This work presents the aerodynamic topology optimisation of high pressure turbine rotor blade tips. Before carrying out the topology optimisation on the blade tip, some initial tip design studies were carried out. The winglet shape was optimised using two different design space setups and parameter limits. The optimum winglet design features the largest overhangs and in the case of unconstrained optimisation proved to have 1.40% greater aerodynamic efficiency. Secondly, a radial basis function based parametrisation was set up to allow the creation of single squealer line using the flat tip blade as a baseline geometry. The optimum case proved to increase efficiency 0.46% compared to the flat tip. After that, a combination of winglet and topology free squealer tips was investigated for topology optimisation. The winglet tip was parametrized as in the winglet only optimisation cases and topology free squealer walls were created using mapping of radial basis function surfaces of different complexities. It is shown that by combining both winglet and novel squealer topology optimisation, better designs of different topologies can be produced.
{"title":"Exploring Topology Optimisation of High Pressure Turbine Blade Tips","authors":"Luka Vincekovic, A. John, N. Qin, S. Shahpar","doi":"10.1115/GT2020-16059","DOIUrl":"https://doi.org/10.1115/GT2020-16059","url":null,"abstract":"\u0000 This work presents the aerodynamic topology optimisation of high pressure turbine rotor blade tips. Before carrying out the topology optimisation on the blade tip, some initial tip design studies were carried out. The winglet shape was optimised using two different design space setups and parameter limits. The optimum winglet design features the largest overhangs and in the case of unconstrained optimisation proved to have 1.40% greater aerodynamic efficiency. Secondly, a radial basis function based parametrisation was set up to allow the creation of single squealer line using the flat tip blade as a baseline geometry. The optimum case proved to increase efficiency 0.46% compared to the flat tip. After that, a combination of winglet and topology free squealer tips was investigated for topology optimisation. The winglet tip was parametrized as in the winglet only optimisation cases and topology free squealer walls were created using mapping of radial basis function surfaces of different complexities. It is shown that by combining both winglet and novel squealer topology optimisation, better designs of different topologies can be produced.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"144 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2020-09-21","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"114762034","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
This paper presents the Large Eddy Simulation (LES) of a Low-Pressure Turbine (LPT) Nozzle Guide Vane (NGV) for different Reynolds (Re) and Mach number (Ma). The analysis is based on a slice of the blade that may be representative of midspan flow where secondary flows, hub and shroud contributions are lower. In LPT, the variation of the Re during the mission of the gas turbine is a well-known effect since its value can vary of a factor four between take-off and cruise. This can induce performance variations due to various phenomena with among them suction side boundary layer separation on the aft portion of the blade due to an adverse pressure gradient and laminar boundary layer that can be maintained due to the relatively low Re in LPT. Similarly, the Ma in the LPT may vary depending on the thrust required from the gas turbine at the considered mission phase. The current paper investigates through numerical simulation the flow representative of a medium-sized LPT with three different Reynolds number Re = 175’000 (cruise), 280’000 (mid-level altitude) and 500’000 (take-off) keeping the same characteristic Mach number Ma = 0.2 and three different Mach number Ma = 0.2, 0.5 and 0.8 keeping the same Reynolds number Re = 280’000. The study focuses on different flow characteristics: pressure distribution around the blade, near-wall flow behavior and wake analysis. This includes the related generation of losses and the effect of Re and Ma on these different phenomena. A special emphasis is given to the generation of loss based on an entropy approach and the redistribution of mean kinetic energy towards turbulent kinetic energy. The results show that the increase of the Re has a destabilizing effect on potential separation while the increase of the Ma has a stabilizing effect. The peak in the TKE downstream of the blade is also moved upstream closer to the trailing edge when the Ma is increased.
