Effect of Non-Axisymmetric Endwall Profiling on Heat Transfer and Film Cooling Effectiveness of a Transonic Rotor Blade

2019 ◽  
Author(s):  
Jinjin Li ◽  
Xin Yan ◽  
Kun He
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Rotor Blade ◽  
Stanton Number ◽  
Total Pressure Loss ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Transonic Rotor

Abstract Effects of non-axisymmetric endwall profiling on total pressure loss, heat transfer and film cooling effectiveness of a transonic rotor blade were numerically investigated. The numerical methods, including the turbulence model and grid sensitivity, were validated with the existing experimental data. To reduce thermal load on endwall, non-axisymmetric endwall profiling near leading edge and at pressure-side corner area were performed with a range of contour amplitudes. Heat transfer and flow fields near the profiled endwalls were analyzed and also compared to the plain endwall configuration. On the profiled endwall, three kinds of cooling-holes, i.e. cylindrical holes, rounded-rectangular holes and elliptical holes, were arranged, and film cooling effect was investigated at three blowing ratios. Results indicate that, with endwall profiling, the area-averaged Stanton number on endwall is reduced by 7.71% and total pressure loss in cascade is reduced by 11.07%. Among three kinds of cooling holes, arrangement of elliptical hole performs the best film cooling effect on profiled endwall. Compared with plain endwall, non-axisymmetric endwall with elliptical cooling holes improves film cooling coverage by 10.87%, reduces the Stanton number by 8.88% and increases the net heat flux reduction performance by 4% at M = 0.7.

Effect of Non-Axisymmetric Endwall Profiling on Heat Transfer and Film Cooling Effectiveness of a Transonic Rotor Blade

2020 ◽  
Vol 142 (5) ◽  
Author(s):  
Jinjin Li ◽  
Xin Yan ◽  
Kun He
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Rotor Blade ◽  
Stanton Number ◽  
Total Pressure Loss ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Transonic Rotor

Abstract Effects of non-axisymmetric endwall profiling on total pressure loss, heat transfer, and film cooling effectiveness of a transonic rotor blade were numerically investigated. The numerical methods, including the turbulence model and grid sensitivity, were validated with the existing experimental data. To reduce the thermal load on endwall, non-axisymmetric endwall profiling near leading edge and at pressure-side corner area was performed with a range of contour amplitudes. Heat transfer and flow fields near the profiled endwalls were analyzed and also compared with the plain endwall configuration. On the profiled endwall, three kinds of cooling holes, i.e., cylindrical holes, rounded-rectangular holes, and elliptical holes, were arranged, and film cooling effect was investigated at three blowing ratios. Results indicate that, with endwall profiling, the area-averaged Stanton number on endwall is reduced by 7.71% and total pressure loss in cascade is reduced by 11.07%. Among three kinds of cooling holes, the arrangement of the elliptical hole performs the best film cooling effect on the profiled endwall. Compared with the plain endwall, non-axisymmetric endwall with elliptical cooling holes improves film cooling coverage by 10.87%, reduces the Stanton number by 8.88%, and increases the net heat flux reduction performance by 4% at M = 0.7.

Investigations of Heat Transfer and Film Cooling Effect on a Worn Squealer Tip

2020 ◽  
Author(s):  
Mingliang Ye ◽  
Xin Yan
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Total Pressure Loss ◽  
Wear Depth ◽  
Original Design ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Cavity Floor

