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VEX design report - TOC

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1. 28V Main Bus Capacitor 1 mF MPPTI MPPT Ref LCL OFF 2 3 Voting 32V 80V PDU A TM TC Control MPPT 2 amp 3 Local Auxiliary Supply A amp B Bus Control Main Error Amplifier 3 x BCDR Pout 3 x 300W E A3 E A2 2 BCR LCL BOR CIL BCR Failure detection BCR Converter Stage Super Buck E Battery 3 Battery 2 Overvoltage Battery 1 BCR Ref Battery Regulation PWM lt BDR LCL OFF Ref p gt id Control BER TCE ORE gt BDR Failure detection 16V 25V lt m Overvoltage 5 a A BDR LCL OFF T TC RTU E Control 5 vercurrent amp V Battery i Telemetry I F BCDR BDR LCL OFF Auxiliary Supply m A amp B Bus Control OFF MEA Figure 7 2 3 PCU Functional Block Diagram T Ref VEX T ASTR TCN 00349 A D Issue 2 Rev 0 Date 06 02 2004 AA ASTRIUM express Page 7 12 E 7 2 4 Main Bus Power Distribution The Venus Express power distribution policy is based on a centralised scheme and is ensured by the Power Distribution Unit PDU One switched protected power line derived from the regulated main power bus is dedicated
2. 10 1 10 1 SOFTWARE COMPONENTS AND ASSOCIATED FUNCTIONS 10 2 2 IL uae dI d VR A A MM APR MEE 10 4 10 3 PROCESSOR MODULE FIRMWARE e tppat tiras u REIN UR M HU CERIS 10 7 TL eee en m T IE 10 9 1041 DMS and AQCMS SY layered DPegkdoWn uuu aaa Ee ai eie uiae Shoe 10 9 1042 Common DMS AOCS SOFTWARE 10 11 135 210 po O nln 10 12 I04 4 AOCNIS Appicaton uuu ayna sukata 10 13 Uo Pii eem 10 14 ID ERE Nae PRP NUM 10 15 22 JI l ABE 52 uu 10 16 IDA IRAN PONDER Al uuu 10 17 PITA SR n aestu apasi ae tan nett 10 18 11 PPOR IPE 11 1 ILI TEMP HIEDAB DII EM MI PME UM 11 1 DL DNE MEMINI UM 11 4 11 3 H W RECONFIGURATION OF CENTRAL COMPUTING AND COMMUNICATIONS 11 6 114 DMS SYSTEM S W IMPLEMENTED 11 9 INE EE BEDA ERREUR 11 11 SUMING Y ULLA ii Sa i 11 13 LL 11 16 E EA EAE A EEEO 11 18
3. Issue 2 Rev 0 _ Date 06 02 2004 72 ASTRIUM r express C Page 53 The Reaction Wheel Assembly RWA includes 4 Reaction Wheels RW implemented on a skewed configuration This configuration identical to Mars Express enables to perform most of the nominal operations of the mission with a 3 wheels configuration among 4 During some critical phases during which the transition to the SAM shall be avoided before Venus Insertion Manoeuvre a 4 wheels configuration may be used under ground request The Reaction wheels provide the AOCS control torques during all the phases of the mission except the trajectory corrections the attitude acquisition and back up modes The Propulsion configuration includes a Main Engine 414 N which is used to perform all the major trajectory changes and 10 N thrusters used for the attitude control and also to produce the thrust during the small trajectory corrections The 10 N thrusters configuration is optimised to perform all the attitude control functions with only 4 redunded thrusters each of them being implemented near a corner of the Z face of the spacecraft The Venus Express propulsion configuration is identical to Mars Express 2 redunded Solar Array Drive Mechanisms SADM are implemented on the Y and Y walls of the spacecraft to control the orientation of the Solar Arrays The SADM position is fixed during the the Venus observation phase requiring no SADM actuation once the selecte
4. sasa sasatuuakusaasuananauakhanqasaqakawssaauawassawansasapawq 3 1 3 1 DREE uu u asss 3 1 3 2 SPACECRAFT CONFIGURATION 3 2 3 3 LAUNCH CONFIGURATION 0 000000000000000 3 6 3 4 DEPLOYED CONFIGURATION S 12 4 THERMAL CONTROL DESIGNN 4 1 4 1 THERMAL CONTROL DESIGN 00000 00008 4 4 2 THERMAL CONTROL CONFIGURATION scscscscscscccccccceececccccecececececececeeeueeeceueceseuseeuseeceees 4 4 4 3 THERMAL CONTROL OVERVIEW 2 2 7200000 4 6 4 3 1 Thermal Conrol RTT ETE TT 4 6 43 2 ee LOR eT a ee ene sa eee nN a ee 4 8 4 3 3 ili n pisi GQ 01 uuu uuu Em m TE 4 10 4 3 4 PEORES TUE AE A OA E 4 11 5 ATTITUDE AND ORBIT CONTROL SYSTENI 5 1 Astrium Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 Date 06 02 2004 Page vi TO ASTRIUM r express 34 IIT au i FR DU UM S 5 2 AOCS HARDWARE ARCHITECTURE J J7 l oon tdi 5 2 22 AUC S MODE ARCHITECTURE u tira NTE ENEE AEAE
5. Figure 3 3 4 Access to skin connectors under fairing Ref VEX T ASTR TCN 00349 EADS Vi Issue 2 Rev 0 Date 06 02 2004 ASTRIUM egenus Page 3 10 HGA 2 s HX me VILI 5 E 41 2 N RW THERMAL RA gt gt gt gt LER I Lj I soar array WING SKIN CONNECTOR 2164 5 Figure 3 3 5 Spacecraft Launch Configuration Outer Dimensions Ref VEX T ASTR TCN 00349 EADS V Issue 2 Rev 0 Date 06 02 2004 ASTRIUM GNUS express Page 3 11 7S 0 Im 0 5 germ qs sj Figure 3 3 6 Spacecraft Launch Configuration Outer Dimensions Ref VEX T ASTR TCN 00349 FADS Issue 2 Rev 0 Date 06 02 2004 ASTRIUM nus 5 Page 3 12 3 4 DEPLOYED CONFIGURATION After spacecraft separation from FREGAT the solar wings are released by pyrocutter firing and deployed thanks to hinge springs The wings are latched at end of deployment Once deployed the wings can rotate around Ys axis each one is moved by a drive mechanism The MAG boom is released and deployed once the spacecraft is in final orbit Release is ensured by a pyrocutter dep
6. TL aA A E E AENEA NT 11 24 1191 Failure Management of Intelligent 11 24 11 9 2 Failure management of CDMS non intelligent modules 11 25 12 RELIABILITY AND REDUNDANCY ARCHITECTURE 12 1 REDUNDANCY uuu S ninan AEA ENN E ET ANEA TEER ETAR 12 1 122 i RF USE IL Um TA 12 2 153 RBDLNDANC o n sido II RCNSRUR CEDEEIERUISEIREUEDEEIETUPEAEPHEOENURI DRE EHI VIDI UT KA 12 4 8 SPACECRAFT BUDGETS nist 13 1 PROPELLANT DUDGETS uuu u Ten MANUI MORE a ws ate pORSA DETUR 13 1 Ce POWER BU I cents 13 4 Lo LLEIED NEC 13 7 J II Mv M MEINE 13 8 KO ONS D CICLOS 13 9 KT E eee nee 13 11 Astrium VEX T ASTR TCN 00349 S 7 2 Rev 0 ASTRIUM exoress 1 7 BBHABERIIY lli PI a EUM UEM M DEUM MM EAP Astrium Ref VEX T ASTR TCN 00349 k V ma S Issue 2 Rev 0 YA p Date 06 02 2004 NERONE SNUS express Bags Li 1 SPACECRAFT OVERVIEW 1 1 DESIGN DRIVERS The major objective of the Venus Express spacecraft design 15 to cope with Venus Express mission requirements with extensive reuse of Mars Ex
7. The helium tank is loaded via the fill amp drain valve The test port is used for pressure regulator performance testing on the ground Ref VEX T ASTR TCN 00349 EADS Issue 2 Rev 0 Date 06 02 2004 ASTRIUM r express Page 6 3 FDV1 ae HPTD PVNO 15 PVNC 01 x K PVNC 02 gt lt 2 5 gt lt LFLVI LFLV2 gt lt 1 2 gt lt 1 s eee eee PVNC 04 PVNC 06 gt lt gt lt FVV9 FVVS FDV10 gt lt gt lt PVNC 09 PVNC 08 TP12 Sep LPTD 03 LPTD 04 PVNO 17 C PVNO 16 fs lt T Jr weanap B I n N PVNC 11 13 AT 1B 2A LIE EISA Mars Express Schematic 30 01 2001 Issue D MAIN TP15 EC TP14 KEY KEY RCT Reaction Control Thruster Dual Valvc 1A to 4A Primary RCTs 1B to Redundant RCTs Low Pressure Transducer TLV Thruster Latch Valve LFLV Low Flow Latch Valve Flow Control Valve Helium Pressurant Tank PYNC Pyrotechnic Valve Normally Closed Fill and Drain Valve PVNO Pyrotechnic Valve Normally Open Fill and Vent Valve NTOI Nitrogen Tetroxide Propellant Tank Test Port 2 Mono Methyl Hydrazine Propellant Filter Tank Figure 6 1 1 Propulsion System Schematics The propulsion schematic is fully identical to Mars Express one Ref VEX T ASTR TCN 00349 2 Issue 2 Rev
8. 06 02 2004 10 18 Each instrument has its own autonomous SW located in the instrument electronic units The command and control of the payloads is performed by the dedicated Payload Management function of the DMS SW The physical interface of the DMS PM with the instruments is the Remote Terminal Unit RTU Data exchange between the payloads and the DMS software are performed by means of packetised TM TC both for commands housekeeping and science telemetry data Commands from the Ground are routed by the DMS software to the payloads through the RTU and the OBDH bus Housekeeping data from all the instruments are transmitted from the RTU to the DMS SW through the OBDH bus Scientific data from low rate payloads PFS ASPERA MAG SPICAV are transmitted from the RTU to the DMS SW through the OBDH bus Scientific data from high rate payloads VIRTIS and are directly transferred to SSMM through TM packets on the IEEE 1355 link Ref VEX T ASTR TCN 00349 EADS Issue 2 Rev 0 Date 06 02 2004 ASTRIUM r express Page 11 1 11 FDIR PRINCIPLES 11 1 FDIR HIERARCHY The Venus Express automated FDIR function as for the other function of S C is highly recurrent from Mars Express automated FDIR function with local adaptations due to the design changes between MEX and VEX The Venus Express automated FDIR function ensures that the anomalies are handled on board with the g
9. Ref VEX T ASTR TCN 00349 EADS Issue 2 Bou c6 Date 06 02 2004 ASTRIUM express pads 4040 Figure 11 4 1 shows the DMS System Layer context with respect to CDMS and AOCMS o Dependable Clock FDIR Safeguard Context SGM DMS Ground 7 TC Decoder TFG i Management CMD o High Power CPDUs o Reconfiguration PROM ol Le L Clock Reference o System Re initialisation MISSION TIME LINE MTL o Ground TC Processing o Autonomous Sequences AOCS Supervision Monitoring AOCS PM and Control o Attitude Management o Ephemeris Data 4 Modes Transitions convergence criteria time out o Autonomous Star Pattern Recognition STR o VOI MEBM Lu a LL lt 0 o x 2 a x gt X gt O x O Z lt 2 3 gt x 2 o D m 9 lt Autonomous MAT9921C Three Axis 4 6 Attitude ED AOCS Units Units Knowledge Figure 11 4 1 DMS System Layer Context 9 Ref VEX T ASTR TCN 00349 P Issue 2 Rev 0 EADS M uL Date 06 02 2004 GNUS express Page 11 11 11 5 SUBSYSTEM LEVEL The DMS S W executes Application Programs TCS RF communications Separation Sequence which include their own FDIR mechanisms The RF communications TT amp C FDIR mechanisms are summarised in Section 11 6 The Po
10. TC Link Monitor TLMAP e TC Link Recovery TLRAP Safe Mode The FDIR on this level relies on essential DMS and AOCMS functions and mechanisms such as to achieve and secure the mission On board time which is permanently available to trigger nominal mission operation events and time stamp any anomaly situation Mission timeline MTL which activate the spacecraft modes and functions according to the mission phase Mission ephemerides updated on board the S C to improve its autonomy especially in case of opposition phase and stored into SGM provide all data necessary to master Earth Sun and Venus position This improvement is a modification with regards to Mars Express S C which ephemerides were updated by the ground These data are used to autonomously re establish ground link and control programmed spacecraft attitude according to the mission phase Attitude measurement with respect to an inertial reference frame is based on star pattern recognition using an on board star catalogue It is performed by the star tracker which supports attitude measurement autonomy since no a priori attitude knowledge 15 necessary and can be used In all mission phases Redundancy management using healthy on board resources Cross strapped access paths allow to confine the reconfiguration to the concerned unit Non volatile memories RAM amp EEPROM SGM allow to restart the spacecraft from a safe context
11. The information bit rate can vary from about 9 bps as a minimum and can be up to about 228 kbps CDMU limitation As a baseline the lowest bit rates will be used in case of emergency whilst the highest ones will be used operationally Coding The telemetry is Reed Solomon coded before being the subject of a convolutional encoding Ref VEX T ASTR TCN 00349 D Issue 2 Rev 0 EADS F Date 06 02 2004 ASTRIUM GNUS express Page 8 8 8 7 RADIO SCIENCE OPERATIONS The Venus Express spacecraft host a radio science experiment so called VeRa Venus Radio science assembly This experiment consists in the TT amp C subsystem whose stability performances are enhanced by the introduction of an Ultra Stable Oscillator USO as illustrated in Figure 8 1 1 The science investigations use the Radio Links between the spacecraft and the Earth in emitting simultaneously not modulated carriers in X band and S band from the S TX and X TX of the same transponder via the High Gain Antenna The Radio Links can operate in two modes Q the One way mode no uplink is involved the stability of the downlink carrier frequencies 15 provided by the USO Q The Two way mode the downlink carrier frequencies are coherent with the uplink whose frequency stability is provided by hydrogen masets dual frequency downlink Two way radio link TWOD Hydrogen Maser X band uplink telecommanding S band downlink
12. VEX T ASTR TCN 00349 e Rev O 06 02 2004 us 5 VENUS EXPRESS l SPACECRAFT DESIGN REPORT 0232 Jacques Barri re Mechanical architect Jean Michel Benvenuto 26 02 04 4 02 26 3 Electrical architect Alain Clochet Payload Manager Bernard Gillot Data Handling architect Prepared by Nicolas Midan Z2blo2 o4 Rams architect Sonia Penalva Thermal architect Patrice Riant AOCS architect Christian Sibilla Software architect Dominique Trillard Operations Architect Verified by Thomas Schirmann system Engineering Manager Robert Osswald roved b App Product Assurance Manager Josian Fabrega Application authorized by Venus Express Project Manager Astrium Ref VEX T ASTR TCN 00349 S Issue 2 Rev 0 ASTRIUM exoress an SUMMARY This document summarises the main design characteristics of the Venus Express spacecraft The detailed description is presented in the Venus Express User Manual Volume 2 Document controlled by Astrium Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 ASTRIUM exoress n aad DOCUMENT CHANGE LOG Revision 31 01 03 _____ First issue for PDR 06 02 04 CDR Issue Astrium Ref VEX T ASTR TCN 00349 2 Issue 2 Rev 0 ASTRIUM r express Dat
13. ee ee a ee FEIHFFISFEFIFIIFFFEFFEI GEFIEFTEREFIUFIKFFEIEFT Spacecraft In Orbit Configuration Figure 3 4 2 Ref VEX T ASTR TCN 00349 gt EADS Issue 2 Rev 0 Date 06 02 2004 ASTRIUM r express Page 3 14 VEX T ASTR TCN 00349 2 0 06 02 2004 4 1 NS ASTRIUM r a express 4 THERMAL CONTROL DESIGN The spacecraft thermal control is in charge of maintaining all spacecraft equipment within their allowed temperature ranges during all mission phases The equipments fall into two categories gt the collectively controlled units for which the heat rejection and heating capabilities design and accommodation are provided by the spacecraft thermal control gt the individually controlled units self provided with their own thermal control features coatings selection heaters insulators for which the spacecraft thermal design controls the thermal interfaces within the required ranges 4 1 THERMAL CONTROL DESIGN APPROACH The thermal control design of Venus Express spacecraft is based on a robust and passive concept with a maximum commonality with Mars Express but some system and design modifications are implemented to cope with the Venus inner orbit and hot environment The two main discrepancies with the Mars Express missions are gt stringent thermal environment with a high solar constant almost 4 times higher than
14. 0 p Date 06 02 2004 enus express Page 6 4 e Low Pressure Gas Side The low pressure gas side comprises 1 a pressure regulator 2 non return valves 3 a pair of low flow latch valves 4 a low range pressure transducer 5 normally closed pyrovalves and 6 test ports and fill amp vent valves This section has a MEOP of 20 bar controlled by the regulator which senses downstream pressure The regulator is the dual series redundant type This design features both a primary and a secondary regulator In the event of failure of the primary regulator 17 bar regulated pressure the secondary regulator will control the system pressure at 17 5 bar Another feature of the regulator is the dynamic flow limiter fitted at its inlet The limiter restricts the rate of rise of downstream pressure in the unlikely event that the firing of a normally closed pyrovalve in the high pressure gas side delivers helium too rapidly for the regulator to respond During main engine firings and over the on orbit life of the spacecraft there exists a potential for propellant vapours to migrate from the propellant tanks toward the pressure regulator To prevent possible mixing of fuel and oxidant vapours a pair of non return valves is fitted in the helium lines to both the fuel and the oxidant sides of the system To increase reliability each pair of non return valves is arranged in series providing two inhibits to prevent mixing of propellant vapours
15. 2 Rev 0 p Date 06 02 2004 ee GNUS CXpress Page 134 13 2 POWER BUDGETS Power budgets are presented for all mission phases Launch amp Early Operation Phase Near Earth Commissioning and Interplanetary Cruise Phases Venus Orbit Insertion Phase Payload Commissioning Routine Operation and Extended Operation Phases Units power consumptions are based either on Mars Express FM measurements for platform units and on conservative assumptions for payloads Heating power budget is based on CDR thermal analyses They demonstrate adequacy of Solar Array sizing with comfortable margins Battery charging after Launch Phase can be performed at maximum charge rate Large allocation can be offered to Payloads in Near Earth Commissioning Phase Solar Array will be very tolerant to pointing error up to 45 approximately once in orbit around Venus Robustness to BDR or battery failure 1s demonstrated in all modes and phases Moreover peak power consumption remains lower than maximum capability of 2 BDR extended to 600W for Venus Express In most modes except science observation Peak power consumption remains lower than PDU maximum capability extended to 750W for Venus Express in all modes Ref VEX T ASTR TCN 00349 FADS Issue 2 Rev 0 Date 06 02 2004 ASTRIUM r express Page 13 5 hemdtoal oo 100 1047 102 1589 12 2026 2012 180 no
16. Cross strappings are implemented so as to improve significantly the spacecraft reliability figure or the operational flexibility but shall not bring additional risks by increased design complexity Ref VEX T ASTR TCN 00349 W S Issue 2 Rev 0 EADS M p Date 06 02 2004 SNUS CXD express Page 12 2 12 2 REDUNDANCY SCHEMES Mission critical functions and system alarms are implemented with a majority voting structure to automatically filter inadvertent and latent failures Vital functions power and essential functions supporting reconfiguration process HPC TC decoder RM SGM RX Transponders are implemented in a hot redundancy structure to provide redundant resources without any configuration commands All other EEE functions are supported by stand by redundancies except for critical phases for which the ground can select a hot stand by redundancy mode to avoid outages due to reconfiguration e g 2 STR 4 RW 2 IMU Hot stand by redundancy mode can also be used as an operational flexibility for some multiple failure cases within a dedicated unit ex possibility to use the AIU in hot stand by mode after the failures of thruster 1N and thruster 2R These different redundancy schemes can be implemented for each unit internally or externally At unit level three types of redundancy packaging are provided Internal redundancy optimised for mass and volume aspects segregation rules betwe
17. 06 02 2004 Page 4 6 ASTRIUM r a express 4 3 THERMAL CONTROL OVERVIEW 4 3 1 Thermal control features The unit temperature control is achieved through the use and the selection of flight proven materials used on numerous spacecraft The key features of the thermal control are presented on Figure 4 3 1 and summarised as follows gt Optical solar reflectors radiators electrically grounded to the aluminium sandwich panel face sheet reject internal heat dissipation toward Space gt Dedicated radiators are provided for the VIRTIS OM integrated to the OM and PFS O payload requiring operation at low temperature gt The platform high dissipative units are mounted on the panels directly behind the radiators to provide a good conductive path from unit to panel Thermal doublers ensure spreading of heat over the radiator areas gt Heat pipes are implemented under the PCU and PDU units to spread the high PCU thermal dissipation gt A high emissivity finish is used inside the spacecraft when required to maximise the radiative heat transfer to the radiators gt Thermal straps are used to connect the Reaction wheels and some payload units needing a dedicated radiator PFS O cryo I F SPICAV SOIR VIRTIS OM coolers VIRTIS ME gt Dedicated radiator paddle for X reaction wheels with OSR deflector to reduce heat load by reflection and IR coupling on the paddle gt Multi Layer Insulation MLI is used to minimise
18. 39 39 33 40 47 42 49 4 D4 38 35 49 56 j 45 59 54 Table 13 2 1 LEOP power budgets Near Earth Commissionning amp Cruise Phases Normal modes K sn np qn nn Comm s Manceuvre Control Acquisition Acquisition DMS total W COM s total W AOCS total W CPS total W Thermal total W Instruments total W 00 00 00 00 Power total W 481 58 58 57 59 60 Table 13 2 2 Near Earth Commissioning Cruise Phases power budgets Ref VEX T ASTR TCN 00349 5 ASTRIUM re express Issue 2 Rev 0 Date 06 02 2004 Page 13 6 AOCStotal 193 1033 133 69 99 1155 75 nss 69 155 Table 13 2 3 Venus Orbit Insertion Phase power budgets Payload Commissioning Routine amp Extended Ops Phases Backports modes Gus cem Fl i 6 526 526 52 6 52 6 52 6 52 6 52 6 AOCGSttal W 83 83 883 893 83 S3 75 155 69 n5 1 7 38 44 33 36 s9 56 47 3 m 49 So ss so 59 s a Table 13 2 4 Venus In orbit Phase power budgets Ref VEX T ASTR TCN 00349 EADS gt Issue 2 Rev 0 p Date 06 02 2004 ASTRIUM express Page 13 7 13 3 LINK BUDGETS 15 6 bps 2000 bps 2000 bps 2000 bps 0 013 AU 0 39 AU 0 43 AU VISUM ourou m 0 15 AU
19. AlU SAS RCS SADM STR RWA IMP reconfiguration Units 9 Figure 11 1 1 Venus Express FDIR Hierarchy Ref VEX T ASTR TCN 00349 EADS Issue 2 Rev 0 F Date 06 02 2004 ASTRIUM egenus express Page 11 2 FDIR Level Order of Magnitude of Time Comment Response D is ose miner _ T TC Recovery several minutes hours Some delays as per Ground programming S C Safe Mode 30 minutes S C Go to SAM to AOCMS lt 20 seconds Ground ops afterwards CDMS Reconfiguration 30 minutes S C Go to SAM to AOCMS lt 30 System Re Initialisation seconds Ground ops afterwards Figure 11 1 2 Order of Magnitude of FDIR Delays vs Levels When mapped onto the hardware and software architecture of Venus Express the FDIR hierarchy can be depicted as in Figure 11 1 3 Ref VEX T ASTR TCN 00349 2 EADS ASTRIUM re express Issue 2 Rev 0 Date 06 02 2004 Page 11 3 Reconfiguration Modules Level 4 reconfiguration modules 1 Reconfigure Level 3 DMS Processors System software 2 RTU power Supplies eeeeeeeeeee ROS O MMM Ue SC Ge ct IH Mmmm III Level 2 S S AOCS S W amp AP software Sloe Level 1 DMS processors Reconfigure SSMM amp SSMM wg users NN 5 Level
20. Sun Pointing Star Acquisition Phase SPP Phase StAP Biased Pointing Switch to SBM no control SUIS BPP for SA deployment Safe Hold Mode E SHM Earth Acquisition EAIP Earth Acquisition EAP Earth Pointing nit EPIP Wheel i Earth Sun Off Loading Pointing EPP MAT 11515 Normal Mode MTL Figure 5 3 3 Attitude acquisition reacquisition sequence Ref VEX T ASTR TCN 00349 E S 2 Issue 2 Rev 0 cr Date 06 02 2004 ASTRIUM l express 5 11 Normal Mode NM The Normal Mode is designed to enable all the nominal operations of the mission except the trajectory corrections It uses Star Tracker measurements and gyros for the attitude measurements and reaction wheels for the control In order to reduce the fuel consumption and the orbit disturbances and to have the best pointing performances during Venus observation phases The Normal mode contains several sub phases used to cover all the functionalities required during the operational mission The Gyro Stellar pointing on Ephemeris Phase GSEP is optimised to ensure to the spacecraft a deterministic and quasi inertial attitude with respect to the Sun and the Earth directions with a pointing accuracy compatible with Earth communication needs HGA 1 or HGA 2 axis pointed towards the Earth and a sun pointing of the Solar arrays This phase 1s used during the cruise phase and in the
21. The potential for propellant vapour migration is particularly relevant to the long cruise to Venus during which the main engine is isolated and the thrusters are fired only intermittently Therefore further protection 15 provided by the addition of a pair of parallel redundant low flow latch valves in the low pressure gas side These allow the pressurisation lines to be closed off for the greater part of the time The latch valves are located upstream of the normally closed pyrovalves This eliminates any risk of debris possibly generated by the firing of the pyrovalves entering the latch valves The low range pressure transducer is used to monitor the pressure in this section In flight and during testing on the ground The purpose of the normally closed pyrovalves is to keep the propellant feed subsystem isolated from the pressurant subsystem until the time comes to bring the propellant tanks up to regulated pressure 17 bar As in the high pressure side of the pressurant subsystem the normally closed pyrovalves are positioned in parallel for redundancy The fill amp vent valves and test ports are used on the ground e g to vent gas during propellant tank filling to pressurise section volumes and to obtain system pressures via ground instrumentation Ref VEX T ASTR TCN 00349 2 Issue 2 Rev 0 p Date 06 02 2004 enus express Page 6 5 6 1 2 Propellant Feed Subsystem The propellant feed subsystem commonly referred to
22. i e not connected if the operations of the S C as for Mars Express incorporate the orientation of the cryogenic panels towards sunlight in case of any identified radiator performance problem linked to icing VEX T ASTR TCN 00349 s 2 Rev 0 06 02 2004 2 10 ASTRIUM l 5 2 3 7 VMC The VMC science payload for Venus Multi spectral Camera 1s devoted to the cartographic imaging of the Venus atmosphere directed towards UV markings O2 airglow and near IR emission of surface and lower atmosphere VMC 15 basically used In connection with VIRTIS but with a higher imaging frequency while its spectral resolution is far lower VIA TIS mosaic 2 hours A 2 6 VMC Camera typical VIRTIS Images for comparison The VMC instrument is a CCD integrated camera of about 1 6 kg mass which houses optics CCD read out circuitry derived from the Mars Express SRC processing electronics and power converter derived from the Rosetta OSIRIS design The specificity of the VMC design lies with its standard cooled CCD Kodak matrix down to 40 C at the lowest optically fed with four miniature optical lenses about 5 mm in diameter and further linked to a high performance highly compact electronics A copper band is attached to the CCD base which is used to extract the heat out of the chip to the Peltier cooler and further on to the close S C wall wall on which th
23. spot shielding 15 to be introduced e And general behavior wrt UV radiation in relation with Sun vicinity in particular as the experiment MLI is concerned the S C MLI solution is to be re conducted here VEX T ASTR TCN 00349 s 2 Rev 0 06 02 2004 2 4 AS ASTRIUM r a express 2 3 2 MAGNETOMETER The Magnetometer experiment or MAG looked as a complement of ASPERA is made of two measurement sensors external to the S C both controlled by a centralised electronics MAGE The design of the sensors of fluxgate type 15 re conducted from the Rosetta design One sensor MAGIS 15 mounted directly on the S C top floor to measure the S C magnetic near proximity field The second MAGOS is fitted to the extremity of a 1 meter deployable CFRP boom to assess the mid proximity field This very specific accommodation allows for a vector combination of the MAGIS and MAGOS measurements to retrieve the undistorted field no magnetic constraint imposed to the S C The boom is attached to the S C top floor for launch and released when around Venus The Magnetometer Boom and One Sensor Unit from Rosetta Lander In addition the potential implementation of the external sensor at the far end of the S C HGA2 structure 1s possible as a fall back option Despite its early impact on the procurement scheme of the new HGA2 through CASA this programmatic approach allows to insure MAGOS
24. x Issue 2 Rev 0 f DS p Date 06 02 2004 AS ASTRIUM enus 5 Page 7 18 5 7 4 2 The Venus Express environment 15 strictly controlled so as to guarantee the S C auto compatibility the Launcher Spacecraft compatibility S C auto compatibility Radiated and Conducted emissions are limited as much as possible in order to offer margin with respect to the susceptibility limits of the units This enables to guarantee the good functioning of the units and payloads once mounted on the S C Compared to Mex no systematic over shielding is applied on the bundles as Venus express S C does not feature any longer the Vensis payload Launcher and Spacecraft compatibility Radiated emissions of units powered during launch are limited to offer margin with respect to launcher susceptibility limits The Spacecraft susceptibility is limited as much as possible in order to offer margin with respect to the Launch vehicle emissions Ref VEX T ASTR TCN 00349 EADS Issue 2 Rev 0 p Date 06 02 2004 ASTRIUM ENUS exoress Page 8 1 8 COMMUNICATIONS 8 1 OVERVIEW The communications with the Earth can be performed either S Band or X Band in accordance with ESA Standards The RF communication S S consists of a redundant set of transponders using S band and X band for the uplink and the downlink Depending on the mission phase the transponder can be routed via RF
25. 0 avionics ee 7 amp platform units with or w o _ Platform units payloads Figure 11 1 3 FDIR Hierarchy Mapping onto Venus Express H W and S W ASTRIUM express Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 Date 06 02 2004 Page 11 4 11 2 GROUND FDIR SUPPORT The Ground constitutes the ultimate level for FDIR management Venus Express supports the ground fault treatment as follows data to allow ground driven fault isolation down to the least reconfigurable item trend analysis long term analysis correlation between multiple parameters high level commanding of CDMU data processing resources processor modules clocks CDMU I O last chance bit checks of data recorded during the execution of autonomous sequences DMS capability to store and execute Master Time Line MTL Short MTL TC files Ground Identified Failed Unit Table GIFUT and Ground Selected Unit Table e g GSUT The GIFUT and GSUT uploaded by the Ground are taken into account by the DMS S W upon system re initialisation to serve ground requests to modify individual entries of the Processor Identified Failed Unit Table PIFUT and Processor Selected Unit Tables PSUT in a real time manner to store and execute ground recovery procedures which may constitute ground answers to on board failures having led to the S C Safe Mode unanticipated or unrecoverable on board fault co
26. AOCS Budgets Technical Note Ref VEX T ASTR TCN 1120 Several error sources which are not properly represented in the simulations are added to the simulation results e The STR spatial noise The STR misalignment e The worst case occultation effect on the drift of the attitude e The impact of ASPERA scan mechanism e The guidance algorithm error The occultation effect takes into account the duration of occultations during Venus Express mission The guidance algorithm error has been updated with respect to Mars Express figures taking into account the new algorithms proposed for Venus Express and the error analysis performed for the Venus mission Note that the ground guidance error 15 not considered A provision for Star Tracker degradation due to radiations effect has been included in the budgets The pointing performances of the antenna during Earth pointing phase are computed in Normal Mode and are also valid for the SHM EPP wheel control after convergence if the gyros configuration has not been changed the budget assumes calibrated drifts The impact of the Aspera scan mechanism is much smaller on Venus Express than on Mars Express due to the absence of Marsis Vensis antenna Ref VEX T ASTR TCN 00349 EADS Issue 2 Rev 0 p Date 06 02 2004 ASTRIUM egenus express Page 13 10 Error source X axis Y axis Z axis VEX simulation results 0 0114 0 0027 0 0039 STR degradation radiations 0 0014 0 0
27. Block Diagram Ref VEX T ASTR TCN 00349 EADS Issue 2 Rev 0 Date 06 02 2004 ASTRIUM r express Page 9 12 Ref VEX T ASTR TCN 00349 E AD S Issue 02 Rev 0 Date 06 02 2004 ASTRIUM express Page 10 1 10 SOFTWARE ARCHITECTURE The purpose of this chapter is to describe the overall Venus Express on board software which as shown below is composed of several software running on processors located on different hardware units This chapter mainly focuses on the interfaces between these software rather than the description of each individual software Such a description is found in the global description of their hosting hardware as shown below DMS application and common Section 1 Chapter 4 Volume 4 Section 7 software Section 1 Chapter 1 Volume 7 Section 3 SSMM software Volume 4 Section 2 AOCS application Volume 4 Section 8 STR Volume 4 Section 5 IMP Volume 4 Section 5 p uut Ref VEX T ASTR TCN 00349 SA Issue 02 Rev 0 Date 06 02 2004 V ASTRIUM r a express tee Page 10 2 10 1 SOFTWARE COMPONENTS AND ASSOCIATED FUNCTIONS The Venus Express on board software manages the payload and the platform It is composed of several software running on different processors The Venus Express software SW is made of the DMS SW and the AOCMS SW together with the CDMU SW Firmware the SSMM SW the STR SW the gyros SW the Transponder SW an
