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.

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.

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.

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.

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.

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.

Large Eddy Simulations of Film-Cooling Flows With a Micro-Ramp Vortex Generator

2011 ◽  
Author(s):  
Aaron F. Shinn ◽  
S. Pratap Vanka
Keyword(s):  
Film Cooling ◽  
Cross Flow ◽  
Vortex Generator ◽  
Vortex Pair ◽  
Large Eddy Simulations ◽  
Wind Tunnel Experiments ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Cooling Flows ◽  
Large Eddy

Large Eddy Simulations were performed to study the effect of a micro-ramp on an inclined turbulent jet interacting with a cross-flow in a film-cooling configuration. The micro-ramp vortex generator is placed downstream of the film-cooling jet. Changes in vortex structure and film-cooling effectiveness are evaluated and the genesis of the counter-rotating vortex pair in the jet is discussed. Results are reported with the jet modeled using a plenum/pipe configuration. This configuration was designed based on previous wind tunnel experiments at NASA Glenn Research Center, and the present results are meant to supplement those experiments. It is found that the micro-ramp improves film-cooling effectiveness by generating near-wall counter-rotating vortices which help entrain coolant from the jet and transport it to the surface. The pair of vortices generated by the micro-ramp are of opposite sense to the vortex pair embedded in the jet.

Secondary Flow Control Using Vortex Generator Jets

2004 ◽  
Vol 126 (4) ◽  
pp. 650-657 ◽  
Author(s):  
R. K. Sullerey ◽  
A. M. Pradeep
Keyword(s):  
Pressure Loss ◽  
Total Pressure ◽  
Static Pressure ◽  
Vortex Generator ◽  
Control Flow ◽  
Secondary Flows ◽  
Total Pressure Loss ◽  
Vortex Generators ◽  
Vortex Generator Jets ◽  
Inflow Conditions

In this paper, results are presented of an experimental investigation into the effectiveness of vortex generator jets in controlling secondary flows in two-dimensional S-duct diffusers. The experiments were performed in uniform and distorted inflow conditions and the performance evaluation of the diffuser was carried out in terms of static pressure recovery and quality of the exit flow. In the case with inflow distortion, tapered fin vortex generators were employed in addition to vortex generator jets to control flow separation that was detected on the wall with inflow distortion. Detailed measurements including total pressure, velocity distribution, surface static pressure, skin friction, and boundary layer measurements were taken at a Reynolds number of 7.8×105. These results are presented in terms of static pressure rise, distortion coefficient, and total pressure loss coefficient at the duct exit. For uniform inflow, the use of vortex generator jets resulted in more than a 30 percent decrease in total pressure loss and flow distortion coefficients. In combination with passive device (tapered fin vortex generators), the vortex generator jets reduce total pressure losses by about 25 percent for distorted inflow conditions. A potential application of this method may include control of secondary flows in turbo machinery.

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.

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