Development and Characterization of the Heated-Anode Cathode Arc Thruster (HA-CAT) by George Lewis Teel

Transcription

1 Development and Characterization of the Heated-Anode Cathode Arc Thruster (HA-CAT) by George Lewis Teel B.S. in Mechanical Engineering, May 2012, The George Washington University A Thesis submitted to The Faculty of The School of Engineering and Applied Science of The George Washington University in partial satisfaction of the requirements for the degree of Master of Science May 18, 2014 Thesis directed by Michael Keidar Professor of Engineering and Applied Science

2 c Copyright 2014 by George Lewis Teel All rights reserved ii

3 Dedication I wish to dedicate this work to my loving and caring family, whom have always been there to support me through all my efforts and eccentric ideas. To my mother Gay and my father John have been incredibly supportive, allowing each of their three children to pursue any goals and careers that we desired. I can t thank them enough for raising me and my two siblings. To my older brother Glenn who has been a huge influence on a majority of hobbies and interests. Growing up with him and being able to join him on some adventures has helped shape who I am today. To my older sister Adelaide who has always looked out for and taken care of me through my ups and downs of school. She has always been there for me and has continued to watch out for me. Without the love from my family, nothing would be possible. The care, the joy, the creativity and humor has always kept me positive allowing me to achieve my goals. iii

4 Acknowledgments I would like to acknowledge Professor Michael Keidar for all of his hard work, his fantastic ideas, and all of his help through my undergraduate and graduate experience. Supporting me through my academic career financially, being there anytime I had a question or concern, I could not have had a better adviser. I would like to acknowledge Alexey Shashurin for all of his help with setting up experiments, providing me with the fundamental knowledge (through multiple iterations), and for being supportive through my undergraduate and graduate experience. Sharing great stories, intellectual discussions and life experiences over lunch always made my day. I would like to acknowledge the amazing work of Zhuang TaiSen, for his incredible work, laying out the path that allowed me to continue on my research. For being inviting and always welcoming into the lab, he was always helpful and was willing to answer any and all of my questions. I would like to acknowledge the whole MPNL group that has been there for me, helped me through all my research problems and supported my ideas. Cheng XiaoQian, Dereck Chiu, Fang XiuQi, Joseph Lukas, Laura Rose and David Scott, you all have been a wonderful research group and helped me succeed through personal and professional developments. I couldn t ask for a better group. I would like to acknowledge the George Washington University Machine Shop Staff, William Rutkowski, Tom Punte, Curtis Stoltzfus, Nicholas Batista and Ryan Mossbarger. Throughout the years of my college experience and through research, I have learned the required skills to work and manufacture the necessary pieces of my research, had great cookouts and many fond memories. All thanks to you guys. I would like to acknowledge the George Washington University and the many Alumni that have donated and helped provided for my educational goals. iv

5 Abstract Development And Characterization Of The Heated-Anode Vacuum Arc Thruster (HA-VAT) A modern approach to satellite based experimentation has evolved from large, multiinstrumented satellites, to cheaper, smaller, almost disposable yet still reliable small spacecrafts. These small satellites are either sent to the International Space Station (ISS) to be dropped out into low earth orbit (LEO), or dropped off as a secondary payload into various orbits. While it is cheap to have small spacecraft accomplishing these missions, the lifetime expectancy is very short. Currently there are no commercialized propulsion systems that exist to keep them flying for prolonged periods of time. Recently researched at the Micro Propulsion and Nanotechnology Lab (MPNL), at the George Washington University (GWU), have been developments of a variety of Vacuum Arc Thrusters (VAT s) dubbed Micro-Cathode Vacuum Arc Thrusters (µcats). µcat s provide an inert electric means of propulsion for small spacecraft. The issue with these µcats has been their efficiency levels and low amounts of thrust that they provide. The µcats can provide µn levels of thrust per pulse. While being proficient for small spacecrafts, an increase in thrust is highly sought for, but the improvements must retain a small footprint and low power consumption. The topic of this thesis is the development and characterization of a new type of µcat. The interest in this new design has been conceptualized based on experiments for plasma coating techniques. By utilizing the physics of evaporation, which has been used to decrease macroparticles (MP s) for thin film deposition, it has been theorized to also be applied to VAT technology. The concept is to increase levels of thrust with the µcat, and provide higher levels of efficiency. This effect can be created without many additional components nor multiple additional loads to the thruster subsystem. v

6 Development of this new mechanic for thruster technology has been investigated through a variety of tests for fundamental proofs of concept. Running in two operations modes, the Heated-Anode Cathode Arc Thruster (HA-CAT), has undergone current efficiency tests, mass measurements, and cross examination through the use of a Scanning Electron Microscope (SEM) and Atomic Force Microscope (AFM). This research hopes to explore an old territory of plasma engineering for future developments with µcat and VAT technology. vi

7 Table of Contents Dedication iii Acknowledgments iv Abstract v List of Figures ix List of Tables xii List of Acronyms List of Symbols xiii xiv Glossary of Terms xv Chapter 1-Introduction Chapter 2-Literature Review Small Spacecraft and Electric Propulsion Mono-propellant Thrusters Electrospray and MEM Thrusters Resistojets Pulse Plasma Thrusters Chapter 3-Vacuum Arc History Physics Cathode Region Cathode Spots Vacuum Arc Thrusters µcats Plasma Heating and Reflection Motivation for Research Chapter 4-HA-CAT The HA-CAT Design Materials : Boron Nitride Materials : Tungsten Materials : Copper Thermocouple Propulsion Equations The HA-CAT Circuit vii

8 4.3 Heating Element Heating Element Equations Chapter 5-Experiments Chamber Scanning Electron Microscope and Energy-dispersive X-ray Spectroscopy Atomic Force Microscope Mass Measurements Mass Experimental Setup Total Ion Current Test Current Experimental Setup Chapter 6-Results Calculations Specific Impulse Ionization Rate Mass Measurement Results Total Ion Current Results Without Heating Element With Heated Element Scanning Electron Microscope Analysis Without Heating Element With Heating Element AFM Analysis Without Heating Element With Heating Element Chapter 7-Conclusion SEM, EDS, and AFM Analysis Efficiency Increase Calculations Future Work References Appendix A Appendix B viii

9 List of Figures 1 Monopropellant Thruster 2D Cross-section Commercialized Monopropellant Thrusters for Cubesats Principle Design of an Electrospray Thruster Micro-machined layering for systems MIT s MEM Thruster Resistojet 2D Cross-section Pulsed Plasma Thruster 2D Cross-section Air Force Research Laboratory s PPT Conventional Vacuum Arc Ion Source System Triggerless Arc Ignition System Cathode to Plasma Breakdown Cathode Spot Formation under Magnetic Field Geometries of µcat s Plasma filtering through magnetic coils Hot Anode Vacuum Arc Metallic Cones for Hypersonic Studies Metallic Cones under Plasma HA-CAT Working Principles D Models of the HA-CAT HA-CAT Testing Thrust to Power Ratio vs ISP µcat Circuit Parts and Diagram HA-CAT s Heating Element ix

10 24 Scanning Electron Microscope Atomic Force Microscopy Machine Mass Measurement Experimental Setup Ion current collection experimental setup with circuit T/P HA-CAT vs µcat Comparison Cold Mode Ion Current Results Cold Mode Ion Current Results Cold Mode Ion Current Results Thermocouple Reading Hot Mode Ion Current Results Hot Mode Ion Current Results Hot Mode Ion Current Results SEM Scan of Cold Anode SEM Scan of Cold Anode SEM Scan of Cold Anode SEM Scan of Cold Anode SEM Scan of Hot Anode SEM Scan of Hot Anode AFM Scan Cold Anode AFM Scan Cold Anode AFM Scan Cold Anode AFM Scan Hot Anode AFM Scan Hot Anode AFM Scan Hot Anode Chamber in MpNL at GWU x

11 49 Laser Cut Part : HA-CAT Holder Laser Cut Part : Welding Wire Holder Laser Cut Parts : Anode Braces xi

12 List of Tables 1 Various satellite classes based on weight ranges Power Required Atomic Weight Ratios of Thick Copper Deposition on Cold Anode 50 xii

13 List of Acronyms AC Alternating Current AFM Atomic Force Microscope AFRL Air Force Research Laboratory DC Direct Current EDS Energy-dispersive X-ray Spectroscopy EP Electric Propulsion GWU The George Washington University HA-CAT Heated Anode Cathode Arc Thruster HAVA Hot Anode Vacuum Arc IES Inductive Energy Storage ISS International Space Station ISP Specific Impulse LEO Low Earth Orbit µcat Micro-Cathode Arc Thruster MEM Microelectromechanical MIT Massachusetts Institute of Technology MP Macroparticle MPNL Micro Propulsion and Nanotechnology Lab NASA National Aeronautics and Space Administration PPT Pulsed Plasma Thruster PPU Power Processing Unit PTFE Polytetrafluoroethyliene SEM Scanning Electron Microscope STRaND-1 Surrey Training, Research and Nanosatellite Demonstrator 1 VARIAC Variable Auto Transformer VAT Vacuum Arc Thruster xiii

