sol-s-astr-tn-0014 solar orbiteremits.sso.esa.int/emits-doc/estec/ao6309_ad3.pdf ·...

Solar Orbiter SOL-S-ASTR-TN-0014

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Company Registration No. 2449259 Registered Office: Gunnels Wood Road, Stevenage, Hertfordshire, SG1 2AS, UK

SOL-S-ASTR-TN-0014 - Solar Orbiter System Design Report - Issue 2.doc

Solar Orbiter System Design Report

CI CODE:

UK EXPORT CONTROL RATING: 9E001/9A004 Rated By: P. D’arrigo

Prepared by:

Date: 28/07/2009

Mark Ayre

Checked by:

Date: 28/07/2009

Paolo D’Arrigo

Approved by:

Date: 28/07/2009

Paolo D’Arrigo

Authorised by:

Date: 28/07/2009

Ivan Ferrario

ESA export licence exception applies, reference HM Customs Tariff Vol 1 Part 4 Para 4.3.11

© Astrium Limited 2009

Astrium Limited owns the copyright of this document which is supplied in confidence and which shall not be used for any purpose other than that for which it is supplied and shall not in whole or in part be reproduced, copied, or communicated

to any person without written permission from the owner.

Astrium Limited Gunnels Wood Road, Stevenage, Hertfordshire, SG1 2AS, England


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EADS Astrium Limited owns the copyright of this document which is supplied in confidence and which shall not be used for any purpose other than that for which it is supplied and shall not in whole or in part be reproduced, copied, or communicated to any person without written permission from the owner.


INTENTIONALLY BLANK


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DOCUMENT CHANGE DETAILS

ISSUE CHANGE AUTHORITY CLASS RELEVANT INFORMATION/INSTRUCTIONS

1 - - Initial Issue

2 - - Updated to include DSR comments

DISTRIBUTION LIST

INTERNAL

EXTERNAL

Paolo D’Arrigo Philippe Kletzkine

Ivan Ferrario Danielle Renton

Eberhard Schutte

Configuration Management

Library


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CONTENTS

SOLAR ORBITER SYSTEM DESIGN REPORT ..............................................................................................1

INTENTIONALLY BLANK..................................................................................................................................2

DOCUMENT CHANGE DETAILS.....................................................................................................................3

CONTENTS .......................................................................................................................................................4

TABLES ............................................................................................................................................................8

1. APPLICABLE AND REFERENCE DOCUMENTS ......................................................................................14 1.1 Reference Documents..........................................................................................................................14 1.2 Normative Applicable Documents ........................................................................................................15

2. SOLAR ORBITER MISSION REQUIIREMENTS........................................................................................16 2.1 Solar Orbiter Mission Requirements ....................................................................................................16 2.2 Lifetime Requirements .........................................................................................................................16 2.3 Space Segment Requirements: Performances....................................................................................17 2.4 Space Segment Requirements: Pointing .............................................................................................21 2.5 Communication Requirements .............................................................................................................23 2.6 Functional Requirements .....................................................................................................................24 2.7 Cleanliness Requirements ...................................................................................................................25 2.8 Launch Requirements ..........................................................................................................................25 2.9 Operations Requirements ....................................................................................................................26 2.10 Ground Station Requirements............................................................................................................26 2.11 Ground Operations Environment........................................................................................................27 2.12 Launch Environment ..........................................................................................................................27 2.13 General System Engineering Requirements......................................................................................27 2.14 Coordinates ........................................................................................................................................28 2.15 Mass Margins .....................................................................................................................................29 2.16 Propulsion Margins.............................................................................................................................30 2.17 Power Margins....................................................................................................................................30 2.18 Data Margins ......................................................................................................................................31 2.19 Communications Margins...................................................................................................................31 2.20 Thermal Margins.................................................................................................................................32 2.21 Heatshield...........................................................................................................................................32 2.22 Software .............................................................................................................................................32 2.23 Instruments.........................................................................................................................................32

3. SOLAR ORBITER SPACECRAFT DESIGN OVERVIEW...........................................................................35 3.1 Design Heritage....................................................................................................................................35 3.2 Design Overview ..................................................................................................................................37 3.3 BepiColombo Technology Reuse.........................................................................................................38

4. SOLAR ORBITER CONFIGURATION ........................................................................................................41 4.1 Introduction...........................................................................................................................................41 4.2 Configurations versus Mission Timeline...............................................................................................42 4.3 Payload Accommodation......................................................................................................................48

4.3.1 Remote Sensing Instruments.....................................................................................................49 4.3.2 In-Situ Instruments .....................................................................................................................51

4.4 Subsystem Accommodation.................................................................................................................57


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5. MISSION ANALYSIS...................................................................................................................................66 5.1 Mission Overview .................................................................................................................................66 5.2 Mission Baseline...................................................................................................................................66

5.2.1 2017 Mission Scenario...............................................................................................................67 5.2.2 2018 Mission Scenario...............................................................................................................69

5.3 Delta V Requirements ..........................................................................................................................74 5.3.1 Launcher Error Correction Manoeuvre.......................................................................................74 5.3.2 Launch Window DSM Penalty....................................................................................................74 5.3.3 Gravity Assist Manoeuvres ........................................................................................................74 5.3.4 Nominal Delta V Budget.............................................................................................................75

6. MISSION OPERATIONS.............................................................................................................................76 6.1 Mission Phases ....................................................................................................................................76

6.1.1 Mission Timeline.........................................................................................................................78 6.1.2 Launch Date Dependent Mission Characteristics ......................................................................84

6.2 Cruise Phase [CP] Overview................................................................................................................84 6.2.1 Cruise Phase Space/Ground Interface and Operations Support...............................................86

6.3 Nominal and Extended Mission Phase [NMP/EMP].............................................................................86 6.3.1 Phase Overview .........................................................................................................................86 6.3.2 Science Window Definition.........................................................................................................86 6.3.3 Space/Ground Interface .............................................................................................................90 6.3.4 Critical Operations......................................................................................................................91 6.3.5 System States and Modes Dictionary ......................................................................................101 6.3.6 System States and Modes Transistion.....................................................................................101

7. MECHANICAL SUBSYSTEM....................................................................................................................104 7.1 Requirements .....................................................................................................................................104 7.2 Design Overview ................................................................................................................................105

7.2.1 Core Structure ..........................................................................................................................105 7.2.2 Outer Structure.........................................................................................................................108 7.2.3 Instrument Boom Design..........................................................................................................113

8. HEATSHIELD ............................................................................................................................................115 8.1 Requirements .....................................................................................................................................115 8.2 Design Heritage..................................................................................................................................116

8.2.1 Heatshield Dimensions ............................................................................................................117 8.3 Thermal Design ..................................................................................................................................119

8.3.1 Design of Main Components....................................................................................................120 8.3.2 Star Brackets Layout ................................................................................................................122 8.3.3 Support Panel...........................................................................................................................123

8.4 Star-Brackets......................................................................................................................................123 8.5 Heatshield Interfaces with SC ............................................................................................................125 8.6 Front shield.........................................................................................................................................127 8.7 High Temperature Heat Barrier ..........................................................................................................127 8.8 Low Temperature Heat Barrier...........................................................................................................128 8.9 MLI on Spacecraft ..............................................................................................................................128 8.10 Stand Offs.........................................................................................................................................128

9. FEEDTHROUGHS, DOORS & MECHANISMS ........................................................................................129 9.1 Common Elements of Feedthrough Design .......................................................................................129

9.1.1 FT Body Design........................................................................................................................129 9.1.2 Interface with the HS Top Layer...............................................................................................130 9.1.3 Interface with the HS Support Panel ........................................................................................132 9.1.4 Materials...................................................................................................................................133

9.2 Feedthrough Design Concepts...........................................................................................................134


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9.3 Doors & Mechanisms .........................................................................................................................139

10. THERMAL CONTROL SUBSYSTEM .....................................................................................................141 10.1 Requirements ...................................................................................................................................141 10.2 Design Overview ..............................................................................................................................141 10.3 Reflection Control .............................................................................................................................144 10.4 Thermal Hardware Design Description ............................................................................................144

10.4.1 Overall Configuration .............................................................................................................145 10.4.2 Multi Layer Insulation (MLI) Blankets .....................................................................................146 10.4.3 Radiators ................................................................................................................................147 10.4.4 Doublers .................................................................................................................................150 10.4.5 Heat Pipes..............................................................................................................................150 10.4.6 Cold Finger Radiators ............................................................................................................151 10.4.7 Heaters...................................................................................................................................153 10.4.8 Thermistors ............................................................................................................................154

11. PROPULSION SUBSYSTEM..................................................................................................................157 11.1 Requirements ...................................................................................................................................157 11.2 Design Overview ..............................................................................................................................158

11.2.1 Pressurisation Subsystem......................................................................................................159 11.2.2 Propellant Subsystem ............................................................................................................160 11.2.3 Temperature monitoring.........................................................................................................160

11.3 Thruster Configuration......................................................................................................................161 11.4 Functional Description......................................................................................................................162

11.4.1 Launch Configuration .............................................................................................................162 11.4.2 Initialization and First Blow-Down Phase...............................................................................163 11.4.3 Repressurization and Second Blow-Down Phase .................................................................163 11.4.4 CPS Monitoring ......................................................................................................................163

11.5 Major Components Design Criteria ..................................................................................................164 11.5.1 Low Pressure NOP Derivation ...............................................................................................164 11.5.2 Propellant Tanks Sizing .........................................................................................................164 11.5.3 Pressurant Tank Sizing ..........................................................................................................165

12. ATTITUDE & ORBIT CONTROL SUBSYSTEM......................................................................................167 12.1 Requirements ...................................................................................................................................167 12.2 Design Overview ..............................................................................................................................167

12.2.1 Pointing Domain Classifications.............................................................................................167 12.2.2 Nominal Operations................................................................................................................168 12.2.3 AOCS Modes .........................................................................................................................170

12.3 AOCS Equipment Overview .............................................................................................................172 12.3.1 Star Tracker............................................................................................................................172 12.3.2 Inertial Measurement Unit ......................................................................................................173 12.3.3 Rate Measurement Unit .........................................................................................................174 12.3.4 Coarse Sun Sensor................................................................................................................174 12.3.5 Reaction Wheels ....................................................................................................................175

13. FAILURE DETECTION ISOLATION & RECOVERY SUBSYSTEM .......................................................176 13.1 Requirements ...................................................................................................................................176 13.2 Design Overview ..............................................................................................................................176

13.2.1 Hard Wired FDIR....................................................................................................................177 13.2.2 APSW Based Safe Mode Implementation .............................................................................182

14. DATA MANAGEMENT SUBSYSTEM.....................................................................................................184 14.1 Design Overview ..............................................................................................................................184

14.1.1 Solid State Mass Memory ......................................................................................................184


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14.1.2 On Board Computer (OBC)....................................................................................................188 14.1.3 Remote Interface Unit (RIU)...................................................................................................190 14.1.4 Failure Control Electronics (FCE) ..........................................................................................192

15. ELECTRICAL POWER SUBSYSTEM ....................................................................................................194 15.1 Requirements ...................................................................................................................................194 15.2 Design Overview ..............................................................................................................................194 15.3 Power Control and Distribution Unit (PCDU) ...................................................................................196

15.3.1 APR and MPPT......................................................................................................................197 15.3.2 Battery Charge Discharge Regulator (BCDR) .......................................................................198 15.3.3 Battery Disconnect Device (BDD) ..........................................................................................198 15.3.4 Equipment Power Distribution Module (LCL/FCL) .................................................................199 15.3.5 Heater Power Distribution Module .........................................................................................200 15.3.6 Pyro Firing Module .................................................................................................................201 15.3.7 Command & Monitoring Module.............................................................................................201 15.3.8 PCDU Layout and Mass Breakdown......................................................................................201

15.4 Battery ..............................................................................................................................................203 15.5 Solar Array Configuration .................................................................................................................203

15.5.1 PVA Design Concept .............................................................................................................205 15.5.2 Solar Array Power Interface ...................................................................................................205 15.5.3 Solar Array Drive Assembly (SADA)......................................................................................206 15.5.4 SADA Operational Modes ......................................................................................................207 15.5.5 SADA Electrical Interfaces .....................................................................................................207

15.6 External Power Interfaces ................................................................................................................208 15.6.1 LCL/FCL Interfaces ................................................................................................................208 15.6.2 Pyro / Thermal Knife Interfaces..............................................................................................209 15.6.3 Heater Interfaces....................................................................................................................210

16. COMMUNICATION SUBSYSTEM ..........................................................................................................211 16.1 Requirements ...................................................................................................................................211 16.2 Design Overview ..............................................................................................................................211

16.2.1 Deep Space Transponder ......................................................................................................213 16.2.2 X/Ka Band TWTA...................................................................................................................213 16.2.3 RFDU .....................................................................................................................................215 16.2.4 Antenna Assembly .................................................................................................................216

17. HARNESS SUBSYSTEM........................................................................................................................221 17.1 Harness Volume Accommodation....................................................................................................221 17.2 Harness Mass...................................................................................................................................222

18. SOLAR ORBITER BUDGETS.................................................................................................................223 18.1 Introduction.......................................................................................................................................223 18.2 Mass Budget.....................................................................................................................................224 18.3 RF-Link Budget.................................................................................................................................228 18.4 Power Budget ...................................................................................................................................231 18.5 Power Mode Summary.....................................................................................................................233 18.6 Delta V and Attitude Control Budget ................................................................................................234 18.7 Pointing Budget ................................................................................................................................237

18.7.1 RPE Budget [POIN30c]..........................................................................................................238 18.7.2 APE Budget [POIN30a]..........................................................................................................241 18.7.3 Instrument Co-alignment Budget ...........................................................................................243

18.8 Science Windows .............................................................................................................................245 18.9 2017 Mission ....................................................................................................................................245 18.10 2018 Mission ..................................................................................................................................247 18.11 Mass Memory Usage Profile ..........................................................................................................249


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APPENDIX: LAUNCHER PERFORMANCE .................................................................................................251 TABLES Table 2.4-1: Solar Orbiter Pointing requirements during operations...............................................................22 Table 2.14-1: Solar Orbiter Physical Reference Frame ..................................................................................28 Table 2.14-2: Solar Orbiter Optical Reference Frame Definition ....................................................................29 Table 3.3-1: BepiColombo equipment re-use for Solar Orbiter at a glance ....................................................40 Table 4.2-1: Solar Orbiter Configurations and their application within the mission timeline ...........................42 Table 4.3-1: Solar Orbiter Payload Complement (note: associated electronic boxes are not included in the

‘number of units’, but number 1 per instrument)......................................................................................48 Table 5.2-1: 04/01/2017 mission summary .....................................................................................................68 Table 5.2-2: 30/07/2018 mission summary .....................................................................................................70 Table 5.3-1: Solar Orbiter nominal Delta V budget for 2017 and 2018 mission scenarios .............................75 Table 6.1-1: Mission timeline for the nominal 2017 scenario (mission phases in green; GAM in grey) .........80 Table 6.1-2: Mission timeline for the backup 2018 scenario (mission phases in green; GAM in grey) ..........81 Table 6.3-1: Projected rotation populations for science window groups for the 1st array design....................97 Table 6.3-2: Percentage of mission time spent in the rolled state ..................................................................99 Table 6.4-1: The Solar Orbiter System Modes..............................................................................................101 Table 6.4-2: List of system mode transitions.................................................................................................103 Table 7.1-1: Solar Orbiter mechanical subsystem key requirements............................................................104 Table 8.1-1: Heatshield principal requirements.............................................................................................116 Table 8.2-1: Protruding element from SC external walls...............................................................................119 Table 10.1-1: TCS principal requirements.....................................................................................................141 Table 10.4-1 : MLI areas ...............................................................................................................................147 Table 10.4-2 : Radiator areas........................................................................................................................148 Table 10.4-3 : Cold finger radiator areas.......................................................................................................151 Table 10.4-4 : Heater distribution ..................................................................................................................154 Table 10.4-5 : Thermistor allocation..............................................................................................................156 Table 11.1-1: Propulsion subsystem key requirements ................................................................................157 Table 11.5-1: Propellant tank Helium temperature evolution ........................................................................166 Table 12.1-1: AOCS principal requirements..................................................................................................167 Table 12.2-1: AOCS resources to mode table ..............................................................................................172 Table 13.1-1: FDIR key requirements ...........................................................................................................176 Table 13.2-1: Solar Orbiter FDIR hierarchy...................................................................................................177 Table 15.1-1: EPS principal requirements.....................................................................................................194 Table 15.3-1: BCDE nominal modes and characteristics..............................................................................198 Table 15.3-2: Mass breakdown .....................................................................................................................202 Table 15.5-1: Solar Orbiter solar array major design data ............................................................................204 Table 15.5-2: 3G28 with TiOx AR coating electrical characteristic ...............................................................205 Table 15.6-1: Solar Orbiter LCL/FCL budget ................................................................................................209 Table 15.6-2: Solar Orbiter pyro and thermal knife interfaces ......................................................................210 Table 16.1-1: TT&C subsystem principal requirements ................................................................................211 Table 16.2-1: Solar Orbiter frequency and bandwidth allocation ..................................................................215 Table 17.2-1: Harness mass as percentage of SC dry mass for various missions.......................................222 Table 18.2-1: Solar Orbiter mass budget (2018 backup mission scenario) for the baseline Atlas launch case

...............................................................................................................................................................226 Table 18.2-2: Solar Orbiter mass budget (2018 backup mission scenario) for the backup Soyuz-Fregat

launch case............................................................................................................................................227 Table 18.3-1: Solar Orbiter downlink budget summary.................................................................................229 Table 18.3-2: Solar Orbiter uplink budget summary......................................................................................229 Table 18.3-3: Solar Orbiter ranging budget summary ...................................................................................230 Table 18.4-1: Power mode descriptions........................................................................................................232 Table 18.4-2: Thermal case definitions and corresponding power modes....................................................233 Table 18.5-1: Solar Orbiter power mode summary (for full breakdown of the power budget, please refer to

the Solar Orbiter budgets document [IR19])..........................................................................................233 Table 18.6-1: Nominal 2017 Mission Scenario..............................................................................................235 Table 18.6-2: Backup 2018 Mission Scenario...............................................................................................236 Table 18.7-1: Solar Orbiter pointing and accuracy requirements during RS-observation.............................237


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Table 18.7-2: Thermal cases corresponding to alignment budgets ..............................................................237 Table 18.7-3: Line-of-Sight RPE pointing budget..........................................................................................238 Table 18.7-4: Line-of-Sight RPE pointing budget..........................................................................................239 Table 18.7-5: Around Line-of-Sight RPE pointing budget .............................................................................240 Table 18.7-6: LoS APE during thermal case D1 (“Hot1”) ..............................................................................241 Table 18.7-7: LoS APE during thermal case D2 (“Hot2”) ..............................................................................241 Table 18.7-8: LoS APE during thermal case D6 (“Cold”) ..............................................................................241 Table 18.7-9: Worst case coalignment budget, corresponding to thermal case D1 (“Hot1”) ........................244 Table 18.9-1: Maximum latitude science window definition for the 2017 launch case..................................245 Table 18.9-2: Perihelion science window definition for the 2017 launch case..............................................246 Table 18.9-3: Minimum latitude science window definition for the 2017 launch case...................................246 Table 18.10-1: Maximum latitude science window definition for the 2018 launch case................................247 Table 18.10-2: Perihelion science window definition for the 2018 launch case............................................247 Table 18.10-3: Minimum latitude science window definition for the 2018 launch case.................................248 Table 18.11-1: 2nd order polynomial fit of Soyuz-Fregat performance data ..................................................252 FIGURES Figure 2.3-1: Science window definition..........................................................................................................18 Figure 2.3-2: SC and Solar Equator angular velocity over the mission duration ............................................21 Figure 2.14-1: Solar Orbiter Physical Reference Frame .................................................................................28 Figure 2.15-1: Mass margin philosophy ..........................................................................................................29 Figure 2.17-1: Power Margin Philosophy ........................................................................................................31 Figure 3.1-1: The evolution of the Solar Orbiter spacecraft design from the initial and second CDF studies,

through the initial EADS Astrium industrial study, to the EADS Astrium design at the end of the Heatshield System Technology Study.....................................................................................................36

Figure 3.2-1: Solar Orbiter detailed functional architecture.............................................................................37 Figure 3.3-1: Solar Orbiter SC with principle design features highlighted.......................................................39 Figure 4.2-1: Pictorial representation of the relationship between mission phases and SC configurations

(*the HTHGA can be stowed for all science windows or just for those windows below 0.3 AU distance to the Sun depending upon the precise operational scenario chosen; number of C_SCIENCE is representative only) .................................................................................................................................42

Figure 4.2-2: C_LEOP configuration, with the arrays, RPW antenna and HTHGA stowed [left] –Y panel view, [right] isometric view – [left] in the Soyuz fairing, [right] isometric view on LVA......................................43

Figure 4.2-3: C_NECP configuration, with the arrays and HTHGA deployed.................................................44 Figure 4.2-4: C_CRUISE configuration, with the arrays, HTHGA, instrument boom and RPW antennas

deployed ..................................................................................................................................................45 Figure 4.2-5: [left] C_SCIENCE configuration, with the arrays, HTHGA, instrument boom and RPW antennas

deployed, and the RS-instrument feedthrough doors open [right] C_SCIENCE sub configuration with HTHGA positioned behind the shadow of the heatshield, for periods in the orbit below 0.28 AU from the Sun ..........................................................................................................................................................46

Figure 4.2-6: C_CONTINGENCY configuration with the MGA deployed. Note that in this figure the HTHGA is shadowed, although it can also remain in the deployed state if the Sun distance is above 0.28 AU..47

Figure 4.3-1: Remote Sensing instrument accommodation ............................................................................49 Figure 4.3-2: Remote sensing instrument FoVs..............................................................................................50 Figure 4.3-3: In Situ instrument accommodation.............................................................................................51 Figure 4.3-4: EPD-STE FoV ............................................................................................................................52 Figure 4.3-5: SWA FoVs..................................................................................................................................53 Figure 4.3-6: EPD-EPT FoVs ..........................................................................................................................54 Figure 4.3-7: EPD-LET FoV ............................................................................................................................55 Figure 4.3-8: EPD-HETn and EPD-SIS FoVs .................................................................................................56 Figure 4.4-1: Solar Orbiter Structure (exploded view).....................................................................................57 Figure 4.4-2: Heatshield accommodation........................................................................................................58 Figure 4.4-3: Remote sensing instrument feedthroughs doors and mechanisms...........................................59 Figure 4.4-4: TCS Accommodation .................................................................................................................60 Figure 4.4-5: Propulsion subsystem accommodation .....................................................................................61 Figure 4.4-6: AOCS and DMS accommodation ..............................................................................................62 Figure 4.4-7: Power subsystem accommodation ............................................................................................63


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Figure 4.4-8: TT&C subsystem accommodation .............................................................................................64 Figure 4.4-9: Harness volume allocation (including connector volumes)........................................................65 Figure 5.2-1: 2017 trajectory viewed from above the ecliptic..........................................................................68 Figure 5.2-2: 2018 trajectory viewed from above the ecliptic..........................................................................70 Figure 5.2-3: Sun, Earth and Venus distance over the baseline 2017 mission scenario................................71 Figure 5.2-4: Solar latitude over the 2017 baseline mission scenario (pink arrows indicate Launch and GAM

events; note that orbital inclination changes occur at GAMs) .................................................................71 Figure 5.2-5: Sun/SC/Earth angle over the 2017 baseline mission scenario..................................................72 Figure 5.2-6: Sun, Earth and Venus distance over the baseline 2017 mission scenario................................72 Figure 5.2-7: Solar latitude over the 2017 baseline mission scenario.............................................................73 Figure 5.2-8: Sun/SC/Earth angle over the 2017 baseline mission scenario..................................................73 Figure 5.3-1: Deterministic DSM as a function of launch day for the [left] 2017 and [right] 2018 launch

windows ...................................................................................................................................................74 Figure 6.1-1: Baseline 2017 mission scenario showing the evolution of Sun distance over the mission, and

the GAM events; the MRD-definitions of the CP, NMP and EMP are also shown..................................77 Figure 6.1-2: Baseline 2017 mission scenario showing the evolution of helio-latitude over the mission, and

the GAM events; the MRD-definitions of the CP, NMP and EMP are also shown..................................77 Figure 6.1-3: Baseline 2017 mission scenario showing the evolution of the Sun/SC/Earth angle over the

mission, and the GAM events; the MRD-definitions of the CP, NMP and EMP are also shown ............78 Figure 6.1-4: Nominal 2017 mission timeline showing start dates of each mission phase and principal events

.................................................................................................................................................................82 Figure 6.1-5: Backup 2018 mission timeline showing start dates of each mission phase and principal events

.................................................................................................................................................................83 Figure 6.2-1: Sun distance and Sun/SC/Earth angle over the course of the CP for the nominal 2017 mission;

also shown are the principal mission events (far-Sun, GAM and conjunction periods) ..........................85 Figure 6.2-2: Sun distance and Sun/SC/Earth angle over the course of the CP for the backup 2018 mission;

also shown are the principal mission events (far-Sun, GAM and conjunction periods) ..........................85 Figure 6.3-1: Sun distance and Sun/SC/Earth angle over the course of the NMP for the nominal 2017

mission; also shown are the principal mission events (science windows, GAM and conjunction periods).................................................................................................................................................................88

Figure 6.3-2: Sun distance and Sun/SC/Earth angle over the course of the NMP for the backup 2018 mission; also shown are the principal mission events (science windows, GAM and conjunction periods).................................................................................................................................................................88

Figure 6.3-3: Sun distance and Sun/SC/Earth angle over the course of the EMP for the nominal 2017 mission; also shown are the principal mission events (science windows and GAM)..............................89

Figure 6.3-4: Sun distance and Sun/SC/Earth angle over the course of the EMP for the backup 2018 mission; also shown are the principal mission events (science windows and GAM)..............................89

Figure 6.3-5: Non-communication periods during the NMP/EMP for the nominal 2017 mission; the large period (circled in blue) is caused by the conjunction, the rest are caused by stowing the HTHGA during periods where the SC is closer than 0.3 AU to the Sun (they can be seen to coincide with the perihelion science windows – blue diamonds) .........................................................................................................90

Figure 6.3-6: Non-communication periods during the NMP/EMP for the backup 2018 mission; the large periods (circled in blue and yellow) ars caused by the conjunction, the rest are caused by stowing the HTHGA during periods where the SC is closer than 0.3 AU to the Sun (they can be seen to coincide with the perihelion science windows – blue diamonds)...........................................................................91

Figure 6.3-7: Back-compatible solar array operational ranges .......................................................................92 Figure 6.3-8: Baseline 2017 mission scenario showing the evolution of Sun distance over the mission, and

the GAM events; the MRD-definitions of the CP, NMP and EMP are also shown; red line denotes preliminary array flipping distance ...........................................................................................................93

Figure 6.3-9: Example incidence profile and temperature for the 1st solar array design for the 2017 baseline mission scenario ......................................................................................................................................94

Figure 6.3-10: Example incidence profile and temperature for the 2nd solar array design for the 2017 baseline mission scenario .......................................................................................................................95

Figure 6.3-11: The first peripassage of the 2017 nominal mission scenario. The red points indicate the beginning and end of the window; because the peripassage is an inflection point, there will be two rotation events for rotation distances below the lower of the start/end distances...................................96

Figure 6.3-12: The transformation in Earth positions in the PRF when a 60 degree roll applied to the SC away from the nominal attitude................................................................................................................98


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Figure 6.3-13: Baseline configuration roll profile for the 2017 mission scenario.............................................99 Figure 6.4-1: Solar Orbiter system mode transitions diagram.......................................................................102 Figure 7.2-1: Solar Orbiter LIR Lay-out .........................................................................................................105 Figure 7.2-2: Exploded view with location of –Y central shear wall ..............................................................106 Figure 7.2-3: Exploded view with location of ±Y±Z shear walls ....................................................................107 Figure 7.2-4: Exploded view with location of 1 tank support panel ...............................................................108 Figure 7.2-5: Exploded view with location of top floor ...................................................................................108 Figure 7.2-6: Top Floor connection with Heatshield......................................................................................109 Figure 7.2-7: Exploded view with location of lower floor ...............................................................................110 Figure 7.2-8: Exploded view with location of ±Z closure panels ...................................................................111 Figure 7.2-9: Exploded view with location of ±Y closure panels ...................................................................112 Figure 7.2-10: Y wall access panel location ..................................................................................................112 Figure 7.2-11: CFRP Tube with Bonded Titanium Fittings............................................................................113 Figure 7.2-12: Hinge mechanism is Titanium and Al. Alloy...........................................................................114 Figure 7.2-13: HDRM Release actuator and Titanium separation interface .................................................114 Figure 8.2-1: Solar Orbiter overall size (I) .....................................................................................................117 Figure 8.2-2: Solar Orbiter overall size (II) ....................................................................................................118 Figure 8.3-1: Heatshield components............................................................................................................120 Figure 8.3-2: Overall HS archtecture, showing support bipod/monopods, corner cut-outs for SWA heads,

Sun sensor cutout, feedthroughs for RS-instruments and star Bracket supports .................................121 Figure 8.3-3: The front shield is oversized to provide HS nominal performance up until the worst-case off-

pointing case..........................................................................................................................................121 Figure 8.3-4: Star-brackets layout .................................................................................................................122 Figure 8.3-5: Heatshield shadow projection over the SC structure...............................................................122 Figure 8.3-6: Heatshield support panel sandwich construction.....................................................................123 Figure 8.4-1: Star bracket accommodation within HS structure....................................................................124 Figure 8.4-2: Star Bracket design and principal dimensions.........................................................................124 Figure 8.5-1: Heatshield Mounting I/F Location ............................................................................................125 Figure 8.5-2: Underside of HS support panel showing blade mounts...........................................................126 Figure 8.5-3: Blade mounts ...........................................................................................................................126 Figure 8.5-4: Conceptual depiction of the central bolt...................................................................................127 Figure 8.7-1: Design and installation of patches ...........................................................................................128 Figure 9.1-1: Corner FT body concept design...............................................................................................129 Figure 9.1-2: View of the FT interface design at top of the EUI FT...............................................................130 Figure 9.1-3: Corner FT concept interface with HS top layer........................................................................131 Figure 9.1-4: View of the FT interface design at HS support panel of the EUI FT........................................132 Figure 9.1-5: Corner FT concept design (II) ..................................................................................................133 Figure 9.2-1: Remote-sensing instruments heat shield feedthroughs, doors and mechanisms. Note STIX

does not have a door and the second PHI door is work underway; note also the interference between the PHI feedthroughs.............................................................................................................................134

Figure 9.2-2: PHI feedthrough showing filter mounting concept ...................................................................135 Figure 9.2-3: Preliminary SPICE feedthrough design ...................................................................................136 Figure 9.2-4: EUI feedthroughs, doors and mechanism design....................................................................137 Figure 9.2-5: METIS feedthrough, door and mechanism design ..................................................................138 Figure 9.2-6: STIX feedthrough baseline design...........................................................................................138 Figure 9.2-7: SWA-HIS feedthrough conceptual design ...............................................................................139 Figure 9.3-1: Door mechanism design ..........................................................................................................140 Figure 10.2-1: Overview of Solo Orbiter Thermal Design Configuration ...............................................143 Figure 10.3-1 : Reflection of solar flux from solar array edge to platform side..............................................144 Figure 10.4-1 : MLI Temperature in View of HGA.........................................................................................145 Figure 10.4-2 : External MLI Blanket Types ..................................................................................................146 Figure 10.4-3: Radiating areas ......................................................................................................................149 Figure 10.4-4: TWT doublers.........................................................................................................................150 Figure 10.4-5 : Heat pipe routing...................................................................................................................151 Figure 10.4-6: Radiating areas ......................................................................................................................152 Figure 10.4-7 : Cold finger LHP routing.........................................................................................................152 Figure 11.2-1: Solar Orbiter CPS Schematic ................................................................................................158


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Figure 11.2-2: Solar Orbiter propulsion system mechanical layout...............................................................159 Figure 11.3-1: [top] Type 1 TCM allowing an unconstrained slew to optimum delta V attitude, [bottom] Type

2 TCM accessing the desired velocity vector by rotating the SC around the sunline to bring the velocity vector into the XY plane of the SC, and then performing a dog-leg manoeuvre (blue arrows).............161

Figure 11.3-2: Solar Orbiter thruster positions and designations..................................................................162 Figure 11.5-1: EADS 10N thruster pressure boxes.......................................................................................164 Figure 11.5-2: Propellant tank pressure and ullage evolution.......................................................................165 Figure 12.2-1: Solar Orbiter AOCS architecture............................................................................................169 Figure 12.2-2: AOCS mode diagram.............................................................................................................170 Figure 12.3-1: Star Trackers mounted on sensor plate together with inertial measurement unit .................173 Figure 12.3-2: The Hemispherical Resonator Gyro SSIRU inertial measurement unit.................................174 Figure 12.3-3: [left] RMU Unit, [right] Coarse Sun Sensor ............................................................................174 Figure 13.2-1: Survival mode simulation for thruster failure, 10Nm over 1s .................................................179 Figure 13.2-2: ARAD functional configuration scheme .................................................................................180 Figure 13.2-3: Survival Mode Electrical Architecture ....................................................................................182 Figure 13.2-4: APSW Safe Mode functionality ..............................................................................................183 Figure 14.1-1: Baseline Electrical Architecture on Solar Orbiter...................................................................184 Figure 14.1-2: Solid State Mass memory Architecture..................................................................................185 Figure 14.1-3: Solid state Mass Memory operational modes........................................................................187 Figure 14.1-4: OBC architecture....................................................................................................................188 Figure 14.1-5: OBC mode diagram ...............................................................................................................189 Figure 14.1-6: RIU architecture diagram .......................................................................................................190 Figure 14.1-7: FCE diagram..........................................................................................................................193 Figure 15.2-1: Solar Orbiter power electrical architecture.............................................................................195 Figure 15.3-1: Function of MPPT ..................................................................................................................197 Figure 15.3-2: APR schematic block diagram ...............................................................................................197 Figure 15.3-3: BCDR schematic block diagram ............................................................................................198 Figure 15.3-4: Battery disconnection device .................................................................................................199 Figure 15.3-5: LCL module block schematic .................................................................................................199 Figure 15.3-6: PCDU TSW arrangement.......................................................................................................200 Figure 15.3-7: Heaters/Thermostats configuration........................................................................................200 Figure 15.3-8: Pyro firing module block schematic .......................................................................................201 Figure 15.3-9: PCDU layout ..........................................................................................................................202 Figure 15.4-1: 18650HC cell..........................................................................................................................203 Figure 15.5-1: Solar Orbiter solar array in deployed condition......................................................................204 Figure 15.5-2: SADM conceptual block diagram...........................................................................................206 Figure 16.2-1: Solar Orbiter TT&C subsystem architecture ..........................................................................212 Figure 16.2-2: X/Ka band TWTA architecture ...............................................................................................214 Figure 16.2-3: RFDU schematic ....................................................................................................................216 Figure 16.2-4: BepiColombo MGA schematic ...............................................................................................217 Figure 16.2-5: HTHGA dual-channel rotary joint ...........................................................................................218 Figure 16.2-6: [left] HTHGA in nominal position [right] HTHGA in close-approach shielded position ..........219 Figure 16.2-7: Isometric view of the Earth position in the SC PRF throughout the mission for the former 2015

launch mission scenario: the yellow sphere represents the Sun position, with the SC XYZ coordinates shown as the red axes-set.....................................................................................................................219

Figure 16.2-8: HTHGA interference regions of the sky for the baseline (2 RPW antenna on –Z SC side: blue – CP; cyan – NMP; green – EMP).........................................................................................................220

Figure 16.2-9: HTHGA elevation rate for the former APM order (elevation/rotation from SC to HGA) throughout the course of the former 2015 launch mission, showing rate-spikes due to mechanism constraints .............................................................................................................................................220

Figure 17.1-1: Harness accommodation .......................................................................................................221 Figure 18.2-1: Mass margin...........................................................................................................................224 Figure 18.4-1: Power Margin Philosophy ......................................................................................................231 Figure 18.6-1: Delta V budget philosophy .....................................................................................................234 Figure 18.7-1: POIN45 interpretation and relationship with POIN30a ..........................................................243 Figure 18.11-1: SSMM usage for the 2017 baseline mission scenario.........................................................250 Figure 18.11-2: SSMM usage for the 2018 backup mission scenario...........................................................250


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Figure 18.11-1: Selected 20 day launch window around the 9/1/2017 launch date (from previous analysis)...............................................................................................................................................................251

Figure 18.11-2: Projected Soyuz-Fregat performance as a function of C3 from Kourou for various declinations: the projected performance for the Solar Orbiter 9/1/2017 launch is 1301kg ...................252


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1. APPLICABLE AND REFERENCE DOCUMENTS

1.1 Reference Documents

The following documents, although not a part of this document, amplify or clarify its contents: Ref Title Document Number IR01 CDF Study Report: Solar Orbiter 2 Report: CDF-25(A)

IR02 Solar Orbiter Assessment Phase Final Executive Report SCI-A/2005/054/NR

IR03 HELEX: Heliophysical Explorers: Solar Orbiter and Sentinels Pre-publication version

IR04 Solar Sentinels/Solar Orbiter Feasibility Study NASA-LWS 462-RPT-0008

IR05 Solar Orbiter Assessment Study Final Report SAS-TN-ASF-012

IR06 Solar Orbiter Assessment Study CCN Final Presentation SAS-HO-ASF-055

IR07 Solar Orbiter Heatshield & System Technology Study System Design Report

SOL-T-ASTR-TN-19

IR08 Solar Array and High Gain Antenna Concept Trade-Off and Feasibility Report

SOLO-ASD-TN-001

IR09 SOLO Ballistic Mission Options Update SOL-S-ASTR-TN-24

IR10 Soyuz to Atlas Environment Delta Definition SOL-S-ASTR-TN-63

IR11 Solar Orbiter Design Changes due to the Combined Atlas launch with Sentinels

SOL-T-ASTR-TN-64

IR12 Solar Orbiter Heatshield/System Technology Study System Level Thermal Analysis

SOL-T-ASTR-TN-67

IR13 Solar Orbiter Design Changes due to the Dedicated Atlas launch and back-compatibility with Soyuz, and integration into the HELEX Concept

SOL-T-ASTR-TN-91

IR14 Heatshield Door/Mechanism Design and Accommodation SOL-S-ASTR-TN-92

IR15 Solar Orbiter System Design Report (TAS-I) SOL-C-AAS-RP-100207183B

IR16 Solar Orbiter System Trade Offs SOL-S-ASTR-TN-0012

IR17 Solar Orbiter Mission Operations and Ground Segment Compliance SOL-S-ASTR-TN-0011

IR18 Solar Orbiter Compliance Matrix SOL-S-ASTR-TN-0013

IR19 Solar Orbiter Budgets SOL-S-ASTR-TN-0010

IR20 Compilation of Engineering Work for Solar Orbiter Phase B1 SOL-F-ASTR-TN-0001

IR21 Solar Orbiter Environment Comparison and Consequences SOL-F-ASTR-TN-0003

IR22 Solar Orbiter Heatshield Thermal Analysis SOL-T-TAS-RP-00001

IR23 Solar Orbiter Heatshield Design Description SOL-T-TAS-RP-000X

IR24 Solar Orbiter Heatshield Trade Offs and Design Concept Proposal SOL-T-TAS-TN-0001

IR25 Solar Orbiter Feedthrough Door & Mechanism Design & Trade off SOL-S-ASTR-TN-0015

IR26 Solar Orbiter Mission Analysis for Launch in 2017-2018 SOL-ESC-RP-GFA-JRC-WP 5XX

IR27 Solar Orbiter DHS Subsystem Design Description SOL.F.ASTR.PR.00002

IR28 Solar Orbiter AOCS Subsystem Design Description SOL.F.ASTR.PR.00001

IR29 Equipment Reuse for Solar Orbiter – Technical Report SOL.F.ASTR.RP.00003

IR30 Procurement Strategy for Solar Orbiter BC-ASD-RP-00019

IR31 Solar Orbiter Power Systema Analysis SOL-U-ASTR-TN-0001

IR32 Solar Orbiter Instrument Proposal Evaluation Report SOL-U-ASTR-TN-0003

IR33 Solar Orbiter Payload Interfaces, Requirements and SOL-U-ASTR-TN-0006


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Accommodation IR34 Solar Orbiter Payload Thermal Requirements and Interface

Definition Document SOL-U-ASTR-TN-0007

IR35 Solar Orbiter Thruster Plume Impingement and Contamination SOL-U-ASTR-TN-0009

IR36 Solar Orbiter CPS Trade-off SOL-U-ASTR-TN-0011

IR37 Feedthroughs Doors and Mechanisms Design Description SOL-S-ASTR-TN-0015

IR38 Solar Orbiter Structure Mechanical Analysis SOL-S-ASTR-TN-0003

IR39 Solar Orbiter Heatshield Sizing SOL-S-ASTR-TN-0006

IR40 Solar Orbiter Fault Tree Analysis SOL-S-ASTR-TN-0008

IR41 Solar Orbiter Heatshield Materials TDA SD-PL-AI-0221

IR42 Solar Orbiter AIT Plan SOL-S-ASTR-PL-0009

IR43 Solar Orbiter Suggested Changes to MRD SOL-S-ASTR-TN-0033

IR44 Solar Orbiter Mission Operations Supporting Analysis SOL-S-ASTR-TN-0034

IR45 Solar Orbiter CPS Design Description SOL-S-ASTR-TN-0021

IR46 Solar Orbiter Delta V and Attitude Control Budget SOL-S-ASTR-TN-0031

IR47 Solar Orbiter Instrument Boom Concept Design Report SOL-S-ASTR-TN-0027

IR48 – [IR45] - -

IR49 Solar Orbiter Definition Study Change Request to ESTEC/Contract no. 21496/08/NL/HB

SOL-EST-CR-01271

IR50 Solar Orbiter Mechanical Design Report SOL-S-ASTR-TN-0017

IR51 Solar Orbiter EPS Design Report SOL.F.ASTR.RP.00006

IR52 Solar Orbiter TT&C Design Report SOL.F.ASTR.RP.00009

IR53 Solar Orbiter TCS Design Description SOL-S-ASTR-TN-0018

1.2 Normative Applicable Documents

The following normative documents contain provisions which, through reference in this text, constitute provisions of this document. For dated references, subsequent amendments to, or revisions of any of these publications do not apply. However, parties to agreements based on this document are encouraged to investigate the possibility of applying the most recent editions of the normative documents indicated below. For undated references, the latest edition of the normative document referred to applies. Ref Title Document Number NR01 Solar Orbiter Mission Requirements Document (4.0) SOL-EST-RS-00049

NR02 Solar Orbiter Spacecraft/System Requirements Specification SOL-S-ASTR-RS-18

NR03 Solar Orbiter Payload Definition Document SOL-EST-SP-00705

NR04 Solar Orbiter Science Requirements Document SCI-SH/2005/100/RGM

NR05 Solar Orbiter Environmental Specification TEC-EES-03-034/JS

NR06 Atlas Launch System Mission Planner’s Guide CLSB-0409-1109

NR07 Soyuz (from Guiana Space Centre) User’s Manual Issue 1, Rev. 0, June 06

NR08 Solar Orbiter: Mission Preparation and Operations Approach (draft) SO-ESC-TN-00001

NR09 Solar Orbiter Heatshield Interface Requirement Document SOL-S-ASTR-ICD-0001

NR10 Solar Orbiter Experiment Interface Document – Part A (EID-A) SOL-EST-IF-50


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2. SOLAR ORBITER MISSION REQUIIREMENTS

This chapter discusses the mission requirements from MRD 4.0 [NR01], providing references where appropriate to design decisions and supporting analyses that attempt to address these requirements. The reader is directed to [IR05, IR07] for discussions of previous MRD editions within the context of previous study phases. Furthermore, several of the requirements listed here are subjected to proposed changes – these suggested changes to the MRD 4.0 are collated and discussed in more detail in [IR43].

2.1 Solar Orbiter Mission Requirements

MISS05 – The Solar Orbiter spacecraft shall be a 3-axes stabilized spacecraft using conventional chemical propulsion for orbit manoeuvres and attitude control. This requirement is a strong design driver for the spacecraft, since it rules out the possibility of electrical propulsion, studied in previous phases. Note that the ability to avoid consideration of EPS is predicated upon the changes made to the transfer trajectory detailed in [IR09], resulting in a zero-deterministic ballistic transfer. MISS10 - Target launch May 2015with backup in 2017. This requirement has actually been superceded during the course of the Phase B1 study due to programmatic problems in the BepiColombo programme from which much of the Solar Orbiter technology is derived. The new baseline launch date for the mission is the original backup (2017), with the new backup launch date in 2018. The characteristics of these two launch date scenarios are detailed in [IR26], and furthermore are discussed within this report (section 6). MISS15 – The maximum mission duration to the end of Nominal mision shall be considered as TBD and the maximum additional duration for the Extended mission shall be considered as TBC. The mission duration is one of the main drivers of the system since it drives both consumable quantities, GS costs, as well as material and degradation effects (solar array, heat Shield). Given the new launch dates, the actual durations of both the Nominal and Extended mission phases are now well known: for the 2017 (new) baseline, the duration to the end of the NMP is 7.5 years and the duration to the end of the EMP is 9.6 years. MISS20 – The solar Orbiter mission shall operate a Store and Forward data architecture. This requirement is a consequence of the operational complexity of the Solar Orbiter mission: a wide variation in Earth distance, Sun/SC/Earth angle, Sun distance, repeated conjunction events, and operationally necessary suspension of communications, due to inability of the HTHGA to survive at Sun distances below 0.28 AU (TBC), all combine to place a requirement for a store and forward data architecture on-board the spacecraft. This feature of the mission has been well understood since the early assessment studies and is fully reflected in the current design, which contains a large SSMM unit to store data prior to download. For a detailed description of the SSMM usage-profile of the mission, as well as associated data-latencies, please refer to [IR17, IR44]. The assumed requirement is that the full data-load generated by the mission (both instruments and housekeeping) should be considered when sizing the SSMM required; however in [NR08] it is stated that conjunction events should not drive the design of the spacecraft. For the current design, the SSMM has been sized to the first assumption; however, if the ESOC document recommendation is followed, then a reduction in the SSMM capacity required can be achieved. Note also that judiscious usage of a 2nd ground station and/or extension of the pass frequency or duration around the SSMM usage profile peak (which occurs at MoL) could potentially very significantly reduce the SSMM requirements of the mission.

2.2 Lifetime Requirements

LIFE05 – All spacecraft consumables and radiation sensitive units shall be sized from the launch until the end of the extended mission lifetime.


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This requirement drives the consumables and radiation sensitive units to be compatible with operation up until the end of the EMP (~9.6 Years for the 2017 baseline, the sizing case). In practise, consumables are not driven by lifetime but rather by manoeuvre events, and in this case it is the 2018 backup mission scenario which is the sizing case due to the presence of a launch window error penalty, needed to maintain the mission compatibility with a 20 day launch window (LAUN15). LIFE10 – During the nominal mission lifetime, the Solar Orbiter Spacecraft performance requirements shall be fully met and include all specified margins. This requirement implies that the design of the SC (apart from consumables and survivability) shall be driven by mission requirements only until the end of the NMP, and that the EMP should not drive the spacecraft design with regards to performance requirements (see below). This is an important requirement because it allows potential relaxation of performance beyond the NMP. LIFE15 – The Solar Orbiter design and build shall be compatible with a 19 month launch delay (launch window locked to the synodic period of Venus). Under the chosen transfer scenario, the SC will target a Venus encounter (in order to perform a GAM) directly after Earth escape. Should the launch window be missed, this necessitates a wait until the relative positions of Earth and Venus (in the ecliptic plane) are repeated. LIFE20 – The spacecraft commissioning phase shall commence after LEOP and last for 2 (3) months. This is understood as the Near Earth Commissioning Phase (NECP) described in the MRD [NR01], where the declaration is 2 months (this should be updated in the next MRD issue). In actuality the payload commissioning is not limited to this 3 months period but will extend during the Cruise Phase (CP) for a TBD period. It is expected that the commissioning of the IS-instruments shall be completed during the NECP, but that residual RS-instrument commissioning shall take place in the early stages of the CP.

2.3 Space Segment Requirements: Performances

PERF05 – The Solar Orbiter Spacecraft shall be capable of accommodating the payload complement as defined in EID-A [NR10] resulting from the selection subsequent to the Solar Orbiter Announcement of Opportunity. The EID-A [NR10] holds a description of the interface definition between the platform and the selected instruments. Payload responses to the AO include the EID-B documents [various], which are a response to the interface definition specified in the EID-A. Currently, although the payload complement has been selected, there is a high degree of missing information and non-compliance within the various EID-B (see [IR33] for a systematic analysis of EID-B compliance with EID-A). As the above requirement states, the design of the Solar Orbiter platform has been and will be conducted according to EID-A and therefore the current non-compliances of payloads are to a great extent ignored. PERF10 – The Solar Orbiter Spacecraft shall provide all necessary resources (power, data handling, and thermal control) to the Payload, as specified in EID-A [NR10]. This requirement is an example of the approach specified by PERF05. As an example of the design route followed, the total payload mass allocation is 156kg including margins, and accordingly a payload mass of 156kg that is used in the design of the Solar Orbiter spacecraft. The EID-A specified payload mass allocation of 180 kg has been reduced to 156 kg through the transfer of payload support equipments (boom, doors and mechanisms, thermal interfaces) to other sections of the spacecraft mass budget, in agreement with the ESA project team. PERF15 - The Solar Orbiter Spacecraft shall provide the Payload with global resources of at least:

• 180 kg of Payload total mass (including payload support elements and 5% overall system margin) • 180 Watts of payload power (including 20% margins and bus at 28V)


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• 100 kbps (kilobits per second) average generation rate for compressed scientific data (including 10% system margin.

This requirement is a summary of the payload envelope as specified in the EID-A [NR10]. See PERF10 for description of current payload mass budget allocation as agreed between the Prime and the ESA Project Team. PERF20 - For each operational orbit, the Solar Orbiter spacecraft shall allow simultaneous operations of the complete payload complement, as well as communications when available, for three continuous periods of 10 days, with a maximum of 2 periods contiguous, each, centred on:

• maximum northern heliographic latitude and • maximum southern heliographic latitude and • perihelion passage.

This requirement provides the definition of the desired science windows of the mission – for a definition of the science windows for both the 2017 and 2018 mission scenarios, using a strict interpretation of this requirement, please refer to [IR17]. It is important to note that under this definition, the early perihelion and minimum latitude windows of the mission overlap to a large extent; as the perihelion of the SC gradually increases over the course of the mission (and accordingly the semi-major axis of the orbit and hence the period), the overlap reduces and eventually disappears. Note that the overlapping science window definition has been used in all the operational analyses based upon science window definitions. These periods are referred to hereafter in the System Design Report as ‘science windows’. Note that the science window is defined as the period during which the entire payload complement (both RS and IS suites) is operating; outside of the science windows only the IS-instrument suite is operating.

