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Magnetically Enhanced Vacuum Arc Thruster (MVAT)

Michael Au, Jochen Schein*, Andrew Gerhan, Kristi Wilson, Benjamin Tang, Mahadevan Krishnan Alameda Applied Sciences Corporation (AASC), San Leandro, CA 94577

The Magnetically Enhanced Vacuum Arc Thruster (MVAT) is a solid propellant micro-thruster that can produce precisely controlled impulse bits. The MVAT is a variant on an earlier thruster (Vacuum Arc Thruster (VAT)). The VAT is a low mass (~500g), high efficiency (~10% measured) and high specific impulse (Isp ~ 1000-3000s) electric propulsion system. By adapting the technology of the vacuum arc thruster and combining with it an external axial magnetic field on the order of 0.1 Tesla, we have developed a thruster that has been shown to provide up to 50% higher thrust-to-power ratio than a conventional vacuum arc thruster. Also, the applied magnetic field can control the plasma plume produced by the thruster and therefore decrease contamination of adjacent optics and sensors on the spacecraft. Such a thruster would be well suited to provide precision control for small, power limited spacecraft for guidance, navigation and control (GN&C) applications.

I. Introduction Development of nano-satellites is presently a strong interest of the USAF, as well as the NASA, DARPA and

MDA 1-3. Spacecraft designs tend towards smaller, less expensive vehicles with distributed functionality. NASA’s future vision for nano-satellites is that they will be reprogrammable, re-configurable and autonomous with small overlapping instruments on inexpensive platforms. Examples of this vision include the Nano-satellite Program and the Orion Formation Experiment. This new trend evokes the same advantages that drive the trend in computing toward distributed and parallel computing. There are already many examples of distributed satellite networks such as the Tracking and Data Relay Satellite System (TDRSS), International Telecommunications Satellite (Intelsat), Global Position System (GPS), Iridium, Globalstar and Space Based Infrared Systems (SBIRS) (high and low). While these are groups of satellites were designed to accomplish a common goal, they are “non-cooperating.” The new wave of USAF constellations will be groups of vehicles that interact and cooperate to achieve mission goals. In such groups, vehicle pointing and positioning will be managed collectively. Fleets may evolve over time, extending and enhancing the overall capabilities. Autonomous vehicles will eliminate the need for extensive ground support. The concept is to replace expensive multi-instrument observatories with low cost, short lead time spacecrafts that can adapt to changing conditions. In addition, this concept can mitigate the risk posed to the overall mission goal in case of equipment failure on board any single spacecraft.

To help achieve the nano-satellites concept, the next generation of micro- and nano-thrusters must be capable of

a wide range of impulse bits from nN to µN at an overall thrust efficiencies of about 10-20%, while maintaining very low total system mass (<500 g). Scaling existing electric propulsion engines such as Xe ion engines and Hall thrusters down to ~1-10 W power levels is not practical. The unavoidable overhead mass of propellant storage tank, flow controls and plumbing in Xe ion engines and the increasing magnetic field with decreasing size in Hall thrusters makes their overall efficiency unacceptably low at these power levels.

The Pulsed Plasma Thruster (PPT) has been shown to be a good candidate for many missions requiring ~µN-s to mN-s impulse bits4. This type of thruster may be scaled down in current and power. However, due to the low propellant utilization efficiency (due to late time ablation from the TeflonTM insulator) and their requirement for high peak operating voltages (~2 kV), PPTs can only achieve a thrust efficiency of less than 10%. In addition, the use of a bulky high voltage capacitor in the Power Processing Unit (PPU) reduces their thrust/mass ratio considerably as the overall power and thrust are scaled down. AASC’s Vacuum Arc Thruster (VAT), which uses a PPU that employs the inductive energy storage, is a low mass alternative that may overcome these limitations and could become the thruster of choice for many micro- and nano-satellite missions.

* Dr. Schein is currently working at the Lawrence Livermore National Laboratory (LLNL).

40th AIAA/ASME/SAE/ASEE Joint Propulsion Conference and Exhibit11 - 14 July 2004, Fort Lauderdale, Florida

AIAA 2004-3618

Copyright © 2004 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

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By adapting the technology of the Vacuum Arc Thruster5-7 and combining with it an external axial magnetic field on the order of 0.1 Tesla, AASC has developed the Magnetically Enhanced Vacuum Arc Thruster (MVAT)8. It has been shown to provide up to 50% higher thrust-to-power ratio than a conventional vacuum arc thruster. Also, the applied magnetic field can reduce the width of the plasma plume produced by the thruster and therefore decrease contamination of adjacent optics and sensors on the spacecraft. This paper presents the principle of operation, preliminary plume measurement, ion velocity measurement, contamination study and performance of the MVAT. The results indicate that the MVAT is a promising candidate in the micro-propulsion arena.

