Turbine Blade Platform Film Cooling With Simulated Swirl Purge Flow and Slashface Leakage Conditions

2016 ◽  
Vol 139 (3) ◽  
Author(s):  
Andrew F Chen ◽  
Chao-Cheng Shiau ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Leading Edge ◽  
Leakage Flow ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Cooling Hole ◽  
Purge Flow ◽  
Endwall Film Cooling ◽  
Film Cooling Holes

The combined effects of inlet purge flow and the slashface leakage flow on the film cooling effectiveness of a turbine blade platform were studied using the pressure-sensitive paint (PSP) technique. Detailed film cooling effectiveness distributions on the endwall were obtained and analyzed. Discrete cylindrical film cooling holes were arranged to achieve an improved coverage on the endwall. Backward injection was attempted by placing backward injection holes near the pressure side leading edge portion. 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. The coolant-to-mainstream mass flow ratios (MFR) were varied from 0.5% and 0.75% to 1% for the purge flow. For the endwall film cooling holes and slashface leakage flow, blowing ratios (M) of 0.5, 1.0, and 1.5 were examined. Coolant-to-mainstream density ratios (DR) that range from 1.0 (close to low temperature experiments) to 1.5 and 2.0 (close to engine conditions) were also examined. The results provide the gas turbine engine designers a better insight into improved film cooling hole configurations as well as various parametric effects on endwall film cooling when the inlet (swirl) purge flow and slashface leakage flow were incorporated.

Turbine Blade Platform Film Cooling With Simulated Swirl Purge Flow and Slashface Leakage Conditions

2016 ◽  
Author(s):  
Andrew F. Chen ◽  
Chao-Cheng Shiau ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Leading Edge ◽  
Suction Side ◽  
Leakage Flow ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Purge Flow ◽  
Endwall Film Cooling ◽  
Film Cooling Holes

The combined effects of inlet purge flow and the slashface leakage flow on the film cooling effectiveness of a turbine blade platform were studied using the pressure sensitive paint (PSP) technique. Detailed film cooling effectiveness distributions on the endwall were obtained and analyzed. The inlet purge flow was generated by a row of equally-spaced cylindrical injection holes inside a single-tooth generic stator-rotor seal. In addition to the traditional 90 degree (radial outward) injection for the inlet purge flow, injection at a 45 degree angle was adopted to create a circumferential/azimuthal velocity component toward the suction side of the blades, which created a swirl ratio (SR) of 0.6. Discrete cylindrical film cooling holes were arranged to achieve an improved coverage on the endwall. Backward injection was attempted by placing backward injection holes near the pressure side leading edge portion. Slashface leakage flow was simulated by equally-spaced cylindrical injection holes inside a slot. 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. The coolant-to-mainstream mass flow ratios (MFR) were varied from 0.5%, 0.75%, to 1% for the inlet purge flow. For the endwall film cooling holes and slashface leakage flow, blowing ratios (M) of 0.5, 1.0, and 1.5 were examined. Coolant-to-mainstream density ratios (DR) that range from 1.0 (close to low temperature experiments) to 1.5 (intermediate DR) and 2.0 (close to engine conditions) were also examined. The results provide the gas turbine engine designers a better insight into improved film cooling hole configurations as well as various parametric effects on endwall film cooling when the inlet (swirl) purge flow and slashface leakage flow were incorporated.

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.

Implementation of Vortex Generator and Ramp to Improve Film Cooling Effectiveness on Blade Endwall

2021 ◽  
Author(s):  
Sadam Hussain ◽  
Xin Yan
Keyword(s):  
Numerical Methods ◽  
Film Cooling ◽  
Vortex Generator ◽  
Leading Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Baseline Case ◽  
Cooling Hole ◽  
Cooling Effects ◽  
Endwall Film Cooling

Abstract With the arrangements of vortex generators (VG) and ramp, film cooling effects on endwall near leading edge were numerically investigated at two blowing ratios (i.e. M = 0.5 and M = 1). To determine suitable numerical methods, mesh independency analysis and turbulence model selection were carried out based on the existing experimental data and LES results. With the numerical methods, flow fields near the leading edge were visualized to illustrate the influence of VG and ramp on coolant coverage on blade endwall. Film cooling effectiveness distributions on endwall and coolant trajectories near leading edge were compared among five different configurations with VG and ramp. The results show that the attachment of coolant on blade endwall is improved with the implement of VG between shaped-hole and leading edge. With the implementation of ramp on endwall between cooling hole and leading edge, the coolant spreads wider on endwall along pitchwise direction than the baseline case. With the implementation of VG and ramp, film cooling effect on endwall near leading edge is significantly improved as compared with the only ramp and only VG cases. Compared with the baseline case, pitchwise-averaged film cooling effectiveness on blade endwall near leading edge is increased by about 9%, and the film cooling effectiveness distributions on endwall along pitchwise direction become much uniform, for the case with both ramp and VG at M = 1.

