Small Hall Effect Thruster with 3D Printed Discharge Channel: Design and Thrust Measurements

aerospace

Article Small Hall Effect Thruster with 3D Printed Discharge Channel: Design and Thrust Measurements

Ethan P. Hopping 1, Wensheng Huang 2 and Kunning G. Xu 3,*

1 Aerospace Engineering Department, Georgia Institute of Technology, Atlanta, GA 30332, USA; [email protected] 2 NASA Glenn Research Center, Cleveland, OH 44135, USA; [email protected] 3 Mechanical and Aerospace Engineering Department, University of Alabama in Huntsville, Huntsville, AL 35899, USA * Correspondence: [email protected]

Abstract: This paper presents the design and performance of the UAH-78AM, a low-power small Hall effect thruster. The goal of this work is to assess the feasibility of using low-cost 3D printing to create functioning Hall thrusters, and study how 3D printing can expand the design space. The thruster features a 3D printed discharge channel with embedded propellant distributor. Multiple materials were tested including ABS, ULTEM, and glazed ceramic. Thrust measurements were obtained at the NASA Glenn Research Center. Measured thrust ranged from 17.2–30.4 mN over a

discharge power of 280 W to 520 W with an anode ISP range of 870–1450 s. The thruster has a similar performance range to conventional thrusters at the same power levels. However, the polymer ABS and ULTEM materials have low temperature limits which made sustained operation difﬁcult.

Keywords: hall thruster; electric propulsion; additive manufacturing

Citation: Hopping, E.P.; Huang, W.; Xu, K.G. Small Hall Effect Thruster with 3D Printed Discharge Channel: 1. Introduction Design and Thrust Measurements. Hall effect thrusters (HET) are a type of electric propulsion device currently used Aerospace 2021, 8, 227. https:// for satellite propulsion and station-keeping. Small HETs have gained interest in the last doi.org/10.3390/aerospace8080227 20 years with the rise of small satellites. There are now multiple companies such as Apollo Fusion, Orbion, Exotrail, Busek, and Sitael SpA, that develop hall HETs. Notably, SpaceX Academic Editor: Igor Levchenko has successfully launched over 1000 small HETs on their Starlink constellation satellites. One of the advantages of small satellites is their significantly lower cost of design and Received: 22 July 2021 production. However, HETs are still traditionally subtractive manufactured using the same Accepted: 12 August 2021 materials as 30 years ago. One source of manufacturing cost and time for HETs are the Published: 15 August 2021 anode and discharge channel assemblies. In most Hall thrusters, the propellant distributor is integrated into the anode assembly. This requires multi-part fabrication and welding to Publisher’s Note: MDPI stays neutral integrate the baffle assemblies, orifices, and other distributor assemblies into the anode. with regard to jurisdictional claims in The discharge channels of most SPT type HETs are made of boron nitride, a hot-pressed published maps and institutional affil- ceramic that must be subtractively machined from a single block. Monolithic boron nitride iations. dimensions are limited by the hot-pressing process, and this poses challenges for the design of large thrusters [1,2]. In addition, the cost of the boron nitride components increases substantially with thruster size. In other aerospace industries, additive manufacturing (AM), or colloquially 3D print- Copyright: © 2021 by the authors. ing, is being used to dramatically reduce the cost of components [3]. Propulsion systems Licensee MDPI, Basel, Switzerland. are well suited to the AM processes due to their complex geometry and low-volume pro- This article is an open access article duction. In this research, we seek to investigate the use of 3D printing to reduce the cost of distributed under the terms and HET fabrication, with a focus on using low-cost and fast turnaround methods. If AM can conditions of the Creative Commons Attribution (CC BY) license (https:// reduce the part count and simplify the design and fabrication of the thruster, not only will creativecommons.org/licenses/by/ it bring down the cost of the device, but also enable agile development and test programs 4.0/). of new thruster designs.

Aerospace 2021, 8, 227. https://doi.org/10.3390/aerospace8080227 https://www.mdpi.com/journal/aerospace Aerospace 2021, 8, x FOR PEER REVIEW 2 of 15

Aerospace 2021, 8, x FOR PEER REVIEW 2 of 15

ditions of the Creative Commons At- Aerospace 2021, 8, 227 will it bring down the cost of the device, but also enable agile development and test2 pro- of 14 ditions of the Creativetribution Commons (CC BY) At- licensewill (http://crea- it bring downgrams the of costnew ofthruster the device, designs. but also enable agile development and test pro- tribution (CC BY) licensetivecommons.org/licenses/by/4.0/). (http://crea- grams of new thruster designs. tivecommons.org/licenses/by/4.0/). 2. Materials and Methods 2. Materials and Methods 2. Materials and2.1. MethodsThruster Design 2.1. Thruster Design 2.1. Thruster DesignThe UAH-78AM is an annular HET designed to fit in CubeSat dimension, 10 × 10 × × × The UAH10-78AM cm.The The is UAH-78AM anthruster annular is is aHET an loose annular designed scaling HET ofto designed thefit inchannel CubeSat to fit inof CubeSatthedimension, P5 HET dimension, [4]10 .× The 1010 ×magnetic 10 10 field cm. The thruster is a loose scaling of the channel of the P5 HET [4]. The magnetic field topology 10 cm. The thrustertopology is a is loose copied scaling from of the the P5. channel It has aof design the P5 discharge HET [4]. Thepower magnetic of 300 –field500 W. Images of is copied from the P5. It has a design discharge power of 300–500 W. Images of the thruster topology is copiedthe thruster from the are P5. provided It has a designin Figure discharge 1, and anpower isometric of 30 0view–500 W.of the Images design of tested at the are provided in Figure1, and an isometric view of the design tested at the NASA Glenn the thruster areNASA provided Glenn in Research Figure 1 , Center and an ( GRCisometric) is provided view of the in Figuredesign 2tested. The at “AM” the designation Research Center (GRC) is provided in Figure2. The “AM” designation stands for additively NASA Glennstands Research for additively Center (GRC manufactured,) is provided and in theFigure “78” 2 is. Thethe diameter “AM” designation of the outer channel in manufactured, and the “78” is the diameter of the outer channel in millimeters. stands for additivelymillimeters. manufactured, and the “78” is the diameter of the outer channel in millimeters.

