• Title/Summary/Keyword: boundary layer flow separation

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Transition Flow Analysis According to the Change of Reynolds Number for Supersonic Launch Vehicle Fairing Expansion Area (초음속 발사체 선두 팽창부의 레이놀즈수 변화에 따른 천이 유동 해석)

  • Shin, Ho-Cheol;Park, Soo-Hyung;Byun, Yung-Hwan
    • 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.

Analysis of Laminar Flow and Heat Transfer in Asymmetric, Sudden Expansion Channel (비대칭급확대채널의 층류유동 및 열전달 해석)

  • Won, Seung-Ho;Maeng, Joo-Sung;Son, Byung-Jin
    • The Magazine of the Society of Air-Conditioning and Refrigerating Engineers of Korea
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    • v.13 no.1
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    • pp.5-13
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    • 1984
  • This analysis of numerical procedure is prediction of laminar flow and heat transfer at two dimension and steady flow in asymmetric sudden expansion channel. At former study, to analyse the flows with separation, the full Navier-Stokes equation is used, but there are many difficulties to analyse, and although significant progress has been made in the development of efficient computational methods for the Navier-Stokes equations, very large computation times are still required. In case of reward-facing flow, boundary-layer equation is used instead of full Navier-Stokes equation to analyse velocity fields, and result of this numerical analysis is good agreement with the given experimental study. In this case, since the computer time required for the boundary-layer calculation is an order of magnitude less than required for the solution of the full Navier-Stokes equation, this boundary-layer model provides a good approximate solution.

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Frequency Effects of Upstream Wake and Blade Interaction on the Unsteady Boundary Layer Flow

  • Kang, Dong-Jin;Bae, Sang-Su
    • Journal of Mechanical Science and Technology
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    • v.16 no.10
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    • pp.1303-1313
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    • 2002
  • Effects of the reduced frequency of upstream wake on downstream unsteady boundary layer flow were simulated by using a Wavier-Stokes code. The Wavier-Stokes code is based on an unstructured finite volume method and uses a low Reynolds number turbulence model to close the momentum equations. The geometry used in this paper is the MIT flapping foil experimental set-up and the reduced frequency of the upstream wake is varied in the range of 0.91 to 10.86 to study its effect on the unsteady boundary layer flow. Numerical solutions show that they can be divided into two categories. One is so called the low frequency solution, and behaves quite similar to a Stokes layer. Its characteristics is found to be quite similar to those due to either a temporal or spatial wave. The low frequency solutions are observed clearly when the reduced frequency is smaller than 3.26. The other one is the high frequency solution. It is observed for the reduced frequency larger than 7.24. It shows a sudden shift of the phase angle of the unsteady velocity around the edge of the boundary layer. The shift of phase angle is about 180 degree, and leads to separation of the boundary layer flow from corresponding outer flow. The high frequency solution shows the characteristics of a temporal wave whose wave length is half of the upstream frequency. This characteristics of the high frequency solution is found to be caused by the strong interaction between unsteady vortices. This strong interaction also leads to destroy of the upstream wake strips inside the viscous sublayer as well as the buffer layer.

The Effects of Nonequilibrium Condensation on Shock/Boundary Layer Interaction

  • Kim, Heuy-Dong;Lee, Kwon-Hee;Toshiaki. Setoguchi
    • Journal of Mechanical Science and Technology
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    • v.15 no.6
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    • pp.788-795
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    • 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.

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The Effect of Nonequilibrium Condensation on Shock/Boundary Layer Interaction (비평형응축이 충격파와 경계층의 간섭에 미치는 영향)

  • Kim, H.D.;Lee, K.H.;Setoguchi, T.
    • Proceedings of the KSME Conference
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    • 2000.11b
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    • pp.544-549
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    • 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.

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Application of the Scaling Law for Swept Shock/Boundary-Layer Interactions

  • Lee, Yeol
    • Journal of Mechanical Science and Technology
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    • v.17 no.12
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    • pp.2116-2124
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    • 2003
  • An experimental study providing additional knowledge of quasi-conical symmetry in swept shock wave/turbulent boundary-layer interactions is described. When a turbulent boundary layer on the flat plate is subjected to interact with a swept planar shock wave, the interaction flowfield far from fin leading edge has a nature of conical symmetry, which topological features of the interaction flow appear to emanate from a virtual conical origin. Surface streakline patterns obtained from the kerosene-lampblack tracings have been utilized to obtain representative surface features of the flow, including the location of the virtual conical origin. The scaling law for the sharp-fin interactions suggested by previous investigators has been reexamined for different freestream Mach numbers. It is noticed that the scaling law reasonably agrees with the present experimental data, however, that the law is not appropriate to estimate the location of the virtual conical origin. Further knowledge of the correlation for the virtual conical origin has thus been proposed.

