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Design and Testing of 1 kW Hall Thruster

Nicholas Ian Denning, Nicholas Alfred Riedel Colorado State University

Advised by: Dr. Binyamin Rubin, Dr. John Williams ndenning@engr.colostate.edu, nriedel@engr.colostate.edu

7 April 2008

Abstract During the last decade there has been an increased interest in the use of electric propulsion on

spacecraft. The high specific impulse of electric propulsion devices enables one to increase the payload mass, life span of spacecraft, or perform complex orbital maneuvers with a minimum amount of propellant. The Hall thruster, one of many electric propulsion devices, has become the most widely used form of electric propulsion due to its high performance and relative simplicity. Hall thrusters are currently used in various space applications including satellite orbit correction and station keeping. A noncontact measurement method was proposed by Rubin, Geulman, and Kapulkin, 20081, enabling real-time thrust estimation, which is currently not possible to do on conventional Hall thrusters. The proposed method uses several miniature magnetic field sensors embedded inside the Hall thruster. These sensors will monitor the magnetic field produced by the plasma current loop in the thruster acceleration channel. Information on magnetic field measurements from the sensors will be used to determine Hall current distribution inside the thruster and to estimate thrust. The sensor data is also useful for thruster design optimization. The Hall thruster that was designed, manufactured, and tested to validate the concept, is described in this paper. This thruster, intended for operation at 1 kilowatt power level, was designed to be easily assembled and disassembled to allow access to the magnetic sensors placed inside. The results of magnetic field and thermal modeling obtained during thruster design, and later validated experimentally, are discussed. The results of the thruster firing tests are also presented.

I. Introduction Although the idea of electric propulsion was introduced around the turn of the century, serious research

didn’t begin in the United States and Russia until the late 1950’s. It wasn’t until 1990 however that this technology was developed enough to be used on modern spacecraft.2 Electric propulsion has become increasingly popular and studied more in depth in the past decade. Although it does not provide thrust levels as high as chemical propulsion, it can prove advantageous for missions where low thrust is not a limiting factor and longer mission duration can be tolerated. The high specific impulse provided by electric propulsion allows spacecraft to travel farther and faster than a similar spacecraft using chemical thrusters. The drawback is that electrical propulsion requires operating the thruster for a longer period of time before the top speed is reached. Electric propulsion systems can also provide a less expensive alternative to the use of chemical thrusters. Currently electric propulsion is used for station keeping, orbit adjustments, and provides the primary source of propulsion for many deep space missions.

Conventional ion thrusters are operated by applying an electric potential between two grids. Gas is then ionized and the ions are accelerated through the grids producing thrust. The biggest drawback of this form of electric propulsion is that the thruster is space charge limited, thus restricting the ion current density as well as the thrust. Hall Effect thrusters are alternative types of ion thrusters which provide higher thrust density. The main advantage is that they involve the formation of quasi-neutral plasma and therefore ion current is not limited by space charge. The induced magnetic field in Hall thrusters is used to slow the flow of electrons to the anode and creates a “halo” of circulating electrons in the discharge chamber. This allows for a more efficient ionization than conventional ion thrusters provide. Hall thrusters operate at high specific impulses (greater than 1500 seconds), moderate thrust levels (100 to 200 mN), and efficiencies

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greater than 50% in the 1 to 10 kW range.3 Due to these characteristics, they are well suited for many varieties of space mission applications.

Hall thruster operation is accomplished by applying a large electrical potential difference between the anode and a cathode as is observed in Figure 1. Small mass flow rates of propellant, usually xenon gas, are fed through both. This creates plasma and allows electrons to flow into the discharge chamber towards the anode. Electromagnetic coils placed in the center and around the perimeter of the thruster are charged and a magnetic field is induced trapping electrons in the discharge chamber. The propellant is continuously fed through the anode into the discharge chamber where collisions occur between the gas and the trapped electrons causing the propellant atoms to ionize. Ions are then accelerated by the electric field and emitted from the discharge chamber. These ions exit at a high velocity when emitted and therefore produce thrust. Once out of the thruster, ions are then neutralized by the flow of electrons from the cathode creating a discharge plume with no net charge.

