• 한국전산유체공학회:학술대회논문집
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• 2010.05a
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• pp.530-533
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• 2010

This paper presents numerical analysis and design optimization of various turbine blade cooling techniques with three-dimensional Reynolds-averaged Navier-Stokes(RANS) analysis. The fluid flow and heat transfer have been performed using ANSYS-CFX 11.0. A fan-shaped hole for film-cooling has been carried out to improve film-cooling effectiveness with the radial basis neural network method. The injection angle of hole, lateral expansion angle of hole and ratio of length-to-diameter of the hole are chosen as design variables and spatially averaged film-cooling effectiveness is considered as an objective function which is to be maximized. The impingement jet cooling has been performed to investigate heat transfer characteristic with geometry variables. Distance between jet nozzle exit and impingement plate, inclination of nozzle and aspect ratio of nozzle hole are considered as geometry variables. The area averaged Nusselt number is evaluated each geometry variables. A rotating rectangular channel with staggered array pin-fins has been investigated to increase heat transfer performance ad to decrease friction loss using KRG modeling. Two non-dimensional variables, the ratio of the eight diameter of the pin-fins and ratio of the spacing between the pin-fins to diameter of the pin-fins selected as design variables. A rotating rectangular channel with staggered dimples on opposite walls are formulated numerically to enhance heat transfer performance. The ratio of the dimple depth and dimple diameter are selected as geometry variables.

• Journal of the Korean Society of Propulsion Engineers
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• v.16 no.3
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• pp.97-103
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• 2012

The combustion chamber and nozzle of a liquid rocket engine should be protected from the high temperature combustion gas generated by the chamber. An upper-stage nozzle extension has a large expansion ratio, therefore, The light-weight refractory materials have been used since the weight impact on the launcher performance is crucial. Gas film cooling method was used before, but was not applicable nowadays. Ablative cooling method and radiative cooling method with niobium alloy, Ni-based superalloy and ceramic based composite have been used to this day.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.505-509
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• 2011

The combustion chamber and nozzle of a liquid rocket engine should be protected from the high temperature combustion gas generated by the chamber. An upper-stage nozzle extension has a large expansion ratio, therefore, The light-weight refractory materials have been used since the weight impact on the launcher performance is crucial. Gas film cooling and ablative cooling methods were used before, but were not applicable nowadays. Radiative cooling method with niobium alloy, Ni-based superalloy and ceramic based composite has been used to this day.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.31 no.1
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• pp.85-96
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• 2003

For a design of rocket engine nozzle, chemical equilibrium analysis which shares the same numerical characteristics with frozen flow analysis can be used as an efficient design tool for predicting maximum thermodynamic performance of the nozzle. 10 this study, a chemical equilibrium flow analysis code was developed for the design of hydrocarbon fueled rocket engines. 10 oder to understand the thermochemical characteristics occurring in a nozzle through the expansion process, such as recombination of chemical components and the accompanying energy recovery, chemical equilibrium flow analysis was carried out for the KSR-III rocket engine nozzles together with frozen flow and non-equilibrium flow analyses. The performance evaluation based on the present KSR-III nozzle flow analyses has provided an understanding of the thermochemical process in the nozzle and additionally, it has confirmed that the newly designed nozzle shape modified to have a reduced exit area ratio is an adequate design for obtaining an increased ground thrust.

• Journal of the Korean Society of Propulsion Engineers
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• v.15 no.4
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• pp.41-47
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• 2011

Turbopump, which is a part of 75 ton open cycle liquid rocket engine has a super sonic impulse turbine. This paper investigated the leading edge thickness effect on the turbine performance experimently. Two rotors were tested with the different leading edge thickness. The ratios (rotor thickness to Pitch) are 1.9 and 1.4 times to 30 ton turbine rotor. As a result, a rotor with 1.4 times ratio had a 1.5% higher efficiency gain than a rotor with 1.9 times ratio. The pressure ratio with the maximum efficiency on the same rotational speed was increased to the full expansion ratio of nozzle.

