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Several internet articles claim that the high Isp of the Soviet RD-180 engine (as used in the Atlas III and V launch vehicles) was (at least partially) due to the fact the RD-180 ran oxidizer-rich, unlike its fuel-rich American contemporaries. These articles provided no references to support the assertion of higher efficiency for oxygen-rich combustion.

EDIT: According to Wikipedia, the combustion chambers of the RD-180 ran fuel-rich. It was their LOX pre-burner which ran oxygen-rich.

EDIT:https://ntrs.nasa.gov/search.jsp?R=19820002372 deals with relative advantages of oxidizer- and fuel-rich pre-burners but not combustion chamber mix ratios.

My understanding is that maximum temperature (and combustion efficiency) is attained with a stoichiometric mix of fuel and oxidizer. However, these high temperatures exceed the service temperature of available engine materials. This relationship is illustrated by data for methane in gas turbines:

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https://www.researchgate.net/publication/267643923_DME_as_a_Potential_Alternative_Fuel_for_Gas_Turbines_A_Numerical_Approach_to_Combustion_and_Oxidation_Kinetics/figures?lo=1

As a result, chemical rocket combustion chambers are designed to run fuel- or oxidizer-rich for durability reasons. Either fuel- or oxidizer-rich conditions will lower combustion temperature, but fuel-rich has the added advantage of lower exhaust gas molecular weight and therefore higher exhaust velocity ( Isp).

As a result, I would expect rocket designers to favor fuel-rich conditions to maximize Isp efficiency.

Fuel-rich engines have coking problems. Oxidizer-rich engines are very hard on metal parts. Both these are disadvantages but do not affect theoretical efficiency. Which strategy is objectively more efficient (e.g.: Isp or TWR)? Are there any successful oxidizer-rich engines?

Woody
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  • " for equal turbine temperatures and pump discharge pressures, an oxidizer-rich cycle has an 87% increase in chamber pressure over a fuel-rich cycle - imply that a smaller oxidizer-rich engine could produce the same thrust as a larger fuel-rich engine." look at Organic Marble's answer https://space.stackexchange.com/a/22950/40489 – blobbymcblobby Oct 10 '23 at 16:14
  • @blobbymcblobby ... unfortunately, Organic Marble's cited link is dead. – Woody Oct 10 '23 at 16:19
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    The RD-180 does not run oxygen-rich. It runs at a O:F ratio of 2.72:1, stoichiometric would be something like 3.4:1 (going by the C[n]H[1.953n] formula from braeunig.us). The "OR" in ORSC refers specifically to the preburners. – Christopher James Huff Oct 10 '23 at 16:19
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    Does this answer your question? What are the advantages of Ox-rich staged combustion? – Russell Borogove Oct 10 '23 at 16:47
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    maximum temperature (and combustion efficiency) is attained with a stoichiometric mix of fuel and oxidizer

    Maximum specific impulse is generally achieved with a fuel-rich mix, because it leaves simpler molecules in the exhaust which hide less energy in interatomic bonds than complex molecules.

