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.

With an increase in rocket motor thrust, the thermal loads on the launch platform by the impingement exhaust jet are significantly increased, this poses serious safety risks to rocket launch. Therefore, a water cooling system should be designed to improve the thermal environment. In this study, the cooling effects of a water spray system with different injection velocities were analyzed through numerical simulation of high precision, and the afterburning effect of the exhaust gas was also considered to improve the calculation accuracy. The results showed that afterburning primarily occurred in the mixing layer, leading to a larger high-temperature region in the impingement plate. The water cooling system realized thermal protection for the launch platform, which was achieved by decreasing the temperature and velocity of impingement jet. With an increase in the aqueous jet velocity, a more significant cooling effect of the water spray system was obtained. However, the guiding performance of the deflector was decreased owing to the large amount of water vapor formed in the channel. It is suggested that the water spray system design should adopt an aqueous jet velocity of 20 m/s as a parameter after considering the cooling effect and guiding performance of the deflector.

The liquid rocket propellant burns inadequately in the combustion chamber; therefore, the exhaust gas reacts with oxygen and the heat that was generated, which in turn affects the thermal environment. By using finite-rate chemical kinetics and discrete ordinates method, the calculation model for the four-engine rocket exhaust flowfield is established, and numerically investigates the afterburning effect on the thermal environment of the engine exhaust plume under different altitudes. The results show that the afterburning reactions have little influence on the Mach number exhaust flow, and the difference between the peak temperatures of reacting and non-reacting flows is less than 5%. However, a significant increase in temperature of the mixing layer can be found, and the afterburning effect on the thermal environment reduces with increases in the flight altitude. At the same altitude, the temperature increment of the reacting flow in the far-field region is higher than that of the near-field region. Moreover, as the flight altitude increases from 15 to 35 km, the increase of the rocket base heat flux caused by the afterburning can vary from 9.4% to 2.7%. The results provide a theoretical basis for the designing of the thermal protection systems of multi-engine rockets.

High-energy gel propellants applied to scramjet engines can effectively enhance combustion chamber efficiency and overall performance, which has a favorable application prospect in the hypersonic vehicle field. In this paper, numerical investigations are carried out on the combustion process of the boron-loaded gel fuel scramjet. Turbulent flow is simulated based on the standard k–ω model, and the motion of boron particles is solved using the discrete phase model. A surface reaction model is used to calculate the combustion process of the boron particles and gel. The comparison between numerical results and experimental data confirms the validity of the multiphase-coupled combustion model for a boron-dodecane gel scramjet engine. Based on this comparison, the variation patterns of engine combustion efficiency, temperature rise efficiency, cavity performance, and mass specific impulse under different boron particle diameters are investigated. The study shows that with increase in the boron particle diameter, the combustion chamber and temperature rise efficiencies initially increase and then decrease, whereas the specific impulse and total pressure recovery coefficients generally show a gradually decreasing trend. When the boron particle diameter is 45 μm, the temperature rise efficiency, combustion efficiency, and mass exchange rate reach their maximum values of 70.19%, 67.22%, and 34.60%, respectively. In summary, these research results provide a reference for the design optimization of boron gel scramjet engines.

During engine operation, the nozzle throat insert components are subjected to high temperatures, pressures, velocities, and erosive gas flow, leading to changes in the nozzle throat diameter, which in turn affect the engine thrust performance. This paper elaborates on the erosion behavior and the mechanism of solid rocket motor throat inserts from three aspects: theoretical study, experimental study, and numerical simulation. In the field of theoretical research, the main focus is on studying empirical formulas, theoretical ablation models, and improved methods for predicting ablation. These studies involve summarizing the characteristics of ablation and proposing simplified ablation models; in terms of the experimental study, the common ablation experimental methods are summarized and the principles, advantages and disadvantages of several types of typical ablation experimental methods are analyzed. In the numerical simulation, more comprehensive ablation models and improved ablation methods are proposed based on the theoretical and experimental contents. Finally, some suggestions are made for the following ablation researches to provide reference and references for future throat insert ablation researches.