Figure 2.3-1: Science window definition

The requirement as expressed above implies uninterrupted provision of the required platform stability and pointing requirements for observations during the science windows (the platform stability and pointing requirements during science windows are given in POIN30). However, in actuality the design of the Solar


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Orbiter SC is highly dependent upon the operational approach chosen for the mission, particularly in and around the science windows, and there are several operational events that are likely to cause interruption to the science mode of the SC during the science windows. These are:

1. Reaction Wheel off-loading: Any net DC SRP-torque experienced during the science windows must be counteracted by the reaction wheels on-board. However, should the HTHGA be exposed to the solar flux during a science window, the magnitude of the net SRP-torque around the SC Y-axis will be such that, potentially, one or more RW offloading events is necessary within the science window due to the wheels becoming saturated. The frequency of such events will be different for each science window, and is a complex function of the SC design and the evolving parameters of the particular science window; consequently some science windows will require no wheel off-loading at all, whilst other windows will require several wheel off-loading events (see [IR44] for a description of the preliminary wheel off-loading analysis for the baseline 2015 mission, the results of which are indicative of the situation for any launch date). These events will necessitate a cessation of RS science observations because the noise introduced by the thrusters will easily violate the required RPE. This problem is the reason that it is recommended to fold the HTHGA behind the heatshield shadow for all science windows, in order to present a torque-balanced SC to the solar flux, and maximise the period between wheel desaturation events. Furthermore, as the reaction wheels become progressively loaded, their rotational frequency will increase, and there may exist frequencies at which coupling to the SC structure occurs, which will serve to further degrade the RPE performance.

2. Repointing: The scientific operations will contain one or more repointing events to allow different areas of the solar disc to be targeted (e.g. the movement from Sun-centre to limb-pointing); the number of repointing events is expected to be a function of the particular scientific interest of the current window; this repointing will necessitate an interruption to the required platform stability.

3. Antenna Movement: If not stowed behind the heatshield shadow, the HTHGA could theoretically be tracking the Earth and providing a communication link during the science windows. However, the mechanical movement of the HTHGA may be too disruptive to the RPE requirement contained in POIN30 (1 arscec/10s LoS, 2 arcsec/10s around LoS), and accordingly would prevent the required platform stability from being achieved (please see [IR19] for the current RPE budget including estimation of the contributions due to HTHGA movement and reaction wheel microvibrations). This, in combination with (1), leads to the recommendation to stow the HTHGA behind the heatshield shadow for all science windows, regardless of Sun distance.

4. Solar Array Rotations. The Solar Array may require1 several discrete rotations within science windows in order to manage the temperature of the array (please refer to [IR44] for a preliminary rotational profile of the Solar Arrays throughout the operational orbit of the mission). Solar Array rotation will easily violate the RPE requirements. As with the Reaction Wheel off-loading, the number of Solar Array rotations required will be a function of the particular parameters of the science window, specifically the range of Sun distances covered by the SC over the duration of the science window.

The management/removal of these sources of disturbance is critical to the successful design of the mission. From the results documented in the operational analysis of the mission [IR17, IR44], It is recommended that the HTHGA is stowed in the SC shadow during all of the science windows. This avoids the continuous contribution of the HTHGA to the RPE budget. It also dramatically improves the symmetry of the SC configuration, reduces the accumulation-rate of angular momentum due to solar radiation pressure, and consequently offers a significant increase in the uninterrupted observation periods. Stowing the dish during all science windows will entail a modest increase in the SSMM capacity requirements, as well as the data latency. Alternatively the approach should be to synchronise the interruptions and possibly coincide dish movements with these events. ORB05 – During the nominal mission, the Solar Orbiter spacecraft shall perform manoeuvres to achieve the following parameters:

1 Although the array design is not yet settled, it is confidently demonstrated in [IR44] that avoiding rotations during science windows will be impossible, particularly the perihelion science windows during which received normal flux changes more rapidly with radial distance from the Sun.


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• Orbital period in 3:2 resonance with Venus • Perihelion radius less than 0.25 AU and greater than 0.2 AU between Venus GAM-V2 and GAM-V3 • Inclination with respect to the solar equator increasing to a goal of 30 degrees.

Two ballistic mission profiles (2017 and 2018) have been selected from the many scenarios previously considered (see [IR26] for a description of the baseline trajectories upon which the SC design is based). The two chosen mission profiles exhibit a maximum Sun distance of 1.484 AU and a minimum Sun distance of 0.2343 AU, with mission duration of 9.6 years up until the end of the EMP – all these sizing parameters are from the baseline 2017 trajectory. The perihelion distance requirement is expressed as a range, and thus there is no specifically expressed spacecraft co-rotation requirement with respect to the Sun. It is important to realise that increasing the perihelion distance towards the upper end of the required range will have very important consequences for the mission:

• An highly-sensitive impact on the maximum encountered normal flux (22.6% normal flux reduction between 0.22 AU and 0.25 AU), corresponding to a reduction of the normal incident flux of the order of 6 kW. This is obviously hugely important from the perspective of heatshield performance, as well as the received flux through the RS-instrument apertures, which in turn drives the thermo-elastic deformation of the SC structure, and the APE/co-alignment performance as required in POIN30 and POIN45. Accordingly a relaxation of perihelion distance could be considered in the event of problems relating to heatshield performance, thermo-elastic deformation, or feedthrough door and mechanism design.

• An impact on pointing requirements (attitude guidance law) • A reduction in the maximum heliographic latitude that can be achieved by the spacecraft. As an

example, an increase in the peridistance from 0.22 to 0.3 would result in an end-of-mission latitude reduction of 5-10 degrees (TBC)

• The ability of the platform to maintain co-rotation with the Sun’s surface. If this requirement shall be taken into account for future mission analyses, a quantitative figure for the co-rotation requirement should be proposed. Note also that the performance accuracy of TCM to be achieved is not specified. The required accuracy in terms of direction and magnitude can be a driver for the SC design and operation definition. The allowable orbit disturbance caused by attitude control operations has also not been specified; currently there is no requirement for force-free attitude control of the SC, and the use of thrusters may induce parasitic forces that shall modify the orbit. This could be detrimental if such activities occurred prior to GAM, or may alter the perihelion distance and accordingly introduce navigation errors in the sun-tracking attitude guidance law. Finally it is worth examining the drivers for the perihelion distance a little more closely. The perihelion distance is often quoted as having been selected in order to achieve a co-rotating vantage point above the Sun’s surface. However, the primary driver of the perihelion distance should be admitted as the desired resolution achieved by the RS-instruments. The velocity of the Sun’s surface at the Solar equator is 7.284x103 km/h; this corresponds to a sidereal rotation period of 25.38 days, and an angular velocity of 2.865x10-6 rads-1. Accordingly even an inertially still SC located in a fixed point would be able to monitor the same region on the Solar disc for a ten-day period. The following figure shows the angular velocity of the SC and the Sun over the mission duration.


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0.000

0.100

0.200

0.300

0.400

0.500

0.600

0.700

0 500 1000 1500 2000 2500 3000 3500 4000 4500

Mission time (days)

SC a

ngul

ar v

eloc

ity (d

eg/h

r)

SC angular velocitySun equator angular velocity

Figure 2.3-2: SC and Solar Equator angular velocity over the mission duration

An average co-rotation of the SC of ~0.3 deg/hr is achieved at the peripassage points; accordingly the integrated mismatch between the Sun surface and the SC over a ten-day period is 70 degrees; from the perihelion vantage point of over 40 solar radii, it is clear that a higher perihelion distance should be acceptable from the point of view of extended observation of the same region of the solar surface.

2.4 Space Segment Requirements: Pointing

POIN05 – The Solar Orbiter spacecraft shall be able to perform orbit and attitude control manoeuvres, necessary for:

• Navigation • Orbit maintenance • Science operations.

This requirement must be considered within the context of the overall mission operations, most crucially the requirement to perform delta V manoeuvres (pre/post GAM corrective and therefore in a priori unknown directions) below the Sun distance below which the reference sun-pointing attitude must be maintained, as defined in POIN10 below – this is hereafter referred to as the Thermal Cut-Off Limit - TCOL). Accordingly the spacecraft must have force authority capable of accessing any direction whilst maintaining the Sun-pointing attitude (note that roll around the Sun-line is permissible). The thruster layout on the spacecraft has been selected to provide this capability. POIN10 – In it’s operational orbit, the Solar Orbiter spacecraft shall be 3-axis stabilised with it’s optical reference axis (+Xopt) pointing to the Sun. During the operational orbit, periodic RS-observations of the Sun shall take place. Furthermore, the entire operational orbit of the Solar Orbiter spacecraft covers a range of Sun distances from 0.23 AU up to nearly 0.8 AU. These distances are all below the TCOL; accordingly it is vital that the spacecraft maintains the reference sun-pointing in order to ensure survival. Thus whilst satisfaction of POIN10 is of course driven by the observation requirements, it is more strongly driven by survivability. A key-component of the Solar


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Orbiter spacecraft design is therefore the HW FDIR [IR28], which ensures the maintenance of sun-pointing in the event of all credible failures. POIN15 – For the needs of science, the spacecraft shall be able to point the Xopt axis to different positions on the Sun within a cone of +/-1.25 degrees of the Sun-centre. At ~0.2 AU the angle subtended by solar disc is 1.3°. The requirement of 1.25° therefore allows pointing of Xopt to any region of the solar disc for distances greater than 0.213 AU. The current baseline perihelion is at 0.2343 AU so this 1.25° is compatible with this baseline. This requirement is of course a contributor to the off-pointing budget of the spacecraft, and accordingly the requirements placed on the heatshield, which must provide adequate protection to all sensitive spacecraft elements. For a description of the resultant heatshield sizing, please refer to [IR39]. POIN20 - For the In-Situ instruments Yopt axis must be within the orbital plane +/- 5 deg TBC. This requirement restricts the attitude of the SC around the sun-line such that the IS-instruments are provided with their desired FoV (e.g. to restrict SWA field of views to their minimum and maintain many of the IS-payloads with a suitable orientation with respect to the ram direction). This pointing constraint has been taken into account for all calculations of Earth position in the Physical Reference Frame (PRF) of the spacecraft for the purposes of communication trade studies (see [IR16]). However it should be noted that there are several situations where it is necessary for this requirement to be violated, primarily related to the occultation of the Earth position by the Solar Arrays, the RPW antennas and the spacecraft body itself (this is partially a function of the length of the HTHGA boom). In these instances a temporary roll of the spacecraft around the sun-line will be required in order to ensure a clean view between the HTHGA and the earth. For a full description of the degree of violation of POIN20 required, please refer to [IR16] which presents an analysis of the roll profile of the spacecraft over the 2017 and 2018 mission scenarios. Because of this unavoidable violation, a change to this requirement is suggested; please refer to [IR43] for a description and justification of this change. POIN30 – The Solar Orbiter spacecraft shall fulfill the pointing accuracy (error) and stability (drift) requirements listed. The requirements are given as single axis rotation angles together with their associated 2-sigma deviation.

Table 2.4-1: Solar Orbiter Pointing requirements during operations

For a summary of the projected performances of the SC in response to these requirements, please refer to [IR19] and the budgets section of this report. Note also that the reported non-compliance of the APE during thermal case D1 (caused by a localised cooling due to the Transponder radiators) could be avoided by leaving the Transponders on during close-approach to the Sun – thereby reproducing the deformation seen during thermal case D2) – this will be verified in future work. It is understood that this requirement is only active within the science windows of the mission, i.e. when the RS-instruments are operating during the NMP/EMP. Outside of science windows only the IS-instruments are active, and the pointing requirements, particularly the RPE, should be substantially relaxed. This is an important point because Reaction Wheel operations shall be substantially simplified for ex-science window periods, through for example allowing the wheels to operate at higher capacities or at


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certain problematic frequencies. No requirements for these periods have been specified as yet; these should be agreed in future phases. It should also be noted that no requirement on attitude measurement is expressed. This requirement in attitude knowledge should be sized as a function of payload requirements in terms of LoS knowledge for science data processing. The current baseline is to rely on the gyro-stellar on-board estimation. The RPE requirement (stability of 1arcsec/10s in the LoS, 2arcsec/10s around the LoS) is the most critical of the requirements listed in the table above, and calls directly for gyro stellar measurement systems combined with relatively low noise actuators – elastic isomer mounted reaction wheels running below capacity in the case of Solar Orbiter. Even so, satisfaction of this requirement is difficult, and may force the suspension of HTHGA movements (and hence communication) during the observation periods. It also has consequences for the relative orientation of multiple STR heads, which will provide improvements in performance partitioned between the two requirements. The fulfilment of this requirement does not allow achievement of the ultimate pointing stability equirements of the most demanding remote sensing instrument (PHI) over exposure duration, and such performances require an image stabilisation system (ISS) implemented at focal plane level. The SC-level RPE requirement will in fact allow a restriction of the control range required for ISS-implementation within the instrument(s) themselves. A resource-sharing approach should be sought which provides the best compromise for partitioning RPE satisfaction responsibilities between platform and payload, which is optimal at a system level. POIN35 – Pointing errors shall be defined in the ESA Handbook of Satellite Pointing Errors. The APE, PDE and RPE requirements may pose constraints on the navigation system on ground, since the spacecraft shall continuously be pointed at the Sun during the 10 day science windows and shall therefore follow an attitude guidance profile computed by ground. If the pointing requirements expressed above are system requirements including also ground computation errors, the partition shall be made between attitude guidance errors and AOCS errors on board the spacecraft. POIN40 – The Remote Sensing instruments require an accurate knowledge of their relative pointings in order to coordinate observations. The instruments shall be hard-mounted according to a joint pointing policy to be agreed with ESA and the Prime Contractor (TBC). POIN45 - remote sensing instruments shall be co-aligned to better than 2 arcmin The requirement is understood as a 3-sigma requirement. The Solar Orbiter budgets document [IR19] demonstrates that this requirement is satisfied for the various thermal cases corresponding to science window observation periods.

2.5 Communication Requirements

COMS05 - The Solar Orbiter Spacecraft shall provide 2-way (Up-link and Down-link) communication with the ground Up-link in X-band and Down-link in X and Ka bands with simultaneous operation. This requirement is driven by the large amount of data to be downloaded by the spacecraft, as well as the widely varying RF-link conditions (including recurring periods where communication is impossible due to conjunction or enforced stowage of the HTHGA) throughout the mission. COMS10 - The spacecraft shall support a data downlink of 200 kbps at a range of 1AU with the New Norcia ground station based on standard ESA link budget calculations. The requirement has changed during the course of the study to a reference data-rate of 150 kbps; this is a reflection of the decision taken in the previous study phase [IR07] to reuse the BepiColombo HTHGA. The new requirement is taken into account in link budgets, and it has been demonstrated that the reduced reference data-rate does not drive the dimensions of the store and forward data architecture (see [IR16]).


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COMS15 - The Solar Orbiter Spacecraft shall provide communications coverage for all phases of the mission, all attitudes of the spacecraft and for both up- and down-link, with the exception of unavoidable communication outages due to planetary alignments. This requirement stipulates an approach to antenna position and type that ensures sufficient coverage throughout the various mission phases. However, the requirement as worded is impossible to meet given the constraints of the mission:

• EGAM pre/post manoeuvres shall take place in a priori unknown attitudes which may well be incompatible with the range of movement of the HTHGA and MGA, and will take place beyond the range of the LGA antennas

• There is an unavoidable need to periodically stow the HTHGA during hot periods close to the Sun. Accordingly we suggest a rewording of the requirement to capture the dependence upon antenna stowage. COMS20 - The Solar Orbiter Spacecraft shall have an onboard data storage capability, to store command files and house-keeping and scientific data during communication outages or when sufficient communications with ground is not possible. This requirement, in conjunction with MISS20, specifies the need for a store and forward architecture. This is needed due to the widely varying RF-link conditions, including periodic outages, throughout the mission. The sizing of the store and forward architecture has been performed for various operational scenarios in [IR16], and the baseline sizing is presented also in [IR44]. In some cases very large data-latencies are associated with the store and forward architecture, and in future phases of the mission design, it will be necessary to develop a data-management scheme in order to provide the best delivery of mission products to the GS. COMS30 - The Solar Orbiter Spacecraft shall provide the necessary on-board functions to allow the Ground Segment to perform doppler and range tracking plus navigation, for all phases of the mission. This places additional requirements for functionality upon the TT&C subsystem. The TT&C subsystem for Solar Orbiter is inherited almost entirely from BepiColombo, which also has the same requirement. The budgets for the TT&C subsystem also include DOR tone and ranging in response to this requirement (see [IR19] and also the budgets section of this document. COMS35 - The spacecraft shall support the use of delta-DOR to perform tracking and navigation and is required within 30 days prior to a GAM. This requirement should also extend to the support of tracking and navigation in the post-GAM period.

2.6 Functional Requirements

FUNC05 - The spacecraft shall autonomously detect separation from the launch vehicle and initiate spacecraft post-launch configuration actions including but not limited to the following steps:

• Switch-on spacecraft transmitter(s) • The spacecraft shall autonomously reduce any remaining body rates • The spacecraft shall autonomously re-orientate itself to achieve the required solar array alignment

with respect to the sun vector and thus terminate any battery discharge This requirement provides a preliminary specification of the sequence of actions to be taken by the spacecraft upon separation from the launch vehicle; this sequence of events is more fully described in [IR17]. FUNC10 - Following separation from the launch vehicle, the spacecraft shall provide all the necessary attitude information to reconstruct the attitude on ground including a record of all events during the separation sequence.


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FUNC20 - In case of spacecraft failures, the separation sequence shall be such that it will autonomously restart safely at the step where the failure occurred. In the event that a major problem occurs during the initialisation sequence it will be necessary for the automatic sequence to continue and achieve an operational configuration. This is because the SC will have a limited lifetime on the battery and without thrusters control the spacecraft attitude may be such that the array is not pointing towards the sun (although this possibility is minimised by correct selection of the separation attitude and spin-rate). FUNC25 - The automatic separation sequence shall always be overridable by ground command after physical separation from the launcher. The initialisation OBCP will be structured to accommodate failures during its execution. Any problems encountered will be recorded using the on-board services and transmitted to the ground at the first opportunity. Once the automatic initialisation activities are completed and the GS has AOS, the GS is able to control final configuration activities. FUNC30 - Spacecraft Arming shall follow Launch authority requirements and shall allow remote connection of the batteries without any Single Point Failures (SPFs). This is a standard requirement.

2.7 Cleanliness Requirements

CLEA05 - The payload shall not be contaminated from particulate, organic or inorganic elements and shall be compatible with the limits specified, including contamination from out-gassing, pre-launch, and during launch, cruise and operations. The EID-A contains recommendations for cleanliness. CLEA10 - The spacecraft shall conform to the magnetic cleanliness limits specified. The Solar Orbiter Cleanliness Requirements and Implementation Plan is TBW.

2.8 Launch Requirements

LAUN05 - The Solar Orbiter Spacecraft shall be designed to be launched by an Atlas V launch vehicle from Kennedy Space Center (KSC), USA. The Soyuz-ST 2-B /Fregat launcher from Kourou is the back-up launcher (TBC). The Atlas V and Delta IV launch vehicles have a significantly greater capacity (both volumetric and mass) than the Soyuz launch vehicle. However, even though it is the backup launch vehicle, the Soyuz is the sizing case for determining allowable mass and volume. An important point to note here is that the actual performance of Soyuz-ST 2-B from Kourou is unknown at present, and some considerable uncertainty exists, particularly for launches into high declinations such as is the case for Solar Orbiter. The current 2018 backup mission scenario from [IR26] specifies a launch negative declination of ~53°, which is extremely high and in the opinion of Astrium Ltd unfeasible for a Kourou launch to deliver a useful mass using the Soyuz vehicle. Performance estimates were produced for this study based on the latest information, giving a PSW of 1301kg (i.e. including adaptor) into the target injection for Solar Orbiter (see [IR16 Appendix]). However, the projected performance for the Soyuz has been agreed with ESA to be 1318kg, a figure taken from the projected performance used in the previous Heatshield and System Technology Study [IR07]. Furthermore the launch environment of the Soyuz-Fregat launcher is comparatively benign (this is a man-rated launch vehicle, and therefore has lower design limits to the forces experienced during launch); the switch to the Atlas/Delta as the baseline has an impact on the structural mass of the SC, because the launch loads experienced are more severe.


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LAUN10 - The Spacecraft maximum launch mass in launch configuration shall not exceed 1250 kg . LAUN15 - The Solar Orbiter Spacecraft shall allow a Launch Window of at least 20 consecutive days. LAUN05, LAUN15 and LAUN10 are incompatible, because they imply a different mass specification; currently we are using an agreed Soyuz performance as specified above. As Soyuz-Fregat performance information from Kourou is updated, this will be reflected in the mass budget for Solar Orbiter. However, it must be stressed that the requirement to maintain back-compatibility with Soyuz-Fregat for the SC design has had a profound impact on the design. LAUN20 - During the integration phases with the launcher, the Solar Orbiter spacecraft shall be compliant with the launcher operations and attitude. This requirement has bearing particularly on the propulsion subsystem design, which assumes vertical integration of the subsystem into the platform (see [IR45] and the propulsion section of this report for a description of the CPS design). LAUN25 - During the launch and ascent phases when attached to the launcher, the Solar Orbiter spacecraft shall be compliant with the attitude profile of the launcher, with respect to the Sun and the Earth. The experience gained on the MEX/VEX missions shall be very useful in this regard. LAUN30 - The spacecraft design and its ancillary equipment shall be compatible with the facilities at the launch site. LAUN35 - Prior to lift-off the spacecraft shall be in an electrically active state. Thus the on-board TT&C and Data Handling subsystems shall be in an operational mode that is able to receive telecommands, handle telemetry packets and perform on-board monitoring functions through an umbilical connection. The Solar Orbiter transponder and OBC shall be powered on whilst connected to the launcher.

2.9 Operations Requirements

SMOC05 - There shall be a Solar Orbiter Mission Operations Centre (MOC) at ESOC that shall be responsible for spacecraft operations. SSOC05 - There shall be a Science Operations Centre (SOC) at ESAC (TBC). Solar Orbiter shall use the common operational model developed during the MEX/VEX and Rosetta missions. 2.10 Ground Station Requirements

GRDS05 - During LEOP the ESA stations at New Norcia and Cebreros will be used. The launch scenario is for a direct injection; consequently the ground station visibility during LEOP will be very good for Solar Orbiter. This shall be specifically analysed in future design phases. GRDS10 - During the first 3 months after launch full daily visibility from New Norcia will be provided to support commissioning operations. Commissioning activities during the NECP are expected, and thus daily communication passes are necessary to support these activities. During the commissioning period, the MOC is responsible for instrument commissioning, in-line with the requirements of the ESOC common operations approach.


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GRDS15 - During cruise phase, science operations will be limited to in-situ instruments only. New Norcia ground station shall support 3 passes per week, each of them with a minimum of 4 hours of telemetry dump. Check-out and completion of remote sensing instruments will be performed during this phase. Although this requirement specifies operation only of the IS-instruments during the CP, it is expected that at some point, RS-instrument teams will request RS-instrument operation during the CP; SSMM-capacity analysis suggests that this will not drive the required SSMM capacity and therefore should be compatible with the mission design as it currently stands. However, we of course only design to the science windows as defined in the PERF20 of the MRD. GRDS20 - During nominal phase, New Norcia station shall provide daily visibility during passes of 6 hours for effective 4 hours telemetry dump. This daily pass has been used as a key input into the store and forward modelling, in order to provide a sizing of the SSMM required to store and forward all the data. GRDS25 - New Norcia and Cebreros stations shall be used to perform delta-DOR during the preparation and execution of gravity assisted maneuvers. Other delta-DOR baselines can be used, if necessary. Around each swing-by event, both pass frequency and duration shall be increased for a few weeks as necessary. Solar Orbiter will periodically employ GAM both during the CP and also in the NMP/EMP; accordingly pass duration and frequency shall need to be increased for these critical periods. It is important to note that for certain trajectories explored (namely the former 2015 baseline), extended conjunction periods (~2 months) exist within which GAMs occur. This is problematic operationally because no communication is available in the month prior to the GAM. Fortunately the 2017 baseline launch scenario does not suffer from this problem. However, it is now confirmed that the 2018 backup scenario, like the original 2015 scenario, also suffers from this problem, with a non-communication period of CHECK days coinciding with VGAM-4. It is therefore recommended that any additional mission analysis activities that occur within the frame of the Solar Orbiter design should take the additional constraint of conjunctions coincident with GAMs into account.

2.11 Ground Operations Environment

GENV05 - Assembly, Integration and Testing (AIT) of the spacecraft shall be performed in a 100,000 -Class clean room. A standard requirement.

2.12 Launch Environment

LENV05 - The launch environment shall comply with the specification of the Atlas V [NR12] and Delta IV [NR17] user’s manual. LENV10 - The Solar Orbiter spacecraft shall be compatible with the launch environment of Soyuz-Fregat’s user manual [NR07]. The more punitive launch environment of the three cases has been used in each case when considering the various environmental parameters (QSL, SRS, acoustic, pressure-decay etc…) imposed upon the spacecraft during the Launch Phase. A summary of the launch environments for the three launchers is given in [IR10]. FENV05 - The Solar Orbiter system shall be compatible with the flight environment specified.

2.13 General System Engineering Requirements


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GSYS05 - System Engineering requirements provided by ECSS-E-00A and the detailed requirements provided by lower level standards defined in ECSS-E-00A shall apply (TBC). GSYS10 - Detailed requirements from the applicable Level 2 and Level 3 ECSS standards shall be tailored according to the characteristics of the mission. GSYS15 - Tailoring of ECSS standards requirements shall be performed. The Solar Orbiter mission design is compliant with ECSS standards.

2.14 Coordinates

COOR05 - The Spacecraft Co-ordinate Systems are axis reference frames physically attached to the spacecraft and shall be right-handed orthogonal triads. A standard requirement. COOR10 - The Solar Orbiter Physical Reference Frame shall be as defined in Table 6-1. Item Definition

Origin Point of intersection of the launcher longitudinal axis (+X_lv) with the separation plane between the launcher and the composite

+X_so Longitudinal axis of Solar Orbiter, from the origin towards Solar Orbiter, positive upwards (launcher in vertical position), coinciding with the +X axis of the launcher (+X_lv)

+Y/Z_so Orthogonal axes completing the orthogonal triad such that Z_so = Y_so*X_so, with the +Y_so axis aligned with the velocity vector of the SC at perihelion and aphelion

Table 2.14-1: Solar Orbiter Physical Reference Frame

This specification of the Solar Orbiter reference frame is presented pictorially in the following figure.

Figure 2.14-1: Solar Orbiter Physical Reference Frame

COOR15 - The Solar Orbiter Optical Reference Frame shall be as defined in Table 6-2.


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Table 2.14-2: Solar Orbiter Optical Reference Frame Definition

2.15 Mass Margins

MASS05 - The Total Mass at launch of the spacecraft shall include an ESA system-level mass margin of >= 20% of the Nominal Mass at launch. MASS10 - The ESA system-level mass margin shall:

• be visible and traceable in the overall mass budget of the spacecraft composite • be maintained as an overall mass margin under agency control • not include any propellant residuals or unused fuel.

MASS15 - The Nominal Mass at launch shall include the design maturity mass margins to be applied at equipment level. The Nominal Mass at launch does not include the ESA system level mass margins. MASS20 - The Basic Mass at launch shall include neither the ESA system-level mass margin, nor the design maturity mass margins to be applied at equipment level and shall represent best current estimate of the equipment. MASS25 - The design maturity margins given in Table 6-3 shall be applied to the Basic Mass at equipment level. The mass margin philosophy described by MASS05 – MASS25 has been implemented from the start in the Solar Orbiter Definition Phase study; the hiearchy of mass margins is shown in the following figure.

Figure 2.15-1: Mass margin philosophy

Some changes have been made (in agreement with the ESA Project Team) within the frame of the ECR to the current study [IR49, and MoM PM7], namely that a 30% system margin shall be applied to the baseline Atlas/Delta launch case, with a 0% system margin to be applied to the backup Soyuz-Fregat launch case. It should be clearly noted that Astrium Ltd do not consider the Soyuz-Fregat launch vehicle to be a viable option for this mission due to the extremely high declination that is currently required for the launch by the mission analysis [IR26] – 20 degrees for the 2017 scenario and -53 degrees for the 2018 scenario. Please see the Appendix for an elaboration of these concerns. An important additional consideration for the mass of BepiColombo units has been followed. This is necessary because the Solar Orbiter mission reuses to a high degree BepiColombo equipment, and much of this equipment is still itself under development. Accordingly the Solar Orbiter Basic Mass of these


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equipments is defined to be their BepiColombo Basic Mass + Equipment Margin. This ensures that any margin-accounted mass growth that occurs in the BepiColombo programme at unit-level is already accounted for in the Solar Orbiter mass budget. This has had the unfortunate effect of significantly contributing to the current non-compliance of the design with the mass requirements.

2.16 Propulsion Margins

PROP05 - The propulsion budget shall contain the following velocity increment margins (covering “On The Flight” dispersions and contingencies such as: Launcher dispersions; Manoeuvre inaccuracies; Navigation errors), which shall be applied to the Effective velocity increment manoeuvres:

• 5 % for trajectory manoeuvres • 100 % for the orbits maintenance manoeuvres, over the specified lifetime • 100 % for the attitude control and angular momentum management manoeuvres.

These margins have been applied to the propulsion budget. PROP10 - When budgets of the manoeuvres concern theoretical values, and do not take into account gravity losses (for instance: impulsive manoeuvres performed by chemical propulsion engines), such gravity losses shall be quantified and added to the specified effective velocity increment. The delta V manoeuvre budgets produced for Solar Orbiter do not suffer from gravity losses, since all manoeuvres are undertaken well away from gravitational bodies. PROP15 - The spacecraft shall be able to perform orbit and attitude control manoeuvres, necessary for orbit maintenance during the specified lifetime. Essentially the spacecraft requires 6-DOF control, most crucially in order to provide the capacity to access random delta V directions whilst maintaining an inertially-fixed platform attitude (in order to perform delta V manoeuvres whilst pointing at the Sun). This key requirement, coupled with no gravity losses, led to the selection of a CPS RCS with no main engine. PROP20 - The propulsion modules of the spacecraft shall be designed and sized for the Maximum Separated Mass (total spacecraft launch mass excluding adaptor). It is clearly understood that the propulsion subsystem sizing of the spacecraft must take into account the full separated wet mass of the spacecraft. PROP25 - The volume of the tanks of the propulsion modules shall be sized for the Maximum Separated Mass plus at least 10 %. This sizing constraint has been implemented in the design of the propulsion subsystem; the selected tanks for the Solar Orbiter spacecraft are OTS from Gaia, in-keeping with the goal of cost-minimisation. These tanks have a considerable capacity surplus to account for any growth in the mass of the spacecraft. PROP30 - For each Gravity Assist Manoeuvre (GAM), an allocation of 15 m/s shall be added to the Delta V budget, to account for preparation and correction of these manoeuvres. The delta V budget [IR19] includes 8 GAM, each allocated 15m/s.

2.17 Power Margins

POWM05 - At equipment level and for conventional electronic units, the following design maturity power margins shall be applied:

• 5 % for “Off-The-Shelf” items (ECSS Category: A/B) • 10 % for “Off-The-Shelf” items requiring minor modifications (ECSS Category: C) • 20 % for new designed / developed items, or items requiring major modifications or re-design

(ECSS Category: D).


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These standard margins have been applied in the design of Solar Orbiter. POWM10 - The total power budget of the spacecraft shall include an ESA system level power margin of 20% of the nominal power requirements of the spacecraft. POWM15 - The ESA system level power margin shall be visible and traceable in the overall power budget of the spacecraft. POWM20 - The nominal power requirements include the power requirements of all spacecraft elements (payload and platform, including their respective design maturity power margins to be applied at equipment level), and does not include the ESA system level power margin. POWM25 - Solar arrays and batteries shall be sized to provide the spacecraft required power, including all specified margins for the appropriate needs:

• from launch until the end of the cruise phase • from end of cruise phase until end of nominal operations • during the extended operations phase.

The power margin philosophy described by POWM05 – POWM25 has been implemented from the start in the Solar Orbiter Definition Phase study; the hiearchy of power margins is shown in the following figure.

Figure 2.17-1: Power Margin Philosophy

An important additional consideration for the power of BepiColombo units has been followed. This is necessary because the Solar Orbiter mission reuses to a high degree BepiColombo equipment, and much of this equipment is still itself under development. Accordingly the Solar Orbiter Basic Power of these equipments is defined to be their BepiColombo Basic Power + Equipment Margin. This ensures that any margin-accounted power growth that occurs in the BepiColombo programme at unit-level is already accounted for in the Solar Orbiter power budget.

2.18 Data Margins

DATA05 - On-board memory (RAM used for code and/or data) shall include a memory margin of at least 50%. DATA10 - On-board processor peak usage shall not exceed 50% of its maximum processing capability. This has to be verified in future phases.

2.19 Communications Margins

COMS70 - Links (up and downlink) budgets and associated margins, for all phases of the mission, shall be computed as defined, including: nominal, adverse, favorable, mean –3 sigma and worst RSS (Root Sum Square) cases. COMS75 - Telecommand and telemetry data rates shall be satisfied with minimum margins as defined, for all phases of the mission, under all cases specified in COMS70.


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The link budgets for Solar Orbiter (see [IR19] and budgets section of this document) have been defined in accordance with these requirements.

2.20 Thermal Margins

TEMP05 - The different temperature ranges (calculated, predicted, design, acceptance, qualification) shall have their respective uncertainties and margins:

• Uncertainties: to be analytically calculated, according to ECSS-E-30 Part 1A, section A.1.3 and fully justified by sensitivity analysis for all items

• Acceptance Margin as shown in Fig 6-4: >5 K (ECSS-E-30 Part 1A, sec. A.2) • Qualification Margin as shown in Fig 6-4: >5 K (ECSS-E-30 Part 1A, sec. A.2).

2.21 Heatshield

HEAT05 – The heatshield sub-system shall be fully testable as a sub-system prior to integration with the spacecraft. This is a key requirement given the criticality of the heatshield as a technology component of the mission. Accordingly, the design of the heatshield (see [IR22, IR23] and the heatshield section of this document) is such that it can be integrated as a complete unit into the spacecraft. The heatshield has a separate support panel structure such that it is a stand-alone item. HEAT10 - For the heat shield sub-system, requirement TEMP05 shall apply with the following levels:

• Uncertainties: > 20 K • Acceptance Margin: > 20 K • Qualification Margin: > 30 K.

This specific margin philosophy has been accounted for in the heatshield design.

2.22 Software

SOFT05 - Spacecraft software shall be coded as Application Programs. SOFT10 - The spacecraft shall support On-Board Control Procedures (OBCP) capability. The foreseen Solar Orbiter design shall use a MTL format with time-tagged telecommands and OBCPs.

2.23 Instruments

INST05 - The instruments shall comply with the EID-A requirements. The mission design has proceeded according to the payload interface definition provided in the EID-A [NR10]. The above requirement ensures that each payload EID-B is eventually compliant with the EID-A, in turn demonstrating compliance with the mission design. INST10 - The Solar Orbiter Spacecraft shall allow the full operations of the In-situ payload throughout the nominal and extended operation phases. This implies completion of the IS-instrument commissioning activities by the end of the NECP. INST15 - The Solar Orbiter spacecraft shall allow the operation of a sub-set of the Payload Instruments during the interplanetary cruise phase when compatible with standard spacecraft operations. This is the same as INST10.


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INST20 - Operation of payload instruments for calibration and validation use shall be possible during the cruise phase. Some commissioning activities (particularly of the RS-suite) are foreseen for the early stages of the CP. INST25 - The Solar Orbiter Spacecraft shall allow the Payload to be operated in the following modes:

• High Data-rate Mode, whereby all instruments of the Remote-Sensing and In-Situ packages can be operated simultaneously and acquire data at their standard resolution

• Medium Data-rate Mode, whereby all instruments of the In-Situ package and a subset of the Remote-Sensing package can be operated simultaneously and acquire data

• Low Data-rate Mode, whereby all instruments of the In-Situ package can be operated and acquire data (remote sensing instruments are in standby/off mode)

• Standby/Off mode, where all instruments are not generating science data • Burst Mode, where selective instruments are in burst mode.

Preliminary collation of the data-rates of these various modes is held within [IR33]. FOPS05 - The Solar Orbiter spacecraft shall perform autonomously nominal operations when ground intervention is not possible. The requirement for autonomy is high on Solar Orbiter, because there are frequent periods where communication with the Earth will not be possible – see [IR17] for a definition of these periods for the baseline 2017 and backup 2018 mission scenarios. Furthermore there is a large single conjunction period in every mission scenario during which no communication will be possible. FOPS10 - The Solar Orbiter spacecraft shall remain safe for a period of at least 30 days (TBC), without ground intervention. The sizing period of non-communication, caused by conjunction with the Sun, is actually 79 days for the 2017 baseline trajectory [IR17]; accordingly it is recommended that this requirement is discussed and reformulated. FOPS15 - Operation of Solar Orbiter during all mission phases shall be possible using either direct communication with Ground or via on-board autonomy and on-board data storage followed by periodic data transfer to Ground. The foreseen Solar Orbiter design shall use a MTL format with time-tagged telecommands and OBCPs. FOPS20 - The satellite shall respond to on-board failures by switching, independent from ground control, to redundant functional path. Where this can be accomplished without risk to satellite safety such switching shall enable the continuity of the mission timeline and performance. In the event that alternative redundant paths do not exist or that the failure effect is too complex to allow autonomous recovery, the satellite shall enter Survival Mode. SOPS05 - Nominal operations of science instruments shall be pre-planned TBD weeks in advance. Minimal planning is expected concerning the IS-instruments which are on at all times once commissioned (there could be cessation of observations in order to allow calibration). RS observations are expected to be more complex for the following reasons:

• Different regions of interest on the Solar disc, which are themselves a function of a priori unknown solar events

• Competing instrument priorities • Door opening and closing sequences.


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No detailed planning has been considered in this phase; once more information on likely instrument scheduling and planning is available, it should be considered immediately within the context of the operational design of the Solar Orbiter mission. SOPS10 - Instrument data shall be available for instrument teams TBD hours after ground receipt. This is a requirement on the SOC. AIVR05 - Verification requirements provided by ECSS E-10-02A shall apply. A standard requirement. AIVR10 - System level spacecraft verifications shall include agreed mission representative test cases prior to and after environmental testing. A standard requirement.


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3. SOLAR ORBITER SPACECRAFT DESIGN OVERVIEW

3.1 Design Heritage

The design presented in this document is the result of a design process extending back to the initial CDF studies in 1999 and 2002 [IR01]. Since the beginning of industrial study, the principal list of documents that supports the Solar Orbiter design as presented in this document is as follows: Assessment Study Final Report [IR05] Assessment Study CCN Final Presentation [IR06] Heatshield & System Technology Study System Design Report [IR07] Heatshield & System Technology Study CCN1 and CCN2 Reports [IR11, IR13] Solar Orbiter System Trade Offs [IR16]. Note that this document does not describe in detail any of the supporting trade studies and/or analysis that has been performed during the study; rather it simply provides a snapshot of the mission and spacecraft design as they currently stand. Please refer to the supporting documentation for more information concerning the rationale behind the design presented. The following figure shows the evolution of the design from the initial CDF study through to the design at the end of the previous study phase.


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Figure 3.1-1: The evolution of the Solar Orbiter spacecraft design from the initial and second CDF

studies, through the initial EADS Astrium industrial study, to the EADS Astrium design at the end of the Heatshield System Technology Study


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3.2 Design Overview

The Solar Orbiter design described in the following sections is a response to the MRD [NR01] and SRS [NR02] requirements. The functional architecture of the Solar Orbiter spacecraft is shown in the following figure.

AOCS

DMS

Originator: H.J. Koenig, EADS Astrium GmbH

Electrical Architecture and Interface Diagram

Doc.No.: SO-ASD-DW-00001 Issue 1 Rev 1

Date: 15.05.2009

Equipments

APM(E) Antenna Pointing Mech. (Electronics)OBC On-Board ComputerCSS Coarse Sun SensorDST Deep Space TransponderFCE Failure Control ElectronicHD Hold-Down (Release Mechanism)HGA High Gain AntennaIMU Inertial Measurement Unit (Gyro’s)

KAT Ka-Band TranslatorLGA Low Gain AntennaLV Latch ValveMGA Medium Gain AntennaPCDU Power Control & Distribution UnitPT Pressure TransducerPV, ISV Pyro Valve, Isolation Pyro ValveRFDA RF Distribution AssemblyRIU Remote Interface Unit

RMU Rate Measurement UnitRW Reaction WheelSADM Solar Array Drive MechanismSADE Drive Mechanism ElectronicsSSMM Solid State Mass MemorySTR Star TrackerTHR ThrusterTWTA Travelling Wave Tube AmplifierWGS Waveguide Switch

SOLO Solar OrbiterSH_IF Sunshield and I/F StructureAOCS Attitude Orbit Control SystemDMS Data Management SystemEPS Electrical Power SubsystemCPS Chemical Propulsion SubsystemTCS Thermal Control Subsystem

Solar OrbiterAPR Array Power RegulatorCPDU Command Pulse Distribution UnitEPC Electronic Power ConverterHPCG High power Command GeneratorMPPT Maximum Power Point TrackerPM Processor ModuleTFG Telemetry Frame Generator

Internal Functions & Interfaces/SignalsComposites & Subsystems Legend:

MIL-1553B Bus Termination

internally redundantMIL-1553B Bus Coupling Transformer (long stub)

Latching Current Limiter (LCL): Current-TM,ON/OFF-Cmd/StatusSwitch (Solid state or relay)

Current Source

Foldback Current Limiter (FCL): Current-TM

DC/DC Converter (galvanic isolation)

C

Z

BC Bus Controller (MIL-1553B)RT Remote Terminal (MIL-1553B)FCV Flow Control Valve Control

UP: upstream, DN: downstreamHTR HeaterSpW SpaceWire

Power and Signal OR-ing (e.g. using diodes)

Test-/Skin-Connector

Z

Skin Conn.

Solar Orbiter Payload

Data R. ~ 13 KpbsMAG

Status/Temp

SpW 01 A+B

Sync A + B

LCL A + B

Cmd/Stat

Data R. ~ 20 KbpsEUI

Status/Temp

SpW 02 A+B LCL A + B

Cmd/Stat

Data R. ~ 4 KbpsEPD

Status/Temp


Cmd/Stat

Data R. ~ 20 KbpsPHI

Status/Temp

Spw 04 A+B LCL A + B

Cmd/Stat

Data R. ~ 0,2 KbpsSTIX

Status/Temp


Cmd/Stat

Data R. ~ 30 KbpsRPW

Status/Temp


Cmd/Stat

Data R. ~ 20 KpbsSoloHI

Status/Temp


Cmd/Stat

Data R. ~ 17 KbpsSpice

Status/Temp


Cmd/Stat

Data R. ~ 14 KbpsSWA

Status/Temp


Cmd/Stat

Data R. ~ 10 KbpsMETIS

Status/Temp


Cmd/Stat

MAGBoom HTR Pwr

Status/Temp

HD1

LCL (A+B)

LCL (A+B)

LCL (A+B)

LCL (A+B)

LCL (A+B)

LCL (A+B)

LCL (A+B)

LCL (A+B)

LCL (A+B)

LCL (A+B)

LCL (A+B)

SSMM

Memory Mod. 1

(256 Gbit)

Out

put-

IF B

Memory Contr. A

Memory Contr. B

PS B

DC

/DC

MC

B

Memory Mod. 2

(256 Gbit)

Memory Mod. 3

(256 Gbit)

Out

put-I

F A

O u t p u t C r o s s -S t r a p

I n p u t C r o s s-S t r a p

DC

/DC

MM

&IO

B

PS A

DC

/DC

MM

&IO

A

DC

/DC

MC

A

Inpu

t-IF

A

Payl

oad

Spa

cew

ire IF

s A

Inpu

t-IF

B

Payl

oad

Spac

ewire

IFs

B

R

R

R

R

R

R

R R

Launcher-IFs EGSE I/F

Communication SystemRFDA

X -Dipl.1

X-Dipl.

2

X-MGA

X-LGA 1WGS

-4

WGS-2

WGS-1

WGS-3

WGS-5

X-LGA 2

MGAPM(3-axis)

Drv

A+B

(3-a

xis)

X-RFI-2

X-RFI-1

Ka-RFI-2

Ka-RFI-1

X/Ka-HGA

X/Ka-Bd

Feed

HGAPM(2-axis)

Shor

t

WGS-6

3 dB Coupler

X-TWTA-2

EPC

X-TWTA-1

EPC

Ka-TWTA-1

EPC

Ka-TWTA-2

EPC

Drv

A+B

(2-a

xis)

LCL(B)

TC/TM

Ka-Coupler

LCL(A)

TC/TM

LCL(B)

TC/TM

LCL(A)

TC/TM

PYR (A+B) HDs

HDs PYR (A+B)

Status (A+B)

Status (A+B)

APME

DC

/DC

DC

/DC RT

N R

RT

N R

HGA Safe Pos.

MGA Safe Pos.

X-TC (Test)

X-TC (Test)

CPS

Tanks

PRESSURE TRANSDUCERs Low Pressure: PT-1 to PT-2High Pressure: PT-3

AOCS Thrusters (T1)10N Dual Seat Solenoid Valve

THR2-1A to -8A (prime)THR2-1B to -8B (red.)

PV-01 to -06

NORMALLYCLOSED

PV-07,-08

NORMALLYOPEN

ISV-01/02

NORMALLYCLOSED

PYR (A+B)

PYR (A+B)

EGSE I/F

(Skin)

RIU

OB

C IF B

OB

C IF A

OBC

IF BO

BC

IF A

Controller A&B

STD I/O

A

EHPANP (24)ANY (80)AN2 (12)AN1 (8)BLD (12)RSA (48)SHP (64)PTA (8)ANTLCL DC/DC

STD

I/O B

EHPANP (24)ANY (80)AN2 (12)AN1 (8)BLD (12)RSA (48)SHP (64)PTA (8)ANTCD/DC

STD I/O

C

CD/DC

ANP (24)ANY (80)AN2 (12)AN1 (8)BLD (12)RSA (48)SHP (64)

CP

S-I/O

A

1 CHTVC1VC

FCVDrivers

LV DriversLVC

LCL DC/DC

LCL DC/DC

CPS

-I/O

B1 CHTVC1VC

FCVDrivers

LV DriversLVC

LCL DC/DC

LCL DC/DC

FCE

Controller A

STD I/O

A

CUR_ABLD

STD I/O

B

CUR_ABLD

RSA

SHP

TFG VC1TFG VC2TFG VC3

TFG VC1TFG VC2TFG VC3

OBC

Processor Module (A)

Processor Module (B)

OBC Mass Memory 2

Service Mode Link

OBC Mass Memory 1SS

MM

A IF

SSM

M-B

IFM

ETI

S In

str.

RIU

-B IF

s

Pro

ram

E

EPR

OM

1P

rora

m

EE

PRO

M 2

EPCCOMMS

BCN R

AOCSBC

N R RIU

-A IF

MET

IS In

str.

SS

MM

-B IF

SS

MM

A IF

RIU

-B IF

sR

IU-A

IF

EPSCOMMS

BC

N R

AOCSBC

N R

DC/DC A (Cold)

DC/DC B (Cold)

S/C Elapsed Time (SCET)

Reconfiguration Module (RM)

Alarm Conditioning

Safeguard Memory (SGM)

High Priority Cmd Generator (HPCG)

(A) (B)OCXO

OCXO

Ana Alarms in (A+B) 3 x 6

AOCS Puls Signal (A+B) 8

HPC (A+B) 64 Cmd

RSA (A+B) 64 Status

Dig Alarms in (A+B) 3 x 6

Puls per Sec (A+B) 8

TC-Decoder & CPDU

TFGs X-Bd/Ka-Bd

(A)DC/DC A (Hot)

TM/TC Bypass

X-TM – 1A / 2AKa-TM – 1A / 2A

X–TC – 1A / 2ASpW VC2 X-Band Nom

SpW VC3 Ka-Band Nom

SpW VC1 X-Band Nom

TC-Decoder & CPDU

TFGs X-Bd/Ka-Bd

(B)DC/DC A (Hot)

TM/TC Bypass

X-TM – 1B / 2BKa-TM – 1B / 2B

X–TC – 1B / 2BSpW VC2 X-Band Red

SpW VC3 Ka-Band Red

SpW VC1 X-Band Red

Controller B

LCL

DC

/DC

LCL

DC

/DC

WG

S-1

A P

os 1

WG

S-1

B P

os 1

WG

S-1

A P

os 2

WG

S-1

B P

os 2

WG

S-1

A S

tatu

sW

GS-

1 B

Sta

tus

WG

S-2

A P

os 1

WG

S-2

B P

os 1

WG

S-2

A P

os 2

WG

S-2

B P

os 2

WG

S-2

A S

tatu

sW

GS-

2 B

Sta

tus

WGS-1 WGS-2

WG

S-3

A P

os 1

WG

S-3

B P

os 1

WG

S-3

A P

os 2

WG

S-3

B P

os 2

WG

S-3

A S

tatu

sW

GS

-3 B

Sta

tus

WGS-3

WG

S-4

A P

os 1

WG

S-4

B P

os 1

WG

S-4

A P

os 2

WG

S-4

B P

os 2

WG

S-4

A S

tatu

sW

GS

-4 B

Sta

tus

WGS-4

WG

S-5

A P

os 1

WG

S-5

B P

os 1

WG

S-5

A P

os 2

WG

S-5

B P

os 2

WG

S-5

A S

tatu

sW

GS

-5 B

Sta

tus

WGS-5

WG

S-6

A P

os 1

WG

S-6

B P

os 1

WG

S-6

A P

os 2

WG

S-6

B P

os 2

WG

S-6

A S

tatu

sW

GS-

6 B

Sta

tus

WGS-6

DST-1

Ka-

Ban

d Tr

ans.

A

X-B

and

Tran

s.A

LCL DC/DC

X-B

and

Rec

eive

r/D

emod

A

FCL DC/DC

Ka-TM 1AKa-TM 1B

RTNR

X-TM 1AX-TM 1B

X-TC 1AX-TC 1B

Discrete TM/TC

DST-2

Ka-

Ban

d Tr

ans.

B

X-B

and

Tran

s.B

LCL DC/DC

X-B

and

Rec

eive

r/D

emod

B

FCL DC/DC

Ka-TM 2AKa-TM 2B

RTNR

X-TM 2AX-TM 2B

X-TC 2AX-TC 2B

Discrete TM/TC

1553B-Bus A+B

1553B-Bus A+B

10 N Dual Seat ValvePrime Branch

10 N Dual Seat ValveRedundant Branch

NT 01

HePressurant

MMH 2

1A 2A 3A 4A 5A 6A 7A 8A 1B 2B 3B 4B 5B 6B 7B 8B

PV 02 PV 01PT3H/P

FW 01

TP 02

TP 03

FW 05

FDV 07

TP 09F 1

FO 01FO 02

F 2

PT 1L/P

PT 2L/P

ISV1AISV2BISV2A ISV1B

TP 10

FDV 08

FW 06

TP 04

PV 04 PV 06 PV 03 PV 05

PV 08 PV 07

NRV 4

NRV 2 NRV 1

NRV 3Orifice

F3

IMU

Red.Elec.

