The Effect of Step Misalignment on Purge Flow Cooling of Nozzle Guide Vane at Transonic Conditions

2020 ◽  
Vol 142 (10) ◽  
Author(s):  
Luke Luehr ◽  
Ridge Sibold ◽  
Shuo Mao ◽  
Wing F. Ng ◽  
Zhigang Li ◽  
...  
Keyword(s):  
Nusselt Number ◽  
Film Cooling ◽  
Guide Vane ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Purge Flow ◽  
Exit Mach Number ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes

Abstract This study describes a detailed investigation on the effects that upstream step misalignment and upstream purge film cooling have on the endwall heat transfer for first stage nozzle guide vanes (NGVs) in a gas turbine at transonic conditions. Endwall Nusselt number and adiabatic film-cooling effectiveness distributions were experimentally measured and compared with flow visualization. Tests were conducted in a transonic linear cascade blowdown facility at an inlet freestream turbulence intensity of 16%, an exit Mach number of 0.85, and an exit Re = 1.5 × 106 based on axial chord. Varied upstream purge blowing ratios (BRs) and a no-blowing case were tested for three different upstream step geometries, the baseline (no misalignment), a span-wise upstream step of +4.86% span, and a step of −4.86% span. Experimentation shows that compared with no-blowing case, the addition of upstream purge film cooling increases the Nusselt number at injection upward of 50% but lowers it in the passage throat by approximately 20%. The backward facing step induces more turbulent mixing between the coolant and mainstream flows, thus reducing film effectiveness coverage and increasing Nusselt number by nearly 40% in the passage throat. In contrast, the presence of a forward step creates a more stable boundary layer for the coolant flow aiding to help keep the film attached to the endwall. Increasing the blowing ratio increases film-cooling effectiveness and endwall coverage up to a certain point, beyond which, the high momentum of the coolant results in poor cooling performance due to jet liftoff.

The Effect of Step Misalignment on Purge Flow Cooling of Nozzle Guide Vane at Transonic Conditions

2019 ◽  
Author(s):  
Luke Luehr ◽  
Ridge Sibold ◽  
Shuo Mao ◽  
Wing F. Ng ◽  
Zhigang Li ◽  
...  
Keyword(s):  
Nusselt Number ◽  
Film Cooling ◽  
Guide Vane ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Purge Flow ◽  
Exit Mach Number ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes

Abstract This study describes a detailed investigation on the effects that step misalignment and upstream purge film cooling have on the endwall heat transfer for 1st stage nozzle guide vanes in a land-based power generating gas turbine at transonic conditions. Endwall Nusselt number and adiabatic film cooling effectiveness distributions were experimentally measured and compared with flow visualization. Tests were conducted in a transonic linear cascade blowdown facility where data were gathered at an exit Mach number of 0.85 with a freestream turbulence intensity of 16% at an exit Re = 1.5 × 106 based on axial chord. Varied upstream purge blowing ratios and a no blowing case were tested for 3 different upstream step geometries, the baseline (no misalignment), a span-wise upstream step of +4.86% span, and a step of −4.86% span. Experimentation shows that the addition of upstream purge film cooling increases the Nusselt number at injection upwards of 50% but lowers it in the throat of the passage by approximately 20% compared to no blowing case. The addition of a backward facing step induces more turbulent mixing between the coolant and mainstream flows, thus reducing film effectiveness coverage and increasing Nusselt number by nearly 40% in the passage throat. In contrast, the presence of a forward step creates a more stable boundary layer for the coolant flow aiding to help keep the film attached to the endwall at higher blowing ratios. Increasing the blowing ratio increases film cooling effectiveness and endwall coverage up to a certain point, beyond which, the high momentum of the coolant results in poor cooling performance due to jet liftoff.

