凹腔前缘角对超声速燃烧室性能的影响  

Effect of cavity leading edge angle on performance of super-sonic combustor

作　　者：贾真[1] 吴迪[1] 朴英[1] 薛梅新[1] 

机构地区：[1]清华大学航天航空学院,北京100084

出　　处：《航空动力学报》2012年第5期993-998,共6页Journal of Aerospace Power

摘　　要：针对带有不同前缘角的凹腔内流动和燃烧过程,分别在冷态和燃烧条件下探讨了前缘角对凹腔内流动损失及阻力特性的影响.研究表明:在壁面垂直喷射的喷口上游和凹腔内部均会形成低速、高温回流区,有利于点火及火焰稳定,燃烧反压通过边界层的亚声速区域上传,形成激波/边界层干扰结构.减小前缘角,可使剪切层分离位置提前,更偏向凹腔内部,导致凹腔后壁面再附激波增强,进而增大了总压损失,降低了总压恢复系数;亦可导致凹腔前、后壁面压差阻力增大,阻力系数上升.进一步认识了凹腔内部流场及稳焰增混机理,进而为优化凹腔结构设计提供依据.The effect of cavity leading edge angle on flow field and cavity drag characteristics in a supersonic combustor under the condition with/without chemical reactions was numerically investigated.It is found that low velocity and high temperature recirculation zones form in cavity and upstream boundary layer of wall orifice,respectively,which are induced by a bow shock produced by the normal fuel jet interacting with the supersonic cross flow and expected to facilitate the fuel ignition and enhance the flame stabilization.A shock/boundary layer interaction phenomenon arises from an intense back pressure fluctuation induced by combustion spreading upstream through subsonic region of boundary layer.Decreasing the leading edge angle moves separation point of free shear layer ahead,which strengthens the reattaching shock at angled back wall and consequently increases the total pressure loss,in other words,decreases the total pressure recovery coefficient as well as the drag coefficient.This investigation provides insight into the characteristics of flow field inside supersonic combustor with/without combustion and corresponding mechanism of fuel-air mixing enhancement,which is crucial for optimization of the cavity structure.

关 键 词：凹腔 前缘角 激波/边界层干扰 总压恢复系数 阻力系数 

分 类 号：V231.2[航空宇航科学与技术—航空宇航推进理论与工程]