Daniele Cirigliano, Herol Lawerence D'Souza, Felix Grimm, Peter Kutne, Manfred Aigner
Micro Gas Turbines (MGTs) are nowadays largely used for electrical and thermal energy production in small buildings and households. Their reliability and compactness allow them to operate for thousands of hours with minimal maintenance. However, the long exposure at high temperatures in combustion chambers can promote creep, which can induce thermal fatigue and potential failure of these components. Creep-induced damage in MGTs has not yet been thoroughly investigated, due to the lack of numerical tools able to model these strongly coupled phenomena. This study presents the development of a Fortran-based subroutine integrated into ANSYS APDL. The code allows for a life assessment based on the Lemaitre-Chaboche creep damage model. Secondary creep and stress relaxation are modeled for the high-temperature resistant alloy Inconel718. A new set of temperature-dependent parameters for the Norton equation is provided, and the method to obtain these parameters from creep rupture tests is outlined. The model is validated and shows good agreement with experimental data. The subroutine correctly reproduces visco-plasticity, stress relaxation and damage under typical MGTs operating temperatures. This model constitutes the foundation of a life-assessment analysis for combustion chambers. The results highlight the impact of temperature and creep on the component’s life and the importance of integrating life assessment analysis into the preliminary design of combustion chambers.
{"title":"Creep-damage modelling for micro gas turbine combustion chambers lifetime prediction","authors":"Daniele Cirigliano, Herol Lawerence D'Souza, Felix Grimm, Peter Kutne, Manfred Aigner","doi":"10.33737/jgpps/163088","DOIUrl":"https://doi.org/10.33737/jgpps/163088","url":null,"abstract":"Micro Gas Turbines (MGTs) are nowadays largely used for electrical and thermal energy production in small buildings and households. Their reliability and compactness allow them to operate for thousands of hours with minimal maintenance. However, the long exposure at high temperatures in combustion chambers can promote creep, which can induce thermal fatigue and potential failure of these components. Creep-induced damage in MGTs has not yet been thoroughly investigated, due to the lack of numerical tools able to model these strongly coupled phenomena. This study presents the development of a Fortran-based subroutine integrated into ANSYS APDL. The code allows for a life assessment based on the Lemaitre-Chaboche creep damage model. Secondary creep and stress relaxation are modeled for the high-temperature resistant alloy Inconel718. A new set of temperature-dependent parameters for the Norton equation is provided, and the method to obtain these parameters from creep rupture tests is outlined. The model is validated and shows good agreement with experimental data. The subroutine correctly reproduces visco-plasticity, stress relaxation and damage under typical MGTs operating temperatures. This model constitutes the foundation of a life-assessment analysis for combustion chambers. The results highlight the impact of temperature and creep on the component’s life and the importance of integrating life assessment analysis into the preliminary design of combustion chambers.","PeriodicalId":53002,"journal":{"name":"Journal of the Global Power and Propulsion Society","volume":"59 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-06-09","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135158644","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}
Lukas Schäflein, Johannes Janssen, Henri Brandies, Peter Jeschke, Stephan Behre
This paper presents an investigation of the aerodynamic influence of rim seal purge flow injection on the main flow in a 1.5-stage turbine with non-axisymmetric end walls and a bowed stator using experimental flow measurements and unsteady RANS simulations. The study focuses on the secondary vortex structures formed in the rotor passages of the 1.5-stage axial turbine rig. Through performance map measurements, it was found that the efficiency gain of the non-axisymmetric configuration is partially eliminated by the injection of purge flow. Numerical investigations, which are supported by detailed flow measurements with five-hole probes and hot-wire probes, revealed that the injection of purge air flow intensifies vortex structures near the hub, thereby generating additional losses. These resulting vortex structures are highly similar both in the axisymmetric baseline and the non-axisymmetric configuration and are the result of jet-like vortices emerging from the cavity. From these findings, it can be concluded that the non-axisymmetric contour and the bowed stator no longer provides any efficiency benefit near the hub. Only the near the casing, where the flow is not affected by the purge flow, the optimized configuration continues to improve the efficiency of the rig by homogenizing the stator outflow and thus reducing the secondary flow structures in the rotor passages.
