固体火箭冲压发动机一体化燃烧特性数值分析  

Numerical Analysis of Integrated Combustion Characteristics of Solid Rocket Ramjet

作　　者：冯钦 林智 邵博 王纪林 FENG Qin;LIN Zhi;SHAO Bo;WANG Ji-lin(Chinese Flight Test Establishment, Xi'an 710089, China;Sichuan Institute of Aerospace Systems Engineering, Chengdu 610199, China;Southwest Technology and Engineering Research Institute, Chongqing 401329, China)

机构地区：[1]中国飞行试验研究院,西安710089 [2]四川航天系统工程研究所,成都610199 [3]西南技术工程研究所,重庆401329

出　　处：《科学技术与工程》2022年第17期7197-7205,共9页Science Technology and Engineering

摘　　要：为研究固体火箭冲压发动机性能,采用计算流体力学方法对包含进气道及补燃室的一体化燃烧流场进行数值分析,研究可燃燃气进口条件、飞行攻角以及进气道与补燃室过渡连接方案对补燃室掺混燃烧的影响。研究结果表明:燃气流量为0.08 kg/s时,燃气射流出现偏移,补燃室两侧壁面温度相差较大,燃气流量为0.3 kg/s时,燃气偏移现象基本消失;随着燃气流量增大,发动机推力增加;攻角增大使得进气道流量系数增大,强化空气与燃气混合燃烧效果,并最终提升发动机推力。进气道与补燃室的过渡连接方式影响进气角度,通过改变过渡连接方式将进气角度从50°增加至90°后,燃气流量为0.3 kg/s时,发动机推力提高10%,但会导致补燃室总压损失增大,发动机比冲降低2%。In order to study the performance of solid rocket ramjet,the integrated combustion flow field including inlet and afterburner was numerically analyzed by using computational fluid dynamics method.The effects of inlet conditions of combustible gas,flight angle of attack and transition connection scheme between inlet and afterburner on the mixed combustion of afterburner were studied.The results show that when the gas flow rate is 0.08 kg/s,the gas jet deviates,and the wall temperature difference between the two sides of the secondary combustion chamber is large.When the gas flow rate is 0.3 kg/s,the gas offset phenomenon basically disappears.With the increase of gas flow,the engine thrust increases.The increase of the angle of attack increases the flow coefficient of the inlet,enhances the effect of air and gas mixed combustion,and finally improves the engine thrust.The transition connection between the inlet and the afterburner affects the intake angle.After the inlet angle is increased from 50°to 90°by changing the transition connection mode,the engine thrust increases by 10%when the gas flow rate is 0.3 kg/s,but the total pressure loss of the secondary combustion chamber increases and the specific impulse of the engine decreases by 2%.

关 键 词：固体火箭冲压发动机 数值模拟 一体化 补燃室 发动机性能 

分 类 号：V228[航空宇航科学与技术—飞行器设计]