

protection category V unrestricted return.

R-PPR-040 Deleted.

5.1.1 Science contamination requirements

R-CON-010 Contaminants or undesired particles (TBD) generated by the spacecraft (e.g.

propellant plume, outgassing, etc.) shall have a density lower than 10

(10

as a

goal) molecules/cm

on the asteroid surface and/or on the collected sample.

Typical unwanted contaminants TBD by the science team. See [AD2]

R-CON-020 The potential contaminants shall be identified, controlled, tracked, documented

and readily identifiable from the asteroid sample material. The tracking of these

contaminants shall start as of the S/C manufacturing process.

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Ref. [AD2] & [AD5]. Relates to the science contamination.

R-CON-030 Witness plates shall be implemented in order to track the possible contaminants

throughout the AIV/AIT and various mission phases and shall be returned to Earth.

E.g. propellant, S/C outgassing, etc. See [AD2].

R-CON-040 The spacecraft elements shall cope with the following sterilization and cleaning

procedures in order to keep low-levels of forward contamination:

 Alcohol cleaning,

 TBD.

Ref. [AD2].

6 PAYLOAD DESCRIPTION AND REQUIREMENTS

6.1 Model payload instruments

Here is a brief description of the model payload instruments. The reader is referred to the Payload

Resources and Requirements Document [RD16] for more detailed data (interfaces, performances,

requirements, etc.). This model payload has been defined by the MarcoPolo-R Science Study Team

such as to fulfil the science requirements defined in [AD2]. Only the core orbital payload is

recalled here. Complementary instruments may also be envisaged but are not part of the model

payload. It will be determined at a later stage (Instrument AO – 2012) by the Science Programme

Committee whether the complementary instruments shall and can be accommodated.

Core Payload

Wide Angle Camera Necessary to image the whole NEA, build the shape model and

retrieve context for the NAC from the science orbit.

Narrow Angle Camera Used for high-resolution science measurement and hazard

mapping from orbit.

Visible/Near Infrared

Spectrometer

Spectrometer used to determine the mineralogy on the NEA

surface.

Mid-Infrared Spectrometer Spectrometer used to determine the mineralogy on the NEA

surface and the surface temperatures.

Radio Science Experiment Radio Science Experiment making use of the communication

system of the spacecraft to accurately determine the gravitational

field of the NEA.

Neutral Particle Analyzer Used to identify the processes related to the exposure to space

environment and collisions.

Close-u

camera Provides the local microsco

ic context of the sam

lin

area.

Table 6-1: List of instruments, all located on the sampling spacecraft

6.2 Payload requirements

R-PLD-010 The model payload as defined in the PRRD [RD16] shall be assumed.

The model payload will be replaced by the selected instruments after the payload

Announcement of Opportunity.

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R-PLD-020 Science instruments shall not be used as baseline GNC sensors.

R-PLD-030 Deleted.

R-PLD-040 Deleted.

R-PLD-050 The Absolute Pointing Error (APE) of the payload, during data acquisition, shall be

less than 0.1 mrad 3-σ, for the yaw, roll and pitch angles (TBC).

R-PLD-060 The Relative Pointing Error (RPE) of the payload shall be less than 0.03 mrad over

1 s 3-σ, deg for the yaw, roll and pitch angles (TBC).

R-PLD-070 The sampling spacecraft shall provide a payload mass allocation of 36 kg (TBC).

Including maturity margins, but no systems margin.

R-PLD-080 During science operations in orbit around the asteroid, the spacecraft shall provide

an operating payload power allocation of 76.8 W (TBC).

Including maturity margins, but no systems margin.

R-PLD-090 During non-science operations in orbit around the asteroid, the spacecraft shall

provide a stand-by payload power allocation of 28.8 W (TBC).

Including maturity margins, but no systems margin.

R-PLD-100 During descent and sampling operations the spacecraft shall provide a power

allocation of 15 W (TBC) to the close-up camera. All other instruments are in off (if

allowed by their thermal requirements) or stand-by mode.

Including maturity margins, but no systems margin. CuC only.

R-PLD-110 The spacecraft shall provide 10.8 W (TBC) of power to the close-up camera when it

is in stand-by mode.

10.8 W is the power consumption of the camera set CCSU, no systems margin.

This power is common to WAC, NAC and CuC and should always only be counted

once.

R-PLD-120 Deleted.

R-PLD-130 Deleted.

R-PLD-140 Deleted.

R-PLD-150 Deleted.

7 TECHNOLOGY REQUIREMENTS

R-TEC-010 Only technologies at a minimum of TRL 5 by the time of the SPC Mission selection

(right after PRR) shall be implemented in the mission design.

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R-TEC-020 Technologies not fulfilling R-TEC-010 may be used in specific cases after approval

by the Agency, if and only if:

 No alternative fulfilling R-TEC-010 has been identified,

 Activities with a low development risk are ongoing/planned to bring the

considered technology up to TRL 5 by the end of the Definition Phase.

The following table defines the generic ESA TRL levels which have to be used in the frame of this

study.

ESA Technology Readiness Levels

Actual system "flight proven" through successful mission

erations

Actual system completed and "flight qualified" through

test and demonstration (ground or flight)

System prototype demonstration in a space environment

System/subsystem model or prototype demonstration in a

relevant environment (ground or space)

Component and/or breadboard validation in relevant

environment

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 7-1: Definition of ESA Technology Readiness Levels and TRL scale

8 LAUNCH SEGMENT REQUIREMENTS

For performances of the launch vehicle it is referred to [RD1].

R-LAS-010 The launch vehicle shall be Soyuz-Fregat 2-1b with the Fregat-MT upper stage.

R-LAS-020 A launch mass margin of 8% shall be considered (TBC).

Recall of [AD13].

Launch mass margin defined as ratio in percent between:

 Launcher performance for the target orbit ([RD1])

and

 Total spacecraft wet mass (incl. all other margins: maturity, system,

propellant, etc.) + launch adapter.

R-LAS-030 Deleted.

R-LAS-040 Launch site shall be CSG (Kourou, French Guyana).

R-LAS-050 Deleted.

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9 OPERATIONS AND MISSION ANALYSIS 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.