28. Control Valve Thrust 1A Fuel Thrust 1A Fuel Thrust 2A Oxydiser Thrust 2A Oxydiser Thrust 2A Fuel Thrust 2A Fuel Latch Control Valve Flow Control Valve Latch Control Valve Flow Control Valve Latch Control Valve Flow Control Valve Thrust 1B Fuel Thrust 1B Fuel Thrust 2B Oxydiser Thrust 2B Oxydiser Thrust 2B Fuel Thrust 2B Fuel Latch Control Valve Flow Control Valve Latch Control Valve Flow Control Valve Latch Control Valve Flow Control Valve Thrust 3A Oxydiser Thrust 3A Oxydiser Thrust 3A Fuel Thrust 3A Fuel Thrust 4A Oxydiser Thrust 4A Oxydiser Latch Control Valve Flow Control Valve Latch Control Valve Flow Control Valve Latch Control Valve Flow Control Valve Thrust 3A Oxydiser Thrust 3A Oxydiser Thrust 3B Fuel Thrust 3B Fuel Thrust 4B Oxydiser Thrust 4B Oxydiser Latch Control Valve Flow Control Valve Thrust 4A Fuel Thrust 4A Fuel Latch Control Valve Flow Control Valve Thrust 4B Fuel Thrust 4B Fuel Figure 12 3 3 Venus Express propulsion reliability block diagram Ref VEX T ASTR TCN 00349 EADS Issue 2 Rev 0 Date 06 02 2004 ASTRIUM 5 Page 12 9 Battery 1 PCU APRI PCU APR2 PCU MEA APCI MPPT APCI MPPT Battery 2 Solar Array Solar Array PCU APRI PCU APR2 PCU MEA Deployment Wing APC2 MPPT EA Deployment Wing APC2 MPPT EA2 Battery 3 PCU APRI PCU A
29. SOYUZ FREGAT ONLY OTA 2 E SECTION A A SOYU AIRING TYPE S F TOP VIEW FREGAT ONLY 03 00 UPDATED WITH DM 2316 VEX BET 1708703 02 00 UPDATED WITH OM 2661 VEX BET 04402703 01 00 FIRST issue WRT OM 2605 vex 8 01 03 Rev Zone Libell Nom Date Dess in v rifi Vis Qual it BET 01 08 03 3395 TBC Format Echelle Project R f rence projet AO Radical iii VENUS EXPRESS astrium concECRAFT SQYUZ FREGAT S LAUNCHER F CE DOCUMENT EST LA PROPRIETE DE ASTRIUM ET NE PEUT ETRE REPRODUIT SANS AUTOR SATION Type N DOCUMENT Ed Rev Folio IRDDT0047618 103 00101 02 11 2 2267 03 10 2003 16 04 19 01 03 04 Figure 1 4 1 Spacecraft coordinate axes J Ref VEX T ASTR TCN 00349 gt Issue 2 Rev 0 EADS p Date 06 02 2004 ASTRIUM enus express Page 1 10 Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 Date 06 02 2004 Page 2 1 MA ASTRIUM snus express 2 SCIENCE PAYLOAD 2 1 THE VENUS EXPRESS MISSION AND PERSPECTIVES The first phase of Venus exploration started in the sixties with the unprecedented series of early Venera Pioneer Venu
30. The SPICAM Instrument Core from Mars Express The SPICAV design features two independent channels e The SPICAM channel Mars Express repeat use of the MEX spare parts operating in both occultation and Nadir pointing modes its only adaptation for VEX lies with its mechanical interface with the new SOIR channel e And the SOIR channel the Sun occultation experiment using a dedicated IR optics detector assembly together with cooling unit located on top of SPICAM This SOIR channel is to be linked to a S C radiator to allow for its proper cooling down VEX T ASTR TCN 00349 S P 2 Rev 0 06 02 2004 ASTRIUM r express 2 35 VeRa The VeRa Venus Radio science Assembly measurements use an enhanced S C RF system In horizontally sounding the Venus atmosphere in directing the HGA towards Earth as for nominal S C communications during specific occultation paths This will occur when the Venus atmosphere region under study 15 optimally placed between the S C and Earth Consequently the VeRa experiment does consist in e Incorporating a USO Ultra Stable Oscillator together with its connecting harness within the S C RF system through a direct link between USO and the TRSP s transponders e And operating the S C in specific conditions making the use of the RF link as a support science radio sounding measurements of the Venus atmosphere The VEX USO design 15 similar optimally
31. W via normal TM or via events The STR respectively the SSMM may also decide to perform internal S W transition to Stand By Mode or Init Mode which is seen from the AOCMS respectively the DMS S W as a specific failure such a unit S W reconfiguration being monitored by the AOCMS respectively the DMS S W As for the non intelligent units the additional surveillance implemented within the AOCMS respectively the DMS S W can be individuated to these units only or larger surveillance which monitor consistency between several units for example monitoring of converters voltages or consistency checks of their outputs with the outputs of other sensors Failure Isolation and Recovery is fully managed by the AOCMS respectively the DMS S W 1 e the decision to switch off a unit and or to switch on its redundancy in case of major failure is taken by the AOCMS respectively the DMS S W only This decision can also lead to reconfiguration to Safe Mode Note that as for the non intelligent units the DMS S W acts as slave of the AOCMS S W relative to the AOCMS intelligent units Ref VEX T ASTR TCN 00349 P Issue 2 Rev 0 EADS m F Date 06 02 2004 11 9 2 Failure management of CDMS non intelligent modules Some CDMU internal modules are accessed only by the DMS S W FDIR of these modules is therefore fully managed by the DMS S W itself For modules which may be used by both the AOCMS S W and the DMS S W these
32. and AOCMS SW are organised in layers Figure 10 4 1 and Figure 10 4 2 The Kernel software layer providing hardware interfaces and basic software services The General Services software layer providing TC services TM services Generic Services equipment management and processing OBCP execution and control services The Sub System software layer providing SSMM management Platform management Thermal Control function RF Communications function Pyrotechnics Payload management AOCMS management and remote PM management on the DMS PM This software layer provides Sensors and Actuators management and SADM management on the AOCMS PM The System software layer providing Mission Time Line management and System Autonomy and FDIR management on the DMS PM and providing AOCMS Algorithms and Modes management and AOCMS Autonomy and FDIR management on the AOCMS PM Ref VEX T ASTR TCN 00349 EADS V Issue 02 Rev O0 Date 06 02 2004 ASTRIUM ENUS express Page 10 10 DMS PROCESSOR MODULE Mission Time Line Autonomy amp FDIR System SW layer Management Management Sub System SW SSMM Platform Payload AOCS Remote PM layer Management Management Management Management Management DMS Equipments General services Management amp Processing SW layer equipments drivers IEEE 1355 Kernel SW layer OBDH Data Bus MAT 10834 Common SW Figure 10 4 1 DMS Software Layered Breakdown AOCS PROC
33. and FDIR information to the DMS SW The AOCMS SW sends mode configuration patch and dump orders to the STR SW The STR SW sends 3 axis attitude STR health status and dump information to the AOCMS SW The AOCMS SW sends configuration orders to the gyros SW The gyros SW sends health status angles and velocities information to the AOCMS SW The SSMM resource is shared by the DMS AOCMS component and the Payload instruments simultaneous data handling concurrent downlink store or routing are directly managed at SSMM unit level TC Time Line Mode Config Reconf OBCP Patch and Dump Orders M5 SW AOCS SW lt o TM Acknowledge Dump Monitoring Info FDIR Info Events log Files TC Equipment Management Dump TC Equipment 9 a SSMM Patch Files Info Files Config Mode TM Equipment and Dump Patch and Dump eee Angles Requests 3 Axis Attitude Velocities Status Health Status Status Dump SSMM SW STR SW GYROS SW Figure 10 2 2 Venus Express on board software internal interfaces Ref VEX T ASTR TCN 00349 A Issue 02 Rev 0 Date 06 02 2004 V ASTRIUM express tee Page 10 7 10 3 PROCESSOR MODULE FIRMWARE Each Processor Module comprises a software stored in PROM called PM Firmware designed to perform PM initialisation PM health status verifications software loading minimum communication handling with a test co
34. and main error voltage Solar Array Power Control When the available array power exceeds the total power demand from the PCU including the battery power charge the Array Power Regulator APR will perform the main bus regulation based on the MEA control line signal The regulator function is a buck type switched regulator which will leave the surplus energy on the array by increasing its input impedance A MPPT function will automatically take over the regulation control of the Regulator when the MEA sional enters the BCR or BDR control domain The MPPT monitors the array voltage and current and controls the Regulator to provide that specific input impedance which will derive the maximum electrical power available on the array The MPPT function finds the maximum power point by oscillating the APR input impedance slightly around the impedance providing the maximum power Each APR function comprises 3 individual Array Power Regulators configured as two out of three hot redundant regulators The active regulators share equally the requested power transfer to the main bus Each of the two solar array wings has its own individual APR function to allow individual tracking of the maximum power point Battery Power Control Each of the three batteries has its own dedicated Battery Charge Discharge Regulator BCDR function in the PCU As the battery voltage is lower than the regulated main bus voltage the BDR 1s a conventional
35. and the X panel for the payload equipments These sides of the spacecraft are the most favourable areas being most of the time protected from the direct sun inputs always for the X side The rest of the spacecraft is insulated with Multi Layer Insulation blankets to minimise the heat exchange and the temperature fluctuations EADS ASTRIUM r express servation Sun direction 7 Sun direction Y Z 0 to 180 2 2 Figure 4 1 1 Sun aspect angle during Venus orbit operations Ref VEX T ASTR TCN 00349 issue 2 2 Rev 0 Date 06 02 2004 Page 4 3 ASTRIUM r express The spacecraft external units Platform and Payload units are thermally decoupled from the spacecraft and provided with their individual radiator when needed The electrical heater system allows raising the temperature of the units above their minimum allowed limits with temperature regulation functions provided either by mechanical device or by the onboard software The main design modifications with regard to MEX consist in gt gt gt gt Using as much as possible low solar absorptance and low ageing coatings Optimising the radiators area and improving their efficiency ITO SSM replaced by OSR Enlarging temperature qualification of some units PDU CDMU WIU Enlarging the X reaction wheels radiator paddle and replacing SSM by OSR In addition OSR deflectors tilted of 2 are fixed arou
36. around Mars The Venus albedo and planet fluxes are imposed by the spacecraft orbit and attitude While the planet IR flux is far lower than around Mars and constant the albedo flux 15 significant during the operation phase around the pericentre gt Using MEX platform as it is Venus inner orbit does not allow keeping a wall in the shadow during the entire mission when the S C 1s pointed to earth To cope with these new constraints the system architecture for the earth telecommunications has been modified to keep the sun direction in an allowable area determined by the thermal requirements of the spacecraft units and subsystems A trade off has been conducted to determine the sun aspect angle limitation on each wall see Figure 4 1 1 The conclusion of this study lead to implementation of a smaller HGA on the VEX top floor in order to restrict the sun illumination possibilities during communication in the X Z quadrant only The VEX bus configuration is very recurrent from MEX Except some payloads that have been changed the structure and units layout is the same than MEX The X shear wall is dedicated to the payloads with the X closure panel accommodating the cryogenic radiators The platform units are mounted on the Y sidewalls and the propulsion subsystem 1s fitted on the Z floor and the X shear wall The heat rejection toward space is performed using radiators mainly on the Y panels for the platform internal units
37. atmosphere vertical composition and the still to be demonstrated surface volcanism Thus a combination of spectrometers spectro imagers and imagers covering a wavelength range from UV to thermal IR along with a full plasma analyser should be able to map and analyse the entire Venus atmosphere from about 200 km or even higher altitude to the surface through the fine atmospheric transparency window Most of the instruments are re using design and or spare hardware originated from either Mars Express or Rosetta program As for Mars Express the Venus Express instrument complement has been confirmed as being highly suitable for such a planetary mission Fitted onto a spacecraft bus designed from the original Mars Express one but thoroughly adapted to the specific Venus thermal environment the science payload will gather from orbit and over a typical 500 Earth day overall mission duration a consistent data set of measurements which will be made available to the science community VEX T ASTR TCN 00349 S 2 Rev 0 06 02 2004 2 2 ASTRIUM express 2 2 THE VENUS EXPRESS SCIENCE PAYLOAD The Venus Express science payload 15 contained In seven instruments with a total mass of slightly less than 90 kg Most of the instruments are re using design and or spare hardware originated from either Mars Express or Rosetta program Principal Investigator Payload Objective ASPERA 3 Kiruna Sweden ili
38. e ROMAP Austria PFS CNR Rome Italy Infrared Fourier Spectroscopy SPICAV Mars Express JL Bertaux SA CNRS Atmospheric spectrometry by Star or SPICAM Verri res France Sun Occultation in the Ultraviolet to Mid Infrared Range Range VeRA Rosetta RSE B Haeusler UniBW Radio Sounding of the Atmosphere Muenchen Germany VIRTIS Rosetta VIRTIS P Drossart ObsPM Atmosphere Surface Meudon France and Spectrographic E from the G Piccioni IASF CNR Ultraviolet and Visible to Mid Rome Italy Infrared Ranges Mars Express W Markiewicz MPAe Ultraviolet and Visible Multi spectral HRSC SRC Lindau Germany Camera and Rosetta OSIRIS The Venus Express Science Payload The VMC instrument 15 the only fully new system although re using some design heritage The other instruments are deeply based on their prime Mars Express or Rosetta background 2 3 SCIENCE PAYLOAD DESIGN While originated from other programs both design and accommodation of the Venus Express science payload has experienced some iteration work from the kick off and subsequent SRR Have been mainly concerned ASPERA MU no more located onto top floor but on Y wall VeRa no more mounted onto shear wall but onto Y wall SPICAV with the SOIR channel thermal coupling and VIRTIS with the definition of tts interface structure Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 Date 06 02 2004 Page 2 3 ASTRIUM l e
39. express 5 12 Trajectory correction modes Main Engine Boost Mode MEBM The Main Engine Boost Mode enables to perform the large trajectory corrections using the 414 N Main Engine The efficiency is greater in this mode for large manoeuvres due to the better specific impulse of the Main Engine and the higher force delivered reducing gravity loss effects During this mode however the disturbance torques due to the Main Engine misalignments are very high and a large attitude depointing transient is observed at the beginning of the manoeuvre This mode is therefore not suitable for small AVs of a few m s The MEBM uses 2 IMUs and the 10N thrusters for attitude control The reaction wheels the STR and the SADE are OFF in MEBM to avoid failure management of these non mandatory equipment that may jeopardise the completion of the manoeuvre The Solar Array orientation is specific for this mode perpendicular to the Thrust and 15 reached in normal mode before the MEBM On Mars Express the late discovery of a large discrepancy between the centre of mass and the Z axis of the Spacecraft lead to a significant increase of the disturbing torques during the Main Engine Boost It has been decided to introduce in the software the capability to control the spacecraft with 8 thrusters The principle of this function is to use both nominal and redundant thruster branches at the same time to double the spacecraft control capacity and be abl
40. figure presents the Wing Block diagram with different sections and strings PANEL 1 4 positive power lines Section 2 Section 1 d Section 4 4 x 3 negative power lines Grounding Figure 7 2 1 1 VEX Wing Block Diagram The SA circuitry is as follows Number of wings 2 Number of panel per wing 2 Numbers of sections per panel 2 Number of strings in parallel per section 6 Number of SCA s in series per string 22 Two redundant bleed resistors each 20 kOhms per panel achieve short circuit protection Each string includes one blocking diode By design each cell provides a by pass diode Ref VEX T ASTR TCN 00349 2 Issue 2 Rev 0 p Date 06 02 2004 ASTRIUM ENUS express Bi 3065 The S A dimensioning has been achieved in order to guarantee 800 W min in earth vicinity and 1100 W min in Venus vicinity The strings are designed in order to cope with a max S A Voc of 80V and a min Vmpp of 32V The cells with a Venus flux have been considered with an efficiency of 25 This lay out leads to a max S A current of 18A wing The following table summarises the SA estimated power values o Pme Power Neat Earth 820 Power BOL Venus 1490 Power EOL Venus 1400 Ref To Issue EA DS p Date ASTRIUM GENUS XPress Page The following figure shows the VEX Power Generation block diagram PCU Array Power Regu
41. guidance function In addition on board ephemeris calculation 15 now offered in order to improve the spacecraft autonomy Propulsion A bi propellant reaction control system 15 used for orbit and attitude manoeuvres by either a 400 N main engine or banks of 10 N thrusters It is the same as Mars Express except the pipe routing had to be modified due to change of pyros valves In addition propellant load must be increased because AV requirement is more stringent than for Mars Express Electrical design The Electrical Power generation 15 performed by Solar Arrays GaAs cells shall be used instead of Silicon cells because they are less sensitive to temperature and to radiation Since they are also more efficient 2 panels per wing are sufficient to meet the power requirements 820W BOL is achieved in Earth vicinity which is the sizing case At least 1380W EOL is then available in Venus vicinity which is much more than the requirement 1100W Once around Venus the spacecraft is thus very tolerant to mispointing of the Solar Array up to 45 Power storage 1s performed by 3 Lithium Ion batteries 24 Ah each as for Mars Express J Ref VEX T ASTR TCN 00349 m 2 Issue 2 Rev 0 _ Date 06 02 2004 ASTRIUM p ENUS express Mostly because of heating budget augmentation PCU modification is needed in order to increase capability from 250W to 300W A standard 28 V regulated main bus is off
42. heat flow from non radiating areas and to minimise the thermal distortions The conductive surfaces of all thermal blanket layers are electrically grounded gt Cold coatings on LVA ring external part Clear Sulphuric Anodisation and on SADM panel flange White paint gt Heaters and thermal blankets on the liquid bi propellant system prevent propellant freezing and enable to optimise propellant management gt Both software and hardware controlled heaters are implemented Appropriate redundancy is included for all heaters thermistors and thermostats to prevent single point failure in the thermal control function gt Low conductive stand offs for the appendages and external payload units minimise heat transfer to the spacecraft main body VEX T ASTR TCN 00349 Ref 06 02 2004 Date EADS Uu 9 gt 2 ASTRIUM Embossed Kapton MLI rarer eee ee eee ee ee ee eee PFEEREFIZFFEFPFKEI KIZFFERFKFIFFIZPFFXPFSXZJZ Pere cee ee PZFFIiIZSFIZSFFSNSgsgIsrISF ee ee ee ee CPPS SES OP PPP PP F PPP ee ee ee ee E Pe eee a Ce ee ee ee SPSS OP ECE EE PPP a ee PET OP PCP E Frere RW OSR deflectors OSR and cells stripe Y walls radiators RW paddle Y walls radiators Virtis OM cryo radiator PFSO cryo radiator Clear sul
43. individual parameters against pre defined thresholds SM or a logical combination of SMs Functional Monitoring FM for which recovery actions are identical The ground has the possibility to enable disable a SM or a FM this can be used to avoid generation of events and out of limit actions for instance after a 1 or even 2 failure of a device Ground Overriding It is possible from the Ground to override switch over that may have been automatically performed on board 1 e the Ground can force the S C to come back to any flight configuration Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 zi Date 06 02 2004 11 3 H W RECONFIGURATION OF CENTRAL COMPUTING AND COMMUNICATIONS The DMS and AOCMS software execute on safe processors as configured by the CDMU Reconfiguration Modules RM and High Power Command Modules HPCM The reconfiguration handling in the CDMS is divided into two parts 1 Monitoring and reconfiguration request function RM 2 Reconfiguration sequencer and HPC generation HPC HPC Reconfig Reconfig sequencer sequencer RM1 RM2 RM3 RM4 Alarm Alarm Alarm Alarm Alive Alive Alive Alive Figure 11 3 1 H W implemented Reconfiguration of CDMU Processors The monitoring and reconfiguration request function 15 located in the RM Each RM contains a watchdog which is periodically re armed by the DMS PM The time out period for each RM differs between 531 ms and 57
44. initial Star Tracker mode management of Mars Express assumed that the sensor remains in Tracking Mode and the FDIR detected a failure when the Tracking of the sensor was lost There was some exceptions to this rule during slew manoeuvres during occultations of the sensor by the planet or during specific AOCS modes when the acceleration seen by the Star Tracker is too high as for instance in OCM The occultations by the planet are managed through 2 different ways depending on the Mode During Normal Mode the occultation are predicted on ground and managed onboard through occultation tables providing the beginning and end dates of these events During Safe and Hold Mode SHM a specific algorithm has been defined able to manage autonomously the occultation periods and trying to restart the Star Tracker in acquisition when the Tracking is lost due to an occultation This attempt will be performed during a delay compatible with the largest duration of the occultation and no failure will be declared by the FDIR during this period This algorithm is necessary to cope with orbit degradations related to thruster pulses in SAM and SHM leading to inaccurate or wrong predictions of occultation tables A modification has been proposed for Venus Express and finally implemented on both Mars Express and Venus Express in order to extend this autonomous occultation management to all AOCS Modes using the Star Tracker 1 e the Normal Mode the Thrust
45. is located on top floor Xs Ys corner the LGA2 under the lower floor on the Xs side The HGA1 antenna orientation has been modified compared to Mars Express for coverage improvement in Venus Express mission the LGA1 15 oriented towards Zs The LGA2 oriention has been kept unchanged from Mars Express 17 5 deg tilted towards Xs an orientation towards Zs would have generated a local interference with the MGSE LGAI HGA2 Cryo Face LGA2 Figure 8 2 1 Antennas accommodation onto the Spacecraft Ref VEX T ASTR TCN 00349 4 Issue 2 Rev 0 EADS Date 06 02 2004 ASTRIUM r exoress Page 83 8 3 TT amp C CONFIGURATIONS ACCORDING TO THE MISSION PHASES In the LEOP phase communications are done in S band via the LGAs When the LEOP phase is completed few days after Launch communications are done in X Band via one The proper HGA 15 selected along the mission depending on the planets configuration to avoid sun illumination on the cryo face Xs as illustrated on the following figure End of Mission HGA switching 2nd Venus day 0 6 0 4 HGA2 selected on HGAI selected on Inferior Conjunction side Inferior i Superior Conjunction side 0 Conjunction 8 0 2 Start Start 39 4 Venus day Venus day 4 0 6 0 8 vol AU Start of scie
46. no RG 0 41 AUno RG 1 2 AU 250 bps 0 039 AU 1 35 AU New Norcia 35m 1 06 AU no RG 1 72 AU no RG Cebreros35m_ NA DSN BWG NA 0 11 AU DSN 70m 1 72 AU no RG 19044 bps 19044 bps 0 013 AU 0 39 AU Kourou 15m 0 43 AU 0 13 AU 0 013 AU no RG 0 45 AU no RG 0 64 AU no RG 0 18 AU no RG 0 039 AU 1 35 AU 38091 bps 19044 bps New Norcia 35m 0 039 AU no RG 1 36 AU no RG 1 72 AU 0 68 AU 45710 bps 19044 bps Cebreros 35m N A 1 72 AU 0 78 AU mpm 38091 bps 19044 bps DSN 34m BWG 1 72 AU 0 69 AU 0 11 AU DSN 70m 1 72 AU N A N A 0 11 AU no RG Note that when not expressly mentioned the link budgets have been checked with ranging mode Table 13 3 1 Venus Express Link Budgets J Ref VEX T ASTR TCN 00349 2 s Issue 2 Rev 0 p Date 06 02 2004 ASTRIUM ENUS express Page 13 8 13 4 LINES BUDGETS The table 13 4 1 provides the PDU power lines allocation Table 13 4 1 PDU Power Lines Budget No spare line is available The table 13 4 2 provides the pyro lines allocation Table 13 4 2 Pyro Lines Budget The detail of lines allocation 15 provided in document Acquisition and command allocation list Ref VEX T ASTR TCN 00349 w V T S Issue 2 Rev 0 p Date 06 02 2004 AS ASTRIUM enus express Page 13 9 13 5 POINTING BUDGETS The pointing performances for Venus Express are computed from Venus Express simulation results They are also detailed in the Venus Express
47. of the system maximising reliability There 15 no appreciable loss of performance because the thrusters are capable of operation over a much wider range of inlet pressures than the main engine Propellant is delivered to the main engine and thrusters the propellant feed subsystem which 15 supplied with helium by the pressurant subsystem Each of these contains pipework with associated fittings and CPS units 6 1 1 Pressurant Subsystem The helium pressurant subsystem 1s commonly referred to as the gas side It may be considered as two sections the high pressure gas side and the low pressure gas side High Pressure Gas Side The high pressure gas side comprises 1 a 35 5 litre helium tank 2 normally open and normally closed pyrovalves 3 a high range pressure transducer 4 a fill amp drain valve and 5 a test port This section has a maximum expected operating pressure MEOP of 276 bar and during all ground operations and through launch it 15 isolated from the pressure regulator by a pair of normally closed pyrovalves These are arranged parallel to each other providing redundancy in the design Helium usage 15 monitored by the high range pressure transducer The purpose of the normally open pyrovalve is to isolate the pressurant tank from the rest of the CPS after the final main engine firing There 15 no need for a redundant normally open pyrovalve since successful tank isolation is not critical to the mission
48. operations of the mission are performed in the Normal Mode which enables all the scientific operations around Venus but also the cruise pointing and all the attitude manoeuvres necessary before and after an orbit control manoeuvre for instance The trajectory correction manoeuvres are performed through 3 Modes The Orbit Control Mode for small trajectory corrections performed with the 10N thrusters The Main Engine Boost Mode for trajectory corrections performed with 415N engine The Braking Mode BM is specifically designed for Aerobraking phase if such a phase is necessary to achieve the final orbit using the force produced by the air drag when passing through the Venus atmosphere at orbit pericentre The Thruster Transition Mode TTM 15 used as a smooth transition between the thruster controlled Modes OCM and BM and the wheel controlled modes Normal Mode Ref VEX T ASTR TCN 00349 FADS gt Issue 2 Rev 0 Date 06 02 2004 ASTRIUM p ENUS express Launch Init From Any Mode except pa MEBM amp SBM Ephemeris Anomaly GSEP Fine Pointing Inertial Phase FPIP Fine Pointing Main Engine Accuracy Boost Mode Phase MEBM FPAP Normal Mode NM MAT 11497L Thruster Transition Mode Mode TTM BM Automatic Attitude acquisition or transient phases TC M
49. re worked to allow for operational flexibility EEPROM replacing PROM and science data production enlargement imaging mode The other module P is kept unchanged Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 Date 06 02 2004 Page 2 6 ASTRIUM L express 2 34 SPICAV The SPICAV instrument for Spectroscopic Investigation of the Characteristics of the Atmosphere of Venus is an imaging spectrometer operating in various wavelength ranges UV 0 12 to 0 32 um near IR 0 8 to 1 5 um SPICAM SIR design and mid IR 2 to 4 4 um SOIR design This payload is primarily designed to perform horizontal sounding direction along the tangent to the planet measurements of the Venus atmosphere either from star or Sun occultation In pointing either specific bright stars from catalogue or the Sun the instrument will acquire significant spectra of these astronomical objects when virtually entering the Venus atmosphere because of the relative movements of S C wrt the planet and will allow for their comparison This will then give clues to analyse the Venus atmosphere content e g water content SO2 volcanic effects HDO escape trap phenomena through differential thus self calibrated measurements In complement the instrument may operate in the Nadir mode SPICAM only thus pointed towards the Venus planet surface for vertical sounding IWAN WAS L N ih REN et
50. reception of ground TC In case no TC has been received during a time period an automatic reconfiguration 15 executed to recover the TC link No on board monitoring on the TM link is implemented since the loss of the TM does not endanger the spacecraft integrity In case of TM loss detected at ground level the recovery actions are under ground conttol s d Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 Date 06 02 2004 ASTRIUM exoress Page 8 6 8 5 UPLINK ON BOARD RECEPTION The communication from the ground station s to the spacecraft is performed in S Band or X Band Uplink Frequencies The frequencies for the uplinks are 2114 335648 MHZ DSN 17 for S Band 7165 780092 MHZ DSN 17 for X Band Ranging signal For ESA ground stations the ranging signal is in accordance with the ESA Ranging standard For the NASA DSN ground stations the ranging signal is in accordance with the DSN handbook Modulation The RF uplink signal which is modulated as NRZ PSK PM on a 16 KHz sinusoidal subcarrier is routed towards a diplexer performing frequency discrimination and then to the Dual Band Transponder input The transponder performs carrier acquisition and demodulation and transmits the extracted signal to the Data Handling for further processing TC bit rates The following telecommand bit rates are handled by the Venus Express Spacecraft as provided by the CDMU design 7 8125 bps 15 6
51. s PH P tees 7 16 FI COMMUNIL ATIONS c 8 1 9 1 EVER IB ee a ene ayo eve 9 1 9 2 ANTENNAS ACCOMMODATION 9 2 8 3 CONFIGURATIONS ACCORDING TO THE MISSION PHASES 8 3 Ot REDUNDANCY AND TDIR PRINCIPLE 8 4 8 4 1 LOA 8 4 8 4 2 QI 8 4 8 4 3 S 8 5 92 DPUNK ON BDOARD RECEPTION oncesi a 8 6 8 6 F Il RR TOS 9 7 9 7 KADIO SCIEBN C B OPERATION 8 8 8 8 IIT NUI NIMM nU ME 9 9 DATA HANDLING ARCHITECTURE 545 9 1 9 1 HPU 9 2 PM IL P O 9 4 9 3 H iri p T RoHS 9 8 Astrium Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 Date 06 02 2004 Page vii ASTRIUM express POLUD TAT MASS MEMORY eee 9 10 10 SOFTWARE ARCHITECTURE uuu uu e
52. solar array Three fold aut two panel arrays 1450 W power at Venus w One axis orientation GaAs triple function Telecommunications Mars Express enhanced with a second X band 1 3 metre diameter fixed HGA SX band 0 3 meter diameter fixed HGA2 X band only Two LGA for omnidirectional coverage wv S X band uplink v S X band downlink v X band TWTA 65 watts RF v Dual S band transponder 5 watts RF j enus express Structure IBI Mars Express primary structure Machine ring from forging v Al honeycomb panels Three shear panels T Outer and shear walls equipment payload mounting Propulsion Mars Express rebuilt v Bi propellant system vw Two 267 litre propellant tanks v 595 kg propellant capacity 540 kg needed One 35 litre pressurant tanks v Pressure regulated and blow down operational modes v Eight attitude thrusters w One 400N main engine Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 Date 06 02 2004 Page 1 6 Attitude contro and measurement Mars Express rebuilt Three axis stabilised using four 12 Nms reaction wheels v Two wide Field of View Star Trackers with autonomous Star Pattern recognition and measurement capabilities Two inertial measurement units including three ring laser gyros and three accelerometers each Two coarse sun sensors Power Mars Express rebuilt with enhanced PCU v Fully regulated 28V b
53. step up regulator design while the BCR 15 a step down regulator The batteries are charged at constant regulated current at 3A until a command selected End of Charge EOC voltage limit is reached The BCR will then maintain the battery at this EOC voltage level EADS ASTRIUM Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 Date 06 02 2004 Page 7 11 u ENUS express VEX PCU Power Limitations Venus express PCU power limitations are as follows output power 2 APR x 3 APC x 250W 1500W 750 W per Wing output power 3 BDR x 300 W 900W 300W per Battery charge current Battery The following figure shows the PCU functional block diagram APR Wing 2 APR Wing 1 P out 250W Solar Array Wing 2 Solar Array Wing 1 APC2 P out 250W 2 3 Hot Redundancy APCI P out 250W x l Filter DC DC Converter Lx T LCL OFF Ref LCL OFF Ref gt BDR Failure detection Overvoltage Overcurrent Zero Demand