{"title":"Aerodynamic influence of rim seal purge flow injection on the main flow in a 1.5-stage axial turbine with nonaxisymmetric end wall contouring","authors":"Lukas Schäflein, Johannes Janssen, Henri Brandies, Peter Jeschke, Stephan Behre","doi":"10.33737/jgpps/162078","DOIUrl":"https://doi.org/10.33737/jgpps/162078","url":null,"abstract":"This paper presents an investigation of the aerodynamic influence of rim seal purge flow injection on the main flow in a 1.5-stage turbine with non-axisymmetric end walls and a bowed stator using experimental flow measurements and unsteady RANS simulations. The study focuses on the secondary vortex structures formed in the rotor passages of the 1.5-stage axial turbine rig. Through performance map measurements, it was found that the efficiency gain of the non-axisymmetric configuration is partially eliminated by the injection of purge flow. Numerical investigations, which are supported by detailed flow measurements with five-hole probes and hot-wire probes, revealed that the injection of purge air flow intensifies vortex structures near the hub, thereby generating additional losses. These resulting vortex structures are highly similar both in the axisymmetric baseline and the non-axisymmetric configuration and are the result of jet-like vortices emerging from the cavity. From these findings, it can be concluded that the non-axisymmetric contour and the bowed stator no longer provides any efficiency benefit near the hub. Only the near the casing, where the flow is not affected by the purge flow, the optimized configuration continues to improve the efficiency of the rig by homogenizing the stator outflow and thus reducing the secondary flow structures in the rotor passages.","PeriodicalId":53002,"journal":{"name":"Journal of the Global Power and Propulsion Society","volume":"26 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-04-28","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"136001579","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}
Lukas Schwerdt, N. Maroldt, Lars Panning‐von Scheidt, J. Wallaschek, J. Seume
To assess the influence of mistuning on the vibration amplitudes of turbomachinery rotors, reduced order models (ROMs) are widely used. A variety of methods are available for single-stage configurations and mostly aeroelastic effects can be taken into account. More recent research focusses on extending these methods to include multiple stages. However, due to the significantly increased computational effort of the aeroelastic simulations when adding more stages to the models, these ROMs are rarely applied with the inclusion of multistage aeroelastic effects. It is therefore desirable to develop reduction methods which minimize the number of these simulations to reduce the computational cost and thereby enable analyses of rotors with multiple stages including aeroelastic effects. In this paper, a cyclic Craig-Bampton reduction method with an a priori interface reduction for multistage rotors is extended with an additional a posteriori interface reduction to reduce the number of aeroelastic simulations necessary for a given accuracy level of the ROM. The interface degrees of freedom between stages are reduced using a modified version of Characteristic Constraint Modes, to yield a more efficient representation of their displacements while retaining their monoharmonic nature. The method is applied to a two-stage axial compressor with full aeroelastic coupling between the stages and its reduced computational effort is demonstrated. Additionally, two sorting methods for the degrees of freedom (DOFs) of the ROM are compared.
