• Journal of the Korean Society of Propulsion Engineers
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• v.9 no.2
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• pp.1-8
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• 2005

The effect of coolants has been studied on the burning properties of low burning rate HTPB/AP composite propellant containing Oxamide or Melamine as coolant for the gas generator. With increasing the content of coolant, the burning rate and the flame temperature could be lowered and the effect on flame temperature was about the same for two coolants. However due to the different thermal decomposition properties of coolant, the burning rate of Melamine propellant was found to abnormally decrease if $200{\mu}m$ AP was partially replaced with $6{\mu}m$ AP.

• Journal of the Korean Society of Propulsion Engineers
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• v.2 no.3
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• pp.54-61
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• 1998

A numerical simulation is attempted for the regenerative cooling heat transfer processes of the liquid propellant rocket engine. The heat transfer from the combustion gases to the thrust chamber wall is called gas side heat transfer. This heat is conducted radially to the coolant through the carbon deposit and metallic wall of thrust chamber Finally, this heat is convected away by the coolant flowing along the passages in the thrust chamber. The equivalence of these three heat fluxes of the above processes is utilized to determine the coolant side wall temperature, gas side wall temperature and the heat flux. When the number and shape(width, height) of coolant passages, the shape(size) of thrust chamber, oxidant and fuel properties, coolant properties, oxidant/fuel mixture ratio, coolant inlet temperature, the thickness of carbon deposit formed along the thrust chamber wall during combustion are given, reasonable radial direction temperature distributions and heat fluxes along the thrust chamber axis are obtained.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.33 no.6
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• pp.71-75
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• 2005

An experimental study was carried out to investigate the effect of film cooling in the lab-scale dump-cooled liquid rocket engine using LOX and kerosene as propellants. The nozzle of the rocket engine was film cooled with water as coolant. A special film cooling adapter was fabricated to introduce the film-coolant into the thrust chamber. The flow rates of film coolant was approximately 15~19 percent of the total propellant. The nozzle heat flux was determined from the measured temperature rise and flow rate of the coolant(water). Large reductions in the nozzle heat flux was resulted when film cooling adapter located directly upstream of the nozzle.

• Journal of the Korean Society of Propulsion Engineers
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• v.22 no.6
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• pp.84-93
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• 2018

Three-dimensional (3D) numerical simulations have been carried out to study how coolant injection conditions influence the cooling efficiency and projectile ejection performance in a mixture-gas ejection system (or gas-steam launch system). The 3D single-phase computational model was verified using a 1D model constructed with reference to the previous research and then a two-phase flow computation simulating coolant injection on to hot gas was performed using a DPM (Discrete Phase Model). As a result of varying the coolant flow rate and number of injection holes, cooling efficiency was improved when the number of injection holes were increased. In addition, the change of the coalescence frequency and spatial distribution of coolant droplets caused by the injection condition variation resulted in a change of the droplet diameter, affecting the evaporation rate of coolant. The evaporation was found to be a critical factor in the design optimization of the ejection system by suppressing the pressure drop while the temperature decreases inside the breech.

• Journal of the Korean Society of Propulsion Engineers
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• v.9 no.4
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• pp.48-54
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• 2005

An experimental study was carried out to investigate the effect of film cooling in the lab-scale liquid rocket engine using liquid oxygen(LOx) and Jet A-1(Jet engine fuel) as propellants. Film coolants(Jet A-1 and water) was injected through the film cooling injector. The outside wall temperature of the combustor and film cooled length were determined for chamber pressure, mixture ratio, and the different geometries(injection angle) with the percent film coolant flow rate. The loss of characteristic velocity was determined for the case of film cooling with water and Jet A-1. As chamber pressure increased, the outside wall temperature increased in the nozzle but unchanged over the 9 percent film coolant flow rate for the combustion chamber used in this study. Characteristic velocity wasn't affected with the mixture ratio over the 9 percent film coolant flow rate.

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.3
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• pp.67-75
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• 2013

The numerical analysis for the verification of preburner's cooling characteristics applying to kerosene-LOx rocket engine has been fulfilled. The distribution of combustion gas properties in primary combustion zone was calculated by the mixture ratio based on head injector arrangement, the properties of oxygen flowing in wall channels as coolant were applied under real-gas conditions, and multi-phase mixing model was employed to calculate the mixing process of primary combustion zone with liquid oxygen which was used for wall cooling. The results of numerical analysis were compared with the experimental results, hence thermo-physical properties in cooling channels and a combustor could be quantitatively identified.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.33 no.3
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• pp.78-85
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• 2005

The purpose of the study was to establish the design procedure and develop the program for designing regenerative cooling system. To obtain the design parameter necessary for the realization of regenerative cooling system, water-cooled regenerative cooling system was designed from suggested procedure. To compare experimental results with a present method of analytically predicting the heat transfer loads, $250kg_{f}$ experimental LRE with water-cooled regenerative cooling system was investigated. Based on the investigation, the good correction between the predicted and measured data was verified. Developed design program can be used to designing Kerosene- cooled regenerative cooling system. The basic experimental data and correlations obtained in this study for 250kgf experimental LRE with water-cooled regenerative cooling system can be directly applicable to the real LRE.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.04a
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• pp.12-15
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• 2011

The most widely used kerosene-LOx liquid-propellant rocket engines in these days have a similar engine schematic to those of the past because of the development cost and the reliability. The efficiency of engines could be increased by the factors such as a cooling method, engine cycles, shape of cooling channels, additional coolant and so on. In this article, it is described that some design ideas for performance enhancement by exchange kerosene with LOx of a coolant.

• Journal of the Korean Society of Propulsion Engineers
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• v.21 no.1
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• pp.72-83
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• 2017

A thermo-fluid dynamic analysis was performed through the numerical simulation of a missile canister. Calculation was made in a fixed analytical volume and fully evaporated water was used as a coolant. To analyze the interaction among the hot gas, coolant, and mixture flow, Realizable $k-{\varepsilon}$ turbulence and VOF(Volume Of Fluid) model were chosen and parametric study was performed with the change of coolant flow rate. It could be found that the pressure on the canister top nonlinearly increased with the increase of coolant flow rate. Temperature and coolant distribution were closely related to the flow behavior in canister. Temperature on the canister bottom indicated a decrease being proportional to coolant flow rate in early times but after a specific time, the temperature increased with the tendency being reversed. In addition, the early part of temperature showed a fluctuating phenomenon because of the overall circulatory flow of mixture gas.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2012.05a
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• pp.106-109
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• 2012

It was designed and tested ignition that an oxidizer rich preburner for a staged combustion cycle liquid rocket engine propelled by kerosene and LOx. Operation conditions of the preburner are about 60 of OF ratio and 20 MPa of combustion pressure. Ignition characteristics were compared by propellants flowrate. As the results, the higher propellants flowrate, the shorter the ignition delay time and the higher ignition stiffness. The ignition delay time was affected by incoming the oxidizer flowrate through the refrigerative cooling channels. The oxidizer flowrate from the cooling channels decreased by inflow of combustion gas during initial ignition. The oxidizer flowrate of the cooling channels increases, it is rapid recovery by cooling effect, eventually the ignition delay time decreases.