Figure 1. Testing of the UAH-78AMUAH-78AM at GRC. Figure 1. Testing of the UAH-78AM at GRC.

Figure 2. 3D rendering of UAH-78AM configuration as tested at GRC and a cross -section of the Figure 2. 3DFigure rendering 2. 3D of rendering UAH-78AMdischarge of channel configurationUAH- 78AMwith embedded configuration as tested propellant at GRC as tested and distributor a at cross-section GRC andand innera ofcross the and- dischargesection outer channelof channelthe walls. with embedded propellantdischarge distributorchannel with and embedded inner and propellant outer channel distributor walls. and inner and outer channel walls. 2.1.1. 3D Printed Discharge Channel and Materials 2.1.1. 3D Printed Discharge Channel and Materials 2.1.1. 3D Printed DischargeThe discharge Channel channel and withMaterials embedded propellant distributor was the focus of the 3D The dischargeprinting.The channel dischargeThis withwas primarilyembedded channel with due propellant embeddedto the lack distributor propellantof accessibl was distributore theprinting focus wasofof ferromagneticthe the 3D focus of themate- 3D printing. Thisrialsprinting. was suchprimarily Thisas iron was due or primarily tolow the carbon lack due of steel. to accessibl the In lack SPTe of printing-type accessible Hall of thrusters, printingferromagnetic of the ferromagnetic discharge mate- channel materials is rials such as ironmadesuch or as oflow iron ceramic carbon or low materials steel. carbon In SPT steel.with-type Inhigh SPT-type Hall secondary thrusters, Hall thrusters,electron the discharge emission the discharge channel (SEE) channel .is The high is made SEE made of ceramicreducesof ceramic materials the materials average with high electron with secondary high temperature secondary electron electronin emissionthe plasma, emission (SEE which). (SEE). The increases high The highSEE ionization SEE reduces effi- reduces the averagethe average electron electron temperature temperature in the in plasma, the plasma, which which increases increases ionization ionization effi- efficiency [5–7]. This means the material selection is limited to dielectrics with high SEE. Polymers provide SEE profiles similar to ceramics, but they also present design challenges due to low melting temperatures.

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ciency [5–7]. This means the material selection is limited to dielectrics with high SEE. Pol- Aerospaceymers2021, 8, 227provide SEE profiles similar to ceramics, but they also present design challenges 3 of 14 due to low melting temperatures. There is little information in the literature on the SEE characteristics of polymers in the incident electron temperatureThere is little range information of Hall in thethrusters. literature This on the is SEElikely characteristics because most of polymers SEE in data is intended forthe scanning incident electron temperature microscope range applications, of Hall thrusters. where This minimum is likely because primary most SEE data is intended for scanning electron microscope applications, where minimum primary electron energy is on the order of 100 eV or more. An estimate for the SEE characteristics electron energy is on the order of 100 eV or more. An estimate for the SEE characteristics of of polymers in the 0polymers–100 eV inrange the 0–100 was eV thus range made was thusby extrapolating made by extrapolating the attenuation the attenuation model model by by Cazaux [8]. The CazauxSEE results [8]. The of SEE the results model of for the modelNylon for-12 Nylon-12 and PTFE and PTFEare presented are presented in inFig- Figure3 ure 3 along with boronalong nitride with boron and nitride Borosil and ceramics Borosil ceramics from by from Goebel by Goebel and and Katz Katz [9] [9.].

Figure 3. 3.Extrapolated Extrapolated secondary secondary electron electron emission emission as a function as ofa function temperatures of temperatures for selected polymers for selected below 100 poly- eV. mers below 100 eV. As the figure shows, the SEEs for both polymers are comparable to the ceramics from As the figure shows,approximately the SEE 40–100s for both eV. At polymers energies below are comparable 40 eV, the ceramics to the have ceramics higher SEE, from but the difference is less than a factor of two. These results indicate that polymer SEE profiles are approximately 40–100comparable eV. At energies to common below HET ceramic 40 eV, materialthe ceramics in the electronhave higher temperature SEE, but range the seen in difference is less thanthrusters. a factor Data of fortwo. ABS These and ULTEM results polymer indicate were that not polymer available, thusSEE could profiles not be are directly comparable to commoncompared, HET but ceramic we expected material their SEEs in the to be electron similar to temperature Nylon and PTFE range as they seen have in similar thrusters. Data for ABSchemical and composition.ULTEM polymer were not available, thus could not be directly The discharge channel was printed in three parts: the rear of the channel with em- compared, but we expecbeddedted propellant their SEE distributor,s to be similar the inner to channelNylon wall,and PTFE and the as outer they channel have sim- wall. The ilar chemical composition.embedded propellant distributor allowed us to use a simple stainless-steel ring as the The discharge anode,channel thus was separating printed propellant in three injection parts: from the therear anode. of the All channel parts were with initially em- printed bedded propellant withdistributor, ABS plastic. the However, inner channel testing revealed wall, and that thethe ABS’s outer relatively channel low wall. glass Th transitione temperature of 105 ◦C limited the lifetime of the channel. The plasma heating caused embedded propellant distributor allowed us to use a simple stainless-steel ring as the an- charring and melting of the channel, especially at the exit plane where the magnetic field ode, thus separatingwas propellant strongest andinjection thus has from the highestthe anode. temperature. All parts Higher were temperature initially printed 3D printable with ABS plastic. However,dielectric materialstesting revealed were considered. that the We ABS’s settled relatively on two materials: low glass ULTEM transition with a glass ◦ temperature of 105 °Ctransition limited temperature the lifetime of ~200 of theC, channel. and a glazed The ceramic. plasma heating caused charring and melting of the Separatechannel, 3D especially printers were at the used exit for plane ULTEM where and ABS the parts.magnetic Part accuracyfield was for the ULTEM printer is ±0.130 mm or better. Accuracy for the ABS 3D printer is more difficult to strongest and thus has the highest temperature. Higher temperature 3D printable dielectric materials were considered. We settled on two materials: ULTEM with a glass transition temperature of ~200 °C, and a glazed ceramic. Separate 3D printers were used for ULTEM and ABS parts. Part accuracy for the UL- TEM printer is ±0.130 mm or better. Accuracy for the ABS 3D printer is more difficult to

