The gas turbine combustion process reaches gas temperatures that exceed the melting temperature of the combustor liner materials. Cooling the liner is critical to combustor durability and is often accomplished with double-walled liners that contain both impingement and effusion holes. The liner cooling is complicated with the interruption of the effusion cooling by large dilution jets that facilitate the combustion process. Given the presence of the dilution jets, it is important to understand the effect that the dilution jet has on the opposing wall in respect to the effusion film. This research includes measurements of the local static pressure distribution for a range of dilution jet momentum flux ratios to investigate the impact that the opposing dilution jet has on the effusion film. The interactions with the effusion cooling were also evaluated by measuring the overall cooling effectiveness across the panel. Measurements show that the opposing dilution jets did impact the liner at dilution jet momentum flux ratios that were greater than 20. The impacts at high momentum flux ratios were indicated through increased local static pressures measured on the surface of the combustor liner. Furthermore, the dilution touchdown decreased the overall cooling effectiveness of the effusion cooling. Results also indicated that the opposing dilution jets changed position on the liner as the dilution jet momentum flux ratio changes.
{"title":"Influence of Opposing Dilution Jets on Effusion Cooling","authors":"M. Creer, K. Thole","doi":"10.1115/gt2021-59429","DOIUrl":"https://doi.org/10.1115/gt2021-59429","url":null,"abstract":"\u0000 The gas turbine combustion process reaches gas temperatures that exceed the melting temperature of the combustor liner materials. Cooling the liner is critical to combustor durability and is often accomplished with double-walled liners that contain both impingement and effusion holes. The liner cooling is complicated with the interruption of the effusion cooling by large dilution jets that facilitate the combustion process. Given the presence of the dilution jets, it is important to understand the effect that the dilution jet has on the opposing wall in respect to the effusion film. This research includes measurements of the local static pressure distribution for a range of dilution jet momentum flux ratios to investigate the impact that the opposing dilution jet has on the effusion film. The interactions with the effusion cooling were also evaluated by measuring the overall cooling effectiveness across the panel.\u0000 Measurements show that the opposing dilution jets did impact the liner at dilution jet momentum flux ratios that were greater than 20. The impacts at high momentum flux ratios were indicated through increased local static pressures measured on the surface of the combustor liner. Furthermore, the dilution touchdown decreased the overall cooling effectiveness of the effusion cooling. Results also indicated that the opposing dilution jets changed position on the liner as the dilution jet momentum flux ratio changes.","PeriodicalId":204099,"journal":{"name":"Volume 5A: Heat Transfer — Combustors; Film Cooling","volume":"71 S2 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2021-06-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"122148618","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}
Double wall cooling, consisting of internal impingement cooling and external film cooling, is believed to be the most advanced technique in modern turbine blades cooling. In this paper, to improve the uniformity of temperature distribution, a flat plate double wall cooling model with gradient diameter of film and impingement holes was proposed, and the heat transfer and flow characteristics were investigated by solving steady three-dimensional Reynolds-Averaged Navier-Stokes (RANS) equations with SST k-ω turbulence model. The influence of gradient diameter on overall cooling effectiveness and total pressure loss was studied by comparing with the uniform pattern at the blowing ratios ranging from 0.5 to 2. For gradient diameter of film hole patterns, results show that −10% film pattern always has the lowest film flow non-uniformity coefficient. The laterally averaged overall cooling effectiveness of uniform pattern lies between that of +10% and −10% film patterns, but the intersection of three patterns moves upstream from the middle of flow direction with the increase of blowing ratio. Therefore, the −10% film pattern exerts the highest area averaged cooling effectiveness, which is improved by up to 1.6% and 1% at BR = 0.5 and 1 respectively compared with a uniform pattern. However, at higher blowing ratios, the +10% film pattern maintains higher cooling effectiveness and lower total pressure loss. For gradient diameter of impingement hole patterns, the intersection of laterally averaged overall cooling effectiveness in three patterns is located near the middle of flow direction under all blowing ratios. The uniform pattern has the highest area averaged cooling effectiveness and the smallest non-uniform coefficient, but the −10% jet pattern has advantages of reducing pressure loss, especially in the laminated loss.
