Hall Effect Thrusters

When the US was beginning to investigate the gridded ion drive, the Soviet Union was investigating the Hall Effect thruster (HET). This is a very similar concept in many ways to the ion drive, in that it uses the electrostatic effect to accelerate propellant, but the mechanism is very different. Rather than using a system of grids that are electrically charged to produce the electrostatic potential needed to accelerate the ionized propellant, in a HET the plasma itself creates the electrostatic charge through the Hall effect, discovered in the 1870s by Edwin Hall. In these thrusters, the backplate functions as both a gas injector and an anode. A radial magnetic field is produced by a set of radial solenoids and and a central solenoid, which traps the electrons that have been stripped off the propellant as it’s become ionized (mostly through electron friction), forming a toroidal electric field moving through the plasma. After the ions are ejected out of the thruster, a hollow cathode that’s very similar to the one used in the ion drives that we’ve been looking at neutralizes the plasma beam, for the same reasons as this is done on an ion drive system (this is also the source of approximately 10% of the mass flow of the propellant). This also provides the electrostatic potential used to accelerate the propellant to produce thrust. Cathodes are commonly mounted external to the thruster on a small arm, however some designs – especially modern NASA designs – use a central cathode instead.

The list of propellants used tends to be similar to what other ion drives use: krypton, argon, iodine, bismuth, magnesium and zinc have all been used (along with some others, such as NaK), but Kr and Ar are the most popular propellants. While this system has a lower average specific impulse (1500-3000 s isp) than the gridded ion drives, it has more thrust (a typical drive system used today uses 1.35 kW of power to generate 83 mN of thrust), meaning that it’s very good for either orbital inclination maintenance or reaction control systems on commercial satellites.

There are a number of types of Hall effect thruster, with the most common being the Thruster with Anode Layer (TAL), the Stationary Plasma Thruster (SPT), and the cylindrical Hall thruster (CHT). The cylindrical thruster is optimized for low power applications, such as for cubesats, and I haven’t seen a high power design, so we aren’t going to really go into those. There are two obvious differences between these designs:

  1. What the walls of the acceleration chamber are made out of: the TAL uses metallic (usually boron nitride) material for the walls, while the SPT uses an insulator, which has the effect of the TAL having higher electron velocities in the plasma than the SPT.
  2. The length of the acceleration zone, and its impact on ionization behavior: The TAL has a far shorter acceleration zone than the SPT (sort of, see Chouieri’s analytical comparison of the two systems for dimensional vs non-dimensional characteristics http://alfven.princeton.edu/publications/choueiri-jpc-2001-3504). Since the walls of the acceleration zone are a major lifetime limitator for any Hall effect thruster, there’s an engineering trade-off available here for the designer (or customer) of an HET to consider.

There’s a fourth type of thruster as well, the External Discharge Plasma Thruster, which doesn’t have an acceleration zone that’s physically constrained, that we’ll also look at, but as far as I’ve been able to find there are very few designs, most of those operating at low voltage, so they, too, aren’t as attractive for nuclear electric propulsion.

Commercially available HETs generally have a total efficiency in the range of 50-60%, however all thrusters that I’ve seen increase in efficiency as the power increases up to the design power limits, so higher-powered systems, such as ones that would be used on a nuclear electric spacecraft, would likely have higher efficiencies. Some designs, such as the dual stage TAL thruster that we’ll look at, approach 80% efficiency or better.

SPT Hall Effect Thrusters

SPT Hall cutaway rough

SPT type thruster

Stationary Plasma Thrusters use an insulating material for the propellant channel immediately downstream of the anode. This means that the electrostatic potential in the drive can be further separated than in other thruster designs, leading to greater separation of ionized vs. non-ionized propellant, and therefore potentially more complete ionization – and therefore thrust efficiency. While they have been proposed since the beginning of research into Hall effect thrusters in the Soviet Union, a lack of an effective and cost-efficient insulator that was able to survive for long enough to allow for a useful thruster lifetime was a major limitator in early designs, leading to an early focus on the TAL.