R-OPS-010 The jettisoning strategy of any element shall ensure collision avoidance with the

sampling spacecraft and the ERC or the asteroid at any stage of the mission with a

TBD margin.

E.g. Soyuz-Fregat stage, ERC, P/L cover, etc.

R-OPS-020 No critical S/C operations shall be performed if the Sun-Earth-S/C angle is lower

than 5

e.g. descent, sampling, local characterization, etc. Data downlink is not

considered to be a critical operation and therefore can be envisaged down to 3

R-OPS-030 No standard S/C operations shall be performed if the Sun-Earth-S/C angle is lower

than 2

R-OPS-040 The ERC shall be released by the sampling spacecraft from the return hyperbolic

trajectory and directly enter the Earth atmosphere.

R-OPS-050 The mission design shall cope with the minimum distances to the Sun during all

mission phases, i.e. coast and thrust arcs and asteroid proximity operations as

specified in [RD1].

R-OPS-060 The mission design shall cope with the maximum distances to the Sun during all

mission phases, i.e. coast and thrust arcs and asteroid proximity operations as

specified in [RD1].

R-OPS-070 The mission design shall cope with the maximum distances to Earth during all

mission phases, cruise and asteroid proximity operations as specified in [RD1].

R-OPS-080 The duration of a Solar conjunction or when the Sun-Earth-S/C angle is lower than

during the transfer to and from the asteroid and during proximity operations

shall be limited to 50 days.

R-OPS-090 Mission analysis shall ensure ERC re-entry velocity and flight path angle such that

heat fluxes during re-entry do not exceed 15 MW/m

(incl. margins as defined in

[AD13]) and total pressure at stagnation point does not exceed 80 kPa.

Re-entry conditions to be defined at 120 km altitude. To remain within a feasible

domain for plasma testing of thermal protection system.

G-OPS-100 Mission analysis should ensure a night re-entry of the ERC.

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

temperature increase after ERC landing.

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R-OPS-110 The duration of the proximity operations about the asteroid shall be at least 180

days.

Mitigation of risk, feasibility of science operations, maximizing rehearsals, etc.

R-OPS-120 The proximity operations shall allow 5 local characterization phases (LCP) before

the first sampling attempt to image the 5 candidate sampling sites at high

resolution and one LCP after a successful sampling attempt above the sampled site.

3 days (TBC) shall be allocated to each LCP.

R-OPS-130 The global characterization orbit or position shall be such that all surface features

(rocks, boulders, steep slopes, crater rims, etc.) larger than 1 meter (TBC) can be

identified and located with a 20 meter (TBC) absolute precision on the asteroid

global scale.

Global mapping of features (+ global shape model).

R-OPS-140 The local characterization campaign shall be such that all surface features larger

than 20 cm (TBC) can be identified and located within a 10 meter (TBC) absolute

precision on the 5 sampling site candidates.

Mapping of features on the sampling site candidates (+ local digital elevation

model for slope estimation).

R-OPS-150 A minimum of 3 sampling attempts shall be foreseen throughout the whole

proximity operations phase.

R-OPS-160 Proximity operations shall be planned so as to avoid collision with the primary

asteroid and its secondary in case of temporary failures of the S/C or the ground

segment.

See also, R-AUT-030 through R-AUT-070, R-PQA-080.

R-OPS-170 The proximity operations shall be such as to fulfil the resolution requirements listed

in [AD2].

R-OPS-180 At least 2 weeks shall be included at the end of the proximity operations for margin

purposes.

R-OPS-190 Deleted.

R-OPS-200 The minimum duration of the global characterization phase before sampling shall

be 35 days (TBC).

R-OPS-210 The proximity operations shall be compatible with a binary body as defined in R-

ENV-070/090/120/130/140/150.

G-OPS-220 Between two sampling attempts one week 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.

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R-OPS-230 During descent a minimum of 5 pictures of the surface shall be taken and shall be

stored in the spacecraft onboard memory for later transmission.

Complementary to minimum TM + outreach.

R-OPS-240 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 TBD.

R-OPS-250 The mission analysis and overall design shall enable asteroid surface touchdown

and sampling operations at any latitude and longitude.

R-OPS-260 The sampling operations on the asteroid shall occur on a fully illuminated location.

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

R-OPS-270 The sampling operations shall be based on a touch and go approach (a few seconds

touchdown, see definition of “touch and go” in 13.2).

R-OPS-280 The descent/touchdown operations shall be such as to prevent the creation of dust

before sampling operations.

R-OPS-290 Deleted.

R-OPS-300 Deleted.

R-OPS-310 The Earth Re-entry Capsule shall be released by the sampling spacecraft TBD hours

before Earth re-entry.

R-OPS-320 The measurements in the visible range shall be performed with a local solar

elevation angle between 30 and 60 degrees. Visible range instruments shall

nevertheless be on even during observations at other Sun aspect angles.

R-OPS-330 The radio science experiment shall be conducted from a “Radio-science” orbit and

allow to fulfil gravity-related science requirements defined in [AD2].

See definition of “Radio-science” orbit in chapter 13.2. Note that, if possible and of

compatible with the instrument measurement requirements, other science

measurements can be performed in parallel.

R-OPS-340 The selected sampling site shall be at least ~ 50 m diameter (TBC) (assuming that

the asteroid surface has at least one such a site).

See definition of sampling site in chapter 13.2. For such a large km-sized body, a

50 m area safe of hazards for a touch and go approach is very likely, see [RD13]

G-OPS-350 The selected sampling site should be free of features larger than TBD cm.

Sampling site minimizing risk. Operational requirement on the sampling site

selection

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10 SPACE SEGMENT REQUIREMENTS

10.1 Spacecraft system requirements

This chapter defines the general system requirements applicable to the sampling spacecraft and

the ERC or the space segment as a whole (general, margin, etc.). They are derived from the science

requirements, the mission safety, the general space design standards or more specific aspects of

the mission applicable at system-level.

R-SYS-010 The spacecraft lifetime shall be 8 years in space and 2 years 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-SYS-020 Deleted.

R-SYS-030 The space segment shall be sized for the mission scenario which is most demanding

(i.e. worst case ΔV, launch mass, thermal and sun illumination environment) as

defined in [RD1], covering the full launch window as defined in [AD13]/R-DV-4.

