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이재훈(Jaehun Lee),정경진(Kyoung Jin Jung),이길태(Kil Tae Lee),강인모(Inmo Kang) 한국전산유체공학회 2009 한국전산유체공학회 학술대회논문집 Vol.2009 No.11
In this study, two and three dimensional low Reynolds number flows are compared. For the two dimensional flow, an airfoil was considered and for the three dimensional low wing and full-body aircraft were considered. Because a flight condition of the aircraft is in a low Reynolds number flow, itl requires reflecting flow transition. In the two dimensional analysis, transition is predicted using en method. In the three dimensional flow, the effect of transition is included using k-w SST turbulence models.
박동훈,심호준,권기정,김양원,이융교 한국항공우주학회 2015 한국항공우주학회 학술발표회 논문집 Vol.2015 No.11
저레이놀즈수 조건의 익형 유동 측정을 위한 PIV 시스템을 구축하였다. 시위 길이 150mm의 DAE51 익형 시험 모델을 제작하고 suction side 주변 및 후류 영역 유동장을 측정하였다. 측정 시스템의 공간분해능이 박리, 천이, 재부착 지점을 판별하여 박리거품 거동 분석에 활용 가능한 수준임을 확인하였다. 박리 영역 내에서 천이에 의해 층류 형태에서 난류 형태로의 속도 프로파일 변화를 확인하였다. 받음각 증가로 박리 영역의 크기 감소 및 상류 이동을 관찰하였다. PIV system is set up for measurement of flow around an airfoil at low Reynolds numbers. A test model of DAE51 airfoil is manufactured. Measurements of flow filed around the airfoil suction side and wake region are carried out. Spatial resolution of the measurement system is found to be adequate for investigation on behavior of separation bubble through specifying the location of separation, transition, and reattachment. Change of velocity profile from laminar to turbulent type due to transition is confirmed in separated region. Decrease of streamwise extent and upstream shift of separated region is observed as angle of attack increases.
박현욱(Hyun Wook Park),이창훈(Changhoon Lee),최정일(Jung-il Choi) 한국전산유체공학회 2013 한국전산유체공학회지 Vol.18 No.3
We develop a novel immersed boundary (IB) method based on implicit direct forcing scheme for incompressible flows. The proposed IB method is based on an iterative procedure for calculating the direct forcing coupled with the momentum equations in order to satisfy no-slip boundary conditions on IB surfaces. We perform simulations of two-dimensional flows over a circular cylinder for low and moderate Reynolds numbers. The present method shows that the errors for estimated velocities on IB surfaces are significantly reduced even for low Reynolds number with a fairly large time step while the previous methods based on direct forcing failed to provide no-slip boundary conditions on IB surfaces.
양재훈(Jae-Hun Yang),장조원(Jo-Won Chang) 한국항공우주학회 2006 韓國航空宇宙學會誌 Vol.34 No.10
저 레이놀즈수에서 NACA 0012 에어포일의 경계층 거동에 관한 연구가 터빈 블레이드와 초소형 비행체에서 적용될 수 있는 경계층을 파악하기 위하여 수행되었다. 레이놀즈수 Re=2.3×10⁴, 3.3×10⁴, 4.8×10⁴과 정적 받음각 α=0˚, 3˚, 6˚에서 경계층을 측정하기 위해 열선 풍속계가 사용되었다. 연구결과는 정적 받음각 0˚에서는 층류 경계층이 에어포일 표면에 부착되며, 정적 받음각 3˚에서는 경계층 층류 분리가 발생된 것을 보여준다. 더욱이 본 연구에서 경계층 재부착 현상은 정적 받음각 6˚에서 레이놀즈수 3.3×10⁴와 4.8×10⁴에서 발생된다. A study on the boundary layer behavior of an NACA 0012 airfoil at low Reynolds numbers was investigated in order to gain knowledge of a boundary layer that might be employed in a turbine blade and MAVs. A hot-wire anemometer was used to measure the boundary layer of an NACA 0012 airfoil at static angles of attack α=0˚, 3˚, and 6°, and Reynolds Numbers Re=2.3×10⁴, 3.3×10⁴, and 4.8×10⁴. The results of this study show that the laminar boundary layer on the airfoil surface is attached to the surface at α=0˚, and the laminar separation of the boundary layer on the airfoil surface occurs at α=3˚. Furthermore, the reattachment of the boundary layer in the present study occurs for the cases of Re=3.3×10⁴ and Re=4.8×10⁴ at α=6˚.
