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Pebblebed NTRs: Solid Fuel, but Different

Hello, and welcome back to Beyond NERVA!

Today, we’re going to take a break from the closed cycle gas core nuclear thermal rocket (which I’ve been working on constantly since mid-January) to look at one of the most popular designs in modern NTR history: the pebblebed reactor!

This I should have covered between solid and liquid fueled NTRs, honestly, and there’s even a couple types of reactor which MAY be able to be used for NTR between as well – the fluidized and shush fuel reactors – but with the lack of information on liquid fueled reactors online I got a bit zealous.

Beads to Explore the Solar System

Most of the solid fueled NTRs we’ve looked at have been either part of, or heavily influenced by, the Rover and NERVA programs in the US. These types of reactors, also called “prismatic fuel reactors,” use a solid block of fuel of some form, usually tileable, with holes drilled through each fuel element.

The other designs we’ve covered fall into one of two categories, either a bundled fuel element, such as the Russian RD-0410, or a folded flow disc design such as the Dumbo or Tricarbide Disc NTRs.

However, there’s another option which is far more popular for modern American high temperature gas cooled reactor designs: the pebblebed reactor. This is a clever design, which increases the surface area of the fuel by using many small, spherical fuel elements held in a (usually) unfueled structure. The coolant/propellant passes between these beads, picking up the heat as it passes between them.

This has a number of fundamental advantages over the prismatic style fuel elements:

  1. The surface area of the fuel is so much greater than with simple holes drilled in the prismatic fuel elements, increasing thermal transfer efficiency.
  2. Since all types of fuel swell when heated, the density of the packed fuel elements could be adjusted to allow for better thermal expansion behavior within the active region of the reactor.
  3. The fuel elements themselves were reasonably loosely contained within separate structures, allowing for higher temperature containment materials to be used.
  4. The individual elements could be made smaller, allowing for a lower temperature gradient from the inside to the outside of a fuel, reducing the overall thermal stress on each fuel pebble.
  5. In a folded flow design, it was possible to not even have a physical structure along the inside of the annulus if centrifugal force was applied to the fuel element structure (as we saw in the fluid fueled reactor designs), eliminating the need for as many super-high temperature materials in the highest temperature region of the reactor.
  6. Because each bead is individually clad, in the case of an accident during launch, even if the reactor core is breached and a fuel release into the environment occurs, the release of either any radiological components or any other fuel materials into the environment is minimized
  7. Because each bead is relatively small, it is less likely that they will sustain sufficient damage either during mechanical failure of the flight vehicle or impact with the ground that would breach the cladding.

However, there is a complication with this design type as well, since there are many (usually hundreds, sometimes thousands) of individual fuel elements:

  1. Large numbers of fuel beads mean large numbers of fuel beads to manufacture and perform quality control checks on.
  2. Each bead will need to be individually clad, sometimes with multiple barriers for fission product release, hydrogen corrosion, and the like.
  3. While each fuel bead will be individually clad, and so the loss of one or all the fuel will not significantly impact the environment from a radiological perspective in the case of an accident, there is potential for significant geographic dispersal of the fuel in the event of a failure-to-orbit or other accident.

There are a number of different possible flow paths through the fuel elements, but the two most common are either an axial flow, where the propellant passes through a tubular structure packed with the fuel elements, and a folded flow design, where the fuel is in a porous annular structure, with the coolant (usually) passing from the outside of the annulus, through the fuel, and the now-heated coolant exiting through the central void of the annulus. We’ll call these direct flow and folded flow pebblebed fuel elements.

In addition, there are many different possible fuel types, which regulars of this blog will be familiar with by now: oxides, carbides, nitrides, and CERMET are all possible in a pebblebed design, and if differential fissile fuel loading is needed, or gradients in fuel composition (such as using tungsten CERMET in higher temperature portions of the reactor, with beryllium or molybdenum CERMET in lower temperature sections), this can be achieved using individual, internally homogeneous fuel types in the beads, which can be loaded into the fuel support structure at the appropriate time to create the desired gradient.

Just like in “regular” fuel elements, these pebbles need to be clad in a protective coating. There have been many proposals over the years, obviously depending on what type of fissile fuel matrix the fuel uses to ensure thermal expansion and chemical compatibility with the fuel and coolant. Often, multiple layers of different materials are used to ensure structural and chemical integrity of the fuel pellets. Perhaps the best known example of this today is the TRISO fuel element, used in the US Advanced Gas Reactor fuel development program. The TRI-Structural ISOtropic fuel element uses either oxide or carbide fuel in the center, followed by a porous carbon layer, a pyrolitic carbon layer (sort of like graphite, but with some covalent bonds between the carbon sheets), followed by a silicon carbide outer shell for mechanical and fission product retention. Some variations include a burnable poison for reactivity control (the QUADRISO at Argonne), or use different outer layer materials for chemical protection. Several types have been suggested for NTR designs, and we’ll see more of them later.

The (sort of) final significant variable is the size of the pebble. As the pebbles go down in size, the available surface area of the fuel-to-coolant interface increases, but also the amount of available space between the pebbles decreases and the path that the coolant flows through becomes more resistant to higher coolant flow rates. Depending on the operating temperature and pressure, the thermal gradient acceptable in the fuel, the amount of decay heat that you want to have to deal with on shutdown (the bigger the fuel pebble, the more time it will take to cool down), fissile fuel density, clad thickness requirements, and other variables, a final size for the fuel pebbles can be calculated, and will vary to a certain degree between different reactor designs.

Not Just for NTRs: The Electricity Generation Potential of Pebblebed Reactors

Obviously, the majority of the designs for pebblebed reactors are not meant to ever fly in space, they’re mostly meant to operate as high temperature gas cooled reactors on Earth. This type of architecture has been proposed for astronuclear designs as well, although that isn’t the focus of this video.

Furthermore, the pebblebed design lends itself to other cooling methods, such as molten salt, liquid metal, and other heat-carrying fluids, which like the gas would flow through the fuel pellets, pick up the heat produced by the fissioning fuel, and carry it into a power conversion system of whatever design the reactor has integrated into its systems.

Finally, while it’s rare, pebblebed designs were popular for a while with radioisotope power systems. There are a number of reasons for this beyond being able to run a liquid coolant through the fuel (which was done on one occasion that I can think of, and we’ll cover in a future post): in an alpha-emitting radioisotope, such as 238Pu, over time the fuel will generate helium gas – the alpha particles will slow, stop, and become doubly ionized helium nuclei, which will then strip electrons off whatever materials are around and become normal 4He. This gas needs SOMEWHERE to go, which is why just like with a fissile fuel structure there are gas management mechanisms used in radioisotope power source fuel assemblies such as areas of vacuum, pressure relief valves, and the like. In some types of RTG, such as the SNAP-27 RTG used by Apollo, as well as the Multi-Hundred Watt RTG used by Voyager, the fuel was made into spheres, with the gaps in between the spheres (normally used to pass coolant through) are used for the gas expansion volume.

We’ll discuss these ideas more in the future, but I figured it was important to point out here. Let’s get back to the NTRs, and the first (and only major) NTR program to focus on the pebblebed concept: the Project Timberwind and the Space Nuclear Propulsion Program in the 1980s and early 1990s.

The Beginnings of Pebblebed NTRs

The first proposals for a gas cooled pebblebed reactor were from 1944/45, although they were never pursued beyond the concept stage, and a proposal for the “Space Vehicle Propulsion Reactor” was made by Levoy and Newgard at Thikol in 1960, with again no further development. If you can get that paper, I’d love to read it, here’s all I’ve got: “Aero/Space Engineering 19, no. 4, pgs 54-58, April 1960” and ‘AAE Journal, 68, no. 6, pgs. 46-50, June 1960,” and “Engineering 189, pg 755, June 3, 1960.” Sounds like they pushed hard, and for good reason, but at the time a pebblebed reactor was a radical concept for a terrestrial reactor, and getting a prismatic fueled reactor, something far more familiar to nuclear engineers, was a challenge that seemed far simpler and more familiar.

Sadly, while this design may end up have informed the design of its contemporary reactor, it seems like this proposal was never pursued.

Rotating Fluidized Bed Reactor (“Hatch” Reactor) and the Groundwork for Timberwind

Another proposal was made at the same time at Brookhaven National Laboratory, by L.P. Hatch, W.H. Regan, and a name that will continue to come up for the rest of this series, John R. Powell (sorry, can’t find the given names of the other two, even). This relied on very small (100-500 micrometer) fuel, held in a perforated drum to contain the fuel but also allow propellant to be injected into the fuel particles, which was spun at a high rate to provide centrifugal force to the particles and prevent them from escaping.

Now, fluidized beds need a bit of explanation, which I figured was best to put in here since this is not a generalized property of pebblebed reactors. In this reactor (and some others) the pebbles are quite small, and the coolant flow can be quite high. This means that it’s possible – and sometimes desirable – for the pebbles to move through the active zone of the reactor! This type of mobile fuel is called a “fluidized bed” reactor, and comes in several variants, including pebble (solid spheres), slurry (solid particulate suspended in a liquid), and colloid (solid particulate suspended in a gas). The best way to describe the phenomenon is with what is called the point of minimum fluidization, or when the drag forces on the mass of the solid objects from the fluid flow balances with the weight of the bed (keep in mind that life is a specialized form of drag). There’s a number of reasons to do this – in fact, many chemical reactions using a solid and a fluid component use fluidization to ensure maximum mixing of the components. In the case of an NTR, the concern is more to do with achieving as close to thermal equilibrium between the solid fuel and the gaseous propellant as possible, while minimizing the pressure drop between the cold propellant inlet and the hot propellant outlet. For an NTR, the way that the “weight” is applied is through centrifugal force on the fuel. This is a familiar concept to those that read my liquid fueled NTR series, but actually began with the fluidized bed concept.

This is calculated using two different relations between the same variables: the Reynolds number (Re), which determines how turbulent fluid flow is, and the friction coefficient (CD, or coefficient of drag, which deptermines how much force acts on the fuel particles based on fluid interactions with the particles) which can be found plotted below. The plotted lines represent either the Reynolds number or the void fraction ε, which represents the amount of gas present in the volume defined by the presence of fuel particles.

Hendrie 1970

If you don’t follow the technical details of the relationships depicted, that’s more than OK! Basically, the y axis is proportional to the gas turbulence, while the x axis is proportional to the particle diameter, so you can see that for relatively small increases in particle size you can get larger increases in propellant flow rates.

The next proposal for a pebble bed reactor grew directly out of the Hatch reactor, the Rotating Fluidized Bed Reactor for Space Nuclear Propulsion (RBR). From the documentation I’ve been able to find, from the original proposal work continued at a very low level at BNL from the time of the original proposal until 1973, but the only reports I’ve been able to find are from 1971-73 under the RBR name. A rotating fuel structure, with small, 100-500 micrometer spherical particles of uranium-zirconium carbide fuel (the ZrC forming the outer clad and a maximum U content of 10% to maximize thermal limits of the fuel particles), was surrounded by a reflector of either metallic beryllium or BeO (which was preferred as a moderator, but the increased density also increased both reactor mass and manufacturing requirements). Four drums in the reflector would control the reactivity of the engine, and an electric motor would be attached to a porous “squirrel cage” frit, which would rotate to contain the fuel.

Much discussion was had as to the form of uranium used, be it 235U or 233U. In the 235U reactor, the reactor had a cavity length of 25 in (63.5 cm), an inner diameter of 25 in (63.5 cm), and a fuel bed depth when fluidized of 4 in (10.2 cm), with a critical mass of U-ZrC being achieved at 343.5 lbs (155.8 kg) with 9.5% U content. The 233U reactor was smaller, at 23 in (56 cm) cavity length, 20 in (51 cm) bed inner diameter, 3 in (7.62 cm) deep fuel bed with a higher (70%) void fraction, and only 105.6 lbs (47.9 kg) of U-ZrC fuel at a lower (and therefore more temperature-tolerant) 7.5% U loading.

233U was the much preferred fuel in this reactor, with two options being available to the designers: either the decreased fuel loading could be used to form the smaller, higher thrust-to-weight ratio engine described above, or the reactor could remain at the dimensions of the 235U-fueled option, but the temperature could be increased to improve the specific impulse of the engine.

There was als a trade-off between the size of the fuel particles and the thermal efficiency of the reactor,:

  • Smaller particles advantages
    • Higher surface area, and therefore better thermal transfer capabilities,
    • Smaller radius reduces thermal stresses on fuel
  • Smaller particles disadvantages
    • Fluidized particle bed fuel loss would be a more immediate concern
    • More sensitive to fluid dynamic behavior in the bed
    • Bubbles could more easily form in fuel
    • Higher centrifugal force required for fuel containment
  • Larger particle advantages
    • Ease of manufacture
    • Lower centrifugal force requirements for a given propellant flow rate
  • Larger particle disadvantages
    • Higher thermal gradient and stresses in fuel pellets
    • Less surface area, so lower thermal transfer efficiency

It would require testing to determine the best fuel particle size, which could largely be done through cold flow testing.

These studies looked at cold flow testing in depth. While this is something that I’ve usually skipped over in my reporting on NTR development, it’s a crucial type of testing in any gas cooled reactor, and even more so in a fluidized bed NTR, so let’s take a look at what it’s like in a pebblebed reactor: the equipment, the data collection, and how the data modified the reactor design over time.

Cold flow testing is usually the predecessor to electrically heated flow testing in an NTR. These tests determine a number of things, including areas within the reactor that may end up with stagnant propellant (not a good thing), undesired turbulence, and other negative consequences to the flow of gas through the reactor. They are preliminary tests, since as the propellant heats up while going through the reactor, a couple major things will change: first, the density of the gas will decrease and second, as the density changes the Reynolds number (a measure of self-interaction, viscosity, and turbulent vs laminar flow behavior) will change.

In this case, the cold flow tests were especially useful, since one of the biggest considerations in this reactor type is how the gas and fuel interact.

The first consideration that needed to be examined is the pressure drop across the fuel bed – the highest pressure point in the system is always the turbopump, and the pressure will decrease from that point throughout the system due to friction with the pipes carrying propellant, heating effects, and a host of other inefficiencies. One of the biggest questions initially in this design was how much pressure would be lost from the frit (the outer containment structure and propellant injection system into the fuel) to the central void in the body of the fuel, where it exits the nozzle. Happily, this pressure drop is minimal: according to initial testing in the early 1960s (more on that below), the pressure drop was equal to the weight of the fuel bed.

The next consideration was the range between fluidizing the fuel and losing the fuel through literally blowing it out the nozzle – otherwise known as entrainment, a problem we looked at extensively on a per-molecule basis in the liquid fueled NTR posts (since that was the major problem with all those designs). Initial calculations and some basic experiments were able to map the propellant flow rate and centrifugal force required to both get the benefit of a fluidized bed and prevent fuel loss.

Rotating Fluidized Bed Reactor testbed test showing bubble formation,

Another concern is the formation of bubbles in the fuel body. As we covered in the bubbler LNTR post (which you can find here), bubbles are a problem in any fuel type, but in a fluid fueled reactor with coolant passing through it there’s special challenges. In this case, the main method of transferring heat from the fuel to the propellant is convection (i.e. contact between the fuel and the propellant causing vortices in the gas which distributes the heat), so an area that doesn’t have any (or minimal) fuel particles in it will not get heated as thoroughly. That’s a headache not only because the overall propellant temperature drops (proportional to the size of the bubbles), but it also changes the power distribution in the reactor (the bubbles are fission blank spots).

Finally, the initial experiment set looked at the particle-to-fluid thermal transfer coefficients. These tests were far from ideal, using a 1 g system rather than the much higher planned centrifugal forces, but they did give some initial numbers.

The first round of tests was done at Brookhaven National Laboratory (BNL) from 1962 to 1966, using a relatively simple test facility. A small, 10” (25.4 cm) length by 1” (2.54 cm) diameter centrifuge was installed, with gas pressure provided by a pressurized liquefied air system. 138 to 3450 grams of glass particles were loaded into the centrifuge, and various rotational velocities and gas pressures were used to test the basic behavior of the particles under both centrifugal force and gas pressure. While some bobbles were observed, the fuel beds remained stable and no fuel particles were lost during testing, a promising beginning.

These tests provided not just initial thermal transfer estimates, pressure drop calculations, and fuel bed behavioral information, but also informed the design of a new, larger test rig, this one 10 in by 10 in (25.4 by 25.4 cm), which was begun in 1966. This system would not only have a larger centrifuge, but would also use liquid nitrogen rather than liquefied air, be able to test different fuel particle simulants rather than just relatively lightweight glass, and provide much more detailed data. Sadly, the program ran out of funding later that year, and the partially completed test rig was mothballed.

Rotating Fluidized Bed Reactor (RBR): New Life for the Hatch Reactor

It would take until 1970, when the Space Nuclear Systems office of the Atomic Energy Commission and NASA provided additional funding to complete the test stand and conduct a series of experiments on particle behavior, reactor dynamics and optimization, and other analytical studies of a potential advanced pebblebed NTR.

The First Year: June 1970-June 1971

After completing the test stand, the team at BNL began a series of tests with this larger, more capable equipment in Building 835. The first, most obvious difference is the diameter of the centrifuge, which was upgraded from 1 inch to 10 inches (25.4 cm), allowing for a more prototypical fuel bed depth. This was made out of perforated aluminum, held in a stainless steel pressure housing for feeding the pressurized gas through the fuel bed. In addition, the gas system was changed from the pressurized air system to one designed to operate on nitrogen, which was stored in liquid form in trailers outside the building for ease of refilling (and safety), then pre-vaporized and held in two other, high-pressure trailers.

Photographs were used to record fluidization behavior, taken viewing the bottom of the bed from underneath the apparatus. While initially photos were only able to be taken 5 seconds apart, later upgrades would improve this over the course of the program.

The other major piece of instrumentation surrounded the pressure and flow rate of the nitrogen gas throughout the system. The gas was introduced at a known pressure through two inlets into the primary steel body of the test stand, with measurements of upstream pressure, cylindrical cavity pressure outside the frit, and finally a pitot tube to measure pressure inside the central void of the centrifuge.

Three main areas of pressure drop were of interest: due to the perforated frit itself, the passage of the gas through the fuel bed, and finally from the surface of the bed and into the central void of the centrifuge, all of which needed to be measured accurately, requiring calibration of not only the sensors but also known losses unique to the test stand itself.

The tests themselves were undertaken with a range of glass particle sizes from 100 to 500 micrometers in diameter, similar to the earlier tests, as well as 500 micrometer copper particles to more closely replicate the density of the U-ZrC fuel. Rotation rates of between1,000 and 2,000 rpm, and gas flow rates from 1,340-1,800 scf/m (38-51 m^3/min) were used with the glass beads, and from 700-1,500 rpm with the copper particles (the lower rotation rate was due to gas pressure feed limitations preventing the bed from becoming fully fluidized with the more massive particles).

Finally, there were a series of physics and mechanical engineering design calculations that were carried out to continue to develop the nuclear engineering, mechanical design, and system optimization of the final RBR.

The results from the initial testing were promising: much of the testing was focused on getting the new test stand commissioned and calibrated, with a focus on figuring out how to both use the rig as it was constructed as well as which parts (such as the photography setup) could be improved in the next fiscal year of testing. However, particle dynamics in the fuidized bed were comfortably within stable, expected behavior, and while there were interesting findings as to the variation in pressure drop along the axis of the central void, this was something that could be worked with.

Based on the calculations performed, as well as the experiments carried out in the first year of the program, a range of engines were determined for both 233U and 235U variants:

Work Continues: 1971-1972

This led directly into the 1971-72 series of experiments and calculations. Now that the test stand had been mostly completed (although modifications would continue), and the behavior of the test stand was now well-understood, more focused experimentation could continue, and the calculations of the physics and engineering considerations in the reactor and engine system could be advanced on a more firm footing.

One major change in this year’s design choices was the shift toward a low-thrust, high-isp system, in part due to greater interest at NASA and the AEC in a smaller NTR than the original design envelope. While analyzing the proposed engine size above, though, it was discovered that the smallest two reactors were simply not practical, meaning that the smallest design was over 1 GW power level.

Another thing that was emphasized during this period from the optimization side of the program was the mass of the reflector. Since the low thrust option was now the main thrust of the design, any increase in the mass of the reactor system has a larger impact on the thrust-to-weight ratio, but reducing the reflector thickness also increases the neutron leakage rate. In order to prevent this, a narrower nozzle throat is preferred, but also increases thermal loading across the throat itself, meaning that additional cooling, and probably more mass, is needed – especially in a high-specific-impulse (aka high temperature) system. This also has the effect of needing higher chamber pressures to maintain the desired thrust level (a narrower throat with the same mass flow throughput means that the pressure in the central void has to be higher).

These changes required a redesign of the reactor itself, with a new critical configuration:

Hendrie 1972

One major change is how fluidized the bed actually is during operation. In order to get full fluidization, there needs to be enough inward (“upward” in terms of force vectors) velocity at the inner surface of the fuel body to lift the fuel particles without losing them out the nozzle. During calculations in both the first and second years, two major subsystems contributed hugely to the weight and were very dependent on both the rotational speed and the pellet size/mass: the weight of the frit and motor system, which holds the fuel particles, and the weight of the nozzle, which not only forms the outlet-end containment structure for the fuel but also (through the challenges of rocket motor dynamics) is linked to the chamber pressure of the reactor – oh, and the narrower the nozzle, the less surface area is available to reject the heat from the propellant, so the harder it is to keep cool enough that it doesn’t melt.

Now, fluidization isn’t a binary system: a pebblebed reactor is able to be settled (no fluidization), partially fluidized (usually expressed as a percentage of the pebblebed being fluidized), and fully fluidized to varying degrees (usually expressed as a percentage of the volume occupied by the pebbles being composed of the fluid). So there’s a huge range, from fully settled to >95% fluid in a fully fluidized bed.

The designers of the RBR weren’t going for excess fluidization: at some point, the designer faces diminishing returns on the complications required for increased fluid flow to maintain that level of particulate (I’m sure it’s the same, with different criteria, in the chemical industry, where most fluidized beds actually are used), both due to the complications of having more powerful turbopumps for the hydrogen as well as the loss of thermalization of that hydrogen because there’s simply too much propellant to be heated fully – not to mention fuel loss from the particulate fuel being blown out of the nozzle – so the calculations for the bed dynamics assumed minimal full fluidization (i.e. when all the pebbles are moving in the reactor) as the maximum flow rate – somewhere around 70% gas in the fuel volume (that number was never specifically defined that I found in the source documentation, if it was, please let me know), but is dependent on both the pressure drop in the reactor (which is related to the mass of the particle bed) and the gas flow.

Ludewig 1974

However, the designers at this point decided that full fluidization wasn’t actually necessary – and in fact was detrimental – to this particular NTR design. Because of the dynamics of the design, the first particles to be fluidized were on the inner surface of the fuel bed, and as the fluidization percentage increased, the pebbles further toward the outer circumference became fluidized. Because the temperature difference between the fuel and the propellant is greater as the propellant is being injected through the frit and into the fuel body, more heat is carried away by the propellant per unit mass, and as the propellant warms up, thermal transfer becomes less efficient (the temperature difference between two different objects is one of the major variables in how much energy is transferred for a given surface area), and fluidization increases that efficiency between a solid and a fluid.

Because of this, the engineers re-thought what “minimal fluidization” actually meant. If the bed could be fluidized enough to maximize the benefit of that dynamic, while at a minimum level of fluidization to minimize the volume the pebblebed actually took up in the reactor, there would be a few key benefits:

  1. The fueled volume of the reactor could be smaller, meaning that the nozzle could be wider, so they could have lower chamber pressure and also more surface area for active cooling of the nozzle
  2. The amount of propellant flow could be lower, meaning that turbopump assemblies could be smaller and lighter weight
  3. The frit could be made less robustly, saving on weight and simplifying the challenges of the bearings for the frit assembly
  4. The nozzle, frit, and motor/drive assembly for the frit are all net neutron poisons in the RBR, meaning that minimizing any of these structures’ overall mass improves the neutron economy in the reactor, leading to either a lower mass reactor or a lower U mass fraction in the fuel (as we discussed in the 233U vs. 235U design trade-off)

After going through the various options, the designers decided to go with a partially fluidized bed. At this point in the design evolution, they decided on having about 50% of the bed by mass being fluidized, with the rest being settled (there’s a transition point in the fuel body where partial fluidization is occurring, and they discuss the challenges of modeling that portion in terms of the dynamics of the system briefly). This maximizes the benefit at the circumference, where the thermal difference (and therefore the thermal exchange between the fuel and the propellant) is most efficient, while also thermalizing the propellant as much as possible as the temperature difference decreases from the propellant becoming increasingly hotter. They still managed to reach an impressive 2400 K propellant cavity temperature with this reactor, which makes it one of the hottest (and therefore highest isp) solid core NTR designs proposed at that time.