14 List of Symbols 1 A Ampere 2 C Celsius 3 cm Centimeter 4 K Kelvin 5 m Meter 6 µ Micro 7 mm Millimeter 8 n Nano 9 N Newton 10 Ω Resistance 11 V Volt 12 W Watt xiv

15 Glossary of Terms CubeSat : Cube Satellite. A nanosatellite with unit size 10 cm by 10 cm by 10cm. Delta-V : Delta-V ( -V) is the change in velocity a propulsion system can achieve in a moment of time. U : A U is a Cube Satellite s Unit of measurement. 1 U is the 10 cm by 10 cm standard. 2 U is 20 cm by 10 cm by 10 cm and stands for 2 units stacked on top of one another. 3 U is 30 cm by 10 cm by 10 cm and is generally 3 U s stacked. The total length is capped at 3 U s for ISS launch, but larger U cubesats exist. An example is a more common size of 6 U s. This configuration is 2, 3 U s side by side making it a 2 x 3 U satellite. xv

16 Chapter 1-Introduction In 2008 at the George Washington Universitys Micro-Propulsion and Nanotechnology Lab, Dr. Michael Keidar and a Doctoral student Zhuang TaiSen started research on micro propulsion for small satellites. Continuing on previous concepts from Dr. Keidar, such as his research at the University of Michigan, Zhuang TaiSen created what he coined as the µcat. The µcat being a type of VAT was developed and analyzed as a propulsion device targeting Cube Satellites (CubeSats), which are in the picosatellite weight classification. Four years were spent developing both the hardware and the thruster technology. A variety of designs were implemented on both the hardware and thruster head to miniaturize and mature them to a point of resembling flight subsystem quality. Turning from simple lab circuitry to a standalone power processing unit (PPU), the important information of the µcat had been characterized through a variety of experiments which proved the thruster s feasibility. Continuing on this research, in early 2013, a new type of CAT has been developed and studied. The objective of this thesis is to describe and characterize the recently developed small satellite propulsion system known as the Heated Anode Cathode Arc Thruster, or simply HA-CAT. This satellite propulsion system was designed utilizing a heated anode mechanism, which has been adapted from a technological variation on conventional cathodic arc jets used for plasma coating. By applying this method of a heated anode, a reflection of plasma can happen due to evaporation, a secondary force of thrust can be produced and a significant reduction in exhaust velocity. This not only provides a thrust increase, but a higher efficiency as well. A construction of a proof of concept has been developed with test apparatuses to gain analytical data. Propulsion equations for thrust have been calculated from this information, and compared to previous designs using the conventional cold anode design. 1

17 This thesis is divided up into various chapters, and their summaries are provided below: Chapter 2 is a literature review of the various types of space propulsion for small micro-satellites. This literature review is a discussion on which systems have been flown, what technology they derive from, and the benefits and problems of each system. With all the variations of propulsion subsystems, what these systems are designed for and why they are inviting to missions is important. This allows a comparison of the diverse systems to the HA-CAT currently under development. Chapter 3 is a literature review specifically about vacuum arc technology. This is the defining mechanism behind the HA-CAT. The history of vacuum arcs, the physics behind them, the triggering mechanisms, and applications are all defined. A review on MPNL s µcat is also discussed which provides background information on how the HA-CAT was conceptualized. Chapter 4 is the chapter on the HA-CAT itself. This chapter supplies information on the setups, the construction, the materials, and circuits which are required to run the HA-CAT. Chapter 5 is a review on the experiments, the experimental setups, the processes, and methods used to obtain the results gathered. Chapter 6 holds the results of the experiments which were described in chapter 5, as well as a discussion about the results obtained. Chapter 7 is a summary of the thesis, and concluding remarks with future ideas and improvements to the design. 2

18 Chapter 2-Literature Review The objective of this chapter is to provide a literature review of current electric propulsion being used and developed for small satellites. This aims to provide the reader an insight to the variety of subsystems. This also will provide a background comparison for the current system being developed and the benefits it can provide to small spacecrafts. 2.1 Small Spacecraft and Electric Propulsion In recent years, small spacecraft have overtaken the satellite community. The old age of large satellites costing millions of dollars is slowing down. Now with the power of modern day fabrication techniques, satellite technology can be miniaturized greatly which opens new doors to the scientific community. Nanosatellites and picosatellites are on the rise and the developed CubeSats are becoming increasingly popular. CubeSats had been developed as a collaborative effort between California Polytechnic State University, San Luis Obispo, and Stanford University s Space System s Development Lab[12]. With this modernization and standardization of a small satellite platform has allowed universities, companies, and space enthusiasts all over the world to explore space related research. Satellite missions have even been started on kickstarter, a website allowing project start-ups and companies to take shape and gain funding[25, 40, 31]. Satellites fall into particular weight categories and are shown in Table 1. Cube- Sats fall under the nanosatellite range. Nanosatellites have mainly been developed to explore new technologies and experimental techniques to space related research. Government agencies and universities alike can utilize these platforms to develop flight hardware which can be tested in space cheaply. National Aeronautics and Space Administration (NASA) Ames Research Center have used cellphones as an on board computer system, and dubbed the project name PhoneSat [16]. With the 3

19 current miniaturization of electronic parts and the availability of new technology, small satellites have been allowed to increase two-fold. Projections from Spaceworks, an aerospace engineering consulting and analysis firm, had estimated 93 nanosatellite launches in the year 2013 [13]. Their projections were spot on as 92 satellites actually launched that year, giving Spacework s projections credibility. This was a 269% increase from 2012 and due to emergent and continued growth, there should be an increase in the number of satellites that are expected to be launched in upcoming years[10]. Satellite Class Femtosatellite Picosatellite Nanosatellite Microsatellite Small Satellite Mass Range g < 1 kg 1-10 kg kg kg Table 1: Various satellite classes based on weight ranges Using these small satellites for experiments has its drawbacks though. Currently there are no standards for types of propulsion for such a small package. This leads to small satellites lifetime expectancy being very short. For example, the three cubesats flown by NASA Ames, the PhoneSats had lifetimes of only one week each after being expelled in Low Earth Orbit (LEO)[16]. In general, the propulsion used for large satellites cannot be shrunk down due to their complicated systems. Currently Electric Propulsion (EP) has become the answer to small spacecrafts for their propulsion needs. EP can run without much effort from the satellite bus, and normally is allowed to directly tap into the on board batteries due to low power requirements. Because of weight and manufacturing complications, not all EP can been shrunk down to nanosatellite specifications and miniaturized. Another issue arises is due to small satellites not being the primary payload on a rocket launch. These problems stem from the various propulsion systems causing 4

20 concerns that arise for the primary payloads and their investors. Due to worries of non-inert gases, combustion and pressure vessels all generally lead to small satellite missions not being flown. Canceled flights have been a common occurrence for the small satellite community, which has caused a focus into propulsion research. These solutions are being looked into by various agencies such as NASA, other cooperate companies such as BUSEK and Aerojet Rocketdyne, and universities all around the world. This wide assortment of developed EP is explained below Mono-propellant Thrusters Mono-propellant thrusters are simply small rocket nozzles which utilize chemical propellant to provide high levels of thrust. Shown in Figure 1 below, the simple design of a mono-propellant thruster is shown. The propellant is fed through a fuel valve which controls the fuel flow rate. The propellant flows through the feed tube into the catalytic layer which converts the propellant to a gas. Then, the gas is ejected from the back of the nozzle to generate thrust. Fuel Valve Injector Catalytic Layer Exhaust Figure 1: Monopropellant Thruster 2D Cross-section Various forms of mono-propellant thrusters exist and have already been flown in larger conventional satellites. Different propellants have been used and the most common propellants are Hydrazine and Hydrogen Peroxide. Both offer a large delta-v with power requirements ranging from 8 to 2 Watts (W), which is an acceptable power range for small satellites. BUSEK and Aerojet 5

21 (a) BUSEK s Thruster (b) Aerojet Rocketdyne s MR-140 Figure 2: Commercialized Monopropellant Thrusters for Cubesats Rocketdyne, both aerospace engineering firms, are working with this technology to miniaturize its use for cubesat standards. An image of both BUSEK s and Aerojet Rocketdyne s thrusters can be seen in Figure 2[35, 34]. These monopropellent rocket nozzles use chemical fuel, pressurized systems, and are combustible. These are the general concerns for rocket launches and primary payload owners. The risks have outweighed the benefits with cubesats and, as of now, none have flown. Perhaps in the future, when cubesats become the primary payloads, these thrusters will get a chance to prove themselves Electrospray and MEM Thrusters Electrospray thrusters are becoming a common researched type of propulsion, as they can manipulate a fluid electromagnetically and provide large forces of thrust with little required energy. Electrosprays work by applying an electric field outside a capillary containing an ionized fluid, with the electrodes extracting the ionized fluid to propel it at high velocities. The working mechanism can be seen in Figure 3. The complex propulsion subsystems require pressurized tanks, series of pressured capillaries carrying the fluids, and an array of valves. BUSEK has been developing a prototype and is stated to provide 0.7 mn levels of thrust with a 9 W power consumption, and an overall system footprint being less 6

