A 300 W Microwave Thruster Design and Performance Testing

A 300 W Microwave Thruster Design and Performance Testing Daniel J. Sullivan * and John F. Kline Research Support Instruments, Princeton, NJ, 08540 S. H. Zaidi and R. B. Miles Department of Mechanical and Aerospace Engineering Princeton University, Princeton, NJ, 08544 An experimental program for the development and performance testing of a microwave electrothermal thruster, the MET-100, is discussed. The thruster has been operated on both helium and nitrous oxide propellants. A pendulum based thrust stand has been used to measure thrust, and I sp and thermal efficiency numbers have been calculated. The effect of buoyancy on the MET-100 operation is examined and its implications considered. The method for determining thrust is described and the design the experiment is presented. A brief history of the MET development is also included. I. Overview HIS paper presents the results of research conducted by Research Support Instruments (RSI) and the Applied T Physics Group in the Department of Mechanical and Aerospace Engineering at Princeton University on the design and performance testing of a low power microwave electrothermal thruster, the MET-100. The MET-100 performance has been evaluated by means of direct thrust measurements. The measurements have been made with the MET-100 operating on both nitrous oxide (N 2 O) and helium (He) propellants. These preliminary results have demonstrated that the MET-100 has a wide operational envelope whereby both its thrust level and specific impulse can be chosen within a relatively wide range. The horizontal orientation of the MET-100 used to obtain thrust measurements introduced a buoyancy force which acted on the plasma discharge. This buoyancy force, while not an issue for space applications, did limit the operational envelope within which direct thrust measurements could be obtained. This buoyancy force is a function of gas density and therefore was more pronounced during nitrous oxide operation than during helium operation. Once thrust measurements where conducted with the horizontal thrust stand, the MET-100 was switched to a vertical orientation. This vertical orientation removed the buoyancy effect and the operational envelope was demonstrated to be substantially increased. However, attempts to directly measure thrust by means of a load cell were unsuccessful. The MET-100 nominally requires about 300 W to operate. The conversion of bus power to microwave power via the 7.5 GHz magnetron is only 50% efficient, so only 150 W of this input power is available to directly heat the flowing propellant gas. This current study has not considered ways to improve on this situation, however, future system studies will examine the feasibility of using the nitrous oxide propellant to provide regenerative cooling for the magnetron. * Principal Scientist, Member AIAA, dsullivan@researchsupport.com. Principal Scientist. Research Staff Member, Member AIAA. Professor, Fellow AIAA Copyright 2004 by the Authors. Published by the, Inc., with permission. 1

II. Experimental Results The MET-100 has been tested using two propellant gases, namely helium and nitrous oxide. The performance was evaluated through direct measurements of mass flow rate, absorbed microwave power, plenum pressure, and thrust. The experimental set-up used for the testing described in this paper is shown in Figure 1. The MET-100 was positioned horizontally and mounted to a mounting flange of the vacuum tank. A support structure was also mounted to the flange. This support structure contained the 7.5 GHz magnetron, the dual-directional power coupler, the power coupling antenna and the sliding short. Two stepper motors connected to this support structure allowed for automated configuration of the MET-100 with respect to the position of the sliding short which determined the cavity length and the depth of insertion into the cavity of the power coupling antenna. The thrust was recorded by means of a thrust catcher which was suspended inside the vacuum tank directly in line with the MET-100 nozzle. The thrust catcher consisted of a perforated steel cylinder open at one end and sealed at the other. The perforated can was surrounded by a series of baffle plates. The exhaust gas would enter the catcher and transfer all of its linear momentum before exiting through the radial holes. The catcher would be displaced by a small amount which corresponded to the total momentum of the exhaust stream. The displacement was sensed by a LVDT sensor. Figure 1: Schematic showing the main components used during testing of the MET-100. The use of stepper motors allowed for repeatable positioning of both the cavity length and the coupling antenna insertion depth. The thrust catcher was suspended by two wires within the vacuum tank and its deflection was recorded by a LVDT sensor. 2

