The Experts below are selected from a list of 264 Experts worldwide ranked by ideXlab platform
Shashimal Banneheke - One of the best experts on this subject based on the ideXlab platform.
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Tip-Leakage Losses in Subsonic and Transonic Blade Rows
Journal of Turbomachinery, 2012Co-Authors: Andrew P. S. Wheeler, Theodosios Korakianitis, Shashimal BannehekeAbstract:In this paper the effect of blade-Exit Mach Number on unshrouded turbine tip-leakage flows is investigated. Previously published experimental data of a high-pressure turbine blade are used to validate a computational fluid dynamics (CFD) code, which is then used to study the tip-leakage flow at blade-Exit Mach Numbers from 0.6 to 1.4. Three-dimensional (3D) calculations are performed of a flat-tip and a cavity-tip blade. Two-dimensional calculations are also performed to show the effect of various squealer-tip geometries on an idealized tip flow. The results show that as the blade-Exit Mach Number is increased the tip-leakage flow becomes choked. Therefore the tip-leakage flow becomes independent of the pressure difference across the tip and hence the blade loading. Thus the effect of the tip-leakage flow on overall blade loss reduces at blade-Exit Mach Numbers greater than 1.0. The results suggest that for transonic blade rows it should be possible to raise blade loading within the tip region without increasing tip-leakage loss
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Tip Leakage Losses in Subsonic and Transonic Blade-Rows
Volume 7: Turbomachinery Parts A B and C, 2011Co-Authors: Andrew P. S. Wheeler, Theodosios Korakianitis, Shashimal BannehekeAbstract:In this paper the effect of blade-Exit Mach Number on unshrouded turbine tip-leakage flows is investigated. Previously published experimental data of a high-pressure turbine blade are used to validate a CFD code, which is then used to study the tip-leakage flow at blade-Exit Mach Numbers from 0.6 to 1.4. Three-dimensional calculations are performed of a flat-tip and a cavity-tip blade. Two-dimensional calculations are also performed to show the effect of various squealer-tip geometries on an idealized tip-flow. The results show that as the blade-Exit Mach Number is increased the tip leakage flow becomes choked. Therefore the tip-leakage flow becomes independent of the pressure difference across the tip and hence the blade-loading. Thus the effect of the tip-leakage flow on overall blade loss reduces at blade-Exit Mach Numbers greater than 1.0. The results suggest that for transonic blade-rows it should be possible to raise blade loading within the tip region without increasing tip-leakage loss.Copyright © 2011 by ASME
Magnus Genrup - One of the best experts on this subject based on the ideXlab platform.
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Conceptual Mean-Line Design of Single and Twin-Shaft Oxy-Fuel Gas Turbine in a Semiclosed Oxy-Fuel Combustion Combined Cycle
Journal of engineering for gas turbines and power, 2013Co-Authors: Majed Sammak, Egill Maron Thorbergsson, Tomas Gronstedt, Magnus GenrupAbstract:The aim of this study was to compare single- and twin-shaft oxy-fuel gas turbines in a semiclosed oxy-fuel combustion combined cycle (SCOC–CC). This paper discussed the turboMachinery preliminary mean-line design of oxy-fuel compressor and turbine. The conceptual turbine design was performed using the axial through-flow code LUAX-T, developed at Lund University. A tool for conceptual design of axial compressors developed at Chalmers University was used for the design of the compressor. The modeled SCOC–CC gave a net electrical efficiency of 46% and a net power of 106 MW. The production of 95% pure oxygen and the compression of CO2 reduced the gross efficiency of the SCOC– CC by 10 and 2 percentage points, respectively. The designed oxy-fuel gas turbine had a power of 86 MW. The rotational speed of the single-shaft gas turbine was set to 5200 rpm. The designed turbine had four stages, while the compressor had 18 stages. The turbine Exit Mach Number was calculated to be 0.6 and the calculated value of AN2 was 40*10^6 rpm^2 m^2. The total calculated cooling mass flow was 25% of the compressor mass flow, or 47 kg/s. The relative tip Mach Number of the compressor at the first rotor stage was 1.15. The rotational speed of the twin-shaft gas generator was set to 7200 rpm, while that of the power turbine was set to 4800 rpm. A twin-shaft turbine was designed with five turbine stages to maintain the Exit Mach Number around 0.5. The twin-shaft turbine required a lower Exit Mach Number to maintain reasonable diffuser performance. The compressor turbine was designed with two stages while the power turbine had three stages. The study showed that a four-stage twin-shaft turbine produced a high Exit Mach Number. The calculated value of AN2 was 38*10^6 rpm^2 m^2. The total calculated cooling mass flow was 23% of the compressor mass flow, or 44 kg/s. The compressor was designed with 14 stages. The preliminary design parameters of the turbine and compressor were within established industrial ranges. From the results of this study, it was concluded that both single- and twin-shaft oxy-fuel gas turbines have advantages. The choice of a twin-shaft gas turbine can be motivated by the smaller compressor size and the advantage of greater flexibility in operation, mainly in the off-design mode. However, the advantages of a twin-shaft design must be weighed against the inherent simplicity and low cost of the simple single-shaft design.