R-SYS-040 In this assessment study, the mission design shall follow the margin philosophy

defined in [AD13]. The margin philosophy and margin depletion scheme will be

firmly defined at a later stage.

R-SYS-050 The space segment shall cope with the integration procedures on a Soyuz-Fregat 2-1

b from CSG in Kourou.

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

R-SYS-060 Deleted.

R-SYS-070 The space segment shall be composed of 2 modules: Sampling spacecraft + Earth

Re-entry Capsule (ERC).

See Figure 13-1.

R-SYS-080 All spacecraft shall be designed according to [RD4], [RD5], [RD6] and [RD7] unless

otherwise specified in this document.

Any conflict between the requirements in this document and the ECSS standards

shall be reported to ESA.

R-SYS-090 The spacecraft system shall accommodate all necessary equipments and sub-

systems, including mass and volume margins as defined in [AD13] and compatible

with launch windows as defined in [RD1].

R-SYS-095 The maximum spacecraft dry mass, respectively wet mass shall be TBD kg, TBD kg.

R-SYS-100 The sampling spacecraft shall accommodate the science payload and enable

fulfilling all primary science measurements.

See [AD2] and payload as defined in [RD16].

R-SYS-110 A design-to-cost and risk minimization mission design approach shall be followed.

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R-SYS-120 The sampling spacecraft shall accommodate and transport the ERC throughout the

whole mission until release upon arrival at Earth.

G-SYS-130 The sampling spacecraft should take images of the separated ERC after release.

R-SYS-140 The design of the sampling spacecraft and ERC shall be robust against damage due

to debris and dust encountered during all proximity operations.

E.g. in-orbit dust/debris, touchdown/sampling on asteroid, etc.

R-SYS-150 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].

R-SYS-160 The main spacecraft shall allow for commissioning and calibration of the orbital

science instruments during the transfer to the NEA, including during fly-bys.

R-SYS-170 Deleted.

R-SYS-180 Deleted.

R-SYS-190 During the touch and go sampling operations, the sampling spacecraft shall

approach the surface with a maximum horizontal velocity of 5 cm.s

-1

(TBC) at

touchdown.

To limit the magnitude of forces on the touchdown/sampling system and to

mitigate toppling.

R-SYS-200 During the touch and go sampling operations, the sampling spacecraft shall

approach the surface with a maximum vertical velocity of 10 cm.s

-1

(TBC) at

touchdown.

R-SYS-210 The sampling spacecraft shall have a maximum attitude of +/-10

with respect to

the local surface at touchdown (TBC).

To mitigate toppling and to limit constraints on spacecraft geometry (e.g. solar

panels).

R-SYS-220 Deleted.

R-SYS-230 Deleted.

R-SYS-240 The touchdown accuracy of the sampling spacecraft with respect to the targeted

sampling site shall be better than 50 meters, 3-σ value (TBC).

R-SYS-250 The sampling spacecraft shall accommodate the ERC in a safe location such that the

risk of collision of the ERC Thermal Protection System with the asteroid surface is

minimized.

R-SYS-260 After touchdown the full set of telemetry data acquired during descent shall be

protected from overwriting until receipt on Earth is confirmed.

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R-SYS-270 The ERC shall be designed such as to limit the total heat flux to 15 MW/m

and

total pressure at stagnation point to 80 kPa.

To remain within a feasible domain for plasma testing of thermal protection

system. Heat fluxes including margins as per [AD13].

R-SYS-280 The ERC shall perform a fully passive re-entry (i.e. no parachute, no propulsion, no

active control).

R-SYS-290 Deleted.

R-SYS-300 The ERC and its sub-systems shall be designed such that the shock loads on the

sample are lower than 2000 g quasi-static load (TBC) for any angle of attack up to

20 degrees (TBC).

Limit adequate for the sample. See [AD2].

G-SYS-310 The ERC and its sub-systems should be designed such that the shock loads on the

sample are lower than 800 g quasi-static load (TBC) for any angle of attack up to 20

degrees (TBC).

R-SYS-320 The ERC shall allow simple access to and recovery of the sample within the capsule

without damaging the sample container seal.

R-SYS-330 The ERC shall be statically stable and its dynamic instabilities, if any, shall be such

that angle of attack variations are limited to 20 degrees (TBC) at all aerodynamic

regimes after the peak of dynamic pressure.

ERC dimensions, inertia, mass and CoM, CoP location. Before peak of dynamic

pressure, angle of attack variations are less critical.

R-SYS-340 The maximum deceleration of the ERC through re-entry shall be 80 g, quasi-static

load (TBC).

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

G-SYS-350 The ERC design should ensure that the sample is never exposed to temperatures

higher than +40

C. For less than one minute, +80

C is acceptable.

See [AD2].

G-SYS-360 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. If implemented, the maximum allowable shocks

on the memory shall be defined and included in R-SYS-300.

R-SYS-370 The sample sealing system shall prevent:

 Less than 0.1 ppm of terrestrial water,

 Dust or liquid particles larger than 0.3 μm,

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to contaminate the sample up to storage of the capsule in the suitable sample

curation facility after return.

See [AD2].

R-SYS-380 The sample sealing system shall withstand a shock load up to 2000 g quasi-static

load (TBC).

See R-SYS-300 and [AD2].

G-SYS-390 The sample sealing system should withstand a shock load up to 800 g quasi-static

load (TBC).

See G-SYS-310 and [AD2].

R-SYS-400 Deleted.

R-SYS-410 The spacecraft design and operations shall transmit back to Earth a minimum total

payload data volume of 122.35 GBit (TBC) before the end of the asteroid proximity

operations.

Without any margins ([AD13] guidelines to be implemented), compressed data

volume.

R-SYS-420 110.8 Gbit (TBC) of science data shall be transmitted back to Earth 2 weeks before

the start of sampling operations (SAM).

Without any margins ([AD13] guidelines to be implemented) , compressed data

volume. Includes data from all phases, including GCP and LCP.

R-SYS-430 TBD Gbit of science data shall be transmitted back to Earth 2 weeks before the start

of LCP operations.

10.2 Platform sub-system requirements

This chapter defines all key S/C sub-system requirements (pointing, thermal, power, data

handling, telecommunications, etc.). They are derived from the system requirements or are

requirements specific to a sub-system derived from space design standards or mission-specific

need.

