scholarly journals Conjugate Heat Transfer Investigation on Swirl-Film Cooling at the Leading Edge of a Gas Turbine Vane

Entropy ◽  
2019 ◽  
Vol 21 (10) ◽  
pp. 1007 ◽  
Author(s):  
Du ◽  
Mei ◽  
Zou ◽  
Jiang ◽  
Xie
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Film Cooling ◽  
Conjugate Heat Transfer ◽  
Leading Edge ◽  
Pressure Side ◽  
Cooling Efficiency ◽  
Cooling Effectiveness ◽  
Turbine Vane ◽  
Cooling Chamber

Numerical calculation of conjugate heat transfer was carried out to study the effect of combined film and swirl cooling at the leading edge of a gas turbine vane with a cooling chamber inside. Two cooling chambers (C1 and C2 cases) were specially designed to generate swirl in the chamber, which could enhance overall cooling effectiveness at the leading edge. A simple cooling chamber (C0 case) was designed as a baseline. The effects of different cooling chambers were studied. Compared with the C0 case, the cooling chamber in the C1 case consists of a front cavity and a back cavity and two cavities are connected by a passage on the pressure side to improve the overall cooling effectiveness of the vane. The area-averaged overall cooling effectiveness of the leading edge () was improved by approximately 57%. Based on the C1 case, the passage along the vane was divided into nine segments in the C2 case to enhance the cooling effectiveness at the leading edge, and was enhanced by 75% compared with that in the C0 case. Additionally, the cooling efficiency on the pressure side was improved significantly by using swirl-cooling chambers. Pressure loss in the C2 and C1 cases was larger than that in the C0 case.

Overall Cooling Effectiveness on a Gas Turbine Vane With Swirl-Film Cooling

2019 ◽  
Author(s):  
Haifen Du ◽  
Danmei Xie ◽  
Wei Chen ◽  
Ziyue Mei ◽  
Jing Zhang
Keyword(s):  
Gas Turbine ◽  
Mass Flow ◽  
Film Cooling ◽  
Conjugate Heat Transfer ◽  
Leading Edge ◽  
Cooling Effectiveness ◽  
Turbine Vane ◽  
Cooling Chamber ◽  
Sst Turbulence Model ◽  
Flow Through

Abstract Numerical calculation of conjugate heat transfer is carried out to study the effect of combined film and swirl cooling at the leading edge of a gas turbine vane with a a cooling chamber inside, in which 3-D steady RANS approach with the k-ω SST turbulence model is used. Two different kinds of coolant chamber configuration (C1 and C2) are selected. In C2, the cooling chamber is composed of a front cavity and a back cavity, and the two cavities are connected by a passage which is divided into 16 segments. The comparative investigations between C1 and C2 cases have been carried out to study the effect of different cooling chambers at M = 0.25, 0.5, 1 and 2. For two cases, overall cooling effectiveness increases with M increasing. In C1 case, with increasing M, differences of mass flow through film holes rows will decrease. The variation of mass flow from holes changes by less than 26.7% at M = 2. However, in C2 case, mass flow through S1 and S2 is significantly larger than that through other film holes rows. Area-averaged overall effectiveness in C2 is larger by 2.5% at M = 0.25 compared to C1 case.

Measurements of Adiabatic Film and Overall Cooling Effectiveness on a Turbine Vane Pressure Side With a Trench

2011 ◽  
Author(s):  
Jason E. Albert ◽  
David G. Bogard
Keyword(s):  
Heat Transfer ◽  
Thermal Conductivity ◽  
Biot Number ◽  
Film Cooling ◽  
Conjugate Heat Transfer ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Blowing Ratio ◽  
Turbine Vane ◽  
Film Cooling Holes