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conceptual mean line design of single and twin shaft oxy fuel gas turbine in a semi closed oxy fuel combustion combined cycle
Proceedings of ASME Turbo Expo 2012: Power for Land Sea and Air. June 11-15 2012 Copenhagen Denmark, 2012Co-Authors: Majed Sammak, Magnus Genrup, Egill Maron Thorbergsson, Tomas GronstedtAbstract:The aim of this study was to compare single- and twin-shaft oxy-fuel gas turbines in a semi-closed oxy-fuel combustion combined cycle (SCOC-CC). This paper discussed the turboMachinery preliminary mean-line design of oxy-fuel compressor and turbine. The conceptual turbine design was performed using the axial through-flow code LUAX-T, developed at Lund University. A tool for conceptual design of axial compressors developed at Chalmers University was used for the design of the compressor. The modeled SCOC-CC gave a net electrical efficiency of 46% and a net power of 106 MW. The production of 95% pure oxygen and the compression of CO2 reduced the gross efficiency of the SCOC-CC by 10 and 2 percentage points, respectively. The designed oxy-fuel gas turbine had a power of 86 MW. The rotational speed of the single- shaft gas turbine was set to 5200 rpm. The designed turbine had four stages, while the compressor had 18 stages. The turbine Exit Mach Number was calculated to be 0.6 and the calculated value of AN2 was 40x10^6 rpm^2m^2. The total calculated cooling mass flow was 25% of the compressor mass flow, or 47 kg/s. The relative tip Mach Number of the compressor at the first rotor stage was 1.15. The rotational speed of the twin-shaft gas generator was set to 7200 rpm, while that of the power turbine was set to 4500 rpm. Twin-shaft turbine designed with five turbine stages to maintain the Exit Mach Number around 0.5. The twin-shaft turbine required a lower Exit Mach Number to maintain reasonable diffuser performance. The compressor turbine was designed with two stages while the power turbine had three stages. The study showed that a four-stage twin-shaft turbine produced a high Exit Mach Number. The calculated value of AN2 was 38x10^6 rpm^2m^2. The total calculated cooling mass flow was 23% of the compressor mass flow, or 44 kg/s. The compressor was designed with 14 stages. The preliminary design parameters of the turbine and compressor were within established industrial ranges. From the results of this study it was concluded that both single- and twin-shaft oxy-fuel gas turbines have advantages. The choice of a twin-shaft gas turbine can be motivated by the smaller compressor size and the advantage of greater flexibility in operation, mainly in off-design mode. However, the advantages of a twin-shaft design must be weighed against the inherent simplicity and low cost of the simple single-shaft design.
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conceptual design of a mid sized semi closed oxy fuel combustion combined cycle
Volume 4: Cycle Innovations; Fans and Blowers; Industrial and Cogeneration; Manufacturing Materials and Metallurgy; Marine; Oil and Gas Applications, 2011Co-Authors: Majed Sammak, Magnus Genrup, Egill Maron Thorbergsson, Tomas Gronstedt, Klas Jonshagen, Marcus Thern, Adrian DahlquistAbstract:This paper presents the study of a mid-sized semi-closed oxy-fuel combustion combined cycle (SCOC-CC) with net power output around 108 MW. The paper describes not only the power balance and the performance of the SCOC-CC, but also the conceptual design of the SCOC turbine and compressor. A model has been built in the commercial heat and mass balance code IPSEpro to estimate the efficiency of semi-closed dual-pressure oxy-fuel combustion combined cycle using natural gas as a fuel. In order to obtain the real physical properties of the working fluids in IPSEpro, the code was linked to the NIST Reference Fluid Thermodynamic and Transport Properties Database (REFPROP). The oxy-fuel turbine was modeled with the in-house Lund University package LUAX-T. Important features such as stage loading, loss modeling, cooling and geometric features were included to generate more accurate results. The oxy-fuel compressor has been modeled using a Chalmers university in-house tool for conceptual design of axial compressors. The conceptual design of the SCOC-CC process has a net efficiency of 47 %. The air separation unit and CO2 compression reduce the cycle efficiency by 10 and 2 percentage points, respectively. A single-shaft configuration was selected for the gas turbine simplicity. The rotational speed chosen was 5200 rpm and the turbine was designed with four stages. All stage preliminary design parameters are within ranges of established industrial axial turbine design limits. The main issue is the turbine Exit Mach Number; the stage must be lightly loaded in terms of pressure ratio to maintain the Exit Mach Number below 0.6. The compressor is designed with 18 stages. The current value of the product of the annulus area and the blade rotational speed squared (AN2) was calculated and found to be 40*10^6.
