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The LANTR entry at Encyclopedia Astronautica describes how a sample nuclear thermal rocket performs at various mixing ratios of LH2/LOX.

It does however stop before reaching the stoichiometric ratio of 8:1. Burning hydrogen in this ratio usually doesn't make sense, as it lowers the specific impulse by not having reacted hydrogen in the exhaust. But in some cases, where your hydrogen supply line is very long, and your oxygen supply line comparatively very short, it does make sense to cram as much oxygen into the propellant tanks. It also likely sets an upper limit for NTR T/W ratio.

Regression suggests that the specific impulse should be comparable to a chemical LH2/LOX (non-stoichiometric) engine, and the thrust only slightly higher than for lower mixing ratios. But there's not enough data, and regression is not to be trusted so close to a fundamental chemical limit.

How does a nuclear thermal rocket perform with an oxygen afterburner at a mixing ratio of 8:1 completely combusting the hydrogen?

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    $\begingroup$ The table in the middle of that article is awful. What is labeled as "$\text{LOX}/\text{LO}_2$ ratio" should be "$\text{LOX}/\text{LH}_2$ ratio". What is labeled as "Tankage mass ratio" is ... confusing. Apparently it's the percentage of the total dry mass of the vehicle that is the tanks (sans propellant) themselves. Hydrogen tanks are massive. $\endgroup$ – David Hammen Sep 28 at 22:01
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    $\begingroup$ Regarding the question at hand, rocket engines rarely burn at the supposed ideal of a stoichiometric ratio. In this case, the ideal mixing ratio (the ratio that maximizes payload mass for a given amount of propellant) is nowhere close to stoichiometric. $\endgroup$ – David Hammen Sep 28 at 22:05

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