A. Performance Calculations Any coherent representation of performance data requires that the terms be accurately defined for the device being tested. This section will provide an explanation of the parameters being used to evaluate the MET-100 performance. The following terms will now be defined: Absorbed power (W) Mass flow rate (mg/s) Plenum pressure (atm) Vacuum Tank Pressure (torr) Specific power (MJ/kg) Thrust (mn) Specific Impulse (s) Thermal efficiency (%) 1. Absorbed Power The MET-100 utilizes microwave power to produce and maintain its plasma discharge. The microwave power is generated by a magnetron, is transmitted though a coaxial transmission line and is introduced into the microwave cavity via an antenna. Microwave power can flow in both directions along a coaxial transmission line and thus any of the power which flows forward along the line from the magnetron (forward power, P f ) which is not absorbed by the plasma discharge due to impedance mismatch is reflected back along the coaxial line (reflected power, P r ) towards the magnetron. This reflected power is prevented from reaching the magnetron by means of a device known as an isolator. The reflected power represents an inefficiency in the system. The forward and reflected power are measured by means of a dual directional power coupler placed in the coaxial transmission line. This device has two output ports and it couples out of the transmission line an amount of power proportional (50 db reduction) to the forward and reflected power present in the transmission line, i.e. if 150 W of forward power is being transmitted by the coaxial line, the coupler will provide a signal of 1.5 mw. The signal is collected by a pair of HP power sensors and displayed by a pair of HP power meters. Thus the power absorbed by the plasma discharge, the absorbed power, P, is defined as: P P f P r (1) The units of power are Watts (W). 2. Mass Flow Rate The mass flow rate is defined as the flow of gas through the nozzle of the MET-100. The mass flow rate was measured using a mass flow meter and the flow rate was controlled using a multiple turn needle value. The mass flow meter and needle valve are connected to the MET-100 through ¼ line and the flow enters the MET-100 via two injection ports which have a diameter of 0.025. The units of mass flow rate are milligrams per sec (mg/s). 3. Plenum Pressure The plenum pressure is defined as the average gas pressure in the section of the MET-100 which contains the plasma discharge. The pressure was monitored using a fast acting pressure transducer. The diameter of the pressure tap was 0.025, and the units of pressure are given in atmospheres (atm). 4. Vacuum Tank Pressure The vacuum tank pressure was measured and recorded during MET-100 operation. The units of pressure are torr (torr). 3

5. Thrust The thrust developed by the MET-100 is a measured quantity. This value provides a means whereby the performance of the thruster can be accurately gauged. The thrust is inferred by capturing the linear momentum of the gas exhausting through the thruster nozzle, this method has been used successfully by other researchers. 1, 2 While it would be more desirable to directly measure the thrust by mounting the MET-100 on a thrust stand of the inverted pendulum type 3,4, such an option was not possible given the time restraints of the Phase I program. Thrust is defined as the force required to accelerate the propellant gas to the exhaust velocity, u e, and can be written as: T d dt ( mue ) m& ue (2) The units of thrust are millinewtons (mn). 6. Specific Power The specific power is defined as a ratio of power to mass flow rate. For the case of the MET-100 thruster, the power level used is the measured quantity of absorbed microwave power. Thus the specific power (P s ) is given as: P s P m& P f P m& r (3) The reader is alerted to that fact that there are other interpretations of specific power and it is important to know how the quantity is defined before making comparisons between electrothermal devices. The choice of absorbed power here, as opposed to bus power or even forward power, allows for a more accurate assessment of the thermal efficiency of the thruster. 