• Proceedings of the KSME Conference
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• 2000.04b
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• pp.514-519
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• 2000

Recently a considerable interest is being concentrated on industrial applications of supersonic Coanda wall jets, but the flow physics are not still understood well. It is of practical importance to evaluate the effectiveness of supersonic Coanda wall jet devices fer such industrial purposes. In the present work, experiments and computations were performed to Set a better understanding of the supersonic Coanda jet physics. The experiments were made using a small blow-down wind tunnel. The operating pressure ratio and the Coanda surface configuration were changed to investigate their influences on the wall jet flows. Two-dimensional Navier-Stokes computations were performed using a TVD finite volume scheme to effectively capture the important wave structures of supersonic Coanda jet flows. Both experimental and computational results showed several important hysterical features of the supersonic Coanda wall jets; the attachment and detachment of supersonic Coanda jet were strongly dependent on the change processes of the operating pressure ratio and the detailed flow configuration.

• 한국연소학회:학술대회논문집
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• 2003.12a
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• pp.159-165
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• 2003

This paper describes numerical efforts to characterize the flame-holding and air-fuel mixing process of model SCRamjet engine combustor, where a hydrogen jet injected into a supersonic cross flow and in a cavity. Combustion phenomena in a model SCRamjet engine, which has been experimentally studied at University of Queensland and Australian National University using a free-piston shock tunnel, was observed around separation region of upstream of the normal injector and inside of cavity. The results show that the separation region and cavity generates several recirculation zones, which increase the fuel-air mixing. Self ignition occurs in the separation-freestream and cavity-freestream interface.

• 한국연소학회:학술대회논문집
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• 2006.04a
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• pp.197-204
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• 2006

The model cavity scramjet engine experiments are carried out using T3 free-piston shock tunnel. Upstream hydrogen fuel is injected before the cavity with different injection pressure. OH planar laser-induced fluorescence is used to investigate the combustion zone and piezoelectric pressure transducers are used to define the pressure rise due to the combustion. Main combustion region is a mixing layer which is between air and fuel. Also high OH fluorescence signal is appeared in the shear layer above the cavity in high equivalence ratio. From the OH signal in the cavity, this fuel injection system can be a role as a flame- holder.

• Explosives and Blasting
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• v.22 no.4
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• pp.23-29
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• 2004

Non-electric detonator was developed to improve the blasting efficiency of electric detonator. It is increasingly utilized in surface and tunnel blasting due to its safety in external electric shock, precise delayed time, and decrease in blasting vibration and noise. The paper describes the detonating system of non-electric detonator, principle operating and planning methods, and case history so that it can be contributed to improve blasting technology.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.10a
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• pp.99-108
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• 2004

In this paper, a small supersonic wind tunnel is designed and built to study the flow characteristics of a supersonic impulse turbine. The flow is visualized by means of a single pass Schlieren system. The supersonic cascade with 2-dimensional supersonic nozzle was tested over a wide range of pressure ratio. Highly complicated flow patterns including shocks, nozzle-cascade interaction and shock boundary layer interactions are observed.

• Advances in aircraft and spacecraft science
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• v.10 no.5
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• pp.457-471
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• 2023

The paper evaluates the aerodynamic coefficients on a blunt-nose re-entry capsule with a conical cross-section followed by a cone-flare body. A computer code is developed to solve three-dimensional compressible inviscid equationsfor flow over a Space Recovery Experiment (SRE) configuration at different flare-cone half-angle at Mach 6 and angle of attack up to 5°, at 1° interval. The surface pressure variation is numerically integrated to obtain the aerodynamic forces and pitching moment. The numerical analysis reveals the influence of flare-cone geometry on the flow characteristics and aerodynamic coefficients. The numerical results agree with wind tunnel results. Increase of cone-flare angle from 25° to 35° results in increase of normal force slope, axial forebody drag, base drag and location of centre of pressure by 62.5%, 56.2% and 33.13%, respectively, from the basic configuration ofthe SRE of 25°.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.45 no.11
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• pp.914-921
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• 2017

The experimental research on a high-altitude environment simulation of space launch vehicle is important for securing independent technologies with launching space vehicles and completing missions. This study selected an altitude of 65 km for the experiment environment where it exceeded Mach number of 6 after the launch of Korean Space Launch Vehicle(KSLV-II). Shock tunnel was used to replicate the flight condition. After flow establishment, in order to confirm aerodynamic characteristics and normal and oblique shockwaves, the flow verification was carried out by measuring stagnation pressure and heat flux of a forebody model, and shockwave stand-off distance of a hemispherical model. In addition, a shock-free technique to recover free-stream condition has been developed and verified. From the results of the three verification tests, it was confirmed that the flow was replicated with the error of about ${\pm}3%$. The error between the slope angle of inclined shockwave of the scaled down transition section model using the shock-free shape and the slope angle of the horizontal plate model, and between the theoretical and the experimental value of the static pressure of the model were confirmed to be 2% and 1%, respectively. As a result, the efficiency of the shockwave cancellation technique has been verified.

• Explosives and Blasting
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• v.41 no.1
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• pp.1-18
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• 2023

Stemming is a process applied to blast holes to prevent gases from escaping during detonation. A stemming material helps confine the explosive energy for longer and increases rock fragmentation. This study developed a stemming material based on a shear-thickening fluid (STF) that reacts to dynamic shock. Two blasting experiments were conducted to Field-verify the performance of the STF-based stemming material. In the first experiment, the pressure inside the blast hole was directly measured based on applying the stemming material. In the second field verification, tunnel blasting was performed, and the blasting results of sand stemming and, that of the STF-based stemming case were compared. The measurement results of the pressure in the blast hole showed that when the STF-based stemming material was applied, the pressure at the top of the blast hole was lower than in the sand stemming case, and the stemming ejection was also lower. The results of the field application verify that the excavation performance of the STF-based stemming case in the tunnel blasting was superior to that of the sand stemming case.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.46 no.10
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• pp.791-797
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• 2018

In this work, the high-altitude environment simulation study was carried out at an altitude of 65 km exceeding Mach number of 6 after the launch of Korean Space Launch Vehicle using a shock tunnel. To minimize the flow disturbance due to the strut support of test model as much as possible, a few different types of strut configurations were considered. Using the configuration with minimum disturbance, the high-altitude environment simulation experiment including a propulsion system with a single-plume, was conducted. From the thruster test through flow visualization, not only a shockwave pattern, but a general flow-field pattern from the mutual interaction between the exhaust plume and the free-stream undisturbed flow, was experimentally observed. The comparison with the computation fluid dynamic(CFD) results, showed a good agreement in the forebody whereas in the afterbody and the nozzle the disagreement was about ${\pm}7%$ due to unwanted shockwave formation emanated from the nozzle-exit.

• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.5
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• pp.31-36
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• 2015

As one of the most promising propulsive systems in the future, the dual-mode scramjet engine has drawn the attention of many researches. Detailed flow features concerned with the isolator play an important role in the dual-mode scramjet system. The 2D numerical method has been used for the dual-mode scramjet with wind tunnel. To validate the ability of the numerical model, numerical results have been compared with the experimental results. Overall pressure distributions show quite good match with the experimental results. Back pressure has been studied for maximum pressure rising. According to the results, pressure distribution of supersonic inlet section is not influenced by back pressure. The shock train is pushed towards upstream as the back pressure increases. The maximum value of back pressure without inlet unstart goes up rapidly and then keeps constant when the isolator length increases. The optimal length of isolator section ($L/H_{th}$) is 8.7 in this model.