II. Principle of Operation (MVAT System)

Two approaches were taken when designing and testing the MVAT. The first approach was to use the inductive energy storage (IES) circuit of the PPU to drive both the vacuum arc and the magnetic field. In the PPU, an inductor is charged through a semiconductor switch. When the switch is opened, a voltage peak (LdI/dt) is produced, which breaks down the carbon coated insulator surface at relatively low voltage levels (≈200V). The typical bulk resistance of the coating on the insulator surface is ~1 kΩ. Porosity on this surface and/or small gaps in the coating generates micro-plasmas by high electric field breakdown. These micro-plasmas expand into the vacuum and allow current to flow directly from the cathode to the anode along a lower resistance plasma discharge path (~10’s of mΩ) than the initial thin coating surface discharge path. The current that was flowing in the solid-state switch (for ≤ 1 µs) is fully switched to the vacuum arc load. Typical currents of ~100 A (for ~100-500 µs) are conducted with voltages of ~25-30 V. Consequently, most of the magnetic energy stored in the inductor is deposited into the plasma pulse. The efficiency of the PPU can be greater than 90%. The basic design of the inductive energy storage circuit is shown in Figure 1.

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In order to produce an axial magnetic field for the MVAT, a coil was wrapped around the thruster head as shown

in Figure 2. To provide a better parts integration, the initial design concept was to use the coil as the inductor of the IES PPU. However, this approach turns out to be problematic because the inductance of the coil was only 60 µH,

Figure. 1: Equivalent circuit of an Inductive Energy Store (IES) PPU for the VAT.

Figure 2: MVAT thruster head

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compared to 270 µH used in the original PPU. This limited the amount of energy that could be stored in the inductor and thus limited the output energy in the plasma pulse. A redesign of the inductor and thruster geometry is currently in progress. For bench testing purposes, a second approach was used, which employed the Pulse Forming Network (PFN) circuits. The PFNs were built to deliver higher currents and therefore higher magnetic field. In order to calculate the expected output from the PFNs, a PSpice model was developed as shown in Figure 3.

A 16-stage PFN was used to drive the vacuum arc thruster. Each inductor in the PFN was 0.4µH and each

capacitor was 0.47µF rated for 2000V. A silicon control rectifier (SCR) (here modeled as a closing switch U2) would switch the energy stored in the PFN to the vacuum arc thruster. The VAT model shown here is assembled using a resistor (R6) for the conductive layer. A breakdown simulation circuit was constructed using two voltage controlled switches (S1 and S2), which are triggered at a voltage of 450V. The arc itself is produced by the voltage source V3, which represents the burning voltage and a series resistance of 51mΩ (R4+R5). The resulting current and waveform traces are shown in Figure 4. The simulation shows a current of 600A for duration of about 25µs. The voltage peaks at the ignition voltage and drops off to the burning voltage of the arc. Figure 5 shows the result of the same PSpice simulation but including a serial coil inductance of 60µH. Due to impedance mismatch, the coil significantly changes the response of the PFN circuit. While this behavior is problematic for the PFN circuit, it will not affect an MVAT using the integrated coil IES PPU.

Figure 3: PSpice model of PFN/VAT interaction

Figure 4: Current and voltage traces of PFN driven VAT

Figure 5: Current and voltage trace for PFN driven MVAT

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III. Plasma Plume Measurement Apparatus

Characterization of MVAT coil and thruster of different configurations is a highly data-intensive process and manual collection of the required data would be very time/labor-intensive. Therefore, a MVAT testing system was created, which automated the data acquisition and thruster firing/movement control process with a Labview program. The system utilizes two stepper motor driven slides mounted on the exterior of orthogonal ports of a vacuum chamber. Stainless steel rods are used to feed the motion into the chamber through feed-throughs. The thruster and the probe are each mounted on the rod inside the chamber (see Figure 6). This apparatus allows for scanning through a range of (R,Z) cylindrical coordinates, specifically 25 cm in R and 38 cm in Z, with 250 µm precision. A 1cm2 copper Faraday probe is used to measure the local (R,Z) ion current. The probe is negatively biased to attract positive ions and repel electrons from the neutral ion plume. The probe could be easily changed to optimize spatial resolution or to better accommodate future thruster or coil configurations. Vacuum arc and plasma ion currents are recorded for a series of pulses taken at each (R,Z) location and averaged to reduce noise in the data associated with changes in the location of arc initiation and variations in ion current due to thruster aging. This system allows for easy comparison of different thruster and coil configurations.