54. suppliers are referenced to the unit housing Q the housing of each unit is locally referenced to the structure Q all return currents flow through dedicated lines wire return policy no current shall be intentionally flown through the spacecraft structure The spacecraft structure is used as low impedance equipotential ground plane The main advantages of DSPG are that it combines the prevention of low frequency interferences provided by Single Point Grounding SPG together with the avoidance of high frequency interferences brought by multiple ground systems Low frequency emissions ate avoided by insulation of the primary power from both the equipment housings and the secondary power supplies prevention of ground loops High frequency interference generated by capacitive coupling with the grounding system is minimized by grounding of the equipment secondary power supplies referenced directly to the structure for each equipment Differential balanced interfaces are used as a rule between different units to offer robustness to common mode between units Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 p Date 06 02 2004 ASTRIUM GNUS Xpress Page 1717 The following figure shows the VEX Distributed Single Point Grounding diagram F 3 TAM AE POWER CONVERTER LI 7 MECHANICAL GROUND E Ni PRIMARY OV F lt gt SECONDARY 0 V Figure 7 4 1 VEX DSPG Diagram Ref VEX T ASTR TCN 00349
55. to each DC DC Power Converter within the users In addition power lines are also dedicated to users which draw directly power from the power bus without any need for a DC DC Converter This is the case of the 10 N Thrusters FCV Flow Control Valve the FCV of the Main Engine coils and of the Latch Valves These lines are routed via the AOCS Interface unit Each power line is switched and protected by means of a Latch Current Limiter LCL An LCL 15 a solid state latching switch which also acts as a protection device in case of over current Should the current through the LCL exceeds the trip off current the device will enter into current limitation In case current limitation continues for more than a given trip off time in the order of 16 ms the LCL will open to isolate the failed unit from the Spacecraft power bus The LCL are actuated using Delayed Memory Load commands DML which are the same as Memory Load Commands but delayed by 100msec within the PDU control module and provide isolated ON OFF status and primary current telemetry It is however not possible for all units to isolate from the power bus for some units in case of overcurrent This is the case for the CDMU and for the Dual Band Transponder Receivers These units shall never be switched off and shall be able to recover autonomously in case of return to normal conditions Primary power is distributed to them through Foldback Current Limiters FCL These are devices identical in es
56. will be ensured as far as possible when the failure is not too critical and if the first objective is not endangered Note that during the Venus insertion manoeuvre the mission continuation is as critical as the Spacecraft safety Hierarchy in the FDIR A classification of the surveillances 15 performed in two ways by their level in the functional chain and by their criticality in a dedicated Mode The level classes of the surveillances are the following Alocal monitoring concerns health checks at unit level A functional monitoring is built from a comparison between several units global monitoring concerns high level vital spacecraft capabilities The reconfiguration actions decided by the onboard Software strongly depend on the level of the surveillance triggered a local monitoring enabling usually a reconfiguration at unit level while a global monitoring triggering often leads to the reconfiguration of several H W subassemblies The criticality of the surveillance in a dedicated mode is described by the capability to continue the current operations or not after reconfiguration Q Non Emergency Surveillance NES enables to continue in the current mode If this surveillance triggers and if redundant units are available the AOCS will stay in the same mode after the adequate reconfiguration Q Emergency Surveillance ES leads to a switch in the safe mode after the reconfiguration Two other classes have
57. 0 06 02 2004 4 5 NS ASTRIUM r express Sun Earth Mars Spacecraft geometry the X side of the Spacecraft is oriented away from the Sun over the complete Venus orbit both during Nadir pointed science phase and Earth pointed communication phase This allows keeping the camera and the spectrometer temperature around 170K and 190K respectively during the Planet observation The connection to the radiators is performed by thermal straps the radiators being themselves decoupled from the rest of the spacecraft using thermal blankets and insulating stand offs Every payload aperture is protected with baffles in order to avoid sun entrance inside the Spacecraft Payload external units like ASPERA and MAG are individually controlled units They are directly exposed to the external environment and they have to withstand larger temperature ranges than the standard units A special care is taken to their accommodation on the spacecraft to provide them the softer thermal environment They are as far as possible insulated from the spacecraft to reduce the interface fluxes Their coatings are selected and trimmed to fulfil the thermal requirements The spacecraft interface temperature has a very limited influence on their thermal behaviour gt CPS The internal propulsion equipments tanks fluid lines valves pressure sensors and pipes are radiatively and conductively isolated from the structure and provided with their own the
58. 00349 EADS Issue 2 Rev 0 Date 06 02 2004 ASTRIUM T 3 3 Top Floor Ys Sidewall 55 Q 1 A Xs Closure Panel Xs Closure Panel Ys Sidewall Xs Ys Figure 3 2 1 Spacecraft Exploded View The Venus Express mechanical bus overall configuration is identical to Mars Express Ref VEX T ASTR TCN 00349 gt FADS Issue 2 Rev 0 Date 06 02 2004 ASTRIUM T exoress Page 3 4 Xs Shearwall Ys Shearwall Figure 3 2 2 Spacecraft Exploded View Core part Ref VEX T ASTR TCN 00349 2 EADS Issue 2 Rev 0 Date 06 02 2004 ASTRIUM Fr 5 Page 3 5 Top Floor Xs Closure Panel vs Xs Ys Sidewall Xs Closure Panel Figure 3 2 3 Spacecraft Exploded View Ref VEX T ASTR TCN 00349 LE S Issue 2 Rev 0 _ Date 06 02 2004 AA ASTRIUM ra eXpIeSS La 36 3 3 LAUNCH CONFIGURATION In launch configuration the solar wings are stowed on the Ys sidewalls thanks to four holddown points same configuration as Mars Express and the MAG boom 15 stowed on the top floor thanks to a launch lock device Spacecraft dimensions in that configuration are depicted in the following pages They are compatible with the spacecraft transport container The spacecraft is planned to be launched on SOYUZ FREGAT same configuration as for Mars Express The adapter between Fregat and the spacecraft is supplied by Astrium The spacecraft dim
59. 014 0 0014 STR spatial noise 0 0075 0 0017 0 0075 STR Payload alignment 0 0337 0 0337 0 0337 TOTAL 0 0367 0 0342 0 0351 Table 13 57 Attitude Estimation Error budget during Nadir observation deg 36 rror source X axis Y axis L Table 13 5 2 Pointing Error budget during Nadir observation deg 36 No occultation Occultation GSEP Error source 0 03 _ Table 13 5 3 Pointing Error budget during Earth pointing phases deg 50 J Ref VEX T ASTR TCN 00349 46 S Issue 2 Rev 0 p Date 06 02 2004 GNUS XPress Page 13 11 13 6 SOFTWARE BUDGETS The best estimates for Venus Express software budgets are the figures measured on Mars Express until Venus Express measurements are available They are reported in table 13 6 1 CDMS Shared Resources SGM EEPROM 64 Kw 12 5 19 6 SGM RAM 64 Kw 23 4 36 6 PROM Cassette 512 Kw 310 7 60 7 DMS SW Physical Memory PM RAM kW 318 8 62 3 CPU Venus OBServation ms 790 9 79 1 Earth VISIbility No STR invest ms 799 2 79 9 96 Earth VISIbility STR diagnosis ms 704 1 70 4 Physical Memory PM RAM 512 kW 174 0 34 0 CPU 1 Hz cycle worse case 1000 ms 733 0 73 3 8Hz cycle worse case 1000 ms 763 0 76 3 Table 13 6 1 Venus Express Software Budgets Ref VEX T ASTR TCN 00349 EADS Issue 2 Rev 0 p Date 06 02 2004 ASTRIUM egenus express Page 13 12 13 7 RELIA
60. 1 lt q S Band Rx 2 P02 X Band Rx 2 X Bd Feeder P04 DIPLX TC to CDMU d Demod 112 P01 P02 P01 S Band Tx 2 ma 4 to CDMU X Band Tx 2 Figure 8 1 1 RF Communications block diagram Ref VEX T ASTR TCN 00349 4 Issue 2 Rev 0 EADS p Date 06 02 2004 ASTRIUM exoress Page 82 82 ANTENNAS ACCOMMODATION High Gain antennas The two fixed high gain antennas HGA1 and HGA2 are accommodated on the spacecraft to allow the Earth pointing of the spacecraft while satisfying the thermal constraints keep the sun direction in the Zs Xs quadrant The is accommodated on the Xs closure panel in identical location as the Mars Express The 5 pointing inclination from the Xs direction has been kept unchanged from Mars Express The antenna diameter has been reduced from 1 6 m Mars Express to 1 3 m The HGA2 is accommodated on the top floor Its pointing direction is 5 from the Xs direction symmetric to HGA1 with respect to Zs axis It is accommodated on the Xs side of the top floor in the aim to not affect the payload fields of view Low Gain antennas The two low gain antennas LGA1 and LGA2 are accommodated on the spacecraft for omnidirectional coverage Both are located in same place as on Mars Express The LGA1
61. 1 7 m width and 1 4 m height reinforced by 3 shear walls and connected to a cylindrical Launch Vehicle Adapter The solar array is composed of two wings providing a symmetrical configuration favourable to aerobraking technique and minimising torques and forces applied on the arrays and the drive mechanisms during the Venus insertion man uvres performed with the main engine Within the overall integrated design of the spacecraft four main assemblies are planned to simplify the development and integration process 1 the Propulsion Module with the core structure 2 the Y lateral walls supporting the spacecraft avionics and the solar arrays 3 the Y X shear wall and the lower and upper floors supporting the payload units and 4 the X lateral walls supporting the High Gain Antenna X and the instruments radiators X Thermal control passive control is kept as for Mars Express However external coatings shall be modified in order to minimize the thermal flux entering the spacecraft In particular black MLI shall be replaced by multi layer Kapton MLI as for Rosetta In addition OSR shall be used on the lateral radiators and on the solar arrays On LVA ring alodine shall be replaced by clear sulphuric anodisation AOCS as for Mars Express Attitude Control is achieved using a set of star sensors gyros accelerometers reaction wheels 10N thrusters Modification of communication strategy lead to minor changes in the
62. 10 NS ASTRIUM r a express 4 3 3 Multilayer insulation The satellite external barrier is a Multi Layer Insulation MLI blanket which completely isolates the Z floor and the X closure panel High Gain Antenna support panel from the cold space environment the sun illumination MLI is also used on Y and X radiator panels and on Z lower floor in order to adjust the radiative area to the units need This adjustment may be performed after thermal balance test if needed without major impact The composition of the blankets number of layers spacers 1s governed by the location on the spacecraft and the temperature level seen by the external layer The Venus environment induces important issues on the MLI materials selection especially for the X and Z walls that may be continuously sun illuminated The first one concerns the outer layer The aim is to select as far as possible a cold coating presenting a low solar absorptance combined to a high emittance and stable to UV and proton radiation low discoloration These considerations led to trade off non usual coatings like Astroquartz Beta cloth and Nextel for the MLI external layers of these 2 walls But development and qualification status of these materials prove to be not advanced enough with respect to schedule to select one of them as the baseline A solution using classical and well known materials 15 preferred and a white coating is foreseen only as patc
63. 2 p d Ref VEX T ASTR TCN 00349 r P Issue 2 Rev 0 uL Date 06 02 2004 TLRAP The TC link monitoring is triggered when no TC has been received in a time period The recovery strategy 15 the following One or two reconfigurations are tried according to the current TT amp C configuration a After each attempt the TC link recovery is checked Ifthe TC link is recovered TLRAP 15 exited Ifthe TC link Is not recovered after the reconfiguration attempts the ULGA mode is entered Ultimate Low Gain Antenna where both LGA s are selected for the communications and the Safe mode 15 entered in this case SAFCAP 15 bypassed Two types of TT amp C configurations are considered by TLRAP Foreseen operational configurations defined as interpretable by TLRAP For these ones two reconfiguration attempts are tried before entering the ULGA mode that offers a 2 failures tolerance These configurations are therefore recommended for use by the Ground Unforeseen operational configurations defined as not interpretable by TLRAP For these ones only one reconfiguration attempt 15 tried by establishing a unique default configuration In the VEX design the interpretable configurations have been modified to take into account the implementation of the HGA2 the configurations established by SAFCAP which are different wrt MEX and the configurations required for Radio Science operati
64. 25 bps 250 bps 1000 bps and 2000 bps As a baseline the lowest bit rates will be used in case of emergency while the highest ones will be used operationally s d Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 Date 06 02 2004 XW ASTRIUM exoress Page 8 7 8 6 DOWNLINK A high data downlink capability is required considering the large data volume generated by the instruments Nevertheless the large spacecraft to Earth distance limits the downlink capacity The downlink of the telemetry data to the ground stations can be performed in either S band or X Band Down link Frequencies The frequencies for the downlink are 2296111111 MHZ DSN 17 for S Band 8419 074073 MHZ DSN 17 for X Band Modulation The telemetry is transmitted to the Dual Band Transponder as PCM PSK PM on a square wave subcarrier for transmitted information rates as high as 22 853 Kbps Two subcarrier frequencies are used 8 192 kHz for low bit rates and 262 144 kKhz for high bit rates For information rate equal or greater than 28 566 Kbps the PCM SP L is directly modulated on the carrier This signal is phase modulated in either S Band or X Band by the Dual Band Transponder The ranging signal in the ranging channel of the transponder directly phase modulates the carrier When simultaneous ranging and telemetry is performed the two signals are added prior to phase modulation of the downlink carrier bit rates
65. 8 ms In steps of 15 6 ms When watchdog timer elapses a reconfiguration request is sent to the reconfiguration sequencer The RM generates the request alternatively between sequencer A and B In order to start the reconfiguration the sequencer has to receive requests from at least two RM When two requests have been asserted to the sequencer the Reset line is asserted to all units within the CDMS This will halt any on going activities in the CDMS The sequencer then starts to execute a sequence of High Power Commands The sequence is stored in a PROM that is read by the sequencer for each reconfiguration Once the sequence is completed duration lt 250 ms the Reset line Reconfiguration Request signals are automatically de asserted and the system 15 re started In case the reconfiguration was 9 Ref VEX T ASTR TCN 00349 A P Issue 2 Rev 0 uL Date 06 02 2004 successful 1 e the RM watchdog timer elapses again a new reconfiguration is started from the other sequencer This means that in case a sequencer is faulty the system will handle this automatically Should one Reconfiguration Request signal be left asserted failure case this does not trigger a new reconfiguration The DMS S W re arms the RM watchdog through the PM Alive signal every 250 ms counting 64 Hz RTC interrupts Not re arming the watchdog is a system level response to a variety of faults which re
66. AS New solar cells mounted on a new ceramic backing Mars Express unit Reaction Wheel 4 bearing Momentum Reaction wheels Telecom Sat and Mars Teldix 12 Nms 0 075 Nm Express Unit SADM 2 Stepper motor with gear Twist capsule Mars Express and Rosetta Kongsberg unit except speed levels Figure 5 2 2 AOCS Harware units J Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 _ Date 06 02 2004 LA ASTRIUM exoress Page 5 5 5 3 AOCS MODE ARCHITECTURE AOCS Mode logic The AOCS includes several modes for attitude acquisition reacquisition purposes for the nominal scientific mission operations and for the orbit control The Attitude acquisition or attitude reacquisition sequence is ensured by 2 modes The Sun Acquisition Mode pointing the X axis of the spacecraft and the Solar Arrays towards the sun Safe Hold Mode completes the acquisition and provides the final 3 axes pointing or HGA 2 axis towards the Earth This attitude acquisition sequence is used nominally after launch and also after a large trajectory correction manoeuvre performed with the Main Engine The same sequence is used 1 case of failure during a Software Safe Mode or a Hardware Safe Mode It is an automatic sequence including all the operations of both Modes except during the first acquisition where the flexibility is let to the ground to introduce 1 or 2 stand by points The nominal routine
67. BILITY BUDGETS The reliability budget of each Venus Express module payload instruments excluded but including spacecraft resources amp interfaces provided to instruments is summarized in table 13 7 1 for the two missions Nominal and Extended Module Nominal mission Extended mission 699 days 1185 days COMS 0 986 0 967 DMS AOCS UL 0 812 POWER 0 992 0 981 PROPULSION 0 9955 0 9955 Table 13 7 1 Venus Express Reliability Budgets These reliability budgets have been derived from Mars Express taking into account following updates Implementation of second High Gain Antenna Venus Express mission phases duration Update of reliability calculation for Propulsion Module Ref VEX T ASTR TCN 00349 FADS Issue 2 Rev 0 Date 06 02 2004 ASTRIUM Anus 5 13 13 Ref VEX T ASTR TCN 00349 EADS ASTRIUM 55 Date 06 02 2004 Page A Issue 2 Rev 0 DISTRIBUTION LIST Overall document Summary Action Information ESTEC Mac Coy D Schmidt R ESOC Accomazzo A EADS ASTRIUM Barriere J Benvenuto JM Billard C Bonnamy O Brugnera E Caspar R Clochet A Cousty JP Etienne C Fabrega J Fournier R Gabilan C Gillot B Harjani D R Hostein B Loche D Midan N Osswald R Penalva S Riant P Schirmann T Sibilla C Trillard D Documentation
68. Battery TBD Subset of commands High Power Commands VEX T ASTR TCN 00349 Rev 0 64 HPCs generation request Generator gt Reconfiguration A VF 4 RMs some internally routed others routed to other equipment Transponders receivers Decoder TC segment transfer p EGSE 2 5 4 Virtual Channel parameters gt 2 PMs redundant CDMU gt Transponder emitters Transfer Frames Generator and bitrate loading gt TM packet transfer Serial a SSMM SSMM branch branch 2 PMs redundant CDMU Figure 9 2 1 CDMU block diagram 1 2 3 and 4 Ref VEX T ASTR TCN 00349 DS Issue 2 Rev 0 _ Date 06 02 2004 ASTRIUM uL A CXD GSS Page 9 6 CDMU Memory Policy Each Processor Module 15 baselined with the three following memory types Start Up Programmable Read Only Memory PROM Electrically Erasable Programmable Read Only Memory EEPROM Random Access Memory RAM The Start Up PROM contains the Firmware and monitor software and also the RAM and EEPROM test software It consists in 32 K of 24 bit words The Start up PROM is enabled after Processor Module reset initialisation and Built In Test BIT When Start Up is completed the Start Up PROM is disabled with a dedicated CPU internal XIO The RAM contains the software to be run i
69. E EER 5 5 5 4 AOC S GENERI PUNCTI ETEEN 5 15 2 3 ACS MODE TST 5 20 5 6 HIGH GAIN ANTENNA Q u i 5 24 PROPULSION SYSTEM ARCHITECTURE 6 1 6 1 DESIGN DESCRIPTION u ul l u au 6 2 O44 GUC REET EERE 80 000 0000080 80 0 0509 92080 90 S ErerenT er 6 2 DI ICH NI MPH 6 5 6 2 RK 00 perenne Mier erect rte ise E nen tr year seer nro Mine nner oor ete 5 6 6 ELECTRICAL AND POWER ARCHITECTURE 7 1 Ta I PP TE e PRUNUS T 7 1 7 2 PLE ad BIC LIU erence ee Un u D La JN lul d 7 4 luu ca IU P P OOU 7 8 Z 2 3 f m i EPEA MERET Sun TTT eee ene eee 7 10 7 2 4 Mam Dus Power SD DUO EER 7 42 7 3 7 14 7 4 l j eT 7 16 7 4 1 LUDUM a apa skp pcre ME 7 16 7 4 2
70. ESSOR MODULE System SW layer AOCS Modes amp Algorithms Mgt AOCS Autonomy amp FDIR Mgt Sub System SW layer General services Equipments Sw layer amp Processing equipments drivers IEEE 1355 Kernel SW layer OBDH Data Bus MAT 10835 5 Common SW Figure 10 4 2 AOCMS SW Layered Breakdown d Ref Issue S Date 10 4 2 Common DMS AOCS SOFTWARE VEX T ASTR TCN 00349 lt 02 Rev 0 06 02 2004 10 11 The DMS SW and the AOCMS SW running on identical processors share a set of general services called the Common SW The Common SW gathers the Kernel SW and the Generic Services The Kernel SW is a set of general services supporting the applicative part of the DMS SW and the AOCMS SW It constitutes the lowest layer of these two software The Kernel SW directly interfaces with the hardware of the hosting processor and provides basic software mechanisms to hide the details of implementation to the high level layers The Generic Services gather the software components in charge of providing general services mainly the services described in the Generic TM TC ICD and the DMS TM TC ICD see Volume 7 to higher level applications lt ASTRIUM Menus Ref VEX T ASTR TCN 00349 Issue 02 Rev 0 Date 06 02 2004 Page 10 12 10 4 3 DMS Application Software The DMS application SW mainly performs the Mission Time Line MTL management the system Aut
71. INIT CONTROL IN CASE IN CASE EEPROM EEPROM NOT VALID NOT VALID GROUND GROUND CURRENT AOCS sw CURRENT MAT 9826 B Figure 9 2 2 Control and Data Management Unit General Memory Map The general memory map of the CDMU allows to cope with erroneous software updates from ground in the Processor Modules EEPROMs In this case the Start Up sequence loads in the Processor Modules memories an uncorrupted copy of the initial Data Handling and AOCS software Ref VEX T ASTR TCN 00349 S Issue 2 Rev 0 _ Date 06 02 2004 MA ASTRIUM ro eXDIeSS 98 9 3 INTERFACE UNITS Interface Unit concept consists in grouping all interface functions with non standard equipment into dedicated units The Remote Terminal Unit RTU is in charge of the interface with Instruments and non AOCS Platform equipment through TTC B 01 standard links The following types of interface are implemented within the RTU 0 O O UD D UD UD Analog acquisitions equipment secondary voltages Serial 16 bits digital acquisitions payloads transponders power subsystem Bi level digital acquisitions Relay status acquisitions Thermistor value acquisitions thermal subsystem High power ON OFF commands Extended high power commands RF switches Memory load commands payloads transponders power subsystem Timer synchronisation pulses STR SSMM The AOCS Interfac
72. MS SW participates to the implementation of the Mission management which sequences nominal and contingency mission phases based on specific sets of AOCMS modes and interfaces with the ground control via the DMS for observability and commandability of the AOCMS equipment the Spacecraft management which administrates commands and allocates all AOCMS resources the system FDIR which rely in its hierarchical approach on the AOCMS sub system FDIR level and the AOCMS equipment FDIR level Ref VEX T ASTR TCN 00349 Issue 02 Rev 0 Date 06 02 2004 ASTRIUM 055 1014 w 10 5 SSMM SOFTWARE The SSMM consists of 2 processor systems The Memory System Supervisor MSS dedicated to the communication with the DMS PM The File and Packet Controller FPC dedicated to the file management on the memory modules and to the data exchange with the instruments and the TFG The SSMM software runs on the micro processor based MSS and the micro controller located in the FPC The main part of the SSMM SW is programmed in C language Parts of the start up function are programmed in Assembler The SSMM software consists in two parts The Initialisation software covering the Init Mode and running in the MSS It is executed in MSS PROM after activation of the SSMM It performs the following main functions initialisation of system controller and control interface hardware tables data etc load nominal softwar
73. PR2 PCU MEA APC3 MPPT EA 1 3 2 3 active 2 3 active 2 3 active 2 3 active BCDRI Control 1 Aux Supply 1 Aux Supply 1 Control 1 Pyro BCDR2 PCU PCU PDU PDU Pyro Control 2 Aux Supply 2 Aux Supply 2 Control 2 BCDR3 2 3 active PDU LCL SADE PDU LCL SADE The 3 batteries are used in hot redundancy however the batteries have been sized such that the power requirements will be met for two batteries surviving from the three So the battery reliability model 1s simplified to a 2 3 active redundancy model Figure 12 3 4 Venus Express Power Supply Subsystem Reliability Block Diagram S Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 Date 06 02 2004 Page 12 10 ASTRIUM express Ref VEX T ASTR TCN 00349 w V Issue 2 Rev 0 Date 06 02 2004 13 SPACECRAFT BUDGETS 13 1 MASS AND PROPELLANT BUDGETS Table 13 1 1 constitutes a synthesis of mass budget for all sub assemblies It was verified that all components of the spacecraft have been considered In this mass budget which is therefore exhaustive It was possible to reduce contingencies on the basis of the latest assessments of all units In particular the structure mass was measured However 16 kg of contingency is still considered in the dry mass It shall be noticed that the total mass of payloads exceeds the specification 88kg Additional contingencies have been considered in order to
74. Sa X band downlink telemetry One way radio link ONED S band downlink USO X band downlink telemetry radolink cdr Ref VEX T ASTR TCN 00349 D Issue 2 Rev 0 EADS F Date 06 02 2004 ASTRIUM express Page 8 9 8 8 RFC COMPATIBILITY Radio Frequency Compatibility RFC is ensured by tight control of potential noise sources emissions and accurate knowledge of receivers susceptibility thresholds Coupling factors between antennas are determined and controlled at System level Compatibility is verified by test between all RF sensitive receivers and all RF power transmitters to prove non couplings into the individual antennas autocomp RF test Coupling Factor 757 AA X Cable Cable Filters Filters etc etc RF Transmitter RF Receiver x Receiver Susceptibility Threshold Transmitter Output Spectrum Ref VEX T ASTR TCN 00349 gt Issue 2 Rev 0 05 p Date 06 02 2004 ASTRIUM genus exoress Page 8 10 a Ref VEX T ASTR TCN 00349 A Issue 2 Rev 0 _ Date 06 02 2004 AA ASTRIUM Liu express 36 9 DATA HANDLING ARCHITECTURE The Data Management System DMS 15 in charge of telecommand distribution to the whole spacecraft telemetry data collection from the whole spacecraft and data storage overall supervision of spacecraft and payload func
75. TL Scientific missions phases DMS Request 1 Modes used for trajectory change manoeuvers Figure 5 3 1 AOCS Modes Diagram 5 ASTRIUM r express J Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 Date 06 02 2004 Page 5 7 The H W configuration in SBM is not fixed and depends on the configuration in the previous mode Before the Venus Orbit Insertion phase the recommended Normal Mode configuration includes 4 wheels and 2 IMUs n out of control closed loop for monitoring purpose or for initialisation ns n wheels in speed control mode nt n wheels in torque control mode 2 Hold the SADM is set in Hold Mode once it has reached the target position Figure 5 3 2 Use of Hardware units in the AOCS Modes Fj Ref VEX T ASTR TCN 00349 4 S 2 Issue 2 Rev 0 L Date 06 02 2004 TA ASTRIUM l express Page 5 8 Attitude acquisition modes Stand By Mode SBM The Stand By Mode is used in Pre launch and Launch phases for general check supervision and during the deployment of the Solar Array It 15 also a mode used during the transient after failure when a switch to the Software or Hardware safe mode 15 necessary This Mode is followed by the SAM upon a DMS command starting the autonomous acquisition sequence up to Earth pointing Only DMS functions are activated in SBM At AOCS level there is no control during this mode and all the H W resources which are not neces
76. X T ASTR TCN 00349 A Issue 2 Rev 0 _ Date 06 02 2004 ASTRIUM express Page 5 9 EPP Earth pointing Phase An autonomous Off Loading of the wheel WOLP wheel Off Loading Phase is activated at this stage when necessary The attitude acquisition sequence including the SAM and the SHM is fully automatic in case of reacquisition after a failure For the initial attitude acquisition two stand by points are implemented in the design gt one at the end of the SAM when the Sun pointing is acquired gt one inside the SHM when the final pointing is acquired with a thruster control just before the switch to a wheel control After each of these stand by points the ground is able to resume the automatic sequence Each of these 2 stand by points are cancelled automatically onboard after a predefined and adjustable duration such that a complete flexibility is provided to the ground to manage the initial sequence Seting the duration to 0 enables to transform initial sequence In an automatic one These Stand By points are also used by the onboard failure management FDIR to avoid a permanent loop between SAM and SHM Ref VEX T ASTR TCN 00349 EADS Issue 2 Rev 0 Date 06 02 2004 ASTRIUM z Page 5 10 Sun Acquisition Mode SAM Rate Reduction Phase RRP Sun Capture Phase SCP Stand By Sun Acquisition Mode SBM Phase SAP Stowed SA Deployed SA
77. a Software safe mode will be entered Sun Acquisition Mode SAM Rate Reduction Phase RRP Sun Capture Phase SCP Stand By Sun Acquisition Mode SBM Phase SAP Stowed SA Sun Pointing Star Acquisition Phase SPP Phase StAP Biased Pointing Switch to SBM no control FEE BPP for SA deployment Safe Hold Mode ERN uisiti SHM Init i Nominal EAIP transition FDIR transition Earth Acquisition EAP Earth Pointing init EPIP Wheel Earth Sun Off Loading Pointing EPP MAT 13260 Normal Mode Figure 11 8 1 AOCS FDIR Transitions ASTRIUM express er A Ref VEX T ASTR TCN 00349 A Issue 2 Rev 0 Date 06 02 2004 E Page 11 20 5 levels in the AOCS FDIR 5 levels of actions are identified in the AOCS FDIR The level 1 corresponds to a reconfiguration without mode change A Typical case is the triggering of a Non Emergency Surveillance during an operational mode The level 2 corresponds to a switch to SAM after a reconfiguration of suspected units The level 2 of the AOCS FDIR leads to a SW Safe Mode The level 3 corresponds to the triggering of a surveillance during the SAM In this case a global reconfiguration is commanded by the Software and the SAM 15 restarted with all redundant units The level 3 of the AOCS FDIR leads to a SW safe mode The level 4 is reached if a surveillance t
78. ads Master Switch WT Era PCU Digital IM usi CIRE Soe eC DR LCL sry etc eactive Suryailjance Enable BDCS Modify Parme Flags DMS PM RAMY PM RAM pis I Decrement AOCMS SHM gt SAM Battery Tramsition counter ria 0 mk S TX s 5 X IX s TWTA s Switch Off urrerntj Go to S C SW Safe Mode Parameters Parameters able Threshold filter SGM EEPROM SGM EEPROM Figure 11 7 1 BCDR FDIR logical diagram able Threshold filter In case of BCDR failure during the eclipse a Main Bus Undervoltage MBU will occur prior to the BARS reaction due to the over load of the remaining BDR s SC consumption higher than 600W This MBU will trigger the under voltage detection UVD of the LCLs which will switch OFF the LCL s After such event the over load conditions are cleared and the Bus voltage rises again to nominal values The System will then automatically restart after an overall reconfiguration The SC will be kept in the same low power consumption mode as commanded by the BDCS triggering until the eclipse exit and the earth acquisition is completed with the SA sun pointed T Ref VEX T ASTR TCN 00349 AD S Issue 2 Rev 0 EADS M uL Date 06 02 2004 GNUS expres Page 11 18 11 8 AOCS FDIR The main role of the FDIR 15 to ensure autonomously the Spacecraft safety in case of failure As a secondary objective the continuation of the mission
