Keiichi Karashima

Enhancement of the leading-edge separation vortices by trailing-edge lateral blowing

The numerical simulations were carried out for the study of trailing-edge lateral blowing for the delta wing

Newly constructed high speed wind tunnel at the Institute of Space and Astronautical Science (ISAS) and related activities

The Institute of Space and Astronautical Science ( I S A S ) built a new high speed wind tunnel facility in 1989 in Sagamihara campus, where the ISAS was reiocated in 1987. Objective of the ,facility construction is to conduct the aerodynamic researches of high speed air andlor space rransporrnlion syst e m , air-breathing propulsion systems and recovery systems as well as basic studies in thefield of high speed aerodynamics. The fucility consists of a set of high pressure air-supply system, a transonic and a supersonic tunnels. Although both wind tunnels may be of conventional blow-down type, the manual procedures in tunnel operations and mea.rurements are highly simplified by making use offully-automatic control s y s t e m to save manpower and driving energy necessary for execution of the wind tunnel experi~ncnts. In this report are presented brief summaries of the facility and its performance, and some of the resulrs obtained in lunnel verification tests as well as a proposed experimental study ;hat will be made in neor future are shown for derr~onxtration.

The Experimental Studies on the Shock Wave/Turbulent Boundary Layer Interaction Region Induced by the Blunt Protuberance.

Three-dimensional shock wave/turbulent boundary layer interaction region induced by a blunt protuberance is experimentally investigated. Experiments are performed under the testing condition of freestream Mach number of 2, 3 and 4 with almost adiabatic wall condition. The primary separation lines locate almost same position irrespective of freestream Mach number, and freestream Mach number independence on the primary separation region is revealed. However, detailed flow structures for each Mach number are different. Pressure begins to rise at slightly ahead of the primary separation line where stream lines begin to deflect outward. As freestream Mach number is increased, the maximum pressure becomes higher but locations, where the pressure begins to rise and reaches to its maximum level, exist at almost same locations respectively regardless of freestream Mach numbers. The oscillating separated shock waves are observed in the interaction region by instantaneous shadow graph technique and significant unsteady flow structure is captured.

Flow Visualization of Secondary Separation and Oscillating Shock Waves in Three-Dimensional Shock Waves-Turbulent Boundary Layer Interaction Region

Secondary separation phenomena in three-dimensional shock wave/turbulent boundary layer interaction regions induced by sharp fin and blunt fins have been investigated. A color oil source method, in which various color oil are supplied through many holes on the interaction surface, are applied for detailed visualization of separated region as well as conventional oil flow technique. Experiments are conducted under the testing conditions of freestream Mach number of 2–4, total pressure of 0. 3-0. 62 MPa and unit Reynolds number of 2. 6-3. 9x107. The results show the new technique is quite useful for detecting the secondary separation. Also quite interesting, complicated features in the separated regions are revealed.

Supersonic Flow Past Axially Symmetric Bodies

An analytical method for superonic flow past axially symmetric bodies is presented and applied to calculate the initial ratio of shock to body curvature and initial surface pressure gradient on pointed bodies of revolution with arbitrary geometry. It is shown that the present method can also be applied to circularcones with such large semi-vertex angle that the flow behind the shock wave is subsonic. As an example, shock wave shape and sonic line are calculated for a cone with semi-vertex angle of 40.57% and at Mach number of 2. It is concluded that the detachment process of shock wave is continuous for circular-cones as well as for plane wedges.