Relevant bibliographies by topics / Aero-thermal model

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  1. Journal articles
  2. Dissertations / Theses
  3. Conference papers

Academic literature on the topic 'Aero-thermal model'

Author: Grafiati
Published: 4 June 2021
Last updated: 16 February 2022

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Journal articles on the topic "Aero-thermal model"

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Ba, Wei, Chunwei Gu, Xiaodong Ren, and Xuesong Li. "Convective cooling model for aero-thermal coupled through-flow method." Proceedings of the Institution of Mechanical Engineers, Part A: Journal of Power and Energy 231, no. 2 (January 6, 2017): 133–44. http://dx.doi.org/10.1177/0957650916685911.

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The aero-thermal coupled phenomenon is significant in the modern cooled turbine, and it is necessary to consider the cooling effect and predict the coolant requirement in the through-flow design. A new cooling model was developed for the aero-thermal coupled through-flow method in this paper to predict the temperatures of both the pressure and suction surfaces of the blade. Based on the given blade temperature limitation rather than the mean blade surface temperature in the formal cooling model, the coolant requirement prediction can be more accurate. The equivalent blade thickness and heat exchange area estimation methods were further developed for blades with different cooling structures, and the estimations were carried out for each calculation station instead of the whole blade. The cooled blade was divided into a few calculation stations, and the heat transfer was studied for each station. Three operating conditions for the NASA-Mark II vane were selected for the verification. The predicted temperatures of both the pressure and suction surfaces agree with the experimental data, and the calculation results for the subsonic conditions are more accurate than the one for the transonic conditions.
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Dai, Huaren, Zhe Chen, Wei Guo, and Ju Wang. "Thermal simulation model of aero-engine blade material forging simulation." Thermal Science 25, no. 4 Part B (2021): 3169–77. http://dx.doi.org/10.2298/tsci2104169d.

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During the high temperature forging process, the thermal parameters such as the temperature field and strain field in the blank have an important influence on the crack damage and micro-structure in the forging. We use the rigid viscoelastic finite element method to carry out the forging process of a heavy aero-engine blade the finite element numerical simulation was carried out to obtain the temperature field, strain field and forging load change law in the forging process with time, and on this basis, combined with the crack damage and repair mechanism and the re?crystallization structure evolution law, an optimization was proposed. The forging process plan. That is, the pre-forging is performed on the basis of the tolerance of the final forging dimension under pressure of 4 mm, the pre-forging temperature is 1160?C, and the final forging temperature is 1120?C. The actual forging process test verifies the feasibility of the process plan, which is the engineering of this process the application lays the scientific foundation.
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Qian, Wei, Yuguang Bai, Xiangyan Chen, and Taojun Lu. "Aero-servo-elastic analysis of a hypersonic aircraft." Journal of Low Frequency Noise, Vibration and Active Control 37, no. 3 (August 23, 2017): 534–53. http://dx.doi.org/10.1177/1461348417725956.

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Aero-servo-elastic analysis of a complex hypersonic aircraft is presented in this paper. A structure geometry was designed and built based on the X-43A vehicle. First, a three-dimensional structural finite element model was proposed with effective two-dimensional elements, which can obtain effective modal analysis results without useless local modes. Second, computational fluid dynamic (CFD) simulation was adopted to find aero-heating distribution of thermal mode via this structure. Aero-heating effect was included to study thermal-modal characteristics of the present structure. Influence due to material characteristic change and thermal stress was studied. After structural finite element analysis was completed, flutter of the present vehicle was investigated. Aero-servo-elastic analysis was then started from the definition of an aero-servo-elastic closed-loop system. In this system, the present aircraft is treated as flexible structure, in which the control sensor on the aircraft received not only rigid motion signal but also elastic vibration signal, and this signal can translate into the deflection signal to form aerodynamic control force through this aero-servo control system, and this force can continually influence aerodynamic force. One of the most important steps for this analysis was computation of unsteady aerodynamic force of the present structure, and the related process was developed based on an effective fitting method. Finally, bode diagrams of pitching, rolling and yawing were investigated, form which the law of aero-servo stability of the X-43A vehicle can be observed and analyzed. It can be found from the results of this paper that effective investigation of aero-servo-elastic characteristics of a complex hypersonic aircraft should be based on accurate structural finite element modeling, modal analysis and flutter analysis. The proposed method in this paper can provide effective analysis process for the design of controller for hypersonic aircraft.
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Ba, Wei, Xuesong Li, Xiaodong Ren, and Chunwei Gu. "Aero-thermal coupled through-flow method for cooled turbines with new cooling model." Proceedings of the Institution of Mechanical Engineers, Part A: Journal of Power and Energy 232, no. 3 (September 25, 2017): 254–65. http://dx.doi.org/10.1177/0957650917731629.

