• Clean Technology
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• v.10 no.4
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• pp.177-187
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• 2004

Dry ice snow jet was produced by high-pressure expansion of liquid carbon dioxide and subsequent acceleration by carrier gas flow. Removal mechanism for coating was not so different from that for contaminating particles on the surface. The removal of coating was quantitatively described by Hutchings' equation. The two parameters obtained from the plot, were used to predict the removal rate or the specific coating area removed by a unit mass of carbon dioxide. Their values also enhanced the reliability of the experimental data and enabled the experimental errors corrected. Hutchings' plots obtained with various values of either stand-off distance or nozzle length tends to meet apparently at a single unique point at which the scar-size enlargement by focusing jet plume was balanced with that by spreading it.

• Journal of Korean Ophthalmic Optics Society
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• v.7 no.1
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• pp.1-8
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• 2002

Various kinds of aerosols such as a mixture of hairspray resin and its solvents, a perfume, a soap water, and ask in-lotion were sprayed onto the contact lens which was mounted on the eye-model made of wood, varying the parameters as the distance between nozzle and contact lens and the spraying time duration. In the case of both the hairspray and the perfume including an ethanol as solvents, the shape of contact lenses has changed tremendously Jess than 1 minute after the exposure to the aerosol particles and then it continued until the lens surfaces were completely flipped over. Driving force for the lens distortion seems to be the expansion coefficient difference between the inner and outer surface of lens, which was caused by the heat of vaporization of ethanol included in the aerosol and subsequent temperature decrease induced on the lens surface.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.37 no.1
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• pp.1-7
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• 2013

An aerodynamic rig test of a radial turbine for an auxiliary power unit (APU) was performed at a high-temperature turbine test facility at the Korea Aerospace Research Institute. The pressure ratio, Mach number, and flow coefficient in the rig test are the same as those under normal engine operation conditions. The design pressure ratio is 3.096, design test speed is 34909 rpm, and turbine inlet temperature is $160^{\circ}C$. The turbine has airfoil-type nozzles, and the diameter of the turbine wheel is 175.74 mm. The turbine map is experimentally measured, and the detailed flow at the turbine inlet is measured. The pressure distribution in the nozzle at both the hub and the shroud sides and the pressure distribution along the shroud casing of the turbine wheel were measured, and this confirmed that the expansion process in the turbine wheel is acceptable.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.33 no.8
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• pp.92-98
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• 2005

Screen can provide any disturbed resistance that affects the change in characteristics of turbulence, velocity and pressure distributions of the flow field, and thus it has been widely used to control the flow. Some previous related studies for compressible flows have limitations such as, considering relatively low-Mach-number flows in the range of 0.3 ∼ 0.7, and not observing the detailed shock structures of the flow fields. An experimental study on highly compressible axi-symmetric supersonic jet flow fields behind wire-gauze screen has thus been carried out. Continuous/instantaneous flow images by Schlieren flow- visualization technique and the information of Pitot pressure/flow-noise measurements of the flow field behind the screen for various jet expansion conditions have been obtained. Effects of various porosity and inclination angles of the screen at the nozzle exit have also been investigated, and the experimental results have been compared to the case with no screen installed.

• Journal of the Korean Society of Propulsion Engineers
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• v.9 no.3
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• pp.85-91
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• 2005

Radiative heating of a liquid rocket base plane due to plume emission is numerically investigated. Calculation of flow and temperature fields around rocket nozzle precedes and thereby realistic plume shape and temperature distribution inside the plume are obtained. Based on the calculated temperature field, radiative transfer equation is solved by discrete ordinate method. With the sample rocket plume, the averaged radiative heat flux reaching the base plane is calculated about 5 kw/m$^{2}$ at the flight altitude of 10.9 km. This value is small compared with radiative heat flux caused by constant-temperature (1500 K) plume emission, but it is not negligibly small. At higher. altitude (29.8km), view factor between the base plane and the exhaust plume is increased due to the increased expansion angle of the plume. Nevertheless, the radiative heating disappears since the base plane is heated to high temperature (above 1000 K due to convective heat transfer.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.05a
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• pp.145-149
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• 2010

Thermal analyses have been performed to study the effect of location of fuel ring and thermal barrier coatings in regenerative cooling channels in a full-scale combustor. For the effective cooling, the fuel ring has better be installed near axial location of the low expansion ratio and low heat flux, and branching of cooling channels is preferable. Also, the radiative cooled nozzle extension is thought to be reasonable for the cooling of combustion walls. Among the possible coatings, $Y_2O_3$ stabilized $ZrO_2$ coating and Ni/Cr coating have been adopted. Compared with Ni/Cr coating which has high oxidation resistance, $Y_2O_3$ stabilized $ZrO_2$ coating, one of ceramic coatings is found to be much effective to sustain the thermal survivability of combustion walls.