• Journal of the Korean Society of Propulsion Engineers
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• v.23 no.3
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• pp.28-34
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• 2019

The thrust control calculation according to the operation of plug nozzle for cold air test and the analysis of the flow characteristics of the pintle shape and operation pressure are performed. The numerical computation was verified by comparing the flow structure and the coefficient of thrust with the experimental data. It was confirmed that the nozzle throat was formed at the design position on each pintle shape, and thrust control up to 1:8 was achieved only by the stroke change. Finally, although the aerospike nozzle is autonomous, it is unfavorable in the under_expansion condition, if it is designed for a very low nozzle pressure ratio.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.30 no.5
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• pp.24-31
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• 2002

The effects of vortex generators, in the form of small delta-shaped tabs or thin tapes at an axi-symmetric supersonic nozzle exit, on the characteristics of supersonic jet flowfields are investigated by Schlieren images and Pitot-tube measurements. Small tabs as small as 1 % of the nozzle exit area can introduce streamwise vortices and produce a significant effect on the jet flowfield downstream of the nozzle. The effect is stronger for the cases of under-expanded jet than over- and perfect-expanded cases, introducing a larger flow entrainment. The effects of the angle of tabs with respect to the flow direction are also investigated, and for over-expanded jet cases, it is found that the tabs bended toward upstream can weaken the interaction strength and remove the Mach disc in the jet flowfield. Introduction of small proturbances inside the nozzle surface by attachment of thin tapes is also found to change the pressure distribution in the circumferential direction of the flowfield. Its effect is also found to be dependent on the jet expansion ratio.

• Journal of the Korean Society of Propulsion Engineers
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• v.4 no.3
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• pp.11-18
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• 2000

When supersonic underexpanded jets are exhausted from the nozzle, complex shock cell configurations such as barrel shock, expansion fan, Mach disc, and exhaust-gas jet boundary are appeared repetitively. The shock cell is smeared by turbulence dissipation and disappeared in long distance from the nozzle. When underexpanded jet is suddenly impinged on a flat plate, it forms very complex flow structure. In this paper, we solve compressible Wavier-Stokes equation adapting finite volume method to obtain jet impingement flow structure and compare calculated data with experimental ones. It is shown that numerical simulation data are in good agreement with experimental one in a short distance between nozzle exit and flat plate and little influence of underexpanded ratio is appeared in jet impingement now distribution.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2008.05a
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• pp.149-152
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• 2008

Effects of chamber configuration on combustion characteristic velocity of full-scale combustion chamber for 30-tonf-class liquid rocket engine were studied. The configurations of combustion chamber are ablative and channel cooling chamber (${\varepsilon}$=3.2) which have detachable mixing head, and single body regenerative cooling chamber which has nozzle expansion ratio of 3.5 and 12, respectively. The combustion chambers have chamber pressure of 53${\sim}$60 bar and propellant mass flow rate of 89 kg/s, and the injectors of all combustion chamber have recess number 1.0 and double-swirl characteristics. The hot firing test results at design point show that the combustion characteristic velocity of the regenerative cooling chamber which has nozzle expansion ratio of 12 is higher than that of other combustion chambers. The reasons for the above result are the increases of combustion pressure and enthalpy of kerosene which is heated due to cooling of the chamber wall before injection into the combustion field.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.05a
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• pp.428-432
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• 2010

A study on the forming characteristics of outer shell structure for thrust chamber nozzle extension has been performed. In order to identify anisotropy of cold rolled sheet metal, three types of tensile specimens according to the direction to the sheet rolling axis were prepared and tested, and Landford's values were obtained using the results and applied to structural analysis. Forming characteristics of the outer shell structure of the nozzle extension are investigated through manufacturing and forming of the full scale outer shell structures, and strain values obtained by the forming processes are compared to the numerical analysis results. The results obtained by this study will be utilized to design forming tools and processes for manufacturing other outer shell structures which have a bigger expansion area ratio.