Dynamic Analysis of Flow Field at the End of Combustor Simulator

Journal of engineering for gas turbines and power, 2005

2005

Performance enhancements and control of heat transfer in high pressure gas turbine vanes and rotors is dependent on understanding the flow and thermal fields approaching the turbine. The flow field exiting the combustor has highly non-uniform pressure and temperature variations in both the radial and circumferential directions as well as high turbulence levels. Several studies have shown significant impact on the overall and secondary flow fields within the turbine due to the inlet profile. The Turbine Research Facility (TRF) at Wright-Patterson Air Force Base has recently added a non-reactive full scale annular combustor simulator to the facility to study these effects. In conjunction with the TRF's experimental efforts to simulate combustor section exit flows, a threedimensional CFD analysis of the newly installed simulator has been undertaken. The analysis aids in the experimental implementation of the simulator and gives further understanding of the simulator's complex internal flow patterns. The goals for the TRF's simulator's is to produce a wide range of profiles to the inlet plane of the vane for evaluation of effects on heat transfer, loss, and loading. The CFD analysis allows an understanding of how those profiles are obtained by tracking the flow through two rows of staggered dilution holes and six rows of staggered film cooling holes on both the ID and OD liners of the main simulator chamber. This enables control as the CFD can guide the experimenter in knowing which liner component influenced the turbine inlet profile shape. Cases can then be run computationally by varying the mass flows and temperatures to tailor the profile to the desired shape prior to running the experiment. These profiles can then be sent through the

Journal of turbomachinery, 2004

The current demands for high-performance gas turbine engines can be reached by raising combustion temperatures to increase power output. High combustion temperatures create a harsh environment that leads to the consideration of the durability of the combustor and turbine sections. This paper presents a computational study of a flow field that is representative of what occurs in a combustor and how that flow field convects through the first downstream stator vane. The results of this study indicate that the development of the secondary flow field in the turbine is highly dependent on the incoming total pressure profile. The endwall heat transfer is also found to depend strongly on the secondary flow field.

The demands for best performance in gas turbine engines can be obtained by increasing combustion temperatures to increase thermal efficiency. Hot combustion temperatures create a harsh environment which leads to the consideration of the durability of the combustor and turbine sections. Improvements in durability can be achieved through understanding the interactions between the combustor and turbine. The flow field at a combustor exit shows non uniformities in pressure, temperature, and velocity in the pitch and radial directions. This inlet profile to the turbine can have a considerable effect on the development of the secondary flows through the vane passage. Presents a computational study of the flow field generated in a non-reacting gas turbine combustor and how that flow field convects through the downstream stator vane. Specifically, the effect that the combustor flow field had on the secondary flow pattern in the turbine was studied.

… of the Institution of Mechanical Engineers, …, 2008

IAEME

The demands for best performance in gas turbine engines can be obtained by increasing combustion temperatures to increase thermal efficiency. Hot combustion temperatures create a harsh environment which leads to the consideration of the durability of the combustor and turbine sections. Improvements in durability can be achieved through understanding the interactions between the combustor and turbine. The flow field at a combustor exit shows non uniformities in pressure, temperature, and velocity in the pitch and radial directions. This inlet profile to the turbine can have a considerable effect on the development of the secondary flows through the vane passage. Presents a computational study of the flow field generated in a non-reacting gas turbine combustor and how that flow field convects through the downstream stator vane. Specifically, the effect that the combustor flow field had on the secondary flow pattern in the turbine was studied

Journal of turbomachinery, 2003

Improved durability of gas turbine engines is an objective for both military and commercial aeroengines as well as for power generation engines. One region susceptible to degradation in an engine is the junction between the combustor and first vane given that the main gas path temperatures at this location are the highest. The platform at this junction is quite complex in that secondary flow effects, such as the leading edge vortex, are dominant. Past computational studies have shown that the total pressure profile exiting the combustor dictates the development of the secondary flows that are formed. This study examines the effect of varying the combustor liner film-cooling and junction slot flows on the adiabatic wall temperatures measured on the platform of the first vane. The experiments were performed using large-scale models of a combustor and nozzle guide vane in a wind tunnel facility. The results show that varying the coolant injection from the upstream combustor liner leads to differing total pressure profiles entering the turbine vane passage. Endwall adiabatic effectiveness measurements indicate that the coolant does not exit the upstream combustor slot uniformly, but instead accumulates along the suction side of the vane and endwall. Increasing the liner cooling continued to reduce endwall temperatures, which was not found to be true with increasing the film-cooling from the liner.

54th AIAA Aerospace Sciences Meeting, 2016

In the present paper some applications of the Italian Aerospace Research Center (CIRA) solver NExT in the CFD simulation of combusting flows in both liquid and hybrid rocket engine configurations are described. A comprehensive numerical model with real-fluid properties and turbulence-chemistry interaction effects was developed in order to predict the combusting flowfield inside a typical rocket combustion chamber. Different internal flow problems reproducing mixing and combustion processes have been investigated, with the aim to verify the implemented models and to validate them by comparing numerical and available experimental data. Typical rocket thrust chamber assemblies have been selected and simulated, including all relevant geometrical features such as injectors, the combustion chamber and the nozzle, to evaluate the solver capability in the simulation of both liquid and hybrid rocket engine combusting flowfields.

2018

Model high-speed combustor on gaseous hydrocarbon fuel, prepared for experiments in T-131 wind tunnel of TsAGI, is presented. Experiments are projected to create an experimental database for validation of calculations and physical models of turbulence and combustion. Geometry of combustor and prepared measurements are described. The main subject of paper is preliminary calculations of this combustor. Numerical methods for 2D and 3D URANS calculations are described. Special attention is given to numerical techniques allowing fast calculations of 3D unsteady flow development in the combustor. Approach to parallel realization of Fractional Time Stepping (FTS) technology is described. One way of In Situ Adaptive Tabulation (ISAT) of kinetic equations solution during the calculation is presented. Possible gasdynamic structure of flow in the combustor with the flame stabilization both in subsonic and in supersonic regime is described. Asymmetrical stationary solution (for the symmetrically-expanding duct) and symmetrical solution with flame oscillations are found and analyzed. PREPARATION OF THE EXPERIMENTAL MODEL Today it is impossible to imagine the creation of perspective aircraft without supplementation of the experiments with numerical simulation. However, modern possibilities to calculate practical flows with combustion in aircraft engines are essentially limited by the huge computer cost for calculation of 3D viscid turbulent flows with finite-rate reactions [1,2] and by the low accuracy of the available models of turbulence, of chemical kinetics, and of turbulence/combustion interaction [3-5]. In 2017, the scientific laboratory "Studies and development of physical models and numerical technologies for description of different combustion regimes in aircraft engines" has been created in Propulsion department of TsAGI under the support of Russian Ministry of education and science. Goals of the laboratory are the development and validation of physically-grounded models for various combustion regimes in air-breathing engines, as well as the creation of special software for use in the cycle of aerodynamic design for new aircraft engines. The laboratory develops and improves physical and mathematical models of turbulent combustion, oriented to calculations in the framework of RANS (Reynolds-Averaged Navier-Stokes) and LES (Large Eddy Simulation). These models are implemented into computer codes, specially adjusted for concrete class of flows to get the best prediction of flow characteristics. Such adjustment is based on experimental data for flows of the considered class. To create the basis for such activities, the new "fire" aerodynamic experiments are prepared in TsAGI. Experiments will be performed on the unique high-speed wind tunnel T-131 (see detailed information about wind