MainElec.

LCL DC/DC

4 Gyro & Acc.

RTNR

RTNR

Discr. TM/TCSync. (tbc)

Discr. TM/TCSync. (tbc)

LCL DC/DC

C(2)C(6)

C(2)C(2)

C(2)C(2)

C(2)C(8)

ZZ

C(2)C(2)

C(2)C(2)

C(2)C(2)

C(2)C(2)

ZZ

C(2)C(4)

C(2)C(4)

C(2)C(4) ZZ

CP

S A

rmin

g (S

kin-

CB

s)

Tank Temp

LCL (PT-x)

PT-x Press

LV-1 to LV-4OPEN CoilCLOSE Coil

Valve Status

LV-xy Open/CloseLV-xy Status

T2-xy US FCV

T2-xy DS FCV

T2-xy Temp

CB Temp

CB HTR

OBC

-A IF

OBC

-B IF

OB

C -A

IF

OB

C -B

IF

Skin-Arm/-Safe

Pyro

Firi

ng (A

+B)

PYR (A+B)

PYR (A+B)

Sol

ar A

rray

Sim

ulat

or I/

F

Internal Functions & Interfaces/Signals

RSA

SHP

STR-1 LCL DC/DC

RT NR

Sync

Discr. TM/TC

STR-2 LCL DC/DC

RT NR

Sync

Discr. TM/TC

STR-3 LCL DC/DC

RT NR

Sync

Discr. TM/TC

EPS

SADE 1

DC

/DC

DC

/DC

RT

N R

RT

N R PCDU

Battery (Li-Ion)Cell redundancy

Driv

e (A

+B)

HTR Cntl B

LCL B1 Sw 1

LCL Bn

Sw nHTR Cntl A

LCL A1 Sw 1

LCL An

Sw n

MB Filter

Ski

n-A

rm(-

Saf

e) tb

c

BDDBatt. Pwr

LCLs / FCLs A

FCL OBC A

FCL DST A

Payload-LCLs A

Platform-LCLs A

Voc, Isc

Driv

e (A

+B)

Earth Reference point close to separation plane

Voc, Isc

BCDR

Protection

MonitorsOFFOFF

OFF

DC/DCConverter

OFF

OFF

DC/DCConverter

FailureDetector

Protection

FailureDetector

OFF

BCDR

Protection

MonitorsOFFOFF

OFF

DC/DCConverter

OFF

OFF

DC/DCConverter

FailureDetector

Protection

FailureDetector

OFF

BCDR

Protection

MonitorsOFFOFF

OFF

DC/DCConverter

OFF

OFF

DC/DCConverter

FailureDetector

Protection

FailureDetector

OFF

BCDR

Protection

MonitorsOFFOFF

OFF

DC/DCConverter

OFF

OFF

DC/DCConverter

FailureDetector

Protection

FailureDetector

OFF

BCDR

Protection

MonitorsOFFOFF

OFF

DC/DCConverter

OFF

OFF

DC/DCConverter

FailureDetector

Protection

FailureDetector

OFF

BCDR

Protection

MonitorsOFFOFF

OFF

DC/DCConverter

OFF

OFF

DC/DCConverter

FailureDetector

Protection

FailureDetector

OFF

APRProtection&

Filter

DC/DCConverter

FailureDetector

MPPT Control

OFF

Voting

OFF OFF

APRProtection&

Filter

DC/DCConverter

FailureDetector

MPPT Control

OFF

Voting

OFF OFF

APRProtection&

Filter

DC/DCConverter

FailureDetector

MPPT Control

OFF

Voting

OFF OFF

APRProtection&

Filter

DC/DCConverter

FailureDetector

MPPT Control

OFF

Voting

OFF OFF

Main Error Amplifier (MEA)DNEL, HK-Cond., Battery Discharge

Detector, Battery Voltage Alarm

Main Error Amplifier (MEA)DNEL, HK-Cond.,

Battery Discharge Detector, Battery Voltage Alarm

APRProtection&

Filter

DC/DCConverter

FailureDetector

MPPT Control

OFF

Voting

OFF OFF

APRProtection&

Filter

DC/DCConverter

FailureDetector

MPPT Control

OFF

Voting

OFF OFF

APRProtection&

Filter

DC/DCConverter

FailureDetector

MPPT Control

OFF

Voting

OFF OFF

APRProtection&

Filter

DC/DCConverter

FailureDetector

MPPT Control

OFF

Voting

OFF OFFSA Pwr A

SA Pwr ASA Pwr B

SA Pwr B

SAD

M

HD

2A

–2B

Solar Array-2

PAN

EL

4

Front

Back

PAN

EL

3

Front

Back

YOKE

SAD

MLCLs / FCLs B

FCL OBC BFCL DST B

Payload-LCLs B

Platform-LCLs B

TM/TC-IF AAux-

Cnv-A

RTN R

Ala

ms

Dis

cr.

TM/T

C

TM/TC-IF BAux-

Cnv-B

RTN R

Alam

sD

iscr

. TM

/TC

Pyro Mod A

Sel N

Sel 1

ARM

ARM

Pyro Mod B

Sel N

Sel 1

ARM

ARM

28 V ± 0,1

HD

1A

–1B

Solar Array-1

PAN

EL2

Front

Back

PAN

EL1

Front

Back

YOKE

SADE 2

DC

/DC

DC

/DC

RT

N R

RT

N R

FCEBCN R

FCEBC

N R

FCERT

N R

FCERT

N R

C(2)C(2)

C(2)C(2)

1553B-Bus A+BZZC

(2)C(2)

C(2)C(2)ZZ

CSS +X-X+y-y

+X-X+y-y

+X-X+y-y

CSS +X-X+y-y

+X-X+y-y

C(2)C(2)

C(2)C(2)

C(2)C(2)

C(2)C(2)

RMU-1LCL DC/DC

Serial TLM+Serial TLM-

Serial CMD+

Analogue ARoxAnalogue ARoyAnalogue ARozAnalogue GND

Serial CMD-Serial SYNC+Serial SYNC-

RMU-2LCL DC/DC

Serial TLM+Serial TLM-

Serial CMD+



RMU-3 LCL DC/DCSerial TLM+Serial TLM-

Serial CMD+



CPSI/O A

FCVDrivers

DC/DC

CPSI/O B

FCVDrivers

DC/DC

SerialI/O

A+B

WDE-1LCL DC/DC

Discr. TM/TC

RTNR

Heater Power

WDE-2LCL DC/DC

Discr. TM/TC

RTNR

Heater Power

WDE-3LCL DC/DC

Discr. TM/TC

RTNR

Heater Power

WDE-4LCL DC/DC

Discr. TM/TC

RTNR

Heater Power

RWA-1RWA-2RWA-3RWA-4

Space Wire

Spa

ce W

ire

Spac

e W

ire

Figure 3.2-1: Solar Orbiter detailed functional architecture

The design drivers for the Solar Orbiter SC come from not only the need to satisfy mission technical and performance requirements, but also to minimise the total cost of the mission. The philosophy is to avoid technology development as far as possible, in order to cost-cap the mission in-keeping with it’s M-class status. The obvious route to achieving this goal is through the use of technologies and engineering design lessons learnt from previous missions. The primary ‘source’ mission is the BepiColombo project, which is a mission with several similarities to Solar Orbiter in terms of the environment. This approach has been explicitly followed for the Solar Orbiter design from the 2nd CDF study [IR01], when the decision was made to reuse the BepiColombo SEPM for the transfer. The design of Solar Orbiter has continued to incorporate BepiColombo technology items as definition has progressed. Furthermore, design heritage from the Express series of missions, which had the goal of rapid and streamlined development, and for which EADS Astrium was responsible, has also featured heavily in the Solar Orbiter design. The principal features of the Solar Orbiter SC are shown in the figure opposite. The SC is sun-pointed during all mission phases after LEOP, the heatshield providing the platform and sensitive equipment with protection from the extremely high levels of solar flux. The heatshield also contains cut-out feddthroughs (with doors) which serve to provide the remote-sensing instruments with their required FoV to the Sun. The SC structure is derived from the MEX/VEX/Aeolus programmes, with internal shear panels providing mounting locations for the remote-sensing instruments and bus units, and external mounting locations on face panels, as well as on a dedicated instrument boom, for in-situ payloads. Two-sided solar arrays provide the capability to produce the required power throughout the mission over the wide range of Sun distances


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experienced, using 1-DOF along their longitudinal axis to allow switching between faces as well as control of the SAA to allow management of the array temperature throughout the mission and in particular during close-approach to the Sun. The PCDU is also taken from the BepiColombo mission. The LIR is located on the opposite face of the structure to the heatshield, such that the heatshield is uppermost when the SC is mated to the launch vehicle. The distribution of the RCS is inspired by the MEX mission, with rear-panel thrusters providing torque and primary thrust control. No main engine is included as the mission does not suffer from gravity losses, and the overall delta V requirements of the mission are comparatively modest. The rear-panel thrusters are complemented by additional thrusters on side panels in order to provide the capability to perform delta V manoeuvres whilst maintaining a Sun-pointing attitude when close to the Sun, a critical capability for the SC. The X/Ka-band TT&C subsystem is taken entirely from BepiColombo, with an articulated HTHGA providing nominal communication with the GS, and BepiColombo MGA and 2 LGAs for use as backup and during the LEOP. The DMS and AOCS subsystems also borrow heavily from BepiColombo.

3.3 BepiColombo Technology Reuse

As mentioned in the previous section, BepiColombo technology features heavily in the design of Solar Orbiter. Accordingly, the specifics of reuse (OTS, modifications) have been specifically studied within the frame of the Solar Orbiter Definition. The following table provides a summary of the BepiColombo technology reuse status at present, including an indication of the design modifications (if any) required to tailor units/technologies to the Solar Orbiter mission. Please refer to [IR29] for more detailed information.


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Figure 3.3-1: Solar Orbiter SC with principle design features highlighted

2-axis steerable HGA for nominal communication

2-sided solar arrays with 1-dof and long yokes

Multi-layer heatshield

Remote-sensing feedthroughs and doors

within heatshield

MEX/VEX structure

Chemical propulsion RCS with no main engine

BepiColombo Reuse Elements: DMS: OBC, RIU, SSMM COMMS: Transponders, RFDU, TWTA, HTHGA, MGA, LGA AOCS: STR, Gyro POWER: SA, PCDU TCS: HTMLI Express Series Reuse Elements: CPS: RCS layout MECH: Structure


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High Gain Antenna2. Dual frequency band feed (FEED) reused

High Gain Antenna3. Antenna pointing Assembly (APA)

BepiColombo HGA mechanisms can be reused, but need to be reviewed since specific high temperature design not needed for Solar Orbiter

High Gain Antenna4. Drive electronics

BepiColombo HGA mechanism drive electronics could be reused, but if no reuse of mechanisms, the drive electronics dedicated to the selected mechanism should be usedInterfaces to OBC (MIL bus) to be maintained if possible

High Gain Antenna5. Wave Guides (WG) reused, but length and support different

Magetometer Boom Mechanisms

BepiColombo boom elements can be reused. Elements may require delta qualification to cover the use with regard to the two hinges design of the boomMechanical qualification of boom assembly required to cover the design changesSpecific zero-g test-rig required

OBCBepiColombo OBC can be reused as is. Potentially reconfiguration sequences in Reconfiguration Module need to be modified to account for different Survival / Save mode implementation.

RIUBepiColombo RIU can be reused with minor changes. Since BepiColombo Specification update still in progress and Solar Orbiter requirements are to be detailed, this is tbc.

SSMBepiColombo SSMM can be reused. Change of the 3 memory modules capacity from 192 Gbit to 256 Gbit required. Availability of the SDRAM's is critical.

FCE No re-use foreseen in Solar Orbiter baseline.

Startracker

BepiColombo star tracker can be reused pending confirmation that it is compatible with the Solar Orbiter environment, see related TDA.Cost improvement can be gained since no extra baffle with shutter is required and the qualified MIL 1553 data bus interface exists

IMU

BepiColombo IMU can be reused as is. Since the performance of this unit is the best actually achievable, and even with this unit the Solar Orbiter pointing requirements may be not be met, the pointing requirements are proposed to be reviewed (i.e. this unit is the SOTA)

Wheels

BepiColombo RW is not proposed for reuse on Solar Orbiter since largest available wheel (68 Nms) is proposed for Solar Orbiter to increase off-loading intervals and to allow to limit the rate. Additionally mounting of the RW's on dampers to minimize the microvibrations - this is a new item not on BepiColombo

Coarse Sun sensor (CSS) No - specific TDA is required for use on Solar Orbiter

RMU No real reuse. BepiColombo RMU as specified can be reused. However no longer procured for BepiColombo, since second IMU used for Survival and Safe Mode

Fine Sun sensor (FSS) No reuse of the BepiColombo Fine Sun Sensor - only coarse sensors used on Solar Orbiter

Table 3.3-1: BepiColombo equipment re-use for Solar Orbiter at a glance


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4. SOLAR ORBITER CONFIGURATION

4.1 Introduction

The configuration of the Solar Orbiter SC is a solution that mediates between the many conflicting drivers and constraints of each of the individual units on board of the SC. All the units on the SC have been iteratively accommodated as knowledge of the design drivers has accumulated – this has been a semi-formal process in which the understanding of the drivers and constraints at a unit-specific level has been developed through various analyses, and this understanding used to arbitrate and resolve conflicts between units. The aim has been to arrive at an optimal configuration given the overall system-level drivers and constraints of the mission. Bear in mind that this means at a local level, several units are placed in less-than-perfect locations. The process of unit accommodation is dominated by the following key considerations:

i. The primacy of thermal considerations in the allocation of units to the various bays of the SC and external surfaces; in particular:

o The interaction between RS-instrument locations, resultant aperture locations in the heatshield, and the resulting performance of the heatshield at a local level

o The thermal interaction between the HTHGA and the heatshield o The combined dissipative and received thermal loads (when observing the Sun) of the RS-

instruments, and their position relative to other high dissipation units o The identification and maintenance of adequate panel radiator area on the various faces of

the SC, heavily constrainted by the presence of ‘in-Sun’ appendages such as the HTHGA and RPW antennas

o The relative positions between thermally critical instruments and their targeted radiator areas

o Thermo-Elastic considerations which lead to shear-wall mounting locations for coalignment and pointing critical RS-instruments, and the conflicting requirement for cooling of these instruments via radiation to cold space from face panels

ii. The operating and environmental requirements of the various payloads which obviously have primacy in defining the configuration, particularly:

o The FoV requirements of all payloads o Magnetic cleanliness requirements of certain key instruments o The combination of the previous two factors in necessitating the use of an instrument boom

for instruments that require it o The necessity to protect all the instruments from the Solar flux during close approach

periods o The conflicting need for a FoV to the Sun for all RS-instruments (except SolOHI) as well as

the SWA IS instrument, leading to a requirement for several feedthrough apertures in the heatshield of the SC

iii. The operational realities of the mission, which often conflict with (i) and (ii), in particular with their effect on the required movements of key SC appendages, including:

o The wide range of HTHGA positions that must be accessed in order to achieve an RF-link with the GS

o The necessity to stow the HTHGA for close-approach periods due to the severity of the thermal environment.

The configuration of the SC is of course a function of the mission timeline, as different configurations are adopted throughout the mission. These configurations involve the deployment and operation of several appendages, namely:

• HTHGA and MGA (in contingency) • Solar arrays • Instrument boom • RPW antennas.

Note that here we present none of the rationale determining where units are accommodated; please refer to the supporting analysis documentation for this.


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4.2 Configurations versus Mission Timeline

The configuration of the Solar Orbiter spacecraft is a function of the mission timeline. The Solar Orbiter SC has four nominal configurations throughout the course of the mission, with an additional contingency configuration for use in failure situations; these are summarised in the following table, along with the periods over which they are used (please note that there is no clean allocation of configurations to phases because configuration changes to new configurations occur before associated phase is entered.

Configuration Designation

Applicable Duration Definition

C_LEOP Launch to TBD days prior to end of LEOP

Instrument Doors RPW Antennas Instrument Boom HTHGA MGA Solar Arrays

Closed Stowed Stowed Stowed Stowed Stowed

C_NECP TBD days prior to end of LEOP to TBD days prior to end of NECP


Closed Stowed Stowed Deployed Stowed Deployed

C_CRUISE TBD days prior to end of NECP to end of EMP (intermittent suspension for C_SCIENCE)


Closed Deployed Deployed Deployed/Shadowed Stowed Deployed

C_SCIENCE Intermittent between start of NMP to end of EMP


Open Deployed Deployed Stowed/Shadowed Stowed Deployed

C_CONTINGENCY - (failure case) Instrument Doors RPW Antennas Instrument Boom HTHGA MGA Solar Arrays

Closed Deployed Deployed Stowed/Shadowed Deployed Deployed

Table 4.2-1: Solar Orbiter Configurations and their application within the mission timeline

Figure 4.2-1: Pictorial representation of the relationship between mission phases and SC

configurations (*the HTHGA can be stowed for all science windows or just for those windows below 0.3 AU distance to the Sun depending upon the precise operational scenario chosen; number of

C_SCIENCE is representative only)


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The SC begins the mission in C_LEOP, with all SC appendages stowed; during LEOP the SC configuration changes to C_NECP, through deployment of the solar array and HTHGA; during the NECP phase of the mission the configuration changes to C_CRUISE through deployment of the remaining payload-related appendages (the RPW antennas and the instrument boom). C_CRUISE is essentially maintained throughout the remainder of the mission, except for the RS-observation science windows during which C_SCIENCE is entered (the instrument doors are opened and the HTHGA is stowed behind the shadow of the heatshield2).

Figure 4.2-2: C_LEOP configuration, with the arrays, RPW antenna and HTHGA stowed [left] –Y

panel view, [right] isometric view – [left] in the Soyuz fairing, [right] isometric view on LVA

2 In the hot cases where the SC is closer than 0.28 AU to the Sun.


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Figure 4.2-3: C_NECP configuration, with the arrays and HTHGA deployed


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Figure 4.2-4: C_CRUISE configuration, with the arrays, HTHGA, instrument boom and RPW antennas

deployed


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Figure 4.2-5: [left] C_SCIENCE configuration, with the arrays, HTHGA, instrument boom and RPW

antennas deployed, and the RS-instrument feedthrough doors open [right] C_SCIENCE sub configuration with HTHGA positioned behind the shadow of the heatshield, for periods in the orbit

below 0.28 AU from the Sun


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Figure 4.2-6: C_CONTINGENCY configuration with the MGA deployed. Note that in this figure the HTHGA is shadowed, although it can also remain in the deployed state if the Sun distance is above

0.28 AU


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4.3 Payload Accommodation

The spacecraft configuration has to accommodate multiple heads and electronic boxes for the RS and IS instruments, with various FoV, temperature (stability), pointing, alignment and environmental requirements. The payload complement for the Solar Orbiter mission consists of the following instruments:

Instrument Acronym Sensor Unit Acronym Number of

Units Remote-Sensing

Polarimetric and Helioseismic Imager PHI 1 Spectral Imaging of the Coronal

Environment SPICE 1

Extreme Ultra-violet Imager EUI 1 Coronagraph COR 1

Spectrometer/Telescope for Imaging X-rays

STIX 1

Solar Orbiter Heliospheric Imager SoloHI 1 In-Situ

Electron Analyzer System EAS 2 Proton-Alpha Sensor PAS 1 Solar Wind Analyzer SWA

Heavy Ion Sensor HIS 1

Antenna ANT 3 Radio and Plasma Wave Analyzer RPW

Search Coil Magnetometer SCM 1

Magnetometer MAG 2

SupraThermal Electron sensor STE 1

Suprathermal Ion Spectrograph SIS 1

Electron Proton Telescope EPT 2

Low Energy Telescope LET 2

Energetic Particle Detector EPD

High Energy Telescope HETn 1

Table 4.3-1: Solar Orbiter Payload Complement (note: associated electronic boxes are not included in the ‘number of units’, but number 1 per instrument)

For a full description of the payload requirements, please refer to the following documents:

• Solar Orbiter Payload Thermal Requirements and Interface Definition [IR33] • Solar Orbiter Payload Interfaces Requirements and Accommodation [IR34] • Solar Orbiter System Trades [IR16].

The instrument accommodation is of course also influenced not just by individual requirements, but also by conflicts and constraints, both with other instruments and also with other elements on the spacecraft. In general terms the accommodation of the payloads has proceeded according to thermal constraints; specifically for the RS-instruments, with the aim to optimise their heat dissipation through high efficiency radiators with large view factors to cold space. Mechanical constraints have also driven the accommodation, with the aim to minimize instruments pointing misalignment due to thermo-elastic and mechanical constraints transferred from the heat shield to the spacecraft; this has led to the decision to mount the three RS-instruments with the tightest coalignment requirements (PHI, EUI & SPICE) together on a common panel. All the RS instruments are mounted on the platform structure shear walls, through iso-static bipods or brackets. These shear walls are as thermally decoupled as possible from the heatshield and from the external walls, which make them largely insensitive to environment changes (distance from the Sun, Solar Arrays orientation…), thus providing the best possible stability for the RS-instruments. The largest aperture instruments receive at perihelion a large amount of solar flux and need therefore large radiator surfaces to keep the illuminated parts (mirrors, baffles) at standard temperature and in some cases at very stable temperature (optical needs). For this reason, these are located on the corners of the satellite, with large available surfaces for radiators at short distance. The specific accommodation of large dissipators


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(primarily COR) has been considered in combination with the accommodation of other high-dissipation units as well as the restrictions on available radiator area caused by the accommodation of surface mounted units such as the EPD-sensor suite and the STR.

4.3.1 Remote Sensing Instruments

The Remote-Sensing instruments accommodation (including eboxes) and their associated FoVs are shown in the following figures.

Figure 4.3-1: Remote Sensing instrument accommodation

SOLOHI

STIX EBOX

EUI EBOX

SPICE

SPICE EBOX

EUI

PHI

PHI EBOX

STIX METIS METIS EBOX

SOLOHI EBOX


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Figure 4.3-2: Remote sensing instrument FoVs

PHI-FOV

STIX-FOV METIS-FOV

EUI-FOV SPICE-FOV

SOLOHI-FOV


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4.3.2 In-Situ Instruments

The accommodation of the various In-Situ instrument units (including eboxes) and their associated FoVs are shown in the following figures.

Figure 4.3-3: In Situ instrument accommodation

EPD-STE

MAG-OBS

RPW-SCM

SWA-EAS

RPW-ANT

SWA-HIS

RPW-ANT EPD-LET

EPD-HETn

EPD-SIS

RPW-ANT

MAG-IBS

EPD-EBOX

RWP-EBOX

SWA-PAS

MAG-EBOX

SWA-EBOX

EPD-EPT


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Figure 4.3-4: EPD-STE FoV


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Figure 4.3-5: SWA FoVs

SWA-EAS FOV

SWA-HIS FOV

SWA-PAS FOV


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Figure 4.3-6: EPD-EPT FoVs


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Figure 4.3-7: EPD-LET FoV


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Figure 4.3-8: EPD-HETn and EPD-SIS FoVs

EPD-SIS FOV

EPD-HETn FOV

EPD-HETn FOV

EPD-SIS FOV


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4.4 Subsystem Accommodation

In addition to the payloads, the SC configuration is composed of the following subsystems:

• Mechanical • Heatshield (including feedthroughs, doors and mechanisms) • Feedthroughs, doors and mechanisms • Thermal control • Propulsion • Attitude and orbit control • Data management • Power • Communications • Harness.

The accommodation of the individual subsystems is illustrated in the following figures. For a complete list of the drivers considered in the placement of units, please refer to [IR16].

Figure 4.4-1: Solar Orbiter Structure (exploded view)

+ Y PANEL

- Y PANEL

- Z PANEL

+ Z PANEL

CORE STRUCTURE

LIR

- X PANEL

+ X PANEL

SOLOHI SUPPORT BRACKET

EPD BRACKETS

SWA-PAS SUPPORT BRACKET

SWA-HIS SUPPORT BRACKET


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Figure 4.4-2: Heatshield accommodation

CORNER BIPOD

SUN SENSOR CUTOUT

MONOPOD BLADE

SWA-PAS CUTOUT

SWA-HIS CUTOUT


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Figure 4.4-3: Remote sensing instrument feedthroughs doors and mechanisms

PHI

STIX

EUI

METIS

SPICE

SWA

SWA


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Figure 4.4-4: TCS Accommodation

HEATPIPES

RADIATOR AREAS

HOT FINGER RADIATOR AREAS


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Figure 4.4-5: Propulsion subsystem accommodation

-X Panel THRUSTERS

+/- Y PANEL THRUSTERS

PROPELLANT TANKS

BENCH

PRESSURANT TANK

PIPEWORK


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Figure 4.4-6: AOCS and DMS accommodation

+X SUN SENSOR

RMU

OBC

SSMM

STAR TRACKERS

FCE

-X SUN SENSOR

GYRO

SADE

RIU

REACTION WHEELS

REACTION WHEELS


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Figure 4.4-7: Power subsystem accommodation

SADE

BATTERY

PCDU

SOLAR ARRAY

SOLAR ARRAY

SADE


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Figure 4.4-8: TT&C subsystem accommodation

EPCs

TRANSPONDERS

TWTAs

LGA

HTHGA

MGA LGA

HTHGA I/F POINTS


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Figure 4.4-9: Harness volume allocation (including connector volumes)


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5. MISSION ANALYSIS

For a full description of the Mission Analysis of the Solar Orbiter mission, please refer to [IR26].

5.1 Mission Overview

The mission objective of Solar Orbiter is principally to reach an orbit with perihelion at less than 0.25 AU and a high inclination of typically over 30 degrees with respect to the ecliptic or more specifically, 35 degrees with respect to the Solar equatorial plane. The transfer can be in principal be accomplished with either high thrust, chemical propulsion system assistance or low thrust, SEP. The selected baseline for the Solar Orbiter mission is to use a ballistic, high thrust transfer (i.e. chemical propulsion). This transfer can be effected principally by a combination of planetary gravity assist manoeuvres and deep space manoeuvres. The mission design drivers can be summarised as:

• To reach the required orbital targets • To maximise the mass in science orbit by controlling deep space manoeuvres and launcher

injection excess hyperbolic speed • To minimise the time taken to reach the operational targets.

However, the final orbit target may be achieved progressively, and the science phase of the mission can be considered as begun after the perihelion target is achieved. Inclination can then be progressively increased via a sequence of gravity assist manoeuvres at Venus.

5.2 Mission Baseline

The mission baseline is currently based on a launch with a Launch Vehicle (LV) to be provided by NASA, but the design and performances of the spacecraft must be compatible with a launch from Kourou using a Soyuz-ST 2-1b + Fregat upper stage (LAUN05 in the MRD [NR01]). To satisfy the mission scientific requirements, the operational orbit shall be a heliocentric orbit with an initial perihelion radius near 0.22 AU (48 to 50 solar radii). During the mission lifetime, the inclination with respect to the solar equator (solar inclination, i

s) should reach a value above 32° (TBC). The direct injection into that

operational orbit is beyond the performance of the LV and of the capabilities of the spacecraft. However, the following considerations lead to the strategy that provides a solution:

• The raising of the solar inclination to the required value could be achieved by repeated Gravity Assist manoeuvres (GAM) with Venus, where the arrival relative velocity of the spacecraft with respect to Venus must be near 18 km/s.

• The direct injection into a trajectory from Earth to Venus arriving at Venus with that relative

velocity will require an escape velocity from the Earth greater than 10 km/s, which is beyond the capabilities of the LV.

• Using a sequence of GAMs with Venus and Earth, and Deep Space Manoeuvres (DSM)

performed with Chemical Propulsion (CP), it is possible to leave the Earth within the capabilities of the LV (escape velocity about 3.65 km/s), and arrive with the required high relative velocity at Venus.

After the Launch and Early Orbit Phase (LEOP), the interplanetary transfer trajectory to reach the operational orbit will include GAMs with Earth and Venus, and any required DSM to arrive at the beginning of the Operation Phase. Of all the planets, Earth and Venus are selected because they allow:

• To find a solution within a reasonable duration of the transfer phase of 3.4 to 4.1 years.

• The maximum distance to the Sun is less than the mean distance of Mars to the Sun.

• The gravitational attraction of Mercury is too weak to justify the added complexity.


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It has been found that the optimal interplanetary transfer trajectory is based on a launch towards Venus where the swing-by conditions are selected to follow a trajectory that returns to the Earth. This Earth swing-by is already performed at a high relative velocity. A selection of the swing-by conditions allow selection of a trajectory that will return to Earth. The duration of this phase will be either an integer number of years (resonant case), or in a non-integer number of years (case where the Earth is at a different location in its orbit). After the second Earth swing-by the conditions are selected to encounter Venus with the required very high relative velocity and initiate the operational phase. The initial solar inclination of the operational orbit will not have the required value, and it will be increased by selecting the orbital period of the operational orbit to be resonant with the orbital period of Venus, and designing a sequence of GAM with Venus that will gradually increase the solar inclination of the operational orbit while maintaining the orbital period. The ratio of resonance is selected as 3:2 (3 orbits of the spacecraft for each 2 orbits of Venus). This results in an orbital period for the operational orbit of about 150 days. For launch in 2017, the resonance ratio of the first operational orbit is selected as 4:3, and changed to 3:2 for the rest of the Venus to Venus phases. About 7 to 9 years after launch the orbit solar inclination will reach a value close to 35°. The optimisation of the transfer trajectories is performed by maximising the mass at the end of the transfer phase (beginning of the operational phase), and maximising the solar inclination reached at the end of the inclination raise sequence. The sequence and the timing of Venus and Earth GAMs are selected to reduce the transfer duration below a reasonable value (~ 4 years).

5.2.1 2017 Mission Scenario

The transfer phase used for a launch in 2017 is described by:

1. Launch in January 2017 with an escape velocity from the Earth of about 3.640 km/s, and declination of the escape velocity of about 21.45°. This value provides a good balance between the launcher performances and the spacecraft on-board propulsion capabilities. In fact, for the optimum launch window of 20 days there is no deterministic DSM required.

2. Corresponding Soyuz/ST performance is ~1300 kg (TBC). Please refer to the Appendix for a

discussion of Astrium Ltd concerns regarding the ability of Soyuz-Fregat from Kourou to achieve the required launch parameters.

3. About 3 months after launch, a Venus GAM, with a pericentre height of more than 4000 km, will

put the spacecraft in a trajectory towards the Earth.

4. 1.4 years later, an Earth GAM puts the spacecraft in a 2 years orbit such that there is another Earth GAM, with a pericentre height of more than 700 km, 2 years later.

5. Less than 6 months after the last Earth swing-by, a Venus GAM, with a pericentre height of 300

km, will put the spacecraft in a 4:3 resonant trajectory with Venus.

6. Less than 1.9 years later the next Venus GAM starts the sequence of 3:2 resonant trajectories that raises the solar inclination.

As mentioned previously, the initial resonance with Venus is 4:3, switching to 3:2 after VGAM-3. The resultant operational orbit has an orbital period of 150 days, corresponding to a semi-major axis of 0.55 AU, a perihelion radius of 0.22 AU, and a solar inclination of 7.7°. A series of Venus GAM every 450 days will increase the solar inclination to the final maximum value.


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Figure 5.2-1: 2017 trajectory viewed from above the ecliptic

Table 5.2-1: 04/01/2017 mission summary


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5.2.2 2018 Mission Scenario

The transfer phase used for a launch in 2018 is described by:

1. Launch in July 2018 with an escape velocity from the Earth of about 3.660 km/s, and declination of the escape velocity of about -52.97°. This value provides a good balance between the launcher performances and the spacecraft on-board propulsion capabilities. Note that it is not possible to maintain a 20-day launch window without incurring a DSM-penalty; 50 m/s of additional delta V is required to satisfy the 20-day launch window requirement.

2. Corresponding Soyuz/ST performance is 1194 kg (TBC). Please refer to the Appendix for a

discussion of Astrium Ltd concerns regarding the ability of Soyuz-Fregat from Kourou to achieve the required launch parameters.

3. About 6 months after launch, a Venus GAM, with a pericentre height of more than 9000 km, will put

the spacecraft in a trajectory towards the Earth. 4. 10.5 months later, an Earth GAM puts the spacecraft in an orbit with a period such that there is

another Earth GAM, with a pericentre height of more than 5300 km, 22.3 months later. 5. Less than 3 months after the last Earth swing-by, a Venus GAM, with a pericentre height of 300 km,

will put the spacecraft in a 3:2 resonant trajectory with Venus. 3:2 resonance with Venus is entered directly in this scenario. The initial operational orbit has an orbital period of 150 days, corresponding to a semi-major axis of 0.55 AU, a perihelion radius of 0.22 AU, and a solar inclination of 10.2°. A series of Venus GAM every 450 days will increase the solar inclination to the final maximum value.


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Figure 5.2-2: 2018 trajectory viewed from above the ecliptic

Table 5.2-2: 30/07/2018 mission summary

Note that in both the above tables, the definition of the end of the NMP is the first perihelion after VGAM-V4. This is not consistent with the definition of the first set of science windows after the GAM stated in the MRD; the end of the NMP should be defined as the end of the last science window in the first operational orbit after VGAM-V4.


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Solar Orbiter 2017 Launch

0.00

0.25

0.50

0.75

1.00

1.25

1.50

1.75

2.00

2.25

May

-16

Sep

-17

Feb-

19

Jun-

20

Oct

-21

Mar

-23

Jul-2

4

Dec

-25

Apr

-27

Sep

-28

Jan-

30

Dis

tanc

e (A

U)

S/C-E S/C-V S/C-Sun

Figure 5.2-3: Sun, Earth and Venus distance over the baseline 2017 mission scenario

Solar Orbiter. Launch 2017

-40

-30

-20

-10

0

10

20

30

40

0 500 1000 1500 2000 2500 3000 3500 4000 4500 5000

Days since launch

Sola

r lat

itude

Figure 5.2-4: Solar latitude over the 2017 baseline mission scenario (pink arrows indicate Launch

and GAM events; note that orbital inclination changes occur at GAMs)


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0

20

40

60

80

100

120

140

160

180

0 500 1000 1500 2000 2500 3000 3500 4000

Days since launch

Sun/

SC/E

arth

ang

le (D

egre

es)

Figure 5.2-5: Sun/SC/Earth angle over the 2017 baseline mission scenario

Solar Orbiter 2018 Launch

0.00

0.25

0.50

0.75

1.00

1.25

1.50

1.75

2.00

2.25

Apr-

18

Aug-

19

Jan-

21

May

-22

Oct

-23

Feb-

25

Jun-

26

Nov

-27

Dis

tanc

e (A

U)

S/C-E S/C-V S/C-Sun

Figure 5.2-6: Sun, Earth and Venus distance over the baseline 2017 mission scenario


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-40

-30

-20

-10

0

10

20

30

40

0 500 1000 1500 2000 2500 3000 3500

Days since launch

Sola

r lat

itude

Figure 5.2-7: Solar latitude over the 2017 baseline mission scenario

Solar Orbiter Launch 2018

0

20

40

60

80

100

120

140

160

180

0 500 1000 1500 2000 2500 3000Days since launch

Sun/

SC/E

arth

ang

le (D

egre

es)

Figure 5.2-8: Sun/SC/Earth angle over the 2017 baseline mission scenario


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5.3 Delta V Requirements

The delta V requirements of the Solar Orbiter mission are modest compared to other interplanetary missions, due to the design decision to pursue a ballistic transfer trajectory with no deterministic delta V component (see [IR03]); accordingly the only contributions to the Solar Orbiter delta V budget are those associated with the correction of trajectory errors, namely:

• Launcher error correction manoeuvre: • Probabilistic error • Deterministic delta V • Launch Window DSM penalty • Pre/post-GAM setup manoeuvre.

5.3.1 Launcher Error Correction Manoeuvre

The launcher error correction manoeuvre will be a response to two sources of error, namely: Probabilistic error due to the underperformance of the launcher: this is computed based on the covariance matrix specified by the launch authority – for Soyuz-Fregat this represents about 12-15 m/s Deterministic delta V needed to compensate for non-optimal flight programmes. This item is due to the fact that, for interplanetary missions, different conditions for V-infinity are required each day of launch. This implies that for every day the launcher has to use a new flight program that produces the desired modulus and declination of the V-infinity. Due to the cost associated with the production and validation of these programs, Arianespace and Starsem usually provide a minimum number of flight programs such that the spacecraft has to perform a deterministic manoeuvre to achieve the required conditions at the escape from the Earth. For VenusExpress this item was 25 m/s.

5.3.2 Launch Window DSM Penalty

Solar Orbiter is required to have a 20-day launch window [NR01]. The trajectory design for Solar Orbiter is documented in [IR03]. By inspecting Error! Reference source not found.the 2017 baseline trajectory, it can be seen that a Launch Window in excess of 20 days exists without incurring any compensation DSM. However for the 2018 trajectory it can be seen that only 13 days is available without incurring a compensation DSM. A DSM delta V budget contribution of 50m/s is required for the 2018 launch in order to comply with the requirement for a 20 day launch window.

Figure 5.3-1: Deterministic DSM as a function of launch day for the [left] 2017 and [right] 2018 launch

windows

5.3.3 Gravity Assist Manoeuvres

As detailed in the MRD [NR01], 15m/s per GAM is allocated for the Solar Orbiter mission, including both pre-GAM setup and post-GAM cleanup.


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5.3.4 Nominal Delta V Budget

Taking the budget items discussed in the preceding sections, the Solar Orbiter ‘clean’ Delta V budget is presented in the following table for both launch cases.

Launch 2017 Best Estimate

Margin/Eff. Factor

Value Source

Deterministic Flight Programme Error 25 0 25 m/s VEX error correction precedent [1]

Probabilistic Launcher Error 15 m/s Provided by [1]

DSM for non-optimal launch date 0 m/s 21 day zero penalty window is possible - taken from [2]

Gravity Assist Manoeuvres 15 m/s/GA 8 120 m/s 15m/s allocation for each GAM [3]

gravity losses 0 m/s Negligible for Solar Orbiter

Total Nominal Mission Delta V 160 m/s Clean nominal mission delta V


Margin/Eff. Factor

Value Source

Deterministic Flight Programme Error 25 m/s VEX error correction precedent [1]


DSM for non-optimal launch date 50 m/s 50m/s penalty required for 20 day LW [2]


gravity losses 0 m/s Negligible for Solar Orbiter

Total Nominal Mission Delta V 210 m/s Clean mission delta V

Table 5.3-1: Solar Orbiter nominal Delta V budget for 2017 and 2018 mission scenarios

The propellant requirements of the mission must of course take into account thruster orientation inefficiencies and constraints on attitude during delta V manoeuvres, as well as of course the need to provide attitude control over the course of the mission – the total effective delta V and attitude control budgets are calculated in [IR46].


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6. MISSION OPERATIONS

Here we provide a brief summary of the main features of the mission operations. The mission operational approach is fully compliant with the common operational model developed by ESOC for interplanetary missions, and is more fully described in [NR08].

6.1 Mission Phases

The Solar Orbiter mission comprises the following principal phases. Pre-launch Phase [PLP]: This phase encompasses all pre-launch operations, starting several weeks before launch. Activities during this phase include the final launch simulations, data flow tests, battery reconditioning (if necessary), filling of the propellant tanks, SVT-2 with the MOC, simulations at ESOC, etc. The pre-launch phase ends with the lift-off. Launch and Early Operations Phase [LEOP]: This phase starts at lift-off and includes the ascent phase, injection into the target escape trajectory, and the separation of the spacecraft from the launch vehicle upper stage. LEOP is concluded once the spacecraft is fully configured for the next phase. Near Earth Commissioning Phase [NECP]: The NECP starts after completion of LEOP. The purpose of this phase is to verify the health and performance of the spacecraft, and checkout the status of the payloads. If outgassing allows, performance of the payload shall also be verified to the maximum extent possible. The in-situ instruments commissioning will be completed and will then remain continuously switched ON. NECP is planned to last for 3 months (note that the MRD v4.0 needs to be updated to say 3 months). Cruise Phase [CP]: The CP starts at the end of NECP and is dedicated to spacecraft monitoring, performance of science using the in-situ instruments and completion of performance verification of the remote sensing payload. Remote sensing instruments will be commissioned after the first Earth Gravity Assist Manoeuvre (EGAM). The RS-suite of instruments will then be operated at least during certain windows around perihelion passages and synoptic observations with NASA’s Inner Heliospheric Sentinels will be performed (TBC). Nominal Mission Phase [NMP]: Full science operations will start after the 2nd Venus Gravity Assist (VGAM-2). Upon VGAM-2 the spacecraft is injected in a 4:3 resonant orbit with Venus, with the eventual goal of a 3:2 resonance in order to increase the inclination of the orbit with respect to the ecliptic. The NMP ends after the science windows of the first spacecraft orbit after Venus Gravity Assist Maneuver VGAM-4. This is 90 months after launch. Note that for the 2015 orbit the 3:2 resonance is entered directly. Extended Mission Phase [EMP]: The extended mission, if approved, will continue until after the science windows of the first spacecraft orbit after Venus Gravity Assist Manoeuvre VGAM-6. This represents an extended mission duration of 6 solar orbits, and will result in an increase of the inclination of the spacecraft orbit with respect to the solar equator. The evolution of some of the key characteristics of the baseline 2017 mission are shown in the following figures, onto which are overlaid the CP, NMP and EMP durations.


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Solar Orbiter 2017 launch

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the mission, and the GAM events; the MRD-definitions of the CP, NMP and EMP are also shown

6.1.1 Mission Timeline

The top-level timelines for the nominal 2017 mission and backup 2018 mission are given in the following tables, which have been constructed using interrogation of trajectories supplied by ESOC, with a time-step of one day, and which are documented in [RD6]. Note that the timelines are based upon the following assumptions:

• NECP of 3 months duration • GAM targeting manoeuvres begin 1 month prior to GAM3 • GAM corrective manoeuvres end 1 week after GAM4 • Each science window is 10 days in duration and placed in strict accordance with the MRD definition

[PERF20 of the MRD]; accordingly, several of the perihelion and minimum latitude science windows are overlapping. It is expected that science window definitions will alter as the mission definition progresses. In particular windows could be moved to avoid overlap, and new windows included in not only the NMP and EMP, but also in the CP

• LEOP lasts 7 days5 – this is a typical duration for Earth-escape LEOP activities and includes launcher error correction.

• Antisun pointing is adopted for SC-Sun distances of above 1.2 AU. The following figures give a graphical representation of the sequence of key mission events.

3 ESOC guidelines state that the additional GS support shall be available up to 2 months prior to the GAM. 4 ESOC guidelines state that the additional GS support shall be available up to 1 month after the GAM. 5 ESOC guidelines state that the LEOP shall last no more than 7 days.


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Mission Phase/Event Start Date End Date Duration (Days)

LEOP 04/01/2017 11/01/2017 7 Launch and early operations 04/01/2017 11/01/2017 7.00 NECP 12/01/2017 09/04/2017 87 Commissioning Activities 12/01/2017 09/04/2017 87.00 CP 10/04/2017 16/02/2021 1410 Pre VGAM1 targeting period 18/03/2017 14/04/2017 28.00 VGAM1 15/04/2017 15/04/2017 1.00 Post VGAM1 correction period 16/04/2017 22/04/2017 7.00 Assume antisun pointing 02/08/2017 10/03/2018 221.00 Pre EGAM1 targeting period 28/07/2018 24/08/2018 28.00 EGAM1 25/08/2018 25/08/2018 1.00 Post EGAM1 correction period 26/08/2018 01/09/2018 7.00 Pre EGAM2 targeting period 28/07/2020 24/08/2020 28.00 EGAM2 25/08/2020 25/08/2020 1.00 Post EGAM2 correction period 26/08/2020 01/09/2020 7.00 Pre VGAM2 targeting period 12/01/2021 08/02/2021 28.00 VGAM2 09/02/2021 09/02/2021 1.00 Post VGAM2 correction period 10/02/2021 16/02/2021 7.00 NMP 17/02/2021 09/07/2024 1239 Science Window 1 (Pass_1_Max_Lat) 07/05/2021 16/05/2021 10.00 Science Window 2 (Pass_1_PeriPassage) 20/06/2021 29/06/2021 10.00 Science Window 3 (Pass_1_Min_Lat) 23/06/2021 02/07/2021 10.00 Science Window 4 (Pass 2_Max_Lat) 22/10/2021 31/10/2021 10.00 Science Window 5 (Pass 2_PeriPassage) 04/12/2021 13/12/2021 10.00 Science Window 6 (Pass 2_Min_Lat) 08/12/2021 17/12/2021 10.00 Science Window 7 (Pass 3_Max_Lat) 09/04/2022 18/04/2022 10.00 Science Window 8 (Pass 3_PeriPassage) 22/05/2022 31/05/2022 10.00 Science Window 9 (Pass 3_Min_Lat) 26/05/2022 04/06/2022 10.00 Science Window 9 (Pass 4_Max_Lat) 25/09/2022 04/10/2022 10.00 Science Window 9 (Pass 4_PeriPassage) 07/11/2022 16/11/2022 10.00 Science Window 9 (Pass 4_Min_Lat) 10/11/2022 19/11/2022 10.00 Pre VGAM3 targeting period 17/11/2022 14/12/2022 28.00 VGAM3 15/12/2022 15/12/2022 1.00 Post VGAM3 correction period 16/12/2022 22/12/2022 7.00 Science Window 13 (Pass 5_Max_Lat) 09/03/2023 18/03/2023 10.00 Science Window 14 (Pass 5_PeriPassage) 05/04/2023 14/04/2023 10.00 Science Window 15 (Pass 5_Min_Lat) 09/04/2023 18/04/2023 10.00 Science Window 16 (Pass 6_Max_Lat) 06/08/2023 15/08/2023 10.00 Science Window 17 (Pass 6_PeriPassage) 01/09/2023 10/09/2023 10.00 Science Window 18 (Pass 6_Min_Lat) 06/09/2023 15/09/2023 10.00 Science Window 19 (Pass 7_Max_Lat) 02/01/2024 11/01/2024 10.00 Science Window 20 (Pass 7_PeriPassage) 28/01/2024 06/02/2024 10.00 Science Window 21 (Pass 7_Min_Lat) 02/02/2024 11/02/2024 10.00 Pre VGAM4 targeting period 10/02/2024 08/03/2024 28.00 VGAM4 09/03/2024 09/03/2024 1.00 Post VGAM4 correction period 10/03/2024 16/03/2024 7.00 Science Window 22 (Pass 8_Max_Lat) 26/05/2024 04/06/2024 10.00 Science Window 23 (Pass 8_PeriPassage) 23/06/2024 02/07/2024 10.00 Science Window 24 (Pass 8_Min_Lat) 30/06/2024 09/07/2024 10.00


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EMP 10/07/2024 11/12/2026 886 Science Window 25 (Pass 9_Max_Lat) 22/10/2024 31/10/2024 10.00 Science Window 26 (Pass 9_PeriPassage) 20/11/2024 29/11/2024 10.00 Science Window 27 (Pass 9_Min_Lat) 27/11/2024 06/12/2024 10.00 Science Window 28 (Pass 10_Max_Lat) 22/03/2025 31/03/2025 10.00 Science Window 29 (Pass 10_PeriPassage) 18/04/2025 27/04/2025 10.00 Science Window 30 (Pass 10_Min_Lat) 25/04/2025 04/05/2025 10.00 Pre VGAM5 targeting period 04/05/2025 31/05/2025 28.00 VGAM5 01/06/2025 01/06/2025 1.00 Post VGAM5 correction period 02/06/2025 08/06/2025 7.00 Science Window 31 (Pass 11_Max_Lat) 09/08/2025 18/08/2025 10.00 Science Window 32 (Pass 11_PeriPassage) 08/09/2025 17/09/2025 10.00 Science Window 33 (Pass 11_Min_Lat) 17/09/2025 26/09/2025 10.00 Science Window 34 (Pass 12_Max_Lat) 05/01/2026 14/01/2026 10.00 Science Window 35 (Pass 12_PeriPassage) 05/02/2026 14/02/2026 10.00 Science Window 36 (Pass 12_Min_Lat) 15/02/2026 24/02/2026 10.00 Science Window 37 (Pass 13_Max_Lat) 04/06/2026 13/06/2026 10.00 Science Window 38 (Pass 13_PeriPassage) 05/07/2026 14/07/2026 10.00 Science Window 39 (Pass 13_Min_Lat) 14/07/2026 23/07/2026 10.00 Pre VGAM6 targeting period 27/07/2026 23/08/2026 28.00 VGAM6 24/08/2026 24/08/2026 1.00 Post VGAM6 correction period 25/08/2026 31/08/2026 7.00 Science Window 40 (Pass 14_Max_Lat) 22/10/2026 31/10/2026 10.00 Science Window 41 (Pass 14_PeriPassage) 20/11/2026 29/11/2026 10.00 Science Window 42 (Pass 14_Min_Lat) 05/12/2026 14/12/2026 10.00 Mission End 11/12/2026 11/12/2026 -

Table 6.1-1: Mission timeline for the nominal 2017 scenario (mission phases in green; GAM in grey)

Mission Phase/Event Start Date End Date Duration (Days)

LEOP 30/07/2018 06/08/2018 7 Launch and early operations 30/07/2018 06/08/2018 7.00 NECP 07/08/2018 02/11/2018 87 Commissioning Activities 07/08/2018 02/11/2018 87.00 CP 03/11/2018 24/12/2021 1149 Pre VGAM1 targeting period 26/12/2018 22/01/2019 28.00 VGAM1 23/01/2019 23/01/2019 1.00 Post VGAM1 correction period 24/01/2019 30/01/2019 7.00 Assume antisun pointing 13/05/2019 15/10/2019 156.00 Pre EGAM1 targeting period 03/11/2019 30/11/2019 28.00 EGAM1 01/12/2019 01/12/2019 1.00 Post EGAM1 correction period 02/12/2019 08/12/2019 7.00 Pre EGAM2 targeting period 02/09/2021 29/09/2021 28.00 EGAM2 30/09/2021 30/09/2021 1.00 Post EGAM2 correction period 01/10/2021 07/10/2021 7.00 Pre VGAM2 targeting period 19/11/2021 16/12/2021 28.00 VGAM2 17/12/2021 17/12/2021 1.00 Post VGAM2 correction period 18/12/2021 24/12/2021 7.00 NMP 25/12/2021 21/07/2024 940 Science Window 1 (Pass_1_Max_Lat) 19/02/2022 28/02/2022 10.00 Science Window 2 (Pass_1_PeriPassage) 27/01/2022 05/02/2022 10.00


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Science Window 3 (Pass_1_Min_Lat) 20/01/2022 29/01/2022 10.00 Science Window 4 (Pass 2_Max_Lat) 19/07/2022 28/07/2022 10.00 Science Window 5 (Pass 2_PeriPassage) 26/06/2022 05/07/2022 10.00 Science Window 6 (Pass 2_Min_Lat) 19/06/2022 28/06/2022 10.00 Science Window 7 (Pass 3_Max_Lat) 16/12/2022 25/12/2022 10.00 Science Window 8 (Pass 3_PeriPassage) 22/11/2022 01/12/2022 10.00 Science Window 9 (Pass 3_Min_Lat) 16/11/2022 25/11/2022 10.00 Pre VGAM3 targeting period 12/02/2023 11/03/2023 28.00 VGAM3 12/03/2023 12/03/2023 1.00 Post VGAM3 correction period 13/03/2023 19/03/2023 7.00 Science Window 13 (Pass 4_Max_Lat) 20/05/2023 29/05/2023 10.00 Science Window 14 (Pass 4_PeriPassage) 24/04/2023 03/05/2023 10.00 Science Window 15 (Pass 4_Min_Lat) 17/04/2023 26/04/2023 10.00 Science Window 16 (Pass 5_Max_Lat) 17/10/2023 26/10/2023 10.00 Science Window 17 (Pass 5_PeriPassage) 21/09/2023 30/09/2023 10.00 Science Window 18 (Pass 5_Min_Lat) 13/09/2023 22/09/2023 10.00 Science Window 19 (Pass 6_Max_Lat) 14/03/2024 23/03/2024 10.00 Science Window 20 (Pass 6_PeriPassage) 18/02/2024 27/02/2024 10.00 Science Window 21 (Pass 6_Min_Lat) 10/02/2024 19/02/2024 10.00 Pre VGAM4 targeting period 06/05/2024 02/06/2024 28.00 VGAM4 03/06/2024 03/06/2024 1.00 Post VGAM4 correction period 04/06/2024 10/06/2024 7.00 Science Window 22 (Pass 7_Max_Lat) 21/08/2024 30/08/2024 10.00 Science Window 23 (Pass 7_PeriPassage) 23/07/2024 01/08/2024 10.00 Science Window 24 (Pass 7_Min_Lat) 12/07/2024 21/07/2024 10.00 EMP 22/07/2024 04/03/2027 956 Science Window 25 (Pass 8_Max_Lat) 18/01/2025 27/01/2025 10.00 Science Window 26 (Pass 8_PeriPassage) 19/12/2024 28/12/2024 10.00 Science Window 27 (Pass 8_Min_Lat) 09/12/2024 18/12/2024 10.00 Science Window 28 (Pass 9_Max_Lat) 16/06/2025 25/06/2025 10.00 Science Window 29 (Pass 9_PeriPassage) 18/05/2025 27/05/2025 10.00 Science Window 30 (Pass 9_Min_Lat) 08/05/2025 17/05/2025 10.00 Pre VGAM5 targeting period 30/07/2025 26/08/2025 28.00 VGAM5 27/08/2025 27/08/2025 1.00 Post VGAM5 correction period 28/08/2025 03/09/2025 7.00 Science Window 31 (Pass 10_Max_Lat) 24/11/2025 03/12/2025 10.00 Science Window 32 (Pass 10_PeriPassage) 26/10/2025 04/11/2025 10.00 Science Window 33 (Pass 10_Min_Lat) 10/10/2025 19/10/2025 10.00 Science Window 34 (Pass 11_Max_Lat) 22/04/2026 01/05/2026 10.00 Science Window 35 (Pass 11_PeriPassage) 24/03/2026 02/04/2026 10.00 Science Window 36 (Pass 11_Min_Lat) 09/03/2026 18/03/2026 10.00 Science Window 37 (Pass 12_Max_Lat) 19/09/2026 28/09/2026 10.00 Science Window 38 (Pass 12_PeriPassage) 21/08/2026 30/08/2026 10.00 Science Window 39 (Pass 12_Min_Lat) 06/08/2026 15/08/2026 10.00 Pre VGAM6 targeting period 23/10/2026 19/11/2026 28.00 VGAM6 20/11/2026 20/11/2026 1.00 Post VGAM6 correction period 21/11/2026 27/11/2026 7.00 Science Window 40 (Pass 13_Max_Lat) 26/02/2027 07/03/2027 10.00 Science Window 41 (Pass 13_PeriPassage) 02/02/2027 11/02/2027 10.00 Science Window 42 (Pass 13_Min_Lat) 10/01/2027 19/01/2027 10.00 Mission End 04/03/2027 04/03/2027 -

Table 6.1-2: Mission timeline for the backup 2018 scenario (mission phases in green; GAM in grey)


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Figure 6.1-4: Nominal 2017 mission timeline showing start dates of each mission phase and principal events


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Figure 6.1-5: Backup 2018 mission timeline showing start dates of each mission phase and principal events


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6.1.2 Launch Date Dependent Mission Characteristics

6.1.2.1 Declination

The fundamental difference between the 2017 and 2018 launches is that the escape declination from the Earth is -ve for the 2018 launch and +ve for the 2017 launch. Accordingly the ‘sense’ of the trajectory is reversed: in the operational orbit, the SC will approach the Sun at +ve helio-latitudes, and consequently the maximum latitude science windows are the first to occur within each set of 3 observation windows – this is a result of the targeted position of Venus for the first VGAM which necessitates a positive declination escape at launch. Beyond this the essential nature of the mission is much the same: 2 EGAMs followed by 2 VGAMs leading to an operational orbit where successive VGAMs are employed to increase the inclination of the orbit.

