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.
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)
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.
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.
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.
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
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.
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:
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.
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.
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.
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.
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!
If you would like to support my work, consider becoming a Patreon supporter at https://www.patreon.com/beyondnerva . Not only do you get early access to blog posts, but I post extra blogs, images from the 3d models I’m working on of both spacecraft and reactors, and more! Every bit helps.
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.
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
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:
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
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
The behavior of the hydrogen as it passes through, and also out of, the liquid fuel was not well understood in practice, and
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
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:
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).
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.
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.
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.
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:
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.
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.
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,
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!
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.
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.
Hello, and welcome back to Beyond NERVA! Today we continue our look into advanced NTR fuel types, by diving into an extended look at one of the least covered design types in this field: the liquid fueled NTR (LNTR).
This is a complex field, with many challenges unique to the phase state of the fuel, so while I was planning on making this a single-part series, now there’s three posts! This first one is going to discuss LNTRs in general, as well as some common problems and challenges that they face. I’ll include a very brief history of the designs, almost all of them dating from the 1950s and 1960s, which we’ll look at more in depth in the next couple posts.
Unfortunately, a lot of the fundamental problems of an LNTR get deep – fast, for a lot of people, but the fundamental concepts are often not too hard to get in the broad strokes. I’m gonna try my best to explain them the way that I learned them, and if there’s more questions I’ll attempt to point you to the references I’ve used as a layperson, but I honestly believe that this architecture has suffered from a combination of being “not terrible, not great” in terms of engine performance (1300 s isp, 19/1 T/W).
With that, let’s get into liquid fueled NTRs (LNTR), their history, and their design!
Basic Design Options for LNTR
LNTRs are not a very diverse group of reactor concepts, partially due to the nature of the fuel and partially because they haven’t been well-researched overall. All designs I’ve found use centrifugal force to contain molten fuel inside a tube, with the central void in the spinning tube being the outlet point for the propellant. The first design used a single, large fuel mass in a single fuel element, but quickly this was divided into multiple individual fuel elements, which became the norm for LNTR through the latest designs. One consequence of this first design was the calculation of the neutronic moderation capacity of the H2 propellant in this toroidal fuel structure, and the authors of the study determined that it was so close to zero that it was worth it to consider the center of the fuel element to be a vacuum as far as MCNP (the standard neutronic modeling code both at the time, and in updated form now) is concerned. This is something worth noting: any significant neutron moderation for the core must come from the reflectors and moderator either integrated into the fuel structure (complex to do in a liquid in many cases) or the body of the reactor, the propellant flow won’t matter enough to cause a significant decrease in neutron velocities.
They do seem to fall into two broad categories, which I’ll call bubblers and radiators. A bubbler LNTR is one where the fuel is fed from the outside of the fuel element, through the molten fuel, and into the central void of the fuel element; a radiator LNTR passes propellant only through the central void along the long axis of the fuel element.
A bubbler has the advantage that it is able to use an incredible amount of surface area for heat transfer from the fuel to the propellant, with the surface area being inversely proportional to the size of the individual bubbles: smaller bubbles, more surface area, more heat transfer, greater theoretical power density in the active region of the reactor. They also have the advantage of being able to regeneratively cool the entire length of the fuel element’s outside surface as a natural consequence of the way the propellant is fed into the fuel, rather than using specialized regenerative cooling systems in the fuel element canister and reactor body. However, bubblers also have a couple problems: first, the reactor will not be operating continuously, so on shutdown the fuel will solidify, and the bubbling mechnaism will become clogged with frozen nuclear fuel; second, the breaching of the bubbles to the surface can fling molten fuel into the fast-moving propellant stream, causing fuel to be lost; finally, the bubbles increase mixing of the fuel, which is mostly good but can also lead to certain chemical components of the fuel being carried at a greater rate by either vaporizing and being absorbed into the bubbles or becoming entrained in the fuel and outgassing when the bubble breaches the surface. In a way, it’s sort of like boiling pasta sauce: the water boils, and the bubbles mix the sauce while they move up, but some chemical compounds diffuse into the water vapor along the way (which ones depend on what’s in the sauce), and unless there’s a lid on the pot the sauce splatters across the stove, again depending on the other components of the sauce that you’re cooking. (the obvious problem with this metaphor is that, rather than the gaseous component being a part of the initial solution they’re externally introduced)/
Radiators avoid many of the problems of a bubbler, but not all, by treating the fuel almost like a solid mass when its under centrifugal force: the propellant enters from the ship end, through the central void in the fuel element, and then out the aft end to enter the nozzle through an outlet plenum. This makes fuel retention a far simpler problem overall, but fuel will still be lost through vaporization into the propellant stream (more on this later). Another issue with radiators is that without the propellant passing all the way through the fuel from the outer to inner diameter, the thermal emissions will not only go into the propellant, but also into the fuel canister and the reactor itself – more efficiently, actually, since H2 isn’t especially good at capturing heat,k and conduction is more efficient than radiation. This requires regenerative cooling both for the fuel canister and the reactor as well most of the time – which while doable also requires a more complex plumbing setup within the reactor body to maintain material thermal limits on even relatively high temperature materials, much less hydrides (which are good low-volume, low-mass moderators for compact reactors, but incredibly thermally sensitive).
As with any other astronuclear design, there’s a huge design envelope to play with in terms of fuel matrix, even in liquid form (although this is more limited in liquid designs, as we’ll see), as well as moderation level, number and size of fuel elements, moderator type, and other decisions. However, the vast majority of the designs have been iterative concepts on the same basic two ideas, with modifications mostly focusing on fuel element dimensions and number, fuel temperature, propellant flow rates, and individual fuel matrix materials rather than entirely different reactor architectures.
It’s worth noting that there’s another concept, the droplet core NTR, which diffuses the liquid fuel into the propellant, then recaptures it using (usually) centrifugal force before the droplets can leave the nozzle, but this is a concept that will be covered alongside the vapor core reactor, since it’s a hybrid of the two concepts.
A (Very) Brief History of LNTR
Because we’re going to be discussing the design evolution of each type of LNTR in depth in the next two posts, I’m going to be incredibly brief here, giving a general overview of the history of LNTRs. While they’re often mentioned as an intermediate-stage NTR option, there’s been a surprisingly small amount of research done on them, with only two programs of any significant size being conducted in the 1960s.
The first proposal for an LNTR was by J. McCarthy in 1954, in his “Nuclear Reactors for Rockets.” This design used a single, large cylinder, spun around the long axis, as both the reactor and fuel element. The fuel was fed into the void in the cylinder radially, bubbling through the fuel mass, which was made of uranium carbide (UC2). This design, as any first design, had a number of problems, but showed sufficient promise for the design to be re-examined, tweaked, and further researched to make it more practical. While I don’t have access to this paper, a subsequent study of the design placed the maximum specific impulse of this type of NTR in the range of 1200-1400 seconds.
This led to the first significant research program into the LNTR, carried out by Nelson et al at the Princeton Aeronautical Engineering Laboratory in 1963. This design changed the single large rotating cylinder into several smaller ones, each rotating independently, while keeping the same bubbler architecture of the McCarthy design. This ended up improving the thrust to weight ratio, specific impulse, power density, and other key characteristics. The study also enumerated many of the challenges of both the LNTR in general, and the bubbler in specific, for the first time in a detailed and systematic fashion, but between the lack of information on the materials involved, as well as lack of both computational theory and modeling capability, this study was hampered by many assumptions of convenience. Despite these challenges (which would continue to be addressed over time in smaller studies and other designs), the Princeton LNTR became the benchmark for most LNTR designs of both types that followed. The final design chosen by the team has a vacuum specific impulse of 1250 s, a chamber pressure of 10 atm, and a thrust-to-weight ratio of about 2:1, with a reactor mass of approximately 100 metric tons.
Studies on the technical details of the most challenging aspect of this design, that of bubble motion, would continue at Princeton for a number of years, including experiments to observe the behavior of the particular bubble form needed while under centrifugal acceleration, but challenges in modeling the two-phase (liquid/gas) interactions for thermodynamics and hydrodynamics continued to dog the bubbler design. It is unclear when work stopper on the bubbler design, but the last reference to it that I can find in the literature was from 1972, in a published Engineering Note by W.L. Barrett, who observed that many of the hoped-for goals were overly optimistic, but not by a huge margin. This is during the time that American astronuclear funding was being demolished, and so it would not be surprising that the concept would go into dormancy at that point. Since the restarting of modest astronuclear funding, though, I have been unable to find any reference to a modern bubbler design for either terrestrial or astronuclear use.
Perhaps the main reason for this, which we’ll discuss in the next section, is the inconveniently high vapor pressure of many compounds when operating in the temperature range of an LNTR (about 8800 K). This means that the constituent parts of the fuel body, most notably the uranium, would vaporize into the propellant, not only removing fissile material from the reactor but significantly increasing the mass of the propellant stream, decreasing specific impulse. This, in fact, was the reason the Lewis Research Center focused on a different form of LNTR: the radiator.
Work on the radiator concept began in 1964, and was conducted by a team headed by R Ragsdale, one of the leading NTR designers ar Lewis Research Center. To mitigate the vapor losses of the bubbler type, the question was asked if the propellant actually had to pass through the fuel, or if radiant heating would suffice to thermalize the hydrogen propellant while minimizing the fuel loss from the liquid/gas interaction zone. The answer was a definite yes, although the fuel temperature would have to be higher, and the propellant would likely need to be seeded with some particulate or vapor to increase its thermal absorption. While the overall efficiency would be slightly lower, only a minimal loss of specific impulse would occur, and the thrust to weight ratio could be increased due to higher propellant flow (only so much propellant can pass through a given volume of bubbler-type fuel before unacceptable splattering and other difficulties would arise). This seems to have reached its conclusion in 1967, the last date that any of the papers or reports that I’ve been able to find, with a final compromise design achieving 1400 s of isp, a thrust-to-core-weight-ratio of 4:1, at a core temperature of 5060 K and a reactor pressure of 200 atm (2020 N/m^2).
However, unlike with the bubbler-type LNTR, the radiator would have one last, minor hurrah. In the 1990s, at the beginning of the Space Exploration Initiative, funding became available again for NTR development. A large conference was held in 1991, in Albuquerque, NM, and served as a combination state-of-research and idea presentation for what direction NTR development should go in, as well as determining which concepts should be explored more in depth. As part of this, presentations were made on many different fundamental reactor architectures, and proposals for each type of NTR were made. While the bubbler LNTR was not represented, the radiator was.
This concept, presented by J Powell of Brookhaven National Lab, was the Liquid Annular Reactor System. Compared to the Lewis and Princeton designs, it was a simple reactor, with only seven fuel elements, These would be spaced in a cylinder of Be/H moderator, and would use a twice-through coolant/propellant system: each cylinder was regeneratively cooled from nozzle-end to ship-end, and then the propellant, seeded with W microparticles, would then pass through the central void and out the nozzle. Interestingly enough, this design did not seem to reference the work done by either Princeton or Lewis RC, so there’s a possibility that this was a new design from first principles (other designs presented at the conference made extensive use of legacy data and modeling). This reactor was only conceptually sketched out in the documentation I’ve found, operated at higher temperatures (~6000 K) and lower pressures (~10 atm) than the previous designs to dissociate virtually all of the hydrogen propellant, and no estimated thrust-to-core-weight ratios.
It is unclear how much work was done on this reactor design, and it also remains the last design of any LNTR type that I’ve been able to come across.
Lessons from History: Considerations for LNTR Design
Having looked through the history of LNTR design, it’s worth looking at the lessons that have been learned from these design studies and experiments, as well as the reasons (as far as we can tell) that the designs have evolved the way they did. I just want to say up front that I’m going to be especially careful about when I use my own interpretation, compared to a more qualified someone else’s interpretation, on the constraints and design philosophies here, because this is an area that runs into SO MANY different materials, neutronics, etc constraints that I don’t even know where to begin independently assessing the advantages and disadvantages.
Also, we’re going to be focusing on the lessons that (mostly) apply to both the bubbler and radiator concepts. The following posts, covering the types individually, will address the specific challenges of the two types of LNTR.
The number of fuel elements in an LNTR is a trade-off.
Advantages to increasing the number of fuel elements
The total surface area available in the fuel/propellant boundary increases, increasing thrust for a given specific impulse
The core becomes more homogeneous, making a more idealized neutronic environment (there’s a limit to this, including using interstitial moderating blocks between the fuel elements to further thermalize the reactor, but is a good rule of thumb in most cases)
Advantages to minimizing the number of fuel elements
The more fuel elements, the more manufacturing headache in making the fuel element canisters and elements themselves, as well as the support equipment for maintaining the rotation of the fuel elements;
depending on the complexity of the manufacturing process, this could be a significant hurdle,
Electronic motors don’t do well in a high neutron flux, generally requiring driveshaft penetration of at least part of the shadow shield, and turbines to drive the system can be so complex that this is often not considered an option in NTRs (to be fair, it’s rare that they would be needed)
The less angular velocity is needed for each fuel element to have the same centrifugal force, due to the larger radius of the fuel element
For a variety of reasons the fuel thickness increases to maintain the same critical mass in the reactor – NOTE: this is a benefit for bubbler-type LNTRs, but either neutral or detrimental to streamer-type NTRs.
Another major area of trade-off is propellant mass flow rates. These are fundamentally limited in bubbler LNTRs (something we’ll discuss in the next post), since the bubbles can’t be allowed to combine (or splattering and free droplets will occur), the more bubbles the more the fuel expands (causing headaches for fuel containment), and other issues will present themselves. On the other hand, for radiator – and to a lesser extent the bubbler – type LNTRs, the major limitation is thermal uptake in the propellant (too much mass flow means that the exhaust velocity will drop), which can be somewhat addressed by propellant seeding (something that we’ll discuss in a future webpage).
Fuel Material Constraints
One fundamental question for any LNTR fuel is the maximum theoretical isp of a design, which is a direct function of the critical temperature (when the fuel boils) and at what rate the fuel would vaporize from where the fuel and propellant interact. Pretty much every material has a range of temperature and pressure values where either sublimation (in a solid) or vaporization (in a liquid) will occur, and these characteristics were not well understood at the time.
This is actually one of the major tradeoffs in bubbler vs radiator designs. In a bubbler, you get the propellant and the maximum fuel temperature to be the same, but you also effectively saturate the fuel with any available vapor. The actual vapor concentrations are… well, as far as I can tell, it’s only ever been modeled with 1960s methods, and those interactions are far beyond what I’m either qualified or comfortable to assess, but I suspect that while the problem may be able to be slightly mitigated it won’t be able to be completely avoided.
However, there are general constraints on the fuels available for use, and the choice of every LNTR has been UC2, usually with a majority of the fuel mass being either ZrC or NbC as the dilutent. Other options are available, potentially, such as 184W-U or U-Si metals, but they have not been explored in depth.
Let’s look at the vapor pressure implications more in depth, since it really is the central limitation of LNTR fuels at temperatures that are reasonable for these rockets.
Vapor Pressure Implications
A study on the vapor pressure of uranium was conducted in 1953 by Rauh et al at Argonne NL, which determined an approximate function of the vapor pressure of “pure” uranium metal (some discussion about the inhibiting effects of oxygen, which would not be present in an NTR to any great degree, and also tantalum contamination of the uranium, were needed based on the experimental setup), but this was based on solid U, so was only useful as a starting point.
W Louis Barret Jr. conducted another study in 1963 on the implications of fuel composition for a bubbler-type LNTR, and the constraints on the potential specific impulse of this type of reactor. The author examined many different fissile fuel matrices in their paper, including Pu and Th compounds:
From this, and assuming a propellant pressure of 10^3 psi, a maximum theoretical isp was calculated for each type of fuel:
Additional studies were carried out on uranium metal and carbon compounds – mostly Zr-C-U, Nb-C-U and 184W-C-U, in various concentrations – in 1965 and 66 by Kaufman and Peters of MANLABS for NASA Lewis Research Center (the center of LNTR development at the time), conducted at 100 atmospheres and ~4500 to ~5500 K. These were low atomic mass fraction systems (0.001-0.02), which may be too low for some designs, but will minimize fissile fuel loss to the propellant flow. Other candidate materials considered were Mo-C-U, B-C-U, and Me-C-U, but not studied at the time.
A summary of the results can be found below:
Perhaps the most significant question is mass loss rates due to hydrogen transport, which can be found in this table:
These values offer a good starting point for those that want to explore the maximum operating temperature of this type of reactor, but additional options may exist. For instance, a high vapor pressure, high boiling point, low neutron absorption metal which will mix minimally with the uranium-bearing fuel could be used as a liquid fuel clad layer, either in a persistent form (meant to survive the lifetime of the fuel element) or as a sacrificial vaporization layer similar to how ablative coatings are used in some rocket nozzles (one note here: this will increase the atomic mass of the propellant stream, decreasing the specific impulse of such a design). However, other than the use of ZrC in the Princeton design study in the inner region of that fuel element design (which was also considered a sacrificial component of the fuel), I haven’t seen anyone discuss this concept in depth in the literature.
A good place to start investigating this concept, however, would be with a study done by Charles Masser in 1967 entitled “Vapor-Pressure Data Extrapolated to 1000 Atmospheres of 13 Refractory Materials with Low Thermal Absorption Cross Sections.” While this was focused on the seeding of propellant with microparticles to increase thermal absorption in colder H2, the vapor-pressure information can provide a good jumping off point for anyone interested in investigating this subject further. The paper can be found here: https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19670030361.pdf.
Author speculation concept:
Another, far more speculative option is available if the LNTR can be designed as a thermal breeder, and dealing with certain challenges in fuel worth fluctuations (and other headaches), especially at startup: thorium. This is because Th has a much lower vapor pressure than U does (although the vapor pressure behavior of carbides in a high temperature, high pressure situation doesn’t seem to have been studied ThO2 and ThO3 outperform UC2 – but oxides are a far worse idea than carbides in this sort of reactor), so it may be possible to make a Th-breeder LNTR to reduce fissile fuel vapor losses – which does nothing for C, or Zr/Nb, but may be worth it.
This requires a couple things to happen: first, the reactor’s available reactivity needs to be able to remain within the control authority of the control systems in a far more complex system, and the breeding ratio of the reactor needs to be carefully managed. There’s a few reasons for this, but let’s look at the general shape of the challenge.
Many LNTR designs are either fast or epithermal designs, with few extending into the thermal neutron spectrum. Thorium breeds into 233U best in the thermal neutron spectrum, so the neutron flux needs to be balanced against the Th present in the reactor in order to make sure that the proper breeding ratio is maintained. This can be adjusted by adding moderator blocks between the fuel elements, using other filler materials, and other options common to NTR neutronics design, but isn’t something that I’ve seen addressed anywhere.
Let’s briefly look at the breeding process: when 232Th is bred into 233U, it goes through a two-week period where the nucleus undergoing the breeding process ends up existing as 233Pa, a strong neutron poison. Unlike the thorium breeding molten salt reactor, these designs don’t have on-board fuel reprocessing, and that’s a very heavy, complex system that is going to kill your engine’s dry mass, so just adding one isn’t a good option from a systems engineering point of view. So, initially, the reactor loses a neutron to the 232Th, which then changes to 233Th before quickly decaying into 233Pa, a strong neutron poison which will stay in the reactor until long after the reactor is shut down (and so waste energy will need to be dealt with, but radiation may/probably is enough to deal with that), and then it’s likely that the next time the engine is started up, that neutron poison has transmuted into an even more fissile material unless you load the fuel with 233U first (233U has a stronger fission capture cross-section than 235U, which in practical effect reduces the fissile requirements by ~33%)!
This means that the reactor has to go through startup, have a reasonably large amount of control authority to continue to add reactivity to the reactor to counterbalance the fission poison buildup of not only 233Pa, but other fission product neutron poisons and fissile fuel worth degradation (if the fuel element has been used before), and then be able to deal with a potentially more reactive reactor (if the breeding ratio has more of a fudge factor due to the fast ramp-up/ramp-down behavior of this reactor, varying power levels, etc, making it higher in effect than ~1.01/4).
The other potential issue is that if you need less fissile material in the core, every atom of fissile is more valuable in the core than a less fissile fuel. If the vapor entrainment ends up being higher than the effective breeding ratio (i.e. the effect of breeding when the reactor’s operating), then the reactor’s going to lose reactivity too fast to maintain. Along these lines, the 233Pa behavior is also going to need to be studied, because that’s not only your future fuel, but also a strong neutron poison, in a not-great neutronic configuration for your fuel element, so there’s a few complications on that intermediate step.
This is an addressable option, potentially, but it’s also a lot of work on a reactor that already has a lot of work needed to make feasible.
Liquid fueled NTRs (LNTRs) show great promise as a stepping stone to advanced NTR development in both their variations, the bubbler and radiator variants. The high specific impulse, as well as potentially high thrust-to-weight ratio, offer benefits for many interplanetary missions, both crewed and uncrewed.
However, there are numerous challenges in the way of developing these systems. Of all the NTR types, they are some of the least researched, with only a handful of studies conducted in the 1960s, and a single project in the 1990s. These projects have focused on a single family of fuels, and those have not been able to be tested under fission power for various neutronic and reactor physics behaviors necessary for the proper modeling of these systems.
Additionally, the interactions between the fuel and propellant in these systems is far more complex than it is in most other fuel types. Only two other types of NTR (the droplet/colloid core and open cycle gas core NTRs) face the same level of challenge in fissile fuel retention and fuel element mass entrainment that the LNTR faces, especially in the bubbler variation.
Finally, they are some of the least well-known variations of NTR in both popular and technical literature, with only a few papers ever being published and only short blurbs on popular websites due to the difficulty in finding the technical source material.
We will continue to look at these systems in the next two blog posts, covering the bubbler-type LNTR in the next one, and the radiator type in the one following that. These blog posts are already in progress, and should be ready for publication in the near term.
If you would like early access to these, as well as all future blog posts and websites, consider becoming a Patron of the page! My Patrons help me be able to devote the time that I need to the website, and provide strong encouragement for me to put out more material as well! You can sign up here: https://www.patreon.com/beyondnerva
Hello, and welcome back to Beyond NERVA! The blog itself has been quiet for a while for a number of reasons, but the website continues to grow! I’ve added extensive sections on radioisotope power sources and many details of their operation. There are now pages for radioisotope power sources in general (which you can find here), the fuel elements and heating units used in these systems (here), the considerations for fuel selection (which can be found here), the various options for the fuel itself (each of these has its own page: 238Pu, 241Am, and 210Po). I also covered the RTGs of the SNAP era (available here), which evolved concurrently with the SNAP-2, SNAP-10, and SNAP-8 reactors that we’ve already looked at in depth (each of those now has its own page, as well as the experimental and development reactors associated with the systems, just click on their names to find them!), as well as the Multi-Hundred Watt RTG that powered the Voyager spacecraft, the General Purpose Heat Source RTG which powered Galileo, Ulysses, and Cassini, and still powers New Horizons, and as this is released a page will be coming online about the Multi-Mission RTG, NASA’s current workhorse RTG! Make sure to check out those pages for in depth information on an ecosystem of power supplies which are fascinating, but often overlooked (including by me) for the flashier, higher-powered fission reactor proposals.
Today, we’re going to look at a particular application of the Multi-Mission RTG, or MMRTG.
This is the newly announced Dragonfly mission to Titan. This mission will only be the second to touch down on the surface of that Saturnian moon, and promises to transform our understanding of the complex hydrocarbon cycles, liquid water (which is under the surface of Titan), and complex organic chemistry that may hold clues to the early atmospheric conditions of Earth.
Congratulations are very much in order for the Johns Hopkins University Applied Physics Laboratory team for their successful mission concept!
This mission will be
primarily based around the quadcopter which has gained the most
attention in the recent announcements. Using eight rotors mounted to
four motor mounts, this system will charge a set of lithium ion
batteries with the output of the MMRTG, and then use this
higher-power-density supply to fly in short hops, first scouting out
the landing point where it will settle for the science gathering
portion after the one it will
be landing on for that flight, and then land at an already-scouted
position on the Titanic surface. There a suite of instruments,
mass spectrometer, a neutron
and gamma spectrometer, a meteorological and seismic sensor suite,
and a camera suite. There
may be others that are yet to
be fully determined but will be refined over the course of the final
mission planning (which can begin now that the mission
has been selected for funding.
on Titan is far easier than on Earth due to the high atmospheric
density (over 4 times higher
than Earth’s at sea level)
and low gravity (1/7
g) of the moon, and as such a
host of proposals for flying probes have been proposed over the
years, from balloons to helicopters to airships to airplanes – and
even a radioisotope thermal rocket proposal!
Furthermore, both surface and lake landers have been proposed.
the team pointed out in their proposal, the Dragonfly concept
combines the benefits of a Curiosity-class lander’s equipment suite
and surface science capability with the mobility of an
aerial movement platform. This provides the scientific advantages of
surface deployment with the ability to relocate the lander at a far
higher rate of travel than a wheeled rover such as Curiosity. This
provides a good balance of the advantages of each design type.
first look at the mission itself, from launch to landing on Titan,
followed by the science goals of the lander, and finally we’ll look
at the power supply relatively briefly, focusing on the thermal
management strategies needed for both the lander and the RTG itself.
All of the information here
is as of July 2019, so much is still subject to change in the next
few years before the launch window.
you’re looking for a particular aspect of the mission, the links
below will jump you to the appropriate section, but as with all
missions, each phase or component informs every other one, so by
skipping ahead you may miss some interesting tidbits about this
Profile Pre-Titan: Launch, Cruise, and EDL
Dragonfly will launch on either an Atlas 541 or equivalent launcher on April 12th, 2025, and conduct a series of flybys of various planets to get out to the Saturnian system. This means that a smaller launcher can be used, and therefore a less expensive one to save more funding for the spacecraft and science teams. However, since this launch vehicle (to my knowledge) hasn’t firmly been pinned down, differences in the actual launcher – as well as any possible onboard non-optimal launch conditions, could change the exact timing of the gravitational assists listed below.
The first gravitational assist will be on April 11th, 2026
from Earth, followed by a Venusian gravitational assist on 4/16/2017,
another Earth gravitational assist on 5/27/2028, and a final Earth
gravitational assist on 9/3/2031. While there are options to
rearrange these gravitational assists, this one was selected due to a
number of orbital mechanical factors. Sadly, Jupiter will be out of
phase with Saturn at the appropriate time, so it’s impossible to
use the large planet’s convenient gravity well to shorten the trip
to the Saturnian system.
After this series of four inner-system gravitational assists,
Dragonfly will have the necessary dV to get to Titan. A mid-course
correction around December 2, 2031 (while the spacecraft is between
Mars and Juptier) will ensure that the spacecraft is oriented
correctly for Titan capture.
During this time, the MMRTG will use its secondary coolant loop (visible as the silver tubes at the base of the RTG’s fins in the above image) combined with a pumped coolant loop similar to what was used on Curiosity’s cruise stage, in order to reject the waste heat from the RTG during the time that it’s enclosed in the heat shield and cruise stage. It’s unclear in the mission documentation whether the RTG will be used to power any instruments during cruise, as Curiosity’s particle detection system was. The instrumentation on Dragonfly is significantly different from Curiosity, and it’s not apparent whether any of the instruments would either be useful in cruise, or perhaps whether operation of the cruise stage would interfere with those experiments that could be useful to the point that the data would simply be too messy or corrupted to bother.
The cruise stage design is also something that I haven’t been able
to find, so it’s possible (although unlikely due to increased
electrical bus complexity) that the MMRTG could be used to power
scientific instruments on the cruise stage itself. While the majority
of the time the spacecraft will be doing the series of gravitational
assists in the inner Solar System, allowing for the use of solar
panels, by the time it is heading out of the inner Solar System the
power available from solar radiation will be dropping off
exponentially. This could likely be handled by battery power on board
the cruise stage, or it could possibly be handled by the MMRTG.
