• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.37 no.11
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• pp.1140-1147
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• 2009

Flow analyses have been performed to investigate the uniformity of propellant flow through the fuel and oxidizer manifolds of a full-scaled gas generator for a pump-fed liquid rocket engines. Injectors were simulated as porous medium layers having equivalent pressure drops. The uniformity of propellants has been analyzed for 3 fuel rings and 3 injector head configurations. The mixture ratio distribution at the exit of injectors has been estimated from the mass flow rates of fuel and oxidizer. The best configuration of fuel ring and injection head was selected through these flow analyses.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.11a
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• pp.261-264
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• 2008

This study experimentally examines combustion characteristics of a hybrid rocket in which solid paraffin is used as a fuel, while oxidizer is pure oxygen. Especially, the experiment investigates the effects of chamber pressure and configuration of fuel grain. The pressure inside the combustion chamber is varied by changing a flow rate of oxidizer. The regression rate is observed to increase as the chamber pressure does. There also exists the effects of shape of fuel grain on thrust. Characteristic of paraffin hybrid rocket changes with shape of fuel grain. When there is a room near the injector, thrust increases. On the other hand, the room near the nozzle does not contribute to thrust increasement.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.595-598
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• 2011

Supersonic combustion phenomena in the main combustor of a dual combustion ramjet (DCR) engine are studied numerically. Since the supersonic combustion is affected significantly by the compressibility effects parametric studies have been carried out for the constant are length and the divergence angle. Numerical studies with fixed inflow condition for different geometric configurations reveals that the supersonic combustion in DCR combustor has the characteristics of lifting flame, where the lifting flame is maintained near the injector tip for the case of long combustor length with small divergence angle, but the lifting height is significantly increase for large divergence angle resulting flame blow-out of the combustor. Therefore, it is concluded that flame stability should be considered sufficiently in the design o DCR combustor.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2003.05a
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• pp.258-261
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• 2003

The optimal design and combustion analysis of the gas generator for Liquid Rocket Engine (LRE) were performed. A fuel-rich gas generator in open cycle turbopump system was designed for 101on1 in thrust with RP-1/LOx combination. The optimal design was done for maximizing specific impulse of main combustion chamber with constraints of combustion temperature and power matching in turbopump system. Results of optimal design show the dimension of length, diameter, and contraction ratio of gas generator. The configuration of the gas generator and the condition for performance which can maximize the objective function were determined and found to meet the design constraints. Also, the combustion analysis was conducted to evaluate the performance of designed chamber and injector of gas generator. And the effect of the turbulence ring was investigated on the mixing enhancement in the chamber.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.75-78
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• 2004

Various jet engines (Turbine engine family and RAM Jet engine) have been developed for high speed aircrafts. but their application to hypersonic flight is restricted by principle problems such as increase of total pressure loss and thermal stress. Therefore, the development of next generation propulsion system for hypersonic aircraft is a very important subject in the aerospace engineering field, SCRAM Jet engine based on a key technology, Supersonic Combustion. is supposed as the best choice for the hypersonic flight. Since Supersonic Combustion requires both rapid ignition and stable flame holding within supersonic air stream, much attention have to be given on the mixing state between air stream and fuel flow. However. the wider diffusion of fuel is expected with less total pressure loss in the supersonic air stream. So. in this study the direction of fuel injection is inclined 30 degree to downstream and the total pressure of jet is controlled for lower penetration height than thickness of boundary layer. Under these flow configuration both streams, fuel and supersonic air stream, would not mix enough. To spread fuel wider into supersonic air an aerodynamic force, baroclinic torque, is adopted. Baroclinic torque is generated by a spatial misalignment between pressure gradient (shock wave plane) and density gradient (mixing layer). A wedge is installed in downstream of injector orifice to induce an oblique shock. The schlieren optical visualization from side transparent wall and the total pressure measurement at exit cross section of combustor estimate how mixing is enhanced by the incidence of shock wave into supersonic boundary layer composed by fuel and air. In this study non-combustionable helium gas is injected with total pressure 0.66㎫ instead of flammable fuel to clarify mixing process. Mach number 1.8. total pressure O.5㎫, total temperature 288K are set up for supersonic air stream.

• Korean Journal of Materials Research
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• v.18 no.3
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• pp.148-152
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• 2008

A new cost-effective atomic layer deposition (ALD) technique, known as Proximity-Scan ALD (PS-ALD) was developed and its benefits were demonstrated by depositing $Al_2O_3$ and $HfO_2$ thin films using TMA and TEMAHf, respectively, as precursors. The system is consisted of two separate injectors for precursors and reactants that are placed near a heated substrate at a proximity of less than 1 cm. The bell-shaped injector chamber separated but close to the substrate forms a local chamber, maintaining higher pressure compared to the rest of chamber. Therefore, a system configuration with a rotating substrate gives the typical sequential deposition process of ALD under a continuous source flow without the need for gas switching. As the pressure required for the deposition is achieved in a small local volume, the need for an expensive metal organic (MO) source is reduced by a factor of approximately 100 concerning the volume ratio of local to total chambers. Under an optimized deposition condition, the deposition rates of $Al_2O_3$ and $HfO_2$ were $1.3\;{\AA}/cycle$ and $0.75\;{\AA}/cycle$, respectively, with dielectric constants of 9.4 and 23. A relatively short cycle time ($5{\sim}10\;sec$) due to the lack of the time-consuming "purging and pumping" process and the capability of multi-wafer processing of the proposed technology offer a very high through-put in addition to a lower cost.

• Journal of the Korean Institute of Gas
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• v.19 no.4
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• pp.22-28
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• 2015

In order to keep the competitiveness of LPG fuel for transportation fuel, the difference in fuel consumption with gasoline and cost for an aftertreatment system should be reduced with continuous development of technology for LPG engine. In the present study, spray-guided type direct injection combustion system, whose configuration is composed of direct injector in the vicinity of spark plug, was employed to realize stable lean combustion. A certain level of nitrogen oxides($NO_x$) emits due to a locally rich mixture regions in the stratified mixture. With the application of EGR system for the reduction of $NO_x$, 15% of $NO_x$ reduction was achieved whereas fuel consumption and hydrocarbon emission increased. By the application of EGR, the combustion speed reduced especially appeared at initial flame development period and peak heat release rates and increasing rates for heat release rate decreased as EGR rate increased due to the dilution effect of intake air.