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IV. Plasma Plume Measurement Results

Plasma plume scans were performed using a co-axial MVAT with the inner rod serving as the cathode. The end position of the magnetic field coil varied from flush with the VAT head to 2 cm overlapping the VAT head. The operating current was 280 A for 45µs full-width-half-maximum (FWHM) as driven by the PFN circuit. The magnetic field coil has 100 turns and the starting position of the scan was set to a distance of Z=1cm away from the Faraday probe. Figure 7 shows a photo of the visible plasma plume without magnetic field confinement, while Figure 8 shows the plume with the magnetic field applied. Note in Figure 8 that the visible plume is confined in the R direction and extended in the Z direction compared to the unconfined plume shown in Figure 7. The following measurement data will confirm this observed phenomenon.

Figure 6: Test Chamber Setup

Figure 7: Photo of the visible plasma plume without confining magnetic field

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4 cm4 cm

The plasma plume measurement results are presented as contour maps in Figures 9-13. Contour lines encircle areas of equal plasma plume intensities. Each data point represents the average of four complete scans, with 10 pulses acquired per (R,Z) location. Figure 9 shows the plasma plume intensity (Iion-peak/Iarc-peak) of a conventional VAT without magnetic field. Figures 10-13 show the plasma plume intensity of the MVAT operated at 0.25 Tesla with varying magnetic coil end position (from flush to 2 cm overlap of the VAT head).

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Figure 9: Plume distribution contour plot (Iion/Iarc) without magnetic field

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Figure 10: Plume distribution contour plot (Iion/Iarc) with 0.25T magnetic field(Magnetic coil flush with VAT head)

Figure 11: Plume distribution contour plot (Iion/Iarc) with 0.25T magnetic field(Magnetic coil overlap VAT head 0.5cm)

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Note that with increasing coil overlap, the confinement of the plasma plume increases; the plume becomes narrower in the R direction and longer in the Z direction. While it appears that for moderate overlaps (0.5 cm and 1 cm) the magnetic lens effect is positive. Further extension of this overlap (2 cm) reduces the amount of plasma and the length of the plume. This reduction is likely due to some of the plasma being intercepted by the coil and was not detected by the Faraday probe. Traces of cathode material deposited on the inside of the coil in this mode of operation attest to this assumption.

The difference in the plasma plume shape between the flush assembly and the moderate overlap lies in the shape

of the magnetic field, which is stronger and more effective inside the coil as shown in this finite element model of Figure 14. For simplicity, the model assumes the region surrounding the inductor was free space; the presence of a thruster head was not considered in this analysis. The model was created assuming an axis-symmetric magnetic

Figure 12: Plume distribution contour plot (Iion/Iarc) with 0.25T magnetic field(Magnetic coil overlap VAT head 1.0 cm)

Figure 13: Plume distribution contour plot (Iion/Iarc) with 0.25T magnetic field(Magnetic coil overlap VAT head 2 cm)

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field and the field strength was set to zero at a distance of 1 m. A relatively modest current of 10 A was used, assuming the inductance of 122 µH. The current density was assumed to be constant at 67 kA/m2. To check the predicted field strength, we used the equation of magnetic field around a coil, B = µοnI, where B is the field strength inside the coil, µo is the permeability of free space, n is the turn-to-length ratio and I is the current. The prediction came to be about 1 T, which is higher than what the model predicted (0.28 T) but expected due to the idealization of the input parameters.

V. Ion Velocity (Time of Flight) Measurement Results

Ion velocity (time of flight) was measured between two Faraday probes. An automated analysis methods were developed, which rely on slope detection method to detect the onset of ion current. Passive filtering was used to reduce high frequency noise and PFN circuit was used to improve the signal-to-noise ratio. Numerous measurements were taken to obtain statistically meaningful data sets.

For the VAT with a Ti cathode, the average ion velocity without an applied magnetic field was measured to be

22,000 m/s. With the magnetic field applied, the average ion velocity increased by about 30% to 29,000 m/s. Further experiments are needed to find a proper explanation for this phenomenon. This is an exciting result as it shows that the specific impulse (Isp) can be influenced by the use of magnetic field to the values that so far only ion thrusters have reached. Note that unlike the ion thruster, the VAT produces quasi-neutral plasma and thus does not require a neutralizer. Measured erosion rate for the titanium cathode was of the order 100µg/C during the experiments. Due to the small amount of mass ablation, no difference could be found between operation with and without magnetic field.

