• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.33 no.8
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• pp.92-98
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• 2005

Screen can provide any disturbed resistance that affects the change in characteristics of turbulence, velocity and pressure distributions of the flow field, and thus it has been widely used to control the flow. Some previous related studies for compressible flows have limitations such as, considering relatively low-Mach-number flows in the range of 0.3 ∼ 0.7, and not observing the detailed shock structures of the flow fields. An experimental study on highly compressible axi-symmetric supersonic jet flow fields behind wire-gauze screen has thus been carried out. Continuous/instantaneous flow images by Schlieren flow- visualization technique and the information of Pitot pressure/flow-noise measurements of the flow field behind the screen for various jet expansion conditions have been obtained. Effects of various porosity and inclination angles of the screen at the nozzle exit have also been investigated, and the experimental results have been compared to the case with no screen installed.

• Journal of Mechanical Science and Technology
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• v.20 no.8
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• pp.1256-1265
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• 2006

The supersonic flows around tandem cavities were investigated by two-dimensional and three-dimensional numerical simulations using the Reynolds-Averaged Navier-Stokes (RANS) equation with the k- ω turbulence model. The flow around a cavity is characterized as unsteady flow because of the formation and dissipation of vortices due to the interaction between the freestream shear layer and cavity internal flow, the generation of shock and expansion waves, and the acoustic effect transmitted from wake flow to upstream. The upwind TVD scheme based on the flux vector split with van Leer's limiter was used as the numerical method. Numerical calculations were performed by the parallel processing with time discretizations carried out by the 4th-order Runge- Kutta method. The aspect ratios of cavities are 3 for the first cavity and 1 for the second cavity. The ratio of cavity interval to depth is 1. The ratio of cavity width to depth is 1 in the case of three dimensional flow. The Mach number and the Reynolds number were 1.5 and $4.5{\times}10^5$, respectively. The characteristics of the dominant frequency between two- dimensional and three-dimensional flows were compared, and the characteristics of the second cavity flow due to the first cavity flow was analyzed. Both two dimensional and three dimensional flow oscillations were in the 'shear layer mode', which is based on the feedback mechanism of Rossiter's formula. However, three dimensional flow was much less turbulent than two dimensional flow, depending on whether it could inflow and outflow laterally. The dominant frequencies of the two dimensional flow and three dimensional flows coincided with Rossiter's 2nd mode frequency. The another dominant frequency of the three dimensional flow corresponded to Rossiter's 1st mode frequency.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.206-212
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• 2011

Hysteresis phenomena in fluid flow systems are frequently encountered in many industrial and engineering applications and mainly appear during the transient processes of change of the pressure ratio. Shock-containing flow field in supersonic nozzles is typically subject to such hysteresis phenomena, but associated flow physics is not yet understood well. In the present study, experimental work has been carried out to investigate supersonic nozzle flows during the transient processes of change in the nozzle pressure ratio. Time-dependent surface wall pressures were measured by a multiple of pressure transducers and the flow field was visualized using a nano-spark Schlieren optical method. The results obtained show that the hysteresis phenomenon is strongly dependent on the nozzle geometry as well as the time scale of the change of pressure ratio.

• International Journal of Aeronautical and Space Sciences
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• v.4 no.1
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• pp.53-62
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• 2003

In this study, nonlinear aeroelastic characteristics of an supersonic missile wing with strong shock interferences are investigated. The missile wing model has a freeplay structural nonlinearity at its pitch axis. To practically consider the effects of freeplay structural nonlinearity, the fictitious mass method is applied to structural vibration analysis based on finite element method. Nonlinear aerodynamic flows with unsteady shock waves are also considered in supersonic flow regions. To solve the nonlinear aeroelastic governing equations including the freeplay effect, a modal-based coupled time-marching technique based on the fictitious mass method is used in the time-domain. Various aeroelastic computations have been performed for the nonlinear wing structure model. Linear and nonlinear aeroelastic analyses have been conducted and compared with each other in supersonic flow regions. Typical nonlinear limit cycle oscillations and phase plots are presented to show the complex vibration phenomena with simultaneous fluid-structure nonlinearities.

• Advances in aircraft and spacecraft science
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• v.9 no.2
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• pp.119-130
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• 2022

Supersonic projectiles like rockets, missiles, or aircraft find various applications in the field of defense. The shape of the wings is mainly designed as wedge shape or delta wings for supersonic vehicles. The study of supersonic flows over the wedges and flat plate delta wings around the large scale of incidence angle is considered in the supersonic projectile. In the present paper, the prime attention is to study the pressure at the nose of the plane wedge over the various Mach number and the various angles of incidence. Ghosh piston theory is used to obtain the pressure distribution analytically, and the results are compared with CFD analysis results. The wedge angle and Mach number are the parameters considered for the research work. The range of wedge angle is 50 to 250, and Mach number is 1.5 to 4.0 are considered for the current research work. The analytical results show excellent agreement with the CFD results. The results show that both the parameters wedge angle and Mach number are influential parameters to vary the static pressure. The static pressure increases with an increase in Mach number and wedge angle.