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predict as this was not a commercial 3D printer. Therefore, factors such as belt backlash and part shrinkage are not taken into account when quoting accuracy. However, axis resolution is 0.01 mm, and since the 3D printing technology is functionally identical to the ULTEM 3D printer, part accuracy is likely comparable. The smallest features in our parts were the propellant distributor orifices, which had a diameter of 0.01 in (0.254 mm). Light sanding was used for part cleanup on polymer components in some areas to improve fit. The glazed ceramic, which is similar to pottery, was produced by an external vendor, i.Materalize. The ceramic was not actually 3D printed, but rather the mold was printed in plastic and then filled with a liquid ceramic mixture. After drying and hardening, the ceramic part undergoes a firing and glazing process to reach the finished state. The firing process naturally induces part shrinkage on the order of 3% of total part size and was accounted for in the design. In addition, minimum feature size was limited to approximately 2 mm by the vendor due to the need for the liquid ceramic to fill all spaces. Due to this large minimum feature size, the propellant distributor was not made in ceramic due to the more stringent tolerances required for the part. The propellant distributor (the rear of the discharge channel) also did not experience significant thermal heating during the initial tests, thus, was kept in ABS. The final thruster that was performance tested consisted of an ABS rear channel/propellant distributor, a glazed ceramic inner wall, and an ULTEM outer wall. This mixture allowed us to test the different materials, mainly the ceramic and ULTEM in a single set of tests.

2.1.2. Magnetic Field Design The UAH-78AM’s magnetic field is based on the design heritage from the P5. The front and back magnetic poles were changed to squares instead of circles to emulate the CubeSat geometry. The magnetic poles were made of low carbon 1018 steel and traditionally manufactured. Five 1018 steel rods were used to connect to front and back poles and complete the magnetic circuit. The magnetic field was simulated in both Finite Element Method Magnetics software, and Ansys to match the P5 fields. An inner electromagnet was wrapped around the center magnetic pole (not shown in Figure2), and an outer electromagnet wrapped on the outside of discharge channel (the copper-colored part in Figure2). To limit independent variables in testing, the inner and outer magnet currents were held fixed at 4.09 A for all tests to generate the desired channel magnetic field.

2.1.3. Cost and Turnaround Time As one of the goals of this work is to reduce the cost and fabrication time of small HETs, Table1 provides a cost breakdown for the UAH-78AM, in USD. All materials for the thruster in the United States can be procured for a total of $300 or less. This cost assumes the availably of 3D printers and other machining equipment.

Table 1. UAH-78AM cost breakdown.

Material Component Cost Notes Carbon Steel Magnetic Circuit $57 Fasteners Magnetic Circuit $30 Magnet Wire Magnetic Circuit $30 Carbon Shim Stock Magnetic Shields $15 Fabricated from stainless steel Stainless Steel Anode $4 washer 3D printed glazed ceramic Inner Channel $21 Quote from manufacturer ULTEM Outer Channel $97 Quote from manufacturer ABS Inner Channel $2 By Volumetric Material Cost Material Total $256 Labor $560 35 $/h for 16 h Total $816

Costs for the ULTEM outer channel and glazed ceramic inner channel are based on quotes directly from vendor. Consequently, these costs are signiﬁcantly inﬂated as Aerospace 2021, 8, 227 5 of 14

compared to the true costs of materials and print time. It is increasingly common for academic institutions to have access to 3D printing services on campus or through business partnerships. These services frequently provide print services at material cost or less, so it is possible that the inner and outer channel components could be procured for much lower cost. ABS 3D printing is so broadly available through professional and hobbyist services that we provide the component price based on volumetric material cost. The most significant labor costs are in fabrication of the magnetic circuit, which is cut using a conventional subtractive machining process. However, significant efforts were made in the design of the thruster to simplify machining operations as much as possible. Machining for the magnetic circuit is dominated by hole processes. While access to CNC machining simplifies manufacturing, all parts could be produced with relative ease using manual machines. We estimate that total labor time for a skilled machinist on magnetic circuit fabrication would be a day or two. However, labor remains the costliest portion of UAH-78AM procurement assuming an hourly rate of $35 per hour. Experience from testing demonstrates that the thruster can be assembled from raw components in a week or less. In addition, the turnaround time for servicing between tests is on the order of a couple of days. In comparison, a first order cost estimate is shown in Table2 below if the UAH-78AM were to be fabrication using traditional manufacturing methods and with a traditional baffled anode that also functions as the propellant distributor.

Table 2. UAH-78AM cost breakdown.