This has various implications for the reactor, including the density of the fissile component of the fuel (as well as the other solid components that make up the pebbles), the void fraction of the reactor (what part of the reactor is made up of something other than fuel, in this particular instance hydrogen within the fuel), and other components, requiring a reworking of the nuclear modeling for the reactor.

An interesting thing to me in the Annual Progress Report (linked below) is the description of how this new critical configuration was modeled; while this is reasonably common knowledge in nuclear engineers from the days before computational modeling (and even to the present day), I’d never heard someone explain it in the literature before.

Basically, they made a bunch of extremely simplified (in both number of dimensions and fidelity) one-dimensional models of various points in the reactor. They then assumed that they could rotate these around that elevation to make something like an MRI slice of the nuclear behavior in the reactor. Then, they moved far enough away that it was different enough (say, where the frit turns in to the middle of the reactor to hold the fuel, or the nozzle starts, or even the center of the fuel compared to the edge) that the dynamics would change, and did the same sort of one-dimensional model; they would end up doing this 18 times. Then, sort of like an MRI in reverse, they took these models, called “few-group” models, and combined them into a larger group – called a “macro-group” – for calculations that were able to handle the interactions between these different few-group simulations to build up a two-dimensional model of the reactor’s nuclear structure and determine the critical configuration of the reactor. They added a few other ways to subdivide the reactor for modeling, for instance they split the neutron spectrum calculations into fast and thermal, but this is the general shape of how nuclear modeling is done.

Ok, let’s get back to the RBR…

Experimental testing using the rotating pebblebed simulator continued through this fiscal year, with some modifications. A new, seamless frit structure was procured to eliminate some experimental uncertainty, the pressure measuring equipment was used to test more areas of the pressure drop across the system, and a challenge for the experimental team – finding 100 micrometer copper spheres that were regularly enough shaped to provide a useful analogue to the UC-ZrC fuel (Cu specific gravity 8.9, UC-ZrC specific gravity ~6.5) were finally able to be procured.

Additionally, while thermal transfer experiments had been done with the 1-gee small test apparatus which preceded the larger centrifugal setup (with variable gee forces available), the changes were too great to allow for accurate predictions on thermal transfer behavior. Therefore, thermal transfer experiments began to be examined on the new test rig – another expansion of the capabilities of the new system, which was now being used rigorously since its completing and calibration testing of the previous year. While they weren’t conducted that year, setting up an experimental program requires careful analysis of what the test rig is capable of, and how good data accuracy can be achieved given the experimental limitations of the design.

The major achievement for the year’s ex[experimentation was a refining of the relationship between particle size, centrifugal force, and pressure drop of the propellant from the turbopump to the frit inlet to the central cavity, most especially from the frit to the inner cavity through the fuel body, on a wide range of particle sizes, flow rates, and bed fluidization levels, which would be key as the design for the RBR evolved.

The New NTR Design: Mid-Thrust, Small RBR

So, given the priorities at both the AEC and NASA, it was decided that it was best to focus primarily on a given thrust, and try and optimize thrust-to-weight ratios for the reactor around that thrust level, in part because the outlet temperature of the reactor – and therefore the specific impulse – was fixed by the engineering decisions made in regards to the rest of the reactor design. In this case, the target thrust was was 90 kN (20,230 lbf), or about 120% of a Pewee-class engine.

This, of course, constrained the reactor design, which at this point in any reactor’s development is a good thing. Every general concept has a huge variety of options to play with: fuel type (oxide, carbide, nitride, metal, CERMET, etc), fissile component (233U and 235U being the big ones, but 242mAm, 241Cf, and other more exotic options exist), thrust level, physical dimensions, fuel size in the case of a PBR, and more all can be played with to a huge degree, so having a fixed target to work towards in one metric allows a reference point that the rest of the reactor can work around.

Also, having an optimization point to work from is important, in this case thrust-to-weight ratio (T/W). Other options, such as specific impulse, for a target to maximize would lead to a very different reactor design, but at the time T/W was considered the most valuable consideration since one way or another the specific impulse would still be higher than the prismatic core NTRs currently under development as part of the NERVA program (being led by Los Alamos Scientific Laboratory and NASA, undergoing regular hot fire testing at the Jackass Flats, NV facility). Those engines, while promising, were limited by poor T/W ratios, so at the time a major goal for NTR improvement was to increase the T/W ratio of whatever came after – which might have been the RBR, if everything went smoothly.

One of the characteristics that has the biggest impact on the T/W ratio in the RBR is the nozzle throat diameter. The smaller the diameter, the higher the chamber pressure, which reduces the T/W ratio while increasing the amount of volume the fuel body can occupy given the same reactor dimensions – meaning that smaller fuel particles could be used, since there’s less chance that they would be lost out of the narrower nozzle throat. However, by increasing the nozzle throat diameter, the T/W ratio improved (up to a point), and the chamber pressure could be decreased, but at the cost of a larger particle size; this increases the thermal stresses in the fuel particles, and makes it more likely that some of them would fail – not as catastrophic as on a prismatic fueled reactor by any means, but still something to be avoided at all costs. Clearly a compromise would need to be reached.

Here are some tables looking at the design options leading up to the 90 kN engine configuration with both the 233U and 235U fueled versions of the RBR:

After analyzing the various options, a number of lessons were learned:

  1. It was preferable to work from a fixed design point (the 90 kN thrust level), because while the reactor design was flexible, operating near an optimized power level was more workable from a reactor physics and thermal engineering point of view
  2. The main stress points on the design were reflector weight (one of the biggest mass components in the system), throat diameter (from both a mass and active cooling point of view as well as fuel containment), and particle size (from a thermal stress and heat transfer point of view)
  3. On these lower-trust engines, 233U was looking far better than 235U for the fissile component, with a T/W ratio (without radiation shielding) of 65.7 N/kg compared to 33.3 N/kg respectively
    1. As reactor size increased, this difference reduced significantly, but with a constrained thrust level – and therefore reactor power – the difference was quite significant.

The End of the Line: RBR Winds Down

1973 was a bad year in the astronuclear engineering community. The flagship program, NERVA, which was approaching flight ready status with preparations for the XE-PRIME test, the successful testing of the flexible, (relatively) inexpensive Nuclear Furnace about to occur to speed not only prismatic fuel element development but also a variety of other reactor architectures (such as the nuclear lightbulb we began looking at last time), and the establishment of a robust hot fire testing structure at Jackass Flats, was fighting for its’ life – and its’ funding – in the halls of Congress. The national attention, after the success of Apollo 11, was turning away from space, and the missions that made NTR technologically relevant – and a good investment – were disappearing from the mission planners’ “to do” lists, and migrating to “if we only had the money” ideas. The Rotating Fluidized Bed Reactor would be one of those casualties, and wouldn’t even last through the 1971/72 fiscal year.

This doesn’t mean that more work wasn’t done at Brookhaven, far from it! Both analytical and experimental work would continue on the design, with the new focus on the 90 kN thrust level, T/W optimized design discussed above making the effort more focused on the end goal.

Multi-program computational architecture used in 1972/73 for RBR, Hoffman 1973

On the analytical side, many of the components had reasonably good analytical models independently, but they weren’t well integrated. Additionally, new and improved analytical models for things like the turbopump system, system mass, temp and pressure drop in the reactor, and more were developed over the last year, and these were integrated into a unified modeling structure, involving multiple stacked models. For more information, check out the 1971-72 progress report linked in the references section.

The system developed was on the verge of being able to do dynamics modeling of the proposed reactor designs, and plans were laid out for what this proposed dynamic model system would look like, but sadly by the time this idea was mature enough to implement, funding had run out.

On the experimental side, further refinement of the test apparatus was completed. Most importantly, because of the new design requirements, and the limitations of the experiments that had been conducted so far, the test-bed’s nitrogen supply system had to be modified to handle higher gas throughput to handle a much thicker fuel bed than had been experimentally tested. Because of the limited information about multi-gee centrifugal force behavior in a pebblebed, the current experimental data could only be used to inform the experimental course needed for a much thicker fuel bed, as was required by the new design.

Additionally, as was discussed from the previous year, thermal transfer testing in the multi-gee environment was necessary to properly evaluate thermal transfer in this novel reactor configuration, but the traditional methods of thermal transfer simply weren’t an option. Normally, the procedure would be to subject the bed to alternating temperatures of gas: cold gas would be used to chill the pebbles to gas-ambient temperatures, then hot gas would be used on the chilled pebbles until they achieved thermal equilibrium at the new temperature, and then cold gas would be used instead, etc. The temperature of the exit gas, pebbles, and amount of gas (and time) needed to reach equilibrium states would be analyzed, allowing for accurate heat transfer coefficients at a variety of pebble sizes, centrifugal forces, propellant flow rates, etc. would be able to be obtained, but at the same time this is a very energy-intensive process.

An alternative was proposed, which would basically split the reactor’s propellant inlet into two halves, one hot and one cold. Stationary thermocouples placed through the central void in the centrifuge would record variations in the propellant at various points, and the gradient as the pebbles moved from hot to cold gas and back could get good quality data at a much lower energy cost – at the cost of data fidelity reducing in proportion to bed thickness. However, for a cash-strapped program, this was enough to get the data necessary to proceed with the 90 kN design that the RBR program was focused on.

Looking forward, while the team knew that this was the end of the line as far as current funding was concerned, they looked to how their data could be applied most effectively. The dynamics models were ready to be developed on the analytical side, and thermal cycling capability in the centrifugal test-bed would prepare the design for fission-powered testing. The plan was to address the acknowledged limitations with the largely theoretical dynamic model with hot-fired experimental data, which could be used to refine the analytical capabilities: the more the system was constrained, and the more experimental data that was collected, the less variability the analytical methods had to account for.

NASA had proposed a cavity reactor test-bed, which would serve primarily to test the open and closed cycle gas core NTRs also under development at the time, which could theoretically be used to test the RBR as well in a hot-fore configuration due to its unique gas injection system. Sadly, this test-bed never came to be (it was canceled along with most other astronuclear programs), so the faint hope for fission-powered RBR testing in an existing facility died as well.

The Last Gasp for the RBR

The final paper that I was able to find on the Rotating Fluidized Bed Reactor was by Ludewig, Manning, and Raseman of Brookhaven in the Journal of Spacecraft, Vol 11, No 2, in 1974. The work leading up to the Brookhaven program, as well as the Brookhaven program itself, was summarized, and new ideas were thrown out as possibilities as well. It’s evident reading the paper that they still saw the promise in the RBR, and were looking to continue to develop the project under different funding structures.

Other than a brief mention of the possibility of continuous refueling, though, the system largely sits where it was in the middle of 1973, and from what I’ve seen no funding was forthcoming.

While this was undoubtedly a disappointing outcome, as virtually every astronuclear program in history has faced, and the RBR never revived, the concept of a pebblebed NTR would gain new and better-funded interest in the decades to come.

This program, which has its own complex history, will be the subject for our next blog post: Project Timberwind and the Space Nuclear Thermal Propulsion program.

Conclusion

While the RBR was no more, the idea of a pebblebed NTR would live on, as I mentioned above. With a new, physically demanding job, finishing up moving, and the impacts of everything going on in the world right now, I’m not sure exactly when the next blog post is going to come out, but I have already started it, and it should hopefully be coming in relatively short order! After covering Timberwind, we’ll look at MITEE (the whole reason I’m going down this pebblebed rabbit hole, not that the digging hasn’t been fascinating!), before returning to the closed cycle gas core NTR series (which is already over 50 pages long!).

As ever, I’d like to thank my Patrons on Patreon (www.patreon.com/beyondnerva), especially in these incredibly financially difficult times. I definitely would have far more motivation challenges now than I would have without their support! They get early access to blog posts, 3d modeling work that I’m still moving forward on for an eventual YouTube channel, exclusive content, and more. If you’re financially able, consider becoming a Patron!

You can also follow me at https://twitter.com/BeyondNerva for more regular updates!

References

Rotating Fluidized Bed Reactor

Hendrie et al, “ROTATING FLUIDIZED BED REACTOR FOR SPACE NUCLEAR PROPULSION Annual Report: Design Studies and Experimental Results, June, 1970- June, 1971,” Brookhaven NL, August 1971 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19720017961.pdf

Hendrie et al, “ROTATING FLUIDIZED BED REACTOR FOR SPACE NUCLEAR PROPULSION Annual Report: Design Studies and Experimental Results, June 1971 – June 1972,” Brookhaven NL, Sept. 1972 https://inis.iaea.org/collection/NCLCollectionStore/_Public/04/061/4061469.pdf

Hoffman et al, “ROTATING FLUIDIZED BED REACTOR FOR SPACE NUCLEAR PROPULSION Annual Report: Design Studies and Experimental Results, July 1972 – January 1973,” Brookhaven NL, Sept 1973 https://inis.iaea.org/collection/NCLCollectionStore/_Public/05/125/5125213.pdf

Cavity Test Reactor

Whitmarsh, Jr, C. “PRELIMINARY NEUTRONIC ANALYSIS OF A CAVITY TEST REACTOR,” NASA Lewis Research Center 1973 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19730009949.pdf

Whitmarsh, Jr, C. “NUCLEAR CHARACTERISTICS OF A FISSIONING URANIUM PLASMA TEST REACTOR WITH LIGHT -WATER COOLING,” NASA Lewis Research Center 1973 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19730019930.pdf

Categories
Development and Testing History Nuclear Thermal Systems

The Nuclear Lightbulb – A Brief Introduction

Hello, and welcome back to Beyond NERVA! Really quickly, I apologize that I haven’t published more recently. Between moving to a different state, job hunting, and the challenges we’re all facing with the current medical situation worldwide, this post is coming out later than I was hoping. I have been continuing to work in the background, but as you’ll see, this engine isn’t one that’s easy to take in discrete chunks!

Today, we jump into one of the most famous designs of advanced nuclear thermal rocket: the “nuclear lightbulb,” more properly known as the closed cycle gas core nuclear thermal rocket. This will be a multi-part post on not only the basics of the design, but a history of the way the design has changed over time, as well as examining both the tests that were completed as well as the tests that were proposed to move this design forward.

Cutaway of simplified LRC Closed Cycle Gas Core NTR, image credit Winchell Chung of Atomic Rockets

One of the challenges that we saw on the liquid core NTR was that the fission products could be released into the environment. This isn’t really a problem from the pollution side for a space nuclear reactor (we’ll look at the extreme version of this in a couple months with the open cycle gas core), but as a general rule it is advantageous to avoid it most of the time to keep the exhaust mass low (why we use hydrogen in the first place). In ideal circumstances, and with a high enough thrust-to-weight ratio, eliminating this release could even enable an NTR to be used in surface launches.

That’s the potential of the reactor type we’re going to be discussing today, and in the next few posts. Due to the complexities of this reactor design, and how interconnected all the systems are, there may be an additional pause in publication after this post. I’ve been working on the details of this system for over a month and a half now, and am almost done covering the basics of the fuel itself… so if there’s a bit of delay, please be understanding!

The closed cycle gas core uses uranium hexafluoride (UF6) as fuel, which is contained within a fused silica “bulb” to form the fuel element – hence the popular name “nuclear lightbulb”. Several of these are distributed through the reactor’s active zone, with liquid hydrogen coolant flowing through the silica bulb, and then the now-gaseous hydrogen passing around the bulbs and out the nozzle of the reactor. This is the most conservative of the gas core designs, and only a modest step above the vapor core designs we examined last time, but still offers significantly higher temperatures, and potentially higher thrust-to-weight ratios, than the VCNTR.

A combined research effort by NASA’s Lewis (now Glenn) Research Center and United Aircraft Corporation in the 1960s and 70s made significant progress in the design of these reactors, but sadly with the demise of the AEC and NASA efforts in nuclear thermal propulsion, the project languished on the shelves of astronuclear research for decades. While it has seen a resurgence of interest in the last few decades in popular media, most designs for spacecraft that use the lightbulb reactor reference the efforts from the 60s and 70s in their reactor designs- despite this being, in many ways, one of the most easily tested advanced NTR designs available.

Today’s blog post focuses on the general shape of the reactor: its basic geometry, a brief examination of its analysis and testing, and the possible uses of the reactor. The next post will cover the analytical studies of the reactor in more detail, including the limits of what this reactor could provide, and what the tradeoffs in the design would require to make a practical NTR, as well as the practicalities of the fuel element design itself. Finally, in the third we’ll look at the testing that was done, could have been done with in-core fission powered testing, the lessons learned from this testing, and maybe even some possibilities for modern improvements to this well-known, classic design.

With that, let’s take a look at this reactor’s basic shape, how it works, and what the advantages of and problems with the basic idea are.

Nuclear Lightbulb: Nuclear Powered Children’s Toy (ish)

Easy Bake Oven, image Wikimedia

For those of us of a certain age, there was a toy that was quite popular: the Easy-Bake Oven. This was a very simple toy: an oven designed for children with minimal adult supervision to be able to cook a variety of real baked goods, often with premixed dry mixes or simple recipes. Rather than having a more normal resistive heating element as you find in a normal oven, though, a special light bulb was mounted in the oven, and the waste heat from the bulb would heat the oven enough to cook the food.

Closed cycle gas core bulb, image DOE colorized by Winchell Chung

The closed cycle gas core NTR takes this idea, and ramps it up to the edges of what materials limits allow. Rather than a tungsten wire, the heat in the bulb is generated by a critical mass of uranium hexafluoride, a gas at room temperature that’s used in, among other things, fissile fuel enrichment for reactors and other applications. This is contained in a fused silica bulb made up of dozens of very thin tubes – not much different in material, but very different in design, compared to the Easy-Bake Oven – which contains the fissile fuel, and prevents the fission products from escaping. The fuel turns from gas to plasma, and forms a vortex in the center of the fuel element.

Axial cross-section of the fuel/buffer/wall region of the lightbulb, Rodgers 1972

To further protect the bulb from direct contact with the uranium and free fluorine, a gaseous barrier of noble gas (either argon or neon) is injected between the fuel and the wall of the bulb itself. Because of the extreme temperatures, the majority of the electromagnetic radiation coming off the fuel isn’t in the form of infrared (heat), but rather as ultraviolet radiation, which the silica is transparent to, minimizing the amount of energy that’s deposited into the bulb itself. In order to further protect the silica bulb, microparticles of the same silica are added to the neon flow to absorb some of the radiation the bulb isn’t transparent to, in order to remove that part of the radiation before it hits the bulb. This neon passes around the walls of the chamber, creating a vortex in the uranium which further constrains it, and then passes out of one or both ends of the bulb. It then goes through a purification and cooling process using a cryogenic hydrogen heat exchanger and gas centrifuge, before being reused.

Now, of course there is still an intense amount of energy generated in the fuel which will be deposited in the silica, and will attempt to melt the bulb almost instantly, so the bulb must be cooled regeneratively. This is done by liquid hydrogen, which is also mostly transparent to the majority of the radiation coming off the fuel plasma, minimizing the amount of energy the coolant absorbs from anything but the silica of the bulb itself.

Finally, the now-gaseous hydrogen from both the neon and bulb cooling processes, mixed with any hydrogen needed to cool the pressure vessel, reflectors of the reactor, and other components, is mixed with microparticles of tungsten to increase the amount of UV radiation emitted by the fuel. This then passes around the bulbs in the reactor, getting heated to their final temperature, before exiting the nozzle of the NTR.

Overall configuration, Rodgers 1972

The most commonly examined version of the lightbulb uses a total of seven bulbs, with those bulbs being made up of a spiral of hydrogen coolant channels in fused silica. This was pioneered by NASA’s Lewis Research Center (LRC), and studied by United Aircraft Corp of Mass (UA). These studies were carried out between 1963 and 1972, with a very small number of follow-up studies at UA completing by 1980. This design was a 4600 MWt reactor fueled by 233U, an isp of 1870 seconds, and a thrust-to-weight ratio of 1.3.

A smaller version of this system, using a single bulb rather than seven, was proposed by the same team for probe missions and the like, but unfortunately the only papers are behind paywalls.

During the re-examination of nuclear thermal technology in the early 1990s by NASA and the DOE, the design was re-examined briefly to assess the advantages that the design could offer, but no advances in the design were made at the time.

Since then, while interest in this concept has grown, new studies have not been done, and the design remains dormant despite the extensive amount of study which has been carried out.

What’s Been Done Before: Previous Studies on the Lightbulb

Bussard 1958

The first version of the closed cycle gas core proposed by Robert Bussard in 1946. This design looked remarkably like an internal combustion firing chamber, with the UF6 gas being mechanically compressed into a critical density with a piston. Coolant would be run across the outside of the fuel element and then exit the reactor through a nozzle. While this design hasn’t been explored in any depth that I’ve been able to determine, a new version using pressure waves rather than mechanical pistons to compress gas into a critical mass has been explored in recent years (we’ll cover that in the open cycle gas core posts).

Starting in 1963, United Aircraft (UA, a subsidiary of United Technologies) worked with NASA’s Lewis Research Center (LRC) and Los Alamos Scientific Laboratory (LASL) on both the open and closed cycle gas core concepts, but the difficulties of containing the fuel in the open cycle concept caused the company to focus exclusively on the closed cycle concepts. Interestingly, according to Tom Latham of UA (who worked on the program), the design was limited in both mass and volume by the then-current volume of the proposed Space Shuttle cargo bay. Another limitation of the original concept was that no external radiators could be used for thermal management, due to the increased mass of the closed radiator system and its associated hardware.

System flow diagram, Rodgers 1972

The design that evolved was quite detailed, and also quite efficient in many ways. However, the sheer number of interdependent subsystems makes is fairly heavy, limiting its potential usefulness and increasing its complexity.

In order to get there, a large number of studies were done on a number of different subsystems and physical behaviors, and due to the extreme nature of the system design itself many experimental apparatus had to be not only built, but redesigned multiple times to get the results needed to design this reactor.

We’ll look at the testing history more in depth in a future blog post, but it’s worth looking at the types of tests that were conducted to get an idea of just how far along this design was:

RF Heating Test Apparatus, Roman 1969

Both direct current and radio frequency testing of simulated fuel plasmas were conducted, starting with the RF (induction heating) testing at the UA facility in East Hartford, CT. These studies typically used tungsten in place of uranium (a common practice, even still used today) since it’s both massive and also has somewhat similar physical properties to uranium. At the time, argon was considered for the buffer gas rather than neon, this change in composition will be something we’ll look at later in the detailed testing post.

Induction heating works by using a vibrating magnetic field to heat materials that will flip their molecular direction or vibrate, generating heat. It is a good option for nuclear testing since it is able to more evenly heat the simulated fuel, and can achieve high temperatures – it’s still used for nuclear fuel element testing not only in the Compact Fuel Element Environment Test (CFEET) test stand, which I’ve covered here https://beyondnerva.com/nuclear-test-stands-and-equipment/non-nuclear-thermal-testing/cfeet-compact-fuel-element-environmental-test/ , but also in the Nuclear Thermal Rocket Environmental Effects Simulator, which I covered here: https://beyondnerva.com/nuclear-test-stands-and-equipment/non-nuclear-thermal-testing/ntrees/ . One of the challenges of this sort of heating, though, is the induction coil, the device that creates the heating in the material. In early testing they managed to melt the copper coil they were using due to resistive heating (the same method used to make heat in a space heater or oven), and constructing a higher-powered apparatus wasn’t possible for the team.

This led to direct current heating testing to achieve higher temperatures, which uses an electrical arc through the tungsten plasma. This isn’t as good at simulating the way that heat is distributed in the plasma body, but could achieve higher temperatures. This was important for testing the stability of the vortex generated by not only the internal heating of the fuel, but also the interactions between the fuel and the neon containment system.

