• Journal of the Korean Society of Propulsion Engineers
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• v.9 no.1
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• pp.1-8
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• 2005

An axisymmetric compressible flow solver accounting for chemical equilibrium has been developed as an analysis tool exclusively suitable for performance prediction and optimum contour design of LRE thrust chamber. By virtue of several features focusing on user-friendliness and effectiveness including automatical grid generation and iterative calculations with changes in design parameters prescribed through only one keyword-type input file, a design engineer can evaluate very fast and easily the influences of various design inputs such as geometrical parameters and operating conditions on propulsive performance. Validations have been carried out for various aspects by detailed comparisons with the result of CEA code, experimental data of JPL nozzle, actual data for two historical engines, and ReTF data for KSR-III.

• Fibers and Polymers
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• v.1 no.2
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• pp.97-102
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• 2000

Pitch precursors were synthesized from coal tar(CT) and pyrolysis fuel oil(PFO, petroleum residue oil) at relatively low temperature of $250^{\circ}$, in the presence of horontrifluorideidiethyletherate complex(BFDE) as a catalyst and nitrobenzene(NB) as a co-catalyst. The softening point, nitrogen content and carbon yield increased with an increase of concentration of NB. The pitch precursors with good spinnability were prepared by removing the volatile components through $N_2$ blowing. The precursor pitches were spun through a circular nozzle, stabilized at $310^{\circ}$ and finally carbonized at $1000^{\circ}$. The optically anisotropic structure formed at the absence of NB was changed into isotropic structure, showing a decrease in size of the flow domain. The hollow carbon fiber could be prepared in the process of stabilization. The results proposed that the morphology of carbon materials could be controlled by changing the concentration of catalyst and/or co-catalyst and/or stabilization condition that affect on the mobility of molecules during carbonization.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.37-42
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• 2010

A design plan was proposed for determining combustor configuration of regenerative- cooled liquid rocket engine in the process of preliminary design. Rocket performance and regenerative cooling results were calculated using the properties of combustion gas estimated in CEA. For required thrust, chamber pressure, atmosphere pressure and propellant mixture ratio the mass flow rate of propellants and combustor performance were predicted using one-dimensional and experimental equations. Finally, determinable plan for contour of combustor were presented through Rao nozzle design method.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.11a
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• pp.825-828
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• 2010

APU for aircraft is operated under severe condition as high altitude and low temperature, and demand high reliability in flight. This study is to be verified of the ignition and the combustion stability of APU under the harsh conditions. The basic data obtained in combustion rig test were directly applied to the altitude test with a engine. That start logic was obtained in ground development test. The results of altitude test show that air swirl injector has good operation and ignition performance at 20kft, hot/cold($-40^{\circ}C$) day.

• Journal of the Korean Society of Propulsion Engineers
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• v.21 no.6
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• pp.49-56
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• 2017

The current study has developed an in-house 3D FEM code in order to model thermoacoustic problems in an annular system and compared the acoustic field calculation results with measured ones from a benchmark combustor. From the comparison of calculation results with the measured data, the current acoustic code could successfully capture the various acoustic mode found in the annular system. In addition, it was found that the transverse waves in the combustor were strongly affected by the nozzle acoustic impedances, as well, the pressure distributions were closely related with the combustor acoustic pressure field.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.20 no.3
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• pp.1061-1073
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• 1996

Vortex tube is a simple device which splits a compressed gas stream into a cold stream and a hot stream without any chemical reactions. The phenomena of energy separation taking place in a vortex tube has been investigated experimentally. Recently, vortex tube is widely used to local cooler of industrial equipments and air conditioner of special purpose. In this study, experimental study on vortex tube efficiency was performed with various cold end orifices and nozzles type. The experimental results indicate that there is an optimum diameter of cold end orifice and nozzle type for the best cooling performance. The variation of the maximum wall temperature along the vortex tube surface provides useful information about the location of the stagnation point of the flow field at the axis of the vortex tube. The similarity relation for the prediction of the temperature of the cold exit air was obtained.

• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.2
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• pp.29-37
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• 2015

This paper presents a numerical investigation on detonation of a kerosene-air mixture in the copper tube and the structural response associated with combustion instability in liquid rocket engine. A single step Arrehnius rate law and Johnson-Cook strength model are used to describe the chemical reaction of kerosene-air mixture detonation and the plastic deformation of the copper tube. The changes of flow field and tube stress which are induced by plastic deformation, are investigated on the different tube thicknesses and nozzle configurations.

• Aerospace Engineering and Technology
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• v.9 no.1
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• pp.17-27
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• 2010

KSLV-I KM plume including alumina particle has been studied using the continuum solver. Alumina particles are assumed to have 7 different diameters, and the specific ratio of the plume gas is assumed to be 1.2, with which the internal nozzle flow characteristics are similar to those of the chemically equilibrium analysis results. The results showed that the expansion angle of the particles is smaller than that of the gas phase, and that the big sized alumina particles are gathered in the plume core and the expansion angles of the big sized particles are smaller than those of the light particles. When the emissivity of the particles are assumed to be 0.1, the radiative heat flux is equivalent to those measured during the flight test of KSLV-I.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.499-505
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• 2004

A Propulsion System of the CRW(Canard Rotor Wing) type UAV(Unmanned Aerial Vehicle) was composed of the turbojet engine to generate the propulsive exhaust gas, and the duct system including straight bent ducts, tip-jet nozzles, a master valve and a variable main nozzle for three flight modes such as lift/landing mode, low speed transition flight mode and high speed forward flight mode. In this study, in order to operate safely the propulsion system, the dynamic Performance behavior of the system was modeled and simulated using the SIMULIN $K^{ }$, which is the user-friendly GUI type dynamic analysis tool provided by MATLA $B^{ }$. In the transient performance model, the inter-component volume model was used. The performance analysis using the developed models was performed at various flight condition, valve angle positions and fuel flow schedules, and these results could set the safe flight mode transition region to satisfy the inlet temperature overshoot limitation as well as the compressor surge margin. Performance analysis results using the SIMULIN $K^{ }$ performance program were compared with them using the commercial program GSP.m GSP.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.05a
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• pp.177-180
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• 2010

It has been studied the effect of mixture ratio and chamber pressure on variations of discharge coefficients. Combustion experiments of bi-liquid swirl coaxial injectors were conducted at fuel-rich conditions with liquid oxygen and kerosene. Using two types of injectors for the experiments, characteristics of the discharge coefficient have been identified from variations in a diameter of the fuel nozzle and a momentum ratio along with the change of a LOx spray angle. It is concluded that discharge coefficients do not vary because of no change of flame structures from the fact that the fuel swirl chamber is completely filled up with fuel flow.