Hello, and welcome back to Beyond NERVA! Today, we finish our look at electric propulsion systems by looking at electrostatic propulsion. This is easily the most common form of in-space electric propulsion system, and as we saw in our History of Electric Propulsion post, it’s also the first that was developed.
I apologize about how long it’s taken to get this blog post published. As I’ve mentioned before, electric propulsion is one of my weak subjects, so I’ve been very careful to try to ensure that the information that I’m giving is correct. Another complication came from the fact that I had no idea how complex and varied each type of drive system is. I have glossed over many details in this blog post on many of the systems, but I’ve also included an extensive list of documentation on all of the drive systems I discuss in the post at the end, so if you’re curious about the details of these systems, please check out the published papers on them!
By far the most common type of electric propulsion today, and the type most likely to be called an “ion thruster,” is electrostatic propulsion. The electrostatic effect was one of the first electrical effects ever formally described, and the first ever observed (lightning is an electrostatic phenomenon, after all). Electrostatics as a general field of study refers to the study of electric charges at rest (hence, “electro-static”). The electrostatic effect is the tendency of objects with a differential charge (one positive, one negative) to attract each other, and with the same charge to repel each other. This occurs when electrons are stripped or added to one material. Some of the earliest scientific experiments involving this effect involved bars of amber and wool – the amber would become negatively ionized, and the wool would be positively ionized, due to the interactions of the very fine hairs of the wool and the crystalline and elemental composition of the amber (for the nitpicky types, this is known as the triboelectric effect, but is still a manifestation of the electrostatic effect). Other experimenters during the 18th and 19th centuries used cat fur instead of wool, a much more mentally amusing way to build an electrostatic charge. However, we aren’t going to be discussing using a rotating wheel of cats to produce an electric thruster (although, if someone feels like animating said concept, I’d love to see it).
There are a number of designs that use electrostatic effects to produce thrust. Some are very similar to some of the concepts that we discussed in the last post, like the RF ionized thruster (a major area of focus in Japan), the Electron Cyclotron Resonance thrusters (which use the same mechanisms as VASIMR’s acceleration mechanism), and the largely-abandoned Cesium Contact thruster (which has a fair amount of similarities with a pulsed plasma or arcjet thruster). Others, such as the Field Emission Electrostatic Thruster (FEEP) and Ionic Liquid Ion Source thruster (also sometimes called an electrospray) thruster, have far fewer similarities. None of these, though, are nearly as common as the electron bombardment noble gas thruster types: the gridded ion (either electron bombardment, cyclotron resonance, or RF ionization) thruster and the Hall effect thruster (which also has two types: the thruster with anode layer and stationary plasma thruster). The gridded ion thruster, commonly just called an ion thruster, is the propulsion system of choice for interplanetary missions, because it has the highest specific impulse of any currently available propulsion system. Hall effect thrusters have lower specific impulse, but higher thrust, making them a popular choice for reaction control systems on commercial and military satellites.
Most electrostatic drives use an ionization chamber or zone, to strip off electrons from an easily-ionized material. These now-positively charged ions are then accelerated toward a negatively charged structure (or accelerated by an electromagnetic field, in some cases), which is then switched off after accelerating the ions, which then are spat out the back of the thruster. Because of the low density of these ion streams, and the lack of an expanding gas, a physical nozzle isn’t used, because the characteristic bell-shaped de Laval nozzle of chemical or thermal engines is absolutely useless in this case. However, there are many ways that this ion stream can be ionized, and many ways that it can be accelerated, leading to a huge variety of design options within the area of electrostatic propulsion.
The first design for a practical electric propulsion system, patented by Robert Goddard in 1917, was an electrostatic device, and most designs, both in the US and the USSR, have used this concept. In the earliest days of electric propulsion design, each went a different way in the development of this drive concept: the US focused on the technically simpler, but materially more problematic, gridded ion thruster, while the Soviet Union worked to develop the more technically promising, but more difficult to engineer, Hall thruster. Variations of each have been produced over the years, and additional options have been explored as well. These systems have traveled to almost every body of in the Solar System, including Pluto and many of the asteroids in the Main Belt, and provide a lot of the station-keeping thrust for satellites in orbit around Earth. Let’s go ahead and look at what the different types are, what their advantages and disadvantages are, and how they’ve been used in the past.