10.2.1 Thermal requirements

R-THE-010 The spacecraft thermal control system shall cope with the thermal needs of the

various spacecraft sub-systems as required, including P/L, at any stage of the

mission as a function of the spacecraft thermal modes, including safe mode.

R-THE-020 The spacecraft thermal control system shall cope with the space environment

throughout the mission (incl. asteroid surface operations).

See [AD4]. Specific attention to be given to asteroid IR flux and closest distance to

the Sun during transfer.

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R-THE-030 Sufficient telemetry and housekeeping information shall be provided for

monitoring of thermal equipment (e.g. radiators, heaters, etc.).

Lessons learns from Rosetta (e.g. need to thoroughly monitor the effect of

pointing/solar aspect angle for the radiators).

R-THE-040 The ERC heat shield thermal protection system material shall be able to withstand

entry conditions defined in R-SYS-270.

Qualification margins for the TPS material are TBD.

10.2.2 Power requirements

R-POW-010 The spacecraft power system shall be made of solar arrays and batteries and shall

cope with the power needs of the various spacecraft sub-systems as required,

including P/L, at any stage of the mission as a function of the spacecraft power

modes, including safe mode.

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

R-POW-020 The spacecraft battery shall be sized for worst case eclipses and the

descent/touchdown/re-ascent phase.

Eclipses due to the asteroid will be avoided for nominal operations but may be

encountered if control of the spacecraft is lost or Earth eclipses during LEOP.

R-POW-030 The sizing of the solar arrays shall allow the S/C to stay on a ”Radio-science” orbit

and safe position as defined in 13.2.

To avoid asteroid escape via solar radiation pressure.

R-POW-040 Deleted.

R-POW-050 The electrical design shall comply with the requirements of [RD6]. Tailoring of

these requirements may be proposed and need to be justified.

10.2.3 GNC/AOCS requirements

R-GNC-010 Deleted.

R-GNC-020 Deleted.

R-GNC-030 Deleted.

R-GNC-040 Asteroid acquisition (i.e. identification of the target) shall be performed one week

(TBC) before the start of proximity operations.

Start of proximity operations defined in chapter 13.2.

R-GNC-050 During sampling operations, the AOCS/GNC system shall prevent total attitude

excursions larger than 30 degrees (TBC) (with respect to the local vertical) and

attitude rates larger than 5 deg/s (TBC).

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For instance to avoid the solar panels to hit the surface and further re-ascent with

a safe attitude.

R-GNC-060 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, a survivable thermal environment and ground communication

for vital equipment.

R-GNC-070 In the event of unavailability of the star trackers, the AOCS/GNC shall have the

capability to maintain and propagate attitude estimation whilst meeting the

relevant pointing requirements.

R-GNC-080 Deleted.

R-GNC-090 Deleted.

G-GNC-100 Asteroid acquisition (i.e. identification of the target) should be performed two

weeks (TBC) before the start of proximity operations..

R-GNC-040 guarantees a minimum detection time.

R-GNC-110 After separation from the launcher upper stage, the AOCS shall:

 Damp out the residual angular rates,

 Bring the S/C into a power safe pointing attitude within a time compatible

with the S/C internal electrical energy capability,

 Maintain the S/C attitude to allow communications with the ground as

planned for LEOP operations.

R-GNC-120 Deleted.

10.2.4 DHS requirements

R-DHS-010 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-DHS-020 The DHS shall be able to receive TC and send TM from/to ground at the same time.

R-DHS-030 The DHS shall be able to command the instruments and equipment onboard.

R-DHS-040 The DHS shall provide reconfiguration capabilities in case of failure detection.

R-DHS-050 The DHS shall manage the redundancy for the relevant sub-systems.

R-DHS-060 During touchdown/sampling operations, the DHS shall provide the capability to

store the scientific and housekeeping data generated by the complete close-up

camera and sampling operations cycle.

R-DHS-070 The DHS system shall support uplink and downlink file transfer.

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R-DHS-080 The DHS shall provide the storage capability such that all science data can be

downlinked (R-SYS-410/420/430) during proximity operations and assuming 2

consecutive downlink windows (i.e. 2 consecutive daily windows) are missed.

R-DHS-090 The DHS system shall be compatible with the maximum data rates of each

instrument as specified in [RD16].

R-DHS-100 The DHS system shall implement system-level failure FDIR of the P/L and

platform.

10.2.5 Propulsion requirements

R-PRO-010 Deleted.

R-PRO-020 The propulsion system shall cope with all operations and associated Delta-V/thrust

requirements incl.: transfer (MCC, DSM, etc.), asteroid proximity operations, initial

fly-bys, gravity field campaign, controlled orbits, local characterization phase,

sampling rehearsals, descent/touchdown and re-ascent phases, etc.

R-PRO-030 The by-products of the propulsion system shall be characterized and reported at

ppb level prior to the mission.

R-PRO-040 The propulsion sub-system shall be designed to be compatible with any operational

S/C attitude. The main engine and any reaction control thruster shall be thermally

qualified for such environmental conditions.

Particularly for the asteroid surface environment.

R-PRO-050 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 10

C above the maximum freezing point of the onboard

propellant.

R-PRO-060 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.

10.2.6 Communications requirements

R-COM-010 Deleted.

R-COM-020 Real-time data shall be provided directly to Earth during descent and sampling

allowing monitoring of the major events. A data rate of 100 bit/s shall be possible

(TBC).

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

descent. 100 bit/s in line with MER data.

R-COM-030 The mission design shall comply with ESA ECSS telecommunication standards

([RD14]).

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R-COM-040 The communication system shall support the two-way Ranging and Doppler

measurements of the S/C throughout all mission phases and ΔDOR if high-

precision navigation is required (e.g. RSE campaign) TBC.

R-COM-050 The link budgets of the spacecraft to ground shall be calculated for a weather

availability of 95%.

R-COM-060 Science data shall be downlinked by the spacecraft in X-band.

R-COM-070 The maximum bit error rate during data downlink shall be better than 10

-5

R-COM-080 Deleted.

R-COM-090 The telecommunication system shall be capable of simultaneously handling

telemetry, ranging and telecommands.

