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Development and Testing Fission Power Systems History Spacecraft Concepts

US Astro-nuclear History part 2: SNAP-8, NASA’s Space Station Power Supply

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.

S8ER Slightly Disassembled
SNAP-8 with reflector halves slightly separated, image DOE

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.

S8ER Assembly 1962
SNAP-8 Experimental Reactor being assembled, 1962

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

S8ER Cutaway
Image DOE

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.

S8ER Cutaway Drawing
S8ER Fuel Element

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.

S8ER Cross SectionThe 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.

S8ER Facility Setup
SNAP-8 Experimental Reactor test facility, Santa Susana Field Site, image DOE

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.

S8ER Core Containment StructureThe 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).

S8ER Drive Mechanism
S8ER Control drum actuator and scram mechanism, image DOE

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:

  1. 5016.5 continuous operations immediately preceeding the shutdown (most at 450 kWt, all at 1200 F or greater)
  2. 12,080 hours of total operations
  3. A total of 5,154,332 kilowatt-hours of thermal energy produced
  4. 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.

S8ER FE Damage Map
Core cross section with location and type of damage to fuel elements, image DOE

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.

S8ER FE Damage Photo

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.

S8ER Post irradiation fuel characteristics

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

S8DR with installed reflector assembly
SNAP 8 Development Reactor post-reflector installation, before being lowered into containment structure. Image DOE

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.

S8DR Cutaway Drawing in test vault
S8DR in Test Vault, image DOE

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.

Control RoomAfter 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).

S8DR FE Damage Map
S8DR Fuel Element Damage map, image DOE

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.

S8DR H loss rate table
Image DOE

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

Artists Depiction
5 kWe Thermoelectric Reactor, artist’s concept cutaway drawing. Image DOE

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.

5 kWe SchematicThe 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.

Blueprint Layout of System
5kWe Thermoelectric power system. Image DOE

Summary TableSummary Table 2

The U-ZrH Reactor: Power for America’s Latest and Greatest Space Stations.

ZrHR Cutaway Drawing
Cutaway drawing, Image DOE

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.

Reference Design Xsec and ElevtionAnother 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.

TE Layout diagramAnother 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.

Radiator Design BaselineAnother 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

OWS with two CSMs, 1966.PNG
Frontispiece for nuclear-powered Saturn V Orbital Workstation (which flew as Skylab), image NASA 

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.

TE Detail FlowBecause 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.

Radiation Shield ZonesOne 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).

Deorbit System ConceptAfter 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

Lunar configuration cutaway
Lunar landing configuration, image DOE

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).

Surface configuration diagram

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!

References

SNAP 8 Summary Report, AEC/Atomics International Staff, 1973 https://www.osti.gov/servlets/purl/4393793

SNAP-8, the First Electric Propulsion Power System, Wood et al 1961 https://www.osti.gov/servlets/purl/4837472

SNAP 8 Reactor Preliminary Design Summary, ed. Rosenberg, 1961 https://www.osti.gov/servlets/purl/4476265

SNAP 8 Reactor and Shield, Johnson and Goetz 1963 https://www.osti.gov/servlets/purl/4875647

SNAP 8 Reactors for Unmanned and Manned Applications, Mason 1965 https://www.osti.gov/servlets/purl/4476766

SNAP 8 Reactor Critical Experiment, ed. Crouter. 1964 https://www.osti.gov/servlets/purl/4471079

Disassembly and Postoperation Examination of the SNAP 8 Experimental Reactor, Dyer 1967 https://www.osti.gov/servlets/purl/4275472

SNAP 8 Experimental Reactor Fuel Element Behavior, Pearlman et al 1966 https://www.osti.gov/servlets/purl/4196260

Summary of SNAP 8 Development Reactor Operations, Felten and May 1973 https://www.osti.gov/servlets/purl/4456300

Structural Analysis of the SNAP 8 Development Reactor Fuel Cladding, Dalcher 1969 https://www.osti.gov/servlets/purl/6315913

Reference Zirconium Hydride Thermoelectric System, AI Staff 1969 https://www.osti.gov/servlets/purl/4004948

Reactor-Thermoelectric System for NASA Space Station, Gylfe and Johnson 1969 https://www.osti.gov/servlets/purl/4773689

Categories
Development and Testing Fission Power Systems History Nuclear Electric Propulsion Test Stands

History of US Astronuclear Reactors part 1: SNAP-2 and 10A

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.

SNAP capabilities 1964
SNAP Reactor Capabilities and Status as of 1973, image DOE

The Beginnings of the US Astronuclear Program: SNAP’s Early Years

Core Cutaway Artist
Early SNAP-2 Concept Art, image courtesy DOE

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.

s2 Toroidal Station
SNAP-2 powered space station concept image via DOE

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

S2 Artist Cutaway Core
SNAP-2 Reactor Cutaway, image DOE

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.

S2 Critical Assembly
SNAP Critical Assembly core, image DOE

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.

S2ER Cross Section
SNAP-2 Experimental Reactor core cros section diagram, image DOE

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.

S2ER Schedule and Timeline

S2DR Core Xsec
SNAP 2 Development Reactor core cross section, image DOE

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.

S2DR Development Timeline

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.

S2 PCS Cutaway Drawing
CRU mercury Rankine power conversion system cutaway diagram, image DOE

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.

s10 FSM Reactor
SNAP-2/10A FSM reflector and drum mechanism pre-test, image DOE

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.

S2 Program History Table

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.

Reactor and Shield Cutaway
SNAP-10A Reactor sketch, image DOE

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.

SNAP Meteorological Satellite
SNAP-10A satellite concept, image DOE

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.

Reflector Ejection Test
Reflector ejection test in progress, image DOE

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.

RFD1
RFD-1 Experimental Payload

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.

RFD1 Flight Path
RFD-1 Mission Profile, image DOE

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.

Scout rocket
Scout sounding rocket, image DOE

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.

RFD-1 Burnup Splice
Film splice of RV burnup during RFD-1, image DOE

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.

S10FSM Vac Chamber
FSEM-3 in vacuum chamber for environmental and vibration tests, image DOE

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.

SNAP_10A_Space_Nuclear_Power_Plant
SNAP-10A F-3 (flight unit for SNAPSHOT) undergoing final checks before spacecraft integration. S10F-4 was identical.

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.

On Orbit Artist
SNAP-10A reactor, artist’s rendering (artist unknown), image DOE

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.

atlas-slv3_agena-d__snapshot__1
SNAPSHOT launch, image USAF via Gunter’s Space Page

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.

LiH FE Cracking
Fuel element post-irradiation. Notice the cracks where the FE would rest in the endcap reflector

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).

Thes

Interim 10A 2
Interim SNAP 10A/2, image DOE

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).

10A2 Table
Interim SNAP 10A/2 Design References, image DOE

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

10B Cutaway System
SNAP=10B Cutaway Diagram, image DOE

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.

10B Basic Table
Image DOE

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.

Yt Getter FE
Fuel element with yttrium getter, image DOE

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.

A-10B Bellows Diagram
Ruthenium vapor bellows design for TCA, image DOE

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.

A-10B Expandable BaseplateThe 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.

A-10B Expandable Baseplate 2While 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.

SNAP Meteorological Satellite
SNAP-10 powered meteorological satellite, image DOE

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.

S2 Cylinder Station
Cylindrical space station, image DOE

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.

s2 Toroidal Station
Hexagonal/Toroid space station. Note the wide radiation shield. Image DOE

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.

Hg Rankine Cutaway Drawing
Power conversion system, SNAP-2

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.

References

SNAP

SNAP Reactor Overview, Voss 1984 http://www.dtic.mil/dtic/tr/fulltext/u2/a146831.pdf

SNAP-2

Preliminary Results of the SNAP-2 Experimental Reactor, Hulin et al 1961 https://www.osti.gov/servlets/purl/4048774

Application of the SNAP 2 to Manned Orbiting Stations, Rosenberg et al 1962 https://www.osti.gov/servlets/purl/4706177

The ORNL-SNAP Shielding Program, Mynatt et al 1971 https://www.osti.gov/servlets/purl/4045094

SNAP-10/10A

SNAP-10A Nuclear Analysis, Dayes et al 1965 https://www.osti.gov/servlets/purl/4471077

SNAP 10 FS-3 Reactor Performance Hawley et al 1966 https://www.osti.gov/servlets/purl/7315563

SNAPSHOT and the Flight Safety Program

SNAP-10A SNAPSHOTProgram Development, Atomics International 1962 https://www.osti.gov/servlets/purl/4194781

Reliability Improvement Program Planning Report for the SNAP-10A Reactor, Coombs et al 1961 https://www.osti.gov/servlets/purl/966760

Aerospace Safety Reentry Analytical and Experimental Program SNAP 2 and 10A Interim Report, Elliot 1963 https://www.osti.gov/servlets/purl/4657830

SNAPSHOT orbit, Heavens Above https://www.heavens-above.com/orbit.aspx?satid=1314

SNAP Improvement Program

Static Control of SNAP Reactors, Birney et al 1966 https://digital.library.unt.edu/ark:/67531/metadc1029222/m2/1/high_res_d/4468078.pdf

SNAP Systems Capabilities Vol 2, Study Introduction, Reactors, Shielding, Atomics International 1965 https://www.osti.gov/servlets/purl/4480419

Progress Report, SNAP Reactor Improvement Program, April-June 1965 https://www.osti.gov/servlets/purl/4467051

Progress Report for SNAP General Supporting Technology May-July 1964 https://www.osti.gov/servlets/purl/4480424/

Categories
Electric propulsion Fission Power Systems History Nuclear Electric Propulsion

Nuclear Electric Propulsion History Part 1: The Soviet Astronuclear Program

Hello, and welcome back to Beyond NERVA, where we’re getting back into issues directly related to nuclear power in space, rather than how that power is used (as we’ve examined in our last three blog posts on electric propulsion)! However, the new Electric Propulsion page is up on the website, including a summary of all the information that we’ve covered in the last three blog posts, which you can find here [Insert Link]! Also, each type of thruster has its own page as well for easier reference, which are all linked on that summary page! Make sure to check it out!