Gridded Ion Drives
This is the best-known of the electric propulsion thrusters of any type, and is often shortened to “ion drive.” Here, the thruster has four main parts: the propellant supply, an ionization chamber, an array of conductive grids, and a neutralizing beam emitter. The propellant can be anything that is able to be easily ionized, with cesium and mercury being the first options, these have largely been replaced by xenon and argon, though.
The type of ionization chamber varies widely, and is the main difference in the different types of ion drive. Particle beams, radio frequency or microwave excitation, in addition to magnetic field agitation, are all methods used in different gridded ion drives over the years and across the different manufacturers. The first designs used gaseous agitation to cause electrons to be stripped, but many higher-powered systems use particle (mostly electron) beams, radio frequency or microwave agitation, or cyclotron resonance to strip the electrons off the atoms. The efficiency of the ionization chamber, and its capacity, define how much propellant mass flow is possible, which is one of the main limiting factors for the overall thrust possible for the thruster.
After being ionized, the gas and plasma are then separated, using a negatively charged grid to extract the positively charged ions, leaving the neutral gas in the ionization chamber to be ionized. In most modern designs, this is also the beginning of the acceleration process. Often, two or three grids are used, and the term “ion optics” is often used instead of “grids.” This is because these structures not only extract and change the acceleration of the plasma, but they also shape the beam of the plasma as well. The amount of charge, and the geometry of these grids, defines the exhaust velocity of the ions; and the desired specific impulse produced by the thruster is largely determined by the charge applied to these screens. Many US designs use a more highly charged inner screen to ensure better separation of the ions, and a charge potential difference between this grid and the second accelerates the ions. Because of this, the first grid is often called the extractor, and the second is called the accelerator grid. The charge potential possible on each grid is another major limitator of the possible power level – and therefore the maximum exhaust velocity – of these thrusters.
These screens also are one of the main limitators for the thruster’s lifetime, since the ions will impact the grid to a certain degree as they’re flowing past (although the difference in charge potential on the plasma in the ionization chamber between the apertures and the structure of the grid tends to minimize this). With many of the early gridded ion thrusters that used highly reactive materials, chemical interactions in the grids could change the conductivity of these surfaces, cause more rapid erosion, and produce other problems; the transition to noble gas propellants has made this less of an issue. Finally, the geometry of the grids have a huge impact on the direction and velocity of the ions themselves, so there’s a wide variety of options available through the manipulation of this portion of the thruster as well.
At the end of the drive cycle, after the ions are leaving the thruster, a spray of electrons is added to the propellant stream, to prevent the spacecraft from becoming negatively charged over time, and thereby attracting some of the propellant back toward the spacecraft due to the same electrostatic effect that was used to accelerate them in the first place. Problems with incomplete ion stream neutralization were common in early electrostatic thrusters; and with the cesium and mercury propellants used in these thrusters, chemical contamination of the spacecraft became an issue for some missions. Incomplete neutralization is something that is still a concern for some thruster designs, although experiments in the 1970s showed that a spacecraft can ground itself without the ion stream if the differential charge becomes too great. In three grid systems (or four, more on that concept later), the final grid takes the place of this electron beam, and ensures better neutralization of the plasma beam, as well as greater possible exhaust velocity.
Gridded ion thrusters offer very attractive specific impulse, in the range of 1500-4000 seconds, with exhaust velocities up to about 100 km/s for typical designs. The other side of the coin is their low thrust, generally from 20-100 microNewtons (lower than average even for electric propulsion, although their specific impulse is higher than average), which is a mission planning constraint, but isn’t a major show-stopper for many applications. An advanced concept, from the Australian National University and European Space Agency, the Dual Stage 4 Grid (DS4G) thruster, achieved far higher exhaust velocities by using a staged gridded ion thruster, up to 210 km/s.
Past and Current Gridded Ion Thrusters
These drive systems have been used on a number of different missions over the years, starting with the SERT missions mentioned in the history of electric propulsion section, and continuing through on an experimental basis until the Deep Space 1 technology demonstration mission – the first spacecraft to use ion propulsion as its main form of propulsion. That same thruster, the NSTAR, is still in use today on the Dawn mission, studying the minor planet Ceres. Hughes Aircraft developed a number of thrusters for station-keeping for their geosynchronous satellite bus (the XIPS thruster).
JAXA used this type of drive system for their Hayabusa mission to the asteroid belt, but this thruster used microwaves to ionize the propellant. This thruster operated successfully throughout the mission’s life, and propelled the first spacecraft to return a sample from an asteroid back to Earth.
