Alberto Amerini, Simone Paccati, Lorenzo Mazzei, Antonio Andreini
Abstract Computational fluid dynamics (CFD) plays a crucial role in the design of cooling systems in gas turbine combustors due to the difficulties and costs related to experimental measurements performed in pressurized reactive environments. Despite the massive advances in computational resources in the last years, reactive unsteady and multi-scale simulations of combustor real operating conditions are still computationally expensive. Modern combustors often employ cooling schemes based on effusion technique, which provides uniform protection of the liner from hot gases, combining the heat removal by means of heat sink effect with liner coverage and protection by film cooling. However, a large number of effusion holes results in a relevant increase of computational resources required to perform a CFD simulation capable of correctly predicting the thermal load on the metal walls within the combustor. Moreover, a multi-physics and multi-scale approach is mandatory to properly consider the different characteristic scales of the several heat transfer modes within combustion chambers to achieve a reliable prediction of aerothermal fields within the combustor and wall heat fluxes and temperatures. From this point of view, loosely coupled approaches permit a strong reduction of the calculation time, since each physics is solved through a dedicated solver optimized according to the considered heat transfer mechanism. The object of this work is to highlight the capabilities of a loosely coupled unsteady multi-physics tool (U-THERM3D) developed at the University of Florence within ansys fluent. The coupling strategy will be employed for the numerical analysis of the TECFLAM effusion cooled swirl burner, an academic test rig well representative of the working conditions of a partially premixed combustion chamber equipped with an effusion cooling system, developed by the collaboration of the Universities of Darmstadt, Heidelberg, Karlsruhe, and the DLR. The highly detailed numerical results obtained from the unsteady multi-physics and multi-scale simulation will be compared with experimental data to validate the numerical procedure.
{"title":"Assessment of a Conjugate Heat Transfer Method on an Effusion Cooled Combustor Operated With a Swirl Stabilized Partially Premixed Flame","authors":"Alberto Amerini, Simone Paccati, Lorenzo Mazzei, Antonio Andreini","doi":"10.1115/1.4056983","DOIUrl":"https://doi.org/10.1115/1.4056983","url":null,"abstract":"Abstract Computational fluid dynamics (CFD) plays a crucial role in the design of cooling systems in gas turbine combustors due to the difficulties and costs related to experimental measurements performed in pressurized reactive environments. Despite the massive advances in computational resources in the last years, reactive unsteady and multi-scale simulations of combustor real operating conditions are still computationally expensive. Modern combustors often employ cooling schemes based on effusion technique, which provides uniform protection of the liner from hot gases, combining the heat removal by means of heat sink effect with liner coverage and protection by film cooling. However, a large number of effusion holes results in a relevant increase of computational resources required to perform a CFD simulation capable of correctly predicting the thermal load on the metal walls within the combustor. Moreover, a multi-physics and multi-scale approach is mandatory to properly consider the different characteristic scales of the several heat transfer modes within combustion chambers to achieve a reliable prediction of aerothermal fields within the combustor and wall heat fluxes and temperatures. From this point of view, loosely coupled approaches permit a strong reduction of the calculation time, since each physics is solved through a dedicated solver optimized according to the considered heat transfer mechanism. The object of this work is to highlight the capabilities of a loosely coupled unsteady multi-physics tool (U-THERM3D) developed at the University of Florence within ansys fluent. The coupling strategy will be employed for the numerical analysis of the TECFLAM effusion cooled swirl burner, an academic test rig well representative of the working conditions of a partially premixed combustion chamber equipped with an effusion cooling system, developed by the collaboration of the Universities of Darmstadt, Heidelberg, Karlsruhe, and the DLR. The highly detailed numerical results obtained from the unsteady multi-physics and multi-scale simulation will be compared with experimental data to validate the numerical procedure.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":null,"pages":null},"PeriodicalIF":0.0,"publicationDate":"2023-04-06","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135836048","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}
Edwin Joseph Munoz Lopez, Alexander Hergt, Sebastian Grund, Volker Gümmer
Abstract The design of compressor blades has been transformed by the advent of optimization algorithms, allowing designers to focus on finding the best optimization strategy for a desired application. However, transonic flow conditions on compressor blades still present considerable modeling challenges, even for a 2D blade section. This paper then focuses on the design of a new state-of-the-art compressor cascade for future test campaigns at the DLR’s Transonic Cascade Wind Tunnel (TGK). For this purpose, a review of the cascades previously tested at the TGK was performed to select a main reference with good efficiency at high loading. The data gathered also informed the optimization strategy applied with the DLR’s optimizer, AutoOpti. The process chain was evaluated with Reynolds Averaged Navier–Stokes CFD simulations using the DLR’s solver, TRACE. The optimization was set to minimize two objective functions: the first one focused on the efficiency at the design point, and the second one focused on the efficiency over the working range. The result is a Pareto front of cascades with a wide variety of design features with an efficiency improvement over the working range of about 24%. This improvement was achieved with a comparable aerodynamic loading. Further analyses were performed to select the “best” cascade for future test campaigns. The significant improvement obtained with respect to the reference and the wide variety of designs observed demonstrates that there is still much to be learned about blade design through optimization; even for 2D cascades and specially in transonic flow.
