Flow Field and Hot Streak Migration Through a High Pressure Cooled Vanes With Representative Lean Burn Combustor Outflow

Abstract

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.

Turbulence intensity measurements across a NGV cooled cascade with representative Lean Burn combustor outflow

“…The three sector configuration was adopted to make the central sector flow field less sensitive to the presence of the lateral walls. Details of the CFD-aided design process and of its experimental verification can be found in the works of Bacci et al (2018b) and Koupper et al (2014).…”

“…The film cooling scheme is made by 8 rows of cylindrical film cooling holes, with four shower head and four pressure side rows. More details on the combustor simulator rig and on NGVs film cooling system can be found in Bacci et al (2018) and Bacci et al (2018b).…”

“…Split-fiber probes were adopted for such a goal. Five hole probe results from the work of Bacci et al (2018b) will be also reported, in order to both provide information regarding the overall flow field at NGV inlet/exit and to assess the capability of the adopted HWA technique to deal with a highly 3D flow field.…”

Unsteady Phenomena at the Combustor-Turbine Interface

“…Bacci et al [24,25] used a test rig with cooled vanes and combustion chamber to investigate hot streaks and swirling inflow to the vanes. It was found that different aerofoils face significantly different temperature fields.…”

“…Cubeda et al [27] compared different numerical methods to the experimental measurements of Bacci et al [25,24], finding that neither RANS of combustion chamber and vanes, nor SAS of the combustion chamber used as a boundary condition to steady RANS gave satisfactory results. RANS simulation gave local adiabatic temperature errors exceeding 150K.…”