Pure Ionic Electrospray Extractor Design ... - Electric Rocket

The 36th International Electric Propulsion Conference, University of Vienna, Austria September 15-20, 2019

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Pure Ionic Electrospray Extractor Design Optimization

IEPC-2019-372

Presented at the 36th International Electric Propulsion Conference University of Vienna • Vienna, Austria

September 15-20, 2019

Robert J. Antypas1 and Joseph J. Wang2

Department of Astronautical Engineering, University of Southern California, Los Angeles, CA, USA

Abstract: The Laboratory of Applied Plasma Dynamics at the University of Southern California has designed, developed and tested an Ion liquid Ion Source (ILIS) operating in the Pure Ionic Regime (PIR) utilizing 1-ethyl-3-methylimidazolium-tetrafluoroborate. The USC Testbed Thruster exploits the repeatability and stability of replaceable 1 cm machined borosilicate emitter and reservoir layers. The workhorse thruster, based on the design of Mike Natisin, employs a convenient modular design which permits the introduction and testing of rapidly developed prototyped parts. Increasing emitter tip density is a critical development to increase the capability and utility of electrospray thrusters. Extractor design and manufacturing is a major limiting factor in the current densification of emitters. At higher density there is simply not enough material between ion grid windows to ensure structural integrity of the extractor at the operating potentials required to emit in the PIR. In an effort to provide the groundwork of the investigation of novel extractor designs the USC Testbed Thruster performance was baselined and is presented here. The USC Testbed Thruster performs well with minimal losses due to grid interception. Four baseline extractor designs were milled with grids of twenty-five 381 μm and 508 μm ion windows.

Nomenclature Vstart = estimated starting potential for emission (V) h = emitter height (μm) Rc = emitter apex radius of curvature (m) ds = emitter apex to extractor grid separation (μm) θ = beam angle measured θeff = effective beam angle θ0 = beam angle offset Vem = emitter voltage (V) qi = species i charge (C) ci = species i particle velocity (m/s) mi = mass of species i (kg) �̇� = mass flow rate (kg/s) �̇� = particle flow rate (particles/s) Va = average beam potential (V) Ib = beam current (A) Iem = emitted current (A) Iex = extractor grid current (A) T = calculated thrust (N) Isp = specific impulse (s) ηprop. = propulsive efficiency (%) j(θ) = collected current density (A/m2)

1 Graduate Student, Department of Astronautical Engineering, [email protected] 2 Professor, Department of Astronautical Engineering, [email protected]


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Lp = distance from the source to the probe (m) ΔV = energy deficit (V)

I. Introduction HE USC electrospray thruster program is based on the design of the U.S. Air Force Research Laboratory’s (AFRL) Air Force Electrospray Thruster II (AFET-II) developed by Mike Natisin1 and based on Dan Courtney’s

novel borosilicate thruster2-4. The USC Testbed Thruster (UTT) is a passively fed electrospray thruster utilizing a porous borosilicate reservoir and emitter layer. It is designed to utilize 1-ethyl-3-methylimidazolium-tetrafluoroborate (EMI-BF4) propellant and operate in the Pure Ionic Regime (PIR). The UTT is a simplified 50% scale AFRL AFET-II, modified to have only 25 emitters on a 1 cm diameter, interchangeable borosilicate emitter disk. The downsizing of the thruster was done in an attempt to fit the thruster in the USC plasma laboratory’s 940-liter test chamber, without saturating the chambers diagnostic and pumping capabilities. At the current pitch the extrapolated thruster tip density is 345 emitters/cm^2. To date, four grid designs have been modeled/constructed and two complete thrusters have been fabricated and tested utilizing one grid design. This report will provide an overview of experimental facilities at USC and the UTT, present results from initial thruster experiments, and highlight future activities to further the advancement of electrospray thruster technology.

