• Journal of the Korean Society of Propulsion Engineers
• /
• v.7 no.3
• /
• pp.45-52
• /
• 2003

Detailed flowfield resulting from the secondary sonic gas injection into a divergent section of supersonic conical nozzle has been numerically investigated. The three-dimensional flowfield associated with the bow-shock/boundary-layer interaction inside the nozzle has been solved by Reynolds-averaged Navier-Stokes equations with an algebraic and $\kappa$-$\varepsilon$ turbulence model. The numerical results have been compared with the experimental results for the identical flow conditions, and it is shown that the comparison is satisfactory Effects of different injection pressures of the secondary jet on the shock/boundary-layer interactions and the overall flow structure inside the nozzle have been investigated. The vortex structures behind the shock interaction and wall pressure variations have also been studied.

• Korea-Australia Rheology Journal
• /
• v.15 no.2
• /
• pp.83-90
• /
• 2003

In cases of the microfluidic channel, the electrokinetic influence on the transport behavior can be found. The externally applied body force originated from the electrostatic interaction between the nonlinear Poisson-Boltzmann field and the flow-induced electrical field is applied in the equation of motion. The electrostatic potential profile is computed a priori by applying the finite difference scheme, and an analytical solution to the Navier-Stokes equation of motion for slit-like microchannel is obtained via the Green's function. An explicit analytical expression for the induced electrokinetic potential is derived as functions of relevant physicochemical parameters. The effects of the electric double layer, the zeta potential of the solid surface, and the charge condition of the channel wall on the velocity profile as well as the electroviscous behavior are examined. With increases in either electric double layer or zeta potential, the average fluid velocity in the channel of same charge is entirely reduced, whereas the electroviscous effect becomes stronger. We observed an opposite behavior in the channel of opposite charge, where the attractive electrostatic interactions are presented.

• Journal of Mechanical Science and Technology
• /
• v.15 no.6
• /
• pp.788-795
• /
• 2001

The effects of nonequilibrium condensation on the shock boundary layer interaction over a transonic bump model were investigated experimentally and numerically. An experiment was conducted using a supersonic indraft wind tunnel. A droplet growth equation was incorporated into two-dimensional Navier-Stokes equation systems. Computations were carried out using a third-order MUSCL type TVD finite-difference scheme with a second-order fractional time step. Computation compared with the experimental results. Nonequilibirum condensation suppressed the boundary layer separation and the pressure fluctuations due to the shock boundary layer interaction. Especially the nonequilbrium condensation was helpful to suppress the high frequency components of the pressure fluctuations.

• Journal of Advanced Marine Engineering and Technology
• /
• v.29 no.6
• /
• pp.637-644
• /
• 2005

This paper describes the numerical simulation of the interaction between longitudinal vortices ("common flow up") and a 3-D turbulent boundary layer over a flat plate To analyze the common flow up Produced from vortex generators. the flow field behind the vortex generators Is modeled by the information that is available from studies on a half-delta winglet. Also. the Reynolds-averaged Navier-Stokes equation for three-dimensional turbulent flows. together with a two-layer turbulence model to resolve the near-wall flow, is solved by the method of AF-ADI. The computational results predict that the boundary layer is thinned in the regions where the secondary flow is directed toward the wall and thickened where it is directed away from the wall Also. the numerical results. such as Reynolds stresses. turbulent kinetic energy and skin friction characteristics generated from the vortex generators . are reasonably close to the experimental data.

• Proceedings of the KSME Conference
• /
• 2000.11b
• /
• pp.544-549
• /
• 2000

The effects of nonequilibrium condensation on the shock boundary layer interaction over a transonic bump model were investigated experimentally and numerically. An experiment was conducted using a supersonic indraft wind tunnel. A droplet growth equation was incorporated into two-dimensional Navier-Stokes equation systems. Computations were carried out using a third-order MUSCL type TVD finite-difference scheme with a second-order fractional time step. Computations compared with the experimental results. Nonequilibirum condensation suppressed the boundary layer separation and the pressure fluctuations due to the shock boundary layer interaction. Especially the nonequilibrium condensation was helpful to suppress the high frequency components of the pressure fluctuations.