Influence of Coolant Density on Turbine Platform Film-Cooling With Stator–Rotor Purge Flow and Compound-Angle Holes

2014 ◽  
Vol 6 (4) ◽  
Author(s):  
Kevin Liu ◽  
Shang-Feng Yang ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Density Ratio ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Fundamental Parameters ◽  
Blowing Ratio ◽  
Purge Flow ◽  
Cylindrical Holes ◽  
Exit Mach Number ◽  
Compound Angle

A detailed parametric study of film-cooling effectiveness was carried out on a turbine blade platform. The platform was cooled by purge flow from a simulated stator–rotor seal combined with discrete hole film-cooling. The cylindrical holes and laidback fan-shaped holes were accessed in terms of film-cooling effectiveness. This paper focuses on the effect of coolant-to-mainstream density ratio on platform film-cooling (DR = 1 to 2). Other fundamental parameters were also examined in this study—a fixed purge flow of 0.5%, three discrete-hole film-cooling blowing ratios between 1.0 and 2.0, and two freestream turbulence intensities of 4.2% and 10.5%. Experiments were done in a five-blade linear cascade with inlet and exit Mach number of 0.27 and 0.44, respectively. Reynolds number of the mainstream flow was 750,000 and was based on the exit velocity and chord length of the blade. The measurement technique adopted was the conduction-free pressure sensitive paint (PSP) technique. Results indicated that with the same density ratio, shaped holes present higher film-cooling effectiveness and wider film coverage than the cylindrical holes, particularly at higher blowing ratios. The optimum blowing ratio of 1.5 exists for the cylindrical holes, whereas the effectiveness for the shaped holes increases with an increase of blowing ratio. Results also indicate that the platform film-cooling effectiveness increases with density ratio but decreases with turbulence intensity.

Application of Unsteady CFD Methods to Trailing Edge Cutback Film Cooling

2014 ◽  
Author(s):  
S. Ravelli ◽  
G. Barigozzi
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Navier Stokes ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Nozzle Guide Vane ◽  
Unsteady Rans ◽  
Exit Mach Number ◽  
Nozzle Guide

The main purpose of this numerical investigation is to overcome the limitations of the steady modeling in predicting the cooling efficiency over the cutback surface in a high pressure turbine nozzle guide vane. Since discrepancy between Reynolds-averaged Navier–Stokes (RANS) predictions and measured thermal coverage at the trailing edge was attributable to unsteadiness, Unsteady RANS (URANS) modeling was implemented to evaluate improvements in simulating the mixing between the mainstream and the coolant exiting the cutback slot. With the aim of reducing the computation effort, only a portion of the airfoil along the span was simulated at an exit Mach number of Ma2is = 0.2. Three values of the coolant-to-mainstream mass flow ratio were considered: MFR = 0.66%, 1.05%, and 1.44%. Nevertheless the inherent vortex shedding from the cutback lip was somehow captured by the URANS method, the computed mixing was not enough to reproduce the measured drop in adiabatic effectiveness η along the streamwise direction, over the cutback surface. So modeling was taken a step further by using the Scale Adaptive Simulation (SAS) method at MFR = 1.05%. Results from the SAS approach were found to have potential to mimic the experimental measurements. Vortices shedding from the cutback lip were well predicted in shape and magnitude, but with a lower frequency, as compared to PIV data and flow visualizations. Moreover, the simulated reduction in film cooling effectiveness toward the trailing edge was similar to that observed experimentally.

Endwall Film Cooling Using the Staggered Combustor-Turbine Gap Leakage Flow

2014 ◽  
Author(s):  
Yang Zhang ◽  
Xin Yuan
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Inlet Temperature ◽  
Local Pressure ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Blind Area ◽  
Nozzle Guide Vane ◽  
Nozzle Guide

