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David G. Bogard - One of the best experts on this subject based on the ideXlab platform.
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Evaluation of Film Cooling Superposition Predictions Using Shaped Holes on the Suction Side of a Blade Model
Volume 5B: Heat Transfer, 2019Co-Authors: Christopher Yoon, Jacob J. Moore, David G. BogardAbstract:Abstract Film cooling is often used for turbine airfoil cooling, and there are numerous studies of the performance of a single row of holes. In actual application there will typically be multiple rows of holes which interact. Consequently there is a need to develop techniques to predict film cooling performance with multiple rows of coolant holes using superposition of single row cooling effectiveness. Although there have been many studies of superposition techniques for predicting film cooling effectiveness with multiple rows of cylindrical holes, there have been very few in which shaped holes were used with a typical turbine airfoil model. In this study, film effectiveness was measured on the Suction Side of a turbine blade model using two rows of shaped coolant holes. Measurements were made with each row independently and with both rows combined. This provided the experimental data for superposition predictions and to evaluate these predictions. Each row had 7-7-7 shaped holes with pitch to diameter ratio of 6, and the two rows were more than 40 diameters apart. The experiments were run using two different upstream blowing ratios, and a wide range of downstream blowing ratios. The superposition predictions of film effectiveness were reasonably accurate when the upstream row of holes were operated at a high blowing ratio with a corresponding smaller film effectiveness (due to jet separation). However, when the upstream coolant holes were operated at the optimum blowing ratio, and hence maximum film effectiveness downstream, the superposition analysis predicted film effectiveness levels slightly lower than actual levels. These results show that there was an interaction between jets that resulted in higher film effectiveness than was accounted for with a superposition prediction.
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Sensitivity of the Overall Effectiveness to Film Cooling and Internal Cooling on a Turbine Vane Suction Side
Journal of Turbomachinery, 2013Co-Authors: Randall P. Williams, David G. Bogard, Thomas E Dyson, Sean D BradshawAbstract:The overall cooling effectiveness for a turbine airfoil was quantified based on the external surface temperature relative to the mainstream temperature and the inlet coolant temperature. This can be determined experimentally when the model is constructed so that the Biot number is similar to that of engine components. In this study, the overall cooling effectiveness was experimentally measured on a model turbine vane constructed of a material deigned to match Bi for engine conditions. The model incorporated an internal impingement cooling configuration. Overall cooling effectiveness and adiabatic film effectiveness were measured downstream of a single row of round holes positioned on the Suction Side of the vane. Experiments were conducted to evaluate the cooling effects of internal cooling alone, and then the combined effects of film cooling and internal cooling for a range of coolant flow rates. While the adiabatic film effectiveness decreased when using high momentum flux ratios for the film cooling, due to coolant jet separation, the overall cooling effectiveness increased at higher momentum flux ratios. This increase was due to increased internal cooling effects. Overall cooling effectiveness measurements were also compared to analytical predictions based on a 1D thermal analysis using measured adiabatic film effectiveness and overall cooling effectiveness without film cooling.