优化算法的出现改变了压气机叶片的设计，使设计人员能够专注于为期望的应用寻找最佳优化策略。然而，压气机叶片的跨声速流动条件仍然存在相当大的建模挑战，即使对于二维叶片截面也是如此。然后，本文重点介绍了在DLR的跨音速叶栅风洞(TGK)进行未来测试的新型最先进的压气机叶栅的设计。为此，对TGK先前测试的叶栅进行了回顾，以选择在高负载下具有良好效率的主要参考。收集到的数据还为DLR的优化器AutoOpti应用的优化策略提供了依据。利用DLR的求解器TRACE，采用Reynolds average Navier-Stokes CFD模拟对流程链进行了评估。优化设置为最小化两个目标函数:第一个目标函数关注设计点的效率，第二个目标函数关注工作范围内的效率。其结果是一个具有多种设计特征的帕累托叶栅，在工作范围内效率提高了约24%。这一改进是在相当的空气动力载荷下实现的。进一步的分析是为了选择未来测试活动的“最佳”级联。相对于参考文献和观察到的各种设计所获得的显著改进表明，通过优化，叶片设计还有很多需要学习的地方;即使对于二维叶栅，特别是在跨音速流动中也是如此。
{"title":"The New Chapter of Transonic Compressor Cascade Design at the DLR","authors":"Edwin Joseph Munoz Lopez, Alexander Hergt, Sebastian Grund, Volker Gümmer","doi":"10.1115/1.4056982","DOIUrl":"https://doi.org/10.1115/1.4056982","url":null,"abstract":"Abstract The design of compressor blades has been transformed by the advent of optimization algorithms, allowing designers to focus on finding the best optimization strategy for a desired application. However, transonic flow conditions on compressor blades still present considerable modeling challenges, even for a 2D blade section. This paper then focuses on the design of a new state-of-the-art compressor cascade for future test campaigns at the DLR’s Transonic Cascade Wind Tunnel (TGK). For this purpose, a review of the cascades previously tested at the TGK was performed to select a main reference with good efficiency at high loading. The data gathered also informed the optimization strategy applied with the DLR’s optimizer, AutoOpti. The process chain was evaluated with Reynolds Averaged Navier–Stokes CFD simulations using the DLR’s solver, TRACE. The optimization was set to minimize two objective functions: the first one focused on the efficiency at the design point, and the second one focused on the efficiency over the working range. The result is a Pareto front of cascades with a wide variety of design features with an efficiency improvement over the working range of about 24%. This improvement was achieved with a comparable aerodynamic loading. Further analyses were performed to select the “best” cascade for future test campaigns. The significant improvement obtained with respect to the reference and the wide variety of designs observed demonstrates that there is still much to be learned about blade design through optimization; even for 2D cascades and specially in transonic flow.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":null,"pages":null},"PeriodicalIF":0.0,"publicationDate":"2023-04-03","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"136329113","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}
Dominik Ade, Johannes Eitenmueller, Sebastian Leichtfuss, Heinz-Peter Schiffer, Christoph Lyko, Gregor Schmid
Abstract Shroudless rotor blades are state-of-the-art in modern high pressure turbines. Tip leakage flow has a crucial impact on turbine efficiency. Specific blade tip designs are a key factor to handle tip leakage losses by controlling tip leakage flow and its re-entry into the rotor passage. Comparative measurements of a cavity squealer type tip and a notch type tip have been conducted at the Large Scale Turbine Rig at Technical University of Darmstadt. The test rig has been operated at the blade tips design point. Experimental data have been acquired at rotor inlet and outlet as well as within the rotor passage. For cavity squealer tips, a tip leakage vortex develops at the suction side as the tip leakage flow is rolled-up and further mixed with main annulus flow. The tip leakage vortex determines the blockage of main annulus flow at the blade tip. The design of the suction side of the notch tip benefits a jet-like re-entry of tip leakage flow into the passage. Results are a tip leakage vortex system with smaller sized vortices and a more homogeneous mass flow redistribution in the outer annulus of the rotor. The zone of affected main annulus flow at the blade tip increases through the dominant tip leakage jet.