Film cooling performance is typically quantified by separating the external convective heat transfer from the other components of the conjugate heat transfer that occurs in turbine airfoils. However, it is also valuable to assess the conjugate heat transfer in terms of the overall cooling effectiveness, which is a parameter of importance to airfoil designers. In the current study, adiabatic film effectiveness and overall cooling effectiveness values were measured for the pressure side of a simplified turbine vane model with three rows of showerhead cooling at the leading edge and one row of body film cooling holes on the pressure side. This was done by utilizing two geometrically identical models made from different materials. Adiabatic film effectiveness was measured using a very low thermal conductivity material, and the overall cooling effectiveness was measured using a material with a higher thermal conductivity selected such that the Biot number of the model matched that of a turbine vane at engine conditions. The theoretical basis for this matched-Biot number modeling technique is discussed in some detail. Additionally, two designs of pressure side body film cooling holes were considered in this study: a standard design of straight, cylindrical holes and an advanced design of “trenched” cooling holes in which the hole exits were situated in a recessed, transverse trench. This study was performed using engine representative flow conditions, including a coolant-to-mainstream density ratio of DR = 1.4 and a mainstream turbulence intensity of Tu = 20%. The results of this study show that adiabatic film and overall cooling effectiveness increase with blowing ratio for the showerhead and pressure side trenched holes. Performance decreases with blowing ratio for the standard holes due to coolant jet separation from the surface. Both body film designs have similar performance at a lower blowing ratio when the standard hole coolant jets remain attached. Far downstream of the cooling holes both designs perform similarly because film effectiveness decays more rapidly for the trenched holes.

Measurements of Adiabatic Film and Overall Cooling Effectiveness on a Turbine Vane Pressure Side With a Trench

2013 ◽  
Vol 135 (5) ◽  
Author(s):  
Jason E. Albert ◽  
David G. Bogard
Keyword(s):  
Heat Transfer ◽  
Thermal Conductivity ◽  
Biot Number ◽  
Film Cooling ◽  
Conjugate Heat Transfer ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Blowing Ratio ◽  
Turbine Vane ◽  
Film Cooling Holes

Film cooling performance is typically quantified by separating the external convective heat transfer from the other components of the conjugate heat transfer that occurs in turbine airfoils. However, it is also valuable to assess the conjugate heat transfer in terms of the overall cooling effectiveness, which is a parameter of importance to airfoil designers. In the current study, adiabatic film effectiveness and overall cooling effectiveness values were measured for the pressure side of a simplified turbine vane model with three rows of showerhead cooling at the leading edge and one row of body film cooling holes on the pressure side. This was done by utilizing two geometrically identical models made from different materials. Adiabatic film effectiveness was measured using a very low thermal conductivity material, and the overall cooling effectiveness was measured using a material with a higher thermal conductivity selected such that the Biot number of the model matched that of a turbine vane at engine conditions. The theoretical basis for this matched-Biot number modeling technique is discussed in some detail. Additionally, two designs of pressure side body film cooling holes were considered in this study: a standard design of straight, cylindrical holes and an advanced design of “trenched” cooling holes in which the hole exits were situated in a recessed, transverse trench. This study was performed using engine representative flow conditions, including a coolant-to-mainstream density ratio of DR = 1.4 and a mainstream turbulence intensity of Tu = 20%. The results of this study show that adiabatic film and overall cooling effectiveness increase with blowing ratio for the showerhead and pressure side trenched holes. Performance decreases with blowing ratio for the standard holes due to coolant jet separation from the surface. Both body film designs have similar performance at a lower blowing ratio when the standard hole coolant jets remain attached. Far downstream of the cooling holes both designs perform similarly because film effectiveness decays more rapidly for the trenched holes.

Heat Transfer Coefficient and Film-Cooling Effectiveness on a Gas Turbine Blade Tip

2002 ◽  
Author(s):  
Jae Su Kwak ◽  
Je-Chin Han
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Gas Turbine ◽  
Film Cooling ◽  
Rotor Blade ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blade Tip