Material Component Cost Notes Carbon Steel Magnetic Circuit $57 Fasteners Magnetic Circuit $30 Magnet Wire Magnetic Circuit $30 Carbon Shim Stock Magnetic Shields $15 Thicker stock material to Anode/Propellant Stainless Steel $50 fabricate the bafﬂed anode Distributor for propellant distribution Estimated from prior Boron Nitride Channel Discharge Channel $1060 purchases of boron nitride for larger thrusters Material Total $1242 35 $/h for 10 days, anode and magnetic circuit Labor $3800 fabrication + $1000 for oriﬁce drilling Total $5042

2.2. Test Facility Tests were conducted in Vacuum Facility 8 at NASA Glenn Research Center. The main chamber of VF-8 has a diameter of 1.5m and a length of 4.5 m. Pumping is provided by four oil-diffusion pumps with a speed of 1.2 × 105 L per second at 10−5 torr [10]. VF-8 features two bell-jars that can be independently isolated from the main chamber using gate valves. The thruster was mounted on an inverted-pendulum thrust stand attached to the vacuum flange of the primary bell jar. The design and operation of this type of thrust stand is well established in the literature, and further details on the design of similar stands at Glenn are provided in [11]. Anode and cathode propellant flow were provided by 100 sccm and 25 sccm mass flow controllers manufactured by Celerity, and all tests were run using Xenon. A BaO cathode was used with a fixed flow rate of 0.5 mg/s for all tests. The cathode was oversized for the anode current required to sustain thruster discharge. Consequently, the discharge current was too low to allow for self-heating, thus requiring the cathode Aerospace 2021, 8, 227 6 of 14

heater to run at half-power during tests to ensure stable operation. Since cathode flow is not optimized, all specific impulses are presented in terms of the anode flow. It is likely that cathode flow could be reduced to 6–7% of anode flow while maintaining stable discharge.

3. Results and Discussion The focus of the UAH-78AM testing was on powers and discharge voltages that could be sustained by small satellites. Furthermore, the low melting temperatures of the polymer materials limited operation at the higher power end due to higher channel wall heating [12]. The testing at GRC focused on identifying stable low-voltage operating points and flow rates and obtaining baseline performance parameters. It was found that 200 V and 1.82 mg/s anode flow produced a stable discharge at our chosen magnet settings. The thruster was run at points barely in the jet mode of hall thruster discharge, where efficiency is significantly improved from a diffuse mode but power requirements remain low [13]. Therefore, once the 200 V operating point was identified, discharge voltage was increased in 20 V increments from 180–260 V, and anode flow was increased in 0.18 mg/s increments between 1.64 and 2.18 mg/s. Initial thrust measurements were taken in 5 s test intervals to limit thruster heating. After the 5 s data were collected for the desired test matrix, the operating time was increased to 15 s to get closer to steady-state behavior. Several thrust measurements were taken over longer durations on the order of 30 s to assess thrust stability in longer tests.

3.1. General Characteristics A characteristic raw thrust curve is presented in Figure4. This thrust measurement is from a 30-s test at a 200 V discharge and anode ﬂow of 1.82 mg/s. Thrust stabilized at approximately 19 mN, but gradually climbed until the 30-s test conclusion. The increase in thrust is likely due to outgassing that increases as the polymer thruster components are heated. Similar behavior is seen in conventional HETs as water vapor and other volatile compounds outgas from the channel walls. The outgassing process is thought to modify the secondary electron behavior of the channel walls in boron nitride thrusters, resulting in artiﬁcially high discharge currents [14]. In an effort to reduce moisture content, we performed a 12-h low-temperature bake-out on the ABS components. The ULTEM outer channel was manufactured in a heated build chamber with temperatures approaching 200 ◦C, so it is expected that the moisture content in this component was low. The outgassing processes may also be more complex in the UAH-78AM than evapora- tion of surface water or other contaminants from exposure to ambient air. What we refer to as outgassing may be a combination of sublimation and/or ablation processes that occur as the channel wall surface is heated beyond the glass transition temperature of the polymer. Chemical processes could also be playing a role, as the polymer gases could decompose into their constituent atoms as they diffuse into the channel. For lack of a better term, we will continue to refer to the process as outgassing; however, the physical process may be more complex than what is encountered in conventional Hall thrusters. Aerospace 2021, 8, x FOR PEER REVIEWAerospace 2021, 8, 227 7 7of of 15 14

Figure 4. ThrustFigure as 4. aThrust function as a function of time of timeat 200 at 200 V Vdischarge discharge voltage voltage and and anode anode ﬂow of flow 1.82 mg/s. of 1.82 mg/s.

The use of polymer materials in the thruster also presented unique challenges for The outgassing processesmeasuring thrust. may The also heating be more of the polymercomplex limited in the testing UAH duration-78AM to 30 s.than Beyond evapo- 30 s, a ration of surface waterfailure or other mode wascontaminants observed where from a hotspot exposure attached to to theambient outer ULTEM air. What discharge we channel refer to as outgassing maywall be a and combination the thruster entered of sublimation a current-limited and/or mode ablation of operation. processes Hotspot formation that occur can be a problem in conventional Hall thrusters [1], but the cause is different as compared as the channel wall surfaceto the UAH-78AM. is heated Hotspots beyond are the especially glass transition troubling with temperature our thruster since of the the hotspotpoly- mer. Chemical processesincreased could the polymeralso be outgassingplaying a rate role, which as encouragedthe polymer continued gases spot could formation decom- and pose into their constituentgrowth. atoms If left unchecked, as they diffuse the hotspot into permanently the channel. damages For the lack discharge of a better channels, term, and the polymer components must be replaced. An image of the spotting behavior is provided we will continue to referin Figure to the5. process as outgassing; however, the physical process may be more complex than whatWithout is encountered the ability to operate in conventional the thruster at steady-state, Hall thrusters. we choose to characterize The use of polymerperformance materials by comparing in the thrustthruster and specificalso presented impulse at fixed unique times challenges after ignition. Thefor ignition event is identified through a derivative approach. The derivative is taken of the measuring thrust. Thethrust heating trace andof the time polymer zero is identified limited as thetesting location duration where the to thrust 30 s. rate Beyond of change 30 s, a failure mode wasexceeds observed 20 mN/s, where as this a behaviorhotspot is attache only seend during to the thruster outer ignition. ULTEM For discharge the 5-s tests, channel wall and thethrust thruster and specific entered impulse a current are measured-limited 4 s aftermode the of ignition operation. event. For Hotspot the 15-s tests,for- measurements are taken 14 s after ignition. We expected thruster thermal conditions to be mation can be a problemsimilar in at fixedconventional times after ignition, Hall enablingthrusters comparison [1], but across the differentcause dischargeis different voltages as compared to the UAHand-78AM. flow rates. Hotspots However, the are heating especially rate and thermaltroubling condition with of theour channel thrust likelyer since varies the hotspot increasedbetween the polymer tests, resulting outgassing in some errorrate inwhich repeatability. encourage We attemptedd continued to characterize spot for- this mation and growth. uncertaintyIf left unchecked, through repeated the tests,hotspot which permanently are discussed later damages in the results. the Calibration discharge and measurement uncertainties associated with the equipment were calculated and reported channels, and the polymeraccording components to the best practices must identified be replaced. in Refs. [15 An,16]. image While thrust of the and spottin ISP uncertaintyg behavior is provided invary Figure slightly 5. with the calibration for each test run, the average thrust uncertainty at 95% confidence is ±0.72 mN, and ISP uncertainty is ±40 s.