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The aero-thermal–coupled phenomenon is significant in modern cooled turbines, and an aero-thermal coupled through-flow method has previously been developed by the authors for considering the influence of heat transfer and coolant mixing in through-flow design. However, the original cooling model is not capable of calculating the distribution of the coolant mass flow rate and pressure loss in complex cooling structures. Therefore, in this paper, a one-dimensional flow calculation for the internal coolant is introduced into the heat transfer calculation to further improve the through-flow cooling model. Based on various empirical correlations, the cooling model can be used to simulate different cooling structures, such as ribbed channels and cooling holes. Three operating conditions were selected for verification of the NASA-C3X vane, which has 10 internal radial cooling channels. The calculated Nusselt number of internal cooling channels strongly agrees with the experimental data, and the predicted blade surface pressure and temperature distributions at mid span are also in good agreement with the experimental data. The convergence history of the meridional velocity and blade surface temperature demonstrates effective convergence properties. Therefore, the aero-thermal–coupled through-flow method with the new cooling model can provide a reliable tool for cooled turbine through-flow design and analysis.
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Yazar, Isil, Tolga Yasa, and Emre Kiyak. "Simulation-based steady-state aero-thermal model for small-scale turboprop engine." Aircraft Engineering and Aerospace Technology 89, no. 2 (March 6, 2017): 203–10. http://dx.doi.org/10.1108/aeat-02-2015-0062.

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Purpose An aircraft engine control system consists of a large scale of control parameters and variables because of the complex structure of aero-engine. Monitoring and adjusting control variables and parameters such as detecting, isolating and reconfiguring the system faults/failures depend on the controller design. Developing a robust controller is based on an accurate mathematical model. Design/methodology/approach In this study, a small-scale turboprop engine is modeled. Simulation is carried out on MATLAB/Simulink for design and off-design operating conditions. Both steady-state and transient conditions (from idle to maximum thrust levels) are tested. The performance parameters of compressor and turbine components are predicted via trained Neuro-Fuzzy model (ANFIS) based on component maps. Temperature, rotational speed, mass flow, pressure and other parameters are generated by using thermodynamic formulas and conservation laws. Considering these calculated values, error calculations are made and compared with the cycle data of the engine at the related simulation conditions. Findings Simulation results show that the designed engine model’s simulation values have acceptable accuracy for both design and off-design conditions from idle to maximum power operating envelope considering cycle data. The designed engine model can be adapted to other types of gas turbine engines. Originality/value Different from other literature studies, in this work, a small-scale turboprop engine is modeled. Furthermore, for performance prediction of compressor and turbine components, ANFIS structure is applied.
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Jafari, Soheil, Ahmed Bouchareb, and Theoklis Nikolaidis. "Thermal Performance Evaluation in Gas Turbine Aero Engines Accessory Gearbox." International Journal of Turbomachinery, Propulsion and Power 5, no. 3 (August 26, 2020): 21. http://dx.doi.org/10.3390/ijtpp5030021.