L. Zhang - One of the best experts on this subject based on the ideXlab platform.
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Numerical Investigation of Aerothermal Characteristics of the Blade Tip and Near-Tip Regions of a Transonic Turbine Blade
Journal of Turbomachinery, 2015Co-Authors: A. Arisi, S. Xue, Hee Koo Moon, L. ZhangAbstract:In modern gas turbine engines, the blade tips and near-tip regions are exposed to high thermal loads caused by the tip leakage flow. The rotor blades are therefore carefully designed to achieve optimum work extraction at engine design conditions without failure. However, very often gas turbine engines operate outside these design conditions which might result in sudden rotor blade failure. Therefore, it is critical that the effect of such off-design turbine blade operation be understood to minimize the risk of failure and optimize rotor blade tip performance. In this study, the effect of varying the Exit Mach Number on the tip and near-tip heat transfer characteristics was numerically studied by solving the steady Reynolds Averaged Navier Stokes (RANS) equation. The study was carried out on a highly loaded flat tip rotor blade with 1% tip gap and at Exit Mach Numbers of MExit = 0.85 (ReExit = 9.75 × 105) and MExit = 1.0 (ReExit = 1.15 × 106) with high freestream turbulence (Tu = 12%). The Exit Reynolds Number was based on the rotor axial chord. The numerical results provided detailed insight into the flow structure and heat transfer distribution on the tip and near-tip surfaces. On the tip surface, the heat transfer was found to generally increase with Exit Mach Number due to high turbulence generation in the tip gap and flow reattachment. While increase in Exit Mach Number generally raises he heat transfer over the whole blade surface, the increase is significantly higher on the near-tip surfaces affected by leakage vortex. Increase in Exit Mach Number was found to also induce strong flow relaminarisation on the pressure side near-tip. On the other hand, the size of the suction surface near-tip region affected by leakage vortex was insensitive to changes in Exit Mach Number but significant increase in local heat transfer was noted in this region.Copyright © 2014 by ASME
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Numerical Investigation of Aerothermal Characteristics of the Blade Tip and Near-Tip Regions of a Transonic Turbine Blade
Volume 5B: Heat Transfer, 2014Co-Authors: A. Arisi, H. K. Moon, S. Xue, L. ZhangAbstract:In modern gas turbine engines, the blade tips and near-tip regions are exposed to high thermal loads caused by the tip leakage flow. The rotor blades are therefore carefully designed to achieve optimum work extraction at engine design conditions without failure. However, very often gas turbine engines operate outside these design conditions which might result in sudden rotor blade failure. Therefore, it is critical that the effect of such off-design turbine blade operation be understood to minimize the risk of failure and optimize rotor blade tip performance. In this study, the effect of varying the Exit Mach Number on the tip and near-tip heat transfer characteristics was numerically studied by solving the steady Reynolds Averaged Navier Stokes (RANS) equation. The study was carried out on a highly loaded flat tip rotor blade with 1% tip gap and at Exit Mach Numbers of MExit = 0.85 (ReExit = 9.75 × 105) and MExit = 1.0 (ReExit = 1.15 × 106) with high freestream turbulence (Tu = 12%). The Exit Reynolds Number was based on the rotor axial chord. The numerical results provided detailed insight into the flow structure and heat transfer distribution on the tip and near-tip surfaces. On the tip surface, the heat transfer was found to generally increase with Exit Mach Number due to high turbulence generation in the tip gap and flow reattachment. While increase in Exit Mach Number generally raises he heat transfer over the whole blade surface, the increase is significantly higher on the near-tip surfaces affected by leakage vortex. Increase in Exit Mach Number was found to also induce strong flow relaminarisation on the pressure side near-tip. On the other hand, the size of the suction surface near-tip region affected by leakage vortex was insensitive to changes in Exit Mach Number but significant increase in local heat transfer was noted in this region.
