45th AIAA Aerospace Sciences Meeting and Exhibit AIAA 2007-0584 8-11 January 2007, Reno, NV

American Institute of Aeronautics and Astronautics

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Characterization of a Hall Effect Thruster Using Thermal Imaging

Capt. James W. Tomaszewski* and Maj. Richard D. Branam† Air Force Institute of Technology, Wright-Patterson AFB, OH 45433

William A. Hargus, Jr‡ and Taylor Matlock§

Air Force Research Laboratory, Space Propulsion Branch, Edwards AFB, CA 93523

A Hall Effect Thruster rejects heat to space and the surrounding spacecraft, thus impacting spacecraft design. In order to characterize the heat rejection, thermal information must be gathered and analyzed. In addition, thruster temperature has an effect on the useable lifetime of the thruster. This paper contains analysis of thruster temperatures obtained using a commercially available FLIR A40M thermographic imager in order to characterize a Busek Inc. 200W Hall Effect Thruster operating in Chamber 6 at the Air Force Research Laboratory at Edwards AFB, CA. This method is non-intrusive in that the thruster is viewed from outside the chamber through a Zinc-Selenide window and the data is output to a computer for further processing. Maximum temperatures observed were 550 C and above on the cathode tip and the thruster body.

Nomenclature U = camera voltage τ = transmissivity ε = emissivity

I. Introduction While many different means of space propulsion are available, the electromagnetic Hall thruster provides a high

performance option in the electric propulsion category1. As such, Hall thruster operation is currently an area of interest. For example, a Hall thruster will undergo testing on TacSat-2, which was launched December 16, 2006. Thruster components, such as the anode cone and the cathode, have been studied using various methods, but thermal imaging data is not known to be available. The goal of this research was to develop a non-intrusive means to obtain and analyze information concerning the thermal characteristics of an operational Hall thruster. To accomplish this, the A40M thermal imaging camera produced by FLIR, Inc was used to record data on a Busek Company, Inc. BHT-200-X3 Laboratory model Hall thruster operating in Chamber 6 at the Air Force Research Lab (AFRL) Electric Propulsion Research Facility at Edwards Air Force Base. Thermacam Researcher software and Matlab were used to correct the raw image data so the temperature shown on the image would be more accurate. Understanding thruster thermal characteristics may provide for better design because temperature has an effect on thruster useable lifetime4. Knowledge of actual thruster thermal performance in a space-like environment would also allow for better satellite design.1

* Student, Engineering Department, 2950 Hobson Way, WPAFB, OH 45433, Member † Associate Professor, Engineering Department, 2950 Hobson Way, WPAFB, OH 45433, Senior Member ‡ Research Engineer, Space Propulsion Branch, Edwards AFB, CA 93523, Senior Member § Research Engineer, Space Propulsion Branch, Edwards AFB, CA 93523, Member “The views expressed in this article are those of the author and do not reflect the official policy or position of the Air Force, Department of Defense or the U.S. Government.”

45th AIAA Aerospace Sciences Meeting and Exhibit8 - 11 January 2007, Reno, Nevada

AIAA 2007-584

Copyright © 2007 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

45th AIAA Aerospace Sciences Meeting and Exhibit AIAA 2007-0584 8-11 January 2007, Reno, NV

American Institute of Aeronautics and Astronautics

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II. Background The means for heat transfer from the thruster are conduction to the spacecraft and radiation to both the spacecraft

and to space. The A40M thermal imager is sensitive to radiation given off by the thruster in the range of 7.5 μm to 13 μm and is capable of measuring temperatures up to 550 C without additional hardware. The screen area consists of 320 x 240 = 76,800 pixels. Each pixel outputs an intensity associated with a specific temperature. The camera software can account for several factors affecting the measured temperature of the object, as shown in Fig. (1)2.