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This paper presents a methodological approach for mathematical modelling and physics-based analysis of accessory gearbox (AGB) thermal behavior in gas turbine aero engines. The AGB structure, as one of the main sources of heat in gas turbine aero engines, is firstly described and its power losses will be divided into load-dependent and no-load dependent parts. Different mechanisms of heat generation are then identified and formulated to develop a toolbox for calculation of the churning, sliding friction, and rolling friction losses between contact surfaces of the AGB. The developed tool is also capable of calculating the heat loss mechanisms in different elements of the AGB, such as gears, bearings, and seals. The generated model is used to simulate and analyze the AGB thermal performance in the different flight phases in a typical flight mission, where the obtained results are validated against publicly available data. The analysis of the results confirms the effectiveness of the proposed method to estimate the heat loss values in the AGBs of gas turbine aero engines and to predict the thermal loads of the AGB in different flight phases. The developed tool enables the gas turbine thermal management system designers to deal with the generated heats effectively and in an optimal way.
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Ji, Fen Zhu, Xiao Xu Zhou, and Mi Tian. "Study on Thermal Management for Cooling System of Aero-Piston Engine." Advanced Materials Research 516-517 (May 2012): 452–56. http://dx.doi.org/10.4028/www.scientific.net/amr.516-517.452.

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A model of thermal management for cooling system of aero-piston engine was presented in this study. The models of main parts in this system were also founded. Based on the measured value of temperature and pressure in the cylinder, the heat transfer coefficient between gas-fired and the cylinder wall was calculated by using the empirical formula. A heat transfer boundary condition between fins and cooling air was determined according to various Reynolds number of the air flow. Moreover, the method of finite element analysis was utilized to calculate the temperature of cylinder block. In the specified working condition of some two-stroke piston engine used in the unmanned aerial vehicle (UAV), the calculation and analysis were made to study on the effect of aircrew speed and flight height on the cylinder block temperature, as well as the effect of cylinder block temperature on airscrew speed by the thermal management model. The calculation results show that, as the flight height rises, the cylinder block temperature increases accordingly when engine power and airscrew speeds are kept constant; however, at the same height, the higher the airscrew speed is, the lower cylinder block temperature will be. The cylinder block temperature should be kept stable by regulating the airscrew speed.
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Javiya, Umesh, John Chew, Nick Hills, and Timothy Scanlon. "Coupled FE–CFD thermal analysis for a cooled turbine disk." Proceedings of the Institution of Mechanical Engineers, Part C: Journal of Mechanical Engineering Science 229, no. 18 (February 18, 2015): 3417–32. http://dx.doi.org/10.1177/0954406215572430.

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This paper presents transient aero-thermal analysis for a gas turbine disk and the surrounding air flows through a transient slam acceleration/deceleration “square cycle” engine test, and compares predictions with engine measurements. The transient solid–fluid interaction calculations were performed with an innovative coupled finite element (FE) and computational fluid dynamics (CFD) approach. The computer model includes an aero-engine high pressure turbine (HPT) disk, adjacent structure, and the surrounding internal air system cavities. The model was validated through comparison with the engine temperature measurements and is also compared with industry standard standalone FE modelling. Numerical calculations using a 2D FE model with axisymmetric and 3D CFD solutions are presented and compared. Strong coupling between CFD solutions for different air system cavities and the FE solid model led to some numerical difficulties. These were addressed through improvement of the coupling algorithm. Overall performance of the coupled approach is very encouraging giving temperature predictions as good as a traditional model that had been calibrated against engine measurements.
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Zheng, Min, Fan Shen, and Pei Luo. "Vibration Fatigue Analysis of the Structure under Thermal Loading." Advanced Materials Research 853 (December 2013): 559–64. http://dx.doi.org/10.4028/www.scientific.net/amr.853.559.

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The fatigue problem of structures under concurrent thermal and vibration loading has not been thoroughly studied even though it is common in applications of aero-engine combustor liners. Here we attempt to explore such a problem using a simplified combustor liner model that is implemented by the commercial finite element software ANSYS Workbench. The modal parameters at various temperatures are calculated and the fatigue behavior under stochastic base excitation and thermal environment are analyzed. The results show that thermal loading not only has an effect on dynamic characteristics but also reduces the vibration fatigue life of the structure.
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Wang, Jiang-Feng, Jia-Wei Li, Fa-Ming Zhao, and Xiao-Feng Fan. "Numerical method of carbon-based material ablation effects on aero-heating for half-sphere." Modern Physics Letters B 32, no. 12n13 (May 10, 2018): 1840011. http://dx.doi.org/10.1142/s0217984918400110.