7. Specific Impulse The specific impulse (I sp ) is defined as the thrust developed per weight of propellant consumed. This is given by: I T ( dw dt) d dt d dt ( mu ) u dm dt dm dt e e sp u ( m g) g g e (4) Note that the exhaust velocity, u e, as defined by both Equations 2 and 4 is a scalar quantity which is the average velocity of the exhaust gas which is moving away from the thruster in a direction perpendicular to the exit plane of the nozzle. 8. Thermal Efficiency The MET-100 thruster converts electrical energy into thermal energy and then adds that thermal energy to a flowing gas. The thermal energy of the propellant is then converted to directed kinetic energy as the gas passes through the thruster nozzle. The efficiency of this energy transfer process is termed thermal efficiency. The power of the exhaust jet arises from the conversion of the thermal energy present in the cold propellant gas plus the thermal energy added by the presence of the plasma discharge to directed kinetic energy. Thus the thermal efficiency can be assessed for this system by evaluating that fraction of the exhaust power which is a result of the thermal energy added by the plasma discharge and dividing this number by the absorbed power: 4

η thermal P P exhaust 1 m& ( ue) 2 P 2 ( u u ) e 2 e, cold 2 P s (5) where u e is the exhaust velocity defined in equation (2) and e cold, is the exhaust velocity that would be developed if the input power was zero and the MET-100 was operated as a cold gas thruster. In order to evaluate Equation 5, it is necessary to evaluate the cold flow term. The velocity of the cold gas can be determined from 1-D isentropic flow theory which gives the exhaust velocity of a gas as a function of gas stagnation temperature: u, u e, cold 2 0 k RGas T k 1 (6) where k and R Gas are the ratio of specific heats and the gas constant, respectively for the propellant gas of interest. For helium at room temperature ( T 0 293 K ): k 1.664 m RGas 2078 2 s K 2 ue 1747 m s (7) For nitrous oxide at room temperature ( T0 293 K ): k 1.302 m RGas 189 2 s K u 2 e 691 m s (8) It is noted here that previous discussions of the MET thruster (refer to Section IV) have used thermal efficiency numbers which have not subtracted out the cold flow term. This is also the case for various other performance studies of other electrothermal thrusters such as arcjets and resistojets. The reader is alerted to this fact when making comparisons between systems. B. Measured Performance Data The MET-100 performance has been evaluated by means of thrust measurements made with a pendulum based thrust catcher. This suspended device catches the linear momentum of the exhaust stream and its deflection, which is linearly proportional to the thrust, is measured using a LVDT sensor. The thrust stand was calibrated using a cold flow of helium. Accurate mass flow measurements and an assumption of exhaust velocity based on 1-D isentropic flow theory allowed the thrust stand deflection as indicated by the LVDT voltage output to be directly correlated to a thrust value. The cold gas exhaust velocity is given by (7) as 1747 m/s and the thrust was then evaluated using Equation (2). The resulting calibration is shown in Figure 2. The cold gas approach allowed direct measurements of thrust up to around 22 mn. The data has been fit to a straight line. During subsequent hot firings of the MET-100, the LDVT output was converted to thrust using this calibration. It should be noted, that this calibration provides a first-order estimate of the actual thrust developed by the MET-100. The three assumptions inherent in this estimate are: 5

