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I cant find anything online, does anyone know what it is?

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Expect the answer to be close to sea-level pressure, as the F-1 is optimized for first stage performance.

The answer can be found easily by using the commonly cited statistics: chamber pressure and nozzle area ratio (expansion ratio).

The isentropic flow equations show that for a given nozzle area ratio there is a fixed supersonic pressure ratio. This ratio is also a function of the ratio of specific heats for the gas. For air, this value is typically taken as 1.4, however for combustion exhaust, this value is typically lower. I will take a value of 1.2 for my analysis (finding a better guess is possible if the O/F ratio, chamber pressure, and propellant formulations are known (which they are, but I am not opening that can of worms for this question)).

Wikipedia says the F-1 has a chamber pressure of 70 bars and an area ratio of 16. This gives a static to total pressure ratio of 0.0068. One more number crunch gives the nozzle exit pressure as 0.48 bars.

This is lower than the sea-level pressure I guessed in the beginning, but typical for first stage engines. 0.48 bar is the pressure at about 25,000 ft altitude, so the first stage likely operates beyond that height.

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A published value for the exit plane pressure can be found in Comprehensive Review of Liquid-Propellant Combustion Instabilities in F-l Engines

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I had done a fun calculation (hence the comments below), but the answer came out way too low.

This is close to the 0.5 bar mentioned in A. McKelvy's answer https://space.stackexchange.com/a/61261/6944

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  • $\begingroup$ Our answers differ by a factor of 10 (unless I made a mistake translating units). Perhaps the precision offered by the "pressure thrust" is not sufficiently high. The pressure thrust is about 10,000 lbf .. ~0.5% of the total thrust. Small uncertainties in the calculation of thrust will result in wild uncertainties in the calculation of exit pressure. I think that is what happened here. Your answer is 0.05 bar, I think this is unphysically low for a first stage engine. $\endgroup$
    – A McKelvy
    Commented Dec 13, 2022 at 20:47
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    $\begingroup$ interesting. If you take the temperature ratio and pull the matching pressure ratio, this gives an exit pressure of 0.08 bar. Much closer to your calculation. But it also is paired with an expansion ratio of 65. There must have been losses or non-ideal processes that lowered the total temp across the nozzle. $\endgroup$
    – A McKelvy
    Commented Dec 13, 2022 at 20:57
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    $\begingroup$ @AMcKelvy I found a published number, it kind of splits the difference. Closer to 0.5 bar though. I'll edit it in. $\endgroup$ Commented Dec 13, 2022 at 21:10
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    $\begingroup$ your sleuthing staggers the mind (as does your fondness for English units). Great find. Sad to see the other spec tables go. $\endgroup$
    – A McKelvy
    Commented Dec 13, 2022 at 21:17
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    $\begingroup$ @AMcKelvy Go here and download the whole glorious treasure trove of F-1 documents. space.stackexchange.com/a/59780/6944 $\endgroup$ Commented Dec 13, 2022 at 21:22

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