In this blog series, we’re going to look at nuclear electric power system reactor cores themselves. While we’ve looked at a number of designs for nuclear thermal reactor cores (insert link for NTR-S page), there are a number of differences in those reactor cores compared to ones that are designed purely for electricity production. Perhaps the biggest one is operating temperature, and therefor core lifetime, but because the coolant doesn’t have to be hydrogen, and because the amount of heat produced doesn’t have to be increased as much as possible (there will be a LOT more discussion on this concept in the next series on power conversion systems), the reactor can be run at cooler temperatures, preventing a large amount of thermally related headaches, which makes far more more materials available for the reactor core, and generally simplifying matters.

Nuclear electric power systems are also unique in that they’re the only type of fission powered electrical supply system that’s ever flown. We’ve mentioned those systems briefly before, but we’ll look at some of them more in depth today, and in the next post as well. While there have been many reactor designs proposed over the years, we’re going to focus on the programs developed by the USSR during the Cold War, since they have the longest operational history of any sort of fission-powered heat source in space.

The United States were the first to fly a nuclear reactor in space, the SNAPSHOT mission in 1963; but, sadly, another American reactor was never placed on a spacecraft. The Soviet program was far longer running, flying reactors almost continuously from 1970-1988, often two spacecraft at once. With the fall of the Berlin Wall, and the end of the Cold War, the Soviet astronuclear program ended, and Russia hasn’t flown another nuclear reactor since then. There was a time, though, in the 1990s, that a mission was on the books (but never funded to a sufficient level to have ever flown) to use US-purchased, Russian-built nuclear reactors for an American crewed moon base!

 

History of Soviet In-Space Fission Power Systems

From the beginning, the Soviet in-space nuclear power designers focused on two different design concepts for their power systems: single-cell thermal fuel elements and multi-cell thermal fuel elements. The biggest difference between the two is how many fuel elements are in each thermal fuel element system: the single cell design uses a single fuel element, while the multi-cell option uses multiple fuel elements, separated by passive spacers, moderation blocks, or thermionic power conversion systems. Both designs were extensively researched, and eventually flown, but the initial research focused on the multi-cell approach with the Romashka, (sometimes translated as “Chamomile,” other times as “Daisy”) reactor. This design type led to the BES-5 (Bouk, or “Beech,” flight reactor), whereas the single cell variation led to the TEU-5 TOPOL (TOPAZ-1), which flew twice, and ENISY (TOPAZ-2) reactor, and was later purchased by the US. We won’t be looking in-depth at the ENISY reactor in this post, despite its close relation to the TOPOL, because later in the blog series we’ll be focusing on it far more, the time the Americans bought two of them (and took out an option on another four) and how it could have powered an American lunar base in the 1990s, had the funding been available.

As is our wont here, let’s begin at the beginning, with one of Korolev’s pet projects and the first in-space reactor design of the USSR: Romashka.

 

Romashka: The Reactor that Started it All

 

Romashka 2
Romashka mockup

The Romashka (daisy or chamomile in English) was a Soviet adaptation of an American idea first developed in the US at Los Alamos Labs in the mid 1950s: in-core thermionic energy conversion. We’ll be looking at thermionics much more in-depth in the next post on power conversion systems, but the short version is that it combines a heat pipe (which we looked at in the Kilopower posts) with the tendency for an incandescent light to develop a static charge on its’ bulb. More on the conversion system itself in the next post; but, for now, it’s worth noting that this is a way to actually stick your power conversion system in the core of your reactor; and, as far as I’ve seen, it’s the only one.

Design on this reactor started in 1957, following a trip by Soviet scientists to Los Alamos (where thermionic energy conversion had been proposed, but not yet tested). The design offered the potential to have no moving parts, no pumps, and only needed conductive cooling from the reactor body for thermal management; all very attractive properties for a reactor that would not be able to be maintained for its lifetime. Work was begun at the Institute of Atomic Energy, I.V. Kurchatov in Moscow, but by the end of the program there were many design bureaus involved in the conceptual design, manufacturing, and testing of this reactor.

A series of disc-shaped uranium carbide (UC2) fuel elements were used in this reactor (90% 235U), with holes drilled through the center, and roughly halfway from the central hole of the disc to the edge of the fuel element. Both of these holes were used to thread the thermionic power conversion system through the core of the reactor. Spacing of the fuel elements was provided by a mixture of beryllium oxide and graphite, which was also used to slightly moderate the neutron spectrum – but the neutron spectrum in the reactor remained in the fast spectrum. Surrounding the reactor core itself, both radially and at the ends of the core, were beryllium reflectors. Boron and boron nitride control rods placed in the radial reflector and base axial reflector were used to maintain reactor control through the use of a hydraulic system, however a large negative thermal reactivity coefficient in the reactor core was also meant to largely control the reactor in the case of normal operations. Finally, the reactor was surrounded by a finned steel casing that provided all heat rejection through passive radiation – no pumps required! The nominal operating temperature of the reactor was meant to be between 1200 C and 1800 C at the center of the core, and about 800 C at the edges of the core at the ends of the cylinder.

Romashka 3
Core undergoing assembly, 1966

Construction and warm-critical tests were completed by April, 1966, and testing began in Moscow. There are some indications that materials incompatibilities in the first Romashka built led to the need to rebuild it with different materials, but it’s unclear what would have been changed (the only other reference, besides on a CIA document, to this is that the thermionic fuel element materials were changed in the reactor, so that may be what occurred – more on that in the direct power conversion post). This reactor underwent about 15,000 hours of testing, and in that time period it produced about 6,100 kWh of electricity at a relatively constant rate of 40 kW of thermal and 500-800 W of electrical power (1.5%-2% energy conversion efficiency). Initial testing (about 1200 hours) only rejected heat into a vacuum chamber using the fins’ radiative cooling capability; and testing of other reactor behavior particulars was carried out, including core self-regulation capability. Later tests (about 14,000 hours) were done using natural convection in a helium environment. During these tests, thermal deformation of the core and the reflector led to a reduction in reactivity, which was compensated for with the control system. By the end of the test cycle, electrical power production had dropped by 25%, and overall reactivity had dropped by 30%. Maximum sustained power production was about 450 W, and 88 amps, if all thermionic converters were activated, and pulsed power of up to 800 W was observed at the beginning of the actively controlled tests.

ustanovka-romashka_1
Reactor being installed in test containment vessel, 1966

Korolev planned to pair this reactor with a plasma pulsed power thruster (based on the time period, possibly a pulsed inductive thruster, or PIT, which we looked at briefly in the second blog post on electric propulsion systems). However, two things conspired to end the Romashka system: Korolev’s death in 1966 meant the loss of its’ most powerful proponent; and the development of the more powerful, more efficient Bouk reactor became advanced enough to make that design available for space travel in the same time frame.

While there were plans to adapt Romashka into a small power plant for remote outposts (the core was known as “Gamma”), the testing program ended in 1966, to be supplanted by the BES-5 “Beech”. The legacy of the Romashka reactor lives on, however, as the first successful design of a thermionic energy conversion system for in-core use, a test-bed for the development and testing of thermionic energy conversion materials (more on that in the first power conversion system post); and it remains the father and grandfather of all Russian in-space reactors to ever fly.

 

Bouk: The Most Flown Nuclear Reactor in History

Buk Cutaway
BES-5 Bouk cutaway diagram, image Rosatom

The Bouk (“Beech”) reactor, also known as the “Buk,” or BES-5 reactor, is arguably the most successful astronuclear design in history. Begun in 1960 by the Krasnya Zvesda Scientific and Propulsion Association, this reactor promised greater power output than the Romashka, at the cost of additional complexity, and requiring coolant to operate. From 1963 to 1969, testing of the fuel elements and reactor core was carried out without using the thermoelectric fuel elements (TFE), which were still under development. From 1968 to 1970, three reactor cores with full TFEs were tested at Baikal; and, with successful testing completed, the reactor design was prepared for launch, integrated into the Upravlenniye Sputnik Aktivny (US-A; in the West, RORSAT, for Radar Ocean Reconnaisance SATellite) spacecraft, designed to use radar for naval surveillance.

DOE Sketch based on KOSMOS 954
LLNL sketch of BES-5 based on KOSMOS-954 wreckage, DOE image via Sven Grahn

Rather than having stacked discs of UC2, the BES-5 used 79 fuel rods made out of uranium (90% enriched, total U mass 30 kg) molybdenum alloy metal, encased in high-temperature steel. NaK was used as a coolant for the reactor, pumped using the energy from 19 of the fuel assemblies to run an electromagnetic fuel pump. Producing over 100 kW of thermal energy, after electric conversion using in-core germanium-silicon thermoelectric power conversion elements (which use the difference in charge potential between two different metals along a boundary to create an electrical charge when a temperature gradient is applied across the join; again, more in a later post), a maximum of 5 kW of electrical energy was available for the spacecraft’s instrumentation. The fact that this core used thermoelectric conversion rather than thermionic is a good indicator that the common use of the term, TOPAZ, for this reactor is incorrect. Reactor control was provided by six beryllium reflector drums that would be slowly lowered through holes in the radial reflector over the reactor’s life to increase the local neutron flux to account for the buildup of neutron poisons.