I. The linear momentum caught by the thrust stand and the resulting displacement is equivalent to the actual thrust developed by the MET-100. II. The cold flow analysis gives a reliable estimate of the cold flow exit velocity. III. The small angle deflection of the thrust stand pendulum remains linear, thereby justifying the forward extrapolation of the calibration curve presented in Figure 2. 60 50 40 Thrust (mn) 30 20 T 20.579V + 3.7598 R 2 0.9975 10 0 0.0 0.2 0.4 0.6 0.8 1.0 1.2 1.4 1.6 1.8 2.0 2.2 2.4 2.6 LVDT Voltage (V) Figure 2: The pendulum based thrust stand measures thrust by deflecting a distance which is directly proportional to the magnitude of linear momentum it captures from the MET-100 exhaust. The deflection is registered by a LVDT sensor (output - 5 V per 0.1 displacement). The thrust stand was calibrated using helium in a cold flow condition. The resulting linear fit is extrapolated to the thrust regime of the hot-fire operation. Figures 3 and 4 on the following two pages provide graphs of the MET-100 performance data for helium and nitrous oxide, respectively. Each set contains three graphs: Thrust as a function of Specific Power o The thrust is evaluated using Figure 2 Specific Impulse as a function of Specific Power o I sp is calculated using the thrust measurement and the mass flow rate by means of Equations 2 and 4. Mass flow rate as a function of absorbed power. o This final graph allows for a determination as to how the specific power was varied during testing, i.e. the value for specific power can be increased by either increasing the amount of absorbed power or decreasing the mass flow rate. The I sp plots are derived from the thrust data and the measured value of the propellant mass flow rate by manipulating Equations 2 and 4: I sp T (9) g m & These plots also indicate the value of the cold flow Isp (178 s for helium and 71 for nitrous oxide) and contain an estimated fit to the data which is determined by manipulation of Equations 4 and 5 to give: 6

60 50 MET-100 nozzle - 0.0157" 40 Thrust (mn) 30 20 10 0 0 2 4 6 8 10 12 14 16 18 20 Specific Power (MJ/kg) 450 400 350 300 Isp (s) 250 200 150 100 50 10% MET-100 nozzle - 0.0157" cold flow Isp 0 0 2 4 6 8 10 12 14 16 18 20 Specific Power (MJ/kg) 18 17 16 Mass Flow Rate (mg/s) 15 14 13 12 11 10 9 8 100 110 120 130 140 150 160 170 180 Absorbed Power (W) Figure 3: Performance data for the MET-100 operating using helium propellant. For all conditions measured, the plasma was steady and well-behaved. The I sp data labeled nozzle - 0.0157 was collected while the thruster was in a vertical orientation and the I sp is inferred from the P s and an assumed efficiency of 10%. 7

100 90 80 70 nozzle - 0.0233" nozzle - 0.0189" nozzle - 0.0157 resistojet Thrust (mn) 60 50 40 30 20 10 0 220 200 180 160 140 0 1 2 3 4 5 6 7 8 Specific Power (MJ/kg) Isp (s) 120 100 80 60 40 20 5% 10% cold flow Isp nozzle - 0.0233" nozzle - 0.0189" nozzle - 0.0157" resistojet 0 0 1 2 3 4 5 6 7 8 Specific Power (MJ/kg) 70 60 nozzle - 0.0189 nozzle - 0.0233 Mass Flow Rate (mg/s) 50 40 30 20 10 0 100 110 120 130 140 150 160 170 180 Absorbed Power (W) Figure 4: Performance data for the MET-100 operating on nitrous oxide. Thrust data was collected while the thruster was operating in a semi-unstable manner. The I sp data labeled nozzle - 0.0157 was collected while the thruster was in a vertical orientation and the I sp is inferred from the P s and an assumed efficiency of 10%. 8

Figure 5: The MET-100 operating on helium propellant showing a steady and compact plasma discharge that is located within the inlet of the nozzle. 2 P η s thermal I sp + Isp, cold g (10) η The values of thermal were determined by iterative manner. The helium performance data was collected while the MET-100 was operating in a well-behaved manner, i.e. the plasma discharge was steady and occupied a compact zone within the inlet of the thruster nozzle; refer to the photo in Figure 5. The thrust stand was used to collect the data set listed as MET-100; the nozzle size was 0.0189. The data listed as nozzle 0.0157 was collected after the thruster was rotated to a vertical orientation and the nozzle size was reduced to 0.0157 ; for this second set of data, the Isp and thrust values were inferred from the measured specific power and the assumed thermal efficiency of 10%. The thermal efficiency was determined by fitting Equation 10 to the measured data. The performance data shows that the MET-100 has a wide operating range where the specific power can be varied from 8 to 20 MJ/kg. The thrust and I sp levels are modified by varying the parameters of absorbed power, nozzle diameter, and mass flow rate, however, it must be noted that the plasma discharge does have stability limits and the controlling parameters must be varied within a specified operating range. It is the goal of future research to map out this parameter space. Figure 6: The MET-100 operating on nitrous oxide propellant in a horizontal operation the nozzle is to the left. The three frames show the unsteady operation of the thruster in which buoyancy forces are acting to move the plasma discharge out of the nozzle inlet to a location above the central axis of the thruster. This behavior is made unsteady by the vertical flow field which exists inside the thruster. 9