{"title":"An improved reduced order model for bladed disks including multistage aeroelastic and structural coupling","authors":"Lukas Schwerdt, N. Maroldt, Lars Panning‐von Scheidt, J. Wallaschek, J. Seume","doi":"10.33737/jgpps/161707","DOIUrl":"https://doi.org/10.33737/jgpps/161707","url":null,"abstract":"To assess the influence of mistuning on the vibration amplitudes of turbomachinery rotors, reduced order models (ROMs) are widely used. A variety of methods are available for single-stage configurations and mostly aeroelastic effects can be taken into account. More recent research focusses on extending these methods to include multiple stages. However, due to the significantly increased computational effort of the aeroelastic simulations when adding more stages to the models, these ROMs are rarely applied with the inclusion of multistage aeroelastic effects. It is therefore desirable to develop reduction methods which minimize the number of these simulations to reduce the computational cost and thereby enable analyses of rotors with multiple stages including aeroelastic effects. In this paper, a cyclic Craig-Bampton reduction method with an a priori interface reduction for multistage rotors is extended with an additional a posteriori interface reduction to reduce the number of aeroelastic simulations necessary for a given accuracy level of the ROM. The interface degrees of freedom between stages are reduced using a modified version of Characteristic Constraint Modes, to yield a more efficient representation of their displacements while retaining their monoharmonic nature. The method is applied to a two-stage axial compressor with full aeroelastic coupling between the stages and its reduced computational effort is demonstrated. Additionally, two sorting methods for the degrees of freedom (DOFs) of the ROM are compared.","PeriodicalId":53002,"journal":{"name":"Journal of the Global Power and Propulsion Society","volume":" ","pages":""},"PeriodicalIF":0.9,"publicationDate":"2023-04-25","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"47607149","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}
Variable pitch low pressure ratio fans could enable higher propulsive efficiency and eliminate the need for heavy thrust reversers. In this paper, the effects of the fan rotor design on reverse thrust capability have been explored by varying key parameters of NASA’s Advanced Ducted Propulsor (ADP) while maintaining cruise point performance. Reverse thrust performance has been assessed using RANS single passage CFD of the variable pitch fan system with an extended domain. This computational approach has been validated using NASA Stage 51B, an experimental variable pitch fan test case. Predicted total pressure and total pressure ratios for this case were found to agree with test data within experimental error, except where large tip region separations occurred at high incidence. Applying individual variations to rotor pitch-to-chord, radial loading distribution, and chordwise camber distribution generated changes to rotor incidence, blockage, and peak surface Mach numbers in reverse operation. An increase in gross reverse thrust of up to 8% was achieved through reductions in rotor pitch-to-chord due to improved loading and lower shock Mach numbers. Controlling section camber distributions was used to move the rotor shock downstream and was found to increase reverse thrust by up to 3%. Increasing rotor loading at the mid span relative to the tip resulted in high tip incidence and a 50% reduction in gross reverse thrust across all rotor speeds. This suggests that tip loaded designs are required for high levels of reverse thrust.
{"title":"Variable pitch fan aerodynamic design for reverse thrust operation","authors":"Tim Williams, Cesare Hall, Mark Wilson","doi":"10.33737/jgpps/160096","DOIUrl":"https://doi.org/10.33737/jgpps/160096","url":null,"abstract":"Variable pitch low pressure ratio fans could enable higher propulsive efficiency and eliminate the need for heavy thrust reversers. In this paper, the effects of the fan rotor design on reverse thrust capability have been explored by varying key parameters of NASA’s Advanced Ducted Propulsor (ADP) while maintaining cruise point performance. Reverse thrust performance has been assessed using RANS single passage CFD of the variable pitch fan system with an extended domain. This computational approach has been validated using NASA Stage 51B, an experimental variable pitch fan test case. Predicted total pressure and total pressure ratios for this case were found to agree with test data within experimental error, except where large tip region separations occurred at high incidence. Applying individual variations to rotor pitch-to-chord, radial loading distribution, and chordwise camber distribution generated changes to rotor incidence, blockage, and peak surface Mach numbers in reverse operation. An increase in gross reverse thrust of up to 8% was achieved through reductions in rotor pitch-to-chord due to improved loading and lower shock Mach numbers. Controlling section camber distributions was used to move the rotor shock downstream and was found to increase reverse thrust by up to 3%. Increasing rotor loading at the mid span relative to the tip resulted in high tip incidence and a 50% reduction in gross reverse thrust across all rotor speeds. This suggests that tip loaded designs are required for high levels of reverse thrust.","PeriodicalId":53002,"journal":{"name":"Journal of the Global Power and Propulsion Society","volume":"83 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-04-07","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135742547","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 hub and casing walls of axial compressors are often modeled as smooth continuous surfaces in CFD simulations, but in real geometries, non-smooth pinches, steps and leakage cavities may exist. In the GPPS first Turbomachinery CFD Workshop, a comprehensive validation and verification campaign of RANS flow solvers was conducted, and all the simulation results consistently over-predicted the total pressure ratio at the rotor exit near the casing and the stator exit near the hub. From a recent examination of the test rig geometry, a pinched casing wall over the rotor and a leakage cavity below the stator were found, which were not considered in the workshop. In this paper, the effects of these endwall geometric uncertainties and errors are analyzed via numerical simulation. When considering the rotor casing pinch of the test geometry, the predicted total pressure ratio and choke mass flow of the compressor stage are smaller than that without the pinch, leading to better agreement with the measured data. When considering a stator hub cavity with a leakage flow rate of about 0.2% of the compressor inlet mass flow, the near-hub total pressure ratio distribution matches slightly better with the experimental data, but the effects on the global compressor stage characteristics are not visible. The relevant mechanisms of these changes in performances are analyzed in detail. The updated geometries and grids will be released to the public as a benchmark test case for turbomachinery CFD validation and verification.