A key technology of gas turbine performance improvement was the increase in the turbine inlet temperature, which brought high thermal loads to the Nozzle Guide Vane (NGV) components. Strong pressure gradients in the NGVs and the complex secondary flow field had made thermal protection more challenging. As for the endwall surface near the pressure side gill region, the relatively higher local pressure and cross flow apparently decreased the film-cooling effectiveness. The aim of this investigation was to evaluate a new design, improving the film-cooling performance in a cooling blind area with upstream staggered slot, simulating the combustor-turbine leakage gap flow. The test cascades model was manufactured according to the GE-E3 nozzle guide vane scaled model, with a scale ratio of 2.2. The experiment was performed under the inlet Mach number 0.1 and the Reynolds number 3.5×105 based on an axial chord length of 78 mm. The staggered slots were positioned upstream of the cascades to simulate the combustor-turbine gap leakage. The Pressure Sensitive Painting (PSP) technique was used to detect the film cooling effectiveness distribution on the endwall surface. Through the investigation, the following results could be achieved: 1) the film-cooling effectiveness on the endwall surface downstream the slot and along the pitchwise direction increased, with the highest parameter at Z/Pitch = 0.6; 2) a larger cooled region developed towards the suction side as the blowing ratio increased; 3) the advantage of the staggered slot was apparent on the endwall surface near the inlet area, while the coolant film was obviously weakened along the axial chord at a low blowing ratio. The influence of the staggered slots could only be detected in the downstream area of the endwall surface at the higher blowing ratio.

Effects of an Axisymmetric Contoured Endwall on a Nozzle Guide Vane: Adiabatic Effectiveness Measurements

2010 ◽  
Author(s):  
A. A. Thrift ◽  
K. A. Thole ◽  
S. Hada
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Cooling Effectiveness ◽  
Flow Rates ◽  
Guide Vanes ◽  
Film Cooling Effectiveness ◽  
Adiabatic Effectiveness ◽  
Improved Performance ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes

Gas turbine designs seek improved performance by modifying the endwalls of nozzle guide vanes in the engine hot section. Within the nozzle guide vanes these modifications can be in the form of an axisymmetric contour as the area contracts from the combustor to the turbine. This paper investigates the effect of axisymmetric endwall contouring on the cooling performance of a film cooled endwall. Adiabatic effectiveness measurements were performed in a planar passage for comparison to a contoured passage whereby the exit Reynolds numbers was matched. For the contoured passage, measurements were performed on both the flat endwall and on the contoured endwall. Fully expanded film cooling holes were distributed on the endwall surface preceded by a two-dimensional slot normal to the inlet axis. Results indicated that the coolant coverage from the upstream leakage slot was spread over a larger area of the contoured endwall in comparison to the flat endwall of the planar passage. Film cooling effectiveness on the flat endwall of the contoured passage showed minimal differences relative to the planar passage results. The contracting endwall of the contoured passage, however, showed a significant reduction with average film cooling effectiveness levels approximately 40% lower than the planar passage at low film cooling flow rates. In the case of all endwalls, increasing leakage and film cooling mass flow rates led to an increase in cooling effectiveness and coolant coverage.

Effects of an Axisymmetric Contoured Endwall on a Nozzle Guide Vane: Adiabatic Effectiveness Measurements

2011 ◽  
Vol 133 (4) ◽  
Author(s):  
A. A. Thrift ◽  
K. A. Thole ◽  
S. Hada
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Cooling Effectiveness ◽  
Flow Rates ◽  
Guide Vanes ◽  
Film Cooling Effectiveness ◽  
Adiabatic Effectiveness ◽  
Improved Performance ◽  
Nozzle Guide ◽  
Nozzle Guide Vanes

Gas turbine designs seek improved performance by modifying the endwalls of nozzle guide vanes in the engine hot section. Within the nozzle guide vanes, these modifications can be in the form of an axisymmetric contour as the area contracts from the combustor to the turbine. This paper investigates the effect of axisymmetric endwall contouring on the cooling performance of a film cooled endwall. Adiabatic effectiveness measurements were performed in a planar passage for comparison to a contoured passage, whereby the exit Reynolds numbers were matched. For the contoured passage, measurements were performed both on the flat endwall and on the contoured endwall. Fully expanded film cooling holes were distributed on the endwall surface preceded by a two-dimensional slot normal to the inlet axis. Results indicated that the coolant coverage from the upstream leakage slot was spread over a larger area of the contoured endwall in comparison to the flat endwall of the planar passage. Film cooling effectiveness on the flat endwall of the contoured passage showed minimal differences relative to the planar passage results. The contracting endwall of the contoured passage, however, showed a significant reduction with average film cooling effectiveness levels approximately 40% lower than the planar passage at low film cooling flow rates. In the case of all endwalls, increasing leakage and film cooling mass flow rates led to an increase in cooling effectiveness and coolant coverage.

