that there is no space charge limitation as there is with ion thrusters, which means that a single MPD thruster may be able to produce orders-of-magnitude more thrust (high thrust per unit area) than any ion thruster. Of course, as more thrust is generated, the power requirement of the thruster will increase proportionally, so that cooling and associated mass penalties may be incurred. However, MPD thrusters are operable at very high temperatures, are capable of high specific impulse, and are structurally small and intrinsically simple devices that should effect a significant cost advantage. Hence, they are attractive for application to SPS. It should be noted that several thousand electric thrusters may be required to transfer the SPS from LEO to its GEO station; therefore, in the interest of minimizing mass, and consequently cost, the propulsion system that requires the least total mass and cost would be preferred. A preliminary analysis indicates that the transfer propulsion cost is minimum when using an MPD propulsion system to transfer a photovoltaic satellite from LEO to GEO. Although selecting the optimum thruster size is not germane to this study, the Boeing MPD is selected as a preliminary baseline. Figure 7-23 presents the design concept of the 100 cm ion thruster. Figure 7-24 shows the MPD thruster design concept proposed by The Boeing Company. This thruster has an anode exit diameter of 10 cm. The outside diameter of the 40 000 A turn magnet, which can be operated in series with the cathode, is approximately 30 cm. When operating at an input power of 587 kW, this thruster does not require cooling equipment. Boeing proposed this thruster concept to serve only as a manufacturing vehicle for estimating cost. There is no technology precedent for this thruster. Figure 7-25 presents the MPD thruster design concept proposed by the J PL for comparison. Propellants such as hydrogen, helium, ammonia, lithium, nitrogen, sodium, potassium, argon, or cesium may be used for MPD electric propulsion with some efficiency variation between each. However, environmental considerations dictate the use of argon as the propellant. Argon also has the property of good density but will have to be tanked under cryogenic conditions. The cryogenic properties of argon are similar to those of oxygen; therefore, the technology for long term storage as well as resupply of cryogenic argon will have to be established. Several concepts for mounting the orbit transfer electric propulsion system on the SPS are under consideration. One concept involves the use of end mounted thruster modules as shown in Figure 7-26. Each module is expected to require 360° of gimbal motion per Earth orbit revolution, such that continuous tangential thrusting can be maintained. A preliminary structural analysis indicates, however, that the SPS bending loads allowed by this concept are
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