Spectral flux from the edge of the fuel body, Rodgers 1972 (will be covered more in depth in another post)

Another concern was determining what frequencies of radiation silicon, aluminum and neon were transparent to. By varying the temperature of the fissioning fuel mass, the frequency of radiation could, to a certain degree, be tuned to a frequency that maximized how much energy would pass through both the noble gas (then argon) and the bulb structure itself. Again, at the time (and to a certain extent later), the bulb configuration was slightly different: a layer of aluminum was added to the inner surface of the bulb to reflect more thermal radiation back into the fissioning fuel in order to increase heating, and therefore increase the temperature of the fuel. We’ll look at how this design option changed over time in future posts.

More studies and tests were done looking at the effects of neutron and gamma radiation on reactor materials. These are significant challenges in any reactor, but the materials being used in the lightbulb reactor are unusual, even by the standards of astronuclear engineering, so detailed studies of the effects of these radiation types were needed to ensure that the reactor would be able to operate throughout its required lifetime.

Fused silica test article, Vogt 1970

Perhaps one of the biggest concerns was verifying that the bulb itself would maintain both its integrity and its functionality throughout the life of the reactor. Silica is a material that is highly unusual in a nuclear reactor, and the fact that it needed to remain not only transparent but able to contain both a noble gas seeded with silica particles and hydrogen while remaining transparent to a useful range of radiation while being bombarded with neutrons (which would change the crystalline structure) and gamma rays (which would change the energy states of the individual nuclei to varying degrees) was a major focus of the program. On top of that, the walls of the individual tubes that made up the bulbs needed to be incredibly thin, and the shape of each of the individual tubes was quite unusual, so there were significant experimental manufacturing considerations to deal with. Neutron, gamma and beta (high energy electron) radiation could all have their effect on the bulb itself during the course of the reactor’s lifetime, and these effects needed to be understood and accounted for. While these tests were mostly successful, with some interesting materials properties of silica discovered along the way, when Dr. Latham discussed this project 20 years later, one of the things he mentioned was that modern materials science could possibly offer better alternatives to the silica tubing – a concept that we will touch on again in a future post.

Another challenge of the design was that it required seeding two different materials into two different gasses: the neon/argon had to be seeded with silica in order to protect the bulb, and the hydrogen propellant needed to be seeded with tungsten to make it absorb the radiation passing through the bulb as efficiently as possible while minimizing the increase in the mass of the propellant. While the hydrogen seeding process was being studied for other reactor designs – we saw this in the radiator liquid fueled NTR, and will see it again in the future in open cycle gas core and some solid core designs we haven’t covered yet – the silica seeding was a new challenge, especially because the material being seeded and the material the seeded gas would travel through was the same as the material that was seeded into the gas.

Image DOE via Chris Casilli on Twitter

Finally, there’s the challenge of nuclear testing. Los Alamos Scientific Laboratory conducted some tests that were fission-powered, which proved the concept in theory, but these were low powered bench-top tests (which we’ll cover in depth in the future). To really test the design, it would be ideal to do a hot-fire test of an NTR. Fortunately, at the time the Nuclear Furnace test-bed was being completed (more on NERVA hot fire testing here: https://beyondnerva.com/2018/06/18/ntr-hot-fire-testing-part-i-rover-and-nerva-testing/ and the exhaust scrubbers for the Nuclear furnace here: https://beyondnerva.com/nuclear-test-stands-and-equipment/nuclear-furnace-exhaust-scrubbers/ ). This meant that it was possible to use this versatile test-bed to test a single, sub-scale lightbulb in a controlled, well-understood system. While this test was never actually conducted, much of the preparatory design work for the test was completed, another thing we’ll cover in a future post.

A Promising, Developed, Unrealized Option

The closed cycle gas core nuclear thermal rocket is one of the most perrenially fascinating concepts in astronuclear history. Not only does it offer an option for a high-temperature nuclear reactor which is able to avoid many of the challenges of solid fuel, but it offers better fission product containment than any other design besides the vapor core NTR.

It is also one of the most complex systems that has ever been proposed, with two different types of closed cycle gas systems involving heat exchangers and separation systems supporting seven different fuel chambers, a host of novel materials in unique environments, the need to tune both the temperature and emissivity of a complex fuel form to ensure the reactor’s components won’t melt down, and the constant concerns of mass and complexity hanging over the heads of the designers.

Most of these challenges were addressed in the 1960s and 1970s, with most of the still-unanswered questions needing testing that simply wasn’t possible at the time of the project’s cancellation due to shifting priorities in the space program. Modern materials science may offer better solutions to those that were available at the time as well, both in the testing and operation of this reactor.

Sadly, updating this design has not happened, but the original design remains one of the most iconic designs in astronuclear engineering.

In the next two posts, we’ll look at the testing done for the reactor in detail, followed by a detailed look at the reactor itself. Make sure to keep an eye out for them!

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References

McLafferty, G.H. “Investigation of Gaseous Nuclear Rocket Technology – Summary Technical Report” 1969 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19700008165.pdf

Rodgers, R.J. and Latham, T.S. “Analytical Design and Performance Studies of the Nuclear Light Bulb Engine” 1972 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19730003969.pdf

Latham, T.S. “Nuclear Light Bulb,” 1992 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19920001892.pdf

Categories
History Nuclear Thermal Systems

Radiator LNTR: The Last of the Line

Hello, and welcome back to Beyond NERVA! Today, we’re finishing (for now) our in-depth look at liquid fueled nuclear thermal rockets, by looking at the second major type of liquid NTR (LNTR): the radiator-type LNTR. If you’re just joining us, make sure to check out the introduction (available here) and the bubbler post (available here) for some important context to understand how this design got here.

Rather than passing the propellant directly through the molten fuel, in this system the propellant would pass through the central void of the fuel element, becoming heated primarily through radiation (although some convection within the propellant flow would occur, overall it was a minor effect), hence the name.

This concept had been mentioned in previous works on bubbler-type LNTRs, and initial studies on the thermalization behavior of the propellant, and conversely fuel cooling behavior, were conducted during the early 1960s, but the first major study wouldn’t occur until 1966. However, it would also extend into the 1990s in its development, meaning that it was a far longer-lived design.

Let’s begin by looking at the differences between the bubbler and radiator designs, and why the radiator offers an attractive trade-off compared to the bubbler.

The Vapor Problem, or Is Homogenization of Propellant/Fuel Temp Worth It?

Liquid fuels offer many advantages for an NTR, including the fact that the fuel will distribute its heat evenly across the volume of the fuel element, the fact that the effective temperature of the fuel can be quite high, and that the fuel is able to be reasonably well contained with minimal theoretical challenges.

The bubbler design had an additional advantage: by passing the propellant directly through the fuel, in discrete enough bundles (the bubbles themselves) that the fuel and the propellant would have the same temperature.

Maximum specific impulse due to vapor pressure, Barrett Jr.

Sadly, there are significant challenges to making this sort of nuclear reactor into a rocket, the biggest one being propellant mass. These types of NTRs still use hydrogen propellant, the problem occurs in the fuel mass itself. As the bubbles move through the zirconium/niobium-uranium carbide fuel, it heats up rapidly, and the pressure drops significantly during this process. This means that all of the components of the fuel (the Zr/Nb, C, and U) end up vaporizing into the bubbles, to the point that the bubble is completely saturated by a mix of these elements in vapor form by the time it exits the fuel body. This is called vapor entrainment.

This is a major problem, because it means that the propellant leaving the nozzle has a far higher mass than the hydrogen that was originally input into the system. While there’s the possibility that a different propellant could be used which would not entrain as much of the fuel mass, but would also be higher molecular mass to start – to the point that the gains might likely outweigh the losses (if you feel like exploring this trade-off on a more technical footing, please let me know! I’d love to explore this more), and it wouldn’t eliminate the entrainment problem.

This led people to wonder if you have to pass the propellant through the fuel in the first place. After all, while there is a thermodynamically appealing symmetry to homogenizing your fuel and propellant temperatures, this isn’t actually necessary, is it? The fuel elements are already annular in shape, after all, so why not use them as a more traditional fuel element for an NTR? The lower surface area would mean that there’s less chance for the inconveniently high vapor pressure of the fuel would be mitigated by the fact that the majority of the propellant wouldn’t come in contact with the fuel (or even the layer of propellant that does interact with the fuel), meaning that the overall propellant molecular mass would be kept low… right?

The problem is that this means that the only method of heating the propellant becomes radiation (there’s a small amount of convection, but it’s so negligible that it can be ignored)… which isn’t that great in hydrogen, especially in the UV spectrum that most of the photons from the nuclear reaction are emitted in. The possibility of using either microparticles or vapors which would absorb the UV and re-emit it in a lower wavelength, which would be more easily absorbed by the hydrogen, was already being investigated in relation to gas core NTRs (which have the same problem, but in a completely different order of magnitude), and offered promise, but was also a compromise: this deliberately increases the molar mass of the propellant one way to minimize the molar mass a different way. This was a design possibility that needed to be carefully studied before it could be considered more feasible than the bubbler LNTR.

The leader of the effort to study this trade-off was one of the best-known fluid fueled NTR designers on the NASA side: Robert Ragsdale at Lewis Research Center (LRC, and we’ll come back to Ragsdale in gas core NTR design as well). There were a collection of studies around a particular design, beginning with a study looking at reactor geometry and fuel element size optimization to not only maximize the thrust and specific impulse, but also to minimize the uranium loss rates of the reactor.

This study concluded that there were many advantages to the radiator-type LNTR over the bubbler-type. The first consideration, minimizing the vapor entrainment problem that was laguing the bubbler, was minimized, but not completely eliminated, in the radiator design. The next conclusion is that the specific impulse of the negine could be maintained, or increased, to 1400 s isp or more. Finally, one of th emost striking was in thrust-to-core-weight ratio, which went from about 1:1 in the Nelson/Princeton design that we discussed in the last post all the way up to 19:1 (potentially)! This is because the propellant flow rate isn’t limited y the bubble velocity moving through the fuel (for more detail on this problem, and the other related constraints, check out the last blog post, here).

These conclusions led to NASA gathering a team of researchers, including Ragsdale, Kasack, Donovan, Putre, and others ti develop the Lewis LNTR reactor.

Lewis LNTR: The First of the Line

Lweis Radiator LNTR, Ragsdale 1967

Once the basic feasibility of the radiator LNTR was demonstrated, a number of studies were conducted to determine the performance characteristics, as well as the basic engineering challenges, facing this type of NTR. They were conducted in 1967/68, and showed distinct promise, for the desired 2000 to 5000 MWt power range (similar to the Phoebus 2 reactor’s power goal, which remains the most powerful nuclear reactor ever tested at 3500 MWt).

Fuel tube cross-section, Putre 1968

As with any design, the first question was the basics of reactor configuration. The LRC team never looked at a single-tube LNTR, for a variety of reasons, and instead focused their efforts on a multi-tube design, but the number and diameter of the tubes was one of the major questions to be addressed in initial studies. Because of this, and the particular characteristics of the heat transfer required, the reactor would have many fuel elements with a diameter of between 1 and 4 inches, but which diameter was best would be a matter of quite some study.

Another question for the study team was what the fuel element temperature would be. As in every NTR design, the hotter the propellant, the higher the isp (all other things being equal), but as we saw in the bubbler design, higher temperatures also mean higher vapor pressure, meaning that mass is lost more easily into the propellant – which increases the propellant mass and reduces the isp, and at some point even cost more specific impulse due to higher mass than is gained with the higher temperature. Because the propellant and the fuel would only interact on the surface of the fuel element, the surface temperature of the fuel was the overriding consideration, and was also explored, in the range of 5000 to 6100 K.

Effect of Reactor Pressure on T/W Ratio and U mass loss ratio in H, Ragsdale 1967

The final consideration which was optimized in this design was engine operating pressures. Because this design wasn’t fundamentally limited by the bubble velocity and void fraction of the molten fuel, the chamber pressure could be increased significantly, leading to both more thrust and a higher thrust-to-weight ratio. However, the trade-off here is that at some point the propellant isn’t being completely thermalized, resulting in a lower specific impulse. This final consideration was explored in the range of 200 to 1000 atm (2020-10100 N/cm2).

The three primary goals were: to maximize specific impulse, maximize thrust-to-weight ratio, and minimize uranium mass loss. They quickly discovered that they couldn’t have their cake and eat it, too: higher temperatures, and therefore higher isp, led to faster U mass loss rates, increasing T/W ratio reduced the specific impulse, and minimizing the U loss rate hurt both T/W and isp. They could improve any one (or often two) of these characteristics, but always at the cost of the third characteristic.

Four potential LNTR configurations, note the tradeoffs between isp, T/W, and fuel loss rates. Ragsdale 1967

We’ll look at many of the design characteristics and engineering considerations of the LRC work in the next section on general design challenges and considerations for the radiator LNTR, but for now we’ll look at their final compromise reactor.

The reactor itself would be made up of several (oddly, never specified) fuel elements, in a beryllium structure, with each fuel element being made up of Be as well. These would be cooled by cryogenic hydrogen moving from the nozzle end to the spacecraft end of the reactor, before flowing back into the central void of the fuel element. As it was fed through the central annulus, it would be seeded with tungsten microparticles to increase the amount of heat the propellant would absorb. Finally, it would be exhausted through a standard De Laval nozzle to provide thrust.

Reference LRC LNTR design characteristics, Putre 1968

The final fuel that they settled on was a liquid ternary carbide design, with the majority of the fuel being niobium carbide (although ZrC was also considered), with a molar mass fraction of 0.02 being UC2. This compromise offered good power density for the reactor while minimizing the vaporization rate of the fuel mass. This would be held in 2 inch diameter, 5 foot long fuel element tubes, with a fuel surface temperature of 5060 K. The propellant would be pressurized to 200 atm in the reactor.

Final LRC LNTR Fuel Characteristics, Putre 1968

This led to a design that struck a compromise between isp, T/W, and U mass loss which was not only acceptable, but impressive: 1400 s isp (on par with some forms of electric propulsion), a T/W ratio (of the core alone) of 4, and a hydrogen-to-uranium flow rate ratio of 50.

They did observe that none of these characteristics were as high as they could be, especially in terms of T/W ratio (which they calculated could go as high as 19!), or isp (with a theoretical maximum of 1660), and the uranium loss was twice the theoretical minimum, but sadly the cost of maximizing any of these characteristics was so high from an engineering point of view that it wasn’t feasible.

Sadly, I haven’t been able to find any documentation on this reactor design – and very few references to it – after February 1968. The exact time of the cancellation, and the reasons why, are a mystery to me. If someone is able to help me find that information it would be greatly appreciated.

LARS: The Brookhaven Design

LARS cross section,

The radiator LNTR would remain dormant for decades, as astronuclear funding was scarce and focused on particular, well-characterized systems (most of which were electric powerplant concepts), until the start of the Space Exploration Initiative. In 1991, a conference was held to explore the use of various types of NTR in future crewed space missions. This led to many proposals, including one from the Department of Energy’s Brookhaven National Laboratory in New York. This was the Liquid Annular Reactor System, or LARS.

A team of physicists and engineers, including Powell, Ludewig, Lazareth, and Maise decided to revisit the radiator LNTR design, but as far as I can tell didn’t use any of the research done by the LRC team. Due to the different design philosophies, lack of references, and also the general compartmentalization of knowledge within the different parts of the astronuclear community, I can only conclude that they began this design from scratch (if this is incorrect, and anyone has knowledge of this program, please get in contact with me!).

LARS was a very different design than the LRC concept, and seems to have gone through two distinct iterations. Rather than the high-pressure system that the LRC team investigated, this was a low-pressure, low-thrust design, which optimized a different characteristic: hydrogen dissociation. This maximizes the specific impulse of the NTR by reducing the mass of the propellant to the lowest theoretically possible mass while maintaining the temperature of the propellant (up to 1600 s, according to the BNL team). The other main distinction from the LRC design was the power level: rather than having a very powerful reactor (3000 to 5000 MWt), this was a modest reactor of only 200 MWt. This leads to a very different set of design tradeoffs, but many of the engineering and materials challenges remain the same.

LARS would continue to us NbC diluted with UC2, but the fuel would not completely melt in the fuel element, leaving a solid layer against the walls of the beryllium fuel element tube. This in turn would be regeneratively cooled with hydrogen flowing through a number of channels in the durm, as well as a gap surrounding the body of the fuel element which would also be filled with cold hydrogen. A drive system would be attached on the cold end of the tube to spin it at an appropriate rate (which was not detailed in the papers). The main changes were in the fuel element configuration, size, and number.

The first iteration of LARS was an interesting concept, using a folded-flow system. This used many small fuel element tubes, arranged in a similar manner to the flow channels in the Dumbo reactor, with the propellant moving from the center of the reactor to the outer circumference, before being ejected out of the nozzle of the reactor. Each layer of fuel elements contained eleven individual tubes, with between 1 and 10 layers of fuel elements in the reactor. As the number of layers increased, the length and radius of the fuel elements decreased.

One of the important notes that was made by the team at this early design date was that the perpendicular fuel element orientation would minimize the amount of fission products that would be ejected from the rocket. I’m unable to determine how this was done, other than if they were solids which would stick to the outside of the propellant flue, however.

Unfortunately, I haven’t been able to discover exactly why this design was abandoned for a more traditional LNTR architecture, but the need to cool the entire exterior of the reactor to keep it from melting seems to possibly be a concern. Reversing the flow, with the hot propellant being in the center of the reactor rather than the external circumference, seems like an easy fix if this was the primary concern, but the discussions of reactor architecture after this seem to pretty much ignore this early iteration. Another complication would be the complexity of the reactor architecture. Whether with dedicated motors, or with a geared system allowing one motor to spin multiple fuel elements, a complex system is needed to spin the fuel elements, which would not only be something which would potentially break down more, but also require far more mass than a simpler system.

The second version of LARS kept the same type of fuel, power output, and low pressure operation, but rather than using the folded flow concept it went with seven fuel elements in a beryllium body. The propellant would be used to cool first the nozzle of the rocket, then the rotating beryllium drum which contained the fuel element, before entering the main propellant channel. The final thermalization of the propellant would be facilitated by the use of tungsten microparticles in the H2, necessary due to the low partial pressure and high transparency of pure H2 (while the vapor pressure issues of any LNTR were acknowledged, the effect that this would have on the thermalization seems to have not been considered a significant factor in the seeding necessity) Two versions, defined by the emissivity of the fuel element, were proposed.

Final two LARS options, f is fuel emissivity, Maise 1999

This design was targeted to reach up to 2000 s isp, but due to uncertainties in U loss rates (as well as C and Nb), the overall mass of the propellant upon exiting the reactor was uncertain, so the authors used a range of 1600-2000 s. The thrust of the engine was approximately 20,000 N, which would result in a T/W ratio of about 1;1 when including a shadow shield (one author points out that without the shield the ratio would be about 3-4/1.

I have been unable to find the research reports themselves for this program (unlike the LRC design), so the specifics of the reactor physics tradeoffs, engineering compromises, actual years of research and the like aren’t something that I’m able to discuss. The majority of my sources are conference papers and journal articles, which occurred in 1991 and 1992, but there was one paper from 1999, so it was at least under discussion through the 1990s (interestingly, that paper discussed using LARS for the 550 AU mission concept, which later got remade into the FOCAL gravitational lens mission: https://www.centauri-dreams.org/2010/11/15/a-focal-mission-into-the-oort-cloud/ ).

This seems to be the last time that LARS has been mentioned in the technical literature, so while it is mentioned as the “baseline” liquid core concept in places such as Atomic Rockets (http://www.projectrho.com/public_html/rocket/enginelist2.php#id–Nuclear_Thermal–Liquid_Core–LARS) it has not been explored in depth since.

Lessons Learned, Lessons to Learn: The Challenges of LNTR

In many ways, the apparent dual genesis of radiator LNTRs offers the ability to look into two particular thought processes in what the challenges are with radiator-type LNTRs. One example of this is what’s discussed more in the “fundamental challenges” sections of the introductory section of the reports: for the LRC team they focus on vapor entrainment minimization, whereas in the BNL presentations it seems quite important to point out that “yes, containing a refractory in a spinning, gas cooled drum is relatively trivial.” This juxtaposition of foci is interesting to me, as an examination of the different engineering philosophies of the two teams, and the priorities of the times.

Wall Construction

Both the LRC and LARS LNTRs ended up with similar fuel element configurations: a high temperature material, with coolant tubes traveling the length of the fuel element walls to regeneratively cool the walls. This material would have to withstand not only the temperature of the fuel element, but also resist chemical attack by the hydrogen used for the regenerative cooling, as well as being able to withstand the mechanical strain of not only the spinning fuel, but also the torque from whatever drive system is used to spin the fuel element to maintain the centripetal force used to contain the fuel element.

Another constant concern is the temperature of the wall. While high temperature loadings can be handled using regenerative cooling, the more heat is removed from the fuel during the regenerative cooling step, but it reduces the specific impulse of the engine. Here’s a table from the LRC study that examines the implications of wall cooling ratio vs specific impulse in that design, which will also apply as a general rule of thumb for LARS:

However, from there, the two designs differed significantly. The LARS design is far simpler: a can of beryllium, with a total of 20% of the volume being the regenerative cooling channel. As mentioned previously, the fuel didn’t become completely molten, but remained solid (and mostly containing the ZrC/NbC component, with very little U). Surrounding the outside of the fuel element can itself was another coolant gap. This would then be mounted to the reactor body with a drive system at the ship end, and a bearing at the hot end. This would then be mounted in the stationary moderator which made up the majority of the internal volume of the reactor core, which was shielded from the heat in the fuel element in a very heterogeneous temperature profile.

The LRC concept on the other hand, was far more complex in some ways. Rather than using a metal can, the LRC design used graphite, which maintains its strength far better than many metals at high temperatures. A number of options were considered to maintain the wall of the can, due not only to the fuel mixture potentially attacking the graphite (as the carbon could be dissolved into the carbide of the fuel element), as well as attack from the hydrogen in the coolant channels (which would be able to be addressed in a similar way to how NERVA fuel elements used refractory metal coatings to prevent the same erosive effects).

The LRC design, since the fuel would be completely molten across the entire volume of the fuel element, was a more complex challenge. A number of options were considered to minimize the wall heating of the fuel element, including:

  • Selective fuel loading
    • A common strategy in solid fuel elements, this creates hotter and cooler zones in the fuel element
      • Neutron heating will distribute the radiative heating past U distribution
    • Convection and fuel mixing will end up distributing the fuel over time
    • May be able to be limited by affecting the temperature and viscosity of the fuel for the life of the reactor
  • Multiple fluids in fuel
    • Step beyond selective loading, a different material may be used as the outer layer of the fuel body, resisting mixing and reducing thermal load on the wall
  • Vapor insulation along exterior of fuel body
    • Using thermally opaque vapor to insulate the fuel element wall from the fuel body
    • Significantly reduces the heating on the outer wall
    • Two options for maintaining vapor wall:
      • Ablative coating on inner wall of fuel element can
      • Porous wall in can (similar to a low-flow version of a bubbler fuel element) pumping vapor into gap between fuel and can
    • Maximum stable vapor-layer thickness based on vapor bubble force balance vs centripetal force of liquid fuel
      • Two phase flow dynamics needed to maintain the vapor layer would be complex

This set of options offer a trade-off: either a simpler option, which sets hard limits on the fuel element temperature in order to ensure the phase gradient in the fuel element (the LARS concept), or the fully liquid, more complex-behaving LRC design which has better power distribution, and a higher theoretical fuel element temperature – only limited by the vapor pressure increase and fuel loss rates in the fuel element, rather than the wall heating temperature limits of the LARS design.

Anyone designing a new radiator LNTR has much work that they can draw from, but other than the dynamics of the actual fuel behavior (which have never gone through a criticality test), the fuel element can design will be perhaps the largest set of engineering challenges in this type of system (although simpler than the bubbler-type LNTR).

Propellant Thermalization

The major change between the bubbler and radiator-type LNTRs is the difference in the thermalization behavior of the propellant: in a bubbler-type LNTR, assuming the propellant can be fed through the fuel, the two components reach thermal equilibrium, so the only thing needed is to direct it out of the nozzle; a radiator on the other hand has a similar flow path to the Rover-type NTRs, once through from nozzle to ship side for regenerative cooling, then a final thermalization pass through the central void of the fuel element.