{"title":"Endwall geometric uncertainty and error on the performance of TUDA-GLR-OpenStage transonic axial compressor","authors":"Kailong Xia, Xiao He, Mingmin Zhu, Fabian Klausmann, Jinfang Teng, Mehdi Vahdati","doi":"10.33737/jgpps/161708","DOIUrl":"https://doi.org/10.33737/jgpps/161708","url":null,"abstract":"The hub and casing walls of axial compressors are often modeled as smooth continuous surfaces in CFD simulations, but in real geometries, non-smooth pinches, steps and leakage cavities may exist. In the GPPS first Turbomachinery CFD Workshop, a comprehensive validation and verification campaign of RANS flow solvers was conducted, and all the simulation results consistently over-predicted the total pressure ratio at the rotor exit near the casing and the stator exit near the hub. From a recent examination of the test rig geometry, a pinched casing wall over the rotor and a leakage cavity below the stator were found, which were not considered in the workshop. In this paper, the effects of these endwall geometric uncertainties and errors are analyzed via numerical simulation. When considering the rotor casing pinch of the test geometry, the predicted total pressure ratio and choke mass flow of the compressor stage are smaller than that without the pinch, leading to better agreement with the measured data. When considering a stator hub cavity with a leakage flow rate of about 0.2% of the compressor inlet mass flow, the near-hub total pressure ratio distribution matches slightly better with the experimental data, but the effects on the global compressor stage characteristics are not visible. The relevant mechanisms of these changes in performances are analyzed in detail. The updated geometries and grids will be released to the public as a benchmark test case for turbomachinery CFD validation and verification.","PeriodicalId":53002,"journal":{"name":"Journal of the Global Power and Propulsion Society","volume":"114 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-03-31","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135821939","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}
Jan Goeing, Hendrik Seehausen, Lennart Stania, Nicolas Nuebel, Julian Salomon, Panagiotis Ignatidis, Friedrich Dinkelacker, Michael Beer, Berend Berend, Joerg Seume, Jens Friedrichs
The effects of real combined variances in components and modules of aero engines, due to production tolerances or deterioration, on the performance of an aircraft engine are analysed in a knowledge-based process. For this purpose, an aero-thermodynamic virtual evaluation process that combines physical and probabilistic models to determine the sensitivities in the local module aerodynamics and the global overall performance is developed. Therefore, an automatic process that digitises, parameterises, reconstructs and analyses the geometry automatically using the example of a real turbofan high-pressure turbine blade is developed. The influence on the local aerodynamics of the reconstructed blade is investigated via a computational fluid dynamics (CFD) simulations. The results of the high-pressure turbine (HPT) CFD as well as of a Gas-Path-Analysis for further modules, such as the compressors and the low-pressure turbine, are transferred into a simulation of the performance of the whole aircraft engine to evaluate the overall performance. All results are used to train, validate and test several deep learning architectures. These metamodels are utilised for a global sensitivity