79. alve function The valves used for the loading of pressurant and propellants at the launch site are positioned at the front of each cluster allowing easy access for personnel wearing SCAPE suits during loading operations The feedlines to the main engine to and from the propellant tank and from the pressurant tank are routed through the Xs shearwall at appropriate locations The feedlines to the thrusters are routed along the Ys and Ys edges of the lower floor of the structure D Ref VEX T ASTR TCN 00349 4 Ds Issue 2 Rev 0 EN Date 06 02 2004 _ ASTRIUM express NTO Tank MMH Tank CPS Units Assembly Pressurant Thruster Tank Module 75 Ys Main Engine Figure 6 2 1 Propulsion Layout inside Spacecraft Structure Figure 6 2 2 Propulsion Layout on Xs shearwall and lower floor F Ref VEX T ASTR TCN 00349 EADS a Issue 2 Rev 0 n Date 06 02 2004 5 lm express Page 6 8 a Pressurant Tank Equipped with Thermal Hardware Propellant Tank Equipped with Thermal Hardware Main Engine Figure 6 2 3 Venus Express CPS Main Units Ref VEX T ASTR TCN 00349 EADS Issue 2 Rev 0 p Date 06 02 2004 ASTRIUM ENUS exoress Bus 271 7 ELECTRICAL AND POWER ARCHITECTURE 71 OVERVIEW The Venus Express electrical architecture is designed to cope with the main design drivers as found on interplane
80. and the AOCS Failure management function AOCS hardware units The Star Tracker STR is the main optical sensor of the AOCS used at the end of the attitude acquisition to acquire the final 3 axes pointing and during almost all the nominal operations of the mission A medium Field Of View 16 4 circular and a sensitivity to Magnitude 5 5 are used to provide a 3 axes attitude measurement with at least 3 stars permanently present in the FOV The STR includes a star pattern recognition function and can perform autonomously the attitude acquisition The Venus Express Star Tracker is produced by Galileo Avionica and the basic design of the hardware is identical to the Mars Express one The robustness to straylight 15 improved for Venus Express through the addition of an internal diaphragm inside the optics The thermal analyses lead also to change the coating of the Star Tracker radiator and baffle and to add a thermal shield on the spacecraft 2 Star Trackers are implemented on the X face of the Spacecraft with an angle of 30 between their optical axes Two Inertial Measurement Units IMU used by each IMU including set of 3 gyros and 3 accelerometers aligned along 3 orthogonal axes The AOCS control can use either the 3 gyros of the same IMU reference solution at the beginning of life or any combination of 3 gyros among the 6 provided by both IMUs For the accelerometers only a full set of accelerometers of one singl
81. ant These provide an additional level of protection to the main engine and thrusters which have filters built into them The low range pressure transducers are used to monitor propellant tank pressures in flight following the opening of the normally closed pyrovalves between the tanks and the pressure transducers Downstream of the filters the pipework divides into separate branches supplying the main engine and the reaction control thrusters In the feedlines to the main engine are the normally open pyrovalves Their purpose is to isolate the engine after its final firing Because engine isolation is not critical to the mission the normally open pyrovalves are not duplicated for redundancy The purpose of the normally closed pyrovalves in the feedlines to the main engine is to allow the main engine to remain isolated until required without compromising the use of the thrusters during Venus transfer Again the normally closed pyrovalves are positioned in parallel for redundancy The main engine is fitted with its own filters and flow control valves FCVs The dual valve thrusters are arranged in pairs primary and redundant Direct switching between the primary and redundant thruster of any pair will be implemented in the unlikely event of failure of any primary thruster Each unit incorporates a filter and a thruster latch valve TLV upstream of a flow control valve FCV providing further redundancy in the system As for th
82. any period multiple of 125 ms Ref VEX T ASTR TCN 00349 EADS Issue 2 Rev 0 Date 06 02 2004 ASTRIUM r express Page 5 19 Active Damping phase phase _r AT puise Wheel torque command Thruster torque command T lee 4 AT wheel b AT seq F Sequence N Sequence N 1 Figure 5 4 3 Wheel Off Loading sequence 4 RW speed surveillance ROR Ge 1 failure detection Forced desat 4 Wheel failure 4 hc On Board off load torq d Wheel id max jee I RWOommnd Friction tore estimation and GTransformation in generation psc wheel frame Wheels n eT Tachometers heel nax Hy AH Wheel id frame Reaction wheel management function Figure 5 4 4 Reaction Wheels management function Ref VEX T ASTR TCN 00349 2 Issue 2 Rev 0 L Date 06 02 2004 V ASTRIUM r express um Page 5 20 5 5 AOCS MODE TRANSITIONS The AOCS mode transitions or the transitions between the modes subphases are usually performed upon ground commands taking into account the associated operational constraints Automatic transitions are however performed autonomously by the AOCS case of complete sequences for attitude acquisition or orbit control manoeuvres Automatic or ground commanded transitions Automatic trans
83. are the SGM and the Central PROM located within the CMM and the Master Clock located within one RM failure detection activities are implemented both within the AOCMS S W and within the DMS S W but responsibility of failure isolation and recovery belongs only to the DMS S W In case of problem the AOCMS S W will simply inform the DMS S W of this problem and it is the DMS S W which shall decide on what to do The AOCMS may access both SGM RAM and EEPROM during initialisation Once the initialisation is completed the AOCMS accesses the SGM through requests sent to the DMS S W which 15 in charge of executing these requests All PM memories are either not sensitive to SEU or SEU protected either directly through EDAC program code or through S W processing rewrite check sum scrubbing double write Ref VEX T ASTR TCN 00349 S Issue 2 Rev 0 f Date 06 02 2004 ASTRIUM genus express Page 11 26 d J Ref VEX T ASTR TCN 00349 Vi Issue 2 Rev 0 L Date 06 02 2004 AN A ASTRIUM enus express Page 12 1 12 RELIABILITY AND REDUNDANCY ARCHITECTURE The Venus Express S C architecture is recurring from Mars Express S C with very few exceptions addition of HGA2 X band only antenna with its associated Diplexer and wave guides Thus the redundancy scheme provides the same Mars Express hardware resources to handle on board failures through an autonomous failure management The redundancy c
84. as deemed necessary to implement a second High Gain Antenna in the opposite direction to the main High Gain Antenna Alternate use of HGAs combined with an optimised attitude guidance law allows to keep the Sun in a narrow area located between X and Z during the steady state Earth pointing phase This solution allows to satisfy with good margins the thermal constraints w r t to both the cold face and the Propulsion face In addition external coatings must be modified w r t Mars Express in order to minimize the thermal flux entering the spacecraft Spacecraft resources sizing 1 e thermal control and power subsystem 15 driven by the characteristics of the observation phase Indeed thermal flux 1s entering the radiators in this phase thus increasing the temperature inside the spacecraft This leads to a time limitation for observation phase In the same way battery discharge occurs in this phase since Solar Arrays cannot be perfectly oriented towards the Sun This also leads to time limitation for this phase mostly in eclipse seasons Spacecraft resources sizing has been done on the basis of a sizing case as defined per Mission Requirement Document It corresponds to 95 minutes of Nadir pointing observation around pericentre Due to the high thermal flux in Venus vicinity it was necessary to enforce the radiators efficiency w r t Mars Express One of the consequence is that more heating is necessary during Cruise phase and d
85. as the liquid side supplies propellant to the main engine and thrusters It comprises 1 a pair of 267 litre propellant tanks 2 normally open and normally closed pyrovalves 3 propellant filters 4 low range pressure transducers 5 main engine 6 reaction control thrusters and 7 test ports and fill amp drain valves This section is pressurised with helium by the low pressure gas side and has a MEOP of 20 bar Propellant is demanded from the fuel and oxidant tanks by the main engine and thrusters at oxidant to fuel mixture ratios of 1 67 and 1 54 respectively The presence of the normally closed pyrovalves allows the propellant feed subsystem downstream of the propellant tanks to remain isolated before flight Thus the tanks may be loaded with simulated propellant for ground testing not envisaged for Venus Express and later with propellant at the launch site Loading of these liquids is performed through the fill amp drain valves during which gas in the tanks is vented out through the fill amp vent valves Another purpose of the normally closed pyrovalves is to isolate the tanks from the rest of the propellant feed subsystem so that proof pressure testing of the pipework without pressurising the propellant tanks may be performed As is the case throughout the CPS the normally closed pyrovalves are positioned in parallel for redundancy Downstream of the pyrovalves are the filters one for fuel and one for oxid
86. been defined for the units that are not mandatory to continue operations the Blocking Surveillances BS and the Off Line Surveillances OLS The triggering of these surveillances does not lead to a mode change Reconfiguration recovery For each unit or module nominal and redundant sets are identified and managed by the S W Out of failures nominal elements are used The link with physical resources A B units 1s ensured through a redundancy table managed by the ground For each surveillance a set of suspected units 1s identified depending of course on the level of the surveillance After surveillance triggering the reconfiguration actions consist first to switch off suspected units and change their nominal redundant status In a second step a switch to the safe mode will be commanded if necessary with redundant units 1f they are also used In this mode EADS Issue 2 Date 06 02 2004 ASTRIUM Page 11 19 Safe Mode In most of the AOCS Modes except the Main Engine Boost Mode when it is not possible to stay in the same mode after a surveillance trigerring Emergency Surveillance the safe mode is entered The mode starts by the Sun Acquisition Mode SAM and the sequence continues autonomously in SHM Ref VEX T ASTR TCN 00349 Rev 0 performing an Earth acquisition Depending on the fact that there is a reconfiguration of the CDMU including a PM reboot or not a Hardware safe mode or
87. beginning on ground decision from MEBM BIP in place of the nominal MELSP or automatically from the MEBFP in case of unrecoverable Main Engine failure detected onboard during the boost The MEBM 15 nominally entered from the Normal Mode during which an orientation of spacecraft slew manoeuvre is performed NM GSP The MEBM can also be entered from the Safe and Hold Mode SHM prior or after the reaction wheels activation in order to shorten the procedure in case of anomaly just before the critical insertion manoeuvre In this case the attitude manoeuvre is done with thrusters in the first phase of the MEBM BIP phase Ref VEX T ASTR TCN 00349 2 Issue 2 Rev 0 f Date 06 02 2004 TA ASTRIUM La express Page 5 13 During an attitude profile programmed by the ground is commanded to the spacecraft This profile includes the attitude profile to be followed during a longer time in case of Main Engine failure The end of the manoeuvre is decided onboard on the basis of the AV estimation derived from accelerometers measurements In order to ensure the required accuracy This strategy requires previously an in flight calibration of the accelerometer biases In case of Back up MELSP the thrust 15 stopped on the basis of a timer The manoeuvre 15 followed by a switch to SAM The following sequence includes an automatic Earth reacquisition in SHM This strategy is necessary due to
88. bility The helium tank boost heating 1s managed that way When the temperature monitoring 1s not able to provide the required detection for reconfiguration the circuits are operated in hot redundancy This 1s the case of most of the CPS lines and thrusters heating lines Ref VEX T ASTR TCN 00349 m di Issue 2 Rev 0 Date 06 02 2004 Page 4 12 ASTRIUM r a 5 F Ref VEX T ASTR TCN 00349 gt Issue 2 Rev 0 Date 06 02 2004 Page 5 1 1 ASTRIUM ka express 5 ATTITUDE AND ORBIT CONTROL SYSTEM 5 1 AOCS BASIC CONCEPTS Attitude manoeuvrability Due to the selection of fixed High Gain Antennas HGAI amp 2 and to the propulsion configuration including a Main Engine the Venus Express mission requires a high level of attitude manoeuvrability for the spacecraft Attitude manoeuvres will be performed Between the observation phase and the Earth communication phase or to reach specific attitudes necessary for science observations SPICAV for instance Before and after each trajectory correction manoeuvre performed either with the Engine or with the 10N thrusters To optimise the Wheel Off Loading through the selection of an adapted attitude for this operation All the attitude manoeuvres of the operational phase are defined on ground using a polynomial description of the Quaternion to be followed by the Spacecraft Attitude estimation an
89. communication phase of the orbit to hold a robust link with the Earth between scientific operations for data transmission It can be also the phase used during long duration solar conjunctions i e with no Earth communication The Fine Pointing Accuracy Phase FPAP is the operational mode used for the scientific mission during the Venus observation It is designed to be able to control the Spacecraft around mission attitude profiles defined by the ground Nadir pointing Earth radio occultation and to ensure the pointing and pointing stability performances necessary for payloads operation The Solar Array orientation is commanded by the ground and 15 fixed during the observation phase This phase 15 also used for attitude transient damping before any thruster controlled mode The Fine Pointing Inertial Phase FPIP controls the Spacecraft attitude around a ground commanded fixed attitude ensuring the pointing and pointing stability performances necessary for the payloads operation with fixed Solar Arrays the Solar Array orientation 1s automatically computed on board at the beginning of the phase then it remains fixed during the observation period This phase is adapted to a period of the mission where inertial observation is required such as SPICAV observations It can also be used before Wheel Off Loading if a specific attitude is required for this operation The Ground Slew Phase GSP is used as a transition between sub phases or mo
90. cope with future evolutions Balancing mass is needed to keep the centre of gravity within the acceptable range w r t Launch Vehicle requirement and Main Engine Boost Maximum balancing mass 19 kg is considered in this mass budget The most probable figure 15 however closer to 12 kg It was considered that only half of apocentre lowering manoeuvre can be performed with the Main Engine which is a worst case assumption derived from the CPS functional constraints This results in a small increase of the propellant budget Finally 2 system margin was added as required Maximum launch mass exceeds by 3 kg the launcher capability at the very end of the launch window 1270 kg on November 25 However structure qualification is still guaranteed with adequate margins Ref VEX T ASTR TCN 00349 EADS Issue 2 Rev 0 Date 06 02 2004 ASTRIUM r express Page 13 2 Co a J J Ref VEX T ASTR TCN 00349 EADS Issue 2 Rev 0 Date 06 02 2004 ASTRIUM Ta express Page 13 3 ee ti ww _ 94 gt N Apocentre lowering Main Engine 641 142 142 A A gt eww j j ww w Qo Co 7 7 Table 13 1 1 Spacecraft Mass and Propellant budget Ref VEX T ASTR TCN 00349 w V A S Issue
91. d control concepts The attitude estimation is based on Star Tracker and gyros ensuring the availability of the measurements in almost any attitude Some constraints have however to be fulfilled the Star Tracker being unable to provide attitude data when the sun or the planet are close to or inside its Field of view Reaction wheels are used for almost all the attitude manoeuvres providing a great flexibility to the Spacecraft and reducing the fuel consumption The angular momentum of the wheels has however to be managed carefully from ground Earth pointing attitude for communication Nadir pointing and P L dedicated profiles for V i or Venus observation Inertial attitude for P L specific observation or wheel off loading MAT 13258 Figure 5 1 1 Manoeuvrability of the Venus Express Spacecraft J Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 _ Date 06 02 2004 72 ASTRIUM r express Page 5 2 5 2 AOCS HARDWARE ARCHITECTURE AOCS Hardware architecture All the sensors and actuators used by the AOCS are connected to the AOCS interface electronics unit AIU through a IEEE 1355 bus for the Star Tracker a MACS bus for the SADM a RS 422 link for the IMU or direct wirings for the SAS and the reaction wheels The Control and Data Management Unit CDMU contains a dedicated processor for the AOCS S W including the processing of the sensors and actuators the estimation and control algorithms
92. d in a 2 out of 3 active redundancy The 4 Reaction Wheels can be used in two redundancy schemes a 3 out of 4 stand by redundancy scheme or a 3 out of 4 active redundancy scheme The Remote Terminal Unit I O boards are used in out of 2 active redundancy In the case one RTU I O board is failed some functions are no more accessible because no cross strapping 1s implemented between the RTU and these users e g the TRSP A is coupled to I O A and TRSP B is coupled to I O B The Venus Express S C reliability block diagram is composed of the following reliability block diagrams Communication Subsystem reliability block diagram Figure 12 3 1 Data Management amp Attitude Orbital Control Management subsystem reliability block diagram Figure 12 3 2 Combined Propulsion Subsystem reliability block diagram Figure 12 3 3 Power Supply subsystem reliability block diagram Figure 12 3 4 Ref VEX T ASTR TCN 00349 ADS 2 Issue 2 Rev 0 E Date 06 02 2004 ASTRIUM express Page 12 6 S Band COMS active X Band RFDU 5 band RFDU S band RFDU S band Coax Switch Coax Switch Diplexeur 1 S band 2 LGA 5 band RFDU Coax SW passive Only for the first 5 days Only for the first 5 days RFDU S band RFDU S band RFDU S band Coax Switch Coax Switch Diplexeur S Band Transponder 1 Transponder 1 S Band Rx S Band Tx passive T
93. d observation attitude 15 reached The SADM uses a stepper motor a gear and a twist capsule technology The SADM motion is defined in the range 7 180 minus margins The Venus Express SADM is identical to the Mars Express unit and also to the Rosetta unit except for the speed levels which are specific to Mars Express and Venus Express The coating of the Venus Express unit has been modified to withstand the Venus thermal environment J Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 EADS p Date 06 02 2004 ASTRIUM enus 5 Page iSd Gyros Acceleros VW 7 7 Star Tracker Control amp Data Management Unit CDMU Solar Array Drive Electronics PS QS Wheel MAT 13259 Integrated Electronics J Duplicated unit Internally Modified unit Redunded Mars Express derived unit Mars Express recurring unit Figure 5 2 1 AOCS Hardware architecture AOCS unit Technology Main characteristics Heritage Supplier Star Tracker CCD detector 16 4 circular FOV Magnitude 5 TBC Rosetta Mars Express Galileo STR depending on straylight modification unit modified for Avionica straylight Gyro accelero B Ring Laser Gyros RLG 3 gyros 3 acceleros per unit Rosetta and Mars Express Honeywell IMU unit Sun Acquisition Solar cells mounted on a pyramid Internal redundancy Derived from Rosetta and TPD TNO Sensor S
94. d the different Payload SW These software components are located on different hardware units and contribute to fulfil different parts of Venus Express mission The DMS SW runs on a dedicated Processor Module PM located on one of the two Control amp Data Management Units CDMU It is made of the Common software and the DMS application software The Common software contains the PM HW interface manager the basic SW services the generic services and the OBCP manager The DMS application software performs the mission management and the DMS functions management The AOCMS SW runs on another CDMU PM It is composed of the Common software quite the same Common SW as for the DMS SW and the AOCMS application software The AOCMS application software performs the AOCMS modes and algorithms management and the AOCMS sub systems management Both DMS SW and AOCMS SW are loaded in the PM RAM and started up by the PM Firmware presented hereafter The DMS SW runs on a dedicated PM and the AOCMS SW runs on another PM belonging to anyone of the two CDMU The two remaining PM are spares which only contain Firmware and are ready to be configured as DMS or AOCMS PM in case of failure Each one of the two CDMU owns two PM and each PM has a software in PROM called the PM Firmware The PM Firmware is automatically activated when the CDMU 15 powered on It initialises the PM and performs PM health status verifications software loading and minimum communicatio
95. de The resources management which handles all the resources needed to achieve AOCMS objectives 1 the star tracker the Sun acquisition sensor the Inertial Measurement Units including gyros and accelerometers the Reaction Wheels the thrusters and the Solar Array Drive Mechanisms This function performs the configuration management and commanding of all these units The processing of sensors output and actuators input which provides the AOCMS modes with specific services allowing the filtering of hardware raw resources measurements or the processing of commands computed by the attitude and control laws The ephemerides propagator continuously running which provides the AOCMS modes with the spacecraft inertial directions to the Earth and Sun The AOCMS modes management which manages the transitions between the AOCMS modes Sun Acquisition Mode Safe Hold Mode Normal Mode Orbit Control Mode Thruster Transition Mode Main Engine Boost Mode Braking Mode The AOCMS algorithms management which performs the attitude estimation and control and the trajectory control in each mode The AOCMS FDIR which manages the FDIR at AOCMS equipment and AOCMS function levels As for the DMS SW this function is based on the monitoring of specific parameters representative of the equipment states and functions completion The AOCMS SW acts at a sub system level in comparison with the DMS SW which manages the system level However the AOC
96. des when an attitude reorientation is necessary between the pointing on ephemeris in GSEP and the observation in FPAP or FPIP before and after transition to the OCM for trajectory corrections in order to orient properly the thrust and before the MEBM The attitude profile is defined by the ground The orientation of the Solar Array 15 fixed during this phase and is adapted to the final attitude and mission during the following phase The Wheel Damping Phase WDP includes a robust control law able to reduce the residual rates and attitude errors when coming from other modes It is therefore the entry point the Normal mode especially usefull when a transition from a thruster controlled mode has to be performed TTM The Wheel Off loading Phase WOLP enables to manage during the Normal mode the wheel angular momentum Thruster pulses are used to reach a target angular momentum during this phase autonomously onboard or upon ground commands This operation is forbidden in some situations where it could have dangerous effects like the slew manceuvres For this reason the WOLP 15 authorised only from the GSEP FPAP and FPIP phases During the observation phase in FPAP the transition to the WOLP is authorised in the S W but should not be used the Wheel Off Loading taking place nominally out of observation period Ref VEX T ASTR TCN 00349 EP S 2 Issue 2 Rev 0 Lr Date 06 02 2004 7 ASTRIUM l
97. dundancy means that a function is ensured by 1 element nominally active among 2 After one failure the second one is used in place of the failed element 1 out of 2 active redundancy or hot redundancy means that a function is ensured by 2 elements nominally active in parallel After one failure the second one still ensures the function 2 out of 3 stand by redundancy or cold redundancy means that a function is ensured by 2 elements nominally active among 3 After one failure the third one is used in place of the failed element 2 out of 3 active redundancy or hot redundancy means that a function is ensured by 3 elements nominally active in parallel In case one element is failed the 2 remaining ones still ensure the function Redundancy policy Consistent with the redundancy schemes most functions are implemented with a one out of two stand by redundancy The main exceptions to this rule are the following The antennas LGA2 HGAI and HGA2 and the front end RF components until the RF switches in the RFDU and the WIU wave guides diplexers RF cables and also the 3 dB Hybrid are not redunded The other items are Single Point Failure free The Dual Band Transponder receiving function is used in out of 2 hot stand by redundancy only one RX is receiving the uplink RF signal but both RX are ON it must be noticed also that the S RX and the X RX of the same transponder cannot operate simultaneous
98. dundant at both actuation electronics and initiator level The purpose of this electronics is to provide necessary means to select a particular firing input power source and firing outlet to fire monitor and control the pyro outlet current to actual pyro devices The necessaty energy is being taken from the batteries whereby battery 1 is dedicated to the pyro primary section and battery 2 to the pyro redundant section The battery 3 is being used as a common spare energy source Each nominal and redundant side of the PDU provides 32 Pyro outputs delivering power to the users The rule to allocate the 32 Pyro lines to one group or to another is that two commands cannot be part of the same group if the erroneous sending of one command instead of the other leads to a mission catastrophic action This allows in particular a safe management of the Propulsion isolation Pyro valves The five following groups are defined on Venus Express Q Group 1 8 firing circuits features all Pyro lines dedicated to the Solar Array wings deployment O Group 2 8 firing circuits features pyro commands for all Pyrotechnic valves used for priming of the Propulsion System O Group 3 9 firing circuits is dedicated to opening Main Engine inhibition Pyrotechnic valves and High Pressure Gas Pyrovalves Group 4 3 firing circuits features the MAG boom deployment command Q Group 5 4 firing circuits is not used The activation of selected pyro requires four ind
99. e 06 02 2004 Page iv PAGE ISSUE RECORD Issue of this document comprises the following pages at the issue shown Astrium Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 Date 06 02 2004 Page v TA ASTRIUM express TABLE OF CONTENTS SPACEC RAPI OYERYIKV i u u 1 1 gt gt pein dea x 1 1 1 2 SPACECRAFT MAIN CHARACTERISTICS ea per adde 1 3 1 3 SPACECRAFT CONFIGURATION 1 5 1 4 SPACECRAFT 1 6 1 5 COORDINATE 1 8 2 LOA uuu u u o u uu Z u usu sss 2 1 2 1 THE VENUS EXPRESS MISSION AND PERSPECTIVES zl 2 2 THE VENUS EXPRESS SCIENCE PAYLOAD cccoscoscescescceccsscsccsccvccsccsecescessascescadsascascessonceees 2 2 2 3 H PTI RTT zd 23 AFL annaa A AER 2 3 2224 RTA GINE Em 2 4 22 2 Pg e e 2 2 2 3 4 I ET u ia s uu 2 6 422 SAT RENT RTT TRENT 2 2 3 6 I EEEE eee Ree Aen ene nn L m Tn 2 8 2 3 7 J a u i o ANAA us 2 10 2 MECHANICAL DESIGN uuu uuu u u
100. e Orbit Control Mode OCM and Thruster Transition Mode TTM for instance The gyro stellar estimator processes gyro and star tracker STR measurements to provide an accurate estimate of the spacecraft attitude It is based on a Kalman filter with constant covariance that allows mixing measurements at different rates 8 Hz for the gyros and 2 Hz for the STR The constant covariance reduces the computer load while ensuring good performances The estimated attitude is a quaternion representing the spacecraft attitude in the J 2000 inertial frame The gyro stellar estimator also estimates the gyros drifts to limit the attitude errors in case of STR measurement absence due for instance to a STR occultation A specific management of the drift estimates 1s implemented for Mars Express and Venus Express taking into account the specific conditions of the scientific mission phase existence of rates due to varying profiles and potential occultation The gyro stellar estimator implements a coherency test between the gyro and STR measurements in order to detect failures that could not be detected at equipment level Reaction wheel Off Loading function The wheel Off Loading function enables to manage the angular momentum of the wheels to a target value through thruster actuations This function 15 completely autonomous during the last phase of the Earth acquisition sequence SHM EPP Earth Pointing Phase During the nominal operations around Ve
101. e mission needs for failure management purposes FDIR for specific mission needs or to have the best preparation and verification of the hardware from ground before specific operations The IMUs configuration can be adapted from the ground during the operational modes leading to a change in the AOCS FDIR actions If a 6 IMU configuration 15 selected for instance the system will be able to react autonomously to some gyro failures in Normal Mode without going to the back up mode SAM This ground capability is especially interesting for some critical mission phases such as the Venus insertion preparation for instance during the insertion manoeuvre itself the 6 axes IMU configuration is locked and managed onboard The IMUS configuration during the safe mode SAM SHM is systematically 6 Ref VEX T ASTR TCN 00349 2 Issue 2 Rev 0 f Date 06 02 2004 TA ASTRIUM express Page 5 23 The wheel configuration during nominal operations of Normal mode includes 3 wheels It is however possible to the ground to set a 4 wheels configuration It allows to avoid the SAM triggering in some cases of wheel failures prior the MEBM transition All these ground flexibilities have an impact on the onboard management of the AOCS Hardware leading to some H W actions which are not performed by the Software but by the ground Ground recommended actions I
102. e IMU is used due to the lower criticality of the accelerometer function and to the availability onboard of an alternative method for the AV measurement pulse counting The Gyros are useful during the attitude acquisition phase for the rate control during the observation phase to ensure the required pointing performances and during the trajectory corrections for the control robustness and failure detection A non mechanical technology is selected to avoid the mechanical sources of failure in flight The Accelerometers are essential during the main trajectory corrections such as the insertion manoeuvre to improve the accuracy of the AV The IMU of Venus Express 15 identical to the Mars Express unit The number of units and the onboard management of the configuration 15 identical to Mars Express Two redunded Sun Acquisition Sensors SAS are implemented on the Spacecraft central body and are used for the pointing of the Sun Acquisition Mode SAM during the attitude acquisition or reacquisition in case of failure The SAS are identical to Mars Express units for what concern their mechanical electrical or functional interfaces New solar cells mounted on a new ceramic backing are used in order to withstand the Venus thermal environment The higher current delivered by the SAS in the Venus environment lead to change the electrical interface with ATU impedance The SAS are provided with customised baffles J Ref VEX T ASTR TCN 00349
103. e MSB and the last bit bit 15 1s the LSB The MSB is transmitted first Word 1 Word 2 Word 3 Word 4 Word 5 Word 6 Word 7 Word 8 Word 9 Word 10 Data acquired and by the AOCS are the last ten words in 8 Hz cycle Data stored in the datapool are the last 8Hz acquisition within a 1 Hz cycle 25 x 10 words 200 Hz 200 Hz 200 Hz 8 Hz acquired by the AOCS at 8 Hz 8 x 10 words 8 Hz 8 Hz 8 Hz stored in the AOCS datapool at 1Hz Figure 10 7 1 IMP data acquisition and storage by the AOCS Ref s Issue Date ASTRIUM express Page 10 8 TRANSPONDER SOFTWARE VEX T ASTR TCN 00349 02 Rev 0 06 02 2004 10 17 The S X band transponder provides the two way link between the S C and the ground terminal via S X band RF signals The Transponder software 15 composed of an RF section and a digital section Its purpose 15 to control the functionalities of the digital section and to execute the signal processing operations required to maintain the forward and return links The Transponder software interfaces with the S C by means of serial commands and digital telemetry The Transponder software 15 not patchable which means that the Transponder is considered as black box at software level gt Ref S Issue TA ASTRIUM UL ieu express Page 10 9 INSTRUMENTS SOFTWARE VEX T ASTR TCN 00349 lt 02 Rev 0