II.USC Facilities

A. Overview The USC Laboratory for Plasma Dynamics is a

plasma research laboratory built around a 940-liter vacuum chamber (Fig.1). Historically the laboratory has utilized this chamber alongside a mesothermal plasma source to investigate solar wind dynamics, spacecraft charging/interactions and lunar dust charging/transport5. In late 2018 the focus of LAPD expanded to include electrospray thruster development and testing. The pumping capacity of the chamber was doubled and the diagnostics were overhauled to specifically enable collection of data relevant to electrospray performance. This overhaul maintained the ability to perform the traditional operations with the plasma source while enabling the expansion of research into the ever-growing field of ILIS electrospray technology. The test chamber has three vacuum pumps and can reach pressures of 2x10-7 Torr. A roughing pump seen in the bottom right of Fig. 1 will take the chamber to 20 mTorr. A newly added T-1000 turbo pump can maintain a chamber pressure of less than 1x10-6 Torr for extended periods of time and during thruster operations. If necessary a gate valve, blocking a cryopump, can be opened to pull the chamber down to 2x10-

7 for approximately 4 hours. This pumping capacity has proven adequate and stable during all electrospray testing.

B. Diagnostic Capability In addition to the doubling of pumping capability a new probe

suit was developed specifically for the electrospray program. A Retarding Potential Analyzer (RPA), Faraday Probe (FP), and Micro Faraday Probe (MFP) were all developed specifically for spatial analysis of a non-neutral electrospray plume. Figure 2 shows the probe suite in its final configuration. A Langmuir probe is included in the suite as a legacy diagnostic device for use with the quasi-neutral plasma source. This probe suit is mounted on a 3-axis traversing system that can scan the full volume of the inside of the test chamber. All probes can be utilized simultaneously but the MFP is situated at the bottom of the suite behind the other probes. This was done in an effort to be able to get the MFP as close to the USC Testbed Thruster, during operation, as possible. The MFP was developed with an inner cup diameter of 1.28 mm. This small diameter inner cup, coupled

T

Figure 1. LAPD 940-liter Chamber

Figure 1. LAPD 940-liter Chamber

a) b) Figure 2. USC Data Acquisition Probe Configuration. a) Top left: RPA. Top Centre: Langmuir Probe. Top Right: FP Bottom Center: MFP. b) Top: MFP. Bottom Left: FP. Bottom Right: RPA.


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with close standoff distances, allows the MFP to resolve the current measurements of individual emission sites. The collection surfaces of the entire instrument suite can be biased to reduce secondary electron emission within individual collection devices.6

III.Thruster Design and Fabrication

A. Thruster Body and Propellant Reservoir The USC Testbed Thruster was designed as a ½ scale AFET-II thruster to take advantage of readily available

materials, namely 1 cm P4 and P5 borosilicate disks. The UTT is 2.54 cm L x 2.54 cm W x 1.27 cm tall. An expanded view of the thruster can be seen in Fig. 3. The UTT consists of 5 main parts; the aluminum body, a PEEK propellant housing, a borosilicate reservoir, a borosilicate emitter layer, a distal electrode, and an extractor grid. There is also a tension spring in the propellant reservoir to ensure contact between the emitter layer and the distal electrode is maintained at a constant pressure. Figure 4 shows an assembled thruster. The body of the thruster is made of aluminum and is connected electrically to the extractor grid. The extractor grid is designed to sit at the top plane of the emitter tips. This reduces ds to zero in an ideal configuration. Variation in emitter height due to manufacturing tolerances can cause ds to vary by approximately 24 µm.

The propellant module sits inside the body and is made of Polyether Ether Ketone (PEEK). The high dielectric strength of the PEEK reduces the bulk and weight of the Thruster and allows for tighter tolerances in manufacturing. The modular design of the propellant reservoir allows for efficient propellant loading operations and quick replacement of faulty or broken components. The propellant and borosilicate disks are encased in the propellant module with the distal electrode. This electrode is a 254 μm thin sheet of stainless steel milled to fully surround the emitter plane.

Borosilicate glass was chosen as the reservoir and emitter substrate due to its nonconductive nature and its inherent resistance to electrochemical degradation2. The selection of borosilicate glass leads to the need for a reliable distal electrode to evenly raise the conductive EMI-BF4 propellant to high potentials.

B. Emitter and Extractor The emitter array and extractor are the most critical components in the UTT. They have the tightest tolerance

requirements of 25 μm. Any miss alignment or discrepancies in manufacturing lead to increased particle impingement and a loss of efficiency. The design of the thruster allows for this inefficiency to be directly measured. During normal operations the body is grounded through the data acquisition system. This allows for a direct measurement of current intercepted by the extractor grid. This mode of operation allows for the intercepted current to be turned into a weighting factor in the design and evaluation of extractor grids. The intercepted current is an important metric in the repeatability of individual thruster operation. It is also the defining metric in the comparison of the efficiency of individual thruster designs.

