• The KSFM Journal of Fluid Machinery
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• 제15권5호
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• pp.27-31
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• 2012

Turbine inlet temperature is steadily increasing to achieve high specific thrust and efficiency of gas turbine engines. Turbine cooling technology is essential to increase turbine inlet temperature. For this study, a small or medium sized aircraft engine of 10,000 lbf class with the turbine inlet temperature of $1,400^{\circ}C$, the engine overall pressure ratio of 32.2, and the bypass ratio of 5 was set as the baseline model and its performance analysis was performed at the design point. The engine has the performance of 10,013 lbf thrust and the specific fuel consumption of 0.362 lbm/hr/lbf. The thrust and the specific fuel consumption of the baseline model were compared with those of similar class engines. Based on these results, the turbine design requirements were assigned. In addition, the parametric analysis of the engine, related to aerodynamic and cooling design of the high pressure turbine, was performed. Based on the baseline model engine, the influence of turbine inlet temperature, cooling flow ratio, and high pressure turbine efficiency variations on the engine performance was analyzed.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• 제43권4호
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• pp.326-333
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• 2015

A turbo fan engine performance analysis program combined with a particle swarm optimization(PSO) has been developed to optimize the major design parameters of the combat aircraft gas turbine engine. The optimized parameters includes bypass ratio, fan pressure ratio, high pressure compression ratio and burner exit temperature. The objective parameters have been determined using a multi-objective function consisting of the net thrust and specific fuel consumption along a weight function. The basic model for the combat aircraft gas turbine engine has been selected as the F404 turbofan engine which is widely used in the combat aircraft, F-18 and Korean high level training aircraft, T-50. The optimal conditions of four parameters have been obtained for various design conditions.

• Journal of computational fluids engineering
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• 제2권2호
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• pp.51-58
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• 1997

A reliable computational solver has been developed for the analysis of three-dimensional inviscid compressible flows around a nacelle of a high bypass ratio turbofan engine, The numerical algorithm is based on the modified Godunov scheme to allow the second order accuracy for space variables, while keeping the monotone features. Two step time integration is used not only to remove time step limitation but also to provide the second order accuracy in a time variable. The multi-block approach is employed to calculate the complex flow field, using an algebraic, conformal, and elliptic method. The exact solution of Riemann problem is used to define boundary conditions. The accuracy of the developed solver is validated by comparing its results around the isolated nacelle in the cruise flight regime with the solution obtained using a commercial code "RAMPANT. "

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 한국추진공학회 2010년도 제35회 추계학술대회논문집
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• pp.833-837
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• 2010

This study aims at finding an optimal exhaust diffuser design of a high altitude testing chamber for a low bypass turbofan engine (F404-402) with thrust pound force of 17,700 and air mass flow rate of 66kg/s ejecting at a speed of Mach 1.66. The final proposed ejector size has better pressure recovery characteristics and targets to reduce operational cost at engine performance testing. Conventional high altitude test chamber layout was adopted and first drawn in two dimensions using Autocad software so as to determine the gas path, the ejector frontal size was then determined from gas dynamics equations considering traditional gas ejection method where both the engine exhaust and cell cooling air are exhausted via the ejector. Modification to a smaller ejector with an alternative secondary cell cooling exhaust port was then performed and modelled in 3D using Solid Works software.

• Journal of the Korean Society of Propulsion Engineers
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• 제19권5호
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• pp.1-14
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• 2015

In case of a gas turbine engine for supersonic operation, the engine have a wide range of operating inlet mass flow rate and required high performance such as thrust and fuel consumption. Therefore, variable system and its optimal control logic are essentially needed. In this work, a method for performance prediction of a gas turbine engine with variable system compressor and its control scheme were developed. Conceptual design of compact acuation system for the operation of the variable system was also conducted. The performance of a low-bypass ratio mixed flow turbofan engine was analyzed, and it was observed that the surge margin of the engine is improved at off-design condition by applying the control scheme.