{"title":"Reynolds and Mach Number Effects on the Flow in a Low-Pressure Turbine","authors":"M. Fiore","doi":"10.1115/GT2020-14325","DOIUrl":"https://doi.org/10.1115/GT2020-14325","url":null,"abstract":"\u0000 This paper presents the Large Eddy Simulation (LES) of a Low-Pressure Turbine (LPT) Nozzle Guide Vane (NGV) for different Reynolds (Re) and Mach number (Ma). The analysis is based on a slice of the blade that may be representative of midspan flow where secondary flows, hub and shroud contributions are lower. In LPT, the variation of the Re during the mission of the gas turbine is a well-known effect since its value can vary of a factor four between take-off and cruise. This can induce performance variations due to various phenomena with among them suction side boundary layer separation on the aft portion of the blade due to an adverse pressure gradient and laminar boundary layer that can be maintained due to the relatively low Re in LPT. Similarly, the Ma in the LPT may vary depending on the thrust required from the gas turbine at the considered mission phase. The current paper investigates through numerical simulation the flow representative of a medium-sized LPT with three different Reynolds number Re = 175’000 (cruise), 280’000 (mid-level altitude) and 500’000 (take-off) keeping the same characteristic Mach number Ma = 0.2 and three different Mach number Ma = 0.2, 0.5 and 0.8 keeping the same Reynolds number Re = 280’000. The study focuses on different flow characteristics: pressure distribution around the blade, near-wall flow behavior and wake analysis. This includes the related generation of losses and the effect of Re and Ma on these different phenomena. A special emphasis is given to the generation of loss based on an entropy approach and the redistribution of mean kinetic energy towards turbulent kinetic energy. The results show that the increase of the Re has a destabilizing effect on potential separation while the increase of the Ma has a stabilizing effect. The peak in the TKE downstream of the blade is also moved upstream closer to the trailing edge when the Ma is increased.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"3 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2020-09-21","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"116278292","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Vahid Iranidokht, Ilias Papagiannis, A. Kalfas, R. Abhari, Shigeki Senoo, Kazuhiro Momma
This paper presents the computational methodology, and experimental investigations accomplished to enhance the efficiency of a turbine stage by applying non-axisymmetric profiling on the rotor hub wall. The experimental setup was a two-stage axial turbine, which was tested at “LISA” test facility at ETH Zurich. The 1st stage was considered to create the flow history for the 2nd stage, which was the target of the optimization. The hub cavity of the 2nd stage was designed with large dimensions as a requirement of a steam turbine. The goal was to optimize the interaction of the cavity leakage flow with the rotor passage flow to reduce the losses and increase efficiency. The computational optimization was completed using a Genetic Algorithm coupled with an Artificial Neural Network on the 2nd stage of the test turbine. Unsteady time-accurate simulations were performed, using in-house developed “MULTI3” solver. Besides implementing all geometrical details (such as hub and tip cavities and fully 3D blade geometries) from the experimental setup into the computational model, it was learned that the unsteady upstream effect could not be neglected. A novel approach was introduced by using unsteady inlet boundary conditions to consider the multistage effect while reducing the computational cost to half. The importance of this implementation was tested by performing a steady simulation on the optimized geometry. The predicted efficiency gain from steady simulations was 4.5 times smaller (and negligible) compared to the unsteady approach. Excluding the cavity geometry was also assessed in a different simulation setup showing 3.9% over-prediction in the absolute efficiency value. Comprehensive steady and unsteady measurements were performed utilizing pneumatic, Fast Response Aerodynamic (FRAP), and Fast Response Entropy (FENT) probes, on the baseline and profiled test cases. The end-wall profiling was found to be successful in weakening the strength of the hub passage vortex by a 19% reduction in the under-over turning. As a result, the blockage was reduced near the hub region leading to more uniform mass flow distribution along the span. The flow angle deviations at the higher span position were also corrected due to better control of the flow angles. Furthermore, the improvements were confirmed by reductions in entropy, Secondary Kinetic Energy, and pressure unsteadiness. The accurate computational implementations led to an excellent agreement between the predicted and measured efficiency gain.
