This page is about a specific design of the radiator type liquid fueled nuclear thermal rocket (LNTR). For more information about the LNTR in general, click here. For information about the radiator-type LNTR, click here.
The leader of the effort to study the trade-off between the bubbler (more info here) and radiator was one of the best-known fluid fueled NTR designers on the NASA side: Robert Ragsdale at Lewis Research Center (LRC, and we’ll come back to Ragsdale in gas core NTR design as well). There were a collection of studies around a particular design, beginning with a study looking at reactor geometry and fuel element size optimization to not only maximize the thrust and specific impulse, but also to minimize the uranium loss rates of the reactor.
This study concluded that there were many advantages to the radiator-type LNTR over the bubbler-type. The first consideration, minimizing the vapor entrainment problem that was laguing the bubbler, was minimized, but not completely eliminated, in the radiator design. The next conclusion is that the specific impulse of the negine could be maintained, or increased, to 1400 s isp or more. Finally, one of th emost striking was in thrust-to-core-weight ratio, which went from about 1:1 in the Nelson/Princeton design that we discussed in the last post all the way up to 19:1 (potentially)! This is because the propellant flow rate isn’t limited by the bubble velocity moving through the fuel.
These conclusions led to NASA gathering a team of researchers, including Ragsdale, Kasack, Donovan, Putre, and others ti develop the Lewis LNTR reactor.
Once the basic feasibility of the radiator LNTR was demonstrated, a number of studies were conducted to determine the performance characteristics, as well as the basic engineering challenges, facing this type of NTR. They were conducted in 1967/68, and showed distinct promise, for the desired 2000 to 5000 MWt power range (similar to the Phoebus 2 reactor’s power goal, which remains the most powerful nuclear reactor ever tested at 3500 MWt).
The three primary goals were: to maximize specific impulse, maximize thrust-to-weight ratio, and minimize uranium mass loss. They quickly discovered that they couldn’t have their cake and eat it, too: higher temperatures, and therefore higher isp, led to faster U mass loss rates, increasing T/W ratio reduced the specific impulse, and minimizing the U loss rate hurt both T/W and isp. They could improve any one (or often two) of these characteristics, but always at the cost of the third characteristic.
We’ll look at many of the design characteristics and engineering considerations of the LRC work in the next section on general design challenges and considerations for the radiator LNTR, but for now we’ll look at their final compromise reactor.
The reactor itself would be made up of several (oddly, never specified) fuel elements, in a beryllium structure, with each fuel element being made up of Be as well. These would be cooled by cryogenic hydrogen moving from the nozzle end to the spacecraft end of the reactor, before flowing back into the central void of the fuel element. As it was fed through the central annulus, it would be seeded with tungsten microparticles to increase the amount of heat the propellant would absorb. Finally, it would be exhausted through a standard De Laval nozzle to provide thrust.
The final fuel that they settled on was a liquid ternary carbide design, with the majority of the fuel being niobium carbide (although ZrC was also considered), with a molar mass fraction of 0.02 being UC2. This compromise offered good power density for the reactor while minimizing the vaporization rate of the fuel mass. This would be held in 2 inch diameter, 5 foot long fuel element tubes, with a fuel surface temperature of 5060 K. The propellant would be pressurized to 200 atm in the reactor.
This led to a design that struck a compromise between isp, T/W, and U mass loss which was not only acceptable, but impressive: 1400 s isp (on par with some forms of electric propulsion), a T/W ratio (of the core alone) of 4, and a hydrogen-to-uranium flow rate ratio of 50.
They did observe that none of these characteristics were as high as they could be, especially in terms of T/W ratio (which they calculated could go as high as 19!), or isp (with a theoretical maximum of 1660), and the uranium loss was twice the theoretical minimum, but sadly the cost of maximizing any of these characteristics was so high from an engineering point of view that it wasn’t feasible.
Sadly, I haven’t been able to find any documentation on this reactor design – and very few references to it – after February 1968. The exact time of the cancellation, and the reasons why, are a mystery to me. If someone is able to help me find that information it would be greatly appreciated.
For more general considerations for the radiator LNTR, check out the Radiator LNTR page.
References
PERFORMANCE POTENTIAL OF A RADIANT-HEAT-TRANSFER LIQUID-CORE NUCLEAR ROCKET ENGINE, Ragsdale 1967 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19670030774.pdf
HEAT- AND MASS-TRANSFER CHARACTERISTICS OF AN AXIAL-FLOW LIQUID-CORE NUCLEAR ROCKET EMPLOYING RADIATION HEAT TRANSFER, Ragsdale et al 1967 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19670024548.pdf
FEASIBILITY OF SUPPORTING LIQUID FUEL ON A SOLID WALL IN A RADIATING LIQUID-CORE NUCLEAR ROCKET CONCEPT, Putre and Kasack 1968 https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19680007624.pdf