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THE PERFORMANCE OF FAN-SHAPED HOLE FILM COOLING ON A GAS TURBINE BLADE AT TRANSONIC CONDITON WITH HIGH FREESTREAM TURBULENCE
50th AIAA Aerospace Sciences Meeting including the New Horizons Forum and Aerospace Exposition, 2012Co-Authors: S. Xue, Hee Koo Moon, Srinath V. Ekkad, L. ZhangAbstract:An experimental investigation was performed to study the cooling performance and convective heat transfer on a film cooled turbine blade surface. A 2D linear cascade model of the first stage turbine rotor blade of a land-based gas turbine was employed in the study. The film cooling configuration on the blade comprises of 2 rows of fan-shaped holes on the pressure side (PS), and 1 row of fan-shaped hole on the suction side (SS). The tests were performed in the Virginia Tech transonic wind tunnel facility, which simulates engine representative conditions of high turbulence intensity at the inlet and transonic Mach Numbers at the Exit, with matching Reynolds Number to that of the real engine. All the tests were performed at inlet turbulence intensity of 12% with integral length scale of 0.26 normalized by cascade blade pitch. Exit Mach Number of 0.67, 0.84, and 1.01 were chosen for the tests. Two combinations of blowing ratios at different rows of cooling holes were tested. (Nominal blowing ratio settings are: suction side injection BR=1.2 and 1.6; pressure side row 1 BR=2.8 and 3.8; pressure side row 2 BR=2.2, and 2.8.) The shock wave/boundary layer interaction effect on Nusselt Number and adiabatic effectiveness was observed on the suction side (SS) with a cascade Exit Mach Number of 1.01. The trends for Nusselt Number and adiabatic effectiveness across the shock impinged region suggest the existence of a separation bubble at the location of shock impingement. The detailed discussion of the shock effect on film cooling is presented.
Anthony Malandra - One of the best experts on this subject based on the ideXlab platform.
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Loss Generation in Transonic Turbine Blading
Journal of Turbomachinery, 2018Co-Authors: Penghao Duan, Choon S. Tan, Andrew Scribner, Anthony MalandraAbstract:The measured loss characteristic in a high-speed cascade tunnel of two turbine blades of different designs showed distinctly different trends with Exit Mach Number ranging from 0.8 to 1.4. Assessments using steady Reynolds-averaged Navier--Stokes equations (RANS) computation of the flow in the two turbine blades, complemented with control volume analyses and loss modeling, elucidate why the measured loss characteristic looks the way it is. The loss model categorizes the total loss in terms of boundary layer loss, trailing edge (TE) loss, and shock loss; it yields results in good agreement with the experimental data as well as steady RANS computed results. Thus, RANS is an adequate tool for determining the loss variations with Exit isentropic Mach Number and the loss model serves as an effective tool to interpret both the computational and the experimental data. The measured loss plateau in blade 1 for Exit Mach Number of 1–1.4 is due to a balance between a decrease of blade surface boundary layer loss and an increase in the attendant shock loss with Mach Number; this plateau is absent in blade 2 due to a greater rate in shock loss increase than the corresponding decrease in boundary layer loss. For Exit Mach Number from 0.85 to 1, the higher loss associated with shock system in blade 1 is due to the larger divergent angle downstream of the throat than that in blade 2. However, when Exit Mach Number is between 1.00 and 1.30, blade 2 has higher shock loss. For Exit Mach Number above an approximate value of 1.4, the shock loss for the two blades is similar as the flow downstream of the throat is completely supersonic. In the transonic to supersonic flow regime, the turbine design can be tailored to yield a shock pattern the loss of which can be mitigated in near equal amount of that from the boundary layer with increasing Exit Mach Number, hence yielding a loss plateau in transonic-supersonic regime.
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Loss Generation in Transonic Turbine Blading
Volume 2A: Turbomachinery, 2017Co-Authors: Penghao Duan, Choon S. Tan, Andrew Scribner, Anthony MalandraAbstract:The measured loss characteristic in a high-speed cascade tunnel of two turbine blades of different designs showed distinctly different trend with Exit Mach Number ranging from 0.8 to 1.4. Assessments using steady RANS computation of the flow in the two turbine blades, complemented with control volume analyses and loss modelling, elucidate why the measured loss characteristic looks the way it is. The loss model categorizes the total loss in terms of boundary layer loss, trailing edge loss and shock loss; it yields results in good agreement with the experimental data as well as steady RANS computed results. Thus RANS is an adequate tool for determining the loss variations with Exit isentropic Mach Number and the loss model serves as an effective tool to interpret both the computational and experimental data. The measured loss plateau in Blade 1 for Exit Mach Number of 1 to 1.4 is due to a balance between a decrease of blade surface boundary layer loss and an increase in the attendant shock loss with Mach Number; this plateau is absent in Blade 2 due to a greater rate in shock loss increase than the corresponding decrease in boundary layer loss. For Exit Mach Number from 0.85 to 1, the higher loss associated with shock system in Blade 1 is due to the larger divergent angle downstream of the throat than that in Blade 2. However when Exit Mach Number is between 1.00 and 1.30, Blade 2 has higher shock loss. For Exit Mach Number above around 1.4, the shock loss for the two blades is similar as the flow downstream of the throat is completely supersonic. In the transonic to supersonic flow regime, the turbine design can be tailored to yield a shock pattern the loss of which can be mitigated in near equal amount of that from the boundary layer with increasing Exit Mach Number, hence yielding a loss plateau in transonic-supersonic regime.