analysis that is able to evaluate the sensitivities and interactions. On the one hand, the results show that the aerodynamics (especially the efficiency <inline-formula><mml:math xmlns:mml="http://www.w3.org/1998/Math/MathML" display="inline" overflow="scroll"><mml:msub><mml:mi>η</mml:mi><mml:mrow><mml:mi>H</mml:mi><mml:mi>P</mml:mi><mml:mi>T</mml:mi></mml:mrow></mml:msub></mml:math></inline-formula> and capacity <inline-formula><mml:math xmlns:mml="http://www.w3.org/1998/Math/MathML" display="inline" overflow="scroll"><mml:msub><mml:mrow><mml:mover><mml:mi>m</mml:mi><mml:mo>˙</mml:mo></mml:mover></mml:mrow><mml:mrow><mml:mi>H</mml:mi><mml:mi>P</mml:mi><mml:mi>T</mml:mi></mml:mrow></mml:msub></mml:math></inline-formula>) are particularly driven by the variation of the stagger angle. On the other hand, <inline-formula><mml:math xmlns:mml="http://www.w3.org/1998/Math/MathML" display="inline" overflow="scroll"><mml:msub><mml:mi>η</mml:mi><mml:mrow><mml:mi>H</mml:mi><mml:mi>P</mml:mi><mml:mi>T</mml:mi></mml:mrow></mml:msub></mml:math></inline-formula> is significantly related to exhaust gas temperature (Tt5), while specific fuel consumption (SFC) and mass flow <inline-formula><mml:math xmlns:mml="http://www.w3.org/1998/Math/MathML" display="inline" overflow="scroll"><mml:msub><mml:mrow><mml:mover><mml:mi>m</mml:mi><mml:mo>˙</mml:mo></mml:mover></mml:mrow><mml:mrow><mml:mi>H</mml:mi><mml:mi>P</mml:mi><mml:mi>T</mml:mi></mml:mrow></mml:msub></mml:math></inline-formula> are related to HPC exit temperature (Tt3). However, it can be seen that the high-pressure compressor has the most significant impact on the overall performance. This novel knowledge-based approach can accurately determine the impact of component variances on overall performance and complement experience-based approach
{"title":"Virtual process for evaluating the influence of real combined module variations on the overall performance of an aircraft engine","authors":"Jan Goeing, Hendrik Seehausen, Lennart Stania, Nicolas Nuebel, Julian Salomon, Panagiotis Ignatidis, Friedrich Dinkelacker, Michael Beer, Berend Berend, Joerg Seume, Jens Friedrichs","doi":"10.33737/jgpps/160055","DOIUrl":"https://doi.org/10.33737/jgpps/160055","url":null,"abstract":"The effects of real combined variances in components and modules of aero engines, due to production tolerances or deterioration, on the performance of an aircraft engine are analysed in a knowledge-based process. For this purpose, an aero-thermodynamic virtual evaluation process that combines physical and probabilistic models to determine the sensitivities in the local module aerodynamics and the global overall performance is developed. Therefore, an automatic process that digitises, parameterises, reconstructs and analyses the geometry automatically using the example of a real turbofan high-pressure turbine blade is developed. The influence on the local aerodynamics of the reconstructed blade is investigated via a computational fluid dynamics (CFD) simulations. The results of the high-pressure turbine (HPT) CFD as well as of a Gas-Path-Analysis for further modules, such as the compressors and the low-pressure turbine, are transferred into a simulation of the performance of the whole aircraft engine to evaluate the overall performance. All results are used to train, validate and test several deep learning architectures. These metamodels are utilised for a global sensitivity analysis that is able to evaluate the sensitivities and interactions. On