22 Taylor Cone Fluid Spray Capillary Electrodes Figure 3: Principle Design of an Electrospray Thruster than 0.5 of a U[11]. Space propulsion is looking for low mass, small volume, and running on a relatively low power budget such as these. Microelectromechanical (MEM) thrusters are under the same concept of Electrosprays, and utilize ionized liquid through micro-machined capillaries to provide thrust through electromagnetic forces. The propellant storage systems can be micro-machined in several thin layers, which can be bonded together to form the various tubes, propellant tanks, and the nozzles required and this technique is shown in Figure 4[17]. Figure 4: Micro-machined layering for systems Massachusetts Institute of Technology (MIT) has been researching MEM thrusters, and has shrunk the thruster system to the size of a quarter which is shown in Figure 5b [26]. These micro-fabricated electrospray devices are 7

23 a densely packed array of individual emitters, allowing more force per square area and shown in Figure 5a. The modules are able to deliver thrusts densities of roughly N/m 2 [26]. Looking into nano and picosatellite range, micro-valves have been created by Meuller et al that use only 0.5 W of energy. This low power consumption is highly appealing to the nano and picosatellite community[28]. (a) Micro-machined Capillaries (b) MIT s MEM Thruster Figure 5: MIT s MEM Thruster The lifetime of these thrusters do not compare to Pulsed Plasma Thrusters (PPT s) or VAT s, which are described in sections below. Shorter mission designs can capitalize on the high thrust levels, but for long missions, the electrosprays and MEM thrusters fall short. MEM thrusters also have higher costs than VAT s and PPT s due to complicated fabrication techniques. This is consequence to expensive production costs, which drive up the subsystem costs. These expenses can cause problems for starting companies or universities looking for cheap propulsion. 8

24 2.1.3 Resistojets Resistojets are simple EP chemical hybrid devices. Shown in Figure 6 one can see that the resistojet consists of a small chamber with an inlet for propellant and a conventional nozzel for exhaust. A heating element of AC or DC power is located in the small chamber which heats the propellant as it enters. Through the nozzle the expelled propellant provides thrust due to thermal and pressure changes. The heating element can increase or reduce the gas flow rate which can vary the specific impulse (ISP). Resistojets can also be used without the heating element but would have a significant decrease in ISP. Fuel Exhaust Heating Element Figure 6: Resistojet 2D Cross-section Current resistojets generally run off of 100 W of power, and consist of an array of pressurized vessels to contain the working fluid, the nozzles to control the flow, and electrical circuits to control the subsystem. Various resistojet fluids can be used, such as: water, nitrous oxide, water/methanol, nitrogen and helium. The most successful and frequently used propellant has been decomposed hydrazine[27]. All these propellants have been looked at as fuel and characterized for small satellites by Lawrence et al[23]. Using water is desirable for a propellant type, especially if these small spacecraft must enter the vicinity of the ISS. Safety is a huge concern when potential hazards could 9

25 cost billions of dollars of permanent damage if something went afoul. Resistojets rely on pressurized vessels, series of mechanical valves, and potentially combustible fuel which all could cause issues on launch. Currently, the only resistojet to date that has launched was a waterbased resistojet that flew on the STRaND-1 cubesat mission. STRaND-1 was a cubesat built in the United Kingdom by the Surrey University s Surrey Space Centre, and is currently still flying. Using PPT s for stationkeeping, the resistojet is dedicated to STRaND-1 s maneuvers[32] Pulse Plasma Thrusters PPT s are generally considered the simplest form of electric propulsion. PPT s are low-power, high-specific impulse, electric thrusters which generally utilize a polytetrafluoroethyliene (PTFE) propellant. Usually Teflon is the chosen PTFE since Teflon is a space grade PTFE with historical use in most space missions. Shown in Figure 7 the reader can see the simplistic operation of the PPT. Ignitor Cathode PPU Teflon Plume Feeding Spring Anode Figure 7: Pulsed Plasma Thruster 2D Cross-section Two electrodes are separated by its propellant, in this case, a Teflon block. An ignition trigger is built near the two plates and discharges an electric spark 10

26 between the two electrodes. Plasma is generated as the Teflon is thermally heated and ablated. Since the ionized teflon is charged, the plasma completes the circuit between the two plates and the current arcs through it. While the two electrodes are charged, Lorentz forces act on the plasma causing it to be ejected at high velocities. Lorentz forces are described in Appendix A. The frequency between the arc pulses is dependent on how long it takes to charge the two electrode plates. Generally, frequency is almost continuous to generate a constant thrust force. A successful development was the µppt developed by the Air Force Research Laboratory (AFRL), which uses a cylindrical design shown in Figure 8 below [21]. As the arc is being generated between electrodes, the Teflon is being eroded away, similar to degradation of a cigar. As mentioned previously, PPT s have flown on the UK based STRaND-1 cubesat[32]. With their low power, high efficiency, and long lifetime, PPT s can be used for a variety of missions. PPT s are a competitor to the VAT s, but have contamination complications. With their plasma plume not fully being ionized, back flux is known to happen and satellites can be coated by the PTFE propellants. Figure 8: Air Force Research Laboratory s PPT 11

27 Chapter 3-Vacuum Arc This chapter will provide a literature review on vacuum arc technology, which includes the history, the physics behind the working mechanism, and the developments with electric propulsion. The various geometries and designs are reviewed to provide a background for the reader of further development. 3.1 History Vacuum arc technology has been developed, researched, and utilized for decades. Back in the early nineteen hundreds, field emissions were accepted as the primary mechanisms for vacuum arc discharges due to the research by W.D. Coolidge[8]. From then, the study into field emissions and the arc discharge continued by Dyke et al. lead to the discovery of joule heating for plasma generation. The heating and melting of the metallic cathode tips were the first studies to research the breakdown of vacuum arcs. The cause of discharge was attributed to the anode breaking down and forming a metal vapor, known as plasma[14, 3]. The electric discharge in a vacuum was then studied by Boyle et al, which lead to the breakdown currents and metal vapors being further characterized[9]. After these initial experiments for characterization, a variety of experiments emerged in order to understand plasma generation within the vacuum arc. A variety of theories and experiments have been explored to describe the phenomenon, but no true clear answer has emerged yet. To dive deeper into this topic would be out of the scope of this thesis, and more can be read through the various sources listed in the references Physics A vacuum arc is an electric discharge, which is the passage of electrical current through a medium that is normally insulating between two electrodes[8]. This discharge is an electric potential drop, 2-3 times the ionization potential, of 12

28 the atoms of the cathode material[19]. There are two main types of vacuum arc discharge, one with a background gas as a medium and one without. A variety of conventional ignition methods exists which are used to ignite arcs: including drawn arc, fuse wire arc, triggered arc, an ignition by laser, and hollow cathode arc[8]. Figure 9: Conventional Vacuum Arc Ion Source System The triggering mechanism the HA-CAT makes use of is a more modern and unique form than the conventional triggering methods. HA-CAT s means of triggering is the triggerless arc. The triggerless arc was designed by Anders et al for plasma film depositions. The principle mechanism behind the triggerless arc initiation is the use of a conductive path between the anode and cathode, such as a metallic coating on a ceramic surface. This provides a means of joule heating of the coating-cathode interface which results in rapid plasma formation once the arc voltage is applied to the electrodes[2]. Shown in Figure 9, a conventional electrode arrangement of a vacuum arc shows the complicated circuitry that is involved. This, however, is a non-ideal apparatus for spacecraft propulsion. Developed by Anders et al in Figure 10, a triggerless 13

29 vacuum arc is displayed which shows the spot initiation along a simplified cathode insulator boundary[2]. Figure 10: Triggerless Arc Ignition System This not only simplifies the overall vacuum arc circuit to a simple triggering device, but it reduces size and complication of the entire vacuum arc system. This method can only continuously work if the plasma re-coats the ceramic boundary between the electrodes. It is known that during the arc discharge a breakdown of the ceramic layer will destroy the insulator itself, ultimately leading to failure of the vacuum arc. Thus, it is an absolute requirement that the plasma rebuilds upon the insulator to redeposit the thin film coating Cathode Region The cathode region is an area on the cathode where cathode spots, explosive plasma generation locales, are formed. The plasma jet that is created, consists of electrons and ions with velocities of hypersonic speeds[19]. Shown in Figure 11 is a diagram describing the cathode region and its various breakdown zones. These zones include the pre-sheath, sheath, plasma zone and plasma flow zone. As cathode spots are formed through thermionic emissions and electron field 14