However, the increased complexity of the electrical system should it
be connected to both the cruise stage as well as the lander may be
sufficient to have the mission planners decide to not go this route.
Descent, and Landing for Dragonfly
Finally, at the close of 2034 (12/30), the spacecraft will reach
Titan, and perform its entry, descent and landing procedures.
For those that remember Curiosity’s landing, this was (rightly)
touted as the “Seven Minutes of Terror,” involving a huge number
of complex and risky maneuvers with a collection of sometimes exotic
craft to land on Gale Crater. These included: a set of hypersonic
parachutes designed for the thin Martian atmosphere, followed by the
use of an eight-engined “sky crane,” which hovered meters off the
surface of Mars to lower the rover on a winch and cable system
(during which the rover’s wheel and bogey driving system deployed
for the first time), followed by placing the rover on the surface of
Mars, disconnecting the cables, and then flying away to crash at some
distance. This was due to the combination of thin atmosphere,
reasonably high gravity, and the need to minimize dust and debris
from collecting on the surface of the rover (to protect the wet and
dry lab sample collection system on the surface of the rover from
being contaminated). It was a stunning series of firsts in an EDL
system, and one which rightly received worldwide attention – and
garnered NASA to repeat the process for Mars 2020.
Titan, on the other hand, is a whole different ballgame. Its thick
atmosphere and low gravity make it a very different EDL environment,
one that is both easier to deal with but also requires a different
set of conditions to be followed. Due to the fact that the gravity is
so low (1/7 of Earth’s), Titan’s sensible atmosphere extends far
higher than even Earth’s does. In fact, even though the atmospheric
density is 4 times that of Earth’s at the surface, the atmospheric
pressure is only slightly higher (1.4x). Combining a far
lighter-weight rover (weight=mass x gravity), with far better lift,
and onboard flight capability, the sky-crane maneuver isn’t
necessary on Titan.
The beginning will look similar: the cruise stage is outwardly very similar to the MSL cruise stage, and will also be ejected as the craft begins to enter the atmosphere. A set of parachutes will slow the craft further, until an optimal velocity is achieved above the surface of Titan. During this time, it’s likely that the radar designed to map the Titanic surface in flight will be used to verify the lander’s location over the surface and find an acceptable landing location for its first touchdown. Based on the typical flight profile of Dragonfly, and depending on the power level in the batteries (which would likely be fully charged) compared to how much power it will take to get to a safe landing location, the next flight’s landing location may be scouted as well.
This video from JHUAPL shows the EDL process for Dragonfly:
The location of the first landing location (as well as all other
subsequent locations) is dependent on communication with Earth.
Dragonfly is designed to communicate directly with Earth, something
which is standard for outer solar system missions due to the lack of
a communications architecture in a useful location (the one exception
is the Huygens probe, also deployed to Titan, which used the Cassini
spacecraft as a communications relay). This places certain limits on
the location that Dragonfly can deploy to on the Titanic surface, and
also where the lander will move across the surface during its ground
science mission – but more on that in the surface mission section.
The mission planners for Dragonfly have selected a similar landing
latitude and season as the Huygens mission’s landing in order to
maximize the knowledge of the atmospheric conditions for the initial,
riskiest portion of the lander’s atmospheric operations. This also
maximizes communications availability with Earth and works well with
orbital mechanical constraints upon entering the Saturnian system and
aerocapture by Titan itself.
After the lander is safely on the surface of Titan, it will deploy
its communications dish, send data back to Earth, and begin surface
science experiments. This will be the start of the Dragonfly surface
mission, which will last a number of years – until either a failure
occurs on the lander which prevents its further operation, or the
MMRTG degrades to the point that communication with Earth is no
longer possible (the science equipment takes less power than the
communications do, so even if Dragonfly isn’t able to fly it can
still provide valuable scientific data – if it can phone home).
This begins the purpose of the whole mission, which will be explored
in the next section
Surface Mission and Flight Profile
Titan is a fascinating place. The coldest location in the Solar
System thanks to its location in the outer solar system, complex
hydrocarbon cycles which in may ways mimic the hydrological cycles of
Earth (but with a complex set of liquid hydrocarbons rather than
water), and a chemical profile that may be similar to Earth’s at
the beginning of the evolution of life on Earth provide a fascinating
place to conduct science missions, with implications reaching not
only far back into Earth’s past, but also have a major impact on
the future of humanity in the Solar System.
The surface mission of Dragonfly largely comes in two major phases:
landed science instruments and communications, and movement. While
some data is collected in flight (mostly imaging), the in depth data
collection and transmission are done on the surface of the moon. This
is important, because while the MMRTG is the best power supply
available for this mission, it isn’t able to provide the power
needed for flight as quickly as needed. This means that a set of
lithium ion batteries, stored in an insulated box that’s heated by
waste heat from the MMRTG, are charged while on the surface, and then
once the desired power level is reached, a new flight can occur.
Let’s begin by looking at the science instruments which will be included on Dragonfly. This list is quoted from “Dragonfly: A Rotorcraft Lander Concept for Scientific Exploration of Titan,” a white paper from Lorentz et al at JHUAPL on January 9th of 2019:
DraMS—Dragonfly Mass Spectrometer (Goddard Space Flight Center). A central element of the payload is a highly capable mass spectrometer instrument, with front-end sample processing able to handle highmolecular-weight materials and samples of prebiotic interest. The system has elements from the highly successful SAM (Sample Analysis at Mars) instrument on Curiosity, which has pyrolysis and gas chromatographic analysis capabilities, and also draws on developments for the ExoMars/MOMA (Mars Organic Material Analyser).
DraGNS—Dragonfly Gamma-Ray and Neutron Spectrometer (APL/Goddard Space Flight Center). This instrument allows the elemental composition of the ground immediately under the lander to be determined without requiring any sampling operations. Note that because Titan’s thick and extended atmosphere shields the surface from cosmic rays that excite gammarays on Mars and airless bodies, the instrument includes a pulsed neutron generator to excite the gamma-ray signature, as also advocated for Venus missions. The abundances of carbon, nitrogen, hydrogen, and oxygen allow a rapid classification of the surface material (for example, ammonia-rich water ice, pure ice, and carbonrich dune sands). This instrument also permits the detection of minor inorganic elements such as sodium or sulfur. This quick chemical reconnaissance at each new site can inform the science team as to which types of sampling (if any) and detailed chemical analysis should be performed.
DraGMet—Dragonfly Geophysics and Meteorology Package (APL). This instrument is a suite of simple sensors with low-power data handling electronics. Atmospheric pressure and temperature are sensed with COTS sensors. Wind speed and direction are determined with thermal anemometers (similar to those flown on several Mars missions) placed outboard of each rotor hub, so that at least one senses wind upstream of the lander body, minimizing flow perturbations due to obstruction and by the thermal plume from the MMRTG. Methane abundance (humidity) is sensed by differential near-IR absorption, using components identified in the TiME Phase A study. Electrodes on the landing skids are used to sense electric fields (and in particular the AC field associated with the Schumann resonance, which probes the depth to Titan’s interior liquid water ocean) as well as to measure the dielectric constant of the ground. The thermal properties of the ground are sensed with a heated temperature sensor to assess porosity and dampness. Finally, seismic instrumentation assesses regolith properties (e.g., via sensing drill noise) and searches for tectonic activity and possibly infers Titan’s interior structure.
DragonCam—Dragonfly Camera Suite (Malin Space Science Systems). A set of cameras, driven by a common electronics unit, provides for forward and downward imaging (landed and in flight), and a microscopic imager can examine surface material down to sand-grain scale. Panoramic cameras can survey sites in detail after landing: in many respects, the imaging system is similar to that on Mars landers, although the optical design takes the weaker illumination at Titan (known from Huygens data) into account. LED illuminators permit color imaging at night, and a UV source permits the detection of certain organics (notably polycyclic aromatic hydrocarbons) via fluorescence.
Engineering systems. Data from the inertial measurement unit (IMU) may be used to recover an atmospheric density profile via the deceleration history during entry. IMU and other navigation data may provide constraints on winds during rotorcraft flight. Additionally, the radio link via Doppler and/or ranging measurements may shed light on Titan’s rotation state, which, in turn, is influenced by its internal structure.”
It’s unclear whether all electrical power will be routed through
these batteries or not. While there are advantages from a power
conditioning point of view (ensuring the correct voltage and wattage,
preventing power dropouts or spikes to sensitive instruments, etc),
it can also cause battery life complications due to the constant
discharging of the batteries themselves – and degradation of
especially the anode in the batteries. It’s unclear which power
conditioning scheme will be used for the always-on systems, such as
the meteorological system, but the high powered systems will likely
draw power from the batteries exclusively.
In order to deploy these instruments, the lander will fly from one location to another, scouting out the location of the landing after the target that it will be landing at on that mission. A good schematic of the flight profile can be seen below:
This ensures that the flight time for the next flight can be
maximized, since a known safe landing location is already known
before the lander takes off for its next flight, as well as providing
margin in flight time for all legs of the mission after the initial
one. It’s unclear if the first flight will also use this profile,
since that’s part of the entry, descent, and landing sequence, but
there’s no apparent reason why it couldn’t if the lander is
ejected from the backshell at moderate altitude and velocity.
Due to the remote nature of this mission, there aren’t many (if
any) options available to use communications hubs off Earth as relays
for Dragonfly, a major contrast to Martian operations where the many
orbiters around the planet also serve as communications satellites
for the various landers and rovers operating on the surface of Mars.
This means that in order for the scientific and engineering data from
Dragonfly to be returned to Earth, and additional commands to be
transmitted, the rover needs line of sight to Earth. This is done via
a deployable high gain antenna, which will be stowed for flight
operations to reduce drag and stress on the antenna itself.
This complicates matters in two ways: if the lander is at too high or
too low a latitude, there’s no line-of-sight available to Earth
from the Titanic surface, meaning that communications is impossible;
second, although the length of the sol (the extraterrestrial
equivalent of a day, in the case of Titan this is almost 16 Earth
days, the same as Titan’s orbital period around Saturn).
While it may be possible to do a hop outside the
communications window, collect scientific data, then return to the
communications window to transmit the data, it also increases the
chance of an unrecoverable failure due to the lack of ability for the
engineering team on Earth to troubleshoot and resolve any potential
The lander isn’t tied to the ability to gather sunlight like a
solar-powered spacecraft is, but at the same time the ability to
communicate with Earth and having the Sun visible are conditions that
pretty closely overlap, so night-time scientific data gathering on
this mission essentially means that all data would need to be stored
on board the spacecraft, and the rate of data collection and data
storage capabilities of Dragonfly aren’t clear from current
documentation. This means that, should the lander need to overnight
(a very real possibility) on Titan, some of the data may need to be
written over in order to make room for more immediately interesting
Whether this is an avoidable circumstance or not is something that
I’ve been unable to determine, but the mission design team have
made provisions for overnighting Dragonfly on Titan, and this may in
fact be required to allow for the time necessary to recharge the
batteries to flight condition. If this is the case, it’s likely
that the mission’s data storage capabilities will be sufficient to
collect all desired data through the Titanic night and transmit them
during daytime surface operations, when line-of-sight with Earth is
Now that we’ve looked at what Dragonfly is going to be doing
on the surface of Titan, let’s look at how it will do it
from a power point of view: the Multi-Mission RTG, or MMRTG.
RTG: NASA and the DOE’s Flagship RTG
is also the first RTG to be built for NASA in decades that was
designed to operate in an atmosphere, a major thermal management
change from the typical spaceborne MHW-RTG and GPHS-RTG systems. RTGs
have of course been designed to reject heat into both atmosphere and
into liquid during the SNAP RTG program, which included naval and air
on those systems here),
but this application hadn’t been used by NASA since the SNAP-19 (the
Viking landers used two of the generators, details
of the non-GPHS elements of the MMRTG are manufactured by Teledyne
(the company that the thermocouple inventors worked for in the 1960s
and 1970s), and Aerojet Rocketdyne. Lockheed Martin and the
Department of Energy provide the systems and materials for the GPHS
modules that fuel the MMRTG.
thermoelectric generator (TEG) assembly of the MMRTG is based on a
legacy design: the lead telluride/telluride arsenic germanium
selenite (PbTe/TAGS85) thermoelectric thermocouple system. This was
first used in the SNAP-19 RTG (more
information available here).
The MMRTG uses an updated configuration, using 768 thermocouples
configured as two series-parallel chains for fault tolerance: any
thermocouple that fails will do so individually, resulting in a
negligible and isolated loss of power that is easy and reliable to
integrate into mission planning.
one significant barrier for the use of these materials is
differential sublimation of the thermoelectric materials themselves.
This is unfortunate, but there are a number of ways to manage this
effect. In the current incarnation, a mix of argon and helium are
used as a cover gas, but other cover gas compositions are also
possible. Additionally, silica insulation is used immediately
surrounding the thermocouples to reduce sublimation rates.
has significant implications for Dragonfly, as we’ll see below.
MMRTG is fascinating for a number of reasons in the context of
thermal management. The two most prominent (upon investigation)
differences between the MMRTG and almost every other off-Earth RTG
radiators are designed to work in both vacuum and atmospheric
MMRTG is able to use a pumped
coolant thermal management system during operation,
the first to do so during a mission with the Mars Science Laboratory
cruise stage. This was not only critical for the cruise stage
thermal management, but also had an impact on the Entry, Descent,
and Landing profile
for a lander or rover.
space as part of a composite spacecraft, such Dragonfly in cruise,
thermal transfer points are mechanically and thermally attached to
the radiator fins of the RTG. The cruise stage for Curiosity had a
total of 23 m of aluminum tubes in a two-split flow configuration
integrated to the cruise stage for this purpose, which when welded to
a 1.5 mm aluminum sheet creates a radiator about 6 m^2 in area.
is integrated into the base of the fins on the radiator, which were
then jettisoned for surface operations (although exactly when and how
this occurs is unclear).
spacecraft interface consists of a mounting bracket which connects to
the spacecraft for mechanical attachment, as well as an electrical
and telemetry connection to the spacecraft through a single
connector. Mechanical integration uses a four bolt mounting
interface. The only telemetry provided is from platinum resistance
thermometers within the RTG.
RTG is only ever integrated onto the spacecraft at the launch pad,
due to nuclear material security concerns, waste heat management
simplification, and radiological safety. Interestingly, the entire
RTG was integrated into Curiosity on the launch pad, after the rest
of the rover and cruise stage was already integrated into the launch
vehicle (a ULA Atlas 541). This will be done with Mars 2020 as well,
with a mass simulator being used in its place. I presume the same
will be done for Dragonfly.
While the MMRTG (more information available here) was designed to handle the Titan environment from the beginning, there are many difference between this version of the MMRTG and those that are (and will be) used on Mars. Additionally, the fact that aerodynamics and mass distribution are a major design criterion for Dragonfly places additional requirements on the use of the system.
A standard MMRTG uses an eight bladed, cylindrical radiator to reject
heat. This is present on both the Mars and deep space version of the
MMRTG, but the Martian version uses a pair of shields on either side
of the rover to both control air flow past the radiator (limiting the
amount of convective heat loss) and to capture waste heat from the
RTG to heat certain temperature-sensitive system components. These
shields don’t touch the radiator, and the configuration of the
radiator is the same as the outer space version of the system
(although the deep space version may be painted black instead of
white for thermal management).
systems that use radiators have a minimum and maximum operating
temperature for the radiator itself. For the MMRTG it’s based on
the temperature at the root of each fin on the radiator. While the
MMRTG was designed with a maximum allowable
temperature of 200 C (!) (with a corresponding loss in conversion
efficiency due to a lower thermal gradient) for Martian or orbital
operations in an absolute worst-case scenario, the system faces the
opposite problem on Titan: the incredibly low surface temperature is
well below the minimum
operating temperature at the fin root temperature of -269 C. This
requires the radiator to be insolated from the exterior environment
to a certain degree, meaning that the heat rejection system needs to
Of course, the fins on the radiator aren’t the most aerodynamic
things around: not only would they cause changes in yaw and roll to
be more difficult, but the upward angle of the RTG’s mounting
causes problems with drag as well.
both of these problems can be addressed together, by taking an idea
that’s already in use on the Curiosity rover, extending and
adapting it for the Titanic environment. On
Mars, as we already mentioned, a pair of shields are used on either
side of the rover. For Dragonfly, those shields are extended to make
a cylindrical housing for the MMRTG, as seen at the back of the
lander. Not only does this ensure that the RTG components are kept at
the appropriate temperature, but also that the lander has improved
aerodynamic conditions. This is especially important because while
lift is easy to gain on Titan, drag is also far more powerful.
and the RTG: Unknown Unknowns in Mission Design and Power Supply
the MMRTG has been performing within the design envelope for the
needs of the Curiosity mission on Mars for years, it has shown some
moderately concerning degradation behaviors that the Dragonfly team
are taking into account when designing the power profile for
JPL “Radioisotope Power Systems Reference Book for Mission
Designers and Planners,” by Young Lee and Brian Barstow at JPL,
covers many of the details of the MMRTG, as well as recommendations
on the design margins that should be adhered to when applying this
power supply to a mission proposal.
first concern is fuel age. While the MMRTG on Curiosity supplied 114
We of power on landing on Mars (and this includes three years in
storage, and four years on vehicle integration, launch, and cruise),
the Reference Book lists the conservative power output for a brand
new MMRTG to be only 107 We on the surface of Mars. This is also a
longer time period than the Mars 2020 rover’s MMRTG, which was
delivered in August of 2018 for a 2020 launch, although when the fuel
was produced is something I have yet to determine. This bodes well
for having sufficient power from the 238Pu fuel for the Dragonfly at
the beginning of its surface mission, nine years after launch.
radioactive decay isn’t the only cause of regular degradation for
the MMRTG. The thermoelectric generator itself, the GeTe/TAGS-85
thermocouples that are the children of those used by the Pioneer 10
and 11 probes, lose their thermoelectric materials (mainly germanium)
over time through sublimation and migration out of the thermocouples
themselves. This is currently reduced by using a cover gas of argon
and helium in a carefully controlled ratio, but sadly the utility of
this seems to be less than ideal on the Martian surface.
Additionally, silicon insulation immediately surrounding the
thermocouples can assist in reducing sublimation, but whether that’s
currently in use or not is unclear.
MMRTG’s output was assumed to degrade between 3.5% and 4.8% per
year, between the decay energy decline in the 238Pu fuel elements and
the sublimation of the materials in the thermocouples themselves.
Sadly, real-world data from
Curiosity shows that the MMRTG on board the rover is degrading at the
top end of that scale: 4.8%. This is a fact to cause worry for a
mission planner, and one that the Dragonfly team at JHUAPL have taken
into account in the mission design.
there are advanced mitigation techniques (such as cladding them in
Al2O3 with an atomic layer deposition process for a highly regular
clad similar to nuclear fuel elements, but far more exacting than
average) have been proposed and demonstrated by the University of
Dayton, it’s unclear whether these mitigation techniques will be
used on Dragonfly due to the unknowns that they can introduce into
of the long cruise phase of the mission, the Dragonfly team assume
that the MMRTG will only be able to provide about 70 We of power once
the mission arrives at Titan. This is sufficient to power the Li ion
batteries onboard the lander (held in a thermal insulative box and
kept at temperature with waste heat from the RTG) for both flight and
energy-intensive testing, as well as provide power for the scientific
instruments that will be running during grounded science
to Lorentz et al at JHUAPL in 2018:
“Although sample acquisition and chemical analysis are somewhat power-hungry activities, they require only a few hours of activity. Science activities that require continuous monitoring, namely meteorological and seismological measurements, although of low power, actually dominate the payload energy budget. Indeed, for these extended periods, the lander avionics are powered down and data acquisition is performed only by the instrument, to maximize the rate of recharge of the battery.”
This means that there is more than enough margin for a years-long successful mission from even an MMRTG degrading at the high end of the degradation curve, ensuring that the power supply will likely not be a cause for significant concern of mission failure.
Titan, Here We Come!
The exploration of Titan has long been a goal, ever since the Voyager spacecraft first sent data back from this fascinating moon. Cassini and Huygens sent back reams of data, but sadly only served to whet our appetite for more.
Now, Dragonfly will provide us far more data, in an incredibly mobile platform, on the composition, chemical processes, and weather on Titan. This will not only increase our knowledge of the moon itself, but early chemistry on Earth and how it could have led to the rise of life in the Solar System.
Sadly, it will be 2034 before we receive data back from Dragonfly on the surface of Titan, but this is not unusual for so distant a location. Until then, we can only follow the mission development, cheer on the team at Johns Hopkins University APL, and wait with bated breath for the first data to be transmitted from this distant, intriguing moon.
For more information on the MMRTG, make sure to check out the new MMRTG page, available here: MMRTG page.
Hello, and welcome back to Beyond NERVA! Today, we take a break from the Topaz International Program to cover a subject that we haven’t touched on at Beyond NERVA yet, and sadly the inspiration is due to the death of one of the true luminaries of astronuclear engineering, Dr. Emanuel “Emil” Skrabek, who passed on March 14th of progressive supranuclear palsy and Parkinson’s disease.
His obituary can be found here: http://www.ruckfuneral.com/obituary/emanuel-andrew-skrabek-phd . May his family have comfort in his passing, and may his memory be eternal. His legacy within astronuclear engineering, and the discoveries that his inventions enabled within (and outside) our solar system, certainly make him one of the true unsung engineering heroes in our race’s ability to reach out beyond our planet and learn about our own solar neighborhood.
Today, we are going to talk about the bread and butter of astronuclear engineering: the radioisotope thermoelectric generator, or RTG. From the dawn of spaceflight, these systems have provided simple, solid state power for missions of all types, from orbiters to landers to rovers, and have enabled incredible science to be done in the far-flung reaches of our solar system – and just recently, beyond.
Simply put, RTGs use the natural radioactive decay of a radioisotope, or radioactive isotope of a material, to produce electricity through the use of a pair (or more, but mostly just two) of materials that, at the place that they meet, produce electricity – assuming that there’s a hot side (where you stick the radioisotope) and a cold side (a radiator). They’re insanely attractive to mission planners for quite a few reasons. First, they’re completely solid state, which means that there are no moving parts to break. Second, they’re well-characterized, meaning that the problems that they face, their effects on a spacecraft, their behavior during launch, and a slew of other factors are well-known. Third, they can provide a heat source for components in a spacecraft, meaning that the freezing cold of space isn’t going to cause mechanical seizing or electronics failure thanks to the waste heat from the generator. Finally, they’re a legacy design that remains incredibly relevant, meaning that we’ve been doing them for a long time but they’re still useful.
This is an incredibly well-documented and widely-used application for in-space nuclear power, so I’m working on a page with more details on this technology, but when it will be complete is still up in the air. Follow me either on FB or Twitter to get notified when it is available!
Radioisotope Thermoelectric Generators: The Fuel
Everything is radioactive, but some things are more radioactive than others, and all fall into a broad set of categories of “radioactivity.” For the purposes of a radioisotope thermoelectric generator, the best option for fuel is an isotope that only undergoes alpha decay (and preferably only decay once) so the fuel needs only minimal radiation shielding to protect the spacecraft from any radiation that could damage the spacecraft’s on-board electronics and other payload. Because of this, and because deep space missions have very long timelines. This led RTG designers to decide to use an isotope of Plutonium, 238Pu, for many of its’ deep space mission.
238Pu is rare in that it only emits alpha radiation, meaning that the fuel pretty much shields itself. It also emits a useful amount of heat for either producing electricity or heat, and over a useful timeframe. There are many other options for radioisotope fuel for RTGs, but of the flown missions the vast majority have used 238Pu for fuel. We wont go into the other fuel types here, but I’m currently working on a page on RTGs which will go into the different options.
Unfortunately, due to a number of organizational and planning difficulties within NASA and the DOE, who supply the 238Pu oxide fuel, the radioisotope itself is becoming scarce. With less fuel available, and more powerful sensors needing support from more powerful transmitters to get all the information back to Earth, the increased use of electric propulsion, and other power requirements, the transition is almost being forced on NASA’s mission planners.
Radioisotope Thermoelectric Generators: The Heat Sink
As with pretty much all other known forms of thermal-to-electric heating, there needs to be a hot side of the system and a cold side. In smaller power units, this heat rejection system is simple: a set of simple metal fins secured to the metal exterior of the power unit. This casing shields the very minor amounts of radiation coming from the fuel during the decay process (if you’re interested about the radiation environment for the payload of an MMRTG, it can be found here: https://trs.jpl.nasa.gov/handle/2014/45778 ), and also provides micrometeorite protection for the power supply. Due to the low power involved, these simple systems are more than sufficient to cool the converter.
Another use for the waste heat is to provide heat to temperature-sensitive sensors and electronics. In fact, the RTG is only one application for a broader category of systems” Radioisotope Heating Units.” which can provide not only heat for components, and electric power for on-board systems, but even propulsion as well, in the form of a radioisotope thermal rocket. This last option isn’t available in current systems, but New Horizons successfully used its RTGs to power maneuvering thrusters in the outer solar system. If it’s able to be used somehow, waste heat isn’t waste yet.
Radioisotope Thermoelectric Generators: The Thermocouple
RTGs use the thermoelectric effect to convert heat to electricity. Thethermoelectric effect occurs when there’s a difference in temperature across the junction of two different metals. This is more properly known as the “Seebeck effect,” after its second discoverer (it was first described by Volta in 1794): Thomas Johann Seebeck in 1821 after his independent rediscovery. The temperature range that the system would be exposed to determines the materials that are best for the particular converter. that is being used. The efficiency depends on a specialized material property known as the “Seebeck coefficient;” a higher Seebeck coefficient means a more efficient power conversion process, and more electricity is generated for a given temperature gradient. This coefficient depends on a number of things, including the microstructure and crystalline structure of the materials being used for each element of the converter, meaning that changing the alloy of the base materials used can have a significant effect on the performance of the converter.
This application was actually widely in use already as a type of sensor called a “thermocouple,” which sends a voltage based on its temperature within a certain range of environments and is one of the most common forms of temperature monitoring in many fields. Designers of early astronuclear systems, such as the designers of the SNAP-3 RTG (which flew for the first time in 1961) scaled this concept, turned it inside out, and placed it on a spacecraft. Many different types of material combinations have been experimented with over the years, mostly based around lead and tellurium, PbTe. Thin strips were alternated around the radius of the fuel canister, with heat being provided by conduction and radiation using wide radiator fins. Another common option is silicon-germanium, a better option at higher temperatures.
Today most designs use lead telluride (PbTe) doped with a special material that Dr. Skrabek invented with Donald Trimmer while working for Teledyne Energy Systems: TAGS-85. In TAGS, the tellurium used in the converter. is carefully doped with silver (Ag) and antimony (Sb), hence TAGS. The most commonly used version of TAGS, TAGS-85, uses (PbTe)85(AgSbTe2)15. By using this combination of metals, the crystalline microstructure of the tellurides by adding materials that differ in one of two properties: the atomic size being either significantly larger or smaller, or the inclusion of something that either does or does not have a localized magnetic moment. The reasons for this are incredibly complex, and are still the focus of ongoing research in the field, with two big goals. The first is decreasing thermal conductivity, thereby enabling better heat retention in the fuel until it’s able to be converted into electricity, and something that can be affected by imperfections in the crystalline structure of a material. The other is to increase the power factor, which is the relationship between the Seebeck coefficient and electrical resistivity, which is where the localized magnetic moments come into play. A paper from 2012 looks at more recent proposed changes to TAGS-85, involving the use of dysprosium as an additional doping agent: https://digitalcommons.unl.edu/cgi/viewcontent.cgi?article=1333&context=usdoepub .
