A Very Low Power Arcjet (VELARC) for Small Satellite Missions

Transcription

1 A Very Low Power Arcjet (VELARC) for Small Satellite Missions IEPC Presented at the 32 nd International Electric Propulsion Conference, Wiesbaden Germany Birk L. Wollenhaupt 1, Alex Hammer 2, Georg Herdrich 3, Stefanos Fasoulas 4, Hans-Peter Röser 5 Institute for Space Systems, University of Stuttgart, 70569, Germany A thermal arcjet thruster system of very low power of less than 300 W is developed for the needs of small satellite systems with low power budgets. This thruster shall expand the arcjet thruster range of the Institute of Space Systems (IRS) ranging currently from 1 kw to 100 kw thruster systems 1. The new thruster is derived from a 1 kw laboratory model and is developed in close reference to former very low power arcjets developed by NASA Lewis Research Center (LeRC) 2,3 and Institute of Space and Astronautical Science (ISAS) 4 in the 1990s. The thruster s performance with ammonia, hydrogen and argon and possible missions are presented within this paper. I. Introduction n the last decade small satellites attracted increasing interest. And the more complex these missions become the I higher is the need of a propulsion system for these satellites with appropriate requirements on mass savings via increased specific impulses. At the moment cold gas systems or resistojets are used to propel these missions. Equipping those satellites with electric propulsion has to respect the limits of an electric power budget below 1 kw. Typical low power thermal arcjets as developed by IRS 1, Aerojet 6, Beijing University 7 or Osaka and Kyushi Institute 8,5,6 have power requirements ranging from 1 to 2 kw. To run a thruster of this class on small satellites requires enormous solar power panels exceeding the usual budgets. The other method is to run these thrusters in a cyclic mode, charging the satellite battery for several hours until enough energy is stored for one hour of operation. However, the mass of these batteries may also consume most of the propellant mass savings achieved by the high specific impulse while the cyclic operation leads to a lowered mean thrust and increased starting erosion. Thus if arcjets shall be utilized for small satellites, thrusters with lower power requirements need to be developed. A thruster development with the objective of a 100 W operational power arcjet is described below. II. Common Arcjet Advantages and Mission Examples Today the general arguments for the usage of arcjets on spacecrafts are the same as 20 years ago. They have a higher specific impulse than cold gas thrusters, chemical thrusters or resistojets and a higher thrust to power ratio than electrostatic or electromagnetic systems fitting to the niche between both systems types. Table 1 shows the performance of a set of thrusters with very low power requirements and similar thrust for better classification of further performance data. Other arguments for thermal arcjet application are the significantly reduced toxicity of ammonia or hydrogen compared to chemical hydrazine thrusters. On the other hand if a hydrazine feed system is already in use aboard the satellite due to chemical main propulsion, hydrazine arcjets can be used to form a dual or even triple mode propulsion system. On storability issues ammonia has a great advantage above many other 1 PhD student, space transportation, wollenhaupt@irs.uni-stuttgart.de 2 Undergraduate student, space transportation 3 Senior scientist, space transportation, herdrich@irs.uni-stuttgart.de 4 Professor, space transportation, fasoulas@irs.uni-stuttgart.de 5 Director of institute, roeser@irs.uni-stuttgart.de 1

