LONG-TERM PERFORMANCE OF VACUUM ARC THRUSTERS

Jochen Schein, Andrew Gerhan, Michael Au, Kristi Wilson and Mahadevan Krishnan Alameda Applied Sciences Corporation, San Leandro, CA 94577

Abstract: Vacuum Arc Thrusters (VAT) have been shown to work over a wide range of operating parameters with a constant, overall system efficiency approaching 10%. The combination of vacuum arc thruster heads with an inductive energy storage (IES) power processing unit (PPU) offers system mass and reliability advantages due to the elimination of energy storage capacitors and high voltages. New geometries and a feed mechanism have been developed to utilize these characteristics for high delta-V missions with constant performance.

Introduction Development of nanosatellites is presently a strong interest of the USAF as well as NASA and DARPA [1-3]. Spacecraft designs are tending towards smaller, less expensive vehicles & distributed functionality. NASA's future vision is one of re-programmable/re-configurable, autonomous systems; small, overlapping instruments; and small, inexpensive nanosatellites. Examples include the University Nanosatellite Program and the Orion Formation Experiment. This new trend evokes the same advantages that drive the trend in computing towards distributed, parallel computing and the internet. There are already examples of distributed satellite networks, such as TDRSS, Intelsat, GPS, Iridium, Globalstar and SBIRS (high and low). While these are groups of satellites designed to accomplish a common goal, they are nevertheless 'non-cooperating'. The next wave of USAF constellations will be groups of interacting vehicles that cooperate to achieve mission goals. In such groups, vehicle pointing and positioning will be managed collectively; fleets will evolve over time, extending and enhancing the overall capabilities of the constellation; and self-controlling vehicles will eliminate the need for extensive ground support. From a programmatic perspective, the concept is to replace multi-instrument observatories with low-cost, short lead-time spacecraft that would allow adaptation to changing conditions. This in turn mitigates the risk that not all formation flying applications provide full programmatic benefits. Tomorrow’s Air Force will rely on a new generation of smaller, highly capable nano and pico-satellites

(having masses of 10 and 1 Kg respectively) that will act singly or collaboratively to accomplish various space missions. (M. Birkan, AFOSR 2002 [4]) For these missions new types of micro- and nano-thrusters are required that offer a wide range of impulse bits from nN-s to µN-s at overall system efficiencies of ~10%, with very low (<1 kg) total thruster and PPU mass. Scaling existing electric propulsion engines such as Xe ion engines and Hall thrusters down to ~1-10W power levels is not practical. The unavoidable overhead mass of propellant tankage, 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. Tanks are also required for cold gas and hydrazine systems thus scaling these systems down is difficult due to problems with small valves. Solid propellant systems like the µPPT and the VAT are currently the only candidates that combine low system mass and reasonable performance. However, long-term reliability has to be demonstrated in order for them to become the propulsion systems of choice. In this paper recent developments of new geometries for the VAT will be discussed leading to a development of a new, low-mass feed mechanism that can provide up to 10g of propellant (100N-s) in a single thruster.

VAT system – Principle of operation

The VAT was designed and built based on an inductive energy storage (IES) PPU and simple thruster head geometry. The IES approach allows for throttleable operation with simple TTL control. In the PPU, an inductor is charged through a semiconductor switch. When the switch is opened, a voltage peak L·di/dt is produced, which breaks down the thin metal film coated anode-cathode insulator surface at relatively low voltage levels (≈200V). The typical resistance of this metal film coated insulator surface is 100 Ω - 1kΩ. The porosity of this surface and/or small gaps in the metal film generate micro-plasmas by high electric field breakdown. These micro-plasmas expand into the vacuum and allow current to

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flow directly from the cathode to the anode along a plasma discharge path with lower resistance (10’s of mΩ) than the initial, thin film, surface discharge path. After initiation the plasma pulse is driven by the remaining energy stored in the inductor. Typical currents of ~100 A (for 100-500 µs) are conducted with anode-cathode voltages of 25-30 V. Consequently, most of the magnetic energy stored in the inductor is deposited into the plasma pulse.

Fig. 1: Equivalent circuit of controllable, Inductive

Energy Store (IES) PPU for the VAT.

The performance of the VAT depends very much on the condition of the metallic thin film. During every ignition a fraction of the film is eroded. This erosion has to be compensated for by re-deposition of cathode material from the vacuum arc plasma. Greater arc currents and longer pulses lead to increased re-deposition per pulse. Over long periods of operation re-deposition is strongly influenced by the geometry of the arc source, and geometries have been found which enhance re-deposition.

Two failure mechanisms that have been observed when operating the VAT are due to insufficient re-deposition rates:

1. The resistance of the thin film increases due to a low rate of re-deposition until the required ignition voltage becomes larger than the hold-off voltage of the semiconductor switch.

2. The resistance of the thin film decreases due to a high rate of re-deposition until the leakage current across the resistive path becomes too large.

Failure of early VAT designs occurred primarily via mechanism 2. In order to understand how the geometry of the arc source can influence the resistance of the thin film the process of re-deposition has to be understood.

The VAT relies on expansion of the plasma driven by high arc-spot plasma pressures. The shape of the plasma expansion follows a cosine law. n= k·I/r2·cos∝,

with n being the plasma density, k represents a constant factor of the order 1013 A-1m-1, I the arc current, r the distance to the arc spot and ∝ the angle of expansion. Significant re-deposition is only possible where a certain amount of the plasma can be intercepted. Flush geometries (used in early VAT designs) – in which the anode and the cathode are located in the same plane will provide very little re-deposition. The plasma density is greatest normal to the insulator surface. Re-deposition at large angles is caused mainly by macroparticles- typically emitted from the cathode at large angles. AASC has developed two geometries that allow for increased rates of re-deposition.