the one hand, the results show that the aerodynamics (especially the efficiency <inline-formula><mml:math xmlns:mml=\"http://www.w3.org/1998/Math/MathML\" display=\"inline\" overflow=\"scroll\"><mml:msub><mml:mi>η</mml:mi><mml:mrow><mml:mi>H</mml:mi><mml:mi>P</mml:mi><mml:mi>T</mml:mi></mml:mrow></mml:msub></mml:math></inline-formula> and capacity <inline-formula><mml:math xmlns:mml=\"http://www.w3.org/1998/Math/MathML\" display=\"inline\" overflow=\"scroll\"><mml:msub><mml:mrow><mml:mover><mml:mi>m</mml:mi><mml:mo>˙</mml:mo></mml:mover></mml:mrow><mml:mrow><mml:mi>H</mml:mi><mml:mi>P</mml:mi><mml:mi>T</mml:mi></mml:mrow></mml:msub></mml:math></inline-formula>) are particularly driven by the variation of the stagger angle. On the other hand, <inline-formula><mml:math xmlns:mml=\"http://www.w3.org/1998/Math/MathML\" display=\"inline\" overflow=\"scroll\"><mml:msub><mml:mi>η</mml:mi><mml:mrow><mml:mi>H</mml:mi><mml:mi>P</mml:mi><mml:mi>T</mml:mi></mml:mrow></mml:msub></mml:math></inline-formula> is significantly related to exhaust gas temperature (Tt5), while specific fuel consumption (SFC) and mass flow <inline-formula><mml:math xmlns:mml=\"http://www.w3.org/1998/Math/MathML\" display=\"inline\" overflow=\"scroll\"><mml:msub><mml:mrow><mml:mover><mml:mi>m</mml:mi><mml:mo>˙</mml:mo></mml:mover></mml:mrow><mml:mrow><mml:mi>H</mml:mi><mml:mi>P</mml:mi><mml:mi>T</mml:mi></mml:mrow></mml:msub></mml:math></inline-formula> are related to HPC exit temperature (Tt3). However, it can be seen that the high-pressure compressor has the most significant impact on the overall performance. This novel knowledge-based approach can accurately determine the impact of component variances on overall performance and complement experience-based approach","PeriodicalId":53002,"journal":{"name":"Journal of the Global Power and Propulsion Society","volume":"163 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-03-13","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"135957411","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}
In this study, a sensor fault diagnostic system to detect/isolate and accommodate faults in sensors from an industrial gas turbine has been developed. The sensor fault diagnostic module is integrated with a gas turbine real-time executable digital-twin (RT xDT) reported in a previous study. The sensor fault diagnostic module of the digital-twin considers analytical sensor redundancy using a reference engine model to provide redundant estimates of measured engine variables. A Software-in-the-Loop (S-i-L) architecture and Hardware-in-the-Loop (H-i-L) facility are constructed to assess the sensor diagnostic module (fault detection/ fault isolation) during failure in sensors from the engine. The results demonstrated that if the discrepancy between virtual measurement (provided by digital-twin) and sensor measurement exceeds the prescribed tolerance levels, the sensor fault diagnostic logic determines the state of switching between the virtual and engine sensor measurements in a dual lane control configuration of the gas turbine control system. The sensor fault detection system implemented in the gas turbine RT xDT can be deployed onto a distributed control system of industrial gas turbines to diagnose sensor deficiencies and ensure continuous and safe operation of the gas turbine. Consequently, the developed system will increase engine availability and reliability by diagnosing engine operational deficiencies before severe failure.