{"title":"Conjugate Heat Transfer Characteristics of Double Wall Cooling With Gradient Diameter of Film and Impingement Holes","authors":"Juan He, Qinghua Deng, Z. Feng","doi":"10.1115/gt2021-59423","DOIUrl":"https://doi.org/10.1115/gt2021-59423","url":null,"abstract":"\u0000 Double wall cooling, consisting of internal impingement cooling and external film cooling, is believed to be the most advanced technique in modern turbine blades cooling. In this paper, to improve the uniformity of temperature distribution, a flat plate double wall cooling model with gradient diameter of film and impingement holes was proposed, and the heat transfer and flow characteristics were investigated by solving steady three-dimensional Reynolds-Averaged Navier-Stokes (RANS) equations with SST k-ω turbulence model. The influence of gradient diameter on overall cooling effectiveness and total pressure loss was studied by comparing with the uniform pattern at the blowing ratios ranging from 0.5 to 2. For gradient diameter of film hole patterns, results show that −10% film pattern always has the lowest film flow non-uniformity coefficient. The laterally averaged overall cooling effectiveness of uniform pattern lies between that of +10% and −10% film patterns, but the intersection of three patterns moves upstream from the middle of flow direction with the increase of blowing ratio. Therefore, the −10% film pattern exerts the highest area averaged cooling effectiveness, which is improved by up to 1.6% and 1% at BR = 0.5 and 1 respectively compared with a uniform pattern. However, at higher blowing ratios, the +10% film pattern maintains higher cooling effectiveness and lower total pressure loss. For gradient diameter of impingement hole patterns, the intersection of laterally averaged overall cooling effectiveness in three patterns is located near the middle of flow direction under all blowing ratios. The uniform pattern has the highest area averaged cooling effectiveness and the smallest non-uniform coefficient, but the −10% jet pattern has advantages of reducing pressure loss, especially in the laminated loss.","PeriodicalId":204099,"journal":{"name":"Volume 5A: Heat Transfer — Combustors; Film Cooling","volume":"206 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2021-06-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"122180207","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}
Discrete film cooling holes are limited by subtractive manufacturing techniques and experience depreciating performance when operating above critical velocity ratios. This study introduces an alternative method of bringing coolant to the surface of the blade via finite strips of porous material interlaced throughout the blade, made possible by advances in additive manufacturing (AM). Both experimental and computational studies were performed on the porous hybrid configuration to characterize downstream and off-wall performance, where experimental adiabatic effectiveness values were achieved using a plastic, fused deposition printed lattice structure. The method of bringing coolant onto the surface of the blade through an additively manufactured porous region experienced downstream adiabatic effectiveness values similar to slots while providing better structural stability. Additionally, the hybrid configuration outperformed shaped film cooling holes by injecting an ultra-thin layer of coolant that was evenly distributed span-wise across the blade. When operating at VRhybrid = 0.052 and L/d = 2 the hybrid configuration produced spatially averaged values 30% greater than the shaped holes while using equivalent coolant mass flow rate. Also, for an L/d = 10, the spatially averaged adiabatic effectiveness, for the hybrid configuration, is a factor of three greater than for shaped film cooling holes, while requiring a five times greater coolant mass flow rate. Finally, the RANS computational model accurately predicted downstream effectiveness values, at low velocity ratios, within experimental uncertainty but showed inaccuracies when predicting off wall effectiveness values and at higher velocity ratios.