Currently, there is no method for measurement of thrust in real time when a Hall thruster is in operation on a spacecraft. A noncontact measurement method was proposed which uses several small magnetic sensors to monitor the magnetic field induced by the azimuthal plasma current.1 In order to validate this theory, a low cost Hall thruster was needed for testing. While this thruster was produced, two separate teams worked on the various components to be used for validation. One team concentrated on the sensor arrays that would be embedded in the thruster. The other team fabricated a thrust stand that enabled thrust measurement and efficiency calculation of the thruster. Careful attention was paid to the thermal characteristics of the thruster so the sensors would not be exposed to higher than allowable temperatures. The thruster designed was to operate at approximately 1 kW power and was to be easily modified for sensor placement.

II. Thruster Design

Dimensioning

The Hall thruster described here is Stationary Plasma Thruster (SPT) type. The main feature of this type is that discharge chamber walls are made of insulating material. Through intensive research of existing Hall thruster documentation, it was decided to use the sizing convention presented at an electric propulsion seminar given at the Massachusetts Institute of Technology in 1991 by Russian Hall thruster designers and published in Gulczinski’s doctoral dissertation at the University of Michigan.2 It was found

Figure 1 - Hall Thruster Operation (Courtesy of www.al.t.u-tokyo.ac.jp)

Figure 2 - Sizing for Hall Thruster2

bm = .3dch [mm] bch = 6+.375bm [mm]

Lc = .32bm [mm] La = 2Lc [mm]

Lch ≥ 1.1La [mm]

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that there exists a relationship between certain dimensions on a Hall thruster which give a maximum efficiency. This relationship is explained using the equations and corresponding picture in Figure 2.

The Russian designed SPT100 runs at approximately 1 kW of power and has a discharge chamber diameter of 100 mm. Taking this as the discharge chamber diameter and using the relationships given above, sizing of the thruster was computed. The calculations and final dimensioning can be found in Appendix A. Discharge Chamber

One of the most important parts of a Hall thruster is the discharge chamber, which is generally made of Boron Nitride or another type of ceramic. This is because ceramics provide a higher sputtering resistance than metals or polymers. Aluminum Oxide (alumina) was chosen due to the fact that it was widely available and cost effective. Most Hall thrusters use a discharge chamber machined from a solid block of ceramic; however, it was decided to use a design in which multiple alumina plates were machined separately and stacked on top of each other for the final assembly.

Like many ceramics, alumina is notoriously brittle. For this reason, the bottom plates were increased in thickness in the radial direction in an effort to prevent cracking and improve the structural integrity of the discharge chamber. The thickness of the top two plates could not be increased as the pole pieces limited the available space.

The multiple plate design was chosen due to the fact that it was easier to machine, more accommodating to sensor placement, and individual broken pieces would be easier to replace. Multiple plates may also provide further research opportunities not possible with a single unit. The drawback of this design was that it required the use of compression plates to hold the machined alumina plates in place and prevent them from shifting in the thruster assembly. To properly align the alumina plates, they were Super-Glued into the correct positions, allowing the discharge chamber to be installed as a unit before compression plates were attached. Once the thruster reaches a moderate operating temperature, the Super-Glue burns out and only the compression plates hold the alumina in place.

In order not to interfere with the magnetics of the thruster, the compression plates were made from a non-magnetic stainless steel. A

Pro-Engineer model of the thruster and how the plates integrate can be seen in Figure 3. This design will also help reduce sputtering of the magnetic pole pieces. If the compression plates themselves become too eroded, they can quickly and easily be replaced.

Thruster Design: Anode

The dimension from the top of the anode to the top of the discharge chamber (Lch) and the dimension from the top of the anode to the top of the discharge chamber (La) both play a large role in determining the efficiency of the thruster. Slight variations in these values allow for fine tuning of the maximum thruster output. For this reason, space has been left below the discharge chamber and below the anode in order to insert stainless steel spacers. The spacers, observed in Figure 4, can be varied in height allowing the thruster output to be adjusted to achieve the maximum efficiency and therefore the optimal power output.