Figure3. USC Thruster V1 Expanded Rendering. Items from bottom to top: Body, Propellent Housing, Compression Spring, Propellent Reservoir, Interface Layer, 25 Emitter Grid, Distal Electrode, Extractor Grid

Figure4. As Built Assembled Thruster


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Four extractors have been conventionally machined with grids of 25 ion windows, measuring 381 μm and 508 μm. Figure 5. Displays the manufactured extractor grids. These grids suffer from the pitfalls of micromachining at such tight tolerances. Figure 5 (a and b) show grids of 508 μm holes with 37.5 μm of material separating the voids. Grids with such tightly packed holes are on the very edge of the tolerance provided by conventionally machining techniques. Figure 5 (d-e) show more conservatively machined grids of 25 381 μm holes. While these grids are easier to manufacture they suffer from higher particle interception due to the reduction in allowable plume angular expansion. A byproduct of the conventional machining process can be minor irregularities in the finished product. These irregularities can be seen in Fig. 5(c) as metal filaments that bridge across individual ion windows. As displayed these filaments would lead to higher current interception and reduced thruster efficiency. More critically, these filaments can lead to a shorting of the thruster if the ends become detached. This shorting will most likely end the life of the thruster if it cannot be removed. To reduce this risk, a process of deburring was utilized to eliminate the filaments from the extractor Fig. 5(f) shows the 381 μm holes after the deburring was performed.

C. Propellant Loading Propellant loading is currently done in

atmosphere with a micro syringe dropper and an Ohaus AP250D scale. The propellant module and the propellant are outgassed for at least 24 hours before the loading operations begin. During the course of this effort, issues with propellant loading of the first thruster led to an optimization of the loading process. Initially the propellant was loaded after the propellant module was fully assembled. The propellant was applied to the emitter surface and subsequently absorbed into the reservoir through capillary forces. This process led to pooling and uneven absorption of propellant into the reservoir layer. This process is displayed in Fig. 6. When loading the propellant through the emitter layer, saturation of the emitters occurred at 24% of the calculated porous maximum allowed. The second thruster was loaded with a modified procedure that proved to show better saturation results. Under this new procedure; the reservoir is filled and weighed before the emitter layer is filled. The distal electrode is installed after all loading operations are completed. This new process enabled an 85% fill rate.

Figure 6. Propellant Loading of First Thruster. Stills from the absorption of the propellant. Saturation and pooling can be seen in the last frame. Yellowing is due to the age of the initial EMI-BF4

Figure 7. Propellant Loading of Second Thruster. Stills from the absorption of the propellant. Even spreading and no pooling can be observed

Figure 5. Extractors. (a) and (b) grids of 25 508 μm holes ,(c) Filament irregularities from the manufacturing process, (d) and (e) grids of 25 381 μm holes, (e) 381 μm holes post deburring process


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IV.Test Results

A. Thruster 1 Testing Thruster 1 was assembled using the 508 μm 25 emitter grid, with the emitter tips centered in the middle of the extractor windows and flush with the bottom plane of the extractor. It was fired in positive ion mode, emitting EMI+, for a duration of only 8 minutes. Startup voltage was approximately 2000 V, 250% of the design Vstart of 800 V. The discrepancies in performance can be attributed to damage to the emitter tips (Fig 8), that happened during installation

of the distal electrode. This damage, alongside the propellant loading problems identified earlier, led to significant shorting and higher operation voltages. Extractor intercepted current was 7% of emitted current. This number is high and would have led to premature failure of the thruster if the testing continued. Considering the severely damaged condition of the emitter tips, the measured intercepted current was lower than what one would expect. An analysis of the startup voltage using Eq. (1) shows that the radius of curvature for the emitters (Rc)

was approximately 20nm. This is significantly more than the design of 5nm and is evidence of the extent of damage to the emitter tips. Figure 9 shows the voltage and current profile of the full duration of the Thruster 1 test set. The test for Thruster 1 consisted of three gradual rises in applied voltage resulting in three thruster startups. The first startup was terminated by an unexpected short at approximately t = 500s. This short is believed to be from liquid bridging between the emitter layer and the extractor grid due to the pooling of propellant after loading operations. The short burnt itself out and the thruster began normal operations for a short period of time during start up 2. A few major shorts in the middle of operations led to the shutdown of the voltage source. The thruster was started for a third time and was subsequently shut down due to the inconsistent operation. Although the test of Thruster 1 was inconsistent and short it showed that the data acquisition system could sufficiently measure the vital parameters needed to investigate future thruster performance, and that the overall design and assembly produced a viable thruster. Measurements taken during Thruster 1 operations were sourced voltage, source current, intercepted current on the extractor, and 3 spatial probe current measurements.