6.1.2.2 Initial Resonance

An additional important difference between the 2018 backup and the 2017 new baseline is that the 2017 transfer involves a 4:3 resonance with Venus after the VGAM-2. This is because, with the nominal Venus approach conditions at the VGAM-2, it is not possible to establish a sufficient heliocentric energy change to reach the lower energy 3:2 resonance. Instead a 4:3 resonance is achieved, and the 3:2 resonance is only entered after VGAM-3. This has an impact on the evolution of the inclination of the SC over the mission; the maximum heliolatitude reached at the end of the EMP (final set of science windows after VGAM-6) is over 34°. Accordingly, the NMP of the 2017 scenario contains 8 peripassages rather than 7. Consequently, for the 2017 science window definition, there are 14 operational orbits in total, rather than the 13 operational orbits of the 2018 mission. See the budgets section of this document for the current definition of the science windows.

6.2 Cruise Phase [CP] Overview

Once the SC has undergone the bulk of commissioning activities during the NECP up until ~3 months after launch, the mission enters the CP. The CP comprises the period between the end of the NECP and the entry into the operational orbit of the SC after VGAM-2, at which point the NMP is entered. The CP combines several operational features that are similar to those of the NMP and EMP, namely multiple gravity assist manoeuvres, and generation of mission products through operation of the In-Situ payload. The critical operations during the CP are:

1. Maintenance of the Sun-pointing attitude at all times except for the two following special cases: • Adoption of anti-sun pointing during aphelion passage above 1.2 AU (TBC) • Re-pointing for pre/post EGAM delta V corrections

2. A GAM sequence VEEV 3. Continuous IS-payload operation 4. Rolling of the SC to facilitate communications 5. Flipping of the Solar Array between the Hot and Cold faces, and rotation of the solar array to

manage the array temperature. The following figures show the principal events during the CP, overlaid onto the Sun distance and Sun/SC/Earth angle of the SC, for the nominal 2017 and backup 2018 trajectories.


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Solar Orbiter 2017 launch - CP

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mission; also shown are the principal mission events (far-Sun, GAM and conjunction periods)

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mission; also shown are the principal mission events (far-Sun, GAM and conjunction periods)


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6.2.1 Cruise Phase Space/Ground Interface and Operations Support

During the CP, New Norcia shall support 3 passes per week, each with a minimum of 4 hours. In addition to nominal SC telemetry (housekeeping and mission product download), these passes shall also support residual commissioning activities of the RS-payload (IS-suite has commissioning priority during the NECP). During the GAMs that occur during this phase, both New Norcia and Cerberos shall be available to aid in preparation and clean-up of each GAM. During these periods (up to 2 months prior and 1 month after the GAM), it is envisaged that both pass frequency and duration shall be increased appropriately. Routine pass activities shall be automated at the control centre as soon as possible after commissioning. During this phase monitoring and maintenance activities in the SC will be performed off-line. The FCT is reduced in size compared to the LEOP and NECP. Operations during this phase will be conducted from the DCR. There will be interruptions to this schedule caused by the following:

• A short conjunction which begins 1006 days after launch and lasts 2 days • Interruptions due to interference with the HTHGA FoV by the SC and SC appendages – these

interruptions can be avoided by rolling the SC, or worked around by modification of the pass schedule (many of the interruptions are of the order of ~few days).

The pass schedule should be defined to take these into account. All nominal communication during the CP is conducted through the HTHGA, which continuously tracks the Earth position throughout the CP (ephemerides-table led tracking is maintained between passes). The HTHGA supports a reference TM bit-rate of 150 kbps at 1 AU. In the event of a serious failure, the MGA is used as a backup, providing a reference TM bit-rate of 2 kbps at 1 AU.

6.3 Nominal and Extended Mission Phase [NMP/EMP]

6.3.1 Phase Overview

These two phases are presented together here because they are essentially operationally identical. The NMP is entered after VGAM-2 and continues until after the first set of science windows after VGAM-4. The EMP is entered after the NMP ends and continues until the end of the first set of science windows after VGAM-6. Throughout both the NMP/EMP, the orbit evolves over time as each successive VGAM is used to increase the inclination of the orbit with respect to the Sun’s equator. The critical operations are:

1. Maintenance of the Sun-pointing attitude at all times, including during GAMs 2. A GAM sequence VVVVV 3. Periodic rolling to facilitate communications 4. Solar Array temperature management through inclination changes and flipping 5. Continuous IS-payload operations 6. Science Window Operations:

1. Reconfiguration of the SC between C_CRUISE to C_SCIENCE 2. RS-Payload Operations.

Due to the criticality of the thermal performance of the heatshield once the operational orbit is reached, a thermal characterisation campaign shall take place once resonant orbit is reached, to confirm/refine the operational envelope of the SC. The NMP is the period during which the closest approach to the Sun occurs: this is 0.2343 AU for the nominal 2017 mission scenario, occuring during the first orbit after VGAM-3. The closest approach distance for the backup 2018 mission is 0.2434 (larger). Another defining feature of the NMP/EMP is the presence of several communication blackouts, which are not present in the earlier phases; these are caused by conjunction events, and also by stowage of the HTHGA during periods of close approach to the Sun.

6.3.2 Science Window Definition

During the NMP/EMP the IS-instruments are continuously operating except during the critical operations surrounding the GAM, in which case they are placed in standby mode. The RS-instruments will operate only


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in the science windows6. The science windows are defined in the budget section of this document, as uninterrupted ten-day periods centred on the following three locations in the operational orbit:

• Point of maximuim helio-latitude • Perihelion passage • Point of minimum helio-latitude.

A strict interpretation of these definitions leads to the observation that the peripassage and minimum latitude science windows overlap each other considerably in the early phases of the operational phase where the lowest perihelion occurs (orbital period is proportional to the semi-major axis of the operational orbit); as the perihelion distance in later passes is increased, the degree of overlap reduces until in the last few orbits the science windows are completely separated. It is expected that the MRD-definition of the science windows shall undergo significant revision as the definition of the mission progresses. It is important to note that this phase of the trajectory of the SC is non-compliant with ESOC recommendations for the backup trajectories, and indeed for all trajectories that start with a -ve declination launch (TBC). For the former 2015 baseline and the current 2018 backup, VGAM-4 occurs during an extended conjunction period. This means that both of the backup trajectories (and by extension all trajectories which begin with a –ve declination launch) are non-compliant with ESOC recommendations. This is shown in Figure 6.3-2 for the 2018 mission scenario; at ~2100 days into the mission, the SC/Sun/Earth angle drops below 5°, coinciding with the VGAM-4. This coincidence will have an impact on the targeted GAM altitude (due to earlier pre-GAM targeting), and consequently a negative effect on the delta V budget (work is ongoing in ESOC to address this issue). Figure 6.3-1 shows the 2017 mission scenario NMP, which also has an extended conjunction period, but this is not coincident with any GAM. The EMP for both the nominal 2017 and backup 2018 missions is shown in Figure 6.3-3 and Figure 6.3-4 respectively; there are no conjunction events during this phase for either mission scenario.

6 If Solar Orbiter operates within the HELEX context, additional observations will occur outside the science windows in support of the Sentinels mission; however these are undefined and therefore not considered here.


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Figure 6.3-1: Sun distance and Sun/SC/Earth angle over the course of the NMP for the nominal 2017

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Solar Orbiter 2018 launch - NMP

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Figure 6.3-2: Sun distance and Sun/SC/Earth angle over the course of the NMP for the backup 2018 mission; also shown are the principal mission events (science windows, GAM and conjunction

periods)


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Solar Orbiter 2017 launch - EMP

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Figure 6.3-3: Sun distance and Sun/SC/Earth angle over the course of the EMP for the nominal 2017

mission; also shown are the principal mission events (science windows and GAM)

Solar Orbiter 2018 launch - EMP

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Figure 6.3-4: Sun distance and Sun/SC/Earth angle over the course of the EMP for the backup 2018

mission; also shown are the principal mission events (science windows and GAM)


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6.3.3 Space/Ground Interface

During the NMP and EMP, daily passes of 5 hours (4 hours effective TM-dump) shall be provided by New Norcia GS. This phase is supported by an increased FCT. Operations are conducted from the DCR. These passes shall support housekeeping and mission product download. There will be interruptions to this schedule caused by the operations of the SC, namely:

• A long-term conjunction lasting ~797 days in the nominal 2017 mission – the SC autonomy must be able to deal with this uninterrupted period of blackout. This is the driving case for the SC autonomy

• A long-term conjunction lasting ~60 days in the backup 2018 mission – the SC autonomy must be able to deal with this uninterrupted period of blackout; this blackout period is particularly problematic because it is coincident with VGAM-4, implying that the associated preparatory manoeuvres for VGAM-4 will have to occur several months before the VGAM, forcing a higher target pericentre and resulting in a possible delta V penalty to the mission – this is a critical issue which must be addressed.

• Periodic close approaches to the Sun, which will enforce stowage of the HTHGA at Sun-distances of below 0.3AU and prevent communications. It is also possible that an operational approach is adopted where the HTHGA is stowed behind the heatshield shadow for all science windows regardless of Sun distance, in which case the periods of blackout shall increase. The period varies but is typically of the order of ~10 days

• Interruptions due to interefence with the HTHGA FoV by the SC and SC appendages – these can be obviated by rolling the SC.

The sequence of operational blackout periods is shown in the figures below, overlaid onto the positions of the science windows and GAMs, for the 2017 and 2018 mission scenarios.

Solar Orbiter 2017 launch - NMP/EMP

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Figure 6.3-5: Non-communication periods during the NMP/EMP for the nominal 2017 mission; the

large period (circled in blue) is caused by the conjunction, the rest are caused by stowing the HTHGA during periods where the SC is closer than 0.3 AU to the Sun (they can be seen to coincide

with the perihelion science windows – blue diamonds)

7 This is assuming a 5° cut-off condition for the SC/Sun/Earth angle; the duration decreases with this angle.


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Solar Orbiter 2018 launch - NMP/EMP

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Figure 6.3-6: Non-communication periods during the NMP/EMP for the backup 2018 mission; the large periods (circled in blue and yellow) ars caused by the conjunction, the rest are caused by

stowing the HTHGA during periods where the SC is closer than 0.3 AU to the Sun (they can be seen to coincide with the perihelion science windows – blue diamonds)

6.3.4 Critical Operations

6.3.4.1 Delta V Manoeuvres

As a consequence of the need to maintain the reference attitude of the SC with the heatshield oriented towards the Sun below 0.7 AU, the TCMs for Solar Orbiter are split into two types:

• Type 1 TCM: these are unconstrained in attitude and allow slew of the SC to the optimum burn attitude and occur above 0.7 AU

• Type 2 TCM: these are constrained by the need to maintain the reference attitude with the heatshield oriented towards the Sun below 0.7 AU from the Sun.

6.3.4.2 Attitude Maintenance

During the NMP/EMP, a Sun-pointing attitude is maintained at all times, with the SC +X axis directed towards the Sun. For the majority of these two phases, the nominal attitude is maintained such that the XY plane of the PRF is parallel to the orbital plane, allowing the –Y unit vector to be maintained as close to the RAM direction as possible.; this is the nominal attitude for the in-situ instrument requirements. However, periodic roll around the sun-line will be required in order to facilitate communication between the Earth and the HTHGA. Roll around the sun-line shall also be used to allow omni-directional delta-V coverage for pre/post VGAM; however the SC shall maintain a Sun-pointing attitude throughout these manoeuvres. Nominal actuation is through the Reaction Wheels, with periodic momentum dumping using thrusters when required; thruster actuation shall also be used for delta V manoeuvres. During science windows, the wheels


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are operating to half-capacity (TBC), and accordingly momentum dumping frequently needs to occur during science windows. When this occurs it forces a suspension of the Science Mode.

6.3.4.3 Solar Array Temperature Management

Flipping

The solar array design is two-sided, with a hot face for close approach periods, and a cold face for periods further away from the Sun. Accordingly there will a periodic requirement for a flipping manoeuvre whenever the spacecraft crosses a certain Sun-distance threshold. Initial design of the 2-sided solar array shows the range of sun distances over which both the hot and cold faces of the solar panel can operate in the following figure:

Figure 6.3-7: Back-compatible solar array operational ranges

Note that there is a region of overlap between the operational range of the hot and cold faces (~few tenths of AU for operational robustness, in the graph above from 0.53 AU down to 0.33 AU). The solar array can theoretically flip at any sun distance within this overlap range. Here we take a preliminary Sun distance for flipping at the mid-point of this range, i.e. 0.4 AU (TBC). This will evolve as the design of the Solar Array is further defined, but serves as a starting point. Now looking at a figure of the SC sun distances over the course of the mission.


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Figure 6.3-8: Baseline 2017 mission scenario showing the evolution of Sun distance over the

mission, and the GAM events; the MRD-definitions of the CP, NMP and EMP are also shown; red line denotes preliminary array flipping distance

From the graph above, we can see that there will be 35 flip events of the Array in total (the first 6 events can be avoided if a flipping distance of less than 0.4 AU is targeted). Flipping events shall always be scheduled to fall outside of science window operations (this is enabled by the substantial overlap range between the hot and cold face array designs). During these flipping periods, the following sequence of activities shall occur:

• The power supply shall automatically switch from the solar arrays to the battery as the power demand exceeds the supply provided by the arrays

• The solar arrays are flip (initial time allocation to this is 15 minutes for battery sizing purposes) • The power supply shall automatically switch from the battery to the solar arrays as the generated

power exceeds the demand. The precise sequence of events is TBD.

Rotation

The temperature of the solar generators shall be managed throughout the large variation of sun distances experienced during the mission by changing the SAA appropriately; this shall take the form of discrete rotation events which shall manage the temperature whilst ensuring the provision of adequate power. This operation is of consequence to the mission during the operational phase of the mission (NMP, EMP) because the mechanical noise generated by array rotation will easily violate the RPE requirement for platform stability required by the RS-payloads during the fine-pointing science windows. Accordingly there is a requirement to avoid or minimise the occurrence of solar array rotations during the science windows. The following figures present example rotation profiles over the operational range of sun distances for two system designs of the solar arrays developed by GmbH. Two designs are presented here to illustrate the effect of SA design on the requirement to rotate. The two designs are:

• A 2-sided array with 37% cell population on the hot side and 100% cell population on the cold side, with a maximum allowable array temperature of 195 degrees C and a maximum allowable SAA of 78 degrees.


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• A 2-sided array with 20% cells population on the hot side and 100% cell population on the cold side, with a maximum allowable array temperature of 165 degrees C and a maximum allowable SAA of 78 degrees.

Shown in the following figure are the array temperature and SAA profiles, and the distance ranges of the science windows for the 2017 and 2018 mission scenarios. These profiles ensures the provision of adequate power throughout the orbit.

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Figure 6.3-9: Example incidence profile and temperature for the 1st solar array design for the 2017

baseline mission scenario


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Figure 6.3-10: Example incidence profile and temperature for the 2nd solar array design for the 2017

baseline mission scenario

Some general observations can be made for:

• The rotation profiles (as shown above) can be optimised for each orbit in order to minimise interactions with science windows, and it is expected that once the design of the solar generator has reached a suitable level of definition, the rotation profile for each orbit can be specified (including flipping of the array).

• The frequency of rotation events climbs as sun distance decreases (unsurprisingly because the rate of change of incident normal flux is increasing with the reciprocal of the square of the distance), and accordingly maximum latitude science windows are able to avoid array rotation events (large sun distance provides good flexibility to position rotation events outside of science windows because the rate of flux change is lower).

• For the 1st array design: o some minimum latitude science windows are likely to suffer one or two rotation events

during their ten-day period, and some perihelion science windows suffer probably at least two rotation events during their ten-day period (there will be two rotation events for some rotation distances because the peripassage science window is an inflection point) – see the following figure.


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Figure 6.3-11: The first peripassage of the 2017 nominal mission scenario. The red points indicate the beginning and end of the window; because the peripassage is an inflection point, there will be

two rotation events for rotation distances below the lower of the start/end distances

o Particularly @ ~0.25 AU, under the current system model, there is a strong requirement for a rotation event to satisfy both power requirements and temperature constraints, and rotation events will be frequent around this region.

• For the 2nd array design: o The much greater constraint on maximum temperature for the array design (165 degrees C

down from 195 degrees C) force the SAA to be much more actively managed during the close-sun periods; the luxury of a low number of discrete rotations is lost because the array cannot become too hot.

o Accordingly the array rotation is now quasi-continuous, and we can expect a large increase in the total number of array rotation events present within the near-sun science windows

• The worst case operational orbit is when the 3:2 resonance is initially entered, because the perihelion is minimised, and accordingly the radial distance travelled during science windows has increased. For the nominal mission, this is during the 2nd half of the NMP science windows; for the backup 2018 mission scenario, this is during the 1st half science windows of the NMP.

To summarise, discrete array rotation events will be present during some of the near-sun science windows (the peripassage and minimum latitude windows). The sensible approach will be to combine these events with other window interrupting events, such as possible HTHGA movements, re-pointing, and reaction wheel off-loading. It is hopefully the case that rotation events can be positioned outside of the maximum latitude window due to the reduced flux change rate at these greater distances. From the figure and from the science window definitions, it can be seen that, in terms of Sun distance, the science windows can be grouped into the following groups:

• NMP_1st half (2017 windows 1 – 12; 2018 windows 1 - 9) • NMP_2nd half (2017 windows 13 – 21; 2018 windows 10 - 18) • EMP_1st half (2017 windows 22 – 30; 2018 windows 19 - 27) • EMP_2nd half (2017 windows 31 – 39; 2018 windows 28 - 36) • EMP_final (2017 windows 40 – 48; 2018 windows 37 - 45).


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We can make a preliminary estimate of the likely number of rotation events per group – this is the appropriate level of accuracy at this stage, given the uncertainties in the solar generator design and performance.

2017 2018 Min_Latitude Peri_Passage Max_Latitude Min_Latitude Peri_Passage Max_Latitude NMP 1st half 1 2 0 0 2 1 NMP 2nd half 2 2-4 0 0 0-2 2 EMP 1st half 1 0 0 0 0-2 1 EMP 2nd half 1 0 0 0 0-2 1 EMP Final 0 0 0 0 0-2 0

Table 6.3-1: Projected rotation populations for science window groups for the 1st array design

Caveat: for the 2nd array design the number of rotation events within near-sun science windows is MUCH higher. This serves to reinforce the conclusion that until array design is known, rotation populations cannot be accurately determined at present. These profiles do not take into account any specific operating requirements of the SAs, but simply ensure satisfaction of the maximmum operating temperature of the arrays whilst providing adequate power. Array temperature or SAA constraints may exist at specific points in the operational orbit due to for example stray-light issues – these may introduce additional constraints, but this is beyond the scope of this phase.

6.3.4.4 SC Rolling

Due to the presence of the various appendages and the body of the SC itself, the nominal RF-link to the Earth via the HTHGA shall periodically become occluded due to occultation of the Earth’s position in the PRF of the SC, when the SC holds the nominal attitude, i.e. sun-pointed with the velocity vector lying in the orbital plane. For at least the periods when the SC body is blocking the LoS of the HTHGA to the Earth, it shall therefore be necessary to roll the SC around the sun-line (PRF X-axis) in order to shift the Earth to a position in the PRF where the HTHGA can access it. An example shift of the Earth positions is shown in the following figure for the formerly considered 1_RPW configuration.


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Figure 6.3-12: The transformation in Earth positions in the PRF when a 60 degree roll applied to the SC away from the nominal attitude

The roll profile of the SC throughout the 2017 nominal mission scenario is shown below, overlaid onto the primary mission events (GAMs, science windows…) and mass memory storage. The FoV interruptions are divided into RPW antenna interruptions (red), which can conceivably be ignored/tolerated depending on the precise form of the interference and signal degradation experienced, and interruptions due to the SC and solar arrays (light blue), which are likely to be complete and will force a roll of the SC, either throughout the period or temporarily for every scheduled pass.


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The table below shows that for the 2017 trajectory, the total proportion of time during the mission that the SC is rolled is ~15%, if roll is used to avoid the RPW antennas. If the RPW interference is tolerated (something that seems possible when the SSMM profile of the mission is considered; rolling for short-duration RPW blockages of only a few days [red spikes in the figure below] can be avoided by electing to not perform passes for the duration, or accepting a degradation in signal), then the roll required just for the solar arrays and SC body is ~8%. These are conservative figures and represent an upper limit.

RPW+Body/Arrays Body/Arrays only2017 14.9% 8.2%

Table 6.3-2: Percentage of mission time spent in the rolled state

Figure 6.3-13: Baseline configuration roll profile for the 2017 mission scenario

6.3.4.5 Science Window Operations

The science window operations of Solar Orbiter are both complex and key to the mission success. Many operational factors are congruent within the science windows, and as such the design of the Solar Orbiter SC is highly dependent upon the operational approach chosen. A key design objective for the Solar Orbiter mission is the provision of long periods of uninterrupted RS-observations of the Sun during the science windows. The original requirement for Solar Orbiter [RD3] is to provide 10 day uninterrupted periods of Remote-Sensing observations centred around 3 locations in the operational orbit. The original design hope was to be able to provide the required platform stability for RPE observations over this period. However, in reality, there are several events that will cause interruption to the science mode of the SC during the science windows; these are: Reaction Wheel off-loading: Any net DC SRP-torque experienced during the science windows must be counteracted by the reaction wheels on-board. However, should the HTHGA be exposed to the solar flux during a science window, the magnitude of the net SRP-torque around the SC Y-axis will be such that one or more RW offloading events is necessary within the science window due to the wheels becoming saturated. The frequency of such events will be different for each science window, and is a complex function of the SC design and the evolving parameters of the particular science window; consequently some science


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windows will require no wheel off-loading at all, whilst other windows will require several wheel off-loading events. These events will necessitate a cessation of the remote-sensing because the noise introduced by the thrusters will easily violate the required RPE. Repointing: The scientific operations will contain one or more repointing events to allow different areas of the solar disc to be targeted (e.g. the movement from Sun-centre to limb-pointing); the number of repointing events is expected to be a function of the particular scientific interest of the current window; this repointing will necessitate an interruption to the required platform stability. Periodically during science windows, subject to planning by the scientists and the SOC, re-pointing of the SC platform to different regions of the solar disc will occur. Because these events are determined by detailed planning in advance of the science window (subject to many drivers such as the position on the Solar disc of interesting phenomena), it is not possible to provide schedule information regarding the timing or frequency of these events. However, it is envisaged that, during a typical ten-day science window period, at least one repointing event shall be required by the payload suite. Note that it is likely to be operationally advantageous to ensure repointing of the platform is coincident with reaction wheel off-loading, HTHGA movement or SA rotation. In this way the number of disturbances to the fine pointing is minimised because these housekeeping activities are performed simultaneously rather than separately. Repointing will necessitate suspension of the science operations as the platform stability requirements will be violated. The RS-instruments will enter standby mode for the duration of the re-pointing manouevre. Entry into standby mode will coincide with door closing for all payloads. Once the doors are closed the repointing manoeuvre shall be performed using the RWs in AOCS mode NM (Sub_mode Sun-pointing). Additionally, desaturation of the wheels may be concurrently performed. Once the required platform RPE stability has been achieved, doors will open and the SC wiill recommence RS operations. It is not envisaged that door-opening shall be problematic to the stability of the platform, as it will be mechanically quiet and of limited duration. Solar Array Rotations: The Solar Array will require several discrete rotations within science windows in order to manage the temperature of the array. Solar Array rotation will violate the RPE requirements. As with the Reaction Wheel off-loading, the number of Solar Array rotations required will be a function of the particular parameters of the science window, specifically the range of Sun distances covered by the SC over the duration of the science window, and just as critically the final design of the solar array. Communications: Periodic communications during science window will be required during above 0.3 AU windows, and could even occasionally require rolling of the SC around the sunline. In principle the design of the mission operations should be so as to combine interrupting events (including communications) to maximise the uninterrupted observation durations.


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System Modes The following section presents the various operational modes and sub-modes of the system (AOCS modes are dealt with in the AOCS modes section).

6.3.5 System States and Modes Dictionary

The dictionary of system modes (functional configurations) is shown in the table below. Note that redundant configurations are not considered as distinct states, but as the same state with different parameters. Mode Designation Description Associated AOCS Mode Off OFF This mode is never used in orbit; it is only used during AIV between

test campaigns and during transport. -

Standby SBY In this mode the basic services are available (communications, data-handling and power distribution) in a default configuration such that the SC can act on and reply to TC. All other subsystems are turned OFF but can be turned on and placed in their various modes by TC. This mode is used for AIV and also during the launch.

SBM

Initialisation INIT Upon detection of the launcher separation this mode is automatically entered from SBY by monitoring the status of the monitor switches on the separation interface (TC can also be used). This mode performs the initialisation of the SC, which brings the SC into a stable attitude and configuration ready for the commissioning activities.

SBM, RRM, SAM

Safe SAFE Functionality in this mode is reduced to a minimum. Criticality in this mode is ensuring the SC safety, primarily by holding the correct attitude.

SAM, SKM

Commissioning COM This mode oversees commissioning of all subsystems including the payloads, which are held in standby mode such that individual commands can be tested. In support of commissioning activities this mode also supports the full functionality of the SC (communications, AOCS, thermal…).

RWO, IPS, FSP (NM)

Manoeuvre MAN This mode manages the periodic manoeuvres required during the mission; it implements the entire manoeuvre sequence, from initial slew (where required) through the burn to the final slew (if required). This mode is by necessity automatic due to the loss of Earth tracking during the manoeuvre for Type 2 TCMs. For Type 1 TCMs communications can be maintained during the manoeuvre.

IPS, OCM

Nominal NOM This mode is used when any of the SC payloads are operating. Accordingly it is used for the majority of the mission after the beginning of the cruise phase. This mode covers the operation of the SC throughout the CP/NMP/EMP, except for manoeuvre or contingency events.

RWO, IPS, FSP (NM)

Table 6.3-3: The Solar Orbiter System Modes

6.3.6 System States and Modes Transistion

The transition between the modes in the table above is expressed in terms of Mode Transition Diagrams (MTD). The diagram defines the allowed transitions between modes, and whether these transitions are automatic or TC-driven. Note that in Figure 6.3-14:

• TC means external telecommand, either real-time or time-tagged • AUTO means automatic events driven by on-board logic • FDIR means automatic events driven by the FDIR • Some modes and transitions are used only for ground AIV – these are shown in lighter colours.

The following Table 6.3-4 lists the transitions between modes.


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OFF

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AUTO(separationdetection)

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TC

Figure 6.3-14: Solar Orbiter system mode transitions diagram


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Transition Event AUTO/TC Event Description OFF to SBM Power ON/OFF Ground AIV only. SBM to OFF Power ON/OFF Ground AIV only. SBM to INIT AUTO (separation detection) This event is driven by the detection of the launcher separation

event using TM from separation interface monitors. A suitable time-delay is used to ensure safe start of the INIT a suitable period after the separation event.

INIT to SBM TC (AIV only) This never takes place during the actual mission, and is TC-driven during AIV activities only.

INIT to COM AUTO This transition takes place when the initialisation of the SC during LEOP is completed, and the SC is ready to enter the commissioning phase of the mission. Once the INIT mode is left there is no possibility of return to the INIT mode afterwards.

COM to NOM TC Upon successful completion of commissioning activities, a TC will command a switch to NOM.

NOM to COM TC TC commands a switch to the COM as required, primarily during the CP when residual commissioning activities will take place.

NOM to MAN TC As commanded by the ground, always as a time-tagged TC (OBCP) for the beginning of the manoeuvre sequence.

MAN to NOM AUTO Upon successful completion of the manoeuvre sequence, an AUTO-commanded transition occurs back to NOM.

All to SAFE FDIR/TC When a non-recoverable failure is detected the system or subsystem FDIR commands a transition to the SAFE mode. Alternatively, a TC can command the transition. Return to NOM is TC-commanded once the problem has been fixed by the ground.

Table 6.3-4: List of system mode transitions


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7. MECHANICAL SUBSYSTEM

For a more complete description of the mechanical subsystem, please refer to [IR50], and [IR47] for a description of the instrument boom design.

7.1 Requirements

The principal requirements placed on the mechanical subsystem are shown in the following table.

Quantity Requirement Comments

Configuration Space environment constraints

Need to support a heatshield (HS) on one face of the spacecraft, (this dictates much of the rest of the configuration)

Configuration Payload constraints

Need to provide stable/robust support for the remote sensing instruments which allows viewing in the sun-direction, (following on from the HS, it means that the stable/robust structures need to be close to the HS face plus there will also need to be holes in this face),

Need to provide stable/robust support for the in-situ instruments which ensures the specified field of views,

Accommodation Deployable appendage accommodation

Need to support deployable appendages (HGA, solar arrays, MGA, Boom, RPWs etc) for both launch and in operational configuration ; this means that there need to be sufficient hard-points accessible on the exterior of the SC for supporting these appendages,

Accommodation Interior unit and payload accommodation

Need to provide sufficient mounting area for platform and payload electronic equipments ; this means that there will need to be a certain available area of flat panels, adequately accessible to space for thermal purposes on which to place all this equipment.

AIV Allow sufficient AIV access Need to ensure that certain panels are removable for AIV access ; this means that not all of the exterior face of the structure will be doing useful structural jobs,

Accommodation Exterior unit and payload accommodation

- Need to provide sufficient mounting area for platform and payload electronic equipments ; this means that there will need to be a certain available area of flat panels, adequately accessible to space for thermal purposes on which to place all this equipment.

Configuration Launch vehicle interface Need to interface with the identified launch vehicles (including launch vehicle adapter interface and fairing limitations) ; this means that there will be a certain face with an adapter interface and that all loads paths from the structure will have to be transferred to this interface.

Cost/Schedule Reuse of existing mechanical design where appropriate

Need to reuse existing design solutions as much as possible. Assessment study identified the Mars Express / Venus Express platform as a suitable solution for Solar Orbiter.

Table 7.1-1: Solar Orbiter mechanical subsystem key requirements


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7.2 Design Overview

The spacecraft configuration consists of a closed box type structure on which all the instruments and accompanying equipment are mounted. Several instruments are mounted on appendages (instrument boom and RPW). Mounted on the top panel (+X) of the SC a heatshield is providing the required shielding from the direct solar illumination. Additionally two propellant tanks are accommodated within the box structure. The structure subsystem essentially consists of

• A core structure • An outer structure • Brackets, joints and attachment devices.

7.2.1 Core Structure

7.2.1.1 Launch Vehicle Interface Ring (LIR)

The Launch Vehicle Adapter Ring (LIR) forms the lower end of the structure and provides the interface to the launcher. It consists of a forged ring made of 7075 T7351 with 200 mm height and 1000 mm external diameter. The LIR is reinforced by two stiffeners that react the propellant tank lower interface loads. The design of the LIR is directly taken from the Mars Express/Venus Express spacecraft with the only possible differences being due to the detailed attachment interfaces of the propellant tanks (depending on the final selection of the tank supplier).

Figure 7.2-1: Solar Orbiter LIR Lay-out

The top flange of the LIR interfaces with the lower floor via 26xMJ6 bolts. Six (6) brackets have been implemented in order to stiffen the whole spacecraft . They connect the shear walls to the LIR. The brackets are made of 2 pieces bolted together. One piece is attached to the LIR, the second piece is supporting the sandwich panel as high as possible via a fork. The panels are locally reinforced with doublers in order to withstand the very high loads (i.e. main load path). The stiffeners are made of 4 pieces riveted together i.e. 2 plates for the flanges and 2 machined profiles for the webs. The webs are locally reinforced with ribs in order to support the lower tank interface. The stiffeners are bolted and riveted to the LIR. The stiffeners are offset by different amounts from the spacecraft centre line due to the need to balance the mass moments due to the differing masses of the propellants in the bipropellant tanks. Four (4) separation spring pads are bolted close to the lower end of the LIR. At the lower end of the LIR, a clamp band that can be pyrotechnically released attaches the LIR to the launcher LVA. The LVA is the standard 937 interface and the interface characteristics are given on the drawing in Annex A.2.

7.2.1.2 Central Shear Wall (-Y)

The position of the -Y central shear wall is shown on the exploded view below.


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Figure 7.2-2: Exploded view with location of –Y central shear wall

The -Y central shear wall gives the spacecraft shear stiffness in the X-Z plane. The face sheet thickness is 0.25 mm and the core height is 20mm. Local reinforcements (i.e. denser core 1/8-5056-0.001P, doublers on both face sheet sides and special inserts) are implemented close to the lower floor, around the LIR brackets and along the ±Y/±Z shear walls. The shear panel design and configuration is essentially the same as MEX/VEX/Aeolus in terms of general construction and panel to panel joints however there will be specific modifications in order to accommodate instrument specific interfaces. The -Y central shear wall supports or provide an interface to:

• The two tanks support panel via inserts • The Helium pressurant tank mounting assembly, • The instrument bays lateral closure panels, • The top floor, • Several of the heatshield connection point with the S/V via edge insert of metallic fitting (TBC) • One hoId-down point of the HGA via a special edge insert (TBC) • The HGA mechanism. • One of the RPW antenna HDRM and mechanism via edge inserts (TBC).

The -Y central shear wall supports the following equipment and boxes:

• The reaction wheel assemblies (on one side of the wheel trihedral interface) • FCE • IMU • SADE (2x) • EPD Ebox • MAG Ebox • RPW Ebox • the CPS sub system.

Cut outs (e.g. for the HGA wave guide, harness, fluid loops, ...) and grounding possibilities are foreseen.

7.2.1.3 ±Z / ±Y Shear Walls

The position of the ±Y/±Z shear walls is shown on the exploded view below.


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Figure 7.2-3: Exploded view with location of ±Y±Z shear walls

These panels have to provide shear stiffness in the X-Y plane. They also contribute to the bending stiffness of the lower floor. The face sheet thickness is 0.50 mm, the core height is 20 mm. In the area close to the lower floor a denser core is required.(i.e. 1/8-5056-0.001P) in order to ensure the integrity of the panel attachment inserts against pull out. Around the LIR bracket doublers are bonded on both face sheets in order to reinforce the panel. In the area of the pockets (necessary for access during panel assembly), doublers are bonded to both face sheets. Reinforcing brackets are also implemented around the lifting points (4 points on the +X/±Y sides) to sustain the local loads in this area during lifting. The shear panels design and configuration is essentially the same as MEX/VEX/Aeolus in terms of general construction and panel to panel joints however there will be specific modifications in order to accommodate instrument specific interfaces. The +Z/-Y shear wall interfaces with PHI and PHI Ebox via potted inserts. Adjustment provision is provided by peel-off shims The +Z/-Y shear wall is attached to the -Y sidewall, the -Y central shear wall, the –Y tank support panels, the top floor and the lower floor via edge inserts (TBC). The +Z/+Y shear wall interfaces with SPICE and EUI as well as their respective Eboxes via potted inserts. Due to high total mass of these instruments, a core 1/8-5056-0.001P is foreseen. For both instruments, adjustment provision is provided by peel-off shims. The +Z/+Y shear wall is attached to the +Y sidewall, the -Y central shear wall, the +Y tank support panels, the top floor and the lower floor via edge inserts (TBC). The -Z/-Y shear wall interfaces with STIX via potted inserts. Adjustment provision is provided by peel-off shims. The -Z/-Y shear wall is attached to the -Y sidewall, the -Y central shear wall, the -Y tank support panels, the top floor and the lower floor via edge inserts (TBC). The -Z/+Y shear wall interfaces with METIS via potted inserts. Adjustment provision is provided by peel-off shims. The -Z/+Y shear wall is attached to the +Y sidewall, the -Y central shear wall, the +Y tank support panels, the top floor and the lower floor via edge inserts (TBC). This panel also features a local extension (towards –X) to provide support to the star trackers externally.

7.2.1.4 Tank Support Panels

The position of the tank floors is shown on the exploded view below.


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Figure 7.2-4: Exploded view with location of 1 tank support panel

Basically, the two tank floors connect the -Y central shear wall to the centre parts of the ±Y/±Z shear walls, closing the high-loaded tank box at the top. The face sheet thickness is 0.50 mm and the core height is 20 mm. They provide the upper interface to the propellant tank. The interface consists of a spherical bearing mounted in a fork attached to inserts. At the insert location, shims are bonded to have a distance of TBD mm between the two legs of the fork in order to be able to procure standard MMS parts. The spherical bearing allows the structure to rotate without introducing stresses in the tank. The tank boss is bonded to the spherical bearing, therefore the tank floor stiffness has been adjusted in order to allow sufficient flexibility for the tank axial displacement while being stiff enough to reach the required first resonance frequency.

7.2.2 Outer Structure

7.2.2.1 Heatshield Mounting Panel (Top floor +X)

The position of the top floor (+X panel) is shown on the exploded view below.

Figure 7.2-5: Exploded view with location of top floor

The top floor closes the top end of the structure. This is a sandwich panel with thin face sheets (0.25 mm) and a 20 mm thick core high density (i.e. 1/8-5056-0.001P). The top floor provides the interface for the heatshield. The top floor connection concept with the Heatshield is shown in the picture below :


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Figure 7.2-6: Top Floor connection with Heatshield

Bi-pods at each of the four corners are used to react the shear loads in the Y/Z plane. Monopods along the shear walls are used to react the loads along the X axis. The concept is designed in order to be able to minimise the load introduced in the top floor while the heatshield is thermally expanding, and to allow for precise positioning of the HS interfaces on the SC. To that effect the design is established to react most of the loads in the ±Y/±Z shear walls, the –Y central shear wall, the ±Z closure panels and the ±Y sidewalls. The HS interfaces are attached to face to face inserts (metallic fitting) of the top panel. The top panel, at each of these interfaces, is attached to the underlying panels via brackets bolted to potted inserts or metallic fitting on the shear walls. The top floor interfaces via potted inserts with the following items (see Appendix A1) :

• SoloHi support bracket, • SWA PAS and SWA HIS support brackets (the top floor is locally extended around the –Z/-Y and –

Z/+Y corners to improve stiffness and thermal stability) • RPW hinges / mechanism.

The top floor interfaces via face to face inserts (metallic fittings) with the following items (see Appendix A1) :

• High Gain Antenna HDRM (2 out of 3) are interfaced with top panel • Solar arrays HDRM • MGA HDRM (TBC).

The top is attached to the ±Y/±Z shear walls, and the –Y central shear wall, and to ±Y sidewalls and ±Z closure panels via edge inserts. Other cut-outs (e.g. for remote sensing equipment pupils, doors harness, …) and grounding possibilities are also provided. Note that in order to facilitate late integration of the METIS instrument (which has its own baffle), it is envisaged to implement a cut-out access panel within the top floor.

7.2.2.2 Lower Floor (-X)

The position of the lower floor panel (-X) is shown on the exploded view below.

bipods

monopods


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Figure 7.2-7: Exploded view with location of lower floor

The lower floor is optimised with respect to stiffness and proper load transfer into the LIR leading to the following characteristics, a face sheet thickness of 0.25 mm and a core height of 20 mm. The lower floor is attached to the LIR via i.e. 26xMJ6 bolts. The lower floor provides the following interfaces:

• Via face to face inserts (metallic fittings) to the:

o 4 thrusters brackets (TBC) o 3 Star Tracker brackets o Umbilical connector bracket o 1 Low Gain Antenna o The power system with the SADM bracket o Instrument boom hinge / mechanism and support beam o High Gain Antenna hinge / mechanism and support beam (TBC) o Medium Gain Antenna hinge / mechanism.

• Via potted inserts to the:

o Fill valves and drain valve brackets o Purging system valve brackets (2x) o Sun sensor (1x) o EPD EPT and EPD HETN in situ instrument brackets. o

The lower floor panel is attached to the ±Y/±Z shear walls, and the –Y central shear wall, and to ±Y sidewalls and ±Z closure panels via edge inserts. Two holes (128 X 108 mm) (TBC) allow attachment of the propellant tanks directly to the stiffeners on the LIR. Six cut-outs allow for direct connection between the LIR and the shear walls (i.e. main load path). Around these cuts-outs a local doubler is required for strength reasons on the top face sheet (TBC/TBD). Around the 26 attachment points a doubler is also required for strength reason on the bottom face sheet (TBC/TBD). Other cut outs (e.g. for pipes, harness,...) and grounding possibilities are provided.

7.2.2.3 ±Z Closure Panels

The position of the closure panels is shown on the exploded view below.


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Figure 7.2-8: Exploded view with location of ±Z closure panels

The closure panels are made of sandwich (face sheet 0.2 mm and core 10 mm high). They close the box structure. They serve as access / removable panels for the remote sensing instruments bays. On the +Z side, the +Z/+Y closure panel provides interface for:

• The instrument boom HDRM and hinge / mechanism • The +Z RPW antenna HDRM.

lt is interfacing with the top panel, lower floor, -Y central shear wall and +Y sidewall via edge inserts or cleats. Grounding possibilities are also provided. The +Z/-Y closure panel provides interface for:

• The Medium Gain Antenna HDRM and hinge / mechanism • SWA-PAS support bracket (panel has local extension to improve stiffness & thermal stability).

lt is interfacing with the top panel, lower floor, -Y central shear wall and -Y sidewall via edge inserts or cleats. Grounding possibilities are also provided. On the -Z side, the -Z/-Y closure panel provides interface for:

• One of High Gain Antenna HDRM (at interface with +X panel) • EPD LET mounting bracket • -Z/-Y RPW antenna HDRM and hinge mechanism (at interface with –Y sidewall) • EPD EPT bracket (at interface with –Y sidewall) • SWA-HIS bracket (panel has local extension to improve stiffness & thermal stability).

lt is interfacing with the top panel, lower floor, -Y central shear wall and -Y sidewall via edge inserts or cleats. Grounding possibilities are also provided The -Z/+Y closure panel provides interface for:

• One of High Gain Antenna HDRM (at interface with +X panel) • One of High Gain Antenna HDRM (at interface with –Y central shear wall) • One of High Gain Antenna HDRM and hinge / mechanism (at interface with -X lower floor) • -Z/+Y RPW antenna HDRM and hinge mechanism (at interface with +Y sidewall).

lt is interfacing with the top panel, lower floor, -Y central shear wall and +Y sidewall via edge inserts or cleats. Grounding possibilities are also provided.


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7.2.2.4 ±Y Sidewalls

The positions of the ±Y sidewalls are shown on the exploded view below :

Figure 7.2-9: Exploded view with location of ±Y closure panels

These panels form the remaining two sides of the closed box. They support several internal equipment and eboxes. They are also used to provide radiator area. They are made of sandwich (face sheet 0.25 mm and core 20 mm high). The ±Y sidewalls are also providing access (by way of cut-out access panel) to the central bay and the propulsion subsystem:

Figure 7.2-10: Y wall access panel location

Each of the ±Y sidewalls provides 4 solar array hold-downs mechanism interface (with local reinforcements using block inserts, doublers, and higher density honeycomb core. The SA HDRM interfaces are located at interface between the ±Y sidewalls and the ±Z / ±Y shear walls. The +Y sidewall (see Appendix A1) interfaces with:

• The communication system with KA-band TWT and EPC’s via potted inserts, • The CDMU and Remote Interface Unit boxes in the +Y/-Z bay, • Two pairs of thrusters, in the +Y central bay, • The star trackers support bracket (via a local panel extension).

Divers other cut-outs (e.g. for pyro, harness,...), MGSE points and grounding are also provided. The -Y sidewall (see Appendix A1) interfaces with:

Access panel within ±Y sidewalls

X ±Y


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• The communication system with KA-band transponder via potted inserts • One LGA antenna • PCDU and batteries in the –Y/-Z bay. • Two pairs of thrusters, in the -Y central bay. • The solid state memory box in the +Z/-Y bay.

Other cut-outs (e.g. for pyro, harness,...), MGSE points and grounding are also provided.

7.2.2.5 Brackets

The reaction wheels, brackets have been designed to minimise dynamic amplification and to have a first frequency above 100Hz hard mounted at the interface. Solar arrays, MGA, HGA, RPW, instrument boom brackets / interface design is TBD. Thrusters, connector brackets, LGA, Star Trackers and IMU brackets design is TBD.

7.2.3 Instrument Boom Design

Several of the instruments are located on an instrument boom which extends in the –X direction within the shadow projected by the heatshield. The selected overall boom design concept is derived from the Swarm programme, consisting of a CFRP tube with bonded interface fittings at the hinge mechanisms and instrument locations. The hinge mechanism and HDRM designs are both derived from Swarm (see [IR47]).

Figure 7.2-11: CFRP Tube with Bonded Titanium Fittings


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Figure 7.2-12: Hinge mechanism is Titanium and Al. Alloy

Figure 7.2-13: HDRM Release actuator and Titanium separation interface


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8. HEATSHIELD

8.1 Requirements


Geometric The Heatshield shall protect the spacecraft and its payload from the sun radiation flux and act as thermal insulation to provide an environment that is suitable for the instruments and SC. Exceptions to this shielding requirements are for appendices as antennas, solar arrays and their jokes, sun sensors and payloads

These requirements drive the shape and the size of the heat shield along the Y and Z axis. Having the SC a squared cross-section, and being unnecessary to protect the appendices of the SC, the HS has a squared cross section.

Geometric The Heatshield shall be sized to provide complete sun-shielding to the entire spacecraft including all equipments and parts of the structure within a 8o half-cone

This requirement sizes the heatshield overspill with respect to the SC structure in tandem with the projected performance of the HW FDIR subsystem. The current performance of the FDIR requires a heatshield protecting the SC withn a 8º half cone.

Thermal The heatshield shall limit the average radiative heat flux to the SC to ±30 W (TBC) during all mission phases. Radiative heat-flow analyses shall consider that the spacecraft body is a blackbody with a temperature of 20oC.

Thermal The Heat Shield shall limit local conductive heat flux at all attachment points to the SC structure to a total of ±15 W (TBC). This includes conductive flux through Heat Shield mounts and through any MLI required by the heatshield on the SC +X face.

These requirements determine the overall thermal design of the heatshield, from gap spacing through to material selection.

The HS could be a flat High Temperature MLI, possibly separated from the SC body by a single small gap to divert to space the heat that has crossed the MLI, support for this MLI could be rods of insulating material.

Functional The Heatshield shall provide apertures through which payload can directly observe the sun.

Functional/mechanical The Heatshield shall provide mechanical support for Feed throughs, mechanisms, thermistors, thermostats and their associated harnesses

The heatshield must support the feedthroughs, doors and mechanisms that allow the required FoV for the RS-instruments to be provided.

