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s HMM Assessment Study Report: CDF-20(A) February 2004 page 404 of 422 Furthermore, the total mass of the systems is one of the key parameters for space reactors, while virtually non-existent for designing classical terrestrial systems (except submarine reactors). To a smaller extent this also holds true for its shape and size. As a result of the above mentioned parameters, space reactor designs show higher reactor core temperatures and operate with much higher enriched 235 U fuel. This leads to lower core dimensions and mass, reduced radiator size and mass and higher efficiencies of the conversion system. In numbers, the enrichment increases from natural-20% for terrestrial up to 93% for space systems, and the core temperature increases from 400-573 K up to 900-1000 K. Technically these changes imply the abandon of the pressurised water reactor design for either liquid metal cooled or gas cooled cores. They furthermore demand the use of materials capable of withstanding for long times high temperatures as well as high radiation (neutron) fluxes. 6.2.3 Approach Given the high complexity of designing space nuclear reactors and especially the highly interrelated subsystem dependencies, it was chosen to not develop a parametrical reactor model with completely open input parameters but to base the assessment on two recently proposed space reactor models. For the present study, the two designs that are taken as reference were recently proposed by European nuclear industry within an <strong>ESA</strong> contract especially for Martian surface power generation purposes. They are both available to a technical design level sufficient for the present study and correspond to the power level as well as lifetime requirements. 6.3 Reactor power subsystems Some aspects of different reactor power system components are provided. The purpose of this small section is to provide some basic elements to understand the choices and implications of the reactor designs proposed. Reactor core • Thermal n core: larger (if small: limited core life due to burn-up), moderator, good negative power coefficient (safety), rather for larger than10 MWe reactors • Fast n core: enriched fuel, smaller, more stable power distr., low burn-up, small negative power coefficient (safety), favourable in 100 kWe range Conversion system Static Thermoelectric: mature, space proven; about 4% (adv.: cascades about7%) Thermionic: mature, in core vs. out core systems, life-time limiting factors, space reactor proven AMTEC: immature, different for ionic and electronic condition, short lifetimes, corrosion problems, 19-25%
s HMM Assessment Study Report: CDF-20(A) February 2004 page 405 of 422 MHD: immature, ionic gas flow in B field, very high T, 30% ThermoPV: immature, high emitter T (larger than 2200K) + low PV temp, 30% Dynamic Brayton cycles: He or He/Xe working gas, no corrosion problems, 1 or 2 cycles, 1500K – 550K, 20% eff. Rankine cycles: state change (evap-condes.), higher eff.; alkali metals (Po, N, toluene) Stirling cycles: kinematic or free piston, 1100K-650K, 23.5% eff., sealing integrity Shielding n shielding: metal hydrides (LiH), usually contained in stainless steel (for structure and protection) γ shielding: Tungsten (alternatives: e.g. depleted Uranium, borated steel, Pb- W-LiH) Shielding is for the considered type of reactor in case of close human operations certainly the most massive subsystem. Requirements dependent on mission design (transfer phase operation? partial (angular) shielding?) Power management Voltage issues (to decrease losses) Cabling to overcome distance to reactor – trade-off between the mass of cabling versus mass of additional shielding in case of closer distances; Chosen European reactor designs The two designs retained by European industry are the liquid metal-cooled thermal core with a conservative thermoelectric conversion unit and thermal radiators as cold well and the slightly more advanced gas-cooled fast neutron particle bed reactor connected to a dynamic Brayton conversion unit using forced convection of Martian atmosphere for waste heat rejection. 1. Liquid metal cooled thermoelectric conversion reactor (LMR) Thermal neutron core ZrH2 moderator Moderator/fuel ratio: 3 NaK coolant (22/78) Thermoelectric conversion + radiator 2. Gas-cooled Brayton-cycle particle bed reactor (GCR) Fast neutron core Particle bed design (1mm diam. fuel particles, 93% enriched U235) He or He/Xe coolant (1 st cycle) Use of Martian CO2 for cooling (2nd cycle) The most fundamental parameters of both designs are listed and compared in Table 6-1.
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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 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 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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Page 419 and 420: s 8 APPENDIX C - ACRONYMS AAA Avion
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