Jan Goeing, H. Seehausen, V. Pak, S. Lueck, J. Seume, J. Friedrichs
In this study, numerical models are used to analyse the influence of isolated component deterioration as well as the combination of miscellaneous deteriorated components on the transient performance of a high-bypass jet engine. For this purpose, the aerodynamic impact of major degradation effects in a high-pressure compressor (HPC) and turbine (HPT) is modelled and simulated by using 3D CFD (Computational Fluid Dynamics). The impact on overall jet engine performance is then modelled using an 1D Reduced Order Model (ROM). Initially, the HPC performance is investigated with a typical level of roughness on vanes and blades and the HPT performance with an increasing tip clearance. Subsequently, the overall performance of the jet engines with the isolated and combined deteriorated domains is computed by the in-house 1D performance tool ASTOR (AircraftEngine Simulation for Transient Operation Research). Degradations have a significant influence on the system stability and transient effects. In ASTOR, a system of differential equations including the equations of motion and further ordinary differential equations is solved. Compared to common ROMs, this enables a higher degree of accuracy. The results of temperature downstream of the high-pressure compressor and low-pressure turbine as well as the specific fuel composition and the HP rotational speed are used to estimate the degree and type of engine deterioration. However, the consideration of the system stability is necessary to analyse the characterisation in more detail. Finally, a simplified model which merges two engines with individual deteriorated domains into one combined deteriorated engine, is proposed. The simplified model predicts the performance of an engine which has been simulated with combined deteriorated components.
本文采用数值模型分析了某大涵道比喷气发动机中单个部件劣化以及杂化部件劣化组合对发动机瞬态性能的影响。为此，利用三维CFD(计算流体动力学)对高压压气机(HPC)和涡轮(HPT)中主要退化效应的气动影响进行了建模和仿真。然后使用1D降阶模型(ROM)对喷气发动机整体性能的影响进行建模。首先，在叶片和叶片的典型粗糙度水平下，研究了HPC性能，并在叶尖间隙增加时研究了HPT性能。随后，使用公司内部的一维性能工具ASTOR (AircraftEngine Simulation for Transient Operation Research)计算具有隔离和组合恶化区域的喷气发动机的整体性能。退化对系统稳定性和瞬态效应有显著影响。在ASTOR中，求解一个微分方程组，包括运动方程和进一步的常微分方程。与普通rom相比，这可以实现更高程度的准确性。利用高压压气机和低压涡轮下游的温度、比燃料成分和马力转速来估计发动机劣化的程度和类型。然而，要更详细地分析表征，必须考虑系统的稳定性。最后，提出了一个简化模型，将两个具有单独退化域的引擎合并为一个组合退化引擎。该简化模型预测了某发动机组合劣化部件的性能。
{"title":"Influence of combined compressor and turbine deterioration on the overall performance of a jet engine using RANS simulation and Pseudo Bond Graph approach","authors":"Jan Goeing, H. Seehausen, V. Pak, S. Lueck, J. Seume, J. Friedrichs","doi":"10.33737/gpps20-tc-140","DOIUrl":"https://doi.org/10.33737/gpps20-tc-140","url":null,"abstract":"In this study, numerical models are used to analyse the influence of isolated component deterioration as well as the combination of miscellaneous deteriorated components on the transient performance of a high-bypass jet engine. For this purpose, the aerodynamic impact of major degradation effects in a high-pressure compressor (HPC) and turbine (HPT) is modelled and simulated by using 3D CFD (Computational Fluid Dynamics). The impact on overall jet engine performance is then modelled using an 1D Reduced Order Model (ROM). \u0000Initially, the HPC performance is investigated with a typical level of roughness on vanes and blades and the HPT performance with an increasing tip clearance. Subsequently, the overall performance of the jet engines with the isolated and combined deteriorated domains is computed by the in-house 1D performance tool ASTOR (AircraftEngine Simulation for Transient Operation Research). Degradations have a significant influence on the system stability and transient effects. In ASTOR, a system of differential equations including the equations of motion and further ordinary differential equations is solved. Compared to common ROMs, this enables a higher degree of accuracy.\u0000 The results of temperature downstream of the high-pressure compressor and low-pressure turbine as well as the specific fuel composition and the HP rotational speed are used to estimate the degree and type of engine deterioration. However, the consideration of the system stability is necessary to analyse the characterisation in more detail. \u0000Finally, a simplified model which merges two engines with individual deteriorated domains into one combined deteriorated engine, is proposed. The simplified model predicts the performance of an engine which has been simulated with combined deteriorated components.","PeriodicalId":53002,"journal":{"name":"Journal of the Global Power and Propulsion Society","volume":null,"pages":null},"PeriodicalIF":0.9,"publicationDate":"2020-09-09","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"44263642","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}