{"title":"Unsteady Steam Turbine Optimization Using High Fidelity CFD","authors":"Vahid Iranidokht, Ilias Papagiannis, A. Kalfas, R. Abhari, Shigeki Senoo, Kazuhiro Momma","doi":"10.1115/GT2020-14674","DOIUrl":"https://doi.org/10.1115/GT2020-14674","url":null,"abstract":"\u0000 This paper presents the computational methodology, and experimental investigations accomplished to enhance the efficiency of a turbine stage by applying non-axisymmetric profiling on the rotor hub wall. The experimental setup was a two-stage axial turbine, which was tested at “LISA” test facility at ETH Zurich. The 1st stage was considered to create the flow history for the 2nd stage, which was the target of the optimization. The hub cavity of the 2nd stage was designed with large dimensions as a requirement of a steam turbine. The goal was to optimize the interaction of the cavity leakage flow with the rotor passage flow to reduce the losses and increase efficiency.\u0000 The computational optimization was completed using a Genetic Algorithm coupled with an Artificial Neural Network on the 2nd stage of the test turbine. Unsteady time-accurate simulations were performed, using in-house developed “MULTI3” solver. Besides implementing all geometrical details (such as hub and tip cavities and fully 3D blade geometries) from the experimental setup into the computational model, it was learned that the unsteady upstream effect could not be neglected. A novel approach was introduced by using unsteady inlet boundary conditions to consider the multistage effect while reducing the computational cost to half. The importance of this implementation was tested by performing a steady simulation on the optimized geometry. The predicted efficiency gain from steady simulations was 4.5 times smaller (and negligible) compared to the unsteady approach. Excluding the cavity geometry was also assessed in a different simulation setup showing 3.9% over-prediction in the absolute efficiency value.\u0000 Comprehensive steady and unsteady measurements were performed utilizing pneumatic, Fast Response Aerodynamic (FRAP), and Fast Response Entropy (FENT) probes, on the baseline and profiled test cases. The end-wall profiling was found to be successful in weakening the strength of the hub passage vortex by a 19% reduction in the under-over turning. As a result, the blockage was reduced near the hub region leading to more uniform mass flow distribution along the span. The flow angle deviations at the higher span position were also corrected due to better control of the flow angles. Furthermore, the improvements were confirmed by reductions in entropy, Secondary Kinetic Energy, and pressure unsteadiness. The accurate computational implementations led to an excellent agreement between the predicted and measured efficiency gain.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"1 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2020-09-21","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"130379614","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
The effect of absolute pressure on deposition is studied in the High-Pressure Deposition Facility at The Ohio State University. Mass flow blockage trends are presented for the case of deposition in a single-wall flat plate effusion cooling geometry. Arizona Road Dust in the ranges of 0–10 μm and 0–3.5 μm is delivered to a 950 K coolant flow at a pressure ratio of 1.03 at absolute pressures ranging from 1 to 15.77 atm. The primary results indicate a non-linear decrease in blockage with increasing absolute pressure. Additional targeted experimental and companion computational fluid dynamics simulations are used to elucidate the relative importance of 3 physical mechanisms responsible for the trend with pressure: (1) the increase in effusion hole discharge coefficient (2) altered particle trajectories due to reduced effective Stokes and (3) altered erosion due to reduced effective Stokes. Results reveal that blockage and sticking rates are minimally affected by the changing velocity field due to the increase in discharge coefficient, thus the increased particle drag effect on particle trajectories and erosion due to changing flow density is the primary candidate. To support these conclusions, mesh morphing simulations of a 0–10 μm test are performed at 1 and 15.77 atm using the OSU Deposition Model, which captures both impact velocity and angle dependencies of deposition. The resulting structures and their unique characteristics are compared to experimental deposits, and the computational and experimental blockage histories support the conclusion that increased drag is the primary mechanism.