R-COM-100 Deleted.

R-COM-110 The telecommunication equipment shall support the RSE as specified in [RD16].

If Ka-band is required for science, this shall be considered as a science instrument.

R-COM-120 The telecommunications system shall be able to downlink all science data as per R-

SYS-410/420/430.

R-COM-130 All images taken by navigation cameras and required to be sent to ground (e.g.

asteroid shape model, local slopes around sampling sites, etc.), if any, shall be

downloaded.

10.2.7 Structure requirements

R-STR-010 The spacecraft structure shall support the launch environment [AD8] with MOS as

defined in [RD5] which will be agreed with the agency.

R-STR-020 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 on the

asteroid).

R-STR-030 The structure directly containing the sample (sample container) shall be able to

withstand the loads (i.e. still fulfilling R-SYS-380) as defined in R-SYS-300.

R-STR-040 The structural stiffness shall guarantee fundamental frequencies of the S/C within

the requirements of the LV to avoid dynamic coupling.

R-STR-050 The structural design shall be based on simple load paths.

R-STR-060 The structural design shall provide a minimum margin of 15% over the minimum

frequencies specified by the Launcher User Manual [AD8] before verification of S/C

dynamic properties by test.

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10.2.8 Mechanism requirements

R-MEC-010 The sampling mechanism shall have the capability to collect hundreds of grams of

material.

See [AD2].

R-MEC-020 The sampling mechanism shall have the capability to collect cm-sized fragments,

plus a large amount (minimum several grams) of small particles (100s of μm-sized

to mms-sized).

See [AD2].

R-MEC-030 The sampling mechanism shall be compatible with asteroid material having the

properties defined in R-ENV-100 and [AD2].

R-MEC-040 The sampling device (or any other part of the spacecraft) shall be equipped with

autonomous means of verifying suitable sampling with a 95% probability (TBC).

Determine whether the sampling operation was successful or not.

G-MEC-050 The sampling device (or any other part of the spacecraft) should have the capability

to provide a rough estimate of the collected sample mass and volume.

R-MEC-060 All mechanisms involved in the SATCS chain shall be cleaned so as to fulfil levels

specified in R-CON-010/R-SYS-370 and according to the standards defined in

R-CON-040.

R-MEC-070 Deleted.

R-MEC-080 All mechanisms shall be protected against debris generated by touchdown and

sampling on the asteroid.

R-MEC-090 The SATCS system shall be able to monitor its torque, thrust and penetration depth

(if applicable).

R-MEC-100 It shall be possible to detect blockage of a sampling unit and release it before re-

ascent in case it gets stuck during the touchdown/sampling operations.

R-MEC-110 Deleted.

G-MEC-120 The design of the sample collection and distribution unit (incl. container) should

ensure that the sample, after collection, is never exposed to temperatures higher

than +40

C. For less than one minute +80

C is acceptable.

Sample temperature and collection process. See [AD2].

G-MEC-130 Through the monitoring of its various sensors’ information as a function of

penetration progress, the SATCS should support the determination/estimation of

the soil’s mechanical properties.

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

etc. To be coupled with R-MEC-090.

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R-MEC-140 The design of the touchdown mechanism system as well as the sampling

mechanism shall be such that total attitude excursions are lower than 30 degrees

(TBC) (with respect to the local vertical) and attitude rates lower than 5 deg/s

(TBC).

See R-GNC-050.

10.3 Autonomy and FDIR requirements

R-AUT-010 Operations of the spacecraft shall be possible during all mission phases via the

execution of pre-programmed sequence, or via onboard autonomy during the

descent and sampling phase. Level of autonomy and detailed spacecraft operations

throughout the various spacecraft operations are TBD.

R-AUT-020 S/C autonomy shall ensure achievement of all mission goals.

R-AUT-030 During asteroid proximity operations, if a ground TC is expected for the S/C to

proceed with its nominal manoeuvre and this TC is not received within a TBD time,

the S/C shall be able to perform a valid transition into an asteroid Collision

Avoidance Mode.

R-AUT-040 During descent/sampling operations on the asteroid, in case of any failure or non-

nominal S/C equipment TM (sub-system or system) the spacecraft shall

autonomously be able to perform a valid transition into an asteroid Collision

Avoidance Mode.

R-AUT-050 After TBD minutes in Collision Avoidance Mode and no ground feedback the

spacecraft shall be able to perform a valid transition into Safe Mode.

R-AUT-060 The S/C shall be able to receive a TC from the ground to switch from Collision

Avoidance Mode to Safe Mode.

R-AUT-070 In any operations other than descent/sampling, in case of major system failure (or

non-recoverable single failure) the spacecraft shall be able to survive in Safe Mode

without the need for ground contact in:

 A “safe position” if during asteroid proximity operations phase,

 On its transfer trajectory in case of LEOP or interplanetary cruise phase.

R-AUT-080 The spacecraft operating modes shall include a system-level Safe Mode that ensures

the following properties are fulfilled:

 Uninterrupted power supply,

 Safe thermal conditions,

 Continuous communication with ground,

 Predictable configuration minimising the onboard power demand and data

traffic.

R-AUT-090 The spacecraft operating modes shall include a system-level Collision Avoidance

Mode that ensures the following properties are fulfilled:

 Autonomously go into a “safe position”,

 Guarantee of no collision with the asteroid before switching to Safe Mode,

 Uninterrupted power supply,

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 Safe thermal conditions,

 Predictable configuration minimising the onboard power demand and data

traffic.

R-AUT-100 When in Safe Mode the spacecraft shall have the capability to transmit a predefined

set of data parameters at a rate of up to TBD b.s

-1

R-AUT-110 The spacecraft shall be able to perform a valid mode transition upon receipt of a

telecommand at all times and in all spacecraft modes.

R-AUT-120 The fault management systems shall be intrinsically fail-safe.

R-AUT-130 The spacecraft shall function autonomously throughout any solar conjunctions on

cruise to and back from the asteroid.

R-AUT-140 The system (or FDIR) shall provide the capability for automatic detection, isolation

and recovery of any credible system or subsystem anomaly.

