Experimental Study of Full Coverage Film Cooling Effectiveness for a Turbine Blade With Compound Shaped Holes

2020 ◽  
Author(s):  
Shuai-qi Zhang ◽  
Cun-liang Liu ◽  
Qi-ling Guo ◽  
Da-peng Liang ◽  
Fan Zhang
Keyword(s):  
Turbine Blade ◽  
Turbulence Intensity ◽  
Mass Flux ◽  
Film Cooling ◽  
Flux Ratio ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Surface ◽  
Cylindrical Holes

Abstract The film coverage of a turbine blade surface is determined by all the film cooling structures. The direct study of full coverage film cooling is relatively rare, especially for related research on turbine blades. In this paper, the pressure-sensitive paint (PSP) measurement technique is used to carry out experiments under different turbulence intensities and mass flux ratios, and the distribution of the film cooling effectiveness on the entire surface is studied in detail. In this study, a basic turbine blade and an improved turbine blade are investigated. The film cooling hole position distribution on the improved blade is the same as that on the basic blade, but the film cooling hole shape on the suction surface and the pressure surface is changed from cylindrical holes to laid-back fan-shaped holes. Both blades have 5 rows of cylindrical holes at the leading edge and 4 rows of film cooling holes on the suction surface and the pressure surface. The leading edge, suction surface, and pressure surface have their own coolant inlet cavities. This kind of design is not only close to the actual working conditions in a flow distribution but also conveniently eliminates the mutual interference caused by the uneven flow distribution between the pressure surface and the suction surface to facilitate the independent analysis of the pressure surface and the suction surface. In this paper, the film cooling effectiveness of two kinds of turbine blades under different turbulence intensities and mass flux ratios is studied. The results show that the average cooling effectiveness of the improved blade is much better than that of the basic blade. The laid-back fan-shaped hole rows improve the cooling effectiveness of the suction surface by 60% to 100% and 50% to 120% on the pressure surface. The increase in turbulence intensity will reduce the cooling effectiveness of the blade surface; however, the effect of the turbulence intensity becomes weaker with an increase in the mass flux ratio. Compared with the multiple rows of cylindrical holes, the cooling effectiveness of the laid-back fan-shaped holes is more affected by the turbulence intensity under the small mass flux ratio.

Experimental Study on Film Cooling Effectiveness of Blade with Chevron Shaped Holes under Wake Influence

2021 ◽  
pp. 1-20
Author(s):  
Jichen Li ◽  
Hui Ren Zhu ◽  
Cun Liang Liu ◽  
Lin Ye ◽  
Zhou Daoen
Keyword(s):  
Mass Flux ◽  
Film Cooling ◽  
Gas Turbines ◽  
Flux Ratio ◽  
Leading Edge ◽  
High Mass ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Surface

Abstract Gas turbines have been widely used. With the continuous improvement of the performance of gas turbines, the turbine inlet temperature has greatly exceeded the heat resistance limit of the turbine blade material, so advanced cooling technology is required. The film cooling effectiveness distribution over the blade under the effect of wake was obtained by Pressure Sensitive Paint (PSP) technique. The test blade has 5 rows of chevron film holes on the pressure side, 3 rows of cylindrical film holes on the leading edge and 3 rows of chevron film holes on the suction side. The mainstream Reynolds number is 130,000 based on the blade chord length, and the mainstream turbulence intensity is 2.7%. The upstream wake was simulated by the spoken-wheel type wake generator. The film cooling effectiveness was measured at three wake Strouhal numbers (0, 0.12 and 0.36) and three mass flux ratios (MFR1, MFR2 and MFR3). The results show that the increase of mass flux ratio leads a decrease of the film cooling effectiveness on the suction surface. In the wake condition, the effect of mass flux ratio is weakened. Wake leads a marked decrease of the film cooling effectiveness over most blade surface except for the surface near leading edge on the pressure surface. In the high mass flux ratio condition, the effect of wake on the film cooling effectiveness is weakened on the suction surface and strengthened on the pressure surface.