{"title":"An Experimental and Computational Investigation of Absolute Pressure Effects on Deposition in an Effusion Cooling Geometry","authors":"Christopher P. Brown, J. Bons","doi":"10.1115/GT2020-15632","DOIUrl":"https://doi.org/10.1115/GT2020-15632","url":null,"abstract":"\u0000 The effect of absolute pressure on deposition is studied in the High-Pressure Deposition Facility at The Ohio State University. Mass flow blockage trends are presented for the case of deposition in a single-wall flat plate effusion cooling geometry. Arizona Road Dust in the ranges of 0–10 μm and 0–3.5 μm is delivered to a 950 K coolant flow at a pressure ratio of 1.03 at absolute pressures ranging from 1 to 15.77 atm. The primary results indicate a non-linear decrease in blockage with increasing absolute pressure. Additional targeted experimental and companion computational fluid dynamics simulations are used to elucidate the relative importance of 3 physical mechanisms responsible for the trend with pressure: (1) the increase in effusion hole discharge coefficient (2) altered particle trajectories due to reduced effective Stokes and (3) altered erosion due to reduced effective Stokes. Results reveal that blockage and sticking rates are minimally affected by the changing velocity field due to the increase in discharge coefficient, thus the increased particle drag effect on particle trajectories and erosion due to changing flow density is the primary candidate. To support these conclusions, mesh morphing simulations of a 0–10 μm test are performed at 1 and 15.77 atm using the OSU Deposition Model, which captures both impact velocity and angle dependencies of deposition. The resulting structures and their unique characteristics are compared to experimental deposits, and the computational and experimental blockage histories support the conclusion that increased drag is the primary mechanism.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"50 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2020-09-21","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"134291694","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
The performance of the gas turbine engines deteriorates under off-design conditions with the change of required power demand. The performance of the engine can be improved by actively controlling the mass flow rate through the engine turbine section using the variable area nozzle turbine (VANT). However, to implement VANT, vanes of the turbine nozzle need to be rotated, which demands part clearance to be provided near both the hub and tip region. In order to keep constant part clearance during vane turning, endwalls are modified to spherical shapes in such a way that distance between vane and upstream as well downstream rotor remains unaffected. Also, vanes are rotated about its pivot, which creates a blockage to the leakage flow in the part clearance. As pivot creates a blockage to the leakage flow, the shape of the pivot need to be selected in such a way that leakage losses can be reduced effectively. Hence, to analyze the effect of different pivot shapes on leakage losses, two different shapes of the pivot i.e., circular and elliptic, are explored in the present study. Also, three vane turning angles are analyzed to observe the effectiveness of these pivot shapes. The effect of the pivot is analyzed by entropy contours and total pressure loss coefficient in the exit plane, and it is found that the elliptic pivot performs better than a circular pivot as an elliptic pivot provides more blockage to the leakage flow and hence minimizing the overall losses.
{"title":"Numerical Investigations on the Effect of Pivot Shapes in Part Clearance Flow Field of Variable Area LP Turbine Nozzle Vane","authors":"Hardikkumar Bhavsar, Chetan S. Mistry","doi":"10.1115/GT2020-15969","DOIUrl":"https://doi.org/10.1115/GT2020-15969","url":null,"abstract":"\u0000 The performance of the gas turbine engines deteriorates under off-design conditions with the change of required power demand. The performance of the engine can be improved by actively controlling the mass flow rate through the engine turbine section using the variable area nozzle turbine (VANT). However, to implement VANT, vanes of the turbine nozzle need to be rotated, which demands part clearance to be provided near both the hub and tip region. In order to keep constant part clearance during vane turning, endwalls are modified to spherical shapes in such a way that distance between vane and upstream as well downstream rotor remains unaffected. Also, vanes are rotated about its pivot, which creates a blockage to the leakage flow in the part clearance. As pivot creates a blockage to the leakage flow, the shape of the pivot need to be selected in such a way that leakage losses can be reduced effectively. Hence, to analyze the effect of different pivot shapes on leakage losses, two different shapes of the pivot i.e., circular and elliptic, are explored in the present study. Also, three vane turning angles are analyzed to observe the effectiveness of these pivot shapes. The effect of the pivot is analyzed by entropy contours and total pressure loss coefficient in the exit plane, and it is found that the elliptic pivot performs better than a circular pivot as an elliptic pivot provides more blockage to the leakage flow and hence minimizing the overall losses.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"1 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2020-09-21","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"128712983","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
The aerodynamic performance of three blade sets that represent the geometric manufacturing constraints of Ceramic Matrix Composite (CMC) blades was measured experimentally in a large-scale transonic turbine blade cascade. The trailing edge thicknesses of CMC blades are anticipated to be significantly larger than those of current state-of-the-art metallic blades. The blades tested in the current study had trailing edge thicknesses of 5%, 7%, and 9% relative to the blade axial chord. The three blade sets were designed with matching throat dimensions, so the blade loading distributions were varied to retain similar overall loading levels. Data were acquired at four Reynolds numbers, covering a factor of six range. All data were acquired at the design isentropic exit Mach number of 0.74. Measurements include blade loading and five-hole probe surveys at two downstream stations. The effects of inlet turbulence intensity were also quantified. Total pressure loss data were integrated to determine overall loss levels for each of the three measured blade passages. Excellent periodicity was noted. For low inlet turbulence levels, losses were surprisingly lower for the thickest trailing edge at low Reynolds numbers, but were highest at the maximum Reynolds number. In general, losses were found to scale well with Reynolds number, although front loading was found to significantly reduce the sensitivity of loss to Reynolds number. For high inlet turbulence intensity, losses were found to scale with trailing edge thickness as expected, and the Reynolds number sensitivity was reduced for all three blade sets. Loss levels at the highest Reynolds number were comparable at low and high inlet turbulence intensity levels.