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FigureFigure 5.5.Image Image ofof hotspothotspot formationformation andand corresponding corresponding decrease decrease in in thrust. thrust.

3.2. ThrustWithout the ability to operate the thruster at steady-state, we choose to characterize performanceFigure6 show by comparing the thrust asthrust a function and specific of anode impulse discharge at fixed voltage times and after anode ignition. mass ﬂow The rateignition for the event 5- and is identified 15-s test runs. through The a general derivative trends approach. are as expected The derivative for HETs, is tak withen thrust of the Aerospace 2021, 8, x FOR PEER REVIEW 9 of 15 increasingthrust trace with and both time discharge zero is voltageidentified and as ﬂow the rate.location The where thrust variedthe thrust from rate 17.2–30.4 of change mN overexceeds the range20 mN/s, of conditions as this behavior tested. is only seen during thruster ignition. For the 5-s tests, thrust and specific impulse are measured 4 s after the ignition event. For the 15-s tests, measurements are taken 14 s after ignition. We expected thruster thermal conditions to be similar at fixed times after ignition, enabling comparison across different discharge voltages and flow rates. However, the heating rate and thermal condition of the channel likely varies between tests, resulting in some error in repeatability. We attempted to characterize this uncertainty through repeated tests, which are discussed later in the results. Calibra- tion and measurement uncertainties associated with the equipment were calculated and reported according to the best practices identified in Ref [15,16]. While thrust and ISP uncertainty vary slightly with the calibration for each test run, the average thrust uncertainty at 95% confidence is ±0.72 mN, and ISP uncertainty is ±40 s.

3.2. Thrust Figure 6 show the thrust as a function of anode discharge voltage and anode mass flow rate for the 5- and 15-s test runs. The general trends are as expected for HETs, with thrust increasing with both discharge voltage and flow rate. The thrust varied from 17.2— (30.4a) mN over the range of conditions tested. (b)

FigureFigure 6.6.Thrust Thrust asas aa functionfunction ofof dischargedischarge voltage voltage and and anode anode mass mass ﬂow flow rate rate from from ( a(a)) the the 5-s 5-s tests tests and and ( b(b)) the the 15-s 15-s tests. tests.

RepeatRepeat tests tests were were done done for for the the 5-s 5- datas data at at 180, 180, 200, 200, and and 220 220 V. V. At At 180 180 V, V, the the 2.00 2.00 mg/s mg/s repeatrepeat tests tests display display a verticala vertical spread spread of approximatelyof approximately 5 mN. 5 mN. The The error error bars bars in the in figurethe figure do notdo account not account for the for differences the differences in thrust in thrust for the for repeat the repeat tests. tests. During During testing, testing, it was it noticed was no- thatticed at 180that V at the 180 thruster V the tookthruster longer took to longer start and to forstart the and discharge for the todischarge settle. It isto suspectedsettle. It is thatsuspected 180 V is that the lower180 V limitis the of lower the jet-mode limit of dischargethe jet-mode with discharge the fixed with magnet the settingsfixed magnet and thesettings 2.00 mg/s and anodethe 2.00 flow mg/s rate, anode thus resultingflow rate, in thus high resulting variability. in high The repeat variability. tests at The 200 repeat and 220tests V fallat 200 within and the220 equipment V fall within error, the indicating equipment that error, the shortindicating duration that teststhe short can produce duration consistenttests can measurementsproduce consistent at voltages measurements above the at limit. voltages At 1.64 above mg/s, the thelimit. thruster At 1.64 could mg/s, not the bethrust starteder could at 180 not or 200be started V without at 180 adjusting or 200 V magnet without settings. adjusting magnet settings. While collecting 15-s data at 220 V and 2.00 mg/s, a spot formed and attached to the outer channel wall. The thruster was turned off after spot formation and allowed to cool before starting again. A few more points were taken after the spot formation and are labeled with open circles in all plots. The damage to the ULTEM outer channel wall associated with spot formation could change the performance of the thruster by increasing anode leakage current. This would limit comparisons with preceding operating points. However, the data is included for completeness since the thruster was able to start and run without further spot formation after the channel had been allowed to cool. Due to the spot formation, 15-s tests were not done at 2.18 mg/s anode flow to prevent further heating of the channel wall as heating rate increases at higher discharge powers, which are associated with higher flow rates [12].

3.3. Anode Specific Impulse

The anode specific impulse (ISP) is shown in Figure 7. The anode ISP is calculated with

ITSP / m a g0  , where ma is the anode mass flow rate and g0 is Earth sea-level gravity.

The ISP ranges from 870–1450 s and increases with discharge voltage as expected. The data also suggest that ISP increases with flow rate; however, the differences are within the measuring uncertainty.

