• Journal of the Korean Society of Industry Convergence
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• v.11 no.2
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• pp.65-70
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• 2008

This study presents gas turbine cooling blade by using experimental and numerical works. The review cover researches related to cooling channels using finite element method in rotating blade. Also, the film cooling device and the heat transfer of the external surface of the blade are included. In addition, several methods to be used for the design of the blade, numerical method and experimental techniques are introduced. This work will contribute to improving the manufacturing of engine and the efficiency of gas turbine engines.

• 한국전산유체공학회:학술대회논문집
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• 2008.03b
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• pp.648-651
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• 2008

Numerical calculations are performed to simulate the film cooling effect of turbine blade tip with squealer rim. Because of high temperature of inside rim, squealer rim is damaged easily. Therefore many various cooling systems were used. The calculations are based on 100,000 Reynolds number in linear cascade model. A blade has 2% tip clearance and 8.4% rim height. The axial chord length and turning angle is 237mm, 126$^{\circ}$. Numerical calculations are performed without and with film cooling. In a film cooling in the cavity, hot spots of cavity were cooled effectively. However hot spots of suction side rim still remains. The CFD results show that the circulation flow in cavity of squealer tip affects the temperature rise of squealer rim. To maintain the blade integrity and avoid the excessive hot spot of blade, rearrangement of cooling hole is needed.

• Proceedings of the KSME Conference
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• 2001.11b
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• pp.30-35
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• 2001

To get higher efficiency of gas turbine, The designer should have more higher turbine inlet temperature (TIT). Today, modem gas turbine having sophisticated cooling scheme has TIT above $1,700^{\circ}C$. In the korea, many gas turbine having TIT above $1,300^{\circ}C$ was imported and being operated, but the gas with high TIT above $1,300^{\circ}C$ in the turbine will give damage to liner of combustor, and blade of turbine and etc. So frequently maintenance for parts enduring high temperature was performed. In this study, the heat transfer analysis of cooling air in the internal cooling channel (network analysis) and temperature analysis of the blade (Finite Element Analysis) in the first stage rotor was conducted for development of the optimal cooling passage design procedure. The results of network analysis and FEM analysis of blade show that the high temperature spot are occured at the leading edge, trailing edge near tip, and platform. so to get more reliable performance of gas turbine, the more efficient cooling method should be applied at the leading edge and tip section. and the thermal barrier coating on the blade surface has important role in cooling blade.

• 제어로봇시스템학회:학술대회논문집
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• 2003.10a
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• pp.237-240
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• 2003

Gas turbines are extensively used in flight propulsion, electrical power generation, and other industrial applications. During its life span, a turbine blade is taken out periodically for repair and maintenance. This includes re-coating the blade surface and re-drilling the cooling holes/channels. A successful laser re-drilling requires the measurement of a hole within the accuracy of ${\pm}0.15mm$ in position and ${\pm}3^{\circ}$ in orientation. Detection of gas turbine blade/vane cooling hole position and orientation thus becomes a very important step for the vane/blade repair process. The industry is in urgent need of an automated system to fulfill the above task. This paper proposes approaches and algorithms to detect the cooling hole position and orientation by using a vision system mounted on a robot arm. The channel orientation is determined based on the alignment of the vision system with the channel axis. The opening position of the channel is the intersection between the channel axis and the surface around the channel opening. Experimental results have indicated that the concept of cooling hole identification is feasible. It has been shown that the reproducible detection of cooling channel position is with +/- 0.15mm accuracy and cooling channel orientation is with +/$-\;3^{\circ}$ with the current test conditions. Average processing time to search and identify channel position and orientation is less than 1 minute.

• Proceedings of the Korean Society for Noise and Vibration Engineering Conference
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• 2007.11a
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• pp.1090-1094
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• 2007

The aim of this study is to estimate the cause of Cooling Tower vibration and eliminate the faults of fan with high vibration in spite of overhaul. The cause of high vibration was that the natural frequency of fan blade coincide with second blade pass frequency. To achieve reduction of Cooling Tower vibration, change motor speed from 1784rpm to 1714rpm, and then the vibration has reduced conspicuously.