The behavior of the MET-100 changed noticeably when the propellant gas was changed from helium to nitrous oxide. The steady plasma which was characteristic of helium operation was replaced by a plasma discharge that exhibited a rotational instability. After a period of testing, it was concluded that the instability was a result of buoyancy whereby the hot, low density plasma volume was being acting on by the surrounding higher density, cold gas resulting in a net upward force. In retrospect, it was realized that this effect had been observed during helium testing, but the effect was not very pronounced owing to the much lower density of helium. Figure 6 shows three sequential video stills that provide some idea as to the extent to which the plasma discharge moved due to this buoyancy effect. The three frames show the unsteady operation of the thruster in which buoyancy forces are acting to move the plasma discharge out of the nozzle inlet to a location above the central axis of the thruster. This behavior is made unsteady by the vertical flow field which exists inside the thruster and the plasma motion resulted in an oscillating pressure within the plenum as the plasma moved in and out of the nozzle throat. This unsteady behavior was reflected in the LVDT signal that was used to monitor the position of the thrust stand. Figure 7 shows three plots which trace the output of the LVDT voltage signal as a function of run time for the MET-100 operating on nitrous oxide. The oscillations of the LVDT signal in the upper plot clearly indicate that the thruster was operating in a very unsteady manner. At any time during the run, the voltage signal varies over a range of 2 V which corresponds to a thrust variation of 41 mn. No attempt was made to manipulate this data to obtain an average thrust value. One technique for stabilizing the plasma relies on increasing the mass flow rate. The increased mass flow rate both increases effect of the vortical flow field and increases the plenum pressure which acts to make the plasma discharge more compact. The middle plot of Figure 7 shows the result of this stabilization attempt. As the mass flow rate was increased, the plasma oscillations began to steady during the middle period of the run. However, the fluctuations at the end of the run preceded the plasma extinction due to excessive plenum pressure. It was possible to obtain useful thrust data from this operation, and this data is labeled nozzle 0.0189 in the plots presented in Figure 4. The bottom plot of Figure 7 was obtained after the nozzle throat diameter was increased from 0.0189 to 0.0233. Increasing the throat diameter allowed the flow rate to be increased to a value such that the vortical flow stabilization effect could become dominant over the buoyancy without the detrimental effect of excessive pressure increase. Of course, it should be pointed out that both solutions represented by the middle and lower plots result in a severely reduced value of specific power. The end potion of the run yielded data which could be resolved into thrust data, and this data is labeled nozzle 0.0233 in the plots presented in Figure 4. The data collected from these tests allowed for estimates of thermal efficiency to be made. While it was not possible to fit the data to a single line as was the case for helium, the I sp vs. P s plot of Figure 4 show that the data falls between the thermal efficiency curves of 5% and 10%. C. Inferred Performance Data The operational envelope of the MET-100 was limited by the buoyancy effect, and the detrimental effects of the buoyancy force where a direct result of the horizontal orientation of the thruster. Thus it was decided to conduct additional testing of the MET-100 by placing it in a vertical orientation, and fabricating a new nozzle with a 0.0157 throat diameter. The thrust stand could not be modified to collect direct thrust data so