{"title":"GAS turbine sensor fault diagnostic system in a real-time executable digital-twin","authors":"Samuel Cruz-Manzo, Vili Panov, Chris Bingham","doi":"10.33737/jgpps/159781","DOIUrl":"https://doi.org/10.33737/jgpps/159781","url":null,"abstract":"In this study, a sensor fault diagnostic system to detect/isolate and accommodate faults in sensors from an industrial gas turbine has been developed. The sensor fault diagnostic module is integrated with a gas turbine real-time executable digital-twin (RT xDT) reported in a previous study. The sensor fault diagnostic module of the digital-twin considers analytical sensor redundancy using a reference engine model to provide redundant estimates of measured engine variables. A Software-in-the-Loop (S-i-L) architecture and Hardware-in-the-Loop (H-i-L) facility are constructed to assess the sensor diagnostic module (fault detection/ fault isolation) during failure in sensors from the engine. The results demonstrated that if the discrepancy between virtual measurement (provided by digital-twin) and sensor measurement exceeds the prescribed tolerance levels, the sensor fault diagnostic logic determines the state of switching between the virtual and engine sensor measurements in a dual lane control configuration of the gas turbine control system. The sensor fault detection system implemented in the gas turbine RT xDT can be deployed onto a distributed control system of industrial gas turbines to diagnose sensor deficiencies and ensure continuous and safe operation of the gas turbine. Consequently, the developed system will increase engine availability and reliability by diagnosing engine operational deficiencies before severe failure.","PeriodicalId":53002,"journal":{"name":"Journal of the Global Power and Propulsion Society","volume":"136 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-03-09","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"136171837","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}
Starting and windmilling simulations with a normal gas turbine performance program require extended compressor and turbine maps which include sub-idle corrected speeds down to say 5–10% of the design value. During such simulations certain specific phenomena which are insignificant in the normal operating range between idle and full power must be considered. For example, while starting a low bypass ratio mixed flow turbofan, flow reversal in the bypass duct can occur. This paper illustrates a general understanding of what happens from when the starter is activated to when stabilized idle operation is reached. Operating lines in the compressor and turbine maps are predicted depending on starter torque, starter power, burner light-up and starter cut-off speed. It is explained why knowing combustor efficiency precisely is not required for that. Simulating engine starting and windmilling is not a magical art. The laws of physics still apply at these somewhat exotic operating conditions.
{"title":"Starting and windmilling simulations using compressor and turbine maps","authors":"Kurzke Joachim","doi":"10.33737/jgpps/159372","DOIUrl":"https://doi.org/10.33737/jgpps/159372","url":null,"abstract":"Starting and windmilling simulations with a normal gas turbine performance program require extended compressor and turbine maps which include sub-idle corrected speeds down to say 5–10% of the design value. During such simulations certain specific phenomena which are insignificant in the normal operating range between idle and full power must be considered. For example, while starting a low bypass ratio mixed flow turbofan, flow reversal in the bypass duct can occur. This paper illustrates a general understanding of what happens from when the starter is activated to when stabilized idle operation is reached. Operating lines in the compressor and turbine maps are predicted depending on starter torque, starter power, burner light-up and starter cut-off speed. It is explained why knowing combustor efficiency precisely is not required for that. Simulating engine starting and windmilling is not a magical art. The laws of physics still apply at these somewhat exotic operating conditions.","PeriodicalId":53002,"journal":{"name":"Journal of the Global Power and Propulsion Society","volume":" ","pages":""},"PeriodicalIF":0.9,"publicationDate":"2023-03-02","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"43153694","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}
Alejandro Castillo Pardo, Tim Williams, Christopher Clark, Nick Atkins, Cesare Hall, Mark Wilson, Raul Vazquez Diaz
Ultra-high bypass ratio turbofans offer significant reductions in fuel and pollution due to their higher propulsive efficiency. Short intakes might lead to a stronger fan-intake interaction, which creates uncertainty in stability at off-design conditions. Due to the prohibitive cost of full-scale experimental testing, subscale testing in wind tunnels is used to understand this behaviour. The low Reynolds number of subscale models results in unrepresentative laminar shock-boundary layer interactions. The boundary layer state thus needs to be conditioned to better represent full-scale transonic fans. This paper proposes the use of an inexpensive and robust flow control method for the suction side of a fan blade. Design guidelines are given for the location and height of the discrete roughness elements used to control the boundary layer state. This paper also presents a rapid experimental validation methodology to ensure and de-risk the application of the boundary layer trip to 3D rig blades. The experimental methodology is applied to a generic aerofoil representative of a fan tip section. The experimental method proves that it is possible to reproduce boundary layers and pressure distributions of a full-scale fan blade on a 1/10 subscale model. The results obtained confirm that the boundary layer trip method successfully promotes transition at the location representative of full-scale blades, avoiding unrepresentative laminar shock wave boundary layer interactions. This highlights the importance of conditioning boundary layers in low Reynolds number fan rig testing.
