hms_warrior

Nuclear Termal Turbo Rocket

14 posts in this topic

Posted (edited)

Stumbled about this on NSFF and since it sounds cool i just drop it here:

https://forum.nasaspaceflight.com/index.php?action=dlattach;topic=43344.0;attach=1438419;sess=0

Basicly it is a combination of a termal nuclear rocket feeding linear aerospike nozzels sitting on the turbine-fans of a turbojet-engine... or something like that XD

Have fun with it ;-)

PS: i want it for KSP!!!

PPS: the guy who designed it was senior engineer for the Raptor engine...

Edited by hms_warrior
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Posted (edited)

Quote

Single gas path rocket fan, ramjet, scramjet and pure rocket combined cycle propulsion system

Four very different modes sounds like an engineering nightmare, especially given the thrust to drag problems of pure scramjets (and that the simpler augmented NTRs were shown to have a negative payload fraction!). That said, wasn't Starship Congress looking for designs out to 2115 or so?

Edited by UmbralRaptor
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Looks cool, downloaded, I'll have to read it sometime :D

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In which I outline the plans for a nuclear-thermal ramrocket SSTO:

 

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The rocket-fans spinning at Mach 1.5 while pushing superheated hydrogen through them sounds like pure engineering hell... and that's before you even start encountering supersonic airflow. 
I think this rocket design's weakest point is the transition to scramjet mode. Even pure scramjet engines haven't been made to work reliably after decades of research, so something that can just switch to scramjet mode as part of four modes is like another circle in Dante's Inferno. 

If we use a hydrogen turbojet up to Mach 3 then an air-augmented oxygen afterburning rocket up to hypersonic speeds, then just switch to rocket mode and exit the atmosphere, we can get the Isp savings needed for a rocket that has 5-10x better payload fraction than chemical rockets. That's revolutionary enough. 

Although the entire concept does not address the main issue with nuclear designs: the fact that they're nuclear. 

 

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I prefer ammonia or just plain water for the working fluid. If you can get the reactor hot enough you can get as much isp as you need, with MUCH better impulse density.

A supersonic-flow ramrocket is MUCH easier to get working than a scramjet.

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1 hour ago, sevenperforce said:

If you can get the reactor hot enough you can get as much isp as you need, with MUCH better impulse density.

Yeah, but whatever your lightbulb's operating temperature, hydrogen will always have better Isp.

Although I do think the other remass materials are unjustly ignored given the reduced tankage size.

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6 minutes ago, DDE said:

Yeah, but whatever your lightbulb's operating temperature, hydrogen will always have better Isp.

Although I do think the other remass materials are unjustly ignored given the reduced tankage size.

The trouble is that at NERVA temperatures, hydrogen is the only fuel that really outperforms chemical hydrolox in terms of isp. Once you boost temperature a good bit higher (my preference is a ceramic-encapsuled molten-uranium pebble-bed), hydrogen is overkill (and very low-thrust).

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2 hours ago, sevenperforce said:

I prefer ammonia or just plain water for the working fluid. If you can get the reactor hot enough you can get as much isp as you need, with MUCH better impulse density.

A supersonic-flow ramrocket is MUCH easier to get working than a scramjet.

I don't understand why you would switch from hydrogen to ammonia. Hydrogen is being used as a fuel here. A majority of the NTTR's thrust at lower velocities comes from combustion of hydrogen in an oxygen atmosphere. Since oxygen propellant would mass 9x the hydrogen propellant but you are getting it free from the atmosphere, there is a massive saving in propellant fraction. Ammonia would not allow these gains.

57 minutes ago, DDE said:

Yeah, but whatever your lightbulb's operating temperature, hydrogen will always have better Isp.

Although I do think the other remass materials are unjustly ignored given the reduced tankage size.

Why do you mention a lightbulb, as in a closed-cycle gaseous core nuclear engine? 

47 minutes ago, sevenperforce said:

The trouble is that at NERVA temperatures, hydrogen is the only fuel that really outperforms chemical hydrolox in terms of isp. Once you boost temperature a good bit higher (my preference is a ceramic-encapsuled molten-uranium pebble-bed), hydrogen is overkill (and very low-thrust).

