Are there any issues in transforming the liquid oxygen into gaseous oxygen before injecting into the combustion chamber?

Gaseous Oxygen will be readily available for combustion and will mix well with fuel for complete combustion. I also did read that Pintle nozzle is effective in two-phase atomisation.

If so, why not transform the liquid into gas before injecting into the combustion chamber? Are there any impediments that are non-intuitive?

Edit: the phase transformation from LOx to GOx could be achieved by using LOx for cooling the nozzle. Would using LOx for nozzle cooling be operationally infeasible?

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    $\begingroup$ My guess is the much lower density of GOx does not allow (or at least makes it pretty hard) to get high enough flow rate? $\endgroup$
    – jkavalik
    Feb 3 '18 at 10:57
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    $\begingroup$ Where would the energy to transform the liquid into gas come from? And what part of the rocket engine cycle would this energy be removed from, and what would be the subsequent effects on performance? Do a quick calculation on the energy required to convert ~950 lbm/sec of LOX at -297 deg F into GOX and you will probably answer your own question. Note: the combustion chamber in existing designs does an excellent job at converting this liquid into gas. $\endgroup$ Feb 3 '18 at 17:45
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    $\begingroup$ Liquid oxygen is needed to store a lot of oxygen in a small and lightweight tank. To transform LOX to GOX, a heat exchanger and a heat source is necessary which would add more mass to the engine. The use of gaseous oxygen may save a little time for complete combustion with fuel, but a little shorter combustion chamber would not compensate the mass of the additional heat exchanger. Unwanted oxidation of the heat exchanger and the injector should be avoided but this is difficult with hot oxygen. $\endgroup$
    – Uwe
    Feb 3 '18 at 21:57
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    $\begingroup$ Just discovered the answer to my question here space.stackexchange.com/q/21374/511 $\endgroup$ Feb 4 '18 at 8:40
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    $\begingroup$ It seems important to note that oxygen rich staged combustion (RD-180, and others) and full flow staged combustion (raptor) do effectively pre-heat the oxygen, by running it through the turbo-pump pre-burner. The hot oxidizer issues mentioned by others are not insurmountable problems. I remember seeing that in the case of the raptor at least, it is heated to the point of being gas injected into the combustion chamber. $\endgroup$
    – Lex
    Feb 4 '18 at 17:23

tl;dr: Gases can take up from hundreds to a thousand times the volume than their liquid counterparts, even at their boiling point. Pipes for these gasses would be too huge to even fit in the rocket.

Because cryogenic propellants are challenging to keep cold in rockets where every bit of weight used for cryogenic storage would mean roughly 20 to 50 times that weight must be removed from the final payload, the greatest use of LOX (by volume) is in the first stages of rockets leaving both the Earth, and its earth-bound LOX production facilities.

This certainly will change in the future, (see Can five refillings of the BFR second stage be useful to get to the Moon? To Mars? All five in Earth orbit?) but the question is not about future rockets.

These engines use propellants at a very, very high rate of mass flow, and so the pipes that carry the LOX are already fairly large. For example using the Wikipedia values for ONE Merlin Engine at sea level: thrust 420,000 Newtons and Isp 275 seconds times 9.8 m/s gives a exhaust velocity of about 2700 m/s. Divide thrust by velocity and you get over 150 kg/second of exhaust, and since every carbon gets two oxygens, that's about 100 kg/sec of LOX for each of the nine engines!

It's a non-trivial fluid dynamics problem to figure out what extreme techniques are necessary to move almost a metric ton of per second of gaseous oxygen through a moderate size pipes. They would certainly have to be extremely large, and they would not likely even fit within the svelt, almost "noodle-like" 3.66 meter diameter F9 body. (See If not constrained by underpasses, etc., would Falcon 9 have been less of a flying noodle? for more on that).

So really, the only place you would like to have this factor of ~1000 expansion happen is in the parts of the rocket that's build for expansion and rapid transport of gas; the combustion chamber itself, and the throat, and nozzle.

  • $\begingroup$ What is the expansion ratio at the compression needed for injection into the combustion chamber? $\endgroup$
    – user20636
    Feb 4 '18 at 12:35
  • $\begingroup$ @JCRM I don't understand your question. If I did, I likely wouldn't know the answer. $\endgroup$
    – uhoh
    Feb 4 '18 at 12:52
  • $\begingroup$ Kg/s of thrust? A thrust rate? What does that mean? Also, what is a "LOX refrigerator"? $\endgroup$ Feb 4 '18 at 13:21
  • $\begingroup$ @OrganicMarble there's nothing wrong with talking about it that way. Thrust (as in the stuff out the nozzle) is mass. If you know the average exhaust velocity and thrust, it gives you mass per second. Mass is conserved (to ppm accuracy, save for Einstein), so mass of thrust per second is mass of propellant per second. Take a few minutes to work through the math a different way, you should get the same answer. $\endgroup$
    – uhoh
    Feb 4 '18 at 16:29
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    $\begingroup$ @OrganicMarble "...the Earth and its LOX refrigerators." That's where they are, on Earth. As far as we know, it's the only place one can obtain rocket-grade LOX in the Solar System. I'll see if I can think of a different wording though. $\endgroup$
    – uhoh
    Feb 4 '18 at 16:51

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