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A numerical method of aerodynamic heating with material thermal ablation effects for hypersonic half-sphere is presented. A surface material ablation model is provided to analyze the ablation effects on aero-thermal properties and structural heat conduction for thermal protection system (TPS) of hypersonic vehicles. To demonstrate its capability, applications for thermal analysis of hypersonic vehicles using carbonaceous ceramic ablators are performed and discussed. The numerical results show the high efficiency and validation of the method developed in thermal characteristics analysis of hypersonic aerodynamic heating.
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Dissertations / Theses on the topic "Aero-thermal model"

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Kocer, Gulru. "Aerothermodynamic Modeling And Simulation Of Gas Turbines For Transient Operating Conditions." Master's thesis, METU, 2008. http://etd.lib.metu.edu.tr/upload/12609642/index.pdf.

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In this thesis, development of a generic transient aero-thermal gas turbine model is presented. A simulation code, gtSIM is developed based on an algorithm which is composed of a set of differential equations and a set of non-linear algebraic equations representing each gas turbine engine component. These equations are the governing equations which represents the aero-thermodynamic process of the each engine component and they are solved according to a specific solving sequence which is defined in the simulation code algorithm. At each time step, ordinary differential equations are integrated by a first-order Euler scheme and a set of algebraic equations are solved by forward substitution. The numerical solution process lasts until the end of pre-defined simulation time. The objective of the work is to simulate the critical transient scenarios for different types of gas turbine engines at off-design conditions. Different critical transient scenarios are simulated for two di®
erent types of gas turbine engine. As a first simulation, a sample critical transient scenario is simulated for a small turbojet engine. As a second simulation, a hot gas ingestion scenario is simulated for a turbo shaft engine. A simple proportional control algorithm is also incorporated into the simulation code, which acts as a simple speed governor in turboshaft simulations. For both cases, the responses of relevant engine parameters are plotted and results are presented. Simulation results show that the code has the potential to correctly capture the transient response of a gas turbine engine under different operating conditions. The code can also be used for developing engine control algorithms as well as health monitoring systems and it can be integrated to various flight vehicle dynamic simulation codes.
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Rey, Villazón Jose María [Verfasser], Arnold [Akademischer Betreuer] Kühhorn, and Klaus [Akademischer Betreuer] Höschler. "Advanced aero engine common preliminary design environment for the automatic construction of secondary air system and thermal models / Jose María Rey Villazón ; Arnold Kühhorn, Klaus Höschler." Cottbus : BTU Cottbus - Senftenberg, 2015. http://d-nb.info/1114284092/34.

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Tsai, Liang-Chih, and 蔡亮至. "Numerical Simulation of Aerodynamic Optical-Dome With Aero-Thermal Radiation Effect In Different Turbulence Models." Thesis, 2012. http://ndltd.ncl.edu.tw/handle/02627558030510268519.

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碩士
國防大學理工學院
機械工程碩士班
100
The purpose of this study is to utilize CFD method to discuss under different turbulence models(DES、k-ε、k-ω) the hypersonic flow across the optical dome with cooling jets, the flow fields associated with geometry derived from the aerodynamic characteristics, the distribution of surface temperature combined with viscous dissipation and radiation effect of aerodynamic heating. Under such flow, due to the friction, the viscous effect near wall causesd no-slip codition, much of kinetic energy diffused to heat energy, hence, its surface and surrounding gases responded to temperature arised. By the cooling technology, the window cooled down against the heating load, meanwhile, the cooling jets would cause complicated shock and shock, shock and boundary interaction to affect flow field surrounding the window, additionally, the strong turbulent fluctuation and sheared effect while cooling jets mixed in main flow may lead to more complex disturbing between the flow properties of velocity, pressure, temperature and dentity to affect the result of real temperature, so we need to discuss in this study. The steady simulation of this study, the appropriate results of turbulence model compared to reference experimaental data of standoff distance validation is the k-ω. As thermal radiation effect introduced, the change of temperature at the stagnation point of nose would be 10.68% descent in DES turbulence model, 0.34% descent in k-ε model, and 6.31% in k-ω model. In order to make sure optical window temperature under 500K request, the optimum mass flow rate of cooling jets with dry air is 0.01kg/s. Further, under transient simulation without cooling, the during time of aerodynamic heating caused optical window’s surface temperature to reach critical temperature(500K) is 12 seconds for DES model, 10 seconds for k-ε model and 14 seconds for k-ω model.
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Conference papers on the topic "Aero-thermal model"

