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There are a lot of articles of gas-core NTR, but mostly are very superficial: only some schematic design and very few data on real technical fasibility.
Now I found this very robust study, where all critical issues like vortex formation, core-propellant heat transfer, propellant seeding, wall and nozzle cooling are properly addressed.
http://ntrs.nasa.gov/archive/nasa/casi. … 003405.pdf
It's to note how simple and practical is the fuel injection system: an uranium bar in a cadmium protected pipe: when a segment of of uranium protrude outside the pipe, it quickly vaporize and mix with the vortex forming the core.
It's all very simple and can be done with existing technologies and materials. I guess almost 10 yr of R&D (if properly founded) and we will have a 1500-2500s Isp rocket, that can explore all the solar system.
Last edited by Quaoar (2014-05-02 16:05:24)
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Thanks, Quaoar, that's one of the reports I never saw decades ago. I saw the Ragsdale Mars mission stuff, and I saw the United Aircraft reports on the stuff they did. A version of that spherical engine is what was proposed, back in about 1969, for the then-planned 1980's manned Mars mission. United Aircraft was of the opinion that 35:1 LH2:U flow ratio was as good as perfect containment, at the U burnup rates they were looking at.
Another version of that same spherical engine operated at higher power, but with a waste heat radiator because they thought regenerative cooling would not be adequate. Vehicle thrust to weight was somewhere between 0.01 and 0.1 gees, but the Isp was 6000 s, not just 2500 s. United Aircraft thought the tradeoff point for radiator-or-not was about 2000-2500 s Isp reactor power levels. Their thinking was even higher engine T/W ratios than NASA, without the radiator, perhaps 30:1.
Point is, this thing could have been ready for test flights back in the early to mid 1980's, if it hadn't all been killed by 1974.
GW
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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Thanks, Quaoar, that's one of the reports I never saw decades ago. I saw the Ragsdale Mars mission stuff, and I saw the United Aircraft reports on the stuff they did. A version of that spherical engine is what was proposed, back in about 1969, for the then-planned 1980's manned Mars mission. United Aircraft was of the opinion that 35:1 LH2:U flow ratio was as good as perfect containment, at the U burnup rates they were looking at.
Another version of that same spherical engine operated at higher power, but with a waste heat radiator because they thought regenerative cooling would not be adequate. Vehicle thrust to weight was somewhere between 0.01 and 0.1 gees, but the Isp was 6000 s, not just 2500 s. United Aircraft thought the tradeoff point for radiator-or-not was about 2000-2500 s Isp reactor power levels. Their thinking was even higher engine T/W ratios than NASA, without the radiator, perhaps 30:1.
Point is, this thing could have been ready for test flights back in the early to mid 1980's, if it hadn't all been killed by 1974.
GW
I found even this interesting work on a Droplet Core Rocket that can reach 2000 s of Isp. In this roket, the uranium droplets are centrifugated and recycled with a liquid lithium flow system: it may be cheeper than Gas Core where the core is vented out after every burn.
https://discover.tudelft.nl/recordview/ … 9920001887
I found also this review about all kind of NTR
http://www.google.it/url?sa=t&rct=j&q=& … 8261,d.ZGU
Most of them are designed in 1960-70. I think that using modern high temperature resistant materials, running at low pressure to achive high hydrogen dissociation and using a particle bed architecture or a FOIL (Fission Fragment Assisted Reactor) even a solid core can reach 1500+ s of Isp.
Last edited by Quaoar (2014-05-03 13:19:55)
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This is good stuff, Quaoar. I saved a copy of the 1991-vintage NASA report. It mentions the droplet core reactor as one of several concepts.
That 1991 NASA report has the results of a panel who put together the outline of a program to focus on an improved NERVA and carry along R&D on some of the better advanced concepts. Their plan would have required a dedicated plume-capture test facility, probably in Nevada at the nuclear test site, and would have brought some version of NERVA on-line for flights in about 2 decades. Could have happened faster at higher funding.
It is worth noting that both the CIA and the KGB got interested in solid core NTR for a while, about that same time. I saw photos of mockups for both agencies not many years after. None of that ever led anywhere, though. There was a fear of a Soviet particle beam weapon with a nuclear energy source under test about that time, based on some fallout that floated over northern Europe from Russia, and a very odd-looking complex at Sary Shagan. It turned out to be their 1990's-vintage nuclear rocket testing.
Myself, I think all these concepts should be pursued at "best possible speed". What with the political fears of nuclear things, and the expense and complexity of a plume-capture test facility, testing experimentally on Earth is a poor option. You cannot test an engine free-falling in space, where every test is a vehicle flight test. I recommend building a nuke engine test facility on the moon. Right now, there is no better compelling reason to go back there. None of NASA's rocket (SLS) and capsule (Orion CEV) efforts are suitable for men-to-Mars (or even an asteroid in-situ "out there"), but they are suitable for men-to-the-moon.
GW
GW Johnson
McGregor, Texas
"There is nothing as expensive as a dead crew, especially one dead from a bad management decision"
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The bulk of reasearch has been done in 60-70thy, and all the rocket has designed for the materials of that times.
With today materials we can build better and better rocket. I imagine some kind of "hairy core" NTR, stable, restartable and easy to handle like a solid core NTR, but with the Isp of a gas core: the core may be a bush of graphite fiber or carbon nanotube, coated with uranium 233 zirconium hafnium ternay carbide that heat hydrogen (or water steam) for nuclear fragmentation, so the propellant can reach more than 5000°K while the fuel is still 3000°K, resulting a more than 1600 s of Isp without the uranium loss of a gas core.
Last edited by Quaoar (2014-05-05 16:45:10)
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