• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.37 no.10
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• pp.984-991
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• 2009

In this study, transonic aeroelastic response analyses have been conducted for the DLR-F4(wing-body) aircraft configuration considering shockwave and flow separation effects. The developed fluid-structure coupled analysis system is applied for aeroelastic computations combining computational structural dynamics(CSD), finite element method(FEM) and computational fluid dynamics(CFD) in the time domain. It can give very accurate and useful engineering data on the structural dynamic design of advanced flight vehicles. For the nonlinear unsteady aerodynamics in high transonic flow region, Navier-Stokes equations using the structured grid system have been applied to wing-body configurations. In transonic flight region, the characteristics of static and dynamic aeroelastic responses have been investigated for a typical wing-body configuration model. Also, it is typically shown that the current computation approach can yield realistic and practical results for aircraft design and test engineers.

• Proceedings of the Korean Institute of Navigation and Port Research Conference
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• v.1
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• pp.45-49
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• 2006

The flow around a foil with waterjet was investigated using the two-frame PIV(CACTUS 3.1) system. After separation, unsteady recirculation & reattachment region was shown as a result at reading edge. Separation area was decreased to 1/3 more by waterjet system with coanda effect. Angle of attack and waterjet velocity was a variable in the experiment. Each parameters was controlled to $0^{\circ}{\sim}35^{\circ}$ and 0[m/s]${\sim}$9.2[m/s]. The separation of flow appearanced at first when the angle of attack is $17^{\circ}{\sim}18^{\circ}$. However, according to grew up of velocity, beginning of the separation was delayed. In this experiment, vortex and separation region was disappeared by blown when each parameters are low level, and separation controlled more certainly.

• Journal of Ocean Engineering and Technology
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• v.7 no.2
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• pp.30-36
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• 1993

Incompressible, unsteady flow around various single rectangular cylinders of side ratios ranging from 0.005 to 2.0 immersed in uniform flow is computed by the vortex tracing me thod. Results with and without a splitter plate pttached to the rear center of the cylinder are compared. The objective of this study is to investigate predictability of the effects of the splitter plate on drag by the method. Without the splitter plate, computed drag coefficients for cylinders of large side ratios are in good agreement with measured values, but are over predicted for those of small side ratios. With the splitter plate, drag coefficient is reduced significantly due to suppression of vortex growing near the base and interaction between the separated shear layers.

• Transactions of the Korean Society of Mechanical Engineers
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• v.6 no.2
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• pp.147-159
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• 1982

내부에 정방형, 외부에 원형인 두 정온동심수평관으로 이루어진 폐쇄공간에서 열부력으로 야기된 대류열전달 현상을 연구하였다. 주어진 비정규적 형상에서의 층류 정상 Boussinesq 유동을 해 석하기 위하여, 지구중력방향에 대하여 가능한 두 개의 대칭형 위치에 관하여, Galerkin 유한요 소법을 사용하였다. 이론의 결과를 확증하기 위하여 실험적으로도 온도측정과 유선의 가시화를 수행하였다. 정방형내관의 직각인 선단은 국소 및 총체열전달에 있어서 소극적인 역할을 하지만, 이들로 인한 경계층 유동의 박리는 발생하지 아니함을 보였다. 이 내관의 상부 수평면상에서는 유동속도와 온도구배가 낮기는 하나 확연한 대칭형의 열상승류(Plume)가 가능하였다. 내관의 벽면들이 지구중력방향에 관하여 .+-.45.deg.의 각을 이룰 경우 과류의 중심부가, Rayleigh수가 6.5*$10^{4}$보다 작을 때는 4개, 이보다 클때는 2개가 폐쇄공간 안에서 발생하였다.