This type of material was first used in the SNAP-19B RTG, for the Nimbus-B satellite which failed to launch. The fuel was recovered, and placed in the Nimbus-3 satellite, which launched April 3, 1969.
This was the first in an impressive list of missions powered by these power supplies: Pioneer 10, the first flyby of Jupiter (launched March 2, 1962, perijove on December 4, 1973); and Pioneer 11, which launched on April 6, 1973, flew by Jupiter on December 2, 1974, and then went on to Saturn on September 1, 1979 (just months before Voyager 1 and 2). The Martian Viking Landers 1 and 2 used a pair of modified SNAP-19’s each as well in their missions.
The next major use of TAG-85 was in the Multi-Mission RTG (MMRTG), which is a true workhorse of American astronuclear engineering, currently powering the Curiosity rover. His legacy will fly again on the Mars 2020 rover, currently under construction at the Jet Propulsion Laboratory.
Current RTG design work is shifting away from the solid-state conversion systems so long favored by NASA, due to its low power conversion efficiency. This is nothing new, but the inherent simplicity of the solid-state systems have dominated the available supply of radioisotope power systems, and the mission needs of NASA, the USAF, and other major customers that using a heat engine to produce electricity has only been explored so far. The upcoming power conversion series will deal with these options in detail.
Thank You, Dr. Skrabek
While RTGs may not be the big, exciting power supplies that we often discuss here on Beyond NERVA, they have literally opened the outer solar system to our understanding, powering the missions that have amazed us all, no matter our level of education, and the knowledge and beauty we’ve all gained is due in part to Dr. Skrabek’s discovery and subsequent work on these systems. The bread and butter power supply for the outer solar system, and one that is powering the most advanced rover ever built by humanity on Mars, is possible thanks to his ideas, and his hard work.
While there is a transition happening to heat engine based RPUs (radioisotope power units, the broader category that a Stirling or Rankine powered RTG would fall under), this does not mean that the traditional RTG is going anywhere any time soon. Their inherent stability, durability, and ruggedness, combined with their ability to power a rover the size of a small truck around Mars, vaporizing rock at a distance with a laser beam to analyze its composition, is not something to be cast aside lightly.
Even if TAGS-85 never flies after Mars 2020 (something I very much doubt), his work will continue to inform us every day about the environment, both past and present, of Mars for years to come. Five of his power supplies (two each on Viking 1 and 2, one on MSL) will, all things being equal, end their days on the Red Planet, with a sixth on its way in another year. His work made us able to open our eyes on the beauty of the outer solar system, showed us Pluto in fascinating detail for the first time, and literally pioneered the path of the Voyager probes at Saturn (Pioneer 11 inserted in the same orbit to verify that particle density was within safe limits), and is now flying out of the solar system in two different directions. One small, but crucial, piece of materials engineering allowed these spacecraft and rovers to do more than they would be able to with other materials, and open our eyes that much more.
Thank you, Dr. Skrabek, for your life’s work. Your memory will live on in all the missions you have, are, and will make possible, and the knowledge that you’ve helped bring humanity as a whole.
Here are some of the pictures that Dr Skrabek helped enable:
Hello, and welcome back to Beyond NERVA! Today, we’re going to return to our discussion of fission power plants, and look at a program that was unique in the history of astronuclear engineering: a Soviet-designed and -built reactor design that was purchased and mostly flight-qualified by the US for an American lunar base. This was the Enisy, known in the West as Topaz-II, and the Topaz International program.
This will be a series of three posts on the system: this post focuses on the history of the reactor in the Soviet Union, including the testing history – which as we’ll see, heavily influenced the final design of the reactor. The next will look at the Topaz International program, which began as early as 1980, while the Soviet Union still appeared strong. Finally, we’ll look at two American uses for the reactor: as a test-bed reactor system for a nuclear electric test satellite, and as a power supply for a crewed lunar base. This fascinating system, and the programs associated with it, definitely deserve a deep dive – so let’s jump right in!
We’ve looked at the history of Soviet astronuclear engineering, and their extensive mission history. The last two of these reactors were the Topaz (Topol) reactors, on the Plasma-A satellites. These reactors used a very interesting type of power conversion system: an in-core thermionic system. Thermionic power conversion takes advantage of the fact that certain materials, when heated, eject electrons, gaining a positive static charge as whatever the electrons impact gain a negative charge. Because the materials required for a thermionic system can be made incredibly neutronically robust, they can be placed inside the core of the reactor itself! This is a concept that I’ve loved since I first heard of it, and remains as cool today as it did back then.
The original Topaz reactor used a multi-cell thermionic element concept, where fuel elements were stacked in individual thermionic conversion elements, and several of these were placed end-to-end to form the length of the core. While this is a perfectly acceptable way to set up one of these systems, there are also inefficiencies and complexities associated with so many individual fuel elements. An alternative would be to make a single, full-length thermionic cell, and use either one or several fuel rods inside the thermionic element. This is the – wait for it – single cell thermionic element design, and is the one that was chosen for the Enisy/Topaz-II reactor (which we’ll call Enisy in this post, since it’s focusing on the Soviet history of the reactor). While started in 1967, and tested thoroughly in the 70s, it wasn’t flight-qualified until the 80s… and then the Soviet Union collapsed, and the program died.
After the fall of the USSR, there was a concerted effort by the US to keep the specialist engineers and scientists of the former Soviet republics employed (to ensure they didn’t find work for international bad actors such as North Korea), and to see what technology had been developed behind the Iron Curtain that could be purchased for use by the US. This is where the RD-180 rocket engine, still in use by the United Launch Alliance Atlas rockets, came from. Another part of this program, though, focused on the extensive experience that the Soviets had in astronuclear missions, and in paricular the most advanced – but as yet unflown – design of the renowned NPO Luch design bureau, attached to the Ministry of Medium Industry: the Enisy reactor (which had the US designation of Topaz-II due to early confusion about the design by American observers).
The Enisy, in its final iteration, was designed to have a thermal output of 115 kWt (at the beginning of life), with a mission requirement of at least 6 kWe at the electrical outlet terminals for at least three years. Additional requirements included a ten year shelf life after construction (without fissile fuel, coolant, or other volatiles loaded), a maximum mass of 1061 kg, and prevention of criticality before achieving orbit (which was complicated from an American point of view, more on that below). The coolant for the reactor remained NaK-78, a common coolant in most reactors we’ve looked at so far. Cesium was stored in a reservoir at the “bottom” (away from the spacecraft) end of the reactor vessel, to ensure the proper partial pressure between the cathode and anode of the fuel elements, which would leak out over time (about 0.5 g/day during operation). This was meant to be the next upgrade in the Soviet astronuclear fleet, and as such was definitely a step above the Topaz-I reactor.
Perhaps the most interesting part of the design is that it was designed to be able to be tested as a complete system without the use of fissile fuels in the reactor. Instead, electrical resistance heaters could be inserted in the thermionic fuel elements to simulate the fission process, allowing for far more complete testing of the system in flight configuration before launch. This design decision heavily influenced US nuclear power plant design and testing procedures, and continues to influence designs today (the induction heating testing of the KRUSTY thermal simulator is a good recent example of this concept, even if it’s been heavily modified for the different reactor geometry), however, the fact that the reactor used cylindrical fuel elements made this process much easier.
So what did the Enisy look like? This changed over time, but we will look at the basics of the power plant’s design in its final Soviet iteration in this post, and the examine the changes that the Americans made during the collaboration in the next post. We’ll also look at why the design changed as it did.
First, though, we need to look at how the system worked, since compared to every system that we’ve looked at in depth, the physics behind the power conversion system are quite novel.
Thermionics: How to Keep Your Power Conversion System in the Core
We haven’t looked at power conversion systems much in this blog yet, but this is a good place to discuss the first kind as it’s so integral to this reactor. If the details of how the power conversion system actually worked don’t interest you, feel free to skip to the next section, but for many people interested in astronuclear design this power conversion system offers the promise to potentially be the most efficient and reliable option available for in-space nuclear reactors geared towards electricity production.
In short, thermionic reactions are those that occur when a material is heated and gives off charged particles. This is something that has been known since ancient times, even though the physical mechanism was completely unknown until after the discovery of the electron. The name comes from the term “thermions,” or “thermal ions.” One of the first to describe this effect used a hot anode in a vacuum: the modern incandescent lightbulb: Thomas Edison, who observed a static charge building up on the glass of his bulbs while they were turned on. However, today this has expanded to include the use of anodes, as well as solid-state systems and systems that don’t have a vacuum.
The efficiency of these systems depends on the temperature difference between the anode and cathode, the work function (or minimum thermodynamic work needed to remove an electron from a solid to a vacuum immediately outside the solid surface) of the emitter used, and the Boltzmann Constant (which relates to the average kinetic energy of particles in a gas), as well as a number of other factors. In modern systems, however, the structure of a thermionic convertor which isn’t completely solid state is fairly standard: a hot cathode is separated from a cold anode, with cesium vapor in between. For nuclear systems, the anode is often tungsten, the cathode seems to vary depending on the system, and the gap between – called the inter-electrode gap – is system specific.
The cesium exists in an interesting state of matter. Solid, liquid, gas, and plasma are familiar to pretty much everyone at this point, but other states exist under unusual circumstances; perhaps the best known is a supercritical fluid, which exhibits the properties of both a liquid and a gas (although this is a range of possibilities, with some having more liquid properties and some more gaseous). The one that concerns us today is something called Rydberg matter, one of the more exotic forms of matter – although it has been observed in many places across the universe. In its simplest form, Rydberg matter can be seen as small clusters of interconnected molecules within a gas (the largest number of atoms observed in a laboratory is 91, according to Wikipedia, although there’s evidence for far larger numbers in interstellar gas clouds). These clumps end up affecting the electron clouds of those atoms in the clusters, causing them to orbit across the nuclei of those atoms, causing a new lowest-energy state for the entire cluster to occur. These structures don’t degrade any faster under radioactive bombardment due to a number of quantum mechanical properties, which brought them to the attention of the Los Alamos Scientific Laboratory staff in the 1950s, and a short time later Soviet nuclear physicists as well.
This sounds complex, and it is, but the key point is this: because the clumps act as a unit within Rydberg matter, their ability to transmit electricity is enhanced compared to other gasses. In particular, cesium seems to be a very good vehicle for creating Rydberg matter, and cesium vapor seems to be the best available for the gap between the cathode and anode of a thermionic convertor. The density of the cesium vapor is variable and dependent on many factors, including the materials properties of the cathode and anode, the temperature of the cathode, the inter-electrode gap distance, and a number of other factors. Tuning the amount of cesium in the inter-electrode gap is something that must occur in any thermionic power conversion system; in fact the original version of the Enisy had the ability to vary the inter-electrode gap pressure (this was later dropped when it was discovered to be superfluous to the efficient function of the reactor).
This type of system comes in two varieties: in-core and out-of-core. The out-of-core variant is very similar to the power conversion systems we saw (briefly) on the SNAP systems: the coolant from the reactor passes around or through the radiation shield of the system, heats the anode, which then emits electrons into the gap, collected by the cathode, and then the electricity goes through the power conditioning unit and into the electrical system of the spacecraft. Because thermionic conversion is theoretically more efficient, and in practice is more flexible in temperature range, than thermoelectric conversion, even keeping the configuration of the power conversion system’s relationship to the rest of the power plant offers some advantages.
The in-core variant, on the other hand, wraps the power conversion system directly around the fissile fuel in the core, with electrical power being conducted out of the core itself and through the shield. The coolant runs across the outside of the thermionic unit, providing the thermal gradient for the system to work, and then exits the reactor. While this increases the volume of the core (admittedly, not by much), it also eliminates the need for more complex plumbing for the primary coolant loop. Additionally, it allows for less heat loss from the coolant having to travel a farther difference. Finally, there’s far less chance of a stray meteor hitting your power conversion system and causing problems – if a thermionic fuel element is damaged by a foreign object, you’re going to have far bigger problems with the system as a whole, since it means that it damaged your control systems and pressure vessel on the way to damaging your power conversion unit!
The in-core thermionic power conversion system, while originally proposed by the US, was seen as a curiosity on their side of the Iron Curtain. Some designs were proposed, but none were significantly researched to the level of being able to be serious contenders in the struggle to gain the significant funding needed to develop as complex a system as an astronuclear fission power plant, and the low conversion efficiency available in practice prevents its application in terrestrial power plants, which to this day continue to use steam turbine generators.
On the other side of the Iron Curtain, however, this was seen as the ideal solution for a power conversion system: the only systems needed for the system to work could be solid-state, with no moving parts: heaters to vaporize the cesium, and electromagnetic pumps to move it through the reactor. Greater radiation resistance and more flexible operating temperatures, as well as greater conversion efficiency, all offered more promise to Soviet astronuclear systems designers than the thermoelectric path that the US ended up following. The first Soviet reactor designed for in-space use, the Romashka, used a thermionic power conversion system, but the challenges involved in the system itself led the Krasnya Zvezda design bureau (who were responsible for the Romasha, Bouk, and Topol reactors) to initially choose to use thermoelectric convertors in their first flight system: the BES-5 Bouk, which we’ve seen before.
Now that we’ve looked at the physics behind how you can place your power conversion system within the reactor vessel of your power plant (and as far as I’ve been able to determine, if you’re looking to generate electricity beyond what a simple sensor needs, this is the only option without going to something very exotic), let’s look at the reactor itself.
Enisy: The Design of the TOPAZ-II Reactor
The Enisy was a uranium oxide fueled, zirconium hydride moderated, sodium-potassium eutectic cooled reactor, which used a single-element thermionic fuel element design for in-core power conversion. The multi-cell version was used in the Topol reactor, where each fuel pellet was wrapped in its own thermionic convertor. This is sometimes called a “flashlight” configuration, since it looks a bit like the batteries in a large flashlight, but this comes at the cost of complexity, mass, and increased inefficiencies. To offset this, many issues are easier to deal with in this configuration, especially as your fuel reaches higher burnup percentages and your fuel swells. The ultimate goal was single-unit thermionic fuel elements, which were realized in the Enisy reactor. While more challenging in terms of materials requirements, the greater simplicity, lower mass, and greater efficiency of the system offered more promise.
The power plant was required to provide 6 kWe of electrical power at the reactor terminals (before the power conditioning unit) at 27 volts. It had to have an operational life of three years, and a storage life if not immediately used in a mission of at least ten years. It also had to have an operational reliability of >95%, and could not under any circumstances achieve criticality before reaching orbit, nor could the coolant freeze at any time during operation. Finally, it had to do all of this in less than 1061 kg (excluding the automatic control system).
Thirty-seven fuel elements were used in the core, which was contained in a stainless steel reactor vessel. These contained uranium oxide fuel pellets, with a central fission gas void about 22% of the diameter of the fuel pellets to prevent swelling as fission products built up. The emitters were made out of molybdenum, a fairly common choice for in-core applications. Al2O3 (sapphire) insulators were used to electrically isolate the fuel elements from the rest of the core. Three of these would be used to power the cesium heater and pump directly, while another (unknown) number powered the NaK coolant pump (my suspicion is that it’s about the same number). The rest would output power directly from the element into the power conditioning unit on the far side of the power plant.
Nine control drums, made mostly out of beryllium but with a neutron poison along one portion of the outer surface (Boron carbide/silicon carbide) surrounded the core. Three of these drums were safety drums, with two positions: in, with the neutron poison facing the center of the core, and out, where the beryllium acted as a neutron reflector. The rest of the drums could be rotated in or out as needed to maintain reactivity at the appropriate level in the core. These had actuators mounted outside the pressure vessel to control the rotation of the drums, and were connected to an automatic control system to ensure autonomous stable function of the reactor within the mission profile that the reactor would be required to support.
The NaK coolant would flow around the fuel elements, driven by an electromagnetic pump, and then pass through a radiator, in an annular flow path immediately surrounding the TFEs. Two inlet and two outlet pipes were used to connect the core to the radiator. In between the radiator and the core was a radiation shield, made up of stainless steel and lithium hydride (more on this seemingly odd choice when we look at the testing history).
The coolant tubes were embedded in a zirconium hydride moderator, which was contained in stainless steel casings.
Finally, a reservoir of cesium was at the opposite end of the reactor from the radiator. This was necessary for the proper functioning of the thermionic fuel elements, and underwent many changes throughout the design history of the reactor, including a significant expansion as the design life requirements increased.
Once the Topaz International program began, additional – and quite significant – changes were made to the reactor’s design, including a new automated control system and an anti-criticality system that actually removed some of the fuel from the core until the start-up commands were sent, but that’s a discussion for the next post.
I saved the coolest part of this system for last: the TISA, or “Thermal Simulators of Apparatus Cores” (the acronym was from the original Russian), heaters. These units were placed in the active section of the thermionic fuel elements to simulate the heat of fission occurring in the thermionic fuel elements, with the rest of the systems and subsystems being in flight configuration. This led to unprecedented levels of testing capability, but at the same time would lead to a couple of problems later in testing – which would be addressed as needed.
How did this design end up this way? In order to understand that, the development and testing process of the Soviet design team must be looked at.
The History of Enisy’s Design
The Enisy reactor started with the development of the thermionic fuel element by the Sukhumi Institute in the early 1960s, which had two options: the single cell and multiple cell variants. In 1967, these two options were split into two different programs: the Topol (Topaz), which we looked at in the Soviet Astronuclear History post, led by the Krasnaya Zvezda design bureau in Moscow, and Enisy, which was headed by the Central Design Bureau of Machine Building in Leningrad (now St. Petersburg). Aside from the lead bureau, in charge of the overall program and system management, a number of other organizations were involved with the fabrication and testing of the reactor system: the design and modeling team consisted of: the Kurchatov Institute of Atomic Energy was responsible for nuclear design and analytics, the Scientific Industrial Association Lutch was responsible for the thermionic fuel elements, the Sukhumi Institute remained involved in the reactor’s automatic control systems design; fabrication and testing was the responsibility of: the Research Institute of Chemical Machine Building for thermal vacuum testing, the Scientific Institute for Instrument Building’s Turaevo nuclear test facility, Kraznoyarsk Spacecraft Designer for mechanical testing and spacecraft integration, Prometheus Laboratory for materials development (including liquid metal eutectic development for the cooling system and materials testing) and welding, and the Enisy manufacturing facility was located in Talinn, Estonia (a decision that would cause later headaches during the collaboration).
The Enisy originally had three customers (the identities of which I am not aware of, simply that at least one was military), and each had different requirements for the reactor. Originally designed to operate at 6 kWe for one year with a >95% success rate, but customer requirements changed both of these characteristics significantly. As an example, one customer needed a one year system life, with a 6 kWe power output, while another only needed 5 kWe – but needed a three year mission lifetime. This longer lifetime ended up becoming the baseline requirement of the system, although the 6 kWe requirement and >95% mission success rate remained unchanged. This led to numerous changes, especially to the cesium reservoir needed for the thermionic convertors, as well as insulators, sensors, and other key components in the reactor itself. As the cherry on top, the manufacture of the system was moved from Moscow to Talinn, Estonia, resulting in a new set of technicians needing to be trained to the specific requirements of the system, changes in documentation, and at the fall of the Soviet Union loss of significant program documentation which could have assisted the Russia/US collaboration on the system.
The nuclear design side of things changed throughout the design life as well. An increase in the number of thermionic fuel elements (TFEs) occurred in 1974, from 31 to 37 in the reactor core, an increase in the height of the “active” section of the TFE, although whether the overall TFE length (and therefore the core length) changed is information I have not been able to find. Additional space in the TFEs was added to account for greater fuel swelling as fission products built up in the fuel pellets, and the bellows used to ensure proper fitting of the TFEs with reactor components were modified as well. The moderator blocks in the core, made out of zirconium hydride, were modified at least twice, including changing the material that the moderator was kept in. Manufacturing changes in the stainless steel reactor vessel were also required, as were changes to the gamma shielding design for the shadow shield. All in all, the reactor went through significant changes from the first model tested to theend of its design life.
Another area with significantly changing requirements was the systems integration side of things. The reactor was initially meant to be launched in a reactor-up position, but this was changed in 1979 to a reactor-down launch configuration, necessitating changes to several systems in what ended up being a significant effort. Another change in the launch integration requirements was an increase in the acceleration levels required during dynamic testing by a factor of almost two, resulting in failures in testing – and resultant redesigns of many of the structures used in the system. Another thing that changed was the boom that mounted the power plant to the spacecraft – three different designs were used through the lifetime of the system on the Russian side of things, and doubtless another two (at least) were needed for the American spacecraft integration.
Perhaps the most changed design was the coolant loop, due to significant problems during testing and manufacturing of the system.
Design Driven by (Expected) Failure: The USSR Testing Program
Flight qualification for nuclear reactors in the USSR at the time was very different from the way that the US did flight qualification, something that we’ll look at a bit more later in this post. The Soviet method of flight qualification was to heavily test a number of test-beds, using both nuclear and non-nuclear techniques, to validate the design parameters. However, the actual flight articles themselves weren’t subjected to nearly the same level of testing that the American systems would be, instead going through a relatively “basic” (according to US sources) workmanship examination before any theoretical launch.
In the US, extensive systems modeling is a routine part of nuclear design of any sort, as well as astronautical design. Failures are not unexpected, but at the same time the ideal is that the system has been studied and modeled mathematically thoroughly enough that it’s not unreasonable to predict that the system will function correctly the first time… and the second… and so on. This takes not only a large amount of skilled intellectual and manual labor to achieve, but also significant computational capabilities.
In the Soviet Union, however, the preferred method of astronautical – and astronuclear – development was to build what seemed to be a well-designed system and then test it, expecting failure. Once this happened, the causes of the failure were analyzed, the problem corrected, and then the newly upgraded design would be tested again… and again, for as many times as were needed to develop a robust system. Failure was literally built into the development process, and while it could be frustrating to correct the problems that occurred, the design team knew that the way their system could fail had been thoroughly examined, leading to a more reliable end result.
This design philosophy leads to a large number of each system needing to be built. Each reactor that was built underwent a post-manufacturing examination to determine the quality of the fabrication in the system, and from this the appropriate use of the reactor. These systems had four prefixes: SM, V, Ya, and Eh. Each system in this order was able to do everything that the previous reactor would be able to do, in addition to having superior capabilities to the previous type. The SM, or static mockup, articles were never built for anything but mechanical testing, and as such were stripped down, “boilerplate” versions of the system. The V reactors were the next step up, which were used for thermophysical (heat transfer, vibration testing, etc) or mechanical testing, but were not of sufficient quality to undergo nuclear testing. The Ya reactors were suitable for use in nuclear testing as well, and in a pinch would be able to be used in flight. The Eh reactors were the highest quality, and were designated potential flight systems.
In addition to this designation, there were four distinct generations of reactor: the first generation was from V-11 to Ya-22. This core used 31 thermionic fuel elements, with a one year design life. They were intended to be launched upright, and had a lightweight radiation shield. The next generation, V-15 to Ya-26, the operational lifetime was increased to a year and a half.
The third generation, V-71 to Eh-42 had a number of changes. The number of TFEs was increased from 31 to 37, in large part to accommodate another increase in design life, to above 3 years. The emitters on the TFEs were changed to the monocrystaline Mo emitters, and the later ones had Nb added to the Mo (more on this below). The ground testing thermal power level was reduced, to address thermal damage from the heating units in earlier non-nuclear tests. This is also when the launch configuration was changed from upright to inverted, necessitating changes in the freeze-prevention thermal shield, integration boom, and radiator mounting brackets. The last two of this generation, Eh -41 and Eh-42, had the heavier radiation shield installed, while the rest used the earlier, lighter gamma shield.
The final generation, Ya-21u to Eh-44, had the longest core lifetime requirement of three years at 5.5 kWe power output. These included all of the other changes above, as well as many smaller changes to the reactor vessel, mounting brackets, and other mechanical components. Most of these systems ended up becoming either Ya or Eh units due to lessons learned in the previous three generations, and all of the units which would later be purchased by the US as flight units came from this final generation.
A total of 29 articles were built by 1992, when the US became involved in the program. As of 1992, two of the units were not completed, and one was never assembled into its completed configuration.
Sixteen of the 21 units were tested between 1970 and 1989, providing an extensive experimental record of the reactor type. Of these tests, thirteen underwent thermal, mechanical, and integration non-nuclear testing. Nuclear testing occurred six times at the Baikal nuclear facility. As of 1992, there were two built, but untested, flight units available: the E-43 and E-44, with the E-45 still under construction.
Nuclear testing revision and development, including fuel loading, radiation and nuclear safety. Studied unstable nuclear conditions and stainless steel material properties, disassembly and inspection. LiH moderator hydrogen loss in test.
Nuclear ground test. ACS startup, steady-state functioning, post-operation disassembly and inspection. TFE lifetime limited to ~2 months due to fuel swelling
Steady state nuclear testing. Significant TFE shortening post-irradiation.
TFE needed redesign, no systems testing. Installed at Turaevo as mockup. Used to establish transport and handling procedures
System incomplete. Used as spacecraft mockup, did not undergo physical testing.
Test stand preparation
Second fabrication stage not completed. Used for some experiments with Baikal test stand. Disassembled in Sosnovivord.
Refabricated at CDBMB. TFE burnt and damaged during second fadrication. Notch between TISA and emitter
# of TFEs
Mechanical, Electrical, Spacecraft integration
Baikal, Krasnoyarsk, Cold Temp Testing
Converted from upright to inverted launch configuration, spacecraft integration heavily modified. First to use 37 TFE core configuration. Transport testing (railroad vibration and shock), cold temperature testing. Electrical testing post-mechanical. Zero power testing at Krasnoyarsk.
Ground control (no ACS)
Nuclear ground test, steady state operation. Leaks observed in two cooling pipes 120 hrs into test; leaks plugged and test continued. Disassembly and inspection.
Prototype Sukhumi ACS
Nuclear ground test, startup using ACS, steady state. Initial leak in EM pump led to large leak later in test. Test ended in loss of coolant accident. Reactor disassembled and inspected post-test to determine leak cause.
Quality not sufficient for flight (despite Eh “flight” designation). Static and torsion tests conducted.
Nuclear ground test, pre-launch simulation. ACS startup and operation. Steady state test. Post-operation disassembly and examination.
Fabrication begin in Estonia, with some changed components. After changes, system name changed to Eh-41, and serial number changed to 17. Significant reactor changes.
Cold temp, coolant flow
Cold temperature testing. No electrical testing. Filled with NaK during second stage of fabrication.
Began life as Eh(?)-39, post-retrofit designation. Transportation (railroad) dynamic, and impact testing. Leak testing done post-mechanical testing. First use of increased shield mass.
Critical component welding failure during fabrication. Unit never used.
# of TFEs
First Gen 4 reactor using modified TFEs. Electrical testing on TFEs conducted. New end-cap insulation on TFEs tested.
6/30/88 (? Unclear what testing is indicated)
Flight unit. First fabrication phase in Talinn completed, second incomplete as of 1994
Flight unit. First fabrication phase in Talinn completed, second incomplete as of 1994
Partially fabricated unit with missing components.
Not many fine details are known about the testing of these systems, but we do have some information about the tests that led to significant design changes. These changes are best broken down by power plant subsystem, because while there’s significant interplay between these various subsystems their functionality can change in minor ways quite easily without affecting the plant as a whole. Those systems are: the thermionic fuel elements, the moderator, the pressure vessel, the shield, the coolant loop (which includes the radiator piping), the radiator coatings, the launch configuration, the cesium unit, and the automatic control system (including the sensors for the system and the drum drive units). While this seems like a lot of systems to cover, many of them have very little information about their design history to pass on, so it’s less daunting than it initially appears.