Thermal The Heat Shield shall limit IR heat transfer to the Remote Sensing instrument suite (PHI, EUI, STIX, EUS, COR) to less than 50 W (TBC). Radiative heat-flow analyses shall consider the remote sensing instruments as blackbodies with a temperature of 20°C.

Of course the feedthroughs are an important heat path, bypassing the HTMLI and injecting heat not only into the instruments via radiation, but also into the support panel of the heatshield itself.

An unobstructed lateral view to space is critically imnportant for the feedthroughs to reject heat radiatively laterally, and this (combined with conduction into the support panel) favours spreading the feedthroughs out from each other and towards the heatshield edges as much as possible (this is of course practically limited by instrument location constraints).

Thermal The Heat Shield shall limit IR heat transfer to the SWA-HIS, SWA-PAS, and Sun Sensors to a total of less than 40 W (TBC). Radiative heat-flow analyses shall consider the SWA and sun sensors as blackbodies with a temperature of 20°C.

This requirement places upper temperature limits on the feedthrough structure itself.

Thermal/mechanical The Heatshield shall ensure that the stability of all Feedthrough interfaces w.r.t the Heatshield interface datum is maintained to within 0.5º (TBC) and 1.0mm (TBC) in any direction and under all operational environments.

This thermo-mechanical requirement drives the design of the support panel, to limit thermal gradients and overall distortion; it places a premium on insulating the MLI support brackets and feedthrougs from the support panel, and using a high-conductivity material for the support panel itself to limit gradients.


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Mass The Heatshield mass (exc. FD&M) shall be no higher than 68kg including all margins.

This is the mass allocation to the heatshield in the overall system.

Mechanical The Heatshield shall be compatible with limit QS accel. Of 35g out-of-plane and 20g in-plane.

The heatshield must be mechanically compatible with the rest of the SC.

Table 8.1-1: Heatshield principal requirements

The heatshield dimension is dependent upon the spacecraft sizing and also the accommodation of sensitive units on the SC exterior. At present the heatshield sizing has occurred under the following directives: The heatshield shall provide shadow protection to all sensitive elements up until an angular excursion from the nominal pointing of 8 degrees, except for those payload units which have FoV requirements which restrict their ability to comply with this performance. Accordingly there are some payload units that will be exposed before the 8 degree maximum excursion is reached. This design directive is therefore under the assumption that these payloads ‘value’ satisfaction of FoV requirements above avoidance of direct illumination. A trade-off is therefore required for each of these payloads between satisfaction of FoV requirements against tolerance of direct illumination during failure cases.

8.2 Design Heritage

The design of the HS is the result of several iterations of design and analysis. Please refer to the following documentation for a history of the Heatshield design:

• Solar Orbiter Heatshield & System Technology System Design Report [IR07] • Solar Orbiter System Design Report [IR15] • Solar Orbiter Design Changes due to the Dedicated Atlas Launch and integration into the HELEX

Concept [IR13] • Solar Orbiter Heatshield Design Description [IR23] • Solar Orbiter Heatshield Design Trade Offs and Design Concept Proposal [IR24] • Solar Orbiter Heatshield Thermal Analysis [IR22]

The latest piece of trade-off work [IR24], resulted in the following synthesised HS design described here; this has the following key features:

• Total height of the HS: 400mm • Main gap including LTHBand HTHB: 265mm • Support panel: 55mm • Secondary gap including MLI on SC: 100mm.

The heatshield dimension is dependent upon the spacecraft sizing and also the accommodation of sensitive units on the SC exterior. At present the heatshield sizing has occurred under the following directives: The heatshield shall provide shadow protection to all sensitive elements up until an angular excursion from the nominal pointing of 8 degrees, except for those payload units which have FoV requirements which restrict their ability to comply with this performance. Accordingly there are some payload units that will be exposed before the 8 degree maximum excursion is reached. This design directive is therefore under the assumption that these payloads ‘value’ satisfaction of FoV requirements above avoidance of direct illumination. A trade-off is therefore required for each of these payloads between satisfaction of FoV requirements against tolerance of direct illumination during failure cases.


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8.2.1 Heatshield Dimensions

The heat shield dimension [2.5mx2.5m] is dependant on the spacecraft size and the function the heat shield is intended to perform:

• For accommodation reasons, the spacecraft size is (X x Y x Z) (1450 mm x 1700 mm x 1685 mm) it is illustrated in the following figures

• The function of the HS as stated in [AD 02] “The Heatshield shall be sized to provide complete sun-shielding to the entire spacecraft including all equipments and parts of the structure within a 8o half-cone”.

The functional requirement is intended to ensure that the design of the HS will provide sun shielding to the SC and externally mounted equipments and appendages in all operating position of the SC and also in case of FDIR events. In that case the off pointing from sun center excursion angle can reach up to 6.5°.

Figure 8.2-1: Solar Orbiter overall size (I)

X

Y

Z

8°

(400) mm

2500 mm x 2500 mm

SPACECRAFT

Support panel

HS upper layer / Sun shield

FDIR max recovery half cône angle

(1450) mm


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Figure 8.2-2: Solar Orbiter overall size (II)

8.2.1.1 Location of Protruding Elements on External Walls

Some elements of the SC (either instrument, or appendages or mechanisms) are protruding from the lateral wall external dimensions. One needs to verify that these elements are not going to be exposed to sun illumination in case of FDIR. The following table lists the payloads and the excursion angles at which they become exposed.

Equipment Distance from

top of the Heatshield

(mm)

Distance from the mounting

wall (mm)

Mounting wall

Shadow margin (*)

(mm)

Max off pointing angle (from sun

center)for no negative margin

SoloHI 890 325 +Y -50 3.5° EPD SIS 1375 325 -Y -120 1.8°

EDP LET 1 1690 180 -Y -20 6° EPD LET 2 1690 215 -Z -47 5.2°

MGA mechanism 1850 100 +Z 45 N/A HET 1850 200 -Y -62 4.8°

EPT 2 1900 100 -Y 31 N/A Solar Array 1995 45 +Y –Y 72 N/A

HGA antenna

YS

ZS

O

XS

(1700)

(1685)

(2500)

(2500)


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mechanism HGA mechanism 2255 80 -Z 8 N/A Instrument Boom

mechanism 2400 35 +Z 32 N/A

(*) Positive value means margin exist between shadow limit and protruding elements location Negative value represent appendages length exposed to the sun (outside of shadow limit)

Table 8.2-1: Protruding element from SC external walls

It can be seen that the current baseline for the HS size allows to protect from direct sun illumination, all appendages (MGA, SA, HGA BOOM) even in FDIR events. However EPD LET 1 and 2, SoloHI, HET as well as EPD SIS are exposed to the sun when off pointing angle of the SC is above the values quoted in the table. It is proposed to address this issue by locally increasing the size of the HS top layer. When including reasonable margin, the local increase would be between 50 and 150 mm for these instruments. Note that the HGA mechanism is very close to be illuminated as well (8mm). FDIR off pointing excursion are limited in duration to very brief time; increasing the HS size should not be necessary for appendages and mechanism, high temperature MLI should be sufficient (TBC).

8.3 Thermal Design

The Heat Shield (HS) occupies a volume on the +X side of the SC, directly in front of the sun. The HS function is to protect the SC from the sun, and to achieve this it consists of a combination of barriers orthogonal to the incident flux, combined with gaps used as lateral paths to space for the heat crossing the barriers.

The HS components are:

1. Insulating layers (listed from the layer exposed to the sun):

a. Front Shield (FS), it is the layer exposed to sun

b. High Temperature Heat Barrier (HTHB), mounted directly behind and in contact with the FS. It is an MLI designed for high temperatures, with the functions to stop all light that might be transmitted by the FS, to make an efficient barrier for heat transmission towards the SC, to create a highly reflective surface in the infrared in the HS main gap.

c. Low Temperature Heat Barrier (LTHB), it is a multilayer insulation laying on the HS support panel, its is highly reflective on its surface towards the main gap. Highly reflective. It is separated from the HTHB by the HS main gap, 245 mm wide.

2. Star-brackets to mount the FS and HTHB on the support panel

3. A support panel used to mount the brackets, the insulating layers and the Feed Through (FT) of the scientific instruments, and to interface with the SC structure. Its thickness is 55 mm, its side towards space is covered by the LTHB, while its side towards SC is used as radiator, with high emissivity, to radiate heat into the secondary gap, 100 mm wide.

4. Blades used to mount the support panel on the SC

5. An MLI mounted on the SC +X panel, it is highly reflective on its surface towards the secondary gap. Such MLI may have an extension to protect the HS from HGA.

6. Thermal stand-offs between star brackets, FT and support panel. 7. Temperature sensors and supports for cables of FT mechanisms


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Figure 8.3-1: Heatshield components

8.3.1 Design of Main Components

The Heatshield consists of a sub-system designed to protect the Solar Orbiter spacecraft (it is attached to +X SC wall), providing shadow to all equipments or parts of SC structures within a half-cone of 8 deg wrt a parallel solar flux in -X direction. The designed architecture presents two main components:

• The multi-layers Front-shield, the front face receiving direct solar light, constituted by several metallic foils;

• The Support-panel, a sandwich construction supporting all Heatshield parts: the Front-shield (held by 10 Titanium Star-brackets), instrument Feedthroughs, mechanisms and doors. The Support-panel is mounted on the spacecraft through Blades-like interfaces.

The following dimensions have been identified in accord to requirements:

• A 2500x2.500mm Frontshield • A 2430x2430mm Support-Panel • Total gap-height of 400mm for heat-rejection to space.

A stand-alone Star-brackets layout, directly mounted on the Support-panel, is the skeleton of the HS, characterised by high degree of modularity: most of the parts are common and replaceable, and the brackets can be translated to facilitate the instrument accommodation. The lateral sides are open, offering to the hot layers a sufficient view-factor-to-space for heat rejection. The Heatshield architecture, here after reported, has been considered and assessed based on the assumption to use the Feedthroughs as Frontshield barrier fixation and support points.


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Figure 8.3-2: Overall HS archtecture, showing support bipod/monopods, corner cut-outs for SWA

heads, Sun sensor cutout, feedthroughs for RS-instruments and star Bracket supports

Global isostatic behaviour, guaranteeing low level of distortion without compromising the heat-protection functionality, has been obtained by fixing loosely the Frontshield on the Star-bracket arms. The Frontshield is larger than the Support-Panel in order to ensure the 8 deg half-cone of sun-shielding. The arms of the Star-Brackets mounted on-side of the panel (as per the following picture) cover this different extension and support the Frontshield border.

Figure 8.3-3: The front shield is oversized to provide HS nominal performance up until the worst-

case off-pointing case


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8.3.2 Star Brackets Layout

The selection of the suitable number of Star-Brackets takes into account the Feedthrough locations and the hanging FrontShield portions because the following constrains have to be satisfied:

• Aacceptable operational Frontshield deflections that could affect HS thermal performance; • Limit the Frontshield mass fraction per each Star-Bracket and Feedthrough.

Two Star-Brackets per HS edge and two others in the middle result an acceptable design solution.

Figure 8.3-4: Star-brackets layout

Figure 8.3-5: Heatshield shadow projection over the SC structure


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8.3.3 Support Panel

In addition to mechanical loads, the support panel has to limit thermal gradients and distortion, ensuring stable mounting of the Feedthroughs when subjected to worst case combinations of ground operations and test, launch and on-orbit environments. The support panel has a sandwich construction with the following materials:

• High conductivity K1100 carbon Fibers • Matrix: M18 • Aluminum Honeycomb 3/16-5056-.0007P.

The Support-Panel surface in the secondary gap (looking in the –X direction at the SC upper wall) is covered with a high emissivity coat. The designed CFRP Sandwich panel have the following attributes:

• Configuration (0°,90°,+45°,-45°,hc/2) SYM • Skin = 4 plies, 0.13 mm Carbon fibers unidirectional-ply, and a total thickness = 0.5mm (each

one) • Core 3/16-5056-.0007P and a total thickness of 44mm.

Figure 8.3-6: Heatshield support panel sandwich construction

8.4 Star-Brackets

Ten Star-Brackets and the Feedthrough groups (by suitable flanges) have to support a Frontshield mass of 12.5 Kg, and the Frontshield mass fraction for each bracket has been assessed to be 1 Kg. The Star-Bracket design has been optimised to achieve a lightweight structure whilst maintaining the necessary stiffness to avoiding low-frequency coupling with first main structural modes of the Heatshield assembly.


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Figure 8.4-1: Star bracket accommodation within HS structure

These brackets are made of Titanium alloy (Ti6Al4V) and its main components are:

• Horizontal ‘Star’ arms • Bracings • Central tube structure.

Riveting eight arms (drawing a ‘star’) on the Vertical-Tube flange, the resultant assembly envelope a surface with a diameter of ≅ 450mm (see the following picture). The arms are C-shaped profile (width ≅ 20mm/length ≅ 185mm) to increase stiffness with respect to out-of-plane/torsion displacement and to be sufficiently rigid in case it would be avoided use of Bracings (using a lighter Frontshield). Wall thickness is ≅ 1.5mm.

Figure 8.4-2: Star Bracket design and principal dimensions

The single Bracing is a tube with flattened ends riveted to the Vertical-tube, at one end, and to the related Horizontal-Arm at the other. This mounting-system jointly to the thin thickness allows flexibility in order to accommodate thermal dilatation of the Horizontal-Arm without excessive overload. The Bracing dimensions are: Axial length ≅ 200mm

O-O-P displ.

I-P displ.


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external diameter = 6mm wall-thickness = 1mm The Vertical-Tube height of 260÷265mm is set to reach the foreseen aperture to space (hot-gap). The cross-section dimensions assure the necessary torsion and lateral stiffness of the overall Star assembly:

• external diameter = 45mm • wall-thickness = 3mm

(Each arm includes two small holes on tip to tie or sew (loosely) the Frontshield as per Frontshield mounting procedure in a next paragraph)

8.5 Heatshield Interfaces with SC

Interface attachments for the Heatshield can lie on any of the areas identified in the following figure.

Figure 8.5-1: Heatshield Mounting I/F Location

The Heatshield is quasi-isostatically fixed on the SC structure. The Heatshield design minimizes interface loads at interfaces due to thermo-elastic and hygrometric (using a CFRP Support-Panel) behavior and, thus, accommodates the effects due to CTE/CME differences with respect to SC structure. The selected baseline HS interfaces are a combination of Blades-like mounts, using the locations on side of the SC upper wall. The two other central locations are reserved for two special calibrated bolts assuring both axial fixing and lateral flexibility.


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Figure 8.5-2: Underside of HS support panel showing blade mounts

The main (corner) Blades mount is a combination of two thin-walled tubes having the longitudinal axes that which intersect each other: The height is 100mm, in order to put in specific place the Support-panel, and the cross-section dimensions are the following:

• External diameter = 45mm • Wall-thickness = 1.5mm

Beside, corner-Blades orientation (local Z axis direction) point toward the geometric centre of the Support-Panel, considered (approximately) as centre of differential expansion. The function of the remaining lateral-Blades is use a simpler “stand offs” not constraining (at least minimizing) the displacement in either global Y or Z direction. The gross dimensions are identical to the corner blades.

Figure 8.5-3: Blade mounts

The major difficulties of the blade design are the reaching of strength (buckling problems) and stiffness requirements without detriment to the isostaticity and the need of assembly precision to avoid pre-loading


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the blades. These issues have been addressed through selecting a high number of mounts and, at the same time, guaranteeing a minimum out-of-plane hard-mounted frequency of HS Support-Panel. The central bolts are laterally flexible items (length ≅ 100mm) used to break longitudinal movement of the Support-Panel, and they have to be mounted from the top by means of suitable cut-outs of the Frontshield. (see Interface Control Document for details).

Figure 8.5-4: Conceptual depiction of the central bolt

8.6 Front shield

The front shield of the HS is a single layer of black Carbon Fabric (by Angeloni - I). However, 2 alternative materials are also candidates for the front shield:

• Alternative 1: KEPLACOAT ® Black (by AHC - D) applied on flexible 50 μm Titanium foils • Alternative 2: “Hybrid” Fabric (by Angeloni - I), Nextel fabric with embedded Titanium wires (Ti CP

ASTM GR1, diameter 0.10 mm).

8.7 High Temperature Heat Barrier

The HTHB is an MLI-pack with materials able to withstand high temperatures, up to 700 C. Reflective layers are separated by a net. Reflective layers:

• Baseline: Nickel foils (20 μm thickness) • Alternative: Titanium foils (18-25 μm thickness).

Separators:

• Baseline: Tissueglass • Alternative 1: Quartzel ® Net (by Angeloni - I) or equivalent ceramic fibre lightweight net • Alternative 2: Fiberglass Spacer (TBC).

Front shield and HTHB are assembled together before installation on the brackets.

Special Insert inside the Support-Panel

to the SC wall


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The final selection of the FS and HTHB materials will be made after the completion of the materials TDA described in [IR41]. The FS layer and the first reflective layer are larger than the remaining reflective layers, to be folded behind them at the edges. Layers are sewn together at edges and in a limited number of points. Layers are not sewn together at holes used for installation of star brackets, sewing will take place during installation of patches. The combination of Front Shield + HTHB is provided with holes in positions corresponding to the interface with the star-bracket. Two additional larger holes are foreseen in the central area to allow the final mounting of the fully integrated HS on the spacecraft main body. Each of the holes require a patch on top of it to maintain the uniformity of the HS properties. Patches are made as the main sheet. All other attachment points of the HS on the SC main body are close to the edges and accessible without need of removal of thermal hardware.

Figure 8.7-1: Design and installation of patches

8.8 Low Temperature Heat Barrier

It is an MLI made with 20 layers double aluminized Kapton, with separator. Face to gap is aluminized to provide high reflectivity in infrared.

8.9 MLI on Spacecraft

MLI is mounted on SC +Xpanel, and makes a wall of the secondary gap. This MLI is as for the LTHB.

8.10 Stand Offs

Stand offs are required between Star Brackets and support panel, and also between FT and support panel (TBC). They are made of glass fibre, with a height of 5 mm.


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9. FEEDTHROUGHS, DOORS & MECHANISMS

The main function of the feedthroughs is to provide the required field of view to the instruments (and sun sensor) while maintaining the thermal integrity of the heat shield and limiting the flux entering the spacecraft. The feedthroughs will also provide a very limited stray-light reduction function. The feedthrough dimensions are derived from the instrument pupil dimensions, UFOV angles, distance from instrument pupil to heat shield top layer, heat shield/feedthrough thermoelastic performance, integration/alignment margins plus provision for stray light suppression vanes. Please refer to IR25] for a detailed description of the feedthrough door and mechanism design.

9.1 Common Elements of Feedthrough Design

9.1.1 FT Body Design

9.1.1.1 Instrument FT

The instrument feedthroughs are (with the exception of the SPICE feedthrough) cylindrical body shaped. The thickness of the cylinder is 0.8 mm (TBC) local reinforcement are defined at the interfaces with the HS support panel and the HS top layer. Local thickening of the FT tube is also necessary at the vanes (when present) for manufacturing reasons.

9.1.1.2 Corner FT

The detail design of the corner FT is TBD. However the concept is illustrated on Figure 9.1-1 and Figure 9.1-5.

Figure 9.1-1: Corner FT body concept design

The corner FT needs to provide interface with the support panel and with the HS top layer. The corner FT need to locally support the HS top layer MLI. The corner FT are made of two plates of titanium riveted at the corner interface between them.

Line of Ti rivets attach the 2 plates constituting the corner FT


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9.1.2 Interface with the HS Top Layer


All of the feedthroughs have a mechanical interface with the HS top layer, the function of this interface is to locally support the HS MLI layer. The design of this interface is then providing the required support without constraining the MLI. The design of this interface is common to all FT (except PHI) only the diameter of the FT is different. The design concept is to hold the HS top layer between two flanges attached to the top part of the feedthrough. These flanges are attached to the FT body by way of 6 M4 screws. The total mass of MLI supported at each FT is low (~0.5 kg) and therefore the interface forces at each screws is low (TBD). The bottom / support flange is made in two parts to allow mounting on the FT body. Each of these parts is maintained in place by 1 screw. When the top / closure flange is put in place, 2 additional screws are attaching each of the bottom / support flange to the FT body.

Figure 9.1-2: View of the FT interface design at top of the EUI FT

The width of the flange is 25 mm. This insures that the MLI is loosely supported without a risk of the MLI ‘slipping’ outside of the interface.

9.1.2.2 Corner FT

The detailed design of the corner FT interface with the top layer is TBD. Two concepts for the interface with the HS top layer are considered. The corner FT cannot provide loose interface for the HS top layer in the same way as the instrument FT do. In the case of the corner FT, the HS top panel MLI is not going ‘around’ the FT but needs to be supported on an edge only. Relocation / extension of starbracket arms is not able to provide correct closure at corner feedthroughs. Therefore, to provide adequate mechanical support and solar flux tightness, the HS top layer MLI needs to be physically attached on corner feedthrough support flange.

Top / closure flange

Bottom / support flange


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Figure 9.1-3: Corner FT concept interface with HS top layer

On Figure 9.1-3 two concepts are illustrated. The current baseline selection is concept B.

Concept A Concept B

HS top layer MLI

Stud Screw Bottom / support

flange

Top / closure flange


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9.1.3 Interface with the HS Support Panel


The FT are attached to the HS support panel by way of 3 or 4 M5 screws (TBC/TBD). The HS support panel is the only mechanical (structural) interface of the FT. All the loads (static, dynamic, thermoelastic, … ) applied to the FT are carried by this interface. The FT at this interface are characterised by a reinforced ring of increased thickness the FT nominal thickness is 0.8 mm. Locally the interface is increased to 1.6 mm.

Figure 9.1-4: View of the FT interface design at HS support panel of the EUI FT

The FT are mounted on the HS support panel from the +X side (upper skin of the support panel) the mounting interface and the screws are then located on the +X side of the support panel as well.

9.1.3.2 Corner FT

The detail design of the corner FT interface with the HS support panel is TBD. However, the corner FT are not a closed shape (i.e. tube) like the instrument FT. The stiffness is then not provided by self supporting structure. Specific support needs to be integrated in the design. The corner FT are then support from behind by brackets, see Figure 9.1-5. The brackets are machined titanium parts. They provide a flange to interface with the corner FT body and a flange to interface with the HS support panel. These brackets are attached (riveted or screwed TBD) to the corner FT body. The brackets and the corner FT body are then mounted on the HS support panel via (TBD) M6 (TBC) screws.

Interface of FT with HS support panel

Increased ring thickness at IF with HS


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Figure 9.1-5: Corner FT concept design (II)

9.1.4 Materials

The feedthrough material is titanium, this material is able to sustain the very hot thermal environment of the Solar Orbiter mission and while providing at the same time a good level of mechanical strength. Titanium is characterised by a relatively low coefficient of thermal expansion (8.6 106 m/mC) which will then reduce any issue related to differential thermal expansion at HS support panel interface (carbon skins) and will also reduce issue of dimensional stability for the PHI feedthrough carrying filters with tight positional requirements. Since titanium is also characterised by a low thermal conductivity (21.9 W/mK) the heat conduction to the HS carbon skins is low, which reduce the temperature at the interface of the FT with the potted inserts, which is mechanical and thermally a favourable situation.

Interface flange with HS support panel

Interface flange with corner FT body

HS support panel


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9.2 Feedthrough Design Concepts Conceptual CAD images of the various feedthrough design are presented in this section. It should be noted that there is a physical conflict between the mounting of the two PHI filters.

Figure 9.2-1: Remote-sensing instruments heat shield feedthroughs, doors and mechanisms. Note STIX does not have a door and the second PHI door is work underway; note also the interference

between the PHI feedthroughs

PHI

STIX

EUI

METIS

SPICE

SWA

SWA


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Figure 9.2-2: PHI feedthrough showing filter mounting concept

Support panel interface

Heat shield top layer interface

PHI filter

Closure flange Door

Door shaft

Door mechanism

Feedthrough


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Figure 9.2-3: Preliminary SPICE feedthrough design


Feedthrough



Feedthrough



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Figure 9.2-4: EUI feedthroughs, doors and mechanism design HRI feedthrough

FSI feedthrough

Doors

Door mechanisms

Doorshaft



Closure flange

HRI feedthrough

FSI feedthrough

Doors

Door mechanisms

Doorshaft



Closure flange


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Figure 9.2-5: METIS feedthrough, door and mechanism design

Figure 9.2-6: STIX feedthrough baseline design

Aspect measurement system apertures

Support panel interface Heat rejecting MLI filters


Feedthrough

Closure

Door

Feedthrough

Door shaft

Door mechanism



Door

Feedthrough

Door shaft

Door mechanism




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Figure 9.2-7: SWA-HIS feedthrough conceptual design

9.3 Doors & Mechanisms

The proposed design solution is shown graphically in the following figure – for more information regarding the door mechanism design, please refer to [IR25].


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Figure 9.3-1: Door mechanism design


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10. THERMAL CONTROL SUBSYSTEM

For a detailed description of the Solar Orbiter TCS, please refer to [IR53]. Note that the TCS design presented does not include the heatshield, which is considered a separate item to the TCS.

10.1 Requirements


SC survival Withstand all thermal environments encountered during the entire life of the mission.

A wide range of sun distances is experienced during the course of the mission, along with several configuration changes that complicate the thermal design.

Unit temperatures Provide the temperature and heat flow environments required of both the payloads and platform equipments at the agreed interfaces.

The high heat fluxes associated with the RS-instruments necessitate a well-specified and robust approach to their thermal management

Verification Be compatible with on ground verification of the performance, and allow demonstration by analysis and test

Implications for heat-pipe locations and orientations, vis-à-vis the allowable test orientations

Table 10.1-1: TCS principal requirements

10.2 Design Overview

The mission goal of Solar Orbiter has many implications for the system and satellite design and the overall operational concept due to the high solar flux experienced at perihelion which is the driver of the entire design. The Solar Orbiter thermal control is based on using a sun pointed, flat heat shield to limit the sun flux on the spacecraft structure. By using this approach the elements behind the heat shield will be in a more benign thermal environment.

All external components shall be shielded from direct solar illumination by the heat shield except for the instruments requiring direct view of the sun and the spacecraft appendages (i.e. the solar arrays, the RPW antennas and the HGA). The heat shield is sized to prevent direct solar illumination on any of the shaded components during nominal pointing and for safe mode events of spacecraft off-pointing up to 6.5 degrees from sun-centre. However, the SC must also withstand reflected solar flux and high IR flux from appendages outside of the heatshield shadow cone. In addition, the remote sensing instruments will all receive additional IR flux from the feedthroughs which allow them to view through the heatshield.

The Solar Orbiter thermal control subsystem (TCS) is responsible for maintaining the equipments of the spacecraft (platform and payload) within their specified temperature limits and ensuring the provision of sufficient heat rejection of the sun exposed payloads. To achieve this, in addition to the function of the heat shield the thermal control subsystem of the platform employs passive thermal control techniques such as black painted interior to maximise internal radiative heat transfer, OSR mirror tile radiator areas of external surfaces of panels to reject heat dissipated by units, heat pipes to link units to radiating surfaces and multi-layer insulation (MLI) to minimise heat losses from other external surfaces, as well as the use of heaters.

The radiators reject heat directly to space, and they are sized so as to maintain the upper temperature of each unit towards the top of their allowable range. The external radiator surfaces OSR mirror tiles to enable efficient rejection of the heat generated by the units. OSR mirror tiles have been selected for the high infrared emissivity and low solar absorptivity of this surface finish. While the heat shield shadows the external radiator surfaces from the sun, OSR mirror tiles have been selected at this phase to minimise the impact of solar flux reflected from the sun exposed external appendages (instruments and antennas). The


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available radiator area on the spacecraft is greater than that required for the spacecraft power, and the temperature can thus be trimmed up or down before launch by varying the exposed surface.

To further improve this heat rejection the majority of the electronics units are mounted to the panels with thermal interface filler to improve conductance between the equipment and the structure. The dissipated heat is rejected via conduction into the panel and subsequent radiation from black painted the high emissivity panels internal surfaces to the external surface of the panel and then to space. However internal radiation to the external panels does not provide sufficient heat rejection capability. To improve the coupling from payloads to radiators heat pipes are employed

Solar Orbiter also travels a great distance from the Sun during transfer. The minimum necessary heater power is applied in the cold cases so that the lower temperature of each unit is maintained towards the bottom of their allowable range. By using the full design temperature range of each unit in this way, the heater power requirement is minimised. However the large radiator area required for near sun operation can then lead to a requirement for large amounts of heater power during transfer to maintain units above their minimum allowable temperature.

The proposed solution to the high heater power requirement is to point the heat shield away from the sun during transfer, thus exposing radiator surfaces to direct solar flux. This would allow the SC to absorb heat, and reduce the heater power requirement. There will exist some minimum sun distance at which the SC can remain anti-sun pointed without exceeding the upper temperature limit of some units.


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Figure 10.2-1: Overview of Solo Orbiter Thermal Design Configuration

Internal Panels: - Black painted to maximise internal radiative heat transfer

Payload Feed Throughs: - provide an unobstructed field of view through the heatshield - internal surface coated using black keplacoat for stray light - managed by rejection of heat radiatively into the heatshield

Platform and Payload Units: Painted black, Mounted directly to panels using thermal interface filler to maximse conductive heat transfer

Battery: - Wrapped with MLI and mounted on feet to ensure internal temperature gradients are minimised

Thrusters: - Individually heater controlled - High temperature MLI surrounding nozzles

Heaters: - Kapton encapsulated heater resistance elements located on items requiring replacement or additional heater during the mission

HGA: - Effect of reflection and emission of heat from antenna to spacecraft minimised by restricting radiator location from –Z face - HGA shadowed by spacecraft at sun distances less than 0.28AU

Heat Shield: - shadows SC body from the solar flux - multiple high temperature and standard MLI layers, gaps, and support structure

Solar Array: - Two sided to manage variation in solar environment - OSR Mirror tiles interspersed with cells to reject excess heat - Rotated away from sun to reduce incident flux

Instrument Boom: - Specified to survive thermal environment without system thermal control

Remote Sensing Payloads: Mounted on shear walls to minises thermo-elastic effect on pointing -coupled to radiators using heat pipes

External MLI: -covering all surface areas not required for radiators - Baselined NEXTEL outer layer on ±Y, ±Z sides, ITO coated Kapton outer layer –X side

External Radiators: - OSR mirror tiles

Thermistors: - Located at TRP of all units for monitoring - Triple Majority Voting (TMV) used for all actively controlled heaters (for redundancy)


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10.3 Reflection Control

The extreme solar flux levels at the near sun distances leads to a particular need to control the reflection from external appendage surfaces illuminated by the sun. Requirements are placed on subsystem configuration and thermal design to ensure that reflections to the platform are minimised.

One example of this need is the edge of the solar array. The angle of the is such that if designed with a flat side, at the near sun distances where the array is angled to the sun, the reflection from the yoke will illuminate the side of the spacecraft including the equipments radiators with significant levels of solar flux.

Figure 10.3-1 : Reflection of solar flux from solar array edge to platform side

Verification of the design features will be carried out during both subsystem and system level analysis and test.

10.4 Thermal Hardware Design Description

The thermal control subsystem consists primarily of common passive means. The following standard techniques are used to the maximum extent to realise the thermal control design: MLI blankets, radiator surfaces, paints, thermal fillers, doublers and electrical heaters, controlled by thermistors. These techniques are well known and therefore increase the system reliability. This allows the thermal design to concentrate on the complex payload thermal requirements.

The TCS hardware is required to fulfil the following functions:

• Control the thermal environment so that temperatures of equipment remain within acceptable ranges and thermal stability requirements are met

• Control the environmental temperatures of externally mounted equipment so that they remain within acceptable temperature ranges

• Control radiative/conductive heat flow both externally and internally

• Control conductive heat flow at unit and sub assembly interfaces

Solar Flux Direction

Solar Array


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• Provide conductive interfaces ensuring acceptably high/low heat fluxes in local areas.

10.4.1 Overall Configuration

The overall configuration of the arrangement of the thermal hardware is driven by the location of the payload units on internal shear walls and the need for verification of the performance of the TCS under gravity in thermal test.

Due to the significant amount of heat that needs to be transported from the internally mounted payloads to the externally mounted radiators, efficient rejection within the structure configuration adopted for Solar Orbiter requires the use of heat pipes. Payloads mounted on shear walls leads to 3D configuration of heat pipes as the surfaces for radiators (±Y external panels) are orientated perpendicular to the shear walls.

The 3D routing of the bent heat pipes creates a complication with respect to the on ground verification of the TCS performance. Heat pipes under gravity are required to function in reflux mode, where the hot end (unit) is at the lower point with respect to gravity. Therefore routing of heat pipes and locations of radiators has been configured to ensure that heat rejection for the payloads is possible in ground test.

The current baseline SC thermal test configuration is with the heat shield facing a solar simulated beam from the horizontal direction (along SC X axis), and the +Z axis pointing up with respect to gravity.

A further restriction to the location of radiating area is the location of the HGA. When deployed at 0.28AU the HGA surface temperatures approach 400°C, radiating a large amount of IR flux to the platform. Figure 10.4-1 shows the distribution of temperature for the outer layer of MLI in view of the deployed HGA at 0.28AU.

Figure 10.4-1 : MLI Temperature in View of HGA

As can be seen in Figure 10.4-1 the temperature of MLI (which has a similar emissivity as the radiators) is in the order of 50°C when exposed to the HGA flux. To reject heat from the platform equipments radiators are required to operate colder than this however as the HGA heat is emitted in IR and the radiators have a high emissivity to reject heat, locating platform radiators in view of the HGA is not viable. Therefore the solar Orbiter thermal configuration has excluded radiators on the –Z side of the platform.


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10.4.2 Multi Layer Insulation (MLI) Blankets

Thermal blankets are used externally to minimise solar heating and radiative heat transfer to space. MLI blankets are installed on all external panels. On the external panels around the radiators, the blankets are used for radiator trimming for the maximum unit temperatures to be coherent with the maximum design temperatures provided by the system. Two types of thermal blankets are currently defined to minimise solar heating and radiative heat loss to space and to isolate particular units within the satellite:

• External MLI blankets: they must have good handling properties to withstand the satellite integration activity without performance degradation. On the ±Y and ±Z sides of the spacecraft where reflection from appendages of high intensity solar flux is possible the outer layer of the blanket is currently base lined Nextel (TBC), use proposed for Bepi-Colombo. On the –X side of the spacecraft where high intensity solar flux illumination is less likely to occur the outer layer is Kapton with a VDA backing to reduce as much as possible the overall weight of the thermal subsystem. Detailed design is to be performed to further define where the high solar flux tolerance materials are required to ensure that the mass is minimised.

• Internal MLI blankets: With low emissivity outer layer (VDA) this is used to cover the battery as required by the supplier and is designed to withstand the temperature range of –40°C to +60°C without degradation of its thermal and mechanical properties.

Nextel Kapton with VDA backing

Figure 10.4-2 : External MLI Blanket Types

All blankets shall be electrically grounded to the structure and shall have a conductive outer layer to ensure ESD requirements are met.

The Solar Orbiter MLI areas are detailed in Table 10.4-1. The estimated areas cover the external surfaces of the spacecraft where radiators are not required. An additional 20% is applied to the calculated radiator area to accommodate the details of individual blanket pack design (multiple packs in each area, accounting for overlap between packs, complex geometry surrounding equipments).

Panel Type MLI Area (including 20% overlap)


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+Y+Z External – Nextel 0.57 +Y External – Nextel 0.52

+Y-Z External – Nextel 0.68 -Y+Z External – Nextel 0.49

-Y External – Nextel 1.18 -Y-Z External – Nextel 0.56 +X External – Nextel 3.39 -X External – Kapton 2.13

LVA ring External – Nextel 2.20 +Z External – Nextel 2.26 -Z External – Nextel 2.93

Prop Tanks Internal 4.00 RCS Lines Internal 1.50

Battery Internal 0.20 External – Nextel 14.77 External – Kapton 2.13 Total

Internal 5.70

Table 10.4-1 : MLI areas

10.4.3 Radiators

The orientation of the spacecraft when on station and throughout the majority of the transfer phases is such that heat shield is sun pointing and radiator surfaces are not exposed to solar fluxes. During normal operation outside of 0.28 AU, the HGA antenna sits below the SC –Z face, in direct solar illumination. Solar flux will be reflected from the antenna onto the –Z face, and high IR heating from the hot dish will also be seen in this area. For these reasons, the –Z face is not suitable for radiator surfaces. The radiator areas required in order to achieve the acceptable upper temperatures for each of the equipments are listed in Table 10.4-2.

Panel Unit Radiator Area (m2) Percentage of Usable Panel Area (%)

EUI 0.1089 SPICE 0.0968 +Y +Z EPC 0.0484

43.0 (excluding area under

SoloHI)

+Y TWTA 0.5080 91.3 (not including access panel)

RIU 0.0726 +Y –Z OBC 0.0907

22.5

SSMM 0.1512 PHI ELEC 0.0907 -Y +Z

PHI OPTICS 0.0726

52.0 (excluding estimated area

under EPD payload) Transponder 1 0.0544

-Y Transponder 2 0.0544

11.5 of total, however radiator integral to access

panel PCDU 0.1270 -Y –Z Battery 0.0121

19.2

EUI & SPICE ELEC 0.0726 STIX 0.0363 IMU 0.1331

-X

FCE 0.0544

26.7


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Panel Unit Radiator Area (m2) Percentage of Usable Panel Area (%)

-X within LVA METIS 0.3496 89.2

(not including area for tank support beams)

SPICE Rad Cold Finger 0.0968 SOLOHI Elec 0.0544

EUI Rad Cold Finger 0.0726 METIS Rad Cold Finger 0.0968

PHI Rad Cold Finger 0.0242 +Y Environment Rads 0.0968

+Z

-Y Environment Rads 0.0968

22.3

Total 2.6722 29.2

Table 10.4-2 : Radiator areas

While the total percentage of the external surfaces used as radiator area is less than a third, there are some significant restrictions in available area. These must be considered in the detailed design of Solar Orbiter:

- Due to the need for a removable panel to access the CPS equipment located on the mid shear wall, the area of the +Y panel above (in the +X direction) the TWTs is not available for the TWT radiator. As it is not possible to route embedded heat pipes to spread the unit dissipation over the access panel surface this area can not be used, thus the radiators percentage of available area is very high, leaving very little margin for growth in the heat rejection from these units

- The transponders located on the –Y panel have been mounted on the CPS access panel so that access can be provided while still accommodating these units in this bay. As a result while the radiator percentage of available area is low if the entire –Y panel is considered, it is not possible to spread the heat from the units over a greater area than the access panel as heat pipes can not pass between access panel and spacecraft structure. At the time of writing the area of access panel was undefined, however it is expected that the transponder radiators will occupy the majority of the access panels external surface

- The –X location of the METIS and STIX radiators is driven by the restriction on heat pipe routing due to the thermal test orientation. The only location for the METIS radiator found to provide a suitable area is within the LVA ring, however occupies almost the entire available area, leaving only a small margin in growth for the heat rejection from this payload

- The +Z radiators used to cool the +Y+Z and –Y+Z bays of the spacecraft are located either side of the spacecraft centre line to avoid an interaction with the MGA. The units located on the middle shear wall require these radiators to reject their dissipation, however due to the location of the MGA this is only possible to do so via the radiation of heat within the spacecraft from the units to the radiator. This is not an efficient method of heat rejection as the units have a reduced view factor to the radiator area and required a large temperature difference to exchange the heat. A more efficient solution will be to relocate the MGA such that the radiators are near the spacecraft centre line and MGA is off to one side. This modification to the configuration is to be investigated in detail in future work.

The location of the radiators is shown in the following figure.


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+Y Wall -Y Wall

-X and +Z Walls

Figure 10.4-3: Radiating areas


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10.4.4 Doublers

The significant dissipation of the TWTs require a large radiator area. Where large radiators are needed heat is transported from the unit to the radiating surface using embedded heat pipes. However due to the high intensity of the heat dissipated over the small interface area of the TWTs, doublers are required to spread this heat across the external panel to provide a good coupling to the embedded heat pipes for the efficient rejection of heat via the external radiators. The location of the doublers is shown in the following figure.

Figure 10.4-4: TWT doublers

10.4.5 Heat Pipes

Heat pipes are used to transport heat from an equipment to its radiator. They are used when the equipment cannot be located directly onto the radiator area such as the remote sensing payloads. Another application is to improve the radiator efficiency by spreading the heat across a larger area when the power concentration is too high. The heat pipes are of two types: surface heat pipes and embedded heat pipes. They are both extruded aluminium alloy pipes with multi re-entrant grooves. High purity ammonia is used as working fluid. The surface heat pipes have a large flange to allow the heat pipes to be bolted to the structure or equipment. This flange increases the contact area between the equipment and the heat pipe and therefore improves the heat transfer. The embedded heat pipes are located inside the honeycomb panels. This improves the conductance between the heat pipe and the radiator surface. On the platform, bent surface heat pipes shall be attached to payload units on the shear walls so that they can be coupled to the external radiator panels. The embedded heat pipes are located with the central section of the ±Y walls to spread heat from the TT&C equipment. To further improve the conductance between the units and the heat pipes, and the heat pipes and the panels, thermal interface filler shall be used where the heat pipes are bolted. The Solar Orbiter heat pipe arrangement uses the bent heat pipes to transfer heat from the equipment and payload hot interfaces to the external surfaces, and the embedded heat pipes to spread heat over the radiator area where needed (usual where the required radiator area is much larger that the area of the bent heat pipe interface. For redundancy anywhere a heat pipe is required a minimum of two are allocated, even when the heat transport capacity of the single pipe is sufficient to remove the heat. The following figure shows the location where heat pipes will be required.

Doubler


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Figure 10.4-5 : Heat pipe routing

Due to the need to locate the payload units on the shear walls to minimise thermo-elastic distortion and maintain the payload co-alignment in science modes, Solar Orbiter requires a large number of heat pipes in a complex 3D arrangement. The heat pipe configuration required brings with it a significant mass penalty (approx 14kg) and complications with respect to verification via test.

10.4.6 Cold Finger Radiators

Separate radiators are required to control the cold fingers of the PHI, SPICE, EUI and METIS. The first is controlled to -10ºC, the others to -60ºC.

Payload Power (W) Target temp (ºC) Area (m2) PHI 5 -10 0.0242

SPICE 5 -60 0.0968 EUI 5 -60 0.0726

METIS 2 -60 0.0988

Table 10.4-3 : Cold finger radiator areas

Currently the areas on the platform foreseen for the location of the cold finger radiators is the +Z side of the platform, away from the solar array and HGA, but in view of one RWP antenna. Details of the accommodation of these radiators are TBC, though sufficient area exists on this panel to accommodate the radiator area. The location of these shall consider the need for this panel to allow for access to the internal bays. Due to the required low operating temperature of the radiators, it is proposed that the radiators are a separate structure thermally decoupled from the spacecraft structure via low conductivity washers.


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Figure 10.4-6: Radiating areas

The cold finger radiators are to be coupled to the payload cold fingers using loop heat pipes. Loop heat pipes have been selected for the flexibility of routing the pipes to the radiator areas (not the same gravity concerns as with capillary heat pipes). The LHP layout for Solar Orbiter is shown below.

Figure 10.4-7 : Cold finger LHP routing


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10.4.7 Heaters

The heater powers are sized by calculating the steady state power required to maintain all units above their minimum temperature limit, plus modelling uncertainty. For the few heater circuits that are actively controlled the actual heater switch-on temperatures are set to 5°C above the lower allowable temperature limit, and the over sizing ensures that the heater is powerful enough to cycle in the coldest flight conditions. e heaters are mounted directly onto the units wherever possible to maximise the efficiency of the heating. These heaters will be bonded onto paint-free areas on the surface of the unit.

The summary of the Solar Orbiter heater circuits and the power demand during the spacecraft modes they operate in is shown below.

Sub

syst

em

Equipment

Inst

alle

d (W

) (20

%

grea

ter t

han

dem

and)

D1

- Nea

r Sun

Sci

ence

(W

)

D2

- 0.2

8 Sc

ienc

e (W

)

D4a

- Fa

r Sun

, Ant

i Su

n Po

intin

g (W

)

D4c

- Fa

r Sun

, Sun

Po

intin

g (W

) D

4d -

Far S

un, A

nti

Sun

Poin

ting,

Saf

e M

ode

(W)

D5

- Lim

it of

Sun

Po

intin

g (W

)

D6

- Far

Sun

Sci

ence

(W

)

D9

- Saf

e M

ode

(doo

rs

clos

ed) (

W)

PHI Optics Casing 9.6 0.0 0.0 0.0 3.0 3.0 1.0 0.0 8.0 PHI Hot Element 9.6 0.0 0.0 3.0 7.5 8.0 7.0 6.5 8.0 PHI Cold Element 8.4 0.0 0.0 7.0 7.0 6.5 6.5 3.0 6.5 PHI Electronics 16.8 0.0 0.0 0.0 5.0 7.0 3.0 0.0 14.0 SPICE Optics Casing 12.0 0.0 0.0 0.0 10.0 0.0 6.0 9.0 9.0 SPICE Hot Element 28.8 0.0 0.0 9.0 24.0 21.0 18.0 19.0 18.5 SPICE Cold Element 13.2 0.0 0.0 11.0 11.0 11.0 11.0 6.0 10.0 SPICE Electronics 25.2 0.0 0.0 0.0 15.0 0.0 9.0 7.5 21.0 EUI Hot Element 4.8 0.0 0.0 0.0 0.0 0.0 4.0 0.0 0.0 EUI Cold Element 12.0 0.0 0.0 9.0 10.0 9.0 10.0 5.0 8.5 EUI Electronics 4.8 0.0 0.0 0.0 4.0 0.0 0.5 0.0 0.0 METIS Hot Element 33.6 0.0 0.0 0.0 9.0 0.0 18.0 4.5 28.0 METIS Cold Element 14.4 0.0 0.0 11.0 12.0 11.0 12.0 6.0 11.0 METIS Electronics 16.8 0.0 0.0 0.0 14.0 0.0 8.0 0.0 14.0 STIX Optics Casing 1.2 0.0 0.0 0.0 0.0 0.0 0.0 0.0 1.0 STIX Electronics 12.0 0.0 0.0 0.0 10.0 0.0 6.0 0.0 9.0 SOLOHI Electronics 10.8 0.0 0.0 6.0 9.0 6.0 5.0 0.0 6.0 SWA Electronics 2.4 0.0 0.0 0.0 0.0 1.0 0.0 0.0 2.0 RPW Electronics 8.4 0.0 0.0 2.0 7.0 6.0 0.0 0.0 6.5

Payl

oad

Inte

rfac

es

MAG Electronics 3.6 0.0 0.0 0.0 3.0 2.0 0.0 0.0 2.5 SPICE 4.2 0.0 0.0 3.5 3.5 3.5 2.6 2.6 2.6 PHI HRT 4.8 0.0 0.0 4.0 4.0 4.0 3.3 3.3 3.3 PHI FDT 4.6 0.0 0.0 3.8 3.8 3.8 3.0 3.0 3.0 EUI 1 5.0 0.0 0.0 4.2 4.2 4.2 0.8 0.8 0.8 EUI 2 4.8 0.0 0.0 4.0 4.0 4.0 0.8 0.8 0.8 METIS 5.5 0.0 0.0 4.6 4.6 4.6 1.0 1.0 1.0 PHI HRT Filter 15.7 0.0 0.0 13.1 0.0 13.1 0.0 0.0 0.0

Feed

Thr

ough

M

echa

nism

s &

Filt

ers

PHI FDT Filter 4.3 0.0 0.0 3.6 0.0 3.6 0.0 0.0 0.0


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Sub

syst

em

Equipment In

stal

led

(W) (

20%

gr

eate

r tha

n de

man

d)

D1

- Nea

r Sun

Sci

ence

(W

)

D2

- 0.2

8 Sc

ienc

e (W

)

D4a

- Fa

r Sun

, Ant

i Su

n Po

intin

g (W

)

D4c

- Fa

r Sun

, Sun

Po

intin

g (W

) D

4d -

Far S

un, A

nti

Sun

Poin

ting,

Saf

e M

ode

(W)

D5

- Lim

it of

Sun

Po

intin

g (W

)

D6

- Far

Sun

Sci

ence

(W

)

D9

- Saf

e M

ode

(doo

rs

clos

ed) (

W)

X band TWTA 1 18.6 0.0 0.0 0.0 0.0 0.0 0.0 15.5 0.0 X band TWTA 2 18.6 0.0 0.0 0.0 4.0 0.0 3.0 15.5 3.0 Ka band TWTA 1 18.6 0.0 0.0 0.0 4.0 0.0 3.0 15.5 3.0 Ka band TWTA 2 18.6 0.0 0.0 0.0 4.0 0.0 3.0 15.5 3.0 X band EPC 2 3.6 0.0 0.0 0.0 3.0 2.0 2.0 0.0 2.5 Ka band EPC 1 3.6 0.0 0.0 0.0 3.0 0.0 2.0 0.0 2.5

Com

ms

Sub

Syst

em

Ka band EPC 2 3.6 0.0 0.0 0.0 3.0 0.0 1.5 0.0 2.5 Star Tracker 1* 6.0 5.0 5.0 0.0 5.0 5.0 5.0 5.0 5.0 Star Tracker 2* 6.0 5.0 5.0 0.0 5.0 5.0 5.0 5.0 5.0 -X Coarse Sun Sensor 1.2 1.0 1.0 0.0 1.0 1.0 1.0 1.0 1.0 A

OC

S

RW1 2.4 0.0 0.0 0.0 0.0 0.0 0.0 0.0 2.0 SSMM 10.8 0.0 0.0 0.0 0.0 5.0 0.0 0.0 9.0 RIU 13.2 0.0 0.0 0.0 0.0 0.0 8.0 0.0 11.0 FCE 9.6 0.0 0.0 0.0 0.0 0.0 0.0 0.0 8.0

Plat

from

Eq

u.

Battery 8.4 0.0 0.0 4.0 7.0 5.0 6.0 6.0 6.5 Prop tank 1 2.6 0.0 0.0 0.5 1.8 1.5 1.5 1.5 2.2 Prop tank 2 3.0 0.0 0.0 1.2 2.4 1.5 2.2 2.0 2.5 Pressurant Tank 0.2 0.0 0.0 0.1 0.2 0.2 0.2 0.2 0.2 Thrusters* 28.8 24.0 24.0 24.0 24.0 24.0 24.0 24.0 24.0

Prop

ulsi

on

CPS Lines* 27.0 22.5 22.5 22.5 22.5 22.5 22.5 22.5 22.5 Total (W) 501.8 57.5 57.5 151.1 274.4 205.0 226.4 207.2 310.4

* Estimated based on other programs.

Table 10.4-4 : Heater distribution

The consumption of heater power varies greatly throughout the mission due to both the operation of the spacecraft and the very large variation in solar distance. One effort to reduce the heater power demand has been to orientate the spacecraft with –X sun pointing (no longer shadowed by the heat shield). The saving in heater power demand can be seen in the comparison of cases D4a and D4c, where 123W is saved in the anti sun orientation at 1.5AU.