2 propellants as it is stored as wet steam at moderate pressures and easily speaking pressurizes itself depending on the ambient temperature. The highest operation time demonstrated for arcjets so far has been above 2000 h for the MR- 510 thrusters 6. A further major increase in life expectancy may be achieved automatically by the reduced number of ignitions due to continuous operation. This is because of the major erosion occurs during ignition and transition from low voltage to high voltage mode. Further, more specific design features have been discusses before in reference 1. Several applications may now be considered for a very low power arcjet. Small satellite developers 11 just set requirements for a thruster system substituting cold gas systems for general attitude and orbit control. Thus, a general small satellite mission concept may envisage a satellite wet mass of around 200 kg with an increased velocity change requirement above 200 m/s. To fulfill these needs the thruster system is granted a power of 200 W and a total propulsion system mass of less than 20 kg (10%). While a typical nitrogen or helium cold gas system with around 75 or 140 s of specific impulse cannot fulfill this mission, a simple arcjet thruster of only 400 s specific impulse reaches the 200 m/s with only 10 kg of propellant leaving 10 kg for the overall thruster system mass or additional payload. With an estimation of the thruster system dry mass to 6 kg ( 0,3 kg for the thruster, 1 kg for the PCU, 1,5 kg for the propellant feed system and additional 3 kg for mounting and tank) a velocity change close to 300 m/s is possible. Increasing the specific impulse to 550 s for ammonia a velocity change of around 400 m/s is possible. With hydrogen at 900 s specific impulse velocity changes of 700 m/s would be possible, unfortunately increased specific tank weight has to be considered. However, if it is be possible to reduce the tank system mass for hydrogen to just 50 % of the tank wet mass the velocity change would still be in the region of 300 m/s. An application as a primary propulsion on a small satellite like the Lunar Mission BW1 concept study 12 may also be possible. With the lowered power consumption the huge dry mass on solar cells and battery (at the moment kg) may be reduced. Admittedly the life time demand of the thruster will raise because of the lower thrust but as the thruster is no longer in need of a cyclic on/off mode for power reasons it may run continuously. This will decrease the erosion dramatically and thus increase the possible live time. Table 1. Performance of very low power arcjets and comparable propulsion systems Performance Name Type Supplier Prop. ESMO- CGS Cold Gas ESMO- Stuttgart / IRS Power (W) Thrust (mn) Isp (s) Efficiency (%) Thrust per Power (mn/kw) He ~ n/a n/a - Monoprop Astrium N 2 H 4 ~ n/a n/a - Resistojet SSTL H 2 O Microjet Tokai Uni N Mini Arcjet LeRC snh LPATS sn Sagami-III Arcjet ISAS 2 H VELARC Arcjet IRS Ar , ,5 275 VELARC Arcjet IRS NH 3 VELARC Arcjet IRS H , , , SPT-35 HET Fakel Xe ADD- Simplex PPT IRS Tefl on , µn-rit Ion astrium Xe 50 0, RITA-10 Ion astrium Xe Ref ,

3 A further mission concept would be the continuous drag compensation of satellites and space infrastructure. With compensation requirements 13 of 0,3 mn/m² in a 400 km orbit and up to 15 mn/m² in a 200 km orbit an arcjet with around 30 mn in thrust can compensate the drag of a small satellite of 2 m² cross section area in 200 km height or a station of 100 m² in 400 km height. Again for missions like these life time requirements of 10 years and above with continuous operation require advanced studies on maximum operation times of arcjets. Although anode erosion is drastically reduced at continuous operation the sublimation effect at the hot cathode has to be studied in more detail. However, with the low mass of a very low power arcjet of less than 300 g a substitution the original thruster with a redundant thruster after 3-5 years would mean no harm to the total mass budget. The last but not least mentioned missions concept may be the station keeping of a small geostationary satellite of around 1000 kg mass as it was already envisaged in the 1990s by ISAS. This may be done very similar to the MR arcjet thrusters run on 6000 kg A2100 geostationary satellites. These satellites fire two thrusters with a thrust of 220 mn at 2 kw each once a week for around 30 to 90 minutes. Two thrusters of 30 mn may do the same on a 1000 kg satellite with very similar operation times. As communication satellites are very stingy on power as it has always to be drawn from the payload the lesser operation time of arcjets compared to other low thrust electric propulsion may be of interest for satellite developers. III. Historic arcjets of lowest power for reference Engaging electric propulsion development one must be aware that nearly everything developable has already been investigated sometime before and should be used as reference to compare one s own achievements. IRS develops thermal arcjets since late 1980s 1. Especially low power arcjets at powers around 1 kw have seen a great variety of models ranging from modular laboratory thrusters up to the flight model ATOS. Also a lot of different propellants have been investigated as ammonia, hydrogen, simulated hydrazine and helium 22. The performance achieved with these thrusters will be compared to the performance of the VELARC thruster in specific values. Helium has already been investigated at the IRS 22 for 1 kw class thrusters with promising efficiencies around 50% but with difficulties in the stability of the thrusters during operation. Arcjets of very low power have first been Figure 1. Arcjet thruster range at IRS (upper line from left: HIPARC 100kW, MARC 10kW, TALOS 1kW, VELARC 300W; lower line from left: ARTUS IM 3 engineering model, ATOS flight model) developed around 1994 to 1996 after the first 2 kw arcjets were brought to service and shortly before the end of general arcjet development in the 1990s. Investigation took place at NASA Lewis Research Center (LeRC) 2,3 and at ISAS 4 in Japan. The aim was the same as for the IRS today, to expand the arcjet thruster range with a thruster system that can be operated on small satellites with low power budget. The advantage was expected to be a higher specific impulse compared to chemical thruster on the one hand and on the other hand a higher thrust and lower complexity compared to other electric thrusters. At LeRC a thruster with a subsonic arc attachment had been developed before achieving powers as low as 100 to 300 W but enormous arc instabilities. The following supersonic arc attachment thruster described below was operated at 200 to 400 W with simulated ammonia and simulated hydrazine gas mixtures. While running better for medium mass flows above 10 mg/s, this thruster also suffered instabilities and efficiency losses at lower mass flows. This was identified to be a problem of high viscous nozzle losses at very low mass flow ratios. It was envisioned that the boundary layer was rapidly increasing at low mass flows prohibiting the Figure 2. NASA LeRC very low power arcjet thruster 2 expansion of the flow. Thus other expansion ratios and nozzle geometries have been investigated. It was found out that a dual cone nozzle solve this problem. The 3