Abstract Wear damage commonly occurs in modern gas turbine rotor blade tip due to relative movements and expansions between rotating and stationary parts. Tip wear has a significant impact on the aerodynamic, heat transfer and cooling performance of rotor blades, thus threatening the economy and safety of whole gas turbine system. Based on a simple linear wear model, this paper numerically investigates the aerodynamic, heat transfer and film cooling performance of a worn squealer tip with three starting-locations of wear (sl = 25%Cax, 50%Cax and 75%Cax) and five wear-depths (wd = 0.82%, 1.64%, 2.46%, 3.28% and 4.10%). Firstly, based on the existing experimental data, numerical methods and grid independence are examined carefully. Then, three dimensional flow fields, total pressure loss distributions, heat transfer coefficients and film cooling effectiveness in worn squealer tip region are computed, which are compared with the original design case. The results show that, with the increase of wear depth and the movement of wear starting-location to the leading edge, the scale and intensity of cavity vortex are increased, which results in the extended high heat transfer area on cavity floor near the leading edge. Wear makes more coolant flow out of the cavity, and reduces the area-averaged film cooling effectiveness at the bottom of cavity, but increases the film cooling effectiveness on pressure-side rim. The increase of wear depth makes more flow leak through the tip gap, thus increasing the scale and intensity of leakage vortex and further increasing the total pressure loss in the tip gap. Compared with the original design case, as the wear depth is increased from 0.82% to 4.10%, the mass-averaged total pressure loss in cascade is increased by 0.3–6.7%, the area-averaged heat transfer coefficient on cavity floor is increased by 1.7–29.1% while on squealer rim it is decreased by 3.1–26.3%, and the area-averaged film cooling effectiveness on cavity floor is decreased by 0.035 at most while on squealer rim it is increased by 0.064 at most.

Turbine Blade Platform Film Cooling With Typical Stator-Rotor Purge Flow and Discrete-Hole Film Cooling

2009 ◽  
Vol 131 (4) ◽  
Author(s):  
Zhihong Gao ◽  
Diganta Narzary ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Total Pressure Loss ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Cooling Hole ◽  
Purge Flow ◽  
Cylindrical Holes ◽  
Film Cooling Holes

This paper is focused on the effect of film-hole configurations on platform film cooling. The platform is cooled by purge flow from a simulated stator-rotor seal combined with discrete-hole film cooling within the blade passage. The cylindrical holes and laidback fan-shaped holes are assessed in terms of film-cooling effectiveness and total pressure loss. Lined up with the freestream streamwise direction, the film holes are arranged on the platform with two different layouts. In one layout, the film-cooling holes are divided into two rows and more concentrated on the pressure side of the passage. In the other layout, the film-cooling holes are divided into four rows and loosely distributed on the platform. Four film-cooling hole configurations are investigated totally. Testing was done in a five-blade cascade with medium high Mach number condition (0.27 and 0.44 at the inlet and the exit, respectively). The detailed film-cooling effectiveness distributions on the platform were obtained using pressure sensitive paint technique. Results show that the combined cooling scheme (slot purge flow cooling combined with discrete-hole film cooling) is able to provide full film coverage on the platform. The shaped holes present higher film-cooling effectiveness and wider film coverage than the cylindrical holes, particularly at higher blowing ratios. The hole layout affects the local film-cooling effectiveness. The shaped holes also show the advantage over the cylindrical holes with lower total pressure loss.

Turbine Blade Platform Film Cooling With Typical Stator-Rotor Purge Flow and Discrete-Hole Film Cooling

2008 ◽  
Author(s):  
Zhihong Gao ◽  
Diganta Narzary ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Total Pressure Loss ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Cooling Hole ◽  
Purge Flow ◽  
Cylindrical Holes ◽  
Film Cooling Holes

This paper is focused on the effect of film hole configurations on platform film cooling. The platform is cooled by purge flow from a simulated stator-rotor seal combined with discrete-hole film cooling within the blade passage. The cylindrical holes and laidback fan-shaped holes are assessed in terms of film cooling effectiveness and total pressure loss. Lined up with the freestream streamwise direction, the film holes are arranged on the platform with two different layouts. In one layout, the film cooling holes are divided into two rows and more concentrated on the pressure side of the passage. In the other layout, the film cooling holes are divided into four rows and loosely distributed on the platform. Four film cooling hole configurations are investigated totally. Testing was done in a five-blade cascade with medium high Mach number condition (0.27 and 0.44 at the inlet and the exit, respectively). The detailed film cooling effectiveness distributions on the platform was obtained using pressure sensitive paint (PSP) technique. Results show that the combined cooling scheme (slot purge flow cooling combined with discrete hole film cooling) is able to provide full film coverage on the platform. The shaped holes present higher film cooling effectiveness and wider film coverage than the cylindrical holes, particularly at higher blowing ratios. The hole layout affects the local film cooling effectiveness. The shaped holes also show the advantage over the cylindrical holes with lower total pressure loss.

