Fundamental Study on Operational Parameters of Diaphragmless Shock Tube

This paper shows influences of initial conditions on a diaphragmless shock tube operation. This facility consists of a driver tube, a driven tube and a damp tank. The driver tube has a circular cross section with diameter of 150 mm and the driven tube, a rectangular cross section (60 mm x 150 mm). The shock tube is operated by using a quick-opening pneumatic piston instead of a diaphragm. For the operation, pressure inside the pneumatic piston (piston pressure) is quickly released by opening a solenoid valve. In this paper, the initial piston pressure was chosen as a parameter to investigate effects on test flow conditions created by the shock tube. As a result, it was found that when the piston pressure at initial condition is large, piston pressure decreased more rapidly than that obtained for a small piston pressure condition, regardless of the pressure ratio of driver and the driven tube. In the condition of a constant initial operational pressure ratio and a different piston pressure, the shock Mach number was almost constant

Fundamental Study on Operational Parameters of Diaphragmless Shock Tube

This paper shows influences of initial conditions on a diaphragmless shock tube operation. This facility consists of a driver tube, a driven tube and a damp tank. The driver tube has a circular cross section with diameter of 150 mm and the driven tube, a rectangular cross section (60 mm x 150 mm). The shock tube is operated by using a quick-opening pneumatic piston instead of a diaphragm. For the operation, pressure inside the pneumatic piston (piston pressure) is quickly released by opening a solenoid valve. In this paper, the initial piston pressure was chosen as a parameter to investigate effects on test flow conditions created by the shock tube. As a result, it was found that when the piston pressure at initial condition is large, piston pressure decreased more rapidly than that obtained for a small piston pressure condition, regardless of the pressure ratio of driver and the driven tube. In the condition of a constant initial operational pressure ratio and a different piston pressure, the shock Mach number was almost constant

Two-dimensional quantitative visualization of isolator shock trains by rainbow schlieren deflectometry

The rainbow schlieren deflectometry is effective in studying quantitatively the density fields in shock-containing free jets at high precision and high spatial resolution. However, there has been no practical application of rainbow schlieren deflectometry for shock trains in a confined duct. Therefore, in the present study, the rainbow schlieren deflectometry is applied to the flow field including a shock train in a constant-area duct where just upstream of the shock train the freestream Mach number is 1.34, the unit Reynolds number is 5.39 × 107 m-1, and the boundary layer displacement thickness is 0.149 mm. As a result, a two- dimensional density field of the shock train is for the first time quantitatively displayed and the fine structure of the shock train is illustrated as a color gradation representation

Two-dimensional quantitative visualization of isolator shock trains by rainbow schlieren deflectometry

The rainbow schlieren deflectometry is effective in studying quantitatively the density fields in shock-containing free jets at high precision and high spatial resolution. However, there has been no practical application of rainbow schlieren deflectometry for shock trains in a confined duct. Therefore, in the present study, the rainbow schlieren deflectometry is applied to the flow field including a shock train in a constant-area duct where just upstream of the shock train the freestream Mach number is 1.34, the unit Reynolds number is 5.39 × 107 m-1, and the boundary layer displacement thickness is 0.149 mm. As a result, a two- dimensional density field of the shock train is for the first time quantitatively displayed and the fine structure of the shock train is illustrated as a color gradation representation

Effect of Stagger on Low-Speed Performance of Busemann Biplane Airfoil

In this study, the low-speed performances of the Busemann biplane were clarified, focusing on the relative contributions of the upper and lower elements to the total aerodynamic characteristics of the biplane. Also, the effects of the staggered biplane, which changes the horizontal distance between two wings in a biplane configuration, were investigated by balance measurements and numerical simulations. The flow velocity was 15 m/s, and the Reynolds number based on the airfoil chord length was 2.1 × 105. In the tests of the integrated biplane wing, the attack angles of the wing elements were varied by a balance system and turntable, which were set in the wind tunnel sidewall. The results show that the lower element generated most of the lift and drag of the Busemann biplane (or the baseline biplane model with no stagger) at high angles of attack. At angles above 20 deg, the contribution of the lower element to total aerodynamic characteristics is almost constant, with 95% of the total lift and 88% of the total drag. The total lift and drag of the baseline model were smaller than the sum of the individual elements that were treated as a single configuration. The increments of lift and drag due to the stagger effects were confirmed, especially at high angles of attack. When the stagger value increases, the high-pressure area near the leading edge of the lower surface of the upper element also increases, which increases the lift and drag of the up-per element. This is the main reason for the increments of total lift and drag of the biplane model. The stagger effects also prevented the leading-edge separation of the lower element in the biplane configuration and increased the lift slopes of the biplane model

防衛大学校高速風洞の最近の話題について

形態: カラー図版ありPhysical characteristics: Original contains color illustrationsレポート番号: SP-23-00

Development of Turbulent Boundary Layer in a Supersonic Nozzle

The effects of velocity profile and local skin friction coefficient on the turbulent boundary-1ayer development in a supersonic nozzle were investigated by using Tucker\u27s method. The calculated results agree well with the experimental data in zero－and favorable pressure gradient flows provided that the pertinent velocity profile parameter and the local skin friction coefficient are selected. Also, it is shown that the turbulent boundary-layer development in the flow of which pressure gradient is almost negligible depends strongly on the local skin friction coefficient and it is almost independent on the velocity profile