4 divergent section close to the throat was kept as it is to avoid influence on the arc attachment. At an expansion ratio of 50-60:1 the divergent cone angle was increased from 15 to 45 to support rapid undisturbed expansion. This increased the efficiency at low mass flows from 17 % to around 30 %. Another discovery was a generally lower performance for simulated hydrazine compared to simulated ammonia with around 40 s of specific impulse difference which is due to the lower molecular mass of ammonia. The mass of the thruster system was very light weight with just 200 g for the thruster and 500 g for the PCU. The second reference thruster was developed at ISAS called SAGAMI-III 4. First experiments were conducted with a relatively big nozzle throat of 0,5 mm based on the design of the low power arcjet SAGAMI-II. But further investigations showed a superior performance regarding efficiencies and stability by a smaller nozzle of 0,3 mm especially for low mass flow rates. Interestingly although LeRC showed lower performance for simulated hydrazine compared to simulated ammonia, ISAS topped the performance of the LeRC ammonia thruster while running on hydrazine. Also the total mass of the thruster was even lower at only 110 g. Figure 3. ISAS Sagami III thruster 4 Another interesting investigation took place later at Tokai University 17 with a thruster consuming only 20 W but with very low specific impulses of only 250 s. Additionally, the current voltage characteristics of this thruster with around 400 V and only 50 ma indicates that this thruster is more of a glow jet type than an arcjet. Thus this thruster is not considered as reference for the following development. IV. VELARC design At IRS 1 the previously investigated power classes range from 1 kw for the ARTUS, ATOS & TALOS thrusters over 10 kw for the MARC thrusters up to 100 kw for the HIPARC thrusters. As a first step to advance to lower power classes the smallest available laboratory model the ARTUS LM 3 was chosen as a platform for very low power tests. It was built in the late 1990s as a mass optimized low power arcjet. This thruster was already equipped with a small nozzle (nozzle 1) with a throat diameter of only 0,4 mm. First tests with this thruster were able to run the thruster stable at power levels down to 300 W. Preliminary results of these tests reached a specific impulse of 370 s and a thrust efficiency of 22 % at a power of 375 W and a mass flow rate of 12,7 mg/s ammonia. Figure 4. VELARC thruster The thruster (Figure 4) follows the usual design guidelines with a supersonic arc attachment nozzle out of tungsten thorium alloy or the newer ones tungsten lanthanum alloy. For insulation boron nitride blocks and aluminum oxide tubes are used. The thruster body is made out of Inconel. As cathode pass through a ceramic insulation part by Friatec is used. For sealing graphite foil is used. The electrode gap can be adapted by a thread at the end of the thruster. There the cathode can be screwed into until contact between cathode and anode is reached and then retracted until the required gap is attained. For design changes on electrode design and configuration the investigations of NASA LeRC 2,3 and ISAS 4 have been used as guideline and for performance comparison. 4