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Grey, Zach, Paul Constantine, and Andrew White. "Enabling aero-engine thermal model calibration using active subspaces." In AIAA Propulsion and Energy 2019 Forum. Reston, Virginia: American Institute of Aeronautics and Astronautics, 2019. http://dx.doi.org/10.2514/6.2019-4329.

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Matney, Andrew, Stephen M. Spottswood, Marc P. Mignolet, Adam J. Culler, and Jack J. McNamara. "Thermal Reduced Order Model Adaptation to Aero-Thermo-Structural Interactions." In 55th AIAA/ASME/ASCE/AHS/ASC Structures, Structural Dynamics, and Materials Conference. Reston, Virginia: American Institute of Aeronautics and Astronautics, 2014. http://dx.doi.org/10.2514/6.2014-0493.

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Braun, James, Guillermo Paniagua, and Francois Falempin. "Aero-Thermal Optimization of Bladeless Turbines." In ASME Turbo Expo 2020: Turbomachinery Technical Conference and Exposition. American Society of Mechanical Engineers, 2020. http://dx.doi.org/10.1115/gt2020-15551.

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Abstract The harnessing of mechanical power from supersonic flows is constrained by physical limitations and substantial aerodynamic losses. Bladeless axial turbines are a viable alternative to extract power in such harsh conditions without restricting the operating conditions. In this paper, we present a shape optimization of the wavy surface of bladeless turbines to maximize the power extraction, while minimizing convective heat fluxes and pressure losses. First, a baseline geometry was defined and an experimental campaign was carried out on the baseline wavy surface of the bladeless turbine at supersonic conditions in the Purdue Experimental Aerothermal Lab. Pressure, heat flux and skin friction measurements were compared with the Reynolds Averaged Navier Stokes results. Afterwards, the evaluation routine which consisted of the blade generation, grid generation, solving, and post-processing was implemented within an evolutionary optimizer with a multi-objective function to maximize the pressure force and minimize heat flux and pressure loss. Finally, a three-dimensional assessment in terms of power, heat load and pressure drop was performed for the best performing geometry with the commercial solver CFD++ of Metacomp. Turbulence closure was provided with the k-omega-SST turbulence model. The annular chamber of the bladeless turbine consisted of an unstructured mesh of approx. 8–10 million grid points.
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Ba, Wei, Longgang Liu, and Hong Liu. "Aero-Thermal Coupled Predictive Model for Preliminary Gas Turbine Blade Cooling Analysis." In ASME Turbo Expo 2018: Turbomachinery Technical Conference and Exposition. American Society of Mechanical Engineers, 2018. http://dx.doi.org/10.1115/gt2018-75089.

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The turbine inlet temperature has been increasing over the years to improve gas turbine efficiency and specific power. Blade cooling technology is essential to keep component temperatures below their critical value, and this makes the aero-thermal coupled phenomenon more significant. Blade life assessment is closely related to blade metal temperature distribution and gradients, and blade cooling analysis is always considered starting from the preliminary design stage. However, traditional blade cooling analysis for preliminary design is always based on external boundary conditions determined by experience, which affects the prediction accuracy as the interaction effect between the main flow and coolant is not considered. In this paper, an aero-thermal coupled blade cooling model is further developed by combining the improved streamline curvature method with a one-dimensional thermo-fluid network. This model is capable of predicting blade surface temperature distribution and internal coolant flow conditions in the preliminary phase of blade cooling design with a limited amount of input information. Experimental data for the NASA C3X profile with film cooling was selected for validation. In addition, a sensitivity analysis was performed on different film cooling mass flow rates to demonstrate the model flexibility for different boundary conditions.
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Ghai, Ramandeep Singh, Kuiying Chen, and Natalie Baddour. "Modelling Thermal Conductivity Of Porous Thermal Barrier Coatings For High-Temperature Aero Engines Using Five Phase Model." In 2018 Canadian Society for Mechanical Engineering (CSME) International Congress. York University Libraries, 2018. http://dx.doi.org/10.25071/10315/35380.

