• 한국전산유체공학회:학술대회논문집
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• 2003.10a
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• pp.243-243
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• 2003

• Proceedings of the KSME Conference
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• 2008.11b
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• pp.2782-2787
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• 2008

The subcavity passive control technique is used in present study. Cavity-induced pressure oscillation has been investigated numerically for a supersonic three-dimensional flow over rectangular cavities at Mach number 1.83 at the cavity entrance. Time dependent supersonic turbulent flow over cavity were examined by using the three-dimensional, mass-averaged Navier-Stokes equations based on a finite volume scheme and large eddy simulation. The results showed that the resultant amount of attenuation of cavity-induced pressure oscillations was dependent on the length and thickness of the flat plate.

• Journal of computational fluids engineering
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• v.11 no.4 s.35
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• pp.62-66
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• 2006

High-speed flight vehicle have various cavities. The supersonic cavity flow is complicated due to vortices, flow separation, reattachment, shock waves and expansion waves. The general cavity flow phenomena includes the formation and dissipation of vortices, which induce oscillation and noise. The oscillation and noise greatly affect flow control, chemical reaction, and heat transfer processes. The supersonic cavity flow with high Reynolds number is characterized by the pressure oscillation due to turbulent shear layer, cavity geometry, and resonance phenomenon based on external flow conditions. The resonance phenomena can damage the structures around the cavity and negatively affect aerodynamic performance and stability. In the present study, we performed numerical analysis of cavities by applying the unsteady, compressible three dimensional Reynolds-Averaged Navier-Stokes(RANS) equations with the ${\kappa}-{\omega}$ turbulence model. The cavity model used for numerical calculation had a depth(D) of 15mm cavity aspect ratio (L/D) of 3, width to spanwise ratio(W/D) of 1.0 to 5.0. Based on the PSD(Power Spectral Density) and CSD(Cross Spectral Density) analysis of the pressure variation, the dominant frequency was analyzed and compared with the results of Rossiter's Eq.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2003.10a
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• pp.108-113
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• 2003

The numerical research has been done for the transverse jet behind a rearward- facing step in turbulent supersonic flow without chemical reaction. The purpose of transverse jet is used to improve mixing of the fuel in the combustor. Two- dimensional unsteady flowfields generated by slot injection into supersonic flow are numerically simulated by the integration of Navier-Stokes equation with two-equation k - $\varepsilon$ turbulence model. Numerical methods are used high-order upwind TVD scheme. Eight cases are computed, comprising slot momentum flux ratios and slot position at downstream of the step. The flow is very similar to the cavity flow, because the jet is like an obstacle. Therefore, the numerical results show the periodic phenomenon.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.25 no.6
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• pp.837-844
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• 2001

Jet cutting technology currently makes use of a generic supersonic gas jet to improve the cutting speed and performance. In order to get a better understanding of the flow characteristics involved in the supersonic jet cutting technology, the axisymmetric Navier-Stokes equations have been solved using a fully implicit finite volume method. Computations have been conducted to investigate some major characteristics of supersonic coaxial turbulent jets. An assistant gas jet has been imposed on the primary gas jet to simulate realistic jet cutting circumstance. The pressure and the temperature ratios of the primary and assistant gas jets are altered to investigate the major characteristics of the coaxial jets. The total pressure and Mach number distributions, shock wave systems, and the jet core length which characterize the coaxial jet flows are strongly affected by the pressure ratio, but not significantly dependent on the total temperature ratio. The assistant gas jet greatly affects the basic flow characteristics of the shock system and the core length of under and over-expanded jets.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.45 no.5
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• pp.367-375
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• 2017

RANS computational analysis was performed on the head of the launch vehicle including the hammerhead nose pairing in the supersonic regime. The two-dimensional axisymmetric analysis was performed by using laminar, fully turbulent and transition models and compared with the experimental data. It was observed that different flow phenomena occurred depending on the Reynolds number. Under the high Reynolds number condition, the boundary layer becomes turbulent, which is not separated from the surface of the launch vehicle. With the low Reynolds number condition, laminar separation bubble was produced due to the separation and reattachment of the boundary layer on the expansion-compression edge of the hammerhead type nose fairing. The three-dimensional computations with the angle of attack showed a fully detached vortical structure due to the laminar separation bubble. It is proved that the turbulent transition should be considered to predict the separation bubble with the Reynolds number.

• Journal of computational fluids engineering
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• v.1 no.1
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• pp.98-111
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• 1996

The compressible laminar and turbulent viscous flows on a slender body in supersonic speed as well as subsonic speed have been numerically simulated at high angle of attack. The steady and time-accurate compressible thin-layer Navier-Stokes code based on an implicit upwind-biased LU-SGS algorithm has been developed and specifically applied at angles of attack of 20, 30 and 40 dog, respectively. The modified eddy-viscosity turbulence model suggested by Degani and Schiff was used to simulate the case of turbulent flow. Any geometric asymmetry and numerical perturbation have not been intentionally or artificially imposed in the process of computation. The purely numerical results for laminar and turbulent cases, however, show clear asymmetric formation of vortices which were observed experimentally. Contrary to the subsonic results, the supersonic case shows the symmetric formation of vortices as indicated by the earlier experiments.

• Journal of computational fluids engineering
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• v.13 no.3
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• pp.1-7
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• 2008

A supersonic base flow is computed to investigate the effect of the eddy viscosity coefficient to the linear ${\kappa}-{\varepsilon}$ turbulence models. Slight modifications to the eddy viscosity coefficient, which are based on the realizability condition, are given to the Launder-Sharma turbulence model so that present models satisfy the realizability condition. Numerical results for supersonic base flow show that turbulence models with the weaky-nonlinear eddy viscosity coefficient can lead to reasonable enhancements in the prediction of the velocity and turbulent kinetic energy profiles.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.20 no.3
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• pp.1083-1095
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• 1996

Shock wave/boundary layer interaction frequently causes the shock wave to oscillate violently and thus the global flow field to unstabilize. In order to stabilize the shock wave system in the diffuser of a supersonic wind tunnel, the present study attempted to control the shock oscillations by using a passive control. A porous wall with the porosity of 19.6% was mounted on a shallow cavity. Experiment was made by means of schlieren optical observation and wall pressure measurements. The flow Mach number just upstream the shock system and Reynolds number based on the turbulent boundary layer thickness were 2.1 and 1.8 * 10$\^$6/, respectively. The results show that the present passive control method on the shock wave/boundary layer interaction in the supersonic diffuser can significantly suppress the oscillations of shock system, especially when the shock system locates at the porous wall.

• Journal of the Korea Institute of Military Science and Technology
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• v.6 no.4
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• pp.124-131
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• 2003

When a secondary gaseous flow is injected vertically into a supersonic flow through circular nozzle, a complicated structure of flow field is produced around the injection area. The interaction between the two streams produces a strong bow shock wane on the upstream side of the side-jet. The results show that bow shock wave and turbulent boundary layer interaction induces the boundary layer separation in front of the side-jet. This study is to analyze the structure of flow fields and distribution of surface pressure on the flat plate according to total pressure ratio using a supersonic cold-flow system and also to study the control force of affected side-jet. The nozzle of main flow was designed to have Mach 2.88 at the exit. The injector has a sonic nozzle with 4mm diameter at the exit of the side-jet. In experiments, The oil flow visualization using a silicone oil and ink was conducted in order to analyze the structure of flow fields around the side-jet. The flow fields are visualized using the schlieren method. In this study, a computational fluid dynamic solution is also compared with experimental results.