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9th World Congress on Mechanical, Chemical, and Material Engineering, MCM 2023 -- 6 August 2023 through 8 August 2023 -- 309509, Strong and precise computational methods are necessary to discover and understand undesirable effects in tanks where liquids slosh, known as sloshing. Sloshing is important for many industrial applications, such as fuel tanks in ships, aircraft, and other transport vehicles. Numerical methods are commonly used in modelling sloshing behaviour, and adaptive mesh refinement (AMR) technology is an effective method used to increase numerical accuracy in sloshing simulations. The primary objective of this research isto conduct Computational Fluid Dynamics (CFD) calculations of sloshing phenomena to establish a methodology for observing the undesirable effects on the relevant system and assess the effectiveness of adaptive mesh refinement by comparing the results of surface impact pressures with experimental case results from literature. A 3D model of a rectangular tank partially filled with water is used to simulate the impact pressure caused by roll motion. The roll motion is based on experimental data and occurs at periods close to the tank's internal wave resonance period. The pressure results are observed from a single monitoring point. Numerical studies are performedusing Star CCM+ software. The Volume of Fluid (VOF) based Eulerian method is utilised to model the free surface flow. The study demonstrates the ability of AMR to accurately model sloshing behaviour and compares fixed and AMR grids at different levels. Finally, the results of fixed and adaptive cartesian grids are compared and verified with corresponding experimental data. The results showed thatthe AMR grid provided higher numerical accuracy, lower computational cost and allowed for more accurate modelling of sloshing behaviour. It is emphasised that the importance of AMR for understanding sloshing behaviour and modelling it accurately. © 2023, Avestia Publishing. All rights reserved.

[No Abstract Available]

International Symposium on Convective Heat and Mass Transfer, CONV 2014 -- 8 June 2014 through 13 June 2014 -- 268949, A gas turbine combustor preliminary design tool that includes all the conceptual and one-dimensional preliminary design process in an automated way has been used for the heat transfer calculations. The flame tube wall temperature characteristics of an experimental through flow combustor with air-blast atomizer has been analysed using the methodology. A small scale gas turbine combustor is used as a test case in order to validate the predictive capability of the newly developed methodology named PreCoDe. Predictions about the mass flow split, flame and combustor liner temperatures for 1 millimetre segments have been performed by using the PreCoDe technique that makes use of analytical and empirical correlations. For the validation of the code, calculation results are compared with the thermal paint and the thermocouple temperature measurements on the combustor liner. As a conclusion, the 1D prediction has revealed good agreement with experimental results and hence good design capability and potentials of the PreCoDe. © 2014, Begell House Inc. All rights reserved., ACKNOWLEDGEMENTS The authors would like to thank colleagues at TEI, Tusas Engine Industries for the contribution of the experimental studies. Numerical studies installation of the atmospheric combustor test rig and conducted experimental studies were a part of the “Development of a Combustion Chamber with Air-blast Fuel-Air Injector” project (00564.STZ.2010-1) and financially supported by the Turkish Republic Ministry of Science, Industry and Technology. The authors would like to thank to the ministry for the invaluable support.

[No Abstract Available]

International Symposium on Convective Heat and Mass Transfer, CONV 2014 -- 8 June 2014 through 13 June 2014 -- 268949, A gas turbine combustor preliminary design tool that includes all the conceptual and one-dimensional preliminary design process in an automated way has been used for the heat transfer calculations. The flame tube wall temperature characteristics of an experimental through flow combustor with air-blast atomizer has been analysed using the methodology. A small scale gas turbine combustor is used as a test case in order to validate the predictive capability of the newly developed methodology named PreCoDe. Predictions about the mass flow split, flame and combustor liner temperatures for 1 millimetre segments have been performed by using the PreCoDe technique that makes use of analytical and empirical correlations. For the validation of the code, calculation results are compared with the thermal paint and the thermocouple temperature measurements on the combustor liner. As a conclusion, the 1D prediction has revealed good agreement with experimental results and hence good design capability and potentials of the PreCoDe. © 2014, Begell House Inc. All rights reserved., ACKNOWLEDGEMENTS The authors would like to thank colleagues at TEI, Tusas Engine Industries for the contribution of the experimental studies. Numerical studies installation of the atmospheric combustor test rig and conducted experimental studies were a part of the “Development of a Combustion Chamber with Air-blast Fuel-Air Injector” project (00564.STZ.2010-1) and financially supported by the Turkish Republic Ministry of Science, Industry and Technology. The authors would like to thank to the ministry for the invaluable support.

