机动飞行条件下固冲发动机补燃室绝热层烧蚀实验研究

The overload effect caused by maneuvering flight can result in the accumulation of combustion gases within afterburning chamber of solid rocket ramjet， leading to a serious erosion area in the thermal insulation layer. To address this issue， a test device is designed and established to simulate the thermal environment of the afterburning chamber during maneuvering flight. Numerical simulation is employed to obtain variations in two-phase flow field parameters. Controllability is achieved for ablation parameters such as combustor pressure， surface temperature of the thermal insulation layer， gas velocity， and angle between the gas and thermal insulation layer. An experimental investigation is conducted on the maneuvering flight conditions of a certain type of solid rocket ramjet using the device， the results demonstrate that the experimental findings satisfy the simulation requirements for serious ablation area in the afterburning chamber. The target pressure is 0.45 MPa， and the measured pressure is 0.467 MPa， resulting in a relative error of 3.78%. Additionally， temperature tests during the experiment yielded values of 1 245 K and 1 341 K， exhibiting relative errors of 5.62% and 5.14% compared to numerical simulation results， respectively. Furthermore， the center temperature of the test specimen is 2 367 K， which corresponds to a relative error of 2.9% when compared to the target temperature. The experimental thermal insulation layer exhibits erosion craters in the direct action area of the gas， with a maximum linear erosion rate of approximately 0.03 mm/s. In the indirect action area， specimen expands， and the crater shapes were influenced by the distribution pattern of the flow field near the specimen.