• Advances in aircraft and spacecraft science
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• v.7 no.6
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• pp.537-551
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• 2020

Investigation of leading edge impingement cooling for first stage rotor blades in an aero-engine turbine, its effect on rotor temperature and trailing edge wake loss have been undertaken in this study. The rotor is modeled with the nozzle for attaining a more accurate simulation. The rotor blade is hollowed in order for the coolant to move inside. Also, plenum with the 15 jet nozzles are placed in it. The plenum is fed by compressed fresh air at the rotor hub. Engine operational and real condition is exerted as boundary condition. Rotor is inspected in two states: in existence of cooling technique and non-cooling state. Three-dimensional compressible and steady solutions of RANS equations with SST K-ω turbulent model has been performed for this numerical simulation. The results show that leading edge is one of the most critical regions because of stagnation formation in those areas. Another high temperature region is rotor blade tip for existence of tip leakage in this area and jet impingement cooling can effectively cover these regions. The rotation impact of the jet velocity from hub to tip caused a tendency in coolant streamlines to move toward the rotor blade tip. In addition, by discharging used coolant air from the trailing edge and ejecting it to the turbines main flow by means of the slot in trailing edge, which could reduce the trailing edge wake loss and a total decrease in the blade cooling loss penalty.

• 유체기계공업학회:학술대회논문집
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• 2001.11a
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• pp.334-338
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• 2001

In order to investigate the effects of various injection hole shapes on the film cooling of turbine blade, three test models having cylindrical and shaped holes were used. A three-dimensional Navier-Stokes code with standard k-$\epsilon$ model was used to compute the film cooling coefficient on the film cooled turbine blade. Over 330,000 grids were used to compute the flow over the blade. Mainstream Reynolds number based on the cylinder diameter was $7.1{\times}10^4$. The turbulence intensity kept at $5.0\%$ for all inlets. The effect of coolant blowing ratio was studied for various blowing ratios. For each blowing ratios, wall temperatures around the surface of test model were calculated. Temperature was visualized using cartesian cut-cell method to obtain traces of the injected secondary air on the test surface, so we could interpret the film effectiveness as temperature distributions.

• Journal of Institute of Convergence Technology
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• v.6 no.1
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• pp.37-41
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• 2016

The serpentine internal passage is located in turbine blade and it shows the variety heat transfer distribution. Especially, the Coriolis force, which is induced by blade rotation, makes different heat transfer distribution of the leading and trailing surfaces of serpentine internal passage. The different heat transfer is one of the reasons why the serpentine cooling passage shows low cooling performance in the rotating condition. So, this study tried to design the advanced the serpentine passage to consideration of the Coriolis force. The design concept of advanced serpentine cooling is maximizing cooling performance using the Coriolis force. So, the flow turns from leading surface to trailing surface in advanced serpentine passage to match the direction of Coriolis force and rotating force. We performed numerical analysis using CFX and compared the existing and advanced serpentine internal passage. This design change is induced the high heat transfer distribution of whole advanced serpentine internal passage surfaces.

• Journal of the Korean Society for Aviation and Aeronautics
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• v.14 no.1
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• pp.15-21
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• 2006

Parametric cycle analysis of a dual-spool, mixed exhaust turbofan engine with turbine blade cooling were described to investigate the effect of turbine blade cooling on the engine performance such as specific thrust and thrust specific fuel consumption. Coolant of low pressure turbine triggers high engine performance loss and cooling effect loss in high pressure turbine. Therefore low pressure turbine coolant should be much more considered for effective design.

• Journal of the Korean Society for Precision Engineering
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• v.20 no.8
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• pp.54-61
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• 2003

The first stage rotating blade of industrial gas turbine is one of the components that is normally run in exposed state at the highest temperature of the combustion gas stream. For this reason superior materials and advanced cooling technology are required to allow higher heat resisting characteristics of the component. The 1st stage blade of a selected commercial gas turbine blade made of directionally solidified Ni-based superalloy has a row of cooling holes on its trailing edge. In most cases, minor cracks have been found at some of the root cooling holes after one cycle operation (24,000 hrs) or even shorter operation time because of the high temperature gradient and the frequently alternating thermal stress. In the repair process, unfortunately, it is usually very difficult to get rid of the damage due to the fact that cracks are initiated at the root cooling hole and propagated deep into the hole. In this study, the feasibility of removing the sidewall cracks in the hole by utilizing EDM drilling has been investigated. Also the criteria of surface integrity for EDM drilling were established to achieve high quality repair as well as machining accuracy.