Influence of Coolant Density on Turbine Platform Film-Cooling With Stator-Rotor Purge Flow and Compound-Angle Holes

2013 ◽  
Author(s):  
Kevin Liu ◽  
Shang-Feng Yang ◽  
Je-Chin Han
Keyword(s):  
Film Cooling ◽  
Density Ratio ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Fundamental Parameters ◽  
Blowing Ratio ◽  
Purge Flow ◽  
Cylindrical Holes ◽  
Exit Mach Number ◽  
Compound Angle

A detailed parametric study of film-cooling effectiveness was carried out on a turbine blade platform. The platform was cooled by purge flow from a simulated stator-rotor seal combined with discrete hole film-cooling. The cylindrical holes and laidback fan-shaped holes were accessed in terms of film-cooling effectiveness. This paper focuses on the effect of coolant-to-mainstream density ratio on platform film-cooling (DR = 1 to 2). Other fundamental parameters were also examined in this study — a fixed purge flow of 0.5%, three discrete-hole film-cooling blowing ratios between 1.0 and 2.0, and two freestream turbulence intensities of 4.2% and 10.5%. Experiments were done in a five-blade linear cascade with inlet and exit Mach number of 0.27 and 0.44, respectively. Reynolds number of the mainstream flow was 750,000 and wad based on the exit velocity and chord length of the blade. The measurement technique adopted was conduction-free pressure sensitive paint (PSP) technique. Results indicated that with the same density ratio, shaped holes present higher film-cooling effectiveness and wider film coverage than the cylindrical holes, particularly at higher blowing ratios. The optimum blowing ratio of 1.5 exists for the cylindrical holes, whereas the effectiveness for the shaped holes increases with increase of blowing ratio. Results also show that the platform film-cooling effectiveness increases with density ratio but decreases with turbulence intensity.

Application of Unsteady Computational Fluid Dynamics Methods to Trailing Edge Cutback Film Cooling

2014 ◽  
Vol 136 (12) ◽  
Author(s):  
Silvia Ravelli ◽  
Giovanna Barigozzi
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Navier Stokes ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Image Velocimetry ◽  
Unsteady Rans ◽  
Exit Mach Number ◽  
Nozzle Guide

The main purpose of this numerical investigation is to overcome the limitations of the steady modeling in predicting the cooling efficiency over the cutback surface in a high pressure turbine nozzle guide vane. Since discrepancy between Reynolds-averaged Navier–Stokes (RANS) predictions and measured thermal coverage at the trailing edge was attributable to unsteadiness, Unsteady RANS (URANS) modeling was implemented to evaluate improvements in simulating the mixing between the mainstream and the coolant exiting the cutback slot. With the aim of reducing the computation effort, only a portion of the airfoil along the span was simulated at an exit Mach number of Ma2is = 0.2. Three values of the coolant-to-mainstream mass flow ratio were considered: MFR = 0.66%, 1.05%, and 1.44%. Nevertheless the inherent vortex shedding from the cutback lip was somehow captured by the URANS method, the computed mixing was not enough to reproduce the measured drop in adiabatic effectiveness η along the streamwise direction, over the cutback surface. So modeling was taken a step further by using the scale adaptive simulation (SAS) method at MFR = 1.05%. Results from the SAS approach were found to have potential to mimic the experimental measurements. Vortices shedding from the cutback lip were well predicted in shape and magnitude, but with a lower frequency, as compared to particle image velocimetry (PIV) data and flow visualizations. Moreover, the simulated reduction in film cooling effectiveness toward the trailing edge was similar to that observed experimentally.