BES-5 Ascent Stage
BES-5 ascent stage cutaway, with core on left and chemical propulsion system on right, Rosatom

One unique aspect to the BES-5 is that the reactor was able to decommission itself at end of life (although this wasn’t always successful) by moving the reactor to a higher orbit and then ejecting the end reflector and fuel assemblies (which were subcritical at time of assembly, and required the Be control rods to be inserted to reach delayed criticality), as well as dumping the NaK coolant overboard. This ensured that the reactor core would not re-enter the atmosphere (although there were two notable exceptions to this, and one late unexpected success). As an additional safety measure following the failure of KOSMOS-954 (more on that below), the reactor was redesigned so that the fuel elements would burn up upon re-entry, diluting the radioactive material to the point that no significant increase in radiation would occur. Over the reactor’s long operational history (31 BES-5 reactors were launched), the lifetime of the reactors was constantly extended, beginning with a lifetime of just 110 minutes (used for radar broadcast testing) to up to 135 days of operational life.

RORSAT_by_Ronald_C._Wittmann,_1982
US-A satellite (with the radiator and ascent stage, but oddly no core), painting by Ronald Wittman 1982

The first BES-5 to be launched was serial number 37 on the KOSMOS-367 satellite on October 3, 1970 (there’s some confusion on this score, with another source claiming it was KOSMOS-469, launched on 25 December 1971). After a very short (110 minute) operational life, the spacecraft was moved into a graveyard orbit and the reactor ejected due to overheating in the reactor core. Three more spacecraft (KOSMOS-402, -469, and 516) were launched over the next two years, with the -469 spacecraft possibly being the first to have the 8.2 GHz side looking radar system that the power plant was selected for. Over time, the US-A spacecraft were launched in parallel, co-planar orbits, usually deployed in pairs with closely attending Russian US-P electronics intelligence satellites (for more on the operational use of the US-A, check out Sven Grahn’s excellent blog on the operational history of the US-A).

Morning Light logo
CNSC/DOE Operation Morning Light logo

The US-A program wasn’t without its failures, sadly, and one led to one of the biggest radiological cleanup missions in the history of nuclear power. On September 18, 1977, a Tsyklon-2 rocket launched from Baikonur Cosmodrome in Khazakhstan carrying the KOSMOS-954 US-A spacecraft on an orbital inclination of 65 degrees. By December, the spacecraft’s orbital maneuvering had become erratic, and Soviet officials contacted US officials that they had lost control of the satellite before they were able to move the reactor core into its’ designated graveyard orbit. On January 24, 1968, the satellite re-entered over Canada, spreading debris over a 600 km long section of the country. Operation Morning Light, the resulting CNES and US DOE program, was able to clear all the debris over several months, in a program that involved hundreds of people from the CNES, DOE, the NEST teams that were then available, and US Military Airlift Command. No fatalities or radiation poisoning cases were reported as a result of KOSMOS-954’s unplanned re-entry, although the remote nature of the re-entry was probably as much of a help as a challenge in this regard. A second KOSMOS spacecraft, KOSMOS-1402, also had its fuel elements re-enter the atmosphere following a failure of the spacecraft to ascend into its graveyard orbit, this time over the North Atlantic. The core re-entered the atmosphere on 23 January 1983, breaking up over the North Atlantic, north of England. No fragments of this reactor were ever recovered, and no significant increase in radioactivity as a result of this unplanned re-entry were detected.

These two incidents caused significant delays in the US-A program, and subsequent redesigns in the reactor as well. However, launches of this system continued until March 14, 1988, with the KOSMOS-1932 mission, which was moved into a graveyard orbit on 20 May, 1988, after a mission time of 66 days. The fate of its’ immediate predecessor, KOSMOS-1900, showed that the additional safety mechanisms for the US-A spacecraft’s reactor were successful: despite an apparent loss of control of the spacecraft, an increasingly eccentric orbit, and the buildup of aerodynamic forces, the reactor core was able to be boosted to a stable graveyard orbit, with the maneuver being completed on 17 October 1988. The main body of the spacecraft re-entered over the Indian Ocean 16 days earlier.

One interesting note on the controversy surrounding these reactor cores’ re-entry into Earth’s atmosphere is that the US planned on doing the exact same thing with the SNAP-10A reactors. The design was supposed to orbit for long enough (on the order of hundreds of year) for the short-lived fission products to decay away, and then the entire reactor would self-disassemble through a combination of mechanical, explosive, and aerodynamic systems; and, as a result, burn up in the upper atmosphere. While the amount of radioactivity that would be added to the atmosphere would be negligible, these accidents showed that this disposal method would not be acceptable; further complicating the American astronuclear program, as well as the one in the USSR. The SNAPSHOT reactor is still in orbit, and is expected to remain there for 2800 years, but considering the fallout of these accidents, retrieval or boosting to a graveyard orbit may be a future mission necessity for this reactor.

The US-A spacecraft demonstrated in-space nuclear fission power, and serial fission power plant production, for over two decades. Despite two major failures resulting in re-entry of the reactor core, the US-A program managed successful operation of the BES-5 reactor for 29 missions, and minimal impact from the two failures. The rest of the BES-5 cores remain parked in graveyard orbits, where they will remain for many hundreds of years until the radioactivity has dropped to natural background radiation.

There is one long-lasting legacy of the BES-5 program on in-orbit space travel, however: the ejected NaK coolant. The coolant remains a cratering hazard for spacecraft in certain orbits, but is not thought to be an object multiplication hazard. It is doubtful that the same core ejection system would be used in a newly designed astronuclear reactor, but this legacy lives on as another example of humanity’s ignorance at the time of a Kessler Syndrome situation.

While this program was not 100% successful, whether from a mission success point of view or from the point of view of it having no ongoing impact from the operations that were carried out, over 25 years of operation of a series of BES-5 reactors remains to this day the most extensive and successful of any astronuclear fission powered design, and it meets or exceeds even the service histories of any RTG design that has been deployed by any country.

 

TOPOL: The Most Powerful Reactor Ever Flown

TOPOL Cutaway
TEU-5 cutaway diagram

The TEU-5 TOPOL (TOPAZ-1) program is the second type of Soviet reactor to fly; and, although it only flew twice, it can be argued to have been even more successful than the BES-5 reactor design. The TEU-5 was the return of the in-core thermionic power conversion system that was first utilized in Romashka; and, just as the Bouk was a step above the Romashka, the Topol was a step beyond that. Thermionic conversion remained more attractive than thermoelectric in terms of wider range of operating capabilities, increased temperature potential, and more forgiving materials requirements, but thermoelectric conversion was able to be readied for flight first. Because of this, and because of the inertia that any flight-tested and more-refined (from a programmatic and serial production sense) program has over one that has yet to fly, the BES-5 flew for over a decade before the TEU-5 would take to orbit.

Despite the different structure, and much higher power, of the TEU-5, the design was able to fulfill the same role of ocean radar reconnaissance; but, initially, it was meant to be a powerful on-orbit TV transmission station. The major advantage of the TEU-5 over the BES-5 is that, due to its higher power level, it wasn’t forced to be in a very low orbit, which increased atmospheric drag, caused the dry mass of the craft to be severely reduced in order to allow for more propellant to be on board, and created a lot of complexity in terms of reactor decommissioning and disposal. Following the KOSMOS-954 and -1402 accidents, the low-flying profile of the US-A satellite was no longer available for astronuclear reactors, and so the orbital altitude increased. TEU-5 offered the capability to get useful image resolution at this higher altitude due to its higher power, and improvements to the (never flown, but ground tested) radar systems.

TOPAZ Core configuration, Bennett
Disgram of multi-cell TFE concept, Bennett 1989

The TOPOL program was begun in the 1960s, under the Russian acronym for Thermionic Experimental Converter in the Active Zone, which translates directly into Topaz in English, but ground testing didn’t begin until 1970. This was a multi-cell thermionic fuel element design similar in basic concept to Romashka, however it was a far more complex design. Instead of a single stack of disc-shaped fuel elements, a “garland” of fuel elements were formed into a thermionic fuel element. The fissile fuel element was surrounded by a thimble of tungsten or molybdenum, which formed the cathode of the thermionic converter, while the anode of the converter was a thin niobium tube; as with most thermionic converters the gap between cathode and anode was filled with cesium vapor. The anode was cooled with pumped NaK, although some sources indicate that lithium was also considered as a coolant for higher-powered versions of the reactor.

BES-5 core cross section
TEU-5 core cross-section, DOE

The differences between the BES-5 and TEU-5 were far more than the power conversion system. Instead of being a fast reactor, the Topaz was designed for the thermal neutron spectrum, and as such used zirconium hydride for in-core moderation (also creating a thermal limitation for the materials in the core; however, hydrogen loss mitigation measures were taken throughout the development process). Rather than using the metal fuels that its predecessor had, or the carbides of the Romashka, the Topol used a far more familiar material to nuclear power plant operators: uranium oxide (UO2), enriched to 90% 235U. This, along with reactor core geometry changes, allowed the amount of uranium needed for the core to drop from 30 kg in the BES-5 to 11.5 kg. NaK remained the coolant, due to its low melting temperature, good thermal conductivity, and neutronic transparency. The cathode temperature in the TEU-5 was in the range of 1500-1800C, which resulted in an electrical power output of up to 10 kW.

Cesium reservoir and regulator
ENISY cesium reservoir, which is very similar to the TEU-5 system, image courtesy DOE

One of the most technically challenging parts of this reactor’s design was in the cesium management system. The metal would only be a gas inside the core, and electromagnetic pumps were used to move the liquid through a series of filters, heaters, and pipes. The purity of the cesium had a large impact on the efficiency of the thermionic elements, so a number of filters were installed, including for gaseous fission waste products, to be evacuated into space.