{"title":"Additively Manufactured Porous Geometries for Hybrid Turbine Cooling","authors":"Nathan D. Fier, D. Bogard","doi":"10.1115/GT2021-58991","DOIUrl":"https://doi.org/10.1115/GT2021-58991","url":null,"abstract":"\u0000 Discrete film cooling holes are limited by subtractive manufacturing techniques and experience depreciating performance when operating above critical velocity ratios. This study introduces an alternative method of bringing coolant to the surface of the blade via finite strips of porous material interlaced throughout the blade, made possible by advances in additive manufacturing (AM). Both experimental and computational studies were performed on the porous hybrid configuration to characterize downstream and off-wall performance, where experimental adiabatic effectiveness values were achieved using a plastic, fused deposition printed lattice structure. The method of bringing coolant onto the surface of the blade through an additively manufactured porous region experienced downstream adiabatic effectiveness values similar to slots while providing better structural stability. Additionally, the hybrid configuration outperformed shaped film cooling holes by injecting an ultra-thin layer of coolant that was evenly distributed span-wise across the blade. When operating at VRhybrid = 0.052 and L/d = 2 the hybrid configuration produced spatially averaged values 30% greater than the shaped holes while using equivalent coolant mass flow rate. Also, for an L/d = 10, the spatially averaged adiabatic effectiveness, for the hybrid configuration, is a factor of three greater than for shaped film cooling holes, while requiring a five times greater coolant mass flow rate. Finally, the RANS computational model accurately predicted downstream effectiveness values, at low velocity ratios, within experimental uncertainty but showed inaccuracies when predicting off wall effectiveness values and at higher velocity ratios.","PeriodicalId":204099,"journal":{"name":"Volume 5A: Heat Transfer — Combustors; Film Cooling","volume":"177 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2021-06-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"121039342","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 turbine rotor inlet temperature of modern aero-engines has continuously increased in order to achieve higher thrust-to-weight ratio and thermal efficiency, which requires higher cooling effectiveness for turbine components. The turbine shroud is exposed to the blade tip leakage flow, and has become a common limiting factor for the turbine stage of advanced aero-engines. The three-dimensional numerical simulation on the unsteady film cooling characteristics of the high-pressure turbine shroud for an aero-engine under the rotor-stator interaction and the high blade rotation speed was conducted. The sliding grid technology was used to realize the relative movement between the turbine blade and the turbine shroud, and the rotor-stator interaction. Effects of the blade rotation, blowing ratios and the film jet direction on the unsteady film cooling performance of the high-pressure turbine shroud were revealed. It is found that the film cooling characteristics of the turbine shroud present an unsteady and periodic phenomenon. The blade tip clearance leakage flow, leakage vortex and mainstream suppression have important effects on the film cooling performance of the high-pressure turbine shroud. More attention should be paid to the insufficient cooling margin of the front row film holes due to coolant jet liftoff from the shroud surface under the high blowing ratio.
{"title":"Unsteady Film Cooling Performance on the High Pressure Turbine Shroud Under Rotor-Stator Interaction for an Aero-Engine","authors":"Zihao Bao, Z. Kou, B. Han, Guang-chao Li","doi":"10.1115/gt2021-58889","DOIUrl":"https://doi.org/10.1115/gt2021-58889","url":null,"abstract":"\u0000 The turbine rotor inlet temperature of modern aero-engines has continuously increased in order to achieve higher thrust-to-weight ratio and thermal efficiency, which requires higher cooling effectiveness for turbine components. The turbine shroud is exposed to the blade tip leakage flow, and has become a common limiting factor for the turbine stage of advanced aero-engines. The three-dimensional numerical simulation on the unsteady film cooling characteristics of the high-pressure turbine shroud for an aero-engine under the rotor-stator interaction and the high blade rotation speed was conducted. The sliding grid technology was used to realize the relative movement between the turbine blade and the turbine shroud, and the rotor-stator interaction. Effects of the blade rotation, blowing ratios and the film jet direction on the unsteady film cooling performance of the high-pressure turbine shroud were revealed. It is found that the film cooling characteristics of the turbine shroud present an unsteady and periodic phenomenon. The blade tip clearance leakage flow, leakage vortex and mainstream suppression have important effects on the film cooling performance of the high-pressure turbine shroud. More attention should be paid to the insufficient cooling margin of the front row film holes due to coolant jet liftoff from the shroud surface under the high blowing ratio.","PeriodicalId":204099,"journal":{"name":"Volume 5A: Heat Transfer — Combustors; Film Cooling","volume":"13 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2021-06-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"131897841","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}
