The H2 arcjet propulsion system is the least efficient among nuclear electric systems. Due to the level of Isp that it delivers this system is also more massive than a chemical one. Because the AV a required for lunar round trip mission is so high, the propellant mass for the arcjet system is very large. A propulsion system with an Isp of 1500 s would not reduce the propellant mass sufficiently over the H2/O2 system. The most efficient systems with respect to the propellant and dry mass are the Ion and MPD systems. They can save up to 50% of the total mass over the chemical propulsion system. Moreover it is important to stress that the ratio [payload I dry mass and propellant] is above 1. These masses strongly depend on the values of Isp and the specific mass of the reactor. The specific mass (kg/kW) reflects the performance of the reactor. The lower the specific mass, the higher the performance. The assumptions made in Figure 8.21 for the arcjet, the MPD and Ion system (1 MW) are: P= 1MW Isp= 5000 s Specific Mass = 10 kg/kW The duration of the transfer will be relatively long (200 to 400 days) because the level of thrust capability with MPD and Ion propulsion systems is low (10 to 100 N). Thus, these missions will be unmanned. The conventional chemical propulsion will be used for the transportation of the crews. These missions could be possible around the years 2020 and beyond. Solar electric OTV requires big solar arrays. Generated electricity can be used to ionize Xe propellant. A 300 KW solar electric power system is considered. Here Figure 8.22 shows the decreased mass resulting from this system over the chemical one. This overall gain is around 50%. The Isp is of the same order of magnitude than for nuclear electric propulsion. Meanwhile the low thrust in comparison with a 1 MW power system will raise the travel duration dramatically (770 days). Figure 8.22 gives an idea of what could be achieved within the next ten to fifteen years. Figure 8.22 An Example of 300 KW Solar Electric OTV 8.7.3 Nuclear Thermal Propulsion Nuclear thermal propulsion is quite similar to chemical propulsion and it will be used for mission profiles from LEO to LLO and back. A nuclear reactor heats an H2 propellant that provides the thrust through the nozzle. Achievable Isp is about 900 s and the reduction in LEO mass is about 20%
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