30 emissions, a release of electrons and ions into the sheath transpire. One can attribute the sheath as an electric field area where the space charges are large enough that the plasma is not considered quasi-neutral[19]. In the sheath, ions and electrons are free to return or be expelled outwards. As the plasma is traveling at hypersonic speeds, it passes through the presheath briefly, which is a region of a weak electric field separating the sheath from the quasi-neutral zone[39]. Once out of the pre-sheath, the plasma then is in the quasi-neutral plasma zone, and can further be characterized later. Plasma Generation Zone Hydrodynamic Flow e e- Atom Emission e- A e- i e- e- e- i i i Electron Emission A Ion Return Electron Return i e- A e- e- A i e- Cathode Sheath Pre-Sheath e- Plasma e- i e- e- i e- e- i e- i i i e- e- e- e- i i e- e- e- i e- i e- i e- e- i Figure 11: Cathode to Plasma Breakdown Two theories of plasma generation are present to explain the defining mechanism: an explosive model written by Mesyats et al [1], and the kineticvaporization model by Beilis et al[5, 4]. A well described model by Keidar, is a combination of the two theories. Starting with the explosive model written by Mesyats, the cathode spots re- 15

31 lease plasma due to explosive tips on the cathode surface. This explosion happens with Ohmic overheating due to high-current densities at the surface irregularities[19]. Once plasma is formed after arc initiation, the kinetic theory written by Beilis is then suggested. Dense plasma is generated by atom ionization near the cathode surface causing kinetic plasma to flow and expand into a plasma jet[19]. Figure 11 also encompasses Keidar s proposed concept. The sheath and pre-sheath are where the plasma bombardments happen. Past the plasma creation zone, the hydrodynamic plasma can then be modeled Cathode Spots An important topic to discuss is the cathode spot formation of the vacuum arc. Being the source of plasma generation during the arc discharge, a localized hot spot generates an explosion of ions and electrons. This explosion forms a nano-sized crater known as a cathode spot. Being the most common studied phenomenon of vacuum arcs, cathode spots still are not fully understood experimentally or theoretically, and remain a controversial topic in the vacuum arc world[8]. These cathode spots are produced in a random, yet organized manner. Shown to sporadically form against the physical norms of what we know today, the cathode generation is usually not at a constant point. Similar to lighting never striking the ground twice, the cathode spots spontaneously spawn at new locations during each pulse. These cathode spots form in a zigzag manner along a diagonal path and erupt at new unheated localized zones. Shown in the Figures 12a and 12b below, the cathode spot formation under a magnetic field can be seen following along diagonal lines. One can notice the cathode spots erupt close to one another, but never at the same spot. 16

32 (a) Cathode Spot Formation (b) Diagonal Cathode Spot Formation Figure 12: Cathode Spot Formation under Magnetic Field These spots undergo another unique phenomenon in that the cathode spots are known to be effected by a retrograde JxB motion under a magnetic field, which means they move in the opposite direction than expected[18]. This phenomenon has been capitalized with they cylindrical µcat design, though ignored for the HA-CAT. The HA-CAT employs the flat plate VAT configuration with no magnetic coil applied. Perhaps future work will include electromagnetic forces. 3.2 Vacuum Arc Thrusters The vacuum arc s natural function creates high velocity, highly ionized plasma through evaporation of the cathode element from discharge. This mechanism is a perfect tool to be harnessed for space propulsion applications[30]. A variety of research has been explored utilizing vacuum arc thrusters such as University of Michigan, University of Illinois, and GWU. What began with large capacitors in a power processing unit (PPU), work with inductors created the inductive energy storage (IES) unit which reduced mass and increased thruster efficiency[33]. Continued work with VAT technology, a magnetic coil was used and the Mag- 17

33 netically enhanced Vacuum Arc Thruster was conceived. Using a magnetic coil, one can control the plasma plume exhaust to reduce contamination, forcing cathode spot rotation leading to uniform erosion and increase efficiency [22]. Recently an increase in development of VAT s world wide has taken place, and multiple papers for developing and characterizing VAT s has happen from South Africa, Japan, and Munich Germany[24, 15, 29]. This shows an increase in interest with vacuum arc technology as papers have been published during the International Electric Propulsion Conference this past year in µcats Micro-cathode arc thrusters are a type of MVAT designed at the GWU s MPNL. Created by Zhuang TaiSen in early 2008, the µcat has been characterized over the past four years. The µcat was designed for nanosatellite class satellite operation, and utilizes a low power PPU. The µcat is the base design of the HA-CAT, and forms the basis for this research. Similar to PPT s, the µcat is based on two electrodes being separated by a ceramic insulator. The difference between PPT s and VAT s is that PPT s use the insulator as its propellant, while the VAT s use one of the electrodes as its propellant, in thµcat s case the cathode. The µcat s propellant has historically been titanium, which has been characterized and is a well known metal for cathodic arcs. The µcat s ignitor electrode is the anode and is composed of copper. The two electrodes are separated by a high temperature nonporous ceramic insulator with a conductive coating. As the PPU charges its capacitor, the arc discharges across the ceramic insulator. Due to thermal heating, the cathode ionizes and plasma is formed. An evolution of the µcat has happened over the past four years with a few design changes in hopes to increase the thruster efficiency. The original design was the ringed electrode configuration and can be seen in Figure 13b below. The anode 18

34 and cathode are the same diameters, forming two cylinders that are stacked on top of each other. The two electrodes being separated by an insulator form a hollow cylinder. As the arc crosses the insulator from the anode, the cathode is eroded and plasma is formed in the hollow center. Due to the natural velocity of the plasma from the cathode spots, and with the help of the magnetic coil, plasma is expelled in a directed manner. This design has the benefit of the plasma re-coating the insulator as it passes by, but suffers losses as it travels farther distances before being expelled. As the cathode erodes away, a spring acts as a feeding mechanism to replenish the cathode for continued use. Cathode Insulator Anode (a) Concentric Electrode Geometry Cathode Insulator Anode (b) Ringed Electrode Geometry Anode Insulator Cathode (c) Flat Plate Geometry Figure 13: Geometries of µcat s To increase efficiency, a new design was conceptualized. The layout was named the concentric electrode configuration and is shown in Figure 13a above. In the concentric electrode design, the anode is in the center of the thruster head. A ceramic insulator being the same length as the anode, separates the two electrodes. The cathode being the largest radius, is outside the ceramic cylinder. Following the same working mechanism, plasma is generated on the cathode and naturally expelled 19

35 at high velocities. The plasma is formed on the outer surface of the thruster head to guarantee more plasma is being expelled than the previous design. This increase in plasma expulsion is improved, but re-coating the ceramic insulator has not been looked into and may be a potential problem. The two µcat designs formed the basis for this thesis research. The driving mechanisms, the plasma generation methods with the triggerless arc, and the experimentation led to this development. The cylindrical design was not used for the HA-CAT however, but the basic flat plate geometry shown above in Figure 13c. Simply two flat plates wedged together by steel supports. Using a boron nitride plate as the insulator. A design overview is discussed in the upcoming chapter. 3.4 Plasma Heating and Reflection Typical conventional cathodic arc jets are developed and used to produce thin films and coatings. These metal vapor plasma jets have high density and adhesive properties. They do produce, however, Macro Particles (MPs) and deposition to the anode surface[7]. Previous methods of filtering these MPs have been through the use of magnetic coils which form a 90 degree turn as shown in Figure 14, or establishing different vacuum arc operation modes[20]. Figure 14: Plasma filtering through magnetic coils 20

36 Beilius et al, discovered with evaporation occurring over an anodes surface, one can eliminate the production of MPs. This heated anode was called the Heated Anode Vacuum Arc (HAVA) and is shown in Figure 15. The HAVA heated the anode surface to the threshold temperature of the cathode s evaporation temperature. This forces plasma to then re-evaporate after being deposited on the anode[7]. Figure 15: Hot Anode Vacuum Arc It has been further characterized that the coated anode, when sufficiently heated, generates a dense plasma plume which reduces MPs and evaporates the MPs themselves[36]. This type of vacuum arc has been previously used as a MP-free plasma source in various applications, but primarily used for thin film deposition[6]. Further research of plasma evaporation studies have been developed by Shashurin et al. The study involved looking into the plasma layer along the boundaries of a hypersonic vehicle[37]. The experiments consisted of generating hypersonic plasma directed at two cones; one cone being made of molybdenum foil while the other cone being made of solid aluminum. Shown in Figure 16 below, the two cones are shown post experiment[37]. The molybdenum cone heated up to sufficient temperatures which caused the re-evaporation of the copper plasma and is noticed to the left in 21

37 Figure 16 while the aluminum cone has been covered in copper and seen in the right in Figure 16[37]. Figure 16: Metallic Cones for Hypersonic Studies In Figure 17 below, the two cones during hypersonic studies are displayed[37]. The reader can notice that the molybdenum foil cone in Figure 17a is noticeably hotter than the aluminum cone, as it reached temperatures above 1000 C. (a) Molybdenum Cone under Plasma (b) Aluminum Cone under Plasma Figure 17: Metallic Cones under Plasma 3.5 Motivation for Research The motivation for this research topic is to increase the thrust of micro thrusters for small satellites. After starting research at the MPNL, and introduction to the µcat 22