SPT Liu et al

SPT Diagram, Liu et al 2010

The SPT has the greatest depth between the gas diffuser (or propellant injector) and the nozzle of the thruster. This is nice, because it gives volume and distance to work with in terms of propellant ionization. The ionized propellant is accelerated toward the nozzle, and the not-yet-ionized portion can still be ionized, even if the plasma component is scooting it toward the nozzle by simply bouncing off the unionized portion like a billiard ball. Because of this, SPT thrusters can have much higher propellant ionization percentages than the other types of Hall effect thruster, which directly translates into greater thrust efficiency. This extended ionization chamber is made out of an electromagnetic insulator, usually boron nitride, although Borosil, a solution of BN and SiO2, is also used. Other types of materials, such as nanocrystalline diamond, graphene, and a new material called ultra-BN, or plasma assisted chemical vapor deposition built BN, have also been proposed and tested.

Worn SPT-30, NASA 1998

SPT-30 Hall thruster post testing, image courtesy NASA

The downside to this type of thruster is that the insulator is eroded during operation. Because the erosion of the propellant channel is the main lifetime limitator of this type of thruster, the longer length of the propellant channel in this type of thruster is a detriment to thruster lifetime. Improved materials for the insulator cavity are a major research focus, but replacing boron nitride is going to be a challenge because there are a number of ways that it’s advantageous for a Hall effect thruster (and also in the other application we’ve looked at for reactor shielding): in addition to it being a good electromagnetic insulator, it’s incredibly strong and very thermally conductive. The only major downside is its’ expense, especially forming it into single, large, complex shapes; so, often, SPT thrusters have two boron carbide inserts: one at the base, near the anode, and another at the “waist,” or start of the nozzle, of the SPT thruster. Inconsistencies in the composition and conductivity of the insulator can lead to plasma instabilities in the propellant due to the local magnetic field gradients, which can cause losses in ionization efficiency. Additionally, as the magnetic field strength increases, plasma instabilities develop in proportion to the total field strength along the propellant channel.

Another problem that surfaces with these sorts of thrusters is that under high power, electrical arcing can occur, especially in the cathode or at a weak point in the insulator channel. This is especially true for a design that uses a segmented insulator lining for the propellant channel.

HERMeS: NASA’s High Power Single Channel SPT Thruster

peterson_hallthruster

Dr. Peterson with TAL-2 HERMeS Test Bed, image courtesy NASA

The majority of NASA’s research into Hall thrusters is currently focused on the Advanced Electric Propulsion System, or AEPS. This is a solar electric propulsion system which encompasses the power generation and conditioning equipment, as well as a 14 kW SPT thruster known as HERMeS, or the Hall Effect Rocket with Magnetic Shielding. Originally meant to be the primary propulsion unit for the Asteroid Redirect Mission, the AEPS is currently planned for the Power and Propulsion Element (PPE) for the Gateway platform (formerly Lunar Gateway and LOP-G) around the Moon. Since the power and conditioning equipment would be different for a nuclear electric mission, though, our focus will be on the HERMeS thruster itself.

HERMeSThis thruster is designed to operate as part of a 40 kW system, meaning that three thrusters will be clustered together (complications in clustering Hall thrusters will be covered later as part of the Japanese RIAJIN TAL system). Each thruster has a central hollow cathode, and is optimized for xenon propellant.

Many materials technologies are being experimented with in the HERMeS thruster. For instance, there are two different hollow cathodes being experimented with: LaB6 (which was experimented with extensively for the NEXT gridded ion thruster) and barium oxide (BaO). Since the LaB6 was already extensively tested, the program has focused on the BaO cathode. Testing is still underway for the 2000 hour wear test; however, the testing conducted to date has confirmed the behavior of the BaO cathode. Another example is the propellant discharge channel: normally boron nitride is used for the discharge channel, however the latest iteration of the HERMeS thruster is using a boron nitride-silicon (BN-Si) composite discharge channel. This could potentially improve the erosion effects in the discharge channel, and increase the life of the thruster. As of today, the differences in plasma plume characterization are minimal to the point of being insignificant, and erosion tests are similarly inconclusive; however, theoretically, BN-Si composite could improve the lifetime of the thruster. It is also worth noting that, as with any new material, it takes time to fully develop the manufacture of the material to optimize it for a particular use.

