高压捕获翼构型亚跨超流动特性数值研究

为研究高压捕获翼布局在亚跨超条件下的流动特性, 选取圆锥?圆台机体组合捕获翼概念构型, 在马赫数0.3 ~ 3速域范围内, 选取典型状态点, 采用数值模拟在 0°攻角条件下进行了计算和分析. 结果表明, 在整个速域范围内, 由于机体与捕获翼在对称面附近的垂向距离最小, 因此二者之间的气动干扰最为明显, 且沿展向逐渐减弱. 同时, 随马赫数增大, 机体与捕获翼间的流场结构明显不同, 具体表现为: 当Ma<0.5时, 未出现流动分离现象, 当Ma>0.5时, 机体后段开始出现明显的流动分离, 由于捕获翼与机体形成先收缩后扩张的等效通道, 捕获翼下表面和机体上表面的压力均先减小后增大; 进入跨声速速域后, 在捕获翼的影响下, 流动分离更加明显, 机体与捕获翼之间开始出现激波, 并且与分离区相互作用, 同时出现激波串, 捕获翼下表面产生明显的压力波动现象, Ma=1.5时, 通道内激波位置基本到达机体尾部, 分离区基本消失; 当Ma>2以后, 整个流场呈现以激波为主导的结构形式, 捕获翼下表面和机体上表面的压力分布逐渐趋于平缓. 

NUMERICAL STUDY ON FLOW CHARACTERISTICS OF HIGH-PRESSURE CAPTURING WING CONFIGURATION AT SUBSONIC,TRANSONIC AND SUPERSONIC REGIME

In order to study the flow characteristics of the high-pressure capturing wing (HCW) configuration at subsonic, transonic and supersonic regime, the conceptual configuration of the conical-cone airframe combined HCW was selected, and the typical state points were selected in the range of Mach number 0.3 to 3. The numerical simulation and analysis were carried out under zero angle of attack condition. The results show that in the speed range studied, since the vertical distance between the airframe and HCW at the symmetry plane was the smallest, the aerodynamic interference was the most obvious and gradually weakened along the span. As the Mach number increased, the flow field structure between the airframe and HCW was obviously different. The specific performance were as follows: when the Mach number was less than 0.5, no flow separation occurred, and when the Mach number was greater than 0.5, the obvious flow separation in the back section of the airframe began to appear. Since HCW and the airframe formed an equivalent channel that first contracted and then expanded, the pressure on the lower surface of HCW and the upper surface of fuselage both decreased first and then increased. After entering the transonic speed domain, under the influence of the HCW, the flow separation became more obvious. The shock wave began to appear between the airframe and HCW, and interacted with the separation zone, and a shock wave train appeared which resulted in significant pressure fluctuations on the lower surface of HCW. When the Mach number was 1.5, the shock wave position between the airframe and HCW reached the tail of the airframe, and separation zone almost disappeared. When Mach number continues to increase, the entire flow field presented a shock-dominated structure and the pressure distribution on the lower surface of HCW and the upper surface of fuselage gradually became flat.