This is a problem for hydrogen propellant, which is largely transparent to the EM radiation coming off the reactor. This thermal transfer accounted for all but 10% of the thermalization effects in the LARS design, and in many of the LRC studies this was completely ignored as negligible, with the convective effects in the propellant mainly being a concern in terms of fuel mass loss and propellant mass increase.

While the fuel mass loss would increase the opacity of the gas (making it absorb more heat), a far better option was available: adding a material in microparticle form to the propellant flow as it goes through the final thermalization cycle. The preferred material for the vast majority of these applications, which we’ll see in the gas cycle NTRs as well, is microparticles of tungsten.

This has been studied in a host of different applications, and will be something that I’ll discuss in depth on a section of the propellant webpage in the future (which I’ll link to here once it’s done), but for the LRC design the target goal for increasing the opacity of the H2 was to achieve between 10,000 and 20,000 cm^2/gm, for a reduction in single-digit percentage of specific impulse due to the higher mass. They pointed out that the simplified calculations used for the fuel mass loss behavior could lead to an error that they were unable to properly address, and which could either increase or decrease the amount of additive used.

The LARS concept used tungsten microparticles as well, and their absorption actually was the defining factor in the two designs they proposed: the emissivity and reflectivity of the fuel in terms of the absorption of the wall and the propellant.

Two other options are available for increasing the opacity of the hydrogen gas.

The first is to use a metal vapor deliberately, as was the paradigm in Soviet gas core design. Here, they used either NaK or Li vapor, both of which have small neutron absorption cross-sections and high thermal capacity. This has the advantage of being more easily mixed with the turbulent propellant stream, as well as being far lower mass than the W that is often used in US designs, but may be less opaque to the EM frequencies being emitted by the fuel’s surface in an LNTR design. I’m still trying to track down a more thorough writeup of the use of these vapors in NTR design at the moment (a common problem in both Soviet and Russian astronuclear literature is a lack of translations), but when I do I’ll discuss it in far more depth, since it’s an idea that doesn’t seem to have translated into the American NTR design paradigm.

As I said, this is a concept that I’m going to cover more in depth in both the gas core and general propellant pages, so with one final – and fascinating – note, we’ll move on to the conclusion.

An Interesting Proposal

The final option is something that Cavan Stone mentioned to me on Facebook a while ago: the use of lithium deuteride (LiD) as a propellant or additive in this design. This is an interesting concept, since Li7 is a fissile material, and is reasonably opaque to the frequencies being discussed in these reactors. The use of deuterium rather than protium also increases the neutron moderation of the propellant, which can in turn increase fissile efficiency of the reactor. The Li will harden the neutron spectrum overall, while the D and Be (in the fuel element can/reactor body) will thermalize the spectrum.

There was a discussion of using LiD as a propellant in NTRs in the 1960s [https://www.osti.gov/biblio/4764043-nuclear-effect-using-lithium-hydride-propellant-nuclear-rocket-reactor-thesis], but sadly I can’t find it anywhere online. If someone is able to help me find it, please let me know. This is a fascinating concept, and one that I’m very glad Cavan brought up to me, but also one that is complex enough that I really need to see an in-depth study by someone far more knowledgeable than me to be able to intelligently discuss the implications of.

Conclusion, or The Future of the Forgotten Reactor

While often referenced in passing in any general presentation on nuclear thermal rockets, the liquid core NTR seems to be the least studied of the different NTR types, and also the least covered. While the bubbler offers distinct advantages from a purely thermodynamic point of view, the radiator offers far more promise from a functional perspective.

Sadly, while both solid and gas core NTRs have been studied into the 20th century, the liquid core has been largely forgotten, and the radiator in particular seems to have gone through a reinvention of the wheel, as it were, between the 1960s NASA design and the 1990s DOE design, with few of the lessons learned from the LRC concept being applied to the BNL design as far as vapor dynamics, thermal transfer, and the like.

This doesn’t mean that the design is without promise, though, or that the challenges that the reactor faces are insurmountable. A number of hurdles in testing need to be overcome for this design to work – but many of the problems that there simply isn’t any data for can be solved with a simple set of criticality and reactor physics tests, something well within the capabilities of most research nuclear programs with the capability to test NTRs.

With the advances in nuclear and two-phase flow modeling, a body of research that doesn’t seem to have been examined in depth for over two decades, and the possibility of a high-isp, moderate-to-high thrust engine without the complications of a gas core NTR (a subject that we’ll be covering soon), the LNTR, and the radiator in particular, offer a combination of promise and ability to develop advanced NTRs as low hanging fruit that few systems are able to offer.

Final Note

With that, we’re leaving the realm of liquid fueled NTRs for now. This is a fascinating field, and one that I haven’t seen much discussion of outside the original technical papers, so I hope you enjoyed it! I’m going to work on getting these posts into a more easily-referenced form on the website proper, and will make a note of that in my blog (and on my social media) when I do! If anyone is aware of any additional references pertaining to the LNTR, as well as its thermophysical behavior, fuel materials options, or anything else relating to these desgins, please let me know, either in the comments or by sending me a message to beyondnerva at gmail dot com.

Our next blog post will be on droplet and vapor core NTRs, and will be covered by a good friend of mine and fellow astronuclear enthusiast: Calixto Lopez. These reactors have fascinated him since he was in school many moons ago, and he’s taught me the majority of what I know about them, so I asked him if he was willing to write that post.

After that, we’re going to move on to the closed cycle gas core NTR, which I’ve already begun research on. There’s lots of fascinating tidbits about this reactor type that I’ve already uncovered, so this may end up being another multiple part blog series.

Finally, to wrap up our discussion of advanced NTRs, we’re going to do a series on the open cycle gas core NTR types. This is going to be a long, complex series on not only the basic physics challenges, but the design evolution of the engine type, as well as discussion on various engineering methods to mitigate the major fuel loss and energy waste issues involved in this type of engine. There may be a delay between the closed and open cycle NTR posts due to the sheer amount of research necessary to do open cycles justice, but rest assured I’m already doing research on them.

As you can guess, this blog takes a lot of time, and a lot of research, to write. If you would like to support me in my efforts to bring the wide and complex history of astronuclear engineering to light, consider supporting me on Patreon: https://www.patreon.com/beyondnerva . Every dollar helps, and you get access to not only early releases of every blog post and webpage, but at the higher donation amounts you also get access to the various 3d models that I’m working on, videos, and eventually the completed 3d models themselves for your own projects (with credit for the model construction, of course!).

I’m also always looking for new or forgotten research in astronuclear engineering, especially that done by the Soviet Union, Russia, China, India, and European countries. If you run across anything interesting, please consider sending it to beyondnerva at gmail dot com.

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References

General References

ANALYSES OF VAPORIZATION IN LIQUID URANIUM BEARING SYSTEMS AT VERY HIGH TEMPERATURES, Kaufman and Peters 1965 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19660002967.pdf

ANALYSIS OF VAPORIZATION OF LIQUID URANIUM, METAL, AND CARBON SYSTEMS AT 9000” AND 10,000” R Kaufman and Peters 1966 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19660025363.pdf

Fundamental Material Limitations in Heat-Exchanger Nuclear Rockets, Kane and Wells, Jr. 1965 https://www.osti.gov/servlets/purl/4610034/

VAPOR-PRESSURE DATA EXTRAPOLATED TO 1000 ATMOSPHERES (1.01~108 N/m2) FOR 13 REFRACTORY MATERIALS WITH LOW THERMAL ABSORPTION CROSS SECTIONS Masser 1967 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19670030361.pdf

Radiator-Specific LNTR References

Lewis Research Center LNTR

PERFORMANCE POTENTIAL OF A RADIANT-HEAT-TRANSFER LIQUID-CORE NUCLEAR ROCKET ENGINE, Ragsdale 1967 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19670030774.pdf

HEAT- AND MASS-TRANSFER CHARACTERISTICS OF AN AXIAL-FLOW LIQUID-CORE NUCLEAR ROCKET EMPLOYING RADIATION HEAT TRANSFER, Ragsdale et al 1967 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19670024548.pdf

FEASIBILITY OF SUPPORTING LIQUID FUEL ON A SOLID WALL IN A RADIATING LIQUID-CORE NUCLEAR ROCKET CONCEPT, Putre and Kasack 1968 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19680007624.pdf

Liquid Annular Reactor System (LARS)

[Paywall] Conceptual Design of a LARS Based Propulsion System, Ludewig et al 1991 https://arc.aiaa.org/doi/abs/10.2514/6.1991-3515

The Liquid Annular Reactor System (LARS) Propulsion, Powell et al 1991 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19910012832.pdf

LIQUID ANNULUS, Ludewig 1992 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19920001886.pdf

[Paywall] The liquid annular reactor system (LARS) for deep space exploration, Maise et al 1999 https://www.sciencedirect.com/science/article/abs/pii/S0094576599000442

Categories
Development and Testing Forgotten Reactors Nuclear Thermal Systems

The Bubbler: Liquid NTRs Without Barriers

Hello, and welcome back to Beyond NERVA! Today, we continue our look at liquid fueled nuclear thermal rockets (LNTRs), with a deep dive into the first of the two main types: what I call the bubbler LNTR.

This potentially attractive form of advanced NTR is a design that has been largely forgotten in the history of NTR designs outside some minor footnotes. Because of this, I felt that it was a great subject for the blog! All of the sources that I can find on the designs are linked at the end of this post, including a couple that are not available digitally, so if you’re interested in a more technical analysis of the concept please check that out!

What is a Bubbler LNTR?

Every NTR has to heat the (usually hydrogen) propellant in some way, which is usually done through (usually thermal) radiation from the fuel’s surface into the propellant.

Bubbles passing through fuel, Nelson 1963

This design, though, changes that paradigm by passing the propellant through the liquid fuel (usually a mix of uranium carbide (UC2) and some other carbide – either zirconium (ZrC) or niobium (NbC). This is done by having a porous outer wall which the propellant is injected through. This is known as a “folded flow propellant path,” and is seen in other NTRs as well, notably the Dumbo reactor from the early days of Project Rover.

In order to keep the fuel in place, each fuel element is spun at a high enough rate to keep the fuel in place using centrifugal force. The number of fuel elements is one of the design choices that varies from design to design, and the overall diameter, as well as the thickness of the fuel layer, is a matter of some design flexibility as well, but on average the individual fuel elements range from about 2 to about 6 inches in diameter, with the ratio between the thickness of the fuel layer and the thickness of the central void where the now-hot propellant passes through to the nozzle being roughly 1:1.

This was the first type of LNTR to be proposed, and was a subject of study for over a decade, but seems to have fallen out of favor with NTR designers in the late 1960s/early 1970s due to fuel/propellant interaction complications and engineering challenges related to the physical structures for injecting the propellant (more on that later).

Let’s look at the history of bubbler LNTR in more depth, and see how the proposals have evolved over time.

History of the Bubbler-type LNTR: The First of its Kind

McCarthy, 1954

Image from Barrett, Jr 1964

The first proposal for a liquid fueled NTR was in 1954, by J McCarthy in “Nuclear Reactors for Rockets” [ed. Note I have been unable to locate this report in digital form, if anyone is able to help me get ahold of it I would greatly appreciate your assistance; the following summary is based on references to this study in later works]: This design was the first to suggest the centrifugal containment of liquid fuel, and was also the first of the bubbler designs. It used a single fuel element as the entire reactor, with a large central void in the center of the fuel body as the propellant flow channel once it left the fuel itself.

This design was fundamentally limited by three factors:

  1. A torus is a terrible neutronic structure, and while the hydrogen propellant in the central void of the fuel would provide some neutron moderation, McCarthy found upon running the MCNP calculations that the difference was so negligible that it could be assumed to be a vacuum; and
  2. Only a certain amount of heat could be removed from the fuel by the propellant based on assumed fuel element geometry, and that cooling the reactor could pose a major challenge at higher reactor powers; and
  3. The behavior of the hydrogen as it passes through, and also out of, the liquid fuel was not well understood in practice, and
  4. the vapor pressure of the fuel’s constituent components could lead to fuel being absorbed in the gas as vapor in both the bubbles and exhausting propellant flow, causing both a loss of specific impulse and fissile fuel. This process is called “entrainment,” and is a (if not the) major issue for this type of reactor.

However, despite these problems this design jump started the design of LNTRs, defined the beginnings of the design envelope for this type of engine, and introduced the concept of the bubbler LNTR for the first time.

The Princeton LNTR, 1963

Princeton LNTR, Nelson et al 1963

The next major design step was undertaken by Nelson et al at Princeton’s Dept. of Aeronautical Engineering in 1963, under contract by NASA. This was a far more in-depth study than the proposal by McCarthy, and looked to address many of the challenges that the original design faced.

Perhaps the most notable change was the shift from a single large fuel element to multiple smaller ones, arranged in a hexagonal matrix for maximum fuel element packing. This does a couple of things:

  1. It homogenizes the reactor more. While heterogeneous (mixed-region) reactors work well, for a variety of reasons it’s beneficial to have a more consistent distribution of materials through the core – mainly for neutronic properties and ease of modeling (this is 1963, MCNP in a heterogeneous core using a slide rule sounds… agonizing).
  2. Given a materially limited, fixed specific impulse (see the Fuel Materials Constraints section for more in-depth discussion on this) NTR, the thrust is proportional to the total surface area of the fuel/propellant interface. By using multiple fuel elements (which they call vortices), the total available surface area increases in the same volume, increasing the thrust without compromising isp (this also implies a greater specific power, another good thing in an NTR).

This was a thermal (0.37 eV) neutron spectrum reactor, fueled by a mix of UC2 and ZrC, varying the dilution level for greater moderation and increased thermal limits. It was surrounded by a 21 cm reflector of beryllium (a “standard reflector”).

From there, the basic geometry of the reactor, from the number of fuel elements and their fueled thickness, to the core diameter and volume (the length was at a fixed ratio compared to the radius), to the shape, velocity, and number of bubbles (as well as vapor entrainment losses of the fuel material) were studied.

This was a fairly limited study, despite its length, due to the limitations of the resources available. Transients and reactor kinetics were specifically excluded from this study, the hydrogen was again replaced with vacuum in calculations, the temperature was assumed to be higher than possible due to vapor entrainment problems (4300 K, instead of 3600 K at 10 atm, 3800 at 30 atm) the chamber pressure was limited to only >1 atm, and age-diffusion theory calculations only give results within an order of magnitude… but it’s still one of the most thorough study of LNTRs I’ve found, and the most researched bubbler architecture. They pointed out the potential benefits of the use of 233U, or a larger but neutronically equivalent volume of 232Th (turning the reactor into a thermal breeder), in order to improve the overall vaporization characteristics, but this was not included in the study.

Barrett LNTR, 1964

The next year, W. Louis Barrett presented a variation of the Princeton LNTR at the AIAA Conference. The main distinction between the two designs was the addition of zirconium hydride in the areas between the fuel elements and the outer reflector, and presented the first results from a study being conducted on the bubble behavior in the fuel (being conducted at Princeton at the time). The UC2/ZrC fuel was the same, as were the number of fuel elements and reactor dimensions. The author concluded that a specific impulse of 1500-1550 seconds was possible, with a T/W of 1 at 100 atm, with thrust not being limited by heat transfer but by available flow area.

Below are the two relevant graphs from his findings: the first is the point at which the fissile fuel itself would end up becoming captured by the passing gas, and the second looks at the maximum specific impulse any particular fissile fuel could theoretically offer. The image for the McCarthy reactor above was from the same paper.

Final Work: Bubbles are Annoying

For this reactor to work, the heat must be adequately transferred from the fuel element to the propellant as it bubbles through the fuel mass radially. The amount of heat that needs to be removed, and the time and distance that it can be removed in, is a function of both the fuel and the bubbles of H2.

Sadly, the most comprehensive study of this has never been digitized, but for anyone who’s able to get documents digitized at Princeton University and would like to help make the mechanics of bubbler-type LNTRs more accessible, here’s the study: Liebherr, J.F., Williams, P.M., and Grey, J., “Bubble Motion Studies for the Liquid Core Nuclear Rocket,” Princeton University Aeronautical Engineering Report No. 673, December 1963. Apparently you can check it out after you can convince the librarians to excavate it, based on their website: https://catalog.princeton.edu/catalog/1534764.

McGuirk 1972

Here, a clear plastic housing was constructed which consisted of two main layers: an outer, solid casing which formed the outer body of the apparatus, and a perforated, inner cylinder, which simulated the fuel element canister. Water was used as the fuel element analog, and the entire apparatus was spun along its long axis to apply centrifugal acceleration to the water at various rotation rates. Pressurized air (again, at various pressures) was used in place of the hydrogen coolant. Stroboscopic photography was used to document bubble size, shape, and behavior, and these behaviors were then used to calculate the potential thermal exchange, vapor entrainment, and other characteristics of the behavior of this system.

One significant finding, based on Gray’s reporting, though, is that there’s a complex relationship between the dimensions, shape, velocity, and transverse momentum of the bubbles and their thermal uptake capacity, as well as their vapor entrainment of fuel element components. However, without being able to read this work, I can only hope someone can make this work accessible to the world at large (and if you’ve got technical knowledge and interest in the subject, and feel like writing about it, let me know: I’m more than happy to have you write a blog post on here on this INSANELY complex topic).

The last reference to a bubbler LNTR I can find is from AIAA’s Engineering Notes from May 1972 by McGuirk and Park, “Propellant Flow Rate through Simulated Liquid-Core Nuclear Rocket Fuel Bed.” This paper brings up a fundamental problem that heretofore had not been addressed in the literature on bubblers, and quite possibly spelled their death knell.

Every study until this point greatly simplified, or ignored, two phase flow thermodynamic interactions. If you’re familiar with thermodynamics, this is… kinda astounding, to be honest. It also leads me to a diversion that could be far longer than the two pages that this report covers, but I won’t indulge myself. In short, two phase flow is used to model the thermal transfer, hydro/gasdynamic properties, and other interactions between (in this case) a liquid and a gas, or a melting or boiling liquid going through a phase change.

This is… a problem, to say the least. Based on the simplified modeling, the fundamental thermal limitation for this sort of reactor was vapor entrainment of the fuel matrix, reducing the specific impulse and changing he proportions of elements in the matrix, causing potential phase change and neutronics complications.

This remains a problem, but is unfortunately not the main thermal limitation of this reactor, rather it was discovered that the amount of thermal rejection available through the bubbling of the propellant through the fuel is not nearly as high as was expected at lower propellant flow rates, and higher flow rates led to splattering of the bubbles bursting, as well as unstable flow in the system. We’ll look at the consequences of this later, but needless to say this was a major hiccup in the development of the bubbler type LNTR.

While there may be further experimentation on the bubbler type LNTR, this paper came out shortly before the cancellation of the vast majority of astronuclear funding in the US, and when research was restarted it appears that the focus had shifted to radiator-type LNTRs, so let’s move on to looking at them.

Bubbler-Specific Constraints

Fuel Element Thickness and Heat Transfer

One of the biggest considerations in a bubbler LNTR is the thickness of the fuel within each fuel canister. The fundamental trade-off is one of mechanical vs thermodynamic requirements: the smaller the internal radius at the fuel element’s interior surface, the higher the angular velocity has to be to maintain sufficient centrifugal force to contain the fuel, btu also the greater time and distance the bubbles are able to collect heat from the fuel.

In the Princeton study, the total volume within the fuel canister was roughly equally divided between fuel and propellant to achieve a comfortable trade-off between fuel mass, reactor volume, and thermal uptake in the propellant. In this case, they included the volume of the propellant as it passed through the fuel to be part of the central annulus’ volume, which eases the neutronic calculations, but also induces a complication in the actual diameter of the central void: as propellant flow increases, the void diameter decreases, requiring more angular momentum to maintain sufficient centrifugal force.

A thinner fuel element, on the other hand, runs into the challenge of requiring a greater volume of propellant to pass through it to remove the same amount of energy, but an overall lower temperature of the propellant that is used. This, in turn, reduces the propellant’s final velocity, resulting in lower specific impulse but higher thrust. However, another problem is that the fluid mixture of the propellant/fuel can only contain so much gas before major problems develop in the behavior of the mixture. In an unpublished memorandum from 1963 (“Some Considerations on the Liquid Core Reactor Concept,” Mar 23), Bussard speculated that the maximum ratio of gas to fuel would be around 0.3 to 0.4; at this point the walls of the bubbles are likely to merge, converting the fuel into a very liquidy droplet core reactor (a concept that we’ll discuss in a future blog post), as well as leading to excess splattering of the fuel into the central void of the fuel element. While some sort of recapture system may be possible to prevent fuel loss, in a classic bubbler LNTR this is an unacceptable situation, and therefore this type of limitation (which may or may not actually be 0.3-0.4, something for future research to examine) intrinsically ties fuel element thickness to maximum propellant flow rates based on volume.

There are some additional limits here, as well, but we’ll discuss those in the next section. While the propellant will gain some additional power through its passage out of the fuel element and toward the nozzle, as in the radiator type LNTR, this will not be as significant as the propellant is entering along the entire length fuel element.

Bubble Dynamics

This is probably the single largest problem that a bubbler faces: the behavior of the bubbles themselves. As this is the primary means of cooling the fuel, as well as thermalizing the propellant, the behavior of these bubbles, and the ability of the propellant stream to control the entirety of the heat generated in the fuel, is of absolutely critical importance. We looked briefly in the last section at the impacts of the thickness of the fuel, but what occurs within that distance is a far more complex topic than it may appear at first glance. With advances in two phase flow modeling (which I’m unable to accurately assess), this problem may not be nearly as daunting as it was when this reactor was being researched, but in all likelihood this set of challenges is perhaps the single largest reason that the bubbler LNTR disappeared from the design literature when it did.

The other effect that the bubbles have on the fuel is that they are the main source of vapor entrainment of fuel element materials in a bubbler, since they are the liquid/gas interface that occurs for the longest, and have the largest relative surface area. We aren’t going to discuss this particular dynamic to any great degree, but the behavior of this interaction compared to inner surface interactions will potentially be significant, both due to the fact that these bubbles are the longest-lived liquid/gas interaction by surface area and are completely encircled by the fuel itself while undergoing heating (and therefore expansion, exacerbated by the decreasing pressure from the centrifugal acceleration gradient). One final note on this behavior: it may be possible that the bubbles may become saturated with vapor during their thermalization, preventing uptake of more material while also increasing the thermal uptake of energy from the fuel (metal vapors were suggested by Soviet NTR designers, including Li and NaK, to deal with the thermal transparency of H2 in advanced NTR designs).

The behavior of the bubbles depends on a number of characteristics:

  1. Size: The smaller the bubble, the greater the surface area to volume ratio, increasing the amount of heat the can be absorbed in a given time relative to the volume, but also the less thermal energy that can be transported by each bubble. The size of the bubbles will increase as they move through the fuel element, gaining energy though heat, and therefore expanding and becoming less dense.
  2. Shape: Partially a function of size, shape can have several impacts on the behavior and usefulness of the bubbles. Only the smallest bubbles (how “small” depends on the fluids under consideration) can retain a spherical shape. The other two main shape classifications of bubbles in the LNTR literature are oblate spheroid and spherical cap. In practice, the higher propellant flow rates result in the largest, spherical cap-type bubbles in the fuel, which complicate both thermal transfer and motion modeling. One consequence of this is that the bubbles tend to have a high Reynolds number, leading to more turbulent behavior as they move through the fuel mass. Most standard two-phase modeling equations at the time had a difficult time adequately predicting the behavior of these sorts of bubbles. Another important consideration is that the bubbles will change shape to a certain degree as they pass through the fuel element, due to the higher temperature and lower centrifugal force being experienced on them as they move into the central void of the fuel element.
  3. Velocity: A function of centrifugal force, viscosity of the fuel, initial injection pressure of the propellant, density of the constituent gas/vapor mix, and other factors, the velocity of a bubble through the fuel element determines how much heat – and vapor – can be absorbed by a bubble of a given size and shape. An increase in velocity also changes the bubble shape, for instance from an oblate spheroid to a spherical cap. One thing to note is that the bubbles don’t move directly along the radius of the fuel element, both oscillation laterally and radially occur as the shape deforms and as centrifugal, convective, and other forces interact with the bubble; whether this effect is significant enough to change the necessary modeling of the system will depend on a number of factors including fuel element thickness, convective and Coriolis behavior in the fuel mass, bubble Reynolds number, and angular velocity of the fuel element,
  4. Distribution: One concern in a bubbler LNTR is ensuring that the bubbles passing through the fuel mass don’t combine into larger conglomerations, or that the density of bubbles results in a lack of overall cohesion in the fuel mass. This means that the distribution system for the bobbles must balance propellant flow rate, bubble size, velocity, and shape, non-vertical behavior of the bubbles, and the overall gas fraction of the fuel element based on the fuel element design being used.

