Leading Edge Film Cooling and the Influence of Shaped Holes at Design and Off-Design Conditions

2007 ◽  
Author(s):  
Guillaume Wagner ◽  
Peter Ott ◽  
Gregory Vogel ◽  
Shailendra Naik
Keyword(s):  
Film Cooling ◽  
Incidence Angle ◽  
Wedge Angle ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Blowing Ratio ◽  
Cylindrical Holes ◽  
Lift Off ◽  
Film Cooling Holes

Transient liquid crystal experiments have been carried out to measure the effectiveness and heat transfer characteristics of leading-edge film cooling for three different film cooling holes configurations at design and off-design incidence angle. The three configurations are based on the same representative leading edge model of a turbine blade, consisting of a symmetrical blunt body with a specific leading edge wedge angle. Film cooling is introduced from two rows of cooling holes, representative of a pressure-side row and a suction-side row. At design incidence, film cooling performances are symmetric. There is a jet lift-off situation and shaped holes significantly improve the film cooling performances because of a better lateral coverage and a reduced coolant momentum at the hole exit. At 5° off-design incidence angle, on the suction side, the situation is similar to that of a 0° incidence but with higher film cooling performances due to a reduced local blowing ratio. At 5° incidence on the pressure side, a beneficial interaction between the jets of the pressure side row appears. For middle and high blowing ratio, the film cooling performances are also better than 0° incidence. At 5° incidence on the pressure side, shaped holes also improve the film cooling performances in comparison to cylindrical holes.

An Experimental Investigation of Flow Characteristics Downstream of Discrete Film Cooling Holes on Turbine Blade Leading Edge

2011 ◽  
Vol 383-390 ◽  
pp. 5553-5560
Author(s):  
Shao Hua Li ◽  
Hong Wei Qu ◽  
Mei Li Wang ◽  
Ting Ting Guo
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Flow Characteristics ◽  
Leading Edge ◽  
Suction Side ◽  
Wall Region ◽  
Pressure Side ◽  
Blowing Ratio ◽  
Curvature Gradient ◽  
Film Cooling Holes

The gas turbine blade was studied on the condition that the mainstream velocity was 10m/s and the Renolds number based on the chord length of the blade was 160000.The Hot-film anemometer was used to measure the two-dimension speed distribution along the downstream of the film cooling holes on the suction side and the pressure side. The conclusions are as follows: When the blowing ratio of the suction side and the pressure side increasing, the the mainstream and the jet injection mixing center raising. Entrainment flow occurs at the position where the blade surface with great curvature gradient, simultaneously the mixing flow has a wicked adhere to the wall. The velocity gradient of the u direction that on the suction side increase obviously, also the level of the wall adherence is better than the pressure side. With the x/d increasing, the velocity u that on the pressure side gradually become irregularly, also the secondary flow emerged near the wall region where the curvature is great. The blowing ratio on the suction side has a little influence on velocity v than that on the pressure side.

The Experimental and Numerical Research for Linear Cascade Film Cooling With Different Hole Shapes

2010 ◽  
Author(s):  
Yi Lu ◽  
Yinyi Hong ◽  
Zhirong Lin ◽  
Xin Yuan
Keyword(s):  
Film Cooling ◽  
Large Scale ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Linear Cascade ◽  
Blowing Ratio ◽  
Cylindrical Holes

Detailed film cooling effectiveness distributions were experimentally obtained on a turbine vane platform within a linear cascade. Testing was done in a large scale five-vane cascade with low freestream Renolds number condition 634,000 based on the axial chord length and the exit velocity. The detailed film-cooling effectiveness distributions on the platform were obtained using pressure sensitive paint technique. Two film-cooling hole configurations, cylindrical and fan-shaped, were used to cool the vane surface with two rows on pressure side, two rows on suction side and three rows on leading edge. For cylindrical holes, the blowing ratio of the coolant through the discrete cooling holes on pressure side and suction side ranged from 0.3 to 1.5 (based on the inlet mainstream velocity) while the blowing ratio ranging from 0.15 to 1.5 on leading edge; for fan-shaped holes, the four blowing ratios were 0.5, 1.0, 1.5 and 2.0. Results showed that average film-cooling effectiveness decreased with increasing blowing rate for the cylindrical holes, while the fan-shaped passage showed increased film-cooling effectiveness with increasing blowing ratio, indicating the fan-shaped cooling holes helped to improve film-cooling effectiveness by reducing overall jet liftoff. Fan-shaped holes improved average film-cooling effectiveness by 93.2%, 287.6% and 489.6% on pressure side, −4.1%, 27.9% and 78.2% on suction side over cylindrical holes at the blowing ratio of 0.5, 1.0 and 1.5 respectively. Numerical results were used to analyze the details of the flow and heat transfer on the cooling area with two turbulence models. Results demonstrated that tendency of the film cooling effectiveness distribution of numerical calculation and experimental measurement was generally consistent at different blowing ratio.