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.

Optimal Arrangement of the Film Cooling Holes Considering the Manufacturing Tolerance for High Pressure Turbine Nozzle

2016 ◽  
Author(s):  
Sanga Lee ◽  
Dong-Ho Rhee ◽  
Kwanjung Yee
Keyword(s):  
High Pressure ◽  
Film Cooling ◽  
Leading Edge ◽  
High Pressure Turbine ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Manufacturing Tolerance ◽  
Cooling Hole ◽  
Film Cooling Holes

In spite of a myriad of researches on the optimal shape of film cooling holes, only a few attempts have been made to optimize the hole arrangement for film cooling so far. Moreover, although the general scale of film cooling hole is so small that manufacturing tolerance has substantial effects on the cooling performance of turbine, the researches on this issue are even scarcer. If it is possible to obtain optimal hole arrangement which not only improve the film cooling performance but also is robust to the manufacturing tolerance, then overall cooling performance of a turbine would become more reliable and useful from the practical point of view. To this end, the present study proposed a robust design optimization procedure which takes the manufacturing uncertainties into account. The procedure was subsequently applied to the film cooling holes on high pressure turbine nozzle pressure side to obtain the robust array shape under the uncertainty of the manufacturing tolerance. First, the array of the holes was parameterized by 5 design variables using the newly suggested shape functions, and 2 representative factors were considered for the manufacturing tolerance of the film cooling hole. Probabilistic process that consists of Kriging surrogate model and Monte Carlo Simulation with descriptive sampling method was coupled with the design optimization process using Genetic Algorithm. Through this, film cooling hole array which shows the high performance, yet robust to the manufacturing tolerance was obtained, and the effects of the manufacturing tolerance on the cooling performance was carefully investigated. As a result, the region where the film cooling effectiveness is noticeable, as well as the maximum width of the variation of the film cooling effectiveness were reduced through optimization, and it is also confirmed that the tolerance of the holes near the leading edge is more influential to the cooling performance because the film cooling effectiveness is more sensitive to the manufacturing tolerance of the leading edge than that of the trailing edge.

Film Cooling Effectiveness Distribution on a Gas Turbine Blade Platform With Inclined Slot Leakage and Discrete Film Hole Flows

2006 ◽  
Author(s):  
Lesley M. Wright ◽  
Zhihong Gao ◽  
Huitao Yang ◽  
Je-Chin Han
Keyword(s):  
Secondary Flow ◽  
Turbine Blade ◽  
Film Cooling ◽  
Leading Edge ◽  
Injection Rate ◽  
Rotor Blades ◽  
Width Ratio ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Film Cooling Holes

A five blade, linear cascade is used to experimentally investigate turbine blade platform cooling. A 30° inclined slot upstream of the blades is used to model the seal between the stator and rotor, and 12 discrete film holes are located on the downstream half of the platform for additional cooling. The film cooling effectiveness is measured on the platform using pressure sensitive paint (PSP). The mainstream Reynolds number is 3.1*105 based on the inlet velocity and the chord length of the scaled high pressure turbine blade. The upstream slot covers 1.5 passages with the coolant exiting the slot at the leading edge of the rotor blades. The length-to-width ratio (ls/w) of the slot is 5.7, and the slot flowrate varies from 0.50% to 2.0% of the mainstream flow. The discrete film cooling holes also have an inclination of 30°, so the length-to-diameter (lf/d) ratio of each hole is 10. The blowing ratio of the coolant through the holes varies from 0.5 to 2.0, based on the mainstream exit velocity. Using PSP it is clear that the film cooling effectiveness on the blade platform is strongly influenced by the platform secondary flow through the passage. Increasing the slot injection rate weakens the secondary flow and provides more uniform film coverage. Increasing the freestream turbulence level was shown to increase film cooling effectiveness on the endwall, as the increased turbulence also weakens the passage vortex. However, downstream, near the discrete film cooling holes, the increased turbulence decreases the film cooling effectiveness (as reported for flat plate film cooling studies). Finally, combining upstream slot flow with downstream discrete film holes should be done cautiously to ensure coolant is not wasted by overcooling regions on the platform.

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

2017 ◽  
Vol 140 (1) ◽  
Author(s):  
Andrew F Chen ◽  
Chao-Cheng Shiau ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Flow Swirl ◽  
Wide Range ◽  
Purge Flow ◽  
Cylindrical Holes ◽  
Density Ratios ◽  
Film Cooling 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. As a continued study, discrete cylindrical holes 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. 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 was 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 (SR) of 0.6. Area-averaged film cooling effectiveness results were compared between cylindrical and fan-shaped holes. 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.