ESA has used different variations of this thruster on multiple different satellites as well, all of which have been radio frequency ionization types. The ArianeSpace RIT-10 has been used on multiple missions, and the Qinetiq T5 thruster was used successfully on the GOCE mission mapping the Earth’s magnetic field.
NASA certainly hasn’t given up on further developing this technology. The NEXT thruster is three times as powerful in terms of thrust compared to the NSTAR thruster, although it operates on similar principles. The testing regime for this thruster has been completed, demonstrating 4150 s of isp and 236 mN of thrust over a testing life of over 48,000 hours, and it is currently awaiting a mission for it to fly on. This has also been a testbed for using new designs and materials on many of the drive system components, including a new hollow cathode made out of LaB6 (a lanthanum-boron alloy), and several new screen materials.
HiPEP: NASA’s Nuclear Ion Propulsion System
Another NASA project in gridded ion propulsion, although one that has since been canceled, is far more germane to the specific use of nuclear electric propulsion: the High Power Electric Propulsion drive (HiPEP) for the Jupiter Icy Moons Observer mission. JIMO was a NEP propelled mission to Jupiter which was canceled in 2005, meant to study Europa, Ganymede, and Callisto (this mission will get an in-depth look later in this blog series on NEP). HiPEP used two types of ionization chamber: Electron Cyclotron Resonance ionization, which combines leveraging the small number of free electrons present in any gas by moving them in a circle with the magnetic containment of the ionization chamber with microwaves that are tuned to be in resonance with these moving electrons to more efficiently ionize the xenon gas; and direct current ionization using a hollow cathode to strip off electrons, which has additional problems with cathode failure and so is the less preferred option. Cathode failure of this sort is another major failure point for ion drives, so being able to eliminate this is a significant advantage, but the microwave system ends up consuming more power, so in less-energy-intensive applications it’s often not used.
One very unusual thing about this system is its’ shape: rather than the typical circular discharge chamber and grids, this system uses a rectangular configuration. The designers note that not only does this make the system more compact to stack multiple units together (reducing the structural, propellant, and electrical feed system mass requirements for the full system), it also means that the current density across the grids can be lower for the same electrostatic potential, reducing current erosion in the grids. This means that the grid can support a 100 kg/kW throughput margin for both of the isp configurations that were studied (6000 and 8000 s isp). The longest distance between two supported sections of grid can be reduced as well, preventing issues like thermal deformation, launch vibration damage, and electrostatic attraction between the grids and either the propellant or the back of the ionization chamber itself. The fact that it makes the system more scalable from a structural engineering standpoint is one final benefit to this system.
As the power of the thruster increases, so do the beam neutralization requirements. In this case, up to 9 Amperes of continuous throughput are required, which is very high compared to most systems. This means that the neutralizing beam has to be both powerful and reliable. While the HiPEP team discuss using a common neutralization system for tightly packed thrusters, the baseline design is a fairly typical hollow cathode, similar to what was used on the NSTAR thruster, but with a rectangular cross section rather than a circular one to accommodate the different thruster geometry. Other concepts, like using microwave beam neutralization, were also discussed; however, due to the success and long life of this type of system on NSTAR, the designers felt that this would be the most reliable way to deal with the high throughput requirements that this system requires.
HiPEP consistently met its program guidelines, for both engine thrust efficiency and erosion studies. Testing was conducted at both 2.45 and 5.85 GHz for the microwave ionization system, and was successfully concluded. The 2.45 GHz test, with 16 kW of power, achieved a specific impulse of 4500-5500 seconds, allowing for the higher-powered MW emitter to be used. The 5.85 GHz ionization chamber was tested at multiple current loads, from 9.7 to 39.3 kW, and achieved a maximum specific impulse of 9620 s, and showed a clear increase in thrust of up to close to 800 mN during this test.
Sadly, with the cancellation of JIMO (a program we will continue to come back to frequently as we continue looking at NEP), the need for a high powered gridded ion thruster (and the means of powering it) went away. Much like the fate of NERVA, and almost every nuclear spacecraft ever designed, the canceling of the mission it was meant to be used on spelled the death knell of the drive system. However, HiPEP remains on the books as an attractive, powerful gridded ion drive, for when an NEP spacecraft becomes a reality.
