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http://hdl.handle.net/2060/19920001876

Low-Pressure Nuclear Rocket Concept, by J.H. Ramstaler

Cliffnotes:

  • Reduced core pressure to allow normally unsustainable exhaust temperatures, which reach into hydrogen dissociation range, resulting in mono-H exhaust; the active zone is spherical rather than cylindrical, with propellant injected right into it. Variable specific impulse capability, thanks to no need for a turbopump-powered injector.
  • The reactor relies entirely on hydrogen as neutron moderator and has no other control system outside of a SCRAM rod; the reaction mass flow thermalizes the neutron flux, providing neutrons for uranium fission - no propellant, no thermal neutrons, the reactor goes subcritical.

As almost always, I'm just reposting from @nyrath.

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Looks nice on paper. Too bad there will be no one brave and rich enough to try building and testing this engine. NASA - the only institution with money and know-how to attempt it prefers to sink good chunk of its budget into the white elephant called SLS.

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Just now, Scotius said:

Looks nice on paper. Too bad there will be no one brave and rich enough to try building and testing this engine. NASA - the only institution with money and know-how to attempt it prefers to sink good chunk of its budget into the white elephant called SLS.

The specialist section of NASASpaceflight is curiously very hostile to NTRs and champions electric motors above all else.

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Then they should work on MagnetoPlasmaDynamic thrusters, or other high thrust (for electrics) engines, instead of ion toys good only for probes. And scientific missions that from the get-go postulate transfer periods counted in several years.

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5 minutes ago, Scotius said:

Then they should work on MagnetoPlasmaDynamic thrusters, or other high thrust (for electrics) engines, instead of ion toys good only for probes.

"High flight-readiness".

"Nuclear reactors for beyond Mars? We'll get back to you on that".

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1 hour ago, DDE said:

http://hdl.handle.net/2060/19920001876

Low-Pressure Nuclear Rocket Concept, by J.H. Ramstaler

Cliffnotes:

  • Reduced core pressure to allow normally unsustainable exhaust temperatures, which reach into hydrogen dissociation range, resulting in mono-H exhaust; the active zone is spherical rather than cylindrical, with propellant injected right into it. Variable specific impulse capability, thanks to no need for a turbopump-powered injector.
  • The reactor relies entirely on hydrogen as neutron moderator and has no other control system outside of a SCRAM rod; the reaction mass flow thermalizes the neutron flux, providing neutrons for uranium fission - no propellant, no thermal neutrons, the reactor goes subcritical.

As almost always, I'm just reposting from @nyrath.

Jeeze, when is NASA going to stop using pounds. He must have been a really old timer.

Thrust = 50 kN at 841 kilograms (Not to bad) as 1250 ISP * 9.8 = 12250 m/s Ve

Quote

The thrust level is too low for Earth escape but it is useful for other manuevers" https://ntrs.nasa.gov/archive/nasa/casi.ntrs.nasa.gov/19920001876.pdf (page 128's bottom sentence)

The question is not so much level is throttleability, can it be shut off.

Here's what we can see, what if we have a controller the size of a flea and the hydrogen tanks are 20% of the total fuel and tank weight.
Its rather simple then to lower the Ve to 9800 (back envelope primary estimate) . Lets just say that we can shed hydrogen tanks as we use some that we only at the end carry one small tank at the core which all other tanks drain into. If we 11300 - 7800 = 3500 dV

3500 dV = 9800  ln (M0/M1)

0.4292F x Engine mass = 360.23 Kgs of hydrogen (liquified I think thats like 3600 liters of hydrogen).
Initial A = 50000N / (360.23 + 841) = 41.62 (4.162 G) results in 84 second of burn time. dΘ/dt of LEO is 0.001230 and the burn would occur over 0.10325 radians (6 degrees, actually less because we lost 30% of the mass at the end) the cosine losses would be less than 0.13%

IOW, yes it can leave Earth orbit. Thus the reason for not exit Earth orbit is do to the nature of the exhaust propellant.  Lets say we place a restriction that its burn span needs to be less that 45 degrees, that means 22.5 on one side and 22.5 on the other side of an assumed hyperbolic periapsis.
Lets assume the average speed is 10,000 m/s and that the dΘ/dt is thus 0.0015, therefore the allowable burn time for exit is 523 seconds, if we set this as our criteria (again speed will increase), 523 Second at 50,000 N = 26150000 N * sec. 2135 kg of fuel used. 427 kg of tanks used, 841 engine

3500 =  12250 ln( (PL + 2135 + 841 + 427)/(PL + 841))

e0.2857 = (PL + 2135 + 841 + 427)/(PL + 841) = 1.33333

k * 0.33333 = 2135 + 427 = 2562

k = 7868

PL = k - 841 = 7047 kg.