     – Russell Borogove Oct 10 '23 at 16:49
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    Thanks for the heads-up! The link in that answer has been fixed. – Organic Marble Oct 10 '23 at 17:11
  • Just for an overview of the Soviet side, this is a nicely illustrated page: https://everydayastronaut.com/soviet-rocket-engines/ – blobbymcblobby Oct 10 '23 at 17:31
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    I seem to remember something about oxy-rich exhaust having a tendency to chew on engine parts or something... is that an issue or am I misremembering/misinformed? – Darth Pseudonym Oct 10 '23 at 19:51
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    @DarthPseudonym you are exactly right. Hot oxygen makes pretty much everything want to burn. The Soviets worked some metallurgical magic to get their engines to work. – Organic Marble Oct 10 '23 at 22:00
  • @ChristopherJamesHuff ... according to https://en.wikipedia.org/wiki/RP-1 , the stoichiometric ratio for LOX/RP1 is 2.56;1 (not 3.4:1) , but you are correct that the RD-180 combustion chambers ran rich at 2.72:1. – Woody Oct 11 '23 at 16:19
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    Soviets also favored hypergols too, running Proton until recently, with 400+ launches. Kerolox development started in the 50s and with early successes, and Soviet methodology (afterall, they still use a variant of the R-7) - ie. continue in the same vein until it doesn't work anymore (look at most of the programs considered successful - iteration after iteration), it seemed natural (for them) that they took this development path. – blobbymcblobby Oct 11 '23 at 17:03
  • @Woody that page doesn't even mention stoichiometric ratios. – Christopher James Huff Oct 11 '23 at 20:38
  • @ChristopherJamesHuff ... my mistake. I quoted the mass oxygen:fuel ratio which is the actual operational ratio. The stoichiometric ratio is the theoretical optimal equivalent. – Woody Oct 11 '23 at 22:33

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Regardless of whether oxidizer-rich engines would be more efficient, the RD-180's efficiency is not due to that, as the RD-180 runs fuel-rich. According to braeunig.us, for each carbon atom in RP-1 there are 1.953 hydrogen atoms, so a stoichiometric ratio would be something like 3.4:1. The RD-180 runs at a O:F ratio of 2.72:1, definitely on the fuel-rich side of stoichiometric.

What distinguishes the RD-180 and similar Russian engines is not that they run oxidizer rich (as they don't), but that they are staged combustion engines. This does several things, but the most significant is that the working fluid used to pump propellant is injected into the combustion chamber to be combusted further and used as reaction mass, so much higher pumping powers can be used, allowing higher chamber pressures than gas generator engines. Their preburners run oxidizer rich, which allows them to do staged combustion with carbon-rich fuels that would otherwise cause coking that would clog the turbine passages and engine injectors.

At the time these engines were developed, the only staged combustion engines developed in the US used fuel-rich preburners with hydrogen fuel which didn't produce coking, and the only kerosene-burning engines used gas-generator or other cycles that were better able to handle the coking issues. Being able to use staged combustion gave kerolox engines like the RD-180 a major performance edge over other kerolox engines, while retaining the thrust and ease of handling advantages of kerosene fuel.

Christopher James Huff
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  • Technically "coking" happens in the liquid fuel and "sooting" happens in the combustion products, but your answer covers it well. https://space.stackexchange.com/a/59099/6944 https://apps.dtic.mil/sti/citations/ADA410860 https://archive.org/download/engine-cycles-2/EngineCycles-2.pdf https://ntrs.nasa.gov/citations/19820002372 – Organic Marble Oct 11 '23 at 23:47
  • @OrganicMarble that's usually the case, but the mixture ratio in a FRSC preburner is far more fuel-rich than that in gas-generator engines. Most of the fuel wouldn't really participate in combustion beyond being heated by it. In a kerolox FRSC, it would probably actually be coking (thermal decomposition and polymerization) instead of sooting (incomplete combustion resulting in carbon deposits). That said, I'm not aware of such a thing being any more than a hypothetical possibility, or maybe a (probably short-lived) laboratory experiment. – Christopher James Huff Oct 12 '23 at 03:08
  • Is a gas generator cycle limited to the achievable chamber pressure, at least in a practical sense to the 1000psi-class engines (F-1, RD-107)? i.e. do "high chamber pressure" and "staged combustion" almost necessarily go together? – Daniel Chisholm Oct 12 '23 at 12:12
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    @DanielChisholm See https://en.wikipedia.org/wiki/Comparison_of_orbital_rocket_engines. In a gas generator cycle, increasing the chamber pressure requires increasing the amount of propellant going to the gas generator + turbine used to pump the propellant. That propellant then gets dumped overboard without producing significant thrust. There is an optimum point where increasing the chamber pressure further reduces effective overall performance. Staged combustion engines don't have this issue. – Christopher James Huff Oct 12 '23 at 13:05