Abstract
Performance measurements were obtained and life-limiting mechanisms were identified on a laboratory-model arcjet thruster designed to operate at a nominal power level of 300 W.
Abstract
The design employed a supersonic-arc-attachment concept and was operated from 200 to 400 W on hydrogen/nitrogen mixtures in ratios simulating fully decomposed hydrazine and ammonia. Power was provided by a breadboard power processor. Performance was found to be a strong function of propellant flow rate. Anode losses were essentially constant for the range of mass flow rates tested. It is believed that the performance is dominated by viscous effects. Significantly improved performance was noted with simulated ammonia operation. At 300 W the specific impulse on simulated ammonia was 410 s with an efficiency of 0.34, while simulated hydrazine provided 370 s specific impulse at an efficiency of 0.27.
INTRODUCTION
In the mid-1980's NASA and industry saw the need for advanced propulsion for large geosynchronous communication satellites (GEO comsats) . At that time, state-oE-art stationkeeping propulsion systems utilized chemical bipropellant thrusters or resistojets, w hich provided a specific impulse of approximately 300 s . Performance of those systems was limi ted by material temperature limits. By utilizing a different heating mechanism to raise the bulk gas temperature to levels exceeding materials limits, arcjets offered the potential of significantly increasing specific impulse and providing a correspondingly large on-board propellant savings.
INTRODUCTION
The fi rs t N AS A low -power arc jet development program was instituted in the 1980's and spurred the development of a commercial 1.8 kW hydrazine arcjet system
INTRODUCTION
• Aerospace Engineer, On-Board Propulsion Branch "Undergraduate, Mechanical Engineering Dept.
INTRODUCTION
with a nominal mISSIOn average (NMA) specific impulse of 502 s.l That system is currently operational on the Telstar 401 spacecraft and baselined on 11 others. The second NASA arcjet program, the Arcjet Technology Development (ATD) Program, sought a step increase of 100 s specific impulse.
INTRODUCTION
To achieve that goal several technology insertions were necessary, including advanced refractory electrode materials and new low-erosion starting techniques.
INTRODUCTION
In a qualifi cation representative test, the ATD program demonstrated over 1000 h of operation at a NMA specific impulse level exceeding 600 s.2 With that successful demonstration, second generation, 600 s-dass technology was transitioned to commercial production and is currently baselined for north-south stationkeeping on a new CEO comsat series. NASA's current arcjet program is targeted at providing advanced propulsion for power-limited satellites. Under the LowPower Arcjet Thruster System (LPATS) program, a 0.5 kW hydrazine arcjet will be developed and demonstrated. The 0.5 kW arcjet is of interest for both primary propulsion for near-Earth science missions and LEO comsat constellations and as auxiliary propulsion for power-limited CEO comsats.
INTRODUCTION
All three of the aforementioned programs have had significant technology assessment portions completed in-house prior to the award of engineering model development contracts. 3-5 NASA's newest focus is on miniaturization of spacecraft to lower cost while maintaining mission goals. On-board propellant represents a major mass fraction of small spacecraft in many missions.
INTRODUCTION
The availability of high performance propulsion systems is mission enhancing and, in many cases, mission enabling.
INTRODUCTION
To support NASA's smallsat program, research leading toward the development of a m iniaturized arcjet system was initiated.
INTRODUCTION
The work is focused on detennining whether arcjets can be scaled to operate reliably below 300 W at performance levels of interest to smallsats. Earlier research demonstrated throttleability from 100 W to 300 W with a subsonic-arc-attachment design; however, voltage instabilities were quite common and may increase the power processor complexity.6 Preliminary research of supersonic-arc-attachment thrusters for the LPATS program demonstrated operation from 300 W to 800 W, but arc instabilities limited lower power operation. 7 A new thruster was designed to operate at nominally 300 W. This paper provides details of the new design and discusses measured performance and observed life issues.
EXPERIMENTAL APPARATUS AND PROCEDURE

Arcjet Thruster
A modular design philosophy was adopted to enable rapid replacement of parts. A cross-sectional schematic of the thruster is provided in Figure 1 . The design was similar to previous 1 kW NASA Lewis 2 laboratory arcjets 3 with the following modifications.
Arcjet Thruster
The anode formerly consisted of two pieces, a W /2 Th02 insert in a molybdenum hOUSing. In the modified design, a single piece of W /2 Th02 was used. The rear-half of the thruster was brazed to insure hermetic integrity . Ca thode / anode electrical isolation was provided by a commercial alumina isolator with tube ends. The ends were connected by modified fluid connectors and sealed using ferrules. Propellant injection was accomplished by radial grooves in the front insulator. Due to space limitations, the helical spring in the 1 kW design was replaced by Belleville washer springs which maintained compression as the parts expanded during operation.
Arcjet Thruster
The nozzle geometry was designed to promote supersonic-are-attachment and is described in Table 1 . Both electrodes were machined from W /2 Th02. The cathode was 1.6 mm in diameter with a 30 0 halfangle conical tip. The arc gap was set by withdrawing the cathode 0.25 mm from contract with the anode. The weight of the thruster was 0.18 kg.