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mean and turbulent velocity profile measurements on the Suction Side of a film cooled turbine vane
ASME Turbo Expo 2013: Turbine Technical Conference and Exposition GT 2013, 2013Co-Authors: John W Mcclintic, David G. Bogard, Thomas E Dyson, Sean D BradshawAbstract:Boundary layer velocity and turbulence profiles were measured on the Suction Side of a scaled up, film-cooled turbine vane airfoil. There have been a number of previous studies of the velocity profile on a turbine vane, but few have taken velocity profile data with film cooling, and none have taken such data on the Suction Side of the vane. Velocity and turbulence profile data were taken at two locations on the Suction Side of the vane — one at a high curvature region and one further downstream in a low curvature region. Data were collected for high (20%) and low (0.5%) mainstream turbulence conditions. For the upstream, high curvature location, velocity and turbulence profiles were found with and without the showerhead blowing and within and outSide of the merged showerhead coolant jet. The data for the low curvature, downstream location was taken with injection from the showerhead alone, a second upstream row of holes alone, and the combination of the two cases. It was found that the presence of an active upstream row of holes thickens the boundary layer and increases urms both within and beyond the extent of the boundary layer. Span-wise variations showed that these effects are strongest within the core of the coolant jets. At the downstream location, the boundary layer velocity profile was most strongly influenced by the row of holes immediately upstream of that location. Finally, turbulence integral length scale data showed the effect of large scale mainstream turbulence penetrating the boundary layer. The increase in turbulence, thickening of the boundary layer, and large scale turbulence all play important roles in row to row coolant interactions and affect the film cooling effectiveness.Copyright © 2013 by ASME
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Sensitivity of the overall effectiveness to film cooling and internal cooling on a turbine vane Suction Side
Volume 4: Heat Transfer Parts A and B, 2012Co-Authors: Randall P. Williams, David G. Bogard, Thomas E Dyson, Sean D BradshawAbstract:The overall cooling effectiveness for a turbine airfoil was quantified based on the external surface temperature relative to the mainstream temperature and the inlet coolant temperature. This can be determined experimentally when the model is constructed so that the Biot number is similar to that of engine components. In this study, the overall cooling effectiveness was experimentally measured on a model turbine vane constructed of a material deigned to match Bi for engine conditions. The model incorporated an internal impingement cooling configuration. Overall cooling effectiveness and adiabatic film effectiveness were measured downstream of a single row of round holes positioned on the Suction Side of the vane. Experiments were conducted to evaluate the cooling effects of internal cooling alone, and then the combined effects of film cooling and internal cooling for a range of coolant flow rates. While the adiabatic film effectiveness decreased when using high momentum flux ratios for the film cooling, due to coolant jet separation, the overall cooling effectiveness increased at higher momentum flux ratios. This increase was due to increased internal cooling effects. Overall cooling effectiveness measurements were also compared to analytical predictions based on a 1D thermal analysis using measured adiabatic film effectiveness and overall cooling effectiveness without film cooling.Copyright © 2012 by ASME
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Heat Transfer Augmentation Downstream of Rows of Various Dimple Geometries on the Suction Side of a Gas Turbine Airfoil
Journal of Turbomachinery, 2010Co-Authors: Jason E. Dees, David G. Bogard, Ronald Scott BunkerAbstract:ABSTRACT Heat transfer coefficients were measured downstream of a row of shaped film cooling holes as well as elliptical, diffuser, and teardrop shaped dimples simulating depressions due to film coolant holes of different shapes. These features were placed on the Suction Side of a simulated gas turbine vane. The dimples were used as approximations to film cooling holes after the heat transfer levels downstream of active fan shaped film cooling holes was found to be independent of film cooling. The effects of the dimples were tested with varying approach boundary layers, freestream turbulence intensity, and Reynolds numbers. For the case of an untripped (transitional) approach boundary layer, all dimple shapes caused approximately a factor of two increase in heat transfer coefficient relative to the smooth baseline condition due to the dimples effectively causing boundary layer transition downstream. The exact augmentation varied depending on the dimple geometry: diffuser shapes causing the largest augmentation and teardrop shapes causing the lowest augmentation. For tripped (turbulent boundary layer) approach conditions, the dimple shapes all caused the same 20% augmentation relative to the smooth tripped baseline. The already turbulent nature of the tripped approach flow reduces the effect that the dimples have on the downstream heat transfer coefficient.
Bengt Sunden - One of the best experts on this subject based on the ideXlab platform.
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effects of the cooling configurations layout near the first stage vane leading edge on the endwall cooling and phantom cooling of the vane Suction Side surface
International Journal of Heat and Mass Transfer, 2018Co-Authors: Bengt SundenAbstract:Abstract Increasing the turbine inlet temperature can enhance the thermal efficiency of a gas turbine. Therefore, modern gas turbines operate at a relatively high level of temperature and endure heavy thermal load. It is important to ensure the modern gas turbine works at a high performance and safe condition. Advanced cooling techniques are implemented in the gas turbine system. In the current study, effects of the cooling configurations layout near the first-stage vane leading edge on the endwall cooling and phantom cooling of the vane Suction Side surface were numerically investigated. Three-dimensional (3D) Reynolds-averaged Navier-Stokes (RANS) equations combined with the shear stress transport (SST) k - ω turbulence model were solved to perform the simulations on basis of validation by comparing the experimental data and computational results. The results indicate that the layout of the cooling configurations has a significant influence on the endwall cooling, but a limited effect on the phantom cooling of the Suction Side surface and the aerodynamic performance. For each type, the endwall cooling and phantom cooling of the Suction Side surface are enhanced with the increase of the blowing ratio ( M ) of the leading edge coolant injection. Meanwhile, the thermodynamic loss is gradually enhanced. Overall, the Type B which has a partly blocked upstream slot achieves the best performance in terms of the coolant mass flowrate, endwall cooling, phantom cooling performance of the Suction Side surface and the aerodynamic performance at M = 1.0 .
