To improve the cooling performance, a novel impingement scheme is proposed to be used in a film-cooled turbine blade leading edge. Different from a normal one, the novel scheme consists of two rows of staggered impinging jets at oblique angles of ±35°, and is thus named as the staggered-oblique impingement scheme. A conjugated numerical investigation is carried out to illustrate the underlying mechanisms of the cooling performance. Three typical jet Reynolds numbers, 6000, 12000, and 18000, are studied using the validated SST k-ω turbulence model. Numerical results show a flow separation within the staggered-oblique impinging jets, which causes the discharge coefficient for the novel impingement scheme lower. Results also reveal a phenomenon difference that two symmetric vortices are induced by each normal impinging jet, while only one vortex appears on the acute angle side along with each staggered-oblique impinging jet. The flow fields of the staggered-oblique impingement scheme create a more uniform heat transfer distribution and a maximum of 23.7% higher area-averaged Nusselt number than the normal impingement scheme. The area-averaged overall cooling effectiveness for the novel scheme is higher than the normal one by a maximum of 4.8%. The uniformity and the enhancement of the overall cooling effectiveness arise from the wall jet being fully developed. The adiabatic cooling effectiveness is similar for both impingement schemes. This indicates that the improvement in overall cooling effectiveness for the staggered-oblique impingement scheme mainly arises from internal heat transfer.
{"title":"Conjugate heat transfer characteristics of a film-cooled turbine blade leading edge with staggered-oblique impinging jets","authors":"Hongye Li, Qiang Du, Qingzong Xu, Guangyao Xu, Haoyang Liu, Dawei Chen","doi":"10.1115/1.4063367","DOIUrl":"https://doi.org/10.1115/1.4063367","url":null,"abstract":"\u0000 To improve the cooling performance, a novel impingement scheme is proposed to be used in a film-cooled turbine blade leading edge. Different from a normal one, the novel scheme consists of two rows of staggered impinging jets at oblique angles of ±35°, and is thus named as the staggered-oblique impingement scheme. A conjugated numerical investigation is carried out to illustrate the underlying mechanisms of the cooling performance. Three typical jet Reynolds numbers, 6000, 12000, and 18000, are studied using the validated SST k-ω turbulence model. Numerical results show a flow separation within the staggered-oblique impinging jets, which causes the discharge coefficient for the novel impingement scheme lower. Results also reveal a phenomenon difference that two symmetric vortices are induced by each normal impinging jet, while only one vortex appears on the acute angle side along with each staggered-oblique impinging jet. The flow fields of the staggered-oblique impingement scheme create a more uniform heat transfer distribution and a maximum of 23.7% higher area-averaged Nusselt number than the normal impingement scheme. The area-averaged overall cooling effectiveness for the novel scheme is higher than the normal one by a maximum of 4.8%. The uniformity and the enhancement of the overall cooling effectiveness arise from the wall jet being fully developed. The adiabatic cooling effectiveness is similar for both impingement schemes. This indicates that the improvement in overall cooling effectiveness for the staggered-oblique impingement scheme mainly arises from internal heat transfer.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":" ","pages":""},"PeriodicalIF":1.7,"publicationDate":"2023-09-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"46831622","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":3,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Jiaxu Yao, Kecheng Zhang, Wenbin He, J. Lei, Jin Xu
Rotating significantly alters the internal cooling of turbine rotor blades by induced Coriolis force and buoyancy force, whose effects are characterized by the nondimensional rotation number (Ro) and buoyancy parameter (Bo). The present work was carried out in a new experimental rig of rotor blade internal cooling to obtain detailed heat transfer distributions when the three nondimensional criterion numbers (i.e., Re, Ro, and Bo) are similar to aero-engine operating conditions. Smooth and ribbed 2-pass internal cooling channels with a 180° tip turn are investigated. The hydraulic diameter is 25.4 mm (1 inch), and the aspect ratio is 2:1. The Reynolds number is fixed at 25000, with the maximum Ro and Bo of 0.316 and 0.272, respectively. The steady-state thermochromic liquid crystal (TLC) technique is used to measure detailed heat transfer distributions in the channel. Steady-state RANS simulations are also employed to resolve the flow characteristics. The effects of rotation on the flow and heat transfer characteristics are studied in this paper. The results show effects of rotation on the heat transfer distribution present apparent spatial discrepancy, especially around the bend region. The significant difference in the influence of rotation is witnessed in the smooth and the ribbed channel.