As of the latest launch estimates, the PPE is scheduled to launch in 2022, and all development work of the AEPS is on schedule to meet the needs of the Gateway.

Nested Channel SPT Thrusters: Increasing Power and Thrust Density

Nested HET

Nested SPT Thruster, Liang 2013

One concept that has grown more popular recently (although it’s far from new) is to increase the number of propellant channels in a single thruster in what’s called a nested channel Hall thruster. Several designs have used two nested channels for the thruster. While there are a number of programs investigating nested Hall effect thrusters, including in Japan and China, we’ll use the X2 as an example, studied at the University of Michigan. While this design has been supplanted by the X3 (more on that below), many of the questions about the operation of these types of thrusters were addressed by experimenting with the X2 thruster. Generally speaking, the amount of propellant flow in the different channels is proportional to the surface area of the emitter anode, and the power and flow rate of the cathode (which is centrally mounted) is adjusted to match whether one or multiple channels are firing. Since these designs often use a single central cathode, despite having multiple anodes, a lot of development work has gone into improving the hollow cathodes for increased life and power capability. None of the designs that I saw used external cathodes, like those sometimes seen with single-channel HETs, but I’m not sure if that is just because of the design philosophies of the institutions (primarily JPL and University of Michigan) that I found while investigating this type of design, and for which I was able to access the papers.

Nested SPT Tradeoffs Liang

Image Liang 2013

There are a number of advantages to the nested-channel design. Not only is it possible to get more propellant flow from less mass and volume, but the thruster can be throttled as well. For higher thrust operation (such as rapid orbital changes), both channels are fired at once, and the mass flow through the cathode is increased to match. By turning off the central channel and leaving the outer channel firing, a medium “gear” is possible, with mass flow similar to a typical SPT thruster. The smallest channel can be used for the highest-isp operation for interplanetary cruise operations, where the lower mass flow allows for greater exhaust velocities.

X2 Different Channel Operation

Ignition sequence for dual channel operation in X2, Liang 2013

A number of important considerations were studied during the X2 program, including the investigation of anode efficiency during the different modes of operation (slight decrease in efficiency during two-channel operation, highest efficiency during inner channel only operation), interactions between the plasma plumes (harmonic oscillations were detected at 125 and 150 V, more frequent in the outer channel, but not detected at 200 V operation, indicating some cross-channel interactions that would need to be characterized in any design), and power to thrust efficiency (slightly higher during two-channel operation compared to the sum of each channel operating independently, for reasons that weren’t fully able to be characterized). The success of this program led to its’ direct successor, which is currently under development by the University of Michigan, Aerojet Rocketdyne, and NASA: the X3 SPT thruster.

X3 on Test Stand

X3 Three Channel Nested SPT on test stand, Hall 2018

The X3 is a 100 kWe design that uses three nested discharge chambers. The cathode for this thruster is a centrally mounted hollow cathode, which accounts for 7% of the total gas flow of the thruster under all modes of operation. Testing during 2016 and 2017 ranging from 5 to 102 kW, 300 to 500 V, and 16 to 247 A, demonstrated a specific impulse range of 1800 to 2650 s, with a maximum thrust of 5.4 N. As part of the NextSTEP program, the X3 thruster is part of the XR-100 electric propulsion system that is currently being developed as a flexible, high-powered propulsion system for a number of missions, both crewed and uncrewed.

LaB6 Cathode

LaB6 Cathode, Hall 2018

While this thruster is showing a great deal of promise, there remain a number of challenges to overcome. One of the biggest is cathode efficiency, which was shown to be only 23% during operation of just the outermost channel. This is a heavy-duty cathode, rated to 120 A. Due to the concerns of erosion, especially under high-power, high-flow conditions, there are three different gas injection points: through the central bore of the cathode (limited to 20 sccm), external flow injectors around the cathode keeper, and supplementary internal injectors.

The cross-channel thrust increases seen in the X2 thruster weren’t observed, meaning that this effect could have been something particular to that design. In addition, due to the interactions between the different magnetic lenses used in each of the discharge channels, the strength and configuration of each magnetic field has to be adjusted depending on the other channels that are operating, a challenge that increases with magnetic field strength.