104. e Unit AIU is dedicated to all AOCS equipment interface It acquires signals from the AOCS sensors and generates actuator commands according to the control law outputs as provided by the Processor module The following interfaces are implemented within the AIU L D O UD UD UD Reaction Wheels command tacho signal acquisition and analogue monitoring signals acquisition Star Tracker configuration programming command and data acquisition Inertial Measurement Unit data acquisition and power switching Sun Sensors current acquisition Thrusters and Main Engine command and current and temperature acquisition Latch Valves command and status acquisition Pressure Transducers power supply and signal acquisition Communication between the PM DMS and the RTU 15 ensured by means of a standard OBDH Data Bus and the communication between the AOCMS and AIU 15 ensured by means of IEEE 1355 links EADS ASTRIUM r express J Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 Date 06 02 2004 Page 9 9 Analog acquisitions Serial digital acquisitions Bi level acquisitions Relay status acquisitions Temperature acquisitions High Power commands Memory load commands synchronisation pulses Star Trackers Inertial Measurement Units Sun sensors solar Array Drive Assembly a Reaction Wheel Assembly Propulsion module OBDH p and Control Interface i 28 V convert
105. e from EEPROM to RAM reduced commands handling transition to Operational Mode The Operational software covering the Operational Mode and Test Mode It runs in the MSS RAM and FPC RAM It performs the following main functions execution and control of telecommands configuration and test of the memory modules control of data flow from instruments and to TFG to and from the Memory Modules failure handling including management of failure log failure recovery creation of event report housekeeping TM packing for all required data Watchdog control In case of fatal failure the SW returns to the Init software to allow for failure investigation we Ref VEX T ASTR TCN 00349 A Vi Issue 02 Rev 0 Date 06 02 2004 ASTRIUM ENUS express Page 10 15 10 6 STAR TRACKER SOFTWARE The STR SW runs on its own 32 bit microprocessor DSP21020 at unit level Its architecture is structured in modules according to different operative modes of the sensor 1 e image acquisition image processing attitude acquisition determination and I O data management It is composed of a generic part and a specific part The STR SW generic part provides basic services to the AOCMS SW for full visibility and investigation health status auto tests programmable window raw video access memory read and write The STR SW specific part gathers all the tasks related to the application modes and include Acquisition and Meas
106. e gas side the test ports are used on the ground L Ref VEX T ASTR TCN 00349 A Issue 2 Rev 0 _ Date 06 02 2004 AA ASTRIUM express 6 2 LAYOUT This section examines physical layout of CPS The propellant tanks are positioned centrally along the spacecraft Ys axis but are offset with respect to the Xs axis The oxidant tank is closer to the centre line to compensate for its greater mass when loaded with NTO This arrangement reduces the shift in the position of the centre of gravity over the mission The tanks are supported on beams under the lower floor hence the propellant outlet pipes have to pass through the beams to emerge under the spacecraft but within the launch vehicle adapter ring A 90 elbow fitting is used to turn the pipes within the available envelope to enable them to bend upwards and pass back through the floor The helium pressurant tank is located in Xs Ys segment of the structure The port boss is mounted through the Ys shearwall and the blind boss is mounted to a dedicated support panel The main engine 15 positioned close to the centre of the lower floor It is mounted on a raised bracket so that the feed pipes to the flow control valves at the top of the engine can run across the upper face of the floor The mounting bracket 1s positioned between the two propellant tanks subsequently the engine valve mechanical couplings are inaccessible after installation of the
107. e instrument is mounted Ref VEX T ASTR TCN 00349 a S V Issue 2 Rev 0 cA Date 06 02 2004 GNUS exc expres Page 9 1 3 MECHANICAL DESIGN 3 1 DESIGN DRIVERS The mechanical design of the Venus Express spacecraft results from the following design drivers a To reuse the Mars Express mechanical bus as far as possible b To take into account the specific constraints of the Venus Express mission c To implement the lessons learnt from Mars Express d To minimize the spacecraft dry mass and optimise the centre of gravity location a The reuse of the Mars Express mechanical bus structure and propulsion system helps in minimising the development risks and securing the very tight schedule of the programme It takes benefit from the qualification status achieved on Mars Express In particular the core structure design remains basically unchanged which allows a qualification approach by similarity The modifications of the secondary structure are strictly limited to the accommodation of the new or modified units Moreover the mechanical environments are identical to Mars Express for most of the units b The main constraints induced on the design by the Venus Express mission are as follows Accommodation of the Venus Express payloads composed of modified payloads ASPERA PFS SPICA V and new payloads VIRTIS MAG VMC and VERA Specific thermal environment with permanent sun illum
108. e to control higher disturbing torques This capability has been also introduced for Venus Express even if a better control of the spacecraft centre of mass 1s expected from a specific action plan The 4 thrusters control remains the baseline The preparation of the Venus Orbit Insertion VOI manoeuvre in Normal Mode has to be performed with a specific configuration defined to reduce the risk of a safe mode before the manoeuvre The recommended configuration includes 4 wheels 2 IMUs and the inhibition of several AOCS surveillances This mode is split in three phases It starts by a boost initialisation phase BIP during which no thrust is created the control 1s ensured by the thrusters in on modulation This phase is dedicated to the switch off of all equipments not mandatory in MEBM It is followed by Liquid Settling Phase LSP during which a low mean acceleration is generated by the 10N thrusters commanded in OFF modulation like during the OCM to reduce the liquid transient in the tanks It 1s followed by the burn itself Burn Firing Phase BFP during which the Main Engine produces the thrust and the attitude control is ensured by the thrusters in on modulation The boost be done with 8 thrusters The Mode is called in this case the back up MELSP Four thrusters are actuated continuously and the four others are OFF modulated to provide both thrust and control torques The back up MELSP can be entered either at the
109. ecessary along Venus Express mission Pointing of the High Gain Antenna HGA 1 or HGA 2 towards the Earth and the Solar Array cells towards the Sun This kind of guidance 15 used during the cruise phase and for communications during the scientific mission phase these two cases corresponding to the AOCS Normal Mode pointing on ephemerides NM GSEP phase For this function the guidance uses the Spacecraft to Earth and the Spacecraft to Sun directions as described in the ephemeris based for Venus Express on Kepler orbit approximation of the planets orbits around the sun This type of guidance is also used in a different way for the Earth acquisition SHM Safe Hold Mode in order to perform the autonomous orientation of the spacecraft towards the Earth The ephemeris data are then used to perform large angle slew manoeuvres with thruster control Attitude profiles this type of guidance is necessary during the observation phase for the Nadir pointing or to follow more specific profiles This function is ensured by an onboard profile description based on Chebychev polynomial the parameters being uploaded from ground This capability enables also to ensure the attitude slew manoeuvres Q Fixed inertial pointing fixed quaternion This type of guidance is used for specific phases of the mission during Orbit Control Mode Thruster Transition Mode or during the scientific mission phase in NM FPIP and NM WDP Three ge
110. ecraft no attitude manoeuvre is computed autonomously in Normal Mode Management of the HGA switching at DMS and AOCS level The HGA switching is performed only 4 times during the mission and is not considered as time critical It is therefore performed from ground through a dedicated TC The onboard software ensures however several tasks autonomously The appropriate storage of parameters in the Safeguard Memory SGM The consistency between DMS data and AOCS data for the selection is performed autonomously by the DMS in order to avoid a critical situation where the RF configuration selected by the DMS 1s different from the HGA pointed towards the Earth by the AOCS The adequate selection of parameters during Safe Mode 15 also ensured by the DMS For this purpose the DMS uses a Selected HGA flag in DMS PM RAM which 1s also stored in SGM and an internal command sent to AOCS for the HGA switching On receipt of the DMS command the AOCS updates its own HGA selected flag and selects the associated directors cosines of the HGA This direction to be pointed will be effective at AOCS level only at the next entry in SHM or NM GSEP as recalled before During a Software or Hardware Safe Mode the DMS also sends this command to the AOCS ensuring that the safe mode is started with the same data at DMS and AOCS level This procedure is defined for the SSCE option of the Guidance which is the basel
111. ed The BCDR FDIR mechanism as illustrated 1s based on the following surveillances The BCDR Anomaly Recovery Surveillance BARS which is a local surveillance that detects any BCDR failure occurring outside eclipses Upon BARS triggering the S C power consumption 1s reduced in a first step by putting the payloads and the SSMM in keep alive or OFF mode Since the overall battery power capability has been reduced to 600W instead of 900W before failure the SC power consumption shall be further reduced during eclipses This 1s the aim of the second monitoring the Battery Discharge Current Surveillance BDCS which 15 therefore enabled by BARS The Battery Discharge Current Surveillance BDCS monitors the TM battery discharge currents of the two safe BDRs to detect the eclipse entry Upon triggering a S W safe mode is commanded to expand the power load shedding initiated by BARS additional units are switched OFF such as the RF emitters TX s and TWT s the AOCS units which are not used in SAM and non critical heaters The SC remains in such low power consumption configuration until the eclipse exit and until the earth acquisition 15 completed with the SA sun pointed Afterwards the telemetry transmission is re started and the non critical heating lines are re powered Ref VEX T ASTR TCN 00349 2 EADS Issue 2 Rev 0 ASTRIUM F Date 06 02 2004 GNUS expres Page 11 17 Disable Enable BARS Disable BARS Execute Paylo
112. ee KNEFFEPREFIJEFPZPFFIREFJJEPFFERFFIESZQZFFISFEZJEJJZEPFEF JEFISFFFEPFFFIEFIZKFPFFZFFSE FEFETSEPFTERTETEFIEPTYETTETF fE Pd FE EEST E PSS FP PESTS EE PS FFISFFFEIEFFEIZSFIZFFIFEEFEIJI 2 06 02 2004 VEX T ASTR TCN 00349 Page 1 5 Ref Issue Date EADS enus express ASTRIUM 1 3 SPACECRAFT CONFIGURATION eee ee ee Fd ff EFPTETEFTEPTETTETE OP ee FIERFFJEFIFFFEEESZIS ZIZ PF FEE VEFPEPPEPFETEFPSPFEOFEPEFPSEPTERFETPEPTPSPFETP FSGS FORESEES EEE SESS FrISPFFEPFFIZPFIZEPFIFFZIIZF CEFIEFFIFFIEFFEEFIAEFIF TETFEFTEFTISEYPFTETEFTEFTTT n FIIFIRIIRIRFPEPPREPREEI ee a a ee MPTP ESP rp PFE SS Pre ra a A FIPIIPPPPEFPEFPRIEXGE FTFFETTYETFFTEFPTETTEFTEFIE ee mr zr r m mm r eee ee F OF PSP ELSES TPE PPE PS PD rig TEE 552 DE ETS EE ET po dik Pee ee E BIST Spacecraft In Orbit Configuration Figure 1 3 1 EADS ASTRIUM 1 4 SPACECRAFT ARCHITECTURE Thermal control Mars Express Principles customized for Venus Controlled heaters Passive amp radiators GaAs cells
113. ef VEX T ASTR TCN 00349 A D Issue 2 Rev 0 Date 06 02 2004 AA ASTRIUM p ex reos E 570 7 2 1 Power Generation Two symmetrical Solar Array wings generate the Venus Express power The solar array consists of two identical low weight deployable wings of 2 panels each and is pointed towards the Sun by means of a one Degree of Freedom Solar Array Drive Mechanism When stowed each wing is clamped to the spacecraft side panel on four hold down points and release mechanisms For deployment four redundant pyrotechnics bolt cutters release each wing individually After deployment the two panels are held in position and in a defined distance to the satellite body by the Inner Yoke and the Outer Yoke The Solar Cell Assemblies SCAs are placed on the 2 panels of a wing and no cells are placed on the outer yoke When stowed the solar cell assemblies located on the outer panel are facing to space Sun The electrical power is transferred to the spacecraft by a harness routed on the rim of the wings onto the connectors of the SADM The chosen cell technology is RWE triple junction GaAs cells GAGET1 ID 160 65x38 with 100 um cover glass thickness allowing to cope with Venus radiation environment In order to decrease temperature OSRs have been introduced on each panel lay out on front and rear side Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 Date 06 02 2004 ASTRIUM r express Pase 205 The following
114. en N amp R functions will be applied to simplify failure propagation avoidance assessment concerns typically multi function units e g AIU External redundancy safest architecture for NOR failure propagation avoidance detrimental to mass amp volume aspects concerns typically single function units e g STR TWTA No redundancy mechanical elements mechanisms are generally not redounded operational reliability of such units is demonstrated through adequate design margins quality control and successful environmental test results Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 EADS Date 06 02 2004 ASTRIUM GNUS XPress Page 12 3 Payload Heater Power Heater Power Heater Power Heater Power 1355 link 28V Regulated Thermal Power Protected lines Pyro Devices Power Supply Communications Solid State Mass Memory USO I F Power Distribution Unit AOCS Interface lt Unit x L riz Solar Analogue Solar Array Drive Assembly Analogue Propulsion Reaction Star Inertial Sun Wheel Tracker Measurement Acquisition Assembly Unit Sensor Attitude and Orbital Control Figure 12 2 1 Venus Express Spacecraft Physical Redundancy Implementation at Unit Level s d Ref VEX T ASTR TCN 00349 A Issue 2 Rev 0 ASTRIUM p Date 06 02 2004 12 3 REDUNDANCY DESCRIPTION Definitions 1 out of 2 stand by redundancy or cold re
115. ensions are compatible with the allowable volume under launcher fairing 3435 mm diameter with a comfortable margin inherited from Mars Express Delta 2 initially alternative launcher When the spacecraft is under fairing access to skin connectors is possible through an access door the fairing same configuration as Mars Express The spacecraft centre of mass in launch configuration including the balance mass 1s as follows 10 5 mm on Xs 10 5 mm on Ys 760 10 mm Zs The unbalance in Xs Ys plane is less than on Mars Express and is compliant with the launcher ICD requirement of 15 mm in both directions The CoG height 15 less than on Mars Express mainly due to removal of Beagle 2 and 15 not an issue Ref VEX T ASTR TCN 00349 EADS Vi Issue 2 Rev 0 Date 06 02 2004 ASTRIUM ENUS exoress Page 3 7 sym 2 apum 2 W Figure 3 3 1 Spacecraft Stowed Configuration Ref 2 EADS VEX T ASTR TCN 00349 Issue 2 Rev 0 Date 06 02 2004 ASTRIUM rm express Page 3 8 T EE a k 7 ku 1 Figure 3 3 2 Spacecraft Stowed Configuration with MLI MLI on Payload not shown Ref VEX T ASTR TCN 00349 2 Issue 2 Rev 0 _ Date 06 02 2004 AS ASTRIUM Lim express Page 39 Figure 3 3 3 Spacecraft Launch Configuration on FREGAT Spacecraft dimensions comply with available volume under fairing Skin Connectors
116. ependent commands Battery source select activation Pyro selection Pyro arming Pyro firing Note that Battery Arm plugs and Pyro Arm plugs are connected just before Flight Otherwise Safe plugs are used as additional mechanical safety barriers during the ground activities EADS ASTRIUM Ref VEX T ASTR TCN 00349 Issue L express The following figure shows the VEX Pyro Electrical chain ARM PLUG CB SKN BAT 2 ARM PLUG CB SKN BAT Battery 3 ARM PLUG CB SKN BAT Battery 1 SOURCE SELECT ARM4 a rsks n nuna cd FIRE SELECT x 3 ARMS SELECT x 4 Figure 7 3 VEX Pyro Electrical Diagram 2 Rev 0 Date 06 02 2004 Page 7 15 ARM SAFE PLUG CB SKN PYR DT ARM SAFE PLUG CB SKN PYR T ARM SAFE PLUG CB SKN PYR DT ARM SAFE PLUG CB SKN PYR D ARM SAFE PLUG CB SKN PYR T SA Pyro CPS Pyro CPS Pyro Spare Pe Ref VEX T ASTR TCN 00349 A D Issue 2 Rev 0 4 Date 06 02 2004 ASTRIUM r 5 Page 7 16 7 4 GROUNDING amp EMC 7 4 44 Grounding The Venus Express selected grounding concept is a Distributed Single Point Grounding DSPG The main characteristics of this concept are the following Q all primary power supplies are referenced in a single point located in the PCU Q primary power supplies are galvanically insulated from secondary ones Q all secondary
117. er Transition Mode TTM and the Orbit Control Mode OCM It is still possible to manage the occultations during the mission as initially foreseen on MEx the specific algorithm being also able to treat occultation tables This modification will increase significantly the robustness of the system to a tracking loss at Star Tracker level which could occur for other reasons than explicitely predicted in the Mars Express initial strategy recalled previously This will also increase the robustness of the sytem to the solar flares since Venus Express radiation analyses showed that the probability to have a tracking loss in this environment is almost equal to 1 In case of tracking loss the current AOCS mode will not be interrupted continuing on gyro measurements and no failure will be declared by the FDIR up to a certain duration which can be selected on ground to be compatible with occultations and Solar Flare duration Ref VEX T ASTR TCN 00349 YA Vi Issue 2 Rev 0 L Date 06 02 2004 W E A ASTRIUM enus express pada 24420 11 9 UNIT LEVEL 11 9 1 Failure Management of Intelligent Units Failure detection of intelligent AOCMS respectively other bus units is managed at 2 levels through some surveillance internal to the units and via additional surveillances implemented within the AOCMS respectively DMS S W Failures detected internally of the units are signalled to the AOCMS respectively DMS S
118. er d Figure 9 3 1 RTU Block diagram IEEE 1355 link TMTC J Module 28 V converter Figure 9 3 2 Block diagram Ref VEX T ASTR TCN 00349 r Issue 2 Rev 0 _ Date 06 02 2004 ASTRIUM UL w GNUS expres Page 9 10 9 4 SOLID STATE MASS MEMORY The SSMM features the following functions Three 4 Gbits Memory Modules MM providing 12 Gbits user capacity In case of failure of one complete memory module the remaining capacity 1s 8 Gbits Two redundant controller paths each path providing One Memory System Supervisor MSS which performs the overall SSMM control and monitoring tasks One PM Interface Controller PIC which provide two bi directional IEEE1355 interface to the DMS processor modules from which it receives the packets housekeeping and science and to which it sends the events and the housekeeping data and some requested packets MTL content dump One User Interface Controller UIC which provides two bi directional IEEE1355 interfaces to the payloads with high data rate VIRTIS VMC two interfaces with the Transfer Frame Generators TFG of the CDMU and the interfaces with the memory modules One File and Packet controller which controls and manages the access to the Memory Modules and performs the file management functions One Input output Communication Switching Matrix CSM One DC DC converter which provides the necessary voltages to t
119. ered to the payload instruments RF planets configuration combined with the need to keep the cold face X away from the Sun lead to implement a second HGA antenna 2 that will be used during approximately one fourth of the mission centred around the inferior conjunction It 1s similar to Rosetta MGA 1 offset antenna 0 3m diameter Only the mechanical support had to be modified Due to its small dimension HGA2 1s X band only Main is very similar to the one of Mars Express with a smaller diameter 1 3m instead of 1 6m because maximum distance is smaller The RF Communications function will transmit X Band telemetry 8 hours per day at rates between 19 and 228 kbps depending of the Venus to Earth Distance An average of 2 Gbits of science data can thus be transmitted to Earth every day A variable telecommand rate of 7 81 to 2000 bps overall is foreseen during up to 8 hour per day Top floor LGA orientation is modified w r t Mars Express in order to take into account the planets configuration Data Handling the existing Mars Express design allows to fulfil Venus Express requirements with no modification Software modifications with Mars Express are mostly limited to RF communication function and AOCS function 0 Rev FFFIEPIZSFFERFNEISPIZFFS EAEE Pe ee Pete ieee te Pee FEFTEPFTERTEPEFISEPTERTFEI a a
120. etween the Main Engine Boost Mode MEBM and the SAM is automatic onboard Fj Ref VEX T ASTR TCN 00349 f S 2 Issue 2 Rev 0 L Date 06 02 2004 TA ASTRIUM l express 5 21 Ground commanded transitions Other transitions between AOCS modes or Modes subphases are commanded by the ground At AOCS Mode level this is for instance the case of the transition between the Normal mode NM and the OCM to start an orbit correction the transition between the NM and the Main Engine Boost Mode MEBM or the transition between the NM and the Braking Mode BM The transition from the Safe Hold Mode and the Normal Mode 15 also under ground control Between the subphases of the Normal Mode the transitions are also managed from ground the various capabilities offered by this Mode being used according to the mission needs The ground commanded transitions can be managed either by TC or by the Mission Time Line MTL This latter possibility enables the ground to build operational sequences including several transitions This is especially useful for instance when a slew manoeuvre to be performed in Normal Mode is necessary before a switch to another mode OCM MEBM BM Some time constraints exist in these sequences in order to ensure mode convergences before transitions Wheels off Loading The wheel Off Loading Phase WOLP is a sub phase of the Safe and Hold Mode SHM and also of the Normal Mode During the SHM th
121. furic anodisation views Venus Express external 3 1 Figure 4 Ref VEX T ASTR TCN 00349 n Issue 2 Rev 0 Date 06 02 2004 N A ASTRIUM an express Page 4 8 4 3 2 Radiators The S C main way of heat rejection are the radiators located on the Y panels for the platform and the X panel for the payload The proposed thermal design uses gt 1 72 m room temperature radiator opened directly on the platform walls that rejects the platform units and some payload units dissipation The available area of panels demonstrates important margins Those areas of the panels in excess w r t the required radiator size are covered by MLI blankets gt 0 22 m of radiator area for PFS O payloads cryogenic interface gt 0 28 m for the X wheel paddle The table provided Figure 4 3 2 presents the radiator areas in the current stage of design Trimming of the radiators 1s possible up to the final integration campaign of the spacecraft radiator size will be adapted 1f required after TB TV test correlations VEX T ASTR TCN 00349 ox Rev 0 f 06 02 2004 ASTRIUM enus express ie Nonna Wawa 7932 21300 37 2 S5MM PCLI PDLU Uso YMC IMU 1 2 SOR COMU 1 Y wal SAD RTU 2 TRSP 1 TRSP 2 TW 2 2 VIRTIS cooler VIRTIS ME IMLI SPICAY PFS Figure 4 3 2 Venus Express radiators definition VEX T ASTR TCN 00349 2 0 06 02 2004 4
122. g RTU availability and allows for a 20 sec lag time for time tagged TC which is about 10 times the RTU reconfiguration delay The DMS S W protects itself from VHF HF NF jitters by introducing time margins greater than worst case jitters determined through tests with actual H W and S W At each 1 Hz cycle the DMS S W OBCP Manager checks that sufficient CPU time is available before launching the execution of OBCP during this cycle If no sufficient time is available the OBCP execution will be re attempted at the next 1 Hz cycle p adt Ref VEX T ASTR TCN 00349 A Vi Issue 2 Rev 0 Date 06 02 2004 w k is E A ASTRIUM GNUS express Bus In case two successive overruns happen despite of the preventive provisions implemented in the DMS S W the overload is considered as confirmed by the DMS S W It consequently provokes a processor halt that in turn triggers the automated CDMS reconfiguration and system re initialisation Fault detection of the DMS S W is managed at 2 levels through some internal surveillances and by 4 watchdogs implemented within the RM but the result 15 the same since what the DMS S W does in case it detects internally a major failure is simply halts thus provoking voluntarily a triggering of the watchdogs Failure recovery and isolation are fully managed by the automatic H W controlled reconfiguration sequence Failure detection of the AOCMS S W is managed on 2 levels by t
123. h 1s attached to the top of the CDMU This approach enables easy PROM content updates on ground since the complete module is exchanged without opening the CDMU The PROM s be replaced by EEPROM s for on ground use The Safeguard Memory is permanently on and contains the nominal context 1 e the data necessary for the Processor Modules to restart automatically in the same mode as they were before the reconfiguration and the Safe Mode context which is the data necessary for the processor modules to restart in Safe Mode The Safeguard Memory is comprised of the following Ref VEX T ASTR TCN 00349 EADS Issue 2 Rev 0 Date 06 02 2004 ASTRIUM r express Page 9 7 SGM RAM in which the nominal context is stored and featuring 64 kwords as implemented using two HX 6228 128kxs SEU immune SRAMs SGM EEPROM in which the Safe Mode context is stored and featuring 64 kwords as implemented using two MEM8129 EEPROMSs EPHEMERIDS PARAMETERS INTERNAL CDMU UPDATE BY GROUND ON DMS PREDEFINED RECONFIGURATION LAST MTL POINT OF DMS i RESTART POINT PROCESSOR AND MODULE EPHEMERIDS CDMU I O COARSE TIME AND CLOCK FOR TM RECOVERY FORBIDEN RECONF ON BOARD ere TIME COPY PC ABOUT ONCE PER DAY ON DMS CONTROL mo START FOI RESTART POINT AND DEFAULT SW AOCS CONTEXT FOR DMS SAVEGUARD AND AOCS FOR CURRENT MODE CONTINUATION SW SW 1Hz INIT
124. he AOCMS S W itself and because the S W could possibly not be correctly working because of the failure by the DMS S W which monitors the correct behaviour of the AOCMS S W Whether an AOCMS S W major failure is detected within the AOCMS S W or by the DMS S W the final result is the same since what the AOCMS S W does in this case 15 to halt and this is considered as a major failure by the DMS S W and what the DMS S W does 15 to halt also As for a DMS S W major failure the final result 15 therefore a triggering of the watchdog monitoring the DMS S W Failure recovery and isolation are fully managed by the automatic H W controlled reconfiguration sequence which follows watchdog triggering Failure protection against irreversible actuators activation due to AOCMS S W error to avoid inadvertent and so irreversible actuator activation H W protection mechanisms are implemented Over speed for the Reaction wheel control and Time out for thrusters control During nominal thrusters activation short duration use of the gyros provides thrusters failure detection Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 p Date 06 02 2004 ASTRIUM egenus express Page 11 13 11 6 TT amp C FDIR SUMMARY The TT amp C FDIR is the major modification between Venus Express S C FDIR and Mars Express S C FDIR The modifications are induced by the design changes of the TT amp C subsystem introduced between MEX and VEX mainly the addition
125. he SSMM internal electronics The controller board is powered in conjunction with the power converter The Memory Modules are switched on by command under control of the MSS Ref VEX T ASTR TCN 00349 d Issue 2 Rev 0 EADS Date 06 02 2004 Page 9 11 ASTRIUM express Main Power N Main Power R PDU N DC DC DC DC HPC ON amp HPC OFF N Converter Converter HPC ON amp HPC OFF R TTT Relay Status Relay Saws R Relay Status N DCCA DCCB Relay Status R Temperature N Temperature R Sec Voltages N Sec Voltages R N Redundant Controller amp Converter PM DMS A R PM DMS B Nominal Controller amp Converter Memory System PM Interface Supervisor Supervisor MSSA Parallel Port Fileand Fileand Parallel Port Use Communicatio Packet Packet Communicatio Interface Switching Controller Controller Switching Controller Matrix Matrix VIRTIS UIC B Control amp Data I F A Module Control Module Control Module Control Memory partition 3 Memory partition 3 Memory partition 3 Memory partition 2 Memory partition 2 Memory partition 2 Memory partition 1 Memory partition 1 Memory partition 1 Memory partition 0 Memory partition 0 Memory partition 0 Latch Up Switch Latch Up Switch Latch Up Switch Memory Module 0 Memory Module 1 Memory Module 2 Figure 9 4 1 Solid State Mass Memory
126. he control 15 performed with thrusters but in a large angular corridor 15 deg in order to reduce the number of thruster actuations necessary to control the spacecraft attitude which is stable around this attitude thanks the aerodynamic forces This mode also allows around the pericentre to off load the wheels if necessary using the aerodynamic torques not the baseline Otherwise they will be maintained at a constant rate during the atmosphere pass The STR is maintained in stand by mode because its implementation has not been optimised with respect to this specific attitude and the SADM 15 controlled in a fixed position Hold during the Braking Mode keeping the Solar array orientation optimised for aerodynamic effects The accelerometer measurements can be sent to the ground to help for ground Navigation purpose age Ref VEX T ASTR TCN 00349 2 Issue 2 Rev 0 f Date 06 02 2004 TA ASTRIUM r express a Page 5 15 5 4 AOCS GENERIC FUNCTIONS The AOCS modes use generic functions for the guidance the attitude estimation and the actuators management At the Software level these functions are common objects used by several modes Guidance The role of the guidance 15 to provide onboard the reference attitude to be followed at each time of the mission by the attitude control and the commanded position of the Solar Array position The analysis of the mission needs shows that 4 types of guidance are n
127. hes on restricted areas These White patches will be added at the very end of the integration phase and the best candidates are Betacloth or Nextel performances still under study Currently an embossed Kapton MLI 15 baselined on all the spacecraft walls and white patches are implemented on the X wall around the wheels deflectors Another issue concerns the internal layer Due to their restricted temperature resistance Mylar is avoided and Dacron 15 partially removed the internal layers are VDA Kapton layers and Dacron is replaced by a tissuglass spacer only for the first 10 layers due to mass constraints The insulation efficiency has been improved by several points of design The Venus Express MLI will be composed of 23 Kapton layers separated with spacers order to increase the efficiency of the ideal blanket compared to a classical one 13 layer Moreover all the interfaces that are the principle sources of heat leakage are carefully designed The overlaps between the MLI will be interleaving overlaps reducing heat loads to the spacecraft Vespel stand offs will be used instead of aluminium stand offs because of their lower conductivity All the grounding points aluminium will be located in the overlapping area and covered by a neighbouring blanket overlap as far as possible If not a dedicated MLI cap will cover them A high temperature titanium MLI is designed to cover the ring cavity around the Mai
128. his period 4 wheels and gyro configurations are selected and several surveillances are inhibited Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 p Date 06 02 2004 ASTRIUM enus exoress Page 11 22 Retrieve Context SBM NO Nominal SAM Separation Sequence or Level 4 Ist reboot due to AOCMS AOCMS boot Yes Level 5 after 2 reboots due to AOCMS flag gt 2 select redundant units select nominal units inhibit all surveillances inhibit transition to SHM Level 3 after any Surveillance triggering switch all units to B and restart SAM SBM SAM sequence managed by the DMS Master switch OFF mm ee ae ae stowed deployed solar solar arrays arrays if authorized go to BPP go to StAP Level2 ES or NES with no redundancy available for at least one Go to SBM upon DMS request suspected unit SBM SAM sequence managed by the DMS Master switch OFF Figure 11 8 2 AOCS FDIR levels Levels 4 amp 5 after any Surveillance triggering with all units already on B request PM reboot Level 1 NES with Redundancy available for all suspected units SHM round TC Normal Mode Other operational modes p adt Ref VEX T ASTR TCN 00349 E Vi Issue 2 Rev 0 L Date 06 02 2004 b A ASTRIUM enus express Page 11 23 Star Tracker mode management The
