Toshihiro Ogawa

Holographic Interferometric Visualization of the Richtmyer-Meshkov Instability Induced by Cylindrical Shock Waves

Results of quantitative holographic interferometric flow visualization of cylindrical interface instability induced by converging cylindrical shock waves are reported. Experiments were conducted in an annular vertical co-axial diaphragmless shock tube, in which cylindrical soap bubbles filled with He, Ne, Air, Ar, Kr, Xe and SF6 were co-axially placed in its test section. Pressure histories at different radii during the shock wave implosion and reflection from the center were measured. Diagnostic method base on double exposure holographic interferometry was applied for the measurement of turbulent mixing zone at the interface. The observed cylindrical interfaces were found to have a higher growth rate of turbulent mixing zone than that of the plane shock / plane interface.

Experimental and Computational Fluid Dynamics Around Supersonic Biplane for Sonic-Boom Reduction

The “supersonic biplane theory” has been proposed in order to reduce the sonic boom in supersonic flight. This theory can be realized at a certain design Mach number, but it cannot be performed at off-design ones. Therefore, some innovations and multidisciplinary design optimization are strongly required to improve the off-design performance. At first, the aerodynamic characteristics around biplanes as a baseline configuration have to be analyzed in order to contribute to the conceptual design of the supersonic biplane. Especially, CFD analyses can produce effective results for the flow phenomena to be captured comprehensively; however, they have to be validated by experimental results. The purpose of this study is to demonstrate experimentally the shock wave interference and cancellation between the wings of biplanes based on the “supersonic biplane theory”. Supersonic and transonic flow fields around the biplanes are also examined by CFD analyses and experimental results. This interdisciplinary data will hopefully assist in the effective conceptual design and multi-objective design exploration (MODE) of the supersonic biplane in the near future.

Wind Tunnel Testing on Start/Unstart Characteristics of Finite Supersonic Biplane Wing

This study describes the start/unstart characteristics of a finite and rectangular supersonic biplane wing. Two wing models were tested in wind tunnels with aspect ratios of 0.75 (model A) and 2.5 (model B). The models were composed of a Busemann biplane section. The tests were carried out using supersonic and transonic wind tunnels over a Mach number range of with angles of attack of 0°, 2°, and 4°. The Schlieren system was used to observe the flow characteristics around the models. The experimental results showed that these models had start/unstart characteristics that differed from those of the Busemann biplane (two dimensional) owing to three-dimensional effects. Models A and B started at lower Mach numbers than the Busemann biplane. The characteristics also varied with aspect ratio: model A () started at a lower Mach number than model B () owing to the lower aspect ratio. Model B was located in the double solution domain for the start/unstart characteristics at , and model B was in either the start or unstart state at . Once the state was determined, either state was stable.

Unsteady drag force measurements over bodies with various configurations in a vertical shock tube

Paper reports the result of unsteady drag force measurements and their comparison with numerical simulation. Experiments were conducted in a vertical shock tube consisting of a 300 mm i.d. and 2.5 m long high-pressure chamber and a 300 mm × 300 mm square low-pressure channel of 5 m in length. Models with various configurations installed with accelerometers in them were suspended along the low-pressure test section and then planar shock waves at shock Mach number of 1.22 in air were loaded on these models. Model shapes were cones with smooth and coarse surface finish, a double cone a sphere, a 2:3 ellipsoid and a cylinder. The time variation of drag forces exerted on the models was compared to that obtained by solving numerically the Euler equations. A good agreement was obtained between the experiments and numerical results. Unsteady drag force over these models shows peak value and gradually decreases to its steady value.

Gun Interior Ballistic Performance with Ammonium Nitrate-Alcohol Propellants

A new type of propellant, a mixture of ammonium nitrate and alcohol, was examined in ballistic experiments. Withprilledammoniumnitrateandethanol,increasedballisticperformancewasobtainedforthemaximumenergy release conditions. From the measured chamber pressures and measured in-tube velocity proe le, the location of energy release and pressure wave motions are documented. This propellant is characterized by low-product molecular mass, high-specie c energy, and high e exibility in controlling its performance both by changing the type of alcohol and by varying the reactants’ molar ratio.