Experimental Study on Film Cooling Effectiveness of Blade With Chevron Shaped Holes Under Wake Influence

2020 ◽  
Author(s):  
Ji-Chen Li ◽  
Hui-Ren Zhu ◽  
Da-Wei Chen ◽  
Dao-En Zhou
Keyword(s):  
Mass Flux ◽  
Film Cooling ◽  
Gas Turbines ◽  
Flux Ratio ◽  
Leading Edge ◽  
High Mass ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Surface

Abstract Gas turbines have been widely used. With the continuous improvement of the performance of gas turbines, the turbine inlet temperature has greatly exceeded the heat resistance limit of the turbine blade material, so advanced cooling technology is required. The film cooling effectiveness distribution over the blade under the effect of wake was obtained by Pressure Sensitive Paint (PSP) technique. The test blade has 5 rows of chevron film holes on the pressure side, 3 rows of cylindrical film holes on the leading edge and 3 rows of chevron film holes on the suction side. The mainstream Reynolds number is 130,000 based on the blade chord length, and the mainstream turbulence intensity is 2.7%. The upstream wake was simulated by the spoken-wheel type wake generator. The film cooling effectiveness was measured at three wake Strouhal numbers (0, 0.12 and 0.36) and three mass flux ratios (MFR1, MFR2 and MFR3). The results show that the increase of mass flux ratio leads a decrease of the film cooling effectiveness on the suction surface. In the wake condition, the effect of mass flux ratio is weakened. Wake leads a marked decrease of the film cooling effectiveness over most blade surface except for the surface near leading edge on the pressure surface. In the high mass flux ratio condition, the effect of wake on the film cooling effectiveness is weakened on the suction surface and strengthened on the pressure surface.

Effects of Axial Row-Spacing for Double-Jet Film-Cooling on the Cooling Effectiveness

2011 ◽  
Author(s):  
Zhan Wang ◽  
Jian-Jun Liu ◽  
Bai-tao An ◽  
Chao Zhang
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Inlet Guide Vane ◽  
Row Spacing ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Surface ◽  
Cylindrical Holes ◽  
Compound Angle

The effects of axial row-spacing for double jet film-cooling (DJFC) with compound angle on the cooling characteristics under different blowing ratios were investigated numerically. First, the flow fields and cooling effectiveness of DJFC on flat plate with different axial row-spacing were calculated. Film-cooling with fan-shaped or cylindrical holes was also calculated for the comparison. The results indicate that a larger axial row-spacing is helpful to form the anti-kidney vortex and to improve the cooling effectiveness. The DJFC was then applied to the suction and pressure surface of a real turbine inlet guide vane. Comparisons of film-cooling effectiveness with the cylindrical and fan-shaped holes were also conducted. The results for the guide vane show that on the suction surface the DJFC with a larger axial row-spacing leads to better film coverage and better film-cooling effectiveness than the cylindrical or fan-shaped holes. On the pressure surface, however, the film-cooling with fan-shaped holes is superior to the others.

Experimental Investigation on Cooling Performance of Impingement-Effusion Full Coverage Film on Suction Surface of Vane

2020 ◽  
Author(s):  
Fan Zhang ◽  
Cunliang Liu ◽  
Shuaiqi Zhang ◽  
Lin Ye ◽  
Bingran Li
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Turbulence Intensity ◽  
Mass Flux ◽  
Film Cooling ◽  
Flux Ratio ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
The Heat Transfer Coefficient

Abstract To study the film cooling performance of impingement-effusion structures, it is important to study their adiabatic film cooling effectiveness. To improve the adiabatic film cooling effectiveness on a vane, some rows of cylindrical effusion holes are changed into fan-shaped holes. This experiment measured the adiabatic film cooling effectiveness of the double-walled system on the suction surface via the pressure-sensitive paint (PSP) technique. The film cooling effectiveness obtained by the PSP technique is coupled with the transient liquid crystal (TLC) technique to determine the heat transfer coefficient. This combination of techniques reduces the time required for the experiment and improves the efficiency of the experiment. The heat transfer coefficient ratio is used to evaluate the level of heating transfer. The net heat flux reduction (NHFR) is used to quantify the net benefit of film cooling. Two experimental vanes’ (A and B) film holes are both arranged in 6 rows of holes. There are 15 holes in each row. Only the positions of the fan-shaped holes are different. The experimental conditions include the mainstream Reynolds number (Re = 151,000) based on the chord length and inlet velocity, the turbulence intensities (Tu = 0.77%, 16.9%), and the mass flux ratios (ṁc/ṁg = 0.4%, 0.8%, 1.6%). The findings show that when the mass flux ratio increases to a point, the film cooling effectiveness does not improve. Increasing the turbulence intensity leads to a decrease in the film cooling effectiveness except for the region after Row 6 on Vane B. Using the coupling of PSP and TLC to determine the heat transfer coefficient can yield credible results. The turbulence intensity and the arrangement of the film holes have obvious effects on the distribution of the heat transfer coefficient ratio. The effects of turbulence intensity, mass flux ratio and hole arrangement on NHFR were studied.