Thermionic Fuel Elements
It should come as no surprise that the thermionic fuel elements (TFEs) were extensively modified throughout the testing program. One of the big problems was short circuiting across the inter-electrode gap due to fuel swelling, although other problems occurred to cause short circuits as well.
Perhaps the biggest change was the change from 31 to 37 TFEs in the core, one of the major changes to minimize fuel swelling. The active core length (where the pellets were) was increased by up to 40 mm (from 335 mm to 375 mm), the inter-electrode gap was widened by 0.05 mm (from 0.45 to 0.5 mm). In addition, the hole through the center of the fuel element was increased in diameter to allow for greater internal swelling, reducing the mechanical stress on the emitter.
The method of attaching the bellows for thermal expansion were modified (the temperature was dropped 10 K) to prevent crystalization of the palladium braze and increase bellows thermal cycling capability after failures on the Ya-24 system (1977-1981).
Perhaps the biggest change was to the materials used in the TFE. The emitter started off as a polycrystaline molybdenum in the first two generations of reactors, but the grain boundaries between the Mo crystals caused brittleness over time. Because of this, they developed the capability to use monocrystalline Mo, which improved performance in the early third generation of reactors – just not enough. In the final version seen in later 3rd generation and fourth generation systems, the Mo was doped with 3% niobium, which created the best available material for the emitter.
There were many other changes during the development of the thermionic fuel elements, including the addition of coatings on some materials for corrosion resistance, changes in electrical insulation type, and others, but these were the most significant in terms of functionality of the TFEs, and their impact on the overall systems design.
The zirconium hydride neutron moderator was placed around the outside of the core. Failures were observed several times in testing, including the Ya-23 test, which resulted in loss of hydrogen in the core and the permanent shutdown of that reactor. Overpower issues, combined with a loss of coolant, led to moderator failure in Ya-82 as well, but in this case the improved H barriers used in the stainless steel “cans” holding the ZrH prevented a loss of hydrogen accident despite the ZrH breaking up (the failure was due to the ZrH being spread more thinly across the reactor, not the loss of H due to ZrH damage).
This development process was one of the least well documented areas of the Soviet program.
Again, this subsystem’s development seems poorly documented. The biggest change, though, seems to be changing the way the triple coating (of chrome, then nickel, then enamel) was applied to the stainless steel of the reactor vessel. This was due to the failure of the Ya-23 unit, which failed at the join between the tube and the end of the tube on one of the TFEs. The crack self-sealed, but for future units the coatings didn’t go all the way to the weld, and the hot CO2 used as a cover gas was allowed to carbonize the steel to prevent fatigue cracking.
The LiH component of the radiation shield (for neutron shielding) seems to not have changed much throughout the development of the reactor. The LiH was contained in a 1.5 mm thick stainless steel casing, polished on the ends for reflectivity and coated black on the outside face.
However, the design of the stainless steel casing was changed in the early 1980s to meet more stringent payload gamma radiation doses. Rather than add a new material such as tungsten or depleted uranium as is typical, the designers decided to just thicken the reactor and spacecraft sides of the LiH can to 65 mm and 60 mm respectively. While this was definitely less mass-efficient than using W or U, the manufacturing change was fairly trivial to do with stainless steel, and this was considered the most effective way to ensure the required flux rates with the minimum of engineering challenges.
The first unit to use this was the E-41, fabricated in 1985, which was also the first unit to be tested in the inverted flight configuration. The heavier shield, combined with the new position, led to the failure of one of the shield-to-reactor brackets, as well as the attachment clips for the radiator piping. These components were changed, and no further challenges occurred with the shield in the rest of the test program.
The NaK coolant loop was the biggest source of headaches dueing the development of the Enisy. A brief list of failures, and actions taken to correct them, is here:
V-11 (July 1971-February 1972): A weld failed at the join between the radiator tubing and collector during thermophysical testing. The double weld was changed to a triple weld to correct the failure mode.
Ya-21 (1971): This reactor seemed to have everything go wrong with it. Another leak at the same tube-to-collector interface led to the welding on of a small sleeve to repair the crack. This fix seemed to solve the problem of failures in that location.
Ya-23 (March 1975-June 1976): Coolant leak between coolant tube and moderator cavity. Both coating changes and power ramp-up limits eliminated issues.
V-71 (January 1981-1994?): NaK leak in radiator tube after 290 hours of testing. Plugged, testing continued. New leak occurred 210 test hours later, radiator examined under x-ray. Two additional poorly-manufactured tubes replaced with structural supports. One of test reactors sent to US under Topaz International.
Ya-81 (September 1980-January 1983): Two radiator pipe leaks 180 hours into nuclear testing (no pre-nuclear thermophysical testing of unit). Piping determined to be of lower quality after switching manufacturers. Post-repair, the unit ran for 12,500 hours in nuclear power operation.
Ya-82 (September 1983 to November 1984): Slow leak led to coolant pump voiding and oscillations, then one of six pump inlet lines being split. There were two additional contributions to this failure: the square surfaces were pressed into shape from square pipes, which can cause stress microfractures at the corners, and second the inlet pump was forced into place, causing stress fracturing at the joint. This failure led to reactor overheating due to a loss-of-coolant condition, and led to the failure of the ZrH moderator blocks. This led to increased manufacturing controls on the pump assembly, and no further major pump failures were noted in the remainder of the testing.
Eh-38 (February 1986-August 1986): This failure is a source of some debate among the Russian specialists. Some believe it was a slow leak that began shortly after startup, while others believe that it was a larger leak that started at some point toward the end of the 4700 hour nuclear test. The exact location of the leak was never located, however it’s known that it was in the upper collector of the radiator assembly.
Ya-21u (December 1987-December 1989): Caustic stress-corrosion cracking occurred about a month and a half into thermophysical testing in the lower collector assembly, likely caused by a coating flaw growing during thermal cycling. This means that subsurface residual stresses existed within the collector itself. Due to the higher-than-typical (by U.S. standards) carbon content in the stainless steel (the specification allowed for 0.08%-0.12% carbon, rather than the less than 0.8% carbon content in the U.S. SS-321), the steel was less ductile than was ideal, which could have been a source of the flaw growing as it did. Additionally, increased oxygen levels in the NaK coolant could have exacerbated the problem more as well. A combination of ensuring that heat treatments had occurred post-forming, as well as ensuring a more oxygen-poor environment, were essential to reducing the chances of this failure happening again.
The only known data poing on the radiator development was during the Ya-23 test, where the radiator coating changed the nuclear properties of the system at elevated temperature (how is unknown). This was changed to something that would be less affected by the radiation environment. The final radiator configuration was a chrome and polymer substrate with an emissivity of 0.85 at beginning of life.
As we saw, the orientation that the reactor was to be launched in was changed from upright to inverted, with the boom to connect the reactor to the spacecraft being side by side inside the payload fairing. This required the thermal cover used to prevent the NaK from freezing to be redesigned, and modified after the V-13 test, when it was discovered to not be able to prevent freezing of the coolant. The new cover was verified on the V-15 tests, and remained largely unchanged after this.
Some of the load-bearing brackets needed to be changed or reinforced as well, and the clips used to secure the radiator pipes to the structural components of the radiator.
Cesium Supply Unit
For the TFEs to work properly, it was critical that the Cs vapor pressure was within the right pressure range relative tot he temperature of the reactor core. This system was designed from first physical principles, leading to a novel structure that used temperature and pressure gradients to operate. While initially throttleable, but there were issues with this functionality during the Ya-24 nuclear test. This changed when it was discovered that there was an ideal pressure setting for all power levels, so the feed pressure was fixed. Sadly, on the Ya-81 test the throttle was set too high, leading to the need to cool the Cs as it returned to the reservoir.
Additional issues were found in the startup subsystem (a single-use puncture valve) used to vent the inert He gas from the interelectrode gap (this was used during launch and before startup to prevent Cs from liquefying or freezing in the system), as well as to balance the Cs pressure by venting it into space at a rate of about 0.4 g/day. The Ya-23 test saw a sensor not register the release of the He, leading to an upgraded spring for the valve.
Finally, the mission lifetime extension during the 1985/86 timeframe tripled the required lifetime of the system, necessitating a much larger Cs reservoir to account for Cs venting. This went from having 0.455 g to 1 kg. These were tested on Ya-21u and Eh-44, despite one (military) customer objecting due to insufficient testing of the upgraded system. This system would later be tested and found to be acceptable as part of the Topaz International program.
Automatic Control System
The automatic control system, or ACS, was used for automatic startup and autonomous reactor power management, and went through more significant changes than any other system, save perhaps the thermionic fuel elements. The first ACS, called the SAU-35, was used for the Ya-23 ground test, followed by the SAU-105 for the Eh-31 and Ya-24 tests. Problems arose, however, because these systems were manufactured by the Institute for Instrument Building of the Ministry of Aviation Construction, while the Enisy program was under the purview of the Ministry of Atomic Energy, and bureaucratic problems reared their heads.
This led the Enisy program to look to the Sukhumi Institute (who, if you remember, were the institute that started both the Topol and Enisy programs in the 1960s before control was transferred elsewhere) for the next generation of ACS. During this transition, the Ya-81 ground nuclear test occurred, but due to the bureaucratic wrangling, manufacturer change, and ACS certification tests there was no unit available for the test. This led the Ya-81 reactor to be controlled from the ground station. The Ya-82 test was the first to use a prototype Sukhumi-built ACS, with nine startups being successfully performed by this unit.
The loss-of-cooling accident potentially led to the final major change to the ACS for the Eh-38 test: the establishment of an upper temperature limit. After this, the dead-band was increased to allow greater power drift in the reactor (reducing the necessary control drum movement), as well as some minor modifications rerouting the wires to ensure proper thermocouple sensor readings, were the final significant modifications before Topaz International started.
The sensors on the Enisy seem to have been regularly problematic, but rather than replace them, they were either removed or left as instrumentation sensors rather than control sensors. These included the volume accumulator sensors on the stainless steel bellows for the thermionic fuel elements (which were removed), and the set of sensors used to monitor the He gas in the TFE gas gap (for fission product buildup), the volume accumulator (which also contained Ar), and the radiation shield. This second set of sensors was kept in place, but was only able to measure absolute changes, not precise measurements, so was not useful for the ACS.
Control Drive Unit
The control drive unit was responsible for the positioning of the control drums, both on startup as well as throughout the life of the reactor to maintain appropriate reactivity and power levels. Like in the SNAP program, these drive systems were a source of engineering headaches.
Perhaps the most recurring problem during the mid-1970s was the failure of the position sensor for the drive system, which was used to monitor the rotational position of the drum relative to the core. This failed in the Ya-20, Ya-21, and Ya-23, after which it was replaced with a sensor of a new design and the problem isn’t reported again. The Ya-81 test saw the loss of the Ar gas used as the initial lubricant in the drive system, and later seizing of the bearing the drive system connected to, leading to its replacement with a graphite-based lubricant.
The news wasn’t all bad, however. The Eh-40 test demonstrated greater control of drum position by reducing the backlash in the kinematic circuit, for instance, and improvements to the materials and coatings used eliminated problems of coating delamination, improving the system’s resistance to thermal cycling and vibrational stresses, and radiator coating issues.
The Eh-44 drive unit was replaced against the advice of one of the Russian customers due to a lack of mandatory testing on the advanced drive system. This system remained installed at the time of Topaz International, and is something that we’ll look at in the next blog post.
A New Customer Enters the Fold
During this testing, an American company (which is not named) was approached about possibly purchasing nearly complete Enisy reactors: the only thing that the Soviets wouldn’t sell was the fissile fuel itself, and that they would help with the manufacturing on. This was in addition to the three Russian customers (at least one of which was military, but again all remain unnamed). This company did not purchase any units, but did go to the US government with this offer.
This led to the Topaz International program, funded by the US Department of Defense’s Ballistic Missile Defense Organization. The majority of the personnel involved were employees of Los Alamos and Sandia National Laboratories, and the testing occurred at Kirtland Air Force Base in Albuquerque, NM.
As a personal note, I was just outside the perimeter fence when the aircraft carrying the test stand and reactors landed, and it remains one of the formational events in my childhood, even though I had only the vaguest understanding of what was actually happening, or that some day, more than 20 years, later, I would be writing about this very program, which I saw reach a major inflection point.
The Topaz International program will be the subject of our next blog post. It’s likely to be a longer one (as this was), so it may take me a little longer than a week to get out, but the ability to compare and contrast Soviet and American testing standards on the same system is too golden an opportunity to pass up.
Hello, and welcome to Beyond NERVA, for our first blog post of the year! Today, we reach the end of the reactor portion of the SNAP program. A combination of the holidays and personal circumstances prevented me from finishing this post as early as I would have liked to, but it’s finally here! Check the end of the blog post for information on an upcoming blog format change. [Author’s note: somehow the references section didn’t attach to the original post, that issue is now corrected, and I apologize, references are everything in as technical a field as this.]
The SNAP-50 was the last, and most powerful, of the SNAP series of reactors, and had a very different start when compared to the other three reactors that we’ve looked at. A fifth reactor, SNAP-4, also underwent some testing, but was meant for undersea applications for the Navy. The SNAP-50 reactor started life in the Aircraft Nuclear Propulsion program for the US Air Force, and ended its life with NASA, as a power plant for the future modular space station that NASA was planning before the budget cuts of the mid to late 1970s took hold.
Because it came from a different program originally, it also uses different technology than the reactors we’ve looked at on the blog so far: uranium nitride fuel, and higher-temperature, lithium coolant made this reactor a very different beast than the other reactors in SNAP. However, these changes also allowed for a more powerful reactor, and a less massive power plant overall, thanks to the advantages of the higher-temperature design. It was also the first major project to move the space reactor development process away from SNAP-2/10A legacy designs.
The SNAP-50 would permanently alter the way that astronuclear reactors were designed, and would change the course of in-space reactor development for over 20 years. By the time of its cancellation in 1973, it had approached flight readiness to the point that funding and time allowed, but changes in launch vehicle configuration rang the death knell of the SNAP-50.
The Birth of the SNAP-50
Up until now, the SNAP program had focused on a particular subset of nuclear reactor designs. They were all fueled with uranium-zirconium hydride fuel (within a small range of uranium content, all HEU), cooled with NaK-78, and fed either mercury Rankine generators or thermoelectric power conversion systems. This had a lot of advantages for the program: fuel element development improvements for one reactor could be implemented in all of them, challenges in one reactor system that weren’t present in another allowed for distinct data points to figure out what was going on, and the engineers and reactor developers were able to look at each others’ work for ideas on how to improve reliability, efficiency, and other design questions.
However, there was another program that was going on at about the same time which had a very different purpose, but similar enough design constraints that it could be very useful for an in-space fission power plant: the Aircraft Nuclear Propulsion program (ANP), which was primarily run out of Oak Ridge National Laboratory. Perhaps the most famous part of the ANP program was the series of direct cycle ramjets for Project PLUTO: the TORY series. These ramjets were nuclear fission engines using the atmosphere itself as the working fluid. There were significant challenges to this approach, because the clad for the fuel elements must not fail, or else the fission products from the fuel elements would be released as what would be virtually identical to nuclear fallout, only different due to the method that it was generated. The fuel elements themselves would be heavily eroded by the hot air moving through the reactor (which turned out to be a much smaller problem than was initially anticipated). The advantage to this system, though, is that it was simple, and could be made to be relatively lightweight.
Another option was what was known as the semi-indirect cycle, where the reactor would heat a working fluid in a closed loop, which would then heat the air through a heat exchanger built into the engine pod. While this was marginally safer from a fission product release point of view, there were a number of issues with the design. The reactor would have to run at a higher temperature than the direct cycle, because there are always losses whenever you transfer heat from one working fluid to another, and the increased mass of the system also required greater thrust to maintain the desired flight characteristics. The primary coolant loop would become irradiated when going through the reactor, leading to potential irradiation of the air as it passed through the heat exchanger. Another concern was that the heat exchanger could fail, leading to the working fluid (usually a liquid metal) being exposed at high temperature to the superheated air, where it could easily explode. Finally, if a clad failure occurred in the fuel elements, fission products could migrate into the working fluid, making the primary loop even more radioactive, increasing the irradiation of the air as it passed through the engine – and releasing fission products into the atmosphere if the heat exchanger failed.
The alternative to these approaches was an indirect cycle, where the reactor heated a working fluid in a closed loop, transferred this to another working fluid, which then heated the air. The main difference between these systems is that, rather than having the possibly radioactive primary coolant come in close proximity with the air and therefore transferring ionizing radiation, there is an additional coolant loop to minimize this concern, at the cost of both mass and thermal efficiency. This setup allowed for far greater assurances that the air passing through the engine would not be irradiated, because the irradiation of the secondary coolant loop would be so low as to be functionally nonexistent. However, if the semi-indirect cycle was more massive, this indirect cycle would be the heaviest of all of the designs, meaning far higher power outputs and temperatures were needed in order to get the necessary thrust-to-weight ratios for the aircraft. Nevertheless, from the point of view of the people responsible for the ANP program, this was the most attractive design for a crewed aircraft.
Both SNAP and ANP needed many of the same things out of a nuclear reactor: it had to be compact, it had to be lightweight, it had to have a VERY high power density and it needed to be able to operate virtually maintenance-free in a variety of high-power conditions. These requirements are in stark contrast to terrestrial, stationary nuclear reactors which can afford heavy weight, voluminous construction and can thus benefit of low power density. As a general rule of thumb, an increase in power density, will also intensify the engineering, materials, and maintenance challenges. The fact that the ANP program needed high outlet temperatures to run a jet engine also bore the potential of having a large thermal gradient across a power conversion system – meaning that high-conversion-efficiency electrical generation was possible. That led SNAP program leaders to see about adapting an aircraft system into a spacecraft system.
The selected design was under development at the Connecticut Advanced Nuclear Engine Laboratory (CANEL) in Middletown, Connecticut. The prime contractor was Pratt and Whitney. Originally part of the indirect-cycle program, the challenges of heat exchanger design, adequate thrust, and a host of other problems continually set back the indirect cycle program, and when the ANP program was canceled in 1961, Pratt and Whitney no longer had a customer for their reactor, despite doing extensive testing and even fabricating novel alloys to deal with certain challenges that their reactor design presented. This led them to look for another customer for the reactor, and they discovered that both NASA and the US Air Force were both interested in high-power-density, high temperature reactors for in-space use. Both were interested in this high powered reactor, and the SNAP-50 was born.
This reactor was an evolution of a series of test reactors, the PWAR series of test reactors. Three reactors (the PWAR-2, -4, and -8, for 2, 4, and 8 MW of thermal power per reactor core) had already been run for initial design of an aircraft reactor, focused on testing not only the critical geometry of the reactor, but the materials needed to contain its unique (at the time) coolant: liquid lithium. This is because lithium has an excellent specific heat capacity, or the amount of energy that can be contained as heat per unit mass at a given temperature: 3.558 J/kg-C, compared to the 1.124 J/kg-C of NaK78, the coolant of the other SNAP reactors. This means that less coolant would be needed to transport the energy away from the reactor and into the engine in the ANP program, and for SNAP this meant that less working fluid mass would be needed transferring from the reactor to the power conversion system. The facts that Li is much less massive than NaK, and that less of it would be needed, makes lithium a highly coveted option for an astronuclear reactor design. However, this design decision also led to needing novel concepts for how to contain liquid lithium. Even compared to NaK, lithium is highly toxic, highly corrosive in most materials and led, during the ANP program, to Pratt and Whitney investigating novel elemental compositions for their containment structures. We’ll look at just what they did later.
SNAP-50: Designing the Reactor Core
This reactor ended up using a form of fuel element that we have yet to look at in this blog: uranium nitride, UN. While both UC (you can read more about carbide fuels here) and UN were considered at the beginning of the program, the reactor designers ended up settling on UN because of a unique capacity that this fuel form offers: it has the highest fissile fuel density of any type of fuel element. This is offset by the fact that UN isn’t the most heat tolerant of fuel elements, requiring a lower core operating temperature. Other options were considered as well, including CERMET fuels using oxides, carbides, and nitrides suspended in a tungsten metal matrix to increase thermal conductivity and reduce the temperature of the fissile fuel itself. The decision between UN, with its higher mass efficiency (due to its higher fissile density), and uranium carbide (UC), with the highest operating temperature of any solid fuel element, was a difficult decision, and a lot of fuel element testing occurred at CANEL before a decision was reached. After a lot of study, it was determined that UN in a tungsten CERMET fuel was the best balance of high fissile fuel density, high thermal conductivity, and the ability to manage low fuel burnup over the course of the reactor’s life.
Perhaps the most important design consideration for the fuel elements after the type of fuel was how dense the fuel would be, and how to increase the density if this was desired in the final design. While higher density fuel is generally speaking a better idea when it comes to specific power, it was discovered that the higher density the fuel was, the lower the amount of burnup would be possible before the fuel would fail due to fission product gas buildup within the fuel itself. Initial calculations showed that there was an effectively unlimited fuel burnup potential of UN at 80% of its theoretical density since a lot of the gasses could diffuse out of the fuel element. However, once the fuel reached 95% density, this was limited to 1% fuel burnup. Additional work was done to determine that this low burnup was in fact not a project killer for a 10,000 hour reactor lifetime, as was specified by NASA, and the program moved ahead.
These fuel pellets needed a cladding material, as most fuel does, and this led to some additional unique materials challenges. With the decision to use lithium coolant, and the need for both elasticity and strength in the fuel element cladding (to deal with both structural loads and fuel swelling), it was necessary to do extensive experimentation on the metal that would be used for the clad. Eventually, a columbium-zirconium alloy with a small amount of carbon (CB-1ZR-0.6C) was decided on as a barrier between the Cb-Zr alloy of the clad (which resisted the high-temperature lithium erosion on the pressure vessel side of the clad) and the UN-W CERMET fuel (which would react strongly without the carburized layer).
This decisions led to an interesting reactor design, but not necessarily one that is unique from a non-materials point of view. The fuel would be formed into high-density pellets, which would then be loaded into a clad, with a spring to keep the fuel to the bottom (spacecraft end) of the reactor. The gap between the top of the fuel elements and the top of the clad was for the release of fission product gasses produced during operation of the reactor. These rods would be loaded in a hexagonal prism pattern into a larger collection of fuel elements, called a can. Seven of these cans, placed side by side (one regular hexagon, surrounded by six slightly truncated hexagons), would form the fueled portion of the reactor core. Shims of beryllium would shape the core into a cylinder, which was surrounded by a pressure vessel and lateral reflectors. Six poison-backed control drums mounted within the reflector would rotate to provide reactor control. Should the reactor need to be scrammed, a spring mechanism would return all the drums to a position with the neutron poison facing the reactor, stopping fission from occurring.
The lithium, after being heated to a temperature of 2000°F (1093°C), would feed into a potassium boiler, before being returned to the core at an inlet temperature of 1900 F (1037°C). From the boiler, the potassium vapor, which is 1850°F (1010°C), would enter a Rankine turbine which would produce electricity. The potassium vapor would cool down to 1118°F (603°C) in the process and return – condensed to its liquid form – to the boiler, thus closing the circulation. Several secondary coolant loops were used in this reactor: the main one was for the neutron reflectors, shield actuators, control drums, and other radiation hardened equipment, and used NaK as a coolant; this coolant was also used as a lubricant for the condensate pump in the potassium system. Another, lower temperature organic coolant was used for other systems that weren’t in as high a radiation flux. The radiators that were used to reject heat also used NaK as a working fluid, and were split into a primary and secondary radiator array. The primary array pulled heat from the condenser, and reduced it from 1246°F (674°C) to 1096°F (591°C), while the secondary array took the lower-temperature coolant from 730°F (388°C) to 490°F (254°C). This design was designed to operate in both single and dual loop situations, with the second (identical) loop used for high powered operation and to increase redundancy in the power plant.
These design decisions led to a flexible reactor core size, and the ability to adapt to changing requirements from either NASA or the USAF, both of which were continuing to show interest in the SNAP-50 for powering the new, larger space stations that were becoming a major focus of both organizations.
The Power Plant: Getting the Juice Flowing
By 1973, the SNAP 2/10A program had ended, and the SNAP-8/ZrHR program was winding down. These systems simply didn’t provide enough power for the new, larger space station designs that were being envisaged by NASA, and the smaller reactor sizes (the 10B advanced designs that we looked at a couple blog posts back, and the 5 kWe Thermoelectric Reactor) didn’t provide capabilities that were needed at the time. This left the SNAP-50 as the sole reactor design that was practical to take on a range of mission types… but there was a need to have different reactor power outputs, so the program ended up developing two reactor sizes. The first was a 35 kWe reactor design, meant for smaller space stations and lunar bases, although this particular part of the 35 kWe design seems to have never been fully fleshed out. A larger, 300 kWe type was designed for NASA’s proposed modular space station, a project which would eventually evolve into the actual ISS.
Unlike in the SNAP-2 and SNAP-8 programs, the SNAP-50 kept its Rankine turbine design, which had potassium vapor as its working fluid. This meant that the power plant was able to meet its electrical power output requirements far more easily than the lower efficiency demanded by thermoelectric conversion systems. The CRU system meant for the SNAP-2 ended up reaching its design requirements for reliability and life by this time, but sadly the overall program had been canceled, so there was no reactor to pair to this ingenious design (sadly, it’s so highly toxic that testing would be nearly impossible on Earth). The boiler, pumps, and radiators for the secondary loop were tested past the 10,000 hour design lifetime of the power plant, and all major complications discovered during the testing process were addressed, proving that the power conversion system was ready for the next stage of testing in a flight configuration.
One concern that was studied in depth was the secondary coolant loop’s tendency to become irradiated in the neutron flux coming off the reactor. Potassium has a propensity for absorbing neutrons, and in particular 41K (6% of unrefined K) can capture a neutron and become 42K. This is a problem, because 42K goes through gamma decay, so anywhere that the secondary coolant goes needs to have gamma radiation shielding to prevent the radiation from reaching the crew. This limited where the power conversion system could be mounted, to keep it inside the gamma shielding of the temporary, reactor-mounted shield, however the compact nature of both the reactor core and the power conversion system meant that this was a reasonably small concern, but one worthy of in-depth examination by the design team.
The power conversion system and auxiliary equipment, including the actuators for the control drums, power conditioning equipment, and other necessary equipment was cooled by a third coolant loop, which used an organic coolant (basically the oil needed for the moving parts to be lubricated), which ran through its own set of pumps and radiators. This tertiary loop was kept isolated from the vast majority of the radiation flux coming off the reactor, and as such wasn’t a major concern for irradiation damage of the coolant/lubricant.
Some Will Stay, Some Will Go: Mounting SNAP-50 To A Space Station
Each design used a 4-pi (a fully enclosing) shield with a secondary shadow shield pointing to the space station in order to reduce radiation exposure for crews of spacecraft rendezvousing or undocking from the space station. This primary shield was made out of a layer of beryllium to reflect neutrons back into the core, and boron carbide (B4C, enriched in boron-10) to absorb the neutrons that weren’t reflected back into the core. These structures needed to be cooled to ensure that the shield wouldn’t degrade, so a NaK shield coolant system (using technology adapted from the SNAP-8 program) was used to keep the shield at an acceptable temperature.
The shadow shield was built in two parts: the entire structure would be launched at the same time for the initial reactor installation for the space station, and then when the reactor needed to be replaced only a portion of the shield would be jettisoned with the reactor. The remainder, as well as the radiators for the reactor’s various coolant systems, would be kept mounted to the space station in order to reduce the amount of mass that needed to be launched for the station resupply. The shadow shield was made out of layers of tungsten and LiH, for gamma and neutron shielding respectively.