Effects of Inlet Swirl on Endwall Film Cooling in Neighboring Vane Passages

2017 ◽  
Author(s):  
Yang Zhang ◽  
Yifei Li ◽  
Xiutao Bian ◽  
Xin Yuan
Keyword(s):  
Film Cooling ◽  
Rotating Flow ◽  
Leakage Flow ◽  
Cooling Effectiveness ◽  
Inlet Flow ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Inlet Swirl ◽  
Endwall Film Cooling ◽  
Film Cooling Holes

The distribution of film cooling effectiveness of endwall film-cooling holes is considered to be periodic between neighboring high pressure turbine passages in most cascade experiments. In reality, because of the difference in the number of combustors and vanes, the flow fields of neighboring passages are completely different. The secondary flow, especially the passage vortex, is dominated by the upstream inlet rotating flow whose relative flow direction is the reverse between the neighboring vane passages. Specifying the direction of rotation to simulate inlet swirl introduces new challenges in film-cooling design. The present experiment compares five groups of endwall film-cooling with anticlockwise rotating flows at inlet at different clocking positions, and the film-cooling effect is analyzed to investigate the effects of inlet rotating flow. The inlet flow condition of neighboring passages is simulated by switching the position of a swirler fan. Hence, different rotating inlet flow conditions in different positions are achieved. The GE-E3 airfoil was used in the cascade rig, with a scaled-up factor of 1.95. The inlet Reynolds number is 1.48 × 105 and the Mach number is 0.07. The effects of the blowing ratio and relative positions of the swirler are investigated in the experiment. Adiabatic film-cooling effectiveness is probed by using pressure-sensitive paint (PSP). The coolant is simulated by nitrogen with which a density ratio of around 1.0 can be achieved. Fan-shaped film-cooling holes are introduced into the endwall surface as well as trailing edge discharge holes. The cooling performance of the combustor-turbine gap leakage flow is not considered. Fan-shaped film-cooling holes are introduced into the endwall surface as well as upstream slot. The cooling performance of the combustor-turbine gap leakage flow is considered in this case. A Pair of nozzle guide vane (NGV) passages are investigated simultaneously by which the film cooling effectiveness can be compared for the same case at the endwall surface. The inlet rotating flow is simulated by an upstream swirler, with five relative positions along the pitchwise direction. According to the experimental results, the inlet rotating flow dominates the film cooling effectiveness distribution at the endwall. The averaged film cooling effectiveness changes substantially with the change in swirler position. The rotating flow at the endwall region mainly interacts with the main flow to modify incidence angle. The influence of the inlet rotating flow is more obvious at the upstream portion. Meanwhile the downstream portion is not as sensitive to rotating flow as the upstream portion.

Effect of Flow Parameter Variations on Full Coverage Film-Cooling Effectiveness for a Gas Turbine Blade

2007 ◽  
Author(s):  
Shantanu Mhetras ◽  
Je-Chin Han ◽  
Ron Rudolph
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Gas Turbines ◽  
Leading Edge ◽  
Suction Side ◽  
Intensity Level ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Mach Numbers ◽  
Film Cooling Holes

The effect of film cooling holes placed along the span of a fully-cooled high pressure turbine blade in a stationary, linear cascade on film cooling effectiveness is studied using the Pressure Sensitive Paint (PSP) technique. Effect of showerhead injection at the leading edge and the presence of compound angled, diffusing holes on the pressure and suction side are also examined. Six rows of compound angled shaped film cooling holes are provided on the pressure side while four such rows are provided on the suction side of the blade. The holes have a laidback and fan-shaped diffusing cross-section. Another three rows of cylindrical holes are drilled at a typical angle on the leading edge to capture the effect of showerhead film coolant injection. The film cooling hole arrangement simulates a typical film cooled blade design used in stage 1 rotor blades for gas turbines used for power generation. A typical blowing ratio is defined for each film hole row and tests are performed for 100%, 150% and 200% of this typical value. Tests are performed for inlet Mach numbers of 0.36 and 0.45 with corresponding exit Mach numbers of 0.51 and 0.68 respectively. The flow remains subsonic in the throat region for both Mach numbers. The corresponding free stream Reynolds number, based on the axial chord length and the exit velocity, are 1.3 million and 1.74 million respectively. Freestream turbulence intensity level at the cascade inlet is 6%. Results show that varying blowing ratios can have a significant impact on film-cooling effectiveness distribution. Large spanwise variations in effectiveness distributions are also observed. Similar distributions were observed for both Mach numbers.

Sign in / Sign up

Export Citation Format

Share Document