38 technology, a well grounded paved road was developed for this research. Learning about previous plasma experiments, interesting ways were conceptualized for a variety of techniques which lead to reflecting plasma. Trying to take this reflective plasma technique from a similar, yet different field, a hope to achieve a new, unique, more efficient thruster design is the goal. This design is similar in power costs, similar in footprint size, and more effective overall by recycling the naturally created byproducts. 23

39 Chapter 4-HA-CAT The conception of the HA-CAT design is to utilize plasmas natural ability of evaporation with the plasma generation of the µcat. With the µcat ignition system generating the quasi-neutral plasma, the initial force of thrust in µn s is generated. While plasma flows in all directions being an ionized gas, the plasma can return to its material solid state since ions are attracted to the nearby surfaces. This creates a coating effect which can be detrimental. While heating the anode with this simple geometry, the plasma will generate a reflecting force as its pushed away due to the high temperatures. This force balance in space would generate a secondary propulsive force to the overall subsystem. 4.1 The HA-CAT Design The HA-CAT utilizes a completely new geometry from the previously developed µcats. The overlaying design is a series of flat plates of boron nitride stacked on top of each other separating the different required layers. The simple design is shown below in Figures 18a. Starting with a heating cavity, the heating element is positioned behind the anode electrode. The anode is separated by another layer of insulating material, which then sits the cathode on the outer surface. If the HA-CAT is running in cold mode, a buildup of copper particles forms along the anode surface. The working principle behind the HA-CAT during hot mode is that the heated anode will evaporate the plasma plume or copper deposition. This will create an additional force of thrust and will increase overall efficiency of the cathode utilization. The working principle can be seen in Figure 18b. The experimental setup consists of various layers in order to replace fractured materials, if required, for consistent operation. Starting with the base layer, a layer of boron nitride keeps the temperatures thermally isolated as much as possible from 24

40 A C A C Copper Deposition (a) HA-CAT s Cold Mode (b) HA-CAT s Hot Mode Figure 18: HA-CAT Working Principles the back end of the HA-CAT. This base layer acts as a holder for the heating element of tungsten wire with dimensions of 25 mm x 30 mm. The next layer of boron nitride has a cutout volume of 5 mm by 4 mm, with a depth of 4 mm. This cavity allows a small surface area for the tungsten foil anode which is discussed more in the radiation section, with the overall dimensions sharing a 25 mm x 30 mm. This cavity also provides a separation between the heating element and the anode in order for the pulses to not be affected by the variable auto transformer (VARIAC). This layer also has two grooves cut into it to allow for a thermocouple to be placed on top of the anode, to provide temperature readings. The tungsten foil anode is then layered next. The anode is sandwiched between a thin plate of Boron Nitride with a small square cutout. This cutout has the conductive coating and is where the arc between the anode and cathode happens. Last is the cathode, which is a small piece of copper that is cut to fit the face plate, and forms an interface with the conductive coating. The whole thruster is wedged between thin steel cutouts to allow the holder to be placed at variable heights. Design documentation can be found in Appendix B in Figures 49, 50 and 51. Figure 19a is a 3D representation of the HA-CAT setup. One can see three layers of boron nitride separating the heating element and tungsten anode. A thin layer 25

41 Cathode (a) 3D Isometric View Insulator Boron Nitride Insulating Layers Anode Heating Element (b) 3D Cutout View Figure 19: 3D Models of the HA-CAT of boron nitride acts as an insulator between the copper cathode and the tungsten anode. A fourth layer of boron nitride is used as support when clamping the device to its holder, and a place for the thermocouple to rest on top of the anode s surface. The colors are the copper cathode, the light blue tungsten anode, and dark gray tungsten wire heating element. The light gray colored pieces are the boron nitride layers. The heating cavity and heating element can be seen in the 3D cutout view in Figure 19b. A photo of the HA-CAT in chamber is shown in Figure 20 with the heating element on during testing. Figure 20: HA-CAT Testing 26

42 4.1.1 Materials : Boron Nitride Boron Nitride is a crystalline chemical compound which consists of equal parts of boron and nitrogen. With boron nitrides incredibly high melting temperature of 3200 K, it makes a perfect candidate for aerospace applications. Boron nitride has a history of applications with hall effect thrusters being used for insulators to the interior walls. Boron nitride has been used as the foundation to the HA-CAT Materials : Tungsten Tungsten, also known as wolfram, is a high density rare earth metal that has a variety of applications in high temperature electrical applications such as light bulbs, vacuum tube filaments, heating elements, and rocket engine nozzles [38]. With its high melting temperature of almost 3700 K, it makes a perfect candidate for both the heating element and anode. The HA-CAT uses a thin sheet of tungsten for the anode, and a small diameter wire for the heating element which is described below Materials : Copper Copper is on the opposite spectrum of tungsten, being very soft and malleable. With low melting temperatures of 1100 C, copper is a great choice as an anode. Due to the elements natural high conductivity, copper has been historically used in electrical components and various thermal applications Thermocouple The thermocouple used in these experiments are a type K 24 Gauge solid lead wire thermocouple. With fiberglass insulation, the thermocouple was placed on the surface of the anode outside the heating chamber to provide proper results. 27

43 Given temperature in mv s, the thermocouple can reach temperatures up to 1370 C. 45 mv s is the proper required oscilloscope readout for the HA-CAT s experiments and was monitored on the anode surface Propulsion Equations The important propulsion parameters stems from Equation 1 which is the force of Thrust generated in a general propulsion system: T = ṁ u e. (1) With T being force of thrust, ṁ being mass of the propellant per unit time being expelled at velocity u e. Having thrust defined opens up a new series of equations. Starting with specific impulse, which can be related to the efficiency of our propulsion system, we get Equation 2: I sp = I t g t+δt 0 t ṁdt (2) The I sp, or specific impulse, is equal to the impulse moment I t, which is described in Equation 3, divided by the gravitational acceleration g 0 by the integral of the mass of the propellant per unit time ṁ over time t to t + δt. The Impulse moment is defined below: I t = t+δt t T dt (3) Impulse Moment I t being equal to the time integral of thrust force T over time t to t + δt. A bodies impulse is the change in the bodies momentum over time. 28

44 With this information I sp can be described as Equation 4: I sp = T ṁg 0 = u e g 0 (4) Which is Thrust T over mass of propellant expelled per unit time ṁ by gravitational acceleration g 0, which is also equal to velocity of the propellent u e over the gravitational acceleration g 0. This is an important relation as it relates the efficiency based on thrust and acceleration of the propellants exhaust. By utilizing this equation, one can measure the velocity of the exit ions to calculate the thrust and efficiency of the thruster. Shown in Equation 5 below, one can calculate the velocity of the exhaust based on thermal temperature which would be the reflected plasmas velocity: u T e = KB T anode m a (5) The thermal velocity u T e is equal to the square root of the Boltzmann s constant K B multiplied by the temperature of the anode T anode, over the molecular mass of the anode material m a. Another necessary equation is total power of a propulsion system which is Equation 6: P T = 1 2 T u2 = 1 2ṁu2 e = 1 T 2 2 ṁ (6) The total power P T is equal to one half of the thrust T by velocity squared u 2, which is equal to one half the mass of propellant per unit time ṁ by velocity of exhaust squared u 2 e, which is also equal to one half by Thrust squared T 2 over mass of propellant per unit time ṁ. Using the this total power equation, one can achieve an equation of efficiency for an EP system which is in Equation 29

45 7: η T = P T P in (7) with the thruster efficiency η T equal to total power P T over the power into the system P in. Therefore, we can arrive at the relation of thrust over power being Equation 8: T P = 2η T g 0 I sp (8) which is the relation of Thrust T over Power P being equal to two times the efficiency η T over the gravity acceleration g 0 by specific impulse I sp. Equation 8 provides an important ratio which can be seen in Figure 21. One can see that the thrust per power decreases significantly as the specific impulse increases, which shows that the higher efficient a propulsion system is the less thrust it can obtain. This creates a trade off for EP, and is the reason why a variety of systems have been developed. The µcat family lies in the ISP range of two to three thousand which provides µn levels of thrust. T T T T T Figure 21: Thrust to Power Ratio vs ISP A concept the HA-CAT is trying to increase the thrust per pulse of the µcat family. One can accomplish this by keeping efficiency the same as the 30

46 previous µcat thrusters and reduce the velocity of the propellant. Following along in Equation 4, reducing the velocity would result in a decrease of I sp. Reducing the I sp would lead to increase the thrust to power ratio shown in Equation The HA-CAT Circuit The previously developed concept of pulsed vacuum arc discharge requires a triggered source of high enough current to induce the joule heating at the cathode-insulator border[22]. In Figure 22a below, the PPU is shown with a voltage source of 25 V. This leads to a capacitor which acts a voltage regulator, and a discharge unit. Passing through the inductor, when the insulated gate bipolar transistor allows current to flow through to the thruster head, a high current wave is generated and the arc discharges. L HA-CAT Anode V C Switch Insulators Cathode AC (a) PPU Circuit Diagram (b) Single Channel µcat Subsystem on the VARIAC Figure 22: µcat Circuit Parts and Diagram In Figure 22b the thruster circuit has been developed by the MPNL at GWU, which is a full 1 channel thruster subsystem. This circuit encompassing voltage regulation, pulse generation, and the PPU. This circuit is being run with 2 sources of 25 V s, and connected to the HA-CAT thruster head into the chamber through Bayonet NeillConcelman (BNC) connectors. 31