all further performance results were inferred from measurements of specific power and an assumed thermal efficiency. The results of this testing is plotted on the I sp vs. P s plots of Figures 3 and 4 and is labeled nozzle 0.0157. The operational envelope of the MET-100 was increased for both propellants. In both cases, the stability of the plasma discharge showed a marked improvement, and in both cases the upper limit of specific power was doubled. Future testing will be directed toward mapping out this operational envelope in greater detail and will include a means of directly measuring the thrust of the vertically mounted MET-100. D. Performance Comparison A search of the available literature provided a nitrous oxide based resistojet system to which the MET-100 could be compared. 5, 6, 7 The references provide a single operating condition for a flight ready system and this value has been plotted on Figure 4. As can be seen, the resistojet has a similar thermal efficiency to that of the MET-100. It should be noted that the specific power utilized for the resistojet was calculated using its specified power level divided by the specified flow rate. 10

5.0 N2O - 0.0189" Nozzle Throat LVDT (V) 4.5 4.0 3.5 3.0 2.5 2.0 1.5 1.0 0.5 0.0 0 100 200 300 400 500 600 700 800 900 1000 Run Time (sec) 4.5 N2O - 0.0189 Nozzle (Increased Flow Rate) 4.0 3.5 3.0 LVDT (V) 2.5 2.0 1.5 1.0 0.5 0.0 0 25 50 75 100 125 150 175 200 225 250 Run Duration (sec) 5.0 N2O - 0.0233" Nozzle Throat LVDT (V) 4.5 4.0 3.5 3.0 2.5 2.0 1.5 1.0 0.5 0.0 0 50 100 150 200 250 300 350 Run Duration (sec) Figure 7: Three plots showing the response of the thrust stand to the thrust developed by the MET- 100 operating on nitrous oxide as a function of run time. The upper plot shows the response to a very unsteady situation no useful performance data could be resolved. The middle plot shows the response as the flow rate was increased. This resulted in an increase in plenum pressure and the plasma began to steady during the middle of the run. The fluctuations at the end of the run preceded the plasma extinction due to excessive plenum pressure. The bottom plot shows that the plasma could be made steady if the flow rate was increased by increasing the nozzle diameter. 11

III. A Brief MET Development History The MET concept has been examined by a number of researchers over the course of the past twenty years. The concept has evolved from an interesting laboratory experiment to a promising thruster which can be scaled to meet a wide range of propulsion requirements. The following paragraphs briefly outline the course of MET development. Note that while an attempt has been made to reference all of the published work it is inevitable that some items may have been omitted. The concept of using a resonant cavity to produce a microwave plasma with the intention of using it as a propulsion device was proposed in 1982 by researches at Michigan State University 8. The subsequent work by this group examined the use of the TM 011 and TM 012 resonant modes while using hydrogen as the working fluid. A limited amount of modeling was also conducted. Most of the experimental work was conducted within quartz 9, 10, 11, 12, 13 vessels at pressures below atmospheric. In 1984, Micci of Penn State began working in parallel on the concept and conducted a number of analytical studies 14, 15. By the late 1980 s, this work had transitioned to an experimental programs which examined both the use of a rectangular waveguide to study the propagation of microwave plasmas 16, 17 and a TM 011 microwave cavity resonator to produce high pressure plasmas using helium and nitrogen 18. A limited amount of spectroscopic analysis of the microwave plasmas was also conducted. 19 A second group at Penn State coupled Maxwell s equations with N-S equations to yield a computational model of the MET thruster. This work greatly aided the experimental program by providing insight in to the various 20, 21, 22 physical processes. By the early 1990 s, the microwave electrothermal thruster had generated some interest within NASA and a research program was undertaken to attempt to scale the experimental TM mode resonant cavity to high power. All the experimental research conducted up to this point had utilized 2.45 GHz magnetrons with maximum power levels of just over 2000 W. The work conducted at NASA-LeRC utilized a high power microwave source that had a maximum output of 30 kw at a frequency of 0.915 GHz. While the experimental design did not lead to high power operation, it did demonstrate that the microwave thruster could be scaled. 