When it came time to replace the core of the reactor at the end of its 10,000 hour design life (which was a serious constraint on the UN fuels that they were working with due to fuel burnup issues), everything from the separation plane back would be jettisoned. This could theoretically have been dragged to a graveyard orbit by an automated mission, but the more likely scenario at the time would have been to leave it in a slowly degrading orbit to give the majority of the short-lived isotopes time to decay, and then design it to burn up in the atmosphere at a high enough altitude that diffusion would dilute the impact of any radioisotopes from the reactor. This was, of course, before the problems that the USSR ran into with their US-A program [insert link], which eliminated this lower cost decommissioning option.
After the old reactor core was discarded, the new core, together with the small forward shield and power conversion system, could be put in place using a combination of off-the-shelf hardware, which at the time was expected to be common enough: either Titan-III or Saturn 1B rockets, with appropriate upper stages to handle the docking procedure with the space station. The reactor would then be attached to the radiator, the docking would be completed, and within 8 hours the reactor would reach steady-state operations for another 10,000 hours of normal use. The longest that the station would be running on backup power would be four days. Unfortunately, information on the exact docking mechanism used is thin, so the details on how they planned this stage are still somewhat hazy, but there’s nothing preventing this from being done.
A number of secondary systems, including accumulators, pumps, and other equipment are mounted along with the radiator in the permanent section of the power supply installation. Many other systems, especially anything that has been exposed to a large radiation flux or high temperatures during operation (LiH, the primary shielding material, loses hydrogen through outgassing at a known rate depending on temperature, and can almost be said to have a half-life), will be separated with the core, but everything that was practicable to leave in place was kept.
This basic design principle for reloadable (which in astronuclear often just means “replaceable core”) reactors will be revisited time and again for orbital installations. Variations on the concept abound, although surface power units seem to favor “abandon in place” far more. In the case of large future installations, it’s not unreasonable to suspect that refueling of a reactor core would be possible, but at this point in astronuclear mission utilization, even having this level of reusability was an impressive feat.
35 kWe SNAP-50: The Starter Model
In the 1960s, having 35 kWe of power for a space station was considered significant enough to supply the vast majority of mission needs. Because of this, a smaller version of the SNAP-50 was designed to fit this mission design niche. While the initial power plant would require the use of a Saturn 1B to launch it into orbit, the replacement reactors could be launched on either an Atlas-Centaur or Titan IIIA-Centaur launch vehicle. This was billed as a low cost option, as a proof of concept for the far larger – and at this point, far less fully tested – 300 kWe version to come.
NASA was still thinking of very large space stations at this time. The baseline crew requirements alone were incredible: 24-36 crew, with rotations lasting from 3 months to a year, and a station life of five years. While 35 kWe wouldn’t be sufficient for the full station, it would be an attractive option. Other programs had looked at nuclear power plants for space stations as well, like we saw with the Manned Orbiting Laboratory and the Orbital Workshop (later Skylab), and facilities of that size would be good candidates for the 35 kWe system.
The core itself measured 8.3 inches (0.211 m) across, 11.2 inches (0.284 m) long, and used 236 fuel elements arranged into seven fuel element cans within the pressure vessel of the core. Six poison-backed control drums were used for primary reactor control. The core would produce up to 400 kW of thermal power. The pressure vessel, control drums, and all other control and reflective materials together measured just 19.6 inches (4.98 m) by 27.9 inches (7.09 m), and the replaceable portion of the reactor was between four and five feet (1.2 m and 1.5 m) tall, and five and six feet (1.5 m and 1.8 m) across – including shielding.
This reactor could also have been a good prototype reactor for a nuclear electric probe, a concept that will be revisited later, although there’s little evidence that this path was ever seriously explored. Like many smaller reactor designs, this one did not get the amount of attention that its larger brother offered, but at the time this was considered a good, solid space station power supply.
300 kWe SNAP-50: The Most Powerful Space Reactor to Date
While there were sketches for more powerful reactors than the 300 kWe SNAP-50 variant, they never really developed the reactors to any great extent, and certainly not to the point of experimental verification that SNAP-50 had achieved. This was considered to be a good starting point for possibly a crewed nuclear electric spacecraft, as well as being able to power a truly huge space station.
The 300 kWe variant of the reactor was slightly different in more than size when compared to its smaller brother. Despite using the same fuel, clad, and coolant as the 35 kWe system, the 300 kWe system could achieve over four times the fuel burnup of the smaller reactor (0.32% vs 1.3%), and had a higher maximum fuel power density as well, both of which have a huge impact on core lifetimes and dynamics. This was partially achieved by making the fuel elements almost half as narrow, and increasing the number of fuel elements to 1093, held in 19 cans within the core. This led to a core that was 10.2 inches (0.259 m) wide, and 14.28 inches (0.363 m) long (keeping the same 1:1.4 gore geometry between the reactors), and a pressure vessel that was 12” (0.305 m) in diameter by 43” (1.092 m) in length. It also increased the thermal output of the reactor to 2200 kWt. The number of control drums was increased from six to eight longer control drums to fit the longer core, and some rearrangement of lithium pumps and other equipment for the power conversion system occurred within the larger 4 pi shield structure. The entire reactor assembly that would undergo replacement was five to six feet high, and six to seven feet in diameter (1.5 m; 1.8 m; 2.1 m).
Sadly, even the ambitious NASA space station wasn’t big enough to need even the smaller 35 kWe version of the reactor, much less the 300 kWe variants. Plans had been made for a fleet of nuclear electric tugs that would ferry equipment back and forth to a permanent Moon base, but cancellation of that program occurred at the same time as the death of the moon base itself.
Mass Tradeoffs: Why Nuclear Instead of Solar?
By the middle of the 1960s, photovoltaic solar panels had become efficient and reliable enough for use in spacecraft on a regular basis. Because of this, it was a genuine question for the first time ever whether to go with solar panels or a nuclear reactor, whereas in the 1950s and early 60s nuclear was pretty much the only option. However, solar panels have a downside: drag. Even in orbit, there is a very thin atmosphere, and so for lower orbits a satellite has to regularly raise itself up or it will burn up in the atmosphere. Another down side comes from MM/OD: micro meteorites and orbital debris. Since solar panels are large, flat, and all pointing at the sun all the time, there’s a greater chance that something will strike one of those panels, damaging or possibly even destroying it. Managing these two issues is the primary concern of using solar panels as a power supply in terms of orbital behavior, and determines the majority of the refueling mass needed for a solar powered space station.
On the nuclear side, by 1965, there were two power plant options on the table: the SNAP-8 (pre-ZrHR redesign) and the SNAP-50, and solar photovoltaics had developed to the point that they could be deployed in space. Because of this, a comparison was done by Pratt and Whitney of the three systems to determine the mass efficiency of each system, not only in initial deployment but also in yearly fueling and tankage requirements. Each of the systems was compared at a 35 kWe power level to the space station in order to allow for a level playing field.
One thing that stands out about the solar system (based on a pair of Lockheed and General Electric studies) is that it’s marginally the lightest of all the systems at launch, but within a year the total system maintenance mass required far outstrips the mass of the nuclear power plants, especially the SNAP-50. This is because the solar panels have a large sail area, which catches the very thin atmosphere at the station’s orbital altitude and drags the station down into the thicker atmosphere, so thrust is needed to re-boost the space station. This is something that has to be done on a regular basis for the ISS. The mass of the fuel, tankage, and structure to allow for this reboost is extensive. Even back in 1965 there were discussions on using electric propulsion for the reboosting of the space station, in order to significantly reduce the mass needed for this procedure. That discussion is still happening casually with the ISS, and Ad Astra still hopes to use VASIMR for this purpose – a concept that’s been floated for the last ten or so years.
Overall, the mass difference between the SNAP-50 and the optimistic Lockheed proposal of the time was significant: the original deployment was only about 70 lbs (31.75 kg) different, but the yearly maintenance mass requirements would be 5,280 lbs (2395 kg) different – quite a large amount of mass.
Because the SNAP-50 and SNAP-8 don’t have these large sail areas, and the radiators needed can be made aerodynamically enough to greatly reduce the drag on the station, the reboost requirements are significantly lower than for the solar panels. The SNAP-50 weighs significantly less than the SNAP-8, and has significantly less surface area, because the reactor operates at a far higher temperature, and therefore needs a smaller radiator. Another difference between the reactors is volume: the SNAP-50 is physically smaller than the SNAP-8 because of that same higher temperature, and also due to the fact that the UN fuel is far more dense than its U-ZrH fueled counterpart.
These reactors were designed to be replaced once a year, with the initial launch being significantly more massive than the follow-up launches, benefitting of the sectioned architecture with a separation plane just at the small of the shadow shield as described above. Only the smaller section of shield remained with the reactor when it was separated. The larger, heavier section, on the other hand, would remain with the space station, as well as the radiators, and serve as the mounting point for the new reactor core and power conversion system, which would be sent via an automated refueling launch to the space station.
Solar panels, on the other hand, require both reboost to compensate for drag as well as equipment to repair or replace the panels, batteries, and associated components as they wear out. This in turn requires a somewhat robust repair capability for ongoing maintenance – a requirement for any large, long term space station, but the more area you have to get hit by space debris, which means more time and mass spent on repairs rather than doing science.
Of course, today solar panels are far lighter, and electric thrusters are also far more mature than they were at that time. This, in addition to widespread radiophobia, make solar the most widespread occurrence in most satellites, and all space stations, to date. However, the savings available in overall lifetime mass and a sail area that is both smaller and more physically robust, remain key advantages for a nuclear powered space station in the future
The End of an Era: Changing Priorities, Changing Funding
The SNAP-50, even the small 35 kWe version, offered more power, more efficiency, and less mass and volume than the most advanced of SNAP-8’s children: the A-ZrHR [Link]. This was the end of the zirconium hydride fueled reactor era for the Atomic Energy Commission, and while this type of fuel continues to be used in reactors all over the world in TRIGA research and training reactors (a common type of small reactor for colleges and research organizations), its time as the preferred fuel for astronuclear designs was over.
In fact, by the end of the study period, the SNAP-50 was extended to 1.5 MWe in some designs, the most powerful design to be proposed until the 1980s, and one of the most powerful ever proposed… but this ended up going nowhere, as did much of the mission planning surrounding the SNAP program.
At the same time as these higher-powered reactor designs were coming to maturity, funding for both civilian and military space programs virtually disappeared. National priorities, and perceptions of nuclear power, were shifting. Technological advances eliminated many future military crewed missions in favor of uncrewed ones with longer lifetimes, less mass, less cost – and far smaller power requirements. NASA funding began falling under the axe even as we were landing on the Moon for the first time, and from then on funding became very scarce on the ground.
The transition from the Atomic Energy Commission to the Department of Energy wasn’t without its hiccups, or reductions in funding, either, and where once every single AEC lab seemed to have its own family of reactor designs, the field narrowed greatly. As we’ll see, even at the start of Star Wars the reactor design was not too different from the SNAP-50.
Finally, the changes in launch system had their impact as well. NASA was heavily investing in the Space Transport System (the Space Shuttle), which was assumed to be the way that most or all payloads would be launched, so the nuclear reactor had to be able to be flown up – and in some cases returned – by the Shuttle. This placed a whole different set of constraints on the reactor, requiring a large rewrite of the basic design. The follow-on design, the SP-100, used the same UN fuel and Li coolant as the SNAP-50, but was designed to be launched and retrieved by the Shuttle. The fact that the STS never lived up to its promise in launch frequency or cost (and that other launchers were available continuously) means that this was ultimately a diversion, but at the time it was a serious consideration.
All of this spelled the death of the SNAP-50 program, as well as the end of dedicated research into a single reactor design until 1983, with the SP-100 nuclear reactor system, a reactor we’ll look at another time.
While I would love to go into many of the reactors that were developed up to this time, including heat pipe cooled reactors (SABRE at Los Alamos), thermionic power conversion systems (5 kWe Thermionic Reactor), and other ideas, there simply isn’t time to go into them here. As we look at different reactor components they’ll come up, and we’ll mention them there. Sadly, while some labs were able to continue funding some limited research with the help of NASA and sometimes the Department of Defense or the Defense Nuclear Safety Agency. The days of big astronuclear programs, though, were fading into a thing of the past. Both space and nuclear power would refocus, and then fade in the rankings of budgetary requirements over the years. We will be looking at these reactors more as time goes on, in our new “Forgotten Reactors” column (more on that below).
The Blog is Changing!
With the new year, I’ve been thinking a lot about the format of both the website and the blog, and where I hope to go in the next year. I’ve had several organizational projects on the back burner, and some of them are going to be started here soon. The biggest part is going to be the relationship between the blog and the website, and what I write more about where.
Expect another blog post shortly (it’s already written, just not edited yet) about our plans for the next year!
I’ve got big plans for Beyond NERVA this year, and there are a LOT of things that are slowly getting started in the background which will greatly improve the quality of the blog and the website, and this is just the start!
Hello, and welcome back to Beyond NERVA! As some of you may have noticed, the website has moved! Yes, we’re now at beyondnerva.com! I’m working on updating the webpage, and am getting the pieces together for a major website redesign (still a ways off, but lots of the pieces are starting to fall into place) to make the site easier to navigate and more user friendly. Make sure to update your bookmarks with this new address! With that brief administrative announcement out of the way, let’s get back to our look at in-space fission power plants.
Today, we’re going to continue our look at the SNAP program, America’s first major attempt to provide electric power in space using nuclear energy, and finishing up our look at the zirconium hydride fueled reactors that defined the early SNAP reactors by looking at the SNAP-8, and its two children – the 5 kW Thermoelectric Reactor and the Advanced Zirconium Hydride Reactor.
SNAP 8 was the first reactor designed with these space stations in mind in mind. While SNAP-10A was a low-power system (at 500 watts when flown, later upgraded to 1 kW), and SNAP-2 was significantly larger (3 kW), there was a potential need for far more power. Crewed space stations take a lot of power (the ISS uses close to 100 kWe, as an example), and neither the SNAP-10 or the SNAP-2 were capable of powering the space stations that NASA was in the beginning stages of planning.
Initially designed to be far higher powered, with 30-60 kilowatts of electrical power, this was an electric supply that could power a truly impressive outpost for humanity in orbit. However, the Atomic Energy Commission and NASA (which was just coming into existence at the time this program was started) didn’t want to throw the baby out with the bath water, as it were. While the reactor was far higher powered than the SNAP 2 reactor that we looked at last time, many of the power system’s components are shared: both use the same fuel (with minor exceptions), both use similar control drum structures for reactor control, both use mercury Rankine cycle power conversion systems, and perhaps most attractively both were able to evolve with lessons learned from the other part of the program.
While SNAP 8 never flew, it was developed to a very high degree of technical understanding, so that if the need for the reactor arose, it would be available. One design modification late in the SNAP 8 program (when the reactor wasn’t even called SNAP 8 anymore, but the Advanced Zirconium Hydride Reactor) had a very rare attribute in astronuclear designs: it was shielded on all sides for use on a space station, providing more than twice the electrical power available to the International Space Station without any of the headaches normally associated with approach and docking with a nuclear powered facility.
Let’s start back in 1959, though, with the SNAP 8, the first nuclear electric propulsion reactor system.
SNAP 8: NASA Gets Involved Directly
The SNAP 2 and SNAP 10A reactors were both collaborations between the Atomic Energy Commission (AEC), who were responsible for the research, development, and funding of the reactor core and primary coolant portions of the system, and the US Air Force, who developed the secondary coolant system, the power conversion system, the heat rejection system, the power conditioning unit, and the rest of the components. Each organization had a contractor that they used: the AEC used Atomics International (AI), one of the movers and shakers of the advanced reactor industry, while the US Air Force went to Thompson Ramo Wooldridge (better known by their acronym, TRW) for the SNAP-2 mercury (Hg) Rankine turbine and Westinghouse Electric Corporation for the SNAP-10’s thermoelectric conversion unit.
1959 brought NASA directly into the program on the reactor side of things, when they requested a fission reactor in the 30-60 kWe range for up to one year; one year later the SNAP-8 Reactor Development Program was born. It would use a similar Hg-based Rankine cycle as the SNAP-2 reactor, which was already under development, but the increased power requirements and unique environment that the power conversion system necessitated significant redesign work, which was carried out by Aerojet General as the prime contractor. This led to a 600 kWe rector core, with a 700 C outlet temperature As with the SNAP-2 and SNAP-10 programs, the SNAP 8’s reactor core was funded by the AEC, but in this case the power conversion system was the funding responsibility of NASA.
The fuel itself was similar to that in the SNAP-2 and -10A reactors, but the fuel elements were far longer and thinner than those of the -2 and -10A. Because the fuel element geometry was different, and the power level of the reactor was so much higher than the SNAP-2 reactor, the SNAP-8 program required its own experimental and developmental reactor program to run in parallel to the initial SNAP Experimental and Development reactors, although the materials testing undertaken by the SNAP-2 reactor program, and especially the SCA4 tests, were very helpful in refining the final design of the SNAP-8 reactor.
The power conversion system for this reactor was split in two: identical Hg turbines would be used, with either one or both running at any given time depending on the power needs of the mission. This allows for more flexibility in operation, and also simplifies the design challenges involved in the turbines themselves: it’s easier to design a turbine with a smaller power output range than a larger one. If the reactor was at full power, and both turbines were used, the design was supposed to produce up to 60 kW of electrical power, while the minimum power output of a single turbine would be in the 30 kWe range. Another advantage was that if one was damaged, the reactor would continue to be able to produce power.
Due to the much higher power levels, an extensive core redesign was called for, meaning that different test reactors would need to be used to verify this design. While the fuel elements were very similar, and the overall design philosophy was operating in parallel to the SNAP-2/10A program, there was only so much that the tests done for the USAF system would be able to help the new program. This led to the SNAP-8 development program, which began in 1960, and had its first reactor, the SNAP-8 Experimental Reactor, come online in 1963.
SNAP-8 Experimental Reactor: The First of the Line
The first reactor in this series, the SNAP 8 Experimental Reactor (S8ER), went critical in May 1963, and operated until 1965. it operated for 2522 hours at above 600 kWt, and over 8000 hours at lower power levels. The fuel elements for the reactor were 14 inches in length, and 0.532 inches in diameter, with uranium-zirconium hydride (U-ZrH, the same basic fuel type as the SNAP-2/10A system that we looked at last time) enriched to 93.15% 235U, with 6 X 10^22 atoms of hydrogen per cubic centimeter.
The biggest chemical change in this reactor’s fuel elements compared to the SNAP-2/10A system was the hydrogen barrier inside the metal clad: instead of using gadolinium as a burnable poison (which would absorb neutrons, then decay into a neutron-transparent element as the reactor underwent fission over time), the S8ER used samarium. The reasons for the change are rather esoteric, relating to the neutron spectrum of the reactor, the particular fission products and their ratios, thermal and chemical characteristics of the fuel elements, and other factors. However, the change was so advantageous that eventually the different burnable poison would be used in the SNAP-2/10A system as well.
The fuel elements were still loaded in a triangle array, but makes more of a cylinder than a hexagon like in the -2/10A, with small internal reflectors to fill out the smooth cylinder of the pressure vessel. The base and head plates that hold the fuel elements are very similar to the smaller design, but obviously have more holes to hold the increased number of fuel elements. The NaK-78 coolant (identical to the SNAP-2/10A system) entered in the bottom of the reactor into a space in the pressure vessel (a plenum), flowed through the base plate and up the reactor, then exits the top of the pressure vessel through an upper plenum. A small neutron source used as a startup neutron source (sort of like a spark plug for a reactor) was mounted to the top of the pressure vessel, by the upper coolant plenum. The pressure vessel itself was made out of 316 stainless steel.
Instead of four control drums, the S8ER used six void-backed control drums. These were directly derived from the SNAP-2/10A control system. Two of the drums were used for gross reactivity control – either fully rotated in or out, depending on if the reactor is under power or not. Two were used for finer control, but at least under nominal operation would be pretty much fixed in their location over longer periods of time. As the reactor approached end of life, these drums would rotate in to maintain the reactivity of the system. The final two were used for fine control, to adjust the reactivity for both reactor stability and power demand adjustment. The drums used the same type of bearings as the -2/10A system.
The S8ER first underwent criticality benchmark tests (pre-dry critical testing) from September to December 1962 to establish the reactor’s precise control parameters. Before filling the reactor with the NaK coolant, water immersion experiments for failure-to-orbit safety testing (as an additional set of tests to the SCA-4 testing which also supported SNAP-8) was carried out between January and March of 1963. After a couple months of modifications and refurbishment, dry criticality tests were once again conducted on May 19, 1963, followed in the next month with the reactor reaching wet critical power levels on June 23. Months of low-power testing followed, to establish the precise reactor control element characteristics, thermal transfer characteristics, and a host of other technical details before the reactor was increased in power to full design characteristics.
The reactor was shut down from early August to late October, because some of the water coolant channels used for the containment vessel failed, necessitating the entire structure to be dug up, repaired, and reinstalled, with significant reworking of the facility being required to complete this intensive repair process. Further modifications and upgrades to the facility continued into November, but by November 22, the reactor underwent its first “significant” power level testing. Sadly, this revealed that there were problems with the control drum actuators, requiring the reactor to be shut down again.
After more modifications and repairs, lower power testing resumed to verify the repairs, study reactor transient behavior, and other considerations. The day finally came for the SNAP-8 Experimental Reactor achieved its first full power, at temperature testing on December 11, 1963. Shortly after, the reactor had to be shut down again to repair a NaK leak in one of the primary coolant loop pumps, but the reactor was up and operating again shortly after. Lower power tests were conducted to evaluate the samarium burnable poisons in the fuel elements, measure xenon buildup, and measure hydrogen migration in the core until April 28, interrupted briefly by another NaK pump failure and a number of instrumentation malfunctions in the automatic scram system (which was designed to automatically shut down the reactor in the case of an accident or certain types of reactor behaviors). However, despite these problems, April 28 marked 60 days of continuous operation at 450 kWt and 1300 F (design temperature, but less-than-nominal power levels).
After a shutdown to repair the control drive mechanisms (again), the reactor went into near-continuous operation, either at 450 or 600 kWt of power output and 1300 F outlet temperature until April 15, 1965, when the reactor was shut down for the last time. By September 2 of 1964, the S8ER had operated at design power and temperature levels for 1000 continuous hours, and went on in that same test to exceed the maximum continuous operation time of any SNAP reactor to date on November 5 (1152 hours). January 18 of 1965 it achieved 10,000 hours of total operations, and in February of that year reached 100 days of continuous operation at design power and temperature conditions. Just 8 days later, on February 12, it exceeded the longest continuous operation of any reactor to that point (147 days, beating the Yankee reactor). March 5 marked the one year anniversary of the core outlet temperature being continuously at over 1200 F. By April 15, when the reactor was shut down for the last time it achieved an impressive set of accomplishments:
5016.5 continuous operations immediately preceeding the shutdown (most at 450 kWt, all at 1200 F or greater)
12,080 hours of total operations
A total of 5,154,332 kilowatt-hours of thermal energy produced
91.09% Time Operated Efficiency (percentage of time that the reactor was critical) from November 22, 1963 (the day of first significant power operations of the reactor), and 97.91% efficiency in the last year of operations.
Once the tests were concluded, the reactor was disassembled, inspected, and fuel elements were examined. These tests took place at the Atomics International Hot Laboratory (also at Santa Susana) starting on July 28, 1965. For about 6 weeks, this was all that the facility focused on; the core was disassembled and cleaned, and the fuel elements were each examined, with many of them being disassembled and run through a significant testing regime to determine everything from fuel burnup to fission product percentages to hydrogen migration. The fuel element tests were the most significant, because to put it mildly there were problems.
Of the 211 fuel elements in the core, only 44 were intact. Many of the fuel elements also underwent dimensional changes, either swelling (with a very small number actually decreasing) across the diameter or the length, becoming oblong, dishing, or other changes in geometry. The clad on most elements was damaged in one way or another, leading to a large amount of hydrogen migrating out of the fuel elements, mostly into the coolant and then out of the reactor. This means that much of the neutron moderation needed for the reactor to operate properly migrated out of the core, reducing the overall available reactivity even as the amount of fission poisons in the form of fission products was increasing. For a flight system, this is a major problem, and one that definitely needs to be addressed. However, this is exactly the sort of problem that an experimental reactor is meant to discover and assess, so in this way as well the reactor was a complete success, if not as smooth a development as the designers would likely have preferred.
It was also discovered that, while the cracks in the clad would indicate that the hydrogen would be migrating out of the cracks in the hydrogen diffusion barrier, far less hydrogen was lost than was expected based on the amount of damage the fuel elements underwent. In fact, the hydrogen migration in these tests was low enough that the core would most likely be able to carry out its 10,000 hour operational lifetime requirement as-is; without knowing what the mechanism that was preventing the hydrogen migration was, though, it would be difficult if not impossible to verify this without extensive additional testing, when changes in the fuel element design could result in a more satisfactory fuel clad lifetime, reduced damage, and greater insurance that the hydrogen migration would not become an issue.
The SNAP-8 Experimental Reactor was an important stepping stone to nuclear development in high-temperature ZrH nuclear fuel development, and greatly changed the direction of the whole SNAP-8 program in some ways. The large number of failures in cladding, the hydrogen migration from the fuel elements, and the phase changes within the crystalline structure of the U-ZrH itself were a huge wake-up call to the reactor developers. With the SNAP-2/10A reactor, these issues were minor at best, but that was a far lower-powered reactor, with very different geometry. The large number of fuel elements, the flow of the coolant through the reactor, and numerous other factors made the S8ER reactor far more complex to deal with on a practical level than most, if any, anticipated. Plating of the elements associated with Hastelloy on the stainless steel elements caused concern about the materials that had been selected causing blockages in flow channels, further exacerbating the problems of local hot spots in the fuel elements that caused many of the problems in the first place. The cladding material could (and would) be changed relatively easily to account for the problems with the metal’s ductility (the ability to undergo significant plastic deformation before rupture, in other words to endure fuel swelling without the metal splitting, cracking, fracturing or other ways that the clad could be breached) under high temperature and radiation fluxes over time. A number of changes were proposed to the reactor’s design, which strongly encouraged – or required – changes in the SNAP-8 Development Reactor that was currently being designed and fabricated. Those changes would alter what the SNAP-8 reactor would become, and what missions it would be proposed for, until the program was finally put to rest.
After the S8ER test, a mockup reactor, the SNAP-8 Development Mockup, was built based on the 1962 version of the design. This mockup never underwent nuclear testing, but was used for extensive non-nuclear testing of the designs components. Basically, every component that could be tested under non-nuclear conditions (but otherwise identical, including temperature, stress loading, vacuum, etc.) was tested and refined with this mockup. The tweaks to the design that this mockup suggested are far more minute than we have time to cover here, but it was an absolutely critical step to preparing the SNAP-8 reactor’s systems for flight test.
SNAP-8 Development Reactor: Facing Challenges with the Design
The final reactor in the series, the SNAP-8 Development Reactor, was a shorter-lived reactor, in part because many of the questions that needed to be answered about the geometry had been answered by the S8ER, and partly because the unanswered materials questions were able to be answered with the SCA4 reactor. This reactor underwent dry critical testing in June 1968, and power testing began at the beginning of the next year. From January 1969 to December 1969, when the reactor was shut down for the final time, the reactor operated at nominal (600 kWt) power for 668 hours, and operated at 1000 kWt for 429 hours.