Jie Wang, Chaomo Zhang, X. Liu, Liming Song, Jun Li, Z. Feng
Aiming at investigating the effects of crossflow and vortex generator on film cooling characteristics of fan-shaped hole, the film cooling performance was measured experimentally by infrared camera. The blowing ratio is fixed at 0.5 and 1.5. The Reynolds number of the mainstream based on the hole diameter remains at 7000 and the inlet Reynolds number of crossflow is 40000. The experimental results show that the film cooling performance becomes better when the blowing ratio increases from 0.5 to 1.5 for each model, and the film cooling performance becomes worse under the influence of crossflow. When the blowing ratio is 1.5, the area-averaged film cooling effectiveness of the fan-shaped hole model with vortex generator decreases by 16.6% because of the influence of crossflow. The combined model always performs better compared with the model without vortex generator under all working conditions. When the blowing ratio becomes 1.5, under the influence of crossflow, the area-averaged film cooling effectiveness of the combined model could increase by 14.8%, compared with the model without vortex generator. To further improve the film cooling performance, the global optimization algorithm based on the Kriging method and the CFD technology are coupled to optimize the combined model under crossflow condition at the high blowing ratio, and the optimized design is verified by experiments. The experimental results show that the area-averaged film cooling effectiveness of the optimized design increases by 17.8% compared with the reference model.
{"title":"Experimental Investigation and Optimal Design on the Film Cooling Performance of Fan-Shaped Hole With Vortex Generator Fed by Crossflow","authors":"Jie Wang, Chaomo Zhang, X. Liu, Liming Song, Jun Li, Z. Feng","doi":"10.1115/gt2021-59144","DOIUrl":"https://doi.org/10.1115/gt2021-59144","url":null,"abstract":"\u0000 Aiming at investigating the effects of crossflow and vortex generator on film cooling characteristics of fan-shaped hole, the film cooling performance was measured experimentally by infrared camera. The blowing ratio is fixed at 0.5 and 1.5. The Reynolds number of the mainstream based on the hole diameter remains at 7000 and the inlet Reynolds number of crossflow is 40000. The experimental results show that the film cooling performance becomes better when the blowing ratio increases from 0.5 to 1.5 for each model, and the film cooling performance becomes worse under the influence of crossflow. When the blowing ratio is 1.5, the area-averaged film cooling effectiveness of the fan-shaped hole model with vortex generator decreases by 16.6% because of the influence of crossflow. The combined model always performs better compared with the model without vortex generator under all working conditions. When the blowing ratio becomes 1.5, under the influence of crossflow, the area-averaged film cooling effectiveness of the combined model could increase by 14.8%, compared with the model without vortex generator. To further improve the film cooling performance, the global optimization algorithm based on the Kriging method and the CFD technology are coupled to optimize the combined model under crossflow condition at the high blowing ratio, and the optimized design is verified by experiments. The experimental results show that the area-averaged film cooling effectiveness of the optimized design increases by 17.8% compared with the reference model.","PeriodicalId":204099,"journal":{"name":"Volume 5A: Heat Transfer — Combustors; Film Cooling","volume":"67 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2021-06-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"127966145","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}
Limited optical access has been a challenge in gas-turbine related researches since the small blade pitch makes it difficult to arrange the camera at a proper viewing angle. In this paper, the application of an endoscopic system in a single-passage wind tunnel is presented. The film cooling effectiveness over the turbine vane’s pressure side with two types of holes was measured using the pressure-sensitive paint (PSP) technique. With the 7-7-7 shaped hole serving as the baseline, the sister shaped hole was compared side-by-side to examine its cooling performance at exit Ma = 0.84. Carbon dioxide (i.e., DR = 1.53) as coolant was discharged into the flow passage through two rows of holes (i.e., 4D spacing between holes and 1.5D spacing between rows) with blowing ratio (M) varied from 0.6 to 2.0. Through the implementation of the homography algorithm, the distorted coolant traces were recovered from the cambered surface. It was found that the film cooling effectiveness of both holes was greatly influenced by the blowing ratio. The sister shaped hole exhibited a relatively high effectiveness distribution at low M but its effectiveness decreased at high M due to the coolant jet detachment. In contrast, the 7-7-7 shaped hole demonstrated significantly higher effectiveness at high M, which can be attributed to the lower momentum flux ratio results of its larger exit area. The endoscope-based PSP technique and the obtained adiabatic effectiveness results may lay the foundation for other investigations and support other CFD studies in the gas turbine community.