The first flight of the TEU-5 was on the KOSMOS-1818 satellite, launched on February 1st, 1987, onto a significantly different orbital trajectory than the rest of the US-A series of spacecraft, despite the fact that superficially it appeared to be quite similar. This was because it was the test-bed of a new type of US-A spacecraft, the US-AM, taking advantage of not only the more powerful nuclear reactor but also employing numerous other technologies. The USSR eventually announced that the spacecraft’s name was the Plasma-A, and was a technology demonstrator for a number of new systems. These included six SPT-70 Hall thrusters for maneuvering and reaction control, and a suite of electromagnetic and sun-finding sensors. Some sources indicate that part of the mission for the spacecraft was the development of a magnetospherically-based navigation system for the USSR. An additional advantage to the higher orbit of this spacecraft was that it eliminated the need for the ascent stage for the reactor core and fuel elements, saving spacecraft mass to complete its’ mission. It had an operational life of 187 days, before the reactor was placed in its graveyard orbit, and the remainder of the spacecraft was allowed to re-enter the atmosphere as its orbit decayed.

The second Plasma-A (KOSMOS-1867) launch was on July 10th, 1987. While the initial flight profile was remarkably similar to the original Plasma-A satellite, the later portions of the mission showed a much larger variation in orbital period, possibly indicating more extensive testing of the thrusters. It was operational for just over a year before it, too, was decommissioned.

Neither of the TEU-5 launches carried radar equipment aboard; but, considering the cancellation of the program also coincided with the fall of the Soviet Union, it’s possible that the increased power output of the TEU-5 would have allowed acceptable radar resolution from this higher orbit (the US-A spacecraft’s orbit was determined by the distance and power requirements of its radar system, and due to the higher aerodynamic drag also significantly limited the lifetime of each spacecraft).

After decommissioning, similar problems with NaK coolant from the reactor core were experienced with the TEU-5 reactors. There is one additional complication from the decommissioning of these larger reactor cores, however, which led to some confusion during the Solar Maximum Mission (SMM) to study solar behavior. Due to the higher operational altitude during the time that the reactor was being operated at full power, and the behavior of the materials that the reactor was made out of, what is often a minor curiosity in reactor physics caused some confusion among some astrophysical and heliophysical researchers: when some materials are bombarded by sufficiently high gamma flux, they will eject electron-positron pairs, which were then trapped in the magnetosphere of the Earth. While these radiation fluxes are minuscule, and unable to adversely affect living tissue, for scientists carefully studying solar behavior during the solar maximum the difference in the number of positrons was not only noticeable, but statistically significant. Both the SMM satellite and one other (Ginga, a Japanese X-ray telescope launched in 1987, which reentered in 1991) have been confirmed to have some instrument interference due to either the gamma wave flux or the resulting positron emissions from the two flown TEU-5 reactors. While this is a problem that only affected a very small number of missions, should astronuclear reactors become more commonly used in orbit, these types of emissions will need to be taken into account for future astrophysical missions.

The Topol program as a whole would survive the collapse of the Soviet Union, but just as with the BES-5, the TEU-5 never flew again after the Berlin Wall came down. KOSMOS-1867 was the last TEU-5 reactor, and the last US-AM satellite, to fly.

 

ENISY, The Final Soviet Reactor

The single-element thermionic reactor concept never went away. In fact, it remained in side-by-side development with the TOPOL reactor, and shared many of the basic characteristics, but was not ready in as timely a fashion as TOPOL was. The program was begun in 1967, with a total of 26 units built.

ENISY was seen to Soviet planners to be the logical extension of the TEU-5 program, and in many ways the reactor designs are linked. While the TEU-5 was designed for high-powered radar reconnaissance, the ENISY reactor was designed to be a communications and TV broadcast satellite. The amount of data that’s able to be transmitted is directly proportional to the amount of power available, and remains one of the most attractive advantages that astronuclear power plants offer to deep space probes (along with propulsion).

We’ll look at this design more in a later post, but it’s important to mention here since it is, in many ways, a direct evolution of the TEU-5. One nice thing about this reactor is that, due to the geometry of the reactor, its non-nuclear components were able to be tested as a unit without fissile fuel. Instead, induction heating units of the same size as the fuel elements could be slid into the location that the fuel rods would be for preflight testing without issues of neutron activation and material degradation due to the radiation flux.

ENISY for NEPSTP
ENISY reactor installation for NEP Space Test Program, DOE

This capability was demonstrated at the 8th US Symposium on Nuclear Energy in Albuquerque, NM, and led to the US purchasing two already-tested units from Russia (numbers V-71 and I-21U), with a buy option taken out on an additional four units, if needed. This purchase included technical information in the fuel elements, and offers of assistance from Russia to help in the fabrication of the fuel elements, but no actual fuel was sold. This reactor design would form the core of the American crewed lunar base concept in the 1990s as part of the Constellation program, as well as the core of a proposed technology demonstration mission deep space probe, but those programs never reached fruition.

We’ll look at this design in our usual depth in a couple blog posts. For now, it’s worth noting that this design reached flight-ready status; but, due to the financial situation of Russia after the collapse of the USSR, the increased availability of high-powered photovoltaic communications satellites, and the lack of funding for an American astronuclear flight test program, this reactor never achieved orbit as its predecessors did.

 

The Legacy of the USSR Astronuclear Program

 

The USSR flew more nuclear power plants than the rest of the world combined, 33 times more to be precise. Their program focused on an area of power generation that continues to hold great promise in the future, and in many ways helped define the problem for the rest of the world: in-core direct power conversion (something we’ll talk more about in the power conversion series). Even the failures of the program have taught the world much about the challenges of astronuclear design, and changed the face of what is and isn’t acceptable when it comes to flying a nuclear reactor in Earth orbit. The ENISY reactor went on to be the preferred power plant for American lunar bases for over a decade, and remains the only astronuclear design that’s been flight-certified by multiple countries.

Russia continues to build on the experience and expertise gained during the Romashka, BES-5, TEU-5, and ENISY programs. A recent test of a heat rejection system that offers far higher heat rejection capacity for its mass than any that has flown to date (a liquid droplet radiator, a concept we’ll cover in the thermal management post that will be coming up in a few months), their focus on high-power Hall thrusters, and their design for an on-orbit nuclear electric tug with a far more powerful reactor than any that we looked at today (1 MWe, depending on the power conversion system, likely between 2-5 MWt) shows that this experience has not been shoved into a closet and left to gather dust, but continues to be developed to advance the future of spaceflight.

 

More Coming Soon!

This post focused on the USSR and Russia’s astronuclear power plant expertise and operational history, a subject that very little has been written about in English (outside a number of reports, mostly focusing on the ENISY/TOPAZ-2 reactor), and is a subject that has long fascinated me. However, the USSR wasn’t the only country focusing on the idea, and wasn’t even the first to fly a reactor, just the most successful at making an astronuclear program.

The next post (which might be split into two due to the sheer number of fission power plant designs proposed in the US) is on the American programs from the same time, the Systems for Nuclear Auxiliary Propulsion, or SNAP, series of reactors (if split, the first post will cover SNAP-2, -10A, SNAPSHOT, -8, and the three reactors that evolved from SNAP-8, with SNAP50/SPUR, SABRE, SP-100, and possibly a couple more, as well as the ENISY/TOPAZ II US-USSR TSET/NEP Space Test Program/lunar base program). While the majority of the SNAP designs that were used were radioisotope thermoelectric generators, the ones that we’ll be focusing on are the fission power plants: the SNAP-2, SNAP-8, SNAP-10A (the first reactor to be launched into orbit), and the SNAP-100/SPUR reactor.

Following that, we’ll wrap up our look at the history of astronuclear electric power plants (the reactors themselves, at least) with a look at the designs proposed for the Strategic Defense Initiative (Reagan’s “Star Wars” program), return to a Russian-designed reactor which would have powered an America lunar base, had the funding for the base been available (ENISY), and the designs that rounded out the 20th century’s exploration of this fascinating and promising concept.

 

We’ll do one last post on NEP reactor cores looking at more recent designs from the last twenty years up to the present time, including the JIMO mission and a look at where Kilopower stands today, and then move on to power conversion systems in what’s likely to be another long series. As it stands that one will have a post on direct energy conversion, one on general heat engines and Stirling power conversion systems, one on Rankine cycle power conversion systems, one on Brayton cycle systems (including the ever-popular, never-realized, supercritical CO2 turbines), one on AMTEC and magnetohydrodynamic power conversion systems (possibly with a couple other non-mechanical heat engines as well), and a wrap up of the concepts, looking at which designs work best for which power levels and mission types. After that, it’ll be on to: heat rejection systems, for another multi-part series; a post on NEP ship and mission design; and, finally, one on bimodal NTR/NEP systems, looking at how to get the thrust of an NTR when it’s convenient and the efficiency of an NEP system when it’s most useful.