Tim Kluge, Iris Lettmann, M. Oettinger, Lars Wein, J. Seume
This paper presents those flow parameters at which coherent structures appear in the blade tip cavities of shrouded turbine blades. To the authors’ knowledge, this is reported for the first time in the open literature. The unsteady flow in a shroud cavity is analysed based on experimental data recorded in a labyrinth seal test rig. The unsteady static wall pressure in the shroud cavity inlet and outlet is measured using time-resolving pressure sensors. Sensors are located at staggered circumferential positions to allow cross-correlation between signals. The unsteady pressure signals are reduced using Fourier analysis and cross-correlation in combination with digital filters. Based on the data, a theory is formulated explaining the phenomena reflected in the measurements. The results suggest that pressure fluctuations with distinct numbers of nodes are rotating in the shroud cavity outlet. Moreover, modes with different node numbers appear to be superimposed, rotating at a common speed in circumferential direction. The pressure fluctuations are not found at all operating points. Further analysis indicates that the pressure fluctuations are present at operating points matching distinct parameters correlating with the cavity flow coefficient. Unsteady RANS simulations predict similar flow structures for the design operating point of the test rig.
{"title":"Unsteady Flow Phenomena in Turbine Shroud Cavities","authors":"Tim Kluge, Iris Lettmann, M. Oettinger, Lars Wein, J. Seume","doi":"10.33737/gpps20-tc-64","DOIUrl":"https://doi.org/10.33737/gpps20-tc-64","url":null,"abstract":"This paper presents those flow parameters at which coherent structures appear in the blade tip cavities of shrouded turbine blades. To the authors’ knowledge, this is reported for the first time in the open literature. The unsteady flow in a shroud cavity is analysed based on experimental data recorded in a labyrinth seal test rig. The unsteady static wall pressure in the shroud cavity inlet and outlet is measured using time-resolving pressure sensors. Sensors are located at staggered circumferential positions to allow cross-correlation between signals. The unsteady pressure signals are reduced using Fourier analysis and cross-correlation in combination with digital filters. Based on the data, a theory is formulated explaining the phenomena reflected in the measurements. The results suggest that pressure fluctuations with distinct numbers of nodes are rotating in the shroud cavity outlet. Moreover, modes with different node numbers appear to be superimposed, rotating at a common speed in circumferential direction. The pressure fluctuations are not found at all operating points. Further analysis indicates that the pressure fluctuations are present at operating points matching distinct parameters correlating with the cavity flow coefficient. Unsteady RANS simulations predict similar flow structures for the design operating point of the test rig.","PeriodicalId":53002,"journal":{"name":"Journal of the Global Power and Propulsion Society","volume":null,"pages":null},"PeriodicalIF":0.9,"publicationDate":"2020-09-09","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"44416580","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. Roy, J. Fridh, James A. Scobie, C. Sangan, G. Lock
This paper investigates flow instabilities inside the cavity formed between the stator and rotor disks of a high-speed turbine rig. The cavity rim seal is of chute seal design. The influence of flow coefficient on the sealing effectiveness at constant purge flow rate through the wheel-space is determined. The effectiveness at different radial positions over a range of purge flow conditions and flow coefficients is also studied. Unsteady pressure measurements have identified the frequency of instabilities that form within the rim seal, phenomena which have been observed in other studies. Frequencies of these disturbances, and their correlation in the circumferential direction have determined the strength and speed of rotation of the instabilities within the cavity. Large scale unsteady rotational structures have been identified, which show similarity to previous studies. These disturbances have been found to be weakly dependent on the purge flow and flow coefficients, although an increased purge reduced both the intensity and speed of rotation of the instabilities. Additionally, certain uncorrelated disturbances have been found to be inconsistent (discontinuous) with pitchwise variation.