{"title":"Investigation on Heat Transfer Distribution of Rotating 2-Pass Smooth and Ribbed Channels under Aero-Engine Simulated Conditions","authors":"Jiaxu Yao, Kecheng Zhang, Wenbin He, J. Lei, Jin Xu","doi":"10.1115/1.4063248","DOIUrl":"https://doi.org/10.1115/1.4063248","url":null,"abstract":"\u0000 Rotating significantly alters the internal cooling of turbine rotor blades by induced Coriolis force and buoyancy force, whose effects are characterized by the nondimensional rotation number (Ro) and buoyancy parameter (Bo). The present work was carried out in a new experimental rig of rotor blade internal cooling to obtain detailed heat transfer distributions when the three nondimensional criterion numbers (i.e., Re, Ro, and Bo) are similar to aero-engine operating conditions. Smooth and ribbed 2-pass internal cooling channels with a 180° tip turn are investigated. The hydraulic diameter is 25.4 mm (1 inch), and the aspect ratio is 2:1. The Reynolds number is fixed at 25000, with the maximum Ro and Bo of 0.316 and 0.272, respectively. The steady-state thermochromic liquid crystal (TLC) technique is used to measure detailed heat transfer distributions in the channel. Steady-state RANS simulations are also employed to resolve the flow characteristics. The effects of rotation on the flow and heat transfer characteristics are studied in this paper. The results show effects of rotation on the heat transfer distribution present apparent spatial discrepancy, especially around the bend region. The significant difference in the influence of rotation is witnessed in the smooth and the ribbed channel.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":" ","pages":""},"PeriodicalIF":1.7,"publicationDate":"2023-08-24","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"47616379","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":3,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
As a large civil gas turbine is cooling down, natural convective flows cause components to cool asymmetrically – the bottom sector cools faster than the top. This can lead to a number of issues that have the potential to damage engine components and affect operability. The ability to predict this cooling cycle of a gas turbine has proven to be extremely difficult, owing to the complex nature of natural convective flow and its dependency on a considerable number of design parameters. An experimental and numerical investigation into the impact of axial ventilation (interaction between the annulus and the external air) and blade rows on the natural convective flow in a large civil gas turbine high-pressure compressor has led to some key discoveries. Axial ventilation caused a 70% increase in the peak top-to-bottom temperature difference in the cooling cycle, when compared to the baseline sealed case. The combinations of four blade rows and axial ventilation caused a 130% increase in peak temperature difference over the baseline case. Numerical simulations illuminated that the root cause of this was the cold air drawn into the lower section of the annulus led to a relatively high heat flux, coupled with a blockage effect on the natural draft in the upper section of the annulus. This study has highlighted the importance and inter-dependency of these effects in defining the level of rotor bow that is observed. Therefore, it is imperative that they are included and sufficiently captured in order for a shutdown.
