The adoption of lean-burn technology in modern aero-engines influences the already critical aerothermal conditions at turbine entry, where the absence of dilution holes preserves the swirl component generated by burners and prevents any control on pattern factor. The associated uncertainty and lack of confidence entail the application of wide safety margins in turbine cooling design, with a detrimental effect on engine efficiency. Computational fluid dynamics (CFD) can provide a deeper understanding of the physical phenomena involved in combustor–turbine interaction, especially with hybrid Reynolds-averaged Navier–Stokes (RANS) large eddy simulation (LES) models, such as scale adaptive simulation (SAS), which are proving to overcome the well-known limitations of the RANS approach and be a viable approach to capture the complex flow physics. This paper describes the numerical investigation on a test rig representative of a lean-burn, effusion cooled, annular combustor developed in the EU Project Full Aerothermal Combustor-Turbine interactiOns Research (FACTOR) with the aim of studying combustor–turbine interaction. Results obtained with RANS and SAS were critically compared to experimental data and analyzed to better understand the flow physics, as well as to assess the improvements related to the use of hybrid RANS-LES models. Significant discrepancies are highlighted for RANS in predicting the recirculating region, which has slight influence on the velocity field at the combustor outlet, but affects dramatically mixing and the resulting temperature distribution. The accuracy of the results achieved suggests the exploitation of SAS model with a view to the future inclusion of the nozzle guide vanes in the test rig.

Modern lean burn aero-engine combustors make use of relevant swirl degrees for flame stabilization. Moreover, important temperature distortions are generated, in tangential and radial directions, due to discrete fuel injection and liner cooling flows respectively. At the same time, more efficient devices are employed for liner cooling and a less intense mixing with the mainstream occurs. As a result, aggressive swirl fields, high turbulence intensities, and strong hot streaks are achieved at the turbine inlet. In order to understand combustor-turbine flow field interactions, it is mandatory to collect reliable experimental data at representative flow conditions. While the separated effects of temperature, swirl, and turbulence on the first turbine stage have been widely investigated, reduced experimental data is available when it comes to consider all these factors together.In this perspective, an annular three-sector combustor simulator with fully cooled high pressure vanes has been designed and installed at the THT Lab of University of Florence. The test rig is equipped with three axial swirlers, effusion cooled liners, and six film cooled high pressure vanes passages, for a vortex-to-vane count ratio of 1:2. The relative clocking position between swirlers and vanes has been chosen in order to have the leading edge of the central NGV aligned with the central swirler. In order to generate representative conditions, a heated mainstream passes though the axial swirlers of the combustor simulator, while the effusion cooled liners are fed by air at ambient temperature. The resulting flow field exiting from the combustor simulator and approaching the cooled vane can be considered representative of a modern Lean Burn aero engine combustor with swirl angles above ±50 deg, turbulence intensities up to about 28% and maximum-to-minimum temperature ratio of about 1.25. With the final aim of investigating the hot streaks evolution through the cooled high pressure vane, the mean aerothermal field (temperature, pressure, and velocity fields) has been evaluated by means of a five-hole probe equipped with a thermocouple and traversed upstream and downstream of the NGV cascade.

Turbine inlet conditions in modern aero-engines employing lean-burn combustors are characterised by highly swirled flow and non-uniform temperature distributions. As a consequence of the lack of confidence in numerical predictions and the uncertainty of measurement campaigns, the use of wide safety margins is of common practice in the design of turbine cooling systems, thus affecting the engine performance and efficiency. Previous experiences showed how only scale-resolving approaches such as Large-eddy and Scale-adapting simulations are capable of overcoming the limitations of RANS, significantly improving the accuracy in the prediction of flow and temperature fields at the combustor outlet. However it is worth investigating the impact of such differences on the aerothermal performance of the NGVs, as to understand the limitations entailed in the current approach for their thermal design. Industrial applications in fact usually rely on 1D, circumferentially-averaged profiles of pressure, velocity and temperature at the combustor-turbine interface in conjunction with Reynolds-averaged Navier-Stokes (RANS) models. This paper describes the investigation performed on an experimental test case consisting of a combustor simulator equipped with NGVs. Three numerical modelling strategies were compared in terms of flow field and thermal loads on the film-cooled vanes: i) RANS model of the NGVs with inlet conditions obtained from a RANS simulation of the combustor; ii) RANS model of the NGVs with inlet conditions obtained from a Scale-Adaptive Simulation (SAS) of the combustor; iii) SAS model inclusive of both combustor and NGVs. The results of this study show that estimating the aerodynamics at the NGV exit does not demand particularly complex approaches, whereas the limitations of standard RANS models are highlighted again when the turbulent mixing is key. High-fidelity predictions of the conditions at the turbine entrance proved to be very beneficial to reduce discrepancies in the estimation of local adiabatic wall temperature of even 100 K. However, a further leap forward can be achieved with an integrated simulation, capable of reproducing the transport of the unsteady fluctuations generated in the combustor up into the turbine, which plays a key role in presence of film cooling. This work therefore points out how keeping the analysis of combustor and NGVs separate can lead to a significantly misleading estimation of the thermal loads and ultimately to a wrong thermal design of the cooling system.