{"title":"Effects of Trailing Edge Thickness and Blade Loading Distribution on the Aerodynamic Performance of Simulated CMC Turbine Blades","authors":"P. Giel, V. Shyam, Paht Juangphanich, J. Clark","doi":"10.1115/GT2020-15802","DOIUrl":"https://doi.org/10.1115/GT2020-15802","url":null,"abstract":"\u0000 The aerodynamic performance of three blade sets that represent the geometric manufacturing constraints of Ceramic Matrix Composite (CMC) blades was measured experimentally in a large-scale transonic turbine blade cascade. The trailing edge thicknesses of CMC blades are anticipated to be significantly larger than those of current state-of-the-art metallic blades. The blades tested in the current study had trailing edge thicknesses of 5%, 7%, and 9% relative to the blade axial chord. The three blade sets were designed with matching throat dimensions, so the blade loading distributions were varied to retain similar overall loading levels. Data were acquired at four Reynolds numbers, covering a factor of six range. All data were acquired at the design isentropic exit Mach number of 0.74. Measurements include blade loading and five-hole probe surveys at two downstream stations. The effects of inlet turbulence intensity were also quantified. Total pressure loss data were integrated to determine overall loss levels for each of the three measured blade passages. Excellent periodicity was noted. For low inlet turbulence levels, losses were surprisingly lower for the thickest trailing edge at low Reynolds numbers, but were highest at the maximum Reynolds number. In general, losses were found to scale well with Reynolds number, although front loading was found to significantly reduce the sensitivity of loss to Reynolds number. For high inlet turbulence intensity, losses were found to scale with trailing edge thickness as expected, and the Reynolds number sensitivity was reduced for all three blade sets. Loss levels at the highest Reynolds number were comparable at low and high inlet turbulence intensity levels.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"70 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2020-09-21","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"132687741","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Fluid Catalytic Cracking (FCC) is a process for the conversion of heavy stocks recovered from different refinery operations into high-octane gasoline, light fuel oils and olefin rich light gases. Zeolite particles are used as catalyst in the process to facilitate the cracking of long hydro-carbon molecules. Because of the contact between catalyst and feedstock, coke deposits on the catalyst active surface reducing its activity, air is used to remove it from the catalyst, in the regenerator. The coke burning produces a hot flow of flue gas with a valuable energy content (700°C – 3bar) that can be recovered using a Hot Gas Expander HGE before releasing the exhausts in the atmosphere. The flue gas is carrying a certain amount of catalyst so the expander hot parts (mainly blades but also the stator vanes) are constantly exposed to an abrasive medium that may compromise the reliability of the asset. Understanding how the solid particle erosion works is one of the challenges that HGEs designer shall deal with. To this purpose, a numerical approach was developed to account for some of the factors that are known to play a major role in erosion phenomena: particle velocity, angle of impact, composition and particle size, as well as the mechanical properties of the target surface at the operating temperature. The methodology was then applied to a cascade blade for which erosion data collected in a stationary erosion test rig are available. The comparison between numerical and experimental data showed a good agreement therefore providing confidence on the proposed methodology. Finally, the numerical procedure has been applied to a real case to show the effect of the different catalyst load on the life of the blades.