{"title":"Flow instability effects related to purge through a gas turbine chute seal","authors":"A. Roy, J. Fridh, James A. Scobie, C. Sangan, G. Lock","doi":"10.33737/gpps20-tc-82","DOIUrl":"https://doi.org/10.33737/gpps20-tc-82","url":null,"abstract":"This paper investigates flow instabilities inside the cavity formed between the stator and rotor disks of a high-speed turbine rig. The cavity rim seal is of chute seal design. The influence of flow coefficient on the sealing effectiveness at constant purge flow rate through the wheel-space is determined. The effectiveness at different radial positions over a range of purge flow conditions and flow coefficients is also studied. Unsteady pressure measurements have identified the frequency of instabilities that form within the rim seal, phenomena which have been observed in other studies. Frequencies of these disturbances, and their correlation in the circumferential direction have determined the strength and speed of rotation of the instabilities within the cavity. Large scale unsteady rotational structures have been identified, which show similarity to previous studies. These disturbances have been found to be weakly dependent on the purge flow and flow coefficients, although an increased purge reduced both the intensity and speed of rotation of the instabilities. Additionally, certain uncorrelated disturbances have been found to be inconsistent (discontinuous) with pitchwise variation.","PeriodicalId":53002,"journal":{"name":"Journal of the Global Power and Propulsion Society","volume":null,"pages":null},"PeriodicalIF":0.9,"publicationDate":"2020-09-09","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"45515777","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}
Mauro Righi, V. Pachidis, L. Könözsy, F. Zhao, M. Vahdati
Surge in modern aero-engines can lead to violent disruption of the flow, damage to the blade structures and eventually engine shutdown. Knowledge of unsteady performance and loading during surge is crucial for compressor design, however, the understanding and prediction capability for this phenomenon is still very limited. While useful for the investigation of specific cases, costly experimental tests and high-fidelity CFD simulations cannot be used routinely in the design process of compressor systems. There is therefore an interest in developing a low-order model which can predict compressor performance during surge with sufficient accuracy at significantly reduced computational cost. This paper describes the validation of an unsteady 3D through-flow code developed at Cranfield University for the low-order modelling of surge in axial compressors. The geometry investigated is an 8-stage rig representative of a modern aero-engine IP compressor. Two deep surge events are modelled at part speed and full speed respectively. The results are compared against high-fidelity, full annulus, URANS simulations conducted at Imperial College. Comparison of massflow, pressure and temperature time histories shows a close match between the low-order and the higher-fidelity methods. The low-order model is shown capable of predicting many transient flow features which were observed in the high-fidelity simulations, while reducing the computational cost by up to two orders of magnitude.
