Sweeping Jet Film Cooling on a Turbine Vane

2019 ◽  
Vol 141 (3) ◽  
Author(s):  
Mohammad A. Hossain ◽  
Lucas Agricola ◽  
Ali Ameri ◽  
James W. Gregory ◽  
Jeffrey P. Bons
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Convective Heat Transfer Coefficient ◽  
Total Pressure Loss ◽  
Pressure Probe ◽  
Blowing Ratio ◽  
Turbine Vane ◽  
Shaped Hole

The cooling performance of sweeping jet film cooling was studied on a turbine vane suction surface in a low-speed linear cascade wind tunnel. The sweeping jet holes consist of fluidic oscillators with an aspect ratio (AR) of unity and a hole spacing of Pd/D = 6. Infrared (IR) thermography was used to estimate the adiabatic film effectiveness at several blowing ratios and two different freestream turbulence levels (Tu = 0.3% and 6.1%). Convective heat transfer coefficient was measured by a transient IR technique, and the net heat flux benefit was calculated. The total pressure loss due to sweeping jet film cooling was characterized by traversing a total pressure probe at the exit plane of the cascade. Tests were performed with a baseline shaped hole (SH) (777-shaped hole) for comparison. The sweeping jet hole showed higher adiabatic film effectiveness than the 777-shaped hole in the near hole region. Although the unsteady sweeping action of the jet augments heat transfer, the net positive cooling benefit is higher for sweeping jet holes compared to 777 hole at particular flow conditions. The total pressure loss measurement showed a 12% increase in total pressure loss at a blowing ratio of M = 1.5 for sweeping jet hole, while 777-shaped hole showed a 8% total pressure loss increase at the corresponding blowing ratio.

Sweeping Jet Film Cooling on a Turbine Vane

2018 ◽  
Author(s):  
Mohammad A. Hossain ◽  
Lucas Agricola ◽  
Ali Ameri ◽  
James W. Gregory ◽  
Jeffrey P. Bons
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Convective Heat Transfer Coefficient ◽  
Total Pressure Loss ◽  
Pressure Probe ◽  
Blowing Ratio ◽  
Turbine Vane ◽  
Shaped Hole

The cooling performance of sweeping jet film cooling was studied on a turbine vane suction surface in a low-speed linear cascade wind tunnel. The sweeping jet holes consist of fluidic oscillators with an aspect ratio (AR) of unity and a hole spacing of Pd/D = 6. Infrared (IR) thermography was used to estimate the adiabatic film effectiveness at several blowing ratios and two different freestream turbulence levels (Tu = 0.3% and 6.1%). Convective heat transfer coefficient was measured by a transient IR technique, and the net heat flux benefit was calculated. The total pressure loss due to sweeping jet film cooling was characterized by traversing a total pressure probe at the exit plane of the cascade. Tests were performed with a baseline shaped hole (777-shaped hole) for comparison. The sweeping jet hole showed higher adiabatic film effectiveness than the 777-shaped hole in the near hole region. Although the unsteady sweeping action of the jet augments heat transfer, the net positive cooling benefit is higher for sweeping jet holes compared to 777 hole at particular flow conditions. The total pressure loss measurement showed a 12% increase in total pressure loss at a blowing ratio of M = 1.5 for sweeping jet hole while 777-shaped hole showed a 8% total pressure loss increase at the corresponding blowing ratio.

Aerodynamic Performance of Suction-Side Gill Region Film Cooling

2010 ◽  
Vol 132 (3) ◽  
Author(s):  
Justin Chappell ◽  
Phil Ligrani ◽  
Sri Sreekanth ◽  
Terry Lucas ◽  
Edward Vlasic
Keyword(s):  
Mach Number ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Suction Side ◽  
Total Pressure Loss ◽  
Pressure Loss Coefficient ◽  
Blowing Ratio ◽  
Exit Mach Number ◽  
Total Pressure Loss Coefficient

The performance of suction-side gill region film cooling is investigated using the University of Utah transonic wind tunnel and a simulated turbine vane in a two-dimensional cascade. The effects of film cooling hole orientation, shape, and number of rows, and their resulting effects on the aerodynamic losses, are considered for four different hole configurations: round axial (RA), shaped axial (SA), round radial (RR), and round compound (RC). The mainstream Reynolds number based on axial chord is 500,000, exit Mach number is 0.35, and the tests are conducted using the first row of holes, or both rows of holes at blowing ratios of 0.6 and 1.2. Carbon dioxide is used as the injectant to achieve density ratios of 1.77–1.99 similar to values present in operating gas turbine engines. Presented are the local distributions of total pressure loss coefficient, local normalized exit Mach number, and local normalized exit kinetic energy. Integrated aerodynamic losses (IAL) increase anywhere from 4% to 45% compared with a smooth blade with no film injection. The performance of each hole type depends on the airfoil configuration, film cooling configuration, mainstream flow Mach number, number of rows of holes, density ratio, and blowing ratio, but the general trend is an increase in IAL as either the blowing ratio or the number of rows of holes increase. In general, the largest total pressure loss coefficient Cp magnitudes and the largest IAL are generally present at any particular wake location for the RR or SA configurations, regardless of the film cooling blowing ratio and number of holes. The SA holes also generally produce the highest local peak Cp magnitudes. IAL magnitudes are generally lowest with the RA hole configuration. A one-dimensional mixing loss correlation for normalized IAL values is also presented, which matches most of the both rows data for RA, SA, RR, and RC hole configurations. The equation also provides good representation of the RA, RC, and RR first row data sets.

