(b) Analyzed magnetic lines Figure 1. Steady state water-cooled MPD thruster.

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1 A. MPD thruster In this study, as one of the In-Space Propulsion projects by JAXA (Japan Aerospace exploration Agency), a practical MPD propulsion system was investigated. We planned to develop MPD thrusters as the main engines which have high thrust, high specific impulse, and high thrust efficiency kw steady-state MPD thrusters have been investigated at Osaka Institute of Technology (OIT). In general, a MPD thruster was equipped solenoid coils to apply magnetic field. However the thrust system of the thruster was complicated because the thruster needed devices to warm cooling water to cool coils in space. In order to make the system of the thruster easier, a MPD thruster with permanent magnets was developed at OIT. Figure 1 shows the water-cooled MPD thruster we developed. The purpose of this study is to develop a MPD thruster system with simple structure for manned Mars exploration 1). (a) 3D model (b) Analyzed magnetic lines Figure 1. Steady state water-cooled MPD thruster. Operations with a rod cathode made of pure tungsten were carried out and a thrust was measured. Figure 2 shows operations with each propellant. Stable operations were confirmed with both cathodes. The results of a thrust of 21.4mN, a specific impulse of 2,907s, and a thrust efficiency of 4.92% could be obtained with H 2 as a propellant. Also results of a thrust of 151mN, a specific impulse of 768s, and a thrust efficiency of 5.18% could be obtained with NH 3. (a) H 2 (b) NH 3 Figure 2. Operations with each propellant. After the operation, severe erosion was confirmed at the cathode tip. This problem has to be solved for practical use. One of the solutions may be to use hollow cathodes. The discharge of these cathodes isn t spot mode but diffusion mode. Reduction of the erosion amount is expected because the maximum cathode surface temperature doesn t close to the melting point of the cathode material. Hollow cathodes made of pure tungsten were designed and used for operations. Figure 3 shows operations with each hollow cathode. Stable operations could be confirmed with both hollow cathodes. However severe erosion didn t improve. Therefore the experimental condition has to be modified. 1

2 (a) Single hollow cathode (b) Multi hollow cathode Figure 3. Operations with each hollow cathode An all radiation-cooled MPD thruster has been designing at OIT 2). The material with high melting point as tungsten, carbon, and TZM has to be used around the electrodes because it was expected the temperature of these will be higher than water-cooled thruster. In order to know the temperature of magnetic circuit in operation, thermal analysis was carried out. Figures 4(a) and 4(b) show 3D model and the result of thermal analysis. According to the result, while the temperature of permanent magnets was over 500 degrees Celsius, the temperature of the radiation panel was around 400 degrees Celsius. 500 degrees Celsius is more than irreversible demagnetization temperature of a SmCo magnet, 350 degrees Celsius. In order to lower temperature of permanent magnets, temperature of the radiation panel must be raised more. (a) 3D model (b) The result of thermal analysis Figure 4. The all radiation-cooled MPD thruster. B. Arcjet thruster 1 kw-class steady-state arcjet thrusters have been investigated 3). It is planned to use for satellite attitude control. We focused on HAN which is low toxicity propellant because it is easy to handle. In addition, a combustion performance with HAN is more than that with hydrazine. Therefore, it is believed that it will become a preferable propellant of next-generation satellite propulsion systems. However, initial ignition and stable operation is difficult with HAN. It is important to improve this point for practical use. Figure 5 shows the plasma plume with HAN decomposed gas. We also focus on water as a propellant, but it is difficult to use it because of freezing in space. To prevent freezing it, we manufactured gas generators to vaporize it. Accordingly, we could observe the short time operation using only water 4.5). The long operation and reduction of the electrodes are future issues. 2

3 Figure 5. The plasma plume with HAN decomposed gas. Figure 6 shows three types of propulsion performance. The propellant flow rate is set 40 mg/s and current rate is set 8 A, respectively the thrusts with the HAN decomposed gas, the hydrazine decomposed gas and pure nitrogen are mn, mn and mn, respectively, with specific impulses of s, 318 s and s at input powers of kw, kw and kw, and the thrust efficiencies are 5.42 %, 5.53 % and 4.68 %, respectively. 3

