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.

Implementation of Rectangular Vortex Generator Pairs to Improve Film Cooling Effectiveness On Transonic Rotor Blade Endwall

2021 ◽  
Author(s):  
Jinjin Li ◽  
Xin Yan ◽  
Kun He ◽  
Richard Goldstein
Keyword(s):  
Film Cooling ◽  
Vortex Generator ◽  
Attack Angle ◽  
Geometrical Parameters ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Cooling Hole ◽  
Endwall Film Cooling ◽  
Film Cooling Holes

Abstract The rectangular vortex generator pairs (RVGPs) are arranged upstream the film cooling holes to achieve a better coolant coverage on endwall near pressure-side corner area. The endwall film cooling effectiveness distributions under transonic flow conditions are numerically calculated for the single RVGP and double rows of RVGPs cases. At first, the effects of three geometrical parameters (i.e. distance between RVGP and cooling hole, height of RVGP and attack angle of RVGP) on endwall film cooling effectiveness are studied with a single hole and RVGP at different mainstream inlet Reynolds numbers and blowing ratios. Then, the double rows of RVGPs are applied to further enhance the overall film cooling effectiveness on blade endwall. The results show that the implementation of RVGPs significantly enhances the film cooling effect on transonic blade endwall at pressure-side corner area. With the increase of RVGP height, the lateral coolant coverage on endwall corner area is improved. However, by decreasing the distance between vortex generator pair and cooling hole, the film cooling effectiveness downstream of the cooling holes is increased. The attack angle of RVGP mainly affects the shape of coolant spreading on endwall surface. The RVGP with optimum dimensions and arrangement is able to suppress the coolant from lifting off the endwall and increase the coolant diffusion near endwall. Compared with no vortex generator case, the area-averaged film cooling effectiveness on endwall with double rows of RVGPs is improved by 13.16%.

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.

Influence of Coolant Density on Turbine Blade Film Cooling at Transonic Cascade Flow Conditions Using the Pressure Sensitive Paint Technique

2021 ◽  
Author(s):  
Izhar Ullah ◽  
Sulaiman M. Alsaleem ◽  
Lesley M. Wright ◽  
Chao-Cheng Shiau ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Density Ratio ◽  
Pressure Side ◽  
Pressure Sensitive Paint ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Pressure Sensitive ◽  
Blade Tip ◽  
Film Cooling Holes

Abstract This work is an experimental study of film cooling effectiveness on a blade tip in a stationary, linear cascade. The cascade is mounted in a blowdown facility with controlled inlet and exit Mach numbers of 0.29 and 0.75, respectively. The free stream turbulence intensity is measured to be 13.5 % upstream of the blade’s leading edge. A flat tip design is studied, having a tip gap of 1.6%. The blade tip is designed to have 15 shaped film cooling holes along the near-tip pressure side (PS) surface. Fifteen vertical film cooling holes are placed on the tip near the pressure side. The cooling holes are divided into a 2-zone plenum to locally maintain the desired blowing ratios based on the external pressure field. Two coolant injection scenarios are considered by injecting coolant through the tip holes only and both tip and PS surface holes together. The blowing ratio (M) and density ratio (DR) effects are studied by testing at blowing ratios of 0.5, 1.0, and 1.5 and three density ratios of 1.0, 1.5, and 2.0. Three different foreign gases are used to create density ratio effect. Over-tip flow leakage is also studied by measuring the static pressure distributions on the blade tip using the pressure sensitive paint (PSP) measurement technique. In addition, detailed film cooling effectiveness is acquired to quantify the parametric effect of blowing ratio and density ratio on a plane tip design. Increasing the blowing ratio and density ratio resulted in increased film cooling effectiveness at all injection scenarios. Injecting coolant on the PS and the tip surface also resulted in reduced leakage over the tip. The conclusions from this study will provide the gas turbine designer with additional insight on controlling different parameters and strategically placing the holes during the design process.