References

General References

http://www.buran.ru/htm/gud%2026.htm?fbclid=IwAR1jt9fsDZ10fHCSo42KUHjGTux8_uIkg43ClPrE1eg5IdQjXyhS2rSAHGY

http://www.proatom.ru/modules.php?name=News&file=print&sid=2740

https://sdelanounas.ru/blogs/29489/?fbclid=IwAR2zftn2RGOk8aU-3m1zGwNBhYMVY2zYFGUGwiorL6LSEFLNe8y-Pt4w_ag

http://elib.biblioatom.ru/text/atomnaya-energiya_t17-5_1964/go,16/?fbclid=IwAR2QzxdvVT5m3Kc3KPcVUmR5ZFrx_Er5d7RmKNFTzFz4k6-Djw_gPnnV6eA

Romashka

http://fti.neep.wisc.edu/neep602/SPRING00/lecture35.pdf

http://nacep.ru/novosti-energetiki/atomnaya-energetika/vysokotemperaturnyj-reaktor-preobrazovatel-romashka.html?fbclid=IwAR2W-9Exgyd63m6NGbVfNixGUzF9FrU2hsZUAMvdb9b75TBHQ6Ukh-EPMIA

http://nacep.ru/novosti-energetiki/atomnaya-energetika/termobatareya-romashka.html

Bouk

RORSAT page, Sven Grahn http://www.svengrahn.pp.se/trackind/RORSAT/RORSAT.html

Morning Light

The Life and Death of KOSMOS 954, Guy Weiss, courtesy Sven Grahn http://www.svengrahn.pp.se/trackind/RORSAT/cosmos954.pdf

History of the the 1035th Technical Operations Group, 1 January – 31 December 1978, via John Greenwald at The Black Vault http://documents.theblackvault.com/documents/accidents/morninglightusaf.pdf

History of the 437 Military Airlift Wing, Manning 1978, courtest John Greenwald at The Black Vault http://documents.theblackvault.com/documents/accidents/morninglight-histories.pdf

CIA Report C06607579, courtesy John Greenwald at The Black Vault http://documents.theblackvault.com/documents/cia/operationmorninglight-cia1.pdf

Topol-1

Gunther’s Space Page, PLASMA-A https://space.skyrocket.de/doc_sdat/plasma-a.htm

 

Categories
Development and Testing Electric propulsion History Non-nuclear Testing Nuclear Electric Propulsion Spacecraft Concepts

Electric Propulsion: The Oldest “Futuristic” Propulsion Possibility

Hello, and welcome back to Beyond NERVA. Today, we are looking at a very popular topic, but one that doesn’t necessarily require nuclear power: electric propulsion. However, it IS an area that nuclear power plants are often tied to, because the amount of thrust available is highly dependent on the amount of power available for the drive system. We will touch a little bit on the history of electric propulsion, as well as the different types of electric thrusters, their advantages and disadvantages, and how fission power plants can change the paradigm for how electric thrusters can be used. It’s important to realize that most electric propulsion is power-source-agnostic: all they require is electricity; how it’s produced usually doesn’t mean much to the drive system itself. As such, nuclear power plants are not going to be mentioned much in this post, until we look at the optimization of electric propulsion systems.

We also aren’t going to be looking at specific types of thrusters in this post. Instead, we’re going to do a brief overview of the general types of electric propulsion, their history, and how electrically propelled spacecraft differ from thermally or chemically propelled spacecraft. The next few posts will focus more on the specific technology itself, its’ application, and some of the current options for each type of thruster.

Electric Propulsion: What is It?

In its simplest definition, electric propulsion is any means of producing thrust in a spacecraft using electrical energy. There’s a wide range of different concepts that get rolled into this concept, so it’s hard to make generalizations about the capabilities of these systems. As a general rule of thumb, though, most electric propulsion systems are low-thrust, long-burn-time systems. Since they’re not used for launch, and instead for on-orbit maneuvering or interplanetary missions, the fact that these systems generally have very little thrust is a characteristic that can be worked with, although there’s a great deal of variety as far as how much thrust, and how efficient in terms of specific impulse, these systems are.

There are three very important basic concepts to understand when discussing electric propulsion: thrust-to-weight ratio (T/W), specific impulse (isp), and burn time. The first is self-explanatory: how much does the engine weigh, compared to how hard it can push, commonly in relation to Earth’s gravity: a T/W ratio of 1/1 means that the engine can hover, basically, but no more, A T/W ratio of 3/1 means that it can push just less than 3 times its weight off the ground. Specific impulse is a measure of how much thrust you get out of a given unit of propellant, and ignores everything else, including the weight of the propulsion system; it’s directly related to fuel efficiency, and is measured in seconds: if the drive system had a T/W ratio of 1/1, and was entirely made out of fuel, this would be the amount of time it could hover (assuming the engine is completely made out of propellant) for any given mass of fuel at 1 gee. Finally, you have burn time: the T/W ratio and isp give you the amount of thrust imparted per unit time based on the mass of the drive system and of the propellant, then the spacecraft’s mass is factored into the equation to give the total acceleration on the spacecraft for a given unit of time. The longer the engine burns, the more overall acceleration is produced.

Electric propulsion has a very poor thrust-to-weight ratio (as a general rule), but incredible specific impulse and burn times. The T/W ratio of many of the thrusters is very low, due to the fact that they provide very little thrust, often measured in micronewtons – often, the thrust is illustrated in pieces of paper, or pennies, in Earth gravity. However, this doesn’t matter once you’re in space: with no drag, and orbital mechanics not requiring the huge amounts of thrust over a very short period of time, the total amount of thrust is more important for most maneuvers, not how long it takes to build up said thrust. This is where the burn time comes in: most electric thrusters burn continuously, providing minute amounts of thrust over months, sometimes years; they push the spacecraft in the direction of travel until halfway through the mission, then turn around and start decelerating the spacecraft halfway through the trip (in energy budget terms, not necessarily in total mission time). The trump card for electric propulsion is in specific impulse: rather than the few hundred seconds of isp for chemical propulsion, or the thousand or so for a solid core nuclear thermal rocket, electric propulsion gives thousands of seconds of isp. This means less fuel, which in turn makes the spacecraft lighter, and allows for truly astounding total velocities; the downside to this is that it takes months or years to build these velocities, so escaping a gravity well (for instance, if you’re starting in low Earth orbit) can take months, so it’s best suited for long trips, or for very minor changes in orbit – such as for communications satellites, where it’s made these spacecraft smaller, more efficient, and with far longer lifetimes.

Electric propulsion is an old idea, but one that has yet to reach its’ full potential due to a number of challenges. Tsiolkovsy and Goddard both wrote about electric propulsion, but because neither was living in a time that it was possible to get into orbit, their ideas went unrealized in their lifetimes. The reason for this is that electric propulsion isn’t suitable for lifting rockets off the surface of a planet, but for in-space propulsion it’s incredibly promising. They both showed that the only thing that matters for a rocket engine is that, to put it simply, some mass needs to be thrown out the back of the rocket to provide thrust, it doesn’t matter what that something is. Electricity isn’t (directly) limited by thermodynamics (except through entropic losses), only by electric potential differences, and can offer very efficient conversion of electric potential to kinetic energy (the “throwing something out of the back” part of the system).

In chemical propulsion, combustion is used to cause heat to be produced, which causes the byproducts of the chemical reaction to expand and accelerate. This is then directed out of a nozzle to increase the velocity of the exhaust and provide lift. This is the first type of rocket ever developed; however, while advances are always being produced, in many ways the field is chasing after more and more esoteric or exotic ways to produce ever more marginal gains. The reason for this is that there’s only so much chemical potential energy available in a given system. The most efficient chemical engines top out around 500 seconds of specific impulse, and most hover around the 350 mark. The place that chemical engines excel though, is in thrust-to-weight ratio. They remain – arguably – our best, and currently our only, way of actually getting off Earth.

Thermal propulsion doesn’t rely on the chemical potential energy, instead the reaction mass is directly heated from some other source, causing expansion. The lighter the propellant, the more it expands, and therefore the more thrust is produced for a given mass; however, heavier propellants can be used to give more thrust per unit volume, at lower efficiencies. It should be noted that thermal propulsion is not only possible, but also common, with electrothermal thrusters, but we’ll dig more into that later.

Electric propulsion, on the other hand, is kind of a catch-all term when you start to look at it. There are many mechanisms for changing electrical energy into kinetic energy, and looking at most – but not all – of the options is what this blog post is about.

In order to get a better idea of how these systems work, and the fundamental principles behind electric propulsion, it may be best to look into the past. While the potential of electric propulsion is far from realized, it has a far longer history than many realize.

Futuristic Propulsion? … Sort Of, but With A Long Pedigree

The Origins of Electric Propulsion

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First Patented Ion Drive, Robert Goddard 1917

When looking into the history of spaceflight, two great visionaries stand out: Konstantin Tsiolkosky and Robert Goddard. Both worked independently on the basics of rocketry, both provided much in the way of theory, and both were visionaries seeing far beyond their time to the potential of rocketry and spaceflight in general. Both were working on the questions of spaceflight and rocketry at the turn of the 20th century. Both also independently came up with the concept of electric propulsion; although who did it first requires some splitting of hairs: Goddard mentioned it first, but in a private journal, while Tsiolkovsky published the concept first in a scientific paper, even if the reference is fairly vague (considering the era, understandably so). Additionally, due to the fact that electricity was a relatively poorly understood phenomenon at the time (the nature of cathode and anode “rays” was much debated, and positively charged ions had yet to be formally described); and neither of these visionaries had a deep understanding of the concepts involved, their ideas being little more than just that: concepts that could be used as a starting point, not actual designs for systems that would be able to be used to propel a spacecraft.

 

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Konstantin Tsiolkovsky, image via Wikimedia

The first mention of electric propulsion in the formal scientific literature was in 1911, in Russia. Konstantin Tsiolkovsky wrote that “it is possible that in time we may use electricity to produce a large velocity of particles ejected from a rocket device.” He began to focus on the electron, rather than the ion, as the ejected particle. While he never designed a practical device, the promise of electric propulsion was clearly seen: “It is quite possible that electrons and ions can be used, i.e. cathode and especially anode rays. The force of electricity is unlimited and can, therefore, produce a powerful flux of ionized helium to serve a spaceship.” The lack of understanding of electric phenomena hindered him, though, and prevented him from ever designing a practical system, much less building one.