As mentioned previously, the final paper on the bubbler I was able to find looked at the challenges of bubble dynamics in a simulated LNTR fuel element; in this case using water and compressed air. Several compromises had to be made, leading to unpredictable behavior of the propellant stream and the simulated fuel behavior, which could be due to the challenges of using water to simulate ZrC/UC2, including insufficient propellant pressure, bubble behavior irregularities, and other problems. Perhaps the most major challenge faced in this study is that there were three distinct behavioral regimes in the two phase system: orderly (low prop pressure), disordered (medium prop pressure), and violent (high prop pressure), each of which was a function of the relationship between propellant flow and centrifugal force being applied. As suspected, having too high a void fraction within the fuel mass led to splattering, and therefore fuel mass loss rates that were unacceptably high, but the point that this violent disorder occurred was low enough that it was not assured that the propellant might not be able to completely remove all the thermal energy from the fuel element itself. If the energy level of each fuel element is reduced (by reducing the fissile component of the fuel while maintaining a critical mass, for instance), this can be compensated for, but only by losing power density and engine performance. The alternative, increasing the centrifugal force on the system, leads to greater material and mechanical challenges for the system.

Adequately modeling these characteristics was a major challenge at the time these studies were being conducted, and the number of unique circumstances involved in this type of reactor makes realistic modeling remain non-trivial; advances in both computational and modeling techniques make this set of challenges more accessible than in the 1960s and 70s, though, which may make this sort of LNTR more feasible than it once was, and restarting interest in this unique architecture.

These constraints define many things in a bubbler LNTR, as they form the single largest thermodynamic constraint on the engine. Increasing centrifugal force increases the stringency for both the fuel element canister (with incorporated propellant distribution system), mechanical systems to maintain angular velocity for fuel containment, maximum thrust and isp for a given design, and other considerations.

Suffice to say, until the bubble behavior, and its interactions with the fuel mass, can be adequately modeled and balanced, the bubbler LNTR would require significant basic empirical testing to be able to be developed, and this limitation was probably a significant contributor to the reason that it hasn’t been re-examined since the early-to-mid 1970s.

The “Restart Problem”

The last major issue in a bubbler-type design is the “restart problem”: when the reactor is powered down, there will be a period of time when the fuel is still molten, requiring centrifugal containment, but the reactor being powered down allows for the fuel to be pressed into the pores of the fuel element canister, blocking the propellant passages.

One potential solution for the single fuel element design was proposed by L. Crocco, who suggested that the fuel material is used for the bubbling structure itself. When powered up, the fuel would be completely solid, and would radiate heat in all directions until the fuel becomes molten [ed. Note: according to Crocco, this would occur from the inner surface to the outer one, but I can’t find backup for that assumption of edge power peaking behavior, or how it would translate to a multi-fuel-element design], and propellant would be able to pass through the inner layers of the fuel element once the liquid/solid interface reached the pre-drilled propellant channels in the fuel element.

Another would be to continue to pass the hydrogen propellant through the fuel element until the pressure to continue pumping the H2 reaches a certain threshold pressure, then use a relief valve to vent the system elsewhere while continuing to reject the final waste heat until a suitable wall temperature has been reached. This is going to both make the fuel element less dense, and also result in a lower fuel element density near the wall than at the inner surface of the fuel element. While this could maybe [ed. Note: speculation on my part] make it so that the fuel is more likely to melt from the inner surface to the outer one, the trapped H2 may also be just enough to cause power peaking around the bubbles, allow chemical reactions to occur during startup with unknown consequences, and other complications that I couldn’t even begin to guess at – but the tubes would be kept clear.

Wall Material Constraints

Other than the “restart problem,” additional constraints apply to the wall material. It needs to be able to handle the rotational stresses of the spinning fuel element, be permeable to the propellant, and able to withstand rather extreme thermal gradients: on one side, gaseous hydrogen at near-cryogenic temperatures (the propellant would have already absorbed some heat from the reactor body) to about 6000 K on the inside, where it comes in contact with the molten fuel.

Also, the bearings holding the fuel element will need to be designed with care. Not only do they need to handle the rather large amount of thermal expansion that will occur in all directions during reactor startup, they have to be able to deal with high rotation rates throughout the temperature range.

The Paths Not (Yet?) Taken

Perhaps due to the early time period in which the LNTR was explored, a number of design options don’t seem to have been explored in this sort of reactor.

One option is neutron moderator. Due to the high thermal gradients in this reactor, ZrH and other thermally sensitive moderators could be used to further thermalize the neutron spectrum. While this might not be explicitly required, it may help reduce the fissile requirements of the reactor, and would not be likely to significantly increase reactor mass.

A host of other options are possible as well, if you can think of one, comment below!

Diffuser LNTR

The other option was brought up by Michael Turner at Project Persephone, in regards to the vapor entrainment and restart problem issues: what if you get rid of the holes in the walls of the fuel element, and the bubbles through the fuel element, altogether? As we saw when discussing Project Rover, hydrogen gets through EVERYTHING, especially hot metals. This diffusion process is done through individual molecules, not through bubbles, meaning that the possibility of vapor entrainment is eliminated. The down side is that the propellant mass flow will be extremely reduced, resulting in a higher-isp (due to the ability to increase fuel temp because the vapor losses are minimized), much-lower-thrust reactor than those designed before. As he points out, this may be able to be mixed with bubbles for a high-thrust, lower-isp mode, if “shutters” on the fuel element outer frit were able to be engineered. Another possible requirement would be to reduce the fissile component density of the fuel to match the power output to the hydrogen flow rates, or to create a hybrid diffuser/radiator LNTR to balance the propellant flow and thermal output of the reactor.

I have not been able to calculate if this would be feasible or not, and am reasonably skeptical, but found it an intriguing possibility.

Conclusion

The bubbler liquid nuclear thermal rocket is a fascinating concept which has not been explored nearly as much as many other advanced NTR designs. The advantage of being able to fully thermalize the propellant to the highest fuel element temperature while maintaining cryogenic temperatures outside the fuel element is a rarity in NTR design, and offers many options for structures outside the fuel elements themselves. After over a decade of research at Princeton (and other centers), the basic research on the dynamics of this type of reactor has been established, and with the computational and modeling capabilities that were unavailable at the time of these studies, new and promising versions of this concept may come to light if anyone chooses to study the design.

The problems of vapor entrainment, fissile fuel loss, and restarting the reactor are significant, however, and impact many area of the reactor design which have not been addressed in previous studies. Nevertheless, the possibility remains that this drive may one day indeed make a useful stepping stone from the solid-fueled NTRs of tomorrow to the advanced NTRs of the decades ahead.

References

A Technical Report on the CONCEPTUAL DESIGN – STUDY OF A LIQUID-CORE NUCLEAR ROCKET, Nelson et al 1963 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19650026954.pdf

The Liquid Core Nuclear Rocket, Grey 1965 (pg 92) https://permalink.lanl.gov/object/tr?what=info:lanl-repo/lareport/LA-03229-MS

Specific Impulse of a Liquid Core Nuclear Rocket, Barrett Jr 1963 https://arc.aiaa.org/doi/abs/10.2514/3.2141?journalCode=aiaaj

Propellant Flow Rate through Simulated Liquid-Core Nuclear Rocket Fuel Bed, McGuirk and Park 1972 https://arc.aiaa.org/doi/abs/10.2514/3.61690?journalCode=jsr

Categories
Nuclear Thermal Systems Test Stands

Fluid Fueled NTRs: A Brief Introduction

 Hello, and welcome back to Beyond NERVA! This is actually about the 6th blog post I’ve started, and then split up when they ran more than 20 pages long, in the last month, and more explanatory material was needed before I discussed the concepts I was trying to discuss (this blog post has also been split up multiple times).

I apologize about the long hiatus. A combination of personal, IRL complications (I’ve updated the “About Me” section to reflect this, but those will not affect the type of content I share on here), and the professional (and still under wraps) opportunity of a lifetime have kept me away from the blog for a while. I want to return to Nuclear Thermal Rockets (NTRs) for a while, rather than continuing Nuclear Electric Propulsion (NEP) power plants, as a fun, still-not-covered area for me to work my way back into writing regularly for y’all again.

This is the first in an extensive blog series on fluid fueled NTRs, of three main types: liquid, vapor, and gas core NTRs. These reactors avoid the thermal limitations of the fuel elements themselves, increasing the potential core temperature to above 2550 K (the generally accepted maximum thermal limit on workable carbide fuel elements), increasing the specific impulse of these rockets. At the same time, structural material thermal limits, challenges in adequately heating the propellant to gain these advantages in a practical way, fissile fuel containment, and power density issues are major concerns in these types of reactors, so we’re going to dig into the weeds of the general challenges of fluid fueled reactors in general in this blog post (with some details on each reactor type’s design envelope).

Let’s start by looking at the basics behind how a nuclear reactor can operate without any solid fuel elements, and what the advantages and disadvantages of going this route are.

Non-Solid Fuels

A nuclear reactor is, at its basic level, a method of maintaining a fission reaction in a particular region for a given time. This depends on maintaining a combination of two characteristics: the number of fissile atoms in a given volume, and the number and energy of neutrons in that same volume (the neutron flux). As long as the number of neutrons and the number of fissile atoms in the area are held in balance, a controlled fission reaction will occur in that area.

Solid Core Fuel Element, image DOE

The easiest way to maintain that reaction is to hold the fissile atoms in a given place using a solid matrix of material – a fuel element. However, a number of things have to be balanced for a fuel element to be a useful and functional piece of reactor equipment. For an astronuclear reactor, there are two main concerns: the amount of power produced by the fission reaction has to be balanced by how much thermal energy the fuel element is able to contain, and the fuel element needs to survive the chemical and thermal environment that it is exposed to in the reactor. (Another for terrestrial reactors is that the fuel element has to contain the resulting fission products from the reaction itself, as well as any secondary chemical pollutants, but this isn’t necessarily a problem for astronuclear reactors, where the only environment that’s of concern is the more heavily shielded payload of the rocket.) 

This doesn’t mean that a reactor has to use a solid fuel element. As the increasingly well known molten salt reactor, as well as various other fluid fueled reactor concepts, demonstrate, the only requirement is the combination of the number of fissile atoms and the required energy level and density of neutrons to exist in the same region of the reactor. This, especially in Russian literature, is called the “active zone” of the reactor core. This can be an especially useful as a term, since the reactor core can contain areas that aren’t as active in terms of fission activity. (A great example of this is the travelling wave reactor, most recently investigated – and then abandoned – by Terrestrial Energy.) But more generally it’s useful to differentiate the fueled areas undergoing fission from other structures in the reactor, such as neutron moderation and control regions in the reactor. The key takeaway is that, as long as there is enough fuel, and the right density of neutrons at the right energy, then a sustained – and controlled – fission reactor has been achieved.

The obvious consequence is that the solid fuel element isn’t required – and in the case of a nuclear thermal rocket, where the efficiency of the rocket is directly tied to the temperature it can achieve, the solid fuel is in fact a major limitation to a designer. The downside to this is that, unlike solids, fluids tend to move, especially under thrust. Because the materials used in a solid fueled rocket are already at the extremes of what molecular bonds can handle, this means that either very clever cooling or very robust containment methods need to be used to keep the rest of the reactor from destroying itself.

Finally, one of the interesting consequences of not having a solid fuel element is that the reactor’s power density (W/m^2) and specific power (W/kg) can be increased in proportion to how much coolant can be used in theory, but in practice it can be challenging to maintain a high power density in certain types of fluid fueled reactors due to the high rate of thermal expansion that these reactors can undergo. There are ways around this, and fluid fueled reactors can have higher power densities than even closely related solid fueled variants, but the fact that fluids are able to expand much more than solids under high temperatures is an effect that should be taken into account. On the other hand, if the fluid expands too much, it can drop the power density, but not necessarily the specific mass of the system.

Types of and Reasons for Fluid Fuels

Fluid fuels fall into three broad categories: liquids, vapors, and gasses. There are intermediate steps, and hybrids between various phase states of fuel, but these three broad categories are useful. While liquid fuels are fairly self-explanatory (a liquid state fissile material is used to fuel the core, often uranium carbide mixed with other carbides, or U-Mo, but other options exist), the vapor and gas concepts are far less straightforward overall. The vapor core has two major variants: discrete liquid droplets, or a low pressure, relatively low temperature gaseous suspension similar to a cloud. The gas core could be more appropriately called a “plasma core,” since these are very high temperature reactors, which either mechanically hold the plasma in place, or use hydrodynamic or electrodynamic forces to hold the plasma in place.

However, they all have some common advantages, so we’ll look at them as a group first. The obvious reason for using non-solid fuel, in most cases, is that they are generally less thermally limited than solid fuels are (with some exceptions). This means that higher core temperatures, and therefore higher exhaust velocity (and specific impulse) can be achieved.

Convection pattern in radiator-type
liquid fuel element, image DOE

An additional benefit to most fluid fueled designs is that the fluid nature of the fuel helps mitigate or eliminate hot spots in the fuel. With solid fuels, one of the major challenges is to distribute the fissile material throughout the fuel as evenly as possible (or along a specifically desired gradient of fissile content depending on the position of the fuel element within the reactor). If this isn’t done properly, either through a manufacturing flaw or migration of the fissile component as a fuel element becomes weakened or damaged during use, then a hot spot can develop and damage the fuel element in both its nuclear and mechanical properties, leaning to a potentially failed fuel element. If the process is widespread enough, this can damage or destroy the entire reactor.

Fluid fuels, on the other hand, have the advantage that the fuel isn’t statically held in a solid structure. Let’s look at what happens when the fuel isn’t fully homogeneous (completely mixed) to understand this:

  1. A higher density of fissile atoms in the fuel results in more fission occurring in a particular volume.
  2. The fuel heats up through both radiation absorption and fission fragment heating.
  3. The fuel in this volume becomes less dense as the temperature increases.
  4. The increased volume, combined with convective mixing of cooler fuel fluids and radiation/conduction from the surface of the hotter region cools the region further.
  5. At the same time, the lower density decreases the fission occurring in that volume, while it remains at previous levels in the “normally heated” regions.
  6. The hot spot dissipates, and the fuel returns to a (mostly) homogeneous thermal and fissile profile.

In practice, this doesn’t necessarily mean that the fuel is the same temperature throughout the element – this very rarely occurs, in fact. Power levels and temperatures will vary throughout the fuel, causing natural vortices and other structures to appear. Depending on the fuel element configuration, this can be either minimized or enhanced depending on the need of the reactor. However, the mixing of the fuel is considered a major advantage in this sort of fuel.

Another advantage to using fluid fuels (although one that isn’t necessarily high on the priority list of most designs) is that the reactor can be refueled more easily. In most solid fueled reactors, the fissile content, fission poison content, and other key characteristics are carefully distributed through the reactor before startup, to ensure that the reactor will behave as predictably as possible for as long as possible at the desired operating conditions. In terrestrial solid reactors, refueling is a complex, difficult process, which involves moving specific fuel bundles in a complex pattern to ensure the reactor will continue to operate properly, with only a little bit of new fuel added with each refueling cycle.

PEWEE Test stand, image courtesy DOE

There were only two refuelable NTR testbeds in the US Rover program: Pewee and the Nuclear Furnace. Both of these were designed to be fuel element development apparatus, rather than functional NTRs (although Pewee managed to hit the highest Isp of any NTR tested in Rover without even trying!), so this is a significant difference. While it’s possible to refuel a solid core NTR, especially one such as the RD-0410 with its discrete fuel bundles, the likely method would be to just replace the entire fueled portion of the reactor – not the best option for ease of refueling, and one that would likely require a drydock of sorts to complete the work. To give an example, even the US Navy doesn’t always refuel their reactors, opting for long-lived highly enriched uranium fuel which will last for the life of the reactor. If the ship needs refueled, the reactor is removed and replaced whole in most cases. This reticence to refuel solid core reactors is likely to still be a thing in astronuclear reactors for the indefinite future, since placing the fuel elements is a complex process that requires a lot of real-time analysis of the particulars of the individual fuel elements and reactors (in Rover this was done at the Pajarito Site in Los Alamos).

Fluid fuels, though, can be added or removed from the reactor using pumps, compressed gasses, centrifugal force, or other methods. While not all designs have the capability to be refueled, many do, and some even require online fuel removal, processing and reinsertion into the active region of the core to maintain proper operation. If this is being done in a microgravity environment, there will be other challenges to address as well, but these have already been at least partially addressed by on-orbit experiments over the decades in the various space programs. (Specific behaviors of certain fluids will likely need to be experimentally tested for this particular application, but the basic physics and engineering solutions have been researched before).

Finally, fluid fuels also allow for easier transport of the fuel from one location to another, including into orbit or another planet. Rather than having a potentially damageable solid pellet, rod, prism, or ribbon, which must be carefully packaged to not only prevent damage but accidental criticality, fluids can be transported with far less risk of damage: just ensure that accidental criticality can’t occur, chemical compatibility between the fluid and the vessel it’s carrying, and package it strongly enough to survive an accident, and the problem is solved. If chemical processing and synthesis is available wherever the fuel is being sent (likely, if extensive and complex ISRU is being conducted), then the fuel doesn’t even need to be in its final form: more chemically inert options (UF4 and UF6 can be quite corrosive, but are easily managed with current materials and techniques), or less fissile-dense options (to reduce the chance of accidental criticality further) can be used as fuel precursors, and the final fuel form can be synthesized at the fueling depot. This may not be necessary, or even desirable, in most cases, but the option is available.

So, while solid fuels offer certain advantages over fluid fuels, the combination of being more delicate (thermally, chemically, and mechanically) combine to make fluid fuels a very attractive option. Once NTRs are in use, it is likely that research into fluid fueled NTRs will accelerate, making these “advanced” systems a reality.

Fuel Elements: An Overview

Now that we’ve looked at the advantages of fluid fuels in general, let’s look at the different types of fluid fuels and the proposals for the form the fuel elements in these reactors would take. This will be a brief overview of the various types of fuels, with more in-depth examinations coming up in future blog posts.

Liquid Fuel

A liquid fueled reactor is the best known popularly, although the most common type (the molten salt reactor) uses either fluoride or chloride salts, both of which are very corrosive at the temperatures an NTR operates at. While I’ve heard arguments that the extensive use of regenerative cooling can address this thermal limitation, this would still remain a major problem for an NTR. Another liquid fuel type, the molten metal reactor, has also been tested, using highly corrosive plutonium fuel in the best known case (the Liquid Annular Molten Plutonium Reactor Experiment, or LAMPRE, run by Los Alamos Scientific Lab from 1957 to 1963, covered very well here).

Early bubbler-type liquid NTR, Barrett 1963

The first proposal for a liquid fueled NTR was in 1954, by J McCarthy in “Nuclear Reactors for Rockets.” This design spun molten uranium carbide to produce centrifugal force (a common characteristic in liquid NTRs of all designs), and passed the propellant through a porous outer wall, through the fuel mass, and into the central void in the reactor before it was ejected out of the nozzle.The main problem with this reactor was that the tube was simply too large to allow for as much heat transfer as was ideal to take place, so the next evolution of the design broke up the single large spinning fuel element up into several thinner ones of the same length, increasing the total surface area available for heating the propellant. This work was conducted at Princeton, and would continue on and off until 1973. These designs I generally call “bubblers,” due to the propellant flow path.

Princeton multi-fuel-element bubbler, Nelson et al 1963

One problem with these designs is that the fuel would vaporize in the low pressure hydrogen environment of the bubbles, and significant amounts of uranium would be lost as the propellant went through the fuel. Not only is uranium valuable, but it’s heavy, reducing the exhaust velocity and therefore the specific impulse. Another issue is that there are hard limits to how much propellant can be passed through the fuel at any given time before it starts to splatter, directly tying thrust to fuel volume. 

In order to combat this, a team at NASA’s Lewis Research Center decided to study the idea of passing the propellant only through the central void in the fuel, allowing radiation to be the sole means of heating the propellant. Additional regenerative cooling structures were needed for this design, and ensuring the propellant got heated sufficiently was a challenge, but this sort of LNTR, the radiator type, became the predominant design. Vapor losses of the uranium were still a problem, but were minimized in this configuration.

It too would be cancelled in the late 1960s, but briefly revived by a team at Brookhaven National Laboratory in the early 1990s for possible use in the Space Exploration Initiative, however this program was not selected for further development.

Despite these challenges, liquid core NTRs have the potential to reach above 1300 s isp, and a T/W ratio of up to 0.5, so there is definite promise in the concept.

Droplet/Vapor Fuel

Picture a spray bottle, the sort used for household plants, ironing, or cleaning products like window cleaner. When the trigger is pulled, there’s a fine spray of liquid exiting the nozzle, which contains a mix of liquid and gas. Using a similar system to mix liquids and gasses is possible in a nuclear reactor, and is called a droplet core NTR. This reactor type is useful in that there’s incredible surface area available for radiation to occur into the propellant, but unfortunately it also means that separating the fuel droplets from the propellant upon leaving the nozzle (as well as preventing the fuel from coating the reactor core walls) is a major hydrodynamics challenge in this type of reactor.

Vapor core NTR, Diaz et al 1992

The other option is to use a vapor as fuel. A vapor is a substance that is in a gaseous state, but not at the critical point of the material – i.e. at standard temperature and pressure it would still be a liquid. One interesting property of a vapor is that a vapor is able to be condensed or evaporated in order to change the phase state of the substance without changing its temperature, which could be a useful tool to use for reactor startup. The downside of this type of fuel is that it has to be in an enclosed vessel in order to maintain the vapor state.

So why is this useful in an NTR? Despite the headaches we’ve just (briefly, believe it or not) discussed in the liquid fuels section, liquid fuel has a major advantage over gaseous fuel (our next section): the liquid phase is far better at containing its constituent parts than the gas phase is, due to the higher interatomic bond strength. At the same time, maintaining a large, liquid body can be a challenge, especially in the context of complex molecular structures in some of the most chemically difficult elements known to humanity (the actinides and transuranics). If the liquid component is small, though, it’s far easier to manage the thermal distribution, as well as offering greater thermal diffusion options (remember, the heat IN the fissile fuel needs to be moved OUT of it, and into the propellant, which is a direct function of available surface area).

The droplet core NTR offers many advantages over a liquid fuel in that the large-scale behavior of the liquid fuel isn’t a concern for reactor dynamics, and the aforementioned high surface area offers awesome thermal transfer properties throughout the propellant feed, rather than being focused on one volume of the propellant.

Vapors offer a middle ground between liquids and gasses: the fissile fuel itself is in suspension, meaning that the individual molecules of fissile fuel are able to circulate and maintain a more or less homogeneous temperature. 

This is another design concept that has seen very little development as an NTR (although NEP applications have been investigated more thoroughly, something that we’ll discuss the application and complications of, for an NTR in the future). In fact, I’ve only ever been able to find one design of each type designed for NTR use (and a series of evolving designs for NEP), the appropriately named Droplet Core Nuclear Rocket (DCNR) and the Nuclear Vapor Thermal Reactor (NTVR).

Droplet Core NTR, Anghaie et al 1992

The DCNR was developed in the late 1980s based on an earlier design from the 1970s, the colloid core reactor. The original design used ultrafine microparticles of U-C-ZR carbide fuel, which would be suspended in the propellant flow. This sort of fuel is something that we’ll look at more when covering gas core NTRs (metal microparticles are one of the fuel types available for a GCNTR), but the use of carbides increases the fuel failure temperature to the point that structural components would fail before the fuel itself would, leading to what could be called an early pseudo-dusty plasma NTR. The droplet core NTR took this concept, and applied it to a liquid rather than solid fuel form. We’ll look at how the fuel was meant to be contained before exiting the nozzle in the next section, but this was the main challenge of the DCNR from an engineering point of view.

