Makoto Mizoguchi

Effect of Aspect Ratio on Aerodynamic Characteristics at Low Reynolds Numbers

The effect of the aspect ratio on the aerodynamic characteristics of rectangular wings is experimentally investigated at Reynolds numbers on the order of 104. The aerodynamic coefficients are analyzed based on visualized flowfields and theoretical results. The characteristics of high-aspect-ratio wings are shown to be affected by leading-edge separation bubbles at low Reynolds numbers. However, the qualitative characteristics are similar when the aspect ratio is 3.0 and above. The characteristics of wings with aspect ratios below 3.0 differ significantly from those of wings with large aspect ratios. In particular, a significant influence is revealed for aspect ratios of 1.0 and below. The effect of the aspect ratio is considered to be caused by wingtip vortices. The discussion also includes the effects of the Reynolds number, the shape of the leading edge, and the thickness ratio.

Effect of Aspect Ratio on Aerodynamic Characteristics of Rectangular Wings at Low Reynolds Numbers

The effect of aspect ratio on the aerodynamic characteristics of wings is investigated at Reynolds numbers on the order of 10. The characteristics of rectangular wings are examined experimentally. Aerodynamic coefficients are analyzed with visualized flow fields and analytical results. The characteristics of high-aspect-ratio wings are shown to be affected by laminar separation bubble at low Reynolds numbers. However, qualitative characteristics are similar when aspect ratio is 3.0 and above. The characteristics of wings below an aspect ratio of 3.0 quite differ from those with large aspect ratios. In particular, the significant influence of aspect ratio reveals for aspect ratios of 1.0 and below. These effects of low-aspect-ratio are caused by wingtip vortices. The discussion also includes the effects of Reynolds number, the shape of leading edge, and thickness ratio.

Static Stall Hysteresis of Low-Aspect-Ratio Wings

The aerodynamic characteristics of rectangular, low-aspect-ratio wings are experimentally investigated at Reynolds numbers ranging from 1.0× 10 to 1.5× 10. The effects of the thickness ratio, the Reynolds number, and the turbulence intensity on the hysteresis loop are analyzed to reveal the occurrence conditions of static stall hysteresis. The hysteresis at stall is pronounced as the thickness ratio decreases at low Reynolds numbers. The hysteresis disappears when the wing has a relatively large thickness. The Reynolds number also affects the stall angle and the span of the hysteresis region. The hysteresis loop enlarges with a decrease in the turbulence intensity, whereas the aerodynamic performance is nearly unchanged.

A Study on Increase of Duration of a Free Piston Shock Tunnel

In the present study, performance of a free-piston shock tunnel driver section, which is equipped with a convergent-divergent diaphragm section, has been investigated analytically, experimentally and numerically. Measurements of shock Mach number, pressure and holding time of the constant pressure in stagnation region have been conducted under the conditions of non-over and over drive operation. Experimental results show qualitatively good agreements with analytical study. And the results show that a disadvantage by using the convergent-divergent diaphragm section has been improved by applying the over drive operation, and duration of constant pressure in the stagnation region is increased. Axisymmetric full Navier-Stokes equations have been solved to study the starting process of the driver section. Numerical results show that vortex and secondary shock wave are generated at the inlet of convergent-divergent diaphragm section, which was not assumed in the analytical study. And the quantitative differences between experimental results and analytical ones are shown to be caused by the complicated flow field. Consequently, investigation of the flow field and increase of duration by use of the convergentdivergent diaphragm section for the small size free piston shock tunnel have been achieved.

Glow Discharge Visualization of Hypersonic Separated Flow past Cylinder/Plate Juncture

In this study, we will try to apply a glow discharge technique to examine the vortical structure around a circular cylinder placed perpendicular to a flat plate in a hypersonic flow. Our purpose is twofold: to demonstrate the capability of the glow technique and to extend our understanding of the flowfield up to hypersonic regime. The experiment is conducted in a gun tunnel. The test condition is a nominal Mach number of 10 with a Reynolds number of 6 ˆ 10 based on the cylinder diameter. The state of the incoming boundary layer at separation is laminar. A comparison of the glow result with a threedimensional Navier-Stokes simulation shows that the density distribution of the flowfield is highly associated with the spatial distribution of glow spectrum, from which we deduce some main flow features. We emphasize that a multiple-horse-shoe-vortex system, which is essentially similar to that in supersonic regime, is formed underneath the separated shear layer. In addition, oil flow and encapsuled thermochromic liquid crystal visualizations augment this observation.

Method for Visualizing Viscous Interaction in Hypersonic Low-Density Flows Using Electrical Discharge

The purpose of this study was to determine whether an electrical discharge method can be used to investigate separated flows around models tested in low-density, short-duration hypersonic facilities. To demonstrate the applicability of the proposed method, the visualization of the vortical structure around a circular cylinder placed perpendicular to a flat plate was attempted at a small Knudsen number, where methods commonly used for locating flow separations are ineffective. The density distribution of the flowfield was shown to be associated with the spatial distribution of the glow spectrum. The foci of the horseshoe vortices where the gas density was low enough appeared as dark regions in the visualization, especially in the viscous regime, allowing them to be easily located. In addition, this paper presents a simple conceptual model developed to describe the mechanism behind the present method.

Ramp Wall Pressure Measurement in Hypersonic Shock/Boundary Layer Interaction

The purpose of this study is to determine whether an unsteady interactive flow in the absence of incoming turbulence is fundamentally possible. The flow unsteadiness was examined based on pressure measurements (focusing on the boundary layer reattachment on the ramp) as well as high-speed schlieren visualization. The experiments were conducted on a flat plate with an aft-mounted ramp in a gun tunnel. The test condition was a nominal Mach number of 10 with a Reynolds number of 2:1 10 5 based on the flat plate chord. The state of the incoming boundary layer at separation was laminar. The unsteady motion of the separation shock, the state of the reattaching boundary layer, and the fluctuations of the wall pressure were mutually dependent.