{"title":"An Industrial Methodology for Erosion Analysis of FCC Expander Blades","authors":"N. Maceli, G. Maffulli","doi":"10.1115/GT2020-14891","DOIUrl":"https://doi.org/10.1115/GT2020-14891","url":null,"abstract":"\u0000 Fluid Catalytic Cracking (FCC) is a process for the conversion of heavy stocks recovered from different refinery operations into high-octane gasoline, light fuel oils and olefin rich light gases.\u0000 Zeolite particles are used as catalyst in the process to facilitate the cracking of long hydro-carbon molecules.\u0000 Because of the contact between catalyst and feedstock, coke deposits on the catalyst active surface reducing its activity, air is used to remove it from the catalyst, in the regenerator. The coke burning produces a hot flow of flue gas with a valuable energy content (700°C – 3bar) that can be recovered using a Hot Gas Expander HGE before releasing the exhausts in the atmosphere.\u0000 The flue gas is carrying a certain amount of catalyst so the expander hot parts (mainly blades but also the stator vanes) are constantly exposed to an abrasive medium that may compromise the reliability of the asset.\u0000 Understanding how the solid particle erosion works is one of the challenges that HGEs designer shall deal with.\u0000 To this purpose, a numerical approach was developed to account for some of the factors that are known to play a major role in erosion phenomena: particle velocity, angle of impact, composition and particle size, as well as the mechanical properties of the target surface at the operating temperature.\u0000 The methodology was then applied to a cascade blade for which erosion data collected in a stationary erosion test rig are available. The comparison between numerical and experimental data showed a good agreement therefore providing confidence on the proposed methodology.\u0000 Finally, the numerical procedure has been applied to a real case to show the effect of the different catalyst load on the life of the blades.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"19 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2020-09-21","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"129355903","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
For unshrouded high pressure turbines, the design of rotor blade tips is a dynamic multi-purpose process. The model should have least aerodynamic and heat transfer losses caused by the flow through the tip gap and the blade. Recent studies show that by the modification of the blade tip, there is some improvement in efficiency and reduced tip gap flow. These include mainly parametric study of cavity tip and winglet tip modifications. Previous studies confirm the benefit of overhang in reducing tip leakage loss and heat transfer by changing the location of the tip leakage vortex away from the blade. But, to the best of authors’ knowledge none of the study reports the effect of casing relative motion on modified winglet and squealer tip. In the present study, novel modified tip blade geometry is introduced named as Top Squealer with Bottom Winglet (TSBW). Tip gap physics and loss generation has been investigated on three other different designs of squealer and winglet geometries and compared with the novel design. These designs are named as Flat Winglet, Cavity Squealer, Top Squealer with Bottom Winglet (new design) and Top Winglet with Bottom Squealer (TWBS). The flat tip rotor blade is considered as the base case for comparison. Three-dimensional computational study using ANSYS CFX 18.2 has been performed in order to examine the effect of casing relative motion on various designs of winglet and squealer tip. Structured mesh is created using ANSYS ICEM 18.2. At the downstream of trailing edge, distinct regions of momentum deficits named as the tip leakage vortex (LV), tip passage vortex (TPV), wakes and hub passage vortex (HPV) has been observed. Wakes formed due to the interaction of scraping vortex (SV), Tip passage vortex (TPV) and the leakage vortex. It has been found that Cavity Squealer and Top Squealer with Bottom Winglet (TSBW)gave the lowest total pressure loss coefficient and lowest tip leakage flow rate. But in casing relative motion case, cavity squealer with bottom wingletout runs the cavity squealer aerodynamic performance. This is due to the enlarged cavity at the tip. The incoming pass over flow got blocked because of enhanced interaction of LV and Scraping vortex (SV) in the tip cavity. As a result, tip leakage losses and tip leakage mass flow rate decreased.