{"title":"Axial Ventilation and Blade Row Effects on Transient Natural Convective Shutdown Cooling in a Gas Turbine","authors":"Daniel D. Fahy, P. Ireland","doi":"10.1115/1.4063246","DOIUrl":"https://doi.org/10.1115/1.4063246","url":null,"abstract":"\u0000 As a large civil gas turbine is cooling down, natural convective flows cause components to cool asymmetrically – the bottom sector cools faster than the top. This can lead to a number of issues that have the potential to damage engine components and affect operability. The ability to predict this cooling cycle of a gas turbine has proven to be extremely difficult, owing to the complex nature of natural convective flow and its dependency on a considerable number of design parameters. An experimental and numerical investigation into the impact of axial ventilation (interaction between the annulus and the external air) and blade rows on the natural convective flow in a large civil gas turbine high-pressure compressor has led to some key discoveries. Axial ventilation caused a 70% increase in the peak top-to-bottom temperature difference in the cooling cycle, when compared to the baseline sealed case. The combinations of four blade rows and axial ventilation caused a 130% increase in peak temperature difference over the baseline case. Numerical simulations illuminated that the root cause of this was the cold air drawn into the lower section of the annulus led to a relatively high heat flux, coupled with a blockage effect on the natural draft in the upper section of the annulus. This study has highlighted the importance and inter-dependency of these effects in defining the level of rotor bow that is observed. Therefore, it is imperative that they are included and sufficiently captured in order for a shutdown.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":" ","pages":""},"PeriodicalIF":1.7,"publicationDate":"2023-08-24","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"49516750","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":3,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
M. Perini, Nicolas Binder, Yannick Bousquet, Eric Schwartz
The present study aims to analyze rotating instabilities that may occur inside shroud cavities above rotors of low pressure turbine configurations. To do so, unsteady simulations on two configurations, one of which being a multistage configuration, at design and off-design conditions were carried out. Unsteady flow structures, uncorrelated from blade passing frequencies and depending on operating points, are identified in every rotor shroud exit cavity under investigation. Similarities regarding flow patterns and interactions with the main flow are observed: hot spots of gas, with different azimuthal periodicity and rotational speed, coming from the shroud, rotating at the interface with the main flow path, and modifying local aerodynamics. The influence on main flow aerodynamics and the origin of these instabilities are then discussed. Last but not least, the study at off-design operating points deepens the analysis and allows us to identify physical parameters driving the instabilities and propose a mechanism for instabilities’ onset. Those phenomena, which are physically sound and in agreement with scaling laws, still need to be experimentally observed. Advanced experiments are currently set up in the community for that purpose.
{"title":"Rotating Instabilities in Shrouded Low Pressure Turbine at Design and Off-Design Conditions","authors":"M. Perini, Nicolas Binder, Yannick Bousquet, Eric Schwartz","doi":"10.1115/1.4063247","DOIUrl":"https://doi.org/10.1115/1.4063247","url":null,"abstract":"\u0000 The present study aims to analyze rotating instabilities that may occur inside shroud cavities above rotors of low pressure turbine configurations. To do so, unsteady simulations on two configurations, one of which being a multistage configuration, at design and off-design conditions were carried out. Unsteady flow structures, uncorrelated from blade passing frequencies and depending on operating points, are identified in every rotor shroud exit cavity under investigation. Similarities regarding flow patterns and interactions with the main flow are observed: hot spots of gas, with different azimuthal periodicity and rotational speed, coming from the shroud, rotating at the interface with the main flow path, and modifying local aerodynamics. The influence on main flow aerodynamics and the origin of these instabilities are then discussed. Last but not least, the study at off-design operating points deepens the analysis and allows us to identify physical parameters driving the instabilities and propose a mechanism for instabilities’ onset. Those phenomena, which are physically sound and in agreement with scaling laws, still need to be experimentally observed. Advanced experiments are currently set up in the community for that purpose.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":" ","pages":""},"PeriodicalIF":1.7,"publicationDate":"2023-08-24","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"45755390","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":3,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
J. Zhang, Cun-liang Liu, Xiying Niu, Wei-jiang Xu, Li Zhang, Xuan Liu
The passage crossflow is one of the important secondary flow features that influence the endwall film cooling characteristic. While the upstream leakage flow can not only control the ingestion of mainstream gas but also provide endwall film cooling effectiveness distribution well. Consequently, this paper employed the pressure sensitive paint technique to investigate the film cooling performance of the endwall for the leakage flow. The effect of the passage crossflow on endwall film cooling is analyzed by numerical methods. Moreover, several crucial parameters that mass flow ratio (MFR), leakage slot inclination angle (a), density ratio (DR), and mainstream Reynolds number (Re) are considered for the study. Results indicate that a crescent-shaped uncooled area can be discovered near the suction side due to the existence of an inverse directional crescent-shaped crossflow region, and the uncooled area disappears gradually with the increase of MFR. Reducing the a, the axial velocity of the leakage flow is increased, which improves the film cooling effectiveness of the endwall for all the MFR cases. The effectiveness of the endwall is grown with the reduction of the DR at the low MFR case, however, the effectiveness is hardly influenced by the DR at the condition of MFR = 1.5%. Moreover, the enhancement of the Re that makes the strength of the secondary flow near the endwall reduce, which results in an increase in the endwall cooling effectiveness for all MFR cases.