Mechanism of Film Cooling with One Inlet and Double Outlet Hole Injection at Various Turbulence Intensities

2018 ◽  
Vol 35 (1) ◽  
pp. 1-9 ◽  
Author(s):  
Guangchao Li ◽  
Yukai Chen ◽  
Zhihai Kou ◽  
Wei Zhang ◽  
Guochen Zhang
Keyword(s):  
Turbulence Intensity ◽  
Film Cooling ◽  
Hole Injection ◽  
Spanwise Direction ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Outlet Hole ◽  
Cylindrical Holes ◽  
High Turbulence

AbstractThe trunk-branch hole was designed as a novel film cooling concept, which aims for improving film cooling performance by producing anti-vortex. The trunk-branch hole is easily manufactured in comparison with the expanded hole since it consists of two cylindrical holes. The effect of turbulence on the film cooling effectiveness with a trunk-branch hole injection was investigated at the blowing ratios of 0.5, 1.0, 1.5 and 2.0 by numerical simulation. The turbulence intensities from 0.4 % to 20 % were considered. The realizable$k - \varepsilon $turbulence model and the enhanced wall function were used. The more effective anti-vortex occurs at the low blowing ratio of 0.5 %. The high turbulence intensity causes the effectiveness evenly distributed in the spanwise direction. The increase of turbulence intensity leads to a slight decrease of the spanwise averaged effectiveness at the low blowing ratio of 0.5, but a significant increase at the high blowing ratios of 1.5 and 2.0. The optimal blowing ratio of the averaged surface effectiveness is improved from 1.0 to 1.5 when the turbulence intensity increases from 0.4 % to 20 %.

Numerical Analysis on the Leading Edge Film Cooling of Bifurcation Holes for Gas Turbine Blade

2021 ◽  
Author(s):  
Zhonghao Tang ◽  
Gongnan Xie ◽  
Honglin Li ◽  
Wenjing Gao ◽  
Chunlong Tan ◽  
...  
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Turbulence Models ◽  
Leading Edge ◽  
Spanwise Direction ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Cylindrical Holes ◽  
Lift Off

Abstract Film cooling performance of the cylindrical film holes and the bifurcated film holes on the leading edge model of the turbine blade are investigated in this paper. The suitability of different turbulence models to predict local and average film cooling effectiveness is validated by comparing with available experimental results. Three rows of holes are arranged in a semi-cylindrical model to simulate the leading edge of the turbine blade. Four different film cooling structures (including a cylindrical film holes and other three different bifurcated film holes) and four different blowing ratios are studied in detail. The results show that the film jets lift off gradually in the leading edge area as the blowing ratio increases. And the trajectory of the film jets gradually deviate from the mainstream direction to the spanwise direction. The cylindrical film holes and vertical bifurcated film holes have better film cooling effectiveness at low blowing ratio while the other two transverse bifurcated film holes have better film cooling effectiveness at high blowing ratio. And the film cooling effectiveness of the transverse bifurcated film holes increase with the increasing the blowing ratio. Additionally, the advantage of transverse bifurcated holes in film cooling effectiveness is more obvious in the downstream region relative to the cylindrical holes. The Area-Average film cooling effectiveness of transverse bifurcated film holes is 38% higher than that of cylindrical holes when blowing ratio is 2.

Film-Cooling on a Gas Turbine Blade Pressure Side or Suction Side With Compound Angle Shaped Holes

2008 ◽  
Vol 131 (1) ◽  
Author(s):  
Zhihong Gao ◽  
Diganta P. Narzary ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Pressure Ratio ◽  
Suction Side ◽  
Pressure Side ◽  
Blade Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Surface ◽  
Compound Angle

The film-cooling effectiveness on the surface of a high pressure turbine blade is measured using the pressure sensitive paint technique. Compound angle laidback fan-shaped holes are used to cool the blade surface with four rows on the pressure side and two rows on the suction side. The coolant injects to one side of the blade, either pressure side or suction side. The presence of wake due to the upstream vanes is simulated by placing a periodic set of rods upstream of the test blade. The wake rods can be clocked by changing their stationary positions to simulate progressing wakes. The effect of wakes is recorded at four phase locations along the pitchwise direction. The freestream Reynolds number, based on the axial chord length and the exit velocity, is 750,000. The inlet and exit Mach numbers are 0.27 and 0.44, respectively, resulting in a pressure ratio of 1.14. Five average blowing ratios ranging from 0.4 to 1.5 are tested. Results reveal that the tip-leakage vortices and endwall vortices sweep the coolant on the suction side to the midspan region. The compound angle laidback fan-shaped holes produce a good film coverage on the suction side except for the regions affected by the secondary vortices. Due to the concave surface, the coolant trace is short and the effectiveness level is low on the pressure surface. However, the pressure side acquires a relatively uniform film coverage with the multiple rows of cooling holes. The film-cooling effectiveness increases with the increasing average blowing ratio for either side of coolant ejection. The presence of stationary upstream wake results in lower film-cooling effectiveness on the blade surface. The compound angle shaped holes outperform the compound angle cylindrical holes by the elevated film-cooling effectiveness, particularly at higher blowing ratios.