129. identical to the one of Rosetta S Band Rx 1 X Band Rx 1 uso L LGA 1 2 0 6 LGA 2 K 2 Ks HGA 2 K 5 2 2 VeRa Experiment with USO is expanding the S C RF System Capabilities Around the Venus planet the operations of VeRa will be unique in the sense the S C HGA1 will have to be directed to Earth for specific S C to Venus conjunctions while no S C data transmission will be allowed for an optimal accuracy of the sounding measurements Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 Date 06 02 2004 Page 2 8 ASTRIUM r express 2 3 6 VIRTIS The instrument Visible and Infra Red and Thermal Imaging Spectrometer 15 a near UV visible and infrared spectro imager from 0 25 to 5 um wavelength range working In various operating modes and covering a large range of observations from pure high resolution spectrometry to spectro imaging The main scientific goals of VIRTIS a key instrument within the science payload are from the atmosphere detailed analysis all layers clouds and markers tracking to any potential surface measurement temperature mapping hot spots including surface atmosphere interaction phenomena meteoro
130. iented to Y directions In order to reduce the radiative coupling between the paddle and the X wall MLI OSR deflectors are implemented on the X wall in the vicinity of the paddle The X reaction wheels thermal dissipation is transferred by means of a thermal strap to the X wall acting as a radiator A thermo switch heating system is installed to compensate the environment changes and or the unit thermal dissipation when non operated gt Payload units The payload units can be divided into two categories gt the internal payloads which thermal control is directly dependant from the S C gt the external payloads that have their own thermal control and are conductively and radiatively decoupled from the S C Most of the internal payload units are collectively controlled in X enclosure They are accommodated on the X shear wall and radiators are implemented on the X closure panel to control the cavity environment temperature The thermal control takes advantage of the transient operation profile For more demanding units like the SPICAV SOIR the VIRTIS cameras and the PFS spectrometer featuring their own thermal control special precautions are taken on the design of their conductive and radiative isolation VIRTIS and PFS are provided with dedicated cryo radiators implemented on the X side of the Spacecraft VIRTIS radiator being integrated to the optical module Whatever the VEX T ASTR TCN 00349 2
131. igure shows the VEX Power Storage block diagram Battery 1 i 28 Series connection of 6 cells to form a string I Pat Charge Regulator 16 strings in to form the battery cell array Bat Discharge Regulator Battery 2 BCDR2 Series connection of 6 cells to form a string l Bat Charge Regulator 4 16 strings in to form the battery cell array Bat Discharge Regulator Battery 3 BCDR3 j Bat Charge Regulator Series connection of 6 cells to form a string Bat Discharge Regulator 16 strings in to form the battery cell array Nom Pyro I Fs Nominal Pyro board Nom Pyro I Fs Redundant Pyro board Figure 7 2 2 VEX Power Storage diagtam _ Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 f nc f D p Date 06 02 2004 AS ASTRIUM ENUS exoress B E V 7 23 5 Power Control The Power Control Unit PCU converts the solar array and battery power inputs into a regulated main bus voltage at 28V 1 The main bus regulation is performed by a conventional three domain control system based on one common and reliable Main Error Amplifier MEA signal which controls the two APRs one per SA wing and the three BCDRs one per battery Power management is supported by an adequate measurement of the power parameters within the PCU This includes array current and voltage BDR output current battery charge and discharge currents total main bus current and voltage
132. ility from the ground In addition it is not possible to transmit simultaneously via both LGAs because they have the same polarisation 8 4 2 communications After completion of the LEOP phase communications are done in X band via the selected HGA Communications in X band are possible only when the Spacecraft is accurately Earth pointed Uplink The X band telecomand 15 recetved in cold redundancy because there is no coupler between the antennas and the X band receivers even if both X band receivers are ON Remark a redundant path might be possible in S band but the Cebreros ground Station which 15 baselined for VEX operations has only X Band capability that means that a backup in S band would require the use of another ground station In addition uplink via LGAs would require the use of the DSN Network for long distances Refer to link budgets for details Downlink The telemetry is transmitted in cold redundancy one X TX emitter feeding one TWT amplifier feeding the selected HGA ae Ref VEX T ASTR TCN 00349 _ A D Issue 2 Rev 0 _ Y Date 06 02 2004 AA ASTRIUM exoress Pus 8 43 principles As explained in the previous sections the Uplink and the Downlink operate in a cold redundancy scheme Due to the criticality of the uplink unrecoverable TC loss means loss of the mission an on board TC link monitoring application is implemented in view of detecting failures that prevent the
133. ination on Zs top floor and Xs faces For that reason accommodation of payloads on the top floor has been avoided as far as possible and RW radiator design on Xs panel has been improved Accommodation of the VIRTIS cryogenic radiator on the non illuminated Xs face of the spacecraft Accommodation of radiators for other payloads on the Ys sidewalls Accommodation of the HGA2 antenna and associated diplexer diameter reduction Accommodation of a new solar array populated with GaAs cells and OSR mirrors Slight changes in SAS and LGA orientations c The lessons learnt from Mars Express are mainly directed towards improving the integration issues In particular during the panel closure operation enlarge the cut outs for harness routing on shearwall edges To increase the distance between CPS piping and some bus units improve waveguide design accessibility attachment flexibility To add cut outs for endoscope Another key lesson learnt from Mars Express is the control of the centre of mass location d The spacecraft dry mass has been minimized in the aim to cope with the launcher capacity The interest for the mission of increasing the propellant mass as far as possible is understood The CoG location is carefully assessed and refined by analysis and measurement along the development Summarising the Venus Express design is a close derivate from the Mars Express one with minimal changes whe
134. ine and especially mandatory at quadrature time The AOCS software ensures autonomously the sign change necessary in this guidance law when changing from one HGA to the other Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 Date 06 02 2004 Page 525 ab E ASTRIUM r express End of Mission HGA 1 IVQRAQYI Start Quadrature 2nd Venus day RR X 0 ad HGA 2 selected on 0 Inferior Conjunction side selected Superior Conjunction side Start 4 Venus day Z 0 6 VOI Start of science Figure 5 6 1 Vex HGA switching strategy Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 Date 06 02 2004 Page 5 26 S ASTRIUM pem express Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 p Date 06 02 2004 enus express Page 6 1 6 PROPULSION SYSTEM ARCHITECTURE The Venus Express propulsion system is based on the bi propellant Mars Express propulsion system with higher propellant mass This chapter gives a description of the design of the VENUS EXPRESS propulsion system including the CPS schematic More detailed information is provided in the CPS Design Report VEX RP 00002 EU ASTR and User s Manual VEX MA 00001 EU ASTR Ref VEX T ASTR TCN 00349 iba E S Issue 2 Rev 0 _ Date 06 02 2004 AA ASTRIUM express 6 1 DESIGN DESCRIPTION The Venus Express CPS is a helium pressurised bipropellant system us
135. ing monomethyl hydrazine MMH as the fuel and mixed oxides of nitrogen with 3 nitric oxide MON 3 as the oxidant also referred to as NTO its main constituent The main engine used for Venus orbit insertion has a thrust of 416 N and a specific impulse of 317 seconds Four pairs of 10 N thrusters 4 primary 4 redundant are provided for trajectory corrections and attitude control reaction wheel unloading These components are the same as used on Eurostar 2000 However the intended use of the main engine on Venus Express with a propellant mass ratio close to 90 between main engine and 10 N thrusters is higher than on Eurostar where this ratio does not exceed 75 This 1s a specificity in the use of the main engine The limitations of main engine use due to propellant tank functional constraints have been identified in the CPS user s manual The CPS is designed to operate in a constant pressure mode during main engine firings for capture manoeuvre and first part of apocentre reduction manoeuvre using a regulated helium supply The latter manoeuvre 1s pursued with main engine in blow down mode Following completion of main engine manoeuvres the regulated helium supply and the main engine are isolated The last part of the apocentre reduction manoeuvre 15 achieved with the 10 N thrusters The thrusters are used in blowdown mode 1 e the system pressure reduces as propellant is consumed This represents a major simplification to the design
136. ion between the DMS and the AOCS the reconfiguration to the Sun Acquisition Mode level 2 or 3 is implemented in three steps 1 The AOCS sends a Safe Mode request to the DMS and switches to the Stand By Mode 2 The DMS answers by sending to the AOCS the AOCS Master Switch Off command switch off of all AOCS equipments except the IMPs 3 The DMS switches the AOCS to the Sun Acquisition Mode at the end of the AOCS Master Switch OFF Ref VEX T ASTR TCN 00349 ADS Issue 2 Rev 0 a uL Date 06 02 2004 GNUS XPress Page 11 21 Specific case of the Main Engine Boost Mode In the MEBM a specific logic is used due the criticality of the insertion manoeuvre Q The number of active surveillances is reduced to the minimum the equipments that are not mandatory for the MEBM completion are off STR SADE and Reaction Wheels Except in the case of the time out triggering the active surveillances are classified as NES leading to continue or restart the manoeuvre after the triggering If a reconfiguration of the AOCS H W is necessary all cases except IMP failures the Main Engine is stopped and a specific procedure is run at DMS level to ensure that the thrust and the AOCS control can be restarted within 10s a thermal constraint at Engine level requires that the motor is not restarted between 10s and 500s after cut OFF If the manoeuvre 15 declared critical by the eround the Main E
137. ion of software contexts and patches and provides SGM access to AOCMS SW through a dedicated TM TC interface The SSMM SW receives data from the DMS SW and directly from some payload instruments high rate science data and participates to the execution of file transfer requests by sending data files to the Transfer Frame Generator TFG VEX T ASTR TCN 00349 Issue 02 Rev 0 EADS L7 Date 06 02 2004 ASTRIUM r express pae 105 TC SW contexts AOCS Sub system Command amp Monitoring AOCS SW PTS STR SW GYROS SW Sub system Command amp Monitoring P F HK TM amp Low Rate Scientific Data TFG Data Files Sub system Command Monitoring amp Low Rate Scientific Data P L High Rate Scientific Data Figure 10 2 1 Venus Express on board software external interfaces Ref VEX T ASTR TCN 00349 E AD S Issue 02 Rev 0 Date 06 02 2004 ASTRIUM r express Page 10 6 Internal interfaces The main components of Venus Express on board software have software interfaces between them through OBDH and IEEE 1355 serial links The DMS SW sends File Management patch and dump requests to the SSMM SW The SSMM SW sends files information memory dumps and SSMM health status to the DMS SW The DMS SW sends Time Line OBCP handling patch dump configuration and reconfiguration orders to the AOCMS SW The AOCMS SW sends acknowledges dumps monitoring
138. ions SADM Commands Ephemeris S C to Earth parameters Ephemeris S C 46 sn dirselions Large angle slews for Propagation Time Earth Sun pointing Autonomous acquisition SHM Attitude Guidance Function Earth pointing HGA misalignments Fixed quaternion lt Normal Ground Commanded Mode ie dco didi iue Quaternion Profiles Attitude q Propagation Guindance Polynomial description Profiles for MAT 13261 observation slews Figure 5 4 2 Guidance function architecture aT Ref VEX T ASTR TCN 00349 A Issue 2 Rev 0 _ Date 06 02 2004 C ASTRIUM r express Reaction wheel management function This function is active in all the modes controlled through wheel torques Normal Mode and Safe Hold Mode at the end of the attitude acquisition sequence but also when the wheels are kept to a constant speed through a specific control loop but not used in the AOCS control as in Orbit Control Mode Thruster Transition Mode or Braking Mode Six states of the wheel configuration are possible with this function depending on the control of the wheels in torques t or in speed s For instance the nominal operation in Normal Mode uses 3 wheels in torques 3t but could sometime require a fourth wheel if a hot redundancy is usefull At During trajectory corrections the configuration includes 3 wheels controlled in speed 3s Intermediate states are necessary between these basic configuratio
139. is operation is automatically commanded onboard on the basis of wheel kinetic momentum criteria when the AOCS has reached the Earth Pointing phase of the mode SHM EPP During the NM this operation is nominally commanded by the ground out of observation phases at a date commanded by the ground In order to limit the orbit disturbances due to the thruster actuations it 1s also possible to perform the wheel Off Loading with a spacecraft attitude defined by the ground in the Fine Pointing Inertial Phase FPIP after a slew manoeuvre As a security the wheel off loading can also be triggered automatically in case of wheel over rate detection from the Fine Pointing Accuracy Fine Pointing Inertial and Gyro stellar Ephemeris Pointing sub phases of the Normal Mode NM FPAP NM FPIP and NM GSEP During these three phases the ground has the capability to inhibit the automatic procedure Between modes where the Wheel Off Loading is authorised and Modes where it 1s inhibited it 1s recommended to the ground to anticipate this inhibition to avoid a thruster pulse just at the time of the transition Ref VEX T ASTR TCN 00349 EP S 2 Issue 2 Rev 0 Lr Date 06 02 2004 7 ASTRIUM l express Page 5 22 Transition validity check For the Mode and Sub phases transitions commanded by the ground some conditions may be required to ensure that the AOCS behaviour is adequate after the transitions The onboard Software pe
140. itions onboard Automatic transitions managed by the onboard Software are implemented when a complete sequence of operations has to be performed without ground intervention involving several AOCS Modes or several AOCS Modes subphases This situation exists for the attitude acquisition sequence either during the nominal sequence or in case of reacquisition after a failure The transitions between all subphases of the Sun Acquisition Mode SAM are automatic onboard The transitions between all subphases of Safe Hold mode are automatic onboard but the transition between Hold Phase HP and Earth Pointing Initialisation Phase EPIP can be inhibited by the ground or the SW in case of failure to avoid wheels use The transition between the SAM and SHM 15 automatic onboard leading to a completely automatic attitude acquisition reacquisition sequence It can be however inhibited by the ground or by the SW in case of failure For both last cases the ground inhibition is active only if the Spacecraft Elapsed Time 1s lower than a given value For the orbit control manoeuvres in OCM which have to be followed by a tranquillisation performed in Thruster Transition Mode TTM automatic transitions are also implemented The transition between the Orbit Control mode and the is automatic onboard The transition between TTM and the Normal mode NM is automatic The transition b
141. l and on the lower floor The main engine is located under the lower floor and orientated in roughly Zs direction while the eight thrusters are located at the four lower corners of the spacecraft The two solar wings are mounted to the Ys sidewalls and can rotate around Ys axis The attachment interfaces are identical to Mars Express Each wing is composed of two panels and a yoke made of two parts There two fixed high gain antennas The antenna is accommodated on the Xs closure panel same attachment interface as Mars Express smaller diameter and the HGA2 antenna is accommodated on top floor and orientated In quasi opposite direction The payloads are accommodated as follows PFS VIRTIS and SPICAV are accommodated on the Xs shearwall with a nadir field of view in Zs direction MAG sensors and deployable boom are accommodated on the top floor MAG electronics is accommodated on the Ys sidewall VMC and VERA are accommodated on Ys sidewall ASPERA is accommodated outside the spacecraft on the Ys sidewall MU and underneath the lower floor IMA Most of the bus electronics are accommodated on the inner side of the Ys sidewalls in the same location as on Mars Express The WIU is gathered in the same cavity as on Mars Express and attached to the sidewall top floor and Ys shearwall The AOCS units are accommodated in the same location as on Mars Express Ref VEX T ASTR TCN
142. lator 1 Array Power Converter 1 MPPT Control 1 MPPT Majority 2 3 Array Power Converter 3 MPPT Control 1 Array Power Converter 2 MPPT Control 1 MEA Array Power Regulator 2 Array Power Converter 1 MPPT Control 2 Array Power Converter 2 MPPT Control 2 MPPT Majority Voting 2 3 Array Power Converter 3 MPPT Control 2 Figure 7 2 1 VEX Power Generation Diagram VEX T ASTR TCN 00349 2 Rev 0 06 02 2004 EI 28v Main Bus MPPT Control 1 Main Error Amplifier 2 3 Voting MPPT Control 2 Ref VEX T ASTR TCN 00349 2 S Issue 2 Rev 0 f p Date 06 02 2004 AS ASTRIUM GNUS express Page 7 8 7 2 2 Power Storage Three batteries supply the spacecraft power when the Solar Array is not illuminated by the sun or in case the power demand 15 higher than what can be generated by the Solar Array The energy is stored within 3 identical batteries of 24 Ah based on low mass Li Ion technology Each battery is built with 16 parallel strings of 6 serial 1 5 Ah battery cells The cells are based on the Sony Hard Carbon typ 18650 which 15 a cylindrical cell Each battery has the following parameters Maximum Battery Voltage 25 2V Minimum Battery Voltage 15V Battery Capacity 24 Ah Battery Energy 518 Wh Ref VEX T ASTR TCN 00349 EADS Issue 2 Rev 0 p Date 06 02 2004 ASTRIUM express pue 278 The following f
143. lemented in two externally redundant Control and Data Management Units CDMU Each CDMU features e two Processor Modules one dedicated to AOCMS software execution the other one to Data Handling software execution The Processor Module design is based on a flexible 16 bit MA3 1750 microprocessor with 1Mword of associated RAM and 512 Kwords of EEPROM e Two reconfiguration modules each one containing an accurate clock function stability better than 3 107 day and a watch dog function which when it triggers sends reconfiguration request to the High Power Command Module One High Power Command Module HPCM containing o decoder which processes the telecommands transmitted by the transponders can generate 64 High Power Commands CPDU commands and transmits the TC segment to the DMS PM o A reconfiguration function which executes an autonomous reconfiguration of the CDMU when it receives at minimum 2 among the 4 reconfiguration requests generated by the 4 reconfiguration modules o A Transfer Frame Generator TFG which contains 3 virtual channels for real time telemetry for telemetry memorised in the SSMM and for the idle frames It provides the possibility to select the convolutional coding and or the Reed Solomon coding The PM DMS delivers the subcarrier clock and the bitrate clock which can vary from 8 bps to 250kbps depending on the Venus to Earth distance on the selected frequency band a
144. logy volcanism The core VIRTIS design 15 fully re conducted from the Rosetta one No change has been recorded yet but the needed adaptation minor indeed but existing of the IEEE 1355 protocol on the high speed link dedicated to science data retrieval The VIRTIS Opto mechanical Unit from Rosetta The instrument measurement unit consists of two independent channels grouped within an Opto Mechanical OM unit The M Channel operating in the 0 3 to 1 um CCD and 1 10 5 um photo conductive HgCdTe spectral ranges linked to a PEM M proximity electronics And the H Channel operating in the 2 to 5 um photo conductive HgCdTe 436x270 pixels detector matrix spectral range with its PEM H proximity electronics These two channels are operated through a centralised system located within a Main Electronics ME unit which fully drives the overall instrument L Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 Date 06 02 2004 Page 2 9 express The critical opto mechanical part of the instrument is fitted to a specific interface plate worked as part of the Spacecraft structure to get its cryo coolers M and H channels well connected through heat pipe system to their specific radiator It is to be noted here that the de icing heaters needed for Rosetta to be operated during the hibernation phase be removed from the VEX design
145. loyment is driven by hinge spring The boom 15 latched at end of deployment and 15 then parallel to Zs axis SA Wing Deployed MAG Boom Deployed Figure 3 4 1 Spacecraft in Orbit Configuration FFF I FPL EES PE PP FIIPLPRPEPISPEIEPRIPEPFRISEER RC PTESSPREPFEEEPEZSPIFPI OF FSP PS EPL EE EE PP DES O TPE T PP EZE PPT PPEP PU E TE eee Pe eee ee Ses iii cee ar 2 z T EP ee O O EPP ee ee ee TET I Fg Pes ee ee ee Pe ee ee o Pee ee ee ee ee ee ee a 0D Y PPE ee ee E E ee ee ee ee ee ee lt N op Pe ee ee 1 Pee EE PEP T ee EPP ee x p gt NO gs UU iu FFT FF e Pee ee A gt PE O L AA 2 k oO EADS ASTRIUM PP a eee ee ee ee VERERIKETERERIERFERTEI ee ee pDrrsFIHRIPDRISPRIDRPEPHRIPEI eee ee ee ee Pe Pe a FETEFRTPTOEFTOEFETEFTOEEFFS E RESELLER JGEREFFRISFFISFISFFEREFI ee
146. ly only one can be selected as active a There are three separate batteries each one being In series with one BCDR PCU used in a 2 out of 3 active redundancy Should one battery or one BCDR fail the 2 remaining ones allow to ensure a graceful degraded mission limitation of operations eclipse The three Array Power Converters of one Array Power Regulator are used in 2 out of 3 active redundancy Should one converter fail the two remaining ones can still provide the power bus with SA wing power but with a power capability reduced by a third There are 4 PM in the whole CDMS 2 CDMU In each CDMU 1PM is DMS configured and the other one is AOCMS configured The DMS PM and AOCMS PM are used in out of 2 stand by redundancy ee Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 p Date 06 02 2004 The On Board Time generation function and the spacecraft monitoring and reconfiguration function of the Control and Data Management Units CDMU are used in 2 out of 3 hot redundancy In particular a System Alarm will lead to a reconfiguration only if it 1s detected by at least two Reconfiguration Modules among the four operational ones This allows to be one failure tolerant during critical phases The TC decoder of CDMU is used in out of 2 hot stand by redundancy both are ON but only is addressed by the ground Telecommand There 3 memory stacks in the SSMM are use
147. n Engine The Titanium foil prevents the MLI overheating during the engine firing It will be black painted FIBA in order to reduce the temperature of the external layer when sun illuminated and then the heat loads on the LVA stiffener and CPS pipes MLI blankets Dacron VDA Mylar are also used internally around the tanks and the CPS components to limit the heater power installed in order to prevent propellant freezing VEX T ASTR TCN 00349 2 Rev 0 06 02 2004 4 11 AA ASTRIUM express 4 3 4 Heating system Electrical heaters are used on Venus Express to prevent excessive cooling of units structures and propellant during the cold phases and eclipses payload non operation partial or total and safe mode They are also used when the payload is fully operational in order to comply with the units minimum temperature limits The heating system consists of 16 nominal heater lines extra 16 identical redundant lines corresponding to a potential 781 W nominal installed power Almost all the heater lines are controlled using bimetallic thermostats with fixed temperature set points in order to allow a large number of controlled areas Two thermostats are used in series to avoid closed circuit failure If the change in temperature set points 15 required a software control is implemented A set of three temperature sensors is dedicated to each circuit to provide an ON OFF regulation type using the on board software capa
148. n handling with a test console an EGSE and another PM The SSMM SW running on its own processor is designed to operate the SSMM and includes two main tasks the Start up amp Initialisation task and the Operational task The STR SW run on its own processor at unit level under the control of the AOCMS SW It is in charge of the Star Tracker STR management and provides three axis autonomous attitude determination to the AOCMS SW The gyros SW runs on a processor at unit level and provides angles and velocities to the AOCMS SW It performs the gyros management under the control of the AOCMS SW The Transponder SW runs on a processor at unit level It interfaces with the DMS SW by means of TM and TC only Several autonomous software are running in stand alone payload instruments under the control of the DMS SW Ref VEX T ASTR TCN 00349 S Issue 02 Rev 0 p Date 06 02 2004 ASTRIUM express pase 2109 Payloads 28 V Regulated Protected Lines Serial IEEE 1355 lines Standard Link Solid State Remote Mass Terminal Memory Unit Control Dual Band Control amp Data Unit t OBDH Data Bus Transponder M on Solar Array Drive CPDU Mechanism Commands Avionics Interface Thrusters Main Engine Reaction Star Cue Sun 4 x 10N 400 N Wheels Drive Tracker Accderos Electronics Acquisition Sensor Figure 10 1 1 Venus Express on boa
149. n necessary for accommodating the Venus Express specific requirements and constraints and implementing the lessons learnt from Mars Express Ref VEX T ASTR TCN 00349 gt Issue 2 0 p Date 06 02 2004 GNUS COXDIESS 0232 3 2 SPACECRAFT CONFIGURATION In the aim to avoid any risk of confusion a unique axis reference system Os Xs Ys Zs has been introduced at spacecraft level It corresponds to the Oa Xa Ya Za system used on Mars Express The Venus Express spacecraft is roughly a cubic box dimensions 1 05 m x 1 7 m x 1 4 m high same as Mars Express The overall configuration of the spacecraft box structure is as follows As part of the core structure four non removable panels which divide the spacecraft in six compartments The lower floor Zs panel The Ys shearwall which gives the spacecraft shear stiffness Xs Zs plane Xs and Xs shearwalls which give the spacecraft shear stiffness in Ys Zs plane As part of the secondary structure The top floor Zs panel which is non removable panel The Ys sidewalls which are opened in horizontal position during spacecraft integration Xs closure panels which close the box and allow access into the spacecraft The propulsion system is accommodated as on Mars Express The two propellant tanks are gathered in the centre part of the core structure and the propulsion units are accommodated on the 5 shearwal
150. n some cases even if a transition validity check and a H W configuration change is already performed by the onboard Software it is however preferable to ask to the ground to anticipate on the Mode transition to make the transition easier this can help to avoid the time loss due for instance to the switching ON sequence of the IMU 15 seconds without valid measurements or the Star tracker Case of the mechanisms The configuration of the Solar Array Drive Mechanism SADM is managed onboard during the attitude acquisition and back up modes During the operational phase the ground has the full capability to manage this unit depending on mission operations A ground intervention is especially necessary in NM FPAP in order to define the final position to be reached by the Solar Array for the further operation observation MEBM The SADM is not locked by the onboard S W Mode transition generic sequence The Mode transition basic sequence includes 3 main steps The validation of the Mode transition request when it is a ground commanded transition During this step the Software tests all the conditions necessary for the switching in the next AOCS Mode The execution of the Hardware configuration change for the considered Mode change This step involves exchanges between the AOCS and the DMS processors and S W at 1 Hz for the request of Hardware units switch OFF and switch ON The exact duration of this seque
151. n the Processor Module DMS or AOCS and all associated data It 1s loaded from the EEPROM or from the non volatile PROM in case of EEPROM failure by the Firmware according to the status defined for the Processor Module The Basic RAM area consists of eight 128 kword banks and each word is 24 bits wide to provide protection by an EDAC The proposed memory circuit 15 the TEMIC M65608 128 K x 8 bit CMOS SRAM manufactured in the their high performance CMOS technology named SCMOS The Start Up PROM and the RAM memory are protected by an EDAC The EDAC 1s a flow through type EDAC with one 24 bit memory port and two 16 bit CPU ports The EEPROM is used to store DMS or AOCS software so that they can be readily accessible to the Processor Module during Start Up The in flight programmable capability allow to introduce software modifications and patches in the EEPROM In this way the most up to date software version will be taken into account during the Start Up sequence The EEPROM area consists of four 128 kword banks and is connected to the internal parallel bus through buffers The memory circuit is MEMS129 128 K x 8 bit 5V EEPROM The Non Volatile Memory contains an uncorrupted copy of the initial DMS and AOCS software which it is not possible to erase or modify These software will be loaded during Start Up sequence if it turns out that EEPROM stored software are not correct The Non Volatile Memory is physically located in a cassette whic
152. nce Figure 8 1 2 VEX HGA selection along the mission The following table sums up the communication along the mission O PI A sC I dis S tz h Payload Commissioning Routine and Extend lt 0 78 AU X Band i Operations sA Ref VEX T ASTR TCN 00349 rA D S Issue 2 Rev 0 4 Date 06 02 2004 AA ASTRIUM T express Page 8 4 8 4 REDUNDANCY AND PRINCIPLES 8 4 1 LGA communications During the LEOP phase the communications are done in S band via the LGAs The nominal configuration is the following LGA1 15 routed to TRSP1 and LGA2 is routed to TRSP2 The communication via one LGA 15 only possible when the Earth direction 15 within its FOV 95 deg around its boresight In Sun pointing attitude rotation around X axis at 0 1 deg s the Earth is within the LGA1 FOV during half of the rotation and within the LGA2 FOV in the other half Q In Earth pointing attitude is Earth pointed for geometrical reasons see Figure 8 2 1 the Earth is within the LGA1 FOV and not in the LGA2 one So communications are recommended via LGA1 Uplink The telecomand is received in cold redundancy most of the time because only one LGA 15 visible from the ground station see above even if both S band receivers are ON Downlink The telemetry is transmitted in cold redundancy only one transponder emitter is ON at the same time It must be routed to the antenna that 15 in visib
153. nce a few seconds depends of course on the number of configuration changes necessary for the mode transition The starting of the next Software Mode as requested as soon as the AOCS S W is informed by the DMS that the last H W configuration change has been transmitted to the RTU The new Mode will start by the initialisation of the new functions and algorithms with all the adapted parameters Y Ref VEX T ASTR TCN 00349 2 Issue 2 Rev 0 f Date 06 02 2004 V ASTRIUM express T Page 5 24 5 6 HIGH GAIN ANTENNA MANAGEMENT Specific need for Venus Express On Venus Express in order to fulfill thermal constraints at spacecraft level 2 HGAs are used for the communications with the Earth The switch from one HGA to another has to be performed around quadratures in order to ensure that is selected on superior conjunction side and HGA2 inferior conjunction side of the mission At AOCS level both and HGA2 direction unit vectors are stored onboard and the AOCS will use the appropriate vector to compute the final attitude of the SHM or the attitude of the Normal Mode GSEP In SHM the adequate HGA parameters will be used when entering in this mode the final attitude and the attitude manoeuvres being computed autonomously by the software In Normal Mode the adequate HGA parameters will be used when entering in GSEP after a ground slew ensuring the correct pointing of the spac