To further explain Fig. (1)2, the radiation sensed is converted to a voltage generated within the camera, which is

used to calculate the temperature of the object as shown Eq. (1)2.

atmlreftotobj UUUU *1*1*1εττ

εε

ετ−

−−

−= (1)

where Uobj is the calculated camera output voltage for a blackbody of temperature Tobj (i.e. a voltage directly

converted into true requested object temperature), Utot is the measured camera output voltage for the actual case, Urefl is the theoretical camera output voltage for a blackbody of temperature Trefl according to the calibration, Uatm is the theoretical camera output voltage for a blackbody of temperature Tatm according to the calibration, ε is the emissivity and τ is the transmissivity. The camera software, Thermacam Researcher, was used to input values for emissivity, transmissivity, and temperature. Thermacam Researcher was used to export the corrected temperature value of each pixel into an output file for post processing. The post processing included different emissivities.

III. Experiment The overall setup for the experiment consisted of the A40M camera outside the vacuum chamber viewing the

operating Hall thruster inside the chamber. The thruster used was the Busek BHT-200-X3 laboratory model, which is a 200W Hall thruster. The thruster used in this experiment was not brand new. As a result, the thruster showed evidence of use in its appearance, such as dark discoloring of the metal, erosion of the Boron Nitride anode, and the Alumina Plasma sprayed portions of the cathode and anode. Since emissivity is, in part, a function of the coating of a surface with another material, such as paint3, this impacted the values used for the emissivity of the various parts of the thruster. The thruster was moved as close as possible to the window in order to maximize the number of pixels covering the thruster. The camera viewed the thruster through a Zinc-Selenide (Zn-Se) window and was connected to a computer via fire wire. There were a total of 5 thermocouples used. In order to compare the temperature value output by the camera to a temperature measured by more conventional means, three thermocouples were attached directly to the thruster using epoxy and one was attached using Kapton tape. Thermocouple #1 was attached to the metal axial support bar in the center of the thruster on the side opposite the camera. Thermocouple #2 was attached to the metal support at the rear of the thruster on the side opposite the

Figure 1. Schematic of thermographic measurement situation.

45th AIAA Aerospace Sciences Meeting and Exhibit AIAA 2007-0584 8-11 January 2007, Reno, NV

American Institute of Aeronautics and Astronautics

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camera. Thermocouple #3 was attached to the inner chamber wall. Thermocouple #4 was attached to the cathode support mount near the thruster/mount interface. Thermocouple #5 was attached to the metal support at the rear of the thruster on the side facing the camera. The thermocouple data was output to a computer. LabView software was used to convert the signal output by the thermocouples to a temperature that was monitored on the computer. The overall setup is shown in Fig. (2).

The vacuum chamber is 1.8 m in diameter and 3 m long with a measured pumping speed of ~ 32,000 l/s (Xe).

The camera was 4.3 cm away from the Zn-Se window. The center line of the thruster was 0.673 m away from the Zn-Se window.

The thruster was connected to the necessary electrical and gas inputs and the chamber door was closed. The

chamber was then brought to near vacuum at ~ 3 x 10-7 Torr. Following thruster start up, it was allowed to run at 70% power (7 sccm). While the experiment was running, analysis of the thermocouple data displayed by LabView indicated a heat up rate of less than 1 degree per hour at approximately 2.5 hrs after start up. At this point, the thruster was considered to be at steady state operation. The thruster was allowed to run at this steady state configuration for an additional hour. The final 30 min of camera steady state data was analyzed.

In order to post process the raw camera data, values were input to the Researcher software and the resulting

temperature value was output to Microsoft Excel and Matlab. Researcher allows for the following values to be input by the user: Object emissivity, distance between camera and object, reflected temperature, external optics temperature and emissivity, and atmospheric temperature and relative humidity.

