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Imagine a rocket that uses a mixture of liquid nitrogen and oxygen as the oxidiser.

It is designed like an oxidiser-rich staged-combustion engine, however there is so much nitrogen in the mixture that all the fuel can be burned in the pre-burner without the gas becoming so hot that it melts the oxidiser pump turbine. The pre-burner then, in effect, becomes the main combustion chamber, and there is no need for a combustion chamber downstream of the oxidiser pump. It may be that there is no need for regenerative cooling channels. The engine is used on very low staging-velocity boosters where high thrust to weight is more important than exhaust velocity in achieving greater efficiency.

How would such an engine be optimised (eg. expansion ratio, isp, pressure) to achieve maximum thrust to weight ratio and minimum production and development cost?

(It's okay for it to be as low perfomance ISP-wise as a steam rocket - though probably with far better mass fraction. Perhaps optimising some-way between the ARCA steam rocket CONCEPT (i.e. on paper not reality) and a conventional chemical rocket)

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  • $\begingroup$ How exactly do you envision reaching a high thrust to weight ratio when you sacrifice a significant part of your chamber pressure to power the turbopumps? Also, routing the entire exhaust flow through the turbine means you'll need a really big turbine and heavy piping. I guess you don't want to choke the flow anywhere before the nozzle, so you'll need really really fat pipes. $\endgroup$ – TooTea Dec 23 '20 at 13:11
  • $\begingroup$ This sounds like effectively a jet engine running off pumped liquefied air. It'd have most of the complexity of a staged combustion engine (except having much less available power to push a much larger mass flow) while being outperformed by pressure-fed boosters (which aren't limited to steam rocket levels of performance). $\endgroup$ – Christopher James Huff Dec 23 '20 at 13:39
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    $\begingroup$ If there's no combustion chamber, what are you pumping into? Factor out the turbopump entirely and use a pressure-fed engine. The suggestion of diluting the oxidizer with nitrogen puts me in mind of the V-2 and Redstone rockets. Those used 75%/25% alcohol/water as fuel; the water didn't contribute to combustion, so it kept combustion temperatures low while contributing mass for high thrust. Specific impulse suffered badly, so once the development of RP-1/RG-1 made regenerative cooling practical, the water-diluted fuels went by the wayside. $\endgroup$ – Russell Borogove Dec 23 '20 at 19:20
  • $\begingroup$ If the gas can not melt the turbine, then it's not hot enough for the nozzle. $\endgroup$ – user3528438 Dec 26 '20 at 0:43
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    $\begingroup$ @Tobe perhaps something closer to what you explained in your question would be using mercury as described in john clarks book "ignition!" to dilute the fuel. it seems very counterintuitive as it is definitely on the heavier side and not helping with exhaust velocity, but despite the drastic decrease in isp, booster could get a higher delta v with this mixture due to it's ludicrous density. I also suspect it'd lower the combustion temperature a great deal, which could allow all of the exhaust product from the combustion chamber to go straight through the turbines. $\endgroup$ – Reuben Farley-Hall Feb 18 at 5:17
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Consider: instead of carrying a third inert propellant mixed in with the oxidizer, you could just carry an excess of fuel. At this point, you're venting most of the gas generator exhaust as rocket exhaust and diverting some to pump the propellants: the system basically reduces to a combustion tap-off cycle engine running extremely fuel-rich.

At that point, optimization would proceed as normal for such an engine...and that optimization process has always ended with much more "traditional" fuel/oxidizer ratios.

Including inert propellants for temperature control might have made sense in the earliest days of rocketry before regenerative cooling techniques were developed (and early fuels included water in large part to keep combustion temperatures down), but technology has developed far past this being a sensible trade today.

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  • $\begingroup$ Temperature control/reduction is a helpful side effect of the idea. The main reason for the idea is to produce large amounts of impulse early and cheaply in the flight where, in conventional rockets, the exhaust velocity >> than the rocket velocity $\endgroup$ – Tobe Feb 17 at 15:53
  • $\begingroup$ Also it seems wasteful using fuel as reaction mass, particularly early in the flight. $\endgroup$ – Tobe Feb 17 at 15:55
  • $\begingroup$ exhaust velocity >> rocket velocity is super inefficient. Maybe it would mean staging at less than 1km per second. $\endgroup$ – Tobe Feb 17 at 15:56
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    $\begingroup$ Purposely reducing your exhaust velocity by adding inert reaction mass may technically be more energy efficient, but all it accomplishes is reducing the total impulse, and that's what really matters in rocketry, not energy efficiency. This is not a useful trade. $\endgroup$ – Christopher James Huff Feb 17 at 17:02
  • $\begingroup$ @ChristopherJamesHuff well, it's not only total impulse that matters. Nitrogen could possibly make sense for getting a Big Dumb Booster cheap and reliable with high TWR. That's normally the domain of solid motors, but a nitrogen-added liquid rocket would burn much cleaner than these. $\endgroup$ – leftaroundabout Apr 30 at 18:53

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