{"title":"Effect of upstream leakage flow on film cooling characteristic of a turbine convex endwall","authors":"J. Zhang, Cun-liang Liu, Xiying Niu, Wei-jiang Xu, Li Zhang, Xuan Liu","doi":"10.1115/1.4063188","DOIUrl":"https://doi.org/10.1115/1.4063188","url":null,"abstract":"\u0000 The passage crossflow is one of the important secondary flow features that influence the endwall film cooling characteristic. While the upstream leakage flow can not only control the ingestion of mainstream gas but also provide endwall film cooling effectiveness distribution well. Consequently, this paper employed the pressure sensitive paint technique to investigate the film cooling performance of the endwall for the leakage flow. The effect of the passage crossflow on endwall film cooling is analyzed by numerical methods. Moreover, several crucial parameters that mass flow ratio (MFR), leakage slot inclination angle (a), density ratio (DR), and mainstream Reynolds number (Re) are considered for the study. Results indicate that a crescent-shaped uncooled area can be discovered near the suction side due to the existence of an inverse directional crescent-shaped crossflow region, and the uncooled area disappears gradually with the increase of MFR. Reducing the a, the axial velocity of the leakage flow is increased, which improves the film cooling effectiveness of the endwall for all the MFR cases. The effectiveness of the endwall is grown with the reduction of the DR at the low MFR case, however, the effectiveness is hardly influenced by the DR at the condition of MFR = 1.5%. Moreover, the enhancement of the Re that makes the strength of the secondary flow near the endwall reduce, which results in an increase in the endwall cooling effectiveness for all MFR cases.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":" ","pages":""},"PeriodicalIF":1.7,"publicationDate":"2023-08-16","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"47656530","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":3,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Matteo Dellacasagrande, Davide Lengani, Daniele Simoni, Marina Ubaldi, Francesco Bertini
Abstract In this work, streamwise oriented riblets were installed on a flat plate exposed to an adverse pressure gradient typical of low-pressure turbine (LPT) blade and, successively, on the suction side of an LPT cascade operating under unsteady flow. Different riblet dimensions and positions have been tested to quantify their effects on the boundary layer transition and on losses. The flat plate experiments allowed the detailed description of the riblet effects on the coherent structures affecting transition, thus providing a rationale for the identification of the optimal riblet geometry once scaled in wall-units. For riblet heights equal to about 20 wall-units, a maximum loss reduction of 8% was observed. Otherwise, for larger riblet dimensions, earlier transition occurs due to enhanced boundary layer instability and losses increase. Interestingly, the streamwise extension of the ribbed surfaces with respect to the transition region was found to play a minor role compared with the riblet dimension. The riblet configurations providing the highest reduction of viscous losses were then tested in the LPT blade cascade for different Reynolds numbers and with impinging upstream wakes. An overall profile loss reduction comparable to that observed in the flat plate case has been confirmed also in the unsteady operation of the turbine cascade. Low sensitivity of the profile losses to the riblet streamwise extension was also observed in the cascade application. This confirms that positive effects in terms of loss reduction can be obtained even when the exact transition position is not known a priori.