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Effects of the swirling coolant jet from the upstream slot on the vane endwall cooling and the vane Suction Side phantom cooling
International Journal of Heat and Mass Transfer, 2018Co-Authors: Bengt SundenAbstract:Abstract In order to obtain a higher thermal efficiency, the inlet temperature of gas turbines is gradually increased. However, this increases the thermal load on the first stage vane endwall surface. Therefore, an advanced cooling technique must be implemented for the endwall cooling to ensure that the gas turbine operates safely. In the current study, effects of swirling coolant flow from the upstream slot on the endwall cooling and vane Suction Side surface phantom cooling were numerically investigated on basis of a validated numerical method. Three-dimensional (3D) Reynolds-averaged Navier–Stokes (RANS) equations combined with the shear stress transport (SST) k - ω turbulence model were solved in the numerical simulations. The results indicate that the endwall cooling and the phantom cooling are significantly influenced by introducing a swirling coolant jet from the upstream slot relative to the baseline. Compared with the baseline case and the negative swirling coolant jet angle, the positive swirling coolant jet angle contributes to increase the overall uniformity of the endwall cooling effectiveness and reduce the hot region along the pressure Side, especially for the case with α = 20°. However, it obtains a relatively low level of the phantom cooling effectiveness on the vane Suction Side. In contrast, the negative swirling coolant jet angle attains a higher level of phantom cooling effectiveness on the vane Suction Side relative to the positive swirling coolant jet angle. The case with α = −20° obtains the largest phantom cooling effectiveness on the vane Suction Side. In addition, the aerodynamic loss is increased within a small range. The largest total pressure loss coefficient is 5% for the case with α = −30° among all investigated cases.
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effects of the mainstream turbulence intensity and slot injection angle on the endwall cooling and phantom cooling of the vane Suction Side surface
International Journal of Heat and Mass Transfer, 2017Co-Authors: Liming Song, Bengt SundenAbstract:In order to obtain better performance, gas turbines always operate with high inlet temperature. This contributes to a high level of thermal load on the first stage vane endwall. To ensure safe operation of a gas turbine within a proper temperature range, the cooling performance of the vane endwall must be further investigated. In the present study, effects of the mainstream turbulence and upstream coolant flow direction on the endwall cooling and vane Suction Side surface phantom cooling were numerically investigated. Three-dimensional (3D) Reynolds-averaged Navier–Stokes (RANS) equations combined with shear stress transport (SST) k-ω turbulence model were solved to conduct the numerical simulations on basis of the validated turbulence model. The calculated results indicate that both the adiabatic cooling effectiveness on the endwall and the phantom cooling effectiveness on the vane Suction Side surface are significantly influenced by slot injection angle. For α=-30°, the coolant injection is driven towards the vane Suction Side, which contributes to the lowest adiabatic cooling effectiveness level on the pressure Side endwall and the highest phantom cooling effectiveness level on the vane Suction Side surface. With the increase of the slot injection angle, the adiabatic cooling effectiveness level on the pressure Side endwall is enhanced significantly. In contrast, the phantom cooling (when the vane Suction Side is cooled by coolant originating from the endwall) of the vane Suction Side surface is reduced significantly. This is because a large slot injection angle leads to a large coolant momentum towards the pressure Side. Moreover, the case with a smaller slot injection angle obtains a slightly higher area-averaged adiabatic cooling effectiveness level around the leading edge due to a relatively larger portion of coolant being confined near the leading edge. In addition, the inlet turbulence intensity has a small impact on the overall endwall cooling and the phantom cooling of the vane Suction Side surface compared to the slot injection angle.