 

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Robert Goddard, image viaWikimedia

The first mention of electric propulsion in history is from Goddard, in 1906, in a private notebook, but as noted by Edgar Choueiri, in his excellent historical paper published in 2004 (a major source for this section), these early notes don’t actually describe (or even reference the use of) an electric propulsion drive system. This wasn’t a practical design (that didn’t come until 1917), but the basic principles were laid out for the acceleration of electrons (rather than positively charged ions) to the “speed of light.” For the next few years, the concept fermented in his mind, culminating in patents in 1912 (for an ionization chamber using magnetic fields, similar to modern ionization chambers) and in 1917 (for a “Method and Means for Producing Electrified Jets of Gas”). The third of three variants was for the first recognizable electric thruster, whichwould come to be known as an electrostatic thruster. Shortly after, though, America entered WWI, and Goddard spent the rest of his life focused on the then-far-more-practical field of chemical propulsion.

 

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Yuri Kondratyuk, image wia Wikimedia

Other visionaries of rocketry also came up with concepts for electric propulsion. Yuri Kondratyuk (another, lesser-known, Russian rocket pioneer) wrote “Concerning Other Possible Reactive Drives,” which examined electric propulsion, and pointed out the high power requirements for this type of system. He didn’t just examine electron acceleration, but also ion acceleration, noting that the heavier particles provide greater thrust (in the same paper he may have designed a nascent colloid thruster, another type of electric propulsion).

 

 

 

 

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Hermann Oberth, image via Wikimedia

Another of the first generation of rocket pioneers to look at the possibilities of electric propulsion was Hermann Oberth. His 1929 opus, “Ways to Spaceflight,” devoted an entire chapter to electric propulsion. Not only did he examine electrostatic thrusters, but he looked at the practicalities of a fully electric-powered spacecraft.

 

 

 

 

 

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Valentin Glushko, image via Wikimedia

Finally, we come to Valentin Glushko, another early Russian rocketry pioneer, and giant of the Soviet rocketry program. In 1929, he actually built an electric thruster (an electrothermal system, which vaporized fine wires to produce superheated particles), although this particular concept never flew.By this time, it was clear that much more work had to be done in many fields for electric propulsion to be used; and so, one by one, these early visionaries turned their attention to chemical rockets, while electric propulsion sat on the dusty shelves of spaceflight concepts that had yet to be realized. It collected dust next to centrifugal artificial gravity, solar sails, and other practical ideas that didn’t have the ability to be realized for decades.

The First Wave of Electric Propulsion

Electric propulsion began to be investigated after WWII, both in the US and in the USSR, but it would be another 19 years of development before a flight system was introduced. The two countries both focused on one general type of electric propulsion, the electrostatic thruster, but they looked at different types of this thruster, reflecting the technical capabilities and priorities of each country. The US focused on what is now known as a gridded ion thruster, most commonly called an ion drive, while the USSR focused on the Hall effect thruster, which uses a magnetic field perpendicular to the current direction to accelerate particles. Both of these concepts will be examined more in the section on electrostatic thrusters; though, for now it’s worth noting that the design differences in these concepts led to two very different systems, and two very different conceptions of how electric propulsion would be used in the early days of spaceflight.

In the US, the most vigorous early proponent of electric propulsion was Ernst Stuhlinger, who was the project manager for many of the earliest electric propulsion experiments. He was inspired by the work of Oberth, and encouraged by von Braun to pursue this area, especially now that being able to get into space to test and utilize this type of propulsion was soon to be at hand. His leadership and designs had a lasting impact on the US electric propulsion program, and can still be seen today.

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SERT-I thruster, image courtesy NASA

The first spacecraft to be propelled using electric propulsion was the SERT-I spacecraft, a follow on to a suborbital test (Program 661A, Test A, first of three suborbital tests for the USAF) of the ion drives that would be used. These drive system used cesium and mercury as a propellant, rather than the inert gasses that are commonly used today. The reason for this is that these metals both have very low ionization energy, and a reasonably favorable mass for providing more significant thrust. Tungsten buttons were used in the place of the grids used in modern ion drives, and a tantalum wire was used to neutralize the ion stream. Unfortunately, the cesium engine short circuited, but the mercury system was tested for 31 minutes and 53 cycles of the engine. This not only demonstrated ion propulsion in principle, but just as importantly demonstrated ion beam neutralization. This is important for most electric propulsion systems, because this prevents the spacecraft from becoming negatively charged, and possibly even attracting the ion stream back to the spacecraft, robbing it of thrust and contaminating sensors on board (which was a common problem in early electric propulsion systems).

The SNAPSHOT program, which launched the SNAP 10A nuclear reactor on April 3, 1965, also had a cesium ion engine as a secondary experimental payload. The failure of the electrical bus prevented this from being operated, but SNAPSHOT could be considered the first nuclear electric spacecraft in history (if unsuccessful).

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ATS (either 4 or 5), image courtesy NASA

The ATS program continued to develop the cesium thrusters from 1968 through 1970. The ATS-4 flight was the first demonstration of an orbital spacecraft with electric propulsion, but sadly there were problems with beam neutralization in the drive systems, indicating more work needed to be done. ATS-5 was a geostationary satellite meant to have electrically powered stationkeeping, but was not able to despin the satellite from launch, meaning that the thruster couldn’t be used for propulsion (the emission chamber was flooded with unionized propellant), although it was used as a neutral plasma source for experimentation. ATS-6 was a similar design, and successfully operated for a total of over 90 hours (one failed early due to a similar emission chamber flooding issue). SERT-II and SCATHA satellites continued to demonstrate improvements as well, using both cesium and mercury ion devices (SCATHA wasn’t optimized as a drive system, but used similar components to test spacecraft charge neutralization techniques).

These tests in the 1960s never developed into an operational satellite that used ion propulsion for another thirty years. Challenges with the aforementioned thrusters becoming saturated, spacecraft contamination issues due to highly reactive cesium and mercury propellants, and relatively low engine lifetimes (due to erosion of the screens used for this type of ion thruster) didn’t offer a large amount of promise for mission planners. The high (2000+ s) specific impulse was very promising for interplanetary spacecraft, but the low reliability, and reasonably short lifetimes, of these early ion drives made them unreliable, or of marginal use, for mission planners. Ground testing of various concepts continued in the US, but additional flight missions were rare until the end of the 1990s. This likely helped feed the idea that electric propulsion is new and futuristic, rather than having its’ conceptual roots reaching all the way back to the dawn of the age of flight.

Early Electric Propulsion in the USSR

Unlike in the US, the USSR started development of electric propulsion early, and continued its development almost continuously to the modern day. Sergei Korolev’s OKB-1 was tasked, from the beginning of the space race, with developing a wide range of technologies, including nuclear powered spacecraft and the development of electric propulsion.

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Early sketch of a Hall effect (TAL) thruster in USSR, image from Kim et al

Part of this may be the different architecture that the Soviet engineers used: rather than having ions be accelerated toward a pair of charged grids, the Soviet designs used a stream of ionized gas, with a perpendicular magnetic field to accelerate the ions. This is the Hall effect thruster, which has several advantages over the gridded ion thruster, including simplicity, fewer problems with erosion, as well as higher thrust (admittedly, at the cost of specific impulse). Other designs, including the PPT, or pulsed plasma thruster, were also experimented with (the ZOND-2 spacecraft carried a PPT system). However, due to the rapidly growing Soviet mastery of plasma physics, the Hall effect thruster became a very attractive system.

There are two main types of Hall thruster that were experimented with: the stationary plasma thruster (SPT) and the thruster with anode layer (TAL), which refer to how the electric charge is produced, the behavior of the plasma, and the path that the current follows through the thruster. The TAL was developed in 1957 by Askold Zharinov, and proven in 1958-1961, but a prototype wasn’t built until 1967 (using cesium, bismuth, cadmium, and xenon propellants, with isp of up to 8000 s), and it wasn’t published in open literature until 1973. This thruster can be characterized by a narrow acceleration zone, meaning it can be more compact.

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E1 SPT-type Hall thruster, image via Kim et al

The SPT, on the other hand, can be larger, and is the most common form of Hall thruster used today. Complications in the plasma dynamics of this system meant that it took longer to develop, but its’ greater electrical efficiency and thrust mean that it’s a more attractive choice for station-keeping thrusters. Research began in 1962, under Alexy Morozov at the Institute of Atomic Energy; and was later moved to the Moscow Aviation institute, and then again to what became known as FDB Fakel (now Fakel Industries, still a major producer of Hall thrusters). The first breadboard thruster was built in 1968, and flew in 1970. It was then used for the Meteor series of weather satellites for attitude control. Development continued on the design until today, but these types of thrusters weren’t widely used, despite their higher thrust and lack of spacecraft contamination (unlike similar vintage American designs).

It would be a mistake to think that only the US and the USSR were working on these concepts, though. Germany also had a diversity of programs. Arcjet thrusters, as well as magnetoplasmadynamic thrusters, were researched by the predecessors of the DLR. This work was inherited by the University of Stuttgart Institute for Space Systems, which remains a major research institution for electric propulsion in many forms. France, on the other hand, focused on the Hall effect thruster, which provides lower specific impulse, but more thrust. The Japanese program tended to focus on microwave frequency ion thrusters, which later provided the main means of propulsion for the Hayabusa sample return mission (more on that below).