VI. Contamination Test Results

The contamination was measured using witness plates. Witness plates were located 5 cm away from the center of a co-axial VAT and MVAT. Three witness plates were used for each test. One was placed directly in front of the thruster. The others were located at an 55o angle on either side of the first plate, as shown in Figure 15. 100,000 pulses were fired with and without the magnetic field at an average peak current of 50A. The resulting deposition can be seen in Figure 16. This photo demonstrates the effect of the magnetic field. While the total amount of material remains fairly constant the deposition is concentrated within a smaller angle, which will reduce contamination of the spacecraft as the thruster can be directed away from sensitive optics.

Figure 14: Simulated Magnetic Field Strength Map

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VII. Thrust Measurements Results Thrust measurements were performed at the JPL µ-thrust stand (GFE). The integrated IES MVAT as well as the

PFN driven MVAT were tested. As mentioned before, the thrust output of the IES MVAT was very small and measurements were extremely noisy. Therefore a decision was made to operate with the PFN circuit described earlier. Measurements were taken with and without magnetic field. To ensure that the generated B-field did not have an influence on the thrust stand, the magnetic field was operated alone and no movement of the thrust stand arm could be detected.

All the thruster electrical connections were fed from outside the vacuum chamber. This introduced about 4 m of

additional cable for each connector, which is equivalent to ≈5 µH of inductance and 0.5Ω of resistance, per cable pair. Current and voltage into the coil and the thruster were measured, which allows for a direct thrust/power measurement. The current and voltage into the system vary significantly due to the mismatch that occurs by adding an inductor, as shown in Figures 17 and 18.

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Figure 15: Contamination Measurement Setup Schematic

Figure 16: Witness Plates Deposition with (upper row) and without (lower row) magnetic field

Figure 17: Measured current (blue) and voltage (red) into the PFN VAT system

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The current pulse becomes wider and the peak current decreases when a coil is added to the circuit. The traces

were multiplied and the area under the resulting curve was computed to obtain a value for the energy into the system. This ratio between the measured impulse and the computed energy was taken and is shown in Figure 19 for the case with (trial #96-115) and without (trial #116-124) magnetic field. The magnetic field during the pulse reaches about 0.1T.

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Trials # 105, 106 and 116 have been performed without the proper cable connection. Therefore, a value of zero is shown. Trials #100 and 110 have shown a discrepancy between coil current and arc current, most likely due to a discharge between the magnetic coil and the thruster. Therefore, these data points have be discounted from the analysis. The remaining data points shows that the thrust-to-power ratio for the MVAT is about 50% higher than the conventional VAT. Base on this result, we are optimistic that a MVAT with an integrated PPU coil design will be able to outperform its VAT predecessors.

VIII. Conclusion

The MVAT has been shown to provide up to 50% higher thrust-to-power ratio than a conventional vacuum arc thruster. The applied magnetic field can reduce the width of the plasma plume produced by the thruster and therefore decrease the contamination of adjacent optics and sensors on the spacecraft. The applied magnetic field has also been shown to increase the ion velocity of the VAT, which is an equivalent to an increase in specific impulse. The next generation of MVAT will incorporate an integrated PPU coil design, which can better utilize the otherwise wasted energy in the inductor coil.

Figure 18: Measured current (blue) and voltage (red) into the PFN MVAT system

Figure 19: Thrust-to-power for PFN VAT and PFN MVAT

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Acknowledgement

This work was sponsored by NASA contract NNC04CA19C. The author would like to thank Dr. John Ziemer of the Jet Propulsion Laboratory (JPL) for his generous support on this project.

References 1M Birkan, Formation Flying and Micro-propulsion Workshop, Lancaster, CA. October 1998. 2J. Dunning and J. Sankovic, 35th JPC, Los Angeles, CA. 1999, AIAA-99-2161. 3R.A. Spores and M. Birkan, 35th JPC, Los Angeles, CA 1999, AIAA-99-2162. 4M. Birkan, AAAF Propulsion Symposium, 2002, 16-355. 5J.Schein, A. Anders, R. Binder, M. Krishnan, J. E. Polk, N. Qi and J. Ziemer “Inductive Energy Storage Driven Vacuum Arc Thruster ” Review of Scientific Instruments, vol. 72, no 3. Feb. 2002. 6Schein, J., Krishnan, M., Ziemer, J., Polk, J “Adding a "Throttle" to a Clustered Vacuum Arc Thruster” AIAA paper 2002-5716, Nanotech, 2002. 7Schein, J., Gerhan, A., Rysanek, F., Krishnan, M. “Vacuum Arc Thruster for Cubesat Propulsion”, IEPC-0276, 28th IEPC, 2003. 8M. Keidar and J. Schein, 40th JPC, Fort Lauderdale, FL, 2004, AIAA 2004-4117.