R-AUT-150 Sampling site targeting and last go/no-go decision during actual descent and

sampling shall be ground-based. Autonomous sampling site re-targeting shall not

be allowed (see R-AUT-040 instead).

10.4 Environmental requirements

The following defines the environment-related requirements that the spacecraft has to comply

with throughout the mission, whether that is the launch, interplanetary, spacecraft, asteroid or

Earth re-entry environment.

R-ENV-010 The spacecraft composite and all its sub-systems, including payload, shall be

compatible with the launch environment as defined in ([AD8]).

R-ENV-020 The mission design shall be compliant with the general space environment (cruise)

requirements defined in [RD4] and the mission-specific (e.g. asteroid) environment

requirements defined in [AD4].

R-ENV-030 Deleted.

R-ENV-040 The sample shall never be exposed to magnetic field larger than 200 µT.

R-ENV-050 The spacecraft shall satisfy the electric and magnetic cleanliness (EMC)

requirements specified by the payload (TBD reference).

R-ENV-060 Deleted.

R-ENV-070 The maximum surface temperature of the sampling site to be considered for the

design of the spacecraft shall be 500 K (TBC).

G-ENV-080 Deleted.

R-ENV-090 The mission design shall be compatible with the rotation period of the primary

asteroid of ~ 3.6 hours +/- 0.1%.

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R-ENV-100 The surface properties which shall be assumed for the design of the sampling

mechanism are as follows:

 No solid rock. Only loose material shall be assumed,

 Grain Size: µm

 up to 3 cm,

 Shape: any (rounded, tabular, elongated, etc.),

 Cohesion: 0.1 – 5 kPa,

 Compressive strength: Up to 30 MPa,

 Bulk density: <1 – 2.2 g.cm

-3

 Angle of internal friction: 20 - 40.

R-ENV-110 Deleted.

R-ENV-120 The asteroid body density shall be assumed to be 1300 ± 600 kg.m

-3

. This also

applies to the secondary body.

See [AD4].

R-ENV-130 A diameter of 1550 m – 2110 m shall be assumed for the primary body.

See [AD4].

R-ENV-140 A diameter of 430 m – 590 m shall be assumed for the secondary body.

See [AD4].

R-ENV-150 A binary with the following properties shall be assumed:

 Secondary orbital semi-major axis: 2.8 + 1.7/-0.7 km,

 Secondary orbital eccentricity: 0.05 ± 0.05,

 Secondary orbital period around primary: 16.15 ± 0.02 hours.

See [AD4].

10.5 AIV and testing requirements

R-AIV-010 The AIV activities of the S/C shall be compliant with the integration requirements

of a Soyuz-Fregat 2-1 b in CSG Kourou.

R-AIV-020 The AIV facilities shall enable the implementation of sterilization and cleaning

procedures as defined in R-CON-040.

10.6 Ground segment

R-GRS-010 The mission shall use and be compatible with the standards of the ESA deep space

network as well as the NASA deep space network.

R-GRS-020 The MarcoPolo-R mission shall be operated by ESA/ESOC.

R-GRS-030 During LEOP, TBD ESA ground stations shall be used for contact with the

spacecraft.

LEOP is defined as the first 72 hours (TBC) after the separation of the spacecraft

from the Fregat upper stage.

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R-GRS-040 After LEOP, the ESA station at TBD location shall be the primary station used for

contact with the spacecraft, with TBD location as backup.

R-GRS-050 The ground segment shall provide for a 24 hour coverage capability for asteroid

descent, touchdown, sampling and local characterization operations.

R-GRS-060 The ground segment shall cope with the data volume defined in R-SYS-

410/420/430.

R-GRS-070 It shall be possible to retrieve the ERC within 4 hours (TBC) after landing on Earth

without the aid of tracking devices onboard the ERC.

Also to be taken into account for ERC landing site selection.

R-GRS-080 It shall be possible to track the ERC re-entry during the whole re-entry operation.

G-GRS-090 It shall be possible to retrieve the ERC within 2 hours (TBC) after landing on Earth

without the aid of tracking devices onboard the ERC.

To limit thermal excursion on the sample.

R-GRS-100 The ground segment shall support the on-orbit calibration of the satellite

autonomous pointing capabilities.

R-GRS-110 The ground segment shall elaborate an operational concept based on multiple

uplink and downlink attempts and a respective timing for the preparation of the

orbit timeline to enable standard operations under superior solar conjunction

conditions with a Sun-Earth-Spacecraft angle as low as 2° (TBC). The required

success rate sets of commands (number of commands in set: 50 TBC) and for

telemetry packets is TBD.

Probability to be calculated for space weather conditions typical for the solar

cycle at the time of the conjunction.

R-GRS-120 The ground segment shall be able to support at least 4 (including rehearsals)

touchdown and sampling attempts during the asteroid proximity phase.

R-GRS-130 The ground segment shall be able to support at least the 5 local characterization

phases before sampling and the last one after sampling (above the sampled site).

R-GRS-140 The ground segment shall provide for a flexible re-planning capability during the

asteroid phase. The manoeuvre planning for new asteroid proximity orbits and new

sampling site targets shall be available within 4 days (TBC) after reception of the

respective request.

R-GRS-150 The ground segment shall construct a global asteroid map and a global Digital

Elevation Model with a 1-m spatial resolution (TBC) of the whole asteroid during

the GCP phase primarily based on the navigation sensor information.

See R-OPS-130.

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R-GRS-160 The ground segment shall construct a local map and a local Digital Elevation Model

with a 20-cm spatial resolution (TBC) of the sampling site candidates during the

LCP phase primarily based on the navigation sensor information.

See R-OPS-140.

R-GRS-170 The ground segment shall be able to obtain the position information relative to the

targeted sampling site sent by the spacecraft as a package of “images” and AOCS

housekeeping data, process this data and upload an appropriate “guidance”

telecommand within one hour (TBC) of reception of the images in order to either

reach the sampling site or to abort the descent.

10.7 Product assurance and RAMS requirements

R-PQA-010 Deleted.

R-PQA-020 Deleted.

R-PQA-030 The mission design shall be free of Single Point Failures, shall be one-failure

tolerant and shall ensure avoidance of failure propagation. .

R-PQA-040 Instrument failures shall not cause failure or damage to other instruments or other

spacecraft subsystems.

