A Numerical Study on the Characteristics of How Heat and Cooling Transfer on the Leading Edge of a Film-Cooling Blade

2011 ◽  
Vol 383-390 ◽  
pp. 3963-3968
Author(s):  
Shao Hua Li ◽  
Li Mei Du ◽  
Wen Hua Dong ◽  
Ling Zhang
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Numerical Study ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Stagnation Line ◽  
Film Cooling Effectiveness ◽  
Pressure Surface

In this paper, a numerical simulation was performed to investigate heat transferring characteristics on the leading edge of a blade with three rows of holes of film-cooling using Realizable k- model. Three rows of holes were located on the suction side leading edge stagnation line and the pressure surface. The difference of the cooling efficiency and the heat transfer of the three rows of holes on the suction side and pressure side were analyzed; the heat transfer and film cooling effectiveness distribution in the region of leading edge are expounded under different momentum rations.The results show that under the same condition, the cooling effectiveness on the pressure side is more obvious than the suction side, but the heat transfer is better on the suction side than the pressure side. The stronger momentum rations are more effective cooling than the heat transfer system.

Numerical Study on the Influence of Trench Width on Film Cooling Characteristics of Double-Wave Trench

2017 ◽  
Author(s):  
Bo-lun Zhang ◽  
Li Zhang ◽  
Hui-ren Zhu ◽  
Jian-sheng Wei ◽  
Zhong-yi Fu
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Numerical Study ◽  
Leading Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Trench Width

Film cooling performance of the double-wave trench was numerically studied to improve the film cooling characteristics. Double-wave trench was formed by changing the leading edge and trailing edge of transverse trench into cosine wave. The film cooling characteristics of transverse trench and double-wave trench were numerically studied using Reynolds Averaged Navier Stokes (RANS) simulations with realizable k-ε turbulence model and enhanced wall treatment. The film cooling effectiveness and heat transfer coefficient of double-wave trench at different trench width (W = 0.8D, 1.4D, 2.1D) conditions are investigated, and the distribution of temperature field and flow field were analyzed. The results show that double-wave trench effectively improves the film cooling effectiveness and the uniformity of jet at the downstream wall of the trench. The span-wise averaged film cooling effectiveness of the double-wave trench model increases 20–63% comparing with that of the transverse trench at high blowing ratio. The anti-counter-rotating vortices which can press the film on near-wall are formed at the downstream wall of the double-wave trench. With the double-wave trench width decreasing, the film cooling effectiveness gradually reduces at the hole center-line region of the downstream trench. With the increase of the blowing ratio, the span-wise averaged heat transfer coefficient increases. The span-wise averaged heat transfer coefficient of the double-wave trench with 0.8D and 2.1D trench width is higher than that of the double-wave trench with 1.4D trench width at the high blowing ratio conditions.

Film-Cooling on a Gas Turbine Blade Pressure Side or Suction Side With Compound Angle Shaped Holes

2008 ◽  
Vol 131 (1) ◽  
Author(s):  
Zhihong Gao ◽  
Diganta P. Narzary ◽  
Je-Chin Han
Keyword(s):  
Turbine Blade ◽  
Film Cooling ◽  
Pressure Ratio ◽  
Suction Side ◽  
Pressure Side ◽  
Blade Surface ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Pressure Surface ◽  
Compound Angle

The film-cooling effectiveness on the surface of a high pressure turbine blade is measured using the pressure sensitive paint technique. Compound angle laidback fan-shaped holes are used to cool the blade surface with four rows on the pressure side and two rows on the suction side. The coolant injects to one side of the blade, either pressure side or suction side. The presence of wake due to the upstream vanes is simulated by placing a periodic set of rods upstream of the test blade. The wake rods can be clocked by changing their stationary positions to simulate progressing wakes. The effect of wakes is recorded at four phase locations along the pitchwise direction. The freestream Reynolds number, based on the axial chord length and the exit velocity, is 750,000. The inlet and exit Mach numbers are 0.27 and 0.44, respectively, resulting in a pressure ratio of 1.14. Five average blowing ratios ranging from 0.4 to 1.5 are tested. Results reveal that the tip-leakage vortices and endwall vortices sweep the coolant on the suction side to the midspan region. The compound angle laidback fan-shaped holes produce a good film coverage on the suction side except for the regions affected by the secondary vortices. Due to the concave surface, the coolant trace is short and the effectiveness level is low on the pressure surface. However, the pressure side acquires a relatively uniform film coverage with the multiple rows of cooling holes. The film-cooling effectiveness increases with the increasing average blowing ratio for either side of coolant ejection. The presence of stationary upstream wake results in lower film-cooling effectiveness on the blade surface. The compound angle shaped holes outperform the compound angle cylindrical holes by the elevated film-cooling effectiveness, particularly at higher blowing ratios.