{"title":"Effects of ribbed surfaces on profile losses of low-pressure turbine blades","authors":"Matteo Dellacasagrande, Davide Lengani, Daniele Simoni, Marina Ubaldi, Francesco Bertini","doi":"10.1115/1.4063049","DOIUrl":"https://doi.org/10.1115/1.4063049","url":null,"abstract":"Abstract In this work, streamwise oriented riblets were installed on a flat plate exposed to an adverse pressure gradient typical of low-pressure turbine (LPT) blade and, successively, on the suction side of an LPT cascade operating under unsteady flow. Different riblet dimensions and positions have been tested to quantify their effects on the boundary layer transition and on losses. The flat plate experiments allowed the detailed description of the riblet effects on the coherent structures affecting transition, thus providing a rationale for the identification of the optimal riblet geometry once scaled in wall-units. For riblet heights equal to about 20 wall-units, a maximum loss reduction of 8% was observed. Otherwise, for larger riblet dimensions, earlier transition occurs due to enhanced boundary layer instability and losses increase. Interestingly, the streamwise extension of the ribbed surfaces with respect to the transition region was found to play a minor role compared with the riblet dimension. The riblet configurations providing the highest reduction of viscous losses were then tested in the LPT blade cascade for different Reynolds numbers and with impinging upstream wakes. An overall profile loss reduction comparable to that observed in the flat plate case has been confirmed also in the unsteady operation of the turbine cascade. Low sensitivity of the profile losses to the riblet streamwise extension was also observed in the cascade application. This confirms that positive effects in terms of loss reduction can be obtained even when the exact transition position is not known a priori.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":"6 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-08-16","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"136337538","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":3,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Bing Qiao, Xiao He, M. Vahdati, Y. Ju, Chuhua Zhang
Reynolds-Averaged Navier-Stokes (RANS) simulation is a routinely used tool for turbomachinery research and development, but it often over-predicts the pressure ratio of centrifugal compressors especially in the high impeller tip Mach number regime. In this paper, the effects of a series of geometrical and numerical uncertainties and errors on the aerodynamic performance of a centrifugal compressor are investigated systematically. The investigated compressor is the NASA CC3 centrifugal compressor under different impeller tip Mach number conditions. The investigated geometrical/numerical factors include the impeller blade fillet, the impeller hub cavity, the impeller running tip clearance, the averaging method in the post-process, the turbulence model, the inlet duct hub rotation, and the inlet turbulence boundary condition. Results show that the uncertainty of the predicted total pressure ratio generally increases with the impeller tip Mach number, and an analytical explanation of such a trend is provided. Among the investigated factors, the averaging method, the impeller blade fillet, the turbulence model and the impeller hub cavity have the most pronounced effects in determining the compressor total pressure ratio. By adopting realistic geometric features, advanced turbulence modeling treatments, and the same averaging method as the experiment, the over-prediction in the total pressure ratio can be alleviated. Detailed flow mechanism analysis with respect to the impeller hub cavity and the turbulence model has been performed. These findings provide valuable guidance for future RANS simulations of centrifugal compressors.