The Birth of Modern Electric Propulsion

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DS1 Mission Patch, Image courtesy JPL

For many people, electric propulsion was an unknown until 1998, when NASA launched the Deep Space 1 mission. DS1 was a technology demonstration mission, part of the New Millennium program of advanced technology testing and experimentation. A wide array of technologies were to be tested in space, after extensive ground testing; but, for the purposes of Beyond NERVA, the most important of these new concepts was the first operational ion drive, the NASA Solar Technology Applications Readiness thruster (NSTAR). As is typical of many modern NASA programs, DS1 far exceeded the minimum requirements. Originally meant to do a flyby of the asteroid 9969 Braille, the mission was extended twice: first for a transit to the comet 19/P Borrelly, and later to extend engineering testing of the spacecraft.

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NSTAR thruster, image courtesy NASA

In many ways, NSTAR was a departure from most of the flight-tested American electric propulsion designs. The biggest difference was with the propellant used: cesium and mercury were easy to ionize, but a combination of problems with neutralizing the propellant stream, and the resultant contamination of the spacecraft and its’ sensors (as well as minimizing chemical reaction complications and growing conservatism concerning toxic component minimization in spacecraft), led to the decision to use noble gasses, in this case xenon. This, though, doesn’t mean that it was a great overall departure from the gridded ion drives of US development; it was an evolution, not a revolution, in propulsion technology. Despite an early (4.5 hour) failure of the NSTAR thruster, it was able to be restarted, and the overall thruster life was 8,200 hours, and the backup achieved more than 500 hours beyond that.

Not only that, but this was not the only use of this thruster. The Dawn mission to the minor planet Ceres uses an NSTAR thruster, and is still in operation around that body, sending back incredibly detailed and fascinating information about hydrocarbon content in the asteroid belt, water content, and many other exciting discoveries for when humanity begins to mine the asteroid belt.

Many satellites, especially geostationary satellites, use electric propulsion today, for stationkeeping and even for final orbital insertion. The low thrust of these systems is not a major detriment, since they can be used over long periods of time to ensure a stable orbital path; and the small amount of propellant required allows for larger payloads or longer mission lifetimes with the same mass of propellant.

After decades of being considered impractical, immature, or unreliable, electric propulsion has come out of the cold. Many designs for interplanetary spacecraft use electric propulsion, due to their high specific impulse and ability to maximize the benefits of the high-isp, low-thrust propulsion regime that these thruster systems excel at.

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Electrothermal arcjet, image courtest Georgia Tech

Another type of electric thruster is also becoming popular for small-sat users: electrothermal thrusters, which offer higher thrust from chemically inert propellants in compact forms, at the cost of specific impulse. These thrusters offer the benefits of high-thrust chemical propulsion in a more compact and chemically inert form – a major requirement for most smallsats which are secondary payloads that have to demonstrate that they won’t threaten the primary payload.

So, now that we’ve looked into how we’ve gotten to this point, let’s see what the different possibilities are, and what is used today.

What are the Options?

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Ion drive schematic, image courtesy NASA

The most well-known and popularized version of electric propulsion is electrostatic propulsion, which uses an ionization chamber (or ionic fluid) to develop a positively charged stream of ions, which are then accelerated out the “nozzle” of the thruster. A stream of electrons is added to the propellant as it leaves the spacecraft, to prevent the buildup of a negative charge. There are many different variations of this concept, including the best known types of thrusters (the Hall effect and gridded ion thrusters), as well as field effect thrusters and electrospray thrusters.

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MPD Thruster concept, image courtesy NASA

The next most common version – and one with a large amount of popular mentions these days – is the electromagnetic thruster. Here, the propellant is converted to a relatively dense plasma, and usually (but not always) magnets are used to accelerate this plasma at high speed out of a magnetic nozzle using the electromagnetic and thermal properties of plasma physics. In the cases that the plasma isn’t accelerated using magnetic fields directly, magnetic nozzles and other plasma shaping functions are used to constrict or expand the plasma flow. There are many different versions, from magnetohydrodynamic thrusters (MHD, where a charge isn’t transferred into the plasma from the magnetic field), to the less-well-known magnetoplasmadynamic (MPD, where the Lorentz force is used to at least partially accelerate the plasma), electrodeless plasma, and pulsed inductive thruster (PIT).

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Electrothermal arcjet, image courtesy Georgia Tech

Thirdly, we have electrothermal drive systems, basically highly advanced electric heaters used to heat a propellant. These tend to be the less energy efficient, but high thrust, systems (although, theoretically, some versions of electromagnetic thrusters can achieve high thrust as well). The most common types of electrothermal systems proposed have been arcjet, resistojet, and inductive heating drives; the first two actually being popular choices for reaction control systems for large, nuclear-powered space stations. Inductive heating has already made a number of appearances on this page, both in testing apparatus (CFEET and NTREES are both inductively heated), and as part of a bimodal NTR (the nuclear thermal electric rocket, or NTER, covered on our NTR page).

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VASIMR operating principles diagram, image courtesy Ad Astra

The last two systems, MHD and electrothermal, often use similar mechanisms of operation when you look at the details, and the line between the two isn’t necessarily clear. For instance, some authors describe the pulsed plasma thruster (PPT), which most commonly uses a solid propellant such as PTFE (Teflon) as a propellant, which is vaporized and occasionally ionized electrically before it’s accelerated out of the spacecraft, as an MHD, while others describe it as an arcjet, and which term best applies depends on the particulars of the system in question. A more famous example of this gray area would be the VASIMR thruster, (VAriable Specific Impulse through Magnetic Resonance). This system uses dense plasma, contained in a magnetic field, but the plasma is inductively heated using RF energy, and then accelerated due to the thermal behavior of the plasma while being contained magnetically. Because of this, the system can be seen as an MHD thruster, or as an electrothermal thruster (that debate, and the way these terms are used, was one of the more enjoyable parts of the editing process of this blog post, and I’m sure one that will continue as we continue to examine EP).

Finally, we come to the photon drives. These use photons as the reaction mass – and as such, are sometimes somewhat jokingly called flashlight drives. They have the lowest thrust of any of these systems, but the exhaust velocity is literally the speed of light, so they have insanely high specific impulse. Just… don’t expect any sort of significant acceleration, getting up to speed with these systems could take decades, if not centuries; making them popular choices for interstellar systems, rather than interplanetary ones. Photonic drives have another option, as well, though: the power source for the photons doesn’t need to be on board the spacecraft at all! This is the principle behind the lightsail (the best-known version being the solar sail): a fixed installation can produce a laser, or other stream of photons (such as a maser, out of microwaves, in the Starwisp concept), which then impact a reflective surface to provide thrust. This type of system follows a different set of rules and limitations, however, from systems where the power supply (and associated equipment), drive system, and any propellant needed are on-board the spacecraft, so we won’t go too much into depth on that concept initially, instead focusing on designs that have everything on-board the spacecraft.

Each of these systems has its’ advantages and disadvantages. Electrostatic thrusters are very simple to build: ionization chambers are easy, and creating a charged field is easy as well; but to get it to work there has to be something generating that charge, and whatever that something is will be hit by the ionized particles used for propellant, causing erosion. Plasmadynamic thrusters can provide incredible flexibility, but generally require large power plants; and reducing the power requirements requires superconducting magnets and other materials challenges. In addition, plasma physics, while becoming increasingly well known, provides a unique set of challenges. Thermoelectric thrusters are simple, but generally provide poor specific impulse, and thermal cycling of the components causes wear. Finally, photon drives are incredibly efficient but very, very low thrust systems, requiring exceedingly long burn times to produce any noticeable thrust. Let’s look at each of these options in a bit more detail, and look at the practical limitations that each system has.

Optimizing the System: The Fiddly Bits

As we’ve seen, there’s a huge array of technologies that fall under the umbrella of “electric propulsion,” each with their advantages and disadvantages. The mission that is going to be performed is going to determine which types of thrusters are feasible or not, depending on a number of factors. If the mission is stationkeeping for a geosynchronous communications satellite, then the Hall thruster has a wonderful balance between thrust and specific impulse. If the mission is a sample return mission to an asteroid, then the lower thrust, higher specific impulse gridded ion thruster is better, because the longer mission time (and greater overall delta-v needed for the mission) make this low-thrust, high-efficiency thruster a far more ideal option. If the mission is stationkeeping on a small satellite that is a piggyback load, the arcjet may be the best option, due to its’ compactness, the chemically inert nature of the fuel, and relatively high thrust. If higher thrust is needed over a longer period for a larger spacecraft, MPD may be the best bet. Very few systems are designed to deal with a wide range of capabilities in spaceflight, and electric propulsion is no different.

There are other key concepts to consider in the selection of an electric propulsion system as well. The first is the efficiency of this system: how much electricity is required for the thruster, compared to how much energy is imparted onto the spacecraft in the form of the propellant. This efficiency will vary within each different specific design, and its’ improvement is a major goal in every thruster’s development process. The quality of electrical power needed is also an important consideration: some require direct, current, some require alternating current, some require RF or microwave power inputs, and matching the electricity produced to the thruster itself is a necessary step, which on occasion can make one thruster more attractive than another by reducing the overall mass of the system. Another key question is the total amount of change in velocity needed for the mission, and the timeframe over which this delta-v can be applied; in this case, the longer timeframe you have, the more efficient your thruster can be at lower thrust (trading thrust for specific impulse).