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influence of the upstream slot geometry on the endwall cooling and phantom cooling of vane Suction Side surface
Applied Thermal Engineering, 2017Co-Authors: Bengt SundenAbstract:Modern gas turbines always operate at a high level of inlet temperature. The current inlet temperature in the aircraft and heavy duty gas turbines is higher than the melting point of the guide vane material. Consequently, advanced cooling schemes must be developed to ensure the safe operation of gas turbines. In the current study, numerical simulations were conducted to investigate the influence of the upstream slot geometry on the endwall cooling and phantom cooling of the vane Suction Side surface. Three-dimensional (3D) Reynolds-averaged Navier–Stokes (RANS) equations combined with the shear stress transport (SST) k-ω turbulence model were solved to conduct the simulations based on the validated turbulence model. The results indicate that the adiabatic cooling effectiveness in the upstream region of the stagnation is significantly increased by introducing the contoured upstream slot. However, the normal upstream slot obtains a relatively high adiabatic cooling effectiveness level in the downstream region of the stagnation. In the present research, the case with normalized amplitude A‾=0.75, initial phase angle φ=45° achieves the largest overall adiabatic cooling effectiveness near the vane leading edge. In contrast, the case with A‾=0.75, φ=30° attains the smallest overall adiabatic cooling effectiveness on the endwall surface. Moreover, the phantom cooling effectiveness on the vane Suction Side surface is relatively small relative to the adiabatic cooling effectiveness on the endwall. The case with the normal upstream slot achieves the largest phantom cooling effectiveness on the vane Suction Side surface compared with the contoured upstream slot. Overall, the contoured upstream slot significantly enhances the endwall cooling effectiveness by rearranging the distribution of the coolant mass flowrate at the slot outlet.
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effects of the layout of film holes near the vane leading edge on the endwall cooling and phantom cooling of the vane Suction Side surface
Numerical Heat Transfer Part A-applications, 2017Co-Authors: Liming Song, Bengt SundenAbstract:In the current research, effects of the layout of film holes near the first-stage vane leading edge on the endwall cooling and phantom cooling of the vane Suction Side surface were numerically studied. The computational results indicate that the case with a positive film-hole angle achieves a higher cooling effectiveness level on the endwall and vane Suction Side surface compared to the case with a corresponding negative film-hole angle. Furthermore, the location of the film hole has a significant influence on the cooling performance of the endwall and vane Suction Side surface. In addition, the case with a smaller distance from film holes to the vane stagnation also attains a slightly higher cooling effectiveness (phantom cooling effectiveness) on the vane Suction Side surface.
Liming Song - One of the best experts on this subject based on the ideXlab platform.
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effects of the mainstream turbulence intensity and slot injection angle on the endwall cooling and phantom cooling of the vane Suction Side surface
International Journal of Heat and Mass Transfer, 2017Co-Authors: Liming Song, Bengt SundenAbstract:In order to obtain better performance, gas turbines always operate with high inlet temperature. This contributes to a high level of thermal load on the first stage vane endwall. To ensure safe operation of a gas turbine within a proper temperature range, the cooling performance of the vane endwall must be further investigated. In the present study, effects of the mainstream turbulence and upstream coolant flow direction on the endwall cooling and vane Suction Side surface phantom cooling were numerically investigated. Three-dimensional (3D) Reynolds-averaged Navier–Stokes (RANS) equations combined with shear stress transport (SST) k-ω turbulence model were solved to conduct the numerical simulations on basis of the validated turbulence model. The calculated results indicate that both the adiabatic cooling effectiveness on the endwall and the phantom cooling effectiveness on the vane Suction Side surface are significantly influenced by slot injection angle. For α=-30°, the coolant injection is driven towards the vane Suction Side, which contributes to the lowest adiabatic cooling effectiveness level on the pressure Side endwall and the highest phantom cooling effectiveness level on the vane Suction Side surface. With the increase of the slot injection angle, the adiabatic cooling effectiveness level on the pressure Side endwall is enhanced significantly. In contrast, the phantom cooling (when the vane Suction Side is cooled by coolant originating from the endwall) of the vane Suction Side surface is reduced significantly. This is because a large slot injection angle leads to a large coolant momentum towards the pressure Side. Moreover, the case with a smaller slot injection angle obtains a slightly higher area-averaged adiabatic cooling effectiveness level around the leading edge due to a relatively larger portion of coolant being confined near the leading edge. In addition, the inlet turbulence intensity has a small impact on the overall endwall cooling and the phantom cooling of the vane Suction Side surface compared to the slot injection angle.
