Yutaka Yamaguchi

Instrumentation and operation of NDA cryogenic wind tunnel

The NDAs (National Defense Academy of Japan) 0.06-m*0.3-m cryogenic wind tunnel was constructed in 1985 for testing transonic airfoils and for other basic research of fluid mechanics. Stainless steel SUS 304 was chosen as the material of the pressure shell, and a centrifugal compressor was chosen as the compressor. External insulation was adopted for the tunnel. Although no information was available on problems of using SUS 304 as the externally insulated tunnel pressure shell at cryogenic temperature and the thermal conductivity of SUS 304 is worse than that of aluminium alloys, only eight thermocouples were installed to monitor the thermal condition of the shell. The original temperature control was achieved by manually controlling the mass flow of liquid nitrogen injected into the tunnel circuit, but that system was found to be inadequate because the settling time of the total temperature took about 15 min in the change of rotational speed of the compressor. The total pressure control systems were modified to simple automatic PID (proportional, integral, derivative) controls. As a result, the pressure control can be achieved almost perfectly, and the temperature control is also greatly improved as the settling time of temperature is greatly reduced.<<ETX>>

Preliminary Study on Diaphragmless Shock Tube for Transonic Airfoil Testing with PDI

In the present study, the diaphragmless shock tube is developed for the transonic and the high-subsonic airfoil testing. The flow visualizations of the shock tube as the wind tunnel is performed by the PDI (point diffraction interferometry) technique. The PDI technique that can obtain the quantitative information is utilized for the visualization. Moreover, the comparison with numerical results by the OpenFOAM is performed. The driver tube of the shock tube is the circular cross section with a diameter of 150 mm. The driven tube is the rectangular cross section with a width of 60 mm and a height of 150 mm. The airfoil model of a NACA0012 as the standard model has a chord length of 40 mm, and aspect ratios are 1.5. The hot gas Mach number behind the incident shock wave is about 0.84, and it is recognized that this facility can perform transonic wind tunnel testings. The flow starting process of the shock tube is visualized, and it is confirmed that the steady flow is established in the test section. And the PDI technique showed the visualization of the shock tube airfoil flow qualitatively. In the high-subsonic flow, the pressure coefficient Cp by PDI is in good agreement with the CFD result quantitatively.

Preliminary Study on Lift Coefficient of Biplane Airfoil in Smoke Wind Tunnel

In the present study, flow visualizations around a biplane airfoil equipped with the leading and trailing edge flaps are performed in a low-speed smoke wind tunnel. And we estimate the lift coefficient on the biplane airfoil utilizing the method based on the smoke line pattern measurements. The results of present study are as follows. It was confirmed that all lift curve slopes were nearly identical with and without uses of the leading edge and trailing edge flaps. When the trailing edge flaps are deflected downward, the lift coefficient increases. The zero lift angle of the present result is not directly proportional to the angle of trailing edge flaps δf. The leading edge deflection increasing angle of flaps increase the stall angle of attack of the biplane airfoil.

Experimental Study on Aerodynamic Characteristics of Blended-Wing-Body by a Wake Integration Method

In the present study, a BWB model that has a large sweepback, wash out, and a reflex wing configuration is designed. And, the aims are to obtain basic data of the aerodynamic characteristic of the BWB in low speed flows using the wake measurement. The wind tunnel in this study is a Gottingen type low-speed wind tunnel. The wind tunnel is a return type with an open test section. The shape of the nozzle exit is an octagon having an inner diameter of 1.0 m. The wake survey system has a five-hole probe to obtain static pressures and a stagnation pressure in the wake survey region. The five-hole probe tube is conically-shaped, and made from stainless steel. The inner diameters of each hole are 0.2 mm, and the total diameter is 3.16 mm. The free-stream velocity is kept to be 25 m/s, and the Reynolds number based on the mean aerodynamic chord length is 2.6×10, respectively. The results are as follows. The section lift coefficient has been generated not only at the wings but also at the center body because the center body was constructed by wing shape. On the section induced drag coefficient, two peaks are seen at the angle of attack 8° and 10°. It seems that the locations of the peaks are at the same locations of the large total pressure loss in the angle of attack 10°. A basic aerodynamic characteristic of BWB was clarified by the present study.