However, looking past just the drive itself, there are quite a few things about the spacecraft itself, and the power supply, that also have to be considered. The first consideration is the power supply available to the drive system. If you’ve got an incredibly efficient drive system that requires a MW to run, then you’re going to be severely limited in your power supply options (there are very few, if any, drive systems that require this high a charge). For more realistic systems, the mass of the power supply, and therefore of the spacecraft, is going to have a direct impact on the amount of delta-v that is able to be applied over a given time: if you want your spacecraft to be able to, say maneuver out of the way of a piece of space debris, or a mission to another planet needs to arrive within a given timeframe, the less mass for a given unit of power, the better. This is an area where nuclear power can offer real benefits: while it’s debatable whether solar or nuclear power is better for low-powered applications in terms of power per unit mass, which is known in engineering as specific power. Once higher power levels are needed, however, nuclear shines: it can be difficult (but is far from impossible) to scale nuclear down in size and power output, but it scales up very easily and efficiently, and this scaling is non-linear. A smaller output reactor and one that has 3 times the output could be very similar in terms of core size, and the power conversion systems used also often have similar scaling advantages. There are additional advantages, as well: radiators are generally speaking smaller in sail area, and harder to damage, than photovoltaic cells, and can often be repaired more easily (once a PV cell get hit with space debris, it needs to be replaced, but a radiator tube designed to be repaired can in many cases just be patched or welded and continue functioning). This concept is known as power density, or power-per-unit-volume, and also has a significant impact on the capabilities of many (especially large) spacecraft. The specific volume of the power supply is going to be a limiting factor when it comes to launching the vehicle itself, since it has to fit into the payload fairing of the launch vehicle (or the satellite bus of the satellite that will use it).

The specific power, on the other hand, has quite a few different implications, most importantly in the available payload mass fraction of the spacecraft itself. Without a payload, of whatever type is needed, either scientific missions or crew life support and habitation modules, then there’s no point to the mission, and the specific power of the entire power and propulsion unit has a large impact on the amount of mass that is able to be brought on the mission.

Another factor to consider when designing an electrically propelled spacecraft is how the capabilities and limitations of the entire power and propulsion unit interact with the spacecraft itself. Just as in chemical and thermal rockets, the ratio of wet (or fueled) to dry (unfueled) mass has a direct impact on the vehicle’s capabilities: Tsiolkovsky’s rocket equation still applies, and in long missions there can be a significant mass of propellant on-board, despite the high isp of most of these thrusters. The specific mass of the power and propulsion system will have a huge impact on this, so the more power-dense, and more mass-efficient you are when converting your electricity into useful power for your thruster, the more capable the spacecraft will be.

Finally, the overall energy budget for the mission needs to be accounted for: how much change in velocity, or delta-v, is needed for the mission, and over what time period this change in velocity can be applied, are perhaps the biggest factors in selecting one type of thruster over another. We’ve already discussed the relative advantages and disadvantages of many of the different types of thrusters earlier, so we won’t examine it in detail again, but this consideration needs to be taken into account for any designed spacecraft.

With each of these factors applied appropriately, it’s possible to create a mathematical description of the spacecraft’s capabilities, and match it to a given mission profile, or (as is more common) to go the other way and design a spacecraft’s basic design parameters for a specific mission. After all, a spacecraft designed to deliver 100 kg of science payload to Jupiter in two years is going to have a very different design than one that’s designed to carry 100 kg to the Moon in two weeks, due to the huge differences in mission profile. The math itself isn’t that difficult, but for now we’ll stick with the general concepts, rather than going into the numbers (there are a number of dimensionless variables in the equations, and for a lot of people that becomes confusing to understand).

Let’s look instead at some of the more important parts of the power and propulsion unit that are tied more directly to the drives themselves.

Just as in any electrical system, you can’t just hook wires up to a battery, solar panel, or power conversion system and feed it into the thruster, the electricity needs to be conditioned first. This ensures the correct type of current (alternating or direct), the correct amount of current, the correct amperage… all the things that are done on Earth multiple times in our power grid have to be done on-board the spacecraft as well, and this is one of the biggest factors when it comes to what specific drive is placed on a particular satellite.

After the electricity is generated, it goes through a number of control systems to first ensure protection for the spacecraft from things like power surges and inappropriate routing, and then goes to a system to actually distribute the power, not just to the thruster, but to the rest of the on-board electrical systems. Each of these requires different levels of power, and as such there’s a complex series of systems to distribute and manage this power. If electric storage is used, for instance for a solar powered satellite, this is also where that energy is tapped off and used to charge the batteries (with the appropriate voltage and battery charge management capability).

After the electricity needed for other systems has been rerouted, it is directed into a system to ensure that the correct amount and type (AC, DC, necessary voltage, etc) of electricity is delivered to the thruster. These power conditioning units, or PCUs, are some of the most complex systems in an electric propulsion systems, and have to be highly reliable. Power fluctuations will affect the functioning of a thruster (possibly even forcing it to shut down in the case of too low a current), and in extreme cases can even damage a thruster, so this is a key function that must be provided by these systems. Due to this, some designers of electrical drive systems don’t design those systems in-house, instead selling the thruster alone, and the customer must contract or design the PCU independently of the supplier (although obviously with the supplier’s support).

Finally, the thermal load on the thruster itself needs to be managed. In many cases, small enough thermal loads on the thruster mean that radiation, or thermal convection through the propellant stream, is sufficient for managing this, but for high-powered systems, an additional waste heat removal system may be necessary. If this is the case, then it’s an additional system that needs to be designed and integrated into the system, and the amount of heat generated will play a major factor in the types of heat rejection used.

There’s a lot more than just these factors to consider when integrating an electric propulsion system into a spacecraft, but it tends to get fairly esoteric fairly quickly, and the best way to understand it is to look at the relevant mathematical functions for a better understanding. Up until this point, I’ve managed to avoid using the equations behind these concepts, because for many people it’s easier to grasp the concepts without the numbers. This will change in the future (as part of the web pages associated with these blog posts), but for now I’m going to continue to try and leave the math out of the posts themselves.

Conclusions, and Upcoming Posts

As we’ve seen, electric propulsion is a huge area of research and design, and one that extends all the way back to the dawn of rocketry. Despite a slow start, research has continued more or less continuously across the world in a wide range of different types of electric propulsion.

We also saw that the term “electric propulsion” is very vague, with a huge range of capabilities and limitations for each system. I was hoping to do a brief look at each type of electric propulsion in this post (but longer than a paragraph or two each), but sadly I discovered that just covering the general concepts, history, and integration of electric propulsion was already a longer-than-average blog post. So, instead, we got a brief glimpse into the most general basics of electrothermal, electrostatic, magnetoplasmadynamic, and photonic thrusters, with a lot more to come in the coming posts.

Finally, we looked at the challenges of integrating an electric propulsion system into a spacecraft, and some of the implications for the very wide range of capabilities and limitations that this drive concept offers. This is an area that will be expanded a lot as well, since we barely scratched the surface. We also briefly looked at the other electrical systems that a spacecraft has in between the power conversion system and the thruster itself, and some of the challenges associated with using electricity as your main propulsion system.

Our next post will look at two similar in concept, but different in mechanics, designs for electric propulsion: electrothermal and magnetoplasmadynamic thrusters. I’ve already written most of the electrothermal side, and have a good friend who’s far better than I at MPD, so hopefully that one will be coming soon.

The post after that will focus on electrostatic thrusters. Due to the fact that these are some of the most widely used, and also some of the most diverse in the mechanisms used, this may end up being its’ own post, but at this point I’m planning on also covering photon drive systems (mostly on-board but also lightsail-based concepts) in that post as well to wrap up our discussion on the particulars of electric propulsion.

Once we’ve finished our look at the different drive systems, we’ll look at how these systems don’t have to be standalone concepts. Many designs for crewed spacecraft integrate both thermal and electric nuclear propulsion into a single propulsion stage, bimodal nuclear thermal rockets. We’ll examine two different design concepts, one American (the Copernicus-B), and one Russian (the TEM stage), in that post, and look at the relative advantages and disadvantages of each concept.

I would like to acknowledge the huge amount of help that Roland Antonius Gabrielli of the University of Stuttgart Institute for Space Studies has been in this post, and the ones to follow. His knowledge of these topics has made this a far better post than it would have been without his invaluable input.

As ever, I hope you’ve enjoyed the post. Feel free to leave a comment below, and join our Facebook group to join in the discussion!

References:

History

A Critical History of Electric Propulsion: The First Fifty Years, Choueiri Princeton 2004 http://mae.princeton.edu/sites/default/files/ChoueiriHistJPC04.pdf

A Method and Means of Producing Jets of Electrified Gas, US Patent 1363037A, Goddard 1917 https://patents.google.com/patent/US1363037A/en

A Synopsis of Ion Propulsion Development Projects in the United States: SERT I to Deep Space 1, Sovey et al, NASA Glenn Research Center 1999 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19990116725.pdf

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

NSTAR Technology Validation, Brophy et al 2000 https://trs.jpl.nasa.gov/handle/2014/13884

Review Papers for Electric Propulsion

Electric Propulsion: Which One for my Spacecraft? Jordan 2000 http://www.stsci.edu/~jordan/other/electric_propulsion_3.pdf

Electric Propulsion, Jahn and Choueiri, Princeton University 2002 https://alfven.princeton.edu/publications/ep-encyclopedia-2001

Spacecraft Optimization

Joint Optimization of the Trajectory and the Main Parameters of an Electric Propulsion Spacecraft, Petukhov et al 2017 https://reader.elsevier.com/reader/sd/D49CFC08B1988AA61C8107737D614C89A86DB8DAE56D09D3E8E60C552C9566ABCBB8497CF9D0CDCFB9773815820C7678

Power Sources and Systems of Satellites and Deep Space Probes (slideshow), Farkas ESA http://www.ujfi.fei.stuba.sk/esa/slidy_prezentacie/power_sources_and_systems_of_satellites_and_deep_space_probes_mk_2.pdf