{"title":"On the Over-Prediction of Centrifugal Compressor Pressure Ratio in the High Impeller Tip Mach Number Regime","authors":"Bing Qiao, Xiao He, M. Vahdati, Y. Ju, Chuhua Zhang","doi":"10.1115/1.4063099","DOIUrl":"https://doi.org/10.1115/1.4063099","url":null,"abstract":"\u0000 Reynolds-Averaged Navier-Stokes (RANS) simulation is a routinely used tool for turbomachinery research and development, but it often over-predicts the pressure ratio of centrifugal compressors especially in the high impeller tip Mach number regime. In this paper, the effects of a series of geometrical and numerical uncertainties and errors on the aerodynamic performance of a centrifugal compressor are investigated systematically. The investigated compressor is the NASA CC3 centrifugal compressor under different impeller tip Mach number conditions. The investigated geometrical/numerical factors include the impeller blade fillet, the impeller hub cavity, the impeller running tip clearance, the averaging method in the post-process, the turbulence model, the inlet duct hub rotation, and the inlet turbulence boundary condition. Results show that the uncertainty of the predicted total pressure ratio generally increases with the impeller tip Mach number, and an analytical explanation of such a trend is provided. Among the investigated factors, the averaging method, the impeller blade fillet, the turbulence model and the impeller hub cavity have the most pronounced effects in determining the compressor total pressure ratio. By adopting realistic geometric features, advanced turbulence modeling treatments, and the same averaging method as the experiment, the over-prediction in the total pressure ratio can be alleviated. Detailed flow mechanism analysis with respect to the impeller hub cavity and the turbulence model has been performed. These findings provide valuable guidance for future RANS simulations of centrifugal compressors.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":" ","pages":""},"PeriodicalIF":1.7,"publicationDate":"2023-08-03","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"44240830","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":3,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Kedar Nawathe, Aaditya R. Nath, Yong Kim, Terrence Simon
Abstract The first-stage gas turbine vane surfaces and endwalls require aggressive cooling. This two-part paper introduces a modified design of the combustor–turbine (C–T) interface, the “close-coupled interface,” that is expected to increase cooling performance of vane passage surfaces. While Part I of the paper describes secondary flows and coolant transport in the passage, this part discusses the effects of the new C–T interface geometry on adiabatic cooling effectiveness of the endwall and vane surfaces. Compared to the traditional C–T interface, the coolant requirement is reduced for the same level of cooling effectiveness on all three surfaces for the new C–T interface design, confirming that it is an improvement over the previous design. The endwall crossflow is reduced by combustor coolant injection with the new interface leading to more pitchwise-uniform cooling of the endwall. For the pressure surface, increasing combustor coolant flowrate directly increases phantom cooling effectiveness and spreading of coolant away from the endwall. With the traditional passage vortex seen in the literature replaced by the impingement vortex of the present design, the suction surface receives less phantom cooling than does the pressure surface. However, cooling performance is still improved over that of the previous C–T interface design.
{"title":"Turbine Vane Passage Cooling Experiments With a Close-Coupled Combustor–Turbine Interface Geometry Part II: Describing the Coolant Coverage","authors":"Kedar Nawathe, Aaditya R. Nath, Yong Kim, Terrence Simon","doi":"10.1115/1.4062802","DOIUrl":"https://doi.org/10.1115/1.4062802","url":null,"abstract":"Abstract The first-stage gas turbine vane surfaces and endwalls require aggressive cooling. This two-part paper introduces a modified design of the combustor–turbine (C–T) interface, the “close-coupled interface,” that is expected to increase cooling performance of vane passage surfaces. While Part I of the paper describes secondary flows and coolant transport in the passage, this part discusses the effects of the new C–T interface geometry on adiabatic cooling effectiveness of the endwall and vane surfaces. Compared to the traditional C–T interface, the coolant requirement is reduced for the same level of cooling effectiveness on all three surfaces for the new C–T interface design, confirming that it is an improvement over the previous design. The endwall crossflow is reduced by combustor coolant injection with the new interface leading to more pitchwise-uniform cooling of the endwall. For the pressure surface, increasing combustor coolant flowrate directly increases phantom cooling effectiveness and spreading of coolant away from the endwall. With the traditional passage vortex seen in the literature replaced by the impingement vortex of the present design, the suction surface receives less phantom cooling than does the pressure surface. However, cooling performance is still improved over that of the previous C–T interface design.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":"18 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-08-02","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"136383688","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":3,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Kedar Nawathe, Aaditya R. Nath, Yong Kim, Terrence Simon
Abstract Due to the proximity of the first-stage gas turbine vanes to the combustor, coolant introduced to the combustor walls interacts with the endwall film coolant and changes the vane passage flow physics. Recent results show that combustor coolant contributes significantly to cooling the endwall and vane surfaces. In this paper, the traditional combustor-turbine interface was modified to improve overall cooling performance. The performance of this new injection cooling scheme on passage fluid dynamics and surface cooling is assessed. The first of this two-part paper reports detailed experimental tests that document secondary flows and coolant transport throughout the vane passage for four combustor coolant flowrates. The experimental facility imitates combustor coolant injection and engine-level turbulence and has a modified transition duct design, called the “close-coupled combustor-turbine interface.” The “impingement vortex” seen in previous studies with combustor cooling appears as the dominant secondary flow. It is observed in the present study over a wide range of flowrates, confirming its tie to the combustor coolant flowrate and not the combustor-turbine interface geometry. It was found, however, that the location and size of the impingement vortex are affected by coolant flowrate. Part II of this paper discusses the impact of the observed secondary flows on cooling vane passage surfaces.