154. nd on the spacecraft mode One Centralised Memory Module containing o A PROM cassette including 512 Kwords of PROM accessible by the 4 PMs and containing the default DMS and AOCMS softwares A SGM containing 64 Kwords of RAM and 64 Kwords of EEPROM Two power supplies one powering one processor module and one reconfiguration module and the other powering the remaining parts of the CDMU OBDH bus B 5 ASTRIUM AIU SSMM EGSE PMs IEEE 1355 links 2 RMs redundant CDMU 2 PMs 4 5 CDMU A PM Alive signals OBDH bus A redundant Genus express Alarms UVD Battery TBD 4 PMs PM Alive signals v v 4 Processor Module j link Reconfiguration amp clocks Module Ref Issue 2 Date 06 02 2004 Page 9 5 Centralised memory module SW source unloading mue w Context management b SGM 64 Kw RAM 64 Kw EEPROM High Power Command Module A g un ad Y signals signals Reconfiguration requests gt link j Processor Module IEEE 1355 links AIU SSMM EGSE PMs Reconfiguration amp clocks Module 4 PM Alive signals 4 RMs 2 PM CDMU 2 RMs redundant CDMU redundant A A PM 4 Alive signals 4 PMs Alarms UVD
155. nd the radiator to reduce the sun reflection and the IR flux received by the paddle and generated by the temperature of the surrounding Implementing heat pipes under the PCU and PDU to reject the high thermal dissipation of the PCU through a larger radiator Reducing heat exchanges through the MLI Two ways of improvement have been implemented The first step consists in changing the external layer coating black coating on MEX to reduce the temperature level of the blankets Embossed Kapton is baselined for all the spacecraft walls and white coating patches will be added on very critical areas where Kapton is still a too hot solution The second step is to increase the number of layers of the blanket Taking advantage of Austrian Aerospace experience on XMM a 23 layer blanket is designed Replacing the alodine treatment of the LVA ring by a clear sulphuric anodisation to minimise the LVA ring temperature level when sun illuminated Changing hot coatings on external units by colder ones e g PFS_S scanner LGA SADAM Mixing cells and OSR on the front side of the solar arrays panel The rear side 15 also completely covered with OSR VEX T ASTR TCN 00349 2 0 06 02 2004 4 4 NS ASTRIUM r a express 4 2 THERMAL CONTROL CONFIGURATION Most of the spacecraft units are collectively controlled inside thermal enclosures created by the spacecraft mechanical architecture in which the heat balance is cont
156. nditions which autonomously triggered the S C Safe Mode to preserve S C integrity See Safe Mode Capability to impose the CDMS PM configuration setting through CPDUs after reset If the Ground selects a not physically working configuration e g DMS PM PMI or PM4 such configuration will be attempted by the CDMS and rejected as non operational provoking a new reconfiguration Capability to disable the safing mechanisms those that trigger the Safe Mode at mission times when on board fault management should not interfere with the autonomous sequence in progress e g orbital burn The ground is of course given the possibility to re enable the safing mechanisms it previously disabled Capability to upload S W patches Application Programs APs or FDIR OBCPs p adt Ref VEX T ASTR TCN 00349 r Vi Issue 2 Rev 0 L Date 06 02 2004 W A ASTRIUM ENUS express Page 11 5 FDIR Data Logging The on board software provides logging of data related to the occurrence and the processing of an anomaly These logs are time stamped with the spacecraft elapsed time SCET Nominally these logs go to the SSMM On the other hand Critical Event Logs CEL are stored in SGM RAM and thus remain accessible to the Ground when the SSMM is off or not used Context Context data required during system initialisation are maintained in SGM RAM or E PROM that they are preserved in case of processor m
157. neric functions have been defined for this purpose at software level the Ground commanded guidance the Onboard Ephemeris propagation the Autonomous Attitude Guidance Function this latter function generating the guidance information necessary either for the fixed Earth pointing or for the Earth acquisition in SHM For Venus Express as for Rosetta the Autonomus Attitude Guidance function provides 2 independent guidance laws for pointing the Earth direction avoiding Sun exposure of sensitive faces of the spacecraft The difference between the 2 laws 1s the spacecraft Y axis orientation it can be set perpendicular to either the ecliptic plane Ecliptic option similar to MEX Guidance or the Sun Spacecraft Earth plane SSCE option For Venus Express the SSCE option is the nominal law to be used by the AOCS and the Ecliptic option 15 a back up law which can be used for very specific situations corresponding to very low values of the angle S C Sun S C Earth Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 f Date 06 02 2004 V ASTRIUM r express Page 5 16 Gyro stellar estimation function The gyro stellar estimation function is common to many AOCS modes It is initialised during the Sun Acquisition Mode SAM to prepare the following Earth acquisition operation SHM Safe Hold Mode It provides accurate attitude estimation during the Normal Mode of course but also in th
158. ng X Band and S Band RX TX S Band configurations featuring the HGA or ultimately a dual LGA mode WIU Load Overheating Surveillance WILOS Surveillance that no TWTA feeds a WIU load to prevent from over heating Should this abnormal condition exists a SW safe mode 15 triggered to passivate the anomaly through the Master Switch Off and establish a safe RF configuration by SAFCAP execution Modification for Venus Express SAFCAP The RF configurations established by SAFCAP depends first on the Earth SC distance flag stored in SGM if the flag indicates Near earth it corresponds to the LEOP Phase SAFCAP establishes a dual LGA S band configuration This is unchanged for VEX because this configuration 15 common between MEX and VEX if the flag indicates Far from Earth in the MEX design SAFCAP establishes a S band configuration with the HGA This is modified for VEX to establish a X band configuration instead of S band because HGA2 is a single X band antenna SAFCAP uses a flag stored in the SGM which indicates which has to be used for communications or HGA2 The start of the telemetry transmission done in SAM StAP in the MEX design has been postponed In SHM after the execution of the attitude manoeuvre from Sun pointing attitude to Earth pointing attitude WILOS This monitoring is deleted on Venus Express S C as there is no more loads in the WIU due to the implementation of HGA
159. ngine is restarted If not the MEBM ends in Back Up MELSP Q When the Main Engine can not be restarted double velocity triggering or shall not be restarted temperature monitoring triggering or when the manoeuvre is declared not critical by the ground the MEBM is ended in Back Up In this phase the 8 thrusters are used to produce both thrust and control torque 4 thrusters are continuously open since the four others are off modulated In case of time out triggering the Main Engine is switched off and the transition to the Sun Acquisition Mode 15 realised The manoeuvres that are not mission critical shall be declared by the ground has not critical because in case of failure the Main Engine restart may generate high torque at the SADM interface Specific cases for Hardware reconfigurations A specific management of the Nominal redundant configurations is necessary for the gyros and wheels due to the specific H W redundancy configuration Q Use of 3 gyros among 6 Q Use of 3 Reaction wheels among 4 A specific algorithm is used onboard to select from the H W IMU channels status and from the ground indications the set of 3 gyros to be used in case of reconfiguration Venus Orbit Insertion VOI prepration Before the Venus Orbit insertion manoeuvre it is necessary to avoid a transition to safe mode which could endanger the mission introducing inacceptable delays in the sequence For t
160. ns in order to spin the wheels for instance 3t 1s This function is also in charge of the generation of wheel torque commands in wheel frame and of the friction torque estimation necessary for compensation and for the failure detection It interfaces also with the Wheel Off Loading function Thruster modulator and selection function The selected amplitude modulator and on time summation algorithms are re used from Mars Express The modulator has only one working phase where the four thrusters can be used to produce a force along the satellite Z axis direction a force ratio P 0 1 is commanded to the modulator to control the 3 axes satellite attitude three torques are commanded to the modulator The modulator working frequency is 8Hz At each step the modulation type used ON modulation or OFF modulation is automatically selected so as to maximise the available torque capacity for attitude control In the case the torque capacity 15 insufficient with respect to the commanded control torque priority is given to the control and the commanded force ratio 15 automatically modified to recover the required torque capacity Moreover in order to limit the actuation delay the attitude control torque 15 always produced at the beginning of the actuation period To limit the number of thrusters ON OFF or to tune the control limit cycle amplitude when using thrusters the modulator output period can be changed to
161. nsole for PM hardware development and maintenance with an EGSE and with the DMS PM The PM Firmware is automatically activated when the CDMU board is powered on The PM Firmware provides the following functions Initialisation of the PM and acquisition of the hardware configuration of the PM DMS AOCMS or spare automatic software patches enabled or not Complete auto tests of the PM and its interfaces IEEE 1355 OBDH BC and RT activated when the CDMU board is powered on including a destructive test for the complete RAM contents Selective auto tests of the PM and its interfaces possibly called during nominal software operation DMS AOCMS to analyse a PM dysfunction during the lt Isolation gt part of a PM FDIR Software loader able to fetch software from the CDMU PROM or EEPROM possibly from the SSMM through IEEE 1355 from the EGSE through IEEE 1355 or from a debugger or test console through RS232 After having checked the loaded software integrity checksums it 1s able to start the loaded software automatically Minimal TM TC to execute requests sent by the current DMS PM via IEEE 1355 or OBDH auto tests housekeeping health status telemetry RAM and EEPROM patch and dump and other critical autonomous functions PROM monitor able to support software debug and hardware investigations The PROM monitor can execute basic instructions sent through the RS232 by the test console such as Load Soft
162. nted Afterwards the telemetry transmission 15 re started and the non critical heating lines are re powered The description is presented in more details in section 11 7 The following actions are included in the reconfiguration sequence initiated by the CDMS RM HPCM PM DMS A ON or OFF PM DMS B ON or OFF PM AOCMS A ON or OFF PM AOCMS B ON or OFF CK A ON or OFF CK B ON or OFF LCB A ON or OFF PMI to OBDH AorB PM2 to OBDH AorB PM3 to OBDH AorB PM4 to OBDH AorB PMI ON or OFF PM2 ON or OFF PM3 ON or OFF PM4 ON or OFF F Ref VEX T ASTR TCN 00349 gt Issue 2 Rev 0 Date 06 02 2004 A ASTRIUM pan express A reconfiguration PROM provides automatically all the possible configurations of the above resources A configuration is selected in accordance with an incremented pointer Also the ground has the possibility to force the pointer value to select a specific configuration AA ASTRIUM express o d Ref VEX T ASTR TCN 00349 TA Issue 2 Rev 0 Date 06 02 2004 Page 11 9 11 4 DMS SYSTEM S W IMPLEMENTED FDIR The DMS System Layer software executes the mission and system level functions System Re Initialisation Mission Time Line MTL Autonomous Sequences e Separation Sequence SSAP including Solar Array SA Deployment e RCS Priming e Venus Orbit Insertion VOI Main Engine Boost Mode MEBM e Safe RF Configuration SAFCAP
163. nus it 15 preferable to command the wheel Off Loading from the ground in Normal Mode GSEP the date being optimised taking into account the mission constraints The Off Loading function manages simultaneously all the wheels It includes several sequences of thruster pulses until angular momentum of each wheel 15 close to the target value This sequence 15 defined by a feed forward 3 axes wheel torque command combined with a thruster pulse The sequence ends with a tranquillisation phase controlled by the wheels in order to damp the dynamic excitation generated by the actuation of thrusters and wheels It must be noticed that it 15 also possible to control the wheels speed In a more classical way through a wheel speed command in some thruster modes Orbit Control Mode Thruster Transition Mode and Braking Mode Ref VEX T ASTR TCN 00349 EADS Issue 2 Rev 0 Date 06 02 2004 ASTRIUM T 5 Page 5 17 Sun Earth ephemeris v propagation Attitude Guidance on v v ephemeris GSEP Large slew V autonomous Guidance Fixed quaternion v v v Guidance FPIP Ground commanded v v v Guidance FPAP GSP Gyro stellar v v v v v v v estimator init Reaction Wheels v Off Loading Reaction Wheels v v v v v management Thrusters v v v v v v management For Wheels Off Loading only Figure 5 4 1 Use of generic functions in the AOCS Modes SADM Commands Computation Specific S A posit
164. o reduce the spacecraft angular rates before using the reaction wheels during the Normal Mode The slew manoeuvre necessary to come back to the nominal operations 1s performed in the Normal Mode after a reduction of the attitude transient due to the end of the TTM Normal Mode Wheel Damping Phase NM WDP Thruster Transition Mode TTM The Thruster Transition Mode ensures a smooth transition from the thruster controlled modes OCM BM and the Normal Mode designed with reaction wheels control The exit from the Thruster Transition Mode is achieved automatically the transition criterion combines an attitude depointing criterion a rate criterion and conditions on the reaction wheels The Thruster Transition Mode uses one Star Tracker 1 or 2 IMUs and thrusters for the torque generation The reaction wheels are commanded during the TTM either at their current speed or at a Y Ref VEX T ASTR TCN 00349 2 Issue 2 Rev 0 f Date 06 02 2004 TA ASTRIUM express Page 5 14 ground uplinked speed level It is therefore possible for the ground to take advantage of this thruster controlled mode for reaction wheel off loading not the baseline The mechanisms SADM have a constant orientation during the mode and are in Hold mode Aerobraking The Aerobraking uses the aerodynamic drag force to provide a deceleration effect during an atmospheric pass and modify the satellite orbit This concept is an alternati
165. oal in a first step to recover from failure in the same operational mode and then in a second step to preserve the spacecraft integrity in case of unforeseen unrecoverable anomalies This second step leads to a spacecraft Safe Mode allowing stable safe attitude and TM TC communications with ground for further diagnosis and recovery TC Figure 11 1 1 shows the hierarchy of the Venus Express fault management The faster the fault response must be the lower level the FDIR mechanism is implemented Also the number of units involved in a reconfiguration may increase with the level of fault management In case of conflicts in terms of fault treatment between two or more FDIR layers or structures the next higher level of FDIR or ultimately the ground will trigger Ground driven diagnostic TM Ground and fault recovery procedures TC CDMS Computing amp Communications Reconfiguration CDMS Reconfiguration Modules Automatic CDMS HM table driven reconfiguration actions DMS S W System FDIR MTL Auto Sequences Supervision Safe Mode Mgt Comm s Recovery Subsystems or application Subsystem Applications FDIR DMS TCS AOCS TTC Subsystem AP recovery programs switching group of units Unit recovery or Avionics platform units DMS AOCS Power Comm s CDMU RTU SSMM PDU PCU RFDU WIU TWTA
166. odule reconfiguration These data are down linked upon request from the ground Critical Data Storage During the execution of autonomous sequences e g Separation Sequence Main Engine Boost the DMS software stores critical data in particular failure reports to check later the execution of the sequence on ground These critical data are stored in the Critical Event Log CEL of the SGM RAM Safe Mode Entry Inhibit safing algorithms that trigger the spacecraft Safe Mode will be disabled by the ground at times of orbital manoeuvres to allow recovery guaranteeing that the burn occurs on time and that fault management does not interfere with the timely execution of the manoeuvre Safe Mode CMDS Reconfiguration Inhibit Rights With the exception of the system re initialisation sequence only the ground has the possibility to enable disable the mechanisms that trigger the CDMS reconfiguration and or the Safe Mode This is done by the Ground directly or through the ground uplinked MTL Short MTL The DMS software has the capability to execute a Short MTL previously stored by the ground in the PM RAM The Short MTL is a SSMM less MTL type of execution and therefore allows S C operations transparency wrt SSMM failure modes during mission critical phases This capability is used by during the VOI MEBM Standard Monitoring Enable Inhibit The DMS S W incorporates a standard monitoring service denoted as S131 to monitor
167. of a second High gain antenna This new HGA is a single X band HGA so called HGA2 The block diagram of the TT amp C subsystem is presented in the following figure p LGA 1 S Band Rx 1 front X Band Rx 1 i I I LGA2 rear X Band Tx 1 DIPL 2 and Ix I K HGA 1 S Band Rx 2 2 X Band Rx 2 S Band Tx 2 X Band Tx 2 Figure 11 6 1 Venus Express TT amp C Subsystem block diagram In violet color is represented the H W modification with regard to Mars Express implementation of 2 WIU Diplexer and its associated Wave Guides AA ASTRIUM express o d Ref VEX T ASTR TCN 00349 TA Issue 2 Rev 0 Date 06 02 2004 Page 11 14 Reminder of the Mars Express FDIR design The following mechanisms are implemented in the Mars Express design Safe RF Configuration Application Program SAFCAP Establishment of an autonomous S Band configuration most of the time featuring the HGA as part of the S C Safe Mode execution HW or SW safe mode TC Link Recovery Application Program TLRAP Detection of the TC link loss and subsequent autonomous recovery of the TC link taking into account the previous RF equipment configuration as an entry for establishing a new configuration The recovery can lead to versatile RF configurations combini
168. oncept is mainly derived from autonomy requirements failure tolerance requirements reuse of Mars Express avionics architecture and reliability figure This concept allows Venus Express spacecraft to be fully one failure tolerant 12 1 REDUNDANCY REQUIREMENTS Autonomy As Venus Express is required to be autonomously one failure tolerant decision taking function without ground intervention on failure event occurrence all failures which endanger the Spacecraft integrity need to be managed on board The FDIR function is in charge of failures management using redundancy resources Most functions are supported by stand by redundancies Hot redundancy and majority voting are used for critical functions main bus regulation reconfiguration module WD To improve availability waiting for powering on hot stand by redundancies can be programmed for critical mission phases Venus orbit insertion manoeuvre Failure tolerance The Venus Express spacecraft is designed to be one failure tolerant which means that each EEE function is redunded as a minimum Specific design rules segregation thermal dissipation control parts redundancy etc is implemented to avoid failure propagation Reuse of Mars Express Avionics architecture The redundancy architecture has been kept as is because of the numerous existing similarities between Mars Express and Venus Express in terms of mission autonomy and failure tolerance constraints Reliability figure
169. onomy management and the system FDIR management It also manages directly or through the AOCMS PM all the equipment and functions necessary to fulfil the spacecraft mission objectives The DMS application SW relies on the services provided by the Common SW Kernel SW Generic services OBCP manager It includes the following functions The SSMM management a DMS SW dedicated function acquires SSMM data stores them In the datapool and monttors them In order to elaborate the SSMM health status The Platform management which administrates commands and allocates all platform bus resources with the exception of the AOCMS Thermal Control System Power Control System RF Control System Pyrotechnics The platform bus 1s composed of non packetised end users So the DMS SW dedicated function which ensures the platform management has to decode TC packets and to route the orders to the platform bus end user as discrete commands Inversely unpacketised TM data from platform bus functions are acquired by the DMS SW on IEEE 1355 link Parameters are updated in the datapool monitored to elaborate platform bus health statuses and telemetry data are packetised for downlink to the ground The Payload management the payload instruments are packetised end users 1 e they are intelligent units providing useful information in telemetry packets and obeying to telecommand packets The DMS SW dedicated function which insures the payload managemen
170. ons The reconfiguration principle 15 unchanged wrt MEX At 1 attempt the X band receiver is changed At2 attempt the receiver is changed from X band to S band The TM chain configuration is unchanged 9 Ref VEX T ASTR TCN 00349 P Issue 2 Rev 0 EADS M uL Date 06 02 2004 GNUS expres Page 11 16 11 7 POWER FDIR SUMMARY The Venus Express Power FDIR 15 identical to the Mars Express one except the FDIR mechanism related to the APR failures The SW implemented Power FDIR includes the 2 following monitorings Battery Discharge Alarm Surveillance BDAS BCDR mechanism based on the Battery Anomaly Reactive Surveillance BARS and the Battery Discharge Current Surveillance BDCS The Battery Discharge Alarm Surveillance BDAS is a global surveillance which monitors any failure out of the power subsystem leading to a battery over discharge condition erroneous pointing over consumption conditions during eclipse erroneous mission programming by the ground Upon BDAS triggering a transition to S C Safe mode with an overall CDMU and avionics reconfiguration is activated leading to a significant power load shedding the Payloads are put in Keep live or OFF mode as well as the AOCS units which are not used in SAM and the RF emitters TX and TWTA are also switched OFF The telemetry transmission is re started in SHM when the Solar Array is properly sun point
171. ors the TFG the VIRTIS and VMC instruments It stores science and global housekeeping telemetry packets MA ASTRIUM express Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 Date 06 02 2004 Page 9 2 9 1 OVERVIEW The Data Handling architecture is organised around the two Control and Data Management Units CDMU They are in charge of controlling ground command reception and execution on board housekeeping and science data telemetry storage and formatting them for transmission The on board data management control laws processing and execution of on board control procedures belongs to their tasks as well Each CDMU features two MA3 1750 Processor Modules each of them processing either Data Management or AOCS software built in failure operational Reconfiguration Module ensures system level FDIR and reconfigures the CDMU as necessary Three other units are part of the DMS architecture The RTU connected to the redundant OBDH bus It is the interface between the DMS processor module and the platform units and the payloads The RTU 15 internally redunded and contains 6 modules 2 redundant core units in charge of processing the interrogations sent by the DMS PM on the OBDH bus 2 I O boards interfacing with the users 2 power supplies delivering the secondary voltages to the cores and I O modules The AOCS Interface Unit AIU 15 dedicated to AOCS equipment It is the interface between the AOCMS processo
172. press design in order to take maximum benefit of the recurrence and minimize development risks As a consequence Venus Express spacecraft is very similar to Mars Express Same system concept body mounted instruments fixed RF antennas and 2 solar arrays mounted on one degree of freedom mechanisms Same structure with only local changes Same propulsion subsystem Same avionics units Same operational concept Earth pointing in steady state in order to allow communication with Earth 8 hours per day and battery charging alternate with Venus observation during specific portions of the orbit However there are specific Venus Express mission features that have lead to design changes mainly regarding thermal control RF communication and power subsystem Science mission new payloads must be accommodated VIRTIS VMC VERA and MAG Two payloads that were design drivers to Mars Express have been removed BEAGLE and MARSIS Venus thermal environment since Venus is much closer to the Sun than Mars 0 72 AU instead of 1 5 AU the thermal flux is four times higher in Venus vicinity than in Mars vicinity 1 e twice higher than on Earth Venus radiation environment it is closely related to the distance to Sun thus quite more stringent for Venus Express than for Mars Express Planets configuration around Mars Earth vector is always within 40 of the Sun vector which helps keeping the cold face away f
173. proper fitting even in case the deployable boom development would not be conclusive Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 Date 06 02 2004 ASTRIUM exoress Page 25 2 3 3 PFS The PFS instrument a Planetary Fourier transform Spectrometer in a quite extended infrared range from 0 9 to 50 microns aims in the vertical optical sounding of the Venus atmosphere Its design relies on an optical module O a pointing unit S to orientate the measurement beam and two electronic boxes for power control distribution and instrument control The PFS opto mechanical design is fully re conducted from the MEX one with some minor but reversible modifications basically located within the optical interferometer The Mars Express PFS Interferometer The optical modifications are aimed for a better scientific exploitation of the NIR transmission window as recently discovered within the dense Venus atmosphere These are limited to A new laser diode for cinematic control operating at 0 9 um instead of 1 2 um e And to anew Short Wave SW detector PbS PbSe sandwich instead of PbS for covering an enhanced wavelength range down to 0 9 um now The module S pointing unit will be slightly adapted at the level of its coatings gold replaced by silvered taping to cope with the new Venus thermal conditions From another hand the module E has been slightly
174. quire the spacecraft to be reconfigured and the science mission to be temporarily suspended The watchdogs in the RM have different elapse times between 531 ms to 578 ms Upon loss of the AOCMS S W heartbeat the DMS S W reacts by non triggering of the RM watchdog 1 e processor halt The heartbeat surveillance consists in checking that the Hz AOCMS S W housekeeping packet 1s periodically received In case one of the Battery Voltage is detected low during several successive acquisitions of the associated PCU SDT data available in the DMS Data Pool the DMS S W causes the watchdog to trigger This in turn provokes a CDMS reconfiguration and a System Re Initialisation with large power shedding 1 e P L Off TX Off AOCMS units off till SAM commanding In case of one Battery BCDR channel failure a specific FDIR mechanism 15 triggered to reduce SC power consumption in order to be compatible with the remaining battery resources In case of triggering the S W makes a first power load shedding by putting the payloads and the SSMM in keep alive or OFF mode At the next eclipse entry a S W safe mode is commanded to expand the power load shedding additional units are switched OFF such as the RF emitters TX s and TWT s the AOCS units which are not used in SAM and non critical heaters The SC remains in such low power consumption configuration until the eclipse exit and until the earth acquisition is completed with the SA sun poi
175. r module and the sensors the actuators and the solar array drive electronics The AIU 15 internally redunded and contains 6 modules 2 interface modules interfacing with the AOCMS processor modules and the IMPs and controlling the generation of internal HPCs and the AIU internal bus 2 TMTC boards which generates the commands to the thrusters and main engine acquires the internal secondary voltages and external AOCMS units telemetry and implements the interface with the reaction wheels assembly The SSMM 1s a file organised 12 Gbits BOL mass memory used to store the Housekeeping and the Science Data collected by the CDMU It also collects directly Science Data from VIRTIS and VMC The SSMM contains three 4 Gbit memory modules two redundant controller paths controlling the interfaces with the CDMU processor modules the Transfer Frame Generators the payloads with high data rate VIRTIS VMC and the Memory Modules each of them being connected to a power supply from which it receives the necessary voltages i Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 E Date 06 02 2004 ASTRIUM U enus express wine os TRSP1 TRSP2 CDMU1 CDMU2 SSMM AOCS P Ls TT amp C TCS EPS Figure 9 1 1 Data Handling Block diagram Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 4 Date 06 02 2004 ASTRIUM p gnus 55 Page 9 4 9 2 CDMU The core of the Data Handling function 15 imp
176. ransponder 2 Transponder 2 S Band Rx S Band Tx WIU X band Tx Isolator 2 HGA X band 2 WIU X band WIU X band Tx 3 dB Hybrid HGA 1 amp HGA2 Diplexer WG Switch X Band Tx WIU X band Tx Isolator X Band Transponder 1 X Band Tx Transponder 2 X Band Tx Transponder 1 X Band Rx active WIU X band Rx WG Switch Transponder 2 X Band Rx Figure 12 3 1 Venus Express Communication reliability block diagram Ref VEX T ASTR TCN 00349 EADS Issue 2 Rev 0 Date 06 02 2004 ASTRIUM SNUS 5 Page 12 7 PDU CDMS CDMS FCL PSI AOCMS CDMS RMI CDMS PDU CDMS CDMS CDMS FCL PS2 PM 2 DMS RM3 TFGA PDU CDMS CDMS CDMS CDMS FCL PS3 PM 3 AOCS RM4 TFGB CDMS HPC A CDMS RTU b PDM Red RTU Moth Board RTU RTU RTU RTU Chain IN STB 1 STB b RTU RTU RTU RTU RTU RTU d amp connect Chain 1R RTU Moth Board SSMM amp connect Power Chain 1 68 Fits Converter RTU SSMM Moth Board Power amp connect Converter Chain 2 306 Fits complex AIU Al AIU AIU AIU SAS 1 SAS 2 i uw m a Elements Nom STB AIU AIU AIU U SAS 1 SAS 2 Elements Red PDU STR STR PDU IMU IMU g LCL EU Nom OH Nom LCL Gyros Nom Acceleros Nom STB PDU STR STR PDU IMU IMU LCL EU Red OH Red LCL Gyros Red Acceleros Red h IMU Gyros Acceleros Processing amp Supply IMU Gyros Acceleros Processing amp S
177. rd software location Ref VEX T ASTR TCN 00349 Issue 02 Rev 0 Date 06 02 2004 r express Page 10 4 10 2 SOFTWARE INTERFACES Interface with the host processor is through the low level layers of the software e g PM HW interface manager of the Kernel layer for the DMS SW and the AOCMS SW To communicate with the external world ground external hardware equipment other software the Venus Express on board software components use OBDH data bus and IEEE 1355 links External interfaces All the ground telecommands are received by the DMS SW through the Packet TC decoder The DMS SW sends housekeeping telemetry and low rate scientific data to the Transfer Frame Generator TFG for transmission to the ground Payload instruments have their own autonomous software located in the instruments electronic units The command and control of the payloads is performed by a dedicated function of the DMS SW which sends commands to the instruments and receives monitoring information and low rate scientific data from them Another dedicated function of the DMS SW sends commands to the Platform equipment and receives their monitoring information The AOCMS SW ensures the same function for the AOCMS equipment e g sensors and actuators to which it sends commands and from which it receives monitoring information The DMS SW interfaces with the Safeguard Memory SGM for the emission and recept
178. red in spacecraft frame and computed from on board ephemeris is tested enabling the gyro stellar estimator initialisation and the transition to the next Mode At the end of the SAM StAP the Spacecraft X axis 15 oriented towards the sun with an angular control on 2 axes and a constant rate on the third axis Spacecraft to sun direction Safe Hold mode SHM The Safe Hold Mode completes the attitude acquisition sequence pointing the HGA 1 or HGA 2 axis towards the Earth The main function of the SHM is to perform a 3 axes attitude acquisition with the Star tracker STR and the gyros from the 2 axes sun pointing of the SAM It relies for this operation on the autonomous attitude acquisition capabilities of the STR The sequence starts in fact at the end of the SAM mode by a switch ON of the STR and a star acquisition phase StAP phase inside the SAM The SHM itself starts by a control based on the STR measurements and cancels the residual rate on X axis EAIP Earth acquisition Init Phase It performs then an autonomous slew manceuvre computed onboard from the actual attitude estimation and the ephemeris data EAP Earth Acquisition Phase ended in the Hold Phase HP Up to this phase the attitude control torques are provided by the thrusters The wheels are then switched ON and spun up EPIP Earth Pointing Initialisation Phase and then used in the attitude control loop once they have reached their target rate gt Ref VE