{"title":"Investigation of the Interaction Between Tip Leakage and Main Annulus Flow in the Large Scale Turbine Rig: Comparison of Different Rotor Tip Geometries","authors":"Dominik Ade, Johannes Eitenmueller, Sebastian Leichtfuss, Heinz-Peter Schiffer, Christoph Lyko, Gregor Schmid","doi":"10.1115/1.4062123","DOIUrl":"https://doi.org/10.1115/1.4062123","url":null,"abstract":"Abstract Shroudless rotor blades are state-of-the-art in modern high pressure turbines. Tip leakage flow has a crucial impact on turbine efficiency. Specific blade tip designs are a key factor to handle tip leakage losses by controlling tip leakage flow and its re-entry into the rotor passage. Comparative measurements of a cavity squealer type tip and a notch type tip have been conducted at the Large Scale Turbine Rig at Technical University of Darmstadt. The test rig has been operated at the blade tips design point. Experimental data have been acquired at rotor inlet and outlet as well as within the rotor passage. For cavity squealer tips, a tip leakage vortex develops at the suction side as the tip leakage flow is rolled-up and further mixed with main annulus flow. The tip leakage vortex determines the blockage of main annulus flow at the blade tip. The design of the suction side of the notch tip benefits a jet-like re-entry of tip leakage flow into the passage. Results are a tip leakage vortex system with smaller sized vortices and a more homogeneous mass flow redistribution in the outer annulus of the rotor. The zone of affected main annulus flow at the blade tip increases through the dominant tip leakage jet.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":null,"pages":null},"PeriodicalIF":0.0,"publicationDate":"2023-04-03","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"136328783","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}
Simone Giaccherini, Lorenzo Pinelli, Michele Marconcini, Roberto Pacciani, Andrea Arnone
Abstract The massive growth of the air traffic during the last years is leading to stricter limitations on the noise emission levels radiated from aircraft engines. To face this issue, the installation of acoustic liners on the intake duct and the exhaust nozzles is a common strategy adopted to properly abate noise emissions coming from the fan, the compressor, the turbine, and the jet. In this context, the aim of the present article is to use high-fidelity large Eddy simulation (LES) to validate a multi-degree-of-freedom (MDOF) extension of the single-degree-of-freedom (SDOF) and double-degree-of-freedom (DDOF) analytical model provided by Hersh for impedance eduction of acoustic liners. First, the results of the original Hersh model are compared with LES calculations performed with the openfoam suite on a single-orifice and single-cavity layout (SDOF). Then the extension of the Hersh model to multicavity (MDOF) geometries by using a recursive formulation is presented. Finally, high-fidelity simulations are carried out for single-orifice and multicavity (MDOF) configurations to validate the method extension and to understand how resonant coupling and acoustic impedance are affected by multicavity resonant elements. The excellent agreement between the high-fidelity results and the analytical predictions for the single-cavity pattern confirms that the Hersh model is a useful formulation for a preliminary design of a SDOF acoustic liner. The model extension to MDOF configurations enables the designers to broaden the design space, and thus, a validated analytical method is strictly necessary to perform sensitivity studies to the multicavity geometrical parameters (i.e., facesheet thickness, cavities depth, porosity). Basically, a multicavity configuration makes the liner element resonate at different frequencies, leading to multiple absorption peaks in the audible range. In this way, the acoustic performance of the liner is extended to a wider frequency range, overcoming the limitations of a traditional SDOF configuration.