• Journal of the Korean Society of Propulsion Engineers
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• v.3 no.4
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• pp.58-66
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• 1999

The flow characteristics of supersonic ejectors is often subject to compressibility, unsteadiness and shock wave systems. The numerical works carried out thus far have been of one-dimensional analyses or some Computational Fluid Dynamics(CFD) which has been applied to only a very simplified configuration. For the design of effective ejector-pump systems the effects of secondary mass flow on the supersonic ejector flow should be fully understood. In the present work the supersonic ejector-pump flows with a secondary mass flow were simulated using CFD. A fully implicit finite volume scheme was applied to axisymmetric compressible Navier-Stokes equations. The standard two-equation turbulence model was employed to predict turbulent stresses. The results obtained showed that the flow characteristics of constant area mixing tube types were nearly independent of the secondary flow rate, but the flow fields of ejector system with the second-throat were strongly dependent on the secondary flow rate due to the effect of the back pressure near the primary nozzle exit.

• Journal of the Korean Society of Propulsion Engineers
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• v.4 no.4
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• pp.78-86
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• 2000

The objective of this research is to understand the characteristics of a nonpremixed, turbulent, hydrogen jet flame which is stabilized in Mach 1.8 coflowing air flows. In order to investigate the flame structure, flame lengths and fuel trajectories were measured by using direct photography, acetone PLIF, Mie scattering techniques, and numerical simulation. Effect of increasing air velocity was investigated when fuel velocity is fixed. The subsonic flame length was decreased drastically, however the supersonic flame length was increased slowly Then the change of flame blow out characteristics was observed as varying fuel nozzle lip thickness. The flame stability can be increased when fuel nozzle lip thickness was increased, which indicates that the minimum fuel lip thickness ratio is required for the stable supersonic flames. Also, it is found that fuel jet is blocked by high pressure zone and low scattering zone is made. Then the fuel that was moving along the recirculation zone had longer residence time within the supersonic flames, which made partially premixed zone.

• Journal of the Korea Academia-Industrial cooperation Society
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• v.17 no.11
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• pp.363-367
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• 2016

A test section of a supersonic wind tunnel was designed for the analysis of flow characteristics over a backward-facing step with Mach 1.0 slot injection in a supersonic flow of Mach 2.5. The cavity flow of a high-speed vehicle is very complex at supersonic speed, so it is necessary to do experiments using supersonic wind tunnels to verify numerical analysis methods. The previous 2D symmetrical nozzle was replaced with an asymmetrical nozzle. The inviscid nozzle contour was designed using Method of Characteristics (MOC), and the boundary layer thickness correction was reflected by experimental data from the wind tunnel. The results were compared with a CFD analysis. The PID control system was changed to be based on the change of tank pressure. This improved the control efficiency, and the run times of supersonic flow increased by about 1 second. The flow characteristics over a backward facing step with slot injection were visualized by a Schlieren device. This equipment will be used for an experimental study of the film cooling effectiveness over a cavity with various velocities, mass flows, and temperatures.

• International Journal of Aeronautical and Space Sciences
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• v.13 no.2
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• pp.188-198
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• 2012

An improvement to the k-${\varepsilon}$ turbulence model is presented and is shown to lead to better agreement with data regarding supersonic base flows. The improvement was achieved by imposing a grid-independent realizability constraint in the Launder-Sharma k-${\varepsilon}$ model. The effects of compressibility were also examined. The numerical results show that the modified Launder-Sharma model leads to some improvement in the prediction of the velocity and turbulent kinetic energy profiles. Compressibility corrections also lead to better agreement in both the turbulent kinetic energy and the Reynolds stress profiles with the experimental data.

• Journal of computational fluids engineering
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• v.10 no.4 s.31
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• pp.32-38
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• 2005

Present study simulates oscillatory supersonic impinging jet flows using the axisymmetric Navier-Stokes code. To capture the salient features of flow oscillation and overcome the divergence during the initial transient period, several tests have been conducted for the grid and time step sizes. The results also show that the effects of the inlet flow condition at the nozzle exit and turbulence on the oscillatory behavior of supersonic impinging jets are negligible. Frequencies of the surface pressure oscillation obtained by the selected numerical method are in good accord with the measured impinging tones for various cases of nozzle-to-plate distance. Two seemingly different staging behaviors with nozzle-to-plate distance and nozzle pressure variations are found to correlate well if the frequency and distance are normalized by the length of the first shock cell.