• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.05a
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• pp.168-172
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• 2008

Hot firing tests of sub-scale combustors were carried out to study the characteristic velocity according to the variation of propellant mass flow and injector arrangement. Test results show that there exists an effective range of relative flow-rate density on the condition of similar combustion pressure and mixture ratio. Numerical analysis has also revealed that the increase of the distance between the outermost injector array and the cylindrical chamber wall with film cooling increases the region of low mixture ratio near combustion chamber wall and it decreases the characteristic velocity of the combustor. Thus, it was confirmed that these two factors play an important part in improving the performance of LRE combustor on a predetermined chamber pressure.

• Transactions of the Korean Society of Mechanical Engineers B
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• v.37 no.3
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• pp.275-283
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• 2013

An injector that uses methane gas ($CH_4$) and liquid oxygen ($LO_x$) as propellants was designed to verify the combustion characteristics of an engine that uses methane, which is one of the next-generation propellants. A swirl/shear coaxial-type injector was used, and flow analysis was performed using Fluent to determine the main design parameters of the injector. A hydraulic test was performed to understand the atomization and spray pattern characteristics of the injector. Next, a combustion test was performed at the design point to understand the ignition and combustion stability. Additional combustion tests were performed according to the O/F ratio to investigate the combustion characteristics and stabilities using the characteristic exhaust velocity ($C^*$) and fluctuation of the chamber pressure. The experimental results showed that the combustion efficiency was greater than 90%, and the pressure fluctuation was lower than 2% under all conditions.

• Journal of the Korea Institute of Military Science and Technology
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• v.9 no.1 s.24
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• pp.117-124
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• 2006

Experiments on film cooling were performed with a small scale rocket engine homing liquid oxygen (LOx) and Jet A-1(jet engine fuel). Film coolants(Jet A-1 and water) were injected through the film cooling injector. Film cooled length and the outside wall temperature of the combustor were determined for chamber pressure, and the different geometries(injection angle) with the flow rates of film coolant. The loss of characteristic velocity due to film cooling was determined for the case of film cooling with water and Jet A-1. As the coolant flow increases, the outside wall temperatures decrease but the decrease in the outside wall temperatures reduced over the 8 percent film coolant flow rate. The efficiency of characteristic velocity was decreased with the Increase of the film coolant flow rate.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.03a
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• pp.560-564
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• 2004

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 10ton$_{f}$ in thrust with RP-1/Lox propellant. The optimal design was done for maximizing specific impulse of main combustion chamber with constraints of combustion temperature and power matching required by turbopump system. Design variables were selected as total mass flow rate to gas generator, O/F ratio in gas generator, turbine injection angle, partial admission ratio, and turbine rotational speed. Results of optimal design show the dimension of length, diameter, and contraction ratio of gas generator. Also, the combustion test was conducted to evaluate the performance of injector and combustion chamber. And the effect of the turbulence ring was investigated on the mixing enhancement in the chamber.r.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2004.10a
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• pp.229-234
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• 2004

The objective of the present study is to conduct model combustion tests for double swirl coaxial injectors to identify their combustion stability characteristics. Gaseous oxygen and mixture of methane and propane have been used as simulant propellants. Two model chambers tuned to the If acoustic resonance mode of a full-scale thrust chamber were manufactured to be used as a combustion cylinder. The main idea of the experiment is that the mixing mechanism is considered as a dominant factor significantly affecting combustion instability in a full-scale thrust chamber. Self-excited dynamic pressure values in a model chamber show different combustion stability zones with respect to a recess number. Upon test results, couplings between combustion conditions and the IT acoustic resonance mode become strengthened with the increase of a recess length.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.6
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• pp.104-109
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• 2004

On this paper study is concentrated on the breakup and atomization characteristics of spray formed by impinging jet injectors(like-doublet) used in liquid rocket engine(LRE). On the process of breakup and atomization, injection velocity and ambient gas pressure are the main parameters, so that these are used as variables that specify the experimental condition. Injection velocity varied from 3m/s to 30m/s and ambient gas pressure changed from 0.1MPa to 4.0MPa with nitrogen gas. As results, measured physical quantities decreased with increasing injection velocity and ambient gas pressure. But the decreasing ratios are different from those of the theory.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.82-85
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• 2009

In the development of Liquid Rocket Engine(LRE) systems, it is essential to understand the spray characteristics which influence mainly the performance and the stability of combustion. The injectors for this study have a recessed Liquid-swirl/Gas-centered jet coaxial type. For the similarity with actual conditions, the experimental conditions are calculated by using the momentum ratio as a matching parameter, and the stimulants of fuel and oxidizer are gaseous nitrogen and water respectively. The spray fields were measured by means of a photographic technique. Moreover, an effect of the momentum ratio has been investigated.

• 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.