Power Supply
A breadboard power processing unit (PPU) designed to operate the arcjet in the 100-400 W range was used in this study and has been described in detail elsewhere. 8 The PPU operated with an input voltage range of 24-35 V and used a full bridge topology. The overall efficiency was 0.92. A highvoltage pulse of 3 kV provided by the PPU was used to ignite the thruster. Efforts are currently underway to develop a new breadboard power processor with flightrepresentative weight designed specifically for the miniature arcjet.
Facility and Test Equipment
All testing was performed in an oildiffusion pumped facility described previously. 9 The maximum facility pressure was 0.03 Pa (2x10-4 torr) and was measured using ionization gages calibrated for air and uncorrected for sensitivity to propellant gases. The thrust stand used is described elsewhere.10 Thrust calibration was done in-situ by loading and unloading weights before, during, and after test runs. The uncertainty in the thrust measurement was determined to be ±0.5 mN and attributable to thermal drift and hysteresis.
Facility and Test Equipment
All the efficiency values reported herein were calculated by dividing the thrust power by the input electrical power. Current was measured using a Hall-effect sensor with built-in calibration capability.
Facility and Test Equipment
Current signals were sent to a digital oscilloscope, and the voltage was measured using 100:1 probes.
Hydrogen
and nitrogen at room temperature in ratios of 3:1 and 2:1 were used to simulate fully decomposed ammonia and hydrazine, respectively. The flow rate was regulated by thermal-conductivitytype flow controllers with a range of 0-2 SLPM and 0-5 SLPM for nitrogen and hydrogen, respectively. Calibration of the flow controllers was done in-situ by flowing the gas into an evacuated cylinder of known volume for a known time period. Using the initial and final pressures and temperatures of the gas and equation of state, the actual flow rate was determined. The uncertainty of the technique has been determined to be ±2%.
RESULTS AND DISCUSSION

Electrical Characteristics
Figures 2a and 2b provide plots of the voltage / current (V / I) characteristics for both simulated ammonia and hydrazine, respectively, over a range of flow rates. In both cases the voltage was stable at all current levels tested (1.5 A to 2.5 A) for the highest flow rate (12 mg/s). Representative operation in this region is provided by the oscillogram shown in Figure 3 . Voltage instabilities began to appear during operation at 10 mg/s at the lower limit of the current range. The instabilities consisted to two types. These were large random step changes in voltage of up to 30 V, noted during ammonia miranda50@example.org A, and high frequency oscillations superimposed on the normal power supply ripple. The high frequency oscillations also occurred at random intervals and appeared at all points located below the region denoted as stable. With the exception of the 2A current level in the hydrazine test matrix, all operating points for both propellants at the lowest flow rate were unstable. The oscillogram provided in Figure 4 shows typical high frequency instabilities. The frequency of the oscillation was 1.6 MHz and had a sawtooth waveform with the voltage minimum of 115 V and a maximum of 125 V. The characteristics of the oscillations are very similar to those noted subsonic-arcattachment thrusters.
Electrical Characteristics
The "restrike" phenomenon documented in subsonic-arcattachment thrusters is due to movement of the anode attachment point. 11
Electrical Characteristics
Performance Performance data for the simulated ammonia and hydrazine propellants are shown in Tables II and III, respectively. Figures 5a and 5b provide plots of specific impulse and efficiency as functions of specific power for simulated ammonia and Figures 6a and 6b show similar data for hydrazine. The same basic trend was observed for operation on both propellants. As is typical with supersonic-arcattachment thrusters, specific impulse steadily increased with specific power for all flow rates tested;5,7 however, at a given specific power level, the specific impulse decreased rapidly with decreasing flow .
Electrical Characteristics
The same effect is also illustrated in the plots of efficiency versus specific power provided in Figure 5b and 6b. Using the ammonia data at a specific power level of 30 MJ /kg as an example, the efficiency dropped from a high of miranda50@example.org mg/ s to miranda50@example.org mg/s and finally to miranda50@example.org mg/ s. The reason for the effect is discussed further on in this section.
Electrical Characteristics
A comparison of performance between the two propellants is provided in Figure 7 . For all three flow rates, both the efficiency and specific impulse were found to be higher for ammonia. Due to the lower molecular weight of ammonia decomposition products when compared to hydrazine, it was expected that the specific impulse would be higher. What was unusual was the large difference in efficiency at a given flow rate. For example, at 300 Wand at 12 mg/s the specific impulse was 410 s with an efficiency of 0.34 for ammonia, while hydrazine provided 370 miranda50@example.org.