A. Object emissivity Raw thermal images are not corrected for emissivity. Corrections for emissivity must be done in post

processing. As noted earlier in this section, the values for emissivity for the parts of a new thruster could not be used. Since the temperature of the steel mount was known from the thermocouple data, the emissivity of the steel was varied within the Researcher software until the temperature displayed matched that of the thermocouple in the same area. The emissivity found was 0.20. Although not ideal, other values obtained from References 1 and 3 were used where available. An emissivity of 0.65 taken from Reference 3 was used for the Alumina portions, which include the cathode tip and the exit plane of the thruster. An emissivity of 0.45 was used for the Boron Nitride cone, as provided by the manufacturer. The raw image data was corrected using the emissivity of 0.20 within the Researcher software and then output to a Matlab (.mat) file. The same raw image was set with an emissivity of 0.65, which is correct only for the tip of the cathode. The same raw image was then set with an emissivity of 0.45, which is correct only for the anode cone section. The data values in the Matlab file which had been corrected with an emissivity of 0.20 was used as a base, since most areas in the image were made of steel. The values from the area

Figure 2. Schematic of experiment setup.

45th AIAA Aerospace Sciences Meeting and Exhibit AIAA 2007-0584 8-11 January 2007, Reno, NV

American Institute of Aeronautics and Astronautics

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that correspond to the area covering the cathode tip were copied from the Matlab file that was corrected for 0.65, and pasted over the corresponding values in the base Matlab file. The same was done with the values from the file using 0.45 as the emissivity corresponding to the anode cone. Thus, the composite image would show the correct values for the three areas of steel, Alumina plasma spray, and Boron Nitride.

B. Distance between camera and object The distance used was the sum of the distances from the camera to the window and the window to the camera

lens. The thickness of the Zn-Se window was also added. The total distance used was 0.719 m.

C. Reflected temperature This value was considered to be the temperature of the inner chamber wall as measured by thermocouple. This

temperature was a constant 19 C during the experiment.

D. External optics temperature and emissivity The values used for the “external optics” were the values related to the Zn-Se window. The transmissivity of the

window is 63 %, ± 2 % over the wavelength range to which the camera is sensitive, therefore an average transmissivity of 63 % was used for the Zn-Se window. The temperature of the window used was 24 C, which may not be correct for the entire thickness of the window. Both of these factors introduce error.

E. Atmospheric temperature and relative humidity Since there was a 4.3 cm thickness of atmosphere between the camera lens and the Zn-Se window, the lab temperature of 24 C and relative humidity of 40 % were used. Researcher calculates the transmissivity based on these values, which for these conditions was 0.99. This situation introduces error since the camera takes these values over the entire distance between the camera and the object.

IV. Results A picture of the thruster next to the raw thermal image output by the Researcher software is shown in Fig. (3).

Figure 3. 200W Thruster and Researcher raw thermal image. 1. Kapton tape covering thermocouple #5. 2. Magnet core. 3. Small screw on cathode. 4. Anode cone. 5. Reflection from rear stand on thruster mount plate. 6. Power wire for anode. 7. Axial support bars.

45th AIAA Aerospace Sciences Meeting and Exhibit AIAA 2007-0584 8-11 January 2007, Reno, NV

American Institute of Aeronautics and Astronautics

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The image was taken from the portion of the experiment when the thruster was operating at steady state. There is no temperature scale on the image since the temperatures would not be accurate without correction for emissivity of the various parts, which can only be done in post processing. Although specific temperatures cannot be stated, the image is still useful for showing which parts of the thruster are hotter than others. Hotter parts, such as the cathode tip and anode cone, are shown in yellow, while cooler parts, such as the axial support bars and cathode support structure, are shown in blue. After using the Researcher software to set the emissivity for the entire image to 0.20, the data was exported to Matlab and plotted, as shown in Fig. (4). The temperatures for the cathode tip and the anode cone are not correct in this image.

Figure (5) shows the result of importing the data from the Researcher software in to Matlab and applying the corrected emissivity for the cathode tip and the anode cone. The resulting composite image is shown in Fig. (5).

Figure 4. Matlab surface plot using imported data from Researcher with image emissivity set to 0.20 and temperature bar (degrees C).