{"title":"Validation of an Analytical Model for the Acoustic Impedance Eduction of Multicavity Resonant Liners by a High-Fidelity Large Eddy Simulation Approach","authors":"Simone Giaccherini, Lorenzo Pinelli, Michele Marconcini, Roberto Pacciani, Andrea Arnone","doi":"10.1115/1.4056984","DOIUrl":"https://doi.org/10.1115/1.4056984","url":null,"abstract":"Abstract The massive growth of the air traffic during the last years is leading to stricter limitations on the noise emission levels radiated from aircraft engines. To face this issue, the installation of acoustic liners on the intake duct and the exhaust nozzles is a common strategy adopted to properly abate noise emissions coming from the fan, the compressor, the turbine, and the jet. In this context, the aim of the present article is to use high-fidelity large Eddy simulation (LES) to validate a multi-degree-of-freedom (MDOF) extension of the single-degree-of-freedom (SDOF) and double-degree-of-freedom (DDOF) analytical model provided by Hersh for impedance eduction of acoustic liners. First, the results of the original Hersh model are compared with LES calculations performed with the openfoam suite on a single-orifice and single-cavity layout (SDOF). Then the extension of the Hersh model to multicavity (MDOF) geometries by using a recursive formulation is presented. Finally, high-fidelity simulations are carried out for single-orifice and multicavity (MDOF) configurations to validate the method extension and to understand how resonant coupling and acoustic impedance are affected by multicavity resonant elements. The excellent agreement between the high-fidelity results and the analytical predictions for the single-cavity pattern confirms that the Hersh model is a useful formulation for a preliminary design of a SDOF acoustic liner. The model extension to MDOF configurations enables the designers to broaden the design space, and thus, a validated analytical method is strictly necessary to perform sensitivity studies to the multicavity geometrical parameters (i.e., facesheet thickness, cavities depth, porosity). Basically, a multicavity configuration makes the liner element resonate at different frequencies, leading to multiple absorption peaks in the audible range. In this way, the acoustic performance of the liner is extended to a wider frequency range, overcoming the limitations of a traditional SDOF configuration.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":null,"pages":null},"PeriodicalIF":0.0,"publicationDate":"2023-04-03","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"136329114","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}
Abstract In this article, an unsteady tip leakage flow instability is identified and investigated for an axial compressor at near-surge conditions. We describe the results of experimental verification of a new compressor developed by improving the blade geometry based on the criterion for the occurrence of this unsteady phenomenon. In a high-speed multistage axial flow compressor having a subsonic high stagger rotor blade, a surge test was carried out by changing the tip clearance. Under a condition of large tip clearance, a drastic decrease in the static pressure rise coefficient near the surge point was observed. At this operating condition, large, unsteady pressure fluctuation at the blade tip was confirmed, and the occurrence of tip leakage vortex breakdown was clarified by unsteady multipoint pressure measurement and detailed unsteady numerical simulations. Due to the blockage effect caused by vortex breakdown of the tip leakage, double leakage and axially reversed flow near the trailing edge were observed. It was found that the vortex breakdown region of the tip leakage vortex propagated in the circumferential direction and caused the rotating instability. In order to investigate the relationship among this unsteady flow phenomenon, tip clearance size, and flow pattern, unsteady calculation was conducted by changing the blade tip stagger and tip clearance size. A new concept of tip clearance of staggered pitch reference was proposed, which makes it possible to include the effect of blade loading on the clearance and clarifies that there exists a threshold at which vortex breakdown occurs/does not occur. On the basis of the aforementioned results, a high-speed multistage improved compressor was designed and manufactured to prevent tip leakage vortex breakdown. A clearance change test using active clearance control technology was conducted, and an increase in the static pressure rise coefficient near the surge point was confirmed for each clearance. The design concept of the improved blade, which suppressed the unsteady tip leakage flow instability, was tested and verified, and the effectiveness of the design guideline in actual gas turbines for power generation was confirmed.