Micro-Shock Waves Generated Inside a Fluid Jet Impinging on Plane Wall

An Nd:glass laser pulse (18 ns, 1.38 J) is focused in a tiny area of about 100-mum diam under ambient conditions to produce micro-shock waves. The laser is focused above a planar surface with a typical standoff distance of about 4 mm, The laser energy is focused inside a supersonic circular jet of carbon dioxide gas produced by a nozzle with internal diameter of 2.9 mm and external diameter of 8 mm, Nominal value of the Mach number of the jet is around 2 with the corresponding pressure ratio of 7.5 (stagnation pressure/static pressure at the exit of the nozzle), The interaction process of the micro-shock wave generated inside the supersonic jet with the plane wall is investigated using double-pulse holographic interferometry. A strong surface vortex field with subsequent generation of a side jet propagating outward along the plane wail is observed. The interaction of the micro-shock wave with the cellular structure of the supersonic jet does not seem to influence the near surface features of the flowfield. The development of the coherent structures near the nozzle exit due to the upstream propagation of pressure waves seems to be affected by the outward propagating micro-shock wave. Mach reflection is observed when the micro-shock wave interacts with the plane wall at a standoff distance of 4 mm, The Mach stem is slightly deflected, indicating strong boundary-layer and viscous effects near the wall. The interaction process is also simulated numerically using an axisymmetric transient laminar Navier-Stokes solver. Qualitative agreement between experimental and numerical results is good.

Visualizing microair blasts using double-exposure holographic interferometry

Micro air blasts are generated over planar and corrugated surfaces, at stand-off distances of 4 and 11 mm, by focusing a 1.38J, 18 ns Nd:Glass laser pulse in a tiny spot of 300 micrometer diameter, under ambient conditions. This triggers the breakdown of air at optical frequencies with subsequent formation of laser plasma, which generates an outward propagating spherical micro-blast wave. The velocity of this wave decreases as time elapses, since the energy deposition is finite, with reference to both space and time. The evolution, propagation and subsequent reflection of the spherical micro- blast wave, from planar and corrugated surfaces are experimentally investigated, using double exposure holographic interferometry. One planar surface and four surfaces with varying degree of surface corrugations are used in this study. Mach reflection of the blast wave is observed at a stand-off distance of 4 mm over a planar obstacle while only regular reflection is visualized at an 11 mm stand-off distance. However only regular reflection is observed at both stand-off distances, in the case of corrugated surfaces.

Computations of Flow Field around an Object Decelerating from Supersonic to Subsonic Velocity

A detached shock wave is formed in front of a blunt object moving with supersonic velocities or an object placed in a supersonic flow. The distance fromthe shock front to the surface of the moving body is called the shock stand-off distance, δ , and is measured along the propagation axis when the object is a solid sphere.

Visualization of weak shock waves emitted from a trombone.

The coalesce of compression waves propagating in pipes into weak shock waves is one of the topics of shock wave dynamics. We have worked for weak shock waves generated in automobile exhaust pipe lines [Sekine et al. (1990)] and train tunnel sonic booms in Japanese high speed train entry into a long tunnel [Takayama et al. (1995)]. Hirschberg et al. (1996) reported, for the first time, the emission of weak shock waves from a trombone blown in ff and Pandya et al. (2003) visualized weak shock waves emitted from brass instruments. So far we understood, these belong to our continuous academic curiosity. It is a highlight of brass instruments to produce dramatizing sounds. In the 631‐bar of the fourth movement of Mahler’s symphony No. 1, fff is specified for a third trombone. We then visualized, by using a set of 1 m diameter schlieren mirrors in schlieren optics and holographic interferometry, weak shock waves emitted from a trombone blown in ff and fff and measured pressure at the muzzle and horn and velocit...

Development of Pressure-Sensitive Paint Technique in a Supersonic Indraft Wind Tunnel and its Application to a Busemann Biplane

Using pressure sensitive paint, the pressure distributions on thin wings of the Busemann biplane model were measured in the small supersonic indraft wind tunnel. The observed phenomenon is so complicated that it is completely different from the prediction of the simple 2-D theory. Referring to Schlieren picture and CFD analysis, it is found that these complex pressure fields are caused by the shock wave generated downstream just behind the wing ridge line. The results of this experiment show that the interference between Busemann biplane wings is sensitive to a small change in flow condition. A care must be taken in a wind tunnel test to realize the design condition of Busemann biplane where the shock waves are completely cancelled with each other.