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.

Effect of Coolant Density on Leading Edge Showerhead Film Cooling Using PSP Measurement Technique

2013 ◽  
Author(s):  
Shiou-Jiuan Li ◽  
Shang-Feng Yang ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Pressure Sensitive ◽  
Cylindrical Holes ◽  
Film Cooling Holes ◽  
Compound Angle ◽  
Low Speed Wind Tunnel

The density ratio effect on leading edge showerhead film cooling has been studied experimentally using the pressure sensitive paint (PSP) mass transfer analogy method. Leading edge model is a blunt body with a semi-cylinder and an after body. There are two designs: seven-row and three-row of film cooling holes for simulating vane and blade, respectively. The film holes are located at 0 (stagnation row), ±15, ±30, and ±45 deg for seven-row design, and at 0 and ±30 for three-row design. Four film holes configurations are used for both test designs: radial angle cylindrical holes, compound angle cylindrical holes, radial angle shaped holes, and compound angle shaped holes. Coolant to mainstream density ratio varies from DR = 1.0, 1.5, to 2.0 while blowing ratio varies from M = 0.5 to 2.1. Experiments were conducted in a low speed wind tunnel with Reynolds number 100,900 based on mainstream velocity and diameter of the cylinder. The mainstream turbulence intensity near leading edge model is about 7%. The results show the shaped holes have overall higher film cooling effectiveness than cylindrical holes, and radial angle holes are better than compound angle holes, particularly at higher blowing ratio. Larger density ratio makes more coolant attach to the surface and increases film protection for all cases. Radial angle shaped holes provides best film cooling at higher density ratio and blowing ratio for both designs.

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.

Numerical Investigation on Film Cooling Effectiveness of Stator Blade with Different Density Ratio

2014 ◽  
Vol 521 ◽  
pp. 104-107
Author(s):  
Ling Zhang ◽  
Quan Heng Jin ◽  
Da Fei Guo
Keyword(s):  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Middle Section ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Stator Blade

The Realizable k-ε turbulence model was performed to investigate the film cooling effectiveness with different blowing ratio 1,1.5,2 and different density ratio 1,1.5,2.The results show that, cooling effectiveness increases with the augment of blowing ratio. On the pressure side, cooling effectiveness increases with the augment of density ratio. On the suction side, with higher density ratio the leading edge cooling increases, the middle section reduces, and the trailing edge cooling effectiveness increases first decreases.

Influence of Film-Hole Shape and Angle on Showerhead Film Cooling Using PSP Technique

2009 ◽  
Vol 131 (6) ◽  
Author(s):  
Zhihong Gao ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Leading Edge ◽  
Stagnation Line ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Cooling Hole ◽  
Cooling Design ◽  
Cylindrical Holes ◽  
Film Cooling Holes ◽  
Compound Angle

The effect of film-hole geometry and angle on turbine blade leading edge film cooling has been experimentally studied using the pressure sensitive paint technique. The leading edge is modeled by a blunt body with a semicylinder and an after-body. Two film cooling designs are considered: a heavily film cooled leading edge featured with seven rows of film cooling holes and a moderately film cooled leading edge with three rows. For the seven-row design, the film holes are located at 0 deg (stagnation line), ±15 deg, ±30 deg, and ±45 deg on the model surface. For the three-row design, the film holes are located at 0 deg and ±30 deg. Four different film cooling hole configurations are applied to each design: radial angle cylindrical holes, compound angle cylindrical holes, radial angle shaped holes, and compound angle shaped holes. Testing was done in a low speed wind tunnel. The Reynolds number, based on mainstream velocity and diameter of the cylinder, is 100,900. The mainstream turbulence intensity is about 7% near of leading edge model and the turbulence integral length scale is about 1.5 cm. Five averaged blowing ratios are tested ranging from M=0.5 to M=2.0. The results show that the shaped holes provide higher film cooling effectiveness than the cylindrical holes, particularly at higher average blowing ratios. The radial angle holes give better effectiveness than the compound angle holes at M=1.0–2.0. The seven-row film cooling design results in much higher effectiveness on the leading edge region than the three-row design at the same average blowing ratio or same amount coolant flow.