{"title":"Investigation of Tip Leakage Vortex Breakdown in a High-Speed Multistage Axial Compressor","authors":"Ryosuke Seki, Toshihiko Azuma, Juniji Iwatani, Akihiro Nakaniwa, Hidetaka Okui, Takanori Shibata","doi":"10.1115/1.4056981","DOIUrl":"https://doi.org/10.1115/1.4056981","url":null,"abstract":"Abstract In this article, an unsteady tip leakage flow instability is identified and investigated for an axial compressor at near-surge conditions. We describe the results of experimental verification of a new compressor developed by improving the blade geometry based on the criterion for the occurrence of this unsteady phenomenon. In a high-speed multistage axial flow compressor having a subsonic high stagger rotor blade, a surge test was carried out by changing the tip clearance. Under a condition of large tip clearance, a drastic decrease in the static pressure rise coefficient near the surge point was observed. At this operating condition, large, unsteady pressure fluctuation at the blade tip was confirmed, and the occurrence of tip leakage vortex breakdown was clarified by unsteady multipoint pressure measurement and detailed unsteady numerical simulations. Due to the blockage effect caused by vortex breakdown of the tip leakage, double leakage and axially reversed flow near the trailing edge were observed. It was found that the vortex breakdown region of the tip leakage vortex propagated in the circumferential direction and caused the rotating instability. In order to investigate the relationship among this unsteady flow phenomenon, tip clearance size, and flow pattern, unsteady calculation was conducted by changing the blade tip stagger and tip clearance size. A new concept of tip clearance of staggered pitch reference was proposed, which makes it possible to include the effect of blade loading on the clearance and clarifies that there exists a threshold at which vortex breakdown occurs/does not occur. On the basis of the aforementioned results, a high-speed multistage improved compressor was designed and manufactured to prevent tip leakage vortex breakdown. A clearance change test using active clearance control technology was conducted, and an increase in the static pressure rise coefficient near the surge point was confirmed for each clearance. The design concept of the improved blade, which suppressed the unsteady tip leakage flow instability, was tested and verified, and the effectiveness of the design guideline in actual gas turbines for power generation was confirmed.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":null,"pages":null},"PeriodicalIF":0.0,"publicationDate":"2023-04-03","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"136329097","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}
Surge is a crucial problem for achieving a good working range of a centrifugal compressor and power cycle safety. Effective and early detection of surge inception is essential to predict and avoid the onset of it. In this paper, experiments have been conducted to investigate the pressure and acoustic characteristics of the centrifugal compressor surge. The compressor prototype used the research consists of an impeller wheel having backward blades, a vaneless diffuser, and a downstream volute constitute. The compressor prototype was instrumented with steady sensors to characterize the performance map, a fast dynamic pressure sensor to measure the surge-induced pressure fluctuation, and thirteen microphones to record the acoustic pressure on the inlet wall and in far field. The transient experimental data were analyzed in time and frequency domains. The surging noise is identified as a chord sound from multiple sources that radiate acoustic impulse synchronously at the compression system surge frequency (below 10 Hz in the current experiment). A previously rarely discussed surge inception is identified from the acoustic spectrogram within the frequency range of 30-85Hz, in which increase of the SPL is detected before the mass flow rate achieves the surge point. The pressure fluctuation corresponding to the inception is too weak to be captured by the normally used dynamic pressure sensors. It suggests that the acoustic measurement may have advantages in surge inception detection and surge prediction.