DS4G: Fusion Research-Inspired, High-isp Drives to Travel to the Edge of the Solar System
The Dual Stage 4 Grid (DS4G) ion drive is perhaps the most efficient electric drive system ever proposed, offering specific impulse well over 10,000 seconds. While there are some drive systems that offer higher isp, they’re either rare concepts (like the fission fragment rocket, a concept that we’ll cover in a future post), or have difficulties in the development process (such as Orion derivatives, which run afoul of nuclear weapons test bans and treaty limitations concerning the use of nuclear explosives in space).
So how does this design work? Traditional ion drives use either two grids (like the HiPEP drive) combining the extraction and acceleration stages in these grids and then using a hollow cathode or electron emitter to neutralize the beam, or use three grids, where the third grid is used in place of the hollow cathode. In either case, these are very closely spaced grids, which has its’ advantages, but also a couple of disadvantages: the extraction system and acceleration system being combined forces a compromise between efficiency of extraction and capability of acceleration, and the close spacing limits the acceleration possible of the propellants. The DS4G, as the name implies, does things slightly differently: there are two pairs of grids, each pair is close to its’ partner, but further apart from the other pair, allowing for a greater acceleration chamber length, and therefore higher exhaust velocity, and the distance between the extraction grid and the final acceleration grid allows for each to be better optimized for their individual purposes. As an added benefit, the plasma beam of the propellant is better collimated than that of a traditional ion drive, which means that the drive is able to be more efficient with the mass of the propellant, increasing the specific impulse even further.
This design didn’t come out of nowhere, though. In fact, most tokamak-type fusion reactors use a device very similar to an ion drive to accelerate beams of hydrogen to high velocities, but in order to get through the intense magnetic fields surrounding the reactor the atoms can’t be ionized. This means that a very effective neutralizer needs to be stuck on the back of what’s effectively an ion drive… and these designs all use four screens, rather than three. Dr. David Fearn knew of these devices, and decided to try and adapt it to space propulsion, with the help of ESA, leading to a 2005 test-bed prototype in collaboration with the Australian National University. An RF ionization system was designed for the plasma production unit, and a 35 kV electrical system was designed for the thruster prototype’s ion optics. This was not optimized for in-space use; rather, it was used as a low cost test-bed for optics geometry testing and general troubleshooting of the concept. Another benefit to this design is a higher-than-usual thrust density of 0.86 mN/cm^2, which was seen in the second phase of testing.
Two rounds of highly successful testing were done at ESA’s CORONA test chamber in 2005 and 2006, the results of which can be seen in the tables above. The first test series used a single aperture design, which while highly inefficient was good enough to demonstrate the concept; this was later upgraded to a 37 aperture design. The final test results in 2006 showed impressive specific impulse (14000-14500 s), thrust (2.7 mN), electrical, mass, and total efficiency (0.66, 0.96, and 0.63, respectively). The team is confident that total efficiencies of about 70% are possible with this design, once optimization is complete.
There remain significant engineering challenges, but nothing that’s incredibly different from any other high powered ion drive. Indeed, many of the complications concerning ion optics, and electrostatic field impingement in the plasma chamber, are largely eliminated by the 4-grid design. Unfortunately, there are no missions that currently have funding that require this type of thruster, so it remains on the books as “viable, but in need of some final development for application” when there’s a high-powered mission to the outer solar system.
Cesium Contact Thrusters: Liquid Metal Fueled Gridded Ion Drives
As we saw in our history of electric propulsion blog post, many of the first gridded ion engines were fueled with cesium (Cs). These systems worked well, and the advantages of having an easily storable, easily ionized, non-volatile propellant (in vapor terms, at least) were significant. However, cesium is also a reactive metal, and is toxic to boot, so by the end of the 1970s development on this type of thruster was stopped. As an additional problem, due to the inefficient and incomplete beam neutralization with the cathodes available at the time, contamination of the spacecraft by the Cs ions (as well as loss of thrust) were a significant challenge for the thrusters of the time.
Perhaps the most useful part of this type of thruster to consider is the propellant feed system, since it can be applied to many different low-melting-point metals. The propellant itself was stored as a liquid in a porous metal sponge made out of nickel, which was attached to two tungsten resistive heaters. By adjusting the size of the pores of the sponge (called Feltmetal in the documentation), the flow rate of the Cs is easily, reliably, and simply controlled. Wicks of graded-pore metal sponges were used to draw the Cs to a vaporizer, made of porous tungsten and heated with two resistive heaters. This then fed to the contact ionizer, and once ionized the propellant was accelerated using two screens.