In modern lean-burn aero-engine combustors, highly swirling flow structures are adopted to control the fuel-air mixing and to provide the correct flame stabilization mechanisms. Aggressive swirl fields and high turbulence intensities are hence expected in the combustor-turbine interface. Moreover, to maximize the engine cycle efficiency, an accurate design of the high-pressure nozzle cooling system must be pursued: in a film-cooled nozzle, the air taken from last compressor stages is ejected through discrete holes drilled on vane surfaces to provide a cold layer between hot gases and turbine components. In this context, the interactions between the swirling combustor outflow and the vane film cooling flows play a major role in the definition of a well-performing cooling scheme, demanding for experimental campaigns at representative flow conditions. An annular three-sector combustor simulator with fully cooled high-pressure vanes has been designed and installed at THT Lab of University of Florence. The test rig is equipped with three axial swirlers, effusion-cooled liners, and six film-cooled high-pressure vanes passages, for a vortex-to-vane count ratio of 1:2. The relative clocking position between swirlers and vanes has been chosen in order to have the leading edge of the central airfoil aligned with the central swirler. In this experimental work, adiabatic film effectiveness measurements have been carried out in the central sector vanes, in order to characterize the film-cooling performance under swirling inflow conditions. The pressure-sensitive paint (PSP) technique, based on heat and mass transfer analogy, has been exploited to catch highly detailed 2D distributions. Carbon dioxide has been used as coolant in order to reach a coolant-to-mainstream density ratio of 1.5. Turbulence and five-hole probe measurements at inlet/outlet of the cascade have been carried out as well, in order to highlight the characteristics of the flow field passing through the cascade and to provide precise boundary conditions. Results have shown a relevant effect of the swirling mainflow on the film cooling behavior. Differences have been found between the central airfoil and the adjacent ones, both in terms of leading edge stagnation point position and of pressure and suction side film coverage characteristics.

No abstract

The characterization of pressure, temperature and velocity fields in turbomachinery flows typically relies on well-proven probe-based technology such as pneumatic probes, hot-wire sensors or thermocouples. These devices have to be introduced into narrow flow channels and by that obstruct part of the duct at the actual measuring position, which can significantly alter the aerodynamic performance of the components under investigation. In this contribution, measurement results of a commercially available five-hole probe with mounted temperature sensor and laser-optical filtered Rayleigh scattering diagnostics are acquired downstream of a nozzle guide vane cascade with lean-burn combustion representative inlet flow distortions. Pressure results obtained by both methods are found to be on a

In order to deepen the knowledge of the interaction between modern lean burn combustors and high pressure turbines, a real scale annular three sector combustor simulator has been assembled at University of Florence, with the goal of investigating and characterizing the generated aerothermal field and the hot streaks transport between combustor exit and the high pressure vanes location. To generate hot streaks and simulate lean burn combustors behavior, the rig is equipped with axial swirlers, fed by main air flow that is heated up to 531 K, and liners with effusion cooling holes that are fed by air at ambient temperature. The three sector configuration is used to reproduce the periodicity on the central sector and to allow to perform measurements inside the chamber, through the lateral walls. Ducts of different length have been mounted on the swirlers, preserving the hot mainflow from the interaction with coolant. Such configurations, together with the one without ducts, have been tested, using different measurement techniques, in order to highlight the differences in the resulting flow fields. First of all, isothermal PIV measurements have been performed on the combustion chamber symmetry plane, to highlight the mixing phenomena between the mainflow and cooling flows. Then a detailed investigation of the mean aerothermal field at combustor exit has been carried out, for nominal operating conditions, by means of a five hole pressure probe provided with a thermocouple, installed on an automatic traverse system. With the aim of analyzing the hot streaks transport and the flow field modification towards the vanes location, such measurements have been performed on two different planes: one located in correspondence of the combustor exit and the further one placed downstream, in the virtual location of the vanes leading edges. Therefore, an experimental database, describing the evolution of the flow field in a combustor simulator with typical traits of modern lean burn chambers, for different injector geometries, has been set up.

Computational Fluid Dynamics (CFD) results are often presented in a deterministic way despite the uncertainties related to boundary conditions, numerical modelling, and discretization error. Uncertainty quantification is the field studying how these phenomena affect the numerical result. With these methods, the results obtained are directly comparable with the experimental ones, for which the uncertainty related to the measurement is always shown. This work presents an uncertainty quantification approach applied to CFD: the test case consists of an industrial prismatic gas turbine vane with standard film cooling shaped holes system on the suction side only. The vane was subject of a previous experimental test campaign which had the objective to evaluate the film cooling effectiveness through pressure-sensitive paint technique. CFD analyses are conducted coherently with the experiments: the analogy between heat and mass transfer is adopted to draw out the adiabatic film effectiveness, solving an additional transport equation to track the concentration of CO2 used as a coolant fluid. Both steady and unsteady simulations are carried out: the first one using a RANS approach with k-ω SST turbulence model the latter using a hybrid LES-RANS approach. Regarding uncertainty quantification, three geometrical input parameters are chosen: the hole dimension, the streamwise inclination angle of the holes, and the inlet fillet radius of the holes. Polynomial-chaos approach in conjunction with the probabilistic collocation method is used for the analysis: a first-order polynomial approximation was adopted which required eight evaluations only. RANS approach is used for the uncertainty quantification analysis in order to reduce the computational cost. Results show the confidence interval for the analysis as well as the probabilistic output. Moreover, a sensitivity analysis through Sobol’s indices was carried out which prove how these input parameters contribute to the film cooling effectiveness, in particular, when dealing with the additive manufacturing process.