Internal flow field ablation is an important issue in thermal protection materials for rocket engines and hypersonic vehicles. In this paper, a new multiphase flow ablation test system, with an Al₂O₃ particle delivery device based on an oxygen-kerosene liquid rocket engine, is designed and manufactured. A general variable-precision modular system simulation method is proposed to analyze the dynamic characteristics of the system. In addition, a unique internal flow field ablation test was performed on the 4D C/C composite simulating the working conditions of the SRM. The results show that the system can provide a wide temperature range (756~3565 K) and pressure range (0.2~4.2 MPa). The multi-disciplinary dynamic variable-precision system simulation method is helpful for more accurate design and test analysis, and the maximum error is less than 5%. The ablation tests show that the line ablation rate of the C/C composite nozzle at 3380 K and 1 MPa is 0.053 mm/s, verifying the combined effect of thermochemical and mechanical ablation. The ablation environment is controllable, which provides an effective way for the ablation test of thermal protection materials. In addition, the variable-precision dynamic simulation method has important reference value for the system design related to liquid rocket engine.

In order to explore the ablation law and characteristics of carbon/carbon composite materials on the surface thermal protection of hypersonic vehicles under the condition of particle erosion, the ablation test of a 4D C/C composite was carried out based on a small plasma simulated ablation test system. Four sets of experiments were designed according to different temperature and particle concentration. After the test, two kinds of ablation rate were measured, and the microscopic ablation morphology was observed by scanning electron microscope (SEM) and the ablation behavior was analyzed. The results show that, under the equivalent test conditions, the rate of jet temperature increases, and the material ablation rate increases. With the increase of particle concentration, the two kinds of ablation rate of materials increased, the maximum mass ablation rate was 0.0363 gs⁻¹, and the line ablation rate was 0.053 mms⁻¹. The temperature affects the thermal erosion rate of the material, while the particle mainly causes mechanical erosion. The particle has a coupling effect of heat transfer and mechanical destruction on the material, and the material ablation is intensified. Moreover, due to the strong corrosion resistance of carbon fiber, the overall corrosion resistance of materials has obvious anisotropy.

For the research demand of reusable LOX/kerosene variable thrust liquid rocket engine, a test system with electric displacement pumps is designed and a multidisciplinary modular dynamic simulation method based on AMESim platform is used to analyze the system. The method comprehensively considers the characteristics of complex components in the engine and realizes the fast module assembly and variable step size solution. Considering the combustion model of thrust chamber, the positive displacement pump model with complex leakage channels, and the cooling jacket heat transfer model, the component dynamic equations are deduced and the final model simulation results reveal that the system has a smooth ignition, stage turning, and shutdown process. The thrust can reach 6900 N in high working condition and the variable thrust ratio is 5 : 1. The dynamic characteristics of the system show that the performance error of main pump components is less than 5%, the maximum average temperature rise of the thrust chamber coolant is about 28°C, and the time of stage adjustment is within 300 ms, which mean the overall design of the system is reasonable. Although the accumulation of LOX before kerosene injection can adversely affect the temperature of the thrust chamber, large pressure pulses do not occur due to the ignitor’s duty flame. Moreover, the pintle injectors based on PID control can effectively stabilize the pressure drop at lower conditions. The system and the simulation method provide important support for the actual engine test and the general LRE dynamic characteristics analysis.

During the development of thermal protection materials for hypersonic vehicles, the simulated ablation test based on thermal plasma has been widely used. In this paper, a small plasma multi-phase flow simulation ablation test system is designed based on the automatic control technology. After the overall design, hardware processing and software development are completed. Then the system parameters are calculated and tested by numerical calculation methods and heat flow densitometers as well as other test instruments. The results show that the system has high enthalpy, high temperature and high heat flux, and solid particles can be injected in it to simulate particle erosion effects. The system can generate plasma jet with continuously adjustable parameters. The stagnation point temperature is 11651 K and the central speed is 1222 m•s⁻¹. The particle addition rate is continuously and adjustable. The maximum speed in the experiment is 280 m•s⁻¹, and the surface temperature is about 2650°C. The work in this paper provides an important way for more inexpensive and efficient ablation tests of thermal protection material. The results provide a useful reference for the selection of the relevant material's ablation test parameters.

The performance of solid rocket ramjets is affected by the ablation resistance of silicone rubber composites in the secondary combustion chamber. To investigate the ablation mechanisms of silicone rubber composites in multiphase flow environments, a new ablation testing methodology, in which particles can be added to the flame, is developed in this investigation. Boron oxide (B2O3) and alumina (Al2O3) particles were used to simulate condensed-phase particles. The experimental results show that because of the high hardness and melting point of Al2O3, the mechanical destruction of the Al2O3 particles was so intense that the surface of silicone rubber composite was completely destroyed, and the matrix was exposed to the surface. The surfaces of the B2O3 particles melted and evaporated in the flame, forming a liquid layer, similar to condensed-phase particles in the secondary combustion chamber of solid ramjet. After the ablation of silicone rubber under the erosion action of B2O3 particles, a char layer and pyrolysis layer were formed, similar to the scenario during the actual ablation of silicone rubber composites. The multiphase flow environment with B2O3 particles can truly reflect the ablative environment experienced by silicone rubber composites.