A Fundamental Study on Aircraft Model by Wake Measurements in Low-Speed Wind Tunnels

In this paper, the drag prediction method (wake integration method) based on the momentum conservation theorem was applied to the aerodynamic investigations on aircraft model. The wake survey of the model is performed to obtain the basic data to evaluate how applicable the technique is to aerodynamic tests in a relatively small low-speed wind tunnel. In this case, it is important to investigate the effect of the measurement interval by the wake survey. The aircraft model is the same geometry model that was used for the correction wind tunnel test previously. The free-stream velocity is kept to be 30 m/s. The Reynolds number based on the mean aerodynamic chord length is

A Modified Point-Diffraction Interferometry for Shock Tube Airfoil Testing

In the present study, a modified point diffraction interferometry was developed for shock tube airfoil flows. This system was made by referring to the idea of the multi-pass technique. And the technique was applied to shock tube airfoil flows to discuss the effect of Mach number and operation gas. The diaphragmless shock tube is used for airfoil testings. The driver tube is the circular cross section of 15 mm in diameter, the driven tube is the rectangular cross section of 60 mm in width and 150 mm in height. At first, the preliminary experiment was done by using the candle with the conventional and the multi-pass optical system for PDI. The fringe obtained by the multi-pass system can confirm more numbers. In the shock tube airfoil testing, the pressure coefficient analyzed by the conventional technique is almost corresponding to the pressure coefficient analyzed by the multi-pass technique. On the effect of the operation gas, in CO2 modes, the weak kink point of the fringe does not appear in the Mach number 0.6 when the Mach number and the Reynolds are adjusted to the same condition of air mode.

Preliminary Study for Quantitative Measurement of Flow Fields by Focusing Schlieren Method

Recently a large-field and a high-brightness focusing schlieren system was developed to visualize three-dimensional density gradients in flow fields. However, a quantitative experiment seems to be still few, and the accumulation of further data is necessary to verify the quantitative measurement technique. In the present study, a diaphragmless shock tube that is expected good stability and efficiency is used to produce basic flows (utilizing a normal shock wave propagating in the test section). The aim is to apply Cook’s theory to calculate the density distribution in the shock tube flow, and discuss the characteristics of the quantitative parameter of the focusing schlieren system. The results are as follows, on the effect of the source grid (utilizing the vertical, horizontal and lattice grid) for the quantitative consideration, the results of the focusing schlieren method observed the cell structure and the shear layer in the free jet. And it seems that the characteristics of the image have same trend as well as the conventional schlieren method. The density across the incident shock wave becomes small as well as the density gradient, when the quantitative parameters β and N are increased. When the parameters are β = 0.1 and N = 3, the density distributions across the incident shock wave almost correspond to the simple theory in the present experimental condition.

Visualization of shock tube airfoil flow with a sharp focusing schlieren method

Visualization of the flow field around the base line model under the flight condition is one of the most valuable and basic steps on the high-speed airplane development. Now we consider a sharp focusing schlieren system that gives a means for observing any cross section of a flow field of perpendicular to the test beam axis of the optical system. This system can be used to examine a complex three-dimensional flow field. As a preliminary study, a sharp focusing schlieren system is designed with referring to the Weinsteins system. And visualization of steady transonic airfoil shock tube flow is performed with this system to investigate the effect of side-wall interference to the shock wave locations on an airfoil surface. The results show that the instrument has a capability for visualizing any cross section in the shock tub airfoil flow. Also, it is pointed out that the main steady shock wave profile focused at the center of the channel is different from the profile focused near the side-wall.

Study on Laser-Induced Acetone Fluorescence in Low-Temperature Gases of Nitrogen and Air

5National Defense Academy,1-10-20 Hashirimizu, Yokosuka, Kanagawa, 239-8686, JAPAN In the present study, the acetone LIF technique with nitrogen as the buffer gas was performed by the static chamber. And the fundamental experiment by using air was performed for the feasibility of the application of quantitative measurements in low temperature flows. The vacuum chamber was attached to achieve vacuum insulating in the test chamber. The size of internal cylindrical chamber is 83.1 mm diameter and 318 mm height. This system is composed of the light source, the optical system, the seeder, the test chamber, and the measuring instruments. Nd:YAG laser with a forth harmonic mode of 266nm is used as the light source . The results are as follows. It is shown that the absorption cross section of acetone shows almost constant value. The fluorescence intensity in both of nitrogen and air slightly increases together with the temperature decrease. And the experimental results are comparatively corresponding to the model curve. Also the flow visualization of the underexpanded air jet was performed by utilizing the acetone PLIF technique. The photograph shows the detail characteristics of the underexpanded jet.

Preliminary Study on the Acetone LIF Method for Quantitative Measurement of Low Temperature Flows

The laser-induced fluorescence method (LIF) has been attracted as an available diagnostic tool for quantitative measurements in wind tunnel testing. Acetone is an ideal tracer material for the LIF method because of low toxicity and its strong fluorescence intensity. Although many studies on the acetone LIF method have been reported, most of these studies are limited to high temperature conditions. In the present study, we investigate the acetone LIF characteristics in low temperature conditions. The 4th harmonic of the Nd:YAG laser (266nm) is used to excite the acetone molecule and the resulting fluorescence intensity is detected by the photomultiplier tube. Nitrogen is selected as a buffer gas. The experimental result reveals that the LIF intensity was almost constant in the present temperature range of 270-300K.