Effects of Aspect Ratio in a Transonic Shock Tube Airfoil Flow

作     者：Masashi KASHITANI Keita MIURA Shinichiro NAKAO Yutaka YAMAGUCHI Masashi KASHITANI;Keita MIURA;Shinichiro NAKAO;Yutaka YAMAGUCHI

作者机构：Department of Aerospace Engineering National Defense Academy1-10-20 Hashirimizu Yokosuka 239-8686 JAPAN Department of Mechanical Engineering Nagasaki Institute of Applied Science 536 Aba-MachiNagasaki 851-0193 JAPAN 

出 版 物：《Journal of Thermal Science》 (热科学学报（英文版）)

年 卷 期：2012年第21卷第5期

页      面：435-440页

核心收录：

学科分类：080704[工学-流体机械及工程] 080103[工学-流体力学] 08[工学] 0807[工学-动力工程及工程热物理] 0801[工学-力学（可授工学、理学学位）] 

主　　题：Shock tube Aerodynamics Transonic flows Aspect ratio Wind tunnel correction 

摘      要：In the present study, the flow visualizations were performed around the NACA 0012 models which differ in aspect ratios. We discussed the effects of the aspect ratio in the test models. Additionally the unsteady, two-dimensional, compressible Euler equations were solved for the NACA 0012 airfoil. Experiments were performed utilizing the conventional gas driven shock tube as the intermittent transonic wind tunnel. The aspect ratios of the models are about 0.86 and 1.5, respectively. The Mach numbers M 2 are about 0.84. The Reynolds numbers of the present experimental conditions were constant that Re based on chord length is about 4.0×10 5 . The results are as follows: in different aspect ratios, the difference of the shock wave location is confirmed though the Mach number and Reynolds number are same. It indicates the different correction Mach number by the effects of the side wall boundary layer though the nominal Mach number measured the same value. Also, on the difference of shock wave location for the effects of the aspect ratio, the tend of CFD shows the qualitative agreement with the result of an experiment.