If each heater identified above is assumed to be a separate heater circuit controlled via a switch in the PCDU, 64 are required for Solar Orbiter. This number includes circuits for each of the payload interfaces, 6 separate CPS line zones and each thruster. It is acknowledged that this a large number of circuits, however it is driven by the large number of payloads on the spacecraft. Future detailed development of Solar Orbiter will address the need for heating with the aim of reducing the demand on the power subsystem.

10.4.8 Thermistors

TBD thermistors are available for monitoring and thermal control. Each of the required active heater circuits shall be controlled by the redundant majority voting system, which requires 3 thermistors per heater circuit


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(195 thermistors in total). Thermistors are installed at appropriate locations within the spacecraft to control the nominal heaters and monitor the performance of the TCS. At least a single thermistor is located at each equipment TRP for monitoring purposes, including tanks and RCS pipework.

The thermistors have been divided into two groups, according to their function, however a thermistor may appear in all the groups depending on the application of the measurements taken:

• Temperature monitoring (TMT) thermistors – these enable the monitoring of the TRPs of each unit. • Heater circuit control (TMC) thermistors – these monitor the temperature of a key point in the

environment of each circuit and provide switching control over and above ground intervention or switching according to power/software mode.

There is 1 TMT thermistor for each unit and 3 TMC thermistors for each of the heater circuits under median selection active thermal control. If three heater control thermistors have been allocated to a unit an additional monitoring thermistor has not been included as this function is provided by the three heater control thermistors. Any units which are heated but do not use median selection for the control of the circuit have two thermistors allocated to them to ensure that heater failure detection is ensured in the event of a failed thermistor. The current thermistor budget for the SC is presented below.

Sub system Equipment Monitoring (TMT) Heater Control (TMC)

Sub-System

Total PHI Optics Casing 1 2 PHI Hot Element 1 2 PHI Cold Element 1 2 PHI Electronics 1 2 SPICE Optics Casing 1 2 SPICE Hot Element 1 2 SPICE Cold Element 1 2 SPICE Electronics 1 2 EUI Optics Casing 1 0 EUI Hot Element 1 2 EUI Cold Element 1 2 EUI Electronics 1 2 METIS Optics Casing 1 0 METIS Hot Element 1 2 METIS Cold Element 1 2 METIS Electronics 1 2 STIX Optics Casing 1 2 STIX Electronics 1 2 SOLOHI OPT 1 0 SOLOHI Electronics 1 2 SWA EAS 1 0 SWA Electronics 1 2 RPW Electronics 1 2 MAG Electronics 1 2 EPD STE 1 0 EPD HETn 1 0

Payl

oad

Inte

rfac

es

EPD Electronics 1 0 67

SPICE 1 2

Feed

Th

rou

gh

Mec

han

ism

s &

Fi

lters

PHI HRT 1 2

24


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Sub system Equipment Monitoring (TMT) Heater Control (TMC)

Sub-System

Total PHI FDT 1 2 EUI 1 1 2 EUI 2 1 2 METIS 1 2 PHI HRT Filter 1 2 PHI FDT Filter 1 2 Ka band transponder 1 0 X band transponder 1 0 X band TWTA 1 1 2 X band TWTA 2 1 2 Ka band TWTA 1 1 2 Ka band TWTA 2 1 2 X band EPC 1 1 0 X band EPC 2 1 2 Ka band EPC 1 1 2 Ka band EPC 2 1 2

Com

ms

Sub

Syst

em

HGA Assembly 5 0 29 SADE1, 2 2 0 PCDU 1 2 Battery 1 2 FCE 1 2 OBC 1 0 SSMM 1 2 Pl

atfo

rm E

qu.

RIU 1 2 18 IMU 1 0 RMU 1 0 CSS -X 1 2 CSS 1 2 Star Tracker 1 1 2 Star Tracker 2 1 2

AO

CS

RW1, 2, 3, 4 4 0 18 Prop tank 1 1 2 Prop tank 2 1 2 Pressurant Tank 1 2 CPS Thrusters 12 24 Pr

opul

sion

CPS Pipes 6 12 63 H/S Heat Shield Interface 10 0 10

Total including 20% margin 275

Table 10.4-5 : Thermistor allocation


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11. PROPULSION SUBSYSTEM

For a full description of the propulsion subsystem, including component selection, please refer to [IR48].

11.1 Requirements

The principal requirements placed on the propulsion subsystem for the Solar Orbiter mission are shown in the following table. Because of the modest total delta V requirement (325m/s in the worst case 2018 mission scenario), coupled with the absence of gravity losses, led to the selection of a chemical propulsion RCS with no main engine.


Authority One high efficiency pure force vector in one direction in the SC PRF

This requirement is to provide high efficiency delta V manoeuvres above the TCOL. The chosen direction is the –X direction of the SC PRF, in keeping with Mars Express heritage

Authority 2D force authority in one plane of the SC PRF

This is necessary in order to allow random force directions to be accessed whilst maintaining the SC +X axis parallel to the Sun-line (i.e. heatshield providing protection to the SC)

Authority 3D torque authority in 3 orthogonal axes. Small parasitic forces may be acceptable (TBD)

Full torque authority is required in order to facilitate wheel off-loading, platform pointing and slewing to delta V attitudes

Lifetime Compatible with an EMP lifetime of 9.6 years

The propulsion subsystem components must be compatible with the mission lifetime up until the end of EMP (9.6 terrestrial years)

Contamination The characteristics of the thrusters and their accommodation on the SC shall not cause deleterious contamination of sensitive elements of the SC, particularly payloads

The thruster positions and orientations must be compatible with the cleanliness requirements of the sensitive payloads and units.

Contamination analysis has demonstrated that a bipropellant system is acceptable, and the thruster orientations have been selected to avoid interaction with any SC appendages or units whilst providing the required control authority

Operations CPS operations after launcher separation to de-tumble the SC and acquire the Sun shall be possible prior to any appendage deployment

The chosen thruster positions and orientations provide omni-directional torque authority for despin (using –X panel thrusters) without conflicting with any stowed appendages

Capacity The propulsion subsystem must be capable of providing:

325 m/s (TBC) of effective delta V in the backup 2018 mission (sizing case)

Effective means that all inefficiencies due to constrained delta V manoeuvres and thruster orientations, plus propellant required for attitude control are taken into account in the overall delta V value.

Compatibility Thruster choice must be compatible with the FDIR system of the mission

Dual seat valve thrusters have been selected in order to avoid the credible possibility of an open thruster failure leading to off-pointing

Table 11.1-1: Propulsion subsystem key requirements


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TP 03

FVV 01

TP 02

HePressurant

PT3H/P

TP 04

PV 02 PV 01

MMH 2

NR V 4 NRV 2 NRV 1 NR V 3

F2

FO-02

PT2L/P

PV 07

NTO1

F1

FO-01

TP 10

FDV 08

TP 09

FDV 07

10N DUAL SEAT VALVEPRIME BRANCH

10N DUAL SEAT VALVEREDUNDANT BRANCH

PT1L/ P

ISV2A ISV2B ISV1A ISV1B

PV 08

Orifice

1A 2A 3A 4A 5A 6A 7A 8A 2B 3B 4 B 5B 6B 7B 8B1B

PV 03 PV 05PV 04 PV 06

FVV 05FVV 06

F3

TP 03

FVV 01

TP 02

HePressurant

PT3H/P

TP 04

PV 02 PV 01

MMH 2

NR V 4 NRV 2 NRV 1 NR V 3

F2

FO-02

PT2L/P

PV 07

NTO1

F1

FO-01

TP 10

FDV 08

TP 09

FDV 07

10N DUAL SEAT VALVEPRIME BRANCH

10N DUAL SEAT VALVEREDUNDANT BRANCH

PT1L/ P

ISV2A ISV2B ISV1A ISV1B

PV 08

Orifice

1A 2A2A 3A3A 4A4A 5A5A 6A6A 7A7A 8A8A 2B2B 3B3B 4 B4 B 5B5B 6B6B 7B7B 8B1B1B

PV 03 PV 05PV 04 PV 06

FVV 05FVV 06

F3

Figure 11.2-1: Solar Orbiter CPS Schematic

The Solar Orbiter Chemical Propulsion System (CPS) is a simple bipropellant helium one shot repressurised system using monomethyl hydrazine (MMH) as the fuel and mixed oxides of nitrogen with 3% nitric oxide (MON-3) as the oxidant. A common propellant storage and feed system supplies the reaction control thrusters (RCT). The design builds on the heritage gained through the Eurostar and ESA programs. The system is designed to operate in two blow-down mode phases with a single repressurization between each phase. During the first stage propellant section is isolated from the pressurant tank by NC pyrovalves PV01 and PV02. With this configuration, propellant tanks pressurised to NOP feed thrusters in blow-down mode. When pressure in the propellant section drops to the thrusters minimum operating pressure, propellant tanks are repressurised to NOP by firing the NC pyrovalves PV01 to PV06. After repressurization, the two NO pyrovalves PV07 and PV08 are fired in order to isolate the pressurant section, leaving the CPS ready for the second blow-down phase.


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11.2.1 Pressurisation Subsystem

The pressurant subsystem comprises a single helium tank, six normally closed pyrovalves, a gas filter, a flow limiter orifice/restrictor, four non-return valves, one high pressure fill and vent valve, one high pressure test port, two low test ports, two low pressure fill and vent valves, a high pressure transducer and two normally open pyrovalves. The pressurisation subsystem components and the pressurant tank are all located on the –Y central shear wall as shown in the following figure.

Figure 11.2-2: Solar Orbiter propulsion system mechanical layout

Pressurant tank

-X Panel THRUSTERS

+/- Y PANEL THRUSTERS

PROPELLANT TANKS

BENCH

PRESSURANT TANK

PIPEWORK


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The pressurant tank provides sufficient helium storage capacity to repressurise the propellant section to NOP after the first blow down phase. A single fill and drain valve (FDV01) upstream of the NC PV01 and PV02 is used to fill and vent the helium pressurant tank, while pressure is monitored by the high pressure transducer PT3. Two non-return valves (NRVs) are positioned in series on the fuel side and two on the oxidiser side of the system preventing propellant or propellant vapour from one side of the system reaching the other during the repressurization. During the first stage NRVs are protected from propellant vapours by four NC pyrovalves, PV03, PV04, PV05 and PV06 which are fired before repressurization. Test port TP02, TP03 and TP04 provides access during testing. Fill and vent valves FVV05 and FVV06 are used for testing and for propellant tank filling. NC pyrovalves PV01 and PV02 isolate the pressurant tank from the propellant tanks and repressurise the propellant section when firing. A flow limiter orifice/restrictor on the inlet would prevent an abrupt repressurisation. A filter is to be fitted before the flow limiter orifice/restrictor to protect it against any particulate contamination (TBD). After repressurisation, the two NO pyrovalves PV03 and PV04 are fired in order to isolate the propellant section reducing any possible leakage though the pressurant section.

11.2.2 Propellant Subsystem

Two propellant tanks, specifically sized to meet the total propellant load requirement for the mission, are fitted in the Solar Orbiter CPS configuration as shown in the following figure. The components downstream of the propellant tanks, isolation valves, pressure transducers, filters, orifices, and service valves, are located on the –Y central shear wall. The location of the fill and drain valves on the –X bottom floor provided access to ground support equipment (GSE) for ground testing, propellant loading and propellant off-loaded if it is necessary. There are two sets of eight thrusters, a prime branch and a redundant branch providing a redundant set of thrusters in each location. Each branch is sufficient to meet the transfer orbit and perform the necessary three-axis stabilisation control. Each propellant tank is accessed via a single fill and drain valve (FDV07 and FDV08) and a single fill and vent valve (FVV05 and FVV06). The mechanical layout of fittings and pipework for these valves will allows gravity fed drainage from the propellant tanks prior to thermal vacuum drying. Filters are fitted downstream of each propellant tank after the isolation valves (NC pyrovalves) to protect CPS components against the possibility of particulate contamination. There are two test ports (TP09 and TP10) providing access to pipework segments no accessible by FDV or FVV. These are for the purposes of evacuations, purging, leak testing, proof pressure testing and component testing. Two low pressure transducers (PT2 and PT3) are used to measure independent propellant tank pressures during the mission. PTs are mounted downstream the isolation valves and they will confirm the successfully opening of the ISVs. Propellant tanks pressure is monitored by strain gauges on the tanks until ISVs are fired. The CPS launch baseline configuration is to allow propellant only between the NC PV3 to PV5 (below NRVs) and ISVs. In the feed lines to the thrusters there are a total of three inhibits between the propellant tank and thrusters, the redundant isolation valves and the dual valve thrusters, composed of one TLV (second inhibit) and a thruster FCV (third inhibit). Tanks access FVV05, FVV06, FDV07 and FDV08 features three mechanical barriers.

11.2.3 Temperature monitoring

Thermistors provide telemetered flight information on the temperature of the CPS. The number of thermistors and their position is TBD.


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11.3 Thruster Configuration

The thruster configuration on Solar Orbiter (shown in the figure) is based upon that of MEX and VEX, with four redundant thrusters (#1 - #4) located on the corners of the –X panel of the SC. These four thrusters provide complete torque control in any direction as well as provide the primary burn authority along the X axis of the SC (all four thrusters firing), and would therefore be sufficient (as is the case with MEX/VEX) to provide the required control authority if the Solar Orbiter SC was not constrained in attitude during Type 2 TCMs – i.e. if any attitude could be targeted to perform delta V manoeuvres.

Figure 11.3-1: [top] Type 1 TCM allowing an unconstrained slew to optimum delta V attitude,

[bottom] Type 2 TCM accessing the desired velocity vector by rotating the SC around the sunline to bring the velocity vector into the XY plane of the SC, and then performing a dog-leg manoeuvre (blue

arrows)


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Figure 11.3-2: Solar Orbiter thruster positions and designations

In order to provide omni-directional force authority in Type 2 TCM cases, the four rear thrusters are complemented by two additional thrusters on the Y panels of the SC (#5 - #8). These additional Y-wall thrusters provide the ability to access certain delta V directions during Type 2 TCMs, when the SC attitude is constrained to remain pointing at the Sun (such that the heatshield can perform it’s function). Essentially during Type 2 TCM, the SC is first rolled to bring the desired velocity vector to within the XY plane of the SC PRF. The required velocity is then accessed through a dog-leg manoeuvre within this plane. Furthermore:

• The –X panel thrusters are tilted outwards by 15 degrees with respect to the longtitudinal axis, which is enough to provide roll torque capacity, whilst not affecting overmuch the thrust efficiency

• The -X panel thrusters are also tilted by 45 degrees with respect to the SC Y axis in order to avoid plume interaction with sensitive SC elements

• The Y panel thrusters are mounted in the middle of the Y panels, with one inclined by 50 degrees with respect to the SC X axis, towards the +X direction. The remaining thruster is normal to the panel.

11.4 Functional Description

The sections below provide an outline of the functional operation of the chemical propulsion system during the mission.

11.4.1 Launch Configuration

The helium tank is pressurised to MEOP at the launch thermal conditions. Temperature and pressure are monitored by one thermistor and the pressure transducer PT1. NC Pyrovalves PV01 and PV02 isolate the pressurant tank from the propellant section avoiding any possibility of premature repressurization during the launch phase and first blow-down stage.


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NRVs and the four NC pyrovalves, PV03, PV04, PV05 and PV06 prevent any possibly of propellant gas mixture. Pyrovalves also provide protection to the NRVs against propellant vapours until repressurisation. The propellant tanks are filled and pressurised to NOP at the maximum launch temperature. Pressure is monitored by strain gauges while temperature is monitored by two thermistors in the bottom and one in the top of each tank. ISVs are closed until initialisation. Propellant feed lines downstream are pressurised with helium at 3 bara. By having a positive pressure differential between the CPS and the environment any potential ingress of contamination prior to and during the launch is eliminated.

11.4.2 Initialization and First Blow-Down Phase

Helium must be vented from the propellant feed lines before they are primed with gas-free propellant. During the initialization process propellant lines below ISVs are vented by opening the prime branch thruster valves. The system is then primed with propellants by firing one ISV in each side. The redundant ISV shall be fired only if PT downstream indicated an unsuccessful prime ISV opening. As a consequence of propellant loading there is a small amount of helium becoming trapped in the pipework just upstream the ISVs which will be released into the lines during priming. Thrusters have been qualified for operation with propellant containing helium bubbles. After initialization, the CPS is ready to delivered propellants to the thrusters in blow-down mode. Orbital injection is performed using four 10N thrusters. The design of the CPS is such that any number of thruster firings can be performed. Limitation only exists in the amount of propellant that can be contained within the tanks, the thrusters themselves, is not limited with regard to operational cycles. The propellant tanks design utilises a PMD to ensure that gas free propellants are delivered to the engines during all mission phases. The device incorporates a reservoir to ensure that a liquid supply is available even when the inertial forces on the spacecraft are such that propellant would normally be driven away from the outlet. The propellants are consumed by the thrusters to perform the necessary attitude control manoeuvres during transfer orbit. During this first stage thrusters are all fed within their acceptance pressure range by a common feed system and therefore the total usable transfer orbit propellant is available for use by any thruster.

11.4.3 Repressurization and Second Blow-Down Phase

After the first blow-down phase, the pressure in the propellant section drops to the thruster minimum operating pressure. The system then is repressurised to NOP by first firing pyrovalves PV03 to PV06 and then firing pyrovalves PV01 and PV02. A flow limiter orifice/restrictor on the inlet would prevent an abrupt repressurization, allowing a gradual build up of pressure in the pipework downstream. Once the pressure in the propellant tanks has stabilise the system is ready for the second blow-down phase, the two NO pyrovalves PV07 and PV08 are fired in order to isolate the pressurant section reducing any possible leakage though the pressurant section, increasing the system reliability. With the propellant tank repressurised to NOP and isolated, the system is ready to perform the necessary attitude control manoeuvres and complete the transfer orbit using the dual-valve 10N thrusters in blow-down mode.

11.4.4 CPS Monitoring

The CPS monitoring facility is provided by the on board computer (OBC). The power distribution unit (PCDU) provides the power to all the CPS electric and electronic components, i.e. pressure transducers, pyrovalves and thruster valves. There is redundancy in the functionality of either the OBC or the PCDU in being able to actuate the CPS components.


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11.5 Major Components Design Criteria

11.5.1 Low Pressure NOP Derivation

As described before, the Solar Orbiter will have two blow-down stages with a one shot repressurisation. During the blow-down stages thrusters will be fed with propellant within their acceptance pressure range, from 18.7 bar to 10.3 bar. The figure below shows the thruster qualification and acceptance boxes for the dual valve thrusters. Assuming a tuned pressure drop, due by the flight orifices (FO-01 and FO-02), between the tanks to the thrusters of 0.3 bar during firing, a NOP of 19.0 bar has been determined.

10N Thrusters Performance Operational Points

E7

E6

E5E4

E3

E2

E1

M1

M2

M3

M4M5

M6

M7

4

5

6

7

8

9

10

11

12

13

14

15

16

17

18

19

20

21

4 5 6 7 8 9 10 11 12 13 14 15 16 17 18 19 20 21Oxidant Pressure [bar]

Fuel

Pre

ssur

e [b

ar] Qualification Pressure Box

Acceptance Test PointsExpanded Operational boxMargin box

Figure 11.5-1: EADS 10N thruster pressure boxes

11.5.2 Propellant Tanks Sizing

The CPS propellant and pressurant tanks sizes are determined using the end of life dry mass and delta V requirements:

a) a ΔV budget in the range of 325 m/s for the most demanding mission case (including margins and thrust efficiency due to thruster configuration).

b) a spacecraft dry mass about 1430 kg with tank configuration composed of two tanks of small diameter, < 800 mm

In addition, the propellant capacity shall meet the following two major functional constraints:

c) During the first blow-down phase the tank ullage shall be enough to delivery propellant to the thrusters within their acceptance pressure range, from 18.7 to 10.3 bar and the remaining propellant shall be enough to complete the mission.

d) After the repressurisation the pressure and ullage shall be enough to delivery all the remaining propellant to the thrusters within their acceptance pressure range to complete the mission.


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e) Pressure in the propellant tank shall be such that the oxidiser boiling point is maintained at least 10°C above the maximum operating temperature limit.

As a result a total of 192.6 kg are required to provide a delta V of 325 m/s and AOCS manoeuvres (16 kg) with an end of life dry mass of 1430 kg. Tanks shall storage 119.8 kg of NTO (83 litres at 20ºC) and 72.8 kg of MMH (83.2 litres at 20º). For a 121 litres tank, an initial ullage of 37.8 litres will delivery 30.0 litres of propellant during the first blow-down, from 19.0 bar to 10.6 bar. After repressurisation, in order to delivery the remaining 53.2 litres of propellant within the thruster acceptance pressure range, the tank ullage of 67.8 litres shall be repressurised to 19 bar. Propellant tank ullage and pressure evolution during the first blow-down stage, repressurization and second blow-down stage are shown in the following figure.

Bipropellant Blowdown Represurized System

9

10

11

12

13

14

15

16

17

18

19

0.0 10.0 20.0 30.0 40.0 50.0 60.0 70.0 80.0 90.0 100.0 110.0 120.0 130.0

Ullage [Litres]

Pres

sure

[bar

] 1st blowdown2nd blowdownRepresurizationInitial UllageTank capacity

Figure 11.5-2: Propellant tank pressure and ullage evolution

11.5.3 Pressurant Tank Sizing

For the Pressurant tank sizing there are two main requirements:

1) Pressurant tank shall contain enough Helium to repressurise the propellant tanks ullage to NOP after the first blow-down stage.

2) He temperature shall be at least 10ºC above the freezing point of the propellant during the repressurisation (-13ºC).

A propellant tank of 13.3 litres, pressurised at 112.0 bar with 0.2284 kg of Helium will increase the propellant tanks ullage of 67.8 litres from 10.6 bar to 19.0 bar. During propellant tanks repressurisation Helium in the pressurant tank will be depressurised and temperature will drop. Using a restrictor with Lohm = 44,000 the depressurisation takes around 2 hours allowing to keep the He temperature well above the propellant freezing temperature.


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The following table presents the 13.3 litres temperature evolution (polytrophic process) during a slow repressurisation of 2 hours. It is important to note that for the calculations 5W tank heaters and a proportional tank mass have been taken into account.

Press [bar]

Temp [K]

Press [bar]

Temp [K]

Press [bar]

Temp [K]

Press [bar]

Temp [K]

112 313.0 88 307.8 64 302.0 40 296.7 111 312.8 87 307.6 63 301.8 39 296.4 110 312.6 86 307.3 62 301.6 38 296.1 109 312.5 85 307.0 61 301.4 37 295.9 108 312.3 84 306.7 60 301.2 36 295.6 107 312.1 83 306.5 59 301.1 35 295.3 106 311.9 82 306.2 58 300.9 34 295.0 105 311.7 81 305.9 57 300.7 33 294.7 104 311.5 80 305.6 56 300.5 32 294.4 103 311.3 79 305.3 55 300.3 31 294.1 102 311.1 78 305.0 54 300.0 30 293.7 101 310.9 77 304.7 53 299.8 29 293.4 100 310.7 76 304.3 52 299.6 28 293.1 99 310.4 75 304.0 51 299.4 27 293.4 98 310.2 74 303.7 50 299.2 26 293.6 97 310.0 73 303.5 49 298.9 25 293.8 96 309.8 72 303.3 48 298.7 24 294.0 95 309.5 71 303.2 47 298.5 23 294.1 94 309.3 70 303.0 46 298.2 22 294.2 93 309.1 69 302.9 45 298.0 21 294.2 92 308.8 68 302.7 44 297.7 20 294.1 91 308.6 67 302.5 43 297.5 19.1 294.2 90 308.3 66 302.3 42 297.2 89 308.1 65 302.2 41 297.0

Table 11.5-1: Propellant tank Helium temperature evolution


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12. ATTITUDE & ORBIT CONTROL SUBSYSTEM

For a detailed description of the AOCS, please refer to [IR28].

12.1 Requirements

The key requirements placed on the Solar Orbiter attitude and orbit control subsystem are presented in the following table.


Autonomy The AOCS shall provide throughout all mission phases an autonomous capability to maintain the required attitude

The AOCS must provide a sufficient degree of autonomy in order to deal with the repeating periods of non-communication that occur throughout the mission

Survivability Maintain attitude to ensure heatshield function

The AOCS must maintain the pointing of the SC to the Sun in order to maintain the function of the heatshield during the operational phase of the mission

Recovery Provide MGA backup communication during failures

Once a serious failure has occurred during the science phase, it must be assumed that thecommunication to ground via the HGA is lost. For back-up communication the MGA link is planned to be used. This requires the AOCS to control the SC rotation about the SC x-axis (sun line) such that the MGA antenna is pointing to the earth.

Pointing Stability As per the table in the MRD The RPE is the key pointing stability requirement during the science windows – this requirement necessitates the use of reaction wheels as the attitude actuators (thrusters are too noisy).

Table 12.1-1: AOCS principal requirements


The Solar Orbiter AOCS design is based on the BepiColombo baseline and aims at a re-use of BepiColombo equipment as far as possible. In fact the only significant functional difference to the BepiColombo implementation is the Survival Mode and the Safe Mode. This implies that the unit FCE will be different for Solar Orbiter and the sensors used in survival mode are coarse sun sensors (CSS) and rate measurement units (RMU) instead of fine sun sensors and a second inertial measurement unit used on BepiColombo.

12.2.1 Pointing Domain Classifications

Allowing small attitude deviations during failure occurrence and recovery, while having hard restrictions on the maximum allowed attitude range, leads to the definition of attitude domains. The attitude domains are defined w.r.t. the sun vector direction in the body frame: Forbidden Domain: Deviations of more than 6.5° from the sun line (sun centre). Allowed contingency Domain for survival: In this domain the attitudes are allowed during failure occurrence and recovery from a failure. In this domain the SC can survive in principle unconstrained in time - within a half cone of 6.5°. Survival Monitor Domain: Pointing within this domain will not trigger the Survival Mode. Attitudes outside this domain will trigger an AOCS Alarm. Half cone: 1.8° (TBC depending on sun sensor performance). Allowed contingency Domain for recovery to nominal mode by ground: This domain adds to the previous domain a constraint as rotation about the sun line such that the MGA is pointed towards the earth. The rated inertial angle need to be maintained by the ground and is assumed to be stored in on-board non-volatile memory tables. The domain size is +/- 5 deg (TBC).


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Operational Domain: The remaining attitudes, which are allowed to be commanded during the nominal science mission: deviation from sun center pointing is within a half cone of 1.3°.

12.2.2 Nominal Operations

From an AOCS viewpoint, the operations start as soon as the spacecraft is powered on the launch vehicle – the OBC will automatically boot up and the AOCS will enter AOCS_Standby Mode_SBM, where no AOCS actions are performed other than FDIR monitoring. When the separation is detected, the DMS will command the AOCS into its first active mode in order to damp out the tip-off rates from the launcher; this is AOCS_Sun Acquisition Mode_Sub Mode (Rate Reduction), and to acquire correct Sun pointing. IMU and thrusters are used for rate damping and Sun sensor information is also used to support Sun acquisition. Then the DMS will command the AOCS to suspend attitude control, initiate deployment of the solar array and, when the deployment is complete, re-activate the AOCS control action to re-acquire Sun pointing. The AOCS LEOP functionality as outlined above assumes sun pointing measurements are available in the +x-axis within a halfcone angle of 15 degrees, as determined by the sensor baffles at the heatshield corners and in the -x-axis direction within a halfcone angle of 60 degree, allowing the use of standard coarse or fine sun sensors. Redundant sensor outputs are available for LEOP in the direction of + and - x-axis, i.e., if not internally redundant, there are 2 sun sensors to be accommodated on each side. The sensors accommodated in the heatshield need to provide in addition to the measurement range required for LEOP the performance and redundancy suitable for Survival mode as specified in the SS TDA. Since BepiColombo will use FSS's, the electrical interfaces for FSS's will be available in the RIU. Interfaces for Coarse SS's will need to be build in specifically for Solar Orbiter. A cost trade may be advisable before a final decision for Solar Orbiter is made. This should consider the outcome of the SS TDA. In order to achieve 3-axis attitude information, star tracker data is required. The star trackers are commissioned in the AOCS_Sun Hold Mode. In this mode, the star trackers and gyros are used for attitude measurement and thrusters are used for attitude control. The necessary re-orientation manoeuvre is performed automatically in this mode to achieve the required 3-axis pointing. The reaction wheels are then commissioned, after which a transition from thrusters to wheel actuators is performed within this mode. This represents the end of the AOCS start up modes. There are further specific requirements to be fulfilled during the mission. One is the required flip to the “reverse” pointing attitude, i.e. Heatshield in the anti-Sun direction. Further on any inertial attitude shall be commandable for orbit control manoeuvres. The AOCS Block Diagram is shown in the following figure.


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CSSCSS

OBC RIU

RMUSTR CSS

RW PROP

Solar Orbiter AOCS Architecture

IMU

FCESADE

APME

AOCS

Figure 12.2-1: Solar Orbiter AOCS architecture


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12.2.3 AOCS Modes

The Solar Orbiter AOCS modes are heavily influenced by the BepiColombo modes. See the figure below for the Solar Orbiter ACOS mode diagram.

Figure 12.2-2: AOCS mode diagram


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Standby Mode [SBM]: This is the entry point of the AOCS diagram. It is triggered after separation from the launcher and also in case of a major system failure. In this mode, the AOCS is actually inactive, as the mode purpose is to perform the system and software initialisation and the solar array deployment. Sun Acquisition Mode [SAM]: This mode occurs during LEOP and starts with the Rate Reduction Sub-Mode, RRM, that reduces the spacecraft rates using gyro data and chemical thrusters. At this time, the attitude is arbitrary and unknown. The rate measurement is based on the 4-axis IMU. As soon as the rates are damped to a low value, the mode aims at pointing a specific spacecraft axis to the Sun in order to ensure a stable and safe attitude. Then performing a rotation around the Sun direction, at a low spin rate, aimed at several purposes (STR field of view clearance, Earth strobing function). The SAM uses a set of Coarse Sun sensors providing a near spherical coverage in order to support a fast and simple Sun acquisition or an optimized attitude for LEOP X-band coverage. Safe Hold Mode [SHM]: Performs a full 3-axis attitude acquisition using star tracker measurements in LEOP and brings the spacecraft from the Sun pointed attitude into its nominal attitude in according to the mission phase. In LEOP the attitude can be preselected to optimize the X-Band link. For this purpose, an autonomous slew function is implemented using thrusters. The Earth link is established with the HTHGA (using on-board stored Earth ephemeris) and reaction wheels are switched on (optionally). Survival Mode [SUM]: This mode is based on a dedicated Failure Control Electronic FCE, and triggers if either the Sun is outside a defined limit or the rate exceeds a predefined limit. This mode has to maintain a correct Sun-pointing even in case of a serious thruster or wheel failure. The SUM uses CSS (looking in nominal Sun direction) and RMUs. This mode is enabled only in case of [NM]_Sun-pointing near the sun (<0.8 AU). Normal Mode [NM]: This mode supports a number of guidance options in the Inertial Pointing Sub-Mode, to be used mainly in cruise phase, including constant or time-varying quaternion profiles, the same Sun guidance options as in SHM, a Sun-Earth guidance option and a slowly spinning Sun-pointed attitude to save fuel in quiet cruise phases. This sub-mode, with the inertial guidance, is also used as a tranquillisation phase with thrusters or wheels at entry into NM from the manoeuvre modes (OCM). . It is used to control the satellite for Orbit Control Manoeuvres and for the “reverse attitude” as whereas the satellite is pointing in the anti sun direction.

Slew Sub-Mode: is used simply to perform ground commanded slews or the seasonal flip-over slews. The actuators used, i.e. wheels or thrusters, are defined by a flag setting. Wheel Off-loading Sub-Mode: can be activated either from ground command or autonomously from the Inertial Pointing or the Nadir Pointing Sub-Mode without changing the guidance law. The reaction wheels are desaturated with the chemical propulsion. Sun-Pointing Sub-Mode: provides the high performance measurement and control performance. The data from 2 APS Star Tracker and the very high performance IMU is combined in a Gyro Stellar Estimator for optimum attitude estimation and pointing stability. The sub-mode allows as well for a pointing to the limb of the sun. This is done by a commandable “Offset table”.

Orbit Control Mode [OCM]: is used for all delta-Vs performed during the mission using thrusters. The attitude guidance can be fixed or slowly varying, the attitude estimation is based on gyro measurements (with or without star tracker), and the ΔV amplitude is controlled through the thruster pulse counting method. The OCM includes a thrust modulation ramp-up function, in order to minimise the initial attitude transient. Once the imparted delta-V is equal to the commanded delta-V, an automatic transition back to normal mode is performed. During the period in OCM, the wheel speeds are either held constant or can be commanded to new values. The use of the various AOCS equipment within the modes is shown in the following table.


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AOCS Mode RMU Sun Sensors IMU STR Reaction Wheels

Thrusters SADE APME

SAM X X X SHM X X X X X

NOM (Sun Pointing)

X X1) X X X

NOM (Slew)

X X X (X) X X

NOM (Inertial pointing)

X X X X X

NOM (Wheel Offloading)

X (X) X X X X

OCM X (X) (X)2) X (X) (X) SUM X X X SFM X X X

Table 12.2-1: AOCS resources to mode table

(*) Optional usage 1): Two STRs are used 2): Reaction wheels in open loop (speed control).

In Operational Modes, any identified unit failure (e.g. wheels, thrusters, gyros, STR) leads to an automatic passivation and reconfiguration to the redundant unit. The nominal mode remains active, and only a transient pointing/stability event may occur (with respect to operational conditions). In case of STR loss, the attitude is propagated using gyro measurements, and a slight pointing performance degradation is observed. The attitude will slowly drifts from its nominal pointing, eventually leading to a system restart (e.g. due to system hardware alarm). However in the case of Sun pointing attitude guidance, the +X Sun sensors can be used to compensate for the gyros drifts, thus avoiding a Sun off-pointing system alarm. In the case where the STR outage duration gets too long to maintain the HGA pointing commensurate with the RF link accuracy constraint, then an automatic switch to a Sun/Earth pointed attitude allowing the MGA to point to the Earth could be entered, in order to recover communications with the ground without entering the survival mode (TBC).

12.3 AOCS Equipment Overview

12.3.1 Star Tracker

The attitude measurement accuracy on-board requires the fusion of measurements of at least two star tracker heads with significant line of sight separation together with the measurements of an inertial measurement unit. The attitude estimation based on a Kalman filter provides precision attitude knowledge whereas direct star tracker measurements augmented by propagation with the inertial measurement unit is used for the attitude control. The star trackers viewing directions have to be optimized with respect to:

• Optimization of the line of sight separation for nominal operation and • Avoidance of sun in the field of view during reverse orientation.

During orbit control manoeuvres transient Sun-blinding of star trackers might occur; however, the attitude estimation concept is robust towards the transient blinding of one or even two star trackers due to propagation with the redundant high precision inertial measurement unit. Three star trackers are accommodated on Solar Orbiter, with one in cold redundancy. The following figure shows a three STR


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accommodation on a common mounting integrated with the IMU; this obviously represents the ideal configuration from a performance perspective, as long as relative angles are sufficiently large – it is therefore proposed to operate the two outer STR as the prime pair, with the central STR the cold-redundant unit.

Figure 12.3-1: Star Trackers mounted on sensor plate together with inertial measurement unit

Each star tracker provides 3-axis attitude data in form of a quaternion. It is able to determine this data from a “lost in space” situation within a few seconds. As a baseline, a star tracker with the new APS (active pixel) technology is selected. This APS type of star tracker has excellent radiation hardness and operational robustness. It delivers high accuracy and update rates even in presence of significant angular rates and is robust towards moon blinding. A standard 26° sun baffle is foreeseen, and the star tracker will never look in sun direction when operating close to the sun (<0.8 AU). Therefore an expensive shutter device - as needed for BepiColombo - is not required.

12.3.2 Inertial Measurement Unit

The inertial measurement unit (IMU) usage is two-fold:

• In combination with the star trackers to improve the attitude estimation • As an angular rate feedback sensor within the attitude control algorithms. Similar to the star trackers

the performance needs are driven by the attitude knowledge requirement. In order to be compliant with the one failure tolerance philosophy, the IMU shall be redundant. This is ensured by a tetrahedral arrangement of four single axis measurement units. Thus any three out of the four units still deliver 3-axes information. However, analysis of the associated 3-axis performance has to apply a form factor to the performance parameter of a single unit to account for the imperfect (non-orthogonal) alignment of the remaining three units.


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Figure 12.3-2: The Hemispherical Resonator Gyro SSIRU inertial measurement unit

Key parameters for the filter based attitude estimation are the angular random walk (ARW), the rate random walk, and also the scale factor stability. The desired compliance with the attitude knowledge accuracy can only be accomplished with a high precision IMU, such as the Hemispherical Resonator Gyro SSIRU shown in the figure. The Hemispherical Resonator Gyro (HRG) sensor is comprised of three simple machined quartz parts – a high-Q vibrating hemispherical resonator, and inner and outer shells which support the resonator. The quartz construction of the HRG is inherently stable, impervious to aging effects, and naturally radiation-hard. Operating in a completely evacuated hermetically sealed case, the HRG sensor is not affected by any phenomena that could make it wear out or limit its life – making it an ideal gyro for space applications.

12.3.3 Rate Measurement Unit

The rate measurement unit (RMU) is a low performance equipment when compared to the IMU, being required to monitor rate for the HW-FDIR system. The control concept does not include any rate integration to obtain an attitude (HW-FDIR attitude monitoring is accomplished using the CSS). Three RMU units shall be integrated in order to allow majority voting.

12.3.4 Coarse Sun Sensor

The use of Coarse Sun Sensors will be investigated within a Solar Orbiter TDA. The concept could be based on the robust and well known Coarse Sun Sensors.

Figure 12.3-3: [left] RMU Unit, [right] Coarse Sun Sensor


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12.3.5 Reaction Wheels

The current baseline is the 68Nms reaction wheel from Teldix – this is the largest reaction wheel available within the European market, selected due to uncertainties in the eventual off-loading frequency of the wheels during science windows – preliminary analysis indicates several desaturation events needed every ten-day window, so the largest wheel has been selected.


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13. FAILURE DETECTION ISOLATION & RECOVERY SUBSYSTEM

For a detailed description of the FDIR, please refer to [IR28].

13.1 Requirements

The key requirements placed on the FDIR subsystem are presented in the following table.


Level 0,1 failures The FDIR must allow response to on-board failures without recovery by switching, ex-G, to the redundant functional path, such that the mission can continue

This requirement for failure-management enables continuity of the mission timeline and performance (in particular the uninterrupted generation of mission products). In the event that redundant paths do not exist or that the failure effect is too complex to allow autonomous recovery, the spaceraft shall enter survival mode

FDIR levels FDIR level definition There should be a clearly demarked separation between redundancy management FDIR and system safety FDIR. These two fundamental FDIR levels should not generate perturbations that effect each other

Subsidiarity For failures that are non-critical, hierarchy should be used to resolve them

For failures whose resolution does not imply safeguarding of system functions, hierarchical steps should be applied before eventually deciding to remove the failed unit from operation

Sun-pointing The Survival mode must maintain sun-pointing to within TBD degrees for all off-pointing failure cases

The criticality of maintaining sun-pointing must be stressed; regardless of failure type, the Survival Mode FDIR must be capable of ensuring continuing functioning of the heatshield

Table 13.1-1: FDIR key requirements


The on-board Solar Orbiter FDIR concept will be based on a hiearchical approach where the FDIR functions are implemented at various levels. A level is defined to be a specific instance reacting to a failure situation. The Solar Orbiter FDIR shall be composed of four levels.

FDIR Level Reacting Instance Triggering Events Actions Performed Level 4 Survival and Safe Mode Sun off-pointing surveillance Initiate AOCS alarm (input to

OBD RM to reset processor and start safe mode SW) If required, start and maintain temporary thruster controlled survival mode control

Level 3 Reconfiguration module HW-generated alarm inputs to the OBC RM

• Watch dog • PM undervoltage • PM CPU (internal

CPU HW errors) • PM other internal

errors

Execution of reconfiguration sequences controlled by RMs

Level 2 CDMU System SW SW generated alarm inputs into the RM;

• AOCS alarm (set by AOCS fail safe monitors)

• SW alarm (due to double EDAC, over/under flow,

Execution of reconfiguration sequences controlled by RMs, e.g. Safe Mode for AOCS alarms


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divide by zero, traps and arriving from CDMU/configuration errors detected by SW monitoring/fnuctions etc)

Level 1 CDMU SW SW implemented unit fail operational monitors Subsystem level SW-based FDIR functions (e.g. AOCS fail operational monitors etc.)

Reconfiguration/switch down performed by APSW

Level 0 Instrument FDIR function Failures requiring immediate instrument reaction faster than 10 seconds

Level 0 Unit internal FDIR Unit internal FDIR functions (e.g. RM ‘fail silent, LCLs etc.)’

Table 13.2-1: Solar Orbiter FDIR hierarchy

13.2.1 Hard Wired FDIR

Of particular criticality for the Solar Orbiter mission is the HW FDIR system corresponding to FDIR Level_4, which is concerned simply with maintaining the attitude of the SC such that the heatshield is providing shadow to all the sensitive elements of the spacecraft at all times for all credible failures. Accordingly the design drivers for this FDIR survival mode detection and control are the failure cases that need to be considered that can lead to off-pointing. The spacecraft inertial rates during the operational phase of the mission are very small, with nominal attitude control provided by the reaction wheels, with periodic offloading effected by the thrusters. The credible failures in this case are:

• Loss of attitude control during Sun-centre to Sun-limb repointing manoeuvres: The worst case of this failure case assumes a loss of control during steering from a Sun centre pointing to the limb

• Closed thruster failure during wheel offloading: Thrusters will be operated below the TCOL for delta V manoeuvres and for wheel desaturation

• Open thruster failure: This has been rendered non-credible by the selection of dual-seat valve thrusters

• Wheels blockage failure: Failure which results in a dumping of momentum to the SC from the wheels.

13.2.1.1 CTF (Closed Thruster Failure)

Within the critical (near) distance to the sun, the thrusters will be operated for wheel desaturation only. During this manoeuvre a TCF may occur. The wheel offloading (WOL) manoeuvre is performed as follows:

• Based on the actual wheel speeds, the steering parameters for the manoeuvre are computed using the actual thrust level of each thruster

• During the WOL manoeuvre, the attitude control is still performed on wheels, but the controller parameters (allowed pointing errors, PD control law) and FDIR parameters are adapted

• The WOL is performed by firing the appropriate thrusters for the minimum period with reasonable ISP, i.e. ~50 ms. This leads to an angular momentum of about 0.25 Nms. Without control reaction this results in a maximum angular rate of ~0.03°/s

• Feed forward control of the wheels will not be applied in order to observe directly the wheels behaviour due to nominal controller operation. This leads to a higher short time pointing error in comparison to feed forward control, but this error remains within acceptable limits. This is the baseline concept. Detailed analysis will be done in the next phase when optimizing the control system


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• For each single manoeuvre, the AOCS computes (via a model) the expected performance (reduction of wheels rates) and checks the reached rates against the limits. In case that one thruster stays closed, the AOCS can detect a TCF within a few seconds (after a few pulses spaced by a few seconds). After detection of a TCF, the WOL manoeuvre is stopped and the redundant thruster branch is activated. With the redundant thruster branch, the WOL manoeuvre is repeated automatically

• The pulse stream period is selected such, that the controller has enough time to damp out the disturbance generated by a single thrust pulse

• Such a thruster closed failure cannot lead to a critical control behaviour (depointing or overrate) because the wheel controller is active and keeps the attitude within the nominal limits. If one thruster stay closed, the attitude error is even less than for the failure free case.

13.2.1.2 Open Thruster Failure

The SC system shall provide the following features:

• Each flow control valve has two seats • Each FCV coil has a separate drive electronic.

The supply electronic includes 2 ASICs to provide the FCV drive outputs. The upstream FCVs are driven by one ASIC and the downstream FCVs by the other ASIC. The upstream ASIC monitors the output lines of the downstream ASIC and the downstream ASIC monitors the output lines of the upstream ASIC. If either ASIC detects a continuous output for longer than about 1.5s on any of the outputs of the ASIC it is monitoring then it inhibits the output from both ASICs and reports this via a status word that can be read at 8Hz. With this system no single electrical error or spurious command can lead to a long false firing of a thruster. The resulting angular momentum reached after an OTF is 7.5 Nms. This leads to a maximum SC rate of 0.9 °/s in case of no control.

13.2.1.3 Wheels Blockage Failure

Wheels blockage can only occur due to a fatal mechanical error of the wheel bearing. This is only credible for a massive bearing over temperature. But the bearing temperature is monitored continuously and checked against limits by standard APSW monitors. If the temperature is above a predefined limit (acceptance temperature limit), the FDIR switches the wheel OFF and configures the AOCS to a 3-wheel configuration without any interruption of the nominal operation. Although such a wheel failure case (RW dumps its momentum into the SC within seconds) is considered as non-credible in agreement with ESA, such a failure has been simulated with the HW based control. Each of these failure-cases have been investigated, and the simulations indicate that the spacecraft sensitive elements are protected for all failuer cases for the baseline heatshield dimensions. The following figure provides an example of the analysis conducted.


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Figure 13.2-1: Survival mode simulation for thruster failure, 10Nm over 1s

13.2.1.4 Functional Design of HW-FDIR

The survival mode FDIR must be able to react quickly to faults leading to off-pointing. Because of this criticality, the proposed implementation is with a Hardware Attitude and Rate Anomaly Detection (ARAD) function. In this way the vulnerability of software-controlled FDIR approaches to OBC reboot timescales is avoided. The ARAD will be activated for the close-approach periods of the mission (i.e. continuously during the NMP/EMP and also for sections of the CP). During the periods when it is activated it will have priority over the OBC. The ARAD will rely on majority voting using three identical sensors; this is necessary because unjustified triggering is not allowed in case the Survival mode control is implemented (and based on these measurements). Because the values of the thresholds will change throughout the mission according to specific situations, the thresholds and mandatory configuration data are stored in dedicated registers and can be updated from the ground. The majority voting stage includes a persistency function in order to avoid false-triggering due to measurement spikes. Both the absolute value and the rate of the angular excursion from the Sun-line are monitored by the ARAD; RMUs monitor the angular rate and sun sensors monitor the angle itself. The sensors of the ARAD are operated in hot redundancy. Control of the Solar Array is not necessary because the angular excursion experienced is small.


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The sensors of the ARAD are operated in hot redundancy. The overall detection function need to be single failure free. The thrusters may be preheated in order to allow for high thruster efficiency from the first firing on. The use of the thruster control interface not selected for nominal mode operation (wheels offloading only in phases where the survival mode is active) is enabled for the survival mode control. The controller algorithms are very simple. Separate interfaces of the Survicval Mode HW are to thruster control electronics in RIU, Alarm line to CDMU, sensor outputs for telemetry and Safe Mode use (TBC), power supply and data Interface for adaptation of pointing and rate thresholds. Ground control to enable / disable the control is provided by HPC1. Control of Solar Array is not necessary because the whole spacecraft is controlled and the Survival Mode will be a transitory mode only. Automatic switch over to APSW Safe Mode after a predefined time (TBC).

Figure 13.2-2: ARAD functional configuration scheme

13.2.1.5 ARAD Electrical Implementation

The electrical implementation has been detailed under consideration of the actual BepiColombo DHS design. The proposed implementation is characterized by the following:

• FCE is controlled via Ground Command only • Ground commands are routed via OBC TM/TC decoder • These commands are required for enabling/disabling of the survival Mode control (HPC1 ground

override) • The ground can load data via the OBC/space wire interface (TBC) into the survival electronics/FCE

SGM • The ground can read data via the OBC/space wire interface (TBC) from the survival electronics/FCE

SGM • All commands controlling the RCS are routed via the FCE • All CSS and RMU analogue signals are converted within FCE • Only two axes information of the RMUs are used for threshold generation and control. The rotation

rate around X (sun direction) is not used • All A/D converted sensor data are routed to both FCE logics (cross strapping) • Both FCE logics operate nominally in hot redundancy • Each FCE uses both, CSS and RMU data • After the majority voting stages only one attitude and one rate vector remains


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• In the thresholding stage the Alarm is generated. An Alarm is generated in case that either the attitude is out of range or the rate is out of range. This alarm set an alarm register which holds its status unless it is reset

• Only in case that Alarms A and B are active, the FCE overtakes control (switching unit); this avoids a failure of one FCE part can lead to a wrong Alarm and Survival Mode triggering

• When enabled, the survival mode controller generates the control signal for the CPS / thrusters which are routed via the Space Wire interface to the RIU

• In case, the FCE overtakes control, the following actions are performed: o The Spacewire router is switched over from OBC control to FCE control of the CPS o The FCE generates a RESET of the OBC o The FCE starts a timer (~60 s). After that time,

the FCE switches over the RCS control back to the OBC to start OBC Safe Mode performs a reset of the of the alarm register

• The OBC receives full Health Status data from FCE via SPW. All RMU data (3 axes • measurement and housekeeping via RIU and SPW) • The ground can command the RMU via OBC and RIU (TBC).

The following data have to be stored within the FCE SGM:

• Calibration polynom for each individual solar cell of the Coarse Sun Sensors: 12 cells, 2 • parameter each: 24 data • Thresholds for each axis of each sensor axis: 3 * 2 axis: 6 parameters • Configuration for the CSS and RMUs (use or not use): 3 * CSS + 3 * RMU: 1 word • Configuration of RCS: Nominal thruster force depending on pressure and use of Branch A • or Branch B: 1 word • Health Status of the two FCE's: 1 word • Time data for FCE operation until take over to OBC: 1 word

The electrical implementation is depicted in the following figure.


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Figure 13.2-3: Survival Mode Electrical Architecture

13.2.2 APSW Based Safe Mode Implementation

There is no need to save/store the accurate attitude data when there is a switch over to survival mode. The required data (orbit, ephemeris, etc.) are stored in an SGM linked with time. The actual time needs to be generated in a safe time module (e.g. time modules triple redundant in PM, and both hot redundant CDMU core modules, e.g. TRR's on Aeolus). In case of solar flares, prohibiting the operation of the STR, the satellite remains in a sun fixed attitude using the IMU for slow rotation about sun line. As soon as a STR provides stable data, the actual position and the desired attitude for MGA coverage is computed. The slew to this attitude is performed with the IMU. The STR supports the hold of the final attitude. The functionality of the APSW based Safe Mode is depicted in the following figure.


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Figure 13.2-4: APSW Safe Mode functionality


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14. DATA MANAGEMENT SUBSYSTEM

For a detailed description of the Solar Orbiter DMS, please refer to [IR27].


The Solar Orbiter Data Management System (DHS) is highly-derived from the BepiColombo DHS and provides all required functionality for operational access and control of the spacecraft including AOCS processing and control, interfaces to units/payload instruments and data storage. The DHS is based on three units: • The On-board Computer [OBC]

• Remote Interface Unit [RIU]

• Solid State Mass Memory [SSMM]

• Failure Control Electronics [FCE].

The overall electrical architecture of the spacecraft, controlled by the DHS, is shown in the following figure.