Finally, the BN insulator was shown to expand in earlier tests to the point that a gap was formed, allowing arcing to occur from the discharge plasma to the body of the thruster. Not only does this mean that the plasma is losing energy – and therefore decreasing thrust – but it also heats the body of the thruster as well.

These challenges are all being addressed, and in the next year the 100-hour, full power test of the system will be conducted at NASA’s Glenn Research Center.

X3 Firing Modes

X3 firing in all anode configurations: 1. Inner only, 2. Middle only, 3. Outer only, 4. Inner and middle, 5. Middle and outer, 6. Inner and outer, 7 Inner, middle and outer; Florenz 2013

TAL Hall Effect Thrusters

Early USSR TAL, Kim et al

Early TAL concept, Kim et al 2007

The TAL concept has been around since the beginning of the development of the Hall thruster. In the USSR, the development of the TAL was tasked to the Central Research Institute for Machine Building (TsNIIMash). Early challenges with the design led to it not being explored as thoroughly in the US, however. In Europe and Asia, however, this sort of thruster has been a major focus of research for a number of decades. Recently, the US has also increased their study of this design as well. Since these designs have (as a general rule) higher power requirements for operation, they have not been used nearly as much as the SPT-type Hall thruster, but for high powered systems they offer a lot of promise.

As we mentioned before, the TAL uses a conductor for the walls of the plasma chamber, meaning that the radial electric charge moving across the plasma is continuous across the acceleration chamber of the thruster. Because of the high magnetic fields in this type of thruster (0.1-0.2 T), the electron cyclotron radius is very small, allowing for more efficient ionization of the propellant, and therefore limiting the size necessary for the acceleration zone. However, because a fraction of the ion stream is directed toward these conduction walls, leading to degradation, the lifetime of these types of thrusters is often shorter than their SPT counterparts. This is one area of investigation for designers of TAL thrusters, especially higher-powered variants.

As a general rule, TAL thrusters have lower thrust, but higher isp, than SPT thrusters. Since the majority of Hall thrusters are used for station-keeping, where thrust levels are a significant design consideration, this has also mitigated in favor of the SPT thruster to be in wider deployment.

High-Powered TAL Development in Japan: Clustered TAL with a Common Cathode

RAIJIN94, Hamada 2017

RAIJIN94, Hamada et al 2017

One country that has been doing a lot of development work on the TAL thruster is Japan. Most of their designs seem to be in the 5 kW range, and are being designed to operate clustered around a single cathode for charge neutralization. The RAIJIN program (Robust Anode-layer Intelligent Thruster for Japanese IN-space propulsion system) has been focusing on addressing many of the issues with high-powered TAL operation, mainly for raising satellites from low earth orbit to geosynchronous orbit (an area that has a large impact on the amount of propellant needed for many satellite launches today, and directly applicable to de-orbiting satellites as well). The RAIJIN94 thruster is a 5 kW TAL thruster under development by Kyushu University, the University of Tokyo, and the University of Mizayaki. Overall targets for the program are for a thruster that operates at 6 kW, providing 360 mN of thrust, 1700 s isp, with an anode mass flow rate of 20 mg/s and a cathode flow rate of 1 mg/s. The ultimate goal of the program is a 25 kW TAL system, containing 5 of these thrusters with a common cathode. Based on mass penalty analysis, this is a more mass-efficient method for a TAL than having a single large TAL with increased thermal management requirements. Management of anode and conductor erosion is a major focus of this program, but not one that has been written about extensively. The limiting of the thruster power to about 5 kW, though, seems to indicate that scaling a traditional TAL beyond this size, at least with current materials, is impractical.