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Effect of Airfoil Shape and Turning Angle on Turbine Airfoil Aerodynamic Performance at Transonic Conditions
Volume 6: Fluids and Thermal Systems; Advances for Process Industries Parts A and B, 2011Co-Authors: Santosh Abraham, Srinath V. Ekkad, Kapil Panchal, Barry J. Brown, Anthony MalandraAbstract:Performance data for high turning gas turbine blades under transonic Mach Numbers is significantly lacking in literature. Performance of three gas turbine airfoils with varying turning angles at transonic flow conditions was investigated in this study. Midspan total pressure loss, secondary flow field and static pressure measurements on the airfoil surface in a linear cascade setting were measured. Airfoil curvature and true chord were varied to change the loading vs. chord for each airfoil. Airfoils A, D and E are designed to operate at different velocity triangles. Velocity triangle requirements (inlet/Exit Mach Number and gas angles) come from 1D and 2D models that include calibrated loss systems. One of the goals of this study was to use the experimental data to confirm/refine loss predictions for the effect of various Mach Numbers and gas turning angles. The cascade Exit Mach Numbers were varied within a range from 0.6 to 1.1. The airfoil turning angle ranges from 120° to 138°. A realistic inlet/Exit Mach Number ratio, that is representative of that seen in a real engine, was obtained by reducing the inlet span with respect to the Exit span of the airfoil, thereby creating a quasi 2D cascade. In order to compare the experimental results and study the detailed flow characteristics, 3D viscous compressible CFD analysis was also carried out.Copyright © 2011 by Siemens Energy Inc.
Andrew P. S. Wheeler - One of the best experts on this subject based on the ideXlab platform.
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Tip-Leakage Losses in Subsonic and Transonic Blade Rows
Journal of Turbomachinery, 2012Co-Authors: Andrew P. S. Wheeler, Theodosios Korakianitis, Shashimal BannehekeAbstract:In this paper the effect of blade-Exit Mach Number on unshrouded turbine tip-leakage flows is investigated. Previously published experimental data of a high-pressure turbine blade are used to validate a computational fluid dynamics (CFD) code, which is then used to study the tip-leakage flow at blade-Exit Mach Numbers from 0.6 to 1.4. Three-dimensional (3D) calculations are performed of a flat-tip and a cavity-tip blade. Two-dimensional calculations are also performed to show the effect of various squealer-tip geometries on an idealized tip flow. The results show that as the blade-Exit Mach Number is increased the tip-leakage flow becomes choked. Therefore the tip-leakage flow becomes independent of the pressure difference across the tip and hence the blade loading. Thus the effect of the tip-leakage flow on overall blade loss reduces at blade-Exit Mach Numbers greater than 1.0. The results suggest that for transonic blade rows it should be possible to raise blade loading within the tip region without increasing tip-leakage loss
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Tip Leakage Losses in Subsonic and Transonic Blade-Rows
Volume 7: Turbomachinery Parts A B and C, 2011Co-Authors: Andrew P. S. Wheeler, Theodosios Korakianitis, Shashimal BannehekeAbstract:In this paper the effect of blade-Exit Mach Number on unshrouded turbine tip-leakage flows is investigated. Previously published experimental data of a high-pressure turbine blade are used to validate a CFD code, which is then used to study the tip-leakage flow at blade-Exit Mach Numbers from 0.6 to 1.4. Three-dimensional calculations are performed of a flat-tip and a cavity-tip blade. Two-dimensional calculations are also performed to show the effect of various squealer-tip geometries on an idealized tip-flow. The results show that as the blade-Exit Mach Number is increased the tip leakage flow becomes choked. Therefore the tip-leakage flow becomes independent of the pressure difference across the tip and hence the blade-loading. Thus the effect of the tip-leakage flow on overall blade loss reduces at blade-Exit Mach Numbers greater than 1.0. The results suggest that for transonic blade-rows it should be possible to raise blade loading within the tip region without increasing tip-leakage loss.Copyright © 2011 by ASME