R-PQA-050 Instrument failures shall not cause the spacecraft to enter a Collision Avoidance or

Safe Mode.

R-PQA-060 Deleted.

R-PQA-070 Where redundancy is employed, the design shall allow operation and verification of

the redundant item/function independent of nominal use.

R-PQA-080 In case of anomalies or failures from which the spacecraft cannot recover

autonomously, the spacecraft shall be able to autonomously perform a valid

transition into:

 Collision Avoidance Mode and later on Safe Mode if during

descent/sampling operation,

 Safe Mode if during any other phase than descent/sampling operations.

R-PQA-090 The overall mission reliability shall be better than 90% (TBC).

R-PQA-100 The design philosophy for all mechanisms shall be as defined in [RD5].

R-PQA-110 The risk of casualty resulting from the ERC re-entry shall be 1:10000 (TBC).

R-PQA-120 Deleted.

R-PQA-130 The communication system shall implement full redundancy except for the main

antenna dishes (e.g. HGA).

R-PQA-140 A highly reliable backup sampling method shall be implemented to ensure a sample

can be collected should the main sampling mechanism fail.

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For the backup sampling method, degraded amount of mass (i.e. TBD grams) are

acceptable

11 COORDINATE SYSTEM

R-COR-010 For the description of the MarcoPolo-R spacecraft and of its major components,

the following S/C Reference Coordinate Frame shall be used:

Figure 11-1: Main spacecraft reference frame

R-COR-020 For the definition of the MarcoPolo-R spacecraft attitude along its orbit, the

following Local Orbital Reference Frame shall be used:

Figure 11-2: Local orbital reference frame

is oriented towards nadir and +X

is oriented along the velocity vector of the spacecraft.

+ Z

+ X

Solar arrays

+ Z

+ X

steroid observing instruments

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12 APPLICABLE AND REFERENCE DOCUMENTS

The Applicable Documents (AD’s) listed below shall be complied with, unless conflicting with the

mission MSRD itself or where specifically stated. Their contents shall be considered as parts of the

requirements of this document. A conflict between these documents and the MSRD shall be

brought to the attention of ESA for clarification.

The published ECSS (European Cooperation for Space Standardisation) space standards

documents quoted in the MSRD shall be used as design references throughout the MarcoPolo-R

assessment phase and are available on the Internet at: www.ecss.nl.

12.1 Applicable documents

[AD1]

Deleted

[AD2]

MarcoPolo-R Science Requirements Document, [MPR-RSSD-RS-001], issue

1.4, Document providing Science objectives and science requirements, including

measurement specifications

[AD3]

Deleted

[AD4]

MarcoPolo-R Environment Document, [SRE-PA/2011.077], issue 1.2

[AD5]

MarcoPolo-R – Assessment of Planetary Protection issues, [SRE-

PA/2011.078], issue 1.0, Document providing recommendations on Planetary

Protection issues relating to the potential NEA targets

[AD6]

MarcoPolo-R Consolidated Report on Mission Analysis, [Reference TBD],

issue 1.0, ESOC, document describing the detailed reference transfers and orbits as

well as navigation analysis to be used in task 2. Note: will be provided by ESA in the

course of task 2

[AD7]

Deleted

[AD8]

Soyuz-Fregat 2-1b from the Guiana Space Centre User’s Manual, issue

2.0, March 2012. Note: The launch vehicle performances to be used in this study are

defined in [AD6] (or [RD1] if [AD6] is not available)

[AD9]

Deleted

[AD10]

Technology Readiness levels handbook, [TEC-SHS/5551/MG/ap], issue 1.6,

September 2008

[AD11]

Deleted

[AD12]

Deleted

[AD13]

MarcoPolo-R margin philosophy document, v1.1, [SRE-F/2012.088]

[AD14]

MarcoPolo-R Mission Operations Assumptions Document, [TBD]

[AD15]

MarcoPolo-R Science Operations Assumptions Document, [TBD]

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12.2 Reference document

The Reference Documents (RD’s) listed below are given as complementary information and

background data related to the MarcoPolo-R Mission or the ESA standards.

[RD1]

MarcoPolo-R Mission Analysis Guidelines (MAG), SRE-F/2012.041, issue

2.0

[RD2]

MarcoPolo-R CDF IFP material

[RD3]

MarcoPolo-R CDF study report, CDF-123(A)

[RD4]

ECSS-E-ST-10 series, available from http://www.ecss.nl

[RD5]

ECSS-E-ST-30, 31, 32, 33 and 35 series, available from http://www.ecss.nl

[RD6]

ECSS-E-ST-20 series, available from http://www.ecss.nl

[RD7]

ECSS-M-ST-10, available from http://www.ecss.nl

[RD8]

MarcoPolo-R: NEA Sample Return Mission proposal, M.A. Barucci et al.

[RD9]

Marco Polo Assessment Study Report, ESA/SRE(2009)3, Dec. 2009

[RD10]

The Marco Polo executive summary report, MP.ASU.ES, issue 1

[RD11]

Marco Polo executive summary, MPL-OHB-TN-014, issue 1

[RD12]

Marco Polo mission executive summary, SD-RP-AI-0673, issue 1

[RD13]

Marco Polo – Technical and programmatic review report, SRE-

PA/2009.069/Marco-Polo, issue 1.0

[RD14]

ECSS-E-ST-50 series, available from http://www.ecss.nl

[RD15]

MarcoPolo-R cost estimate template, Excel file

[RD16]

MarcoPolo-R Payload Resources and Requirements Document, [SRE-

F/2012.089], issue 1.3, July 2012

[RD17]

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'

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13 ACRONYMS AND TERMINOLOGY

13.1 Acronyms

AD Applicable Document

AIV (AV) Assembly, Integration and Verification

AME Absolute Measurement Error

AO Announcement of Opportunity

AOCS Attitude and Orbit Control System

APE Absolute Pointing Error

AUT Autonomy requirement

BSR Baseline Selection Review

CAM Collision Avoidance Mode

CDF Concurrent Design Facility

CoG Centre of Gravity

CoM Centre of Mass

COM Communication Requirement

CoP Centre of Pressure

COTS Commercial Off-The-Shelf

CSG Centre Spatial Guyanais (Kourou, French Guyana)