47 4.3 Heating Element The heating element was fabricated using tungsten wire of 0.02 inch in diameter and is shown below in Figure 23. Wrapped around a small screw to form the coiled section, the length of the coiled section is roughly inches (5 mm) in length, with an outer diameter of inches (3.5 mm) and an inner diameter of 0.1 inches (2.5 mm). Protruding from the back of the HA-CAT, the heating element is directly connected to the VARIAC. Figure 23: HA-CAT s Heating Element Heating Element Equations The necessary power required to heat the tungsten anode by radiation is given in equation 9 below: P = A (ɛ) (σ)t 4. (9) P is the power required, A is the area of the heated element, ɛ is the emissivity constant of tungsten, σ is the Stefan-Boltzmann constant which is 5.67E 8, and T being temperature. Because we need the anode to reach the melting temperature of copper, we use 1100 C, which means we set our T as 1373 K. Shown in table 2 below, one can see the power required P, for a variety of areas A, with the area being in units of mm 2. 32

48 T = 2000K A (mm 2 ) P (W) Table 2: Power Required Over an area from 1 mm to 6 mm squared, the necessary powers required have been calculated. For the HA-CAT, we have an heated cavity of roughly 5mm 5mm. This leaves us with a power requirement of 0.9 Watts. Utilizing the units of a Watt which is A 2 Ω with A being amperes and Ω being resistance, the required current can be calculated. The required 0.9 Watts with the heating element resistance of roughly 0.2 Ω gives the required current as 2.21 A. Using ohms law shown in Appendix A as Equation 21, the required voltage can be calculated with the resistance of the wire used, and current calculated previously from the power used. Theoretically, this gives us approximately a voltage of 0.2 V. 33

49 Chapter 5-Experiments The various experiments took place in the MPNL research facilities. All experiments took place in the chamber described below, and most of the data was collected and analyzed in the MPNL Cleanroom through the various machines. 5.1 Chamber The vacuum chamber is the conventional cylindrical shaped chamber, and is 45 cm in diameter and roughly 64 cm in length. All experimentation ran at pressures of 10 4 or 10 5 Torr. Utilizing a roughing pump, into a diffusion pump, allows for pressures this low to be reached. A photo can be seen in Appendix B in Figure Scanning Electron Microscope and Energy-dispersive X-ray Spectroscopy Utilizing a Scanning Electron Microscope (SEM) with an Energy-dispersive X-ray Spectroscopy (EDS), a distribution of what materials present on a sample can be found. The SEM scans objects by beaming them with a series of electrons. While the electrons interact with an objects surface, the excited electrons reach the X-ray spectrum and data can be collected by the signals given off. Shown in Figure 24, the SEM consists of an electron beam which thermally emits electrons from a tungsten cathode. This tungsten cathode is commonly known as an electron gun. Under vacuum the beam of electrons from the cathode is sent through multiple lenses towards the subject. These lenses focus the electron beam, which then scatters the surface of the object under examination. The electrons scattered can be scanned by X-ray spectrometers, and these signals are converted to voltage readings to allow analysis of different materials. 34

50 Electrode Lenses Magnetic Coils X-Ray Detector 2nd Electron Detector Sample Figure 24: Scanning Electron Microscope The SEM with EDS has been used to scan the anode surfaces. In Cold mode operation, copper particles should be present along the surface as the plasma coats the surface. In Hot modes, copper should not be noticeable as the plasma should be reflected. 5.3 Atomic Force Microscope Using an Atomic Force Microscopy machine, or AFM, high resolution photos of the plasma deposition can be found. An AFM is a type of scanning probe microscopy technique which can scan variations on a surface macroscopically. Shown in Figure 25 the AFM consists of a cantilevered beam with a tip probe at the end of the beam. This tip probe is nanometers in size, and this cantilevered contraption is usually made of silicon or silicon nitride. According to Hooke s law, which is Equation 22 in Appendix A. If a force is provided to this cantilevered beam through the microscope, and the spring constant k is known for its material, a displacement can be measured through the use of its deflection. Shown in Figure 25, one can see that a laser is being used with the use of photo diodes to measure the 35

51 deflection of the cantilevered beam when in contact with a surface. Photodiode Laser Cantilever Sample Figure 25: Atomic Force Microscopy Machine An AFM can be operated in a variety of modes, for which two are the most common, contact and non-contact. Contact mode has the probe dragged across the surface and can be thought of as more of a brute force method. Through bending due to variations on a surface, and the AFM can measure the deflection and compare this deflection with the feedback signal, which attempts to keep the cantilever at a constant height. This method is reliable, but can cause scraping to a surface which may not be desired for some applications. The other common method is non-contact mode where the probe utilizes a tapping method. This tapping method works with an oscillating motion by a piezoelectric element mounted at the tip holder. This works well for fluids or biological scans as it doesn t damage the surface. If not calibrated right, this method can cause errors. By tapping, one reduces the possibility of breaking the tip off the probe during scanning, as contact mode may get stuck on a surface element if large enough. Since there are two modes of operation for the HA-CAT, the experiments need to be altered slightly. For cold operation, thermal tape can be applied to the surface of the anode. By covering a quarter of the anode surface, after a series of pulses, the thermal tape can be removed. This provides a differential between the clean material and coated material, and can be analyzed at specified points with the AFM. 36

52 In cold operation mode, we should see more mass deposited on the anode. Due to the geometry of the thruster head the plasma will coat without reflecting outside to the glass collector. In hot operation mode, we should have the exact opposite effect. In theory, no plasma should be on the surface of the anode. Since being heated to its melting temperature, the copper plasma should be completely evaporated. Most of the plasma should be collected on the glass slide outside and the height can be measured in an AFM machine. In hot operation mode, the thermal tape cannot be used on the anode, so a variety of points have been arbitrarily located around the arc location to gain a range of heights. 5.4 Mass Measurements A vital piece of information for characterizing the HA-CAT is obtaining the mass measurements to calculate thruster efficiency and thrust. For mass measurements, a metallic cup was used to capture the released plasma plume. As the plasma is discharged from the cathode, parts of the plasma plume is sent both forwards and backwards from the cathode spot. When running both the heated and unheated modes the variation on mass gain should be noticeable. While the heating element is on, plasma is re-evaporated from the anode surface and reflected back in the desired direction and more mass will be gained from the collector. Working with equations below, which are erosion rate calculations, one can calculate the mass expectancy if obtaining 100% of the ablated material. E r = m = 3 13 µg t I arc C (10) In Equation 10 above, the erosion rate E r is equal to the change in mass m over the change in time t by arc current I arc, which for copper is equal to 3 to 13 37

53 micrograms per coulomb µg C [8]. This relation is a constant value for a material which can be beneficial for calculations. Since the Erosion rate wont change, parameters can be set to collect the data that is required. In this case, the mass eroded is the necessary variable to be calculated. Through the experiments, a t of roughly 3600 seconds was the operating time of the thruster head. This t needs to be multiplied by the duty cycle, shown in Equation 11, of the system since the thruster was pulsing and not constantly on. Duty Cycle = Pulse Duration Total Period = 190µS 1.545S (11) The Duty Cycle is calculated by the pulse duration of 190 µs over a total period of S. This gives a duty cycle of 1.23E4. Multiplying the t by the duty cycle gives a t of S. Multiply by the arc current I arc which was running at 60 A, the Charge Q equals C. The last step is to multiply the erosion rate of 10 µg by C to calculate an erosion of µg which is 2.65E 4 g milligrams is a small amount, and the reader must be aware that this is 100 percent erosion collection. Due to the HA-CAT s geometry, plasma is free to escape in all directions which can cause uncertainties in mass measurements Mass Experimental Setup The thruster was attached to vertical steel holder, cut out from the laser cutter at the GWU Machine shop. This holder provides two holes for the heating element to pass through the holder. The metal holder itself is attached to the center of and acrylic base plate which separate any form of connections from the chamber walls. This helps provide a floating thruster head with no grounding issues. A metallic cup was attached to an L beam holder in front of the HA-CAT exhaust. 38

54 Shown in Figure 26 is the equivalent circuit with thruster PPU and VARIAC A C PPU Figure 26: Mass Measurement Experimental Setup VARIAC setup. The VARIAC and heating element are connected through 10 gauge welding wire. Due to the small resistance of the heating element, 0.2 Ω, the connections needed to be a lower resistance. In a circuit, objects with higher resistance heat up, which is difficult with such a small resistance heating element, which is why the welding wire has been used. 5.5 Total Ion Current Test The total ion current with the arc current ratio is an important parameter to have to calculate the efficiency of a thruster. With the form in Equation 12 below, the relation of a percentage of efficiency is given. η ion η arc = efficiency out efficiency in = Total efficiency (12) Typical VAT s utilizing titanium cathodes have total efficiency of around 8%[8]. With the magnetic coil the µcat family of thrusters were able to reach an efficiency of roughly 3.5% [41]. 39