{"title":"Experimental Investigations of Centrifugal Compressor Surge Noise","authors":"Ben Zhao, T. Zhou, Ce Yang","doi":"10.1115/1.4062244","DOIUrl":"https://doi.org/10.1115/1.4062244","url":null,"abstract":"\u0000 Surge is a crucial problem for achieving a good working range of a centrifugal compressor and power cycle safety. Effective and early detection of surge inception is essential to predict and avoid the onset of it. In this paper, experiments have been conducted to investigate the pressure and acoustic characteristics of the centrifugal compressor surge. The compressor prototype used the research consists of an impeller wheel having backward blades, a vaneless diffuser, and a downstream volute constitute. The compressor prototype was instrumented with steady sensors to characterize the performance map, a fast dynamic pressure sensor to measure the surge-induced pressure fluctuation, and thirteen microphones to record the acoustic pressure on the inlet wall and in far field. The transient experimental data were analyzed in time and frequency domains. The surging noise is identified as a chord sound from multiple sources that radiate acoustic impulse synchronously at the compression system surge frequency (below 10 Hz in the current experiment). A previously rarely discussed surge inception is identified from the acoustic spectrogram within the frequency range of 30-85Hz, in which increase of the SPL is detected before the mass flow rate achieves the surge point. The pressure fluctuation corresponding to the inception is too weak to be captured by the normally used dynamic pressure sensors. It suggests that the acoustic measurement may have advantages in surge inception detection and surge prediction.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":null,"pages":null},"PeriodicalIF":1.7,"publicationDate":"2023-03-30","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"44048989","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. Lepicovsky, D. Šimurda, R. Kielb, P. Šidlof, M. Štěpán
A lack of reliable experimental data on transonic blade flutter in real turbomachines hampers further improvement of computational design predictions for off-design operation regimes of newly built machines. Acquiring unsteady pressure distributions on blades in real turbomachines is practically impossible. The goal of this work is to explore if an approximate unsteady pressure distributions can be created experimentally in a simple aerodynamic tunnel by composing a sequence of blade surface steady pressures acquired for gradually varying blade incidence angle offsets. An essential condition for such an approximation is that the dynamic pressure component induced by the blade motion is substantially smaller than the flow pattern changes caused by the varying interblade channel geometry. A dedicated test facility, called the Blade Flutter Module (BFM), has been built and used for this purpose. The BFM is a linear cascade consisting of five transonic airfoils that can be operated either in a static or a dynamic regime. For the dynamic operation, any of the blades can be oscillated at frequencies of up to 400 Hz with the maximum angular amplitude of three degrees. The obtained results confirm that within the range of the test conditions, the proposed compounded quasi-dynamic approach exhibits similar characteristics to dynamically acquired unsteady blade pressures. This is true for a test range of a maximum inlet Mach number of 1.09, maximum blade oscillating frequency of 100 Hz, and measurement of unsteady pressure distributions on a blade suction surface. The corresponding blade reduced frequency is 0.21.
{"title":"Quasi-dynamic approximation of unsteady pressure distribution for transonic airfoils in flutter","authors":"J. Lepicovsky, D. Šimurda, R. Kielb, P. Šidlof, M. Štěpán","doi":"10.1115/1.4062181","DOIUrl":"https://doi.org/10.1115/1.4062181","url":null,"abstract":"\u0000 A lack of reliable experimental data on transonic blade flutter in real turbomachines hampers further improvement of computational design predictions for off-design operation regimes of newly built machines. Acquiring unsteady pressure distributions on blades in real turbomachines is practically impossible. The goal of this work is to explore if an approximate unsteady pressure distributions can be created experimentally in a simple aerodynamic tunnel by composing a sequence of blade surface steady pressures acquired for gradually varying blade incidence angle offsets. An essential condition for such an approximation is that the dynamic pressure component induced by the blade motion is substantially smaller than the flow pattern changes caused by the varying interblade channel geometry. A dedicated test facility, called the Blade Flutter Module (BFM), has been built and used for this purpose. The BFM is a linear cascade consisting of five transonic airfoils that can be operated either in a static or a dynamic regime. For the dynamic operation, any of the blades can be oscillated at frequencies of up to 400 Hz with the maximum angular amplitude of three degrees. The obtained results confirm that within the range of the test conditions, the proposed compounded quasi-dynamic approach exhibits similar characteristics to dynamically acquired unsteady blade pressures. This is true for a test range of a maximum inlet Mach number of 1.09, maximum blade oscillating frequency of 100 Hz, and measurement of unsteady pressure distributions on a blade suction surface. The corresponding blade reduced frequency is 0.21.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":null,"pages":null},"PeriodicalIF":1.7,"publicationDate":"2023-03-21","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"47479544","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}
Shuo Mao, Daniel Van Hout, Kai Zhang, Jin Woo Lee, Wing Ng, Hongzhou Xu, M. Fox, Jun Li
This paper presents an experimental and computational study on implementing a dual cavity slashface cooling scheme on the thermal performance for the first stage nozzle guide vane with an axisymmetric, converging endwall. An upstream dual-row, staggered cylindrical hole cooling scheme provided purged coolant. The study was conducted under representative engine conditions of Maexit = 0.85 and Reexit,Cax = 1.5×106. Data were collected and analyzed using infrared thermography to map the endwall heat transfer performance throughout the passage. A flow visualization study was employed to gather qualitative insights into the endwall flow field. In addition, a complimentary CFD study was carried out to understand the endwall flow ingestion and egress behavior near the slashface. Results indicate that the dual-plenum slashface scheme leads to a unique ingestion-egression-ingestion-egression pattern. The ingestion and egression suppress the upstream coolant attachment at mid-passage near the suction side but create a favorable coolant coverage downstream of the slashface tail-end. The study also compared the dual-plenum design with the single-plenum design and concluded that the dual-plenum slashface is superior in film cooling performance.
