• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2010.05a
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• pp.349-352
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• 2010

A study is analyzed on the transient flow of supersonic diffuser and performed on the of supersonic diffuser with Computational Fluid Dynamic. The flow field of supersonic diffuser is calculated using Axisymmetric two-dimensional Navier-Stokes equation with $k-{\epsilon}$ turbulence model. The transient simulation is compared in terms of mach number and temperature of vacuum chamber according to the chamber pressure of starting transient on Liquid rocket engine.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.867-868
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• 2011

Transient Analysis on heat transfer of rocket engine combustion chamber and wall temperature variation was carried out, especially, calculations of LOx/kerosene rocket engine with/without fuel film-cooling were conducted. Convective and radiative heat flux inside combustion chamber wall were calculated by the empirical equations for rocket engine combustion, and conduction of wall interior was calculated by numerical method with 2D axisymmetric grid. In this calculations the transient variations of wall temperature, the location changes of peak temperature and so on affected by film-cooling were analyzed.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.321-324
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• 2011

A computational analysis of nozzle flow characteristics and plume structure using Reynolds-averaged Navier-Stokes equations with $k-{\omega}$ SST turbulence model was conducted to examine performance of the supersonic nozzle employed in a small liquid-rocket engine for ground firing test. Computed results and experimental outcome of 2-D converging-diverging nozzle flow were compared for verifying the computational capability as well as the turbulence model validity. Numerical computations of 2-D axisymmetric nozzle flow was carried out with the selected model. As a result, flow separation with backflow appeared around the nozzle exit. This investigation was reported as a background data for the optimal nozzle design of small liquid-propellant rocket engine for ground test.

• Journal of the Korean Society of Propulsion Engineers
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• v.7 no.4
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• pp.46-52
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• 2003

This paper presents the numerical thermal analysis for regeneratively cooled rocket thrust chambers. An integrated numerical model incorporates computational fluid dynamics for the hot-gas thermal environment, and thermal analysis for the liner and coolant channels. The flow and temperature fields in rocket thrust chambers is assumed to be axisymmetric steady state which is presumed to the combustion liner. The heat flux computed from nozzle flow is used to predict the temperature distribution of the combustion liner As a result, we present the wall temperature of combustion liner and the temperature change of coolant.

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

An axisymmetric compressible flow solver accounting for chemical equilibrium has been developed as an analysis tool exclusively suitable for performance prediction and optimum contour design of LRE thrust chamber. By virtue of several features focusing on user-friendliness and effectiveness including automatical grid generation and iterative calculations with changes in design parameters prescribed through only one keyword-type input file, a design engineer can evaluate very fast and easily the influences of various design inputs such as geometrical parameters and operating conditions on propulsive performance. Validations have been carried out for various aspects by detailed comparisons with the result of CEA code, experimental data of JPL nozzle, actual data for two historical engines, and ReTF data for KSR-III.

• Journal of the Korean Society of Propulsion Engineers
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• v.19 no.3
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• pp.65-72
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• 2015

In present study, a supersonic inlet for dual mode ramjets or RBCC/TBCC engines with a wide range of flight Mach numbers is designed. A conical variable inlet configuration is chosen for the inlet design. Geometric relations with angles of compression cones and conical shock waves are used for the design of the inlet configuration. The performance of the supersonic inlet is confirmed by the numerical analysis. The capture area ratio is maintained around 100% from Mach 3 to 8 conditions.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.32 no.8
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• pp.29-36
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• 2004

Supersonic inlet buzz can be defined as unstable subcritical operation associated with fluctuating internal pressures and a shock pattern oscillating about the inlet entrance. The flow pulsations could result in flameout in the combustor or even structural damage to the engine. An experimental study was conducted to investigate the phenomenon of supersonic inlet buzz on axisymmetric, external-compression inlet. An inlet model with a cowl lip diameter of 30mm was tested at a free stream Mach number of 2.0. Subcritical instability was investigated by considering the frequency of pressure pulsation and shock wave structure at the inlet entrance. The results obtained show that total pressure recovery ratios were varied from 0.42 to 0.78, and capture area ratio from 0.34 to 0.98. The frequency of the subcritical flow increased with decrease in capture area ratios. Frequency was measured at $224{\sim}240Hz$.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.487-491
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• 2009

A study is analyzed on the design factor of Center-body diffuser and performed on conceptual design of Center-body diffuser with Computational Fluid Dynamic. The flow field of Center-body diffuser is calculated using Axisymmetric two-dimensional Navier-Stokes equation with $k-{\omega}$ turbulence model. The center-body diffuser is compared with second throat exhaust diffuser in terms of starting pressure, the degree of vacuum pressure, the design factor.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2003.05a
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• pp.9-12
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• 2003

A cooling performance analysis has been made in the 8-channel calorimeter based on sub-scale KSR-III engine. Three-dimensional heat transfer analysis in cooling channels has been performed using the heat flux distribution through the chamber wall predicted from axi-symmetric compressible flow inside the combustion chamber. The heat flux distribution was verified against the published literature. Presented for the development and operation of the calorimeter are the coolant pressure drop, coolant temperature rise and the maximum chamber wall temperature.

• Journal of the Korean Society of Propulsion Engineers
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• v.7 no.3
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• pp.8-14
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• 2003

A cooling performance analysis has been made in the 8-channel calorimeter based on sub-scale KSR-III engine. Three-dimensional heat transfer analysis in cooling channels has been performed using the heat flux distribution through the chamber wall predicted from axi-symmetric compressible flow inside the combustion chamber. The heat flux distribution is verified against the published literature. Presented for the development and operation of the calorimeter are the coolant pressure drop, coolant temperature rise and the maximum chamber wall temperature. Required coolant flow rate is determined for given chamber pressure. Cooling performance is also predicted for temperature dependant coolant properties.