Electrical Characteristics
An effort was made to determine the reason for the large effect of propellant hydrogen fraction and total flow rate on efficiency. Figure 8 details the energy deposition in an arcjet, and a more detailed description of arcjet thermal characteristics can be found elsewhere. 12 Of the energy lost to the anode, Qanode, some is recovered by regenerative heating of the incoming gas, Qregen, and the remainder is radiated away, Qrad· Qregen is simply the mass flow rate times the enthalpy increase of the incoming gas. Hydrogen has a significantly larger heat capacity than nitrogen. With an increased hydrogen fraction, cold incoming simulated ammonia should absorb more heat. An estimate of the amount of heat absorbed by the incoming gas was made using hydrogen and nitrogen thermodynamic properties and assuming both that the nozzle was isothermal and that the incoming gas reached the nozzle temperature. The results, provided in Table IV, indicate that the fraction of energy recovered regeneratively is not a strong function propellant mixture or of flow rate.
Electrical Characteristics
Since the effects are not accounted for by a regenerative effect, the next logical assumption is that the anode losses were both greater for hydrazine than ammonia and increase with decreasing flow rate. The following shows that not to be true. Estimates made for Qrad combined with Qregen allowed Oanode to be determined. The results presented in Table IV show that at a given specific power level, the anode losses were essentially independent of flow rate and were estimated to be approximately 0.19 to 0.23 of the input 4 power. Interestingly, a similar fraction was found to be lost in a larger 1 kW thruster.1 2 In light of the above results, the cause for the efficiency difference between ammonia and hydrazine is unclear; however, it is suspected that the decrease in efficiency with flow rate may be due to increased viscous losses in the nozzle. If viscous effects are a major loss, optimizing the geometry may provide improved performance.
Electrical Characteristics
The testing sequence was such that the ammonia data were obtained, followed by the hydrazine, and then the ammonia data were repeated. Data were taken over the course of several days, and the final performance data matched the initial. Also, the mass flow rate for each test run was held constant, and data were obtained at constant current levels in random order.
Life Issues
Endurance testing of the thruster was beyond the scope of this study; however, anode erosion issues wi th the current design appear life limiting. Post test examination of the nozzle showed severe erosion in the constrictor. Photomicrographs of both ends of the constrictor are provided in Figure 9 . Material was removed from the converging section and deposi ted downstream, probably during start-up. Radial cracking was also noted, and the throat had an irregular shape. Advanced tungsten alloys such as W /4Re/0.35HfC could increase the yield strength and possibly prevent some of the deformation. That material was not used in this study due to fabrication issues, availability, and cost.
Life Issues
Life issues concerning the cathode were also identified. On several occasions following shutdown, the thruster could not be restarted and was found to be electrically shorted. Deformation of the cathode is noted in the post-test photomicrograph shown in Figure 10 . The original cathode had a sharp tip machined, and after operation the tip had receded with a small molted nodule remaining. Molten cathode material was found to bulge out near the tip and is suspected to be the cause of the short. Also, approximately 1 mm upstream of the shoulder, molten spots were noted along the circumference of the cathode. Propellant was injected radially toward the cathode and stagnated at that point. It is thought that at arc ignition, the higher local pressure caused arc attachment in that area.
Life Issues
Design changes need to be implemented to reduce the effects of cathode deformation to increase long-life reliabili ty.
CONCLUDING REMARKS
A laboratory model miniaturized arcjet was designed and tested at power levels from 200-400 Wand had a mass of 0.18 kg. Operating characteristics were obtained on nitrogen/hydrogen mixtures simulating fully decomposed ammonia and hydrazine. Voltage instabilities increased with a decrease in flow rate and consisted mainly of high frequency oscillations; however, large step changes in voltage were also noted at some operating points below 300W.
CONCLUDING REMARKS
Ammonia performance was substantially higher than hydrazine. At the nominal power level of 300 W, ammonia produced a specific impulse of 410 s with an efficiency of 0.34 while similar operation on hydrazine provided 370 s specific impulse at an efficiency of 0.27. The reason for the large difference is not currently understood. For both propellants at a given specific power level the performance degraded dramatically with decreased flow rate. For example, the efficiency for ammonia at a specific power level of 30 MJ /kg dropped from miranda50@example.org mg/s to miranda50@example.org mg/ s. The fraction of energy lost to the anode and the fraction of energy recovered regeneratively were determined to be relatively constant with flow rate. It is believed that viscous losses in the nozzle increase significantly with the decreased flow rates, causing the loss in performance.
CONCLUDING REMARKS
Although performance was repeatable at the beginning and end of the study, posttest examination of the electrodes showed erosion to be severe and probably life-5 limiting. Erosion of the anode may be decreased through the use of advanced refractory alloys and improved starting techniques. Cathode deformation caused shorting of the thruster and needs to be further addressed.
CONCLUDING REMARKS
Steady operation at 300 W has been successfully demonstrated at performance levels Significantly improved over chemical and resistojet systems . Development efforts must now focus on systems issues including mass reduction and long-life operation.
7.
Pencil, E.J., et al., "Subkilowatt Arcjet Development," Proceedings of the 24th International Electric Propulsion Conference, Sept. 1995. to be published. 

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