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While collecting 15-s data at 220 V and 2.00 mg/s, a spot formed and attached to the outer channel wall. The thruster was turned off after spot formation and allowed to cool before starting again. A few more points were taken after the spot formation and are labeled with open circles in all plots. The damage to the ULTEM outer channel wall associated with spot formation could change the performance of the thruster by increasing anode leakage current. This would limit comparisons with preceding operating points. However, the data is included for completeness since the thruster was able to start and run without further spot formation after the channel had been allowed to cool. Due to the spot formation, 15-s tests were not done at 2.18 mg/s anode ﬂow to prevent further heating of the channel wall as heating rate increases at higher discharge powers, which are associated with higher ﬂow rates [12].

3.3. Anode Speciﬁc Impulse

The anode specific impulse (ISP) is shown in Figure7. The anode ISP is calculated with . . ISP = T/ mag0 , where ma is the anode mass flow rate and g0 is Earth sea-level gravity. The ISP ranges from 870–1450 s and increases with discharge voltage as expected. The Aerospace 2021, 8, x FOR PEER REVIEW 10 of 15 data also suggest that ISP increases with flow rate; however, the differences are within the measuring uncertainty.

(a) (b)

FigureFigure 7. 7.(a )(a Anode) Anode specific specific impulse impulse as as a functiona function of of discharge discharge voltage voltage and and anode anode mass mass flow flow rate rate for for 5-s 5- tests;s tests; (b ()b Anode) Anode specificspecific impulse impulse for for 15-s 15- tests.s tests.

3.4.3.4. Anode Anode Efficiency Efficiency . 2 2 The thrust anode efficiency, calculated from ηa = T / 2maPd , where Pd isd the dis- The thrust anode efficiency, calculated from aa TP/2 m d , where P is the discharge power is shown in Figure8. The anode efficiency neglects additional power sources charge power is shown in Figure 8. The anode efficiency neglects additional power such as the cathode and magnets. Anode efficiency increased with discharge voltage and sources such as the cathode and magnets. Anode efficiency increased with discharge volt- varied from 27.8–42.2%, with a few exceptions. In the 5-s tests at 1.82 mg/s, the efficiency age and varied from 27.8–42.2%, with a few exceptions. In the 5-s tests at 1.82 mg/s, the was slightly lower at 260 V compared to 240 V. The same occured in the 15-s tests at efficiency was slightly lower at 260 V compared to 240 V. The same occured in the 15-s 2.00 mg/s. The latter measurement was taken after the spot formation, as indicated by tests at 2.00 mg/s. The latter measurement was taken after the spot formation, as indicated the open circles. Thus, the loss in efficiency may be a product of increased anode leakage by the open circles. Thus, the loss in efficiency may be a product of increased anode leak- current due to the damage to the outer channel wall. age current due to the damage to the outer channel wall.

(a) (b) Figure 8. (a) Anode efficiency as a function of discharge voltage and anode mass flow rate for 5-s tests; (b) anode efficiency for the 15-s tests.

The effect of flow rate on efficiency is unclear from the data due to the measurement uncertainty. In the 15-s tests, it appears that the 1.64 mg/s flow rate resulted in higher efficiencies at discharge voltages above 200 V, while 2.00 mg/s was more efficient at the lower voltages. The 5-s data suggest that 2.18 mg/s flow resulted in the highest anode efficiency at all operating points except 180 V. No pattern is clearly discernable from these results, and the measurement uncertainties limit the significance of any identified trends with respect to flow rate.

Aerospace 2021, 8, x FOR PEER REVIEW 10 of 15

(a) (b) Figure 7. (a) Anode specific impulse as a function of discharge voltage and anode mass flow rate for 5-s tests; (b) Anode specific impulse for 15-s tests.

3.4. Anode Efficiency

2 d The thrust anode efficiency, calculated from aa TP/2 m d , where P is the discharge power is shown in Figure 8. The anode efficiency neglects additional power sources such as the cathode and magnets. Anode efficiency increased with discharge voltage and varied from 27.8–42.2%, with a few exceptions. In the 5-s tests at 1.82 mg/s, the efficiency was slightly lower at 260 V compared to 240 V. The same occured in the 15-s tests at 2.00 mg/s. The latter measurement was taken after the spot formation, as indicated Aerospace 2021, 8, 227 10 of 14 by the open circles. Thus, the loss in efficiency may be a product of increased anode leakage current due to the damage to the outer channel wall.

(a) (b)

FigureFigure 8. 8.(a ()a Anode) Anode efficiency efficiency as as a a function function of of discharge discharge voltage voltage and and anode anode mass mass flow flow rate rate for for 5-s 5- tests;s tests; (b ()b anode) anode efficiency efficiency for the 15-s tests. for the 15-s tests.

TheThe effect effect of of flow flow rate rate on on efficiency efficiency is is unclear unclear from from the the data data due due to to the the measurement measurement uncertainty.uncertainty. In In the the 15-s 15-s tests, tests, it it appears appears that that the the 1.64 1.64 mg/s mg/s flow flow rate rate resulted resulted in in higher higher efficienciesefficiencies at at discharge discharge voltages voltages above above 200 200 V, V, while while 2.00 2.00 mg/s mg/s was was more more efficient efficient at at the the lowerlower voltages. voltages. The The 5-s 5-s data data suggest suggest that that 2.18 2.18 mg/s mg/s flow flow resulted resulted in in the the highest highest anode anode efficiency at all operating points except 180 V. No pattern is clearly discernable from these Aerospace 2021, 8, x FOR PEER REVIEWefficiency at all operating points except 180 V. No pattern is clearly discernable from these11 of 15 results,results, and and the the measurement measurement uncertainties uncertainties limit limit the the significance significance of of any any identified identified trends trends withwith respect respect to to flow flow rate. rate.

