Application of a CFD-Based Film Cooling Model to a Gas Turbine Vane Cascade With Cylindrical and Shaped Hole Endwall Film Cooling

2011 ◽  
Author(s):  
Tilman auf dem Kampe ◽  
Stefan Vo¨lker
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
Cooling Flows ◽  
Turbine Vane ◽  
Mesh Resolution ◽  
Shaped Hole ◽  
Node Count ◽  
Endwall Film Cooling ◽  
Cooling Model ◽  
Near Wall

This paper presents the application of a CFD-based film cooling model to a gas turbine vane cascade test rig. The experimental investigations feature aerodynamic and endwall film cooling measurements on a first stage gas turbine vane in a linear cascade. An extended version of a previously developed cylindrical hole film cooling model has been employed, which now includes modeling of shaped hole cooling flows. The computational domain extends approximately one axial chord length upstream of the leading edge and downstream of the trailing edge of the vane. Adjacent solid parts are included by means of a conjugate heat transfer analysis to account for conduction effects. A hybrid mesh with resolved boundary layers and high spatial mesh resolution in the near-wall region is being used. This meshing approach ensures that the near-wall mesh resolution requirements of the film cooling model are satisfied, while maintaining a manageable total node count. Results obtained using the film cooling model are compared to surface distributions of film cooling effectiveness from the experimental cascade. Due to the moderate node count (≈ 3.5 × 106), CFD calculations including film cooling flows can be performed at comparatively low computational cost. The film cooling model, which previously had been validated against flat plate measurement data and applied to single cooling hole configurations only, is therefore shown to be a viable tool for the thermal design of gas turbine components with film cooling.

Turbine vane endwall film cooling from mid-chord or downstream rows and upstream coolant injection

2019 ◽  
Vol 133 ◽  
pp. 247-255 ◽  
Author(s):  
Nian Wang ◽  
Chao-Cheng Shiau ◽  
Je-Chin Han ◽  
Hongzhou Xu ◽  
Michael Fox
Keyword(s):  
Film Cooling ◽  
Coolant Injection ◽  
Turbine Vane ◽  
Endwall Film Cooling

scholarly journals Study on the Turbine Vane and Blade for a 1500°C Class Industrial Gas Turbine

1993 ◽  
Author(s):  
S. Amagasa ◽  
K. Shimomura ◽  
M. Kadowaki ◽  
K. Takeishi ◽  
H. Kawai ◽  
...  
Keyword(s):  
Gas Turbine ◽  
Film Cooling ◽  
High Efficiency ◽  
Development Program ◽  
Test Results ◽  
Directionally Solidified ◽  
Barrier Coating ◽  
Turbine Vane ◽  
Nickel Base ◽  
Industrial Gas Turbine

This paper describes the summary of a three year development program for the 1st stage stationary vane and rotating blade for the next generation, 1500°C Class, high efficiency gas turbine. In such a high temperature gas turbine, the 1st turbine vane and blade are the most important hot parts. Full coverage film cooling (FCFC) is adopted for the cooling scheme, and directionally solidified (DS) nickel base super-alloy and thermal barrier coating (TBC) will be used to prolong the creep and thermal fatigue life. The concept of the cooling configuration, fundamental cascade test results and material test results will be presented.

Design of Full-Scale Endwall Film Cooling of a Turbine Vane

2019 ◽  
Vol 142 (2) ◽  
Author(s):  
Jian Liu ◽  
Wei Du ◽  
Guohua Zhang ◽  
Safeer Hussain ◽  
Lei Wang ◽  
...  
Keyword(s):  
High Temperature ◽  
Film Cooling ◽  
Pressure Coefficient ◽  
Pressure Side ◽  
Junction Region ◽  
Film Cooling Effectiveness ◽  
Coefficient Distribution ◽  
Turbine Vane ◽  
Endwall Film Cooling ◽  
Film Cooling Holes

Abstract Endwall film cooling is a significant cooling method to protect the endwall region and the junction region of endwall and a turbine vane, where usually a relatively high temperature load exists. This work aims to find the optimized arrangement of film cooling holes on the endwall and improve the film cooling in some difficult regions on the endwall, such as pressure side-endwall junction region. Several ideas for film cooling hole arrangement design are proposed, based on the pressure coefficient distribution, the streamline distribution, and the heat transfer coefficient (HTC) distribution, respectively. Four specified designs are built and compared. The results are obtained by numerical calculations with a well-validated turbulence model, the k–ω shear stress transport (SST) model. From this work, the designs based on the pressure coefficient distribution (designs 1 and 2) force the flow from the pressure side to the suction side (SS), especially in design 2, which adopts compound angle holes. The designs based on pressure coefficients have benefit in the cooling of the SS but give worse coolant coverage on the pressure side. In addition, designs 1 and 2 have little influence on the original pressure field. The design based on the streamline distributions (design 3) has larger coolant coverage on the endwall and provides good coolant coverage on the endwall and pressure side junction region. The design based on the HTC distribution provides large overall film cooling effectiveness on both the pressure side and the SS. More film cooling holes are placed on the high temperature regions, which is more effective in practice.