CFD, Computational Fluid Dynamics, tools have become an integral part of product design and analysis of gas turbines, rocket engines and thrusters. In order to model the chemical reacting flow in the combustion chambers, a CFD solver dealing with combustion of multi species reactions is developed. The solver must also be able to cope with the three flow regimes, subsonic, transonic and supersonic flow that occur in many flow problems. This is namely called a solver for all flow speeds. This solver also utilizes all resources of computers by being CPU parallel and GPU accelerated. A third party Open Source matrix solver ViennaCL is integrated to the Solver to solve linear system of equations. © ICCFD 2018.

In this study, non-adiabatic flamelet combustion model have been conducted for resolving finite rate chemistry effects on University of Virginia’s Scramjet Facility Dual-Mode Scramjet Combustor Configuration A using large-eddy simulation(LES). Numerical results are compared with coherent anti-stokes raman spectroscopy measurements (CARS). Flamelet structure is defined with mixture fraction, scalar dissipation rate and heat loss ratio (enthalpy defect). In addition, Reynolds – averaged Navier- Stokes(RANS) based computational grid investigation is presented with three different grid resolutions. The result of this study, the effectiveness and accuracy of non-Adiabatic flamelet combustion model, is investigated. © 2017 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

Small scale turbojet engines are used in miniature UAVs and target drones. Combustor design at this scale is a challenge due to reduced volume for atomization, mixing and subsequent combustion. Furthermore correct estimation for the engine lifetime is critical in the design phase. In terms of lifetime hot-section components are much more critical. Exit temperature profile, which is commonly quantified using radial temperature distribution factor or using overall temperature distribution factor, is critical for the lifetime determination of critical hot-section engine components such as the first stage inlet guide vane of the turbine. This study presents the viability of the CFD approach for combustor design in terms of comparisons with experimental results put a spotlight to the areas that should improve in order to develop better CFD methods and practices. Results indicate that significant variability might occur even with the slightest manifold design modification. Also injector-to-injector flow rate variability has an effect on the exit plane temperature distribution causing hot-spots. CFD predictions agree with experimental results within engineering accuracy. However, experimental verification is quite necessary to reach a final judgement on the combustor design. © 2014 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved., 564.STZ.2010-1

The design of an aero engine combustion chamber is a difficult process in order to meet engine performance requirements. One of the most important problems in the design and development of an aero engine combustion chamber is to achieve uniform temperature pattern in circumferential direction and a radial temperature profile in traverse direction with a peak in the middle. The exit temperature profile affects the turbine nozzle guide vane and the overall engine life time. Moreover, Large Eddy Simulation(LES), is a substantial technique for prediction of reacting flow simulations that contain turbulent mixing and combustion process. Hence, in this study LES of non-premixed reacting flow in a full annular RQL, Rich-burn, quick-Quench, Lean-burn, combustor is performed to investigate the effects of fuel distribution on combustion chamber exit temperature for both uniform and non-uniform fuel distribution for fuel injectors. © 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved., Seventh Framework Programme: AAT-2011-RTD-1-284859-ESPOSA Seventh Framework Programme

While simulating and analyzing air-vehicle motion during flight, it is not possible to use the nonlinear model directly without appropriate initial conditions. The initial condition stands for a point about which air-vehicle flight condition must not change abruptly. This point is referred to as the trim condition. In this paper, classical Newton-Raphson method will be used first to find the trim conditionby solving a system of nonlinear algebraic equations. When the starting point is not close to a solution, the Newton-Raphson algorithm behaves erratically. Sometimes, components of the unknown inputs and states or the Jacobian will result in numerical instability. Global convergence is then addressed, which is the issue of trying to force convergence to a solution from a remote starting point. The Newton-Raphson method can be made more robust by using line search technique. A comparative study is carried out by incorporating different methods for solving the nonlinear equations in F-16 aircraft trim. The attainability of various trim conditions is also investigated. © 2020, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.