• Proceedings of the Korean Society for Noise and Vibration Engineering Conference
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• 1996.10a
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• pp.65-70
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• 1996

엔진소음을 소음특성에 따라 분류하면 공력소음(Aerodynamic Noise), 연소소음(Combustion Noise), 기계적인 소음(Mechanical Noise)으로 나눌 수 있으며 소음원의 종류에 따라 분류하면 배기계소음(Exhaust System Noise)으로 나눌 수 있으며 소음원의 종류에 따라 분류하면 배기계소음(Exhaust System Noise), 흡기계소음(Intake System Noise), 냉각계소음(Cooling System Noise), 엔진표면소음(Engine System Noise)등으로 분류할 수 있다. 이러한 여러소음중 엔진 내부의 유동에 의한 흡배기계통으로의 소음방출은 자동차 실 내외 소음의 중요한 문제로 대두되는데, 이를 줄이기 위해 그 동안 소음기 등의 서브시스템의 형태와 그 위치조정에 관한 연구가 수행되어 왔다. 그러나 이것이 비용 또는 성능에 영향을 미치므로 본질적인 소음원을 규명해 내는 것이 필요하게 되었다. 흡배기계의 소음은 엔진의 흡입, 배기행 정시 피스톤의 운동에 의해 팽창 및 압축파 형태의 압력파(pressure wave)로 발생하게 되고, 밸브근방에서는 유동의 박리(separation)에 의해 발생하게 된다. 소음기 등의 서브시스템에서도 유동의 박리에 의해 발생하게 되며 특히 배기행정시 발생하는 압력파는 비선형영역에 있게된다. 흡기소음은 배기에 비해 그 크기가 작아서 그동안 등한시 되어왔으나 이것이 소비자의 불평요인으로 작용하므로써 이에 대한 연구도 활발히 수행되어야 한다. Bender, Bramer[1]는 흡배기계 소음의 외부 방사에 관하여 전반적으로 기술하였고 Sierens등[2]은 흡기계에서 1차원 MOC(Method of Characteristics)방법으로 비정상 유동해석을 하고 실험결과와 비교하였다. J.S.Lamancusa 등[3]은 흡기 소음원을 실험을 통해 예측하였고, 흡기소음도 비선형 거동을 보인다고 밝혔다. Yositaka Nishio 등[4]은 새로운 흡기실험장치를 고안하여 공명기(resonator)의 위치 변화에 의한 저소음 흡기계를 설계 초기단계에서부터 적용하려 하였다. 일반적으로 흡배기계의 복잡한 형상 때문에 대부분 실험을 통해 문제를 해결하려 하였고, 수치해석은 피스톤의 운동을 배제한 단순화한 흡배기계의 정상상태 유동해석이 주를 이루어왔다. Taghaui and Dupont 등[5]은 KIVA코드를 사용하여 흡기포트와 연소실 그리고 밸브의 움직임을 동시에 고려한 수치해석을 도입하였다. 하지만 이들이 밸브의 운동을 고려하기 위해 사용한 이동격자는 격자점은 시간에 따라 변화하지만 그 격자의 수가 일정하게 유지되어 있어서 밸브의 완전개폐를 해석할 수가 없다. 강희정[6]은 단일 실린더와 단일 배기밸브를 갖는 문제로 단순화하여 피스톤과 밸브의 움직임을 고려하므로써 배기행정 후 소음이 어떻게 전파해 나가는가를 연구하였다. 본 연구에서도 최소밸브간격과 최대밸브간격 사이에서만 계산이 가능하나 흡기의 경우는 밸브가 닫힐 때 생기는 압력파가 중요하므로 실린더와 밸브사이에 벽면조건을 주어 밸브의 개폐를 모사하였다.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.34 no.4
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• pp.53-62
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• 2006

Two-dimensional steady flowfields generated by transverse injection jets into a supersonic mainstream are numerically simulated. Fine-scale turbulence effects are represented by a k-${\omega}$ SST two-equation closure model which includes $y^+$ effects on the turbulence model. Solution convergence is evaluated by using Grid Convergence Index(GCI), a measure of uncertainty of the grid convergence. Comparison is made with experimental data and other turbulence models in term of surface static pressure distributions, the length of the upstream separation region, and the penetration height. Results indicate that the k-${\omega}$ SST model correctly predicts the mean surface pressure distribution and the upstream separation length for low static pressure ratios. However, the numerical predictions become less consistent with experimental results as the static pressure ratio increases. All these results are taken within 1% error band of grid convergence.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.31 no.6
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• pp.1-7
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• 2003