Clustered HETs, Miyasaka 2013

Side by side HETs for testing with common cathode, Miyazaka et al 2013

There are challenges with this design paradigm, however, which also will impact other clustered Hall designs. Cathode performance, as we saw in the SPT section is a concern, especially if operating at very high power and mass flow rates, which a large cluster would need. Perhaps a larger consideration was plasma oscillations that occurred in the 20 kHz range when two thrusters were fired side by side, as was done (and continues to be done) at Gifu University. It was found that by varying the mass flow rate, operating at a slightly lower power, and maintaining a wider spacing of the thruster heads, the plasma flow instabilities could be accounted for. Experiments continue there to study this phenomenon, and the researchers, headed by Dr. Miyasaka, are confident that this issue can be managed.

Dual Stage TAL and VHITAL

Two Stage TALOne of the most interesting concepts investigated at TsNIIMash was the dual-stage TAL, which used two anodes. The first anode is very similar to the one used in a typical TAL or SPT, which also serves as an injector for the majority of the propellant and provides the charge to ionize the propellant. As the plasma exits this first anode, it encounters a second anode at the opening of the propellant channel, which accelerates the propellant. An external cathode is used to neutralize the beam. This design demonstrated specific impulses of up to 8000s, among the highest (if not the highest) of any Hall thruster to date. The final iteration during this phase of research was the water-cooled TAL-160, which operated at a power consumption from 10-140 kW.

VHITAL Propellant Reservoir

VHITAL bismuth propellant feed system, Sengupta et al 2007

Another point of interest with this design is the use of bismuth as the propellant for the thruster. As we’ll see below, propellant choice for an electrostatic thruster is very broad, and the choice of propellant you use is subject to a number of characteristics. Bismuth is reasonably inexpensive, relatively common, and storable as a solid. This last point is also a headache for an electrostatic thruster, since ionized powders are notorious for sticking to surfaces and gumming up the works, as it were. In this case, since bismuth has a reasonably low melting temperature, a pre-heater was used to resistively heat the bismuth, and then an electromagnetic pump was used to propel it toward the anode. Just before injection into the thruster, a vaporization plug of carbon was used to ensure proper mass flow into the thruster. As long as the operating temperature of the thruster was high enough, and the mass flow was carefully regulated, this novel fueling concept was not a problem.

VHITAL 160 at TsNIIMash

VHITAL mounted to test bracket at TsNIIMash, Sangupta et al 2007

This design was later picked up in 2004 by NASA, who worked with TsNIIMash researchers to develop the VHITAL, or Very High Isp Thruster with Anode Layer, over two and a half years. While this thruster uses significantly less power (36 kW as opposed to up to 140 kW), many of the design details are the same, but with a few major differences: the NASA design is radiatively cooled rather than water cooled, it added a resistive heater to the base of the first anode as well, and tweaks were made to the propellant feed system. The original TAL-160 was used for characterization tests, and the new VHITAL-160 thruster and propellant feed system were built to characterize the system using modern design and materials. Testing was carried out at TsNIIMash in 2006, and demonstrated stable operation without using a neutralizing cathode, and expected metrics were met.

While I have been able to find a summary presentation from the time of the end of the program in the US, I have been unable to find verified final results of this program. However, 8000 s isp was demonstrated experimentally at 36 kW, with a thrust of ~700 mN and a thrust efficiency of close to 80%.

If anyone has additional information about this program, please comment below or contact me via email!

Hybrid and Non-Traditional Hall Effect Thrusters: Because Tweaking Happens

PLAS40 Sketch

PLa

As we saw with the VHITAL, the traditional SPT and TAL thrusters – while the most common – are far from the only way to use these technologies. One interesting concept, studied by EDB Fakel in Russia, is a hybrid SPT-TAL thruster. SPT thrusters, due to their extended ionization chamber lined with an insulator, generally provide fuller ionization of propellant. TAL thrusters, on the other hand, are better able to efficiently accelerate the propellant once it’s ionized. So the designers at EDB Fakel, led by M. Potapenko, developed, built, and tested the PlaS-40 Hybrid PT, rated at up to 0.4 kW, and proposed and breadbox tested a larger (up to 4.5 kW) PlaS-120 thruster as well, during the 1990s (initial conception was in the early 90s, but the test

PLAS40

PLaS40 Thruster

model was built in 1999). While fairly similar in outward appearance to an SPT, the acceleration chamber was shorter. The PlaS-40 achieved 1000-1750 s isp and a thrust of 23.5 mN, while the PlaS-120 showed the capability of reaching 4000 s isp and up to 400 mN of thrust (these tests were not continued, due to a lack of funding). This design concept could offer advances in specific impulse and thrust efficiency beyond traditional thruster designs, but currently there isn’t enough research to show a clear advantage.