{"title":"Three-dimensional low-order surge model for high-speed axial compressors","authors":"Mauro Righi, V. Pachidis, L. Könözsy, F. Zhao, M. Vahdati","doi":"10.33737/gpps20-tc-138","DOIUrl":"https://doi.org/10.33737/gpps20-tc-138","url":null,"abstract":"Surge in modern aero-engines can lead to violent disruption of the flow, damage to the blade structures and eventually engine shutdown. Knowledge of unsteady performance and loading during surge is crucial for compressor design, however, the understanding and prediction capability for this phenomenon is still very limited. While useful for the investigation of specific cases, costly experimental tests and high-fidelity CFD simulations cannot be used routinely in the design process of compressor systems. There is therefore an interest in developing a low-order model which can predict compressor performance during surge with sufficient accuracy at significantly reduced computational cost.\u0000This paper describes the validation of an unsteady 3D through-flow code developed at Cranfield University for the low-order modelling of surge in axial compressors. The geometry investigated is an 8-stage rig representative of a modern aero-engine IP compressor. Two deep surge events are modelled at part speed and full speed respectively. The results are compared against high-fidelity, full annulus, URANS simulations conducted at Imperial College. Comparison of massflow, pressure and temperature time histories shows a close match between the low-order and the higher-fidelity methods. The low-order model is shown capable of predicting many transient flow features which were observed in the high-fidelity simulations, while reducing the computational cost by up to two orders of magnitude.","PeriodicalId":53002,"journal":{"name":"Journal of the Global Power and Propulsion Society","volume":null,"pages":null},"PeriodicalIF":0.9,"publicationDate":"2020-09-09","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"46700654","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 dependence of the aerodynamic stability of fan blades with amplitude and nodal diameter of potential perturbations associated with the presence of pylons is studied. The analysis is conducted using a novel block-wise spatial Fourier decomposition of the reduced-passages to reconstruct the full-annulus solution. The method represents very efficiently unsteady flows generated by outlet static pressure non-uniformities. The explicit spatial Fourier approximation is exploited to characterize the relevance of each nodal diameter of outlet perturbations in the fan stall process, and its nonlinear stability is studied in a harmonic by harmonic basis filtering the nonlinear contribution of the rest. The methodology has been assessed for the NASA rotor 67. The maximum amplitude of the downstream perturbation at which the compressor becomes unstable and triggers a stall process has been mapped. It is concluded that the fan stability dependence with the amplitude of the perturbation is weaker than in the case of intake distortion. For perturbations with an odd number of nodal diameters, the nonlinear stability analysis leads to the same conclusions as to the small amplitude linear stability analysis. However, if the perturbations have an even number nodal diameters, the flow exhibits a supercritical bifurcation and have a stabilizing effect.
{"title":"IMPACT OF DOWNSTREAM POTENTIAL PERTURBATIONS ON THE NONLINEAR STABILITY OF A GENERIC FAN","authors":"D. Romera, R. Corral","doi":"10.33737/gpps20-tc-144","DOIUrl":"https://doi.org/10.33737/gpps20-tc-144","url":null,"abstract":"The dependence of the aerodynamic stability of fan blades with amplitude and nodal diameter of potential perturbations associated with the presence of pylons is studied. The analysis is conducted using a novel block-wise spatial Fourier decomposition of the reduced-passages to reconstruct the full-annulus solution. The method represents very efficiently unsteady flows generated by outlet static pressure non-uniformities. The explicit spatial Fourier approximation is exploited to characterize the relevance of each nodal diameter of outlet perturbations in the fan stall process, and its nonlinear stability is studied in a harmonic by harmonic basis filtering the nonlinear contribution of the rest. The methodology has been assessed for the NASA rotor 67. The maximum amplitude of the downstream perturbation at which the compressor becomes unstable and triggers a stall process has been mapped. It is concluded that the fan stability dependence with the amplitude of the perturbation is weaker than in the case of intake distortion. For perturbations with an odd number of nodal diameters, the nonlinear stability analysis leads to the same conclusions as to the small amplitude linear stability analysis. However, if the perturbations have an even number nodal diameters, the flow exhibits a supercritical bifurcation and have a stabilizing effect.","PeriodicalId":53002,"journal":{"name":"Journal of the Global Power and Propulsion Society","volume":null,"pages":null},"PeriodicalIF":0.9,"publicationDate":"2020-09-09","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"46792894","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}
