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s Z (m) x 10 -1 4 0 1 2 3 4 Impulsive Hohmann Transfer Applied Delta V Station Keeping Forced Motion Station Keeping Forced Motion Free Drift Trajectory in CRF (height vs. X) -16 -14 -12 -10 -8 -6 -4 -2 0 2 x 10 4 5 X (m) Figure 4-110: Intermediate rendezvous HMM Assessment Study Report: CDF-20(A) February 2004 page 374 of 422 4.5.2.4.2 Terminal rendezvous Next, the crew commands a forced approach to the next hold point located at 100 meters in front of the target and also in V-bar. This new hold point will be the base of a station keeping manoeuvre that will allow the crew to check again the status of the subsystem and the availability of the docking port, etc. Finally, a new forced approach is commanded and the MAV reaches the docking port of the Orbiter that establishes the structural latching of all required elements. 4.5.2.4.3 Transfer of crew and dedocking Finally the structural latching occurs. The crew is transferred from the MAV to the orbiter. After that the MAV is ejected. Then there is a check out of all orbiter subsystems. Once this is done, the departure to Earth can begin. 4.5.2.5 Baseline design The MAV final stage is a three-axis stabilised spacecraft. The attitude determination main sensors are the star tracker (STR) and the inertial measurement unit (IMU). During Sun acquisition and safe modes, Sun acquisitions sensor and Earth infrared sensors are principally used. The two STRs that will be operated in parallel for some phases of the mission.
s HMM Assessment Study Report: CDF-20(A) February 2004 page 375 of 422 The attitude control actuators are a set of reaction wheels assembly (RWA), 4 clusters of 5 hydrazine thrusters plus 4 clusters of 2 hydrazine thrusters of 10 N thrust each, and two ASTRIUM engines acting with 550 N force each. The RWA includes 4 Reaction Wheels (RWL) implemented on a skewed configuration. This configuration enables to perform most of the nominal operations of the mission with a 3 RWL configuration among 4. The Reaction wheels provide the GNC control torques during all the phases of the mission except the trajectory corrections, the attitude acquisition and back up modes. Each Reaction Wheel provides an angular momentum storage capacity of 12 Nms and a motor torque of ± 75 mNm in the speed range from + 6000 rpm to - 6000 rpm. Under nominal conditions, the RWA uses a 3-wheel configuration. The fourth wheel is used during critical mission phases to avoid any mission outage if one wheel fails and minimises the reconfiguration duration. During nominal modes hydrazine thrusters are used to perform wheel off- loading, while during safe mode they become the attitude control actuators. The 10 N thrusters for attitude control are all located in the second stage of the orbiter. A detailed analysis of the CoG and the controllability of the orbiter has not yet been done. However, the present configuration based on the ATV looks feasible in this first iteration. The LIDAR expected features are the following: • Validity range of the measurements shall be between 1 metre and 10000 metres • Field of view (FoV) shall be of no less than 10 degrees • The weight shall be of no more than 10 kg. • The average power consumption shall be of no more than 20 Watts. Even with the specifications above, the power consumption and weight values are slightly increased, according to the latest news about the LIDAR development process. The CAMERA features to be tested shall be as follows: • The validity range of the measurements shall be between 0 metres and 5000 metres to the target • The field of view (FoV) shall be of no less than 20 degrees • The weight shall be of no more than 2 kg. • The average power consumption shall be of no more than 3 Watts. In the same way as with the LIDAR, slightly more conservative values have been taken. Table 4-53 summarizes some of the characteristics of the GNC equipment.
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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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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 3.3.3.1 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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Page 419 and 420: s 8 APPENDIX C - ACRONYMS AAA Avion
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