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s HMM Assessment Study Report: CDF-20(A) February 2004 page 186 of 422 (see [RD19],[RD20] & [RD21]) data are considered for the design. The electrical mounting, the cells, their coverglasses and the substrate are the only part taken into account in the solar panels mass budgets for the power subsystem. The other components of the panels (deployment mechanisms, motor, yoke,..) are integrated in the mechanical subsystem. For the multi-junction cells, coverglasses of 150 µm are assumed (see Figure 3-43). For the thin film cells, a value of 600 g/m 2 is assumed (see [RD22]) Option 1: 140 µm (kg/m 2 Option 2: 100 µm ) (kg/m 2 ) Triple junction cells 0.720 0.514 150 µm coverglass 0.397 0.397 Coverglass adhesive 0.067 0.067 Interconnects 0.013 0.013 Cell adhesive 0.213 0.213 Bus/wire/diodes 0.307 0.307 Substrate 50 µm kapton 1.144 1.144 Total 2.878 2.878 Figure 3-43: Mass budget for multi-junction cells options 15 panels of 1 metre make up one polar wing. Changing the cells and the substrate will require a new qualification. The length of the panels will be computed for this study. To cope with a mechanical failure during the folding/unfolding manoeuvres, the solar wings are sized taking into account one panel loss. The solar panels are mounted with a driving mechanism. Except during eclipses and manoeuvres, the solar panels are Sun pointed. The sizing mode for the solar panels is the “Orbiting on Mars” when the eclipse is the longest (41 minutes) and the solar irradiance at its minimum. In that case, the solar arrays are designed for being able to fully recharge the battery before the start of the next eclipse. A failure of a complete solar panel is also taken into account for this computation. 3.3.4.3.2 Power storage The power storage has to supply power during the eclipses on LEO and in Mars orbit. But the sizing cases are the manoeuvre phases in which the solar panels may be stowed up to 6 hours. Alternatively, there are other interesting storage technologies: • Secondary batteries. The most efficient are the Li-Ion cells with round trip efficiency around 94%. The specific mass nowadays is around 100 Wh/kg. Improvement to 150 Wh/kg for 2015 is expected (cf [RD3],[RD4] and [RD5]). • Regenerative fuel cells. The most advance ones are the PEM providing electricity and water by combining Oxygen and Hydrogen. On the other hand, the charge is performed by electrolysed water. The efficiency is estimated to be about 50 to 60%. The regenerative fuel cells may be interesting for high energy requirements such as in this mission, but their poor efficiency involves a huge increase of the solar generator size. Since this part is already a critical point, the use of secondary batteries is preferred.
s 3.3.4.3.3 Power conditioning and distributing HMM Assessment Study Report: CDF-20(A) February 2004 page 187 of 422 To limit the harness losses a higher voltage should be chosen, typically 120V. Nevertheless, this voltage cannot be too high otherwise there is a risk of electrocution of the crewmembers and also to avoid plasma interactions on the solar wings. Due to the large amount of power units, a regulated topology shall be selected. 90% is the efficiency assumed for the Battery Charger Regulator. 85% is the efficiency assumed for the Battery Discharge Regulator. For safety reasons, the Power Conditioning and Distributing has to be double-failure tolerant: • First failure: the spacecraft shall remain fully operational • Second failure: the spacecraft shall still be operational with possible limitations on the mission. For safety, an architecture with separated and identical power systems with crosstraps to the different units connected on the bus is proposed. Each power system is composed of a dedicate solar panel, a battery module, a power conditioning and distributing unit. Compared to the power required, one power system is added in the design to cope with possible failures. 3.3.4.4 Baseline design 3.3.4.4.1 Budgets Solar wings Solar wings 2 power module 2 power module Figure 3-44: Power system overall architecture There are two separate power modules in each of the two nodes, as shown in Figure 3-44. Each power module includes:
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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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Page 141 and 142: 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 193 and 194: s Configuration Length (m) Total ma
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Page 213 and 214: s Angular separation SEM > 10º SEM
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s Item Nr. Mass [kg] Margin [%] Mas
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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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