 

Gridded Hall concept

Early EKB Fakel concept for SPT thruster with “magnetic screens.” Kim et al 2007

Another interesting hybrid design was a gridded Hall thruster, researched by V. Kim at Fakel in 1973-1975. Here, again, an SPT-type ionization chamber was used, and the screens were used to more finely control the magnetic lensing effect of the thruster. This was an early design, and one that was used due to the limitations of the knowledge and technology to do away with the grids. However, it’s possible that a hybrid Hall-gridded ion thruster may offer higher specific impulse while taking advantage of the more efficient ionization of an SPT thruster. As we saw with both the DS4G and VHITAL, increasing separation of the ionization, ion extraction, and acceleration portions of the thruster allows for a greater thrust efficiency, and this may be another mechanism to do that.

One design, out of the University of Michigan, modifies the anode itself, by segmenting it into many different parts. This was done to manage plasma instabilities within the propellant plume, which cause parasitic power losses. While it’s unclear exactly how much efficiency can be gained by this, it solves a problem that had been observed since the 1960s close to the anode of the thruster. Small tweaks like this may end up changing the geometry of the thruster significantly over time as optimization occurs.

Other modifications have been made as well, including combining discharge chambers, using conductive materials for discharge chambers but retaining a dielectric ceramic in the acceleration zone of the thruster, and many other designs. Many of these were early ideas that were demonstrated but not carried through for one reason or another. For instance, the metal discharge chambers were considered an economic benefit, because the ceramic liners are the major cost-limiting factor in SPT thrusters. With improved manufacturing and availability, costs went down, and the justification went away.

There remains an incredible amount of flexibility in the Hall effect thruster design space. While two stage, nested, and clustered designs are the current most advanced high power designs, it’s difficult to guess if someone will come up with a new idea, or revisit an old one, to rewrite the field once again.

Sources

Fundamental Difference Between the Two Variants of Hall Thruster: SPT and TAL http://alfven.princeton.edu/publications/choueiri-jpc-2001-3504

History of the Hall Thrusters Development in the USSR, Kim et al 2007 http://erps.spacegrant.org/uploads/images/images/iepc_articledownload_1988-2007/2007index/IEPC-2007-142.pdf

Recent Progress and Perspectives on Space Electric Propulsion Systems Based on Smart Nanomaterials, Levchenko et al 2018 https://www.ncbi.nlm.nih.gov/pmc/articles/PMC5830404/pdf/41467_2017_Article_2269.pdf

SPT

Testing of the PPU Mk 3 with the XR-5 Hall Effect Thruster, Xu et al, Aerojet Rocketdyne 2017 https://iepc2017.org/sites/default/files/speaker-papers/iepc-2017-199.pdf

AEPS and HERMeS

Overview of the Development and Mission Application of the Advanced Electric Propulsion System (AEPS), Herman et al 2017 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20180001297.pdf

13kW Advanced Electric Propulsion Flight System Development and Qualification, Jackson et al 2017 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20180000353.pdf

Wear Testing of the HERMeS Thruster, Williams et al 2017 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20170000963.pdf

Performance, Stability, and Plume Characterization of the HERMeS Thruster with Boron Nitride Silica Composite Discharge Channel, Kamhawi et al 2017 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20180000687.pdf

Hollow Cathode Assembly Development for the HERMeS Hall Thruster, Sarver-Verhey et al 2017 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20170001280.pdf

Nested HET

Constant-Power Performance and Plume Effects of a Nested-Channel Hall-Effect Thruster, Liang et al University of Michigan 2011 http://pepl.engin.umich.edu/pdf/IEPC-2011-049.pdf

The Combination of Two Concentric Discharge Channels into a Nested Hall Effect Thruster, Liang PhD Thesis, University of Michigan 2013 http://pepl.engin.umich.edu/pdf/2013_Liang_Thesis.pdf