3.5.3.5. Thrust Thrust to to Power Power Ratio Ratio Finally,Finally the, the thrust-to-power thrust-to-power ratio ratio is is shown shown in in Figure Figure9. 9 The. The thrust thrust to to power power ratio ratio is is generallygenerally seen seen to to decrease decrease with with increasing increasing discharge discharge voltage, voltage, with with a a peak peak at at low low voltages. voltages. AsAs with with anodeanode efﬁciency,efficiency, uncertaintiesuncertainties limit the the ability ability to to draw draw conclusions conclusions about about the the re- relationshiplationship between between anode anode flow ﬂow and and thrust thrust to to power power ratio. ratio.

(a) (b)

FigureFigure 9. 9.(a ()a Thrust-to-power) Thrust-to-power ratio ratio as as a a function function of of discharge discharge voltage voltage and and anode anode mass mass ﬂow flow rate rate for for 5-s 5-s tests; tests; ( b(b)) thrust-to- thrust-to- powerpower ratio ratio for for the the 15-s 15- test.s test.

3.6.3.6. Repeatability Repeatability WhileWhile measurement measurement uncertainty uncertainty due due to to equipment equipment error error is is quantiﬁed, quantified, a a more more signif- signif- icanticant challenge challenge with with transient transient testing testing is is verifying verifying that that our our methodology methodology for for comparing comparing performanceperformance at at different different operating operating points points is valid. is valid. The transient The transient performance performance of the thruster of the thruster is sensitive to the thermal condition of the channel; therefore, we chose to compare performance at points where the thermal condition of the channel should be similar by looking at a fixed time after ignition. However, the heating rate and wall power losses in Hall thrusters also vary depending on the discharge voltage, since this affects electron temperature and consequently channel wall losses [9,12]. This effect could be a source of error since it could cause the outgassing to vary between tests. Furthermore, the vacuum environment limits heat dissipation mechanisms, causing the channel to take a long time to return to ambient temperature after a test. In order to minimize the amount of propellant used for running the cathode, tests were run in sequences of four or five with five minutes between tests to allow the thruster to cool. For the 5-s tests, the pauses were adequate to allow the channel to cool between tests. How- ever, the pauses were not adequate during the 15 s tests, and the cumulative heating from running at higher discharge powers eventually led to spot formation on the outer channel wall. In an effort to characterize the impact channel temperature may have, we repeated tests at several operating points. Figure 10 shows four repeat thrust measurements at 200 V and 1.82 mg/s, which was the most-tested operating point.

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is sensitive to the thermal condition of the channel; therefore, we chose to compare performance at points where the thermal condition of the channel should be similar by looking at a fixed time after ignition. However, the heating rate and wall power losses in Hall thrusters also vary depending on the discharge voltage, since this affects electron temperature and consequently channel wall losses [9,12]. This effect could be a source of error since it could cause the outgassing to vary between tests. Furthermore, the vacuum environment limits heat dissipation mechanisms, causing the channel to take a long time to return to ambient temperature after a test. In order to minimize the amount of propellant used for running the cathode, tests were run in sequences of four or five with five minutes between tests to allow the thruster to cool. For the 5-s tests, the pauses were adequate to allow the channel to cool between tests. However, the pauses were not adequate during the 15 s tests, and the cumulative heating from running at higher discharge powers eventually led to spot formation on the outer channel wall. In an effort to characterize the impact channel temperature may have, we repeated Aerospace 2021, 8, x FOR PEER REVIEW 12 of 15 tests at several operating points. Figure 10 shows four repeat thrust measurements at 200 V and 1.82 mg/s, which was the most-tested operating point.

Figure 10. Repeat tests at 200 V and 1.82 mg/s anode mass flow. Figure 10. Repeat tests at 200 V and 1.82 mg/s anode mass ﬂow.

TheThe initial initial thermal thermal condition condition ofof thethe channelchannel waswas differentdifferent forfor eacheach ofof thethe repeatrepeat tests.tests. ForFor example, example, Test Test 2 2 was was the the last last of of a a sequence sequence of of four four 5-s 5-s tests, tests, so so the the channel channel temperature temperature shouldshould havehavebeen been elevated elevated above above ambient ambient conditions. conditions. TestsTests 1–3 1–3 show show strong strong agreement agreement andand convergeconverge onon aa thrustthrust ofof approximately approximately 19.519.5 mN.mN. TestTest 4 4 is is noticeably noticeably lower lower than than the the firstfirst three three tests. tests. However, However, test test 4 4 is is a a final final 30-s 30-s run, run, and and the the thrust thrust measurement measurement eventually eventually convergesconverges on on a a value value within within 0.2 0.2 mN mN of of the the first first three three tests. tests. The The behavior behavior suggests suggests that that the the outgassingoutgassing contributioncontribution toto thrustthrust approachesapproaches aa steady-statesteady-state valuevalue thatthat is is stable stable over over our our shortshort duration duration teststests before before spotting spotting occurs. occurs. SimilarSimilar behaviorbehavior waswas seen seen with with other other repeat repeat tests,tests, providingproviding somesome confidenceconfidence thatthat transient transient thruster thruster performance performance cancan be be compared compared acrossacross operating operating points. points.

3.7.3.7. ComparisonComparison withwith Other Other Low-Power Low-Power Hall Hall Thrusters Thrusters TableTable3 3provides provides a a comparison comparison of of the the performance performance parameters parameters collected collected from from the the UAH-78AMUAH-78AM toto otherother HallHall thrusters of similar power power levels. levels. The The data data are are provided provided not not to tocompare compare thruster thruster performance, performance, but but to todemonstrate demonstrate that that performance performance measurements measurements of the of theUAH UAH-78AM-78AM are are comparable comparable to to experimental experimental measurements measurements of of other 300–500 300–500 WW Hall Hall thrusters.thrusters. Anomalous Anomalous performance performance results results would would indicate indicate that that the the unconventional unconventional design design features of the UAH-78AM were changing the performance in a manner that requires further testing and research to understand.

Table 3. Comparison of Thrusters in the UAH-78AM Power Class at Similar Discharge Voltages.