55 5.5.1 Current Experimental Setup The setup is almost the same as the mass measurement experimental setup. A large, half spherical metallic ion collector is attached to the base facing the HA-CAT and partially encompassing the whole setup. This ion collector was negatively bias to -75 volts to draw in the ions, since ions are positively charged. Measurements were taken over an 9.4 Ω resistor. The circuit of the HA-CAT Ion collection apparatus is shown in Figure 27 with the VARIAC included. VARIAC A C PPU Oscilloscope Figure 27: Ion current collection experimental setup with circuit Through the thruster line in from PPU, a current transistor probe is used to measure the arc current going through the thruster. The apparatus was then placed inside the chamber, which was pumped down to torr. 40

56 Chapter 6-Results This chapter provides an overview of all the results collected from the experiments previously described in Chapter 5. Analysis and discussion can be found in their respective sections. 6.1 Calculations Specific Impulse A very important result that has been achieved is a massive ISP reduction. Going back to Equation 4 in Chapter 4 is the equation for ISP: I sp = u e g 0 (13) For the exit velocity of u e, using the thermal velocity which is Equation 5: u T e = KB T anode m a (14) With the Boltzmann Constant being equal to , the temperature of the anode being the temperature reached during the experiment which was 1375 K, and the molecular mass of copper, ( ) which is coppers atomic mass times Avagadros number, the thermal velocity comes out to m. s Plugging this velocity into Equation 4 above: I sp = u T e = g = s (15) This is a major reduction in ISP, considering previous µcat s are in the ISP range 2000 to We can look back at the chart in Chapter 4 and 41

57 compare with the previous curves. Picking an efficiency of 0.2, since VAT s reach these efficiencies, we can compare the Thrust to Power ratios in Figure 28: T Figure 28: T/P HA-CAT vs µcat Comparison For the same base VAT concept, the HA-CAT can gain an enormous increase in thrust through the use of the heating element. Shown above moves the HA-CAT T P ratio of N, from 2 W 10 5 to 0.08 N. A major increase for W little extra power consumption. This concept can lead to potentially three modes of operation: 1) Thrusting with no heating element, 2) Thrusting with heating element, 3) No thrusting with heating element only, and evaporating the coated anode to provide thrust. Options 2 and 3 gain large benefits with higher thrust levels and can be used for maneuvers if necessary, or work as a normal µcat Ionization Rate Another important piece of information is knowing the ionization rate, which tells how much the plasma is ionized. The plasma consists mainly of neutrals, which means a neutral charged plasma, and is a benefit for satellite 42

58 propulsion. Neutral plasma has low chances, if any, for any form of contamination. Shown in Equation 16 below is the ionization rate: α = Γ i Γ a (16) This relation is the flux of ions Γ i over the total flux Γ a. Where flux is given as Equation 17: Γ i = n i u i = I i es (17) The flux is equal to the density of ions in the plasma n i multiplied by the velocity of ions u i, and also equal to the ion current I i, over the charge of an electron e, multiplied by the surface area S, which in this case is the anodes surface area taken as 0.01m 2. This provides a flux of A. Cm 2 The flux for arc current can be taken using the erosion rate which is Equation 10. Flux becomes: Γ a = E ri arc m a S (18) The erosion rate E r for copper is , while the arc current I arc, is 36 A, over the atomic mass of copper m a as kg, by the surface area of the ion collector S of m 2. This provides a total flux of A Cm 2. The ionization rate is thus: α = Γ i Γ a = = (19) This can be taken as α << 1 and gives us a low ionization rate. This allows a conclusion to be made that the plasma consists mostly of neutrals, and based on the evaporation of the copper. 43

59 6.2 Mass Measurement Results The Mass measurements did not provide adequate results. Comparing both the calculated mass loss per experiment, and the mass collected through the metallic cup, unsatisfactory results were measured. The calculations gave 2E 4 mass variations, but what was measured was a wide range of masses ranging from 8E 4 to 1E 5. Some measurements had mass gained after the tests were finished. The author feels best to characterize the HA-CAT mathematically for mass loss, until more suitable results can be obtained. 6.3 Total Ion Current Results Without Heating Element The first method was running the HA-CAT on cold mode. Pulsing into an ion collector which was negatively biased, the HA-CAT provided very low efficiencies. Shown in Figure 29, 30 and 31, are the data collected from the 4 Channel Oscilloscope. This data was put into MATLAB to provide adequate plotting with two Y-axes to show the Arc Current vs Ion Voltage. The data collected above shows current on the arc current y-axis (left) and voltage on the ion voltage y-axis (right). It is important to note that the arc current measurements were actually voltage measurements, but since the reading was taken through a current transformer probe, a direct 1 V to 1 A relation can be used. Hoping to not create confusion, the author labeled the arc voltage readings as current due to the ability to directly correlate the two. Therefore, these graphs read current next to voltage. The ion current can be calculated by taking the peak points on the curves, and using equation 21 as shown in Appendix A. The efficiencies of 29, 30 and 31, are 0.081%, 0.059% and 0.064% respectively. These numbers are very low and prove that 44

60 the geometry requires the heating element for the desired efficiency increase. With a majority of the plasma coating the anode, small amounts of plasma are escaping. In comparison with previous efficiencies characterized by Zhuang, µcat s with no magnetic coil, these are comparable results. Zhuang reported efficiencies for the µcat s of 0.06% which is very similar to these HA-CAT cold mode outputs[41]. A good sign the overall device is working properly, and reasonable results are being obtained through the experimental setup Arc Current Ion Current Arc Current (A) Ion Current (A) Time (S) x 10 5 Figure 29: Cold Mode Ion Current Results Arc Current Ion Current Arc Current (A) Ion Current (A) Time (S) x 10 5 Figure 30: Cold Mode Ion Current Results 2 45

61 60 50 Arc Current Ion Current Arc Current (A) Ion Current (A) Time (S) x 10 5 Figure 31: Cold Mode Ion Current Results With Heated Element With the heating element on the HA-CAT, the current collection readings were gathered through the same method as above. The data gathered from the oscilloscope was transferred, compiled, and graphed through MATLAB. Shown in Figures 33, 34 and 35 are few results with noticeable improvements to ion current collection. The thermocouple readings for these sets of data are displayed in Figure 32. This is a reading of roughly 41 mv, which reads roughly 1000 C and is near the desired temperature Voltage (V) Time (S) x 10 4 Figure 32: Thermocouple Reading 46

62 The efficiencies gathered above are a large improvement over previous gathered data. In Figure 33, 34 and 35, efficiencies are calculated as 0.836%, 0.478% and 0.515% respectively. This is on the order of magnitude roughly 10 times higher than the cold anode. A Major leap in efficiency. With the heated anode concept and applications of a magnetic field, a greater efficiency could be reached making the HA-CAT a likely candidate for space applications Arc Current Ion Current Arc Current (A) Ion Current (A) Time (S) x 10 4 Figure 33: Hot Mode Ion Current Results Arc Current Ion Current Arc Current (A) Ion Current (A) Time (S) x 10 4 Figure 34: Hot Mode Ion Current Results 2 47

63 60 50 Arc Current Ion Current Arc Current (A) Ion Current (A) Time (S) x 10 4 Figure 35: Hot Mode Ion Current Results Scanning Electron Microscope Analysis Without Heating Element Using the SEM and EDX to achieve scans of the anode surface, and copper deposition analysis, various anodes at arc ignition spots have been analyzed. Examples below are scans of a cold anode after 10,000 pulses, and small trace amounts of copper have been found. Figures 38 and 39 show both the image at a resolution of 50 to 60 µm. As the EDX scans the surface, the yellow peaks are formed displaying which materials are present and are shown on the right of each figure. (a) SEM Scan 1 (b) SEM Materials Graph 1 Figure 36: SEM Scan of Cold Anode 1 To give an extreme example, a thorough test of a cold anode HA-CAT of roughly 10 trials at 10,000 pulses each trial was undertaken. This provided a 48

64 (a) SEM Scan 2 (b) SEM Materials Graph 2 Figure 37: SEM Scan of Cold Anode 2 large deposition of copper that had been caked onto the cold anode. A variety of locations starting near the arc ignition location and linearly outwards have been scanned in order to gauge the copper deposition distribution. Results of two distributions starting at the arc discharge location and then farther away are shown in Figures 38 and 39, which provide further proof the defining mechanism is working properly and large traces of copper have been found. (a) SEM Scan 3 (b) SEM Materials Graph 3 Figure 38: SEM Scan of Cold Anode 3 In Table 3 the copper distribution can be seen in Figures 38 and 39. The anode is shown to have a copper weight consists of roughly 30% being near the arc discharge location. This much copper deposition is a large indication that the device is working properly. Looking back at Equations 10 above, this is present more clearly due to the large sum of pulses, roughly 100,000 pulses. 49