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effects of the layout of film holes near the vane leading edge on the endwall cooling and phantom cooling of the vane Suction Side surface
Numerical Heat Transfer Part A-applications, 2017Co-Authors: Liming Song, Bengt SundenAbstract:In the current research, effects of the layout of film holes near the first-stage vane leading edge on the endwall cooling and phantom cooling of the vane Suction Side surface were numerically studied. The computational results indicate that the case with a positive film-hole angle achieves a higher cooling effectiveness level on the endwall and vane Suction Side surface compared to the case with a corresponding negative film-hole angle. Furthermore, the location of the film hole has a significant influence on the cooling performance of the endwall and vane Suction Side surface. In addition, the case with a smaller distance from film holes to the vane stagnation also attains a slightly higher cooling effectiveness (phantom cooling effectiveness) on the vane Suction Side surface.
Sean D Bradshaw - One of the best experts on this subject based on the ideXlab platform.
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Sensitivity of the Overall Effectiveness to Film Cooling and Internal Cooling on a Turbine Vane Suction Side
Journal of Turbomachinery, 2013Co-Authors: Randall P. Williams, David G. Bogard, Thomas E Dyson, Sean D BradshawAbstract:The overall cooling effectiveness for a turbine airfoil was quantified based on the external surface temperature relative to the mainstream temperature and the inlet coolant temperature. This can be determined experimentally when the model is constructed so that the Biot number is similar to that of engine components. In this study, the overall cooling effectiveness was experimentally measured on a model turbine vane constructed of a material deigned to match Bi for engine conditions. The model incorporated an internal impingement cooling configuration. Overall cooling effectiveness and adiabatic film effectiveness were measured downstream of a single row of round holes positioned on the Suction Side of the vane. Experiments were conducted to evaluate the cooling effects of internal cooling alone, and then the combined effects of film cooling and internal cooling for a range of coolant flow rates. While the adiabatic film effectiveness decreased when using high momentum flux ratios for the film cooling, due to coolant jet separation, the overall cooling effectiveness increased at higher momentum flux ratios. This increase was due to increased internal cooling effects. Overall cooling effectiveness measurements were also compared to analytical predictions based on a 1D thermal analysis using measured adiabatic film effectiveness and overall cooling effectiveness without film cooling.
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mean and turbulent velocity profile measurements on the Suction Side of a film cooled turbine vane
ASME Turbo Expo 2013: Turbine Technical Conference and Exposition GT 2013, 2013Co-Authors: John W Mcclintic, David G. Bogard, Thomas E Dyson, Sean D BradshawAbstract:Boundary layer velocity and turbulence profiles were measured on the Suction Side of a scaled up, film-cooled turbine vane airfoil. There have been a number of previous studies of the velocity profile on a turbine vane, but few have taken velocity profile data with film cooling, and none have taken such data on the Suction Side of the vane. Velocity and turbulence profile data were taken at two locations on the Suction Side of the vane — one at a high curvature region and one further downstream in a low curvature region. Data were collected for high (20%) and low (0.5%) mainstream turbulence conditions. For the upstream, high curvature location, velocity and turbulence profiles were found with and without the showerhead blowing and within and outSide of the merged showerhead coolant jet. The data for the low curvature, downstream location was taken with injection from the showerhead alone, a second upstream row of holes alone, and the combination of the two cases. It was found that the presence of an active upstream row of holes thickens the boundary layer and increases urms both within and beyond the extent of the boundary layer. Span-wise variations showed that these effects are strongest within the core of the coolant jets. At the downstream location, the boundary layer velocity profile was most strongly influenced by the row of holes immediately upstream of that location. Finally, turbulence integral length scale data showed the effect of large scale mainstream turbulence penetrating the boundary layer. The increase in turbulence, thickening of the boundary layer, and large scale turbulence all play important roles in row to row coolant interactions and affect the film cooling effectiveness.Copyright © 2013 by ASME
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Sensitivity of the overall effectiveness to film cooling and internal cooling on a turbine vane Suction Side