Numerical investigation of blowing ratio, density ratio and axial position of film holes on the vane endwall film cooling effectiveness with upstream step

2021 ◽  
pp. 095441002110165
Author(s):  
Qingzong Xu ◽  
Qiang Du ◽  
Pei Wang ◽  
Xiangtao Xiao ◽  
Jun Liu
Keyword(s):  
Secondary Flow ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Density Ratio ◽  
Total Pressure Loss ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Endwall Film Cooling

The aerothermal performance of interrupted slot and film holes was numerically investigated. Previous study indicates that the interrupted slot performs better compared to the conventional slot. In the meanwhile, the step formed along with the interrupted slot affects the film cooling characteristics. In this article, a row of film holes is arranged downstream of the step, and the mass flow rate for the interrupted slot is constant at 1%. Blowing ratio (BR) from 0.5 to 1.5 and density ratio from 1 to 2 were studied for the film holes. Endwall film cooling effectiveness distribution indicates that film cooling is easily affected by the secondary flow inside passage and the upstream step. Coolant traces are split into two parts due to the effects of step vortex and transverse flow. For different density ratios, increasing BR shows a different trend of film cooling effectiveness due to the variation of coolant momentum. The coolant jet is easily affected by the secondary flow when its momentum is low, but tends to liftoff when its momentum is too high. As a result, it is better to position the film holes far away from the upstream step. The total pressure loss coefficient distribution at the passage exit indicates that the coolant injection increases the total pressure loss. But density ratio has smaller effect on the loss variation. Besides, two axial positions of cooling holes were studied to improve the endwall cooling performance. Without the effect of step vortex, the film effectiveness of cooling holes is improved.

scholarly journals Prediction of Vane Surface Film Cooling Effectiveness Using Compressible Navier-Stokes Procedure and k-ε Turbulence Model With Wall Function

1995 ◽  
Author(s):  
Yoshitaka Fukuyama ◽  
Fumio Otomo ◽  
Minoru Sato ◽  
Yuichi Kobayashi ◽  
Hiroyuki Matsuzaki
Keyword(s):  
Turbulence Model ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Total Pressure Loss ◽  
Navier Stokes ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Turbine Vane ◽  
Near Wall

A numerical prediction has been performed on the film cooling effectiveness and the total pressure loss of the actual turbine vane geometry. The Navier-Stokes code used in this study is an implicit, cell-centered, finite volume code with k-ε turbulence models. The convection term was stabilized by the variable order up-winding scheme. The film cooling injection has been simulated by adding the prescribed flux terms at the vane surface. The k and ε near wall distribution functions were developed based on the experimental and the DNS results in the literature. The wall functions for k and ε can be used with the selected low Reynolds number version of k-ε turbulence model irrespective of the distance between the wall and the first grid point. This combination would result in lower computational costs, since, near wall grid number can be reduced significantly. Based on the study, the Navier-Stokes predictions were performed on the actual turbine vane geometry. Also, the comparisons were made with the experimental total pressure loss distribution behind the vane row and the mid-span film-cooling effectiveness distribution for single and double row injection cases.

Blowing Ratio Effect on Film Cooling Performance for a Row Holes Installed in Different Trench Configurations

2020 ◽  
Vol 28 ◽  
pp. 65-75
Author(s):  
Khadidja Boualem ◽  
Abbes Azzi
Keyword(s):  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Total Pressure Loss ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Loss Coefficient ◽  
Blowing Ratio ◽  
Ansys Cfx ◽  
Cooling Hole

In the present study, a numerical analysis was performed to evaluate the performance of cooling hole embedded in different trenched designs (triangular trench, semi-cylindrical trench and corrugated trench) in improving the film cooling efficiency over a flat plate. These concepts are compared to the rectangular trenched and the traditional cylindrical hole. The commercial software ANSYS CFX 18 was used to conduct a series of required numerical calculations. The centerline and laterally averaged film cooling effectiveness and total pressure loss coefficient for the five cases are analyzed at three blowing ratios, M=0.5, M=1.0 and M=1.5. Results show a uniform coverage is obtained by hole installed in trench. The main result obtained in this paper that the cooling hole with corrugated trench enhance the film cooling effectiveness with less total pressure loss. The main result of this study reveals that the jet installed in the trenches yield a better film cooling effectiveness especially at higher blowing ratios (M≥1).

