Electrostatic 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)
Xenon 131.293 1170.4 2.942 (BP) 161.4 165.051 1200
Krypton 83.798 1350.8 2.413 (BP) 115.78 119.93 75
Bismuth 208.98 703 10.05 (MP) 544.7 1837 29
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
Argon 39.792 1520.6 1.395 (BP) 83.81 87.302 5
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
Magnesium 24.304 737.7 1.584 (MP) 923 1363 6
Cadmium 112.414 867.8 7.996 (MP) 594.22 1040 5

 

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. 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.

Sources

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

1 reply »

Leave a Reply

Fill in your details below or click an icon to log in:

WordPress.com Logo

You are commenting using your WordPress.com account. Log Out /  Change )

Google+ photo

You are commenting using your Google+ account. Log Out /  Change )

Twitter picture

You are commenting using your Twitter account. Log Out /  Change )

Facebook photo

You are commenting using your Facebook account. Log Out /  Change )

Connecting to %s