The SNAP-8 Development Reactor (S8DR) was installed in the same facility as the S8ER, although it operated under different conditions than the S8ER. Instead of having a cover gas, the S8DR was tested in a vacuum, and a flight-type radiation shield was mounted below it to facilitate shielding design and materials choices. Fuel loading began on June 18, 1968, and criticality was achieved on June 22, with 169 out of the 211 fuel elements containing the U-ZrH fuel (the rest of the fuel elements were stainless steel “dummy” elements) installed in the core. Reactivity experiments for the control mechanisms were carried out before the remainder of the dummy fuel elements were replaced with actual fuel in order to better calibrate the system.
Finally, on June 28, all the fuel was loaded and the final calibration experiments were carried out. These tests then led to automatic startup testing of the reactor, beginning on December 13, 1968, as well as transient analysis, flow oscillation, and temperature reactivity coefficient testing on the reactor. From January 10 to 15, 1969, the reactor was started using the proposed automated startup process a total of five times, proving the design concept.
1969 saw the beginning of full-power testing, with the ramp up to full design power occurring on January 17. Beginning at 25% power, the reactor was stepped up to 50% after 8 hours, then another 8 hours in it was brought up to full power. The coolant flow rates in both the primary and secondary loops started at full flow, then were reduced to maintain design operating temperatures, even at the lower power setting. Immediately following these tests on January 23, an additional set of testing was done to verify that the power conversion system would start up as well. The biggest challenge was verification that the initial injection of mercury into the boiler would operate as expected, so a series of mercury injection tests were carried out successfully. While they weren’t precisely at design conditions due to test stand limitations, the tests were close enough that it was possible to verify that the design would work as planned.
After these tests, the endurance testing of the reactor began. From January 25 to February 24 was the 500-hour test at design conditions (600 kWt and 1300 F), although there were two scram incidents that led to short interruptions. Starting on March 20, the 9000 hour endurance run at design conditions lasted until April 10. This was followed by a ramp up to the alternate design power of 1 MWt. While this was meant to operate at only 1100 F (to reduce thermal stress on the fuel elements, among other things), the airblast heat exchanger used for heat rejection couldn’t keep up with the power flow at that temperature, so the outlet temperature was increased to 1150 F (the greater the temperature difference between a radiator and its environment, the more efficient it is, something we’ll discuss more in the heat rejection posts). After 18 days of 1 MWt testing, the power was once again reduced to 600 kWt for another 9000 hour test, but on June 1, the reactor scrammed itself again due to a loss of coolant flow. At this point, there was a significant loss of reactivity in the core, which led the team to decide to proceed at a lower temperature to mitigate hydrogen migration in the fuel elements. Sadly, reducing the outlet temperature (to 1200 F) wasn’t enough to prevent this test from ending prematurely due to a severe loss in reactivity, and the reactor scrammed itself again.
The final power test on the S8ER began on November 20, 1969. For the first 11 days, it operated at 300 kWt and 1200 F, when it was then increased in power back to 600 kWt, but the outlet temperature was reduced to 1140F, for an additional 7 days. An increase of outlet temperature back to 1200 F was then dialed in for the final 7 days of the test, and then the reactor was shut down.
This shutdown was an interesting and long process, especially compared to just removing all the reactivity of the control drums by rotating them all fully out. First, the temperature was dropped to 1000 F while the reactor was still at 600 kWt, and then the reactor’s power was reduced to the point that both the outlet and inlet coolant temperatures were 800 F. This was held until December 21 to study the xenon transient behavior, and then the temperatures were further reduced to 400 F to study the decay power level of the reactor. On January 7, the temperature was once again increased to 750 F, and two days later the coolant was removed. The core temperature then dropped steadily before leveling off at 180-200F.
Once again, the reactor was disassembled and examined at the Hot Laboratory, with special attention being paid to the fuel elements. These fuel elements held up much better than the S8ER’s fuel elements, with only 67 of the 211 fuel elements showing cracking. However, quite a few elements, while not cracked, showed significant dimensional changes and higher hydrogen loss rates. Another curiosity was that a thin (less than 0.1 mil thick) metal film, made up of iron, nickel, and chromium, developed fairly quickly on the exterior of the cladding (the exact composition changed based on location, and therefore on local temperature, within the core and along each fuel element).
The fuel elements that had intact cladding and little to no deformation showed very low hydrogen migration, an average of 2.4% (this is consistent with modeling showing that the permeation barrier was damaged early in its life, perhaps during the 1 MWt run). However, those with some damage lost between 6.8% and 13.2 percent of their hydrogen. This damage wasn’t limited to just cracked cladding, though – the swelling of the fuel element was a better indication of the amount of hydrogen lost than the clad itself being split. This is likely due to phase changes in the fuel elements, when the UzrH changes crystalline structure, usually due to high temperatures. This changes how well – and at what bond angle – the hydrogen is kept within the fuel element’s crystalline structure, and can lead to more intense hot spots in the fuel element, causing the problem to become worse. The loss of reactivity scrams from the testing in May-July 1969 seem to be consistent with the worst failures in the fuel elements, called Type 3 in the reports: high hydrogen loss, highly oval cross section of the swollen fuel elements (there were a total of 31 of these, 18 of them were intact, 13 were cracked). One interesting note about the clad composition is that where there was a higher copper content due to irregularities in metallography there was far less swelling of the Hastelloy N clad, although the precise mechanism was not understood at the time (and my rather cursory perusal of current literature didn’t show any explanation either). However, at the time testing showed that these problems could be mitigated, to the point of insignificance even, by maintaining a lower core temperature to ensure localized over-temperature failures (like the changes in crystalline structure) would not occur.
The best thing that can be said about the reactivity loss rate (partially due to hydrogen losses, and partially due to fission product buildup) is that it was able to be extrapolated given the data available, and that the failure would have occurred after the design’s required lifetime (had S8DR been operated at design temperature and power, the reactor would have lost all excess reactivity – and therefore the ability to maintain criticality – between October and November of 1970).
On this mixed news note, the reactor’s future was somewhat in doubt. NASA was certainly still interested in a nuclear reactor of a similar core power, but this particular configuration was neither the most useful to their needs, nor was it exceptionally hopeful in many of the particulars of its design. While NASA’s reassessment of the program was not solely due to the S8DR’s testing history, this may have been a contributing factor.
One way or the other, NASA was looking for something different out of the reactor system, and this led to many changes. Rather than an electric propulsion system, focus shifted to a crewed space station, which has different design requirements, most especially in shielding. In fact, the reactor was split into three designs, none of which kept the SNAP name (but all kept the fuel element and basic core geometry).
A New Life: the Children of SNAP-8
Even as the SNAP-8 Development Reactor was undergoing tests, the mission for the SNAP-8 system was being changed. This would have major consequences for the design of the reactor, its power conversion system, and what missions it would be used in. These changes would be so extensive that the SNAP-8 reactor name would be completely dropped, and the reactor would be split into four concepts.
The first concept was the Space Power Facility – Plumbrook (SPT) reactor, which would be used to test shielding and other components at NASA’s Plum Brook Research Center outside Cleveland, OH, and could also be used for space missions if needed. The smallest of the designs (at 300 kWt), it was designed to avoid many of the problems associated with the S8ER and S8DR; however, funding was cut before the reactor could be built. In fact, it was cut so early that details on the design are very difficult to find.
The second reactor, the Reactor Core Test, was very similar to the SPF reactor, but it was the same power output as the nominal “full power” reactor, at 600 kWt. Both of these designs increased the number of control drums to eight, and were designed to be used with a traditional shadow shield. Neither of them were developed to any great extent, much less built.
A third design, the 5 kWe Thermoelectric Reactor, was a space system, meant to take many of the lessons from the SNAP-8 program and apply them to a medium-power system which would apply both the lessons of the SNAP-8 ER and DR as well as the SNAP-10A’s experience with thermoelectric power conversion systems to a reactor between the SNAP-10B and Reference Zirconium Hydride reactor in power output.
The final design, the Reference Zirconium Hydride Reactor (ZrHR), was extensively developed, even if geometry-specific testing was never conducted. This was the most direct replacement for the SNAP-8 reactor, and the last of the major U-ZrH fueled space reactors in the SNAP program. Rather than powering a nuclear electric spacecraft, however, this design was meant to power space stations.
The 5 kWe Thermoelectric Reactor: Simpler, Cleaner, and More Reliable
The 5 kWe Thermoelectric Reactor (5 kWe reactor) was a reasonably simple adaptation of the SNAP-8 design, intended to be used with a shadow shield. Unsurprisingly, a lot of the design changes mirrored some of the work done on the SNAP-10B Interim design, which was undergoing work at about the same time. Meant to supply 5 kWe of power for 5 years using lead telluride thermoelectric convertors (derived from the SNAP-10A convertors), this system was meant to provide power for everything from small crewed space stations to large communications satellites. In many ways, this was a very different departure from the SNAP-8 reactor, but at the same time the changes that were proposed were based on evolutionary changes during the S8ER and S8DR experimental runs, as well as advances in the SNAP 2/10 core which was undergoing parallel post-SNAPSHOT design evolution (the SNAP-10A design had been frozen for the SNAPSHOT program at this point, so these changes were either for the followon SNAP-10A Advanced or SNAP-10B reactors). The change from mercury Rankine to thermoelectric power conversion, though, paralleled a change in the SNAP-2/10A origram, where greater efficiency was seen as unnecessary due to the constantly-lower power requirements of the systems.
The first thing (in the reactor itself, at least) that was different about the design was that the axial reflector was tapered, rather than cylindrical. This was done to keep the exterior profile of the reactor cleaner. While aerodynamic considerations aren’t a big deal (although they do still play a minute part in low Earth orbit) for astronuclear power plants, everything that’s exposed to the unshielded reactor becomes a radiation source itself, due to radiation scattering and material activation under neutron bombardment. If you could get your reactor to be a continuation of the taper of your shadow shield, rather than sticking out from that cone shape, you can make the shadow shield smaller for a given reactor. Since the shield is often many times heavier than the power system itself, especially for crewed applications, the single biggest place a designer can save mass is in the shadow shield.
This tapered profile meant two things: first, there would be a gradient in the amount of neutron moderation between the top and the bottom of the reactor, and second, the control system would have to be reworked. It’s unclear exactly how far the neutronics analysis for the new reflector configuration had proceeded, sadly, but the control systems were adaptations of the design changes that were proposed for the SNAP-10B reactor: instead of having the wide, partial cylinder control drums of the original design, large sections (235 degrees in total) of the reflector would be slid up or down around the core containment vessel to control the amount of reactivity available. This is somewhat similar to the SNAP-10B and BES-5 concepts in its execution, but the mechanism is quite different from a neutronics perspective: rather than capturing the unwanted neutrons using a neutron poison like boron or europium, they’re essentially vented into space.
A few other big changes from the SNAP-8 reference design when it comes to the core itself. The first is in the fuel: instead of having a single long fuel rod in the clad, the U-XrH fuel was split into five different “slugs,” which were held together by the clad. This would create a far more complex thermal distribution situation in the fuel, but would also allow for better thermal stress management within the hydride itself. The number of fuel elements was reduced to 85, and they came in three configurations: one set of 27 had radial fins to control the flow that spiralled around the fuel element in a right-hand direction, another set of 27 had fins in the left-hand direction, and the final 31 were unfinned. This was done to better manage the flow of the NaK coolant through the core, and avoid some of the hydrodynamic problems that were experienced on the S8DR.
The U-ZrH Reactor: Power for America’s Latest and Greatest Space Stations.
The Reference ZrH Reactor was begun in 1968, while the S8DR was still under construction. Because of this increased focus on having a crewed space station configuration, and the shielding requirement changes, some redesign of the reactor core was needed. The axial shield would change the reactivity of the core, and the control drums would no longer be able to effectively expose portions of the core to the vacuum of space to get rid of excess reactivity. Because of this, the number of fuel elements in the core were increased from 211 to 295. Another change was that rather than the even spacing of fuel elements used in the S8DR, the fuel elements were spaced in such a way that the amount of coolant around each fuel element was proportional to the amount of power produced by each fuel element. This means that the fuel elements on the interior of the core were wider spaced than the fuel elements around the periphery. This made it far more unlikely that local hot spots will develop which could lead to fuel element failures, but it also meant that the flow of coolant through the reactor core would need to be far more thoroughly studied than was done on the SNAP 8 reactor design. These thermohydrodynamic studies would be a major focus of the ZrHR program.
Another change was in the control drum configuration, as well as the need to provide coolant to the drums. This was because the drums were now not only fully enclosed solid cylinders, but were surrounded by a layer of molten lead gamma shielding. Each drum would be a solid cylinder in overall cross section; the main body was beryllium, but a crescent of europium alloy was used as a neutron poison (this is one of the more popular alternatives to boron for control mechanisms that operate in a high temperature environment) to absorb neutrons when this portion of the control drum was turned toward the core. These drums would be placed in dry wells, with NaK coolant flowing around them from the spacecraft (bottom) end before entering the upper reactor core plenum to flow through the core itself. The bearings would be identical to those used on the SNAP-8 Development Reactor, and minimal modifications would be needed for the drum motion control and position sensing apparatus. Fixed cylindrical beryllium reflectors, one small one along the interior radius of the control drums and a larger one along the outside of the drums, filled the gaps left by the control drums in this annular reflector structure. These, too, would be kept cool by the NaK coolant flowing around them.
Surrounding this would be an axial gamma shield, with the preferred material being molten lead encased in stainless steel – but tungsten was also considered as an alternative. Why the lead was kept molten is still a mystery to me, but my best guess is that this was due to the thermal conditions of the axial shield, which would have forced the lead to remain above its melting point. This shield would have made it possible to maneuver near the space station without having to remain in the shadow of the directional shield – although obviously dose rates would still be higher than being aboard the station itself.
Another interesting thing about the shielding is that the shadow shield was divided in two, in order to balance thermal transfer and radiation protection for the power conversion system, and also to maximize the effectiveness of the shadow shields. Most designs used a 4 pi shield design, which is basically a frustrum completely surrounding the reactor core with the wide end pointing at the spacecraft. The primary coolant loops wrapped around this structure before entering the thermoelectric conversion units. After this, there’s a small “galley” where the power conversion system is mounted, followed by a slightly larger shadow shield, with the heat rejection system feed loops running across the outside as well. Finally, the radiator – usually cylindrical or conical – completed the main body of the power system. The base of the radiator would meet up with the mounting hardware for attachment to the spacecraft, although the majority of the structural load was an internal spar running from the core all the way to the spacecraft.
While the option for using a pure shadow shield concept was always kept on the table, the complications in docking with a nuclear powered space station which has an unshielded nuclear reactor at one end of the structure were significant. Because of this, the ZrHR was designed with full shielding around the entire core, with supplementary shadow shields between the reactor itself and the power conversion system, and also a second shadow shield after the power conversion system. These shadow shields could be increased to so-called 4-pi shields for more complete shielding area, assuming the mission mass budget allowed, but as a general rule the shielding used was a combination of the liquid lead gamma shield and the combined shadow shield configuration. These shields would change to a fairly large extent depending on the mission that the ZrHR would be used on.
Another thing that was highly variable was the radiator configuration. Some designs had a radiator that was fixed in relation to the reactor, even if it was extended on a boom (as was the case of the Saturn V Orbital Workshop, later known as Skylab). Others would telescope out, as was the case for the later Modular Space Station (much later this became the International Space Station). The last option was for the radiators to be hinged, with flexible joints that the NaK coolant would flow through (this was the configuration for the lunar surface mission), and those joints took a lot of careful study, design, and material testing to verify that they would be reliable, seal properly, and not cause too many engineering compromises. We’ll look at the challenges of designing a radiator in the future, when we look at heat rejection systems (at this point, possibly next summer), but suffice to say that designing and executing a hinged radiator is a significant challenge for engineers, especially with a material at hot, and as reactive, as liquid NaK.
The ZrHR was continually being updated, since there was no reason to freeze the majority of the design components (although the fuel element spacing and fin configuration in the core may have indeed been frozen to allow for more detailed hydrodynamic predictability), until the program’s cancellation in 1973. Because of this, many design details were still in flux, and the final reactor configuration wasn’t ever set in stone. Additional modifications for surface use for a crewed lunar base would have required tweaking, as well, so there is a lot of variety in the final configurations.
The Stations: Orbital Missions for SNAP-8 Reactors
At the time of the redesign, three space stations were being proposed for the near future: the first, the Manned Orbiting Research Laboratory, (later changed to the Manned Orbiting Laboratory, or MOL), was a US Air Force project as part of the Blue Gemini program. Primarily designed as a surveillance platform, advances in photorecoinnasance satellites made this program redundant after just a single flight of an uncrewed, upgraded Gemini capsule.
The second was part of the Apollo Applications Program. Originally known as the Saturn V Orbital Workshop (OWS), this later evolved into Skylab. Three crews visited this space station after it was launched on the final Saturn V, and despite huge amounts of work needed to repair damage caused during a particularly difficult launch, the scientific return in everything from anatomy and physiology to meteorology to heliophysics (the study of the Sun and other stars) fundamentally changed our understanding of the solar system around us, and the challenges associated with continuing our expansion into space.
The final space station that was then under development was the Modular Space Station, which would in the late 1980s and early 1990s evolve into Space Station Freedom, and at the start of its construction in 1998 (exactly 20 years ago as of the day I’m writing this, actually) was known as the International Space Station. While many of the concepts from the MSS were carried over through its later iterations, this design was also quite different from the ISS that we know today.
Because of this change in mission, quite a few of the subsystems for the power plant were changed extensively, starting just outside the reactor core and extending through to shielding, power conversion systems, and heat rejection systems. The power conversion system was changed to four parallel thermoelectric convertors, a more advanced setup than the SNAP-10 series of reactors used. These allowed for partial outages of the PCS without complete power loss. The heat rejection system was one of the most mission-dependent structures, so would vary in size and configuration quite a bit from mission to mission. It, too, would use NaK-78 as the working fluid, and in general would be 1200 (on the OWS) to 1400 (reference mission) sq. ft in surface area. We’ll look more at these concepts in later posts on power conversion and heat rejection systems, but these changes took up a great deal of the work that was done on the ZrHR program.
One of the biggest reasons for this unusual shielding configuration was to allow a compromise between shielding mass and crew radiation dose. In this configuration, there would be three zones of radiation exposure: only shielded by the 4 pi shield during rendezvous and docking (a relatively short time period) called the rendezvous zone; a more significant shielding for the spacecraft but still slightly higher than fully shielded (because the spacecraft would be empty when docked the vast majority of the time) called the scatter shield zone; and the crewed portion of the space station itself, which would be the most shielded, called the primary shielded zone. With the 4 pi shield, the entire system would mass 24,450 pounds, of which 16,500 lbs was radiation shielding, leading to a crew dose of between 20 and 30 rem a year from the reactor.
The mission planning for the OWS was flexible in its launch configuration: it could have launched integral to the OWS on a Saturn V (although, considering the troubles that the Skylab launch actually had, I’m curious how well the system would have performed), or it could have been launched on a separate launcher and had an upper stage to attach it to the OWS. The two options proposed were either a Saturn 1B with a modified Apollo Service Module as a trans-stage, or a Titan IIIF with the Titan Trans-stage for on-orbit delivery (the Titan IIIC was considered unworkable due to mass restrictions).
After the 3-5 years of operational life, the reactor could be disposed of in two ways: either it would be deorbited into a deep ocean area (as with the SNAP-10A, although as we saw during the BES-5’s operational history this ended up not being considered a good option), or it could be boosted into a graveyard orbit. One consideration which is very different from the SNAP-10A is that the reactor would likely be intact due to the 4 pi shield, rather than burning up as the SNAP-10A would have, meaning that a terrestrial impact could lead to civilian population exposures to fission products, and also having highly enriched (although not quite bomb grade) uranium somewhere for someone to be able to relatively easily pick up. This made the deorbiting of the reactor a bit pickier in terms of location, and so an uncontrolled re-entry was not considered. The ideal was to leave it in a parking orbit of at least 400 nautical miles in altitude for a few hundred years to allow the fission products to completely decay away before de-orbiting the reactor over the ocean.
Nuclear Power for the Moon
The final configuration that was examined for the Advanced ZrH Reactor was for the lunar base that was planned as a follow-on to the Apollo Program. While this never came to fruition, it was still studied carefully. Nuclear power on the Moon was nothing new: the SNAP-27 radioisotope thermoelectric generator had been used on every single Apollo surface mission as part of the Apollo Lunar Surface Experiment Package (ALSEP). However, these RTGs would not provide nearly enough power for a permanently crewed lunar base. As an additional complication, all of the power sources available would be severely taxed by the 24 day long, incredibly cold lunar night that the base would have to contend with. Only nuclear fission offered both the power and the heat needed for a permanently staffed lunar base, and the reactor that was considered the best option was the Advanced ZrH Reactor.
The configuration of this form of the reactor was very different. There are three options for a surface power plant: the reactor is offloaded from the lander and buried in the lunar regolith for shielding (which is how the Kilopower reactor is being planned for surface operations); an integral lander and power plant which is assembled in Earth (or lunar) orbit before landing, with a 4 pi shield configuration; finally an integrated lander and reactor with a deployable radiator which is activated once the reactor is on the surface of the moon, again with a 4 pi shield configuration. There are, of course, in-between options between the last two configurations, where part of the radiator is fixed and part deploys. The designers of the ZrHR decided to go with the second option as their best design option, due to the ability to check out the reactor before deployment to the lunar surface but also minimizing the amount of effort needed by the astronauts to prepare the reactor for power operations after landing. This makes sense because, while on-orbit assembly and checkout is a complex and difficult process, it’s still cheaper in terms of manpower to do this work in Earth orbit rather than a lunar EVA due to the value of every minute on the lunar surface. If additional heat rejection was required, a deployable radiator could be used, but this would require flexible joints for the NaK coolant, which would pose a significant materials and design challenge. A heat shield was used when the reactor wasn’t in operation to prevent exessive heat loss from the reactor. This eased startup transient issues, as well as ensuring that the NaK coolant remained liquid even during reactor shutdown (frozen working fluids are never good for a mechanical system, after all). The power conversion system was exactly the same configuration as would be used in the OWS configuration that we discussed earlier, with the upgraded, larger tubes rather than the smaller, more numerous ones (we’ll discuss the tradeoffs here in the power conversion system blog posts).
This power plant would end up providing a total of 35.5 kWe of conditioned (i.e. usable, reliable power) electricity out of the 962 kWt reactor core, with 22.9 kWe being delivered to the habitat itself, for at least 5 years. The overall power supply system, including radiator, shield, power conditioning unit, and the rest of the ancillary bits and pieces that make a nuclear reactor core into a fission power plant, ended up massing a total of 23,100 lbs, which is comfortably under the 29,475 lb weight limit of the lander design that was selected (unfortunately, finding information on that design is proving difficult). A total dose rate at a half mile for an unshielded astronaut would be 7.55 mrem/hr was considered sufficient for crew radiation safety (this is a small radiation dose compared to the lunar radiation environment, and the astronauts will spend much of their time in the much better shielded habitat.
Sadly, this power supply was not developed to a great extent (although I was unable to find the source document for this particular design: NAA-SR-12374, “Reactor Power Plants for Lunar Base Applications, Atomics International 1967), because the plans for the crewed lunar base were canceled before much work was done on this design. The plans were developed to the point that future lunar base plans would have a significant starting off point, but again the design was never frozen, so there was a lot of flexibility remaining in the design.
The End of the Line
Sadly, these plans never reached fruition. The U-ZrH Reactor had its budget cut by 75% in 1971, with cuts to alternate power conversion systems such as the use of thermionic power conversion (30%) and reactor safety (50%), and the advanced Brayton system (completely canceled) happening at the same time. NERVA, which we covered in a number of earlier posts, also had its funding slashed at the same time. This was due to a reorientation of funds away from many current programs, and instead focusing on the Space Shuttle and a modular space station, whose power requirements were higher than the U-ZrH Reactor would be able to offer.
At this point, the AEC shifted their funding philosophy, moving away from preparing specific designs for flight readiness and instead moving toward a long-term development strategy. In 1973 head of the AEC’s Space Nuclear Systems Division said that, given the lower funding levels that NASA was forced to work within, “…the missions which were likely to require large amounts of energy, now appear to be postponed until around 1990 or later.” This led to the cancellation of all nuclear reactor systems, and a shift in focus to radioisotope thermoelectric generators, which gave enough power for NASA and the DoD’s current mission priorities in a far simpler form.
Funding would continue at a low level all the way to the current day for space fission power systems, but the shift in focus led to a very different program. While new reactor concepts continue to be regularly put forth, both by Department of Energy laboratories and NASA, for decades the focus was more on enhancing the technological capability of many areas, especially materials, which could be used by a wide range of reactor systems. This meant that specific systems wouldn’t be developed to the same level of technological readiness in the US for over 30 years, and in fact it wouldn’t be until 2018 that another fission power system of US design would undergo criticality testing (the KRUSTY test for Kilopower, in early 2018).
More Coming Soon!
Originally, I was hoping to cover another system in this blog post as well, but the design is so different compared to the ZrH fueled reactors that we’ve been discussing so far in this series that it warranted its own post. This reactor is the SNAP-50, which didn’t start out as a space reactor, but rather one of the most serious contenders for the indirect-cycle Aircraft Nuclear Propulsion program. It used uranium carbide/nitride fuel elements, liquid lithium coolant, and was far more powerful than anything that weve discussed yet in terms of electric power plants. Having it in its own post will also allow me to talk a little bit about the ANP program, something that I’ve wanted to cover for a while now, but considering how much more there is to discuss about in-space systems (and my personal aversion to nuclear reactors for atmospheric craft on Earth), hasn’t really been in the cards until now.
This series continues to expand, largely because there’s so much to cover that we haven’t gotten to yet – and no-one else has covered these systems much either! I’m currently planning on doing the SNAP-50/SPUR system as a standalone post, followed by the SP-100 and a selection of other reactor designs. After that, we’ll cover the ENISY reactor program in its own post, followed by the NEP designs from the 90s and early 00s, both in the US and Russia. Finally, we’ll cover the predecessors to Kilopower, and round out our look at fission power plant cores by revisiting Kilopower to have a look at what’s changed, and what’s stayed the same, over the last year since the KRUSTY test. We will then move on to shielding materials and design (probably two or three posts, because there’s a lot to cover there) before moving on to power conversion systems, another long series. We’ll finish up the nuclear systems side of nuclear power supplies by looking at heat sinks, radiators, and other heat rejection systems, followed by a look at nuclear electric spacecraft architecture and design considerations.
A lot of work continues in the background, especially in terms of website planning and design, research on a lot of the lesser-known reactor systems, and planning for the future of the web page. The blog is definitely set for topics for at least another year, probably more like two, just covering the basics and history of astronuclear design, but getting the web page to be more functional is a far more complex, and planning-heavy, task.
I hope you enjoyed this post, and much more is coming next year! Don’t forget to join us on Facebook, or follow me on Twitter!
Hello, and welcome to Beyond NERVA! Today we’re going to look at the program that birthed the first astronuclear reactor to go into orbit, although the extent of the program far exceeds the flight record of a single launch.
Before we get into that, I have a minor administrative announcement that will develop into major changes for Beyond NERVA in the near-to-mid future! As you may have noticed, we have moved from beyondnerva.wordpress.com to beyondnerva.com. For the moment, there isn’t much different, but in the background a major webpage update is brewing! Not only will the home page be updated to make it easier to navigate the site (and see all the content that’s already available!), but the number of pages on the site is going to be increasing significantly. A large part of this is going to be integrating information that I’ve written about in the blog into a more topical, focused format – with easier access to both basic concepts and technical details being a priority. However, there will also be expansions on concepts, pages for technical concepts that don’t really fit anywhere in the blog posts, and more! As these updates become live, I’ll mention them in future blog posts. Also, I’ll post them on both the Facebook group and the new Twitter feed (currently not super active, largely because I haven’t found my “tweet voice yet,” but I hope to expand this soon!). If you are on either platform, you should definitely check them out!