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.

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.

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.

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.

Large-Eddy Simulation of Film Cooling Performance Enhancement Using Vortex Generator and Semi-Sphere

2020 ◽  
Author(s):  
Wen Wang ◽  
Jiahuan Cui ◽  
Shaoxing Qu
Keyword(s):  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Vortex Generator ◽  
Vortex Pair ◽  
Total Pressure Loss ◽  
Cooling Efficiency ◽  
Blowing Ratio ◽  
Large Eddy ◽  
Film Cooling Efficiency

Abstract Film cooling is an essential cooling method to prevent high-pressure turbine blade from melting down due to the high inlet temperature. In order to improve the film cooling efficiency, several flow control methods have been proposed. In this paper, large-eddy simulations are performed to study the effectiveness of a vortex generator (VG) and a semi-sphere installed downstream of the cooling jet. Before the detailed analyses, the numerical framework is validated against the available experimental data. Both the laminar and turbulent approaching boundary layers are considered. The turbulent boundary layer is generated by a numerical plasma actuator. After validation, the influence of VG and semi-sphere on the film cooling efficiency at various blowing ratios are analyzed. It is found that a counter-rotating vortex pair (CVP) is formed downstream and its strength increases with the blowing ratio in the configuration without VG/semi-sphere. When the VG is installed, it produces another vortex pair that rotates in the reverse direction of the CVP, which reduces the CVP strength and increases the lateral diffusion of the coolant. As a result, the film cooling efficiency is greatly improved, especially at a higher blowing ratio. For the case with a semi-sphere, the film cooling efficiency is also improved, especially at low–medium blowing ratios. However, it is not as effective as the VG in terms of enhancing cooling efficiency. In addition, the total pressure loss is calculated to examine the aerodynamic penalty associated with the VG and semi-sphere. It is found that the total pressure loss increased by only 1% due to the VG or semi-sphere, within the range of blowing ratio investigated in the current study. Considering the overall performance and the feasibility of being applied in practice, a semi-sphere installed downstream of the cooling hole is a promising method to improve the cooling efficiency.

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).

Film Cooling and Aerodynamic Performance on Multi-Cavity Squealer Tip of a Turbine Blade

2021 ◽  
Author(s):  
Feng Li ◽  
Zhao Liu ◽  
Zhenping Feng
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Pressure Loss ◽  
Total Pressure ◽  
Aerodynamic Performance ◽  
Total Pressure Loss ◽  
Cooling Performance ◽  
Pressure Loss Coefficient ◽  
Blade Tip ◽  
Total Pressure Loss Coefficient

Abstract The blade tip region of the shroud-less high-pressure gas turbine is exposed to an extremely operating condition with combined high temperature and high heat transfer coefficient. It is critical to design new tip structures and apply effective cooling method to protect the blade tip. Multi-cavity squealer tip has the potential to reduce the huge thermal loads and improve the aerodynamic performance of the blade tip region. In this paper, numerical simulations were performed to predict the aerothermal performance of the multi-cavity squealer tip in a heavy-duty gas turbine cascade. Different turbulence models were validated by comparing to the experimental data. It was found that results predicted by the shear-stress transport with the γ-Reθ transition model have the best precision. Then, the film cooling performance, the flow field in the tip gap and the leakage losses were presented with several different multi-cavity squealer tip structures, under various coolant to mainstream mass flow ratios (MFR) from 0.05% to 0.15%. The results show that the ribs in the multi-cavity squealer tip could change the flow structure in the tip gap for that they would block the coolant and the leakage flow. In this study, the case with one-cavity (1C) achieves the best film cooling performance under a lower MFR. However, the cases with multi-cavity (2C, 3C, 4C) show higher film cooling effectiveness under a higher MFR of 0.15%, which are 32.6%%, 34.2%% and 41.0% higher than that of the 1C case. For the aerodynamic performance, the case with single-cavity has the largest total pressure loss coefficient in all MFR studied, whereas the case with two-cavity obtains the smallest total pressure loss coefficient, which is 7.6% lower than that of the 1C case.

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