Heat Transfer Coefficients and Film Cooling Effectiveness on the Squealer Tip of a Gas Turbine Blade

2003 ◽  
Vol 125 (4) ◽  
pp. 648-657 ◽  
Author(s):  
Jae Su Kwak ◽  
Je-Chin Han
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Turbine Blade ◽  
Film Cooling ◽  
Rotor Blade ◽  
Transfer Coefficients ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Heat Transfer Coefficients ◽  
Blowing Ratio

Experimental investigations were performed to measure the detailed heat transfer coefficients and film cooling effectiveness on the squealer tip of a gas turbine blade in a five-bladed linear cascade. The blade was a two-dimensional model of a first stage gas turbine rotor blade with a profile of the GE-E3 aircraft gas turbine engine rotor blade. The test blade had a squealer (recessed) tip with a 4.22% recess. The blade model was equipped with a single row of film cooling holes on the pressure side near the tip region and the tip surface along the camber line. Hue detection based transient liquid crystals technique was used to measure heat transfer coefficients and film cooling effectiveness. All measurements were done for the three tip gap clearances of 1.0%, 1.5%, and 2.5% of blade span at the two blowing ratios of 1.0 and 2.0. The Reynolds number based on cascade exit velocity and axial chord length was 1.1×106 and the total turning angle of the blade was 97.9 deg. The overall pressure ratio was 1.2 and the inlet and exit Mach numbers were 0.25 and 0.59, respectively. The turbulence intensity level at the cascade inlet was 9.7%. Results showed that the overall heat transfer coefficients increased with increasing tip gap clearance, but decreased with increasing blowing ratio. However, the overall film cooling effectiveness increased with increasing blowing ratio. Results also showed that the overall film cooling effectiveness increased but heat transfer coefficients decreased for the squealer tip when compared to the plane tip at the same tip gap clearance and blowing ratio conditions.

Conjugate Heat Transfer Characteristics of Double Wall Cooling With Gradient Diameter of Film and Impingement Holes

2021 ◽  
Author(s):  
Juan He ◽  
Qinghua Deng ◽  
Zhenping Feng
Keyword(s):  
Heat Transfer ◽  
Pressure Loss ◽  
Total Pressure ◽  
Flow Direction ◽  
Turbine Blades ◽  
Total Pressure Loss ◽  
Advanced Technique ◽  
Cooling Effectiveness ◽  
Uniform Pattern ◽  
Double Wall

Abstract Double wall cooling, consisting of internal impingement cooling and external film cooling, is believed to be the most advanced technique in modern turbine blades cooling. In this paper, to improve the uniformity of temperature distribution, a flat plate double wall cooling model with gradient diameter of film and impingement holes was proposed, and the heat transfer and flow characteristics were investigated by solving steady three-dimensional Reynolds-Averaged Navier-Stokes (RANS) equations with SST k-ω turbulence model. The influence of gradient diameter on overall cooling effectiveness and total pressure loss was studied by comparing with the uniform pattern at the blowing ratios ranging from 0.5 to 2. For gradient diameter of film hole patterns, results show that −10% film pattern always has the lowest film flow non-uniformity coefficient. The laterally averaged overall cooling effectiveness of uniform pattern lies between that of +10% and −10% film patterns, but the intersection of three patterns moves upstream from the middle of flow direction with the increase of blowing ratio. Therefore, the −10% film pattern exerts the highest area averaged cooling effectiveness, which is improved by up to 1.6% and 1% at BR = 0.5 and 1 respectively compared with a uniform pattern. However, at higher blowing ratios, the +10% film pattern maintains higher cooling effectiveness and lower total pressure loss. For gradient diameter of impingement hole patterns, the intersection of laterally averaged overall cooling effectiveness in three patterns is located near the middle of flow direction under all blowing ratios. The uniform pattern has the highest area averaged cooling effectiveness and the smallest non-uniform coefficient, but the −10% jet pattern has advantages of reducing pressure loss, especially in the laminated loss.

Sign in / Sign up

Export Citation Format

Share Document