scholarly journals Film Cooling on a Gas Turbine Rotor Blade

1991 ◽  
Author(s):  
Kenichiro Takeishi ◽  
Sunao Aoki ◽  
Tomohiko Sato ◽  
Keizo Tsukagoshi
Keyword(s):  
Mass Transfer ◽  
Film Cooling ◽  
Rotor Blade ◽  
Turbine Rotor ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Turbine Rotor Blade ◽  
Pressure Surface ◽  
Rotating Blade

The film cooling effectiveness on a low-speed stationary cascade and the rotating blade has been measured by using a heat-mass transfer analogy. The film cooling effectiveness on the suction surface of the rotating blade fits well with that on the stationary blade, but a low level of effectiveness appears on the pressure surface of the rotating blade. In this paper, typical film cooling data will be presented and film cooling on a rotating blade is discussed.

Turbine Blade Platform Film Cooling With Fan-Shaped Holes Under Simulated Swirl Purge Flow and Slashface Leakage Conditions

2017 ◽  
Author(s):  
Andrew F. Chen ◽  
Chao-Cheng Shiau ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Leakage Flow ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Flow Swirl ◽  
Wide Range ◽  
Parametric Effects ◽  
Purge Flow ◽  
Cylindrical Holes

The combined effects of upstream purge flow, slashface leakage flow, and discrete hole film cooling on turbine blade platform film cooling effectiveness were studied using the pressure sensitive paint (PSP) technique. Detailed adiabatic film cooling effectiveness distributions on the platform were obtained and analyzed. As a continued study, discrete cylindrical holes [1] were replaced by laidback fan-shaped (10-10-5) holes which generally provide better film coverages on the endwall. Experiments were done in a five-blade linear cascade with an average turbulence intensity of 10.5%. The inlet and exit Mach numbers were 0.26 and 0.43, respectively. The inlet and exit mainstream Reynolds numbers based on the axial chord length of the blade were 475,000 and 720,000, respectively. A wide range of parameters were evaluated in this study. The coolant-to-mainstream mass flow ratio (MFR) was varied from 0.5%, 0.75%, to 1% for the upstream purge flow. For the platform film cooling holes and slashface gap, average blowing ratios (M) of 0.5, 1.0, and 1.5 were examined. Coolant-to-mainstream density ratios (DR) that range from 1 (close to low-temperature experiments) to 1.5 (intermediate DR) and 2 (close to engine conditions) were also examined. Purge flow swirl effect was studied particularly at a typical swirl ratio of 0.6. The results provide the gas turbine engine community a better insight into various parametric effects on turbine blade platform film cooling with fan-shaped holes when the upstream swirl purge flow and slashface leakage flow were presented. Area-averaged film cooling effectiveness results were compared between cylindrical and fan-shaped holes under various parametric conditions. The results indicate that the fan-shaped holes provide superior film coverage than cylindrical holes for platform film cooling especially at higher blowing ratios and momentum flux ratios.

Influence of Coolant Density on Turbine Blade Platform Film-Cooling

2012 ◽  
Vol 4 (2) ◽  
Author(s):  
Diganta P. Narzary ◽  
Kuo-Chun Liu ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Flow Rate ◽  
Turbulence Intensity ◽  
Film Cooling ◽  
Density Ratio ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Mainstream Flow ◽  
Purge Flow

Detailed parametric study of film-cooling effectiveness was carried out on a turbine blade platform of a five-blade linear cascade. The parameters chosen were freestream turbulence intensity, upstream stator-rotor purge flow rate, discrete-hole film-cooling blowing ratio, and coolant-to-mainstream density ratio. The measurement technique adopted was temperature sensitive paint (TSP) technique. Two turbulence intensities of 4.2% and 10.5%; three purge flows between the range of 0.25% and 0.75% of mainstream flow rate; three blowing ratios between 1.0 and 1.8; and three density ratios between 1.1 and 2.2 were investigated. Purge flow was supplied via a typical double-toothed stator-rotor seal, whereas the discrete-hole film-cooling was accomplished via two rows of cylindrical holes arranged along the length of the platform. The inlet and the exit Mach numbers were 0.27 and 0.44, respectively. Reynolds number of the mainstream flow was 7.5 * 105 based on the exit velocity and chord length of the blade. Results indicated that platform film-cooling effectiveness decreased with turbulence intensity, increased with purge flow rate and density ratio, and possessed an optimum blowing ratio value.

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