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

2011 ◽  
Vol 134 (1) ◽  
Author(s):  
Shantanu Mhetras ◽  
Je-Chin Han ◽  
Ron Rudolph
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Gas Turbines ◽  
Leading Edge ◽  
Intensity Level ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Throat Region ◽  
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 technique. The effect of showerhead injection at the leading edge and the presence of compound angled, diffusing holes on the pressure and suction sides 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. An optimal target blowing ratio is defined for each film hole row, and tests are performed for 100%, 150%, and 200% of this target 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 freestream Reynolds numbers, based on the axial chord length and the exit velocity, are 1.3×106 and 1.74×106, respectively. Freestream turbulence intensity level at the cascade inlet is 6%. The 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.

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.

Film-Cooling Effectiveness From Shaped Film Cooling Holes for a Gas Turbine Blade

2008 ◽  
Author(s):  
Shantanu Mhetras ◽  
Je-Chin Han ◽  
Ron Rudolph
Keyword(s):  
Mach Number ◽  
Turbine Blade ◽  
Film Cooling ◽  
Gas Turbines ◽  
Leading Edge ◽  
Intensity Level ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Throat Region ◽  
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. 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 holes are drilled 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 minimum blowing ratio is defined for each film hole row and tests are performed for 1.0x, 1.33x, 1.67x, 2.0x and 2.67x of this minimum value. Tests are performed for an inlet Mach number of 0.36 with a corresponding exit Mach number of 0.51. The flow remains subsonic in the throat region. The corresponding free stream Reynolds number, based on the axial chord length and the exit velocity, is 1.3 million. Turbulence intensity level at the cascade inlet is 5% with an integral length scale of around 5cm. 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. Results also show that the effectiveness magnitudes from superposition of effectiveness data from individual rows are comparable with that from full coverage film cooling.

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.

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.

Effect of Superposition on Spanwise Film-Cooling Effectiveness Distribution on a Gas Turbine Blade

2006 ◽  
Author(s):  
Shantanu Mhetras ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Pressure Ratio ◽  
Suction Side ◽  
Axial Direction ◽  
Intensity Level ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Full Coverage ◽  
Film Cooling Holes

The effect of film cooling holes placed along the span of high pressure turbine blade in a 5 bladed linear cascade on film cooling effectiveness is studied using the Pressure Sensitive Paint (PSP) technique. Four rows of film cooling holes are provided on the pressure side while two such rows are provided on the suction side of the blade. Around 22 cylindrical holes with a diameter of 0.65mm are drilled in each row at a compound angle of 45° to the blade span in the radial direction and at 45° in the axial direction. Film cooling effectiveness over the entire blade region is determined from full coverage film cooling with coolant blowing from all rows and from each individual row. The effect of superposition of film cooling effectiveness from each individual row is then compared with full coverage film cooling. The coolant is injected at four different average blowing ratios of 0.6, 0.9, 1.2 and 1.5. The free stream Reynolds number, based on the axial chord length and the exit velocity, is 750,000 and the inlet and the exit Mach numbers are 0.27 and 0.44, respectively resulting in a blade pressure ratio of 1.14. Turbulence intensity level at the cascade inlet is 6% with an integral length scale of around 5cm. Results show that the effectiveness magnitudes from superposition of effectiveness data from individual rows are comparable with that from full coverage film cooling. Varying blowing ratios can have a significant impact on film-cooling effectiveness distribution with a blowing ratio of 0.6 showing highest effectiveness immediately downstream of the holes.

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.