5 A. Nozzle design For the anode three different designs have been tested. The original design (nozzle 1) of the low power arcjet was already a small nozzle of 0,4 mm in diameter and 0,3 mm in length compared to usual low power arcjets of Ø 0,6 x 0,6 mm nozzles. The exit cone had an angle of 30. To scale the anode configuration according to the power a second nozzle of 0,3 mm in diameter and 0,15 mm in length has been fabricated. A smaller throat diameter was not possible due to current machining limits, but will be approached later. A third nozzle then included the dual cone principle, as described above. This should enable a faster expansion downstream by the highly diverging cone and thus reducing the friction losses of the nozzle close to the exit, while at the same time not altering the arc settling point close to the throat in the normal cone part. This principle was investigated before by NASA LeRC 3 and should increase the performance at very low mass flow rates. B. Injector design As injector two different designs have been tested. The original design consists of two injector holes directly bored into the tungsten nozzle. No swirl was introduced to the gas. This design suffered from three problems. First of Figure 5. Nozzle designs all, by including the injector into the nozzle, nozzle and injector could not be 2 & 3 exchanged separately. By incorporating the injector to the easily manufacturable boron nitride (BN) insulator the modular character of the thruster model could be enhanced. Next only two boreholes without a swirl produced stable areas of low dynamic pressure between these holes. In these areas the arc can get stuck in low voltage mode without the gas flow forcing it out of the convergent nozzle area. This leads to high erosion and a failed arcjet starting. And last the ignition voltage of the original injector was very close to the maximum achievable voltage of the power supply. This led to failed Figure 6. Injector design ignitions frequently. The new injector (Figure 6) was designed with respect to the knowledge gained at the low power arcjet project TALOS in the last years. Four holes have been implemented in the insulator opening close to the cathode and with a tangential offset for introduction of a swirl. Now the ignition and transition to high voltage mode is very reliable at around 3 kv. V. Test facility The test facility used is the same as for the 1 kw arcjet class. It is build up in a cylindrical vacuum test chamber of 1,2 m in diameter and 2 m in length. The 3-stage pumping system is capable of generating a pressure of 1 Pa before test and around 3 Pa at arcjet operation. The thrust is measured by a pendulum thrust balance with a displacement sensor measuring the deflection of the pendulum. The balance is calibrated by three very well known weights at 0, 50 and 100 mn. An option to calibrate the pendulum with more sampling points by an accurate force transducer was considered but didn t perform well as the pendulum always bounces off from the transducer leading to a highly alternating sensor signal. The accuracy of the thrust balance is around 1 mn considering linearity and repeatability. A small zero drift of up to 2 mn was observed and the data was corrected accordingly. The mass flow is regulated by a pressure regulator and an orifice and rechecked with a thermal mass flow controller. First experiments with the VELARC thruster showed unacceptable deviation of the mass flow of at very low values. Thus the mass flow had to be recalibrated for subsequent tests with a recent calibrated thermal mass flow controller. For increased transition speed between low voltage mode and high voltage mode an afterburner valve with a second orifice is included downstream of the pressure regulator increasing the mass flow and thereby the dynamic pressure in combustion chamber for a short pulse. This pushes the arc through the nozzle after ignition but reduces the pressure again before it is blown out. The remainder of the test facility consists of a pyrometer for temperature measurement on the nozzle surface, a thermocouple for temperature measurement at the mounting flange and a data scan unit for sensor data recording. VI. VELARC Performance The following section presents the performance results of the VELARC thruster run with ammonia, hydrogen and argon for the nozzles 2 and 3. Unfortunately, for helium no stable operation could be achieved, thus no data is 5

6 presented here. Further experiments on helium will increase the electrode gap as it was done by Rybakov 22 to acquire stable operation. Argon was only investigated with nozzle 2 as the exit velocities are just too low to be considered furthermore, although the power level of argon was very low and easily met the 100 W objective of this study. A. Absolute Performance Figure 7. Absolute performance Figure 7 shows the absolute thrust above absolute power of the VELARC thruster. The aim of 100 W could not be reached for hydrogen and ammonia. Lowest stable operation was achieved at 230 W for ammonia and 290 W for hydrogen. With argon the 100 W aim is no problem due to the very low voltage (see FIGURE 8), but due to the low specific impulse it is not considered as success. Having a look the current-voltage-characteristics (FIGURE 8) it seems as if the reason for the power limit is a minimum current. At a current lower than 4 A the thruster became unstable or just switched off independently of the gas used. Compared to LeRC and ISAS operation conditions this is still a relatively high current for very low power arcjets. Additionally the voltage for VELARC operation is very low with voltages ranging from 80 V down to 55 V. Thus it may seem as if the thruster was run a low voltage mode. But this can be excluded as a clear plasma plume could be seen as an indication for a high voltage mode and a lower voltage mode has been observed during transition. Two design criteria alter the voltage and may be the reason for this behavior: the electrode gap and the dynamic pressure of the gas. Thus the reason for the low voltage may be a smaller electrode gap as in the reference systems, but the gap determining method and the gap distance is the same as at LeRC. Indeed Sankovic admits 3 that the gap of prior tests 2 may have been higher than declared but correcting this altered the voltage only from 180 down to 120 V while we are measuring voltages of only 60 to 80 V. The second reason, a too low pressure can also be neglected as the feed line pressure close to the thruster was very high with up to 4 bars at 10 mg/s. Further experiments will now abandon design guidelines of the reference systems and will test a higher electrode gap for higher voltage. This may also reduce the anode losses due to a reduced current level and thus may increase the efficiency. Next the reason for the current limit has to be found out and the current has to be reduced further at higher voltages, else the low power requirement can t be fulfilled with this thruster system. 6