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Ba, Wei, and Xiaodong Ren. "Aero-Thermal Coupled Throughflow Method With Cooling Model Based on Flow Network Analysis." In ASME Turbo Expo 2017: Turbomachinery Technical Conference and Exposition. American Society of Mechanical Engineers, 2017. http://dx.doi.org/10.1115/gt2017-63614.

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The aero-thermal coupled phenomenon is significant in the modern cooled turbine, and it is necessary to consider the cooling effect in the throughflow design phase. A new cooling model based on flow network analysis for the aero-thermal coupled throughflow method was developed to consider the film cooling effect and predict the blade wall temperature downstream of the film cooling holes. The flow network analysis is introduced into the cooling model to determine the flow rate of each cooling hole. The mixing loss caused by film cooling is investigated as local total pressure loss, and the heat transfer influence caused by film cooling is considered by the film cooling effectiveness estimated by empirical correlation. The blade heat transfer downstream the film cooling holes is calculated from pressure and suction surfaces separately, based on main flow parameters calculated by the streamline curvature method. The experimental data of the C3X profile is selected for the cooling model verification. The film cooling flow rate calculated by the flow network analysis agrees well with the experimental data, and the calculated temperatures of both the pressure and suction surfaces downstream the film cooling holes are also in accordance with the experimental data. Therefore, aero-thermal coupled throughflow method with this cooling model can be a powerful tool for preliminary design of cooled turbine.
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Elmi, Carlo Alberto, Ignazio Vitale, Hauke Reese, and Antonio Andreini. "Multi-Objective Optimization of Aero Engine Combustor Adopting an Integrated Procedure for Aero-Thermal Preliminary Design." In ASME Turbo Expo 2021: Turbomachinery Technical Conference and Exposition. American Society of Mechanical Engineers, 2021. http://dx.doi.org/10.1115/gt2021-58945.

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Abstract The preliminary design of an aero-engine combustor is a multidisciplinary process that involves an extensive and systematic analysis of the design space. Simulation-driven approaches, in which several design configurations are numerically analyzed, may lead to heterogeneous models interacting with each other, sharing miscellaneous information within the process. Iterative and user-defined approaches, moreover, are inefficient when multiple and conflicting requirements are in place. To rely on integrated design methodologies has been demonstrated to be beneficial, especially if adopted in a structured approach to design optimization. In this paper, the application of the Combustor Design System Integration (DSI) to the definition of an optimal combustor preliminary configuration will be presented. Given a combustor baseline design, the multi-objective optimization problem has been defined by targeting an optimal distribution for temperature profiles and patterns at the combustor’s exit. Dilution port characteristics, such as hole number and dimension as well as the axial position of the row have been selected as design variables. To guarantee a water-tight design process while minimizing the user effort, the DSI tools were included in a dedicated framework for driving the optimization tasks. Here, a proper CFD domain for RANS, constituted by the flame tube region extended to the dilution port feeds, was adopted for imposing the air split designed for the combustor. Concerning a “complete” combustor sector, this allows a reduction in the computational effort while still being representative for its aero-thermal behavior. The optimization task was performed using a Response Surface Method (RSM), in which multiple, specific combustor configurations were simulated and the CFD result elaborated to build a meta-model of the combustor itself. Finally, the suitability of the resulting optimized configuration has been evaluated through an “a posteriori” analysis for thermal conditions and emission levels (NOx and CO). A lean combustion concept developed by Avio Aero with the aim of the homonymous EU research project, the NEWAC combustor, has been considered as test case.
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Sun, Zixiang, Dario Amirante, John W. Chew, and Nicholas J. Hills. "Coupled Aero-Thermal Modeling of a Rotating Cavity With Radial Inflow." In ASME Turbo Expo 2015: Turbine Technical Conference and Exposition. American Society of Mechanical Engineers, 2015. http://dx.doi.org/10.1115/gt2015-42609.