An experimental study was conducted for evaluating the feasibility of using silica aerogel as thermal insulator for combustor liners. Aerogels are a superior material for minimizing heat flux to the metal structure of the combustion liner due to their low thermal conductivity. In this study, a conical natural gas fired swirling-flame combustor was utilized for reproducing the combustion environment. The silica aerogel blanket was attached to the inner side of a perforated combustor liner. Temperature distribution on the outer side of the combustion liner was measured using a calibrated IR camera. To create a protective cooling film over the aerogel surface, cooling air was supplied from the back side of the perforated metal liner and was allowed to penetrate the silica aerogel blanket to be discharged to the combustor. As the combustor was operated at a fixed equivalence ratio of 0.83, cooling air flow rates were varied to evaluate the effectiveness of transpiration cooling on the aerogel blanket as various cooling flow rates. The measured evolution of temperature distribution confirmed thermal equilibriums for every test condition with transpiration cooling. The measured temperature distribution of metal liner demonstrated superior thermal insulation of aerogel blanket under the protection of cooling film with a temperature difference as high as 1580 K between combustion products temperature and the metal liner temperature on the back side. In addition, silica aerogel samples were examined before and after the combustion tests to understand their material degradation exposing to a typical gas turbine combustor environment using high-resolution scanning electron microscope (SEM). Test results suggest multiple degradation mechanisms to the silica aerogel blanket samples from the combustion tests. Improvements can be made to the silica aerogel blankets for a more resilient thermal insulator, for example, by replacing glass fibers in silica aerogels.
{"title":"Experimental Evaluation of Using Silica Aerogels as the Thermal Insulator for Combustor Liners","authors":"Yeongmin Pyo, T. Robertson, Sean Yun, Zekai Hong","doi":"10.33737/gpps20-tc-46","DOIUrl":"https://doi.org/10.33737/gpps20-tc-46","url":null,"abstract":"An experimental study was conducted for evaluating the feasibility of using silica aerogel as thermal insulator for combustor liners. Aerogels are a superior material for minimizing heat flux to the metal structure of the combustion liner due to their low thermal conductivity. In this study, a conical natural gas fired swirling-flame combustor was utilized for reproducing the combustion environment. The silica aerogel blanket was attached to the inner side of a perforated combustor liner. Temperature distribution on the outer side of the combustion liner was measured using a calibrated IR camera. To create a protective cooling film over the aerogel surface, cooling air was supplied from the back side of the perforated metal liner and was allowed to penetrate the silica aerogel blanket to be discharged to the combustor. As the combustor was operated at a fixed equivalence ratio of 0.83, cooling air flow rates were varied to evaluate the effectiveness of transpiration cooling on the aerogel blanket as various cooling flow rates. \u0000\u0000The measured evolution of temperature distribution confirmed thermal equilibriums for every test condition with transpiration cooling. The measured temperature distribution of metal liner demonstrated superior thermal insulation of aerogel blanket under the protection of cooling film with a temperature difference as high as 1580 K between combustion products temperature and the metal liner temperature on the back side. In addition, silica aerogel samples were examined before and after the combustion tests to understand their material degradation exposing to a typical gas turbine combustor environment using high-resolution scanning electron microscope (SEM). Test results suggest multiple degradation mechanisms to the silica aerogel blanket samples from the combustion tests. Improvements can be made to the silica aerogel blankets for a more resilient thermal insulator, for example, by replacing glass fibers in silica aerogels.","PeriodicalId":53002,"journal":{"name":"Journal of the Global Power and Propulsion Society","volume":null,"pages":null},"PeriodicalIF":0.9,"publicationDate":"2020-09-09","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"49015917","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}
Experimental studies have shown that tip injection upstream of the rotor can extend its operational range when subjected to circumferential inlet distortion. Typically, injectors are placed uniformly around the annulus. However, such arrangement consumes a large amount of high-pressure air and decreases the overall efficiency of the compression system. The aim of this paper is to minimise the amount of the injected air by determining the most effective circumferential location for the injector. In this study, NASA stage 35 was used as the test case. The experiment was conducted with a circumferential total pressure distortion of 120 degrees. In the first part of this paper, numerical simulations were compared against the experimental data and good match was obtained. In the second part, tip injection at three different positions were tested: the clean flow region (Position 1), the distorted region (Position 2) and the border between the clean and distorted regions (Position 3). It was found that a mild injection (0.66% of the main flow) at Position 2 and Position 3 can extend the stall margin by 1.8% and 2.7%, respectively. No obvious improvement was observed for the injection at Position 1. With a larger injection of 1.5% of main flow at Position 3, the stall margin improved further with no efficiency loss.