The anode was machined out of non-magnetic stainless steel so it would not affect the magnetic field. Eighteen evenly spaced holes, .016” in diameter, were placed to allow propellant to flow into the discharge chamber. The hole size and quantity were chosen so that the sum of the area of the discharge holes was less than the area of the inlet gas line. This promotes a uniform discharge flow.

Once dimensions for the thruster were achieved a Pro-Engineer model was generated, a cross section of which can be seen in Figure 5.

Figure 4 - Anode Spacers

Figure 3 - Thruster Compression Plates

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Figure 5 - Cross Section of Thruster Design

III. Preliminary Analysis

Magnetic Analysis

Although much of the thruster dimensioning had been determined by the equations presented above, there was still a great deal of work to be done with the overall magnetic field. In order to achieve the greatest thruster efficiency, it was ideal to have the magnetic field focused between the pole pieces, with the area of the discharge chamber as magnetically neutral as possible. To achieve these properties, thin rings of magnetic steel were placed between the discharge chamber and the coils. These magnetic shields are common in most Hall thrusters.

Since the thruster was designed for interchangeability and variability, it was necessary for the magnetic shields to be removable. Instead of welding the shields into place, as was originally purposed, it was decided they could be bolted onto the thruster base plate. To allow room to drill and tap, the steel thickness was set at .125”. This left only the height of the shield to be determined. The shield height has a dramatic effect not only on the magnetic fields in the discharge chamber, but also the strength of the field between the pole pieces. If the shield is too tall and placed too close to a pole piece, much of the magnetic flux created by the coils is drawn to the shield and not transmitted between the two poles as desired. Conversely, if the shields are too short, the magnetic field between the pole pieces is stronger but there is more magnetic interference in the discharge chamber.

To determine the proper shield height, a number of tests were conducted using the two-dimensional magnetic modeler Finite Element Method Magnetics (FEMM). The height of the shield began at 1.25” and increased by .0625” increments to 1.6875”. To simulate the load applied to the thruster, the center core was modeled to have 300 coils of wire with a load of 2 Amps and each of the outer cores had 150 coils with a load of 1 Amp. (It was necessary to run more current through the center coil to produce a more symmetric magnetic field). The simulations were run for each height and the resulting magnetic fields were mapped. Plots of magnetic field intensity through the center of the discharge chamber were also constructed. Using the results in Appendix B, comparisons were then made and the screen height that provided the greatest balance of characteristics (1.5625”) was selected.

Figure 6 - 3D Magnetic Magnitudes

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With a screen height chosen, the next step was to verify the 2D results with a 3D analysis. The entire magnetic system of the thruster was modeled using the analysis program Maxwell 3D. The same loads applied to the coils in the 2D analysis were applied to the 3D model. The resulting magnitude plot (Figure 6) and vector field (Figure 7) were then compared to the fields and magnitudes previously estimated. It turned out that both the magnitude and field shapes predicted by the FEMM program were almost identical to the results of the Maxwell analysis. It was concluded that the data gathered from the 2D analysis was valid, and would provide an accurate prediction of the magnetic characteristics.

Thermal Analysis

The next issue for consideration was the thermal characteristics of the thruster. In most Hall Effect thrusters, temperatures can easily exceed 400 °C. Since the sensors to be placed inside the thruster have a temperature limit of 300 °C it was necessary to get an approximation for the temperatures that would be experienced during operation. Researchers in France have observed a maximum temperature in a similar 1kW thruster at approximately 600 °C 4. In these thrusters, the anode acts as the primary source of heat; the thermal addition from the coils is negligible. Using this information, the thruster was modeled using the program ePhysics. To generate a conservative model, the anode was set to a constant temperature of 800 °C while the rest of the thruster began at 30 °C. The material properties of the model were set to match the materials to be used on the actual thruster. A vacuum condition was also modeled to allow conduction and radiation as the only modes of heat transfer.

This simulation provided results that were similar to findings observed with actual thrusters at other institutions. The maximum temperature at the anode was very conservative, around 750 °C. Most of the thruster was a little above 150 °C. These results, shown in Figure 8, served as a general model until actual temperatures were measured when the thruster was operational.