4 Figure 6. Three types of propulsion performance kw steady-state arcjet thrusters have also been studied 6). It is planned to use for orbital transfer. Since hydrazine can be identified with monopropellant and bipropellant system, it is used as a propellant for arcjet thrusters. Hydrogen and ammonia are also investigated for high-performance and long-time operation. Figure 7 shows red heat of the radiation-cooled anode just after 10 minutes operation. The surface temperature of the anode was measured by a radiation thermometer (TASCO JAPAN corp.). It was 1,098 K with the thermal emissivity of carbon, In addition, the radiation energy loss at the anode was calculated with the surface temperature. As a result, the radiation heat loss at the anode was roughly calculated to be 13.18% on the total input power to the arcjet thruster. We are designing an all radiation-cooled arcjet thruster system. Figure 7. The red heat of the anode. C. Pulsed Plasma Thruster The project of Osaka Institute of Technology Electric-Rocket-Engine onboard Small Space Ship (PROITERES) was started at OIT in ). In 1st PROITERES, a nano-satellite with electrothermal-acceleration-type Pulsed Plasma Thrusters (PPTs) was successfully launched by the Indian PSLV C-21 rocket on September 9th, The main mission is to change an orbital altitude orbits as a powered flight by PPT systems for 1 km, and to observe Kansai 4

5 district in Japan with a high-resolution camera. Furthermore, the project of the 2nd PROITERES was started in 2010, as shown in Fig. 8. The 2nd PROITERES satellite aims at the powered flight with longer distance, i.e. changing km in altitude on near-earth orbits, than that of the 1st PROITERES. In order to achieve the main mission of the 2nd PROITERES satellite, a total impulse of 5,000-10,000 Ns is required. For the reason, we cannot use the PPT system for the 1st PROITERES satellite. The PPT by the numerical calculation and the experiment is performed development. The input energy of the PPT system was enlarged from 2.43 to J for the improvement of a thrust performance. The numerical calculation, we 400 shots operation Figure 8. The flight image of the 2nd PROITERES satellite. by changing cavity length and diameter was performed for high-power PPT onboard the 2nd PROITERES satellite. Tables 1 and 2 show the condition of numerical calculation by changing cavity length and diameter. Figure 9 show the results of numerical calculation 9). Table 1. Conditions of numerical calculation by changing cavity length. Cavity room Nozzle 10.0 mm 2.4, 3.0, 4.0, 5.0 mm 19.0 mm 20.0 mm Capacitance 19.5 µf Charging voltage 1,800 V Inductance 0.35 µh Resistance Ω Table 2. Conditions of numerical calculation by changing cavity diameter. Cavity room Nozzle 10.0, 15.0, 20.0, 25.0 mm 4.0 mm 19.0 mm 20.0 mm Capacitance 19.5 µf Charging voltage 1,800 V Inductance 0.35 µh Resistance Ω shot 400 shots Impulse bit, μns 1400 Impulse bit, μns shot 400 shots Cavity length, mm (a) Changing cavity length Figure 9. Results of numerical calculation Cavity diameter, mm (b) Changing cavity diameter 5

6 The experiment, we measured initial performances of this PPT system with 350 shot. The experimental PPT is shown in Fig. 10. Experimental condition is shown in Table 3. We measured performance characteristics by changing discharge room length from 20 to 50 mm in order to determine an optimum discharge room configuration 10). (a) Thruster overview (b) Operation (discharge room length: 40mm) (c) Schematic view Figure 10. Experimental PPT. Table 3. Experimental conditions. Discharge room 4 mm 20, 25, 30, 35, 40, 45, 50 mm Nozzle 20 mm 18 mm Capacitance 19.5 µf Charging voltage 1,800 V Input energy J Figures 11 and 12 show the results of this experiment. The impulse bit increased with longer discharge room length. It is caused by the mass shot increased because the discharge room length gets longer. On the other hand, the specific impulse decreased. It is shown the performance approaches that of the electromagnetic PPT with decreasing discharge room length. This PPT achieved initial performances of the impulse bit of 1,794-2,600 µns, the specific impulse of s and the thrust efficiency of % at J with discharge room lengths between 20 mm and 50 mm. We developed Multi- discharge-room type PPT of longtime operation, as shown Fig