Out

put-

IF B

DC

/DC

MC

B

Out

put-I

F A

DC

/DC

MM

&IO

B

DC

/DC

MM

&IO

A

DC

/DC

MC

A

Inpu

t-IF

A

Pay

load

Spa

cew

ire IF

s A

Inpu

t-IF

B

Payl

oad

Spa

cew

ire IF

s B

Drv

A+B

(3-a

xis)

Shor

t

Drv

A+B

(2-a

xis)

DC

/DC

DC

/DC

OBC

IF BO

BC IF A

OBC

IF BO

BC IF A

STD I/O

AS

TD I/O

BSTD

I/O C

CP

S-I/O

AC

PS-

I/O B

STD I/O

ASTD

I/O B

SSM

M A

IFS

SMM

-B IF

ME

TIS

Inst

r.

RIU

-B IF

s

Pro

ram

E

EP

RO

M 1

Pro

ram

E

EP

RO

M 2

RIU

-A IF

ME

TIS

Inst

r.SS

MM

-B IF

SSM

M A

IFR

IU-B

IFs

RIU

-A IF

LCL

DC

/DC

LCL

DC

/DC

WG

S-1

A P

os 1

WG

S-1

B P

os 1

WG

S-1

A P

os 2

WG

S-1

B P

os 2

WG

S-1

A S

tatu

sW

GS

-1 B

Sta

tus

WG

S-2

A P

os 1

WG

S-2

B P

os 1

WG

S-2

A P

os 2

WG

S-2

B P

os 2

WG

S-2

A S

tatu

sW

GS

-2 B

Sta

tus

WG

S-3

A P

os 1

WG

S-3

B P

os 1

WG

S-3

A P

os 2

WG

S-3

B P

os 2

WG

S-3

A S

tatu

sW

GS

-3 B

Sta

tus

WG

S-4

A P

os 1

WG

S-4

B P

os 1

WG

S-4

A P

os 2

WG

S-4

B P

os 2

WG

S-4

A S

tatu

sW

GS

-4 B

Sta

tus

WG

S-5

A P

os 1

WG

S-5

B P

os 1

WG

S-5

A P

os 2

WG

S-5

B P

os 2

WG

S-5

A S

tatu

sW

GS

-5 B

Sta

tus

WG

S-6

A P

os 1

WG

S-6

B P

os 1

WG

S-6

A P

os 2

WG

S-6

B P

os 2

WG

S-6

A S

tatu

sW

GS

-6 B

Sta

tus

Ka-

Ban

d Tr

ans.

A

X-B

and

Tran

s.A

X-B

and

Rec

eive

r/D

emod

A

Ka-

Ban

d Tr

ans.

B

X-B

and

Tran

s.B

X-B

and

Rec

eive

r/D

emod

B

CP

S A

rmin

g (S

kin-

CB

s)

OB

C -A

IF

OB

C -B

IF

OBC

-A IF

OB

C -B

IF

Pyro

Firi

ng (A

+B)

Sol

ar A

rray

Sim

ulat

or I/

F

DC

/DC

DC

/DC

Driv

e (A

+B)

Ski

n-A

rm(-S

afe)

tbc

Driv

e (A

+B)

SAD

M

HD

2A

–2B

PAN

EL

4

PAN

EL

3

SAD

M

Alam

sD

iscr

. TM

/TC

Ala

ms

Dis

cr.

TM/T

C

HD

1A

–1B

PAN

EL2

PAN

EL1

DC

/DC

DC

/DC

Spac

e W

ire

Spa

ce W

ire

Figure 14.1-1: Baseline Electrical Architecture on Solar Orbiter

14.1.1 Solid State Mass Memory

The Solid State Mass Memory (SSMM) records data emitted from Payload Instruments and passes the data, on request, to the OBC (On-board Computer) and to the X-band and Ka-band downlink. In the standard design (Bepi Colombo Design), the SSMM controls 9 instruments. In order to keep the number of interfaces to the instruments the same (allowing unmodified resuse of the interface), one instrument (METIS) is routed via the OBC SpW to the SSMM. The data exchange between the instruments is handled via the SSMM. The architecture of the SSMM is shown in the following figure.


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Mass Memory Array

512 Gbit(EOL)

Internal Redundant

Memory Controller A

Memory Controller B

Output Routing

A

Output Routing

B

SpWRoutingFunction

A

SpWRoutingFunction

B

Payload Instruments

nominal

Payload Instrumentsredundant

123456789

123456789

OBC A OBC B

OBC A OBC B

Science and Non-

Science TM Packets

Science andNon-Science TM Packets

Replay

Replay

X- Band TFGA (VC1)

X- Band TFGA (VC2)

Ka- Band TFGA (VC3)

Power Converter A

Power Converter B

Non Science T MPackets

Non Science

TM Packets

SpW link

A & B Cross- strapping

EGSE Quick Load / Dump Link

EGSE Quick Load / Dump Link

Main Power

EGSE Power

Main Power

EGSE Power

Data Retention

Data Retention

X- Band TFGB (VC1)

X- Band TFGB (VC2)

Ka- Band TFGB (VC3)

TC Packets

TC Packets

Option

SpW Interfaces Cross- strapping

Figure 14.1-2: Solid State Mass memory Architecture

The SSMM consists of the following functional blocks:

• Space Wire Router A and B interfacing with the 9 instruments main and redundant inputs/outputs • Mass Memory Array consisting of 3 modules each 256 = 768 Gb BoL. For EoL, 512 Gb are

considered assuming one of the 3 module will degrade; this is in-line with the memory requirements of the mission as analysed – see the budgets section for the sizing SSMM-usage profiles, indicating 491 Gb (2017) and 333 Gb (2018) peak usage at MoL. Typical usage is more like 200 Gb.

• Power Converter of A and B • Memory Controller A and B • Output Router, interfacing with the OBC.

14.1.1.1 Solid State Mass Memory External Interfaces

The Solid State Mass Memory provides the following external interfaces: • 28 Space Wire Interfaces

o 18 Space Wire Interfaces to 9 Instruments (Non + Red) o 4 Space Wire Interfaces (Nom + Red) with OBC, for communication o 6 Space Wire Interfaces (Nom + Red) with OBC for X/Ka-band Telemetry.

• DC 28 V Power Interface nominal and redundant • EGSE SW fast load and dump Interface for access to nominal and redundant Controller.

14.1.1.2 Functional Description

The SSMM provides the following basic functionality:

• Behave as a disc unit to the users and support a simple (flat) filing system • Support random access • Manage free space and automatically mark bad areas as unusable • Make information available on request about free space, files stored and bad memory blocks • Allow multiple read/write access simultaneously for all possible data sources and destinations


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• Allow the Ground, via dedicated telecommands, to create, delete, rename and copy files • Receive Science and Non-Science TM Packets from each of the nine Payload Instruments, and

store them in the Packet stores • Receive TM Data Packets from the OBC, and store them in the Packet stores • Forward data from the Packet stores to the telemetry interfaces of the OBC TFGs (Transfer Frame

Generators VC1, VC2 and VC3) • Process telecommands from the OBC • Route TC packets from the OBC to the Payload Instruments • distribute SpaceWire time code to payloads • Route identified non-science packets from the Payload Instruments to the OBC • Generate SSMM internal telemetry, route it to the OBC and store it in Packet stores.

The SSMM maintains an internal reference time which is updated and synchronised to the Spacecraft Elapsed Time (SCET) of the OBC each second from the OBC SCET by the SpaceWire time code protocol. The reference time is maintained in the SSMM by 48 bit. The SSMM synchronises the instruments on SCET. The SSMM provides periodic housekeeping of the following form:

• Analogue Telemetry for Current Voltage and Temperature • General Status Telemetry for HW/SW status, EDAC status, EEPROM erease/write cycles • Status of storage and retrieval operations on the file for each created packet store:

o number of packets stored o generation time of the next packet to be retrieved o generation time of the last packet stored o position of the read and write pointers o status of storage and retrieval on the file (e.g. parallel bounded read operation on this file) o number of memory blocks allocated to the file o any other file pointers as provided by the design

• Mass Memory Array Status Telemetry o Partition/Module Power Configuration o Number of EDAC errors (correctable and uncorrectable) detected by the scrubbing function

per partition/memory module based on failed memory blocks o Uncorrectable detected failures o Address of uncorrectable errors.

14.1.1.3 Operational Modes

The operational modes of the SSMM are shown in the following figure.


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WDReset

OFF

BIT

INIT OPERA TEST

SERVICEConfig / Test

Memory

Config / TestComplete

OPERA

INIT

BIT complete or

time-out

Slave MCor

EEPROM CRC Error

BIT

PWR ON

TC

TC

Auto

OFF(Red. MC)

O/O Cmd

Reboot(Main MC)

Figure 14.1-3: Solid state Mass Memory operational modes

OFF mode: OFF mode is the mode in which the Memory Controller is after deactivation of the spacecraft primary power bus. BIT mode: In BIT (built-in test) mode, the SSMM Memory Controller and all input/output interface sections are tested. The SSMM enters BIT mode only from OFF or INIT mode. INIT mode: In INIT mode data storage is inhibited. The SSMM enter INIT mode when a Goto INIT TC is received . The file system is affected in INIT mode In INIT mode the SSMM executes and acknowledges telecommands from the OBC and nominal housekeeping TM data are not suspended. Modification (patch) of the SSMM application software and upload/download of Mass Memory Array contents via the test connector is possible. The transition from INIT mode to OPERA or BIT mode is performed on receipt of the appropriate telecommand. OPERAtional mode: The SSMM enter the operational mode on receipt of an telecommand when in INIT mode. In Operational mode the SSMM parameter tables and Mass Memory Array will be configured. The SSMM is fully operational and perform all routing, storage and retrieval functions and transmit data packets. The SSMM exit Operational mode on receipt of an INIT telecommand. Mass Memory Array TEST mode: In TEST mode, the Mass Memory Array will be configured and tested. The SSMM enters TEST mode only from OPERA mode. After the specified configuration and testing is complete, the SSMM return to operational mode autonomously. It is possible to exit TEST mode via the telecommand stopping any memory configuration. During TEST mode, the SSMM ongoing operations are not affected. Service Mode: The SSMM performs an autonomous transition to the SERVICE mode upon occurrence of a fatal error (e.g. non-correctable error in operational software program or data RAM, runtime error, watchdog error, etc.) during any mode. In this configuration both memory controllers are powered and active, one performing nominal operations and the other in Service mode. All functionality of the INIT mode are possible in Service mode with the exception, that all memory controller I/O functions are inhibited in order not to disturb the ongoing operation of the nominal memory controller. In Service mode the memory controller provides housekeeping telemetry and is time synchronised to the OBC.


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14.1.2 On Board Computer (OBC)

The Solar Orbiter Control and Data Management Unit (OBC) is the central control unit for all onboard data handling activities, Attitude and Orbit Control System (AOCS) and the management of the Payloads. Data management functions mainly consist of command distribution, telemetry acquisition and timing facilities during all phases of the mission. The OBC architecture is based on dual-redundant Mil-1553B buses, Command/Monitoring with SpaceWire interfaces for the RIU, FCE, SSMM and EGSE. Furthermore the OBC performs monitoring functions and depending on detected failures, provide safe system reconfiguration capabilities.

14.1.2.1 OBC Architecture

The OBC architecture is shown in the following figure.

Transfer Frames Generator

Ka-Band

Transfer Frames Generator

X-Band

Safe Guard Memory EEPROM

Bank 1PowerGuarded RAM EEPROM

Bank 2

TC Decoder and High Power Command Function

TC Decoder

Command Pulse Distribution Unit

CPDU I/F

= Cold Redundant

ProcessorModule

1

ProcessorModule

2Reconfiguration

Module

SCET

ReconfigurationModule

SCETHPCs:

some internally routedothers routed to other

equipments

Transponders receivers(X-Band )

Transponder emitters(Ka-Band )

EGSE

Transponder emitters(X-Band )

EGSE

Communicationlink

Communicationlink

2 RMs

Context Management

Subset of commands

generation request

Virtual Channel parameters and bitrate loading

Virtual Channel parameters and bitrate loading

Reconfigurationrequest

SSMMVC2

SSMMVC1

SSMMVC3

Alarms(Bus undervoltage , Battery

discharge , Battery low -voltage , Temerature)

SSMM,MPO-RIU,MTM-RIU,

Test

SpW

lin

ksSp

W l

inks

CLKsignals

MIL

Bus

C

MIL

Bus

B

MIL

Bus

A

VCO

Com

mun

icat

ion

link

DC /DC Converter

A

DC /DC Converter

B

Alarms(Bus undervoltage , Battery

dischare, Battery low -voltage , Temperature )

EGSE

Payloads emergency HPCs

CLCW

UARTTest I /F

UARTTest I /F

ProgramEEPROM 1

Program EEPROM 2

LCL FCL

LCL FCL

OCXO

OCXO

= FCL Powered

SpWMatrix

SpWMatrix

Com

mun

icat

ion

links

OBC MassMemory 1

OBC MassMemory 2

to PM2

to PM1

SSMM,MPO-RIU,MTM-RIU,

Test

Separation and

Deployment Switches

Separation and

Deployment Switches

Status I /F

Status I /F

Alarm I/F

Alarm I/F

Figure 14.1-4: OBC architecture

The OBC consists of the following redundant functional blocks and services and interfaces:

• ERC 32 Processor Module (Processor Module) • On-board time generation, synchronisation • MIL-STD-1553B buses • Transponder interface functions (TIF) • Switch status interfaces • Alarm Interfaces • OBC Mass Memory (OMM) • UART RS-422 link for interfacing to EGSE for access to processor control and S/W fast load/dump


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• SpaceWire links Modules • Reconfiguration Function • Safe Guard Memory • DC/ DC Converters

The following OBC functions are hot redundant: • TC Decoder function • OBC Mass Memory • Safe Guard Memory (SGM) • Reconfiguration Module • Spacecraft Elapsed Timer (SCET) • Oven-stabilised oscillators (OCXO) • DC/DC Converters associated with hot redundant functions.

The OBC modes are shown in the following figure.

OFF

Reset/ Init.

FCL power applied

Service

OBCPower Up

Ground Nominal

PM ONcommand

PM-Reset

PM isoperational

Build-in Test

Ground strappresent

Time out

Self test

Test complete

PM is selectedto service-mode

PM OFFcommand

Sef test

Test complete

Figure 14.1-5: OBC mode diagram

Off Mode: The OBC Off Mode is defined as when no power is applied to the OBC primary power inputs. The OBC Off Mode is encountered when the OBC HW is in storage, transport conditions or if the Power Conversion and Distribution Unit (PCDU) is not supplied by its own primary sources. Power Up Mode: The OBC shall enter the Power Up Mode upon connection of primary power via FCL to the corresponding primary power inputs of the OBC.


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Reset/Init Mode: The OBC enters Reset/Init Mode upon reception of a Processor Module Core power on command, detection of an error in the Processor Module Core or by reception of an external Processor Module Reset. A Processor Module Reset by the reception of CPDU commands is possible. This commanding shall only be possible by an arm and subsequent reset command. Nominal Mode: The Nominal Mode will automatically be entered if the PM is selected as operational at the end of the successfully passed Reset/Init Mode with the Central Software executing from RAM. In Nominal Mode the hot and cold redundant functions of the OBC shall be powered. All OBC internal parts including the I/O interface functions and processing resources of the Processor Module shall be fully available. Service Mode: The OBC enters PM Service Mode, when the alternate processor is already active. The Service Mode shall provide support for Central Software ground and in orbit maintenance (loading/dumping of RAM areas) and testing capability of the processor module functions. Ground Mode: In ground mode the OBC can be boot from UART interface.

14.1.3 Remote Interface Unit (RIU)

The RIU is used as a stand-alone unit in the Solar Orbiter DHS. The RIU represents the central interface unit of the satellite.

14.1.3.1 RIU Architecture

The architecture of the RIU is shown in the following figure.

Figure 14.1-6: RIU architecture diagram


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The RIU consists of the following functional blocks:

• A common bank of I/O modules, configurable in hot or cold redundancy to allow any combination of external units or I/O interfaces. Each of the modules have separate power supplies

o AOCS & STD I/F A, B, C o CPS I/O A,B (FCV Drivers and LV Drivers)

• Controller A, B • SpaceWire Interface A, B.

14.1.3.2 RIU External Interfaces

AOCS & STD Interfaces

• Monitoring/Control Thermistor (type YSI-44907) (ANY) or Fenwell 15K: 220 Thermistor acquisition interfaces of type YSI-44907 (A single point failure will not lead to a loss of more than 32 interfaces)

• Control Thermistors (PT-100) (ANP): 88 Thermistor acquisition interface of type PT100 (A single point failure will not lead to a loss of more than 4 interfaces)

• Analogue voltage (-5V to 5V) (AN1): 32 Analogue acquisition interfaces of type -5V to + 5V are provided (A single point failure will not lead to a loss of more than 4 interfaces)

• Analogue voltage (0 to 5V) (AN2): 20 Analogue acquisition interface of type 0V to + 5V are provided (A single point failure will not lead to a loss of more than 4 interfaces)

• Relay Status Acquisition (RSA): 140 Relay status acquisition interfaces are provided (A single point failure will not lead to a loss of more than 4 interfaces)

• Bi-Level Digital Telemetry Acquisition: 32 Bi-Level Digital acquisition interfaces are provided (A single point failure will not lead to a loss of more than 4 interfaces)

• Standard High Power (SHP) ON/OFF: 192 Standard high power interfaces are provided, 96 nominal and 96 redundant (cold redundant).

• FSA TBD • CSA TBD • CPS Pressure Transducer Acquisition (PTA) • CPSTemperature Acquisition NTC 15K (ANT).

Chemical Propulsion System (CPS)

The RIU provides the following elements interfaces to the CPS function:

• Latch Valve Control and Monitoring (LVC): 20 Latch valves Control and Monitoring interfaces are provided, 10 for main and 10 for redundant. Each CPS function provides one LVC OPEN and one LVC CLOSE pulse on separate lines for each of the latch valves. The CPS functions ensures that, in response to either a LVC OPEN command or a LVC CLOSE command, it is not possible to energise the latch valve coils continuously for more than 1 second

• Flow Control Valve Control and Monitoring (FCV): The CPS function provides the means by which to start each of the 8 (TBD) main and 8 (TBD) redundant Thruster FCVs from either control module in hot or cold redundancy.


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14.1.4 Failure Control Electronics (FCE)

The FCE has been introduced as a highest level (Level_4) FDIR instance in order to safeguard the Spacecraft sun-pointing in serious failures. In order to guarantee that Sun off-pointing angular constraints are met at all times, the FCE supervises the attitude of the satellite using:

• Heatshield-mounted Coarse Sun Sensor (CSS) internally triple redundant, which measures the deviation from the sun vector

• 3 Rate Measurement Units (RMU), which measure the rate of the satellite. The FCE monitoring function will be activated by ground command when the SC goes below 0.8 AU (TBD) Sun distance, when a pointing loss will endanger the SC. In normal condition although the FCE is activated, it will not be active unless a serious failure occurs which causes an unacceptable depointing. This mode where the FCE is active and where it takes over attitude control by thrusters is called the Survival Mode. One major requirement on the FCE is a simple design and high reliability with respect to unintentional thruster firing. The FCE must never disturb the attitude control performed by the OBC during nominal case, therefore all functions leading to a thruster activation within the FCE must be safe. It is however acceptable that in case of a failure inside the FCE the unit fails silent. This leads to the following design features:

• Simple architecture by two separate FCE branches. Both branches must detect a hazardous depointing and initiate attitude control by thrusters.

• Functionally decoupled from the OBC controlled AOCS Modes. • The thruster firing is controlled by four enables:

o Thruster interface module power ON/OFF o Alarm command from FCE branch A o Alarm command from FCE branch B o Thruster branch selection (enable/disable) by OBC/Ground HPC

• No software, simple control algorithm for each operational mode, located in a PROM (TBC). • No interrupt controlling • Safeguard memory with EDAC control (TBC). • Applying of Majority voting (2 out of 3) by triple redundant sensors. Majority voting is simple since

applied to alarm sugnals derived from threshold exceedings of individual measurements.


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Figure 14.1-7: FCE diagram


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15. ELECTRICAL POWER SUBSYSTEM

For a detailed description of the EPS, please refer to [IR51].

15.1 Requirements


Performance The power system shall provide all the power required by the SC throughout the various phases of the mission.

Summary sizing cases:

~1062W peak power at operational mode

(see budgets section for specific sizing cases)

The power subsystem must provide the power required throughout the mission

Thermal The solar array must be capable of generating adequate power down to the perihelion distance from the Sun of 0.2343 AU

The solar array design is TBD, pending the resolution of design issues on the BepiColombo mission, from which the array design shall be largely derived

Survivability The EPS (particularly the solar panels) must be able to tolerate the Solar Orbiter environment

The solar array design is TBD, pending the resolution of design issues on the BepiColombo mission, from which the array design shall be largely derived

Programmatic The EPS must reuse BepiColombo technology to the utmost extent

The PCDU, battery and SA for Solar Orbiter will be derived from BepiColombo

Table 15.1-1: EPS principal requirements


The Solar Orbiter EPS consists of the following elements:

Two solar array wings (each wing with two panels) double sided equipped with solar cells: o The hot side is equipped by a mixture of solar cells and OSR in order to cope with the near

sun environment. o The cold side is equipped only by solar cells, in order to provide sufficient power for far

distances from the sun. One Power Conditioning and Distribution Unit, including:

o Redundant solar array power regulation chain o Power Distribution Modules o Pyro Commanding Modules o Battery Charge and Discharge Regulators o Redundant Telecommand and Telemetry Interface via 1553 data bus o Redundant alarm signals o Battery Disconnect Device (BDD)

One Battery pack Two Solar Array Driving Mechanism (SADM) with relevant redundant Electronic (SADE) to operate

the Solar Arrays. The detailed architecture is shown in the following figure.


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Figure 15.2-1: Solar Orbiter power electrical architecture


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15.3 Power Control and Distribution Unit (PCDU)

The PCDU converts the power provided from all power sources (e.g. Solar Array, Battery, simulators) into a 28 V ± 0,5% fully regulated power bus for consumption by the various units and payloads on the spacecraft. The PCDU performs the following tasks:

Provides a reliable regulated 28V power bus conditioned from the energy sources Distribute the 28V bus to all power users via protected outlets:

o Scientific instruments o Spacecraft equipment o Distributes the 28V bus to the heaters via protected outlets o Provide sufficient status monitoring and command capability of operations and diagnostics

of the Electrical Power System. Provide firing commands to the onboard pyrotechnic (Thermal Knives for Solar Array) or NEA

devices via the necessary safety inhibits and monitors of the firing current. Provides Alarm signals to DMS Provides EGSE power interface and Launch support.

The PCDU regulation function is independent from other functions either within the PCDU or externally and operates in a conventional three domain mode setup: sunlight, charging and discharging.

• Domain 1: Sunlight Solar power > Bus load + charge demand

APR's are in linear mode BCR Battery charging Regulation

• Domain 2: Charging Solar power > Bus load, but less than bus load plus charge demand

APR's are in MPPT mode (BCR) • Domain 3: Discharging

Solar Power < Bus load BDR Battery Discharge Regulation.

The conventional three domain control system is based on one common, reliable Main Error Amplifier (MEA) signal. The MEA is fully redundant using triple majority voting circuitry. The output signal of the MEA is sent to the Array Power Regulator (APR), the Battery Discharge Regulator (BDR) and to the Battery Charge Regulator (BCR).


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15.3.1 APR and MPPT

The input voltage from the Solar Array is converted by a step-down converter to a regulated primary power bus voltage of 28V performed by a control loop with the capability to increase the APR input impedance to left SA surplus energy on the SA itself. The APR includes also the MPPT function which monitors the regulator input parameter (current and voltage) and controls the regulator to provide a specific input impedance that allows obtaining maximum power from the SA. The Maximum Power Point is obtained by oscillating the APR input impedance around the impedance providing the maximum power with an accuracy better than 1%. The MPPT function takes over the regulation control when the system power exceeds the maximum SA power availability.

Figure 15.3-1: Function of MPPT

The figures above depict the mode of operation and the principle circuit diagram of the tracker. In point P1 the S/H circuit 1 stores the array voltage signal Vmax -ΔV. The control voltage Vcontrol of the integrator ramps down to increase the loading of the array, Thus the array voltage decreases until the actual array voltage is equal to the stored voltage Vmax -ΔV. At this point P2 the comparator 1 sets the flip flop in the opposite position. The S/H circuit 2 stores the current signal Imax - ΔI. The control voltage Vcontrol of the integrator now ramps up to decrease the load of the array, the array current decreases until the acutal array current is equal to the stored current Imax - ΔI.

Figure 15.3-2: APR schematic block diagram

The BepiColombo design uses 8 identical APR modules to condition the SA power. Each APR module is able to provide 250W. The APRs operate in 7 out of 8 hot redundancy and can deliver up to 2000W to the power bus. The number of APR modules required for Solar Orbiter is TBD.


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15.3.2 Battery Charge Discharge Regulator (BCDR)

The battery charge/discharge electronics (BCDE) of the PCU is designed to operate the battery autonomously and majority voted in the nominal modes shown in the following table.

Mode Current Voltage Limits

Discharge 0 ≤ Idischarge ≤ 40 A VBus = 28 V 18 V ≤ VBat ≤ 25,2 V

Full Charge Level 1 Icharge = 50 A Level 2 Icharge = 70 A

VBus = 28 V 15 V ≤ VEOC ≤ 25,2 V

Taper charge 0 ≤ Icharge ≤ 30 A VBus = 28 V 15 V ≤ VEOC ≤ 25,2 V

Current Limited Charge Icharge ≤ 0,6 A VBus = 28 V 15 V ≤ VEOC ≤ 25,2 V

Table 15.3-1: BCDE nominal modes and characteristics

The BCDRs are implemented to interface the 28V power bus with the battery. The BCDR module consists of two power converters, the Battery Charge Regulator (BCR) and the Battery Discharge Regulator (BDR) realized on the same module. Since the BCDR function is designed for a battery voltage lower than the regulated main bus voltage, the BCR (upper DC/DC converter) is a step-down regulator, whilst the BDR (lower DC/DC converter) is a conventional step-up regulator.

Figure 15.3-3: BCDR schematic block diagram

Each BCDR module is sized to provide up to 300 W maximum during battery discharge and up to 9 A during battery charge. It is possible to switch ON and OFF independently the charge and the discharge functions. Since the Lithium Ion battery is used, the BCR implement charging starting with current limited charge that continues unit the End of Charge (EOC) voltage is reached, followed by a constant voltage charge (taper charge) to the EOC voltage limit. The battery EOC voltage level selection is available by 8 commandable steps, graduated by equal increments of 0.12 V. In the BepiColombo design, the BCDR's operate as 6 out of 7 hot redundant, the battery discharge function is sized to 1800 W maximum. The battery charge function is sized to 54 A (6 x 9A). For Solar Orbiter 4 out of 5 BCDRs would be sufficient to cope with the power demand, but no change to the BepiColombo design is required but removal of exceeding number of BCDR's may be envisaged for mass saving.

15.3.3 Battery Disconnect Device (BDD)


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Figure 15.3-4: Battery disconnection device

The Battery Disconnection Device performs the battery disconnection in order to obtain a complete SC OFF during battery connection. The Battery Disconnection Device is based on the usage of two relays, and the battery disconnection can only be performed via Standard High Power commands from EGSE. The Battery connection can be performed either via SHP from Ground Interface or via 1553 data link from DMS interface. Relay status is provided to the OBC via 1553 data link and to the EGSE via RSA discrete lines.

15.3.4 Equipment Power Distribution Module (LCL/FCL)

Figure 15.3-5: LCL module block schematic

Each module provides 16 switched and protected output lines from the main bus to the spacecraft and payload equipment. These lines are protected by Latching Current Limiters (LCL) or Foldback Current Limiter (FCL). The selection between LCL or FCL function and the current limitation classes must be selected before module integration. In order to fulfill the demand of 52 nominal and redundant LCL/FCL, seven Power Distribution Modules are provided.


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15.3.5 Heater Power Distribution Module

The PCDU heater power distribution module contains 24 switchable outputs with three upstream LCLs organized in three groups were one LCL is followed by eight switches. The power is provided to the heaters via outlet group which are LCL protected. Each outlet is switchable individually via dedicated TSW (Transistor Switches). The TSW are arranged in group of eight and they are protected by two types of LCL of either 3A or 5A maximum. All TSW and all LCLs are commandable ON/OFF via 1553 data bus.

LCL TSW

Heater

Heater

TSW

TSW Heater

TP

TP

TP

Figure 15.3-6: PCDU TSW arrangement

The downstream heater switch is implemented by a Mosfet with ON/OFF control by telecommand and switch status telemetry. A single failure can cause a lost of a group of heaters but a downstream switch failure cannot cause a heater being permanently supplied since the protection LCL can be commanded OFF.

Figure 15.3-7: Heaters/Thermostats configuration

The PCDU provides an additional output power line for each heater power outlet in order to allow thermostatic protection.


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15.3.6 Pyro Firing Module

The Pyro Firing Module consists of three main functions, an arming switch function, a DC/DC conversion stage for generation of the regulated current and a switch matrix for the selection of the Pyro device to be fired. Three individual commands have to be executed in the Pryo module to allow pyro firing:

The arming command will turn on the battery line input to the DC/DC conversion stage Pyro selection command will connect the desired pyro device to the current generator Pyro firing command will initiate the generation of pyro fire current.

All switching functions are implemented as solid state functions. The current regulation stage provides a regulated current of 5 A into the selected pyro device. The duration is limited by a timer to 50 ms. The current generation will terminate when the pyro firing occur or when the timer period is elapsed. The PCDU allows actuation of three initiator types:

NSI-type igniter of 1 A 1 W Shape Memory Allow or equivalent heating device (thermal knife) Low Shock Release Unit (LSRU) either with NSI-type equivalent electrical activation, or alternative

with constant current of 400 mA over a time of 100 … 500 ms.

Figure 15.3-8: Pyro firing module block schematic

In order to provide up to 48 main and 48 redundant Pyrotechnics actuator, six Pyro Firing Modules are installed inside the PCDU. The BepiColombo Solar Array release system is expected to be a Thermal Knife system. Therefore the pyro module will be updated to be compatible with such a device. For Solar Orbiter the use of same actuators for the SA is assumed. Solar Orbiter will require 16 (8 nominal / 8 redundant) thermal knife drivers for the SA. The release mechanisms for the HGMA and MGMA are also Thermal Knife systems and the PCDU provides the required interfaces.

15.3.7 Command & Monitoring Module

The PCDU contains two command and monitoring modules providing hot redundant (i.e. both are powered and active) access to the PCDU telecommand and telemetry. The external interface is based on the dual 1553 Remote Terminal interface.

15.3.8 PCDU Layout and Mass Breakdown


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Figure 15.3-9: PCDU layout

Description Quant. Module Mass [g] Mass [g] Array Power Regulator 8 515 4120 Battery Charge/Discharge 7 553 3871 Equipment Power Distribution 7 525 3675 Pryo Firing 6 500 3000 Heater Power Distribution 4 450 1800 Command & Monitoring 2 340 680 Back Plane 1 2125 2125 Housing 1 5312,5 5312,5 Unit Total 24583,5

Table 15.3-2: Mass breakdown


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15.4 Battery

The Battery has been sized in order to provide power during the following two phases:

LEOP phase of 120 minutes with the following contributions: o 10 minutes 75 Watt o 1 hour 484 Watt o 1 hour 525 Watt

This leads to an energy demand of 75/6 + 484 + 528 = 1024,5 Wh Considering a 15 min Solar Array flipping and a power demand of 1138 Watt during this time, the

energy demand is 285 Wh. Considering both phases, a namplate energy of 1160 Wh BOL is considered sufficient, taking also into account that one string fails. The selected cell for the battery is the 18650HC lithium-ion cells.

Figure 15.4-1: 18650HC cell

For BepiColombo a cell configuration 6s62p is used in order to achieve 2009 Wh. This battery is specially created for the BepiColombo mission and maintains design maturity through similarity to previous proven batteries. Downscaling of the BepiColombo design to Solar Orbiter battery design would results in a configuration of 6s36p.

15.5 Solar Array Configuration

The use of the BepiColombo Solar Array technology / design for the Solar Orbiter environment will need to be proven by a future TDA. Ther selection of a different technology may even be required. The Solar Array thermal analysis and sizing analysis is provided in SOL.F.ASTR.TN.00001. In the following, the Solar Orbiter SA description is based on BepiColombo design. The solar array consists of 2 wings located on the -Y and +Y side of the spacecraft. Each solar array wing consists of 2 panels. Each panel is double sided covered by photo voltaic cells.

The "cold" side of the SA consists of 100% solar cells. The "hot" side of the SA will consist of 30 to 40% solar cells and the rest will be optical reflectors

(OSR).


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The Solar Array is designed for a high temperature environment and high temperature range of -150 oC to + 230 oC (TBC).

Figure 15.5-1: Solar Orbiter solar array in deployed condition

The main Solar Array characteristics are: Four double sided populated panels (two panels for each wing). Each wing consists of a yoke, including root hinge between the inner panels of the SADM of 1300

mm length (TBC). Each wing provides flight proven hinges between the two panels and between the inner panel and

the yoke Each wing has 4 fight proven hold-downs with Thermal Knife devices. Micro-switches for detection of successful deployment Shielded cable duct at two panel sides and shields at the other panel edges for thermal protection. Magnetic compensated strings. Temperature sensors on each panel.

UNIT DATA Solar Array Wings per spacecraft

Operating Voltage Panels / Wing Sections / Panel Strings in parallel/section Solar cells in series/string

2 with double sided photo voltaic assembly Vmp by MPPT 33V < Vmp < 160 V 2 1 19 (TBC) 46 (TBC)

Panel substrate Face skin material Core material Isolation

Carbon fibre reinforced cyanate Carbon fibre reinforced cyanate honycomb core Kapton HN

Solar Cells Designation Type Size Total quantity /array

Azur 3G28 with AIO AR coating Triple junction GaInP2/GaAs/Ge 40 mm x 40 mm x 150 μm with cropped corner TBD

Bypass diode External Diode in cropped corner (Preliminary developed under ESA contract)

Coverglass Material Coating Size

Qioptiq ceria doped coverglass CMX 100 AR 40 mm x 40 mm x 100 μm with cropped corner

Optical Solar Reflector (OSR)

Material Coating Size

CMG 150 Ag/Nichrome 17 mm x 40 mm x 150 μm

Blocking Diode Silcon Carbide diode Actually in pre development under BepiColombo program

Table 15.5-1: Solar Orbiter solar array major design data


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15.5.1 PVA Design Concept

Extensive investigation on the Photo Voltaic Assembly (PVA) have been performed by the BepiColombo project. The cells considered as design baseline for BepiColombo are 3G28 cells, manufactured by Azur Space. They are Triple Junction cells (GaInP2/GaAs/Ge on Ge substrate) with 28% efficiency (at 20 degC). Since the presence of TiOx antireflective coating in combination with Silicone-based cover glass adhesive provides performance degradation during high temperature and UV illumination, a new AlOx AR antirefective coating has been developed and is currently under test. In the frame of BepiColombo, a pre qualification activity (UV-HT test) was defined to demonstrate the possibility of a qualification for 200°C with some time limited excursion to 230°C. Table below shows the characteristic of standard 3G28 cells with TiOx AR coating at room temperature.

Item GaInP2/GaAs/Ge Solar Cell with Al2O3 AR Manufacturer AZUR SPACE Solar Power GmbH Germany Type GaInP2/GaAs/Ge with Al2O3 AR Designation/Type BC TJ cell Junction Triple Junction AR Coating Al2O3 Contact System front side Au/Ag/AuGe weldable Contact System rear side AuGe/Ag/Au weldable Cell Dimension 40,0 {mm] x 40.0 [mm] One Cropped corner Yes (13,5 [mm] x 13,5 [mm] /2) Cell Area 15,088 cm2 Cell Thickness 150 μm Bypass Diode Electrical SCA Data Isc 0,239 [A] Imp 0,225 [A] Vmp 2,360 [V] Ioc 2,667 [V]

Table 15.5-2: 3G28 with TiOx AR coating electrical characteristic

A backup option using Dual junction cells (GaInP2/GaAs on GaAs substrate) is now under development at Azur Space and preliminary evaluation is currently performed on samples by EADS Ottobrunn under BepiColombo contract. No further activities for Dual junction cells are foreseen under the BepiColombo project.

15.5.2 Solar Array Power Interface

In the present design, the solar array area is 8.4 m2, which means 2.1 m2 for each of the four panels (TBC). The packing factor on both sides of the SA panel is 85% of the panel area. The electrical losses are 9%. The individual array strings of both sides of the panel are combined at panel level. The panel wiring is combined to form wing level wiring at the Solar array drive Mechanism and transferred to the PCDU via two redundant power harness to separate input power connectors on the PCDU. Both solar array wings can be independently controlled. The MPPT is sensitive to differences of generated power of the solar cell strings, which provide more than one power maximum. This could require to split the 8 APRs in two parts, 4 APRs for each wing. In order to keep the PCDU commonality with BepiColombo, matched solar array wings with respect to size, temperature and attitude are required. A voltage drop of 3V from the PCDU input to the main bus regulation is acceptable for achieving full regulation performance. Therefore with a 28 V main bus, the minimum operating array voltage at the PCU APR input must not be less than 31 V. The minimum power requested at the PCDU output is specfied by:


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818 W for a spacecraft-sun distance of 1.5 AU (600 W/m² solar flux) 1138 W for a spacecraft-sun distance from 0.5 to 0.22AU (27900 W/m² solar flux) The PCDU efficiency is better than 93%.

*note that the current power demand, as reported in the budgets section, is 743W – we take a 10% margin. The maximum Power transferred from the solar array can be 2000 W. Assuming a SA voltage of 35 V results in a maximum array current of 28 A per wing. Individual monitoring cells per SA panel are used to measure the Open-circuit voltage (Voc) and the short circuit current (Isc) versus solar cell temperature. The Voc and Isc measurement accuracy is better than ± 1%. The measurements are provided in the EPS telemetry data.

15.5.3 Solar Array Drive Assembly (SADA)

The SADA comprises:

SADM Solar Array Drive Mechanism SADE Solar Array Drive Electronics

Each solar array wing has its own SADA.

Figure 15.5-2: SADM conceptual block diagram

Solar Array Drive Mechanism (SADM) includes the following major assemblies:

Housing structure from Aluminium Alloy, Titanium or Steel according to the BepiColombo design. The motor gear unit consisting of a stepper motor and a planetary gear stage which is connected

with its output shaft to the main deployment hinge. Redundant position transducer (potentiometer) Redundant electrical power and signal harness and connectors Thermistor for temperature reading Electrical power and signal transmission between fixed and rotary part by means of slip rings.


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The Solar Array Drive Electronics include the following functions:

Motor driver Acquisition of position sensor and reference indicator signals Primary power bus interface Telecommand/Telemetry interface (MIL 1553 bus) Discrete I/Fs (e.g. SHP, RSA).

The Solar Array Drive Mechanism allows continuous rotation around the Y-axis. For Solar Orbiter only stepwise rotation is foreseen. The SADM can operate at high temperature (up to 150 oC) induced by the sum of the following effects:

Coupling to the hot solar array Heating by ohmic losses in the high power electrical transfer lines.

15.5.4 SADA Operational Modes

The SADA supports the following operational modes: Stand-by Mode Hold Mode Drive Mode

o Slow Mode o Fast Mode.

In Stand-by mode, the SADE supports only the monitoring of the SADM parameters. In Hold mode, the SADE command the SADM motors that no relative motion of the solar array wing with respect to Spacecraft body will occur. Slow Mode During this mode, the SADE controls the solar array orientation in normal operations for BepiColombo. The solar array follows the required sun alignment provided by the AOCS with minimal impact on the spacecraft RPE, with a total Pointing Random Error of < 720 arcsec by:

Maximum speed 0,5 o/sec Acceleration 0,002 o/sec2 Jerk 0,0005o/sec3

Fast Mode The purpose of this mode is to allow the solar array to follow rapidly alignment profile coming from the AOCS. In the fast mode the SADM is able to rotate the solar Array by an angular rate of 6 o/sec – this provides a full pi rotation of the array in 30 seconds – well within the assumed 15 minutes of battery power provision for array-flipping assumed in the battery sizing.

15.5.5 SADA Electrical Interfaces

Each SADM porives the following connections with the Solar Array: 18 AWG#22 TP cables to connect the SA with the power tracks 4 AWG#22 TP cables to connect the SA with the engineering string tracks 24 AWG#22 TP cables to connect the SA with the signal tracks 2 AWG22 wires to connect the SA with the ground tracks

The SADM interfaces to the Spacecraft electrical harness by connectors located on the SADM and with the solar array by pig tails.


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15.6 External Power Interfaces

15.6.1 LCL/FCL Interfaces

LCL ID

LCL class

Nom I(A)

Trip I (A)

Load (W) Unit / Instrument

Instruments 1 Inst_EPD_NOM 2 Inst_EUI_NOM 3 Inst_MAG_NOM 4 Inst_METIS_NOM 5 Inst_PHI_NOM 6 Inst_RPW_NOM 7 Inst_SoloHI_NOM 8 Inst_SPICE_NOM 9 Inst_STIX_NOM 10 Inst_SWA_NOM DHS 11 DHS_OBC_PROC_NOM 12 (FCL) DHS_OBC_DEC_NOM 13 DHS_SSMM_CONTROL_NOM 14 DHS_SSMM_MEM_IO_NOM 15 DHS_RIU_CONTROLLER_NOM 16 DHS_RIU_STDIO_1 17 DHS_RIU_STDIO_2 18 DHS_RIU_FCV_NOM 19 DHS_RIU_LV_NOM 20 DHS_FCE_CONTROL_NOM 21 DHS_FCE_FCV_NOM COMMS 22 TTC_DST1_X_KA_TRANSM 23 (FCL) TTC_DST1_RECEIVER 24 TTC_TWTA_X_1 25 TTC_TWTA_Ka_1 26 TTC_APME_NOM AOCS 27 AOCS_IMU_NOM 28 AOCS_STR1 29 AOCS_RMU1 30 AOCS_RMU2 31 AOCS_WDE1 32 AOCS_WDE2 EPS 33 EPS_SADE_1_NOM 34 EPS_SADE_2_NOM CBS 35 CBS PRESURE_TRANS_1 LCL ID

LCL class

Nom I(A)

Trip I (A)


Instruments 1 Inst_EPD_RED 2 Inst_EUI_RED 3 Inst_MAG_RED 4 Inst_METIS_RED


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LCL ID

LCL class

Nom I(A)

Trip I (A)


5 Inst_PHI_RED 6 Inst_RPW_RED 7 Inst_SoloHI_RD 8 Inst_SPICE_RD 9 Inst_STIX_RED 10 Inst_SWA_RED DHS 11 DHS_OBC_PROC_RED 12 (FCL) DHS_OBC_DEC_RED 13 DHS_SSMM_CONTROL_RED 14 DHS_SSMM_MEM_IO_RED 15 DHS_RIU_CONTROLLER_RED 16 DHS_RIU_STDIO_3 17 DHS_RIU_FCV_RED 18 DHS_RIU_LV_RED 19 DHS_FCE_CONTROL_RED 20 DHS_FCE_FCV_RED COMMS 21 TTC_DST2_X_KA_TRANSM 22 (FCL) TTC_DST2_RECEIVER 23 TTC_TWTA_X_2 24 TTC_TWTA_Ka_2 25 TTC_APME_RED AOCS 26 AOCS_IMU_RED 27 AOCS_STR2 28 AOCS_STR3 29 AOCS_RMU3 30 AOCS_WDE3 31 AOCS_WDE4 EPS 34 EPS_SADE_1_RED 35 EPS_SADE_2_RED CBS 36 CBS PRESURE_TRANS_2

Table 15.6-1: Solar Orbiter LCL/FCL budget

15.6.2 Pyro / Thermal Knife Interfaces

Pyro (Non) ID Unit /Instrument Pyro (Red)

ID Unit /Instrument

1 EPS_SA1_1_NOM 1 EPS_SA1_1_RED 2 EPS_SA1_2_NOM 2 EPS_SA1_2_RED 3 EPS_SA1_3_NOM 3 EPS_SA1_3_RED 4 EPS_SA1_4_NOM 4 EPS_SA1_4_RED 5 EPS_SA2_1_NOM 5 EPS_SA2_1_RED 6 EPS_SA2_2_NOM 6 EPS_SA2_2_RED 7 EPS_SA2_3_NOM 7 EPS_SA2_3_RED 8 EPS_SA2_4_NOM 8 EPS_SA2_4_RED 9 COMMS_HGA_1_NOM 9 COMMS_HGA_1_RED 10 COMMS_HGA_2_NOM 10 COMMS_HGA_2_RED 11 COMMS_HGA_3_NOM 11 COMMS_HGA_3_RED 12 COMMS_MGA_1_NOM

12 COMMS_MGA_1_RED


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13 COMMS_MGA_2_NOM 13 COMMS_MGA_2_RED 14 CPS_ All-Fire Current 14 CPS_ All-Fire Current

Table 15.6-2: Solar Orbiter pyro and thermal knife interfaces

15.6.3 Heater Interfaces

48 transistor switches, ~200 heaters in total, allocation TBD.


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16. COMMUNICATION SUBSYSTEM

For a detailed description of the TT&C subsystem, please refer to [IR52].

16.1 Requirements


Performance Provide a reference TM data-rate of 150 kbps @ 1 AU.

This has been reduced from 200 kbps, allowing a reuse of the BepiColombo reflector size (1.1m).

Functional 2-way (TM/TC) communication with the ground TC X TM X/Ka

This requirement specifies the use of both X and Ka bands.

Functional Communication coverage for all mission phases for all attitudes of the SC (except planetary alignment outages)

This requirement is impossible to meet as worded: outages shall be caused by HTHGA shadowing below 0.28 AU, and by the limitations of HTHGA/MGA coverage when combined with Sun-pointing requirements. LGA coverage is also not omni-directional, and a spin-separation must be used during LEOP.

Functional On board storage to deal with changing communication conditions

Solar Orbiter experiences a high variance in Earth distances, conjunctions, dish shadowing etc, and therefore a store and forward data handing system is necessary

Table 16.1-1: TT&C subsystem principal requirements


The Solar Orbiter telecommunication subsystem is a complete reuse, from an RF-perspective, of the TT&C subsystem developed for BepiColombo. It consists of a redundant set of transponders using X-Band for the uplink, and X-Band and Ka-Band for the downlink. Depending on the mission phases, the transponders can be routed via RF switches to different antennas. The telecommunication subsystem provides hot redundancy for the receiving function and cold redundancy for the transmitting function. The baseline communications subsystem architecture is shown in the following figure. It is highly recurrent from the BepiColombo one and is composed of the following elements:

• The nominal and the redundant X/X/Ka-Transponders (up- and down-link in X-band and only down link in Ka-band); both downlinks can be active simultaneously

• One fully redundant X-TWTA, 35 W RF with 55% efficiency

• One fully redundant Ka-TWTA, 35 W RF with 50% efficiency. The option of having a non redundant Ka band TWTA to save a few kg, once envisaged on BepiColombo, is not retained in the baseline

• A two-axes steerable X & Ka bands 1.1m Reflector High Gain Antenna (HGA), with its Antenna Pointing Mechanism (APM) & associated waveguide rotary joints. It is derived from the Bepi-Colombo Antenna

• Two X-band LGAs with hemispherical coverage each (0 dBi) for near Earth TM transmission (up to ca 0.54 AU distance)

• One X-band 1-axis articulated MGA (type TBD) to provide a contingency communication capability in case of HGA failure

• One X-Band & one Ka-Band 3 dB Hybrid Couplers

• A Radio Frequency Distribution Unit / WaveGuide Interface (RFDU/WUI) which includes five X-Band & one Ka-Band waveguide switches, two X-Band diplexers, two X-Band & two Ka-Band waveguide isolators and an Antenna Pointing Mechanism Electronics (APME), controlling the HGA & MGA pointing mechanisms; the RFDU elements will be recurring from Bepi Colombo. Circulators protect the TWTA output amplifiers from reflections.

There are some key mechanical differences between the Solar Orbiter and BepiColombo designs:


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• The HTHGA reflector is mounted on a long boom • The requirements placed on the HTHGA and MGA APM are substantially different:

o The elevation ranges of the HTHGA and MGA APM are different for Solar Orbiter o Both the HTHGA and MGA mechanisms are permanently shadowed by the heatshield in

the case of Solar Orbiter o The MGA mechanism for Solar Orbiter requires only 1-axis.

Nevertheless it is recommended to reuse the BepiColombo mechanisms to ensure high commonality of the electrical and RF-interfaces for the TT&C subsystem. The detailed link budgets for the Solar Orbiter TT&C are presented in [IR19], with summary tables also presented in the budgets section of this document.

Figure 16.2-1: Solar Orbiter TT&C subsystem architecture


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16.2.1 Deep Space Transponder

The Solar Orbiter DST is identical to the BepiColombo DST. he Deep Space Transponders (Nominal and Redundant) are the main units within the Communication Subsystem. Each DST consists of:

• X-Band Receiver, both are hot redundant. • X-Band Transmitter, both are cold redundant • Ka-Band Transmitter, both are cold redundant.

ECSS-E-50-05 Standard for Deep Space operations category B spacecraft, these are:

• X-Band Reception (Earth to Space): 7145 - 7190 MHz • X-Band Transmission (Space to Earth) 8400 - 8450 MHz • Ka-Band Transmission (Space to Earth) 31800 - 32300 MHz.

16.2.2 X/Ka Band TWTA

The X/Ka-Band TWTA's are identical to those used for BepiColombo. The functionality of the X-TWTAs is to amplify the X/Ka Band Transponder Downlink signal up to the required power and send it to the antennas via the Radio Frequency Distribution Assembly (RFDA). Each TWTA consists of the following three items:

• Electrical Power Conditioner (EPC) • Travelling Wave Tube (TWT) • High Voltage Harness.


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Figure 16.2-2: X/Ka band TWTA architecture

2 X-Band TWTAs and 2 Ka-Band TWTAs (nominal + redundant) are provided. The TWTAs operate in cold redundancy, and are cross coupled such that:

• Each of the X-Band TWTA input is coupled via a 3-dB coupler to both DST X-Band Transmitter

outputs • Each of the Ka-Band TWTA input is coupled via a 3-dB coupler to both DST Ka-Band Transmitter

outputs. The TWTA features the following operational modes:

• OFF Mode • STAND-BY Mode (EPC ON and limited to low-voltage filament heating) • OPERATIONAL Mode (High Voltage ON) • OPERATIONAL Mode Low Power.

The TWTA amplifies the input signal up to a power of 35W (45,5 dBm). The output power of each TWTA can be reduced by 20dB, via a dedicated telecommand. 2 different operational modes are possible:


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• Full Power Mode • Reduced Power Mode (by 20 dB).

In both operational modes the input power to the TWTA shall be in the range -40 ÷ 0 dBm.

16.2.3 RFDU

The RDFU (Radio Frequency Distribution Unit) consists of the following items:

• X-Band o X-Band waveguides (WR112) o 5 Waveguide switches o 1 Diplexer to separate up/down-link o Circulators o 1 3-dB coupler

• Ka-Band o Ka-Band waveguide (WR 28) o 1 Waveguide switch o 2 Circulators o 1 Bandpass Filter o 1 3 dB coupler.

