• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.46 no.5
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• pp.359-367
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• 2018

The vortex ring state that occurs during the descending flight of a rotorcraft generates a circulating flow like a donut near the rotating surface, and it often causes a rotorcraft fall due to loss of thrust. In this paper, we have physically identified the flow field in the vortex ring state of the quadcopter, one of the types of unmanned aerial vehicles. The descending flight of the quadcopter was simulated in a 1m subsonic wind tunnel of the Korea Aerospace Research Institute(KARI) and the Particle Image Velocimetry(PIV) was used for the flow field measurement. The induced velocity in the hovering state is estimated by using the momentum theory and the test was carried out in the range of descent rate at which the vortex ring condition could be caused. The development and the direction of the vortex ring were confirmed by the measurement of the flow field according to not only the descent rate but also propeller separation distance. In addition, the results of the study show the vortex ring state can be predicted sufficiently by measuring the flow velocity around the quadcopter.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.25 no.12
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• pp.1821-1830
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• 2001

The flow around free end of a finite circular cylinder (FC) embedded in an atmospheric boundary layer has been investigated experimentally. The experiments were carried out in a closed-return type subsonic wind tunnel with varying aspect ratio of the finite cylinder mounted vertically on a flat plate. The wakes behind a 2-D cylinder and a finite cylinder located in a uniform flow were measured for comparison. Reynolds number based on the cylinder diameter was about Re=20,000. A hot-wire anemometer was employed to measure the wake velocity and the mean pressure distributions on the cylinder surface were also measured. The flow past the FC free end shows a complicated three-dimensional wake structure and flow phenomenon is quite different from that of 2-D cylinder. The three-dimensional flow structure was attributed to the downwashing counter rotating vortices separated from the FC free end. As the FC aspect ratio decreases, the vortex shedding frequency decreases and the vortex formation length increases compared to that of 2-D cylinder. Due to the descending counter-rotating twin-vortex, near the FC free end, regular vortex shedding from the cylinder is suppressed and the vortex formation region is hardly distinguished. Around the center of the wake, the mean velocity for the FC located in atmospheric boundary layer has large velocity deficit compared to that of uniform flow.

• Proceedings of the KSME Conference
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• 2004.11a
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• pp.1670-1675
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• 2004

A thermal vapor compressor in which the subsonic/supersonic flow appears simultaneously, has been accurately designed through the CFD analysis for the various shape parameters such as the primary nozzle shape, converging duct shape. mixing tube diameter, and so on. The performance of the developed thermal vapor compressor has been experimentally verified to be installed in a Multi Effect Desalination(MED) plant as an important element, In this paper, the experimental results for Various boundary conditions(motive pressure, suction pressure, and discharge pressure) are presented in comparing with CFD results. The two results show a good agreement with each other within 3.5 % accuracy with regard to the entrainment ratio.

• Transactions of the Korean Society for Noise and Vibration Engineering
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• v.14 no.1
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• pp.50-55
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• 2004

In this study, slit-tones by plane impinging jet are investigated experimentally over the whole subsonic flow range, especially at low speeds, in order to obtain the instability behaviour of impinging plane jet. Slit-tones are generated at low speeds associated with laminar shear layer instability as well as at high speeds associated with turbulent instability. Most of low-speed slit-tones are induced by symmetric mode instability unless the slit is not so wide, in which case antisymmetric modes are induced like edge-tones. It is found that the frequencies at low speeds ate controled by the unstable condition of the vortex at the nozzle exit and its pairings by which the frequencies are decreased by half. In the case of symmetric modes related with low-speed slit-tones, frequencies lower than those associated with one-step pairings are not found.

• Transactions of the Korean Society of Mechanical Engineers
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• v.18 no.7
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• pp.1808-1817
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• 1994

Unsteady aerodynamic analysis of an aircraft is done using a frequency domian 3-D panel method. The method is based on an unsteady linear compressible lifting surface theory. The lifting surface is placed in a flight patch, and angle of attack and camber effects are implemented in upwash. Fuselage effects are not considered. The unsteady solutions of the code are validated by comparing with the solutions of a hybrid doublet lattice-doublet point method and a doublet point method for various wing configurations at subsonic and supersonic flow conditions. The calculated results of dynamic stability derivatives for aircraft are shown without comparision due to lack of available measured data or calculated results.