Turbine Vane Endwall Film Cooling Effectiveness of Different Purge Slot Configurations in a Linear Cascade

2019 ◽  
Author(s):  
Gunther Müller ◽  
Christian Landfester ◽  
Martin Böhle ◽  
Robert Krewinkel
Keyword(s):  
Film Cooling ◽  
Density Ratio ◽  
Region Of Interest ◽  
Experimental Tests ◽  
Secondary Flows ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Turbine Vane ◽  
Density Ratios ◽  
Endwall Film Cooling

Abstract This study is concerned with the film cooling effectiveness of the flow issuing from the gap between the NGV and the transition duct on the NGV endwall, i.e. the purge slot. Different slot widths, positions and injection angles were examined in order to represent changes due to thermal expansion as well as design modifications. Apart from these geometric variations, different blowing ratios (BR) and density ratios (DR) were realized to investigate the effects of the interaction between secondary flow and film cooling effectiveness. The experimental tests were performed in a linear scale-1 cascade equipped with four highly loaded turbine vanes at the Institute of Fluid Mechanics and Fluid Machinery of the University of Kaiserslautern. The mainstream flow parameters were, with a Reynolds number of 300,000 and a Mach number (outlet) of 0.6, set to meet real engine conditions. By using various flow conditioners, periodic flow was obtained in the region of interest (ROI). The adiabatic film cooling effectiveness was determined by using the Pressure Sensitive Paint (PSP) technique. In this context, nitrogen and carbon dioxide were used as tracer gases realizing two different density ratios DR = 1.0 and 1.6. The investigation was conducted for a broad range of blowing ratios with 0.25 ≤ BR ≤ 1.50. In combination with 10 geometry variations and the aforementioned blowing and density ratio variations 100 single operating points were investigated. For a better understanding of the coolant distribution, the secondary flows on the endwall were visualized by oil dye. The measurement results will be discussed based on the areal distribution of film cooling effectiveness, its lateral spanwise as well as its area average. The results will provide a better insight into various parametric effects of gap variations on turbine vane endwall film cooling performance — notably under realistic engine conditions.

Single and Multiple Row Endwall Film-Cooling of a Highly Loaded First Turbine Vane With Variation of Loading

2013 ◽  
Vol 136 (6) ◽  
Author(s):  
Martin Kunze ◽  
Konrad Vogeler ◽  
Michael Crawford ◽  
Glenn Brown
Keyword(s):  
Flow Field ◽  
Secondary Flow ◽  
Film Cooling ◽  
Wall Material ◽  
Measured Temperature ◽  
Single Row ◽  
Wall Flow ◽  
Endwall Film Cooling ◽  
Highly Loaded ◽  
Near Wall

This paper reports endwall film-cooling investigations with single and multiple rows of fan-shaped film holes using temperature-sensitive paint (TSP). The experiments are carried out in a six-bladed linear cascade based on the geometry of a highly loaded gas turbine first vane. The film effectiveness performance of the cooling rows is investigated under the influence of enhanced near-wall secondary flow. Tests are conducted at three different loading conditions changing the profile incidence. Film-cooling injection is established at elevated coolant density ratios of 1.4 using heated carbon dioxide. Due to the finite thermal conductivity of the wall material, the heat conduction effects observed in the measured temperature fields are assessed by a newly developed data analysis based on a finite element thermal analysis and tracking algorithms along CFD-computed near-wall surface streamlines. The results showed that the coolant trajectories are visibly influenced revealing the intense interaction between the film jets and the near-wall flow field. These effects are certainly enhanced with higher incidence leading to increased streamwise coolant consumption and reduced wall coverage. At the cascade inlet, the film-cooling injection is significantly affected by the near-wall flow field showing distinct over- and undercooled regions. Due to the enhanced deflection and mixing of the film jets injected from a single row, area-averaged film effectiveness and wall coverage decreases about 9 and 11%, respectively. With adding more cooling holes to this endwall area, the influence of the enhanced secondary flow becomes more pronounced. Hence, larger reduction in film effectiveness of 23% and wall coverage with 28% is observed. For single row injection at the airfoil pressure side, the stronger secondary flow motion with intensified streamwise mixing leads to a visibly decreased endwall coverage ratio of about 38% and maximum flow path reduction of about 41%. In this case, film effectiveness is found to be reduced up to 47% due to the small amount of coolant injected through this row. This effect is significantly smaller when more cooling rows are added showing an almost constant cooling performance for all incidence cases.

Deposition Near Film Cooling Holes on a High Pressure Turbine Vane

2011 ◽  
Vol 134 (4) ◽  
Author(s):  
Weiguo Ai ◽  
Nathan Murray ◽  
Thomas H. Fletcher ◽  
Spencer Harding ◽  
Scott Lewis ◽  
...  
Keyword(s):  
Film Cooling ◽  
Elevated Temperatures ◽  
Coal Ash ◽  
Test Facility ◽  
Barrier Coating ◽  
Turbine Vane ◽  
Turbine Component ◽  
Shaped Hole ◽  
Density Ratios ◽  
Film Cooling Holes

Deposition on film-cooled turbine components was studied in an accelerated test facility. The accelerated deposition facility seeds a natural-gas burning combustor with finely ground coal ash particulate at 1180°C and 180 m/s (M=0.25). Both cylindrical and shaped holes, with and without thermal barrier coating, were studied over a range of blowing ratios from 0.5 to 4.0. Coolant density ratios were maintained at values from 2.1 to 2.4. Deposition patterns generated with the cylindrical film cooling holes indicated regions of low deposition in the path of the coolant with heightened deposition between film holes. This distinctive pattern was more accentuated at higher blowing ratios. Optical temperature measurements of the turbine component surface during deposition showed elevated temperatures between coolant paths. This temperature nonuniformity became more accentuated as deposition increased, highlighting a mechanism for deposition growth that has been documented on in-service turbines as well. The shaped-hole components exhibited little or no deposition in the region just downstream of the holes due to the distributed coolant film. Close cylindrical hole spacing of 2.25d displayed similar behavior to the shaped-hole configuration.

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