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s HMM Assessment Study Report: CDF-20(A) February 2004 page 204 of 422 3.3.6.1.8.1 CFDP design concepts It is expected that a pair of CFDP entities, which have files to exchange, may at any given moment be unable to communicate; for example, a spacecraft orbiting Mars may be on the far side of the planet, unable to transmit to Earth. Therefore, if transmission of a file from Earth to a Mars-orbiting spacecraft is interrupted when the spacecraft passes behind the planet, the CFDP entities at both ends of the transmission can react in two ways: 1. Storing their outbound protocol data units (PDUs) – possibly in non-volatile memory, to assure continued service even in the event of an unplanned system reset – until the spacecraft reemerges and transmission can resume. This approach does not affect CFDP transactions, which can continue to produce PDUs regardless of the link unavailability. 2. Freezing, at CFDP level, all the file transfer transaction related to the temporarily unavailable link and resuming it at the next communication opportunity. Both events, (freeze and resume), are triggered by an external module that holds knowledge of the link state. A collateral benefit of this approach is that it largely insulates user applications from the state of the communication system: an instrument can record an observation in a file and “transmit” it (that is, submit it to CFDP for transmission) to Earth immediately without considering whether or not physical transmission is currently possible. This deferred transmission mechanism, sequestering outbound data management and transmission-planning functions within CFDP, can enormously simplify flight and ground software and thereby reduce mission costs. The large signal propagation delays that characterize interplanetary transmission limit the usefulness and efficiency of the retransmission strategies commonly used in terrestrial protocols (especially sliding-window ones). For this reason, CFDP’s retransmission model is one of concurrent transmission: data PDUs for multiple files are transmitted as rapidly as possible, one after another, without waiting for acknowledgment, and requests for retransmission are handled asynchronously as they are received. As a result, portions of multiple files may be in transit concurrently. CFDP is scalable, designed for reuse without modification in any number of communication environments as well. No specific direct interface to radio hardware, or even to any specific linklayer protocol, is mandated in the Recommendation. Instead, a minimal abstract underlying “unit data transfer” or UT layer service is assumed to be available for CFDP’s use, enabling CFDP to be run on top of a wide range of services as UDP/IP on the Internet and the Packet services of CCSDS Telemetry [RD39] and Telecommand [RD40] in space. 3.3.6.1.8.2 CFDP outlook CFDP is a stable international standard that can reduce mission operations cost and risk by enabling reliable file transfer and remote file system management over interplanetary distances. The application of emerging delay-tolerant networking technology to Interplanetary Internet operations, and specifically to the use of CFDP in complex mission configurations, will further enhance CFDP’s usefulness and value to deep-space exploration missions.
s 3.3.6.2 Baseline design HMM Assessment Study Report: CDF-20(A) February 2004 page 205 of 422 3.3.6.2.1 Requirements and design drivers The THM integrated avioncis has been evaluated taking into account all the function normally requested of a command and control module in a complex spacecraft. Moreover, the data rate has been dimensioned assuming the need for almost continuous audio and video link with the ground station and the need of an internet-style connection. 3.3.6.2.2 Baseline design The DCDMS (Distributed Control and Data Management Systems) is implemented as several distributed units called Control and Data Management Units (CDMU). Every unit exploits the following modular functions: • processors modules (PM) each including the digital interfaces required to communicate with the platform units. Any number of processor can be powered at a given time, depending from the workload (a similar mechanism is used on multiprocessing computer farms and is totally managed by the application software) • telemetry transfer frame generators (TFG), directly interfaced with trasponders. • reconfiguration modules (RM), two of which are always powered, each containing a clock function and a reconfiguration function. A configuration mechanism allows choosing one of the modules as “master clock”. The other ones act as spares or as backup processors. • distributed memory modules (DMM), each containing VRAM and NVRAM modules, the latter to be used as safeguard memory. A specific application checks the synchronization of the content of safeguard memories. The overall amount of available mass memory shall be of the order of several terabytes, even only for astronaut’s personal use. 3.3.6.2.3 Budgets The DHS mass and power budget has been evaluated as shown in the following Table 3-39: Property Value DHS+harness mass, inhabited modules 650 Kg Percentage of DHS+harness mass, unmanned modules 5 % Power, per module, manned 2200 W Power, per module, unmanned 400 W 3.3.7 Communications 3.3.7.1 Requirements and design drivers Table 3-39: DHS mass and power budgets • The vehicle should support Tracking, Telemetry and Command (TT&C) communications during all mission phases and any attitude. • Two-way ranging and Doppler capabilities are required during all mission phases.
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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 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 450.00 400.00 350.00 300.00 250.0
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s Surface Surface Container MAV Con
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s Bio-Lock Masses: Core Samples No.
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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 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 4.5.5 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 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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