固体火箭发动机喷管喉部瞬态热流的测量技术  

Measuring technique for transient heat flux at inner surface of throat insert of solid rocket motor nozzle

作　　者：李宏顺[1,2] 张俭[1] 谭建伟[1] 杜瑞兵[1] 王刚[1] 张海龙[1] 

机构地区：[1]中国航天科工集团第六研究院41所,内蒙古呼和浩特010010 [2]武汉工程大学理学院,湖北武汉430074

出　　处：《武汉工程大学学报》2011年第12期94-98,共5页Journal of Wuhan Institute of Technology

摘　　要：改进了美国Nanmac公司的自更新快速响应热电偶,并将改进后的热电偶用于测量固体火箭发动机喷管喉衬内壁面上的瞬态温度和热流密度.测量是在一小型试验发动机上进行的,其喷管喉径为18mm.分别采用黑火药、双基推进剂和复合推进剂进行了点火试验,试验结果表明本文的测量方法是可行的.The ＂self renewing＂ fast response thermocouple developed by the Nanmac Corporation, USA, was improved,and employed to measure the transient temperature and heat flux density at the inner surface of the nozzle throat insert of a solid rocket motor. Measurements were conducted with a laboratory scale motor,whose inner diameter of throat insert was 18 mm. Ignition tests were conducted with black powder, double base propellant and composite propellant respectively. The experimental results show that the present measuring technique is feasible.

关 键 词：固体火箭发动机 喷管 传热 测量 

分 类 号：V435[航空宇航科学与技术—航空宇航推进理论与工程]