Bump and Trench Modifications to Film-Cooling Holes at the Vane-Endwall Junction

2008 ◽  
Vol 130 (4) ◽  
Author(s):  
N. Sundaram ◽  
K. A. Thole
Keyword(s):  
Film Cooling ◽  
Large Scale ◽  
Leading Edge ◽  
High Heat ◽  
Junction Region ◽  
High Heat Transfer ◽  
Turbine Vane ◽  
Cooling Hole ◽  
Adiabatic Effectiveness ◽  
Film Cooling Holes

The endwall of a first-stage vane experiences high heat transfer and low adiabatic effectiveness levels because of high turbine operating temperatures and formation of leading edge vortices. These vortices lift the coolant off the endwall and pull the hot mainstream gases toward it. The region of focus for this study is the vane-endwall junction region near the stagnation location where cooling is very difficult. Two different film-cooling hole modifications, namely, trenches and bumps, were evaluated to improve the cooling in the leading edge region. This study uses a large-scale turbine vane cascade with a single row of axial film-cooling holes at the leading edge of the vane endwall. Individual hole trenches and row trenches were placed along the complete row of film-cooling holes. Two-dimensional semi-elliptically shaped bumps were also evaluated by placing the bumps upstream and downstream of the film-cooling row. Tests were carried out for different trench depths and bump heights under varying blowing ratios. The results indicated that a row trench placed along the row of film-cooling holes showed a greater enhancement in adiabatic effectiveness levels when compared to individual hole trenches and bumps. All geometries considered produced an overall improvement to adiabatic effectiveness levels.

Bump and Trench Modifications to Film-Cooling Holes at the Vane Endwall Junction

2007 ◽  
Author(s):  
N. Sundaram ◽  
K. A. Thole
Keyword(s):  
Film Cooling ◽  
Large Scale ◽  
Leading Edge ◽  
High Heat ◽  
Junction Region ◽  
High Heat Transfer ◽  
Turbine Vane ◽  
Cooling Hole ◽  
Adiabatic Effectiveness ◽  
Film Cooling Holes

The endwall of a first stage vane experiences high heat transfer and low adiabatic effectiveness levels because of high turbine operating temperatures and formation of leading edge vortices. These vortices lift the coolant off the endwall and pull the hot mainstream gases towards it. The region of focus for this study is the vane-endwall junction region near the stagnation location where cooling is very difficult. Two different film-cooling hole modifications, namely trenches and bumps, were evaluated to improve the cooling in the leading edge region. This study uses a large-scale turbine vane cascade with a single row of axial film-cooling holes at the leading edge of the vane endwall. Individual hole trenches and row trenches were placed along the complete row of film-cooling holes. Two-dimensional semi-elliptically shaped bumps were also evaluated by placing the bumps upstream and downstream of the film-cooling row. Tests were carried out for different trench depths and bump heights under varying blowing ratios. The results indicated that a row trench placed along the row of film-cooling holes showed a greater enhancement in adiabatic effectiveness levels when compared to individual hole trenches and bumps. All geometries considered produced an overall improvement to adiabatic effectiveness levels.

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.

Effects of Deposits on Film Cooling of a Vane Endwall Along the Pressure Side

2008 ◽  
Vol 130 (4) ◽  
Author(s):  
N. Sundaram ◽  
M. D. Barringer ◽  
K. A. Thole
Keyword(s):  
Film Cooling ◽  
Large Scale ◽  
Secondary Flows ◽  
Pressure Side ◽  
Surface Deposition ◽  
Hot Gas ◽  
Turbine Vane ◽  
Adiabatic Effectiveness ◽  
Film Cooling Holes ◽  
Effect Of Surface

Film cooling is influenced by surface roughness and depositions that occur from contaminants present in the hot gas path, whether that film cooling occurs on the vane itself or on the endwalls associated with the vanes. Secondary flows in the endwall region also affect the film-cooling performance along the endwall. An experimental investigation was conducted to study the effect of surface deposition on film cooling along the pressure side of a first-stage turbine vane endwall. A large-scale wind tunnel with a turbine vane cascade was used to perform the experiments. The vane endwall was cooled by an array of film-cooling holes along the pressure side of the airfoil. Deposits having a semielliptical shape were placed along the pressure side to simulate individual row and multiple row depositions. Results indicated that the deposits lowered the average adiabatic effectiveness levels downstream of the film-cooling rows by deflecting the coolant jets toward the vane endwall junction on the pressure side. Results also indicated that there was a steady decrease in adiabatic effectiveness levels with a sequential increase in the number of rows with the deposits.