• Transactions of the Korean hydrogen and new energy society
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• v.30 no.6
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• pp.593-600
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• 2019

The combustion characteristics of methane/hydrogen pre-mixed flame have been investigated with swirl stabilized flame in a laboratory-scale pre-mixed combustor with constant heat load of 5.81 kW. Hydrogen/methane fuel and air were mixed in a pre-mixer and introduced to the combustor through a burner nozzle with different degrees of swirl angle. The effects of hydrogen addition and swirl intensity on the combustion characteristics of pre-mixed methane flames were examined using particle image velocimetry (PIV), micro-thermocouples, various optical interference filters and gas analyzers to provide information about flow velocity, temperature distributions, and species concentrations of the reaction field. The results show that higher swirl intensity creates more recirculation flow, which reduces the temperature of the reaction zone and, consequently, reduces the thermal NO production. The distributions of flame radicals (OH, CH, C₂) are dependent more on the swirl intensity than the percentage of hydrogen added to methane fuel. The NO concentration at the upper part of the reaction zone is increased with an increase in hydrogen content in the fuel mixture because higher combustibility of hydrogen assists to promote faster chemical reaction, enabling more expansion of the gases at the upper part of the reaction zone, which reduces the recirculation flow. The CO concentration in the reaction zone is reduced with an increase in hydrogen content because the amount of C content is relatively decreased.

• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.3
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• pp.47-53
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• 2015

Design parameters required for Methane/oxygen bipropellant small-rocket-engine were derived through a theoretical performance analysis. The theoretical performance of the rocket engine was analyzed by using CEA and optimal propellant mixture ratio, characteristic length, and optimal expansion ratio were calculated by assuming chemical equilibrium. A coaxial-type swirl injector was chosen because of its outstanding atomization performance and high combustion efficiency compared to other types of injector and also a bell nozzle with 80% of its full length was designed. The rocket engine configuration with 1.72 MPa of chamber pressure, 0.18 kg/s in total propellant mass flow, and O/F ratio of 2.7 was proposed as a ground-firing test model.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.11a
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• pp.137-142
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• 2008

Ignition characteristics of combustion chamber with LOx lead cyclogram for liquid rocket engine were described. The combustion chamber has chamber pressure of 60 bar, propellant mass flow rate of 89 kg/s, and nozzle expansion of 12. Cold flow test to determine the filling time of propellant for cyclogram with LOx lead supply, ignition test to check the ability to ignite starting fuel from the ignitor, low pressure combustion test to check the propagation of flame into main fuel-oxidizer mixture from starting fuel and the main combustion stage, and design point combustion test to check the combustion performance were performed. Ignition and combustion tests with LOx lead supply were successfully performed and the stable cyclogram of start sequence for combustion chamber was developed.

• Journal of the Korean Society of Propulsion Engineers
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• v.18 no.5
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• pp.54-61
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• 2014

In this study, a plume exhausted from rocket nozzle is investigated by using an unstructured 2-dimensional axisymmetirc DSMC code at various altitude. The small back-pressure to total-pressure ratio($P_b/P_o$) and large $P_b/P_o$ represent low and high altitude condition, respectively. At low altitude, the plume shows a typical complicated structure (e.g. Mach disk) of underexpanded jet while the high altitude plume experiences plain expansion. The various features of exhaust plume is discussed including density, translational/rotational temperature, Mach number and Knudsen number. The results shows that even at 20 km altitude where the freestream Knudsen number is small as $1.5{\times}10^{-5}$, the transitional and rarefied flow regimes can occur locally within the plume. It confirms the necessity of DSMC computation at low altitude.