• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.04a
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• pp.309-317
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• 2011

In this study, two-stage hypersonic scramjet vehicle was designed for the flight condition of Mach number 6. In order to launch at sea level and Mach number 0, two stage concept was applied. The first stage of the vehicle is rocket-powered and is mounted under the second stage. The second stage is scramjet-powered propulsion system and has wing. The suggested mission scenario is to deliver 0.2 ton payload to the range less of 2000km. For the first step of conceptual design, trajectory of air vehicle was calculated by 3-DOF trajectory code. Based on the result of trajectory code, scramjet engine design and mass estimation were performed by non-equilibrium nozzle flow code and NASA's HASA model, respectively. In order to find best solution, all step of designing process was iterated until they were converged.

• 유체기계공업학회:학술대회논문집
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• 2001.11a
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• pp.303-308
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• 2001

A rotating blade of steam turbines is designed using blade design system. To minimize the design time. quasi three dimensional flow analysis code is adopted to calculate blade section. The blade section lies on a streamline determined by previous steam turbine design procedures. The blade design system makes a transform of streamline coordinates, (m, r$\theta$), to (m', $\theta$) coordinates and all design procedure except 3 dimensional stack-up is performed in the coordinates. Each designed blade section is stacked-up and whole 3 dimensional blade can be modified by correcting 2D section, repeatly. The full 3D numerial analysis for the one stage including designed rotating blade will be performed later

• Proceedings of the Korean Society for Noise and Vibration Engineering Conference
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• 2011.10a
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• pp.687-692
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• 2011

Experimental analysis has been carried out on noise reduction of subsonic jet flow with chevron nozzle to investigate relationships between geometry variation of chevron nozzles and jet noise reduction. Including base nozzle, seven nozzles are tested at Mach number 0.9. Parametric variables are chevron count, chevron length, and chevron shape. From these experiment, the more chevron count increase, the more jet noise reduction gain. Varying the chevron length, short chevron nozzle increase the jet noise at certain direction all the more. Chevron shape is also considered as important factor to reduce the jet noise.

• International Journal of Fluid Machinery and Systems
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• v.6 no.4
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• pp.170-176
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• 2013

Counter Rotating Turbine (CRT) is an axial turbine with a nozzle followed by a rotor and another rotor that rotates in the opposite direction of the first one. Axial spacing between blade rows plays major role in its performance. Present work involves computationally studying the performance and flow field of CRT with axial spacing of 10, 30 and 70% for different mass flow rates. The turbine components are modeled for all the three spacing. Velocity, pressure, entropy and Mach number distributions across turbine stage are analyzed. Effect of spacing on losses and performance in case of stage, Rotor1 and Rotor2 are elaborated. Results confirm that an optimum axial spacing between turbine components can be obtained for the improved performance of CRT.

• 한국전산유체공학회:학술대회논문집
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• 2005.10a
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• pp.287-292
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• 2005

The flowfields generated by gaseous slot injection into a supersonic flow at a Mach number of 3.75 and a Reynolds number of $2.07{\times}10^7$ are simulated numerically. Fine-scale turbulence effects are represented by a two-equation(k-w SST model) closure model which includes $y^+$ effects on the turbulence model. Grid convergence index(GCI) is also considered to provide a measure of uncertainty of the grid convergence. Comparison is made with experimental data and other turbulence model in term of surface static pressure distributions, the length of the upstream separation region, and the penetration height. Results indicate that the k-w SST model correctly predicts mean surface pressure distribution and upstream separation length. However, it is also observed that the numerical simulation over predicts the pressure spike and penetration height compared with experimental data. All these results are taken within $1\%$ error band of grid convergence.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2000.04a
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• pp.8-8
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• 2000

In this paper are presented main power and gasdynamic characteristics of C-l6VK hypersonic test cell of Research Test Center of CIAM. Gasdynamic adjustment of the C-l6VK test cell was carried out with the working section constructed on scheme of Ramjet/scramjet test in free stream. Gasdynamic adjustment was conducted stage by stage in tile following sequence. First, check and preparation of all technical systems and checking measuring system. Than determination of the characteristics of test cell on cold (without the heating of air at entrance) regime and determination of the characteristics of test cell on regimes with the heating of air. Finally determination of tile characteristics of test cell with the loading of the working part by object. On tile final stage of gasdynamic adjustment two experiments with tile axisymmetric Scramjet model loaded into the working part of test cell were conducted. The first experiment was conducted with the purpose of determination of flow parameters with the object leaded into the working part and verification of experiment cyclogram. The second experiment was conducted with injection of hydrogen into the combustion chamber of object, that is tile conditions on test cell simulated Scramjet flight Mach number M = 6. Such methodology of gasdynamic adjustment allows to determine influence of experimental object on flow parameters in the working part at different conditions of experiment (with the burning in combustion chamber of object and without the homing), and also to compare flow characteristics in the object duct.