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Figure 9. Source Voltage and Current of Thruster 1 Initial Test. Emitter source voltage (Vem), emitted current (Iem)), and extractor intercepted current (Iex) from the full test campaign of thruster 1

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Figure 8. Thruster 1 Emitter Profile. Irregular emitter tips and a row of clipped emitters


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B. Thruster 2 testing Thruster 2 was assembled using the 508 μm 25 emitter

grid, with an expected distance to the emitter tips of 0m Fig.10. It was fired in negative ion mode, emitting BF4-. The thruster was run intermittently over 3 days for over 2.5 hours of emission time. Thruster 2 showed a 420% improvement in performance and intercepted current over Thruster 1. The thruster operated with an average extractor interception of 1.66% and an initial high of 4.5%. The third operation preformed the most remarkable with a 0.34% average interception over 30 min of 1700 volt emission. Utilizing equations 2-7 below, thruster performance was established over the current operational time of 2.5 hours.

The paper on “Comparing Direct and Indirect Thrust Measurements from Passively Fed Ionic Electrospray Thrusters.3” used a Time of Flight mass spectrometer to determine the percentage of monomers and dimers in the plume of a passively fed borosilicate glass electrospray. Utilizing EMI-BF4 Dan Courtney found a 44% monomer and dimer population with 4% of trimers and 4% of larger mass constituents. This information was used to make performance estimates of the UTT across its operational voltage range of 1300V-1750V Fig 11. With 44% monomer and dimer constituents the Thrust was found to be between 74nN and 2µN and the ISP was estimated to be between 1650 s and 2350 s. This Isp

measurement over estimates performance as it does not account for neutral mass loss from electrochemical breakdown of EMI-BF41.

C. Plume investigation Multiple Faraday Probe and Micro Faraday

Probe measurments were made of Thruster 2’s plume during operation. Figure 12 and 13 show Micro Faraday probe scans of the emission surface of the thruster. Scans wer performed at stand off distances of 4” and 0.75”. Limitations in the 3 axis traversing system and the probe configuration initially led to a minimum scan distance of 4 inches from the truster face. As seen in Figure 12 this scan distance was too far away from the truster to resolve individual emitter sites. For clarification, an image of the thruster has been overlayaid on top of the sensor heatmaps to display the relative position to the scan area.

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Figure 10. Thruster 2 Emitter Profile

Figure 11. Startup Source Voltage v Current of Thruster 2. Initial startup at 1200 V with stable startup of 1300 V and maximum operation of 1750 V

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Figure 13 illustrates better emitter site emission resolution but the plume appears to be saturated even at this

distance. A Closer map with smaller steps sizes is warranted.

Faraday probe data was collected across two of the Thruster 2 operation cycles. Heatmaps of the horizontal plane

with respect to the thruster has been shown in Fig. 14 and 15. These heatmaps are renderings of the plume charge in

the Y (chamber horizontal axis WRT the thruster) as Z is moved away from the thruster. The thruster position relative to the scan area would be centered on the x axis starting at 4 inches away from the thruster. Analysis of the first rendering was initially thought to be erroneous. A second map was taken along with a scan of the vertical plane with respect to the distance from the thruster Fig 16. Although the magnitude is slightly different due to different operating parameters of the thruster, a secondary anode is consistently found at 6 inches from the thruster plane. The XZ plume scan in Fig 16. shows the same secondary anode at a Z distance of 6 inches from the thruster. It is believed that the leading cause of error in these probe scans is the chamber environment. The chamber is a non-coated stainless-steel chamber that is susceptible to very high secondary electron/ion emission under high energy negative ion bombardment7. This diagnostic anomaly seems to be a consequence of the experimental set up and may be unavoidable

Figure 12. Heat Map of MFP with 4” Dis. To Thruster Face. Thruster position is relative to scanning probe location and scaled for distance

Figure 13. Heat Map of MFP with .75” Dis. To Thruster Face. Thruster position is relative to scanning probe location and scaled for distance

Figure 15. FP Heat map of the Horizontal plane. Z distance is distance to the thruster, Y distance is from the center of the thruster face (2.5”)

Figure 14. Second FP Heat map of the Horizontal plane. Z distance is distance to the thruster, Y distance is from the center of the thruster face (2.5”)


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with the current chamber. There is an ongoing effort to model the chamber environment to provide consistency in data analysis across different operating conditions and to further investigate this phenomenon.

