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s 3.3.5.6 Electrical architecture Figure 3-52 shows the following electrical architecture which applies to all stages: EMAs PDU - ECU EMAs TVC8 EMAs TVC… TVC1 PDU - ECU Battery PDU - ECU Battery Battery IMU CSS “ Stage” ECU ACS Figure 3-52: Electrical architecture RCS CMG IMU OBC STR HMM Assessment Study Report: CDF-20(A) February 2004 page 196 of 422 The main intelligence shall reside in the habitation module, where the principal attitude sensors (Inertial Measurement Unit, IMU and Star Tracker, STR) shall be mounted as well. This main intelligence shall control directly the actuators mounted on the habitation module: i.e. Thrusters (RCS) and Control Moment Gyros (CMG). OBC will control the other actuators via several remote units, the "stage" ECU (Electronic Control Units). During the propelled phase the OBC shall control (via the ECU) also the Thrust Vector Control system (TVC). The inner control of the actuators of each TVC shall be realized through an additional box (PDU-ECU) mounted on the individual motor. The stage ECUs shall be in charge of the control of the stage when not mounted on the main vehicle (before assembly or after separation, for de-orbiting), and shall work as routers when connected with the main OBC. To accomplish the secondary function of controlling the stand alone stages, the ECU shall make use of local sensors: IMU and Coarse Sun Sensor (CSS). The IMU can be used (if needed) also during the control of the main vehicle to control the structural flexibility, which has not been considered in this study, but which might be a design issue in the future. 3.3.6 Data handling 3.3.6.1 General consideration This section summarizes the requirements applicable to avionic systems that will fly on a long lasting human mission with target launch date 2033. Several assumptions shall be made to generate a consequent design. From the DHSs point of view a human mission to Mars is extremely challenging considering presently available space technologies. Major issues that require attention are: 3.3.6.1.1 Computing power
s HMM Assessment Study Report: CDF-20(A) February 2004 page 197 of 422 The present and near-future European space computers are based on the SPARC RISC architecture (ERC32 then LEON2-FT). More powerful rad-hard Power-PCs are currently available, but under US ITAR control. The roadmap for the first generation of LEON processor is now stabilized : flight models in 0.18 µm ATMEL technology should be available during the first quarter of 2005, and should fulfill the needs of the first Aurora missions. Nevertheless, note that that the US processor available today already provide more processing power than the LEON will do in 2005: the first generation of radiation-hardened PowerPC 750 from BAE-Systems in 0.25 µm technology provides 240 to 300 MIPS, and a 370 MIPS version in 0.18 µm is foreseen before 2007. To lower the gap with the US products, ways to improve of the internal LEON architecture should be studied. The power PC 750 architecture is far more complex than the LEON architecture: it includes several independent processing units that allow executing more than one instruction per cycle. Moreover it also supports level 2 cache. This kind of architecture is needed to support higher-level operating systems that can guarantee soft real-time performances like Linux. The use of COTS processors (mainly Power PC line of products) shall be assessed for noncritical payloads. The availability of commercial SOI process (see [RD28]), may boost safe use of high power chips in space, but dedicated development (especially for ASICs) is needed. 3.3.6.1.2 Maintenance, availability The availability of space qualified electronic components has decreased in the latest years. Moreover, the evolution of rad-hard technologies after 0.13 Pm is very difficult to predict: it is not yet known whether or not the hardening techniques at design level will compensate the increasing SEE sensitivity (see Figure 3-53). Consequently, the gap between commercially available technologies and radiation-hardened technologies may either decrease or increase. Figure 3-53: Comparison of present roadmaps for space qualified and commercial microprocessors
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s HMM Assessment Study Report: CDF-
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s STUDY TEAM HMM Assessment Study R
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s T A B L E O F C O N T E N T S HMM
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s L I S T O F F I G U R E S HMM Ass
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s L I S T O F T A B L E S HMM Asses
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s 1 INTRODUCTION 1.1 Background HMM
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s 2 GENERAL ARCHITECTURE 2.1 Study
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s 2.2.4.2 Accelerations HMM Assessm
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s 2.2.4.4 Temperature and relative
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s Figure 2-10: Trajectory Overview
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s 2.4.3.5 TEI and planetary protect
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s Mars Excursion Vehicle Transfer H
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s 2.7.5.1 Mission elements dry mass
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s 2.7.5.8 MEV release The MEV is re
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s Total mission time (days) 1200 10
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s stack 9 Propellant mass of the ne
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s The core supporting structure has
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s 2.7.7.1.3 Mars excursion module S
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s 2.7.7.1.4 Earth return capsule HM
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s Mission Phase Description Event s
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s Mission Phase Description Event s
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s 2.7.10 Mission performance Table
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s Days on Martian surface 450 400 3
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s Maximum manoeuvre duration 6 mont
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s % of loss from baseline 20 18 16
