固体火箭发动机喷管喉衬需要承受高温、高压燃气的烧蚀、冲刷,其喉径变化率在整个工作过程中极有可能是非线性的,目前国内外在喷管烧蚀方面研究主要集中在数值模拟分析及喷管本体烧蚀试验方面,缺乏对喷管烧蚀过程与发动机总体性能之间联系的研究。文中基于差分进化算法和径向基函数神经网络,建立了一种基于固体火箭发动机试验测试曲线的发动机喷管喉径烧蚀率计算方法。基于固体火箭发动机地面静止试验结果,获得了发动机喷管喉径烧蚀率随发动机工作时间的变化曲线。研究结果表明,发动机喷管喉径烧蚀率在发动机工作前期较小,随着发动机工作时间增长而逐渐增大,喉径在发动机工作过程中会逐渐升温,从而导致喉衬烧蚀率逐渐增大,在一段时间后喉衬温度变化趋于稳定,喉衬烧蚀率保持在0.2 mm/s,而随着喉衬表层烧蚀完毕后底层漏出,烧蚀率又会逐渐升高。

Solid rocket engine nozzle throat liner needs to withstand the ablation, erosion of high temperature and high-pressure gas and it nozzle throat diameter ablation rate has a high probability being non-linear during the whole working process. At present, research about nozzle ablation both domestic and abroad is mainly focused on numerical simulation analysis and ablation test of the nozzle body, however, there's a lack of investigation on the link between the nozzle ablation process and the function of solid rocket motor. This paper is based on the differential evolution algorithm and radial basis function neural network, and established a method for calculating the ablation rate of the nozzle throat diameter based on the test curve of solid rocket motor. Based on the ground static test results of solid rocket motor, the curve of nozzle throat diameter ablation rate changing with motorworking time was obtained. The results showed that the ablation rate of the nozzle throat diameter was small in the early stage of the motor operation, and gradually increased with the increase of the working time, as the temperature of the nozzle throat increased, this caused the ablation rate increased, after a period of time, the temperature of the throat liner became constant, and the nozzle ablation rabe is maintained at 0.2 mm/s. And after the ablation of the surface of throat liner was fully completed, the inner surface was exposed; the ablation rate would increase again.