{"title":"Investigation of the tip injection for stall control in a transonic compressor with inlet distortion","authors":"Wenqiang Zhang, M. Vahdati","doi":"10.33737/gpps20-tc-22","DOIUrl":"https://doi.org/10.33737/gpps20-tc-22","url":null,"abstract":"Experimental studies have shown that tip injection upstream of the rotor can extend its operational range when subjected to circumferential inlet distortion. Typically, injectors are placed uniformly around the annulus. However, such arrangement consumes a large amount of high-pressure air and decreases the overall efficiency of the compression system. The aim of this paper is to minimise the amount of the injected air by determining the most effective circumferential location for the injector.\u0000In this study, NASA stage 35 was used as the test case. The experiment was conducted with a circumferential total pressure distortion of 120 degrees. In the first part of this paper, numerical simulations were compared against the experimental data and good match was obtained. In the second part, tip injection at three different positions were tested: the clean flow region (Position 1), the distorted region (Position 2) and the border between the clean and distorted regions (Position 3). It was found that a mild injection (0.66% of the main flow) at Position 2 and Position 3 can extend the stall margin by 1.8% and 2.7%, respectively. No obvious improvement was observed for the injection at Position 1. With a larger injection of 1.5% of main flow at Position 3, the stall margin improved further with no efficiency loss.","PeriodicalId":53002,"journal":{"name":"Journal of the Global Power and Propulsion Society","volume":null,"pages":null},"PeriodicalIF":0.9,"publicationDate":"2020-09-09","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"46251189","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}
Stavros Vouros, Mavroudis D. Kavvalos, Smruti Sahoo, K. Kyprianidis
Hybrid-electric propulsion has emerged as a promising technology to mitigate the adverse environmental impact of civil aviation. Boosting conventional gas turbines with electric power improves mission performance and operability. In this work the impact of electrification on pollutant emissions and direct operating cost of geared turbofan configurations is evaluated for an 150-passenger aircraft. A baseline two-and-a-half-shaft geared turbofan, representative of year 2035 entry-into-service technology, is employed. Parallel hybridization is implemented through coupling a battery-powered electric motor to the engine low-speed shaft. A multi-disciplinary design space exploration framework is employed comprising modelling methods for multi-point engine design, aircraft sizing, performance and pollutant emissions, mission and economic analysis. A probabilistic approach is developed considering uncertainties in the evaluation of direct operating cost. Sensitivities to electrical power system technology levels, as well as fuel price and emissions taxation are quantified at different time-frames. The benefits of lean direct injection are explored along short-, medium-, and long-range missions, demonstrating 32% NOx savings compared to traditional rich-burn, quick-mix, lean-burn technologies in short-range operations. The impact of electrification on the enhancement of lean direct injection benefits is investigated. For hybrid-electric powerplants, the take-off-to-cruise turbine entry temperature ratio is 2.5% lower than the baseline, extending the corresponding NOx reductions to the level of 46% in short-range missions. This work sheds light on the environmental and economic potential and limitations of a hybrid-electric propulsion concept towards a greener and sustainable civil aviation.