23 The experimental work conducted to date and the computational model yielded insight that allowed for the development of a new thruster design. This thruster design retained the TM011 resonant mode but increased the length/diameter ratio which led to increased plasma stability. The new design also removed the need for a quartz pressure vessel and positioned the microwave coupling antenna to be coincident with the central axis of the cavity. These changes led to a large increase in MET performance, and this design concept has remained in use with little 24, 25, 26 modification. After the development of the new thruster configuration, experimental focus switched to making the thruster design more compact. This led to the development of a fixed cavity design which had the microwave magnetron coupled directly to the thruster. 27 All previous experimental configurations had connected the thruster to the power source using some type of microwave waveguide configuration and had left some means whereby the length of the cavity and the insertion depth of the coupling antenna could be adjusted. This work was successful and demonstrated that the MET thruster could be packaged into a compact system. As the development work entered the late 1990 s, the interest in the MET branched out into industry and research was conducted by Physical Sciences Inc., Research Support Instruments, Inc., and the Aerospace Corporation. Much of this experimental work was either funded internally or through the SBIR/STTR program and aside from the water propellant work conducted by Aerospace 1,2 little of the work was published. It was during this period that the MET thruster was scaled to low power at a frequency of 7.5 GHz. This work was pursued separately but in tandem by both Penn State 28 and RSI. It is this background work which has resulted in the thruster design which has been discussed in this paper. A variation of the MET concept was also briefly pursued experimentally and computationally at Princeton University. In this concept two microwave cavities where coupled together in an attempt to add additional microwave power to the flow in the expansion section of the nozzle. 29 IV. Acknowledgments This work was funded under two separate Phase I SBIR awards (DARPA Phase I SBIR SB021-020 and NASA Phase I SBIR E2-05). Funding from the Defense Advanced Research Projects Agency and the NASA Johnson Space Center is greatly appreciated. 12

V. References 1 Diamant, K. D., Brandenburg, J. E., Cohen, R. B., Kline, J. F., Performance measurements of a water fed microwave electrothermal thruster, July 2001, AIAA-2001-3900. 2 Diamant, K. D., Cohen, R. B., Brandenburg, J. E., High power microwave electrothermal thruster performance on water, 38 th AIAA/ASME/SAEASEE Joint Propulsion Conference, 7-10 July, 2002, Indianapolis, IN, AIAA-2002-3662. 3 Haag. T. W., Curran, F., Arcjet starting reliability: a multistart test on hydrogen/nitrogen mixtures, 19 th AIAA/DGLR/JSASS International Electric Propulsion Conference, 11-13 May, 1987, Colorado Springs, CO, AIAA-1987-1061. 4 Haag, T. W., Design of a thrust stand for high power electric propulsion devices, 25 th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 10-12 July, 1989, Monterey, CA, AIAA-1989-2829. 5 Surrey Satellite Technology LTD., Nitrous oxide resistojet, http://www.sstl.co.uk/documents /Subsys_NitrousRJet_HQ.pdf, July 2004. 6 Lawrence, T. J., et. al., Performance testing of a resistojets thruster for small satellite applications, AIAA-1998-3933. 7 Gibbon, D., Baker, A., Development of a 50-100 millinewton level thrusters for low cost small spacecraft, 38 th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 7-10 July, 2002, Indianapolis, IN, AIAA-2002-4150. 8 Morin, T., Chapman, R., Filpus J., Hawley, M., Kerber, R., Asmussen, J., Measurements of the energy distribution and thrust for microwave plasma coupling of electrical energy to hydrogen for propulsion, 16 th AIAA/JSASS/DGLR International Electric Propulsion Conference, 17-19 November, 1982, New Orleans, LA, AIAA-82-1951. 