{"title":"Upstream Jet Cooling and Dual Cavity Slashface Leakage Cooling on a Transonic Nozzle Guide Vane Endwall","authors":"Shuo Mao, Daniel Van Hout, Kai Zhang, Jin Woo Lee, Wing Ng, Hongzhou Xu, M. Fox, Jun Li","doi":"10.1115/1.4062180","DOIUrl":"https://doi.org/10.1115/1.4062180","url":null,"abstract":"\u0000 This paper presents an experimental and computational study on implementing a dual cavity slashface cooling scheme on the thermal performance for the first stage nozzle guide vane with an axisymmetric, converging endwall. An upstream dual-row, staggered cylindrical hole cooling scheme provided purged coolant. The study was conducted under representative engine conditions of Maexit = 0.85 and Reexit,Cax = 1.5×106. Data were collected and analyzed using infrared thermography to map the endwall heat transfer performance throughout the passage. A flow visualization study was employed to gather qualitative insights into the endwall flow field. In addition, a complimentary CFD study was carried out to understand the endwall flow ingestion and egress behavior near the slashface. Results indicate that the dual-plenum slashface scheme leads to a unique ingestion-egression-ingestion-egression pattern. The ingestion and egression suppress the upstream coolant attachment at mid-passage near the suction side but create a favorable coolant coverage downstream of the slashface tail-end. The study also compared the dual-plenum design with the single-plenum design and concluded that the dual-plenum slashface is superior in film cooling performance.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":null,"pages":null},"PeriodicalIF":1.7,"publicationDate":"2023-03-21","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"44607048","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}
Bryan Leicht, Brian T. Bohan, F. Schauer, R. Kemnitz, L. Rueschhoff, Benjamin Lam, James W. Kemp, W. Costakis
Compact military-grade jet engines offer many potential applications, including use in remotely piloted vehicles, but can be expensive to use for research and development purposes. A study aimed at increasing the power and thrust output of an inexpensive commercial compact engine found a material limitation issue in the turbomachinery. To gain the additional power, hotter turbine inlet temperatures were required. This temperature increase exceeded the limit of current uncooled metal turbine rotors but could be achieved through turbine rotors made from ceramics, such as silicon nitride, which would allow an increase in the thrust and power output by a factor of 1.44. Current ceramic turbine manufacturing methods are costly and time consuming for rapid prototyping, but recent breakthroughs in ceramic additive manufacturing have allowed for cheaper methods and faster production which are beneficial for use in research and development when designs are being rapidly changed and tested. This research demonstrated, through finite element analysis, that a silicon nitride turbine rotor could meet the increased turbine inlet temperature conditions to provide the desired thrust and power increase. Further, as a proof of concept, an additively manufactured drop-in replacement alumina turbine rotor was produced for the JetCat P400 small-scale engine in a manner that was cost-effective, timely, and potentially scalable for production. This compact engine was used to demonstrate that a cost-effective ceramic turbine could be manufactured. At the time of publication, the desired ceramic material, silicon nitride, was not available for additive manufacturing.
