• Journal of the Korean Society of Radiology
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• v.9 no.7
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• pp.509-513
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• 2015

In this study, we evaluated the thermal denaturation characteristics of reusable NIPAM tissue mimicking (TM) Phantom by measuring the thermal sensitivity. And the changes of acoustic characteristic and thermal denaturation shape in NIPAM TM phantom according to the number of re-use time and re-use period were observed. With the result, as the sonication time is increased, the sound velocity of NIPAM phantom was decreased by 100 m/s and the attenuation was increased slightly. However, the changes according to the re-use period was not observed. In the thermal denaturation shape and size observation by ultrasound sonicaton, the remarkable changes have not been confirmed. With the result of this study, NIPAM Phantom was considered appropriate to evaluate and predict the effect of therapeutic ultrasound by in repeated sonication test.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.43 no.3
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• pp.224-231
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• 2015

An experimental study on aerodynamic coefficients of a lambda wing configuration was performed at the low speed wind tunnel of Agency for Defense Development. The main purpose of this study was to investigate the effects of sideslip angle on various aerodynamic coefficients. In the case of $0^{\circ}C$ sideslip angle, nose-up pitching moment rapidly increases at a specific angle of attack. This unstable pitching moment characteristic is referred to as pitch break or pitch up. As the sideslip angle increases, the pitch break is found to be generated at a higher angle of attack. Rolling moment is found to show similar behavior pattern to 'pitch break' style with angle of attack at non-zero sideslip angles. This trend gets severer at greater sideslip angles. Yawing moment also shows substantial variation of the slope and the unstable directional stability with sideslip angles at higher angles of attack. These characteristics of the three moments clearly implies the difficulty of the flight control which requires efficient control augmentation system.

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.6
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• pp.81-88
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• 2013

The component based propulsion modeling and simulation of an dual ramjet engine using Taylor-Maccoll flow equation and quasi 1-D combustor model. The subsonic and supersonic intake were modeled with Taylor-Maccoll flow having $25^{\circ}$ cone angle, the gas generator which transfers a pre-combustion gas into supersonic combustor was developed using Lumped model, and the determination of the size of nozzle throat of a gas generator was described. A quasi 1-D model was applied to model a supersonic combustor and the variation of temperature and pressure inside combustor were presented. Furthermore, the thrust and specific impulse applying fuel regulation by pressure recovery ratio and equivalence ratio were derived.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• v.y2005m4
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• pp.394-398
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• 2005

Ramjet engine have been usually operated on Mach $1.5\sim3$ as the vehicle of supersonic cruising engine and studied about the higher performance above Mach 4. The research of Duel mode Scramjet engine which have duel operating mode of ramjet/Scramjet are in progress actively nowadays. This paper suggests the effect the flow characteristics and the effects of back pressure, angle of attack, angle of yow on the supersonic air intake on mach 4 through the Schlieren/Oil flow visualization, and pressure measurement on experimental model.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.10
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• pp.112-117
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• 2004

This work is on numerical and experimental studies on the new type igniter using aerodynamic energy. The aerodynamic igniter consists of a nozzle and a resonance tube. The supersonic jet from a nozzle coming into the resonance tube generates pressure oscillation between the nozzle and the resonator. This oscillation changes the kinetic energy to thermal energy in the resonator under a certain condition. In this study, sonic and supersonic nozzles were tested in two different resonators, results has been compared. And geometrical optimum values of a supersonic nozzle has been suggested to reduce aerodynamic loss and friction in the expanded surface of the nozzle.

• Journal of Biomedical Engineering Research
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• v.19 no.1
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• pp.9-18
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• 1998

In this paper, we present the degradation of focusing induced by velocity inhomogeneity in human tissue. For simulation, the fatty layer which is the major factor of degradation for its lower velocity, is modeled as a uniform velocity perturbation layer. And we simulate the degradation of resolution resulting from change of beam path due to refraction and the time delay due to velocity difference. We show that focusing error can be compensated for considering the velocity inhomogeneity only. The proposed compensation method can be operated in real time in the presently used digital focusing systems.

• The Journal of the Acoustical Society of Korea
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• v.24 no.2
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• pp.87-96
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• 2005

In matched field processing (MEP), the observed acoustic field data is basically correlated with the replica produced by the modeling. therefore the results of source localization and correlation is limited by the mismatch of the environment and sensor location. In this paper. the effects of mismatch in environment and system on the bias in estimating the source location are investigated in the context of source localization. In the Pekeris waveguide, the simulation shows that the mismatches in environment and system, can cause a significant biases in the source localization and a degradation in MFP correlation. Mismatch caused by uncertainties in array tilt and depth, bottom depth, bottom sound speed, etc. causes degradation in source localization performance.

• Journal of the Korean Society of Propulsion Engineers
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• v.3 no.1
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• pp.86-94
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• 1999

Measured shock wave effects were investigated by changing shock strength and position with particular emphasis on the stability limits of hydrogen-air jet flames. For this purpose, a supersonic nonpremixed, jet-like flame was stabilized along the axis of a Mach 2.5 wind tunnel, and wedges were mounted on the sidewall in order to interact oblique shock waves with the flame. This experiment was the first reacting flow experiment interacting with shock waves. Schilieren visualization pictures, wall static pressures, and flame stability limits were measured and compared to corresponding flames without shock-flame interaction. Substantial improvements in the flame stability limits were achieved by properly interacting the shock waves with the flameholding recirculation zone. The reason for the significant improvement in flame stability limits is believed to be the adverse pressure gradient caused by the shock, which can elongate the recirculation zone.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.05a
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• pp.199-202
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• 2008

The advantage of ultrasonic burning rate measurement of solid propellant is measuring burning rates with wide range of pressure in a single test. In the ultrasonic method, instantaneous thickness of solid propellants as function of pressure or time were measured using time of flight(TOF) of ultrasonic signals. So, uncertainties of the measured burning rates by ultrasonic method have to evaluate with variation of pressure, TOF and initial propellant thickness. In this study, we evaluated uncertainties of ultrasonic method for measuring burning rates on the types 317 and the 318 propellants.

• Composites Research
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• v.15 no.1
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• pp.1-8
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• 2002

In this study, the flutter analysis for a rectangular box wing and an actual fighter wing with composite shin, aluminum spar and aluminum rib has been conducted. A conservative 3D wing-box model of an actual wing is modeled by MSC/PATRAN and the corresponding free vibration analysis has been performed by MSC/NASTRAN. The finite elements of membrane, rod and shear panel are used. Using the practical ply angles, various composite laminates are composed and analysed. The DLM code which is linear aerodynamic theory in frequency domain is applied to calculate unsteady aerodynamic pressure in subsonic flow region and the V-g and p-k methods are applied to obtain the solution of aeroelastic governing equation in frequency domain.