The detailed distributions of heat transfer coefficient and film cooling effectiveness on a gas turbine blade tip were measured using a hue detection based transient liquid crystal technique. Tests were performed on a five-bladed linear cascade with blow down facility. The blade was a 2-dimensional model of a first stage gas turbine rotor blade with a profile of the GE-E3 aircraft gas turbine engine rotor blade. The Reynolds number based on cascade exit velocity and axial chord length was 1.1 × 106 and the total turning angle of the blade was 97.7°. The overall pressure ratio was 1.32 and the inlet and exit Mach number were 0.25 and 0.59, respectively. The turbulence intensity level at the cascade inlet was 9.7%. The blade model was equipped with a single row of film cooling holes at both the tip portion along the camber line and near the tip region of the pressure-side. All measurements were made at the three different tip gap clearances of 1%, 1.5%, and 2.5% of blade span and the three blowing ratios of 0.5, 1.0, and 2.0. Results showed that, in general, heat transfer coefficient and film effectiveness increased with increasing tip gap clearance. As blowing ratio increased, heat transfer coefficient decreased, while film effectiveness increased. Results also showed that adding pressure-side coolant injection would further decrease blade tip heat transfer coefficient but increase film effectiveness.

Conjugate heat transfer modeling of a turbine vane endwall with thermal barrier coatings

2019 ◽  
Vol 123 (1270) ◽  
pp. 1959-1981 ◽  
Author(s):  
Xing Yang ◽  
Zhenping Feng ◽  
Terrence W. Simon
Keyword(s):  
Heat Transfer ◽  
Thermal Barrier Coatings ◽  
Film Cooling ◽  
Conjugate Heat Transfer ◽  
Thermal Protection ◽  
Design Stage ◽  
Thermal Barrier ◽  
Cooling Effectiveness ◽  
Barrier Coatings ◽  
Turbine Vane

ABSTRACTAdvanced cooling techniques involving internal enhanced heat transfer and external film cooling and thermal barrier coatings (TBCs) are employed for gas turbine hot components to reduce metal temperatures and to extend their lifetime. A deeper understanding of the interaction mechanism of these thermal protection methods and the conjugate thermal behaviours of the turbine parts provides valuable guideline for the design stage. In this study, a conjugate heat transfer model of a turbine vane endwall with internal impingement and external film cooling is constructed to document the effects of TBCs on the overall cooling effectiveness using numerical simulations. Experiments on the same model with no TBCs are performed to validate the computational methods. Round and crater holes due to the inclusion of TBCs are investigated as well to address how film-cooling configurations affect the aero-thermal performance of the endwall. Results show that the TBCs have a profound effect in reducing the endwall metal temperatures for both cases. The TBC thermal protection for the endwall is shown to be more significant than the effect of increasing coolant mass flow rate. Although the crater holes have better film cooling performance than the traditional round holes, a slight decrement of overall cooling effectiveness is found for the crater configuration due to more endwall metal surfaces directly exposed to external mainstream flows. Energy loss coefficients at the vane passage exit show a relevant negative impact of adding TBCs on the cascade aerodynamic performance, particularly for the round hole case.

Film Cooling Jet Injection Effect in Heat Transfer Coefficient Augmentation for the Pressure Side Cooling of Turbine Vane

2014 ◽  
Author(s):  
Hossein Nadali Najafabadi ◽  
Matts Karlsson ◽  
Mats Kinell ◽  
Esa Utriainen
Keyword(s):  
Heat Transfer ◽  
Heat Transfer Coefficient ◽  
Transfer Coefficient ◽  
Film Cooling ◽  
Density Ratio ◽  
Leading Edge ◽  
Suction Side ◽  
Pressure Side ◽  
Time Resolved ◽  
Turbine Vane

Improving film cooling performance of turbine vanes and blades is often achieved through application of multiple arrays of cooling holes on the suction side, the showerhead region and the pressure side. This study investigates the pressure side cooling under the influence of single and multiple rows of cooling in the presence of a showerhead from a heat transfer coefficient augmentation perspective. Experiments are conducted on a prototype turbine vane working at engine representative conditions. Transient IR thermography is used to measure time-resolved surface temperature and the semi-infinite method is utilized to calculate the heat transfer coefficient on a low conductive material. Investigations are performed for cylindrical and fan-shaped holes covering blowing ratio 0.6 and 1.8 at density ratio of about unity. The freestream turbulence is approximately 5% close to the leading edge. The resulting heat transfer coefficient enhancement, the ratio of HTC with to that without film cooling, from different case scenarios have been compared to showerhead cooling only. Findings of the study highlight the importance of showerhead cooling to be used with additional row of cooling on the pressure side in order to reduce heat transfer coefficient enhancement. In addition, it is shown that extra rows of cooling will not significantly influence heat transfer augmentation, regardless of the cooling hole shape.