The NVTR was a compromise design based on NERVA fuel element development with a different fissile fuel carrier. Here, the fuel (in the form of UF4) is contained within a carbon-carbon composite fuel element in sealed channels, with interspersed coolant channels to manage the thermal load on the fuel element. While significant thrust-to-weight ratio improvements were possible, and (in advanced NTR terms) modest specific impulse gains were possible, the design didn’t undergo any significant development. We’ll cover containment in the next section, and other options for architectures as well.

Gas Fuel

Finally, there are gas core NTRs. In these, the fuel is in gaseous form, allowing for the highest core temperatures of any core configuration. Due to the very high temperatures of these reactors, the uranium (and in general the rest of the components in the fuel) become ionized, meaning that a “plasma core” is as accurate a description as a “gas core” is, but gas remains the convention. The fuel form for a gas core NTR has a few variants, with the most common being UF6, or metal fuel which vaporizes as it is injected into the core. Due to the high temperatures of these reactors, the UF6 will often break down as all of the constituent molecules become ionized, meaning that whatever structures will come in contact with the fuel itself (either containment structures or nozzle components) must be designed in such a way to prevent being attacked by high temperature fluorine ions and hydrofluoric acid vapors formed when the fluorine ions come in contact with the propellant.

Containing the gas is generally done in one of three ways: either by compressing the gas mechanically in a container, by holding the gas in the middle of the reactor using the gas pressure from the propellant being injected into the core, or by using electromagnets to contain the plasma similarly to how a spherical tokamak operates. The first concept is a closed cycle   gas core (CCGCNTR, or GC-C), while the second two are called open cycle gas core NTRs (OCGCNTR or GC-O), because while the first one physically contains the fuel and prevents fission products, unburned fuel, and the previously mentioned free fluorine from exiting in the exhaust plume of the reactor, the open cycle’s largest problem in designing a workable NTR is that the vast majority (often upwards of 90%) if the uranium ends up being stripped away from the plasma body before it undergoes fission – a truly hot radioactive mess which you don’t want to use anywhere near anything sensitive to radiation and an insanely inefficient use of fissile material. There are many other designs and hybrids of these concepts, which we’ll cover in the gas core NTR series, and will look briefly at the containment challenges below.

Fluid Fuel Elements: Containment Strategies

Fluid fuels are, well, fluid. Unlike with a solid fuel element, as we’ve looked at in the past, a fluid has to be contained somehow. This can be in a sealed container or by using some outside force to keep it in place.

Another issue with fluid fuels can be (but isn’t always) maintaining the necessary density to achieve the power requirements for an NTR (or any astronuclear system, for that matter). All materials expand when heated, but with fluids this change can be quite dramatic, especially in the case of gas core NTRs. Because of this, careful design is required in order to maintain the high density of fissile fuel necessary to make a mass-efficient rocket engine possible.

This leads to a rather obvious conclusion: rather than the fuel element being a physical object, in a fluid fueled NTR the fuel element is a containment structure. Depending on the fuel type and the reactor architecture, this can take many forms, even in the same type of fuel. This will be a long-ish review of the proposed fuel containment strategies, and how they impact the performance of the reactors themselves.

One thing to note about all of these reactor types is that 235U is not required to be the fissile component in the fuel, in fact many gas core designs use 233U instead, due to the lower requirements for critical mass. (According to most Russian literature on gas core NTRs, this  reduces the critical mass requirements by 1/3). Other options include using 242mAm, a stable isomer of 242Am, which has the lowest critical mass of any fissile fuel. By using these types of fuels rather than the typical 235U, either less of the fuel mass needs to be fissile (in the case of a liquid fueled NTR), or less fuel in general is needed (in the case of vapor/gas core NTRs). This can be a double-edged sword in some systems with high fuel loss rates (like an open cycle gas core), which would require more robust and careful fuel management strategies to prevent power transients due to fuel level variations in the active zone of the reactor, but the overall reduction in fuel requirements means that there’s less fuel that can be lost. Many other fissile fuel types also exist, but generally speaking either short half-lives, high spontaneous fission rates, or expense in manufacture have prevented them from being extensively researched.

Let’s look at each of the design types in general, with a particular focus on gas core NTRs at the end.

Liquid FE

For liquid fuels, there’s one universal option for containing the fuel: by spinning the fuel element. However, after this, there’s two main camps on how a liquid fueled NTR interacts with the propellant. The original design, first proposed in the 1950s and researched at least through the 1960s, proposed the use of either one or several spinning cylinders with porous outer walls (frits), which would be used to inject the propellant into the reactor’s active region. For those that remember the Dumbo reactor, this may be familiar as a folded flow NTR, and does two things: first, it allowed the area surrounding the fuel elements to be kept at very low temperatures, allowing the use of ZrH and other thermally sensitive materials throughout the reactor, and second it increases the heat transfer area available from the fuel to the propellant. Experiments (using water as a uranium analog) were conducted to study the basics of bubble behavior in a spinning fluid to estimate fuel mass loss rates, and the impact of evaporation or vaporization of various forms of uranium (including U metal, UC2, and others) were conducted. 

This concept is the radiator type LNTR. Here, rather than the folded flow used previously, axial flow is used: the H2 is used as a coolant for reactor structures (including the nozzle) passing from the nozzle end to the ship end, and then injected through the central void of each of the fuel elements before exiting the nozzle. This design reduces the loss of fuel mass due to bubbling in the fuel, but adds an additional challenge of severely reducing the amount of surface area available for heat transfer from the fuel to the propellant. In order to mitigate this, some designs propose to seed the propellant with microparticles of tungsten, which would absorb the significant about of UV and X rays coming off the fuel, and turn it into IR radiation which is more easily absorbed by the H. At the designed operating temperatures, this reactor would dissociate the majority of the H2 into monatomic hydrogen, increasing the specific impulse significantly.

In all these designs, there is no solid clad between the fuel itself and the propellant, because this means that the hottest portion of the fuel element would be limited by how high the temperature can reach before melting the clad. Some early LNTR designs used a mix of molten UC2 and ZrC/NbC as a fuel element, with the ZrC meant to migrate to the upper areas of the fuel element and not only provide neutron moderation but reduce the amount of erosion from the propellant. It may be possible to use a liquid metal clad as a barrier to prevent mass erosion of the fissile fuel in a metal fueled reactor as well, and possibly even add some neutron moderation for the fuel element itself. However, the material would need to have not only a very high boiling point, high thermal conductivity, low reactivity to both hydrogen and the fuel, and low neutron capture cross section, it would also need to have a high vapor pressure in order to prevent erosion from the propellant flow (although I suppose adding additional clad during the course of operation would also be an option, at the cost of higher propellant mass and therefore lost specific impulse).

Droplet/Vapor FE

Now let’s look at the vapor core NTR.

NVTR fuel element, Diaz et al 1992

Containing the UF4 vapor in the NVTR vapor core NTR is done by using a sealed tube embedded in a fuel element, which is then surrounded by propellant channels to carry away the heat. Two configurations were proposed in the NTVR concept: the first used a large central cavity, sealed at both ends, to contain the vapor, and the second design dispersed the fuel cylinders in an alternating hexagonal pattern throughout the fuel element. The second option provides a more even thermal distribution not only within the fuel element itself, but across the entire active zone of the reactor core.

Droplet core NTRs are very different in their core structure. Rather than having multiple areas that the fissile fuel is isolated in, the droplet core sprays droplets of fissile fuel into a large cylinder, which is spun to induce centrifugal force. The fuel is kept away from the walls of the reactor core using a collection of high-pressure H2 jets, injecting the propellant into the fuel suspension and maintaining hydrostatic containment on the fuel. The last section of the reactor core, instead of using hydrogen, injects a liquid lithium spray to bind with the uranium, which is then carried to the walls of the reactor due to the lack of tangential force. The fuel is then recirculated to the top of the reactor vessel, where it is once again injected into the core.

This hydrostatic equilibrium concept is very similar to how many gas core NTRs operate (which we’ll look at below), and has proven to be the biggest Achilles’ Heel of these sorts of designs. While it may be theoretically possible to do this (the lower temperatures of the droplet core allow for collection and recirculation, which may provide a means of fissile fuel loss reduction), many of the challenges of the droplet core are very similar to that of the open cycle gas core, a far more capable engine type.

Gas Core

Gas core containment is possibly the most complex topic in this post, due to the sheer variety of possible designs and extreme engineering requirements. We’ll be discussing the different designs in depth in upcoming blog posts, but it’s worth doing an overview of the different designs, their strengths and weaknesses, here.

Closed Cycle

One half of the lightbulb configuration, McLafferty et al 1968

The simplest design to describe is the closed cycle gas core, which in many ways resembles a vapor core NTR. In most iterations, a sealed cylinder with a piston at one end (similar in many ways to the piston in an automobile engine), is filled with UF6 gas. This is compressed in order to reach critical geometry, and fission occurs in the cylinder. The walls of the cylinder are generally made out of quartz, which is transparent to the majority of the radiation coming off the fissioning uranium, and is able to resist the fluorination from the gas (other options include silicon dioxide, magnesium oxide, and aluminum oxide). Additionally, while the quartz will darken under the heat, the radiation actually “anneals” the quartz to keep it transparent, and coolant is run through the cylinder to maintain the material within thermal limits; a vortex is induced during fission which, when properly managed, also keeps the majority of the uranium (now in a charged state) from coming in contact with the walls of the chamber as well, reducing thermal load on the material. Some designs have used pressure waves in place of the piston to induce fission, but the fluid-mechanical result is very similar. This results in a lightbulb-like structure, hence the common nickname “nuclear lightbulb.” One variation mentioned in Russian literature also uses a closed uranium loop, circulating the fissile fuel to minimize the fission product buildup and maintain the fissile density of the reactor.

The main advantage to these types of designs is that all fission products and particle radiation are contained within the bulb structure, meaning that fission product and radiation release into the environment is eliminated, with only gamma and x-ray radiation during operation being a concern. However, due to the fact that there’s a solid structure between the fuel element and the propellant, this engine is thermally limited more than any other gas core design, and its performance in both thrust and specific impulse suffers as a result.

Open Cycle

The next very broad category is an open cycle gas core. Here, there is usually no solid structure between the fissioning uranium and the propellant, meaning that core temperatures can reach astoundingly high temperatures (sometimes limited only by the melting temperature of the materials surrounding the active reactor zone, such as reflectors and pressure vessel). Sadly, this also means that actually containing the fuel is the single largest challenge in this type of reactor, and the exhaust tends to be incredibly radioactive as a result, On the plus side, this sort of rocket can achieve isp in the tens of thousands of seconds (similar to or better than electric propulsion), and also achieve high thrust.

Perhaps the easiest way to make a pure open cycle gas core NTR is to allow the fuel and the propellant to fully mix, similarly to how the droplet core NTR was done, and either ensure all (or most) of the fissile fuel is burned before leaving the rocket nozzle. Insanely radioactive, sure, but with a complete mixing of the fissioning atoms and the propellant the theoretically most efficient transfer of energy is possible. However, the challenge of fully fissioning the fuel in such a short period of time is significant, and I can’t find any evidence of significant research into this type of gas core reactor.

Due to the challenges of burning the fissile fuel completely enough during a single pass through the reactor, though, it is generally considered required to maintain a more stable fissile structure within the reactor’s active region. Maintaining this sort of structure is a challenge, but is generally done through gasdynamic effects: the propellant injected into the reactor is used to push the fuel back into the center of the reactor. This involves the use of a porous outer wall of the reactor, where the hydrogen propellant is inserted at a high enough pressure and evenly enough spaced intervals to counterbalance both the tendency of the plasma to expand until it’s not able to undergo fission and the tendency of the fuel to leave the nozzle before being burned.

Soviet-type Vortex Stabilized open cycle, image Koroteev et al 2007

The next way is to create a low pressure stagnant area in the center of the core, which will contain the fissile fuel. In order to maintain this type of pressure differential, a solid structure is usually needed, generally made out of a high temperature refractory metal. In a way this is a hybrid closed/open cycle gas core (even though the plasma isn’t in direct contact with the structure of the reactor itself), because the structure itself is key to generating this low pressure zone necessary for maintaining this plasma body fuel element. This type of NTR has been the focus of Russian gas core research since the 1970s, and will be covered more in the future.

Spherical gas core diagram, image NASA

As I’m sure most of you have guessed, fuel containment is a very complex and difficult problem, and one that’s had many solutions over the years (which we’ll cover in a future post). Most recent gas core NTR designs in the US are based on the spherical gas core. Here, the plasma is held in the center of the active zone using jets of propellant from all sides. This is generally called a porous wall gas core NTR, and while it takes advantage of any vortex stabilization that may occur in the fuel, it does not rely on it; in many ways, it’s a lot like an indoor skydiving arena with air jets blowing from all sides. This design, first proposed in the 1970s, uses high pressure propellant to contain the fuel in the reactor, and in many designs the flow can be adjusted to deal with the engine being under thrust, pushing the fuel toward the nozzle in traditional design configurations. Most designs suffer from massive erosion of the fuel by shear forces from the propellant eroding the fuel from the outside edge, but in some conceptual sketches this can be gotten around using non-traditional nozzle configurations which have a solid structure along the main thrust axis of the rocket. (More on that in a future post. I’m still trying to track down the sources to fully explain that pseudo-aerospike concept).

Hybrid gas core diagram, Beveridge 2017

The most promising designs as far as fuel loss rates minimize the amount of plasma required to maintain the reaction. This is what’s known as a hybrid solid-gas NTR, first proposed by Hyland in the 1970s, and also one of the designs which has been most recently investigated by Lucas Beveridge. Here, the fissile fuel is split between two components: the high-temperature plasma fuel is used for final heating of the propellant, but isn’t able to sustain fission independently. Instead, a sphere of solid fuel encases the outside of the active zone of the reactor. This minimizes the amount of fuel that can be easily eroded while ensuring that a critical mass of fissile material is contained in the active region of the reactor. This really is less complicated than it sounds, but is difficult to summarize briefly without delving into the details of critical geometry, so I’ll try to explain it this way: the interior of the reactor is viewed by the neutrons in the reactor as a high-density low temperature fuel area, surrounding a low density high temperature fuel area, with the coolant/moderator passing through the high density area and flowing around the low density area, making a complete reactor between these parts while minimizing how much of the low density fuel is needed and therefore minimizing the fuel loss. I wish I was able to make this more clear in less than a couple pages, but sadly I’m not that good at summarizing in non-technical terms. I’ll try and do better on the hybrid core post coming in the future.

All of these designs suffer from massive fuel loss, leading to highly radioactive exhaust and incredibly inefficient engines which are absurdly expensive to operate due to the amount of highly enriched fissile fuel needed. (Because everything going into the reactor needs to fission as quickly as possible, every component of the fuel itself needs to undergo fission as easily as possible.) This is the major Achilles heel of this NTR type: despite the massive potential promise, the fuel loss, and radioactive plume coming off these reactors, make them unusable with current engineering.

There’s going to be a lot more that I’m going to write about this type of NTR, and I skipped a lot of ideas, and variations on these ideas, so expect a lot more in the coming year on this subject.

Cooling the Reactor/Heating the Propellant

Finally there’s cooling, which usually comes in one of two varieties:

  1. cooling using the propellant, as in most NTR designs that we’ve seen, to reject all the heat from the reactor
  2. cooling in a closed loop, as is done in an NEP system
Hybrid gas core with secondary cooling diagram, Beveridge 2017

While the ideal situation is to reject all the heat into the propellant, which maximizes the thrust and minimizes the dry mass of the system, this is the exception in many of these systems, rather than the norm. There’s a couple reasons for this: containing a fluid with fast-moving (or high pressure) hydrogen is challenging because the gas wants to strip away the mass that it comes in contact with (far easier in a fluid than a solid), H2 is insanely difficult to contain at almost any temperature, and these reactors are designed to achieve incredibly high temperatures which can outstrip the available heat rejection area that the reactor designs allow.

Complicating the issue further, hydrogen is mostly transparent to the radiation that a nuclear reactor puts off (mostly in the hard UV/X/gamma spectrum), meaning that it takes a lot of hydrogen to reject the heat produced in the reactor (a common complaint in any gas-cooled reactor, to be fair), and that hydrogen doesn’t get heated that much on an atom-by-atom basis, all things considered.

There’s a way around this, though, which many designs, from LARS on the liquid side to basically every gas core design I’ve ever seen use: microparticle or vapor seeding. This is a form of hybrid propellant, which I mention in my NTR propellants page. Basically, a metal is ground incredibly fine (or is vaporized), and then included in the propellant feed. This captures the high-wavelength photons (due to its higher atomic mass, and greater opacity to those wavelengths as a result), which are re-emitted at a lower frequency which is more easily absorbed by the propellant. While the US prefers to use tungsten microparticles in their designs, the USSR and Russia have also examined two other types of metals: lithium and NaK vapor. These have the advantage of being lower mass, impacting the overall propellant mass less, and also far easier to control fluid insertion rates (although microparticles can act as fluidized materials due to their small size, and maintain suspension in the H2 propellant well). This is a subject that I’ll cover in more depth in the future in the gas core NTR post.

(Side note: I’ve NEVER seen data on non-hydrogen propellant in a liquid-fueled NTR, but this problem would be somewhat ameliorated by using a higher atomic mass fuel, but which one is used will determine both how much more radiation would be directly absorbed, and what kind of loss in specific impulse would accompany this substitution. Also, using other elements/molecules would significantly change the neutronic structure and hydrodynamic behavior of the reactor, a subject I’ve never seen covered in any paper.)

Sadly, in many designs there simply isn’t the heat capacity to remove all of the reactor’s thermal energy through the propellant stream. Early gas core NTRs were especially notorious for this, with some only able to reject about 3% of the reactor’s thermal energy into the propellant. In order to prevent the reactor and pressure vessel from melting, external radiators were used – hence the large, arrowhead-shaped radiators on many gas core NTR designs.

This is unfortunate, since it directly affects the dry mass of the system, making it not only heavier but less power efficient overall. Fortunately, due to the high temperatures which need to be rejected, advanced high temperature radiators can be used (such as liquid droplet radiators, membrane radiators, or high temperature liquid metal radiators) which can reject more energy in less mass and surface area.

Another example, one which I’ve never seen discussed before (with one exception) is the use of a bimodal system. If significant amounts of heat are coming off the reactor, then it may be worth it to use a power conversion system to convert some of the heat into electricity for an electric propulsion system to back up the pure thermal system. This is something that would have to be carefully considered, for a number of reasons:

  1. It increases the complexity of the system: power conversion system, power conditioning system, thrusters, and support subsystems for each must be added, and each needs extensive reliability testing.
  2. It will significantly increase the mass of the system, so either the thrust needs to be significantly increased or the overall thrust efficiency needs to offset the additional dry mass (depending on the desire for thrust or efficiency in the system).
    1. Knock on mass increases will be extensive, with likely additions being: an additional primary heat loop, larger radiators for heat rejection, main truss restructuring and brackets, additional radiation shielding for certain radiation sensitive components, possible backup power conditioning and storage systems, and many other subsystem support structures.
  3. This concept has not been extensively studied; the only example that I’ve seen is the RD-600, which used a low power mode with an MHD that the plasma passed directly through in a closed loop system (more on this system in the future); this is obviously not the same type of system being discussed here. The only other similar parallel is with the Werka-type dusty plasma fission fragment rocket, which uses a helium-xenon Brayton turbine to provide about 100 kWe for housekeeping and system electrical power. However, this system only rejected less than 1% of the total FFRE waste heat.
    1. The proper power conversion system needs to be selected, thruster selection is in a similar position, and other systems would go through similar selection and optimization processes would need to be done. This is made more complex due to the necessity to match the PCS and thermal management of the system to the reactor, which has not been finalized and is currently very inefficient in terms of fissile material. If a heat engine is used, the quality of the heat reduces, meaning larger (and heavier) radiators are needed, as well.

Fluid Fuels: Promises of Advanced Rockets, but Many Challenges to Overcome

As we’ve seen in this brief overview of fluid fueled NTRs, the diversity in advanced NTR designs is broad, with an incredible amount of research having been done over the decades on many aspects of this incredibly promising, but challenging, propulsion technology. From the chemically challenging liquid fuel NTR, with several materials and propellant feed challenges and options, to the reliable vapor core, to the challenging but incredibly promising gas core NTR, the future of nuclear thermal propulsion is far more promising than the already-impressive solid core designs we’ve examined in the past.

Coming up on Beyond NERVA, we will examine each of these types in detail in a series of blog posts, and the information both in this post and future posts will be adapted into more-easily referenced web pages. Interspersed with this, I will be working on filling in details on the Rover series of engines and tests on the webpage, and we may also cover some additional solid core concepts that haven’t been covered yet, especially the pebble-bed designs, such as Timberwind and MITEE (the pebble-bed concept is also sometimes called a fluidized bed, since the fuel is able to move in relation to the other pellets in the fueled section of the reactor in many designs, so can be considered a hybrid system in some ways).

With the holiday season, life events, and concluding the project which has kept me from working as much as I would have liked on here in the coming months, I can’t predict when the next post (the first of three on liquid fueled NTRs) will be published, but I’ve already got 7 pages written on that post, six on the next (bubblers), and 6 on the final in that trilogy (radiator LNTR) with another 4 on vapor cores, and about 10 pages on the basic physics principles of gas core reactor physics (which is insanely complex), so hopefully these will be coming in the near future!

As ever, I look forward to your feedback, and follow me on Twitter, or join the Beyond NERVA Facebook page, for more content!

References

This is just going to be a short list of references, rather than the more extensive typical one, since I’m covering all this more in depth later… but here’s a short list of references:

Liquid fuels

“Analysis of Vaporization of Liquid Uranium, Metal, and Carbon Systems at 9000 and 10000 R,” Kaufman et al 1966 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19660025363.pdf

“A Technical Report on Conceptual Design Study of a Liquid Core Nuclear Rocket,” Nelson et al 1963 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19650026954.pdf

“Performance Potential of a Radiant Heat Transfer Liquid Core Nuclear Rocket Engine,” Ragsdale 1967 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19670030774.pdf

Vapor and Droplet Core

“Droplet Core Nuclear Reactor (DCNR),” Anghaie 1992 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19920001887.pdf

“Vapor Core Propulsion Reactors,” Diaz 1992 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19920001891.pdf

Gas Core

“Analytical Design and Performance Studies of the Nuclear Light Bulb Engine,” Rogers et al 1973 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19730003969.pdf

“Open Cycle Gas Core Nuclear Rockets,” Ragsdale 1992 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19920001890.pdf

“A Study of the Potential Feasibility of a Hybrid-Fuel Open Cycle Gas Core Nuclear Thermal Rocket,” Beveridge 2017 https://etd.iri.isu.edu/ViewSpecimen.aspx?ID=439

Categories
Electric propulsion Electrothermal Thrusters MPD Thrusters Spacecraft Concepts

Electric Propulsion Part 1: Thermal and Magnetoplasmadynamic Thrusters

Hello, and welcome back to Beyond NERVA! My apologies for the delay in this post, electric propulsion is not one of my strong points, so I spent a lot of extra time on research and in discussion with people who are more knowledgeable than I am on this subject. Special thanks to both Roland A. Gabrielli and Mikkel Haaheim for their invaluable help, not only for extensively picking their brains but also for their excellent help in editing (and sometimes rewriting large sections of) this post.

Today, we continue looking at electric propulsion, by starting to look at electrothermal and magnetoplasmadynamic (MPD) propulsion. Because there’s a fair bit of overlap, and because there are a lot of similarities in design, between these two types of thruster, we’ll start here, and then move to electrostatic thrusters in the next post.

As we saw in the last post, there are many ways to produce thrust using electricity, and many different components are shared between the different thruster types. This is most clear, though, when looking at thermal and plasma-dynamic thrusters, as we will see in this post. I’ve also made a compromise on this post’s structure: there are a few different types of thruster that fall in the gray area between thermal and MPD thrusters, but rather than writing about it between the two thruster types, one will be left for last: VASIMR, the Variable Specific Impulse Magnetoplasma Rocket. This thruster has captured the public’s imagination like few types of electric propulsion ever have; and, sadly, this has bred an incredible amount of clickbait. At the end of this post, I hope to lay to rest some of the misconceptions, and look at not only the advantages of this type of thruster, but the limitations as well. This will involve looking a little bit into mission planning and orbital mechanics, an area that we haven’t addressed much in this blog, but I hope to keep it relatively simple.