Volume 4: Heat Transfer Parts A and B, 2012Co-Authors: Randall P. Williams, David G. Bogard, Thomas E Dyson, Sean D BradshawAbstract:The overall cooling effectiveness for a turbine airfoil was quantified based on the external surface temperature relative to the mainstream temperature and the inlet coolant temperature. This can be determined experimentally when the model is constructed so that the Biot number is similar to that of engine components. In this study, the overall cooling effectiveness was experimentally measured on a model turbine vane constructed of a material deigned to match Bi for engine conditions. The model incorporated an internal impingement cooling configuration. Overall cooling effectiveness and adiabatic film effectiveness were measured downstream of a single row of round holes positioned on the Suction Side of the vane. Experiments were conducted to evaluate the cooling effects of internal cooling alone, and then the combined effects of film cooling and internal cooling for a range of coolant flow rates. While the adiabatic film effectiveness decreased when using high momentum flux ratios for the film cooling, due to coolant jet separation, the overall cooling effectiveness increased at higher momentum flux ratios. This increase was due to increased internal cooling effects. Overall cooling effectiveness measurements were also compared to analytical predictions based on a 1D thermal analysis using measured adiabatic film effectiveness and overall cooling effectiveness without film cooling.Copyright © 2012 by ASME
Fangpan Zhong - One of the best experts on this subject based on the ideXlab platform.
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A Novel Suction-Side Winglet Design Philosophy for High-Pressure Turbine Rotor Tips
Journal of Turbomachinery, 2017Co-Authors: Chao Zhou, Fangpan ZhongAbstract:Winglet tips are promising candidates for future high-pressure turbine rotors. Many studies found that the design of the Suction-Side winglet is the key to the aerodynamic performance of a winglet tip, but there is no general agreement on the exact design philosophy. In this paper, a novel Suction-Side winglet design philosophy in a turbine cascade is introduced. The winglets are obtained based on the near-tip flow field of the datum tip geometry. The Suction-Side winglet aims to reduce the tip leakage flow particularly in the front part of the blade passage. It is found that on the casing endwall, the pressure increases in the area where the winglet is used. This reduces the tip leakage flow in the front part of the blade passage and the pitchwise pressure gradient on the endwall. As a result, the size of the tip leakage vortex reduces. A surprising observation is that the novel optimized winglet tip design eliminates the passage vortex and results in a further increasing of the efficiency. The tip leakage loss of the novel winglet tip is 18.1% lower than the datum cavity tip, with an increase of tip surface area by only 19.3%. The spanwise deflection of the winglet due to the centrifugal force is small. The tip heat load of the winglet tip is 17.5% higher than that of the cavity tip. Numerical simulation shows that in a turbine stage, this winglet tip increases the turbine stage efficiency by 0.9% mainly by eliminating the loss caused by the passage vortex at a tip gap size of 1.4% chord compared with a cavity tip.
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A Novel Suction Side Winglet Design Method for High Pressure Turbine Rotor Tips
Volume 2B: Turbomachinery, 2016Co-Authors: Chao Zhou, Fangpan ZhongAbstract:Winglet tips are promising candidates for future high pressure turbine rotors. Many studies found that the design of the Suction Side winglet is the key to the aerodynamic performance of a winglet tip, but there is no general agreement on the exact design method. In this paper, a novel Suction Side winglet design method will be introduced. The winglets are obtained based on the near tip flow field of the datum tip geometry. The Suction Side winglet aims to reduce the tip leakage flow particularly in the front part of the blade passage. It is found that on the casing endwall, the pressure increases in the area where the winglet is used. This reduces the tip leakage flow in the front part of the blade passage and the pitchwise pressure gradient on the endwall. As a result, the size of the tip leakage vortex reduces. A surprising observation is that the novel winglet tip design eliminates the scraping vortex and results in a further increasing of the efficiency. The tip leakage loss of the novel winglet tip is 23% lower than the datum cavity tip, with an increase of tip surface area by only 20%. The spanwise deflection of the winglet due to the centrifugal force is small. Numerical simulation shows that in a turbine stage, this winglet tip increases the turbine stage efficiency by 0.9% at a tip gap size of 1% span compared with a cavity tip.