A pebble bed reactor might allow core temperatures to rise from about 3000K to 4500K. Solid materials cannot survive any higher. According to this Root Mean Square calculator, if we take the completely dissociated H2->H atoms at 1g/mol, then at 3000K a NERVA would reach 8.65km/s exhaust velocity while a pebble-bed would increase this by 22% to 10.6km/s. This is not a significant enough increase to consider moving away from hydrogen because it is 'overkill'. 

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Posted (edited)

2 hours ago, MatterBeam said:

I don't understand why you would switch from hydrogen to ammonia. Hydrogen is being used as a fuel here. A majority of the NTTR's thrust at lower velocities comes from combustion of hydrogen in an oxygen atmosphere. Since oxygen propellant would mass 9x the hydrogen propellant but you are getting it free from the atmosphere, there is a massive saving in propellant fraction. Ammonia would not allow these gains.

http://astronautix.com/l/loxammonia.html

Methane would also combust quite readily.

2 hours ago, MatterBeam said:

Why do you mention a lightbulb, as in a closed-cycle gaseous core nuclear engine?

Because I assumed that @sevenperforce's claims were solidly out of the solid-core reactor territory (pun not intended) in terms of core tmperature.

Edited by DDE
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4 hours ago, MatterBeam said:

I don't understand why you would switch from hydrogen to ammonia. Hydrogen is being used as a fuel here. A majority of the NTTR's thrust at lower velocities comes from combustion of hydrogen in an oxygen atmosphere. Since oxygen propellant would mass 9x the hydrogen propellant but you are getting it free from the atmosphere, there is a massive saving in propellant fraction. Ammonia would not allow these gains.

 

As @DDE pointed out, ammonia burns quite enthusiastically with atmospheric oxygen. But remember that the atmosphere is 70% nitrogen, so for any airbreathing engine, the majority of the thrust is actually coming from the momentum of the ejected nitrogen. That's why even a NTTR using water as a working fluid would be able to have tremendous thrust simply by compressing air, mixing it with nuclear-heated steam, and ejecting it.

4 hours ago, MatterBeam said:

A pebble bed reactor might allow core temperatures to rise from about 3000K to 4500K. Solid materials cannot survive any higher. According to this Root Mean Square calculator, if we take the completely dissociated H2->H atoms at 1g/mol, then at 3000K a NERVA would reach 8.65km/s exhaust velocity while a pebble-bed would increase this by 22% to 10.6km/s. This is not a significant enough increase to consider moving away from hydrogen because it is 'overkill'. 

But it is enough to get the isp of denser propellants up into SSTO territory and overcome hydrogen's low TWR.

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4 hours ago, DDE said:

http://astronautix.com/l/loxammonia.html

Methane would also combust quite readily.

Methane? CH4. Ammonia. NH3. 

48 minutes ago, sevenperforce said:

As @DDE pointed out, ammonia burns quite enthusiastically with atmospheric oxygen. But remember that the atmosphere is 70% nitrogen, so for any airbreathing engine, the majority of the thrust is actually coming from the momentum of the ejected nitrogen. That's why even a NTTR using water as a working fluid would be able to have tremendous thrust simply by compressing air, mixing it with nuclear-heated steam, and ejecting it.

But it is enough to get the isp of denser propellants up into SSTO territory and overcome hydrogen's low TWR.

Ammonia is hard to use as a fuel. 
It has 6.3 times less energy per kg than hydrogen and needs a catalyst to burn in air. That's normally possible in a rocket... but not at hypersonic speeds where the catalyst will melt. Although... maybe having the ammonia sperheated by the nuclear reactor beforehand might make this a non-issue.

The fact that the reaction products of an ammonia-oxygen flame are similar to the atmosphere at the engine inlet doesn't quite matter - the only thing which affects Isp is temperature, molar mass and proper expansion. An ammonia/oxygen flame mixed with with atmospheric gasses will burn colder and have higher molar mass than the same with hydrogen instead of ammonia. The small disadvantage in atmospheric flight becomes a massive one in the exoatmospheric regime, where ammonia's molar mass of 17 is compared to hydrogen's 2.

If we factor in dissociation of ammonia, the picture still isn't pretty. Complete dissociation of 1 mol ammonia gives us 1 mol of nitrogen and 3 mols of hydrogen, for an average molar mass of 4.25. This means and exhaust velocity less than half that of hydrogen. When the NTTR breaks out of the atmosphere at Mach 10, it might still have to produce 6km/s of deltaV to reach orbit. With dissociated ammonia at 3000K, it will have an exhaust velocity of 4.2km/s and needs a mass ratio of 4.18. With hydrogen at 3000K, it has an exhaust velocity of 8.65km/s and needs a mass ratio of 2.