Thruster VD (V) Power (W) Thrust (mN) Anode ISP (s) Anode Efficiency (%) MaSMi [17] 200–250 160–747 8.8–33 775–1321 21–29 SPT-50 [18] 199–282 210–389 12.9–18.9 1160–1524 35–41 BHT-200 [19] 200–275 200 11.5–12.5 - 35–42 ACE [20] 400 18 1300 Aurora [21] 300 100–300 5.7–19 1080–1460 30–44 UAH-78AM 180–260 280–520 17–30 870–1450 27–42

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features of the UAH-78AM were changing the performance in a manner that requires further testing and research to understand.

Table 3. Comparison of Thrusters in the UAH-78AM Power Class at Similar Discharge Voltages.

Anode Anode I Thruster VD (V) Power (W) Thrust (mN) SP Efﬁciency (s) (%) MaSMi [17] 200–250 160–747 8.8–33 775–1321 21–29 SPT-50 [18] 199–282 210–389 12.9–18.9 1160–1524 35–41 BHT-200 [19] 200–275 200 11.5–12.5 - 35–42 ACE [20] 400 18 1300 Aurora [21] 300 100–300 5.7–19 1080–1460 30–44 UAH-78AM 180–260 280–520 17–30 870–1450 27–42

4. Conclusions A main goal of this work was to assess whether low-cost 3D printing with polymers or glazed ceramics can be used in the fabrication of Hall effect thrusters. In addition to reduced fabrication costs, 3D printing has to potential of open up the design space, as was done for liquid rocket engines. The ability to print complex geometries or internal passages could allow novel designs and configurations of Hall thrusters to be built. It is not known at this time what those designs may be, but the possibility is now there. The low cost and fast turnaround time of the UAH-78AM also shows a path toward a rapid design and test cycle for Hall thrusters. Changes to the design of a thruster to optimize the performance can be easily tested in the lab over multiple iterations at minimal cost of materials. While the data collected cover only short duration testing, our results demonstrate that the UAH-78AM is capable of operating within normal jet-mode Hall discharge with performance comparable to other thrusters of a similar power and size. Therefore, by the most basic definition, the UAH-78AM is a fully functioning Hall Thruster, and 3D printing and other low-cost AM technologies can be used to build Hall thrusters. However, for a Hall thruster to be useful, it must be capable of sustaining a continuous discharge for satellite propulsion. In the case of the UAH-78AM, current test durations are too short to collect plume data or steady-state thrust and temperature data using conventional methods. The only measurements that can be collected with the UAH-78AM are transient, since the steady state thermal operating condition for the thruster is beyond the temperature limits of the materials used for the channel. While transient data might be insightful for baselining the performance of a thruster, steady state data is ultimately needed for the development of flight hardware. Another important consideration is that thruster design deficiencies of the UAH-78AM could be contributing to the short test durations. Channel wall heating and erosion in Hall effect thrusters are dependent on the magnetic field topology in the channel. Magnetically shielded field topologies have been found to reduce channel wall heating by reducing the contact between the plasma and the wall [22,23]. Our magnetic circuit design was based on an unshielded field topology for the purpose of simplicity; however, the unshielded design only increases thermal losses to the channel walls as compared to shielded designs. Performance comparisons of the UAH-78AM with other thrusters of similar power levels and sizes were given in Table3, two of which have been flown. The BHT-200 was flown onboard the TACSAT-2 (2006), FalconSat-5 (2010), and FaclonSat-6 (2018) missions. It was slated to be flown on the Iodine Satellite (iSat) mission [24]. The SPT-50 was flown on the Canopus-V mission in 2012. This showed that the AM thruster, though it has some non- standard materials, was able to reach similar performance metrics as flight thrusters. The operational lifetime is a limitation, as mentioned, but this can be solved as AM technology advances. Small Hall thrusters have also been flown on small satellite missions including the ExoMG thruster (2020), and most notably SpaceX’s krypton small Hall thruster used Aerospace 2021, 8, 227 13 of 14

in their Starlink constellation of satellites. However, no data on the Starlink thruster is currently available. Future revisions of the thruster may explore applications for 3D printing in the design of the magnetic circuit. However, at this time, 3D printing of magnetic components is still an emerging technology with limited availability. Furthermore, many 3D printed magnet processes currently are focused on polymer-bonded magnets [25,26]. Bonded magnets generally have lower maximum energy product than metallic magnets [27], thus requiring larger magnets to produce the required magnetic field intensities. Other future work may include the use of high temperature 3D printable materials. Of special interest are the ceramic or metal impregnated PLA filaments that have become commercially available in the last few years. These filaments have 40–80% by mass of a filler material such as clay, zirconium silicate, copper, and iron. If the printed part is put into a debinding and sintering process, it may be possible to produce a fully ceramic or metal Hall thrust component while still using a desktop 3D printer.

Author Contributions: Conceptualization, E.P.H. and K.G.X.; methodology, E.P.H., W.H. and K.G.X.; software, E.P.H.; validation, E.P.H., W.H. and K.G.X.; formal analysis, E.P.H.; investigation, E.P.H.; re- sources, W.H. and K.G.X.; data curation, E.P.H.; writing—original draft preparation, E.P.H.; writing— review and editing, W.H. and K.G.X.; visualization, E.P.H.; supervision, W.H. and K.G.X.; project administration, K.G.X.; funding acquisition, W.H. and K.G.X. All authors have read and agreed to the published version of the manuscript. Funding: E.P.H. was funded by the Alabama Space Grant Fellowship. The work at NASA GRC was supported by the LERCIP program. Data Availability Statement: The supporting data will be provided upon request to the corresponding author. Acknowledgments: The authors would like to thank the staff at the UAH Propulsion Research Center and NASA GRC for assisting in the work and making it possible. Conﬂicts of Interest: The authors declare no conﬂict of interest.

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