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Abstract When a solid rocket engine is ignited, the throat lining of the nozzle is prone to chemical ablation owing to high-temperature gas erosion, resulting in thrust loss. In this paper, a coupled fluid–solid model for thermochemical ablation on the nozzle wall is established based on the multi-component Navier–Stokes equations, SST k - ω turbulence model, finite-rate chemical reaction model on the nozzle wall, variable transport properties of the nozzle material, and the heat conduction equation. Compared with the experimental data, the maximum error of the calculated ablation rate was 4.37%, validating the effectiveness of the model. Subsequently, the effects of different combustion chamber components, pressures, and temperatures on the ablation rate of the carbon–carbon (C/C) throat lining were studied. The results indicate that the temperature at the nozzle throat was the highest, resulting in the maximum ablation rate. As the Al mass fraction at the nozzle inlet increased, the thermochemical ablation rate of the nozzle decreased with a lower oxidizer mass fraction. The inlet pressure and temperature of the nozzle were positively correlated with the ablation rate, with the temperature having a more significant impact than the pressure. These findings provide theoretical guidance for the thermal protection design of rocket engine nozzles.
Wang et al. (Wed,) studied this question.