Therefore you could carry on the NTR  7047kg of PL (i.e. more than a flea on a rocket engine) to Earth escape. This is a bit of an over estimate, because a substantial amount of the burn would be at non-optimal velocity due to separation from minimum orbit before burn completion. However it the rocket can be throttled on and off completely, then the orbit can be achieved in a number of kicks. This of course is not the limit of the rocket, because we could increase the number of kicks and increase payload. But I single NTR of this type would hardly transport a crew, at least 30T of space craft, to Mars Orbit.

 

 

 

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Heck on some of these things I start to worry if I'll see them in MY lifetime... and I just turned 20.  It's not like they are that hard, the government just doesn't seem that interested and when it is interested it doesn't seem to stay so for long. I was looking forward to the Ares 1 so much...  still suffering from cancelation heartbreak

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Quote

I might mention that NERVA demonstrated that you could operate with reactivity control fixed. The drums were fixed and could run a complete startup, full power hold, and complete shutdown on reactivity feedack (no control drum movement).

This just blew my mind. I knew this from theoretical advanced design, but never realized it was actually tested. But it begs a question - if I understand it correctly, these designs are generaly cooled by pumping hydrogen through zone after "active" burn until fuel elements cool down. But if this increases reactivity, that kind of defeats the idea? How is that thing supposed to shut down? (I can see neutron poison injection in one of diagrams, but that looks like emergency feature only).

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2 minutes ago, radonek said:

This just blew my mind. I knew this from theoretical advanced design, but never realized it was actually tested. But it begs a question - if I understand it correctly, these designs are generaly cooled by pumping hydrogen through zone after "active" burn until fuel elements cool down. But if this increases reactivity, that kind of defeats the idea? How is that thing supposed to shut down? (I can see neutron poison injection in one of diagrams, but that looks like emergency feature only).

Possibly it can do startup, full-power burn, and shutdown without rotating the control drums...but then you need to rotate the control drums to lower the reactivity of the core for subsequent cool-down phases.

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17 hours ago, radonek said:

if I understand it correctly, these designs are generaly cooled by pumping hydrogen through zone after "active" burn until fuel elements cool down. But if this increases reactivity, that kind of defeats the idea? How is that thing supposed to shut down? (I can see neutron poison injection in one of diagrams, but that looks like emergency feature only).

I suspect that the flow rate needed to keep the core from melting due to residual heat is far less than what's needed as a moderator.  It does imply that you can't just "shut off" thrust, however -- minimum cooling flow will still produce considerable thrust, albeit only a small fraction of normal operation.
 

The other big thing with these LH2 designs is: LH2 isn't even space storable.  You can't fly this engine to, say, Jupiter, because most of your capture remass will boil off during your transfer.  So, TWR is too low for launch, almost too low for transfer from LEO, and the engine can't be used after a long flight.  And for the LPNTR design, there's really no point to trying to run it on something with longer tank life like water or ammonia -- the Isp won't even be close.

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Technically you could retain the hydrogen with a double walled tank in which the second wall is used to recapture hydrogen. You would have to carry helium along as a refrigerant to and a sizable heat exchanger to drop heat from compressed helium into space, solar panels to drive the compressor. It all gets kind of iffy with current solar panels beyond the asteroid belt due to the lack of power.

The hypothesis here arrives at the problem is that more refined the thrust generation the lower the power, bomby like things such as SFRBs are great at producing power but have lackluster ISP. Photon drives which have pentultimate ISP have almost no thrust, and ION drives that are intermediate have minimal thrust.

The problem NTR is to leave orbit it needs help, the engine is not the best at shutting down. On and ION drive I can reduce off power heat simply by tilting solar panels. The second problem is the place you want to use it the most, that critical last pass of LEO as it heads into deep space, politically is not agreeable.

X -ION drive, cannae, photon . . . .very low trust high efficiency engines . . . .Need alot of electric power . . . . . power supply not available (solar bulky/heavy and limited by sun, fission inefficient limited by heat).

Y- NTR drive . . . .heavy. . . . .low thrust . . . . .moderately efficient  . . . . . . . . Need alot of bulky hydrogen, may have heat problems when attenunated.

Z -Hydrolox drive . . . . light weight . . . moderate/low thrust . . . . . inadequate deep space efficiency . . . . Need alot of bulky hydrogen.

 

 

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