Highly swirled non-premixed methane-air combustion is studied using Unsteady RANS. The increased CPU power at present times makes RANS solutions a viable design methodology for industrial applications. LES, Large Eddy Simulation, is still a bit far away from being the routine approach as a design tool in industry. The confined non-premixed TECFLAM S09C flame is investigated because of its similarity with gas turbine engine combustors. A block structured hexahedral computational mesh is used for the whole domain for higher numerical accuracy. A first-order accurate Euler Implicit technique is used for the temporal discretization of the transient terms. Realizable k-? turbulence model is employed in order to account for the turbulent flow effects on the flow field and chemical reactions. Results of fast and two step reactions are compared with finite rate chemistry. Results show that, the best solution is obtained by using Finite Rate Eddy Dissipation Model. The infinitely fast chemistry approach is not capable of predicting the reaction delay that is clearly observed in the experiments. Instead, the fast chemistry approach show reaction zones in the close vicinity of the swirler where the fuel-air mixing is not well achieved for the reactions to take place in such a short distance. The flow field is directly affected by the heat release rate that is determined by the fuel air mixing and combustion model. © 2014 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

A computational tool has been developed to compute transient surface temperature utilizing basic flight parameters such as altitude, Mach number and angle of attack. An explicit finite difference technique is used to discretize the governing differential equations that account for convection and radiation heat transfer with non-reactive chemistry. Different approaches are applied to two different surface types; thermally thin and thermally thick. The Biot number criterion between the wall and the surrounding air is considered to determine the surface type. Transport properties of air are calculated at Eckert's reference temperature. Boundary layer transition is taken into account by considering local Mach number and Reynolds number. Heat transfer coefficients are calculated by use of flat plate approaches. Effect of angle of attack is taken into account through modified Newtonian theory. The available X-15 flight data for two different flight trajectories and HIFiRE-5 flight data are used to validate the prediction tool. Vehicles with different geometries are modeled using CFD simulations and results of different configurations are compared with the computed temperatures. Predicted results for surface temperatures are found to be in good agreement with measured flight data and simulation results. The validation shows that the methodology developed in the present study could be useful in predicting aerodynamic heating loads during conceptual and preliminary design phases. © 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All right reserved.

Simulating a gas turbine combustor is a challenging issue due to complexity of both flow and reaction kinetics. URANS, Unsteady RANS, is an interesting approach for simulating transient reacting flows in combustors. Studies clearly indicate the superiority of LES, Large Eddy Simulation, over URANS computations. However, it is also well known that URANS computations have surely CPU overhead advantages over the more CPU costly LES approaches. In this study, URANS computations are carried out on a highly swirled non-premixed combustion chamber by applying four mesh densities from 0.3M to 11.5M block-structured hexahedral cells. TECFLAM S09 methane-air case with 0.9 swirl number is investigated due to its similarity to stationary gas turbine combustors. Simulations showed that URANS is unable to capture the complexity of swirling effects as good as LES predictions even on a computational mesh with 2.7M elements. The strong inner recirculation zone is predicted with a good degree of accuracy with LES computations while it cannot be captured by URANS approach. © 2015 Begell House, Inc.

The fast development in the field of aeronautics dictates development of more powerful aircraft engines. Constant increase in overall pressure ratios and operating temperatures are essential for the required higher power and thrust.. However, higher Turbine Entry Temperatures reduce the operation life of engine components. Cooling of engine components is one of the vital issues for development of efficient engines in terms of engine life and cost. Gas turbine cooling technologies have been investigated for more than 70 years and it has substantially changed in the course of new aircraft engine development. In this study, a nozzle guide vane (NGV) is designed and an efficient cooling system is investigated for a commercial aircraft engine NGV leading edge. The engine is developed for “AIAA Undergraduate Team Engine Design Competition” and it has extremely a high overall pressure ratio of 60. The cooling of NGV, Nozzle Guide Vane, is one of the most critical design issues in the whole engine design cycle. In the present study film cooling of an NGV is investigated using computational fluid dynamics (CFD). Parametric study of different cooling hole configurations with a study of different turbulence sub models are studied. The parametric study of film cooling includes the cooling hole types, hole numbers and positions. © 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All right reserved.