{"title":"Enabling the Potential of Hybrid Electric Propulsion Through Lean-Burn-Combustion Turbofans","authors":"Stavros Vouros, Mavroudis D. Kavvalos, Smruti Sahoo, K. Kyprianidis","doi":"10.33737/gpps20-tc-86","DOIUrl":"https://doi.org/10.33737/gpps20-tc-86","url":null,"abstract":"Hybrid-electric propulsion has emerged as a promising technology to mitigate the adverse environmental impact of civil aviation. Boosting conventional gas turbines with electric power improves mission performance and operability. In this work the impact of electrification on pollutant emissions and direct operating cost of geared turbofan configurations is evaluated for an 150-passenger aircraft. A baseline two-and-a-half-shaft geared turbofan, representative of year 2035 entry-into-service technology, is employed. Parallel hybridization is implemented through coupling a battery-powered electric motor to the engine low-speed shaft. A multi-disciplinary design space exploration framework is employed comprising modelling methods for multi-point engine design, aircraft sizing, performance and pollutant emissions, mission and economic analysis. A probabilistic approach is developed considering uncertainties in the evaluation of direct operating cost. Sensitivities to electrical power system technology levels, as well as fuel price and emissions taxation are quantified at different time-frames. The benefits of lean direct injection are explored along short-, medium-, and long-range missions, demonstrating 32% NOx savings compared to traditional rich-burn, quick-mix, lean-burn technologies in short-range operations. The impact of electrification on the enhancement of lean direct injection benefits is investigated. For hybrid-electric powerplants, the take-off-to-cruise turbine entry temperature ratio is 2.5% lower than the baseline, extending the corresponding NOx reductions to the level of 46% in short-range missions. This work sheds light on the environmental and economic potential and limitations of a hybrid-electric propulsion concept towards a greener and sustainable civil aviation.","PeriodicalId":53002,"journal":{"name":"Journal of the Global Power and Propulsion Society","volume":null,"pages":null},"PeriodicalIF":0.9,"publicationDate":"2020-09-09","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"45208940","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 combustor-turbine interface in a gas turbine is characterised by complex, highly unsteady flows. In a combined experimental and large eddy simulation (LES) study including realistic combustor geometry, the standard model of secondary flows in the nozzle guide vanes (NGV) is found to be oversimplified. A swirl core is created in the combustion chamber which convects into the first vane passages. Four main consequences of this are identified: variation in vane loading; unsteady heat transfer on vane surfaces; unsteadiness at the leading edge horseshoe vortex, and variation in the position of the passage vortex. These phenomena occur at relatively low frequencies, from 50–300 Hz. It seems likely that these unsteady phenomena result in non-optimal film cooling, and that by reducing unsteadiness designs with greater cooling efficiency could be achieved. Measurements were performed in a high speed test facility modelling a large industrial gas turbine with can combustors, including nozzle guide vanes and combustion chambers. Vane surfaces and endwalls of a nozzle guide vane were instrumented with 384 high speed thin film heat flux gauges, to measure unsteady heat transfer. The high resolution of measurements was such to allow direct visualisation in time of large scale turbulent structures over the endwalls and vane surfaces. A matching LES simulation was carried out in a domain matching experimental conditions including upstream swirl generators and transition duct. Data reduction allowed time-varying LES data to be recorded for several cycles of the unsteady phenomena observed. The combination of LES and experimental data allows physical explanation and visualisation of flow events.