Plasma Oscillation Effects on Nested Hall Thruter Operation and Stability, McDonald et al University of Michigan 2013 http://pepl.engin.umich.edu/pdf/IEEE-2013-2502.pdf

Investigation of Channel Interactions in a Nested Hall Thruster Part 1, Georgin et al University of Michigan 2016 http://pepl.engin.umich.edu/pdf/JPC_2016_Georgin.pdf

Investigation of Channel Interactions in a Nested Hall Thruster Part 2, Cusson et al University of Michigan 2016 http://pepl.engin.umich.edu/pdf/JPC_2016_Cusson.pdf

X3 NHT

The X3 100 kWclass Nested Hall Thruster: Motivation, Implementation, and Initial Performance; Florenz, University of Michigan doctoral thesis http://pepl.engin.umich.edu/pdf/2014_Florenz_Thesis.pdf

First Firing of a 100-kW Nested Channel Hall Thruster, Florenz et al University of Michigan, 2013 http://www.dtic.mil/dtic/tr/fulltext/u2/a595910.pdf

High-Power Performance of a 100-kW Class Nested Hall Thruster, Hall et al University of Michigan 2017 http://pepl.engin.umich.edu/pdf/IEPC-2017-228.pdf

Update on the Nested Hall Thruster Subsystem for the NextSTEP XR-100 System, Jorns et al University of Michigan 2018 http://pepl.engin.umich.edu/pdf/AIAA-2018-4418.pdf

Multichannel Hall Effect Thruster patent, McVey et al, Aerojet Rocketdyne https://patents.google.com/patent/US7030576B2/en

MW-Class Electric Propulsion Systems Designs for Mars Cargo Transport; Gilland et al 2011 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20120001636.pdf

TAL

An Overview of the TsNIIMASH/TsE Efforts under the VHITAL Program, Tverdokhlebov et al, TsNIIMASH, 2005 http://erps.spacegrant.org/uploads/images/images/iepc_articledownload_1988-2007/2005index/141.pdf

Thrust Performance in a 5 kW Class Anode Layer Type Hall Thruster, Yamamoto et al Kyushu University 2015 https://www.jstage.jst.go.jp/article/tastj/14/ists30/14_Pb_183/_pdf

Investigation of a Side by Side Hall Thruster System, Miyasaka et al 2013 http://erps.spacegrant.org/uploads/images/images/iepc_articledownload_1988-2007/2013index/1vrun4h9.pdf

Modeling of a High Power Thruster with Anode Layer, Keidar et al University of Michigan 2004 https://deepblue.lib.umich.edu/bitstream/handle/2027.42/69797/PHPAEN-11-4-1715-1.pdf;sequence=2

Design and Performance Evaluation of Thruster with Anode Layer UT-58 For High-Power Application, Schonherr et al, University of Tokyo 2013 http://erps.spacegrant.org/uploads/images/images/iepc_articledownload_1988-2007/2013index/k97rps25.pdf

Very High Isp Thruster with Anode Layer (VHITAL): An Overview, Marrese-Reading et al, JPL 2004 https://trs.jpl.nasa.gov/handle/2014/40499

The Development of a Bismuth Feed for the Very High Isp Thruster with Anode Layer VHITAL program, Marrese-Reading et al, JPL 2004 https://trs.jpl.nasa.gov/handle/2014/37752

Hybrid

Characteristic Relationship Between Dimensions and Parameters of a Hybrid Plasma Thruster, Potapenko et al 2011 http://erps.spacegrant.org/uploads/images/images/iepc_articledownload_1988-2007/2011index/IEPC-2011-042.pdf

Measurement of Cross-Field Electron Current in a Hall Thruster due to Rotating Spoke Instabilities, McDonald et al 2011 https://core.ac.uk/download/pdf/3148750.pdf

Metallic Wall Hall Thruster patent, Goebel et al 2016 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20160012003.pdf

History of the Hall Thrusters Development in the USSR, Kim et al 2007 http://erps.spacegrant.org/uploads/images/images/iepc_articledownload_1988-2007/2007index/IEPC-2007-142.pdf