氢燃料超音速燃烧室实验研究  

AN EXPERIMENTAL INVESTIGATION OF HYDROGEN-AIR SUPERSONIC COMBUSTION

作　　者：刘陵[1] 张榛[2] 唐明[2] 刘敬华[3] 杨丽新[3] 王兴甫 

机构地区：[1]西安西北工业大学,东118710072 [2]西北工业大学 [3]航空航天部三十一研究所

出　　处：《航空动力学报》1991年第3期267-270,共4页Journal of Aerospace Power

基　　金：国家自然科学基金资助项目

摘　　要：本文介绍超音速燃烧冲压发动机燃烧室实验研究 ,模型燃烧室呈突扩台阶和扩张形 ,用电弧加热空气。燃烧室入口Ma =2 1,总温 12 0 0K ,总压 7× 10 5Pa。氢气由壁面沿垂直或平行于超音速空气流的方向喷入 ,实现了超音速稳定燃烧 。The experimental results of Hydrogen fueled supersonic combustors are considered.Tests were conducted with arc heated air at inlet Mach number 2 1,total temperature 2100K,total pressure 7×10 5Pa.The axisymmetric combustor with inlet of 4 6cm inner diameter has been investigated with hydrogen injected normally or parallelly to air stream from a ring of equally spaced holes on the wall.The combustor comprised assemblies of cylindrical and diverging conical sections on which are mounted static pressure taps and thermocouples.The total pressure profiles were measured by a pressue probe rake just downstream of the combustor exit.Each fuel injection persisted about 8 seconds to obtain steady state operiating conditions.The preudo one dimensional flow analysis of supersonic combustor has been improved and employed to yield a description of the bulk processes in the supersonic combustor and to explain the characters of the measured pressure distributions.The results of a series of tests show that the combution efficiency of the transverse injection ahead of a rearward facing step is higher than behind it,and the combustion efficiency of transverse injection is higher than that of the downstream injection,and it is also evident that the total pressure loss is maximum in the case of transverse injection ahead of a rearward facing step and is minimum in the case of parallel injection.

关 键 词：燃烧室 流体损失 试验 航空发动机 

分 类 号：V235.113[航空宇航科学与技术—航空宇航推进理论与工程]