V.Future Work The second thruster’s continued operation will be utilized to further baseline performance of the conventional extractor grids. All 4 grids will be fired and baselined on the same thruster. This will eliminate the variability of emitter construction from the analysis of extractor performance. Modeling of the chamber and the non-neutral plasma environment will be performed. This model is needed to understand the formation of secondary anodes and the total flux of secondary electrons and ions. The formation of secondary anodes in the non-monotonic plume needs to be investigated and characterized further in an effort to properly baselined thruster performance in a confined space. The high energy of emitted ions and the small size of the stainless-steel chamber has been shown to lead to up to 4 times the secondary electron and ion emission of other metallic substraits7. This secondary electron emission will compound the formation of secondary anodes when emitting BF4-. A TOF mass spectrometer will be built to verify mass distribution of the emitted plume.3 This TOF will be utilized to characterize performance of individual extractor designs. With a proper TOF Isp estimations will be more realistic. A TOF mass spectrometer will allow for the investigation of the suitability of other novel ionic liquids.

VI.Conclusion The University of Southern California’s Laboratory of Applied Plasma Dynamics has a thriving electrospray

development program. This program in combination with a well-established, physics based, plasma modeling and simulation heritage provide for a unique environment where electrospray physics can be thoroughly investigated and optimized. Heatmap comparison of Micro Faraday Probe measurements from 4” and 0.75” away shows promise as a primary diagnostic tool in resolving the individual emitter/extractor combined performance. Increasing emitter tip density is a critical development needed to increase the capability and utility of electrospray thrusters. USC’s focus on extractor design and quantification of performance will push the boundaries of the major limiting factor in the manufacturing of densified emitters.

Acknowledgments This work was conducted with the technical expertise and best practices provided by the In-Space Propulsion

Branch at the Air Force Research Laboratory. A special thanks to; Mike Holmes, Mike Natisin, Dan Eckhardt, and Justin Koo. This research was made possible by the DOD Science Mathematics and Research for Transformation (SMART) Scholarship program.

Figure 16. FP Heat map of the Vertical plane. Z distance is distance to the thruster, X distance is from the center of the thruster face (4.55”)


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References 1 Natisin, M., Zamora, H. Holmes, M., Eckhardt, D. “Performance of a Fully Conventionally Machined Liquid-Ion Electrospray Operated in PIR,” 36th International Electric Propulsion Conference, IEPC, University of Vienna, Vienna, Austria. September 15-20, 2019 2 Coffman, C., Perna, L., li, H., and Lozano, P. “On the Manufacturing and Emission Characteristics of a Novel Borosilicate Electrospray Source.” 49th AIAA/ASME/SAE/ASEE Joint Propulsion Conference. 2013 3 Courtney, D., Dandavino, S., and Shea, H. “Comparing Direct and Indirect Thrust Measurements from Passively Fed Ionic Electrospray Thrusters.” Journal of Propulsion and Power 32.2 (2016): 392–407 4 Courtney, D., “Ionic liquid ion source emitter arrays fabricated on bulk porous substrates for spacecraft propulsion,” PhD Thesis, Department of Aeronautics and Astronautics, MIT, 2011. 5 Yu, W., Wang, J., & Chou, K. (2015). Laboratory Measurement of Lunar Regolith Simulant Surface Charging in a Localized Plasma Wake. IEEE Transactions on Plasma Science, 43(12), 4175–4181. https://doi.org/10.1109/TPS.2015.2492551 6 Zhang, Z., Tang, H., Zhang, Z., Wang, J., and Cao, S. “A Retarding Potential Analyzer Design for keV-Level Ion Thruster Beams.” Review of Scientific Instruments 87.12 (2016): 123510. 7 Atomic and Ionic Impact Phenomena on Metal Surfaces. Springer-Verlag, 2013. Pg 262, 227 Print.

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