Electrothermal Propulsion

This is, to put it simply, the use of electric heaters to energize  a propellant and to produce thrust by expanding it. In the most primitive and low energy thrusters,   this is done with a Laval nozzle as in chemical and other thermal engines. This can be an inefficient use of energy (although this is definitely not always the case), however it can produce the most thrust for the same amount of power out of any of the systems that will be discussed today (debatably, depending on the systems and methods used). It is something that has been used since the 1960s, and continues to be used today for small-sat propulsion systems.

There are a number of ways to use electricity to make heat, and each of these methods are used for propulsion. We’ll look at them in turn: resistance heating, induction heating, and arc heating are all used by different designs. Each have their advantages and disadvantages, some of the concepts used for each thruster type are used in other types of thrusters also, and we’ll look at each in turn.

Resistojets

Primex resistojet, Choueiri
Primex hydrazine fueled resistojet, Chouieri

Using electricity to produce heat is something that everyone is familiar with. Space heaters and central heating are the obvious examples, but any electrical circuit produces heat due to electrical resistance within the system; this is why incandescent light bulbs and the computer that you’re reading this on get hot. This is resistive heating (or joule or Ohmic heating), and in a propulsion application this is called a “resistojet,” or an “electro-thermal thruster.” Often, this is used as a second stage for a chemical reaction – in this case the use of hydrazine propellant that undergoes catalysis to force chemical decay into a more voluminous gas. This two-stage approach is something that we’ll see again with a different heating method later in this post.

The first use of resistojets was with the Vela military satellites (first launched in 1963, canceled in 1985), which used a suite of instruments to detect nuclear tests from space, using a BE-3A AKM thruster (which I can’t find anything but the name of, if someone has documentation, please leave it in a comment below). The Intelsat-V program also used resistojet thrusters, and it has become a favored station-keeping option for smallsats. The reason for this is that the thrust is produced thermally, with no need for chemically reactive components, which is often pretty much a requirement for smallsats, which generally speaking are secondary payloads for larger spacecraft, and as such need to be absolutely safe for everything around them in order to get permission to be placed on the launch vehicle.

One of the main advantages of the resistojet is that they can achieve very high thrust efficiencies of up to 90%. Resistojets are primarily limited by two factors: first, the heat resistance of the Ohmic elements themselves; and second, the thermal transfer capacity of the system. As we have seen in NTRs, the ability to remove heat needs to be balanced with the heat produced, and the overall system needs to provide enough thrust to be useful. In the case of propelling a spacecraft on interplanetary missions, this is unlikely to come out to a useful result; however, for station-keeping with a high thrust requirement, it proves to be useful as shown in figure ## naming a few examples of satellites with EP. Exhaust velocities of about 3500 m/s are possible with decomposed hydrazine monopropellant, at about 80% efficiency. According to the ESA, specific impulse on this type of system is between 150 and 700 s of isp depending on the propellant, which is the bottom of the electric propulsion range.

Induction Thermal Thrusters

Another option for electrothermal thrusters is to use induction heating. Induction heating occurs when a high frequency alternating current is passed through a coil. Because of this, the induced magnetic field in the surroundings is swinging rapidly, stirring polar particles (particles that have a distinct plus and minus pole even if they’re electrically neutral overall) in the field. This can result even in ripping molecules apart (dissociation) and knocking electrons out of their orbitals (ionization). The charged remnants are ions and electrons, forming a plasma. Because of this, the device is called an “inductive plasma generator” or IPG. Plasma are even more susceptible to this high frequency heating. In purely thermal IPG based thrusters, Laval nozzles are used for expansion, once more, but magnetic nozzles, as explained later, can augment the performance on top of what a physical nozzle can provide. This principle is something that we’ve already seen a lot in this blog (both CFEET and NTREES operate through induction heating), and is used in one concept for bimodal nuclear thermal-electric propulsion, in the Nuclear Thermo-Electric Rocket (NTER) concept by Dr. Dujarric at ESA. This principle is also used in several different sort of electric thrusters, such as the Pulsed Induction Thruster (PIT); this is not a thermal thruster, though, so we’ll look more at it later.

This is a higher-powered system if you want significant thrust, due to the current required for the induction heater; so it’s not one that is commonly in use for smaller satellites like most of these systems. One other limitation noted by the NTER team is that supersonic induction is not an area that has been studied, and in most cases heating supersonic gasses don’t actually make them travel faster (called frozen energy), so during heating it’s necessary to make sure the propellant velocity remains subsonic.

IPG is also one of the foci of research at the Institute of Space Systems of the University of Stuttgart, studying both space and terrestrial applications, grouped in figure ## below, demonstrating the versatility of the concept: Generating a plasma without contact, the propellant cannot damage e.g. electrodes. This allows a near arbitrary selection of gasses as propellant, and therefore viable in-situ resource utilization concepts. Even space station wastes could be fed to such a thruster. Eventually, this prompted research on IPG based waste treatment for terrestrial communities. At the Institute of Space Systems, IPG also serve for the emulation of planetary atmospheres in re-entry experimentation in plasma wind tunnels.

IPG tech tree
Applications for inductive plasma generators investigated at the Insitute of Space Systems USTUTT and affiliations. Gabrielli 2018

Which type of heating is used is generally a function of the frequency of energy used to cause the oscillations, and therefore the heat. Induction heating, as we’ve discussed before in context of testing NTR fuel elements, usually occurs between 100 and 500 kHz. Radio Frequency heating occurs between 5 and 50 MHz. Finally, microwave heating occurs above 100 MHZ, although GHz operational ranges are common for many applications, like domestic microwaves which are found in most kitchens.

RF Electrothermal Thruster

RF thrusters operate via dielectric heating, where a material that has magnetic poles is oscillated rapidly in an electromagnetic field by a beam of radio waves, or more properly the molecules flip orientation relative to the field as the radio waves pass across them, causing heat to transfer to adjacent molecules. One side effect of using RF for heating is that these have very long wavelengths, meaning that the object being heated (in this case the propellant) can be heated more evenly throughout the entire mass of propellant than is typically possible in a microwave heating device.

While this is definitely a viable way to heat a propellant, this mechanism is more commonly used in ionization chambers, where the oscillating orientation of the dielectric molecules causes electrons of adjacent molecules to be stripped off, ionizing the propellant. This ionized propellant is then often accelerated using either MPD or electrostatic forces. We’ll look at that later, though, it’s just a good example of the way that many different components of these thrusters are used in different ways depending on the configuration and type of the thrusters in question.

Microwave Thermal Thrusters

 

MeT Clemens 2008
Microwave Electrothermal Thruster Diagram, Clemens 2008

Finally, we come to the last major type of electrothermal thruster: the microwave frequency thruster. This is not the Em-drive (nor will that concept be covered in this post)! Rather, it’s more akin to the microwave in your home: either radio frequencies or microwaves are used to convert the propellant, often Teflon (Polyfluorotriethylene, PFTE), into a plasma, which causes it to expand and accelerate out of a nozzle. This is most commonly done with microwaves rather than the longer wavelength radio frequencies due to a number of practical reasons.

 

Microwave thermal thrusters have been demonstrated with a very wide range of propellants, from H2 and N2 to Kr, Ar, Xe, PTFE, and others, at a wide variety of power levels. Due to the different power levels and propellant masses, specific impulse and thrust vary wildly. However, hydrogen-based thruster concepts have been demonstrated to have a specific impulse of approximately 1000 s with 54 kN of thrust.

An interesting option for this type of thruster is to not have your power supply on-board your spacecraft, and instead have a beam of microwaves hit a rectifier, or microwave antenna, which is then directed into the propellant. This has a major advantage of not having to have your power supply, electric conversion system, and microwave emitters weighing down your spacecraft. The beam will diverge over time, growing wider and requiring larger and larger collectors, but this may still end up being a major mass savings for quite a few different applications. Prof. Komurasaki at University of Tokyo is a major contributor to research in this concept, but this isn’t going to be something that we’re going to delve too deeply into in this post.

Electrothermal: What’s it Good For?

As we’ve seen, these systems aren’t much, if any, more efficient than a nuclear thermal system in terms of specific impulse, and the additional mass of the power conversion and heat rejection systems make them less attractive than a purely nuclear thermal system. So why would you use them?

There are a number of current applications, as has been mentioned in each of the concepts. They offer a fair bit of thrust for the system mass and complexity, a wide array of propellant options, and a huge range of sizes for the thrusters as well (including systems that are simply too small for a dedicated reactor for a nuclear thermal rocket).

Some designs for nuclear powered space stations (including Von Braun’s inflatable torus space station in the 1960s) use electrothermal thrusters for reaction control systems, partly due to their relatively high thrust. This could be a very attractive option, especially with chemically inert, inexpensive propellants such as PTFE that don’t require cryogenic propellants . They could also be used for orbital insertion burns, as they offer advantages toward thrust capability rather than efficiency, due to their simplicity and relatively low dry mass. For instance, an electric spacecraft on an interplanetary mission may use an electrothermal system to leave low Earth orbit on an interplanetary insertion burn, and then another drive system is used for the interplanetary portions of the mission; the drive system may be staged, discarding the now-burned-out drive system, or the propellant tankage (if any) could be discarded after use to minimize mass, and at orbital insertion at the destination the system could be activated again (this is, of course, not necessary, but in some cases may be advantageous, for instance for crewed missions, where the living beings on board don’t want to spend a couple months climbing out of Earth’s gravity well if they can avoid it).

Overall, electrical and thrust efficiency can be high, which makes these systems attractive for spacecraft. However, as a sole method of propulsion for interplanetary missions, this type of system DOES leave a lot to be desired, due to the generally low specific impulse of these types of thrusters, and in practice is not something that would be able to be used for this type of mission. Electric propulsion’s advantages for spaceflight are in high exhaust velocities, high specific impulse, and continuous use resulting in high spacecraft velocities, and thrust is generally secondary.

Arcjets – The First Middle Ground Between Thermal and MPD

These aren’t the only ways to produce heat from electricity, though. The first option we will discuss in the gray area between thermal and magnetically based propulsion, arc heating, is a very interesting option. Here, a spark, or arc, of electricity is sustained between two electrodes. This heating is virtually the same way an arc welder operates. This has the advantage that you aren’t thermally limited by your resistor for your highest thermal temperature: instead of being limited to about a thousand Kelvin by the melting point of your electric components, you can use the tens of thousands of Kelvin from the plasma of the arc – meaning more energy can be transferred, over a shorter time, within the same volume, due to the greater temperature difference. In most modern designs, the positive electrode – the anode – is at the throat of  the nozzle. However, this arc also erodes your electrodes, carrying ablated and vaporized and even plasmified bits of their material., So there’s a limitation to how long this sort of thruster can operate before the components have to be replaced. The propellant isn’t just heated by the arc itself, but also conduction and convection from the heated structural components of the thruster.

Arcjet diagram, Gabrielli
Schematic of an arcjet (Institute of Space Systems, USTUTT). F: thrust, ce: exhaust velocity,m: propellant mass flow (feed).

Arcjets have been studied by NASA since the 1950s, however they didn’t become commonly used in spacecraft until the 1990s. Several companies, including Lockheed Martin and others, offer a variety of operational arcjet thrusters. As with the resistojet, the chemical stability is excellent for piggyback payloads, and they offer better efficiencies than a resistojet. They are higher-powered systems, though; often higher than your average satellite power bus (sometimes in the kW range), necessitating custom power supplies for most spacecraft (in a nuclear-powered spacecraft, this would obviously be less of an issue). Arcjets offer much higher exhaust velocities compared to resistojets, generally 3500-5000 m/s for hydrazine decomposition drives similar to what we discussed above, and up to 9000 m/s using ammonia, and are also able to scale by both scale and power efficiently. In these systems, though, the propellant doesn’t necessarily need to be a gas or liquid at operational temperature: some thrusters have used polytetrafluoroethelene (PFTE, Teflon) as a propellant.

This type of propellant is also very common in a sub-type of arcjet thruster, one that doesn’t use an internal cathode: the pulsed plasma thruster. Here, the PFTE propellant block is brought into contact with a cathode on one side of the thruster, and an anode on the other. The electric charge arcs across the gap, vaporizing the propellant, and pushing the propellant back slightly, The arc and resulting plasma continue to the end of the thrust chamber, and then the propellant (usually loaded on a spring mechanism) is brought back to the point that the propellant can be vaporized. This type of thruster is very common for small spacecraft, since it’s incredibly simple and compact, although certain designs can have engineering challenges with the spring mechanism and the lifetime of the cathode and anode.

Arcjets can also be combined with magnetoplasmadynamic thruster acceleration chambers, since arc heating is also a good way to create plasma. The pulsed plasma thruster often uses electric arcs for charging their propellant, for instance. This mechanism is also used in magnetoplasmadynamic (MPD) thrusters, which is why we haven’t placed them with the rest of the thermal thrusters.

In fact, there’s more in common between an arcjet and an MPD thruster than between other thermal designs. The cathode and anode of an arcjet are placed in exactly the same configuration as most designs for an MPD (with some exceptions, to be fair). The exhaust itself is not only vaporized, but ionized, which – like with the RF or MW thrusters – lends itself to adding electromagnetic acceleration.

Self Field MPD
Comparison of self-field MPD and arcjet thruster geometry. Top half of the diagram is MPD, bottom half arcjet, in same scale. (Institute of Space Systems, USTUTT)

VASIMR: the VAriable Specific Impulse Magnetoplasma Rocket

Coauthor Mikkel Haaheim

VASIMR 3d image Bering et al 2014
VASIMR 3d diagram, Bering et al 2014

As mentioned earlier, the difference between MPD and thermal thrusters is a very gray area, and, even more than our previous examples, the VASIMR engine shows how gray this area is: the propellant is plasma of various types (although most designing and testing have focused on argon and xenon, other propellants could be used). This propellant is first ionized in what’s typically a separate chamber, and then fed into a magnetically confined chamber, with RF heating. This then is accelerated, and that thrust is then directed out of a magnetic nozzle.

VASIMR is the stuff of clickbait. Between the various space and futurism groups I’m active in on Facebook, I see terribly written, poorly understood, and factually incorrect articles on this design. This has led me to avoid the concept for a long time, and also puts a lot of pressure on me to get the thruster’s design and details right.

VASIMR isn’t that different from many types of electrothermal thrusters; after all, the primary method of accelerating the propellant, imparting thrust, and determining the specific impulse of the thruster is electrically generated RF heating. The fact that the propellant is a plasma also isn’t just an MPD concept, after all: the pulsed plasma thruster, and in fact most arcjets, produce plasma as their propellants as well. This thruster really demonstrates the gray area between thermal and MPD thrusters in ways that are unique in electric propulsion.

VASIMR Schematic Bering et al 2014
VASIMR system sketch, Bering et al 2014

Since the characteristics of plasma did not play a vital role in the working principles of the previous thermo-electric thrusters, we should briefly discuss the concept. The energy of plasma is so high that electrons are no longer tied to their atoms, which then become ions. Both electrons and ions are charged particles whizzing around in a shared cloud, the plasma. Despite being neutral to the outward due to containing the same amount of negative as of positive charges, the plasma is interacting with magnetic fields. These interactions present themselves for various magnetohydrodynamic applications, ranging from power generators in terrestrial power plants over magnetic plasma bottles to propulsion.

In VASIMR, these forces push the hot plasma away from the walls, protecting both the walls from damaging heat loads and the plasma from crippling quenching. This allows VASIMR to have a very hot medium for expansion. While this puts VASIMR among MHD thrusters, it would not yet be a genuine “plasma thruster,” if it was not for the magnetic nozzle, which adds electromagnetic components to the forces generating the thrust. Among these components, the most important is the Lorentz-force, which occurs when a charged particle moves through a magnetic field. The Lorentz-force is orthogonal to the local magnetic field line as well as to the particle’s trajectory.

Despite the incredible amount of overblown hype, VASIMR is an incredibly interesting design. Dr. Franklin Chang Diaz, the founder of Ad Astra Rocket Company, became interested in the concept while he was a PhD candidate at MIT for applied plasma physics. Before he was able to pursue this concept, though, his obsession with space led him to become an astronaut, and he flew seven times on the Space Transport System (Space Shuttle), spending a total of over 66 days on orbit. Upon retiring from NASA, he founded the Ad Astra Rocket Company to pursue the idea that had captured his imagination during his doctoral thesis; refined by his understanding of aerospace engineering and propulsion from his time at NASA. Ad Astra continues to develop the VASIMR thruster, and consistently meets its deadlines and budgetary requirements (as well as the modeling expectations of Ad Astra), but the concept is complex in application, and as with everything in aerospace, the development process takes a long time to come to fruition.

VX-200 Prototype
Ad Astra VX-200 Prototype, image courtesy Ad Astra

After the end of several rounds of funding from NASA, and a series of successful tests of their VX-100 prototype, Ad Astra continued to develop the thruster privately. Their newer VX-200 thruster is designed for higher power, and with better optimization of several of its components. Following additional testing, the engine is currently going through another round of upgrades to prepare for a 100-hour test firing of the thruster. Ad Astra has been criticized for its development schedule, and the problems that they face are indeed significant, but so far they’ve managed to meet every target that they’ve set.

The main advantage of this concept is that it eliminates both friction and erosion between the propellant and the body of the thruster. This also reduces the thermal load on the thruster, because, since there’s no physical contact, conduction can’t occur, and the amount of heat that’s absorbed by the thruster is limited to radiation (which is limited by the surface area of the plasma and the temperature difference between that plasma and the thruster body). This doesn’t mean that there’s not a need to cool the thruster in most cases, it does mean that more heat is kept within the plasma, and in fact, by using regenerative cooling (as most modern chemical engines do) it’s possible to increase the efficiency of the thruster.

Another major advantage, and one that may be unique to VASIMR, is the first part of the acronym: VAriable Specific Impulse. Every staged rocket has variable specific impulse, in a way: most first-stage boosters have very low specific impulse compared to the upper stages (although, in the case of the boosters, this is due to both the atmospheric pressure and the need to impart a large amount of thrust over a limited timespan), and there are designs that use different propulsion systems with different specific impulse and thrust characteristics to optimize their usefulness for particular mission profiles (such as bimodal thermal-electric nuclear rockets, the subject of our next blog series after our look at electric propulsion), but VASIMR offers the ability to vary its’ exhaust velocity by changing the temperature it heats the propellant to. This, in turn, changes the specific impulse, and therefore its’ thrust. This is where the “30 Day Round Trip to Mars” clickbait headlines come into play: by continuously varying its’ thrust and isp depending on where it is in terms of the interplanetary transfer maneuver, VASIMR is able to optimize the trip time in ways that few, if any, other contemporary propulsion types can. However, the trip time is highly dependent on available power, and trip times on the order of 90 days require a power source of 200 MW, and the specific power of the system becomes a major concern. To explain this in detail gets into orbital mechanics far more deeply than I would like in this already very long blog post, so we’ll save that discussion for another time.

So how does VASIMR actually work, what are the requirements for efficient operation, and how does it have these highly unusual capabilities? In many ways, this is very similar to a typical RF thruster: a gas, usually argon, is injected into a quartz plenum, and then run through a helicon RF emitter. Because of the shape of the radio waves produced, this causes a cascading ionization effect within the gas, converting it into a plasma, but the electrons aren’t removed, like in the case of more familiar electrostatic thrusters (the focus of our next blog post). This excitation also heats the plasma to about 5800K. The plasma then moves to a second RF emitter, designed to heat the plasma further using an ion cyclotron emitter. This type of RF emitter efficiently heats the plasma to the desired temperature, which is then directed out of the back of the thruster. Because all of this is occurring at very high temperatures, the entire thruster is wrapped in superconducting electromagnets to contain the plasma away from the walls of the thruster, and the nozzle used to direct the thrust is magnetic as well. Because there are no components physically in contact with the plasma after it becomes ionized, there are no erosion wear points within the thruster, which extends the lifetime of the system. By varying the amount of gas that is fed into the system while maintaining the same power level, the gas will become ionized to different levels, and the amount of heating that occurs will be different, meaning that the exhaust velocity will be higher, increasing the specific impulse of the engine while reducing the thrust. This is perhaps the most interesting part of this propulsion concept, and the reason that it gets so much attention. Other systems that use pulsed thrust rather than steady state are able to vary the thrust level without changing the isp (such as the pulsed induction thruster, or PIT) by changing the pulse rate of the system, but these systems have limits as to how much the pulse rate can be varied. We’ll look at these differences more in a later blog post, though.

 

Thrust Efficiency charts, Bering et al 2014
Thrust efficiency vs RF power and isp, Bering et al 2014

Many studies have looked at the thrust efficiency of the VASIMR. Like many electric propulsion concepts, it becomes more efficient as more power is applied to the system; in addition, the higher the specific impulse being used, the more efficiently it uses the electrical power available. The current VX-200 prototype is a 212 kW input, 120 kW thrust system, far more powerful than the original VX-10, and as such is more efficient. Most estimates of average efficiency seem to suggest a minimum of 60% thrust efficiency (the amount of efficiency increases with power input), increasing to 90% for higher-isp functioning. However, given the system’s sensitivity to available power level, and the fact that it’s not clear what the final flight thruster’s power availability will be, it’s difficult to guess what a flight system’s thrust efficiency will be.

 

VASIMR is currently upgrading their VX-200 thruster for extended operations. As of this point, problems with cooling various components (such as the superconducting electromagnets) have led to shorter lifetimes than are theoretically possible, although to be fair the lack of cooling problems come down to cooling systems not being installed. Additionally, more optimization is being done on the magnetic nozzle. One of the challenges with using a magnetic nozzle is that the plasma doesn’t want to “unstick” from the magnetic field lines used to contain the propellant. While this isn’t a major challenge for the thruster the way that the thermal management problems are, it is a source of inefficiency in the system, and so is worth addressing.

There’s a lot more that we could go into on VASIMR, and in the future we will come back to this concept; but, for the purposes of this article, it’s a wonderful example of how gray the area between thermal and MPD thrusters are: the propellant ionization and magnetic confinement of the heated plasma are both virtually identical to the applied field MPD thruster (more on that below), but the heating mechanism and thrust production are virtually identical to an RF thruster.

Let’s go ahead and look at what happens if you use magnetic fields instead of heat to accelerate your propellant, but keep most of the systems we’ve described identical in function: the magnetoplasmadynamic thruster.

 

Magnetoplasmadynamic Thrusters

Coauthor: Roland A. Gabrielli, IRS

NASA MPD concept
Self-field MPD Thruster Concept, image courtesy NASA

Magnetoplasmadynamic thrusters are a high-performance electric propulsion concept; and, as such, offer greater thrust potential than the electrostatic thrusters that we’ll look at in the next blog post. They also tend to have higher power requirements. Therefore they have not been used as a dedicated thruster on operational spacecraft to date, although they’ve been researched since the 1960s in the USSR, the USA, Western Germany, Italy, and Japan. Only a few demonstrators have flown on both Russian and Japanese experimental satellites. They remain an attractive and cost efficient option for high-thrust electric propulsion including Mars transfer engines.

So far in this article, we have discussed electric thrusters which are principally thermal thrusters: The propellant runs into a reaction chamber, tanks heat and expands through a nozzle. This is as true for VASIMR, resistojets, thrusters based on inductive plasma generators, and also for arcjets. Yet, VASIMR introduces a different set of physics for thrusters, magnetohydrodynamics (MHD). This term designates the harnessing of fluids (hence ‘hydro’) with the forces (hence ‘dynamic’) emerging from magnetic fields (hence ‘magneto’). In order to effectively use magnetic forces, the fluid has to be susceptible to them, and its particles should somehow be electrically polar or even charged. The latter case, plasma, is the most common in electric propulsion.

Since the characteristics of plasma did not play a vital role in the working principles of the previous thermo-electric thrusters, we should briefly discuss the concept. The energy of plasma is so high that electrons are no longer tied to their atoms, which then become ions. Both electrons and ions are charged particles whizzing around in a shared cloud, the plasma. Despite being neutral to the outward due to containing the same amount of negative as of positive charges, the plasma interacts with magnetic fields. These magnetohydrodynamic interactions present themselves for various applications, ranging from power generators in terrestrial power plants over magnetic plasma bottles to propulsion.