7 Figure 8. Current voltage characteristics compared to LeRC and ISAS thruster B. Specific Performance Figure 9 shows the specific impulse over specific power for the nozzles 2 and 3 for all three gases. As usual the gases line up depending on their atomic weight with argon achieving lowest specific impulse around 150 s, ammonia with medium specific impulse of s and hydrogen with high specific impulse of s. Unfortunately the efficiencies of argon and ammonia are very low at around 20 to 25 % and also hydrogen only scratches the 30 % while typical hydrogen thrusters would achieve efficiencies of over 40 % as depicted in Figure 10. Especially at lower mass flow rates the efficiency drops further. It even seems as if the efficiency is a direct function of the absolute mass flow rate (Figure 12). A reason for this efficiency drop was thought to be the viscous losses in the nozzle dominant for low mass flows as observed by Sankovic 3. Nozzle 3 with the new design adopted from LeRC is intended to solve this problem. Comparing the data of nozzle 2 and nozzle 3 run with ammonia as propellant an actual slight increase in performance can be observed for lower mass flows. Although this is not as significant as shown by LeRC, the principle seems to work. However, for hydrogen and argon the data base isn t large enough to make a prediction of the effect of the new nozzle. Compared to the reference thrusters of LeRC and ISAS (Figure 11) the VELARC thruster still needs improvement. Both reference thrusters achieve the 30% and above easily. The only advantage of the VELARC thruster are the higher specific energies above 60 MJ/kg achievable at stable operation generating high specific impulses regardless the low efficiency. If it will be possible to increase the efficiency these specific impulses will increase further. Anyhow it has to be admitted that the influence of these high specific energies on lifetime expectance have to be investigated and may generate requirements on high durable tungsten alloys 23 like W-4Re-HfC (also known as unobtainium) or again a reduction of the specific energies. For hydrogen even higher specific energies are almost no problem for thruster durability making it a perfect option for high specific impulses. Its high specific impulses above 850 s are very comparative to other electric propulsion systems as low power Hall Effect Thrusters (HET) like SPT-35. This fact supplemented with a relatively high thrust to power ratio of 60 mn/kw may make hydrogen driven very low power arcjets a good option for the missions discussed above, if storability problem of hydrogen in orbit can be solved. 7

8 Figure 9. Specific performance of the VELARC thruster Figure 10. Specific performance compared to low power arcjets from IRS 8

9 Figure 11. Specific performance compared to very low power thruster from LeRC 2,3 and ISAS 4 Figure 12. Efficiency depending on mass flow 9