Figure 7 - 3D Magnetic Vector Field

Figure 8 - Thermal Analysis Results

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IV. Equipment

Although the general concepts behind the operation of a Hall thruster are relatively simple, there is a large amount of hardware required for testing. All testing must be performed in a high vacuum environment, with enough space and a high enough pumping speed such that the gas emitted in the discharge plume doesn’t have a chance to return to the thruster for re-ionization. For this reason a steel vacuum chamber, 5 feet in diameter and 15 feet in length, was used for the tests. To pump the chamber down from atmospheric pressure, an Edwards GV250 dry mechanical pump was used, which was later assisted by an Edwards EH-1200 mechanical booster pump when the pressure reached 30 Torr. When the chamber pressure reached 20 mTorr with only the mechanical pumps running, two Varian HS-20 diffusion pumps were turned on. This allowed the chamber to reach a baseline pressure in the high 10-6 Torr range.

An argon high pressure cylinder was connected to a network of stainless steel gas lines and used to supply fuel to the anode and cathode, as pictured in Figure 9. The purity of the argon was 99.998%. Flow of the gas was regulated using two Brooks 5850E mass flow controllers; the anode controller rated for 200 sccm argon and the cathode controller rated for 50 sccm argon. The mass flow controllers were powered and managed using two Brooks 5896 power supplies.

Power was provided to all the separate components through a number of power supplies. The cathode heater used a Sorensen DCS40-25E, while the keeper was powered by a Spellman SL-1200. A Gold Tool DPS-5050 power supply provided the current for the inner coil, and a TDK-Lambda ZUP10-20 powered the outer coils. The supply used for the main discharge was a custom built unit, capable of 600V/15A output. A wiring schematic is provided in Figure 10.

Test information was collected using a HP/Agilent 34970A data acquisition unit, and was sent to a PC

equipped with LabVIEW. The LabVIEW program was used to send the data to an Excel file to allow for easy processing. The data recorded through this system included chamber pressure, anode fuel flow rate, cathode fuel flow rate, discharge voltage, inner magnet voltage, outer magnet voltage, and readings from 5 thermocouples.

The final thruster design resulted in an overall diameter of 8 inches, with a height of approximately 2 inches. The inner core was wrapped with 22 gauge magnet wire allowing for 850±45 turns, while each outer core was wrapped with 18 gauge magnet wire allowing for 200±10 turns. The overall weight of the

Figure 9 - Fuel Supply Schematic

Figure 10 - Wiring Diagram

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thruster was about 12 pounds. The pole pieces and base plate were cut from ASTM A36 steel, while the cores and shields were constructed from AISI 1018 steel. The anode and compression plates were machined from stainless steel. Figure 11and Figure 12 display the final assembly of the thruster. A hollow cathode with a porous tungsten insert impregnated with low work function material was used to provide electrons both for discharge and for plume neutralization.

V. Testing and Results

Magnetic Mapping

Once the design and manufacture of the thruster was finished, it was possible to begin testing. Before the thruster was fired in the vacuum, a mapping of the magnetic fields was completed. This was done using a 2-axis linear actuation system and a Gauss-meter. Since the alumina discharge chamber, anode, and compression plates were made of non-magnetic materials, they were not installed on the thruster. This allowed the magnetic probe to take readings in a wider range of locations and permitted a broader understanding of the fields produced. The readings were taken in the axial direction at 1 mm increments; from the base plate to 10 cm away. This was done at 5 different radial locations between the inner and outer pole pieces. A LabVIEW program collected the probe positions and magnetic field strengths and wrote the data to an Excel file.

With a desired power output of 1 kW, the ideal field strength would be 150 Gauss at the center of the acceleration channel between the pole pieces. After further computer simulations were performed knowing the actual number of turns in the coils, it was determined that this optimal performance would be achieved with 1 Amp of current to the inner coil and 2 Amps to the outer coils. The models showed that this excitation would provide the same field shapes seen in the preliminary analysis, with the correct field strength. Using this information, the thruster magnetic system was mapped.

The data gathered from mapping was used to plot the magnetic field lines and intensities, shown in Figure 13. It was observed that the actual field strength was about 35 Gauss less than computer simulations predicted. Efforts were made to increase the strength of the magnetic field by varying the current provided to the coils. The results of these trials can be seen in Figure 14. Magnetic field lines and intensities for each case of current variation can be found in Appendix C. Since the field strength was relatively unresponsive when the current was increased, it was believed that the center core was reaching magnetic saturation.