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.

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 Prediction of Film Cooling and Heat Transfer on the Leading Edge of a Rotating Blade With Two Rows Holes in a 1-1/2 Turbine Stage at Design and Off Design Conditions

2005 ◽  
Author(s):  
Huitao Yang ◽  
Hamn-Ching Chen ◽  
Je-Chin Han ◽  
Hee-Koo Moon
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Leading Edge ◽  
Suction Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Rotating Speed ◽  
Rotating Blade

Numerical simulations were performed to predict the film cooling effectiveness and the associated heat transfer coefficient on the leading edge of a rotating blade in a 1-1/2 turbine stage using a Reynolds stress turbulence model together with a non-equilibrium wall function. Simulations were performed for both the design and off-design conditions to investigate the effects of blade rotation on the leading edge film cooling effectiveness and heat transfer coefficient distributions. It was found that the tilt stagnation line on the leading edge of rotor moves from the pressure side to the suction side, and the instantaneous coolant streamlines shift from the suction side to the pressure side with increasing rotating speed. This trend was supported by the experimental results. The result also showed that the heat transfer coefficient increases, but film cooling effectiveness decreases with increasing rotating speed. In addition, the unsteady characteristics of the film cooling and heat transfer at different time phases, as well as different rotating speeds, were also reported.

Heat Transfer and Film Cooling of Blade Tips and Endwalls

2011 ◽  
Vol 134 (4) ◽  
Author(s):  
S. Naik ◽  
C. Georgakis ◽  
T. Hofer ◽  
D. Lengani
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Film Cooling ◽  
Leading Edge ◽  
Pressure Side ◽  
High Pressure Turbine ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blade Tip

This paper investigates the flow, heat transfer, and film cooling effectiveness of advanced high pressure turbine blade tips and endwalls. Two blade tip configurations have been studied, including a full rim squealer and a partial squealer with leading edge and trailing edge cutouts. Both blade tip configurations have pressure side film cooling and cooling air extraction through dust holes, which are positioned along the airfoil camber line on the tip cavity floor. The investigated clearance gap and the blade tip geometry are typical of that commonly found in the high pressure turbine blades of heavy-duty gas turbines. Numerical studies and experimental investigations in a linear cascade have been conducted at a blade exit isentropic Mach number of 0.8 and a Reynolds number of 9×105. The influence of the coolant flow ejected from the tip dust holes and the tip pressure side film holes has also been investigated. Both the numerical and experimental results showed that there is a complex aerothermal interaction within the tip cavity and along the endwall. This was evident for both tip configurations. Although the global heat transfer and film cooling characteristics of both blade tip configurations were similar, there were distinct local differences. The partial squealer exhibited higher local film cooling effectiveness at the trailing edge but also low values at the leading edge. For both tip configurations, the highest heat transfer coefficients were located on the suction side rim within the midchord region. However, on the endwall, the highest heat transfer rates were located close to the pressure side rim and along most of the blade chord. Additionally, the numerical results also showed that the coolant ejected from the blade tip dust holes partially impinges onto the endwall.

Turbine Blade Tip Film Cooling With Blade Rotation: Part I — Tip and Pressure Side Coolant Injection

2015 ◽  
Author(s):  
Onieluan Tamunobere ◽  
Sumanta Acharya
Keyword(s):  
Heat Transfer ◽  
Film Cooling ◽  
Suction Side ◽  
High Heat ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
High Heat Transfer ◽  
Blade Tip

This is the first in a two-part series of an experimental film cooling study conducted on the tip of a turbine blade with a blade rotation speed of 1200 RPM. In this part of the study, the coolant is injected from the blade tip and pressure side (PS) holes, and the effect of the blowing ratio on the heat transfer coefficient and film cooling effectiveness of the blade tip is investigated. The blade has a tip clearance of 1.7% of the blade span and consists of a cut back squealer rim, two cylindrical tip holes and six shaped pressure side holes. The stator-rotor-stator test section is housed in a closed loop wind tunnel that allows for the performance of transient heat transfer tests. Measurements of the heat transfer coefficient and film cooling effectiveness are done on the blade tip using liquid crystal thermography. These measurements are reported for the no coolant case and for blowing ratios of 1.0, 1.5, 2.0, 3.0 and 4.0. The heat transfer results for the no coolant injection show a region of high heat transfer on the blade tip near the blade leading edge region as the incident flow impinges on that region. This region of high heat transfer extends and stretches on the tip as more coolant is introduced through the tip holes at higher blowing ratios. The cooling results show that increasing the blowing ratio increases the film cooling effectiveness. The tip film cooling profile is such that the tip coolant is pushed towards the blade suction side thereby providing better coverage in that region. The shift in coolant flow profile towards the blade suction side as opposed to the pressure side in stationary studies can primarily be attributed to the effects of the blade relative motion.