{"title":"Boundary layer control for low Reynolds number fan rig testing","authors":"Alejandro Castillo Pardo, Tim Williams, Christopher Clark, Nick Atkins, Cesare Hall, Mark Wilson, Raul Vazquez Diaz","doi":"10.33737/jgpps/158035","DOIUrl":"https://doi.org/10.33737/jgpps/158035","url":null,"abstract":"Ultra-high bypass ratio turbofans offer significant reductions in fuel and pollution due to their higher propulsive efficiency. Short intakes might lead to a stronger fan-intake interaction, which creates uncertainty in stability at off-design conditions. Due to the prohibitive cost of full-scale experimental testing, subscale testing in wind tunnels is used to understand this behaviour. The low Reynolds number of subscale models results in unrepresentative laminar shock-boundary layer interactions. The boundary layer state thus needs to be conditioned to better represent full-scale transonic fans. This paper proposes the use of an inexpensive and robust flow control method for the suction side of a fan blade. Design guidelines are given for the location and height of the discrete roughness elements used to control the boundary layer state. This paper also presents a rapid experimental validation methodology to ensure and de-risk the application of the boundary layer trip to 3D rig blades. The experimental methodology is applied to a generic aerofoil representative of a fan tip section. The experimental method proves that it is possible to reproduce boundary layers and pressure distributions of a full-scale fan blade on a 1/10 subscale model. The results obtained confirm that the boundary layer trip method successfully promotes transition at the location representative of full-scale blades, avoiding unrepresentative laminar shock wave boundary layer interactions. This highlights the importance of conditioning boundary layers in low Reynolds number fan rig testing.","PeriodicalId":53002,"journal":{"name":"Journal of the Global Power and Propulsion Society","volume":"133 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-02-08","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"136176844","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
This paper presents an overview of the most relevant fuel cell types and identifies the most promising options for application in propulsion systems for commercial electrified aviation. The general design, operating principles and main characteristics of polymer electrolyte membrane, alkaline, direct methanol, phosphoric acid, molten carbonate and solid oxide fuel cells are described. Evaluation criteria are derived from aviation-specific requirements for the application of fuel cells in electrified aircraft. Based on these criteria, the presented fuel cell types are evaluated by means of a weighted point rating. The results of this evaluation reveal the high potential for application of solid oxide, low-temperature and high-temperature polymer electrolyte membrane fuel cells. Design challenges of all fuel cell types are being emphasised, for instance, concerning cold start, cooling and supply of pressurised air.
{"title":"Review of fuel cell technologies and evaluation of their potential and challenges for electrified propulsion systems in commercial aviation","authors":"Stefan Kazula, Stefanie de Graaf, Lars Enghardt","doi":"10.33737/jgpps/158036","DOIUrl":"https://doi.org/10.33737/jgpps/158036","url":null,"abstract":"This paper presents an overview of the most relevant fuel cell types and identifies the most promising options for application in propulsion systems for commercial electrified aviation. The general design, operating principles and main characteristics of polymer electrolyte membrane, alkaline, direct methanol, phosphoric acid, molten carbonate and solid oxide fuel cells are described. Evaluation criteria are derived from aviation-specific requirements for the application of fuel cells in electrified aircraft. Based on these criteria, the presented fuel cell types are evaluated by means of a weighted point rating. The results of this evaluation reveal the high potential for application of solid oxide, low-temperature and high-temperature polymer electrolyte membrane fuel cells. Design challenges of all fuel cell types are being emphasised, for instance, concerning cold start, cooling and supply of pressurised air.","PeriodicalId":53002,"journal":{"name":"Journal of the Global Power and Propulsion Society","volume":"130 1","pages":"0"},"PeriodicalIF":0.0,"publicationDate":"2023-02-08","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"136174943","PeriodicalName":null,"FirstCategoryId":null,"ListUrlMain":null,"RegionNum":0,"RegionCategory":"","ArticlePicture":[],"TitleCN":null,"AbstractTextCN":null,"PMCID":"","EPubDate":null,"PubModel":null,"JCR":null,"JCRName":null,"Score":null,"Total":0}
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