Design of a gas turbine combustor is a challenging process. It contains many stages including estimation of liner wall temperatures in order to determine the lifetime of the engine. Therefore, accurate prediction of the liner wall temperature is an essential part of the gas turbine combustor design. The liner is heated by convection and radiation of the combustion gases inside the flame tube and it is cooled by convection to the air in annulus and radiation to the outer casing. Many studies have been performed with Conjugate Heat Transfer (CHT) methodology neglecting the effect of radiation on the liner temperature due to its complexity and limited computational power sources. The Effect of radiative heat transfer on combustor liner temperature using CHT methodology is studied in the present work. Radiation from luminous flame which is caused by formation of soot particles in the flame is not accounted for and radiation from non-luminous flame due to combustion gases, CO2 and H2O, and surface radiation is investigated. Effect of turbulent Schmidt/Prandtl number is also investigated in this paper. © 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.

Strong mixing of fuel and air is one of the most important parameter ín combustion design criteria. Air-blast atomizer has two important functions in combustion chamber. One of them is to atomize the liquid fuel, and the second one is to enable good mixing of the fuel and air in the flametube by creating a swirling and re-circulating flow in the primary zone. This strong mixing can be achieved by swirler channels of the air-blast atomizers. In this study, the effects of the co- and counter-swirl configurations of a double swirler airblast atomizer on flow field, temperature distribution and exit emissions in an annular turbojet combustor is investigated by CFD. ANSYS-Fluent software is used for CFD simulations. Flamelet combustion model was used for modelling of the non-premixed combustion. Simulation results of temperature distribution, equivalence ratio and NOx emissions are presented and discussed. © 2015, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.

CFD, Computational Fluid Dynamics, tools have become an integral part of product design and analysis of gas turbines, rocket engines and thrusters. In order to model the chemical reacting flow in the combustion chambers, a CFD solver dealing with combustion of multi species reactions is developed. The solver must also be able to cope with the three flow regimes, subsonic, transonic and supersonic flow that occur in many flow problems. This is namely called a solver for all flow speeds. This solver also utilizes all resources of computers by being CPU parallel and GPU accelerated. A third party Open Source matrix solver ViennaCL is integrated to the Solver to solve linear system of equations. © ICCFD 2018.

In this study, non-adiabatic flamelet combustion model have been conducted for resolving finite rate chemistry effects on University of Virginia’s Scramjet Facility Dual-Mode Scramjet Combustor Configuration A using large-eddy simulation(LES). Numerical results are compared with coherent anti-stokes raman spectroscopy measurements (CARS). Flamelet structure is defined with mixture fraction, scalar dissipation rate and heat loss ratio (enthalpy defect). In addition, Reynolds – averaged Navier- Stokes(RANS) based computational grid investigation is presented with three different grid resolutions. The result of this study, the effectiveness and accuracy of non-Adiabatic flamelet combustion model, is investigated. © 2017 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

The design of an aero engine combustion chamber is a difficult process in order to meet engine performance requirements. One of the most important problems in the design and development of an aero engine combustion chamber is to achieve uniform temperature pattern in circumferential direction and a radial temperature profile in traverse direction with a peak in the middle. The exit temperature profile affects the turbine nozzle guide vane and the overall engine life time. Moreover, Large Eddy Simulation(LES), is a substantial technique for prediction of reacting flow simulations that contain turbulent mixing and combustion process. Hence, in this study LES of non-premixed reacting flow in a full annular RQL, Rich-burn, quick-Quench, Lean-burn, combustor is performed to investigate the effects of fuel distribution on combustion chamber exit temperature for both uniform and non-uniform fuel distribution for fuel injectors. © 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved., Seventh Framework Programme: AAT-2011-RTD-1-284859-ESPOSA Seventh Framework Programme

Small scale turbojet engines are used in miniature UAVs and target drones. Combustor design at this scale is a challenge due to reduced volume for atomization, mixing and subsequent combustion. Furthermore correct estimation for the engine lifetime is critical in the design phase. In terms of lifetime hot-section components are much more critical. Exit temperature profile, which is commonly quantified using radial temperature distribution factor or using overall temperature distribution factor, is critical for the lifetime determination of critical hot-section engine components such as the first stage inlet guide vane of the turbine. This study presents the viability of the CFD approach for combustor design in terms of comparisons with experimental results put a spotlight to the areas that should improve in order to develop better CFD methods and practices. Results indicate that significant variability might occur even with the slightest manifold design modification. Also injector-to-injector flow rate variability has an effect on the exit plane temperature distribution causing hot-spots. CFD predictions agree with experimental results within engineering accuracy. However, experimental verification is quite necessary to reach a final judgement on the combustor design. © 2014 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved., 564.STZ.2010-1

