• Proceedings of the KSME Conference
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• 2001.06d
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• pp.989-994
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• 2001

Numerical simulations of high temperature catalytic combustion have been performed for the application to a gas turbine combustor. Dependences of inlet temperature and pressure on the distributions of temperature and species concentrations were investigated using plug flow model with detailed homogeneous and heterogeneous chemistries of methane-air mixtures. Honeycomb typecombustor deposited with Pt catalyst of 100mm in length and 26mm in diameter is used. The results show that rapid increase of temperature profile occurs earlier with the increase of inlet temperature and the decrease of inlet pressure. The condition which catalytic combustion is stabilized exists at certain range of inlet temperature and pressure. The state of catalytic combustion is also confirmed by the distributions of species concentration.

• Journal of the Korean Society of Propulsion Engineers
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• v.17 no.1
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• pp.97-102
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• 2013

A prestudy on expendable turbine engine for high-speed vehicle was conducted. After two possible mission profiles were established to decide the engine requirements, design point analysis was performed with the values of design parameter which were obtained from similar class engines, references, etc. The results showed that specific net thrust and specific fuel consumption with turbine inlet temperature of 3,600 R are 2,599.4 ft/s and 1.483 lb/(lb*h) respectively at the flight condition of sea level, Mach 1.2. It was also found that major design parameters for determining maximum net thrust were turbine inlet temperature for low supersonic and transonic flight speed and compressor exit temperature for high supersonic flight speed from the results of performance analysis on the two possible mission profiles. In addition, simple turbojet engine with an axial compressor, a straight annular combustor, an one stage axial turbine and a fixed throat area converge-diverge exhaust nozzle was proposed as the configuration of simple low cost lightweight turbine engine.

• Proceedings of the Korean Society of Propulsion Engineers Conference
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• 2011.11a
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• pp.629-634
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• 2011

A prestudy on expendable turbine engine for high-speed vehicle was conducted. The two possible mission profiles were established to decide the engine requirements and Design Point, and Design Point analysis was performed with the values of design parameter which were obtained from similar class engines and technical references. The results showed that Specific Net Thrust is 2599.4 ft/s and Specific Fuel Consumption is 1.483 lb/($lb^*h$) at the flight condition of Sea Level, Mach 1.2. It was also found through the performance analysis on the two possible mission profiles that major design parameters for determining Net Thrust were Turbine Inlet Temperature for low supersonic flight speed and Compressor Exit Temperature for high supersonic flight speed. In addition, simple turbojet engine with axial compressor, straight annular combustor, axial turbine and fixed throat area converge-diverge exhaust nozzle was proposed as the configuration of simple low cost light engine.

• Transactions of the Korean Society of Mechanical Engineers A
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• v.39 no.11
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• pp.1145-1151
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• 2015

High pressure components of a gas turbine engine must operate for a long life under severe conditions in order to maximize the performance and minimize the maintenance cost. Enhanced cooling design, thermal barrier coating techniques, and nickel-base superalloys have been applied for overcoming them and furthermore, material modeling, finite element analysis, statistical techniques, and etc. in design stage have been utilized widely. This article aims to evaluate the effects on the low cycle fatigue life of the high pressure turbine nozzle caused by different turbine inlet temperature profiles and installation conditions and to investigate the most favorable operating condition to the turbine nozzle. To achieve it, the structural analysis, which utilized the results of conjugate heat transfer analysis as loading boundary conditions, was performed and its results were the input for the assessment of low cycle fatigue life at several critical zones.

• Journal of the Korean Society for Aeronautical & Space Sciences
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• v.44 no.12
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• pp.1062-1070
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• 2016

The heat transfer and flow characteristics of gas turbine blade tip were investigated in this paper by using the conjugate heat transfer analysis. The rotor inlet boundary condition profile which was taken from the first stage nozzle outlet was used to analyse. The profile contained the velocity and temperature information. This study presents the influence of tip clearance about aerodynamic loss, heat transfer coefficient and film cooling effectiveness with the squealer tip designed blade model which tip clearance variation range from 1% to 2.5% of span. Results showed that the aerodynamic loss and the heat transfer coefficient were increased when the tip clearance was increased. Especially when the tip clearance was 2% of the span, the average heat transfer coefficient on the tip region was increased obviously. The film cooling effectiveness of tip region was increasing with decreasing of the tip clearance. There was high film cooling effectiveness at cavity and near tip hole region.

• Journal of the Korean Society of Propulsion Engineers
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• v.21 no.3
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• pp.49-55
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• 2017

In this study, gas turbine engine inlet duct was designed to satisfy uniform flow at aerodynamic interface plane (AIP). Haack-series was selected as nose cone profile and duct outer radius($r_o$) was designed to satisfy to match with area change rate between the nose cone and outer duct wall by the 1-D sizing. The design object of the inlet duct wall profile which has the gradual area change rate was uniform Mach number in the core flow region and minimum boundary later thickness at the both inner nose wall and outer duct wall. The flow characteristics inside the inlet duct was evaluated using CFD. The static pressure distribution at the AIP showed uniform pattern within 0.16%. Based on Mach number profile, the boundary layer thickness was 2% of channel height. Kiel temperature rake location was decided less than 100 mm in front of nose cone where the Mach number is less than 0.1 in order to maximize the temperature probe recovery rate.

• The KSFM Journal of Fluid Machinery
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• v.18 no.2
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• pp.60-66
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• 2015

Conjugate heat transfer analysis was performed to investigate the flow and cooling performance of the high pressure turbine nozzle of gas turbine engine. The CHT code was verified by comparison between CFD results and experimental results of C3X vane. The combination of k-${\omega}$ based SST turbulence model and transition model was used to solve the flow and thermal field of the fluid zone and the material property of CMSX-4 was applied to the solid zone. The turbine nozzle has two internal cooling channels and each channel has a complex cooling configurations, such as the film cooling, jet impingement, pedestal and rib turbulator. The parabolic temperature profile was given to the inlet condition of the nozzle to simulate the combustor exit condition. The flow characteristics were analyzed by comparing with uncooled nozzle vane. The Mach number around the vane increased due to the increase of coolant mass flow flowed in the main flow passage. The maximum cooling effectiveness (91 %) at the vane surface is located in the middle of pressure side which is effected by the film cooling and the rib turbulrator. The region of the minimum cooling effectiveness (44.8 %) was positioned at the leading edge. And the results show that the TBC layer increases the average cooling effectiveness up to 18 %.