CuC Close-up Camera

DEM Digital Elevation Model

DHS Data Handling System

DOR Differential One-way Ranging

DSM Deep Space Manoeuvre

ECSS European Cooperation for Space Standardisation

ENV Environment requirement

ERC Earth Re-entry Capsule

ESA European Space Agency

ESTEC European Space research and TEchnology Center

FCP Far global Characterization Phase

FEM Finite Element Model

FMECA Failure Modes, Effects and Criticality Analysis

GA Gravity Assist

GCP Global Characterization Phase

GNC Guidance, Navigation and Control

GRS Ground Segment Requirement

HGA High Gain Antenna

IR Infrared

ITT Invitation To Tender

LAS Launch Segment Requirement

LCP Local Characterization Phase

LEOP Low Earth Orbit Phase

MCC Mid-Course Correction

M-class Medium Class

MCR Mission Consolidate Review

MEC Mechanism Requirement

MIS Mission Requirement

MOS Margin Of Safety

MSRD Mission Requirements Document

NEA Near Earth Asteroid (also sometimes referred to as Near-Earth Objects)

NPA Neutral Particle Analyser

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OPS Operational requirement

P/L Payload

PLD Payload Requirement

POW Power Requirement

PP Planetary Protection

PPR Planetary protection Requirement

ppb particle per billion (10

-9

)

ppm particle per million

PRM Programmatic Requirement

PRO Propulsion System Requirement

PQA Product/Quality Assurance Requirement

RAMS Reliability, Availability, Maintainability and Safety

RD Reference Document

RF Radio Frequency

RPE Relative Pointing Error

RSE Radio Science Experiment

RSP Radio Science Phase

SAM Descent/Sampling Phase

SATCS Sample Acquisition, Transfer and Containment System

SRE-F Advanced Studies and Technology Preparation Office

S/C Spacecraft

SCI Science Requirement

STD Study requirement

SEP Solar Electric Propulsion

SPF Single Point Failure

SRP Solar Radiation Pressure

SS Sub-System

SSAC Space Science Advisory Committee

STD Study Requirement

STR Star Tracker

STR Structure and Mechanical Requirement

SYS System Requirement

TBC To Be Confirmed

TBD To Be Determined

THE Thermal Requirement

TPS Thermal Protection System

TRL Technology Readiness Level

TRS Technology Reference Study

VIS Visible

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13.2 Terminology

The following spacecraft architecture was selected at the Baseline Selection Review of the

assessment phase in June 2012:

Figure 13-1: Baseline spacecraft architecture (note that the sealing function of the SATCS

might also be performed within the ERC)

The role of each element is briefly described below.

 Spacecraft Composite: The combination of the sampling spacecraft and the Earth re-entry

capsule, stacked together from launch through asteroid surface and proximity operations and

cruises phases up until the release of the ERC. Note: the Spacecraft Composite does not

include the launch vehicle adaptor.

 Sampling Spacecraft: Module used for outbound and inbound transfer to and from the

asteroid, orbital operations around the asteroid, descent, sampling and re-ascent operations to

and from the asteroid surface as well as release of the Earth re-entry capsule. Sometimes also

referred to as the main spacecraft in the assessment phase. Note: the Sampling Spacecraft does

include the spacecraft mechanical interface to the LV adaptor (but not the adaptor itself).

 Earth Re-Entry Capsule (ERC): The capsule which performs the re-entry through the Earth

atmosphere. The retrieved sample container is located inside the ERC.

In addition, the following definitions/clarifications apply:

 Sample Acquisition, Transfer and Containment System (SATCS): 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).

 Touch And Go: Sampling strategy defined by the following characteristics:

- Sampling operations during GNC guided impact,

- Surface contact by sampling mechanism only and any possible skirt around it to

mitigate dust and debris impact,

- The impact energy may or not be used for sampling/ascent,

- The sampling contact with the asteroid is in the order of seconds (e.g. maximum 5-10

seconds).

 Payload: The payload is defined as the ensemble of physical element from the input signal

collection up to the digital electronic signal production.

 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.

Spacecraft

Composite

Sampling

Spacecraft

ERC

SATCS

Payload

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 Absolute measurement error: Angular separation between the actual and measured

generalised pointing vectors of the S/C.

 Horizontal or Lateral touchdown velocity: The component of the touchdown velocity parallel

to the average local slope at the asteroid sampling site.

 Vertical touchdown velocity: The component of the touchdown velocity perpendicular to the

average local slope at the asteroid sampling site.

 Sampling site (candidate): Safe area operationally (on an illuminated part such as to enable

vision-based navigation and with hazards compatible with the S/C capability) and

scientifically suitable for sample collection, defined on the asteroid surface by the ground

teams.

 Touchdown accuracy: Downrange major axis of the touchdown ellipse.

 Touchdown ellipse: Ellipse centred on the sampling site which envelops all possible

touchdown 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.

 Proximity operations: The asteroid proximity operations are arbitrarily defined at the start of

the close asteroid approach, i.e. at a distance of 500 km from the asteroid and end when the

SEP system is switched on and the S/C leaves the asteroid system. See Figure 2-2.

 Radio-science orbit: In the context of MarcoPolo-R, a Radio-science orbit is an asteroid-bound

orbit, favourable to radio science measurements, where the S/C does not require any orbit and

attitude manoeuvres in order to remain above 3 km distance to the surface at any time, during

four consecutive communication windows of 8 hours.

 Safe position: In the context of MarcoPolo-R, a safe position is an eclipse-free location where

the spacecraft does not require any orbit and attitude manoeuvres during at least one week in

order to remain at least 1 km distance away from the surface of the asteroid surface (and its

secondary if applicable). Such a position can be either an asteroid-bound orbit or formation

flying point or a heliocentric orbit, out of the asteroid influence (however close enough to allow

return to asteroid in limited time and manoeuvre budget).

 Failure: The termination of the ability of an item to perform a required function (e.g. solar

arrays not producing electric current following string failure, sampling mechanism stuck and

not able to perform sampling).

14 LIST OF TBC AND TBD

14.1 List of TBC

To be written

Table 14-1: List of TBC

14.2 List of TBD

To be written

Table 14-2: List of TBD