65 (a) SEM Scan 4 (b) SEM Materials Graph 4 Figure 39: SEM Scan of Cold Anode 4 With small pulse amounts it is harder to see a copper distribution such as in Figures 36 and 37, however, this case is an extreme. Atomic Weight Ratios Sample Reference Weight% Atomic% Figure Figure Table 3: Atomic Weight Ratios of Thick Copper Deposition on Cold Anode If this copper deposited anode heats up to the melting temperature, all of this would be evaporated providing thrust to the HA-CAT. This concept can also be used as an alternative firing mode, which could provide a constant thrust for short duration using low power With Heating Element Scans of the heating element produce expected results. Little to no traces of copper can be found. The EDX can find materials if their percentage is larger than 1%. Therefore, copper is less than 1% on these anodes or none at all. Shown below in Figures 40 and 41 are scans of one of few heated anodes scanned, and graphs of materials as well. 50

66 (a) SEM Scan 5 (b) SEM Materials Graph 5 Figure 40: SEM Scan of Hot Anode 1 (a) SEM Scan 6 (b) SEM Materials Graph 6 Figure 41: SEM Scan of Hot Anode 2 Further proving the heating concept, no copper has been found on the heated anodes, solidifying expectations. 6.5 AFM Analysis Without Heating Element Using the AFM to scan the surface of the cathode elements is the final method of analysis in this thesis. In Figures 42, 43 and 44 below, a series of contact scans have been taken over an anode in cold operation mode. By covering half the anode with a piece of thermal tape, scanning the same sample can provide results of with and without copper coating. Shown in Figure 42 is the sample that was scanned and the reader can see a square section which had 51

67 been covered with thermal tape. Figure 43 is the scan of which the thermal tape was covering. One can notice the smooth surface with no copper particles present. Moving towards the boundary layer between the thermal tape and arc discharge location, copper particles are starting to populate the surface and is noticeable in Figure 44. The last scan is the sample taken near the arc location and the copper particles are quite noticeable with a large height variation. Further proof that the thruster head geometry is working as it should, with a film beginning to form on the cold anode. Figure 42: AFM Scan Cold Anode 1 Figure 43: AFM Scan Cold Anode 2 52

68 Figure 44: AFM Scan Cold Anode With Heating Element Utilizing the AFM has proven more difficult with the heated anodes. The scans taken and shown below in Figure 47, 46 and 45 do not tell very much whether the anode is coated with copper or not. One can notice that the scans taken show an incredible amount of height variation. The author is attributing to the wide variety of high based on thermal heating and displacement of the surface molecules. One can notice, however, in Figures 46 and 47 that depressions exists on the anode surface. These areas could be anode spots, which are the same concept as cathode spots but take place on the anode s surface if certain conditions are met. A deviation in height surrounds these depressions appear to form craters and supports an anode spot concept. Figure 45: AFM Scan Hot Anode 1 53

69 Figure 46: AFM Scan Hot Anode 2 Figure 47: AFM Scan Hot Anode 3 54

70 Chapter 7-Conclusion This chapter is a review of the information discovered above. Conclusion remarks are made about the thesis and process of the author, and future works is then discussed. 7.1 SEM, EDS, and AFM Analysis Using the various machines to further analyze the samples was where the fun began. In depth analysis could be made with close up observation to the µm, and the samples could be scanned using X-rays for material type. Proving the conception works, plasma deposition was found on the cold anode with the AFM and copper elements were also found with the SEM and EDS. The difficult data for the AFM was the heated anode, as the surface had wide range in height values. This gives problems with the definitive proof of copper deposition. The SEM and EDS could then be used instead by scanning the surface for copper particles is the key backup. Doing a variety of scans on multiple hot anodes, no copper was discovered. All these methods prove the concept works physically, but the more important parameter comes from the current testing. 7.2 Efficiency Increase The efficiency increase through the ion current tests were noticeable. Starting around 0.06% efficient, the cold operation mode matched the same efficiency as the previous µcat designs with no magnetic coil influences. Then with heated operation mode, an increase tenfold to 0.8% efficiency was reached. The thought of adding a magnetic coil could potentially increase the thruster efficiency well beyond µcat efficiency levels. More studies should be looked into with magnetic coil operation and potentially reach record breaking levels. 55

71 7.3 Calculations The ionization calculations can confirm the thruster is working as it should. With low ionization rates, the thrust is basically based on the evaporation of its use of propellant. Leading into the reduced ISP, the thermal velocity of the HA-CAT s plasma is considerably slower than the typical µcat s plasma exhaust velocity. This heavy reduction in ISP greatly increases the Thrust to Power ratio by magnitudes higher, reaching levels of 0.08 N W from roughly N W assuming a 0.2% efficient thruster. Potentially the HA-CAT runs with a higher efficiency and with low powers of ISP, which means greater levels of thrust can be obtained. 7.4 Future Work Future work would be to dig deeper and continue to analyze the HA-CAT. The author suggests spending more time creating a better thruster design. Starting with thermal tape to help assemble the thruster lead to odd deposition to parts inside of the chamber. These parts gained a sticky adhesive from the thermal tape as it all melted away. A new design was implemented, but rather cumbersome to piece together for constant testing. If a concrete design were developed, it would allow for easy fastening and heating element swapping. More experiments with quicker turnover could have been completed. One large oversight was the current necessary to heat the anode. The power equations that were calculated were to heat the heating element, which consisted of a small diameter tungsten wire. This took a considerably long amount of time to heat up the anode a small fraction of what it needed to be. During experiments, larger currents were used to heat up the element, and this would need to be re-evaluated as an overall design choice. Higher wattage would be needed to get the HA-CAT working quicker, else long durations would be required. Alternatively one could run simulation work looking into the heated anode oper- 56

72 ation, and the feasibility of heat transfer from the heating element. Time taken to heat a small volume would be the main study and would prove invaluable for the HA-CAT s further development. Another problem arose with the heating element breaking after each run. This would require constant heating elements to be made. With runs that were done consecutively after one another, the heating element was fine without fracturing. However, the titanium wire became brittle after exposure to such high heats and made resetting testing quite difficult. For a thruster head in space, this most likely wouldn t cause an issue, but for repetitive experiments it was not ideal. With efficiency through arc and ion current characterized, a variety of experiments need to be investigated. Plasma velocity should be measured to further characterize and compare calculations. Contamination tests could be done, which would show whether the plume exhaust caused issues for the satellite or not. Though by using the µcat family of design, which consists of mainly neutral plasma, contamination should not be an issue. However, further analysis could be useful for characterization efforts. Crafting a thrust stand and physically measuring the two thrust pulses would also be invaluable proof that the HA-CAT is working. Measuring the exact thrust has been a long standing goal for the MpNL, to further prove the theories behind conception. This technology has a potential for development but will face many more engineering feats. Keeping the heat contained without spreading to the rest of the spacecraft will be huge problem. Once materials start to break down and become brittle, issues arise. With the increase in thrust it may be worth looking into a design which will allow for prolonged lifetime. Also a satellite with a constant power drain will be required to run the heating element for hours at a time, potentially days if the mission requires it. With the heating element being such low resistivity, having that heat up 57

73 over other parts of the circuit will also be an intellectual design challenge. This concept can be taken as a mission design option. Perhaps thrusting under cold mode for typical operation and thrusting, but when higher maneuvers are required the hot mode can be activated. If the satellite bus requires heating in deep space, perhaps the hot mode can be used for thermal management. These are a few ideas how the HA-CAT can be applied to a variety of applications for small spacecrafts. The author believes this technology could be improved and matured, but many things need to be thoroughly investigated to make it satellite ready and safe. The concept is proven to be more ideal than previous designs. The HA-CAT can be run in cold mode, and still compare with other competing VAT technology. Or the HA-CAT can run in hot mode, and be more efficient with a trade off of constant power drain. The HA-CAT is a flexible thruster subsystem that could benefit from more thought to the engineering. A mature product will take some time, and technological challenges will need to be tackled before a polished product will come to fruition. The author hopes this thesis may be a starting point to new conceptions for VAT technology, and to induce a variety of designs and inspirations. 58

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79 Appendix A Lorentz force is the combination of electric and magnetic forces on a point charge due to electromagnetic fields and is shown in the Equation 20 below: F = q (E + v B) (20) F is the Lorentz force created by the electric and magnetic cores. q is the particles charge as it moves with a velocity of v. E is the electric field and B is the magnetic field. Shown below is Ohms law in Equation 21 below: V = I R (21) V is the Volts being equal to the current I multiplied by resistance R. This equation has been used to calculate the power required for the heating element. Hooks Law is a relation of force applied from a spring and shown in equation 22 below: F = k X (22) Some force F is equal to a spring constant k displaced by some distance X. This relation is frequently used in mechanical systems and is one of the basic laws required for mechanical engineering. 64

80 Appendix B Figure 48: Chamber in MpNL at GWU. 65