Figure 13 - Magnetic Field Lines and Intensities. The Black U-Shape Indicates the Walls of the Discharge Chamber.

(Inner Coil: 1A , Outer Coil: 2A)

Figure 12 - Thruster Assembly Figure 11 - Rear of Thruster

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Simulations in Maxwell 3D showed the core was close to saturation, though not reaching it. However, the computer model uses an ideal B-H curve for the steel. Magnetic properties can vary greatly even in similar metals, thus it is likely that the actual B-H curve yields a lower saturation point for this particular sample.

Although the magnetic field was lower than desired, it was decided that the thruster would still be able to operate. The decreased magnetic peak would only lead to a slight drop in efficiency.

1 2 3 4 5 6 7 8 9 100

20

40

60

80

100

120

Axial coordinate,cm

Br,

gaus

s

inner coil - 0.5A, outer coil - 1 Ainner coil - 0.5A, outer coil - 2 Ainner coil -1A, outer coil - 1 Ainner coil - 1A, outer coil - 2 Ainner coil - 1A, outer coil - 3 Ainner coil - 1.5A, outer coil - 2 Ainner coil - 1.5A, outer coil - 3 A

Anodelocation

Exitplane

Testing in Vacuum

Once the magnetic mapping was complete, the entire thruster was assembled and placed into the vacuum chamber. Electric and fuel lines were connected as previously described and the chamber was pumped down. When the chamber had reached its baseline pressure, the cathode heater was turned on along with the cathode gas flow. As soon as the cathode finished conditioning, the heater was turned off and the keeper was turned on. The gas supply to the anode was activated and a large electric potential was applied to the anode. This resulted in the start of the thruster. When the discharge had stabilized, power was provided to the coils and the thruster was completely operational. The operational thruster can be seen in Figure 15 and Figure 16.

Figure 14 - Changes in Field Strength Corresponding to Changes in Applied Current

Figure 16 - Front View of Thruster Operation Figure 15 - Side View of Thruster Operation

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Table 1 provides data from the modes tested. It is believed that higher power levels will be reached as

testing continues.

Table 1 - Modes of Thruster Operation

Cathode Flow

Rate (sccm)

Anode Flow Rate

(sccm)

Discharge

Voltage (V)

Dicharge Current

(A)

Discharge Power

(W)

5 55 154 2.02 311

5 60 146 2.79 407

5 100 125 4.95 618

Thermal

The thruster was active in vacuum for a period of approximately three hours. It demonstrated stable operation in a range of modes as described in Table 1. Thermocouples were placed at 4 separate locations on the thruster and 1 location on the stand in order to monitor temperatures. Figure 17 shows a plot of each temperature during the operation of the thruster. Due to manipulation of flow rates and voltages, the thruster experienced unstable modes twice during operations (after approximately 1 hour and then again after 3 hours). During these times the thruster was inactive for a short period of time and a drop in temperatures resulted. Although steady state was not reached, it can be approximated from the trends that no temperature measured will exceed approximately 180°C. The coil temperatures should not be greater than 150°C, which is well below the maximum operating temperature of the magnet wire (200°C). These readings correspond well to the preliminary thermal analysis, though more data will be gathered before final placements of sensors are determined.

Figure 17 - Plot of Thruster Temperatures

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VI. Difficulties Encountered Some minor difficulties arose during vacuum testing of the thruster. On the first testing attempt,

plasma was formed behind the thruster and electrical arcing occurred. This prevented the stabilization of discharge and therefore inhibited continual thruster operation. After a brief investigation, it was found that a gas line fitting was leaking and allowed the formation of plasma behind the thruster. The faulty o-ring was replaced and all exposed gas line was wrapped with Teflon tape to prevent arcing.

Another problem was encountered when an attempt was made to restart the thruster after operation for extended periods of time. After troubleshooting the situation, it was concluded that the magnetic remanence in the cores and pole pieces prevented the flow of electrons from the cathode to the anode, thus preventing restart of the thruster. To remedy the problem a procedure was implemented in which the leads of the coils were reversed on the power supplies and a very low current was provided for a short period of time in order to induce an opposing magnetic field. This canceled out the remnant magnetic field and allowed for successful start up of the thruster.