9 Chapman, R., Finzel, M., Hawley, M., Measurements of energy distribution and wall temperature in flowing hydrogen microwave plasma systems, 18 th AIAA/JSASS/DGLR International Electric Propulsion Conference, 1985, AIAA-1985-2052. 10 Frasch, L. L., Griffin, J. M., Asmussen, J., Coupling and eigenfrequencies for microwave electrothermal thruster discharges, 19 th AIAA /DGLR/JSASS International Electric Propulsion Conference, 11-13 May, 1987, Colorado Springs, CO, AIAA-1987-1012. 11 Whitehair, S., Frasch, L. L., Asmussen, J., Experimental performance of a microwave electrothermal thruster with high temperature nozzle materials, 19 th AIAA /DGLR/JSASS International Electric Propulsion Conference, 11-13 May, 1987, Colorado Springs, CO, AIAA-1987-1016. 12 Haraburda, S., Hawley, M., Investigations of microwave plasmas applications in electrothermal thruster systems, 25 th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 10-12 July, 1989, Monterey, CA, AIAA- 1989-2378. 13 Haraburda, S., Hawley, M., Theoretical nozzle performance of a microwave electrothermal thruster using experimental data, 28 th AIAA/SAE/ASME/ASEE Joint Propulsion Conference, 6-8 July, 1992, Nashville, TN, AIAA-1992-3110. 14 Micci, M. M., Prospects for microwave heated propulsion. 20 th AIAA/SAE/ASME Joint Propulsion Conference, 11-13 June, 1984, Cincinnati, OH, AIAA-1984-1390. 15 Durbin, M. R., Micci, M. M., Analysis of propagating microwave heated plasmas in hydrogen, helium and nitrogen, 19 th AIAA /DGLR/JSASS International Electric Propulsion Conference, 11-13 May, 1987, Colorado Springs, CO, AIAA-1987-1013. 16 Mueller, J., Micci, M., Investigation of propagation mechanism and stabilization of a microwave heated plasma, 25 th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 10-12 July, 1989, Monterey, CA, AIAA-1989-2377. 17 Mueller, J., Micci, M., Microwave electrothermal thrusters using waveguide heated plasmas, 21 st AIAA /DGLR/JSASS International Electric Propulsion Conference, 18-20 July, 1990, Orlando, FL, AIAA-1990-2562. 18 Balaam, P., Micci, M., Investigation of free-floating nitrogen and helium plasmas generated in a microwave resonant cavity, 25 th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 10-12 July, 1989, Monterey, CA, AIAA-1989-2380. 19 Balaam, P., Micci, M. M., The stabilization and spectroscopic study of microwave generated resonant cavity plasmas, 21 st AIAA /DGLR/JSASS International Electric Propulsion Conference, 18-20 July, 1990, Orlando, FL, AIAA-1990-2635. 20 Venkateswaran, S., Merkle, C., Micci, M., Analytical Modeling of microwave absorption in a flowing gas, 21 st AIAA Fluid Dynamics, Plasmas Dynamics and Lasers Conference, 18-20 June, 1990, Seattle, WA, AIAA-1990-1611. 13

21 Venkateswaran, S., Merkle, C., Numerical investigation of bluff body stabilized microwave plasmas, 22 nd AIAA Fluid Dynamics, Plasmas Dynamics and Lasers Conference, 24-26 June, 1991, Honolulu, HI, AIAA-1991-1503. 22 Schwer, D. A., Venkateswaran, S., Merkle, C. L., Analysis of microwave-heated rocket engines for space propulsion, 29 th AIAA/SAE/ASME/ASEE Joint Propulsion Conference, 28-30 June, 1993, Monterey, CA, AIAA- 1993-2105. 23 Power, J. L., Sullivan, D. J., Preliminary investigation of high power microwave plasmas for electrothermal thruster use, 29 th AIAA/SAE/ASME/ASEE Joint Propulsion Conference, 28-30 June, 1993, Monterey, CA, AIAA- 1993-2106. 24 Sullivan, D. J., Micci, M. M., Performance testing and exhaust plume characterization of the microwave arcjet thruster, 30 th AIAA/ASME/SAEASEE Joint Propulsion Conference, 27-29 June, 1994, Indianapolis, IN, AIAA- 1994-3127. 25 Sullivan, D. J., Kline, J., Philippe, C., Micci, M. M., Current status of the microwave arcjet thruster, 31 st AIAA/ASME/SAEASEE Joint Propulsion Conference, 10-12 July, 1995, San Diego, CA, AIAA-1995-3065. 26 Sullivan, D. J., Development and performance characterization of a microwave electrothermal thruster prototype, Ph.D. Dissertation, Aerospace Engineering Department, Penn State, Jan. 1996. 27 Kline, J., Thrust measurements of a microwave electrothermal thruster, M.S. Thesis, Aerospace Engineering Department, Penn State, Aug. 1996. 28 Nordling, D., Souliez, F., Micci, M., M., Low-power microwave arcjet testing, AIAA-1998-3499. 29 Chiravalle, V. P., Miles, R. B., Choueiri, E. Y., A non-equilibrium numerical study of a microwave electrothermal thruster, 38 th AIAA/ASME/SAE/ASEE Joint Propulsion Conference, 7-10 July, 2002, Indianapolis, IN, AIAA-2002-3663. 14