179. rforms therefore a validity check when the transition 1s requested by the ground If the validity conditions are not fulfilled the transition 1s reyected and the AOCS stays In the current Mode or subphase The transition will have to be commanded again by the ground after analysis of the situation and the Software event On board ground sharing for the equipment configuration management Onboard management of the AOCS equipment Usually an AOCS Mode requires some H W resources which are absolutely mandatory for the achievement of the Mode objectives In this kind of situation the onboard Software manages autonomously the H W equipment of the Mode through a validity check at the Mode transition that the required resources for the next Mode are available through an automatic switch ON switch OFF of the required unit and an automatic change of the desired configuration number of units or H W functions H W modes This management principle is the one applied in most of the cases for all the AOCS units It is considered in this case that the considered H W 15 locked by the AOCS Mode This principle prevents from unexpected ground telecommands for instance if the ground attempts to switch off a unit that 1s locked the ON state an AOCS mode the ground TC will be rejected Flexibility let to the ground for some AOCS units For some AOCS units the ground has the capability to adapt the Hardware configuration to th
180. riggers after the previous step in SAM In this case it is assumed that the anomaly cannot be managed by the AOCS alone A HW Safe mode with a processor reboot is then performed by the DMS under AOCS request After the reboot the AOCS restarts in SAM with all the nominal units The level 5 corresponds to the triggering of a surveillance in SAM after the level 4 The AOCS reboot counter is incremented and a second HW safe mode processor reboot is commanded The SAM is restarted with all redundant units All surveillances are inhibited and the transition between SAM and SHM is inhibited this is the ultimate back up Mode These 5 steps enable to scan all the possible configurations of the H W starting of course at the level 1 and 2 by the simplest actions which should cover the most probable failures at AOCS level At level 2 or 3 the back up mode termination 1s the Earth Pointing phase EPP of the SHM wheel controlled phase in which the Spacecraft waits for ground orders But as some transitions to this mode may be forbidden because of a lack of redundancy after a second failure for instance the final mode may be the phase of the SHM that allows communications with Earth the SC being controlled by thrusters instead of wheels after a second wheel failure for instance the Star Acquisition Phase of the SAM ultimate back up mode after a second STR failure for instance In order to ensure the synchronisat
181. rmal control The Mars Express CPS individual thermal control principle is kept as much as possible and broaden to all the CPS pipes The main engine and the thrusters have their thermal coupling with the spacecraft tailored to meet their thermal requirement while preserving the spacecraft thermal behaviour They are provided with individual electrical heaters sized to maintain these external units within the acceptable temperature range accounting for wide change in radiative environment The tanks thermal control is sized to respect the various requirements related to the different phases of the mission cruise VOI in orbit Controlled heaters are bonded on the tanks structures For the particular case of pressurant tank a software control 15 necessary to cope with the VOI requirements gt External appendages Some units are externally accommodated on the spacecraft antennas or AOCS sensors for example It means that they have some space exposed areas which allows them to reject dissipation using their own structure or a dedicated radiator mounted on its side 1f necessary In order to reduce the interface fluxes with the rest of the spacecraft these units are generally thermally isolated from the support structure by stand offs or insulation washers and MLI blankets They are provided with an individual thermal control designed at unit level radiators heaters MLI Ref VEX T ASTR TCN 00349 Issue 2 Rev 0 Date
182. rolled by proper sizing of heat rejecting radiators and heating power implementation This allows maintaining the unit temperatures to acceptable levels The heat transfer from the units to the radiators is performed by conduction when unit base plate is attached to the radiator honeycomb panels and by radiation The units and the panels have a black finish to maximise heat transfer inside the thermal enclosures gt Platform units The platform units are directly accommodated on Y walls which are acting as radiators These walls are covered by OSR with an embossed Kapton MLI trimming to reduce solar and albedo entrances The units are conductively cooled dissipation is transferred from the baseplate or the foot bracket of the unit to the radiator via the supporting honeycomb panel The conductive couplings are improved by means of thermal interface fillers and aluminium doublers For the special case of the PCU heat pipes are needed to conduct the dissipation to a large radiator All the internal units and the sidewalls are black painted in order to homogenise the temperature of the cavities The wheels thermal interface conductively cooled at baseplate interface combined with the geometrical accommodation constraints prevents from mounting the wheels directly on a radiator panel The X reaction wheels thermal control principle 15 the same as MEX The two wheels are connected by means of a thermal strap to an extra radiator paddle radiator or
183. rom the Sun Since Venus 1s an inner planet there 1s no longer such convenient property Distances to Earth Venus maximum distance to Earth is smaller than Mars maximum distance to Earth 1 72 AU instead of 2 7 AU Venus gravity it is quite bigger than Mars gravity 0 81 Earth gravity instead of 0 11 One of the consequence is that more AV is needed for injection which leads to propellant mass increase Finally operational orbit duration is driven by the tank capability and happens to be Y Ref VEX T ASTR TCN 00349 2 Issue 2 Rev 0 D Date 06 02 2004 TA ASTRIUM express Page 12 Wawa much longer than for Mars Express 24h instead of approximately 7h In addition spacecraft velocity at pericentre 1s also much bigger about 9 km s instead of 4 km s It was possible to accommodate all new payloads with no major change w r t Mars Express structure The biggest challenge was to implement VIRTIS that has very stringent thermal requirements infra red detectors must be kept at very low temperature It was achieved by coupling VIRTIS to a dedicated radiator located on the cold face of the spacecraft as well as for PFS This cold face needs to be kept permanently away from the Sun Beside the accommodation of new payloads the main driver to the spacecraft design is obviously the thermal design Considering the planets configuration and the need to keep the cold face away from Sun it w
184. s and Vega missions These allowed for a first and basic description of the conditions prevailing in the atmosphere and the near environmental sphere of the planet or even at its surface from very time limited measurements obtained from landers or balloons The more intrusive and intensive radar imaging with the late Venera or Magellan orbiters combined with Galileo or Cassini fly by images have greatly expanding our knowledge in the geology and geophysics fields They have revealed hidden behind a curtain of dense clouds an exotic planetary world which 15 still full of mysteries But many of the questions raised on the processes sustaining these conditions remain unsolved e Is it possible to explain by in situ measurements of the various plasma species from energetic neutrals to ions and electrons the drastic evolution of the terrestrial atmosphere of Venus into this wild world of carbon dioxide and sulphuric acid micron size droplets How to explain the global atmospheric circulation as seen from the typical ultra violet and infrared markers from various poorly known constituents In which the deep atmospheric layer shows a zonal and retrograde super rotation 20 times the planet one with velocities decreasing from up to 120 m s at the cloud tops down to almost 0 near the surface e In fully exploiting the recently discovered near infrared windows in the high and middle atmospheres can we close the gap between the low
185. s contained in SC LV separation plane and oriented toward the High Gain Antenna side of the spacecraft The Zs axis is coincident with the launcher X1 axis It represents the SC line of sight toward Venus during science operation The Ys axis is contained in SC LV separation plane and oriented so as to complete the right handed co ordinate system It is therefore parallel to the solar array plane In launch configuration the Ys Venus Express spacecraft axis is located in the Y 1 Soyuz Z1 Soyuz quadrant at 45 deg from Y 1 Soyuz axis The Ob Xb Yb Zb Reference Frame 15 structure related and is no more used at S C or operations level Ref VEX T ASTR TCN 00349 2 Rev 0 Issue EADS Date 06 02 2004 ASTRIUM enus express Page 1 9 01 02 03 04 05 06 07 08 09 10 11 12 XI Zs xf SECTION C C Xlva LAUNCH VEHICULE ADPATOR LVA FREGAT IN SATELL TE INTERFACE PLANE PLAN lt lt o 750 E INTERFAC PLANE 0 Xs Ys ZS VEX SPACECRAFT YL ZL LAUNCHER ENCE AXIS v Ylva Zlva LVA ENCE AXIS Zf FREGAT CE AXIS PT SECTION B B
186. sary in the following mode SAM are switched OFF The AIU and the IMUs are therefore kept ON except during launch phase where the IMUs are OFF Sun Acquisition Mode SAM The Sun Acquisition Mode performs a first attitude acquisition to orient the X and the Solar Array cells towards the sun It 15 used for the initial attitude acquisition after launch for nominal attitude re acquisition after Solar Arrays deployment for the nominal re acquisition after the Main Engine Boost Mode and also after a failure during Software or Hardware safe mode It uses gyros IMUs and Sun Acquisition Sensors SAS for the attitude measurements and thrusters for the control This mode starts by a reduction of the Spacecraft rates RRP Rate Reduction Phase The sun is then acquired close to the X Z half plane corresponding to the global Field Of View of the 2 SASs SCP Sun Capture Phase A third phase orients the Spacecraft in such a way that the X axis points to the sun SAP Sun Acquisition Phase If the solar array is not yet deployed the Mode continues with a Sun Pointing Phase SPP and a biased Pointing phase BPP If the Solar array is deployed a preparation to the next mode is performed in the Star acquisition phase StAP During this phase the Autonomous Attitude Acquisition is commanded to the star tracker outside the control loop Once a 3 axis attitude is provided by the star tracker its consistency with the Sun direction measu
187. sence to LCL except that they do not feature ON OFF switching capability and that overcurrent will never lead to disconnection when the trip off time is exceeded The current limiting function is of the foldback type meaning that the unit voltage will decrease as the current demand increases above the limitation threshold Six FCL are baselined in the Venus Express PDU which are allocated to nominal and redundant Dual Band Transponder Receivers and CDMU 2 DC DC Converters are implemented in each CDMU Seven LCL classes are defined on Venus Express in order to cope with a wide range of nominal currents while ensuring an efficient protection A total of 78 LCLs and 6 FCLs are implemented within the PDU VEX PDU output power capability is limited to 750W compared to Mex PDU limitation 650W Ref VEX T ASTR TCN 00349 EADS Issue 2 Rev 0 p Date 06 02 2004 ASTRIUM express Page 7 13 The following figure shows the PDU block diagram PCU Main Bus Power LCL 78 S C users RTU N i HPC Reset MLC ON OFF HPC 1 FCLs 6 MLC SDT CDMU N 1 HPC Priority S C users PYRO SA 2 Propulsion MAG Boom MLC FIRE MLC ARM MLC SELECT Figure 7 2 4 VEX PDU block diagram T Ref VEX T ASTR TCN 00349 AD Issue 2 Rev 0 Date 06 02 2004 AA ASTRIUM express Page 2714 E 7 5 PYRO DEVICES The Venus Express Pyro function is included in the PDU and 15 fully re
188. switches RFDU WIU to different antennas Two Low Gain Antennas LGA allowing an omni directional reception and hemispherical emission in S Band One dual band High Gain Antenna 1 allowing high rate TM emission and TC reception in S Band and X Band One single band offset Antenna HGA2 allowing high rate TM emission and TC reception in X Band The Dual Band Transponder performs the demodulation of the up link signal before routing the resulting bit flow to the Data Handling CDMU The stored TM within the SSMM 15 routed through the CDMU then modulated in either S Band or X Band within the Dual Band Transponder which also performs S Band signal amplification with 5 W output power X Band signal amplification is performed using a 65 W Travelling Wave Tube Amplifier TWTA LGA1 front S Bd Feeder TRSPD 1 DIPLI Py S Band Rx X Band Rx 1 TC to CDMU Demod S Band Tx 1 poz bang sx lt TM to CDMU P10 X Band Tx 1 P04 il Frequency Reference USO Internal TCXO P04 P04 VERA USO P05 P04 TRSPD 2 Frequency Reference USO Internal TCXO S Bd Feeder LGA2 rear 1 5 Bd Feeder Attenuator 1 X Bd Feeder o Hybrid P03 P02 Wo 1 lt 112 PO P02 Pol Ke lt TWTA 2 Attenuator 2 P04 P0
189. t elaborates and sends TC packets to the payloads It receives TM packets coming from the payloads updates datapool parameters according to packets content Payloads monitoring is performed using Service 12 or by specific OBCP to be developed by PIs The AOCMS management provides the DMS SW with the knowledge of the AOCMS current configuration and state The remote or Service Mode PM management provides the DMS SW with the knowledge of the remote PM state The Mission management which sequences the nominal and contingency mission phases and interfaces with the ground control for units observability and commandability the FDIR management Based on a hierarchical approach the FDIR is handled at two levels by the DMS SW at DMS sub system level with the monitoring of equipment health statuses and the management of local reconfigurations and at system level with the monitoring of the current functions to be fulfilled and the management of functional modes reconfigurations VEX T ASTR TCN 00349 Issue 02 Rev 0 Date 06 02 2004 AA ASTRIUM express Page 10 13 10 4 4 AOCMS Application Software The AOCMS application SW manages the AOCMS modes according to a mission dependent modes transition logic It also manages the different AOCMS sensors and actuators used In AOCMS modes and mission phases The AOCMS specific functions rely on the services provided by the Common SW Kernel SW Generic services and inclu
190. tanks For this reason testing of the main engine sub assembly is conducted prior to installation of the oxidant tank to the spacecraft structure The testing verifies that the engine pipe connections are leak tight The four thruster modules two thrusters per module are positioned below the lower floor at the corners Their feedlines pass locally through the floor to emerge within the envelope of the thruster mounting brackets The CPS units are mounted on the Xs face of the Xs shearwall and on the Zs side of the lower floor in the Xs Ys and Xs Ys segments This arrangement facilitates the use of a jig for manufacture The layouts of the fuel and oxidant supply assemblies are similar to one another as are the layouts of the low pressure gas assemblies feeding them The high pressure gas side completes the shearwall layout The unit layout has been modified with respect to Mars Express in the aim to accommodate the CONAX pyrovalves and the POLYFLEX non return valves The 15 fill drain vent valves are mounted at the lower floor accessible from underneath the spacecraft They are grouped in five clusters of three valves located along the Xs edge of the structure Each cluster 15 directly associated with one of the five assemblies 1 e liquid sides for fuel and oxidant low pressure gas sides feeding fuel and oxidant and high pressure gas side The positioning of the valves within each cluster follows the same pattern with regard to v
191. tary missions Q The first one is the need for a high autonomy due to the absence of real time control of the spacecraft and at mission critical phases such as Venus orbit insertion and eclipses Q Secondly the spacecraft shall be able to cope with a highly variable environment Sun to Venus distance impact on Solar flux Earth to Venus distance changing attitude etc and with optimised resources to cope with the launch mass restriction The following diagram gives an overview of the VEX electrical architecture Payload Heater Power Heater Power Heater Power Heater Power 1355 link 28V Regulated Thermal Power Protected lines Pyro Devices Communications ZR Power Supply Solid State Mass Memory RemoteTerminal USO I F Unit Power Distribution Unit Bis AOCS Interface ty Unit x Data Handling Solar Array Drive Assembly MACS Analogue Analogue Propulsion Analogue Reaction Star Inertial Sun Wheel Tracker Measurement Acquisition Assembly Unit Sensor Attitude and Orbital Control Figure 7 1 VEX Electrical Architecture ge Ref VEX T ASTR TCN 00349 A D Issue 2 0 4 Date 06 02 2004 AA ASTRIUM expres DAE 7 2 ELECTRICAL POWER Two Solar Arrays wings equipped with triple junction Gas cells generate electrical power The Solar Arrays are oriented towards the Sun by a Solar Array Drive Mechanism SADM During eclipses three Li Ion bat
192. teries supply the required power Power management and regulation is performed by the Power Control Unit PCU providing a controlled 28Volts main bus voltage The use of a Maximum Power Point Tracker MPPT avoids to oversize the solar array in order to cope with both near Earth and Venus orbit conditions It allows working at the Solar Array maximum power point while three Battery Charge and Discharge Regulators BCDR are in charge of the battery management controlled by an Error Amplifier Control Loop MEA The resulting 28 V regulated power bus is distributed to all spacecraft users by a Power Distribution Unit PDU featuring Latch Current limiters LCL The PDU 1s also responsible for Pyro commands generation whereby the necessary energy is drawn from the batteries Ref VEX T ASTR TCN 00349 EADS Issue 2 Rev 0 ASTRIUM p Date 06 02 2004 GNUS express Page 7 3 The following figure shows the VEX Power Subsystem From From To RTU Power Distribution Unit PDU Power Control Unit PCU TM TC CDMU Transponder Receivers Array Power 28V Main Bus Regulator 1 p Spacecraft Bus i t Array Power dix Payload Y Solar Regulator 2 Instruments Thermal Control Heaters FEE m Pyro Y Propulsion Interface Payload MAG Li Ion Li Ion Li Ion Battery 1 Battery 2 Battery 3 Figure7 2 VEX Power Subsystem diagram R
193. the possible use of back up MELSP with various final attitudes It 1s the most robust one Orbit Control Mode OCM The Orbit Control Mode enables to perform the small trajectory corrections using the 10 N thrusters The uses 1 Star tracker or 2 IMPs and thrusters The reaction wheels are commanded during the OCM either at their current speed or at a ground uplinked speed level It is therefore possible for the ground to take advantage of this mode using thrusters for reaction wheel off loading not the baseline The mechanisms SADM have a constant orientation during the mode This mode uses the thrusters located on Z face of the Spacecraft to generate the thrust and also to control the attitude through an OFF modulation command It starts by a Liquid Settling Phase LSP during which a lower mean acceleration 1s generated to reduce the liquid transient in the tanks It 1s followed by the burn itself Burn Firing Phase BFP The 15 preceded by an orientation of the spacecraft slew manoeuvre performed during the Normal Mode NM GSP before entering the orbit control mode The end of the manoeuvre is decided onboard on the basis of the AV estimation derived from accelerometers measurements or thruster ON time counting The first method is recommended for large delta V manoeuvres and requires previously an accelerometer in flight calibration All manoeuvres are followed by a tranquillisation phase in TTM allowing t
194. tions and health Timing functions including distribution of time and synchronisation information The DMS 15 based on a dual processor architecture embedding standard communication links such as a standard OBDH bus and high rate IEEE 1355 serial data links The OBDH bus 15 data route for data acquisition of platform units and payloads with a low data rate and for commands distribution via the RTU IEEE 1355 links are used between the CDMU processor and SSMM the CDMU processor and the AIU and between the payloads with high data rate VIRTIS VMC and the SSMM DMS includes 4 identical Processor Modules PM located in 2 CDMUs Two processor modules are dedicated to the DMS and two to the AOCMS The PM selected for the DMS function 15 in charge of the Platform subsystems management Communications Power Thermal The PM selected as the computer is in charge of acquisition and control of all sensors SAS STR IMP actuators wheels main engine thrusters and Solar Array Drive Electronics SADE through the AOCMS Interface Unit AIU Each CDMU includes a TC decoder for the telecommands decoding and the direct high power commands generation and a Transfer Frame Generator TFG for the telemetry generation incorporating TM packets from the CDMU itself and from the SSMM The Solid State Mass Memory SSMM is used for data storage including 12 Gbits of memory at BOL It is coupled to the two DMS process
195. upply Figure 12 3 2 Venus Express DMS amp AOCMS reliability block diagram Ref VEX T ASTR TCN 00349 EADS Issue 2 Rev 0 Date 06 02 2004 ASTRIUM SNUS 5 Page 12 8 PVNCI HE tank Pressure Transducer Fill amp Drain Flow valve EIE M Pyro Valve PVNC2 Pyro Valve PVNC3 a PVNCS active active Pressure Transducer active 1 active Pressure Regulator Pressure Regulator a a b Pyro Valve PVNC6 Filter Fl a Fill amp Vent Valve Latch Control Valve Pyro Valve FDV11 b Fill amp Vent Valve Latch Control Valve Pyro Valve FVV8 leakage LV2 PVNC9 Pyro Valve PVNC8 Pressure Transducer 2 tank Fill amp Drain Flow valve PT3 leakage FDV10 F2 PT4 leakage Pyro Valve Pyro Valve PVNC11 Pyro Valve PVNC12 Main Engine Main Engine Pyro Valve PVNO3 Flow Orifice Filter Flow Control Valve Main Engine Main Engine Main Engine Main Engine Pyro Valve PVNC13 Pyro Valve PVNC14 Main Engine Main Engine Latch Control Valve Flow Control Valve Thrust 1A Oxydiser Thrust 1A Oxydiser Pyro Valve PVNO2 Flow Orifice Filter Flow Control Valve Main Engine Main Engine Main Engine Main Engine Latch Control Valve Flow Control Valve Thrust 1B Oxydiser Thrust 1B Oxydiser Latch Control Valve Flow Control Valve Latch Control Valve Flow Control Valve Latch Control Valve Flow
196. urement mode provides three axis attitude restitution to AOCMS SW without initial information from the AOCMS SW it performs an initial mapping an automatic pattern recognition selects stars and automatically tracks them Mapping mode provides magnitude and co ordinates of all targets present in the field of view for ground investigation backup mode Calibration mode gives the AOCMS SW the possibility to change default parameters The STR SW is a slave to the AOCMS application SW which can always through equipment TC switch off or reset it change the current mode modify parameters dump or patch data and code In the same way the AOCMS SW asks for TM data and is free of reading them interrupting a dialogue asking for other TM data without corrupting the nominal execution of the mission modes TM TC is managed as a high priority task by the STR SW Ref VEX T ASTR TCN 00349 E ADS Issue 02 Rev 0 Date 06 02 2004 ASTRIUM Linn eXQreSS mu 1 1016 10 7 IMP SOFTWARE The IMP SW also runs at unit level It concerns the gyros only and provides the AOCMS SW with spacecraft angular rate As the STR SW the gyro SW is a slave to the AOCMS application SW which performs the configuration commanding and surveillance of this equipment Serial data is output from the IMU to the AIU via the RS 422 signal interface in blocks of ten words at 200 Hz Each word consists of 16 bits where the first bit bit 0 is th
197. uring Earth pointing phase As a consequence minor changes of the power subsystem have been deemed necessary in order to cope with more stringent requirements In particular increase of the Battery Discharge Regulator capability from 250W to 300W was implemented Finally one of the major design change regards the Solar array It was proven that use of Silicon cells as for Mars Express is not suitable for Venus Express due to the fact that Venus thermal environment leads to a very wide temperature range for solar cells thus to a wide voltage range that would not be compatible with Power Control Unit GaAs cells will be used instead since they are much less sensitive to temperature as well as to radiation environment Each solar array wing 15 twice smaller than for Mars Express 2 panels per wing instead of 4 due to the fact that thermal flux is higher and that GaAs cells are more efficient As a conclusion it was possible to cope with Venus Express mission requirements while keeping the changes w r t Mars Express to the minimum J Ref VEX T ASTR TCN 00349 2 2 Issue 2 Rev 0 p Date 06 02 2004 AS ASTRIUM enus express Page 1 3 1 2 SPACECRAFT MAIN CHARACTERISTICS Main characteristics of the baseline design are presented hereafter Mechanical design Mars Express structure concept has been fully reused with only local changes The core structure is a honeycomb parallelipedic box sizing about 1 7 m length
198. us v Maximum Power Point Tracking 900 watts peak power for Venus operations 3 Li ion batteries 24 2x42 State Power Controller for power distribution and protection Data handling T Mars Express rebuilt w ESA class 1 OBDH architecture O four 3 1750 processors for AOCS amp OMS D reconfiguration modules with majority voting capability O transfer frame generation up to 228 kbps 12 Gbits Solid State Mass Memory v IEEE 1355 links for high data rates transfers On board software Mars Express update v P55 05 0 standard ADA based Separate AOCS and DMS functions Modular design v Top level object oriented design Figure 1 4 1 Spacecraft Architecture w 2x32 Solid State Power Controller for pyralines Two Command amp Data Management Units providing Ref VEX T ASTR TCN 00349 EADS ASTRIUM express Issue 2 Rev 0 Date 06 02 2004 Page 1 7 Thermal Control Heater Power Attitude and Orbit Control Mex Rebuilt New equipment EN Mex modified Ros modified MAT13257 Figure 1 4 2 Spacecraft Electrical Architecture Ref VEX T ASTR TCN 00349 V m S Issue 2 Rev 0 p Date 06 02 2004 enUS express Page 1 8 1 5 COORDINATE SYSTEM The origin of the spacecraft Reference Frame named Os is located at the separation plane between the spacecraft and the adapter at the centre of the interface diameter of 937 mm The Xs axis i
199. ve solution to propulsive methods for achieving the transition from the high elliptical insertion orbit into the observation orbit If an Aerobraking phase 15 necessary to reach the final mission orbit or to change the orbit the necessary operations will include on each orbit a phase with an Earth pointing near the apocentre and a phase with a specific pointing adapted to aerodynamic pressure near the pericentre during the drag pass For this latter function the Braking Mode BM 15 used to control the Spacecraft around an attitude which is stable with respect to aerodynamic disturbances During the Aerobraking phase the ground 1s in charge of the overall sequence of operations on each orbit described through the Mission Time Line MTL The Navigation is also a ground responsibility in order to ensure the efficiency of the Aerobraking and define appropriate apocentre manoeuvres in when necessary to ensure that the S C remains in the domain for which it has been designed less than 0 3 N m Braking Mode BM The Braking Mode uses gyros only 1 or 2 IMUs for the attitude estimation owing to the small duration of the drag pass and 10N thrusters for the control Just before the atmospheric pass which lasts about 400 seconds the satellite is aligned with the aerodynamic frame in Normal Mode the manoeuvre 15 performed m NM GSP and satellite maintained in FPAP until the transition to the Braking Mode is commanded T
200. ware Set Break Point Set Memory or Register Examine Break Point Examine Memory or Register Start Stop This PROM monitor connected to a simple console or to a TLD debugger is useful for PM prototype validation AIV thermal tests campaigns and launch campaign if EGSE link 15 not available and PM boards maintenance All the anomalies detected by the PM Firmware are logged into a synthetic table in RAM health status Detection of anomalies does not prevent from starting the loaded software except if all the software loading attempts failed The active software DMS SW or AOCMS SW will be in charge of checking the health status and conduct appropriate PM FDIR if necessary EADS ASTRIUM rem express Ref VEX T ASTR TCN 00349 Issue 02 Rev 0 Date 06 02 2004 Page 10 8 BOOT UP INIT PM INIT SW START LOADER SSMM EGSE SVF TEST CONSOLE TESTER SERVICES PM MEMORY TM TC IEEE 1355 HW interfaces OBDH DELAYS Figure 10 3 1 PM Firmware architecture T 9 d Ref C S Issue T ASTRIUM UL sii enus express pass 10 4 DMS AND AOCS SOFTWARE VEX T ASTR TCN 00349 lt 02 Rev 0 06 02 2004 10 9 Both the DMS and AOCS software are composed of the so called Common SW common to DMS and AOCS and an applicative part called the DMS application SW and the AOCMS application SW 10 4 1 DMS and AOCMS SW layered breakdown Both the DMS
201. wer FDIR is summarised in Section 11 7 The AOCMS FDIR is summarised in Section 11 8 Typical related errors of the DMS and AOCMS S W are RTC error overflow overrun overload The DMS S W implements specific mechanisms to protect itself and the SSMM against overload conditions that would create S W crashes Overload conditions would typically occur upon ground programming exceeding the max operational limits such as the number of TC with the same SCET in the MTL q 24 TC sec or the number of P L TM polling during the 8 second polling period etc Also it might happen in case the DMS is busy with a RTU reconfiguration while TC continue to accumulate in the DMS S W buffers Finally jitters of the VHF sequencer 64 Hz HF sequencer 8 Hz and NF sequencer 1 Hz could result in an 1 Hz cycle overrun whenever cyclic task time allocations are computed irrespective of the worst case jitters The DMS S W protects itself against burst of TC originating from the ground uplinked MTL as follows 1 if the number of MTL TC with the same SCET is larger than 24 then the DMS S W executes the first 24 TC in the current 1 Hz cycle and spreads the remaining TC over the next Hz cycles Should the burst of TC last for several seconds the DMS S W continue to behave the same way 1 The MTL execution FDIR will trigger if and only if a TC is 20 second overdue The DMS S W queues up ground TC during RTU reconfiguration suspends the MTL execution pendin
202. xpress 2 3 1 The ASPERA experiment for Analyzer of Plasma and EneRgetic Atoms consists of two plasma measurement untts located on two different walls of the Spacecraft structure e The Main Unit MU provides electron and neutral sensors namely ELS NPI and NPD together with scanning one axis mechanism and central experiment electronics this unit is placed on the Y wall of S C the scanner rotation axis being parallel to the Y direction e The Ion Measurement Assembly IMA complements the plasma measurements through a non rotating design large FOV the unit is placed on the S C bottom floor ASPERA Main Unit MU and Ion Measurement Assembly IMA from Mars Express Compared with Mars Express no major design change 15 to be noticed Anyway to cope with the Venus thermal environment to cope with the direct and significant planet flux when the nadir instruments are operated a new location for the MU has been defined on the Y wall but finally not impacting the Mars Express defined mechanical environment e and some thermal adaptation has been worked out for the sensors e g enlarged radiating surfaces use of new MLI and coatings In addition and specific to Venus Express more ASPERA analyses are to be concluded on e High energy radiations effects on EEE parts both IMA and MU being not significantly protected by S C structure as are the optical instruments some limited

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