Turbine Endwall Film Cooling With Pressure Side Radial Holes

2013 ◽  
Author(s):  
Yang Zhang ◽  
Xin Yuan
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Pressure Side ◽  
Local Pressure ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Blowing Ratio ◽  
Nozzle Guide Vane ◽  
Cylindrical Holes ◽  
Nozzle Guide

A key technology of gas turbine performance improvement was the increase in the turbine inlet temperature, which brought high thermal loads to the nozzle guide vane (NGV) components. Strong pressure gradients in the NGVs and the complex secondary flow field had made thermal protection more challenging. As for the endwall surface near the side gill pressure region, the relatively higher local pressure and cross flow apparently decreased the film-cooling effectiveness. The aim of this investigation was to evaluate a new design, improving the film-cooling performance in a cooling blind area with radial cylindrical holes on the pressure side. The test cascades model was manufactured according to the GE-E3 nozzle guide vane scaled model, with a scale ratio of 2.2. The experiment was performed under the inlet Mach number 0.1 and the Reynolds number 3.5×105 based on an axial chord length of 78 mm. Four rows of staggered radial film-cooling holes were placed at the pressure side gill region. The diameter of the cylindrical holes was 1 mm and the length was 5 d, with a hole space of 6 d. The spanwise angle of the cooling holes was 35 ° and the radial angle was 90 °. Three blowing ratios were chosen as the test conditions in the experiment, M = 0.7, M = 1.0 and M = 1.3. The film-cooling effectiveness was probed using PSP (pressure sensitive painting) technology and the post processing was performed by means of a mass and heat transfer analogy. Through the investigation, the following results could be achieved: 1) the film-cooling effectiveness on the endwall surface near the pressure side gill region increased, with the highest parameter at X/Cax = 0.3; 2) a double-peak cooled region developed towards the suction side as the blowing ratio increased; 3) the advantage of the pressure side radial cooling holes was apparent on the endwall surface near the gill region, while the coolant film was obviously weakened along the axial chord at a low blowing ratio. The influence of the pressure film cooling could only be detected in the downstream area of the endwall at the higher blowing ratio.

Pressure Side and Cutback Trailing Edge Film Cooling in a Linear Nozzle Vane Cascade at Different Mach Numbers

2012 ◽  
Vol 134 (5) ◽  
Author(s):  
Giovanna Barigozzi ◽  
Antonio Perdichizzi ◽  
Silvia Ravelli
Keyword(s):  
Film Cooling ◽  
Guide Vane ◽  
Pressure Side ◽  
Trailing Edge ◽  
Cooling Effectiveness ◽  
Film Cooling Effectiveness ◽  
Mach Numbers ◽  
Nozzle Guide ◽  
Cooling Air ◽  
Momentum Boundary Layer

Tests on a specifically designed linear nozzle guide vane cascade with trailing edge coolant ejection were carried out to investigate the influence of trailing edge bleeding on both aerodynamic and thermal performance. The cascade is composed of six vanes with a profile typical of a high pressure turbine stage. The trailing edge cooling features a pressure side cutback with film cooling slots, stiffened by evenly spaced ribs in an inline configuration. Cooling air is ejected not only through the slots but also through two rows of cooling holes placed on the pressure side, upstream of the cutback. The cascade was tested for different isentropic exit Mach numbers, ranging from M2is = 0.2 to M2is = 0.6, while varying the coolant to mainstream mass flow ratio MFR up to 2.8%. The momentum boundary layer behavior at a location close to the trailing edge, on the pressure side, was assessed by means of laser Doppler measurements. Cases with and without coolant ejection allowed us to identify the contribution of the coolant to the off the wall velocity profile. Thermochromic liquid crystals (TLC) were used to map the adiabatic film cooling effectiveness on the pressure side cooled region. As expected, the cutback effect on cooling effectiveness, compared to the other cooling rows, was dominant.

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