• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.05a
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• pp.355-358
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• 2006

A linear acoustic analysis is conducted to examine bituning of acoustic resonators for acoustic damping in a combustion chamber of liquid rocket engine. Bituned resonators are tuned to the two principal modes, the first tangential(1T) and the first radial(1R) modes. First, the acoustic-damping effect of monotuned resonators is investigated. The damping capacity is quantified by damping factor as a function of the number of the resonators monotuned to 1T or 1R mode. Next, the damping characteristics of the bituned resonators are investigated. From the numerical data, the number of resonators, to be tuned to 1T and 1R modes, respectively, can be selected properly.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.05a
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• pp.288-293
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• 2006

The combustion performances and characteristics of the subscale liquid rocket combustion chamber are discussed in this paper. Subscale combustion chamber is composed of mixing head, ablative cooling cylinder, and water cooling nozzle. The mixing head has eighteen coaxial swirl injectors and one center coaxial swirl injector for ignition. The mixing heads employing the injectors of low different recess length are considered in this paper. The results of the firing test, comparison of performance, and characteristics of static and dynamic pressures of the four different mixing heads are described. The characteristics of combustion at design and of design points are also discussed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.05a
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• pp.151-155
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• 2006

Elastic-viscoplastic structural analysis is performed for regenerative cooling chamber of liquid rocket thrust chamber using Bodner-Partom visco-plastic model. Strain rate test was also conducted for a copper alloy at various temperatures in order to get material constants of visco-plastic model used in the structural analysis. Material constants of visco-plastic model were obtained from strain rate test results and visco-plsstic model was incorporated into finite element program, Marc, by means of user subroutine. The structural analysis results indicate that the deformation of cooling channel is mostly caused by thermal loading rather than pressure loading and confirmed structural stability of the cooling channel under operating condition.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.05a
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• pp.161-164
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• 2006

A design program has been developed to predict film cooling performance in a liquid rocket engine combustion chamber. A thermal protecting effect of low mixture ratio gas has been analysed by CFD. A one-dimensional film cooling model based on the CFD results has been implemented in the previously developed design program of regenerative cooling. The predicted heat flux at the nozzle throat ranges from -16% to +28% when it is compared to the published measured data. The throat heat flux reduces by 36% when film cooling of 10% of fuel mass flow rate is applied.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.358-362
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• 2005

An effective control of the axial thrust of a turbopump is one of the critical issues for its operational stability. In order to assure the stability of a turbopump-type fuel pump for a liquid rocket engine, an axial thrust measurement system was developed and a series of axial thrust tests were performed in water environment. In the tests, the axial thrust of the fuel pump at the design flowrate satisfied the axial force condition of the bearing of the pump. Also, it was found that by using orifices with different geometries in the secondary flow passage the overall axial thrust of the pump could be controlled.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2005.11a
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• pp.220-223
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• 2005

System design parameters of gas generator cycle liquid rocket engines were investigated and compared in the present study. Characteristic velocity of combustor, pressure drop of combustor injector, exit pressure of pump, pump efficiency and specific power of turbine were considered as a system design parameter. The result shows the characteristic velocity is in the range of 1700-1770 m/s, pressure drop of combustor injector, 4-10 bar, pump exit pressure ratio to combustion pressure, 120-230%, pump efficiency, 60-80%, specific power of turbine, $0.28-0.58MW{\cdot}s/kg$.

• Journal of the Korean Society of Propulsion Engineers
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• v.10 no.4
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• pp.69-76
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• 2006

Elastic-viscoplastic structural analysis has been performed for regenerative cooling chamber of liquid rocket thrust chamber using Bodner-Partom visco-plastic model. Strain rate test was conducted for a copper alloy at various temperatures in order to get material constants of visco-plastic model used in the structural analysis. Material constants of visco-plastic model were obtained from strain rate test results and visco-plastic model was incorporated into finite element program, Marc, by means of a user subroutine. The structural analysis results indicated that the deformation of cooling channel is mostly caused by thermal loading rather than pressure loading and confirmed structural stability of the cooling channel under the operating condition.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.821-826
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• 2011

The turbopump+gas generator (TP+GG) coupled test for the liquid rocket engine development was performed. By comparing the results of a engine startup transient analysis with this test results, the verification of the analysis model was performed. From this, as to the analysis of the engine startup, the method calculating the pressure ratio of the turbine during the initial stage of startup was improved. And a fact that the transient heat transfer phenomenon between the working fluid and the solid parts of turbine effects to the calculation of turbine pressure ratio and consequentially to the startup analysis was revealed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.127-134
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• 2011

A genetic algorithm (GA) has been employed to optimize the major design variables of the liquid rocket engine. Pressure of the main combustion chamber, nozzle expansion ratio and O/F ratio have been selected as design variables. The target engine has the open gas generator cycle using the LO2/RP-1 propellant. The gas properties of the combustion chamber have been obtained from CEA2 and the mass has been estimated using reference data. The objective function has been set as multi-objective function with the specific impulse and thrust to weight ratio using the weight method. The result shows about 4% improvement of the specific impulse and 23% increase of the thrust to weight ratio. The Pareto frontier line has been also obtained for various thrust requirements.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2006.11a
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• pp.196-200
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• 2006

This paper describes the current development status of the major subsystems, turbopump and gas generator, for a turbopump-fed liquid oxygen-kerosene rocket engine system. As a secondary stage of the liquid rocket engine development test, turbopump-gas generator powerpack tests are planned. The schematics of the test hardware and the test facility for the TP+GG coupled test are presented. The results of a preliminary analysis for operating regimes of the TP+GG coupled test are also presented.