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s 2.7.15 Sensitivity analysis HMM A
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s 2.7.15.4 Influence of the mass of
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s Parameters used: • No Shuttle
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s 2.8.3.3 Launch 3- Front node Figu
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s 2.10.1.4 Communications HMM Asses
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s 2.10.2.2.2 LEO assembly HMM Asses
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s Basic RF link Four 70-m Ka-band s
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s 3 TRANSFER VEHICLE 3.1 Systems…
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s Operational Requirements It shall
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s Trans Mars Injection Module Mars
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s Figure 3-2: Parallel configuratio
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s Figure 3-6: Global dimensions com
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s Front Node 3.2.3.1.3 EVA systems
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s Trans Mars Injection (three stage
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s Figure 3-16: Recommendations for
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s Factors to be considered Impacts
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Page 149 and 150: s Figure 3-24: Baseline design back
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Page 153 and 154: s Module part 3............= 0 m 3
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Page 213 and 214: s Angular separation SEM > 10º SEM
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s 3.4.4 Mechanisms HMM Assessment S
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s 4 MARS EXCURSION VEHICLE 4.1 Syst
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s Figure 4-2: Option 2: Vertical co
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s 4.3 Surface habitation module Fig
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s Figure 4-11: Section through the
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s 4.3.1.5 Baseline design HMM Asses
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s 4.3.1.5.3 Top level Figure 4-20:
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s 4.3.2.3 Baseline design HMM Asses
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s Schedule in hours for the most ac
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s 4.3.3.6 Waste production HMM Asse
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s Figure 4-26: Total pressure vs. O
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s ECLS system mass: Figure 4-27: Ma
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s Figure 4-28: Mars albedo (L), ars
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s sun flux (W/m2) 450 400 350 300 2
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s Insulation structure (external) F
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s cryocooler heat lift [W] cryocool
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s • a duty cycle value (duration
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s Figure 4-48: MAV Power Inputs HMM
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s 4.3.5.4.2 Power storage HMM Asses
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s Figure 4-52: Regenerative Fuel Ce
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s Figure 4-54: Solar irradiance on
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s 450.00 400.00 350.00 300.00 250.0
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s 4.3.6.1 Budgets HMM Assessment St
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s Surface Surface Container MAV Con
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s 4.3.9.3 Baseline design 4.3.9.3.1
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s Bio-Lock Masses: Core Samples No.
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s Figure 4-72: Entry Velocity (L) a
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s 2 3.9 233 923.1 347 923 12.3 352
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s Figure 4-85: Communications MEV/M
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s 4.4.5.3.2 GNC equipment HMM Asses
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s 4.4.5.4 Control laws generation H
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s Finally, the Figure 4-97 shown th
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s velocity (m/sec) altitude (km) 50
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s 4.5 Mars Ascent Vehicle 4.5.1 Tra
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s Term Value Unit Radius of equator
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s 4.5.2.1 Requirements and design d
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s Inclination (deg) 47.5 47 46.5 46
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s 4.5.3 Structures HMM Assessment S
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s 4.5.5 Thermal HMM Assessment Stud
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s 4.5.5.3 Baseline thermal design H
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s Crew Ingress/Egress Hatch: HMM As
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s Element 3: Mars Ascent Vehicle HM
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s 4.5.7.5 Budgets Characteristic Va
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s PER DAY PER MISSION DRINKING WATE
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s 5 OVERALL CONCLUSIONS HMM Assessm
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s 6 APPENDIX A - MARTIAN SURFACE NU
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s Figure 6-1: Example of buried rea
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s 7 APPENDIX B - REFERENCES [RD1] C
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s [RD26] IES2, Phase A0: Final Repo
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s [RD85] Mars Transportation Enviro
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s 8 APPENDIX C - ACRONYMS AAA Avion
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s MLI Multi-Layer Insulation MMH Mo
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