{"title":"Unsteady Phenomena at the Combustor-Turbine Interface","authors":"Faisal Shaikh, B. Rosic","doi":"10.33737/gpps20-tc-150","DOIUrl":"https://doi.org/10.33737/gpps20-tc-150","url":null,"abstract":"The combustor-turbine interface in a gas turbine is characterised by complex, highly unsteady flows. In a combined experimental and large eddy simulation (LES) study including realistic combustor geometry, the standard model of secondary flows in the nozzle guide vanes (NGV) is found to be oversimplified. A swirl core is created in the combustion chamber which convects into the first vane passages. Four main consequences of this are identified: variation in vane loading; unsteady heat transfer on vane surfaces; unsteadiness at the leading edge horseshoe vortex, and variation in the position of the passage vortex. These phenomena occur at relatively low frequencies, from 50–300 Hz. It seems likely that these unsteady phenomena result in non-optimal film cooling, and that by reducing unsteadiness designs with greater cooling efficiency could be achieved.\u0000\u0000Measurements were performed in a high speed test facility modelling a large industrial gas turbine with can combustors, including nozzle guide vanes and combustion chambers. Vane surfaces and endwalls of a nozzle guide vane were instrumented with 384 high speed thin film heat flux gauges, to measure unsteady heat transfer. The high resolution of measurements was such to allow direct visualisation in time of large scale turbulent structures over the endwalls and vane surfaces.\u0000\u0000A matching LES simulation was carried out in a domain matching experimental conditions including upstream swirl generators and transition duct. Data reduction allowed time-varying LES data to be recorded for several cycles of the unsteady phenomena observed. The combination of LES and experimental data allows physical explanation and visualisation of flow events.","PeriodicalId":53002,"journal":{"name":"Journal of the Global Power and Propulsion Society","volume":null,"pages":null},"PeriodicalIF":0.9,"publicationDate":"2020-09-09","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"44319742","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}
Pub Date : 2020-07-15DOI: 10.32743/2658-4077.2020.7.16.340
Ruslana Marockina
The article considers the semantic transformation of the term “leader” over the course of five decades of the twentieth century. The article touches on the interdisciplinary field of linguistic anthropology and history. The term "leader" was undoubtedly one of the most relevant terms in the USSR, which allows the use of rich source material to study it. The ability of people to follow the change in political sentiments in the country reflected in the vocabulary of the language, preserved in historical documents. The author of the article gives several interpretations of the meaning of the term “leader”, indicated in the domestic explanatory dictionaries, and, citing examples from history, explains the reasons for the semantic transformation of this term in the periods from 1917 to the 60s. ХХ century. Ключевые слова: лингвистическая антропология, общественно-политический термин, смысловая трансформация термина «вождь», В. Ленин, И. Сталин, Н. Хрущёв.
{"title":"CONSIDERATION OF SEMANTIC TRANSFORMATIONS OF THE TERM \"LEADER\" IN HISTORICAL DOCUMENTS USING LINGUISTIC ANTHROPOLOGY","authors":"Ruslana Marockina","doi":"10.32743/2658-4077.2020.7.16.340","DOIUrl":"https://doi.org/10.32743/2658-4077.2020.7.16.340","url":null,"abstract":"The article considers the semantic transformation of the term “leader” over the course of five decades of the twentieth century. The article touches on the interdisciplinary field of linguistic anthropology and history. The term \"leader\" was undoubtedly one of the most relevant terms in the USSR, which allows the use of rich source material to study it. The ability of people to follow the change in political sentiments in the country reflected in the vocabulary of the language, preserved in historical documents. The author of the article gives several interpretations of the meaning of the term “leader”, indicated in the domestic explanatory dictionaries, and, citing examples from history, explains the reasons for the semantic transformation of this term in the periods from 1917 to the 60s. ХХ century. Ключевые слова: лингвистическая антропология, общественно-политический термин, смысловая трансформация термина «вождь», В. Ленин, И. Сталин, Н. Хрущёв.","PeriodicalId":53002,"journal":{"name":"Journal of the Global Power and Propulsion Society","volume":null,"pages":null},"PeriodicalIF":0.9,"publicationDate":"2020-07-15","publicationTypes":"Journal Article","fieldsOfStudy":null,"isOpenAccess":false,"openAccessPdf":"","citationCount":null,"resultStr":null,"platform":"Semanticscholar","paperid":"79293250","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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