10 VII. Conclusion A thruster running stable on 300 W with ammonia and hydrogen has been developed. Unfortunately by decreasing the current to less than 3,5 A the thruster got instable and shut off after a minute. Thus the initial objective of developing a 100 W arcjet was only achieved for argon as propellant due the low bow voltage. Especially as it seems as higher voltages are needed for increased efficiency this current restriction has to be further investigated and somehow circumvented. The thrust efficiency of all propellants is low at around 20 % and drops with decreasing mass flow. This is still low compared to the reference thrusters of same or higher power. A try to increase the efficiency by incorporating a dual cone nozzle was partly successful, but not sufficient. However the specific impulses achieved for ammonia and hydrogen ranging between 400 to 520 s and 700 to 860 s respectively are already high enough to consider this thruster as an option for the small satellite missions. The next steps of VELARC development should be the investigation of operation stability and efficiency at increased electrode gap. This may also help to increase the stability of the VELARC thruster run with helium. Due to the low frozen flow losses of helium and its low molecular weight high efficiencies and high specific impulses are expected for helium operation. When a solution of the current limit problem is found the power should be further reduced to 100 W to meet the requirements of small satellites. Then mass reduction and increase of usage comfort shall complete the development program. Acknowledgements The funding of the PhD of Birk Wollenhaupt by Carl-Zeiss-Stiftung is greatly acknowledged. References 1 Wollenhaupt, B., Herdrich, G., et. al.: Overview about Thermal Arcjet Thruster Development, ISTS-2011-b-51, 28 th ISTS, Okinawa, Japan, Sankovic, J.M, Jacobson, D.T.: Performance of Miniaturized Arcjet, AIAA , 31 st Joint Propulsion Conference & Exhibit, San Diego, USA, July Sankovic, J.M., Hopkins, J.B.: Miniaturized Arcjet Performance Improvement, AIAA , 32 nd Joint Propulsion Conference & Exhibit, Buena Vista, USA, July Ogiwara, K., Hosoda, S., et. al.: Development and Testing of a 300 W Class Arcjet, IEPC-95-17, 24 th IEPC, Moscow, Russia, Herdrich, G., Bauder, U., Bock, D., Eichhorn, C., Fertig, M., Haag, D., Nawaz, A., Lau, M., Schönherr, T., Stindl, T., Röser, H.-P., Auweter-Kurtz, M.: Activities in Electric Propulsion Development at IRS, Trans. JSASS Space Tech. Japan, Vol. 7 (2009), pp.tb_5-tb_14 6 Lichtin, D.A., Chilelli, N.V., Henderson, J.B., Rauscher, R.A., Young, K.J., McKinnon, D.V., Bailey, J.A., Roberts, C.R., Zube, D.M., Fisher, J.R. : AMC-1 (GE-1) Arcjets at 12-plus Years On-Orbit, AIAA , Joint Propulsion Conference & Exhibit, Denver, USA, August Tang, H., Zhang, X., Liu, Y., Wang, H., Shi, C.: Experimental Study of Startup Characteristics and Performance of a Low-Power Arcjet, AIAA , Journal of Propulsion and Power, Vol. 27, No.1, January Hiratsuka, S., Onoe, K., Tahara, H., Yoshikawa, T., Suzuki, H., Uematsu, K.: Thrust Performance and Thermal Analysis of a Low Power Arcjet Thruster, IEPC , 26 th Int. Electric Propulsion Conference, Kitakyushu, Japan, October Masuda, I., et. al.: Performance Characteristics of Direct-Current Arcjet Thrusters Using Hydroxyl-Ammonium- Nitrate Propellant, ISTS-2011-b-49, 28 th ISTS, Okinawa, Japan, Kakami, A., Beeppu,S., Maiguma, M., Tachibana, T.: Performance Characteristics of a DME Propellant Arcjet Thruster, ISTS-2011-b50, 28 th ISTS, Okinawa, Japan, Roemer, S.: TET-1/X. Die Rolle von EP für Kleinsatelliten, 4th German Workshop on Electric Propulsion, Leipzig, Germany, Bock, D.,Herdrich, G.,et.al.: An Advanced Ammonia Propellant Feed System for the Thermal Arcjet TALOS, AIAA , 43rd JPCE, Cincinnati, USA, Montenbruck, O., Gill, E.: Satellite Orbits, 1 st ed., Springer-Verlag, 2000, Chap ESMO Propulsion Team Stuttgart, PSGF PDR Documentation, Stuttgart, Germany, Astrium Website ( 16 SSTL Website ( 17 Horisawa, H., Kimura, I., et.al.: DC Plasma-Jet Microthrusters, IEPC , 27 th IEPC, Pasadena, USA, Nawaz, A., Albertoni, R., Herdrich, G., Auweter-Kurtz, M.: Thrust Efficiency Optimization of a Pulsed Plasma Thruster, 26 th ISTS, Hamamatsu, Japan, DGLR-Expert Committee R1.3, Präsentation des Status von EP in Deutschland, DGLR Report, Leipzig, Germany, June,

11 21 Riehle, M., Laux, T., Huber, F., Kurtz, H.L., Auweter-Kurtz, M.: Performance Evaluation of Regeneratively Cooled 1, 10 & 100 kw Arcjets, IEPC , 26 th International Electric Propulsion Conference, Kitakyushu, Japan, October Rybakov, A., Auweter-Kurtz, M., Kurtz, H., Riehle, M.: Investigations of a 1 kw Class Arcjet Thruster with Helium as Propellant, AIAA , 38 th Joint Propulsion Conference & Exhibit, Indianapolis, USA, July Lichon, P.G., Sankovic, J.M.: Development and Demonstration of a 600 Second Mission Average Arcjet, IEPC , 23 th IEPC, Seattle, USA,