As we’ll see in the propellant section, after looking at Hall Effect thruster, Cs (as well as other metals, such as barium) could have a role to play in the future of electric propulsion, and looking at the solutions of the past can help develop ideas in the future.
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:
- 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.
- 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
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.
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.
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
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.
This 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
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.
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.
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.
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.
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.
TAL Hall Effect Thrusters
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
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.
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
One 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.
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.
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
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
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.
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.
Propellants: Are the Current Propellant Choices Still Effective For High Powered Missions?
One of the interesting things to consider about these types of thrusters, both the gridded ion and Hall effect thrusters, is propellant choice. Xenon is, as of today, the primary propellant used by all operational electrostatic thrusters (although some early thrusters used cesium and mercury for propellants), however, Xe is rare and reasonably expensive. In smaller Hall thruster designs, such as for telecommunications satellites in the 5-10 kWe thruster range, the propellant load (as of 1999) for many spacecraft is less than 100 kg – a significant but not exorbitant amount of propellant, and launch costs (and design considerations) make this a cost effective decision. For larger spacecraft, such as a Hall-powered spacecraft to Mars, the propellant mas could easily be in the 20-30 ton range (assuming 2500 s isp, and a 100 mg/s flow rate of Xe), which is a very different matter in terms of Xe availability and cost. Alternatives, then, become far more attractive if possible.
Argon is also an attractive option, and is often proposed as a propellant as well, being less rare. However, it’s also considerably lower mass, leading to higher specific impulses but lower levels of thrust. Depending on the mission, this could be a problem if large changes in delta-vee are needed in a shorter period of time, The higher ionization energy requirements also mean that either the propellant won’t be as completely ionized, leading to loss of efficiency, or more energy is required to ionize the propellant
The next most popular choice for propellant is krypton (Kr), the next lightest noble gas. The chemical advantages of Kr are basically identical, but there are a couple things that make this trade-off far from straightforward: first, tests with Kr in Hall effect thrusters often demonstrate an efficiency loss of 15-25% (although this may be able to be mitigated slightly by optimizing the thruster design for the use of Kr rather than Xe), and second the higher ionization energy of Kr compared to Xe means that more power is required to ionize the same amount of propellant (or with an SPT, a deeper ionization channel, with the associated increased erosion concerns). Sadly, several studies have shown that the higher specific impulse gained from the lower atomic mass of Kr aren’t sufficient to make up for the other challenges, including losses from Joule heating (which we briefly discussed during our discussion of MPD thrusters in the last post), radiation, increased ionization energy requirements, and even geometric beam divergence.
This has led some designers to propose a mixture of Xe and Kr propellants, to gain the advantages of lower ionization energy for part of the propellant, as a compromise solution. The downside is that this doesn’t necessarily improve many of the problems of Kr as a propellant, including Joule heating, thermal diffusion into the thruster itself, and other design headaches for an electrostatic thruster. Additionally, some papers report that there is no resonant ionization phenomenon that facilitates the increase of partial krypton utilization efficiency, so the primary advantage remains solely cost and availability of Kr over Xe.
|Atomic Mass (Ar, std.)||Ionization Energy (1st, kJ/mol)||Density (g/cm^3)||Melting Point (K)||Boiling Point (K)||Estimated Cost ($/kg)|
|Mercury||200.592||1007.1||13.534 (at STP)||234.32||629.88||500|
|Cesium||132.905||375.7||1.843 (at MP)||301.7||944||>5000|
|Sodium||22.989||495.8||0.927 (at MP) 0.968 (solid)||370.94||1156.09||250|
|Potassium||39.098||418.8||0.828 (MP) 0.862 (solid)||336.7||1032||1000|
|NaK||Varies||Differential||0.866 (20 C)||260.55||1445||Varies|
|Iodine||126.904||1008.4||4.933 (at STP)||386.85||457.4||80|
Early thrusters used cesium and mercury for propellant, and for higher-powered systems this may end up being an option. As we’ve seen earlier in this post, neither Cs or Hg are unknown in electrostatic propulsion (another design that we’ll look at a little later is the cesium contact ion thruster), however they’ve fallen out of favor. The primary reason always given for this is environmental and occupational health concerns, for the development of the thrusters, the handling of the propellant during construction and launch, as well as the immediate environment of the spacecraft. The thrusters have to be built and extensively tested before they’re used on a mission, and all these experiments are a perfect way to strongly contaminate delicate (and expensive) equipment such as thrust stands, vacuum chambers, and sensing apparatus – not to mention the lab and surrounding environment in the case of an accident. Additionally, any accident that leads to the exposure of workers to Hg or Cs will be expensive and difficult to address, notwithstanding any long term health effects of chemical exposure to any personnel involved (handling procedures have been well established, but one worker not wearing the correct personal protective equipment could be constantly safe both in terms of personal and programmatic health) Perfect propellant stream neutralization is something that doesn’t actually occur in electrostatic drives (although as time goes on, this has consistently improved), leading to a buildup of negative charge in the spacecraft; and, subsequently, a portion of the positive ions used for propellant end up circling back around the magnetic fields and impacting the spacecraft. Not only is this something that’s a negative impact for the thrust of the spacecraft, but if the propellant is something that’s chemically active (as both Cs and Hg are), it can lead to chemical reactions with spacecraft structural components, sensors, and other systems, accelerating degradation of the spacecraft.
