• Journal of the Korean Society of Propulsion Engineers
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• v.24 no.1
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• pp.47-57
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• 2020

The present study investigates to improve the film cooling performance on the vane endwall. Numerical simulation was conducted to analyze film cooling characteristics on the vane endwall. Six different hole arrangements were designed considering flow characteristics on the endwall. The results showed that the film cooling effectiveness was low on the pressure side and nozzle throat in the base case, because coolant was deflected from the pressure side to the suction side. On the other hand, when the holes were installed near the pressure side, the film cooling effectiveness was enhanced on the pressure side and nozzle throat, because the coolant was less affected by cross flow. Therefore, the film cooling effectiveness increases about 16% compared to the base hole arrangement.

• International Journal of Fluid Machinery and Systems
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• v.3 no.2
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• pp.150-159
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• 2010

Laidback fan shaped film-cooling hole is formulated numerically and optimized with the help of three-dimensional numerical analysis, surrogate methods, and the multi-objective evolutionary algorithm. As Pareto optimal front produces a set of optimal solutions, the trends of objective functions with design variables are predicted by hybrid multi-objective evolutionary algorithm. The problem is defined by four geometric design variables, the injection angle of the hole, the lateral expansion angle of the diffuser, the forward expansion angle of the hole, and the ratio of the length to the diameter of the hole, to maximize the film-cooling effectiveness compromising with the aerodynamic loss. The objective function values are numerically evaluated through Reynolds- averaged Navier-Stokes analysis at the designs that are selected through the Latin hypercube sampling method. Using these numerical simulation results, the Response Surface Approximation model are constructed for each objective function and a hybrid multi-objective evolutionary algorithm is applied to obtain the Pareto optimal front. The clustered points from Pareto optimal front were evaluated by flow analysis. These designs give enhanced objective function values in comparison with the experimental designs.

• Journal of the Korea Academia-Industrial cooperation Society
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• v.19 no.9
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• pp.69-76
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• 2018

The present study analyzed the effect of film hole position of 45 degree ribbed cooling channel on film cooling performance of gas turbine blades. We also investigated the influence of the ribs under the fixed blowing ratio. Three-dimensional numerical model was constructed and extensive simulation was conducted using the commercial code (Fluent ver. 17.0) under steady-state condition. Base on the simulation results, We investigated the cooling effectiveness, flow velocity, streamline, and pressure coefficient. Moreover, We analyzed the effect of cooling hole position on ejection of the secondary flow caused by the rib structure. From the results, It was found that internal flow of the cooling channel forms a vortex pair in the counterclockwise from the top side, and clockwise from the bottom side. For the channels with ribs, the vortex flow generated by the ribs caused a higher pressure difference near the hole outlet, resulting in at least 12% higher cooling effectiveness than the channel without ribs. Additionally, when the hole is located on the left side of the ribbed channel (Rib-Left), it can be found that the secondary flow generated by the ribs hits against wall surface near the hole to form a flow in the direction of the hole inclination angle. Therefore, It is considered that the region where the cooling gas discharged to the blade surface stays in the main flow boundary layer is wider than the other cases. In this case, The largest pressure coefficient difference was observed near the outlet of the hole, and as a result, the discharge of the cooling gas was accelerated and the cooling efficiency was slightly increased.

• 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 the Korean Society of Propulsion Engineers
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• v.12 no.5
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• pp.67-72
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• 2008

The film cooling effectiveness on a flat plate measured by pressure sensitive paint technique. Six film cooling hole were fabricated on a flat plate with 30 degree angle with respect to the surface and three blowing ratios of 0.5, 1, and 2 were tested. Results showed that PSP technique successfully evaluated the distribution of film cooling effectiveness and showed similar results with references. The film cooling effectiveness near the film cooling holes was higher for lower blowing ratio case. As the blowing ratio was increased, the film cooling effectiveness near the film cooling hole decreased due to the lift off of the coolant. At far downstream, the film cooling effectiveness for higher blowing ratio was higher due to the coolant reattachment.

• 유체기계공업학회:학술대회논문집
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• 2002.12a
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• pp.298-303
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• 2002

Dispersion of coolant jets in a film cooling flow field is the result of a highly complex interaction between the film cooling jets and the mainstream. In order to investigate the effect of blowing ratios on the film cooling of turbine blade, cylindrical body model was used. Mainstream Reynolds number based on the cylinder diameter was $7.1{\times}10^4$. The effect of coolant flow rates was studied for blowing ratios of 0.7, 0.9, 1.2 and 1.5, respectively. The temperature distribution of the cylindrical model surface is visualized by infrared thermography (IRT). Results show that the film-cooling performance could be significantly improved by the shaped injection holes. For higher blowing ratio, the spanwise-diffused injection holes are better due to the lower momentum flux away from the wall plane at the hole exit.

• 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.

• The KSFM Journal of Fluid Machinery
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• v.17 no.6
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• pp.52-58
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• 2014

In the present work, design optimization of film-cooling hole array on the pressure side of high pressure turbine nozzle was conducted. There are four rows of fan-shaped film cooling holes on the nozzle pressure side surface and each row has a straight array of holes in the spanwise direction for baseline model. For design optimization, hole distributions in streamwise and spanwise directions for three rows of holes except first row are parameterized as a 2nd-order shape function. Three-dimensional compressible RANS equations are used for flow and thermal analysis around the nozzle surface and optimization technique using Design of Experiment, Kriging surrogate model and Genetic Algorithm is used. The results shows that averaged adiabatic wall temperature at the whole nozzle surface decreases about 2.7% and averaged film cooling effectiveness at the pressure side of nozzle increased about 8.2%.

• Journal of the Korean Society of Propulsion Engineers
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• v.25 no.4
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• pp.71-77
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• 2021

Several cooling techinques have been studied for protecting gas turbine blades from hot gas. In terms of film-cooling techniques, various shapes of film cooling holes have been studied including fan shaped holes, which are used on gas turbine blades. However, owing to increasing demands on smaller gas turbines, a research on film-cooling holes on thin walls is required. This study was conducted at blowing ratios of 1 and 2, using numerical analysis. Through the numerical analysis, the effect of geometrical parameters on the effectiveness of fan-shaped hole film cooling was studied. Moreover, optimization was performed on three geometrical parameters: metering length, lateral expansion angle and forward expansion angle. As a result, we realized that the optimal fan-shaped holes on each blowing ratio were found to have very similar geometry and cooling performance.

• The KSFM Journal of Fluid Machinery
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• v.6 no.3 s.20
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• pp.21-27
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• 2003

Dispersion of coolant jets in a film cooling flow field is the result of a highly complex interaction between the film cooling jets and the mainstream. In order to investigate the effect of blowing ratios on the film cooling of a turbine blade, cylindrical body model is used. Mainstream Reynolds number based on the cylinder diameter is $7.1{\times}10^4$. The effects of coolant flow rates are studied for blowing ratios of 0.7, 1.0, 1.3 and 1.7, respectively. The temperature distribution of the cylindrical model surface is visualized with infrared thermography (IRT). Results show that the film cooling performance could be significantly improved by the shaped injection holes. For higher blowing ratio, the spanwise-diffused injection holes are better due to the lower momentum flux away from the wall plane at the hole exit.