In VASIMR, these forces push the hot plasma away from the walls, protecting both the walls from damaging heat loads and the plasma from cooling so rapidly that thrust is lost. This allows VASIMR to have a very hot medium for expansion. While this puts VASIMR among MHD thrusters, it would not yet be a genuine “plasma thruster,” if it was not for the magnetic nozzle, which adds electromagnetic components to the forces generating the thrust. Among these components, the most important is the Lorentz-force, which occurs when a charged particle moves through a magnetic field. The Lorentz-force is at right angles to both the local magnetic field line as well as to the particle’s trajectory.

There are two main characteristics of an MPD thruster:

  1. The plasma constitutes a substantial part of the medium, which imparts a significant integral Lorentz force,
  2. the integral Lorentz force has a relevant contribution towards the exhaust direction.

The electromagnetic contribution is the real distinction from the previous thermal approaches, as the kinetic energy of the jet is not only gained from undirected heating, but also from a very directed acceleration. The greater the electric discharge, and the more powerful the magnetic field, the more the propellant is accelerated, therefore the more the exhaust velocity is increased. Besides the Lorentz force, there are also minor electromagnetic effects, like a “swirl” and Hall acceleration (which will be looked at in the next blog post), but the defining electromagnetic contribution of MPD thrusters is the Lorentz force. Since the latter applies on a plasma, this type of thrusters is called magnetoplasmadynamic (MPD) thrusters.

The Lorentz force contribution is also the way magnetic nozzles work:the forces involved can be broken into three parts: along the thruster axis, toward the thruster axis, and at right angle to both of these around the axis. The first part adds to the thrust, the second pushes the plasma towards to the centre, the third generates a swirling effect, both contributing to the thrust as to spreading the arc into a radial symmetry.

 

X16 Plasma
X-16 Plasma plume with argon propellant, ISS USTUTT

There are various ways to build MPD thrusters, with differing different propellants, geometries, methods of plasma and magnetic field generation, and operation regimes, stationary or pulsed. The actual design depends mostly on the available power. The core of the most common architecture for stationary MPD thrusters is the arc plasma generator, which makes MPD thrusters seem fairly similar to thermal arcjets. But that’s just in appearance, as you can in fact build MPD thrusters almost completely free of a thermal contribution to the thrust, as evidenced by the German Aerospace Center’s (DLR) X-16, or the PEGASUS thruster that we’ll look at later in this post.

 

These types of thruster (technically known as stationary MPD thrusters with arc generation) differ most noticeably is the way the magnetic field is generated:

  • Applied-field (AF) MPD thrusters, endowed with either and torus of permanent magnets, or a Helmholtz coil placed around the jet.
  • Self-field (SF) MPD thrusters, which generate their magnetic field by induction around the current travelling in the arc.

Note that arc generator based AF-MPD thrusters also experience (to a minor extent) self-field effects. A schematic of an SF-MPD thruster is shown below, illustrating the conceptual differences between arcjets, and MPD thrusters (the top half is an MPD, the bottom half is an arcjet, note the difference in throst length and nozzle size). The most crucial difference is the contact of the arc with the anode. While a very long arc is undesirable in arcjets (for very important design reasons which we don’t have time to go into here), in the MPD thruster this is crucial to provide the thruster with sufficient Lorentz force. Moreover, the longer the oblique leg of the arc, the more of the Lorentz force will point out of the thruster. This effect means that relatively large anode diameters are the norm with MPD thrusters of this type. Therefore, simple arcjets and other electro-thermal thrusters tend to be more slender than most arc based MPD thrusters. The anode diameter may however not be too large, as the arc will be more resistive with increasing lengths, entailing more and more energy losses.

Self Field MPD
Schematic opposing an SF-MPD thruster above the dash-dotted axis to a simple Arcjet below (Institute of Space Systems, USTUTT). F: thrust, ce: exhaust velocity,m: propellant mass flow (feed). Note how the arc pushes out far of the nozzle exit. The dashed lines j indicate the current in the MPD thruster’s arc, and the fat lines B the induced magnetic field. Its circle would be within a plane vertical to the thruster axis. The thin arrows show the local direction of both lines. It is to these arrows that the Lorentz-force FLor is at right angles.

In stationary arc based MPD thrusters, the choice of propellant is mainly dictated by the ease of ionisation which tends to be more important than molar mass, which is what causes the preference for hydrogen in thermal thrusters. This shift is the more pronounced, the more the Lorentz-force contribution outweighs the thermal contribution. Consequently, many arc based MPD experiments are run with noble gasses, like Helium, or Neon; while Xenon is often discussed in pure development, it is rarely considered for missions due to its cost. The most important noble gas for MPD is thus Argon. Other easily ionised substances are liquid alkali metals, commonly Lithium, which enables very good efficiencies. However, the complicated propellant feed system and the risk of depositions in that case is a serious drawback. Nevertheless, there is still a very large field for hydrogen or ammonia as propellants.

The major lifetime restricting component in arc based MPD are the cathode and – to a minor extent – anode lifetimes. These will erode over time which is caused by the plasma arc. The arc will gnaw at the metals due to electron emission, sublimation and other mechanisms. Depending on the quality of the design and the material, this will be significant after a few hundred hours of operation, or a tens of thousands. Extending their life is a challenge, because the plasma behavior will change depending on a number of factors determined by the plasma and the system in question. To add to the complexity, the geometry of the electrodes, affected by the erosion, is one of them. Because of this, some designs have easily replaceable cathodes, others (like the Pegasus which we’ll cover below) just swap out the drive: the original design for the SEI that Pegasus was proposed for actually had seven thrusters on board, run in series as the cathode wore out on each one.

AF-MPD – The Lower-Power Option

 

Japanese AF-MPD, permanent magnet
Japanese AF-MPD concept with 0.1 T permanent magnet

Depending on the available power, the arc current in MPD may or may not be intense enough to induce a significant magnetic self-field. At the lower end of the power scale, this is definitely breaking the MPD thruster principle. Because of this, in lower powered systems an external magnet  is required to create the magnetic field, which is why it’s called an applied field MPD. In general, these systems  range from 50 to 500 kW of electric power, although this is far from a hard limit. The advantage of applied field MPD thrusters over a self-field types (more on the self-field later) is that the magnetic fields can be manipulated independent of the amount of charge running through the cathode and anode, which can mean longer component life for the thruster. There are two main approaches to provide for an external field: The first is a ring of a permanent magnet around the volume occupied by the arc; the second is the placement of a Helmholtz coil instead (an electromagnet whose coil wraps around the lengthwise axis of the thruster, sometimes using superconductors). At the lower ending of the power range, the permanent magnet may be the better option because it doesn’t consume what little electricity you have, while the electromagnets are more interesting at the upper end.

 

All these solutions do require cooling, and the requirements are important the more powerful the magnet is. This cooling can be achieved passively at the lower ending of the power range (given enough free volume). For mid level power, the cold propellant itself can provide the cooling prior to running alongside the hot anode and entering the plasma generator. Using cold propellant for cooling the thrusters is called regenerative cooling (a mainstay of current chemical and nuclear thermal engines). The most performant magnets for AF-MPD, superconducting coils, must be brought to really low temperatures, and this tends to require an additional, secondary coolant cycle, including an own refrigeration system, with pumps, compressors, and radiators.

 

ISS SX-3 AF-MPD Helmholtz coil
Recent development at the Institute of Space Systems, Stuttgart: SX 3 prior experimentation. The outer flange covers a Helmholtz coil.

The nice thing about the electromagnets is that it’s possible to tune the strength of the field in a certain range. If the coil degrades over time, more electricity (and coolant, due to increased electrical resistance) can be pumped through. This isn’t an option for a permanent magnet. However, the magnetic field generation equipment is one of the lifetime limiting components of this type of thruster, so it’s worth considering.

 

There’s not really a limit to how much power you use in an applied field MPD thruster, and especially with a Helmholtz coil you can theoretically tune your drive system in a number of interesting ways, like increasing the strength to constrict the plasma more if there’s a lower-mass stream. Something happens once the plasma has enough charge going through it, though: the unavoidable self field contribution increases. .  Besides increasing the complexity of the determination of the field topology, the self field is an advantage. At sufficient power, you can get away without coil or magnets, making the system lighter, simpler, and less temperature-sensitive. This is why most very high powered systems use the self-field type of MPD.

Before we look at this concept in the next subsection, let us have a look at current developments from over the world. Table ## summarises a few interesting AF-MPD thrusters, both the performance parameters thrust F, exhaust velocity c_e, thrust efficiency η_T, electric (feed) power P_e and jet power P_T, and design,  like anode radius r_A, cathode radius r_C, arc current I, magnetic field B and the propellant. Recent AF-MPD-thruster development was conducted by Myers in the USA, by MAI (the Moscow Aviation Institute) in Russia, at the University of Tokyo in Japan, and SX 3 in Germany at the Institute of Space Systems, Stuttgart. The types X 9 and X 16 in table ## are the IRS’ legacy from the German Aerospace Center (X 9, X 16).

Thruster Pro-
pellant
r_A / mm r_C / mm I / A B / T F / mN c_e/ km/s η_T / % P_e / kW P_T / kW
Myers Ar 25 6.4 1000 0.12 1400 14 22 44.5 9.8
MAI Li 80 22.5 1800 0.09 2720 33.6 44.1 103.5 45.7
U Tokyo H2 40? 4 200 0.1 50 55.6 19.3 7.2 1.4
SX 3 Ar 43 6 450 0.4 2270 37.9 58 74 42.9
X 16 Ar 20 3 80 0.6 251 35.9 38.8 11.6 4.5
X 16 Xe 20 3 80 0.6 226 25.1 29.6 9.6 2.84
X 9 Ar 20 5 1200 0.17 2500 20.8 28.1 93 26.1

Design parameters and experimental performance data from various AF-MPD thrusters from over the world. Gabrielli 2018

SX-3 plume (argon propellant), IRS
Visual plume of SX 3 at the Institute of Space Systems, Stuttgart. 

Russian 100 kWe AF-MPD
Russian 100 kWe lithium AF-MPD thruster.

Self-Field MPD: When Power Isn’t a Problem

In the previous section, we looked at low and medium powered MPD thrusters. At those power levels, an external field had to be applied to ensure a powerful enough magnetic field is applied to the plasma to generate the Lorentz force. Even though it wasn’t enough to impart enough thrust, there was always a self field contribution, albeit a weak, almost negligible one. The cause of the self field contribution is the induction of a magnetic field around the arc due to the current carried. You can get an idea of the direction of a magnetic field with the “right fist rule” by closing your right fist around the generating current, with your thumb pointing towards the cathode. Your fingers will then curl in the direction of the magnetic field. To get the direction of the Lorentz force, all you have to do in the next step is aligning your right hand again. This time, your thumb has to point in the direction of the magnetic field, and – at right angle – your index finger into the direction of the current. At right angle to both fingers, the middle finger will point in the direction of the Lorentz-force. (Note that you can also use the latter three-fingers-rule to study the acceleration in AF-MPD thrusters.)

The strength of the induced self field will depend on the current. The stronger the current is, the stronger the magnetic field will be, and, in turn, the Lorentz acceleration. As a consequence, given a sufficient current, the self field will be effective enough to provide for a decent Lorentz acceleration.

The current depends on the available electric power put into the arc generator, making the applied field obsolete from certain power levels up. This reduces complications arising from using an external magnet, and provides good efficiencies and attractive performance parameters. For example, at 300 kWe, and with an arc current of almost 5 kA (compare to AF-MPDs currents ranging from 50 A to 2 kA) DT2, an SF-MPD thruster developed at the Institute of Space Systems in Stuttgart, can provide a thrust of approximately 10 N at an exhaust velocity of 12 km/s, with a thrust efficiency of 20%. The performance possibilities have many people considering the technology as a key technology for rapid, man rated interplanetary transport, in particular to Mars. In this use case, SF-MPD thrusters may even be competitive with VASIMR, weighing possible shortcomings in efficiency up with a significantly simpler construction and, hence, much smaller cost. However, lacking current astronuclear sources of sufficient power, the development is stagnant, and awaiting disruption on the power source side.

DT-2 Plume Argon
DT 2 in operation at the Institute of space systems, Stuttgart.

 

MPD Cutaway high res loose pin
Simplified model of DT 2. ISS USTUTT design, image BeyondNERVA

Another example of a “typical” self-field high powered MPD thruster application (since, like all types of electric propulsion, the amount of power applied to the thruster defines the operational parameters) is that seen in the PEGASUS drive, an electric propulsion system developed for the Space Exploration Initiative (SEI) for an electric propulsion mission to Mars. Committed research on this concept began in the mid-1980s, and was meant for a mission in the late-1990s to early 2000s, but funding for SEI was canceled, and the development has been on hold  ever since. Perhaps the most notable is the shape, which is fairly typical of nozzles designed for a concept we discussed briefly earlier in the post: the sinuous curvature of the nozzle profile is designed to minimize the amount of thermal heating that occurs within the plasma, so if a nozzle has this shape it means that the thermal contribution to the thrust is not only not needed, but is detrimental to the performance of the thruster.

 

Thruster Components
PEGASUS drive system schematic, Coomes et al 1993

 

A number of novel technologies were used in this design, and as such we’ll look at it again a couple of times during this series: first for the thruster, then for its power conversion system, and finally for its heat rejection system.

Nozzle xsection
PEGASUS MPD Thruster, Coomes et al

Pulsed Inductive Thrusters

Pulsed inductive thrusters (PIT) are a type of thruster that has many advantages over other MPD thrusters. The thrusters don’t need an electrode, which is one of the major causes of wear in most thrusters, and they also are able to maintain their specific impulse over a wide range of power levels. This is because the thruster isn’t a steady-state thruster, like many other forms of thruster that are commonly in use; instead a gaseous propellant is sprayed in brief jets onto a flat induction coil, which is then discharged for a very brief period from a bank of capacitors (usually in the nanosecond range), causing the gas to become ionized and then accelerated through the Lorentz force. The frequency of pulses is dependent on the time it takes to charge the capacitors, so the more power that is available, the faster the pulses can be discharged. This directly affects the amount of thrust that’s available from the thruster, but since the discharges and volume of gas are all the same, the Lorentz force applied – and therefore the exhaust velocity of the propellant and the isp – remain the same. Another advantage of the inductive plasma generation is the wide variety of propellants available, from water to ammonia to hydrazine, making it attractive for possible in-situ propellant use with minimal processing. In fact, one proposal by Kurt Polzin at Marshall SFC uses the Martian atmosphere for propellant, making refueling a Mars-bound interplanetary spacecraft a much easier proposition.

PIT Schematic, NuPIT
Schematic of PIT operation. Image on left is gas flow, image on right is magnetic fields. Frisbee, 2005

This gives a lot of flexibility to the system, especially for interplanetary missions, because additional thrust has distinct advantages when escaping a gravity well (such as Earth orbit), or orbital capture, but isn’t necessary for the “cruise” phase of interplanetary missions. Another nice thing about it is, for missions that are power-constrained, many thruster types have variation in specific impulse, and therefore the amount of propellant needed for the mission, depending on the amount of power available for propulsion when combined with other electricity requirements, like sensors and communications. For the PIT, this just means less thrust per unit time, while the isp remains the same. This isn’t necessarily a major advantage in all mission types, but for some it could be a significant draw.

PIT was one of the proposed propulsion types for Project Prometheus (which ended up using the HiPEP system that we’ll discuss in the next blog post), known as NuPIT. This thruster offered thrust efficiency of greater than 70%, and an isp of between 2,000-9,000 seconds, depending on the specific design that was decided upon (the isp would remain constant for whatever value was selected), using a 200 kWe nuclear power plant (which is on the lower end of what a crewed NEP mission would use), with ammonia propellant. Other propellants could have been selected, but they would have affected the performance of the thruster in different ways. An advantage to the PIT, though, is that its breadth of propellant options are far wider than most other thruster types, even thermal rockets, because if there’s chemical dissociation (which occurs to a certain degree in most propellants), anything that would become a solid doesn’t really have a surface to deposit onto effectively, and what little residue builds up is on a flat surface that doesn’t rely on thermal conductance or orifice size for its’ functionality, it’s just a plate to hold the inductive coil.

NuPIT Characteristics, Frisbee 2003
NuPIT pulsed inductive thruster characteristics, Frisbee 2005

For a “live off the land” approach to propellant, PIT thrusters offer many advantages in their flexibility (assuming replacement of the gaseous diffuser used for the gas pulses), predictable (and fairly high) specific impulse, and variable thrust. This makes them incredibly attractive for many mission types. As higher powered electrical systems become available, they may become a popular option for many mission applications.

We’ll return to PIT thrusters in a future post, to explore the implications of the variable thrust levels on mission planning, because that’s a very different topic than just propulsion mechanics. It does open fascinating possibilities for unique mission profiles, though, in some ways very similar to the VASIMR drive.

More to follow!

Thermo-electric and MPD thrusters cover a wide range of low to high power thruster options, and offer many unique capabilities for mission planners and spacecraft architects. With the future availability of dense, high-powered fission power systems for spacecraft, these systems may show that they offer unique capabilities, not just for short missions or reaction control systems, but also for interplanetary missions as well. Some will need to wait until these power sources are available to be used, but others are already in use on operational satellites, and have shown dozens of years of efficient and effective operation.

The next post will complete our look at electric propulsion systems with a look at electrostatic thrusters, including gridded ion drives, Hall effect thrusters, and other forms of electric propulsion that use differences in electric potential to accelerate an ionized propellant. These have been in use for a long time, and are far more familiar to many people, but there are some incredible designs that have yet to be flown that extend the capabilities of these systems beyond even the very efficient systems in use today.

Another thank you to Roland Gabrielli and Mikkel Haaheim for their invaluable help on this blog post. Without them, this wouldn’t be able to be nearly as comprehensive or accurate as it is.

Again, I apologize that this blog post has taken so long. Unfortunately, I reached the point that I typically decide to split one blog post into two several times in this post, and actually DID split it a couple times. Much of this information, as well as a lot of it from the next post on electrostatic thrusters, was originally going to be part of the last post, but once that one reached 25 pages in length I decided to split it between history and summary, and this occurred once again in writing this post, separating the electrostatic thrusters from the thermal and MPD thrusters. The latter two concepts also almost got their own blog posts, but as we’ve seen, the two share key features and  so it made sense to keep them together. The electrostatic thruster post is already coming along well, and I hope that it won’t take as long for me to write as this one did… sadly, I can’t promise that, but I’m trying.

Sources

Electrothermal

An Analysis of Current Propulsion Systems, Weebly website http://currentpropulsionsystems.weebly.com/electrothermal-propulsion-systems.html

Resistojet

Vela spacecraft, Gunters’ Space Page profile https://space.skyrocket.de/doc_sdat/vela.htm

Alta Space Systems Resistojet page: https://web.archive.org/web/20130604101644/http://www.alta-space.com/index.php?page=resistojet

Induction Thermal Thruster

 

Microwave/RF Thermal Thruster

Coaxial Microwave Electrothermal Thruster Performance in Hydrogen, Richardson et al, Michigan State 1968 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19950005171.pdf

Microwave Electrothermal Thruster webpage, Princeton http://alfven.princeton.edu/research/past/met

The Microwave Thermal Thruster Concept, Parkin et al, CalTech https://authors.library.caltech.edu/3304/1/PARaipcp04b.pdf

Microwave Electro-thermal Thruster patent, Rayethon 1999 https://patents.google.com/patent/US5956938

Fourth Symposium on Beamed Energy Propulsion, ed. Komurasaki and Yabe, 2005 https://sciencedocbox.com/Physics/70705799-Beamed-energy-propulsion.html

Arcjet

Arc-Jet Thruster for Space Propulsion, Wallner et al 1965 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19650017046.pdf

Aerojet MR-510 hydrazine arcjet page, Astronautix: http://www.astronautix.com/m/mr-510.html

University of Stuttgart ISS PTFE PFTE Arcjet Mission concept webpage: https://web.archive.org/web/20140318021932/

http://www.elringklinger.de/en/germany-land-of-ideas-elringklinger-drives-satellite

VASIMR

High Power Electric Propulsion with VASIMR Technology, Chiand-Diaz et al 2016

http://www.unoosa.org/documents/pdf/psa/hsti/CostaRica2016/2-4.pdf

VX-200 Magnetoplasma Thruter Performance Results Exceeding 50% Thruster Efficiency, Longmier et al 2011 https://www.researchgate.net/publication/228977378_VX-200_Magnetoplasma_Thruster_Performance_Results_Exceeding_Fifty-Percent_Thruster_Efficiency

Improved Efficiency and Throttling Range of the VX-200 Magnetoplasma Thruster, Longmier et al 2014 http://www.adastrarocket.com/Ben-JPP-2014.pdf

Low Thrust Trajectory Analysis (A Survey of Missions using VASIMR For Flexible Space Exploration, Ilin et al 2012 http://www.adastrarocket.com/VASIMR_for_flexible_space_exploration-2012.pdf

Nuclear Electric Propulsion Mission Scenarios using VASIMR, Chiang-Diaz et al 2012 https://www.lpi.usra.edu/meetings/nets2012/pdf/3091.pdf

MPD

Steady State MPD

Magnetic Nozzle Design for High-Power MPD Thrusters, Hoyt Tethers, Unlimited 2005 http://www.tethers.com/papers/IEPC05_HoytNozzlePaper.pdf

Applied Field MPD

Applied-Field MPD Thruster Geometry Effects, Myers Sverdup 1991 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19910017903.pdf

Performance of an Applied Field MPD Thruster, Paganucci et al 2001 http://erps.spacegrant.org/uploads/images/images/iepc_articledownload_1988-2007/2001index/2002iepc/papers/t12/132_2.pdf

Mayer, T., Gabrielli, R. A., Boxberger, A., Herdrich, G., and Petkow, D.,: “Development of Analytical Scaling Models for Applied Field Magnetoplasmadynamic Thrusters,” 64th International  Astronautical Congress, International Astronautical Federation, Beijing, September 2013.

Myers, R. M., “Geometric Scaling of Applied-Field Magnetoplasmadynamic Thrusters,” Journal of Propulsion and Power, Vol. 11, No. 2, 1995, pages 343–350.

Tikhonov, V. B., Semenikhin S. A., Brophy J.R., and Polk J.E., “Performance of 130 kW MPD Thruster with an External Magnetic Field and Li as a Propellant”, International Electric Propulsion Conference, IEPC 97-117, Cleveland, Ohio, 1997, pp. 728-733.

Boxberger, A., et al.. “Experimental Investigation of Steady-State Applied-Field Magnetoplasmadynamic Thrusters at Institute of Space Systems”, 48th AIAA/ASME/SAE/ASEE Joint Propulsion Conference & Exhibit, Atlanta, Georgia, 2012.

Boxberger, A., and G. Herdrich. “Integral Measurements of 100 kW Class Steady State Applied-Field Magnetoplasmadynamic Thruster SX3 and Perspectives of AF-MPD Technology.” 35th International Electric Propulsion Conference. 2017.

Pegasus Drive

The Pegasus Drive: A Nuclear Electric Propulsion System for the Space Exploration Initiative; Coomes and Dagle, PNL 1990 https://www.osti.gov/servlets/purl/6399282

A Low-Alpha Nuclear Electric Propulsion System for Lunar and Mars Missions; Coomes and Dagle, PNL 1992 https://www.osti.gov/servlets/purl/10116111

MPD Thruster Performance Analysis Models; Gilland and Johnson NASA GRC, 2007 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20070032052.pdf

Self-Field MPD

On the Thrust of Self-Field MPD Thrusters, Choueiri 1997 https://alfven.princeton.edu/publications/choueiri-iepc-1997-121

Pulsed Inductive Thruster (PIT)

The PIT Mark V Pulsed Inductive Thruster, Dailey et al 1993 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19930023164.pdf

The Nuclear Electric Pulsed Inductive Thruster (NuPIT) Mission Analysis for Prometheus, Frisbee et al 2005 https://trs.jpl.nasa.gov/bitstream/handle/2014/38357/05-1846.pdf

Pulsed Inductive Thruster Using Martian Atmosphere as Propellant, Polzin 2012 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/20120015307.pdf