Influence of Coolant on Cooling Performance Sensitivity of Internally Convective Turbine Vane

2021 ◽  
Author(s):  
Thanapat Chotroongruang ◽  
Prasert Prapamonthon ◽  
Rungsimun Thongdee ◽  
Thanapat Thongmuenwaiyathon ◽  
Zhenxu Sun ◽  
...  
Keyword(s):  
Heat Transfer ◽  
Gas Turbine ◽  
Film Cooling ◽  
Internal Cooling ◽  
Cooling Effectiveness ◽  
Cooling Performance ◽  
Turbine Engine ◽  
Performance Sensitivity ◽  
Turbine Vane ◽  
Cooling Air

Abstract Based on the Brayton cycle for gas-turbine engines, the high thermal efficiency and power output of a gas-turbine engine can be obtainable when the gas-turbine engine operates at high turbine inlet temperatures. However, turbine components e.g., inlet guide vane, rotor blade, and stator vane request high cooling performance. Typically, internal cooling and film cooling are two effective techniques that are widely used to protect high thermal loads for the turbine components in a state-of-the-art gas turbine. Consequently, the high thermal efficiency and power output can be obtained, and the turbine lifespan can be prolonged, also. On top of that, a comprehensive understanding of flow and heat transfer phenomena in the turbine components is very important. As a result, both experiments and simulations have been used to improve the cooling performance of the turbine components. In fact, the cooling air used in the internal cooling and film cooling is partially extracted from the compressor. Therefore, variations in the cooling air affect the cooling performance of the turbine components directly. This paper presents a numerical study on the influence of the cooling air on cooling-performance sensitivity of an internally convective turbine vane, MARK II using the computational fluid dynamics (CFD)/conjugate heat transfer (CHT) with the SST k-ω turbulence model. Result comparisons are conducted in terms of pressure, temperature, and cooling effectiveness under the effects of the inlet temperature, mass flow rate, turbulence intensity, and flow direction of the cooling air. The cooling-performance sensitivity to the coolant parameters is shown through variations of local cooling effectiveness, and area and volume-weighted average cooling effectiveness.

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.

Overall and Adiabatic Effectiveness Values on a Scaled Up, Simulated Gas Turbine Vane

2013 ◽  
Vol 135 (5) ◽  
Author(s):  
Jason E. Dees ◽  
David G. Bogard ◽  
Gustavo A. Ledezma ◽  
Gregory M. Laskowski
Keyword(s):  
Heat Transfer ◽  
Surface Temperature ◽  
Gas Turbine ◽  
Film Cooling ◽  
Conjugate Heat Transfer ◽  
Flux Ratio ◽  
Boundary Layer Transition ◽  
Data Set ◽  
Turbine Vane ◽  
Adiabatic Effectiveness

Recent advances in computational power have made conjugate heat transfer simulations of fully conducting, film cooled turbine components feasible. However, experimental data available with which to validate conjugate heat transfer simulations is limited. This paper presents experimental measurements of external surface temperature on the suction side of a scaled up, fully conducting, cooled gas turbine vane. The experimental model utilizes the matched Bi method, which produces nondimensional surface temperature measurements that are representative of engine conditions. Adiabatic effectiveness values were measured on an identical near adiabatic vane with an identical geometry and cooling configuration. In addition to providing a valuable data set for computational code validation, the data shows the effect of film cooling on the surface temperature of a film cooled part. As expected, in nearly all instances, the addition of film cooling was seen to decrease the overall surface temperature. However, due to the effect of film injection causing early boundary layer transition, film cooling at a high momentum flux ratio was shown to actually increase surface temperature over 0.35 < s/C < 0.45.

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