7 Figure 11. Impulse bit and mass shot vs discharge room length. Figure 12. Specific impulse and thrust efficiency vs discharge room length. Figure 13. Multi-discharge-room type PPT. D. Hall thruster Three types of Hall thruster have been investigated at OIT 11-13). Figure 14 shows the Anode-Layer-Type Hall thruster (TAL-type) and the Magnetic-Layer-Type Hall thruster (SPT-Type) and the Cylindrical-Type Hall Thruster (CHT-type). The high-power TAL and SPT types are studied for manned Mars exploration, and the CHT-type will be equipped in the 3rd PROITERES nano-satellite for powered-flight to Moon. In order to achieve all missions, these Hall thrusters was measured the performance. Figure 15 shows operations each Hall thrusters. Figures 16 and 17 show SPT-type and CHT-type of propulsion performance. SPT-type was achieved a thrust of mn, specific impulse of s, and a thrust efficiency of % with an input voltage of 300-1,000 V, and a hollow cathode flow rate was increased at discharge voltage of V in gray area as shown Fig. 16. CHT-type was achieved a thrust of mn, a specific impulse of s, and a thrust efficiency of % with an input power of W. TAL-type has been searched stable operations now. 10)-13) 7

8 (a) TAL-type (b) SPT-type (c) CHT-type Figure 14. Hall thrusters. (a) SPT-type (b) CHT-type Figure 15. Operations each Hall thrusters. (a) Specific impulse vs discharge voltage (b) Thrust efficiency vs discharge voltage Figure 16. SPT-type of propulsion performance. 8