The Systems for Nuclear Auxiliary Propulsion, or SNAP program, was a major focus for a wide range of organizations in the US for many decades. The program extended everywhere from the bottom of the seas (SNAP-4, which we won’t be covering in this post) to deep space travel with electric propulsion. SNAP was divided up into an odd/even numbering scheme, with the odd model numbers (starting with the SNAP-3) being radioisotope thermoelectric generators, and the even numbers (beginning with SNAP-2) being fission reactor electrical power systems.
Due to the sheer scope of the SNAP program, even eliminating systems that aren’t fission-based, this is going to be a two post subject. This post will cover the US Air Force’s portion of the SNAP reactor program: the SNAP-2 and SNAP-10A reactors; their development programs; the SNAPSHOT mission; and a look at the missions that these reactors were designed to support, including satellites, space stations, and other crewed and uncrewed installations. The next post will cover the NASA side of things: SNAP-8 and its successor designs as well as SNAP-50/SPUR. The one after that will cover the SP-100, SABRE, and other designs from the late 1970s through to the early 1990s, and will conclude with looking at a system that we mentioned briefly in the last post: the ENISY/TOPAZ II reactor, the only astronuclear design to be flight qualified by the space agencies and nuclear regulatory bodies of two different nations.
The Beginnings of the US Astronuclear Program: SNAP’s Early Years
Beginning in the earliest days of both the nuclear age and the space age, nuclear power had a lot of appeal for the space program: high power density, high power output, and mechanically simple systems were in high demand for space agencies worldwide. The earliest mention of a program to develop nuclear electric power systems for spacecraft was the Pied Piper program, begun in 1954. This led to the development of the Systems for Nuclear Auxiliary Power program, or SNAP, the following year (1955), which was eventually canceled in 1973, as were so many other space-focused programs.
Once space became a realistic place to send not only scientific payloads but personnel, the need to provide them with significant amounts of power became evident. Not only were most systems of the day far from the electricity efficient designs that both NASA and Roscosmos would develop in the coming decades; but, at the time, the vision for a semi-permanent space station wasn’t 3-6 people orbiting in a (completely epic, scientifically revolutionary, collaboratively brilliant, and invaluable) zero-gee conglomeration of tin cans like the ISS, but larger space stations that provided centrifugal gravity, staffed ‘round the clock by dozens of individuals. These weren’t just space stations for NASA, which was an infant organization at the time, but the USAF, and possibly other institutions in the US government as well. In addition, what would provide a livable habitation for a group of astronauts would also be able to power a remote, uncrewed radar station in the Arctic, or in other extreme environments. Even if crew were there, the fact that the power plant wouldn’t have to be maintained was a significant military advantage.
Responsible for both radioisotope thermoelectric generators (which run on the natural radioactive decay of a radioisotope, selected according to its energy density and half-life) as well as fission power plants, SNAP programs were numbered with an even-odd system: even numbers were fission reactors, odd numbers were RTGs. These designs were never solely meant for in-space application, but the increased mission requirements and complexities of being able to safely launch a nuclear power system into space made this aspect of their use the most stringent, and therefore the logical one to design around. Additionally, while the benefits of a power-dense electrical supply are obvious for any branch of the military, the need for this capability in space far surpassed the needs of those on the ground or at sea.
Originally jointly run by the AEC’s Department of Reactor Development (who funded the reactor itself) and the USAF’s AF Wright Air Development Center (who funded the power conversion system), full control was handed over to the AEC in 1957. Atomics International Research was the prime contractor for the program.
There are a number of similarities in almost all the SNAP designs, probably for a number of reasons. First, all of the reactors that we’ll be looking at (as well as some other designs we’ll look at in the next post) used the same type of fissile fuel, even though the form, and the cladding, varied reasonably widely between the different concepts. Uranium-zirconium hydride (U-ZrH) was a very popular fuel choice at the time. Assuming hydrogen loss could be controlled (this was a major part of the testing regime in all the reactors that we’ll look at), it provided a self-moderating, moderate-to-high-temperature fuel form, which was a very attractive feature. This type of fuel is still used today, for the TRIGA reactor – which, between it and its direct descendants is the most common form of research and test reactor worldwide. The high-powered reactors (SNAP 2 and 8) both initially used variations on the same power conversion system: a boiling mercury Rankine power conversion cycle, which was determined by the end of the testing regime to be possible to execute, however to my knowledge has never been proposed again (we’ll look at this briefly in the post on heat engines as power conversion systems, and a more in-depth look will be available in the future), although a mercury-based MHD conversion system is being offered as a power conversion system for an accelerator-driven molten salt reactor.
SNAP-2: The First American Built-For-Space Nuclear Reactor Design
The idea for the SNAP-2 reactor originally came from a 1951 Rand Corporation study, looking at the feasibility of having a nuclear powered satellite. By 1955, the possibilities that a fission power supply offered in terms of mass and reliability had captured the attention of many people in the USAF, which was (at the time) the organization that was most interested and involved (outside the Army Ballistic Missile Agency at the Redstone Arsenal, which would later become the Goddard Spaceflight Center) in the exploitation of space for military purposes.
The original request for the SNAP program, which ended up becoming known as SNAP 2, occurred in 1955, from the AEC’s Defense Reactor Development Division and the USAF Wright Air Development Center. It was for possible power sources in the 1 to 10 kWe range that would be able to autonomously operate for one year, and the original proposal was for a zirconium hydride moderated sodium-potassium (NaK) metal cooled reactor with a boiling mercury Rankine power conversion system (similar to a steam turbine in operational principles, but we’ll look at the power conversion systems more in a later post), which is now known as SNAP-2. The design was refined into a 55 kWt, 5 kWe reactor operating at about 650°C outlet temperature, massing about 100 kg unshielded, and was tested for over 10,000 hours. This epithermal neutron spectrum would remain popular throughout much of the US in-space reactor program, both for electrical power and thermal propulsion designs. This design would later be adapted to the SNAP-10A reactor, with some modifications, as well.
SNAP-2’s first critical assembly test was in October of 1957, shortly after Sputnik-1’s successful launch. With 93% enriched 235U making up 8% of the weight of the U-ZrH fuel elements, a 1” beryllium inner reflector, and an outer graphite reflector (which could be varied in thickness), separated into two rough hemispheres to control the construction of a critical assembly; this device was able to test many of the reactivity conditions needed for materials testing on a small economic scale, as well as test the behavior of the fuel itself. The primary concerns with testing on this machine were reactivity, activation, and intrinsic steady state behavior of the fuel that would be used for SNAP-2. A number of materials were also tested for reflection and neutron absorbency, both for main core components as well as out-of-core mechanisms. This was followed by the SNAP-2 Experimental Reactor in 1959-1960 and the SNAP 2 Development Reactor in 1961-1962.
The SNAP-2 Experimental Reactor (S2ER or SER) was built to verify the core geometry and basic reactivity controls of the SNAP-2 reactor design, as well as to test the basics of the primary cooling system, materials, and other basic design questions, but was not meant to be a good representation of the eventual flight system. Construction started in June 1958, with construction completed by March 1959. Dry (Sept 15) and wet (Oct 20) critical testing was completed the same year, and power operations started on Nov 5, 1959. Four days later, the reactor reached design power and temperature operations, and by April 23 of 1960, 1000 hours of continuous testing at design conditions were completed. Following transient and other testing, the reactor was shut down for the last time on November 19, 1960, just over one year after it had first achieved full power operations. Between May 19 and June 15, 1961, the reactor was disassembled and decommissioned. Testing on various reactor materials, especially the fuel elements, was conducted, and these test results refined the design for the Development Reactor.
The SNAP-2 Development Reactor (S2DR or SDR, also called the SNAP-2 Development System, S2DS) was installed in a new facility at the Atomics International Santa Susana research facility to better manage the increased testing requirements for the more advanced reactor design. While this wasn’t going to be a flight-type system, it was designed to inform the flight system on many of the details that the S2ER wasn’t able to. This, interestingly, is much harder to find information on than the S2ER. This reactor incorporated many changes from the S2ER, and went through several iterations to tweak the design for a flight reactor. Zero power testing occurred over the summer of 1961, and testing at power began shortly after (although at SNAP-10 power and temperature levels. Testing continued until December of 1962, and further refined the SNAP-2 and -10A reactors.
A third set of critical assembly reactors, known as the SNAP Development Assembly series, was constructed at about the same time, meant to provide fuel element testing, criticality benchmarks, reflector and control system worth, and other core dynamic behaviors. These were also built at the Santa Susana facility, and would provide key test capabilities throughout the SNAP program. This water-and-beryllium reflected core assembly allowed for a wide range of testing environments, and would continue to serve the SNAP program through to its cancellation. Going through three iterations, the designs were used more to test fuel element characteristics than the core geometries of individual core concepts. This informed all three major SNAP designs in fuel element material and, to a lesser extent, heat transfer (the SNAP-8 used thinner fuel elements) design.
Extensive testing was carried out on all aspects of the core geometry, fuel element geometry and materials, and other behaviors of the reactor; but by May 1960 there was enough confidence in the reactor design for the USAF and AEC to plan on a launch program for the reactor (and the SNAP-10A), called SNAPSHOT (more on that below). Testing using the SNAP-2 Experimental Reactor occurred in 1960-1961, and the follow-on test program, including the Snap 2 Development reactor occurred in 1962-63. These programs, as well as the SNAP Critical Assembly 3 series of tests (used for SNAP 2 and 10A), allowed for a mostly finalized reactor design to be completed.
The power conversion system (PCS), a Rankine (steam) turbine using mercury, were carried out starting in 1958, with the development of a mercury boiler to test the components in a non-nuclear environment. The turbine had many technical challenges, including bearing lubrication and wear issues, turbine blade pitting and erosion, fluid dynamics challenges, and other technical difficulties. As is often the case with advanced reactor designs, the reactor core itself wasn’t the main challenge, nor the control mechanisms for the reactor, but the non-nuclear portions of the power unit. This is a common theme in astronuclear engineering. More recently, JIMO experienced similar problems when the final system design called for a theoretical but not yet experimental supercritical CO2 Brayton turbine (as we’ll see in a future post). However, without a power conversion system of useable efficiency and low enough mass, an astronuclear power system doesn’t have a means of delivering the electricity that it’s called upon to deliver.
Reactor shielding, in the form of a metal honeycomb impregnated with a high-hydrogen material (in this case a form of paraffin), was common to all SNAP reactor designs. The high hydrogen content allowed for the best hydrogen density of the available materials, and therefore the greatest shielding per unit mass of the available options.
Testing on the SNAP 2 reactor system continued until 1963, when the reactor core itself was re-purposed into the redesigned SNAP-10, which became the SNAP-10A. At this point the SNAP-2 reactor program was folded into the SNAP-10A program. SNAP-2 specific design work was more or less halted from a reactor point of view, due to a number of factors, including the slower development of the CRU power conversion system, the large number of moving parts in the Rankine turbine, and the advances made in the more powerful SNAP-8 family of reactors (which we’ll cover in the next post). However, testing on the power conversion system continued until 1967, due to its application to other programs. This didn’t mean that the reactor was useless for other missions; in fact, it was far more useful, due to its far more efficient power conversion system for crewed space operations (as we’ll see later in this post), especially for space stations. However, even this role would be surpassed by a derivative of the SNAP-8, the Advanced ZrH Reactor, and the SNAP-2 would end up being deprived of any useful mission.
The SNAP Reactor Improvement Program, in 1963-64, continued to optimize and advance the design without nuclear testing, through computer modeling, flow analysis, and other means; but the program ended without flight hardware being either built or used. We’ll look more at the missions that this reactor was designed for later in this blog post, after looking at its smaller sibling, the first reactor (and only US reactor) to ever achieve orbit: the SNAP-10A.
SNAP-10: The Father of the First Reactor in Space
At about the same time as the SNAP 2 Development Reactor tests (1958), the USAF requested a study on a thermoelectric power conversion system, targeting a 0.3 kWe-1kWe power regime. This was the birth of what would eventually become the SNAP-10 reactor. This reactor would evolve in time to become the SNAP-10A reactor, the first nuclear reactor to go into orbit.
In the beginning, this design was superficially quite similar to the Romashka reactor that we’ll examine in the USSR part of this blog post, with plates of U-ZrH fuel, separated by beryllium plates for heat conduction, and surrounded by radial and axial beryllium reflectors. Purely conductively cooled internally, and radiatively cooled externally, this was later changed to a NaK forced convection cooling system for better thermal management (see below). The resulting design was later adapted to the SNAP-4 reactor, which was designed to be used for underwater military installations, rather than spaceflight. Outside these radial reflectors were thermoelectric power conversion systems, with a finned radiating casing being the only major component that was visible. The design looked, superficially at least, remarkably like the RTGs that would be used for the next several decades. However, the advantages to using even the low power conversion efficiency thermoelectric conversion system made this a far more powerful source of electricity than the RTGs that were available at the time (or even today) for space missions.
Within a short period, however, the design was changed dramatically, resulting in a design very similar to the core for the SNAP-2 reactor that was under development at the same time. Modifications were made to the SNAP-2 baseline, resulting in the reactor cores themselves becoming identical. This also led to the NaK cooling system being implemented on the SNAP-10A. Many of the test reactors for the SNAP-2 system were also used to develop the SNAP-10A. This is because the final design, while lower powered, by 20 kWe of electrical output, was largely different in the power conversion system, not the reactor structure. This reactor design was tested extensively, with the S2ER, S2DR, and SCA test series (4A, 4B, and 4C) reactors, as well as the SNAP-10 Ground Test Reactor (S10FS-1). The new design used a similar, but slightly smaller, conical radiator using NaK as the working fluid for the radiator.
This was a far lower power design than the SNAP-2, coming in at 30 kWt, but with the 1.6% power conversion ratio of the thermoelectric systems, its electrical power output was only 500 We. It also ran almost 100°C cooler (200 F), allowing for longer fuel element life, but less thermal gradient to work with, and therefore less theoretical maximum efficiency. This tradeoff was the best on offer, though, and the power conversion system’s lack of moving parts, and ease of being tested in a non-nuclear environment without extensive support equipment, made it more robust from an engineering point of view. The overall design life of the reactor, though, remained short: only about 1 year, and less than 1% fissile fuel burnup. It’s possible, and maybe even likely, that (barring spacecraft-associated failure) the reactor could have provided power for longer durations; however, the longer the reactor operates, the more the fuel swells, due to fission product buildup, and at some point this would cause the clad of the fuel to fail. Other challenges to reactor design, such as fission poison buildup, clad erosion, mechanical wear, and others would end the reactor’s operational life at some point, even if the fuel elements could still provide more power.
The SNAP-10A was not meant to power crewed facilities, since the power output was so low that multiple installations would be needed. This meant that, while all SNAP reactors were meant to be largely or wholly unmaintained by crew personnel, this reactor had no possibility of being maintained. The reliability requirements for the system were higher because of this, and the lack of moving parts in the power conversion system aided in this design requirement. The design was also designed to only have a brief (72 hour) time period where active reactivity control would be used, to mitigate any startup transients, and to establish steady-state operations, before the active control systems would be left in their final configuration, leaving the reactor entirely self-regulating. This placed additional burden on the reactor designers to have a very strong understanding of the behavior of the reactor, its long-term stability, and any effects that would occur during the year-long lifetime of the system.
At the end of the reactor’s life, it was designed to stay in orbit until the short-lived and radiotoxic portions of the reactor had gone through at least five product half-lives, reducing the radioactivity of the system to a very low level. At the end of this process, the reactor would re-enter the atmosphere, the reflectors and end reflector would be ejected, and the entire thing would burn up in the upper atmosphere. From there, winds would dilute any residual radioactivity to less than what was released by a single small nuclear test (which were still being conducted in Nevada at the time). While there’s nothing wrong with this approach from a health physics point of view, as we saw in the last post on the BES-5 reactors the Soviet Union was flying, there are major international political problems with this concept. The SNAPSHOT reactor continues to orbit the Earth (currently at an altitude of roughly 1300 km), and will do so for more than 2000 years, according to recent orbital models, so the only system of concern is not in danger of re-entry any time soon; but, at some point, the reactor will need to be moved into a graveyard orbit or collected and returned to Earth – a problem which currently has no solution.
The Runup to Flight: Vehicle Verification and Integration
1960 brought big plans for orbital testing of both the SNAP-2 and SNAP-10 reactors, under the program SNAPSHOT: Two SNAP-10 launches, and two SNAP-2 launches would be made. Lockheed Missiles System Division was chosen as the launch vehicle, systems integration, and launch operations contractor for the program; while Atomics International, working under the AEC, was responsible for the power plant.
The SNAP-10A reactor design was meant to be decommissioned by orbiting for long enough that the fission product inventory (the waste portion of the burned fuel elements, and the source of the vast majority of the radiation from the reactor post-fission) would naturally decay away, and then the reactor would be de-orbited, and burn up in the atmosphere. This was planned before the KOSMOS-954 accident, when the possibility of allowing a nuclear reactor to burn up in the atmosphere was not as anathema as it is today. This plan wouldn’t increase the amount of radioactivity that the public would receive to any appreciable degree; and, at the time, open-air testing of nuclear weapons was the norm, sending up thousands of kilograms of radioactive fallout per year. However, it was important that the fuel rods themselves would burn up high in the atmosphere, in order to dilute the fuel elements as much as possible, and this is something that needed to be tested.
Enter the SNAP Reactor Flight Demonstration Number 1 mission, or RFD-1. The concept of this test was to demonstrate that the planned disassembly and burnup process would occur as expected, and to inform the further design of the reactor if there were any unexpected effects of re-entry. Sandia National Labs took the lead on this part of the SNAPSHOT program. After looking at the budget available, the launch vehicles available, and the payloads, the team realized that orbiting a nuclear reactor mockup would be too expensive, and another solution needed to be found. This led to the mission design of RFD-1: a sounding rocket would be used, and the core geometry would be changed to account for the short flight time, compared to a real reentry, in order to get the data needed for the de-orbiting testing of the actual SNAP-10A reactor that would be flown.
So what does this mean? Ideally, the development of a mockup of the SNAP-10A reactor, with the only difference being that there wouldn’t be any highly enriched uranium in the fuel elements, as normally configured; instead depleted uranium would be used. It would be launched on the same launch vehicle that the SNAPSHOT mission would use (an Atlas-Agena D), be placed in the same orbit, and then be deorbited at the same angle and the same place as the actual reactor would be; maybe even in a slightly less favorable reentry angle to know how accurate the calculations were, and what the margin of error would be. However, an Atlas-Agena rocket isn’t a cheap piece of hardware, either to purchase, or to launch, and the project managers knew that they wouldn’t be able to afford that, so they went hunting for a more economical alternative.
This led the team to decide on a NASA Scout sounding rocket as the launch vehicle, launched from Wallops Island launch site (which still launches sounding rockets, as well as the Antares rocket, to this day, and is expanding to launch Vector Space and RocketLab orbital rockets as well in the coming years). Sounding rockets don’t reach orbital altitudes or velocities, but they get close, and so can be used effectively to test orbital components for systems that would eventually fly in orbit, but for much less money. The downside is that they’re far smaller, with less payload and less velocity than their larger, orbital cousins. This led to needing to compromise on the design of the dummy reactor in significant ways – but those ways couldn’t compromise the usefulness of the test.
Sandia Corporation (which runs Sandia National Laboratories to this day, although who runs Sandia Corp changes… it’s complicated) and Atomics International engineers got together to figure out what could be done with the Scout rocket and a dummy reactor to provide as useful an engineering validation as possible, while sticking within the payload requirements and flight profile of the relatively small, suborbital rocket that they could afford. Because the dummy reactor wouldn’t be going nearly as fast as it would during true re-entry, a steeper angle of attack when the test was returning to Earth was necessary to get the velocity high enough to get meaningful data.
The Scout rocket that was being used had much less payload capability than the Atlas rocket, so if there was a system that could be eliminated, it was removed to save weight. No NaK was flown on RFD-1, the power conversion system was left off, the NaK pump was simulated by an empty stainless steel box, and the reflector assembly was made out of aluminum instead of beryllium, both for weight and toxicity reasons (BeO is not something that you want to breathe!). The reactor core didn’t contain any dummy fuel elements, just a set of six stainless steel spacers to keep the grid plates at the appropriate separation. Because the angle of attack was steeper, the test would be shorter, meaning that there wouldn’t be time for the reactor’s reflectors to degrade enough to release the fuel elements. The fuel elements were the most important part of the test, however, since it needed to be demonstrated that they would completely burn up upon re-entry, so a compromise was found.
The fuel elements would be clustered on the outside of the dummy reactor core, and ejected early in the burnup test period. While the short time and high angle of attack meant that there wouldn’t be enough time to observe full burnup, the beginning of the process would be able to provide enough data to allow for accurate simulations of the process to be made. How to ensure that this data, which was the most important part of the test, would be able to be collected was another challenge, though, which forced even more compromises for RFT-1’s design. Testing equipment had to be mounted in such a way as to not change the aerodynamic profile of the dummy reactor core. Other minor changes were needed as well, but despite all of the differences between the RFD-1 and the actual SNAP-10A the thermodynamics and aerodynamics of the system were different in only very minor ways.
Testing support came from Wallops Island and NASA’s Bermuda tracking station, as well as three ships and five aircraft stationed near the impact site for radar observation. The ground stations would provide both radar and optical support for the RFD-1 mission, verifying reactor burnup, fuel element burnup, and other test objective data, while the aircraft and ships were primarily tasked with collecting telemetry data from on-board instruments, as well as providing additional radar data; although one NASA aircraft carried a spectrometer in order to analyze the visible radiation coming off the reentry vehicle as it disintegrated.
The test went largely as expected. Due to the steeper angle of attack, full fuel element burnup wasn’t possible, even with the early ejection of the simulated fuel rods, but the amount that they did disintegrate during the mission showed that the reactor’s fuel would be sufficiently distributed at a high enough altitude to prevent any radiological risk. The dummy core behaved mostly as expected, although there were some disagreements between the predicted behavior and the flight data, due to the fact that the re-entry vehicle was on such a steep angle of attack. However, the test was considered a success, and paved the way for SNAPSHOT to go forward.
The next task was to mount the SNAP-10A to the Agena spacecraft. Because the reactor was a very different power supply than was used at the time, special power conditioning units were needed to transfer power from the reactor to the spacecraft. This subsystem was mounted on the Agena itself, along with tracking and command functionality, control systems, and voltage regulation. While Atomics International worked to ensure the reactor would be as self-contained as possible, the reactor and spacecraft were fully integrated as a single system. Besides the reactor itself, the spacecraft carried a number of other experiments, including a suite of micrometeorite detectors and an experimental cesium contact thruster, which would operate from a battery system that would be recharged by electricity produced by the reactor.
In order to ensure the reactor would be able to be integrated to the spacecraft, a series of Flight System Prototypes (FSM-1, and -4; FSEM-2 and -3 were used for electrical system integration) were built. These were full scale, non-nuclear mockups that contained a heating unit to simulate the reactor core. Simulations were run using FSM-1 from launch to startup on orbit, with all testing occurring in a vacuum chamber. The final one of the series, FSM-4, was the only one that used NaK coolant in the system, which was used to verify that the thermal performance of the NaK system met with flight system requirements. FSEM-2 did not have a power system mockup, instead it used a mass mockup of the reactor, power conversion system, radiator, and other associated components. Testing with FSEM-2 showed that there were problems with the original electrical design of the spacecraft, which required a rebuild of the test-bed, and a modification of the flight system itself. Once complete, the renamed FSEM-2A underwent a series of shock, vibration, acceleration, temperature, and other tests (known as the “Shake and Bake” environmental tests), which it subsequently passed. The final mockup, FSEM-3, underwent extensive electrical systems testing at Lockheed’s Sunnyvale facility, using simulated mission events to test the compatibility of the spacecraft and the reactor. Additional electrical systems changes were implemented before the program proceeded, but by the middle of 1965, the electrical system and spacecraft integration tests were complete and the necessary changes were implemented into the flight vehicle design.
The last round of pre-flight testing was a test of a flight-configured SNAP-10A reactor under fission power. This nuclear ground test, S10F-3, was identical to the system that would fly on SNAPSHOT, save some small ground safety modifications, and was tested from January 22 1965 to March 15, 1966. It operated uninterrupted for over 10,000 hours, with the first 390 days being at a power output of 35 kWt, and (following AEC approval) an additional 25 days of testing at 44 kWt. This testing showed that, after one year of operation, the continuing problem of hydrogen redistribution caused the reactor’s outlet temperature to drop more than expected, and additional, relatively minor, uncertainties about reactor dynamics were seen as well. However, overall, the test was a success, and paved the way for the launch of the SNAPSHOT spacecraft in April 1965; and the continued testing of S10F-3 during the SNAPSHOT mission was able to verify that the thermal behavior of astronuclear power systems during ground test is essentially identical to orbiting systems, proving the ground test strategy that had been employed for the SNAP program.
SNAPSHOT: The First Nuclear Reactor in Space
In 1963 there was a change in the way the USAF was funding these programs. While they were solely under the direction of the AEC, the USAF still funded research into the power conversion systems, since they were still operationally useful; but that changed in 1963, with the removal of the 0.3 kWe to 1 kWe portion of the program. Budget cuts killed the Zr-H moderated core of the SNAP-2 reactor, although funding continued for the Hg vapor Rankine conversion system (which was being developed by TRW) until 1966. The SNAP-4 reactor, which had not even been run through criticality testing, was canceled, as was the planned flight test of the SNAP-10A, which had been funded under the USAF, because they no longer had an operational need for the power system with the cancellation of the 0.3-1 kWe power system program. The associated USAF program that would have used the power supply was well behind schedule and over budget, and was canceled at the same time.
The USAF attempted to get more funding, but was denied. All parties involved had a series of meetings to figure out what to do to save the program, but the needed funds weren’t forthcoming. All partners in the program worked together to try and have a reduced SNAPSHOT program go through, but funding shortfalls in the AEC (who received only $8.6 million of the $15 million they requested), as well as severe restrictions on the Air Force (who continued to fund Lockheed for the development and systems integration work through bureaucratic creativity), kept the program from moving forward. At the same time, it was realized that being able to deliver kilowatts or megawatts of electrical power, rather than the watts currently able to be produced, would make the reactor a much more attractive program for a potential customer (either the USAF or NASA).
Finally, in February of 1964 the Joint Congressional Committee on Atomic Energy was able to fund the AEC to the tune of $14.6 million to complete the SNAP-10A orbital test. This reactor design had already been extensively tested and modeled, and unlike the SNAP-2 and -8 designs, no complex, highly experimental, mechanical-failure-prone power conversion system was needed.
SNAPSHOT consisted of a SNAP-10A fission power system mounted to a modified Agena-D spacecraft, which by this time was an off-the-shelf, highly adaptable spacecraft used by the US Air Force for a variety of missions. An experimental cesium contact ion thruster (read more about these thrusters on the Gridded Ion Engine page) was installed on the spacecraft for in-flight testing. The mission was to validate the SNAP-10A architecture with on-orbit experience, proving the capability to operate for 9 days without active control, while providing 500 W (28.5 V DC) of electrical power. Additional requirements included the use of a SNAP-2 reactor core with minimal modification (to allow for the higher-output SNAP-2 system with its mercury vapor Rankine power conversion system to be validated as well, when the need for it arose), eliminating the need (while offering the option) for active control of the reactor once startup was achieved for one year (to prove autonomous operation capability); facilitating safe ground handling during spacecraft integration and launch; and, accommodating future growth potential in both available power and power-to-weight ratio.