{"title":"An Experimental Study of Turbine Vane Film Cooling Using Endoscope-Based PSP Technique in a Single-Passage Wind Tunnel","authors":"Kechen Wang, Hongyi Shao, Xu Zhang, D. Peng, Yingzheng Liu, Wenwu Zhou","doi":"10.1115/gt2021-60050","DOIUrl":"https://doi.org/10.1115/gt2021-60050","url":null,"abstract":"\u0000 Limited optical access has been a challenge in gas-turbine related researches since the small blade pitch makes it difficult to arrange the camera at a proper viewing angle. In this paper, the application of an endoscopic system in a single-passage wind tunnel is presented. The film cooling effectiveness over the turbine vane’s pressure side with two types of holes was measured using the pressure-sensitive paint (PSP) technique. With the 7-7-7 shaped hole serving as the baseline, the sister shaped hole was compared side-by-side to examine its cooling performance at exit Ma = 0.84. Carbon dioxide (i.e., DR = 1.53) as coolant was discharged into the flow passage through two rows of holes (i.e., 4D spacing between holes and 1.5D spacing between rows) with blowing ratio (M) varied from 0.6 to 2.0. Through the implementation of the homography algorithm, the distorted coolant traces were recovered from the cambered surface. It was found that the film cooling effectiveness of both holes was greatly influenced by the blowing ratio. The sister shaped hole exhibited a relatively high effectiveness distribution at low M but its effectiveness decreased at high M due to the coolant jet detachment. In contrast, the 7-7-7 shaped hole demonstrated significantly higher effectiveness at high M, which can be attributed to the lower momentum flux ratio results of its larger exit area. The endoscope-based PSP technique and the obtained adiabatic effectiveness results may lay the foundation for other investigations and support other CFD studies in the gas turbine community.","PeriodicalId":204099,"journal":{"name":"Volume 5A: Heat Transfer — Combustors; Film Cooling","volume":"55 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2021-06-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"116930819","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}
A vortex generator was incorporated into a wind tunnel to investigate more realistic inflow conditions for film cooling test rigs. The flow field signals are sampled numerically by probes in LES simulations and experimentally by using 1D hot-wire measurements to determine turbulence quantities. The LES shows that the turbulence is anisotropic which cannot be detected by the 1D hot-wire. Furthermore, the integral length scale which shall provide insight into the sizes of the turbulent eddies is determined using two approaches. The first uses the one probe at two times correlation method and could be evaluated from the numerical and experimental probes. The second correlation method exploits the spatial resolution in the LES domain by using the two probes at one time approach. Both methods show combustor-like turbulence length scales downstream of the vortex generator if the triple decomposition is applied onto the velocity signal.