A while back on the Facebook group I asked the members about the use of these propellants, and an interesting discussion developed (primarily between Mikkel Haaheim, my head editor and frequent contributor to this blog, and Ed Pheil, who has extensive experience in nuclear power, including the JIMO mission, and is currently the head of Elysium Industries, developing a molten chloride fast reactor) concerning the pros and cons of using these propellants. Two other options, with their own complications from the engineering side, were also proposed, which we’ll touch on briefly: sodium and potassium both have low ionization energies, and form a low melting temperature eutectic, so they may offer additional options for future electrostatic propellants as well. Three major factors came up in the discussion: environmental and occupational health concerns during testing, propellant cost (which is a large part of what brings us to this discussion in the first place), and tankage considerations.
As far as cost goes, this is listed in the table above. These costs are all ballpark estimates, and costs for space-qualified supplies are generally higher, but it illustrates the general costs associated with each propellant. So, from an economic point of view, Cs is the least attractive, while Hg, Kr, and Na are all attractive options for bulk propellants.
Tankage in and of itself is a simpler question than the question of the full propellant feed question, however it can offer some insights into the overall challenges in storing and using the various propellants. Xe, our baseline propellant, has a density as a liquid of 2.942 g/cm, Kr of 2.413, and Hg of 13.53. All other things aside, this indicates that the overall tankage mass requirements for the same mass of Hg are less than 1/10th that of Xe or Kr. However, additional complications arise when considering tank material differences. For instance, both Xe and Kr require cryogenic cooling (something we discussed in the LEU NTP series briefly, which you can read here [insert LEU NTP 3 link]. While the challenges of Xe and Kr cryogenics are less difficult than H2 cryogenics due to the higher atomic mass and lower chemical reactivity, many of the same considerations do still apply. Hg on the other hand, has to be kept in a stainless steel tank (by law), other common containers, such as glass, don’t lend themselves to spacecraft tank construction. However, a stainless steel liner of a carbon composite tank is a lower-mass option.
The last type of fluid propellant to mention is NaK, a common fast reactor coolant which has been extensively studied. Many of the problems with tankage of NaK are similar to those seen in Cs or Hg: chemical reactivity (although different particulars on the tankage), however, all the research into using NaK for fast reactor coolant has largely addressed the immediate corrosion issues.
The main problem with NaK would be differential ionization causing plating of the higher-ionization-energy metal (Na in this case) onto the anode or propellant channels of the thruster. While it may be possible to deal with this, either by shortening the propellant channel (like in a TAL or EDPT), or by ensuring full ionization through excess charge in the anode and cathode. The possibility of using NaK was studied in an SPT thruster in the Soviet Union, but unfortunately I cannot find the papers associated with these studies. However, NaK remains an interesting option for future thrusters.
Solid propellants are generally considered to be condensable propellant thrusters. These designs have been studied for a number of decades. Most designs use a resistive heater to melt the propellant, which is then vaporized just before entering the anode. This was first demonstrated with the cesium contact gridded ion thrusters that were used as part of the SERT program. There (as mentioned earlier) a metal foam was used as the storage medium, which was kept warm to the point that the cesium was kept liquid. By varying the pore size, a metal wick was made which controlled the flow of the propellant from the reservoir to the ionization head. This results in a greater overall mass for the propellant tankage, but on the other hand the lack of moving parts, and the ability to ensure even heating across the propellant volume, makes this an attractive option in some cases.
