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s HMM Assessment Study Report: CDF-20(A) February 2004 page 34 of 422 Figure 2-12: Geometry During Mars Phase Figure 2-13 shows the Earth-Sun-spacecraft geometry for the transfer to Earth. Here, there are no superior conjunctions. Figure 2-13: Geometry During Mars-Earth Transfer 2.4.3.2 Earth escape and planetary protection The TV will be assembled in LEO. To minimise gravity losses, allow a staged solution and enhance operational safety, the TMI burn is split into three parts. The first two insert into eccentric orbits, the third and final one into the hyperbolic escape. After each burn, the spent stages are jettisoned and perform a controlled reentry. Otherwise they would reenter at some undetermined time and could endanger inhabited regions. The required ∆v for de-orbiting is budgeted. The final, escape burn must be biased such that the spent third stage does not hit Mars. After separation, the TV is retargeted on the course to Mars. This adds to the TV manoeuvre budget. 2.4.3.3 MOI and planetary protection The assumption is that MOI inserts into a 4-sol orbit (approx 500x96 000 km). The MOI stage is jettisoned and remains in this orbit, the TV continues to a circular 500 km orbit. It was shown that the orbit of the MOI stage can be expected to decay within 10 years. A lifetime of over 50 years can be achieved by raising the pericentre altitude to 3000 km. This adds to the MOI stage budget. 2.4.3.4 The Mars orbit phase After MOI, the orbit is circularized. The Mars orbit phase is spent in a 500 km circular orbit. This orbit will remain stable for over 50 years. The operations taking place during this phase are discussed in separate chapters.
s 2.4.3.5 TEI and planetary protection HMM Assessment Study Report: CDF-20(A) February 2004 page 35 of 422 If TEI is broken into phases to allow for a staged approach, then it must be ensured that the jettisoned first stage does not impact the surface of Mars within 50 years. The escape manoeuvre must be biased such that the escape stage does not hit the Earth. The THM must then be redirected to a trajectory closer to the Earth. This adds to the THM budget. 2.4.3.6 Final retargeting and Earth arrival Designing this phase implies striking a balance between maximizing protection of the Earth and maximizing crew safety. The strategy assumed for the return trajectory is shown in Figure 2-14. The TEI dispersions are corrected with the retargeting manoeuvre seven days after Mars escape. The trajectory still is fail-safe and Earth-avoiding. Any failure at this time will lead the THM past the Earth. Sixty days prior to the nominal arrival, at a distance of 18 million km, if the spacecraft is still functional, a retargeting is performed. Optionally, at this time the longitude and latitude of the touchdown point can be changed. One day prior to arrival, the crew moves to an Earth entry capsule and separates from the THM. Figure 2-14: Earth Return Trajectory Strategy There are two options for the final phase: • At the retargeting 60 days before arrival, the THM can be targeted for an entry orbit. Then, after the entry capsule is separated one day before entry, the THM must be moved to a trajectory with a slightly higher perigee so that is passes by the Earth • Alternatively, the retargeting can send the THM on a trajectory that passes close to the Earth without entering the atmosphere. Then, after the entry capsule is separated, it must navigate itself to an entry trajectory 2.4.4 Budgets 2.4.4.1 Earth escape window • Escape window: 08/04/2033 – 28/04/2033 • Hyperbolic escape velocity: 3.022 – 3.200 km/s
Page 1 and 2: s HMM Assessment Study Report: CDF-
Page 3 and 4: s STUDY TEAM HMM Assessment Study R
Page 5 and 6: s T A B L E O F C O N T E N T S HMM
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
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Page 47 and 48: s 2.7.5.1 Mission elements dry mass
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Page 51 and 52: s Total mission time (days) 1200 10
Page 57 and 58: 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
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Page 77 and 78: s Days on Martian surface 450 400 3
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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 HMM Assessment Study Report: CDF-
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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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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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