Effect of Coolant Density on Leading Edge Showerhead Film Cooling Using the Pressure Sensitive Paint Measurement Technique

2013 ◽  
Vol 136 (5) ◽  
Author(s):  
Shiou-Jiuan Li ◽  
Shang-Feng Yang ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Pressure Sensitive Paint ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Pressure Sensitive ◽  
Cylindrical Holes ◽  
Film Cooling Holes ◽  
Compound Angle

The density ratio effect on leading edge showerhead film cooling has been studied experimentally using the pressure sensitive paint (PSP) mass transfer analogy method. The leading edge model is a blunt body with a semicylinder and an after body. There are two designs: seven-row and three-row of film cooling holes for simulating a vane and blade, respectively. The film holes are located at 0 (stagnation row), ±15, ±30, and ±45 deg for the seven-row design, and at 0 and ±30 for the three-row design. Four film hole configurations are used for both test designs: radial angle cylindrical holes, compound angle cylindrical holes, radial angle shaped holes, and compound angle shaped holes. The coolant to mainstream density ratio varies from DR = 1.0, 1.5, to 2.0 while the blowing ratio varies from M = 0.5 to 2.1. Experiments were conducted in a low speed wind tunnel with Reynolds number 100,900 based on mainstream velocity and diameter of the cylinder. The mainstream turbulence intensity near the leading edge model is about 7%. The results show the shaped holes have an overall higher film cooling effectiveness than the cylindrical holes, and the radial angle holes are better than the compound angle holes, particularly at a higher blowing ratio. A larger density ratio makes more coolant attach to the surface and increases film protection for all cases. Radial angle shaped holes provide the best film cooling at a higher density ratio and blowing ratio for both designs.

The Effects of Inlet Reynolds Number, Exit Mach Number and Incidence Angle on Leading Edge Film Cooling Effectiveness of a Turbine Blade in a Linear Transonic Cascade

2015 ◽  
Author(s):  
Cong Liu ◽  
Hui-ren Zhu ◽  
Zhong-yi Fu ◽  
Run-hong Xu
Keyword(s):  
Reynolds Number ◽  
Mach Number ◽  
Film Cooling ◽  
Incidence Angle ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Exit Mach Number

This paper experimentally investigates the film cooling performance of a leading edge with three rows of film holes on an enlarged turbine blade in a linear cascade. The effects of blowing ratio, inlet Reynolds number, isentropic exit Mach number and off-design incidence angle (i<0°) are considered. Experiments were conducted in a short-duration transonic wind tunnel which can model realistic engine aerodynamic conditions and adjust inlet Reynolds number and exit Mach number independently. The surface film cooling measurements were made at the midspan of the blade using thermocouples based on transient heat transfer measurement method. The changing of blowing ratio from 1.7 to 3.3 leads to film cooling effectiveness increasing on both pressure side and suction side. The Mach number or Reynolds number has no effect on the film cooling effectiveness on pressure side nearly, while increasing these two factors has opposite effect on film cooling performance on suction side. The increasing Mach number decreases the film cooling effectiveness at the rear region mainly, while at higher Reynolds number condition, the whole suction surface has significantly higher film cooling effectiveness because of the increasing cooling air mass flow rate. When changing the incidence angle from −15° to 0°, the film cooling effectiveness of pressure side decreases, and it presents the opposite trend on suction side. At off-design incidence of −15° and −10°, there is a low peak following the leading edge on the pressure side caused by the separation bubble, but it disappears with the incidence and blowing ratio increased.

Effect of a Cutback Squealer and Cavity Depth on Film-Cooling Effectiveness on a Gas Turbine Blade Tip

2008 ◽  
Vol 130 (2) ◽  
Author(s):  
Shantanu Mhetras ◽  
Diganta Narzary ◽  
Zhihong Gao ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Suction Side ◽  
Pressure Side ◽  
Trailing Edge ◽  
Cavity Depth ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Blade Tip ◽  
Film Cooling Holes

Film-cooling effectiveness from shaped holes on the near tip pressure side and cylindrical holes on the squealer cavity floor is investigated. The pressure side squealer rim wall is cut near the trailing edge to allow the accumulated coolant in the cavity to escape and cool the tip trailing edge. Effects of varying blowing ratios and squealer cavity depth are also examined on film-cooling effectiveness. The film-cooling effectiveness distributions are measured on the blade tip, near tip pressure side and the inner pressure side and suction side rim walls using pressure sensitive paint technique. The internal coolant-supply passages of the squealer tipped blade are modeled similar to those in the GE-E3 rotor blade with two separate serpentine loops supplying coolant to the film-cooling holes. Two rows of cylindrical film-cooling holes are arranged offset to the suction side profile and along the camber line on the tip. Another row of shaped film-cooling holes is arranged along the pressure side just below the tip. The average blowing ratio of the cooling gas is controlled to be 0.5, 1.0, 1.5, and 2.0. A five-bladed linear cascade in a blow down facility with a tip gap clearance of 1.5% is used to perform the experiments. The free-stream Reynolds number, based on the axial chord length and the exit velocity, was 1,480,000 and the inlet and exit Mach numbers were 0.23 and 0.65, respectively. A blowing ratio of 1.0 is found to give best results on the pressure side, whereas the tip surfaces forming the squealer cavity give best results for M=2. Results show high film-cooling effectiveness magnitudes near the trailing edge of the blade tip due to coolant accumulation from upstream holes in the tip cavity. A squealer depth with a recess of 2.1mm causes the average effectiveness magnitudes to decrease slightly as compared to a squealer depth of 4.2mm.

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