While the threshold for mission success was set at 90 days, for Atomics International wanted to prove 1 year of capability for the system; so, in those 90 days, the goal was that the entire reactor system would be demonstrated to be capable of one year of operation (the SNAP-2 requirements). Atomics International imposed additional, more stringent, guidelines for the mission as well, specifying a number of design requirements, including self-containment of the power system outside the structure of the Agena, as much as possible; more stringent mass and center-of-gravity requirements for the system than specified by the US Air Force; meeting the military specifications for EM radiation exposure to the Agena; and others.
The flight was formally approved in March, and the launch occurred on April 3, 1965 on an Atlas-Agena D rocket from Vandenberg Air Force Base. The launch went perfectly, and placed the SNAPSHOT spacecraft in a polar orbit, as planned. Sadly, the mission was not one that could be considered either routine or simple. One of the impedance probes failed before launch, and a part of the micrometeorite detector system failed before returning data. A number of other minor faults were detected as well, but perhaps the most troubling was that there were shorts and voltage irregularities coming from the ion thruster, due to high voltage failure modes, as well as excessive electromagnetic interference from the system, which reduced the telemetry data to an unintelligible mess. This was shut off until later in the flight, in order to focus on testing the reactor itself.
The reactor was given the startup order 3.5 hours into the flight, when the two gross adjustment control drums were fully inserted, and the two fine control drums began a stepwise reactivity insertion into the reactor. Within 6 hours, the reactor achieved on-orbit criticality, and the active control portion of the reactor test program began. For the next 154 hours, the control drums were operated with ground commands, to test reactor behavior. Due to the problems with the ion engine, the failure sensing and malfunction sensing systems were also switched off, because these could have been corrupted by the errant thruster. Following the first 200 hours of reactor operations, the reactor was set to autonomous operation at full power. Between 600 and 700 hours later, the voltage output of the reactor, as well as its temperature, began to drop; an effect that the S10-F3 test reactor had also demonstrated, due to hydrogen migration in the core.
On May 16, just over one month after being launched into orbit, contact was lost with the spacecraft for about 40 hours. Some time during this blackout, the reactor’s reflectors ejected from the core (although they remained attached to their actuator cables), shutting down the core. This spelled the end of reactor operations for the spacecraft, and when the emergency batteries died five days later all communication with the spacecraft was lost forever. Only 45 days had passed since the spacecraft’s launch, and information was received from the spacecraft for only 616 orbits.
What caused the failure? There are many possibilities, but when the telemetry from the spacecraft was read, it was obvious that something badly wrong had occurred. The only thing that can be said with complete confidence is that the error came from the Agena spacecraft rather than from the reactor. No indications had been received before the blackout that the reactor was about to scram itself (the reflector ejection was the emergency scram mechanism), and the problem wasn’t one that should have been able to occur without ground commands. However, with the telemetry data gained from the dwindling battery after the shutdown, some suppositions could be made. The most likely immediate cause of the reactor’s shutdown was traced to a possible spurious command from the high voltage command decoder, part of the Agena’s power conditioning and distribution system. This in turn was likely caused by one of two possible scenarios: either a piece of the voltage regulator failed, or it became overstressed because of either the unusual low-power vehicle loads or commanding the reactor to increase power output. Sadly, the cause of this system failure cascade was never directly determined, but all of the data received pointed to a high-voltage failure of some sort, rather than a low-voltage error (which could have also resulted in a reactor scram). Other possible causes of instrumentation or reactor failure, such as thermal or radiation environment, collision with another object, onboard explosion of the chemical propellants used on the Agena’s main engines, and previously noted flight anomalies – including the arcing and EM interference from the ion engine – were all eliminated as the cause of the error as well.
Despite the spacecraft’s mysterious early demise, SNAPSHOT provided many valuable lessons in space reactor design, qualification, ground handling, launch challenges, and many other aspects of handling an astronuclear power source for potential future missions: Suggestions for improved instrumentation design and performance characteristics; provision for a sunshade for the main radiator to eliminate the sun/shade efficiency difference that was observed during the mission; the use of a SNAP-2 type radiation shield to allow for off-the-shelf, non-radiation-hardened electronic components in order to save both money and weight on the spacecraft itself; and other minor changes were all suggested after the conclusion of the mission. Finally, the safety program developed for SNAPSHOT, including the SCA4 submersion criticality tests, the RFT-1 test, and the good agreement in reactor behavior between the on-orbit and ground test versions of the SNAP-10A showed that both the AEC and the customer of the SNAP-10A (be it the US Air Force or NASA) could have confidence that the program was ready to be used for whatever mission it was needed for.
Sadly, at the time of SNAPSHOT there simply wasn’t a mission that needed this system. 500 We isn’t much power, even though it was more power than was needed for many systems that were being used at the time. While improvements in the thermoelectric generators continued to come in (and would do so all the way to the present day, where thermoelectric systems are used for everything from RTGs on space missions to waste heat recapture in industrial facilities), the simple truth of the matter was that there was no mission that needed the SNAP-10A, so the program was largely shelved. Some follow-on paper studies would be conducted, but the lowest powered of the US astronuclear designs, and the first reactor to operate in Earth orbit, would be retired almost immediately after the SNAPSHOT mission.
Post-SNAPSHOT SNAP: the SNAP Improvement Program
The SNAP fission-powered program didn’t end with SNAPSHOT, far from it. While the SNAP reactors only ever flew once, their design was mature, well-tested, and in most particulars ready to fly in a short time – and the problems associated with those particulars had been well-addressed on the nuclear side of things. The Rankine power conversion system for the SNAP-2, which went through five iterations, reached technological maturity as well, having operated in a non-nuclear environment for close to 5,000 hours and remained in excellent condition, meaning that the 10,000 hour requirement for the PCS would be able to be met without any significant challenges. The thermoelectric power conversion system also continued to be developed, focusing on an advanced silicon-germanium thermoelectric convertor, which was highly sensitive to fabrication and manufacturing processes – however, we’ll look more at thermoelectrics in the power conversion systems series of blog posts, just keep in mind that the power conversion systems continued to improve throughout this time, not just the reactor core design.
On the reactor side of things, the biggest challenge was definitely hydrogen migration within the fuel elements. As the hydrogen migrates away from the ZrH fuel, many problems occur; from unpredictable reactivity within the fuel elements, to temperature changes (dehydrogenated fuel elements developed hotspots – which in turn drove more hydrogen out of the fuel element), to changes in ductility of the fuel, causing major headaches for end-of-life behavior of the reactors and severely limiting the fuel element temperature that could be achieved. However, the necessary testing for the improvement of those systems could easily be conducted with less-expensive reactor tests, including the SCA4 test-bed, and didn’t require flight architecture testing to continue to be improved.
The maturity of these two reactors led to a short-lived program in the 1960s to improve them, the SNAP Reactor Improvement Program. The SNAP-2 and -10 reactors went through many different design changes, some large and some small – and some leading to new reactor designs based on the shared reactor core architecture.
By this time, the SNAP-2 had mostly faded into obscurity. However, the fact that it shared a reactor core with the SNAP-10A, and that the power conversion system was continuing to improve, warranted some small studies to improve its capabilities. The two of note that are independent of the core (all of the design changes for the -10 that will be discussed can be applied to the -2 core as well, since at this point they were identical) are the change from a single mercury boiler to three, to allow more power throughput and to reduce loads on one of the more challenging components, and combining multiple cores into a single power unit. These were proposed together for a space station design (which we’ll look at later) to allow an 11 kWe power supply for a crewed station.
The vast majority of this work was done on the -10A. Any further reactors of this type would have had an additional three sets of 1/8” beryllium shims on the external reflector, increasing the initial reactivity by about 50 cents (1 dollar of reactivity is exactly break even, all other things being equal; reactivity potential is often somewhere around $2-$3, however, to account for fission product buildup); this means that additional burnable poisons (elements which absorb neutrons, then decay into something that is mostly neutron transparent, to even out the reactivity of the reactor over its lifetime) could be inserted in the core at construction, mitigating the problems of reactivity loss that were experienced during earlier operation of the reactor. With this, and a number of other minor tweaks to reflector geometry and lowering the core outlet temperature slightly, the life of the SNAP-10A was able to be extended from the initial design goal of one year to five years of operation. The end-of-life power level of the improved -10A was 39.5 kWt, with an outlet temperature of 980 F (527°C) and a power density of 0.14 kWt/lb (0.31 kWt/kg).
e design modifications led to another iteration of the SNAP-10A, the Interim SNAP-10A/2 (I-10A/2). This reactor’s core was identical, but the reflector was further enhanced, and the outlet temperature and reactor power were both increased. In addition, even more burnable poisons were added to the core to account for the higher power output of the reactor. Perhaps the biggest design change with the Interim -10A/2 was the method of reactor control: rather than the passive control of the reactor, as was done on the -10A, the entire period of operation for the I-10A/2 was actively controlled, using the control drums to manage reactivity and power output of the reactor. As with the improved -10A design, this reactor would be able to have an operational lifetime of five years. These improvements led the I-10A/2 to have an end of life power rating of 100 kWt, an outlet temperature of 1200 F (648°C), and an improved power density of 0.33 kWt/lb (0.73 kWt/kg).
This design, in turn, led to the Upgraded SNAP-10A/2 (U-10A/2). The biggest in-core difference between the I-10A/2 and the U-10A/2 was the hydrogen barrier used in the fuel elements: rather than using the initial design that was common to the -2, -10A, and I-10A/2, this reactor used the hydrogen barrier from the SNAP-8 reactor, which we’ll look at in the next blog post. This is significant, because the degradation of the hydrogen barrier over time, and the resulting loss of hydrogen from the fuel elements, was the major lifetime limiting factor of the SNAP-10 variants up until this point. This reactor also went back to static control, rather than the active control used in the I-10A/2. As with the other -10A variants, the U-10A/2 had a possible core lifetime of five years, and other than an improvement of 100 F in outlet temperature (to 1300 F), and a marginal drop in power density to 0.31 kWt/lb, it shared many of the characteristics that the I-10A/2 had. SNAP-10B: The Upgrade that Could Have Been
One consistent mass penalty in the SNAP-10A variants that we’ve looked at so far is the control drums: relatively large reactivity insertions were possible with a minimum of movement due to the wide profile of the control drums, but this also meant that they extended well away from the reflector, especially early in the mission. This meant that, in order to prevent neutron backscatter from hitting the rest of the spacecraft, the shield had to be relatively wide compared the the size of the core – and the shield was not exactly a lightweight system component.
The SNAP-10B reactor was designed to address this problem. It used a similar core to the U-10A/2, with the upgraded hydrogen barrier from the -8, but the reflector was tapered to better fit the profile of the shadow shield, and axially sliding control cylinders would be moved in and out to provide control instead of the rotating drums of the -10A variants. A number of minor reactor changes were needed, and some of the reactor physics parameters changed due to this new control system; but, overall, very few modifications were needed.
The first -10B reactor, the -10B Basic (B-10B), was a very simple and direct evolution of the U-10A/2, with nothing but the reflector and control structures changed to the -10B configuration. Other than a slight drop in power density (to 0.30 kWt/lb), the rest of the performance characteristics of the B-10B were identical to the U-10A/2. This design would have been a simple evolution of the -10A/2, with a slimmer profile to help with payload integration challenges.
The next iteration of the SNAP-10B, the Advanced -10B (A-10B), had options for significant changes to the reactor core and the fuel elements themselves. One thing to keep in mind about these reactors is that they were being designed above and beyond any specific mission needs; and, on top of that, a production schedule hadn’t been laid out for them. This means that many of the design characteristics of these reactors were never “frozen,” which is the point in the design process when the production team of engineers need to have a basic configuration that won’t change in order to proceed with the program, although obviously many minor changes (and possibly some major ones) would continue to be made up until the system was flight qualified.
Up until now, every SNAP-10 design used a 37 fuel element core, with the only difference in the design occurring in the Upgraded -10A/2 and Basic -10B reactors (which changed the hydrogen barrier ceramic enamel inside the fuel element clad). However, with the A-10B there were three core size options: the first kept the 37 fuel element core, a medium-sized 55-element core, and a large 85-element core. There were other questions about the final design, as well, looking at two other major core changes (as well as a lot of open minor questions). The first option was to add a “getter,” a sheath of hydrophilic (highly hydrogen-absorbing) metal to the clad outside the steel casing, but still within the active region of the core. While this isn’t as ideal as containing the hydrogen within the U-ZrH itself, the neutron moderation provided by the hydrogen would be lost at a far lower rate. The second option was to change the core geometry itself, as the temperature of the core changed, with devices called “Thermal Coefficient Augmenters” (TCA). There were two options that were suggested: first, there was a bellows system that was driven by NaK core temperature (using ruthenium vapor), which moves a portion of the radial reflector to change the core’s reactivity coefficient; second, the securing grids for the fuel elements themselves would expand as the NaK increased in temperature, and contract as the coolant dropped in temperature.
Between the options available, with core size, fuel element design, and variable core and fuel element configuration all up in the air, the Advanced SNAP-10B was a wide range of reactors, rather than just one. Many of the characteristics of the reactors remained identical, including the fissile fuel itself, the overall core size, maximum outlet temperature, and others. However, the number of fuel elements in the core alone resulted in a wide range of different power outputs; and, which core modification the designers ultimately decided upon (Getter vs TCA, I haven’t seen any indication that the two were combined) would change what the capabilities of the reactor core would actually be. However, both for simplicity’s sake, and due to the very limited documentation available on the SNAP-10B program, other than a general comparison table from the SNAP Systems Capability Study from 1966, we’ll focus on the 85 fuel element core of the two options: the Getter core and the TCA core.
A final note, which isn’t clear from these tables: each of these reactor cores was nominally optimized to a 100 kWt power output, the additional fuel elements reduced the power density required at any time from the core in order to maximize fuel lifetime. Even with the improved hydrogen barriers, and the variable core geometry, while these systems CAN offer higher power, it comes at the cost of a shorter – but still minimum one year – life on the reactor system. Because of this, all reported estimates assumed a 100 kWt power level unless otherwise stated.
The idea of a hydrogen “getter” was not a new one at the time that it was proposed, but it was one that hadn’t been investigated thoroughly at that point (and is a very niche requirement in terrestrial nuclear engineering). The basic concept is to get the second-best option when it comes to hydrogen migration: if you can’t keep the hydrogen in your fuel element itself, then the next best option is keeping it in the active region of the core (where fission is occurring, and neutron moderation is the most directly useful for power production). While this isn’t as good as increasing the chance of neutron capture within the fuel element itself, it’s still far better than hydrogen either dissolving into your coolant, or worse yet, migrating outside your reactor and into space, where it’s completely useless in terms of reactor dynamics. Of course, there’s a trade-off: because of the interplay between the various aspects of reactor physics and design, it wasn’t practical to change the external geometry of the fuel elements themselves – which means that the only way to add a hydrogen “getter” was to displace the fissile fuel itself. There’s definitely an optimization question to be considered; after all, the overall reactivity of the reactor will have to be reduced because the fuel is worth more in terms of reactivity than the hydrogen that would be lost, but the hydrogen containment in the core at end of life means that the system itself would be more predictable and reliable. Especially for a static control system like the A-10B, this increase in behavioral predictability can be worth far more than the reactivity that the additional fuel would offer. Of the materials options that were tested for the “getter” system, yttrium metal was found to be the most effective at the reactor temperatures and radiation flux that would be present in the A-10B core. However, while improvements had been made in the fuel element design to the point that the “getter” program continued until the cancellation of the SNAP-2/10 core experiments, there were many uncertainties left as to whether the concept was worth employing in a flight system. The second option was to vary the core geometry with temperature, the Thermal Coefficient Augmentation (TCA) variant of the A-10B. This would change the reactivity of the reactor mechanically, but not require active commands from any systems outside the core itself. There were two options investigated: a bellows arrangement, and a design for an expanding grid holding the fuel elements themselves.
The first variant used a bellows to move a portion of the reflector out as the temperature increased. This was done using a ruthenium reservoir within the core itself. As the NaK increased in temperature, the ruthenium would boil, pushing a bellows which would move some of the beryllium shims away from the reactor vessel, reducing the overall worth of the radial reflector. While this sounds simple in theory, gas diffusion from a number of different sources (from fission products migrating through the clad to offgassing of various components) meant that the gas in the bellows would not just be ruthenium vapor. While this could have been accounted for, a lot of study would have needed to have been done with a flight-type system to properly model the behavior. The second option would change the distance between the fuel elements themselves, using base plate with concentric accordion folds for each ring of fuel elements called the “convoluted baseplate.” As the NaK heated beyond optimized design temperature, the base plates would expand radially, separating the fuel elements and reducing the reactivity in the core. This involved a different set of materials tradeoffs, with just getting the device constructed causing major headaches. The design used both 316 stainless steel and Hastelloy C in its construction, and was cold annealed. The alternative, hot annealing, resulted in random cracks, and while explosive manufacture was explored it wasn’t practical to map the shockwave propagation through such a complex structure to ensure reliable construction at the time. While this is certainly a concept that has caused me to think a lot about the concept of a variable reactor geometry of this nature, there are many problems with this approach(which could have possibly been solved, or proven insurmountable). Major lifetime concerns would include ductility and elasticity changes through the wide range of temperatures that the baseplate would be exposed to; work hardening of the metal, thermal stresses, and neutron bombardment considerations of the base plates would also be a major concern in this concept.
These design options were briefly tested, but most of these ended up not being developed fully. Because the reactor’s design was never frozen, many engineering challenges remained in every option that had been presented. Also, while I know that a report was written on the SNAP-10B reactor’s design (R. J . Gimera, “SNAP 1OB Reactor Conceptual Design,” NAA-SR-10422), I can’t find it… yet. This makes writing about the design difficult, to say the least.
Because of this, and the extreme paucity of documentation on this later design, it’s time to turn to what these innovative designs could have offered when it comes to actual missions.
The Path Not Taken: Missions for SNAP-2, -10A
Every space system has to have a mission, or it will never fly. Both SNAP-2 and SNAP-10 offered a lot for the space program, both for crewed and uncrewed missions; and what they offered only grew with time. However, due to priorities at the time, and the fact that many records from these programs appear to never have been digitized, it’s difficult to point to specific mission proposals for these reactors in a lot of cases, and the missions have to be guessed at from scattered data, status reports, and other piecemeal sources.
SNAP-10 was always a lower-powered system, even with its growth to a kWe-class power supply. Because of this, it was always seen as a power supply for unmanned probes, mostly in low Earth orbit, but certainly it would also have been useful in interplanetary studies as well, which at this point were just appearing on the horizon as practical. Had the SNAPSHOT system worked as planned, the cesium thruster that had been on board the Agena spacecraft would have been an excellent propulsion source for an interplanetary mission. However, due to the long mission times and relatively fragile fuel of the original SNAP-10A, it is unlikely that these missions would have been initially successful, while the SNAP-10A/2 and SNAP-B systems, with their higher power output and lifetimes, would have been ideal for many interplanetary missions.
As we saw in the US-A program, one of the major advantages that a nuclear reactor offers over photovoltaic cells – which were just starting to be a practical technology at the time – is that they offer very little surface area, and therefore the atmospheric drag that all satellites experience due to the thin atmosphere in lower orbits is less of a concern. There are many cases where this lower altitude offers clear benefits, but the vast majority of them deal with image resolution: the lower you are, the more clear your imagery can be with the same sensors. For the Russians, the ability to get better imagery of US Navy movements in all weather conditions was of strategic importance, leading to the US-A program. For Americans, who had other means of surveillance (and an opponent’s far less capable blue-water navy to track), radar surveillance was not a major focus – although it should be noted that 500 We isn’t going to give you much, if any resolution, no matter what your altitude.
One area that SNAP-10A was considered for was for meteorological satellites. With a growing understanding of how weather could be monitored, and what types of data were available through orbital systems, the ability to take and transmit pictures from on-orbit using the first generations of digital cameras (which were just coming into existence, and not nearly good enough to interest the intelligence organizations at the time), along with transmitting the data back to Earth, would have allowed for the best weather tracking capability in the world at the time. By using a low orbit, these satellites would be able to make the most of the primitive equipment available at the time, and possibly (speculation on my part) have been able to gather rudimentary moisture content data as well.
However, while SNAP-10A was worked on for about a decade, for the entire program there was always the question of “what do you do with 500-1000 We?” Sure, it’s not an insignificant amount of power, even then, but… communications and propulsion, the two things that are the most immediately interesting for satellites with reliable power, both have a linear relationship between power level and capability: the more power, the more bandwidth, or delta-vee, you have available. Also, the -10A was only ever rated for one year of operations, although it was always suspected it could be limped along for longer, which precluded many other missions.
The later SNAP-10A/2 and -10B satellites, with their multi-kilowatt range and years-long lifespans, offered far more flexibility, but by this point many in the AEC, the US Air Force, NASA, and others were no longer very interested in the program; with newer, more capable, reactor designs being available (we’ll look at some of those in the next post). While the SNAP-10A was the only flight-qualified and -tested reactor design (and the errors on the mission were shown to not be the fault of the reactor, but the Agena spacecraft), it was destined to fade into obscurity.
SNAP-10A was always the smallest of the reactors, and also the least powerful. What about the SNAP-2, the 3-6 kWe reactor system?
Initial planning for the SNAP-2 offered many options, with communications satellites being mentioned as an option early on – especially if the reactor lifetime could be extended. While not designed specifically for electric propulsion, it could have utilized that capability either on orbit around the Earth or for interplanetary missions. Other options were also proposed, but one was seized on early: a space station.
At the time, most space station designs were nuclear powered, and there were many different configurations. However, there were two that were the most common: first was the simple cylinder, launched as a single piece (although there were multiple module designs proposed which kept the basic cylinder shape) which would be finally realized with the Skylab mission; second was a torus-shaped space station, which was proposed almost a half a century before by Tsiolkovsky, and popularized at the time by Werner von Braun. SNAP-2 was adapted to both of these types of stations. Sadly, while I can find one paper on the use of the SNAP-2 on a station, it focuses exclusively on the reactor system, and doesn’t use a particular space station design, instead laying out the ground limits of the use of the reactor on each type of station, and especially the shielding requirements for each station’s geometry. It was also noted that the reactors could be clustered, providing up to 11 kWe of power for a space station, without significant change to the radiation shield geometry. We’ll look at radiation shielding in a couple posts, and look at the particulars of these designs there.
Since space stations were something that NASA didn’t have the budget for at the time, most designs remained vaguely defined, without much funding or impetus within the structure of either NASA or the US Air Force (although SNAP-2 would have definitely been an option for the Manned Orbiting Laboratory program of the USAF). By the time NASA was seriously looking at space stations as a major funding focus, the SNAP-8 derived Advanced Zirconium Hydride reactor, and later the SNAP-50 (which we’ll look at in the next post) offered more capability than the more powerful SNAP-2. Once again, the lack of a mission spelled the doom of the SNAP-2 reactor.
The SNAP-2 reactor met its piecemeal fate even earlier than the SNAP-10A, but oddly enough both the reactor and the power conversion system lasted just as long as the SNAP-10A did. The reactor core for the SNAP-2 became the SNAP-10A/2 core, and the CRU power conversion system continued under development until after the reactor cores had been canceled. However, mention of the SNAP-2 as a system disappears in the literature around 1966, while the -2/10A core and CRU power conversion system continued until the late 1960s and late 1970s, respectively.
The Legacy of The Early SNAP Reactors
The SNAP program was canceled in 1971 (with one ongoing exception), after flying a single reactor which was operational for 43 days, and conducting over five years of fission powered testing on the ground. The death of the program was slow and drawn out, with the US Air Force canceling the program requirement for the SNAP-10A in 1963 (before the SNAPSHOT mission even launched), the SNAP-2 reactor development being canceled in 1967, all SNAP reactors (including the SNAP-8, which we’ll look at next week) being canceled by 1974, and the CRU power conversion system being continued until 1979 as a separate internal, NASA-supported but not fully funded, project by Rockwell International.
The promise of SNAP was not enough to save the program from the massive cuts to space programs, both for NASA and the US Air Force, that fell even as humanity stepped onto the Moon for the first time. This is an all-too-common fate, both in advanced nuclear reactor engineering and design as well as aerospace engineering. As one of the engineers who worked on the Molten Salt Reactor Experiment noted in a recent documentary on that technology, “everything I ever worked on got canceled.”
However, this does not mean that the SNAP-2/10A programs were useless, or that nothing except a permanently shut down reactor in orbit was achieved. In fact, the SNAP program has left a lasting mark on the astronuclear engineering world, and one that is still felt today. The design of the SNAP-2/10A core, and the challenges that were faced with both this reactor core and the SNAP-8 core informed hydride fuel element development, including the thermal limits of this fuel form, hydrogen migration mitigation strategies, and materials and modeling for multiple burnable poison options for many different fuel types. The thermoelectric conversion system (germanium-silicon) became a common one for high-temperature thermoelectric power conversion, both for power conversion and for thermal testing equipment. Many other materials and systems that were used in this reactor system continued to be developed through other programs.
Possibly the most broad and enduring legacy of this program is in the realm of launch safety, flight safety, and operational paradigms for crewed astronuclear power systems. The foundation of the launch and operational safety guidelines that are used today, for both fission power systems and radioisotope thermoelectric generators, were laid out, refined, or strongly informed by the SNAPSHOT and Space Reactor Safety program – a subject for a future web page, or possibly a blog post. From the ground handling of a nuclear reactor being integrated to a spacecraft, to launch safety and abort behavior, to characterizing nuclear reactor behavior if it falls into the ocean, to operating crewed space stations with on-board nuclear power plants, the SNAP-2/10A program literally wrote the book on how to operate a nuclear power supply for a spacecraft.
While the reactors themselves never flew again, nor did their direct descendants in design, the SNAP reactors formed the foundation for astronuclear engineering of fission power plants for decades. When we start to launch nuclear power systems in the future, these studies, and the carefully studied lessons of the program, will continue to offer lessons for future mission planners.
More Coming Soon!
The SNAP program extended well beyond the SNAP-2/10A program. The SNAP-8 reactor, started in 1959, was the first astronuclear design specifically developed for a nuclear electric propulsion spacecraft. It evolved into several different reactors, notably the Advanced ZrH reactor, which remained the preferred power option for NASA’s nascent modular space station through the mid-to-late 1970s, due to its ability to be effectively shielded from all angles. Its eventual replacement, the SNAP-50 reactor, offered megawatts of power using technology from the Aircraft Nuclear Propulsion program. Many other designs were proposed in this time period, including the SP-100 reactor, the ancestor of Kilopower (the SABRE heat pipe cooled reactor concept), as well as the first American in-core thermionic power system, advances in fuel element designs, and many other innovations.
Originally, these concepts were included in this blog post, but this post quickly expanded to the point that there simply wasn’t room for them. While some of the upcoming post has already been written, and a lot of the research has been done, this next post is going to be a long one as well. Because of this, I don’t know exactly when the post will end up being completed.
After we look at the reactor programs from the 1950s to the late 1980s, we’ll look at NASA and Rosatom’s collaboration on the TOPAZ-II reactor program, and the more recent history of astronuclear designs, from SDI through the Fission Surface Power program. We’ll finish up the series by looking at the most recent power systems from around the world, from JIMO to Kilopower to the new Russian on-orbit nuclear electric tug.
After this, we’ll look at shielding for astronuclear power plants, and possibly ground handling, launch safety, and launch abort considerations, then move on to power conversion systems, which will be a long series of posts due to the sheer number of options available.
These next posts are more research-intensive than usual, even for this blog, so while I’ll be hard at work on the next posts, it may be a bit more time than usual before these posts come out.