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在实验研究的基础上,提出固体火箭发动机液体喷射熄火模型。该模型综合考虑了固体推进剂的瞬态燃烧、射流换热、液滴蒸发和发动机内弹道等耦合作用,成功地实现了对液体喷射熄火过程(临界参数和熄火用液量)的理论预示。理论研究发现液体喷射瞬变燃烧存在着临界喷射压降。当喷射压降大于该临界值时,熄火才能实现。随着推进剂能量的升高,临界喷射压降增加。随着喷射压降的增加,熄火用液量和降压速率分别下降和升高,其变化率逐渐减小。熄火用液量不存在最小值,因而在工程设计中,必须合理选择喷液量和喷射压强这两个设计参数。理论预示与实验结果吻合良好。
Based on the experimental study, the liquid jet flameout model of the solid rocket motor was proposed. The model comprehensively considered the transient propulsion of solid propellant, the heat transfer in jet, the droplet evaporation and the in-vehicle ballistic coupling. The theoretical predictions of liquid injection flameout (critical parameters and liquidus) were successfully achieved. Theoretical studies have found that there is a critical jet pressure drop in transient combustion of liquid jet. When the jet pressure drop is greater than the critical value, flameout can be achieved. As the propellant energy increases, the critical injection pressure drop increases. With the increase of injection pressure drop, the amount of fluid extinction and the depressurization rate decreased and increased, respectively, and the rate of change decreased gradually. Extinguishing liquid volume does not exist minimum, so in engineering design, we must choose the appropriate spray volume and injection pressure of these two design parameters. The theoretical predictions are in good agreement with the experimental results.