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Flow and heat transfer in an aero-engine compressor disc cavity with radial inflow has been studied using computational fluid dynamics (CFD), large eddy simulation (LES) and coupled fluid/solid modelling. Standalone CFD investigations were conducted using a set of popular turbulence models along with 0.2° axisymmetric and a 22.5° discrete sector CFD models. The overall agreement between the CFD predictions is good, and solutions are comparable to an established integral method solution in the major part of the cavity. The LES simulation demonstrates that flow unsteadiness in the cavity due to the unstable thermal stratification is largely suppressed by the radial inflow. Steady flow CFD modelling using the axisymmetric sector model and the Spalart-Allmaras turbulence model was coupled with a finite element (FE) thermal model of the rotating cavity. Good agreement was obtained between the coupled solution and rig test data in terms of metal temperature. Analysis confirms that use of a small radial bleed flow in compressor cavities is effective in reducing thermal response times for the compressor discs and that this could be applied in management of compressor blade clearance.
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Gu, Chunwei, Haibo Li, Wei Ba, and Xiaodong Ren. "An Aero-Thermal Coupled Throughflow Method for Convective Cooled Turbines." In ASME Turbo Expo 2016: Turbomachinery Technical Conference and Exposition. American Society of Mechanical Engineers, 2016. http://dx.doi.org/10.1115/gt2016-56563.

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Cooling technologies have been widely applied to protect the turbine blades at high inlet temperature, which also makes the aero-thermal coupled phenomenon more remarkable. Nevertheless, the aero-thermal phenomenon had not been considered in traditional throughflow methods and led to challenges of cooled turbine design. This paper proposes a new cooling model for the aero-thermal coupled throughflow method which was first proposed by the same author. The cooling model considers the variation of the temperature caused by air cooling both along the stream and span direction to improve the heat flux calculation accuracy. The impacts of heat transfer on mainstream enthalpy and entropy are further studied in this paper. The equivalent blade thickness and the estimation method of the heat exchange area were also introduced into the cooling model. The cooling model is validated with experimental data of the Mark II profile. This paper applies the method in the design of a high-pressure axial turbine, of which the first stator is cooled with convective cooling. With the prescribed blade temperature limitation, the flow field properties and the coolant requirement are predicted. The three dimensional CHT analysis is performed to validate the aerothermal coupled throughflow method, and the aerodynamic parameter predicted by the throughflow method is in accordance with the 3-D CHT analysis.
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Amirante, Dario, Nicholas J. Hills, and Paolo Adami. "A Multi-Fidelity Aero-Thermal Design Approach for Secondary Air Systems." In ASME Turbo Expo 2020: Turbomachinery Technical Conference and Exposition. American Society of Mechanical Engineers, 2020. http://dx.doi.org/10.1115/gt2020-15922.

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Abstract The paper presents a multi-disciplinary approach for aero-thermal and heat transfer analysis for internal flows. The versatility and potential benefit offered by the approach is described through the application to a realistic low pressure turbine assembly. The computational method is based on a run time code-coupling architecture that allows mixed models and simulations to be integrated together for the prediction of the sub-system aero-thermal performance. In this specific application the model is consisting of two rotor blades, the embedded vanes, the inter-stage cavity and the solid parts. The geometry represents a real engine situation. The key element of the approach is the use of a fully modular coupling strategy that aims to combine (1) flexibility for design needs, (2) variable level of modelling for better accuracy and (3) in memory code coupling for preserving computational efficiency in large system and sub-system simulations. For this particular example Reynolds Averaged Navier-Stokes (RANS) equations are solved for the fluid regions and thermal coupling is enforced with the metal (conjugate heat transfer). Fluid-fluid interfaces use mixing planes between the rotating parts while overlapping regions are exploited to link the cavity flow to the main annulus flow as well as in the cavity itself for mapping of the metal parts and leakages. Metal temperatures predicted by the simulation are compared to those retrieved from a thermal model of the engine, and the results are discussed with reference to the underlying flow physics.
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