Ammonia will mean you need a rocket twice as massive!

If we consider pebble-bed reactors at 4500K, the mass ratios are reduced to 3.2 and 1.76, which is still pretty much means the ammonia rocket is twice as heavy as the hydrogen rocket. 

However, further research shows that ammonia dissociates into N2 and H2 at about 600K, however N2 only completely dissociates at 30000K and H2 at 10000K. The real dissociated molar mass of an ammonia gas at 3000-4500K temperatures is very likely to be between 8.5 and 4.85 g/mol, leading to frankly terrible exhaust velocities of 2.9km/s at worst and 3.9km/s at best (3000K). 

Use this calculator: http://calistry.org/calculate/kineticTheoryVelocityCalculator

Why would hydrogen give a low TWR?

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10 hours ago, MatterBeam said:

Ammonia is hard to use as a fuel. 
It has 6.3 times less energy per kg than hydrogen and needs a catalyst to burn in air. That's normally possible in a rocket... but not at hypersonic speeds where the catalyst will melt. Although... maybe having the ammonia sperheated by the nuclear reactor beforehand might make this a non-issue.

Yeah, combustion of the ammonia would definitely be a non-issue.

10 hours ago, MatterBeam said:

The fact that the reaction products of an ammonia-oxygen flame are similar to the atmosphere at the engine inlet doesn't quite matter

Right, why would it?

You miss the point; I wasn't saying anything about the reaction products of an ammonia-oxygen flame.

Rather, I was pointing out that for any airbreathing engine, the majority of the thrust comes not from atmospheric oxygen, but from inert atmospheric nitrogen. For orbital aspirations, a ramrocket is not significantly less efficient than a ramjet, and it's a lot easier to achieve.

10 hours ago, MatterBeam said:

- the only thing which affects Isp is temperature, molar mass and proper expansion.

True isp, perhaps. But effective isp is the reciprocal of thrust-specific fuel consumption, which is where you have to take into account the mass flow at the inlet.

Even if you are merely pumping water into your nuclear core and mixing the resultant superheated steam-exhaust with the airflow, with no combustion whatsoever, the increase in thrust is massive. You don't have to worry about fine conditions for supersonic combustion, either. A denser exhaust will mix more effectively with the airflow as well; the only downside is that it may only be useful to a slightly lower Mach number (though honestly body heating is a bigger problem).

10 hours ago, MatterBeam said:

An ammonia/oxygen flame mixed with with atmospheric gasses will burn colder and have higher molar mass than the same with hydrogen instead of ammonia. The small disadvantage in atmospheric flight becomes a massive one in the exoatmospheric regime, where ammonia's molar mass of 17 is compared to hydrogen's 2.

Water/ammonia/whatever will have better in-atmosphere performance than hydrogen, up to its actual exhaust velocity. Secondary combustion contributes only a very small amount of additional energy. Thrust is linearly proportional to propellant density but only proportional to the square root of energy.

10 hours ago, MatterBeam said:

When the NTTR breaks out of the atmosphere at Mach 10, it might still have to produce 6km/s of deltaV to reach orbit. With dissociated ammonia at 3000K, it will have an exhaust velocity of 4.2km/s and needs a mass ratio of 4.18. With hydrogen at 3000K, it has an exhaust velocity of 8.65km/s and needs a mass ratio of 2.

Ammonia will mean you need a rocket twice as massive!

Dry mass, not wet mass, is the great determining factor for SSTO concepts, particularly for airbreathing concepts which need to carry a heavy fan to orbit and back. Liquid hydrogen masses 71 kg per cubic meter while liquid ammonia masses over 680 kg per cubic meter. This means that even though a liquid hydrogen SSTO may only need to hold half as much propellant mass, the tanks must be 4.8x larger!

10 hours ago, MatterBeam said:

Why would hydrogen give a low TWR?

Hydrogen's high specific impulse means it has a lot of energy but not very much thrust. Couple that with its very high tankage mass and you've got a uniquely low TWR. Some more on that...

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Posted (edited)

On 7/16/2017 at 10:25 AM, MatterBeam said:

the main issue with nuclear designs: the fact that they're nuclear

That's a marketing problem

Edited by Nothalogh
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