VII. Future

With the successful test of the thruster, the design proved to be functional. To improve the magnetic field strength, a replacement core will be machined with a larger diameter center. This increased area should bring the steel away from its saturation point and allow a greater magnetic flux. Adjustments to anode and exit plane height will be made in efforts to improve efficiency. Further tests will also be conducted in order to obtain current-voltage curves. With these tests finished, an optimal operating condition will be found. Thermal characteristics will continue to be investigated until steady-state temperatures are verified and a more complete thermal model can be made. Sensors will then be integrated into the thruster in order to measure the magnetic fields induced by the Hall current. Measurements taken by the thrust stand will be incorporated and in turn validate that the non-intrusive thrust measurement is possible.

VIII. References

1. B.Rubin, M.Guelman, and A.Kapulkin, “Magnetic Sensing of Azimuthal Current in Hall Thruster: In-Flight Diagnostic Potential”, Journal of Propulsion and Power, 2008 vol.24 no.1, p.118. 2. F.S.Gulczinski III, “Examination of the Structure and Evolution of Ion Energy Properties of a 5 kW Class Laboratory Hall Effect Thruster at Various Operational Conditions”, University of Michigan Doctoral Dissertation, 1999. 3. N.Gascon, W.Scott Crawford, R.L.Corey, M.A. Cappelli, “Coaxial Hall Thruster with Diamond Inner Channel Wall”, 42nd AIAA Joint Propulsion Conference & Exhibit, 2006. 4. S.Mazouffre, P.Echegut, M. Dudeck, “A Calibrated Infrared Imaging Study on the Steady State Thermal

Behavior of Hall Effect Thrusters,” Plasma Sources Science and Technology, 2007 vol.16, p.13-22.

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APPENDIX A – Sample Thruster Calculations The dimensions for the thruster were determined using the following equations. Knowing that a

discharge chamber diameter of 100 mm would yield approximately a 1 kW power output, the other dimensions were calculated as seen below:

Due to the fact that the above equations are used to find the dimensions in millimeters, the values were

converted to inches to keep consistent tooling on the project. When converted to English units the values were corrected slightly to allow for more rounded dimensions to aid in machining. A summary of these values can be seen in Table A1 below.

Table A1 - Thruster Dimensions

Thruster

Dimension[mm] [inch]

[inch]

(corrected)

dch 100 3.937 3.875

bm 30 1.1811 1.1875

Lc 17.25 0.6791325 0.6875

La 19.2 0.755904 0.75

Lch 21.12 0.8314944 0.8125

dch 100:=

bm .3 dch⋅:= bm 30=

bch 6 .375 bm⋅+:= bch 17.25=

Lc .32 bm⋅:= Lc 9.6=

La 2 Lc⋅:= La 19.2=

Lch 1.1 La⋅:= Lch 21.12=

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APPENDIX B - Finite Element Method Magnetics (FEMM) Results The following are the plots from the FEMM analysis used to determine the proper shield height. The

plots on the left show the strength of the magnetic field down the center of the discharge chamber. The figures on the right show the resulting magnetic field lines. Using these results, a shield height of 1.5625 was chosen, as it displayed the best combination of properties.

Shield Height: 1.25”

Shield Height: 1.3125”

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Shield Height: 1.375”

Shield Height: 1.4375”

Shield Height: 1.5”

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Shield Height: 1.5625”

Shield Height: 1.625”

Shield Height: 1.6875”

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Appendix C – Magnetic Field Mapping Results The following are plots of magnetic field lines and intensities for all cases of current variations tested.

The black U-shape represents the walls of the discharge chamber of the thruster in reference to fields.

Inner coil – 1A, Outer coil – 2 A

Inner coil – 1A, Outer coil – 1 A

Inner coil – 1A, Outer coil –3 A

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Inner coil – 0.5A, Outer coil – 1 A

Inner coil – 0.5A, Outer coil – 2 A

Inner coil – 1.5A, Outer coil – 2 A

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Inner coil – 1.5A, Outer coil – 3 A