{"title":"Analysis of LES and 1D Hot-Wire Data to Determine Actively Generated Main Flow Turbulence in a Film Cooling Test Rig","authors":"L. Fischer, M. Straußwald, M. Pfitzner","doi":"10.1115/gt2021-58661","DOIUrl":"https://doi.org/10.1115/gt2021-58661","url":null,"abstract":"\u0000 A vortex generator was incorporated into a wind tunnel to investigate more realistic inflow conditions for film cooling test rigs. The flow field signals are sampled numerically by probes in LES simulations and experimentally by using 1D hot-wire measurements to determine turbulence quantities. The LES shows that the turbulence is anisotropic which cannot be detected by the 1D hot-wire. Furthermore, the integral length scale which shall provide insight into the sizes of the turbulent eddies is determined using two approaches. The first uses the one probe at two times correlation method and could be evaluated from the numerical and experimental probes. The second correlation method exploits the spatial resolution in the LES domain by using the two probes at one time approach. Both methods show combustor-like turbulence length scales downstream of the vortex generator if the triple decomposition is applied onto the velocity signal.","PeriodicalId":204099,"journal":{"name":"Volume 5A: Heat Transfer — Combustors; Film Cooling","volume":"33 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2021-06-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"127006407","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}
D. Peral, A. Zaid, Christoph Benninghoven, Silvia Araguás-Rodríguez, D. Kluß, S. Karagiannopoulos, Robert Krewinkel, J. Feist
The requirement for reduced emissions and the growing demand on gas turbine efficiency are in part met through increasing firing temperatures. However, development budgets leave only limited time for dedicated thermal testing. Consequently, manufacturers are seeking novel temperature measurement technologies to validate new engine designs. This paper will demonstrate how a new temperature mapping technology can be utilized for non-dedicated (multi-cycling) testing while still delivering high-resolution temperature data in a non-dedicated test on a combustor of an industrial gas turbine. Typically, thermocouples are used to monitor the temperature during tests, but they only provide one data point. Colour changing thermal paints are used to deliver measurements over complete surfaces, but they require dedicated testing with short-duration exposure, necessitating dismantling and re-assembling the engine for further testing. Thermal History Coatings (THC) present an alternative solution to providing high-density temperature information. This coating permanently changes consistent with the maximum temperature of exposure during test. A laser-based instrumentation technique is then used to obtain temperatures. The maximum temperature profile of the surface can be determined through a customized calibration. Given the complex cooling system of a combustor, the high temperatures and the long-time exposure, this case offers a unique possibility for the testing of the coating under real engine conditions. The coated region covered the external surface of the can. Highly significant is the number of measurement points in excess of 7,000 (2 × 2 mm resolution, which enables advanced analysis. This provides insight into the impact of local features, e.g. the region adjacent to a cooling hole. The temperature profile is compared to a CFD-CHT model and thermocouple measurements for the calibration of cooling pre-design methods.
{"title":"High-Resolution Thermal Profiling of a Combustor in a Non-Dedicated Test Using Thermal History Coatings","authors":"D. Peral, A. Zaid, Christoph Benninghoven, Silvia Araguás-Rodríguez, D. Kluß, S. Karagiannopoulos, Robert Krewinkel, J. Feist","doi":"10.1115/gt2021-59687","DOIUrl":"https://doi.org/10.1115/gt2021-59687","url":null,"abstract":"\u0000 The requirement for reduced emissions and the growing demand on gas turbine efficiency are in part met through increasing firing temperatures. However, development budgets leave only limited time for dedicated thermal testing. Consequently, manufacturers are seeking novel temperature measurement technologies to validate new engine designs. This paper will demonstrate how a new temperature mapping technology can be utilized for non-dedicated (multi-cycling) testing while still delivering high-resolution temperature data in a non-dedicated test on a combustor of an industrial gas turbine.\u0000 Typically, thermocouples are used to monitor the temperature during tests, but they only provide one data point. Colour changing thermal paints are used to deliver measurements over complete surfaces, but they require dedicated testing with short-duration exposure, necessitating dismantling and re-assembling the engine for further testing.