LOW-SPEED AERODYNAMIC CHARACTERISTIC OF TRANSITION FLOW OVER THE NACA0012 (NACA0012 천이 유동의 저속 공력 특성 해석)

  • Jeon, Sang-Eon;Park, Soo-Hyung;Kim, Sang-Ho;Byun, Yung-Hwan;Jung, Kyung-Jin;Kang, In-Mo
    • Journal of computational fluids engineering
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    • v.15 no.3
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    • pp.1-8
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    • 2010
  • Laminar separation bubble and transitional flow over the NACA0012 are investigated at a moderate range of Reynolds numbers. A Reynolds-Averaged Navier-Stokes code is coupled with an empirical transition model that can predict transition onset points and the length of transition region. Without solving the boundary layer equations, approximated e-N method is directly applied to the RANS code and iteratively solved together. The computational results are compared with the experimental data for the NACA0012 airfoil. Results of transition onset point and the length are compared well with experimental data and Xfoil prediction. The present RANS results show at high angles of attack better agreement with experimental data than Xfoil results using the boundary layer equations.

Flow Visualization of an Unsteady Airfoil at Low Reynolds Numbers (저 레이놀즈수에서 비정상 에어포일의 흐름 가시화)

  • Kim, Dong-Ha;Chang, Jo-Won
    • Journal of the Korean Society of Visualization
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    • v.4 no.2
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    • pp.51-58
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    • 2006
  • A boundary layer visualization was carried out in order to investigate the influence of Reynolds number on an oscillating airfoil. An NACA 0012 airfoil is sinusoidally pitched at the quarter chord point with oscillation amplitude of ${\pm}6^{\circ}$. A smoke-wire technique was employed to visualize the boundary layer and the near-wake. The freestream velocities are 1.98, 2.83 and 4.03m/s and corresponding chord Reynolds numbers are $2.3{\times}10^4,\;3.3{\times}10^4$, and $4.8{\times}10^4$, respectively. As the reduced frequency of K=0.1 is fixed, the corresponding frequency of an airfoil was adjusted in each case. The results reveal that the point at which the shear stress in an unsteady boundary layer separation disappears does not correspond with the position of the breakdown of the boundary layer, and that the breakdown of the boundary layer occurs further downstream.

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Compressor Cascade Flow Analysis by Using Upwind Flux Difference Splitting Method (풍상차분법을 이용한 압축기 익렬유동 해석)

  • 권창오;송동주;강신형
    • Transactions of the Korean Society of Mechanical Engineers
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    • v.18 no.3
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    • pp.653-661
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    • 1994
  • In this paper the CSCM type upwind flux difference splitting Navier-Stokes method has been applied to study the ARL-SL19 supersonic/transonic compressor cascade flow. H-type grid was chosen for its simplicity in applying cyclic tridiagonal matrix algorithm along with conventional slip/no-slip boundary conditions. The thin-layer algebraic model of Baldwin-Lomax was employed for the calculation of turbulent flows. The test case inlet Mach No. was 1.612 and inlet/exit pressure ratio($P_2/P_1$) was 2.15. The results were compared with experimental results from current method were compared well in suction surface with the experiments and other computational results; however, not well in pressure surface. It might be due to the complex flowfields such as shock/boundary layer interaction, turbulence, and flow separation, etc. In the future, a proper turbulence modelling and adaptive grid system will be studied to improve the solution quality.

Numerical Simulation of the Aeroacoustic Noise in the Separated Laminar Boundary Layer

  • Park, Hyo-Won;Young J. Moon;Lee, Kyu-Jung
    • Journal of Mechanical Science and Technology
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    • v.17 no.2
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    • pp.280-287
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    • 2003
  • The unsteady flow characteristics and the related noise of separated incompressible laminar boundary layer flows (Re$\sub$$\delta$/* = 614, 868, and 1,063) are numerically investigated. The characteristic lines of the wall pressure are examined to identify the primary noise source, related with the unsteady motion of the vortex at the reattachment point of the separation bubble. The generation and propagation of the vortex-induced noise in the separated laminar boundary layer are computed by the method of Computational Aero-Acoustics (CAA), and the effects of Reynolds number, Mach number and adverse pressure gradient strength are examined.