Highly swirled non-premixed methane-air combustion is studied using Unsteady RANS. The increased CPU power at present times makes RANS solutions a viable design methodology for industrial applications. LES, Large Eddy Simulation, is still a bit far away from being the routine approach as a design tool in industry. The confined non-premixed TECFLAM S09C flame is investigated because of its similarity with gas turbine engine combustors. A block structured hexahedral computational mesh is used for the whole domain for higher numerical accuracy. A first-order accurate Euler Implicit technique is used for the temporal discretization of the transient terms. Realizable k-? turbulence model is employed in order to account for the turbulent flow effects on the flow field and chemical reactions. Results of fast and two step reactions are compared with finite rate chemistry. Results show that, the best solution is obtained by using Finite Rate Eddy Dissipation Model. The infinitely fast chemistry approach is not capable of predicting the reaction delay that is clearly observed in the experiments. Instead, the fast chemistry approach show reaction zones in the close vicinity of the swirler where the fuel-air mixing is not well achieved for the reactions to take place in such a short distance. The flow field is directly affected by the heat release rate that is determined by the fuel air mixing and combustion model. © 2014 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved.

While simulating and analyzing air-vehicle motion during flight, it is not possible to use the nonlinear model directly without appropriate initial conditions. The initial condition stands for a point about which air-vehicle flight condition must not change abruptly. This point is referred to as the trim condition. In this paper, classical Newton-Raphson method will be used first to find the trim conditionby solving a system of nonlinear algebraic equations. When the starting point is not close to a solution, the Newton-Raphson algorithm behaves erratically. Sometimes, components of the unknown inputs and states or the Jacobian will result in numerical instability. Global convergence is then addressed, which is the issue of trying to force convergence to a solution from a remote starting point. The Newton-Raphson method can be made more robust by using line search technique. A comparative study is carried out by incorporating different methods for solving the nonlinear equations in F-16 aircraft trim. The attainability of various trim conditions is also investigated. © 2020, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.

The fast development in the field of aeronautics dictates development of more powerful aircraft engines. Constant increase in overall pressure ratios and operating temperatures are essential for the required higher power and thrust.. However, higher Turbine Entry Temperatures reduce the operation life of engine components. Cooling of engine components is one of the vital issues for development of efficient engines in terms of engine life and cost. Gas turbine cooling technologies have been investigated for more than 70 years and it has substantially changed in the course of new aircraft engine development. In this study, a nozzle guide vane (NGV) is designed and an efficient cooling system is investigated for a commercial aircraft engine NGV leading edge. The engine is developed for “AIAA Undergraduate Team Engine Design Competition” and it has extremely a high overall pressure ratio of 60. The cooling of NGV, Nozzle Guide Vane, is one of the most critical design issues in the whole engine design cycle. In the present study film cooling of an NGV is investigated using computational fluid dynamics (CFD). Parametric study of different cooling hole configurations with a study of different turbulence sub models are studied. The parametric study of film cooling includes the cooling hole types, hole numbers and positions. © 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All right reserved.

Design of a gas turbine combustor is a challenging process. It contains many stages including estimation of liner wall temperatures in order to determine the lifetime of the engine. Therefore, accurate prediction of the liner wall temperature is an essential part of the gas turbine combustor design. The liner is heated by convection and radiation of the combustion gases inside the flame tube and it is cooled by convection to the air in annulus and radiation to the outer casing. Many studies have been performed with Conjugate Heat Transfer (CHT) methodology neglecting the effect of radiation on the liner temperature due to its complexity and limited computational power sources. The Effect of radiative heat transfer on combustor liner temperature using CHT methodology is studied in the present work. Radiation from luminous flame which is caused by formation of soot particles in the flame is not accounted for and radiation from non-luminous flame due to combustion gases, CO2 and H2O, and surface radiation is investigated. Effect of turbulent Schmidt/Prandtl number is also investigated in this paper. © 2016, American Institute of Aeronautics and Astronautics Inc, AIAA. All rights reserved.