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.

Transonic Turbine Vane Endwall Film Cooling Using PSP Measurement Technique

2019 ◽  
Author(s):  
Chao-Cheng Shiau ◽  
Izzet Sahin ◽  
Izhar Ullah ◽  
Je-Chin Han ◽  
Alexander V. Mirzamoghadam ◽  
...  
Keyword(s):  
Shock Wave ◽  
Film Cooling ◽  
Density Ratio ◽  
Cross Flow ◽  
Flow Conditions ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Cooling Flow ◽  
Turbine Vane ◽  
Endwall Film Cooling

Abstract This work focuses on the parametric study of film cooling effectiveness on turbine vane endwall under various flow conditions. The experiments were performed in a five-vane annular sector cascade facility in a blowdown wind tunnel. The controlled exit isentropic Mach numbers were 0.7, 0.9, and 1.0, from high subsonic to transonic conditions. The freestream turbulence intensity is estimated to be 12%. Three coolant-to-mainstream mass flow ratios (MFR) in the range 0.75%, 1.0%, and 1.25% are studied. N2, CO2, and Argon/SF6 mixture were used to investigate the effects of density ratio (DR), ranging from 1.0, 1.5 to 2.0. There are 8 cylindrical holes on the endwall inside the passage. Pressure-sensitive paint (PSP) technique was used to capture the endwall pressure distribution for shock wave visualization and obtain the detailed film cooling effectiveness distributions. Both the high-fidelity effectiveness contour and the laterally (spanwise) averaged effectiveness were measured to quantify the parametric effect. This study will provide the gas turbine designer more insight on how the endwall film cooling effectiveness varies with different cooling flow conditions including shock wave through the endwall cross-flow passage.

The Effects of Axisymmetric Convergent Contouring and Blowing Ratio On Endwall Film Cooling and Vane Pressure Side Surface Phantom Cooling Performance

2021 ◽  
Author(s):  
Bo Bai ◽  
Zhigang Li ◽  
Jun Li ◽  
Shuo Mao ◽  
Wing Ng
Keyword(s):  
Numerical Method ◽  
Film Cooling ◽  
Side Surface ◽  
Operating Conditions ◽  
Constant Density ◽  
Pressure Side ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Endwall Film Cooling

Abstract In this paper, a detailed numerical investigation on the endwall film cooling and vane pressure side surface phantom cooling was performed, at the simulated realistic gas turbine operating conditions (high inlet freestream turbulence level of 16 %, exit Mach number of 0.85 and exit Reynolds number of 1.7×106). Based on a double coolant temperature model, a novel numerical method for the predictions of adiabatic wall film cooling effectiveness was proposed. This numerical method was validated by comparing the predicted results with experimental data of endwall Nusselt number, endwall film cooling effectiveness and near endwall flow visualization. The results indicate that the present numerical method can accurately predict endwall thermal load distributions and endwall film cooling distributions, and vane surface phantom cooling distributions. The endwall heat transfer coefficient, endwall film cooling effectiveness, phantom cooling effectiveness of the vane pressure side surface and total pressure loss coefficients (TPLC) were predicted and compared for two endwall contouring shapes (flat endwall and axisymmetric convergent contoured endwall) at three different blowing ratios (low blowing ratio of BR=1.0, design blowing ratio of BR=2.5 and high blowing ratio of BR=3.5) with a constant density ratio of DR=1.2, based on the present novel numerical method.

Sign in / Sign up

Export Citation Format

Share Document