A more recent design that we also discussed (the VHITAL) uses bismuth propellant for a TAL thruster, a NASA update of a Soviet TsNIIMash design from the 1970s (which was shelved due to the lack of high-powered space power systems at the time). This design uses a reservoir of liquid bismuth, which is resistively heated to above the melting temperature. An argon pressurization system is used to force the liquid bismuth through an outlet, where it’s then electromagnetically pumped into a carbon vaporization plug. This then discharges into the anode (which in the latest iteration is also resistively heated), where the Hall current then ionizes the propellant. It may be possible with this design to use multiple reservoirs to reduce the power demand for the propellant feed system; however, this would also lead to greater tankage mass requirements, so it will largely depend on the particulars of the system whether the increase in mass is worth the power savings of using a more modular system. This propellant system was successfully tested in 2007, and could be adapted to other designs as well.
Other propellants have been proposed as well, including magnesium, iodine, and cadmium. Each has its’ advantages and disadvantages in tankage, chemical reactivity limiting thruster materials considerations, and other factors, but all remain possible for future thruster designs.
For the foreseeable future, most designs will continue to use xenon, with argon being the next most popular choice, but as the amount of propellant needed increases with the development of nuclear electric propulsion, it’s possible that these other propellant options will become more prominent as tankage mass, propellant cost, and other considerations become more significant.
Electrospray thrusters use electrically charged liquids as a propellant. They fall into three main categories: colloid thrusters, which accelerate charged droplets dissolved in a solvent such as glycerol or formamide; field emission electric propulsion (FEEP) thrusters, which use liquid metals to produce positively charged metal ions; and, finally, ionic liquid ion source (ILIS) thrusters, which use room temperature molten salts to produce a beam of salt ions.
All types of electrospray end up demonstrating a phenomenon known as a Taylor cone, which occurs in an electrically charged fluid when exposed to an electrical field. If the field is strong enough, the tip of the cone is extruded to the point that it breaks, and a spray of droplets from the liquid is emitted. This is now commonly used in many different industrial applications, and the advances in these fields have made the electrospray thruster more attractive, as have a focus on volume of propulsion systems. Additionally, the amount of thrust produced, and the thrust density, is directly proportional to the density of emitters in a given area. Recent developments in nanomaterials fabrication have made it possible to increase the thrust density of these designs significantly. However, the main lifetime limitation of this type of thruster is emitter wear, which is dependent on both mass flow rates and any chemical interactions between the emitters and the propellant.
The vast majority of these systems focus on cube-sat propulsion; but one company, Accion Systems, has developed a tileable system which could offer high-powered operation through the use of dozens of thrusters arrayed in a grid. Their largest thruster (which measures 35mm by 35 mm by 16 mm, including propellant) produces a total of 200,000 N of impulse, a thrust of 10 mN, at an isp of 1500 s. While their primary focus is on cubesats, the CEO, Natalya Bailey, has mentioned before that it would be possible to use many of their TILE drive systems in parallel for high-powered missions.
One of the biggest power demands of an electrostatic engine of almost any type is the ionization cost of the propellant. Depending on the mass flow and power, different systems are used to ionize the propellant, including electron beams, RF ionization, cyclotron resonance, and the Hall effect. What if we could get rid of that power cost, and instead use all of the energy accelerating the propellant? Especially in small spacecraft, this is very attractive, and it may be possible to scale this up significantly as well (to the limits of the electrical charge that is able to be placed on the screens themselves). Some fluids are ionic, meaning that they’re positively charged, reasonably chemically stable, and easily storable. By replacing the uncharged propellant with one that carries an electric charge without the need for on-board ionization equipment, mass, volume, and power can be conserved. Not all electrospray thrusters use an ionic liquid, but ones that do offer considerable advantages in terms of energy efficiency, and possibly can offer greater overall thruster efficiency as well. I have yet to see a design for a gridded ion or Hall effect thruster that utilizes these types of propellants, but it may be possible to do so.
With that, we come to the end of our overview of electric thrusters. While there are some types of thruster that we did not discuss, they are unlikely to be able to be used in high powered systems such as would be found on an NEP spacecraft. When I began this series of blog posts, I knew that electric propulsion is a very broad topic, but the learning process during writing these three posts has been far more intense, and broad, than I was expecting. Electric propulsion has never been my strong suit, so I’ve been even more careful than usual to stick to the resources available to write these posts, and I’ve had a lot of help from some very talented people to get to this point.