In this study, steady/unsteady aerodynamic characteristic for three dimensional symmetric wing was investigated numerically using Vortex Panel Method. This program utilized linearly varying vortices in x and y directions distributed on the wing surface and was applied to the incompressible potential. flow around a three dimensional wing Separation and deformation of the wake are not considered. The comparison between NACA Airfoil Data and the computed results showed excellent agreement. πus method was applied to unsteady wings undergoing both sudden pitch-up and constant rate pitching motion. In the unsteady flow analysis, a formation and a time-dependent locations of Starting Vortices are considered and the effect of Starting Vortices on aerodynamic characteristic of the wing was calculated. The present method can be extended to apply for more complicated cases such as pitching, flapping and rotating wing analysis.

• The Journal of the Acoustical Society of Korea
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• v.40 no.4
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• pp.261-269
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• 2021

Without any validation of the incompressible assumption, most of previous studies on cavitation flow and its noise have utilized numerical methods based on the incompressible Reynolds Average Navier-Stokes (RANS) equations because of advantage of its efficiency. In this study, to investigate the effects of the flow compressibility on the Tip Vortex Cavitation (TVC) flow and noise, both the incompressible and compressible simulations are performed to simulate the TVC flow, and the Ffowcs Williams and Hawkings (FW-H) integral equation is utilized to predict the TVC noise. The DARPA Suboff submarine body with an underwater propeller of a skew angle of 17 degree is targeted to account for the effects of upstream disturbance. The computation domain is set to be same as the test-section of the large cavitation tunnel in Korea Research Institute of Ships and Ocean Engineering to compare the prediction results with the measured ones. To predict the TVC accurately, the Delayed Detached Eddy Simulation (DDES) technique is used in combination with the adaptive grid techniques. The acoustic spectrum obtained using the compressible flow solver shows closer agreement with the measured one.

• The Journal of the Acoustical Society of Korea
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• v.42 no.3
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• pp.250-261
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• 2023

In this study, the noise contributions by the duct, stator and rotor, which are the propulsor components, are evaluated to identify the flow noise source in cavitation and non-cavitation conditions on pumpjet propulsion and the noise levels in both conditions are compared. The unsteady incompressible Reynolds averaged Navier-Stokes (RANS) equation based on the homogeneous mixture assumption is applied on the suboff submarine hull and pumpjet propeller in the cavitation tunnel, and the Volume of Fluid (VOF) method and Schnerr-Sauer cavitation model are used to describe the two-phase flow. Based on the flow simulation results, the acoustic analogy formulated by Ffowcs Williams and Hawkings (FW-H) equation is applied to predict the underwater radiated noise. The noise contributions are evaluated by using the three types of impermeable integral surface on the duct, stator and rotor, and the two types of permeable integral surface surrounding the propulsor. As a result of noise prediction, the contribution by the stator is insignificant, but it affects the generation of flow noise source due to flow separation in the duct and rotor, and the noise is predominantly radiated into the upward and right where the flow separations are. Also, the noise is radiated into the thrust direction due to pressure fluctuation between suction and pressure sides on the rotor blades, and the it can be seen that the cavitation effect into the noise can be considered through the permeable integral surface.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.34 no.5
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• pp.12-18
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• 2006

A new steady approach has been developed to predict the transonic buffet onset of a high speed aircraft. In this paper, the flow is assumed to be steady for the buffet onset. The present study involves the analysis of a distinct change in the variation of various static aerodynamic parameters. These distinct changes indicates the onset of transonic buffet. Among the various aerodynamic parameters considered in this study, the variation in the center of pressure has shown to provide a clearest indicator of transonic buffet onset. This new steady approach can be applied to predict the transonic buffet onset for airfoils with shock induced separation bubble and for large swept wings with small aspect ratios. Good agreements have been obtained compared with unsteady wind tunnel buffet test data. Based on the results obtained the new steady approach, it can be newly suggested that the distinct slope changes of the center of pressure curve can be used as an indicator of buffet onset for the steady experimental method on a full aircraft configuration.