9 (a) Specific impulse vs discharge voltage (b) Thrust efficiency vs discharge voltage Figure 17. CHT-type of propulsion performance. References 1 Suzuki, T., Koyama, N., Sugiyama, Y., Sakoda, H. and Tahara, H., Performance Characteristics of Steady-State MPD Thrusters with Permanent Magnets and Multi Hollow Cathodes for Manned Mars Exploration, 30th International Symposium on Space Technology and Science34th International Electric Propulsion Conference and 6th Nano-satellite Symposium, IEPC /ISTS-2015-b-197, Hyogo, Japan, Sugiyama, Y., Koyama, N., Suzuki, T., Sakoda, H. and Tahara, H., Thermal Characteristics of Radiation-Cooled Steady- State MPD Thrusters with Permanent Magnets and Multi Hollow Cathodes for In-Space Propulsion, 30th International Symposium on Space Technology and Science34th International Electric Propulsion Conference and 6th Nano-satellite Symposium, IEPC /ISTS-2015-b-198, Hyogo, Japan, Fukutome, Y., Shiraki, S., Matsumoto K, Inoue, F., Tahara1, H., Nogawa, Y., and Momozawa, A., Performance Characteristics of Low-Power Arcjet Thrusters Using Low-Toxicity Propellants of HAN, 30th International Symposium on Space Technology and Science 34th International Electric Propulsion Conference and 6th Nano-satellite Symposium, IEPC and ISTS-2015-b-229, Hyogo, Japan, Shiraki, S., Fukutome, Y., Inoue, F., Matsumoto, K., Tahara, H., Nogawa, Y., and Momozawa, A., Performance Characteristics of Low-Power Arcjet Thruster Systems with Gas Generators for Water, 30th International Symposium on Space Technology and Science 34th International Electric Propulsion Conference and 6th Nano-satellite Symposium, IEPC /ISTS-2015-b-230, Hyogo, Japan, Nogawa, Y. and Tahara, H., Water Electrical Propulsion System Combined with Manned Space Mission, 30th International Symposium on Space Technology and Science 34th International Electric Propulsion Conference and 6th Nano-satellite Symposium, IEPC /ISTS-2015-b-33, Hyogo, Japan, Inoue, F., Fukutome, Y., Shiraki, S., Matsumoto K., and Tahara1, H., Performance and Thermal Characteristics of High- Power Hydrogen Arcjet Thrusters with Radiation-Cooled Anodes for In-Space Propulsion, 30th International Symposium on Space Technology and Science 34th International Electric Propulsion Conference and 6th Nano-satellite Symposium, IEPC and ISTS-2015-b-231, Hyogo, Japan, Kamimura, T., Nishimura, Y., Ikeda, T., and Tahara, H., R&D and Final Operation of Osaka Institute of Technology 1st PROITERES Nano-Satellite with Electrothermal Pulsed Plasma Thrusters and Development of 2nd and 3rd Satellites, 30th International Symposium on Space Technology and Science 34th International Electric Propulsion Conference and 6th Nanosatellite Symposium, IEPC /ISTS-2015-b-209, Hyogo, Japan, Kojima, Y., Kamimura, T., Nishimura, Y., Ikeda, T., Fujita, R., Tahara, H., and OIT PROITERES Team, R&D and Final Operation of Osaka Institute of Technology 1st PROITERES Nano-Satellite with Electric Rocket Engines and Development of 2nd and 3rd Satellites, 30th International Symposium on Space Technology and Science 34th International Electric Propulsion Conference and 6th Nano-satellite Symposium, NSAT-2015-f-13, Hyogo, Japan, Fujita, R., Muraoka, R., Kanaoka, K., Huanjun, C., Tanaka, M., Tahara, H. and Wakizono, T., Flowfield Simulation and Performance Prediction of Electrothermal Pulsed Plasma Thrusters Onboard Osaka Institute of Technology PROITERES Nano- 9

10 Satellite Series, 30th International Symposium on Space Technology and Science 34th International Electric Propulsion Conference and 6th Nano-satellite Symposium, IEPC /ISTS-2015-b-207, Hyogo, Japan, Kanaoka, K., Fujita, R., Muraoka, R., Tahara, H. and Wakizono, T., Research and Development of High-Power Electrothermal Pulsed Plasma Thruster Systems for Osaka Institute of Technology 2nd PROITERES Nano-Satellite, 30th International Symposium on Space Technology and Science 34th International Electric Propulsion Conference and 6th Nanosatellite Symposium, IEPC /ISTS-2015-b-22, Hyogo, Japan, Kagota, T., Takahata, Y., Kakuma, T., Nishida, M., Ikeda, T. and Tahara, H., Performance Characteristics of High-Power, High-Specific-Impulse Anode-Layer-Type Hall Thrusters for In-Space Propulsion, 30th International Symposium on Space Technology and Science34th International Electric Propulsion Conference and 6th Nano-satellite Symposium, IEPC /ISTS-2015-b-153, Hyogo, Japan, Takahata, Y., Kakuma, T., Kagota, T., Nishida, M., Ikeda, T. and Tahara, H., Research and Development of High-Power, High-Specific-Impulse Magnetic-Layer-Type Hall Thrusters for Manned Mars Exploration, 30th International Symposium on Space Technology and Science34th International Electric Propulsion Conference and 6th Nano-satellite Symposium, IEPC /ISTS-2015-b-151, Hyogo, Japan, Kakuma, T., Ikeda, T., Nishida, M., Kagota, T., Takahata, Y. and Tahara, H., Research and Development of Low-Power Cylindrical-Type Hall Thrusters for Nano/Micro Satellites, 30th International Symposium on Space Technology and Science34th International Electric Propulsion Conference and 6th Nano-satellite Symposium, IEPC /ISTS-2015-b-302, Hyogo, Japan, Joint Conference of 30th ISTS, 34th IEPC and 6th NSAT,