{"title":"MANUFACTURING A CERAMIC TURBINE ROTOR FOR A COMPACT JET ENGINE","authors":"Bryan Leicht, Brian T. Bohan, F. Schauer, R. Kemnitz, L. Rueschhoff, Benjamin Lam, James W. Kemp, W. Costakis","doi":"10.1115/1.4062124","DOIUrl":"https://doi.org/10.1115/1.4062124","url":null,"abstract":"\u0000 Compact military-grade jet engines offer many potential applications, including use in remotely piloted vehicles, but can be expensive to use for research and development purposes. A study aimed at increasing the power and thrust output of an inexpensive commercial compact engine found a material limitation issue in the turbomachinery. To gain the additional power, hotter turbine inlet temperatures were required. This temperature increase exceeded the limit of current uncooled metal turbine rotors but could be achieved through turbine rotors made from ceramics, such as silicon nitride, which would allow an increase in the thrust and power output by a factor of 1.44. Current ceramic turbine manufacturing methods are costly and time consuming for rapid prototyping, but recent breakthroughs in ceramic additive manufacturing have allowed for cheaper methods and faster production which are beneficial for use in research and development when designs are being rapidly changed and tested. This research demonstrated, through finite element analysis, that a silicon nitride turbine rotor could meet the increased turbine inlet temperature conditions to provide the desired thrust and power increase. Further, as a proof of concept, an additively manufactured drop-in replacement alumina turbine rotor was produced for the JetCat P400 small-scale engine in a manner that was cost-effective, timely, and potentially scalable for production. This compact engine was used to demonstrate that a cost-effective ceramic turbine could be manufactured. At the time of publication, the desired ceramic material, silicon nitride, was not available for additive manufacturing.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":null,"pages":null},"PeriodicalIF":1.7,"publicationDate":"2023-03-13","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"44266411","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}
D. Fellows, Vincent Iskandar, D. Bodony, R. Mcgowan, Sang-Guk Kang, Aaron J. Pope, Chol-Bum M. Kweon
Aircraft intermittent combustion engines often incorporate turbochargers adapted from ground-based applications to improve their efficiency and performance. These turbochargers can operate at off-design conditions and experience blade failures brought on by aerodynamic-induced blade resonances. A reduced-order model of the aeroelastic response of general fluid-structural configurations is developed using the Euler-Lagrange equation informed by numerical data from uncoupled computational fluid dynamic (CFD) and computational structural dynamic calculations. The structural response is derived from a method of assumed-modes approach. The unsteady fluid response is described by a modified version of piston theory that approximates the local transient pressure fluctuation in conjunction with steady CFD solution data. The reduced-order model is first applied to a classical panel flutter scenario and found to predict a flutter boundary that compares favorably to the boundary identified by existing theory and experimental data. The model is then applied to the high-pressure turbine of a dual-stage turbocharger. The model predictions are shown to reliably determine the lack of turbine blade flutter, and rudimentary damping comparisons are performed to assess the ability of the model to ascertain the susceptibility of the turbine to forced response. Obstacles associated with the current experimental state of the art that impinge upon further numerical validation are discussed.
{"title":"A Method for Predicting the Aeroelastic Response of Radial Turbomachines","authors":"D. Fellows, Vincent Iskandar, D. Bodony, R. Mcgowan, Sang-Guk Kang, Aaron J. Pope, Chol-Bum M. Kweon","doi":"10.1115/1.4057062","DOIUrl":"https://doi.org/10.1115/1.4057062","url":null,"abstract":"\u0000 Aircraft intermittent combustion engines often incorporate turbochargers adapted from ground-based applications to improve their efficiency and performance. These turbochargers can operate at off-design conditions and experience blade failures brought on by aerodynamic-induced blade resonances. A reduced-order model of the aeroelastic response of general fluid-structural configurations is developed using the Euler-Lagrange equation informed by numerical data from uncoupled computational fluid dynamic (CFD) and computational structural dynamic calculations. The structural response is derived from a method of assumed-modes approach. The unsteady fluid response is described by a modified version of piston theory that approximates the local transient pressure fluctuation in conjunction with steady CFD solution data. The reduced-order model is first applied to a classical panel flutter scenario and found to predict a flutter boundary that compares favorably to the boundary identified by existing theory and experimental data. The model is then applied to the high-pressure turbine of a dual-stage turbocharger. The model predictions are shown to reliably determine the lack of turbine blade flutter, and rudimentary damping comparisons are performed to assess the ability of the model to ascertain the susceptibility of the turbine to forced response. Obstacles associated with the current experimental state of the art that impinge upon further numerical validation are discussed.","PeriodicalId":49966,"journal":{"name":"Journal of Turbomachinery-Transactions of the Asme","volume":null,"pages":null},"PeriodicalIF":1.7,"publicationDate":"2023-03-08","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"47077572","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}
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