The RFDA is designed to route the frequencies selected for Solar Orbiter mission as follows:

Table 16.2-1: Solar Orbiter frequency and bandwidth allocation

The RFDU provides the functionality to connect the various components of the TT&C subsystem according to the needs of the mission, in summary:

• Connecting the Ka-Band Transmitter to the HGA Antenna • Connecting the X-Band Transmitter via TWTA to the selected antenna (HGA, MGA or LGAs) • Connecting the X-Band Receivers via TWTA to the selected antenna (HGA, MGA or LGAs).

The downlink connection between X-Band DST Transmitter output to the TWTA input and the uplink between WGS-5 to DST Receiver are performed by coax cables. All other RF connections are performed by waveguides.


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Figure 16.2-3: RFDU schematic

16.2.4 Antenna Assembly

The Solar Orbiter Antenna Assembly consists of 4 Antennas:

• 2 Low Gain Antenna (LGA) • Medium Gain Antenna (MGA) • 1 High Gain Antenna (HGA).

The bandwidth of the antennas is given by the frequency range:

• X-Band Reception (LGA, MGA, HGA): 7145 - 7190 MHz • X-Band Transmission (LGA, MGA, HGA) 8400 - 8450 MHz • Ka-Band Transmission (HGA) 31800 - 32300 MHz.

All antennas are RHC polarized.

16.2.4.1 LGA

The LGA's are identical to those used on BepiColombo. They are accommodated on opposite sides of the Solar Orbiter Spacecraft body in order to provide a nearly omnidirectional coverage for the X-Band up and


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downlink. The coverage is unavoidably disturbed by the heat shield which prevents an omnidirectional coverage by ~10 degrees (this is the effect due to geometrical blockage – diffraction effects have not been considered, although it should be appreciated that any diffraction of the signal will serve to reduce the blocked region). Each antenna provides a uniform gain within the bandwidth taking into account the uplink and downlink bandwidth plus Doppler shift. Each LGA comprises:

• Helix radiating element • Radom (protection against thermal stress and mechanical damage) • Support structure (mounting bracket).

The functionality of the LGA's is to allow transmission and reception of X-Band TM/TC during LEOP and NECP.

16.2.4.2 MGA

The MGA is used as the contingency antenna during the mission once the SC is away from the near-Earth region. The MGMA (Medium Gain Antenna Mechnical Assembly) consists of the following parts:

• RF components (feed, reflectors and waveguides) • Antenna boom • Three Hold Down and Release Mechanisms (HDRM) • APA (Antenna Pointing Assembly) consisting: • pointing mechanism (APM) for one axis located at the spacecraft interface • APME (Antenna Pointing Mechanism Electronics) • Single-axis joint for X-Band.

The feed and the reflector are identical to BepiColombo. All other components are TBD.

Figure 16.2-4: BepiColombo MGA schematic

16.2.4.3 HTHGA

The High Gain Antenna is the main antenna of the Solar Orbiter and it is for the high datarate downlink in X and/or Ka Band during the science phase. The High Gain Antenna Major Assembly (HGAMA) consists of:

• ARA (Antenna Reflector Assembly o Antenna Reflector Dish with diameter of 1100 mm o Sub reflector with diameter of 148 mm


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o Struts • RFC (Radio Frequency Chain)

o Dual Band Antenna Feed horn including diplexing OMT's and polarizer o Waveguide runs including all transitions and flanges

• APA (Antenna Pointing Assembly) o APM Antenna Pointing Mechanism

Two-axial motor for rotation of the Antenna reflector in Azimuth and Elevation. Antenna boom and Flanges Rotary Joint for two axis and two frequencies

o APME (Antenna Pointing Mechanism Electronics) which drives the motor.

Figure 16.2-5: HTHGA dual-channel rotary joint

For Solar Orbiter the APM is located at the foot point of the antenna boom in the shadow of the heat shield. The ARA Main Reflector, Sub Reflector, Struts, Feeder and the rotary joints will be most likely identical to BepiColombo. All other components are TBD. The rotary joint figure above is a preliminary design for BepiColombo and is suitable too for Solar Orbiter. Two orthogonal arranged rotary joint with the associated motors provide Elevation and Azimuth actuation. The rotary joint has an 'L' configuration for the X-Band channel and a 'U' configuration for Ka-Band. The sequence of the joint (rotation/elevation from SC to HGA) has been selected on Solar Orbiter in order to avoid ‘pointing shocks’ (spikes in the required elevation and rotation rates) due to mechanism constraints in achieving the Earth position, which would be present were the order to mechanisms reversed. The main difference between the BepiColombo and Solar Orbiter HTHGA is the Solar Orbiter ARA is located at the top of a boom (1220mm TBC from the APM to the centre of the dish). The boom is necessary to provide the antenna with a better FoV of the Earth, which undergoes a large variation in position in the PRF of the SC throughout the mission. Even with this position away from the SC body, and with 2-DOF, the antenna is unable to track the Earth throughout the mission without imposing periodic rolls on the SC attitude (this is presented in the Mission Operations section of this document), due to occultation of the HTGHA FoV by SC appendages and the body of the SC. The HTHGA is placed on the –Z panel due to a slight advantage in accessing the Earth from this position throughout the mission.


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Figure 16.2-6: [left] HTHGA in nominal position [right] HTHGA in close-approach shielded position

Figure 16.2-7: Isometric view of the Earth position in the SC PRF throughout the mission for the former 2015 launch mission scenario: the yellow sphere represents the Sun position, with the SC

XYZ coordinates shown as the red axes-set


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Solar Orbiter 2017 launch

-60.00

-50.00

-40.00

-30.00

-20.00

-10.00

0.00

10.00

20.00

30.00

40.00

50.00

60.00

-180

-160

-140

-120

-100 -8

0

-60

-40

-20 0 20 40 60 80 100

120

140

160

180

Azimuth in the XY plane (0° = towards +X in Sun direction, +ve towards +Y) in deg

elev

atio

n w

.r.t.

XY p

lane

in d

eg (0

° = in

XY

plan

e, +

ve to

war

ds +

Z)

+Y Solar Array -Y Solar Array

+Y RPW Antenna -Y RPW Antenna

SC Body

Figure 16.2-8: HTHGA interference regions of the sky for the baseline (2 RPW antenna on –Z SC

side: blue – CP; cyan – NMP; green – EMP)

Figure 16.2-9: HTHGA elevation rate for the former APM order (elevation/rotation from SC to HGA) throughout the course of the former 2015 launch mission, showing rate-spikes due to mechanism

constraints


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17. HARNESS SUBSYSTEM

Detailed definition of the harness design is not commensurate with a Phase B1 study. However, a top-down definition of the harness has been conducted in order to provide confidence in the SC design, and allow the impact of the harness to be properly accounted for in the SC design. The impact of the harness on the SC design is two-fold (volumetric and mass accommodation). Both of these aspects are described in the following sections.

17.1 Harness Volume Accommodation

The volumetric accommodation of the harness has proceeded by considering the primary harness routings required. The requirement for connection between the various units and payloads has been analysed, allowing the primary harness routing volumes to be defined; these are shown in the following figure, for the payload. Power, TT&C and miscellaneous harnesses. The same procedure has been followed for the propulsion subsystem pipework, but this is not shown here. The harness and pipework routing exercise has alllowed the impact of these ‘soft’ subsystems on the likely dimensions of the SC, as well as unit accommodation, to be properly taken into account at this early stage.

Figure 17.1-1: Harness accommodation


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17.2 Harness Mass

The harness mass is typically a significant percentage of the dry mass of the SC (typically between 6 and 10% of the dry mass). The allocation of harness mass for Solar Orbiter is 6.2% of the dry mass of the SC, which is identical to the harness mass as dry mass percentage for the Mars Express SC, which has a similar number of instruments (8), and a similar structure and size. The following table summarises a short review of the harness mass for recent missions.

Bepi-Colombo MPO 8.32% source: BepiColombo mass budget 20080331 Bepi-Colombo MTM 3.49% source: BepiColombo mass budget 20080331 Mars Express 6.3% source: mass budget correspondance from Steve Kemble Venus Express 5.6% source: mass budget correspondance from Steve Kemble Aeolus 7.83% source: mass budget correspondance from John Brewster

Table 17.2-1: Harness mass as percentage of SC dry mass for various missions

Accordingly we took the average of the BepiColombo MPO, MEX, VEX and Aeolus, which comes at 7.0%. This number will be revised as the maturity of the design progresses.


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18. SOLAR ORBITER BUDGETS

18.1 Introduction

This section provides an abbreviated set of the budgets that are currently tracked for the Solar Orbiter SC. These are the following:

• Mass (inc. propellant loading) • RF-link • Power • Delta V and Attitude Control • Pointing

o APE o RPE o RS-instrument coalignment

• Science Window definition • Mass Memory budget and profile.

For the complete set of budgets, including explicit declarations of the methodology and definitions used, please refer to [IR19].


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18.2 Mass Budget

The mass budget is driven by the need to satisfy the requirements LAUN05 and LAUN10 in the MRD [NR01]. It has been computed according to the requirements MASS05, MASS10, MASS15, MASS20, and MASS25 in the MRD, and the additional requirements specified in the ECR [IR49], according to the definitions shown in the following figure.

Figure 18.2-1: Mass margin

The ECR requirements specify a 30% system margin for the baseline launch case using the Atlas launcher from KSC, and a 0% system margin for the backup launch case on Soyuz-Fregat from Kourou was requested at PM7 of the study. Note also that for those BepiColombo units under development, the margins applied according to the figure above are applied in addition to the unit margins used within the BepiColombo programme. The mass budgets presented are for the worst-case 2018 mission scenario (325 m/s effective delta V is used to determine the propellant loading) and are shown in Table 18.2-1: Solar Orbiter mass budget (2018 backup mission scenario) for the baseline Atlas launch case


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Solar Orbiter Estimated Mass

System NO. Unit Name Basic Unit Mass

Maturity margin Nominal Mass

Design Mass

[kg] [%] [kg] [kg] [kg]

Total Dry (w/o adapter) 933.7 14.5% 135.3 1100.9 1100.9

System Margin 0.0% 0.0

Total Dry with Margin (w/o adapter) 1100.9

Consumables

Consumables subtotal 145.3 5.5% 8.0 153.3 153.3

Total Wet with Margin (w/o adapter) 1254.2

Adapter Mass 41.0

Total Launch Mass 1295.2

Target Launch Mass (Soyuz) 1318.0

Difference 22.8

Table 18.2-2. The capacity of the Atlas 401 (2500kg) is estimated from the Launch Vehicle User’s Manual (a precise estimate will require consultation with the Atlas launch authority due to insufficient information in the User’s Manual). The capacity of Soyuz-Fregat (1318kg) is as specified to be used by the ESA project team within the frame of the current study, and should be regarded as a preliminary estimate only, due to uncertainties in the projected performance of Soyuz-Fregat from Kourou into high declinations.


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Design Mass

[kg] [%] [kg] [kg] [kg]

Structure

Structure Subtotal 209.3 13.0% 27.2 236.5 236.5

Thermal

Thermal Subtotal 36.2 25.0% 9.0 45.2 45.2

Communications

Communication Subtotal 102.6 0.0% 0.0 111.9 111.9

DHS

DHS Subtotal 29.3 9.1% 2.7 32.0 32.0

AOCS

AOCS Subtotal 67.8 11.5% 7.8 75.6 75.6

Power

Power Subtotal 136.0 5.7% 7.7 166.4 166.4

CPS

Propulsion Subtotal 50.6 10.7% 5.4 56.0 56.0

Payload

Payload Subtotal 124.8 25.0% 31.2 156.0 156.0

Harness

Harness Subtotal 79.2 25.0% 19.8 99.0 99.0

Misc

Misc Subtotal 14.0 25.0% 3.5 17.5 17.5

Feedthroughs

Feedthroughs Subtotal 17.9 25.0% 4.5 22.4 22.4

Heatshield

Heatshield Subtotal 66.0 25.0% 16.5 82.5 82.5




Consumables



Adapter Mass 41.0


Target Launch Mass (Atlas) 2500.0

Difference 833.6

Table 18.2-1: Solar Orbiter mass budget (2018 backup mission scenario) for the baseline Atlas launch case


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Design Mass

[kg] [%] [kg] [kg] [kg]




Consumables



Adapter Mass 41.0


Target Launch Mass (Soyuz) 1318.0

Difference 22.8

Table 18.2-2: Solar Orbiter mass budget (2018 backup mission scenario) for the backup Soyuz-Fregat launch case


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18.3 RF-Link Budget

The RF-link budgets for Solar Orbiter are driven by the requirements COMS10 and COMS55 in the MRD [NR01]; note that the requirement COMS10 has been updated in agreement with the ESA project team for a TM total data-rate target of 150 kbps at 1 AU. The downlink, uplink and ranging budgets are shown in the following tables respectively.

Ground station

SC Antenna

Bitrate [Bps]

SC- Earth [AU]

Mod. In. [rad]

SC EIRP

[dBW] Margin

[dB]

Downlink X - Band

LGA 8000 0,007 1,25

LGA 500 0,028 1,25 Kourou 15m

LGA 10 0,195 1,25

Nom 3,4 3Sig 1,9 RSS 3,0

LGA 8000 0,033 1,25

LGA 500 0,127 1,25 New

Norcia 35m

LGA 10 0,9 1,25

No 13,4 Ad 13,4 Fa 19,1

Nom 3,1 3Sig 1,3 RSS 2,5

MGA 8000 0,1 1,25

MGA 500 0,39 1,25 Kourou 15m

MGA 75 1,0 1,25

Nom 3,5 3Sig 2,4 RSS 2,7

MGA 8000 0,5 1,25

MGA 2000 1,0 1,25 New

Norcia 35m

MGA 400 2,0 1,0

No 36,3 Ad 35,3 Fa 37,1

Nom 3,4 3Sig 2,4 RSS 2,5

HGA 8000 0,55 1,25

HGA 2000 1,1 1,25 Kourou 15m

HGA 500 2,0 1,25

Nom 3,1 3Sig 2,0 RSS 2,6

HGA 150000 0,82 1,25

HGA 100000*) 1,0 1,25 Cebreros 35m

HGA 25000 2,0 1,25

Nom 3,1 3Sig 2,5 RSS 2,6

HGA 150000 0,72 1,25

HGA 79.000**) 1,0 1,25 New

Norcia 35m

HGA 20.00 2,0 1,25

No 50,7 Ad 50,6 Fa 50,6

Nom 3,1 3Sig 2,0 RSS 2,5

Downlink Ka - Band

HGA 150.000 0,68 1,25

HGA 72.000**) 1,0 1,25 New

Norcia 35m

HGA 18.000 2,0 1,25

No 58,5 Ad 58,3 Fa 59,6

Nom 3,1 3Sig 2,7 RSS 2,6


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Ground station

SC Antenna

Bitrate [Bps]

SC- Earth [AU]

Mod. In. [rad]

SC EIRP

[dBW] Margin

[dB]

HGA 150.000 0,8 1,25

HGA 120.000*) 1,0 1,25 Cebreros

35m HGA 20.000 2,0 1,25

No 58,5 Ad 58,3 Fa 59,6

Nom 3,2 3Sig 2,8 RSS 2,7

*) New Norcia Groundstation: 79 Kbps (X-Band) + 72 Kbps (Ka-Band ) = 151 Kbps **) Cebreros Groundstation: 100 Kbps (X-Band) + 120 Kbps (Ka-Band ) = 220 Kbps

Table 18.3-1: Solar Orbiter downlink budget summary

Ground station

SC Antenna

Bitrate [Bps]

SC- Earth [AU]

Mod. Index [rad]

SC G/T

[dB/K] TC Margin

[dB]

Uplink X - Band LGA 4000 0,006 1,4

LGA 250 0,025 1,4 Kourou 15m

LGA 7,6 0,06 1,0

Nom 3,6 3Sig 1,0 RSS 2,5

LGA 4000 0,1 1,4

LGA 250 0,4 1,4 New

Norcia 35m

LGA 7,6 1,4 0,7

No -28,1 Ad -28,6 Fa -22,8

Nom 3,4 3Sig 1,0 RSS 2,2

MGA 4000 0,07 1,4

MGA 250 0,28 1,4

Kourou 15m

MGA 7,6 0,97 0,7

Nom 3,2 3Sig 1,7 RSS 2,1

MGA 4000 1,29 1,4

MGA 2000 1,8 1,4

New Norcia 35m

MGA 1000 2,0 1,4

No -7,3 Ad -7,8 Fa -6,8

Nom 3,1 3Sig 1,6 RSS 2,0

HGA 4000 0,38 1,4

HGA 250 1,4 1,2

Kourou 15m

HGA 16 2,0 0,5

Nom 3,0 3Sig 1,7 RSS 2,0

Cebreros 35m HGA 4000 2,0 1,4

Nom 13,8 3Sig 12,5 RSS 12,8

New Norcia 35m

HGA 4000 2,0 2,0

No 7,3 Ad 7,3 Fa 7,6

Nom 12,7 3Sig 11,6 RSS 11,7

Table 18.3-2: Solar Orbiter uplink budget summary


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Uplink Bitrate [Bps]

Uplink Mod Index

[rad]

TC Margin

[dB]

Down link

Bitrate [Bps]

Down link

Mod Ind. [rad]

TM Margin

[dB]

Rang. Margin

[dB]

RNG Noise Accuracy

[m]

X-Band Uplink, X-Band Downlink Ranging via HGA, Distance 2 AU, Groundstation New Norcia

4000 TC 0,7 RNG 0,8

Nom 6,9 3Sig 5,7 RSS 5,8

8000 TM 0,8 RNG 0,7

Nom 3,1 3Sig 2,2 RSS 2,3

Nom 9,8 3Sig 7,8 RSS 8,1

Nom 0,2 Adv 0,3 Fav 0,2

X-Band Uplink, Ka-Band Downlink Ranging via HGA, Distance 2 AU, Groundstation New Norcia

4000 TC 0,5 RNG 1,0

Nom 3,4 3Sig 2,1 RSS 2,3

8000 TM 1,2 RNG 0,2

Nom 4,8 3Sig 4,2 RSS 4,2

Nom 9,8 3Sig 6,4 RSS 6,8

Nom 1,6 Adv 2,3 Fav 1,0

Table 18.3-3: Solar Orbiter ranging budget summary


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18.4 Power Budget

The power budget has been calculated according to the requirements POWM05, POWM10, POWM15, POWM20, POWM25 and POWM30 in the MRD, according to the definitions shown in the following figure.

Figure 18.4-1: Power Margin Philosophy

Because of the wide range of Sun distances experienced over the mission, there are 13 defined power modes for Solar Orbiter. Many of the power modes are identical from the perspective of unit function on-board the SC, but differ in heater power due to the varying thermal environments encountered. The 13 power modes are described in Table 18.4-1. Table 18.4-2 lists the thermal cases studied and marks their corresponding power modes in the generation of heater power requirements. The following table presents the power budgets by mode.

Mode Description Applicable Sun Distance

(AU) Power (W)

Mode 1

Standby (Launcher) Mode: This is the mode of the SC when on the launcher. The OBC, RIU, PCDU and Transponder (Rx only) are ON drawing power from the Battery. 1.0 156

Mode 2

Rate Reduction Mode: This mode corresponds to the Rate Reduction Mode of the AOCS: during this mode the Propulsion Subsystem, PCDU, OBC, RIU, IMU and X-band communications are operating. 1.0 460

Mode 3

Sun Acquisition Mode: This mode corresponds to the Sun Acquisition Mode of the AOCS: during this mode the Propulsion Subsystem, PCDU, SADE, OBC, RIU, CSS, IMU and X-band communications are operating. 1.0 478

Mode 12

Cruise Mode: This mode is identical to the Operational Mode (7), but with a different range of applicable Sun distances. 0.8 to 1.2 818

Mode 4

Safe Hold mode: This mode corresponds to the Safe Hold Mode of the AOCS: during this mode the Propulsion Subsystem, PCDU, SADE, OBC, RIU, CSS, STR, IMU, HTHGA Mechanism and X-bannd communications are operating. 0.8 to 1.2 730

Mode 11

Anti-Sun Mode: This mode corresponds to the anti-sun period in the CP, when the SC is greater than 1.2 AU from the Sun (TBC): durng this mode the PCDU, SADE, DHS Subsystem, AOCS Subsystem, and Communication subsystem (X+Ka band) and IS-instrument payload are operating. 1.2 to 1.5 756

Mode 15

Anti-Sun Safe Mode: This mode is identical to Mode (4) except with a different heater power due to the altered Sun distance. 1.2 to 1.5 641

Mode 7

Operational Mode: This mode corresponds to the normal Sun-pointing mode of the SC outside of the science windows (i.e. no RS-instrument operation): during this mode the Propulsion Subsystem, the PCDU, SADE, DHS Subsystem, AOCS Subsystem, and Communication subsystem (X+Ka band) and IS-instrument payload are operating. 0.234 to 0.8 925

Mode 5

Sun Keeping Mode: This mode corresponds to the Survival Mode of the AOCS: during this mode the Propulsion Subsystem, PCDU, SADE, OBC, RIU, 0.234 to 0.8 839


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CSS, IMU, RMU, FCE and X-band communications are operating.

Mode 14

Venus Flyby Mode: This mode represents the Venus Flyby manouevres: During this mode the Propulsion subsystem and payloads are switched off. 0.72 750

Mode 8

Science with Comms: This mode corresponds to the Fine Sun Pointing AOCS mode, when the platform is stabilised for operation of the Remote Sensing payload, and the SC is above 0.28 AU and therefore the HTHGA is deployed and active: during this mode the PCDU, SADE, DHS Subsystem, AOCS Subsystem, Communication Subsystem (X+Ka band) and IS-instrument payload are operating. 0.28 to 0.8 993

Mode 10

Science No Comms: This mode corresponds to the Fine Sun Pointing AOCS mode, when the platform is stabilised for operation of the Remote Sensing payload, and the SC is below 0.28 AU and therefore the HTHGA is shielded behind the SC: during this mode the PCDU, SADE, DHS Subsystem, AOCS Subsystem, and IS-instrument payload are operating. 0.234 to 0.28 722

Mode 13

Cold Science Case: This mode is identical to the Science with Comms Mode (10), except that the heater power is different due to the greater Sun distance. 0.8 832

Table 18.4-1: Power mode descriptions

Design Cases

Sun Distance

(AU) HGA

deployed RPW

deployed Solar Array

Angle (°) SC

Temp SC

pointing Power

Budget Case Purpose Notes

D1 0.2343 N Y 78 40 Sun Mode 10 Radiator sizing

Fine Sun pointing, non-comms: Science, perihelion case, dish behind heatshield (HOT OP CASE) - 2017 baseline, interpolated from ESOC supplied data

D2 0.28 Y Y 75 40 Sun Mode 8 Radiator sizing

Fine Sun pointing, with comms: Science, Limiting case with HGA exposed to Sun (i.e. minimum distance at which HGA is out in flux) (HOT OP CASE)

D3 DELETED

D4a 1.5 Y Y 0 10 Anti-Sun Mode 11 Heater sizing (ANTI_SUN HEATER SIZING)

Anti-Sun mode (COLD CASE)

D4b DELETED

D4c 1.5 Y Y 0 10 Sun Mode 11 Alt. Heater sizing (SUN-POINTING HEATER SIZING)

As a comparator against case D4a to assess benefit of AntiSun strategy


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D4d 1.5 Y Y 0 10 Anti-Sun Mode 15

Alt. Heater sizing (ANTISUN-POINTING HEATER SIZING)

Sizing heaters of survival. Solar array sizing case.

D5 1.2 Y Y 0 10 Sun Mode 12 Heater sizing (NON-SCIENCE)

Limit of Sun-pointing case (COLD CASE) - max heater power required before pointing anti-sun

D6 0.8 Y Y 70 30 Sun Mode 13 Heater sizing (SCIENCE)

Fine Sun pointing aphelion: science cold case, max heater power required during science

D9 0.8 y y 70 30 Sun Mode 5 Heater sizing (SAFE)

Sun-keeping mode: safe mode heater power

Table 18.4-2: Thermal case definitions and corresponding power modes

18.5 Power Mode Summary

Mode Description Sun Distance

(AU) Power (W) Mode 1 Standby (Launcher) Mode 1.0 156 Mode 2 Rate Reduction Mode 1.0 460 Mode 3 Sun Acquisition Mode 1.0 478 Mode 12 Cruise Mode 0.8 to 1.2 893 Mode 4 Safe Hold mode 0.8 to 1.2 854 Mode 11 Anti-Sun Mode 1.2 to 1.5 743 Mode 15 Anti-Sun Safe Mode 1.2 to 1.5 691 Mode 7 Operational Mode 0.234 to 0.8 1062 Mode 5 Sun Keeping Mode 0.234 to 0.8 826 Mode 14 Venus Flyby Mode 0.72 762 Mode 8 Science with Comms 0.28 to 0.8 993 Mode 10 Science No Comms 0.234 to 0.28 722 Mode 13 Cold Science Case 0.8 944

Table 18.5-1: Solar Orbiter power mode summary (for full breakdown of the power budget, please refer to the Solar Orbiter budgets document [IR19])


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18.6 Delta V and Attitude Control Budget

The delta V and attitude control budget has been calculated to satisfy the requirement PROP15, according to the requirements PROP05 and PROP10 in the MRD. The following definitions are used:

• Nominal Mission Delta V: Basic delta V with no margins applied, or consideration of attitude constraints and thruster inefficiency factors

• Nominal Effective Delta V: Basic delta V with accounting for attitude constraints and thruster inefficiency factors

• Total Effective Delta V: Nominal effective delta V with 5% margin applied.

Figure 18.6-1: Delta V budget philosophy

Table 18.6-1 and Table 18.6-2 present the delta V and attitude control budgets for the nominal 2017 and backup 2018 mission scenarios respectively.


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Margin/Eff. Factor Value Source

Deterministic Flight Programme Error 25 0 25 m/s VEX error correction precedent [1]


DSM for non-optimal launch date 0 m/s 21 day zero penalty window is possible - taken from [2]


Gravity losses 0 m/s Negligible for Solar Orbiter

Total Nominal Mission Delta V 160 m/s Clean nominal mission delta V

AOCS longitudinal thrusters inefficiency 70 m/s 3% 2 m/s Thr. Eff. For longitudinal manoeuvres [4]

AOCS transverse thrusters inefficiency 90.0 m/s 107% 96 m/s Worst-case Thr. Eff. For transverse manoeuvres [4]

Total thrusters configuration effect 98 m/s Calculated for current thruster configuration

Nominal Effective Delta V 258 m/s Nominal effective mission delta V

Total Effective Delta V margin 258 m/s 5% 13 m/s As required by [3]

Total Effective Delta V 271 m/s Total effective mission delta V

Propellant mass for attitude control 8.0 kg 100% 16.0 kg Estimate based on mission angular impulse [4]

Table 18.6-1: Nominal 2017 Mission Scenario


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Margin/Eff. Factor Value Source

Deterministic Flight Programme Error 25 m/s VEX error correction precedent [1]


DSM for non-optimal launch date 50 m/s 50m/s penalty required for 20 day LW [2]


Gravity losses 0 m/s Negligible for Solar Orbiter

Total Nominal Mission Delta V 210 m/s Clean mission delta V

AOCS longitudinal thrusters inefficiency 120 m/s 3% 4 m/s Thr. Eff. For longitudinal manoeuvres [4]

AOCS transverse thrusters inefficiency 90.0 m/s 107% 96 m/s Worst-case Thr. Eff. For transverse manoeuvres [4]

Total thrusters configuration effect 100 m/s Calculated for current thruster configuration

Nominal Effective Delta V 310 m/s Nominal effective mission delta V

Total Effective Delta V margin 310 m/s 5% 15 m/s As required by [3]

Total Effective Delta V 325 m/s Total effective mission delta V

Propellant mass for attitude control 8.0 kg 100% 16.0 kg Estimate based on mission angular impulse [4]

[1] Email correspondance with [email protected], 01/04/2009.

[2] SOL-ESC-RP_GFA-JRC-WP 5XX (draft) - Solar Orbiter Mission Analysis for Launch in 2017-2018.

[3] SOL-EST-RS-00049 Issue 4 - Solar Orbiter Mission Requirements Document. [4] SOL-S-ASTR-TN-00XX (draft) - Solar Orbiter Delta V and Attitude Control Budgets. Type 2 TCMs are conducted using a dog-leg manoeuvre to access a priori unknown thrust directions – transverse thrusters inefficiency accounts for the inefficiencies associated with this.

Table 18.6-2: Backup 2018 Mission Scenario


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18.7 Pointing Budget

The following section presents the budgets for the various pointing requirements presented in POIN30, and the coalignment requirement POIN45 of the MRD. The requirements are reproduced in the following table for convenience. It should be noted that these requirements are considered active only during the science windows of the mission when the RS-payloads are operating.

Table 18.7-1: Solar Orbiter pointing and accuracy requirements during RS-observation

The requirements are interpreted as complete, i.e. the APE requirement applies over the entire alignment-chain, from the STR LoS through to the reference optical LoS of the instrument. Accordingly, budget allocations are included for instrument misalignment. For the APE and Co-alignment budgets, three budgets are presented, corresponding to the three thermal cases D1, D2 and D6 (see Table 18.4-2 for a specification of the thermal cases studied), which have been identified as the extreme thermal cases of the operational phase of the mission, during which the alignment requirements of POIN30 are applicable. These are:

D1 (“Hot1”) 0.2343 AU HTHGA folded, all instruments operating D2 (“Hot2”) 0.28 AU HTHGA exposed, all instruments operating D6 (“Cold”) 0.8 AU HTHGA exposed, all instruments operating

Table 18.7-2: Thermal cases corresponding to alignment budgets


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18.7.1 RPE Budget [POIN30c]

18.7.1.1 Line of Sight Y-Axis

Solar Orbiter AOCS Relative Pointing Error Type

LOS Pointing Stability over time (10 sec)

Comments

LOS Y-Axis Confidence level 95,5 % Proposed SRS Requirement (arcsec) 2.40 Error contributors (in arcsec):

Micrometeorids

B 0.00 negligible (impacts causing 0.1"/s

rate change happen 0.03 times/day Solar pressure noises R 0.00 Negligible during this period. Wheels Microvibrations (3000 rpm) not part of AOCS Budget R 0.00 HGA pointing mechanism R 0.10 Simulation Result Solar Array rotation H 0.00 No rotation during measurement Orbit Estimation H 0.00 AOCS error: bandwidth = 0.015Hz, gyro stellar estimator B 1.80 Simulation Result AOCS Wheel torque noise (included in AOCS error) R 0.00 AOCS internal thermal deformations H 0.10 Assumption Total (linear sum) 2.00 arcsec Proposed SRS Requirement (arcsec) 2.40 arcsec

Table 18.7-3: Line-of-Sight RPE pointing budget


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18.7.1.2 Line of Sight Z-Axis



Comments

LOS Z-Axis Confidence level 95,5 % Proposed SRS Requirement (arcsec) 1.56 Error contributors (in arcsec):

Micrometeorids

B 0.00 negligible (impacts causing 0.1"/s

rate change happen 0.03 times/day Solar pressure noises R 0.00 Negligible during this period. Wheels Microvibrations (3000 rpm) not part of AOCS Budget R 0.00 HGA pointing mechanism R 0.10 Simulation Result Solar Array rotation

H 0.00 No rotation during measurement Orbit Estimation H 0.00 AOCS error: bandwidth = 0.015Hz, gyro stellar estimator B 1.10 Simulation Result AOCS Wheel torque noise (included in AOCS error) R 0.00 AOCS internal thermal deformations H 0.10 Assumption Total (linear sum) 1.30 arcsec Proposed SRS Requirement (arcsec) 1.56 arcsec

Table 18.7-4: Line-of-Sight RPE pointing budget


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18.7.1.3 Around Line-of-Sight



Comments

Around LOS

Confidence level 95,5 % Proposed SRS Requirement (arcsec) 3.50 Error contributors (in arcsec):

Micrometeorids

B 0.00

negligible (impacts causing 0.1"/s rate change happen 0.03

times/day Solar pressure noises R 0.00 Negligible during this period. Wheels Microvibrations (3000 rpm) not part of AOCS budget R 0.00 HGA pointing mechanism R 0.10 Simulation Result Solar Array rotation

H 0.00 No rotation during measurement Orbit Estimation H 0.00 AOCS error: bandwidth = 0.015Hz, gyro stellar estimator B 1.40 Simulation Result AOCS Wheel torque noise included in AOCS error R 0.00 AOCS internal thermal deformations H 0.10 Assumption Total (linear sum) 1.60 arcsec Proposed SRS Requirement (arcsec) 3.50 arcsec

Table 18.7-5: Around Line-of-Sight RPE pointing budget


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18.7.2 APE Budget [POIN30a]

18.7.2.1 APE LoS

Thermal Case D1 (“Hot1”)

Frequency Class Error Component

Error Type Class

APE Contribut

ion Units Data source

Total APE 196.0 arcsec

APE Requirement 120.0 arcsec

Budget Margin (Absolute) -76.0 arcsec

Budget Margin (Percentage) -63.3%

Table 18.7-6: LoS APE during thermal case D1 (“Hot1”)

Thermal Case D2 (“Hot2”)


Error Type Class

APE Contribut




Budget Margin (Absolute) 30.6 arcsec

Budget Margin (Percentage) 25.5%

Table 18.7-7: LoS APE during thermal case D2 (“Hot2”)

Thermal Case D6 (“Cold”)


Error Type Class

APE Contribut




Budget Margin (Absolute) -25.3 arcsec

Budget Margin (Percentage) -21.1%

Table 18.7-8: LoS APE during thermal case D6 (“Cold”)

Note that the reported non-compliance of the APE during thermal case D1 (caused by a localised cooling due to the TWTA radiators) could be avoided by leaving the TWTAs on during close-approach to the Sun, or locally heating the area – thereby reproducing the deformation seen during thermal case D2. This will be verified in future work. Also note that the non-compliance of the APE during thermal case D6 should be viewed in the context of the suggested change to the MRD reported in [IR43], whereby the requirement is interpreted to be relaxable as the SC moves away from the Sun (the requirement is driven by the occulter-sizing on the METIS instrument). If the recommended updated APE requirement at ~0.8 AU is taken (~7-8 arcminutes), then the APE is easily compliant also for thermal case D6.


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18.7.2.2 APE Around LoS

The APE performance around the LoS has yet to be calculated. However, given that TED contributions are not expected to exceed 2-3 arcmin, the 20’ APE requirement around X is expected to be easily met.


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18.7.3 Instrument Co-alignment Budget

This budget presents the contributions to the RS-instrument co-alignment requirement POIN45 of within 2 arcminutes. The requirement is interpreted as an allowable misalignment of 2 arcmins between any two of the RS-instruments LoS (SPICE, STIX, METIS, EUI …). Accordingly, in combination with POIN30a, this implies that any one of the instruments can be misaligned from the STR LoS by up to 4 arcmins. The following figure pictorially represents the interpretation of the requirements.

Figure 18.7-1: POIN45 interpretation and relationship with POIN30a

In addition to the misalignment caused by TED, the following budget items are also included:

• Gravity release • Launch slippage • AIT alignment (including error internal to instrument) • Co-alignment.

When these budget items are added to the TED (quadratically), the total worst-case co-alignment budget (corresponding to Case D1) is as shown in Table 18.7-9.

Commanded Attitude

METIS

STIX EUI

SPICE

PHI

2’ 2’

2’

2’

2’

2’ 2’

2’

2’

2’ 2’


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Frequency

Class Error Component Error Type Class

APE Contribution Units Data source

Systematic

Gravity Release (1g-0g) U 20.0 arcsec ASU estimate (TBC)

Launch Slippage (STR to METIS slippage) B 20.0 arcsec ASU estimate (TBC)

Ageing U 0.0 Waiting on ASU Estimate (TBC)

Typical AIT alignment U 33.0 arcsec Typical achieved by AIT

Worst-Case TED between any two instruments 49.0

Systematic Error Subtotal 92.5 arcsec

Allocation 120.0 arcsecs

Table 18.7-9: Worst case coalignment budget, corresponding to thermal case D1 (“Hot1”)


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18.8 Science Windows

The following tables present the characteristics of each of the science windows for the 2017 and 2018 mission scenarios, according to a strict interpretation of requirement PERF20 of the MRD. There are two important features to note:

• Several of the early minimum latitude and perihelion science windows are overlapping; as the perihelion of the SC increases, the overlap reduces and eventually disappears

• In the 2017 case a 4:3 resonance is entered after VGAM-2, such that 4 SC orbits are prescribed before the next VGAM – there are thus 14 operational orbits in total for the 2017 mission. This is in comparison to the backup 2018 mission in which a 3:2 resonance is entered directly after VGAM-2 and which therefore contains 13 operational orbits in total.

Colour coding:

• Dark Grey: NMP science windows • Light Grey: EMP science windows.

18.9 2017 Mission

Date (midpoint) Day (midpoint) SC-E distance (AU) SC-S dist. (AU) S-SC-E (°) Solar Latitude (°) Min Max Min Max Beg. End Begin End

07/05/2021 1584.0 0.819 0.827 0.712 0.800 22.770 23.360 7.700 7.640 22/10/2021 1752.0 0.262 0.354 0.717 0.804 104.710 116.830 7.690 7.650 09/04/2022 1921.0 1.179 1.289 0.712 0.800 7.240 7.090 7.700 7.640 24/09/2022 2089.0 1.424 1.495 0.717 0.804 156.010 159.860 7.690 7.660 09/03/2023 2255.1 0.558 0.614 0.497 0.627 16.330 23.680 16.660 16.310 06/08/2023 2405.1 0.738 0.762 0.494 0.625 81.350 80.650 16.670 16.270 02/01/2024 2554.1 1.223 1.351 0.506 0.634 57.860 70.160 16.630 16.400 26/05/2024 2698.5 0.474 0.524 0.510 0.624 32.590 28.630 25.230 24.680 23/10/2024 2848.5 0.849 0.871 0.508 0.622 126.310 141.440 25.250 24.630 22/03/2025 2998.5 1.187 1.306 0.505 0.620 14.530 21.220 25.270 24.580 09/08/2025 3138.9 0.400 0.436 0.519 0.615 81.110 76.210 31.260 30.450 05/01/2026 3287.9 0.965 0.981 0.528 0.621 57.220 69.710 31.150 30.720 04/06/2026 3437.9 1.147 1.252 0.525 0.620 39.280 34.420 31.180 30.650 22/10/2026 3578.3 0.390 0.393 0.504 0.585 116.830 127.060 34.550 33.610

Table 18.9-1: Maximum latitude science window definition for the 2017 launch case


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19/06/2021 1627.0 0.468 0.596 0.257 0.298 25.730 95.970 -2.830 -7.280 04/12/2021 1795.0 0.667 0.962 0.257 0.292 122.730 44.180 -2.290 -7.430 22/05/2022 1964.0 0.929 0.985 0.257 0.298 47.130 125.380 -2.780 -7.300 07/11/2022 2133.0 0.512 0.846 0.257 0.304 85.730 16.970 -3.270 -7.140 04/04/2023 2281.1 0.466 0.750 0.234 0.285 80.100 163.060 -0.190 -16.550 01/09/2023 2431.1 0.889 0.979 0.234 0.288 39.050 34.190 -0.730 -16.480 28/01/2024 2580.1 0.555 0.898 0.234 0.277 147.250 98.050 1.360 -16.710 23/06/2024 2726.5 0.448 0.729 0.270 0.305 18.700 78.000 -1.840 -25.310 20/11/2024 2876.5 0.952 0.998 0.270 0.307 104.360 41.330 -2.500 -25.360 18/04/2025 3025.5 0.579 0.867 0.270 0.299 68.010 136.620 0.140 -25.080 08/09/2025 3168.9 0.458 0.710 0.316 0.337 36.220 17.140 -2.950 -29.320 05/02/2026 3318.9 1.024 1.035 0.316 0.339 155.560 123.130 -3.600 -29.550 05/07/2026 3468.9 0.577 0.799 0.316 0.340 13.480 54.950 -4.250 -29.770 20/11/2026 3607.3 0.504 0.704 0.359 0.372 103.490 68.050 0.920 -26.280

Table 18.9-2: Perihelion science window definition for the 2017 launch case


23/06/2021 1631.0 0.478 0.693 0.257 0.354 54.440 114.210 -6.520 -5.630 08/12/2021 1799.0 0.793 1.050 0.257 0.346 89.220 26.060 -6.140 -5.870 26/05/2022 1968.0 0.899 0.973 0.257 0.353 77.870 148.910 -6.490 -5.660 10/11/2022 2136.0 0.429 0.746 0.257 0.345 63.440 5.300 -6.100 -5.900 09/04/2023 2286.1 0.583 0.892 0.236 0.363 131.680 133.080 -13.280 -13.140 06/09/2023 2436.1 0.950 0.997 0.237 0.366 5.460 51.000 -13.650 -12.980 03/02/2024 2586.1 0.394 0.693 0.238 0.369 139.240 64.140 -14.000 -12.810 30/06/2024 2733.5 0.637 0.912 0.280 0.394 61.100 104.440 -22.480 -21.230 27/11/2024 2883.5 0.992 1.012 0.281 0.396 57.710 18.420 -22.810 -21.030 25/04/2025 3032.5 0.414 0.665 0.276 0.385 117.110 147.180 -21.370 -21.830 18/09/2025 3178.9 0.710 0.940 0.337 0.432 17.140 36.190 -29.320 -28.030 15/02/2026 3328.9 1.030 1.034 0.339 0.434 123.130 78.320 -29.550 -27.840 14/07/2026 3477.9 0.415 0.598 0.333 0.426 51.050 81.090 -28.520 -28.570 05/12/2026 3622.3 0.810 0.990 0.400 0.480 52.940 32.030 -33.390 -32.630

Table 18.9-3: Minimum latitude science window definition for the 2017 launch case


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18.10 2018 Mission

Date (midpoint) Day (midpoint) SC-E distance SC-S dist. S-SC-E (°) Solar Latitude (°) Min Max Min Max Beg. End Begin End

19/02/2022 1300.0 1.202 1.335 0.485 0.614 5.130 10.550 10.060 10.080 19/07/2022 1450.0 0.754 0.779 0.488 0.616 111.270 90.260 10.080 10.070 16/12/2022 1600.0 0.521 0.574 0.491 0.619 51.440 53.320 10.090 10.060 20/05/2023 1755.0 1.201 1.321 0.511 0.625 55.860 44.910 20.090 20.020 17/10/2023 1905.0 0.833 0.864 0.513 0.627 103.510 105.130 20.120 20.000 14/03/2024 2054.0 0.463 0.510 0.503 0.618 14.060 8.830 19.990 20.080 21/08/2024 2214.0 1.167 1.273 0.533 0.626 135.550 117.320 27.530 27.250 18/01/2025 2364.0 0.947 0.975 0.535 0.628 29.630 33.650 27.570 27.210 16/06/2025 2513.0 0.391 0.423 0.527 0.621 76.340 62.680 27.410 27.350 24/11/2025 2674.0 1.075 1.183 0.520 0.598 66.680 72.740 31.790 31.270 22/04/2026 2823.0 1.035 1.046 0.513 0.592 40.350 29.750 31.570 31.480 19/09/2026 2973.0 0.367 0.368 0.515 0.593 123.460 127.010 31.630 31.430 26/02/2027 3133.0 0.952 1.083 0.479 0.551 17.610 19.520 32.820 32.200

Table 18.10-1: Maximum latitude science window definition for the 2018 launch case


27/01/2022 1277.0 0.604 0.938 0.243 0.290 103.500 26.830 -9.580 4.050 26/06/2022 1427.0 0.859 0.975 0.244 0.293 81.730 177.900 -9.460 4.280 22/11/2022 1576.0 0.447 0.692 0.243 0.283 45.900 22.380 -9.870 3.310 24/04/2023 1729.0 0.598 0.897 0.271 0.303 146.150 114.350 -18.540 5.090 21/09/2023 1879.0 0.918 0.994 0.271 0.305 15.180 70.180 -18.300 5.550 18/02/2024 2029.0 0.425 0.668 0.271 0.308 119.000 49.820 -18.050 5.990 23/07/2024 2185.0 0.611 0.849 0.315 0.339 67.930 121.060 -21.980 4.980 19/12/2024 2334.0 1.000 1.030 0.315 0.334 53.510 10.920 -23.410 2.810 18/05/2025 2484.0 0.439 0.684 0.315 0.335 125.180 154.520 -23.070 3.360 26/10/2025 2645.0 0.595 0.779 0.362 0.377 12.530 34.170 -19.440 6.720 24/03/2026 2794.0 1.069 1.074 0.362 0.374 126.180 86.620 -21.180 4.670 21/08/2026 2944.0 0.488 0.691 0.362 0.374 53.360 86.620 -20.760 5.180 02/02/2027 3109.0 0.550 0.717 0.388 0.398 55.150 31.280 -10.580 14.250

Table 18.10-2: Perihelion science window definition for the 2018 launch case


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20/01/2022 1270.0 0.401 0.711 0.244 0.356 148.160 77.630 -9.170 -6.650 19/06/2022 1420.0 0.951 1.007 0.244 0.352 40.860 110.440 -9.240 -6.370 16/11/2022 1570.0 0.563 0.870 0.243 0.349 72.020 17.250 -9.320 -6.080 17/04/2023 1722.0 0.416 0.691 0.272 0.365 98.330 168.250 -19.220 -13.340 13/09/2023 1871.0 0.985 1.019 0.274 0.376 21.860 25.070 -18.740 -15.060 10/02/2024 2021.0 0.606 0.892 0.274 0.373 159.680 103.880 -18.860 -14.650 12/07/2024 2174.0 0.391 0.588 0.329 0.418 28.760 63.070 -26.250 -23.730 09/12/2024 2324.0 1.030 1.040 0.327 0.416 84.600 53.510 -26.370 -23.410 08/05/2025 2474.0 0.684 0.926 0.326 0.413 75.850 125.180 -26.490 -23.070 10/10/2025 2629.0 0.378 0.500 0.395 0.472 35.380 19.990 -31.090 -29.450 09/03/2026 2779.0 1.063 1.068 0.394 0.471 128.710 138.600 -31.190 -29.240 06/08/2026 2929.0 0.799 0.985 0.393 0.469 20.620 39.140 -31.280 -29.020 10/01/2027 3086.0 0.363 0.415 0.458 0.529 96.510 82.960 -32.930 -31.560

Table 18.10-3: Minimum latitude science window definition for the 2018 launch case


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18.11 Mass Memory Usage Profile

The mass memory profiles are presented here for the baseline operational scenario (stowage of the HTHGA below 0.3 AU.) for the 2017 baseline mission scenario, and the 2018 backup mission scenario. The analysis was performed under the following assumptions:

• Data-rates: o 6 kbps for SC housekeeping, which is a constant data-rate o 23 kbps for IS suite observations, which are considered to be permanently collecting data o 77 kbps for RS suite baseline observations, which occur at the points in the trajectory as

defined by the science window schedules given in the previous section o ¼ turbo coding.

• TT&C capability: o 150 kbps @ 1AU reference data rate, scaled with distance and rounded (downwards) to the

nearest available bit-rate supported by the transponder o 4 hour communication window available per day, subject to the following constraints:

A minimum visibility of either New Norcia or Cerbreros of at least 10º above the local horizon

Greater than 5º S/SC/E angle (scintillation effects assumed to prevent communication below this).

The baseline sizing requirement of the mass memory is therefore 491Gb, occurring at MoL in the baseline 2017 mission. It should be noted that appropriate use of a 2nd ground station during this critical period will substantially reduce the SSMM peak usage.


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491 Gbits

0

100

200

300

400

500

600

0 500 1000 1500 2000 2500 3000 3500 4000

Day Of Mission

Stor

ed T

M (G

Bits

)

0

50

100

150

200

250

300

350

400

450

500

TM D

ata

rate

(kbp

s)

Stored TM (Gb)Min LatitudeMax LatitudePerihelionBit Rate (kbps)

Figure 18.11-1: SSMM usage for the 2017 baseline mission scenario

333 Gbits

0

50

100

150

200

250

300

350

0 500 1000 1500 2000 2500 3000 3500

Day Of Mission

Stor

ed T

M (G

Bits

)

0

50

100

150

200

250

300

350

400

450

500

TM D

ata

rate

(kbp

s)

Stored TM (Gb)Min LatitudeMax LatitudePerihelionBit Rate (kbps)

Figure 18.11-2: SSMM usage for the 2018 backup mission scenario


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APPENDIX: LAUNCHER PERFORMANCE

Currently, the Solar Orbiter mission is required to maintain SC compatibility with the Soyuz-Fregat launch vehicle from Kourou (LAUN05 in the MRD [NR01]). The declination/V_infinity required for the injection by the baseline mission analysis [IR26] are 21.45°/3.64km/s and -52.97°/3.66km/s for the 2017 and 2018 scenarios respectively; the Soyuz PSW performance for these required launch conditions are quoted in [IR26] as being 1300kg (TBC) and 1194kg (TBC). It is a primary concern of Astrium Ltd that these projected launch capacities will not be achievable by Soyuz, in particular in the case of the high declination 2018 mission scenario. We determine SF performance from Kourou from this V_inf requirement using a 2nd order polynomial fit to performance data supplied by ESOC – this is the latest information regarding Soyuz-Fregat performance from Kourou into escape trajectories as a function of varying declination. A required V_infinity of 3.76 km/s is used from previous Mission Analysis in order to satisfy the 20-day launch window requirement (LAUN15).

Figure 18.11-1: Selected 20 day launch window around the 9/1/2017 launch date (from previous

analysis)

Taking the latest data from ESOC, we can see that compliance with LAUN15 implies an all-up launch mass of the SC of 1301.2kg8 (including the adaptor). Note that this capability is projected for only +ve declinations at injection. It is not clear that Soyuz-Fregat is able to deliver any useful mass into –ve declinations from Kourou due to restrictions on the 3rd stage drop-zone. The relevance to Solar orbiter is clear: The current mission analysis for Solar Orbiter [IR26] states a requirement for a declination of -52.97 degrees for the 2018 launch scenario – Astrium Ltd has no confidence that Soyuz-Fregat will be able to deliver any useful mass to these injection conditions.

8 The projected Soyuz-Fregat capability for the selected mission scenario from the previous phase [IR07] was 1261.7kg (this figure included the LVA mass and a 20 day Launch Window penalty).


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Figure 18.11-2: Projected Soyuz-Fregat performance as a function of C3 from Kourou for various

declinations: the projected performance for the Solar Orbiter 9/1/2017 launch is 1301kg

20 degrees x y 0 0 2075 5 25 1750 10 100 1500 15 225 1290 20 400 1080 c2 0.771428571 c1 -64.42857143 b 2067.571429 x 14.1376 y 1310.893 25 degrees x y 0 0 1975 5 25 1690 10 100 1430 15 225 1240 20 400 1050 c2 0.742857143 c1 -60.85714286 b 1974.142857 x 14.1376 y 1262.245 21 degrees x 14.1376 y 1301.163

Table 18.11-1: 2nd order polynomial fit of Soyuz-Fregat performance data