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s Number of stacks 18 17 16 15 14 13 12 Number of stacks required 3 months commissioning time 1 month between launches 11 10 3 months commissioning time 3 months between launches 0 500 1000 Boil-off rate (kg/month) 1500 2000 Figure 2-26: Number of stacks required as a function of launcher rate and boil-off rate HMM Assessment Study Report: CDF-20(A) February 2004 page 58 of 422 The analysis led to a boil-off requirement rate of 500 kg/month. Figure 3-31 shows that the maximum number of stacks in the worst case is 13. Cryogenic propulsion systems could be extended to other mission phases, i.e. MOI, if we take into account that the time from departure to the MOI is around 6 months and the assembly time is already more than 54 months, and also that the interplanetary environment is more benign than the LEO environment from the thermal point of view. However, for the design case, the conservative case of only cryogenic TMI has been taken. 2.7.6.4 Return approach Several strategies can be adopted for the return to Earth. The crew will return from Mars in the THM and ingress in the ERC to perform the atmospheric entry descent and landing. You could insert the THM into Earth orbit so it can be reused in a next mission. If not, it has to be discarded so that it does not impact the Earth. Besides, the ERC with the crew inside can also be inserted into an Earth orbit allowing the crew transfer to an orbital space station (e.g. ISS) before landing, or it can perform a direct entry. Table 3-16 shows a preliminary analysis of these possibilities: THM and THM discarded, THM discarded, ERC inserted ERC inserted ERC direct entry Mass to LEO (tonne) 4892 1813 1336 Table 2-15: Mass to LEO for the different strategies on Earth return As shown in Table 3-16, parking the THM and/ or ERC in Earth orbit is too expensive in terms of mass to LEO, as the propellant to perform the manoeuvre has to be taken from Earth to Mars and back (other techniques as aerocapture/aerobraking have not been analysed for this manoeuvre). Planetary protection issues imply, no contaminated vehicle should remain in Earth orbit, and this is the case of the THM, as its exterior will be contaminated by the Mars Ascent Vehicle (MAV). In the case of the ERC, which will be also contaminated, the problem is overcome by the sterilization arising from high temperatures during reentry.
s HMM Assessment Study Report: CDF-20(A) February 2004 page 59 of 422 Therefore, direct entry of the ERC is selected as it presents a lower complexity and scientifically reduces the mass to LEO. 2.7.6.5 Orbit insertion and acquisition around Mars Several technologies or strategies can be used for the orbit insertion around Mars. Preliminary analysis have been performed for these cases: • Propulsive manoeuvre, using storable bipropellant technology. • Aerobraking, no configuration changes assumed, orbit insertion performed by means of a propulsive manoeuvre, storable technology, 120 m/s added for orbit corrections • Aerocapture, no configuration changes assumed, no heat shield included, 120 m/s added for orbit corrections Propulsive Aerobraking Aerocapture Orbit insertion duration (days)
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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
Page 7 and 8: s HMM Assessment Study Report: CDF-
Page 9 and 10: s L I S T O F F I G U R E S HMM Ass
Page 15 and 16: s L I S T O F T A B L E S HMM Asses
Page 19 and 20: s 1 INTRODUCTION 1.1 Background HMM
Page 21 and 22: s 2 GENERAL ARCHITECTURE 2.1 Study
Page 25 and 26: s 2.2.4.2 Accelerations HMM Assessm
Page 27 and 28: s 2.2.4.4 Temperature and relative
Page 33 and 34: s Figure 2-10: Trajectory Overview
Page 35 and 36: s 2.4.3.5 TEI and planetary protect
Page 45 and 46: s Mars Excursion Vehicle Transfer H
Page 47 and 48: s 2.7.5.1 Mission elements dry mass
Page 49 and 50: s 2.7.5.8 MEV release The MEV is re
Page 51 and 52: s Total mission time (days) 1200 10
Page 57: s stack 9 Propellant mass of the ne
Page 65 and 66: s The core supporting structure has
Page 67 and 68: s 2.7.7.1.3 Mars excursion module S
Page 69 and 70: s 2.7.7.1.4 Earth return capsule HM
Page 71 and 72: s Mission Phase Description Event s
Page 73 and 74: s Mission Phase Description Event s
Page 75 and 76: s 2.7.10 Mission performance Table
Page 77 and 78: s Days on Martian surface 450 400 3
Page 83 and 84: s Maximum manoeuvre duration 6 mont
Page 85 and 86: s % of loss from baseline 20 18 16
Page 87 and 88: s 2.7.15 Sensitivity analysis HMM A
Page 89 and 90: s 2.7.15.4 Influence of the mass of
Page 93 and 94: s Parameters used: • No Shuttle
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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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s Figure 3-24: Baseline design back
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s Figure 3-26: Baseline design priv
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s Module part 3............= 0 m 3
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s 3.3.2.1 Requirements and design d
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s flux [W/m2] 250 200 150 100 50 0
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s Figure 3-38: Power required to ma
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s 3.3.4.3 Assumptions and trade-off
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s 3.3.4.3.3 Power conditioning and
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s 3.3.5 AOCS HMM Assessment Study R
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s Configuration Length (m) Total ma
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s 3.3.5.5 ACS for TMI 1 st HMM Asse
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s 3.3.6.2 Baseline design HMM Asses
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s 5 Reduce 2dB pointing precision t
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s 70-m antenna 70-m antenna 3.3.7.5
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s Angular separation SEM > 10º SEM
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s Figure 3-63: Communications MEV/M
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s Element 1: Transfer Habitation Mo
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s Radiation Shielding Shielding Req
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s The propulsion system and its cha
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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 4.3.1.5.3 Top level Figure 4-20:
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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 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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