김동하(Dong Ha Kim),장조원(Jo Won Chang) 한국가시화정보학회 2006 한국가시화정보학회지 Vol.4 No.2
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 ±6°. A smoke-wire technique was employed to visualize the boundary layer and the near-wake. The freestream velocities are 1.98, 2.83 and 4.03㎧ and corresponding chord Reynolds numbers are 2.3×10⁴, 3.3×10⁴, and 4.8×10⁴, 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.
저 레이놀즈수에서 진동하는 에어포일의 비정상 경계층 측정
김동하(Dong-Ha Kim),장조원(Jo-Won Chang) 한국항공우주학회 2006 韓國航空宇宙學會誌 Vol.34 No.12
진동하는 에어포일에서 비정상 경계층의 거동을 조사하기 위하여 실험적 연구가 수행되었다. 가로세로비가 2.7인 NACA 0012 에어포일은 시험부에 수직으로 설치되었고, 1/4 시위에서 조화 피칭운동을 한다. 에어포일의 진동 진폭은 -6˚에서 +6˚까지 변화하며 평균 받음각은 0˚ 이다. 표변에 부착되는 프로브(글루온 프로브)가 경계층 표면 유동를 측정하기 위하여 이용되었다. 측정은 자유흐름속도는 1.98, 2.83, 4.03m/s에서 수행되었고, 시위길 이를 근거로 한 레이놀즈수는 각각 2.3×10⁴, 3.3×10⁴, 4.8×10⁴이다. 에어포일의 무차원 진동수를 모든 경우에서 0.1로 고정하였다. 비정상 경계층에서 최소 전단력의 위치와 경계층 붕괴의 위치는 레이놀즈수 2.3×10⁴와 3.3×10⁴ 사이에서 크게 다르게 나타난다. An experimental study was carried out to examine the behavior of the unsteady boundary layer. An NACA 0012 airfoil with aspect ratio of 2.7 was set vertically in a test section, which is sinusoidally pitched about the quarter chord. The oscillating amplitude is from -6˚ to +6˚ and the mean angle of attack is 0˚. Surface mounted probes (Glue-on probes) were employed to measure the surface flow of the boundary layer. Measurements were made at free-stream velocities of 1.98, 2.83, and 4.03㎧, and the corresponding Reynolds numbers based on the chord length were 2.3×10⁴, 3.3×10⁴ and 4.8×10⁴, respectively. The reduced frequency is fixed as 0.1 in all cases. The results show that the surface position of minimum shear stress and of boundary layer break-down can be discerned in the Reynolds number between 2.3×10⁴ and 3.3×10⁴.
A STUDY ON THE LOW REYNOLDS NUMBER AIRFOILS FOR THE DESIGN OF THREE DIMENSIONAL WING
K.J. Jung(정경진),J. Lee(이재훈),J.H. Kwon(권장력),I.M. Kang(강인모) 한국전산유체공학회 2009 한국전산유체공학회 학술대회논문집 Vol.2009 No.4
In this study, a generic airfoil designed by the inverse method was evaluated with several candidate airfoils as a first step. Each airfoil was compared with respect to aerodynamic performance to meet the requirement of HALE(high altitude long endurance) aircraft. The second step was to optimize the candidate airfoil using the couple of optimization formulations to down select an optimum airfoil. For the analysis of low Reynolds number 2D flow, Drela's MSES was used. After comparing the aerodynamic results, the best airfoil was chosen to construct the baseline 3D wing. The Navier-Stokes code was used to evaluate the overall aerodynamic performance of designed wing with other wings. The results show that the designed wing has the best performance compared with other wings.
소산률방정식의 개선을 통한 저레이놀즈수 k-ε모형의 개발
송경,유근종,조강래 창원대학교 공작기계기술연구센터 1999 연구업적집 Vol.1 No.1
Recent methods of developing k-εmodel have been carried out with the aid of DNS data to include the effect of near wall. Though these methods opened new way of turbulence modelings, newly developed turbulence models of its kind can't predict properly the turbulent flows with various Reynolds numbers and various geometric conditions. As a remedy for these shortcomings, new k.-ε model is presented here which includes the improved dissipation rate equation and the improved damping function for eddy viscosity model. The new dissipation rate equation was modeled based on the energy spectrum and magnitude analysis The damping function for eddy viscosity was also formulated on the ground of distribution of distribution of dissipation rate length scales near a wall and the DNS data. The new K-ε model was applied to the fully developed turbulent flows in a channel and a pipe with a wide range of Reynolds numbers. prediction results show that the present model represents properly the turbulence properties in all turbulent regions over a wide range of Reynolds numbers.