Comparison of RANS and DES Modeling Against Measurements of Leading Edge Film Cooling on a First-Stage Vane

2016 ◽  
Author(s):  
S. Ravelli ◽  
G. Barigozzi
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Leading Edge ◽  
Spanwise Direction ◽  
Intensity Level ◽  
Edge Region ◽  
Cooling Effectiveness ◽  
Stagnation Line ◽  
Film Cooling Effectiveness ◽  
Adiabatic Effectiveness

The performance of a showerhead arrangement of film cooling in the leading edge region of a first stage nozzle guide vane was experimentally and numerically evaluated. A six-vane linear cascade was tested at an isentropic exit Mach number of Ma2s = 0.42, with a high inlet turbulence intensity level of 9%. The showerhead cooling scheme consists of four staggered rows of cylindrical holes evenly distributed around the stagnation line, angled at 45° towards the tip. The blowing ratios tested are BR = 2.0, 3.0 and 4.0. Adiabatic film cooling effectiveness distributions on the vane surface around the leading edge region were measured by means of Thermochromic Liquid Crystals technique. Since the experimental contours of adiabatic effectiveness showed that there is no periodicity across the span, the CFD calculations were conducted by simulating the whole vane. Within the RANS framework, the very widely used Realizable k-ε (Rke) and the Shear Stress Transport k-ω (SST) turbulence models were chosen for simulating the effect of the BR on the surface distribution of adiabatic effectiveness. The turbulence model which provided the most accurate steady prediction, i.e. Rke, was selected for running Detached Eddy Simulation at the intermediate value of BR = 3. Fluctuations of the local temperature were computed by DES, due to the vortex structures within the shear layers between the main flow and the coolant jets. Moreover, mixing was enhanced both in the wall-normal and spanwise direction, compared to RANS modeling. DES roughly halved the prediction error of laterally averaged film cooling effectiveness on the suction side of the leading edge. However, neither DES nor RANS provided the expected decay of effectiveness progressing downstream along the pressure side, with 15% overestimation of ηav at s/C =0.2.

Numerical modeling of innovative film cooling hole schemes

2021 ◽  
Author(s):  
Siavash Khajehhasani
Keyword(s):  
Flat Plate ◽  
Turbine Blade ◽  
Film Cooling ◽  
Leading Edge ◽  
Lateral Spread ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Film Cooling Effectiveness ◽  
Mainstream Flow ◽  
Hole Geometry

A numerical investigation of the film cooling performance on novel film hole schemes is presented using Reynolds-Averaged Navier-Stokes analysis. The investigation considers low and high blowing ratios for both flat plate film cooling and the leading edge of a turbine blade. A novel film hole geometry using a circular exit shaped hole is proposed, and the influence of an existing sister holes’ technique is investigated. The results indicate that high film cooling effectiveness is achieved at higher blowing ratios, results of which are even greater when in the presence of discrete sister holes where film cooling effectiveness results reach a plateau. Furthermore, a decrease in the strength of the counter-rotating vortex pairs is evident, which results in more attached coolant to the plate’s surface and a reduction in aerodynamic losses. Modifications are made to the spanwise and streamwise locations of the sister holes around the conventional cylindrical hole geometry. It is found that the spanwise variations have a significant influence on the film cooling effectiveness results, while only minor effects are observed for the streamwise variations. Positioning the sister holes in locations farther from the centerline increases the lateral spreading of the coolant air over the plate’s surface. This result is further verified through the flow structure analysis. Combinations of sister holes are joined with the primary injection hole to produce innovative variant sister shaped single-holes. The jet lift-off is significantly decreased for the downstream and up/downstream configurations of the proposed scheme for the flat plate film cooling. These schemes have shown notable film cooling improvements whereby more lateral distribution of coolant is obtained and less penetration of coolant into the mainstream flow is observed. The performance of the sister shaped single-holes are evaluated at the leading edge of a turbine blade. At the higher blowing ratios, a noticeable improvement in film cooling performance including the effectiveness and the lateral spread of the cooling air jet has been observed for the upstream and up/downstream schemes, in particular on the suction side. It is determined that the mixing of the coolant with the high mainstream flow at the leading edge of the blade is considerably decreased for the upstream and up/downstream configurations and more adhered coolant to the blade’s surface is achieved.

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.

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