• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.44 no.4
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• pp.333-342
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• 2016

Liquid rocket propulsion is a system that produces required thrust for satellites and space launch vehicles by using chemical reactions of a liquid fuel and a liquid oxidizer. Monomethylhydrazine/dinitrogen tetroxide, liquid hydrogen/liquid oxygen and RP-1/liquid oxygen are typical combinations of liquid propellants commonly used for the liquid rocket propulsion system. The objective of the present study is to investigate useful design and performance data of liquid rocket engine by conducting a numerical analysis of thermochemical reactions of liquid rocket propellants. For this, final products and chemical compositions of three liquid propellant combinations are calculated using equilibrium constants of major elementary equilibrium reactions when reactants remain in chemical equilibrium state after combustion process. In addition, flame temperature and specific impulse are estimated.

• Proceedings of the Korea Concrete Institute Conference
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• 2008.04a
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• pp.889-892
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• 2008

This study, having combined and displaced fly ash known as admixture material that delays hydration reaction with coarse particle cement("CC" hereinafter) collected in particle classification method during ordinary portland cement("OPC" hereinafter), reviewed the hydration heat characteristics affecting the concrete. To reduce hydration heat, the study plain-mixed which used 100% OPC for WB 50% level 1, displaced CC at level 3 of 25%, 50% and 75% for OPC, and by displacing FA with admixture material at level 5 of 0%, 10%, 20%, 30% and 40%, experimented totally 16 batches. As a result of experiment, in the case of flow, the more CC displacement rate increased, the more it tended to decrease, and the more FA displacement rate increased, the more it decreased. As for simple adiabatic temperature rise by the CC and FA displacement rates, it decreased as displacement rate increased, and particularly in the case of FA40, temperature rise amount, $5.8{\sim}7.4^{\circ}C$, was very low. Compressive strength decreased in proportion to displacement rate, however strength reduction increment was shown to decrease with age progress.

• Journal of the Korea Institute of Building Construction
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• v.14 no.2
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• pp.156-162
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• 2014

This study aims to develop a rapidly hardening type of concrete to achieve the removal of form intensity (more than 10MPa) using the method of curing at room temperature in order to solve some economic environmental problems by omitting the steam curing process involved in producing PC (Precast Concrete). Therefore, this study evaluated a rapidly hardening cement containing a high amunt of C3S, which is very responsive in expressing early intensity, and a rapidly hardening type of concrete which uses some hardening accelerator to increase thehydration reaction of $C_3S$. The results of the experiment on concrete using some hardening accelerator are asfollows. In the slump flow experiment for identifying the liquidity and the air test, the desired values were met. The compression strength showed rapid expression response by 12 hours, and met the desired value within 6~9 hours. Its drying shrinkage value and Autogenous shrinkage value were measured as below ($-754.5{\times}10^{-6}$),and satisfied the requirements. In addition, in the Semi-Adiabatic Temperature Test, it was found that the concrete rose to its peak temperature within 24 hours and then its temperature dropped.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2003.05a
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• pp.91-93
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• 2003

A comprehensive numerical study is carried out to investigate for the understanding of the flow evolution and flame development in a supersonic combustor with normal injection of ncumally injecting hydrogen in airsupersonic flows. The formulation treats the complete conservation equations of mass, momentum, energy, and species concentration for a multi-component chemically reacting system. For the numerical simulation of supersonic combustion, multi-species Navier-Stokes equations and detailed chemistry of H2-Air is considered. It also accommodates a finite-rate chemical kinetics mechanism of hydrogen-air combustion GRI-Mech. 2.11[1], which consists of nine species and twenty-five reaction steps. Turbulence closure is achieved by means of a k-two-equation model (2). The governing equations are spatially discretized using a finite-volume approach, and temporally integrated by means of a second-order accurate implicit scheme (3-5).The supersonic combustor consists of a flat channel of 10 cm height and a fuel-injection slit of 0.1 cm width located at 10 cm downstream of the inlet. A cavity of 5 cm height and 20 cm width is installed at 15 cm downstream of the injection slit. A total of 936160 grids are used for the main-combustor flow passage, and 159161 grids for the cavity. The grids are clustered in the flow direction near the fuel injector and cavity, as well as in the vertical direction near the bottom wall. The no-slip and adiabatic conditions are assumed throughout the entire wall boundary. As a specific example, the inflow Mach number is assumed to be 3, and the temperature and pressure are 600 K and 0.1 MPa, respectively. Gaseous hydrogen at a temperature of 151.5 K is injected normal to the wall from a choked injector.A series of calculations were carried out by varying the fuel injection pressure from 0.5 to 1.5MPa. This amounts to changing the fuel mass flow rate or the overall equivalence ratio for different operating regimes. Figure 1 shows the instantaneous temperature fields in the supersonic combustor at four different conditions. The dark blue region represents the hot burned gases. At the fuel injection pressure of 0.5 MPa, the flame is stably anchored, but the flow field exhibits a high-amplitude oscillation. At the fuel injection pressure of 1.0 MPa, the Mach reflection occurs ahead of the injector. The interaction between the incoming air and the injection flow becomes much more complex, and the fuel/air mixing is strongly enhanced. The Mach reflection oscillates and results in a strong fluctuation in the combustor wall pressure. At the fuel injection pressure of 1.5MPa, the flow inside the combustor becomes nearly choked and the Mach reflection is displaced forward. The leading shock wave moves slowly toward the inlet, and eventually causes the combustor-upstart due to the thermal choking. The cavity appears to play a secondary role in driving the flow unsteadiness, in spite of its influence on the fuel/air mixing and flame evolution. Further investigation is necessary on this issue. The present study features detailed resolution of the flow and flame dynamics in the combustor, which was not typically available in most of the previous works. In particular, the oscillatory flow characteristics are captured at a scale sufficient to identify the underlying physical mechanisms. Much of the flow unsteadiness is not related to the cavity, but rather to the intrinsic unsteadiness in the flowfield, as also shown experimentally by Ben-Yakar et al. [6], The interactions between the unsteady flow and flame evolution may cause a large excursion of flow oscillation. The work appears to be the first of its kind in the numerical study of combustion oscillations in a supersonic combustor, although a similar phenomenon was previously reported experimentally. A more comprehensive discussion will be given in the final paper presented at the colloquium.