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s HMM Assessment Study Report: CDF-20(A) February 2004 page 78 of 422 be allowed. In this case northern latitudes are preferred as the dust storms coincide with the winter. 2.7.12 Abort Options In the event of a failure or emergency that precludes completion as planned, the mission will have to be aborted. Two categories of failures have been considered: • Failures in equipment/subsystems • Failures in manoeuvres/operations The first category is dealt with at subsystem design level by making the design fail safe. For the second category, mission-abort scenarios have been considered. In the event of a mission abort, the original mission objectives shift to a safe return of the crew, as soon as possible and within the constraints dictated by the system design. Any defined abort scenario must be consistent with all budgets imposed for the various subsystems, e.g., propellant, thermal, structural, power or ECLSS consumables. Firstly, the abort options for each phase of the mission require study. Then the cost (in terms of the budget of each relevant subsystem) must be quantified for each identified option. The outcome is to know what can conceivably be done to save the crew at each point of the mission. The options for a mission abort strongly depend on the selected scenario. Here, the study is limited to the baseline scenario as studied for the 2033 launch opportunity. First, the options per mission phase were analysed. Then, the abort cost for each option was quantified and listed. 2.7.12.1 Phase-by-phase analysis The mission phases were studied chronologically, starting from Earth escape and ending with the return trajectory from Mars to Earth. 2.7.12.1.1 Abort during Earth escape The Earth escape sequence is split into three manoeuvres, each of which further raises the apogee until a hyperbolic orbit with the required orbital parameters is achieved. Any of these burns can be prematurely terminated. For over 90% of the total manoeuvre duration, abort leaves the spacecraft still in a wide elliptical orbit around the Earth. The crew will have to wait until the next perigee to be able to then board the entry capsule and return to Earth’s surface. 2.7.12.1.2 Fast abort during Mars transfer The fast option requires a large DSM to insert the spacecraft into a trajectory that intersects the Earth’s orbit. The nominal mission parameters are shown for comparison with the abort transfer characteristics in Table 2-27: • Earth escape date: 8 April 2033 • Nominal Mars arrival date: 24 October 2033 • Nominal Earth arrival date: October - November 2035 • Nominal mission duration (maximum): 962 days • Nominal Earth arrival velocity (maximum): 3.052 km/s
s HMM Assessment Study Report: CDF-20(A) February 2004 page 79 of 422 Table 2-27: Fast abort transfer characteristics The fast-abort transfer characteristics are shown in Table 2-27. The abort manoeuvre cost, despite rising for late abort, is always within the mission budget, as all propellant reserved for MOI, orbit circularization and TEI can be used for this purpose. The remaining ∆V capability after TMI is 5.2 km/s, once the MEV is discarded. The duration is counted starting from Earth escape and increase from 1 year to 1.5 years if the abort decision is delayed. In terms of life support systems, there would be no problem as the system is designed for a nominal lifetime that is longer than the duration of the abort mission in any case. The time spent in deep-space can exceed the nominal duration, but in no case should the radiation limits be exceeded. The shaded area indicates the potential problem with a late fast abort, the Earth arrival velocity is significantly larger than that in the nominal mission. This situation occurs 75 days after departure. One option here is to use the remaining ∆V capabilities to reduce the arrival velocity at Earth. If not, the ERC should be redesigned to cope with entry velocities as high as 17 km/s. This will require further study. 2.7.12.1.3 Slow abort during Mars transfer The slow option retargets for a modified Mars swing-by. After that, another deep-space manoeuvre is required for retargeting to Earth arrival. Abort decision date 1 May 2033 1 August 2033 1 October 2033 Abort decision date ∆V [km/s] Earth arrival date Earth arrival velocity [km/s] DSM 1 22 October 2033 23 October 2033 24 October 2033 ∆V -1 [km/s] Mars swingby DSM 2 ∆V-2 [km/s] Total ∆V [km/s] Earth arrival date Arrival hyperbolic arrival velocity [km/s] 0.899 2.110 3.010 5.280 0.895 25 21 2.111 3.005 28 5.284 0.894 October 2033 January 2035 2.110 3.004 September 2035 Table 2-28: Slow abort transfer dependencies 5.287 Duration [d] 1 April 2033 2.335 8 April 2034 1.9 365 1 May 2033 2.447 13 April 2034 2.8 370 1 June 2033 2.695 27 April 2034 4.6 384 1 July 2033 2.910 13 May 2034 6.4 400 1 August 2033 3.117 29 May 2034 8.1 416 1 September 2033 3.323 19 June 2034 9.5 431 1 October 2033 3.523 27 June 2034 10.9 445 Duration [d] 904
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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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Page 51 and 52: s Total mission time (days) 1200 10
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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 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 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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