• Advances in aircraft and spacecraft science
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• v.10 no.2
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• pp.141-158
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• 2023

In current research work, the aerodynamics performance of a newly designed large flying V aircraft is numerically investigated. Three Flying V configurations, with V-angles of 50°, 70° and 90° that represent the minimum, moderate, and maximum configurations respectively, were designed and modeled to assess their aerodynamic performance at cruise flight conditions. The unstructured mesh was developed using ICEM CFD and Ansys-Fluent was used as an aerodynamic solver. The developed models were numerically simulated at cruise flight conditions with a Mach number equal to 0.15. K-ω SST turbulence model was chosen to account for flow turbulence.The authors performed steady flow simulations.The results obtained from the experimentation reveal that the maximum main angle configuration of 90° had the highest C_Lmax value of 0.46 compared to other configurations. While the drag coefficient remained the same for all three configurations, the 50° V-angle configuration achieved the maximum stall angle of 35°. With limited stall delay benefits, the flying V possesses no sufficient stability, due to the flow separation detected at whole elevon and winglet suction side areas at AoA equal and higher than 30°.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2009.11a
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• pp.573-577
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• 2009

Dynamic behavior simulation of supersonic engine was performed and PI control algorithm was studied for the buzz control in the inlet and the thrust control. Firstly, required thrust was tracked according to the fuel flow control and then inlet pressure was regulated through the nozzle throat area control so that the buzz margin has the positive all the time. The control was performed according to the change of flight Mach number, altitude and angle of attack. The proportional gain and the integral gain for regulating the buzz margin was induced and simulated. In the results, it was confirmed and satisfied that control target in the operating area was changed the angle of attack from $0^{\circ}$ to $10^{\circ}$ at the flight Mach number of 2.1~3.0.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.36 no.8
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• pp.774-779
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• 2008

An experimental study on the micro-supersonic jet flow fields has been carried out. A sonic nozzle of 440 ${\mu}m$-exit diameter and a Laval nozzle of 800 ${\mu}m$ exit diameter with the nozzle exit Mach number 2.0 were fabricated by stretching a micro Pyrex glass tube for the present experiments. Schlieren flow visualization and Pitot pressure distribution of the jet flow field were obtained. Representative characteristics of the jet flow fields such as, supersonic length, jet core length, similarity of the velocity field, and jet spreading rates, have been observed. All the results were compared to previous observations of larger supersonic jets of higher Reynolds numbers, and it was found that overall characteristics of the micro supersonic jet are qualitatively similar as those of the higher Reynolds number jets, except the jet core length and the jet spreading rate.

• Journal of computational fluids engineering
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• v.18 no.2
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• pp.66-71
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• 2013

In this paper, the supersonic flows around space launch vehicles have been numerically simulated by using a 3-D RANS flow solver. The focus of the study was made for investigating plume-induced flow separation(PIFS). For this purpose, a vertex-centered finite-volume method was utilized in conjunction with 2nd-order Roe's FDS to discretize the inviscid fluxes. The viscous fluxes were computed based on central differencing. The Spalart-Allmaras model was employed for the closure of turbulence. The Gauss-Seidel iteration was used for time integration. To validate the flow solver, calculation was made for the 0.04 scale model of the Saturn-5 launch vehicle at the supersonic flow condition without exhaust plume, and the predicted results were compared with the experimental data. Good agreements were obtained between the present results and the experiment for the surface pressure coefficient and the Mach number distribution inside the boundary layer. Additional calculations were made for the real scale of the Saturn-5 configuration with exhaust plume. The flow characteristics were analyzed, and the PIFS distances were validated by comparing with the flight data. The KSLV-1 is also simulated at the several altitude conditions. In case of the KSLV-1, PIFS was not observed at all conditions, and it is expected that PIFS is affected by the nozzle position.