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.

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.

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.

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.

Unsteady Wake and Coolant Density Effects on Turbine Blade Film Cooling Using PSP Technique

2010 ◽  
Author(s):  
Akhilesh P. Rallabandi ◽  
Shiou-Jiuan Li ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Suction Side ◽  
Suction Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Unsteady Wake ◽  
Lift Off ◽  
Film Cooling Holes

The effect of an unsteady stator wake (simulated by wake rods mounted on a spoke wheel wake generator) on the modeled rotor blade is studied using the Pressure Sensitive Paint (PSP) mass transfer analogy method. Emphasis of the current study is on the mid-span region of the blade. The flow is in the low Mach number (incompressible) regime. The suction (convex) side has simple angled cylindrical film-cooling holes; the pressure (concave) side has compound angled cylindrical film cooling holes. The blade also has radial shower-head leading edge film cooling holes. Strouhal numbers studied range from 0 to 0.36; the exit Reynolds Number based on the axial chord is 530,000. Blowing ratios range from 0.5 to 2.0 on the suction side; 0.5 to 4.0 on the pressure side. Density ratios studied range from 1.0 to 2.5, to simulate actual engine conditions. The convex suction surface experiences film-cooling jet lift-off at higher blowing ratios, resulting in low effectiveness values. The film coolant is found to reattach downstream on the concave pressure surface, increasing effectiveness at higher blowing ratios. Results show deterioration in film cooling effectiveness due to increased local turbulence caused by the unsteady wake, especially on the suction side. Results also show a monotonic increase in film-cooling effectiveness on increasing the coolant to mainstream density ratio.

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 Unsteady Wake on Film-Cooling Effectiveness Distribution on a Gas Turbine Blade With Compound Shaped Holes

2007 ◽  
Author(s):  
Diganta P. Narzary ◽  
Zhihong Gao ◽  
Shantanu Mhetras ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Pressure Ratio ◽  
Suction Side ◽  
Intensity Level ◽  
Pressure Side ◽  
Wake Effect ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Cylindrical Holes

The effect of fan-shaped, laid-back compound angled cooling holes placed along the span of a fully-cooled high pressure turbine blade in a 5-blade linear cascade on film cooling effectiveness is studied using the Pressure Sensitive Paint (PSP) technique. Four rows of shaped film cooling holes are provided on the pressure side while two such rows are provided on the suction side of the blade. Three rows of cylindrical holes are drilled at 30° to the surface on the leading edge to capture the effect of showerhead film coolant injection. The coolant is injected at four different average blowing ratios of 0.3, 0.6, 0.9 and 1.2. Presence of wake due to upstream vanes is studied by placing a periodic set of rods upstream of the test blade. The wake is generated using 4.8mm diameter rods. The wake rods can be clocked by changing their stationary positions in front of the test blade to simulate a progressing wake. Effect of wake is recorded at four phase locations with equal intervals. 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 fan-shaped, laid-back compound angled holes produce uniform and wide coolant coverage on the suction side except for those regions affected by the passage and tip leakage vortices. The advantage of compound shaped hole design is seen from the higher effectiveness values on the suction side compared to that of the compound cylindrical holes. The presence of a stationary upstream wake can result in lower film cooling effectiveness on the blade surface. Variation of blowing ratio from 0.3 to 1.2 show more or less uniform increment in effectiveness increase on the pressure side, whereas on the suction side, the increment shows signs of saturation beyond M = 0.6.

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.

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