{"title":"Turbine Vane Passage Cooling Experiments With a Close-Coupled Combustor-Turbine Interface Geometry—Part I: Describing the Flow","authors":"Kedar Nawathe, Aaditya R. Nath, Yong Kim, Terrence Simon","doi":"10.1115/1.4056439","DOIUrl":"https://doi.org/10.1115/1.4056439","url":null,"abstract":"Abstract Due to the proximity of the first-stage gas turbine vanes to the combustor, coolant introduced to the combustor walls interacts with the endwall film coolant and changes the vane passage flow physics. Recent results show that combustor coolant contributes significantly to cooling the endwall and vane surfaces. In this paper, the traditional combustor-turbine interface was modified to improve overall cooling performance. The performance of this new injection cooling scheme on passage fluid dynamics and surface cooling is assessed. The first of this two-part paper reports detailed experimental tests that document secondary flows and coolant transport throughout the vane passage for four combustor coolant flowrates. The experimental facility imitates combustor coolant injection and engine-level turbulence and has a modified transition duct design, called the “close-coupled combustor-turbine interface.” The “impingement vortex” seen in previous studies with combustor cooling appears as the dominant secondary flow. It is observed in the present study over a wide range of flowrates, confirming its tie to the combustor coolant flowrate and not the combustor-turbine interface geometry. It was found, however, that the location and size of the impingement vortex are affected by coolant flowrate. Part II of this paper discusses the impact of the observed secondary flows on cooling vane passage surfaces.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":"31 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-08-02","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"136383685","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":3,"RegionCategory":"工程技术","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
Samuel Altland, Xiang Yang, Karen A. Thole, Robert F. Kunz, Stephen T. McClain
Abstract Design for cooling effectiveness in turbine blades relies on accurate models for dynamic losses and heat transfer of internal cooling passages. Metal additive manufacturing (AM) has expanded the design space for these configurations, but can give rise to large-scale roughness features. The range of roughness length scales in these systems makes morphology resolved computational fluid dynamics (CFD) impractical. However, volumetric roughness models can be leveraged, as they have computational costs orders of magnitude lower. In this work, a discrete element roughness model (DERM), based on the double-averaged Navier–Stokes equations, is presented and applied to additively manufactured rough channels, representative of gas turbine blade cooling passages. Unique to this formulation of DERM is a generalized sheltering-based treatment of drag, a two-layer model for spatially averaged Reynolds stresses, and explicit treatment of dispersion. Six different AM rough surface channel configurations are studied, with roughness trough to peak sizes ranging from 15% to 60% nominal channel passage half-width, and the roughness Reynolds number ranges from Rek = 60 to 300. DERM predictions for spatially and temporally averaged mean flow quantities are compared to previously reported direct numerical simulation results. Good agreement in the mean velocity profiles, stress balances, and drag partitions are observed. While DERM models are typically calibrated to specific deterministic roughness morphologies at comparatively small roughness Reynolds numbers, the present more generalized DERM formulation has wider applicability. Here, it is demonstrated that the model can accommodate random roughness of large scale, typical of AM.
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