Bi-Level Thruster (BLT) While the BLT was introduced and successfully tested earlier, [5] performance tests had shown that the main failure mechanism for this design was the lack of re-deposition of cathode material on the insulator (mechanism 2). Thruster lifetime has been significantly improved by re-designing the BLT geometry. Both the cathode and the anode have been moved forward with respect to the insulator surface to produce a “cavity” with increased surface area for re-deposition. Additionally the insulator is given rounded edges to further increase the surface area and to provide a small “shadow region” on the insulator with no direct line-of-sight to the cathode in an attempt to avoid failure mechanism 1 (figure 2). The BLT geometry also has to be adapted to the average energy in the plasma pulse. Higher energies require thicker insulators and vice versa. As a result of these changes thruster lifetime has increased compared to the lifetime of an original BLT. Over 1,000,000 pulses can now be fired with a BLT thruster without failure. Figure 3 shows the directed ion current output of a re-designed BLT thruster. Note that the output, which is a measure of the thruster performance, does not degrade significantly over the entire performance period.

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Figure 2. Front View of a re-designed Bi-Level Thruster after

a long-term performance test (From top: Holder / Worn Ti-Cathode / Recessed insulator /

Cu-Anode / Insulator)

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Figure 3. A snapshot of the performance test;

Ring Geometry Another new geometry has been developed in which the plasma is produced inside a tube like structure. The first prototype of this thruster was assembled using Ti metal disks. Two and three electrode variations have been successfully tested. The first disk acts as the cathode, which is separated from an anode by an insulating disk coated with a metallic thin film. Optionally additional insulating material can be applied and a third electrode can be added. In the three electrode design the center electrode acts as an “ignition anode” and the main anode would then have to be separated electrically by a ~30Ω resistor (figure 4).

Figure 4: VAT ring geometry

The inner surface of the tube is coated with a conductive graphite thin-film between the cathode and the nearest anode. When the ignition voltage is applied an initial arc is formed between the “ignition anode” and the cathode across the conductive layer inside the tube. In the three electrode design the anode attachment commutates to the main anode driven by the voltage drop across the resistor. By doing this, the plasma is directed more towards the center of the tube and away from the conductive layer. When the plasma is established, metal deposition takes place on the location opposite the arc spot. Although this does not “heal” the damage caused by the initial ignition it produces another ignition spot at a different location on the cathode ring on subsequent firings. Consequently the cathode will get eroded homogenously. Even though the arc spot and thereby the location of the thrust producing plasma changes with every pulse the thrust vector remains constant due to the “ignition anode”/main anode configuration. Control of re-deposition rates is achieved by varying the current, dielectric thickness, and the inner diameter of the thruster.

Feed mechanism AASC has developed a highly effective feed mechanism for the ring geometry. By replacing the cathode ring with a thin walled tube the amount of propellant used has been increased significantly (figure 5 –again with optional “ignition anode’).

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Figure 5: Feed mechanism for new thruster geometry During operation of the thruster the cathode material close to the insulator is eroded. Due to the re-deposition process the preferred cathode attachment moves along the cathode/insulator interface and homogenous erosion takes place. When the part of the tube closest to the insulator is eroded sufficiently the spring pushing on the tube’s back end forces the tube to move forward until it is flush with the insulator surface. Through manipulation of the thickness of the insulator between the anode and cathode the lifetime of the feed mechanism thruster has been extended to more than 4,300,000 pulses. Experiments suggest that increasing the length of the cathode tube would incrementally increase the lifetime of the thruster. While encouraging results have been obtained with the present feed mechanism geometry, it might become cumbersome for long missions where a lot of propellant material will have to be used. Another design could solve this problem: By replacing the tube with a large number of tiny metal balls more appropriate methods of material storage might be

employed. In order to do this a ceramic guide will have to be constructed, leading the replacement balls to the right location, but even this will be possible by using the force of a simple spring. Testing of the feed mechanism The feed mechanism was fabricated with an inner diameter of 6 mm. A picture of the assembled thruster is shown in figure 6.

Figure 6: Assembled lab thruster head with feed mechanism. The final version will feature a smaller thruster body to save mass. After 3 days of operation about half of the tube was eroded as shown in the following figures.

Figure 7: new cathode (10 mm long)

Laboratory Thruster body

anode

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Figure 8: Cathode after one day of operation

Figure 9: cathode after three days of operation

Conclusion

Thruster head geometries have been developed that enable controlled re-deposition of a metallic thin film placed on top of an insulator separating anode and cathode of a vacuum arc thruster, which is necessary for efficient operation. Reliable operation has been shown. A spring-fed feed mechanism has been developed enabling a low mass method to supply larger amounts of propellant without sacrificing performance.

Acknowledgement This work was sponsored by USAF, SBIR Grant F29601-02-C-0016 References:

1. M. Birkan, Formation Flying and Micro-

propulsion Workshop, Lancaster, CA, Oct. 1998. 2. J. Dunning and J. Sankovic, in 35th Joint

Propulsion Conf., Los Angeles, CA, 1999. AIAA-99-2161.

3. R. A. Spores and M. Birkan, in 35th JPC., Los Angeles, CA, 1999. AIAA paper 99-2162.

4. M. Birkan, AAAF propulsion symposium, paper 16-355 (2002).

5. J. Schein et al. RSI 73 (2), 2002