\u0000 Thermal History Coatings (THC) present an alternative solution to providing high-density temperature information. This coating permanently changes consistent with the maximum temperature of exposure during test. A laser-based instrumentation technique is then used to obtain temperatures. The maximum temperature profile of the surface can be determined through a customized calibration. Given the complex cooling system of a combustor, the high temperatures and the long-time exposure, this case offers a unique possibility for the testing of the coating under real engine conditions. The coated region covered the external surface of the can. Highly significant is the number of measurement points in excess of 7,000 (2 × 2 mm resolution, which enables advanced analysis. This provides insight into the impact of local features, e.g. the region adjacent to a cooling hole.\u0000 The temperature profile is compared to a CFD-CHT model and thermocouple measurements for the calibration of cooling pre-design methods.","PeriodicalId":204099,"journal":{"name":"Volume 5A: Heat Transfer — Combustors; Film Cooling","volume":"282 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2021-06-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"126852364","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}
Shoaib Ahmed, B. Wahls, S. Ekkad, Hanjie Lee, Y. Ho
One of the most effective ways to cool the combustor liner is through effusion cooling. Effusion cooling (also known as full coverage effusion cooling) involves uniformly spaced holes distributed throughout the combustor liner wall. Effusion cooling configurations are preferred for their high effectiveness, low-pressure penalty, and ease of manufacturing. In this paper, experimental results are presented for effusion cooling configurations for a realistic swirl driven can combustor under reacting (flame) conditions. The can-combustor was equipped with an industrial engine swirler and gaseous fuel (methane), subjecting the liner walls to engine representative flow and combustion conditions. In this study, three different effusion cooling liners with spanwise spacings of r/d = 6, 8, and 10 and streamwise spacing of z/d = 10 were tested for four coolant-to-main airflow ratios. The experiments were carried out at a constant main flow Reynolds number (based on combustor diameter) of 12,500 at a total equivalence ratio of 0.65. Infrared Thermography (IRT) was used to measure the liner outer surface temperature, and detailed overall effectiveness values were determined under steady-state conditions. The results indicate that decreasing the spanwise hole-to-hole spacing (r/d) from 10 to 8 increased the overall cooling effectiveness by 2–5%. It was found that reducing the spanwise hole-to-hole spacing further to r/d = 6 does not affect the cooling effectiveness implying the existence of an optimum spanwise hole-to-hole spacing. Also, the minimum liner cooling effectiveness on the liner wall was found to be downstream of the impingement location, which is not observed in existing literature for experiments done under non-reacting conditions.
{"title":"Effect of Spanwise Hole to Hole Spacing on Overall Cooling Effectiveness of Effusion Cooled Combustor Liners for a Swirl Stabilized Can Combustor","authors":"Shoaib Ahmed, B. Wahls, S. Ekkad, Hanjie Lee, Y. Ho","doi":"10.1115/gt2021-59312","DOIUrl":"https://doi.org/10.1115/gt2021-59312","url":null,"abstract":"\u0000 One of the most effective ways to cool the combustor liner is through effusion cooling. Effusion cooling (also known as full coverage effusion cooling) involves uniformly spaced holes distributed throughout the combustor liner wall. Effusion cooling configurations are preferred for their high effectiveness, low-pressure penalty, and ease of manufacturing. In this paper, experimental results are presented for effusion cooling configurations for a realistic swirl driven can combustor under reacting (flame) conditions. The can-combustor was equipped with an industrial engine swirler and gaseous fuel (methane), subjecting the liner walls to engine representative flow and combustion conditions. In this study, three different effusion cooling liners with spanwise spacings of r/d = 6, 8, and 10 and streamwise spacing of z/d = 10 were tested for four coolant-to-main airflow ratios. The experiments were carried out at a constant main flow Reynolds number (based on combustor diameter) of 12,500 at a total equivalence ratio of 0.65. Infrared Thermography (IRT) was used to measure the liner outer surface temperature, and detailed overall effectiveness values were determined under steady-state conditions. The results indicate that decreasing the spanwise hole-to-hole spacing (r/d) from 10 to 8 increased the overall cooling effectiveness by 2–5%. It was found that reducing the spanwise hole-to-hole spacing further to r/d = 6 does not affect the cooling effectiveness implying the existence of an optimum spanwise hole-to-hole spacing. Also, the minimum liner cooling effectiveness on the liner wall was found to be downstream of the impingement location, which is not observed in existing literature for experiments done under non-reacting conditions.","PeriodicalId":204099,"journal":{"name":"Volume 5A: Heat Transfer — Combustors; Film Cooling","volume":"4 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2021-06-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"122106618","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}