{"title":"Effects of Casing Relative Motion on Aerodynamic Performance With Modified Designs of Winglet and Squealer Tip","authors":"A. Garg, S. Rawat, B. Prasad","doi":"10.1115/GT2020-15003","DOIUrl":"https://doi.org/10.1115/GT2020-15003","url":null,"abstract":"\u0000 For unshrouded high pressure turbines, the design of rotor blade tips is a dynamic multi-purpose process. The model should have least aerodynamic and heat transfer losses caused by the flow through the tip gap and the blade. Recent studies show that by the modification of the blade tip, there is some improvement in efficiency and reduced tip gap flow. These include mainly parametric study of cavity tip and winglet tip modifications. Previous studies confirm the benefit of overhang in reducing tip leakage loss and heat transfer by changing the location of the tip leakage vortex away from the blade. But, to the best of authors’ knowledge none of the study reports the effect of casing relative motion on modified winglet and squealer tip. In the present study, novel modified tip blade geometry is introduced named as Top Squealer with Bottom Winglet (TSBW). Tip gap physics and loss generation has been investigated on three other different designs of squealer and winglet geometries and compared with the novel design. These designs are named as Flat Winglet, Cavity Squealer, Top Squealer with Bottom Winglet (new design) and Top Winglet with Bottom Squealer (TWBS). The flat tip rotor blade is considered as the base case for comparison. Three-dimensional computational study using ANSYS CFX 18.2 has been performed in order to examine the effect of casing relative motion on various designs of winglet and squealer tip. Structured mesh is created using ANSYS ICEM 18.2. At the downstream of trailing edge, distinct regions of momentum deficits named as the tip leakage vortex (LV), tip passage vortex (TPV), wakes and hub passage vortex (HPV) has been observed. Wakes formed due to the interaction of scraping vortex (SV), Tip passage vortex (TPV) and the leakage vortex. It has been found that Cavity Squealer and Top Squealer with Bottom Winglet (TSBW)gave the lowest total pressure loss coefficient and lowest tip leakage flow rate. But in casing relative motion case, cavity squealer with bottom wingletout runs the cavity squealer aerodynamic performance. This is due to the enlarged cavity at the tip. The incoming pass over flow got blocked because of enhanced interaction of LV and Scraping vortex (SV) in the tip cavity. As a result, tip leakage losses and tip leakage mass flow rate decreased.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"6 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2020-09-21","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"115371190","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
G. Xia, G. Kalitzin, Jin Lee, G. Medic, P. SharmaOm
Accurate prediction of thermal field in high pressure turbines is a critical aspect of aerodynamic and durability design. This is particularly true when the flow at turbine inlet exhibits large gradients in temperature, both radially and circumferentially. In other words, in the presence of hot streaks from the combustor. In the numerical study presented in this paper, coupled high-fidelity eddy-resolving simulations of a combustor and a turbine are used to study the differences in the temperature profile at the exit of the first vane and the heat flux on the first blade, resulting from different positioning, or clocking, between the combustor fuel nozzles and turbine vanes. The resolved unsteadiness and turbulence from the combustor impacts mixing and secondary flow in the high pressure turbine. Temperature profiles from both actual combustor CFD simulations, as well as and modulated profiles with more pronounced variation, or pattern factor, are used at the turbine inlet. A threshold of the pattern factor that brings the benefit of clocking is identified. Clocking positioning between the combustor and vanes was studied for the most benefit.