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.

Assessment of Steady State PSP and Transient Ir Measurement Techniques for Leading Edge Film Cooling

2005 ◽  
Author(s):  
Zhihong Gao ◽  
Lesley M. Wright ◽  
Je-Chin Han
Keyword(s):  
Heat Transfer ◽  
Steady State ◽  
Film Cooling ◽  
Measurement Techniques ◽  
Turbine Blades ◽  
Leading Edge ◽  
Cooling Effectiveness ◽  
Stagnation Line ◽  
Film Cooling Effectiveness ◽  
Film Cooling Holes

Film cooling is commonly used on the leading edge of turbine blades to protect the blade surface from hot mainstream gases in the turbine. Obtaining detailed film cooling effectiveness distributions on the leading edge can be challenging. This paper considers two measurement techniques which can be applied to the leading edge (modeled by a cylinder) to obtain detailed distributions of the film effectiveness. A steady state pressure sensitive paint (PSP) technique and a transient infrared (IR) thermography technique are used to obtain detailed film cooling effectiveness distributions on the cylinder. The cylinder, 7.62 cm in diameter, is placed in a low speed wind tunnel, with the mainstream flow having a Reynolds number of 100,900 (based on the cylinder diameter). The cylinder has two rows of film cooling holes located at ±15° from the cylinder’s stagnation line. The pitch-to-diameter ratio of the film holes is 4, and holes are inclined 30° in spanwise direction. PSP continues to show promise for film cooling effectiveness measurements. Detailed distributions can be obtained near the film cooling holes because this technique relies on mass transfer rather than heat transfer. In order to reduce the error caused by conduction in heat transfer experiments, transient measurement techniques are favorable. Transient IR measurements are taken, and film cooling effectiveness is determined on the cylinder’s surface. Although the effect of conduction is reduced with the transient IR technique (compared to a steady state heat transfer experiment), heat conduction through the cylinder has not been eliminated (or even minimized). Without correction, the results obtained from transient heat transfer experiments must be used cautiously. For this reason, PSP is developing a niche within the gas turbine community for detailed film cooling effectiveness measurements.

Heat Transfer and Film Cooling of Blade Tips and Endwalls

2010 ◽  
Author(s):  
S. Naik ◽  
C. Georgakis ◽  
T. Hofer ◽  
D. Lengani
Keyword(s):  
Heat Transfer ◽  
High Pressure ◽  
Film Cooling ◽  
Leading Edge ◽  
Pressure Side ◽  
High Pressure Turbine ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blade Tip

This paper investigates the flow, heat transfer and film cooling effectiveness of advanced high-pressure turbine blade tips and endwall. Two blade tip configurations have been studied, including a full rim squealer and a partial squealer with a leading edge and trailing edge cut-out. Both blade tip configurations have pressure side film cooling, and cooling air extraction through dust holes which are positioned along the airfoil camber line on the tip cavity floor. The investigated clearance gap and the blade tip geometry are typical of that commonly found in the high pressure turbine blades of heavy-duty gas turbines. Numerical studies and experimental investigations in a linear cascade have been conducted at a blade exit isentropic Mach number of 0.8 and a Reynolds number of 9 × 105. The influence of the coolant flow ejected from the tip dust holes and the tip pressure side film holes has also been investigated. Both the numerical and experimental results showed that there is a complex aero-thermal interaction within the tip cavity and along the endwall. This was evident for both tip configurations. Although, the global heat transfer and film cooling characteristics of both blade tip configurations were similar, there were distinct local differences. The partial squealer exhibited higher local film cooling effectiveness at the trailing edge but also low values at the leading edge. For both tip configurations, the highest heat transfer coefficients were located on the suction side rim within the mid-chord region. However on the endwall, the highest heat transfer rates were located close to the pressure side rim and along most of the blade chord. Additionally, the numerical results also showed that the coolant ejected from the blade tip dust holes partially impinges onto the endwall.

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