• Proceedings of the Korean Society for Noise and Vibration Engineering Conference
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• 2001.05a
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• pp.420-425
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• 2001

Theoretical analysis of noise reduction by a membrane-duct system is presented. When acorn waves propagate in the membrane-duct, the membrane is also excited and its motion is coup with interior medium. It has been shown that propagating waves with supersonic wave speed exist beyond a certain critical frequency that is determined from the mass ratio of the me and the fluid. Also found are subsonic waves which couple strongly wi th the membrane a provide a powerful mechanism of energy dissipation. Existence of an exterior medium alter dispersion characteristics. It provides additional mass loading and reduces the subsort speed further. The effect of mean flow speed on dispersion characteristics is also consider results show that the membrane-duct system can be applied to diminish and absorb 1 frequency noise in duct instead of passive muffler, such as a simple expansion chamber absorption material.

• 한국전산유체공학회:학술대회논문집
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• 1999.11a
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• pp.29-35
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• 1999

A two-dimensional Euler code based on flux vector splitting scheme has been developed to simulate the behavior of supersonic shock wave over the cylinder. AF+ADI scheme was used for time integration. The sliding multiblock technique was implemented to handle the relative motion of the moving cylinder and the stationary tunnel. The code is validated with a problem of subsonic flow around a Naca-0012 airfoil. The Computation results show complex phenomena of the propagation of shock waves and the reflection as expansion wave at tunnel exit.

• Proceeding of EDISON Challenge
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• 2012.04a
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• pp.113-116
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• 2012

본 연구에서는 마하수 변화에 따른 층류유동 변화를 살펴보았다. 해석 프로그램은 EDISON_CFD를 이용하고, EDISON_CFD에서 사용한 수치기법과 Scheme에 대해서 언급한다. CFD기법을 이용하여 해석한 결과를 경계층조건의 이론 해석방법인 Blasius Boundary Layer와 비교하였다. 각 요소마다 해석한 결과를 통하여 층류 경계층의 특성을 살펴보았다. 그 결과 마하수 증가에 따른 평판의 온도 증가와 밀도 감소가 경계층을 선형적으로 증가시키는 것을 보았다. 또한 마하수 증가에 따른 점성계수의 증가를 살펴봄으로서 층류유동에서 마하수의 증가는 점성에 의한 운동량 확산을 증가시킨다는 것을 보았다.

• Journal of The Korean Astronomical Society
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• v.37 no.4
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• pp.205-210
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• 2004

The interaction of disk galaxies with intergalactic winds has been invoked as a possible mechanism of the generation of galactic warps. Here we discuss conditions under which intergalactic flows can be relevant for warping field galaxies. Constraints include the heating of the outer disk, the level of asymmetry in the vertical distribution of the volume gas density, the angular frequency of the warp, the symmetry of galactic warps amplitude between the approaching and receding sides of the galaxy, and the speed of the intergalactic flow whether subsonic or supersonic. These constraints are discussed in this paper in reference to the proposal of Lopez-Corredoira et al. that warps can be a natural consequence of accretion flows onto the disk.

• Journal of the Korean Society of Propulsion Engineers
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• v.6 no.2
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• pp.9-17
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• 2002

For turbulent spray combustion flows a coupled numerical procedure was developed, This method was discretized by using generalized curvilinear coordinates to handle complex geometries. The preconditioning and eigenvalue rescaling techniques were employed to provide efficient convergences over a wide range of subsonic Mach numbers. The accuracy was validated by simulating the laminar cavity flow. The film cooling effect of a liquid rocket engine (KSR-III) were investigated by a spray combustion analysis. The film cooling showed a negative effect on the combustion efficiency. In the combustion chamber wall, the film cooling effect was revealed to be promoted by the production of fuel rich zone.