S. Cubeda, L. Andrei, L. Innocenti, F. Paone, Lorenzo Cocchi, A. Picchi, B. Facchini
In the recent years Additive Manufacturing (AM) methods, such as the Direct Metal Laser Melting (DMLM) technology, are getting more and more attractive and feasible for the realization of components and subcomponents of gas turbines. In particular, they are receiving much attention since, on one hand, the manufacturing of complex 3D geometries is allowed and, on the other, manufacturing and delivery times can be cut down. At the current state of the art, although AM is entering and spreading within modern gas turbines at fast pace, to the authors’ knowledge only few applications have yet been commercialized relatively to cooling holes, due to the intrinsic difficulties associated with such a critical feature. Lately, Baker Hughes is studying the possibility to manufacture film-cooling holes via the DMLM technology in order to exploit the flexibility of such innovative manufacturing method and hence eliminate additional processes and lead time. From the open literature it is known that additively manufactured holes can have a more irregular shape and higher roughness than traditional ones, which may lead not only to a reduction in coolant flow but more importantly to a decay of the film-cooling adiabatic effectiveness. For this reason, a test campaign has been conducted in collaboration with the University of Florence (Italy) with the objective of characterizing the performance (minimum passage diameter, flow check and adiabatic effectiveness) of AM vs traditional cylindrical holes on simple-geometry coupons built upon different construction angles. Results were then analyzed in order to fully compare the performance of AM vs traditional film-cooling holes at different operating regimes. In addition, selected holes were inspected through tomography in order to reveal the microscopic characteristics of lateral and outlet surfaces and get a further appreciation of the two different technologies. Ultimately the dependency of AM holes performance on print angles is sought with the purpose of characterizing the impact of such manufacturing technology on film-cooling holes design.
{"title":"Flow Check and Adiabatic Effectiveness Measurements on Traditionally Versus Additively Manufactured Film-Cooling Holes","authors":"S. Cubeda, L. Andrei, L. Innocenti, F. Paone, Lorenzo Cocchi, A. Picchi, B. Facchini","doi":"10.1115/gt2021-59638","DOIUrl":"https://doi.org/10.1115/gt2021-59638","url":null,"abstract":"\u0000 In the recent years Additive Manufacturing (AM) methods, such as the Direct Metal Laser Melting (DMLM) technology, are getting more and more attractive and feasible for the realization of components and subcomponents of gas turbines. In particular, they are receiving much attention since, on one hand, the manufacturing of complex 3D geometries is allowed and, on the other, manufacturing and delivery times can be cut down. At the current state of the art, although AM is entering and spreading within modern gas turbines at fast pace, to the authors’ knowledge only few applications have yet been commercialized relatively to cooling holes, due to the intrinsic difficulties associated with such a critical feature.\u0000 Lately, Baker Hughes is studying the possibility to manufacture film-cooling holes via the DMLM technology in order to exploit the flexibility of such innovative manufacturing method and hence eliminate additional processes and lead time. From the open literature it is known that additively manufactured holes can have a more irregular shape and higher roughness than traditional ones, which may lead not only to a reduction in coolant flow but more importantly to a decay of the film-cooling adiabatic effectiveness. For this reason, a test campaign has been conducted in collaboration with the University of Florence (Italy) with the objective of characterizing the performance (minimum passage diameter, flow check and adiabatic effectiveness) of AM vs traditional cylindrical holes on simple-geometry coupons built upon different construction angles.\u0000 Results were then analyzed in order to fully compare the performance of AM vs traditional film-cooling holes at different operating regimes. In addition, selected holes were inspected through tomography in order to reveal the microscopic characteristics of lateral and outlet surfaces and get a further appreciation of the two different technologies. Ultimately the dependency of AM holes performance on print angles is sought with the purpose of characterizing the impact of such manufacturing technology on film-cooling holes design.","PeriodicalId":204099,"journal":{"name":"Volume 5A: Heat Transfer — Combustors; Film Cooling","volume":"2 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2021-06-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"122480545","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}
京公网安备 11010802042870号
京ICP备2023020795号-1
扫码关注我们
求助内容：
应助结果提醒方式：