45th AIAA Aerospace Sciences Meeting and Exhibit AIAA 2007-0584 8-11 January 2007, Reno, NV

American Institute of Aeronautics and Astronautics

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In Fig. (5), the neck down area of the cathode shows a sudden increase in temperatures, rather than the expected decrease from conduction along the cathode from the tip to the base, which can be seen along the cathode tip itself. Although not included in this report, the view looking directly into the exit plane does show the expected continuous decrease in temperature from the cathode tip to the cathode mount along the top of the steel portion of the cathode. This suggests that the higher temperatures on the side seen in Fig (5) are due to some interaction with the plume. Maximum temperatures of 550 C are shown at the cathode tip, the portion of the thruster body nearest the cathode, and at the center line of the thruster body near the exit plane. The actual temperatures may be higher since the camera is limited to 550 C without additional hardware. The vertical ring of the thruster near the exit plane appears hotter than the steel axial support bar to which it is attached and does not appear to be conducting heat to the axial support bars. This would suggest it is either thermally isolated or the applied emissivity of 0.20 is incorrect for the ring. The high temperatures seen in this area of the thruster may be from high temperature Xenon neutrals interacting with the thruster surface. The anode cone is of interest as it can be studied to gain insight in to the process of insulator erosion, which is a determining factor for Hall thruster lifetime5. The erosion rate of the cone may be a function of temperature. Failure of this insulator may cause degraded performance or failure if the plasma shunts through the magnetic circuit until the Curie temperature is exceeded5. The cone seen in Fig (5) is somewhat obstructed by the attachment nut for the axial support bar. However, the area of the cone that can be seen has a maximum temperature of 380 C.

V. Conclusion Thermal imaging is a non-intrusive way to measure the temperature of an operating thruster. The raw image can

be used to generally show if a particular area is heating or cooling as the thruster operates. However, to generate the correct temperatures, rather than qualitative data, the emissivity of each area of concern must be known for the particular thruster in question. Gathering data to determine the emissivity of each part of the thruster would be time consuming and somewhat difficult due to the operating temperatures of the thruster. It would be helpful if the

Figure 5. Composite image with corrected emissivity for steel, Alumina plasma spray, and Boron Nitride with temperature bar (degrees C).

45th AIAA Aerospace Sciences Meeting and Exhibit AIAA 2007-0584 8-11 January 2007, Reno, NV

American Institute of Aeronautics and Astronautics

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emissivity of particular parts of a particular thruster was tracked as the thruster accumulated hours. Once this was done, another user could simply apply the values based on the hours of use. Maximum temperature on thruster could be above 550 C at the cathode tip and on the ring near the exit plane. The maximum temperature seen on the anode cone was 380 C.

Acknowledgments The Air Force Institute of Technology would like to thank the Air Force Space Propulsion Lab at Edwards Air

Force Base for material, financial, and personnel support in completing this project. Capt. James W. Tomaszewski would like to thank Dr. William Hargus for sponsoring this work, Taylor Matlock for his extensive help in getting this project done, and Garrett Reed and Lt. Jared Eckholm for setting up the experiment and operating the vacuum chamber. Thanks to Michael Huggins at the Air Force Space Propulsion Lab for providing funding.

References

1Larson, Wiley J. and Wertz, James R., Space Mission Analysis and Design, 3nd ed., Microcosm Press, El Segundo, California, 2005, pp. 702-708.

2Thermovision A40M Operator’s Manual, FLIR Systems, Pub No. 1557813, Rev A72, Issued Oct 29,2004 3Siegel, Robert and Howell, John R., Thermal Radiation Heat Transfer, 4th ed, Taylor & Francis, New York, New York,

2002, Chapter 4. 4Van Nord, J., Kamhawi, H., McEwen, H.K., “Characterization of a High Current, Long Life Hollow Cathode” NASA TM-

214095, 2006 5Hargus, William A., “A Diagnostic for Hall Thruster Boron Nitride Insulator Erosion”, Air Force Research Lab, Edwards

AFB, CA, 2004