I was initially planning on writing a post about the power conditioning units that are used to prepare the power provided by the power supply to these thrusters, but the more I researched, the less these systems made sense to me – something that I’ve been assured isn’t uncommon – so I’m going to skip that for now.
Instead, the next post is going to look at the power conversion systems that nuclear electric spacecraft can use. Due to the unique combination of available temperature from a nuclear reactor, the high power levels available, and the unique properties of in-space propulsion, there are many options available that aren’t generally considered for terrestrial power plants, and many designs that are used by terrestrial plants aren’t available due to mass or volume requirements. I’ve already started writing the post, but if there’s anything writing on NEP has taught me, it’s that these posts take longer than I expect, so I’m not going to give a timeline on when that will be available – hopefully in the next 2-3 weeks, though.
After that, we’ll look more in depth at thermal management and heat rejection systems for a wide range of temperatures, how they work, and the fundamental limitations that each type has. After another look at the core of an NEP spacecraft’s reactor, we will then look at combining electric and thermal propulsion in a post on bimodal NTRs, before moving on to our next blog post series (probably on pulse propulsion, but we may return to NTRs briefly to look at liquid core NTRs and the LARS proposal).
I hope you enjoyed the post. Leave a comment below with any comments, questions, or corrections, and don’t forget to check out our Facebook group, where I post work-in-progress visuals, papers I come across during research, and updates on the blog (and if you do, don’t feel shy about posting yourself on astronuclear propulsion designs and news!).
Gridded Ion Thrusters
MIT Open Courseware Astronautics Course Notes, Lecture 10-11: Kaufmann Ion Drives https://ocw.mit.edu/courses/aeronautics-and-astronautics/16-522-space-propulsion-spring-2015/lecture-notes/MIT16_522S15_Lecture10-11.pdf
NSTAR Technology Validation, Brophy et al 2000 https://trs.jpl.nasa.gov/handle/2014/13884
The High Power Electric Propulsion (HiPEP) Ion Thruster, Foster et al 2004 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20040139476.pdf
Dual stage 4 Grid thruster page, ASU: https://physics.anu.edu.au/cpf/sp3/ds4g/
Dual Stage Gridded Ion Thruster ESA page: http://www.esa.int/gsp/ACT/projects/ds4g_overview.html
The NASA Evolutionary Xenon Thruster: The Next Step for U.S. Deep Space Propulsion, Schmidt et al NASA GRC 2008 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20080047732.pdf
Deflectable Beam Linear Strip Cesium Contact Ion Thruster System, Dulgeroff et al 1971 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19710027900.pdf
Hall Effect Thrusters
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
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
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
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
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
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
High Power Hall Thrusters; Jankovsky et al, 1999
Energetics of Propellant Options for High-Power Hall Thrusters, Kieckhafer and King, Michigan Technological University, 2005 http://aerospace.mtu.edu/__reports/Conference_Proceedings/2005_Kieckhafer_1.pdf
A Performance Comparison Of Xenon and Krypton Propellant on an SPT-100 Hall Thruster, Nakles et al 2011 http://erps.spacegrant.org/uploads/images/images/iepc_articledownload_1988-2007/2011index/IEPC-2011-003.pdf
Evaluation of Magnesium as Hall Thruster Propellant; Hopkins, Michigan Tech, 2015 https://pdfs.semanticscholar.org/0520/494153e1d19a46eaa0a63c9bc5bd466c06eb.pdf
Modeling of an Iodine Hall Thruster Plume in the Iodine Satellite (ISAT), Choi 2017 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20170001565.pdf
Electrospray Thruster lecture notes, MIT Open Courseware 2015 https://ocw.mit.edu/courses/aeronautics-and-astronautics/16-522-space-propulsion-spring-2015/lecture-notes/MIT16_522S15_Lecture20.pdf
Application of Ion Electrospray Propulsion to Lunar and Interplanetary Missions, Whitlock and Lozano 2014 http://ssl.mit.edu/files/website/theses/SM-2014-WhitlockCaleb.pdf
Preliminary Sizing of an Electrospray Thruster, Rostello 2017 http://tesi.cab.unipd.it/56748/1/Rostello_Marco_1109399.pdf
Accion Systems TILE Data sheet https://beyondnerva.files.wordpress.com/2018/10/bd6da-tileproductfamilycombineddatasheet.pdf