{"title":"Hybrid RANS/LES Simulation of Combustor/Turbine Interactions","authors":"G. Xia, G. Kalitzin, Jin Lee, G. Medic, P. SharmaOm","doi":"10.1115/GT2020-14873","DOIUrl":"https://doi.org/10.1115/GT2020-14873","url":null,"abstract":"\u0000 Accurate prediction of thermal field in high pressure turbines is a critical aspect of aerodynamic and durability design. This is particularly true when the flow at turbine inlet exhibits large gradients in temperature, both radially and circumferentially. In other words, in the presence of hot streaks from the combustor. In the numerical study presented in this paper, coupled high-fidelity eddy-resolving simulations of a combustor and a turbine are used to study the differences in the temperature profile at the exit of the first vane and the heat flux on the first blade, resulting from different positioning, or clocking, between the combustor fuel nozzles and turbine vanes. The resolved unsteadiness and turbulence from the combustor impacts mixing and secondary flow in the high pressure turbine. Temperature profiles from both actual combustor CFD simulations, as well as and modulated profiles with more pronounced variation, or pattern factor, are used at the turbine inlet. A threshold of the pattern factor that brings the benefit of clocking is identified. Clocking positioning between the combustor and vanes was studied for the most benefit.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"13 6","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2020-09-21","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"133135588","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
One of major safety requirements from current airworthiness regulations is that the probability of hazardous engine effects should not occur exceed 10−7 per engine flight hour even in the event of component failure. Service experience of aeroengines indicates that turbine blade fracture is a common fault whose probability is far more than 10−7 per engine flight hour. It is obvious that overall engine system will be affected by blade failure. So, aerodynamic performance investigation in the event of one blade fracture failure has been assessed in the current study. With ANSYS-CFX, numerical model of GE-E3 (Energy Efficient Engine) high pressure turbine was established according to literature data. By comparing surface Mach number distribution at mid-span of vane in the first stage obtained numerically and experimentally, the most efficient turbulence model, i.e., the SST k-ω model, was identified. Based on the model, the 3-dimensional flow simulations under two configurations, full wheel geometry GE-E3 high pressure turbine without and with one blade fracture failure have been achieved. The following conclusions were drawn from 3-dimensional simulations: firstly, as for GE-E3 high pressure turbine, the effect of single turbine blade failure on turbine characteristics is slight; secondly, with blade loading coefficient as a criterion which is used for judging whether blade is affected, five blades which are significantly affected can be identified, and the surface pressure distributions of these five affected blades alter to varying degrees, accord-ingly, these film outflow static pressure characteristics alter as well; thirdly, after turbine blade fails, airflow accelerates violently along the suction side of downstream blade closest to failed blade and separates, however, air flow can not expand efficiently along the pressure side of upstream blade nearest to failed blade.
{"title":"Aerodynamic Performance Investigation of Turbine in the Event of One Blade Primary Fracture Failure","authors":"Shenmin Zhang, S. Ding, Tian Qiu","doi":"10.1115/GT2020-15152","DOIUrl":"https://doi.org/10.1115/GT2020-15152","url":null,"abstract":"\u0000 One of major safety requirements from current airworthiness regulations is that the probability of hazardous engine effects should not occur exceed 10−7 per engine flight hour even in the event of component failure. Service experience of aeroengines indicates that turbine blade fracture is a common fault whose probability is far more than 10−7 per engine flight hour. It is obvious that overall engine system will be affected by blade failure. So, aerodynamic performance investigation in the event of one blade fracture failure has been assessed in the current study. With ANSYS-CFX, numerical model of GE-E3 (Energy Efficient Engine) high pressure turbine was established according to literature data. By comparing surface Mach number distribution at mid-span of vane in the first stage obtained numerically and experimentally, the most efficient turbulence model, i.e., the SST k-ω model, was identified. Based on the model, the 3-dimensional flow simulations under two configurations, full wheel geometry GE-E3 high pressure turbine without and with one blade fracture failure have been achieved. The following conclusions were drawn from 3-dimensional simulations: firstly, as for GE-E3 high pressure turbine, the effect of single turbine blade failure on turbine characteristics is slight; secondly, with blade loading coefficient as a criterion which is used for judging whether blade is affected, five blades which are significantly affected can be identified, and the surface pressure distributions of these five affected blades alter to varying degrees, accord-ingly, these film outflow static pressure characteristics alter as well; thirdly, after turbine blade fails, airflow accelerates violently along the suction side of downstream blade closest to failed blade and separates, however, air flow can not expand efficiently along the pressure side of upstream blade nearest to failed blade.","PeriodicalId":388234,"journal":{"name":"Volume 2B: Turbomachinery","volume":"32 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2020-09-21","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"130793972","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
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