摘要
利用FLUENT发展了一种瞬态流-固-热一体化一步耦合三维有限元分析方法,并用该方法分析了一种采用无铝PET推进剂的小推力长工作时间(130 s)固体火箭发动机喷管温度场分布情况。分析结果表明,从喷管中心到喷管表面温度逐渐降低,喷管喉部中心温度达到2200℃以上,喉衬温度为1481~2000℃,隔热层温度为134~1864℃,壳体温度为134~234℃;喷管中心轴线上各点的温度开始时急剧增加,20 s后基本保持不变,从喷管出口到喷管入口,沿喷管中心轴线温度逐渐增加;喷管壳体表面温度随工作时间增加而增加,随离喷管出口的距离增加先增加而后逐渐降低;喷管喉部壳体表面温度最高,达220℃左右;推进剂燃温增加,喷管金属壳体表面温度呈线性增加;换热系数增加,金属壳体表面温度近似线性降低;燃烧室压强变化对金属壳体表面温度影响不大;隔热材料EPDM、5-Ⅲ和GFM降低金属壳体表面温度的效果突出。选用合适的隔热材料,做好喷管喉部位置的热防护是降低喷管壳体表面温度的有效方法。发动机试车结果与仿真结果吻合,验证了仿真分析方法的正确性。该瞬态流-固-热耦合一体化一步分析法,克服了过去流-固-热耦合方法步骤多、绝热假设不符合实际情况的局限性,可为合理设计固体发动机喷管、控制喷管温度提供依据。
By using Fluent three⁃dimension finite element software,a new transient flow⁃structure⁃thermal coupling integration one⁃step simulation method was developed,and the nozzle temperature of a low⁃thrust and long⁃combustion⁃time(130 s)solid rocket motor(SRM)with alminum⁃free PET propellant,was simulated by this method.The analysis results show that the temperature decrea⁃ses from center to surface of nozzle.The temperature at the center of nozzle throat reaches more than 2200℃,the temperature on the throat insert is 1481~2000℃,the temperature of thermal insulated layer is 134~1864℃,the temperature of the case is 134~234℃.The temperature along the axis of the nozzle increases fast at the beginning,and remains almost constant after 20 s.The temperature along the axis of the nozzle gradually increases from the exit to the inlet of nozzle.The temperature on the surface of nozzle increases with the increase of combusting time,and increases at first and then decreases with the increases of the distance from the nozzle out⁃let.The temperature on the metal case surface of nozzle throat is the highest,which reaches about 220℃.The temperature on the metal case surface increases linearly with the increases of the combustion temperature of propellant,and decreases opproximately lin⁃early with the increases of heat transfer coefficient.The change of chamber pressure has little influence on the temperature on the metal case surface.The thermal insulation materials EPDM,5⁃III and GFM have outstanding effect on the reducing the surface tem⁃perature on metal case.It is an effective method to reduce the surface temperature of nozzle case by selecting the appropriate insula⁃tion material and conducting a good thermal protecion of the nozzle throat.The SRM test sesults are in good agreement with the simu⁃lation results,which verifies the correctness of the simulation analysis method.The transient fluid⁃solid⁃thermal coupling integrated one⁃step analysis method overcomes the limitations of th
作者
邓康清
王鹍鹏
余小波
向进
朱雯娟
王相宇
杨育文
汤亮
张琪敏
DENG Kangqing;WANG Kunpeng;YU Xiaobo;XIANG Jin;ZHU Wenjuan;WANG Xiangyu;YANG Yuwen;TANG Liang;ZHANG Qimin(Key Laboratory of Aerospace Chemical Power Technology,Xiangyang441003,China;Hubei Institute of Aerospace Chemistry Technology,Xiangyang441003,China)
出处
《固体火箭技术》
CAS
CSCD
北大核心
2023年第4期507-513,共7页
Journal of Solid Rocket Technology
关键词
固体火箭发动机
喷管
长工作时间
流-热-固耦合
温度场分布
solid rocket motor
nozzle
long⁃combustion⁃time
flow⁃structure⁃thermal coupling
temperature field distribution