• Bulletin of the Korean Space Science Society
• /
• 2004.10b
• /
• pp.328-331
• /
• 2004

Shock wave boundary layer interactions can make flows around a vehicle be very high pressure and temperature due to pass shock waves in small areas of the hypersonic vehicle. These phenomena can affect a critical problem in the design of hypersonic vehicles. To research the effect of shock wave boundary layer interactions, comer flows were studied in this paper using numerical studies with the NSMB (Navier-Stokes Multi Block) solver and then comparing corresponding numerical results with experimental data of the Huston High Speed Flow Field Workshop II. The mach number of flows is 12.3 in comer flows. The comparison with the computational result is presented based on diverse numerical schemes. Good agreement is obtained.

• Transactions of the Korean Society of Mechanical Engineers
• /
• v.16 no.10
• /
• pp.1915-1927
• /
• 1992

A three-dimensional Navier-Stokes code has been developed for analysis of viscous flows through turbomachinery blade rows or other internal passages. The Navier-Stokes equations are written in a cartesian coordinate system, then mapped to a general body-fitted coordinate system. Streamwise viscous terms are neglected and turbulent effects are modeled using the baldwin-Lomax model. Equations are discretized using finite difference method on the stacked C-type grids and solved using LU-ADI decomposition scheme. calculations are made for a two-dimensional cascade in a transonic wind-tunnel to see the infuence of the endwalls. The flow pattern of the three-dimensional flow near the endwall is found to be different from that of the two-dimensional flow due to the existence of the endwalls.

• Journal of Navigation and Port Research
• /
• v.27 no.5
• /
• pp.473-478
• /
• 2003

Numerical study was made on the flow characteristics around a circular pipeline between parallel walls. The incompressible Navier-Stokes equations were solved by using a third-order upwind differential scheme. When the distance near a wall is small enough, the vortex shedding is almost completely suppressed because of the interaction with the wall boundary layer separation. This study aims to clarify the characteristics of the vortex shedding regime as the body approaches a wall as Reynolds number varies. The feature of separated vorticity dynamics is analyzed at different conditions with particular attention to the interaction between the pipeline wake and the induced separation on the plane walls.

• Proceedings of the Korean Society of Propulsion Engineers Conference
• /
• 2004.03a
• /
• pp.426-432
• /
• 2004

Large-Eddy Simulation (LES) is applied for the simulation of compressible flat plate boundary with Reynolds number up to 5 X 10$^{5}$ . Numerical examples include shock/boundary layer interaction and boundary layer transition, aiming future application to the analysis of transonic fan/compressor cascades. The present LES code uses hybrid com-pact/WENO scheme for the spatial discretization and compact diagonalized implicit scheme for the time integration. The present code successfully predicted the bypass transition of subsonic boundary layer. As for supersonic turbulent boundary layer, mean and fluctuation velocity of the attached boundary, as well as the evolution of the friction coefficient and the displacement thickness both upstream and downstream of the separation region are all in good agreement with experiment. The separation point also agreed with the experiment. In the simulation of the shock/laminar boundary layer interaction, the dependence of the transition upon the shock strength is reproduced qualitatively, but the extent of the separation region is overpredicted. These numerical examples show that LES can predict the behavior of boundary layer including transition and shock interaction, which are hardly managed by the conventional Reynolds-averaged Navier-Stokes approach, although there needs to be more effort before achieving quantitative agreement.

• Journal of Mechanical Science and Technology
• /
• v.15 no.5
• /
• pp.646-655
• /
• 2001

This paper reports the effects of nozzle exit boundary layer swirl on the instability modes of underexpanded supersonic jets emerging from plane rectangular nozzles. The effects of boundary layer swirl at the nozzle exit on thrust and mixing of supersonic rectangular jets are also considered. The previous study was performed with a 30°boundary layer swirl (S=0.41) in a plane rectangular nozzle exit. At this study, a 45°boundary layer swirl (S=1.0) is applied in a plane rectangular nozzle exit. A three-dimensional unsteady compressible Reynolds-Averaged Navier-Stokes code with Baldwin-Lomax and Chiens $\kappa$-$\xi$ two-equation turbulence models was used for numerical simulation. A shock adaptive grid system was applied to enhance shock resolution. The nozzle aspect ratio used in this study was 5.0, and the fully-expanded jet Mach number was 1.526. The \"flapping\" and \"pumping\" oscillations were observed in the jets small dimension at frequencies of about 3,900Hz and 7,800Hz, respectively. In the jets large dimension, \"spanwise\" oscillations at the same frequency as the small dimensions \"flapping\" oscillations were captured. As reported before with a 30°nozzle exit boundary layer swirl, the induction of 45°swirl to the nozzle exit boundary layer also strongly enhances jet mixing with the reduction of thrust by 10%.