The aerothermal interaction of the combustor exit flow on the first vane row has been examined at the Large Scale Turbine Rig (LSTR) at Technische Universität Darmstadt (Darmstadt, Germany). A baseline configuration of axial inflow and a variation of swirling combustor inflow have been studied. The nozzle guide vane (NGV) featured endwall cooling, airfoil film cooling and a trailing edge slot ejection as well as NGV-rotor wheel space purge flow. CO₂ is injected for coolant flow tracing. The results are compared to five hole probe (5HP) measurements. The experiments for the baseline configuration are accompanied by numerical simulations using a passive scalar tracking method to validate the results and study the propagation of the coolant flow. The endwall coolant injection is detected to influence the pressure losses in the NGV. It has an impact on the Trailing Edge (TE) coolant ejection as well. For swirling combustor inflow, increased NGV pressure losses and increased mixing of Rear Inner Discharge Nozzle (RIDN) coolant and main flow is observed. An influence of the clocking position of the swirler to the vane is detected. Additional losses within the NGV row can be assigned to the swirler by means of flow tracing.

Knowing the flow conditions at the combustor turbine interface is a key asset for an efficient cooling design of high-pressure turbines. However, measurements and numerical predictions of combustor exit conditions are challenging due to the extreme temperatures and complex flow patterns in modern combustors. Even the time-averaged flow fields at the combustor exit which are commonly used as inlet condition for simulations of the turbine are therefore subject to uncertainty. The goal of this paper is to illustrate how aleatory uncertainties in the magnitude and position of residual swirl and hot spots at the combustor exit affect uncertainties in the prediction of cooling and heat load of the first nozzle guide vane. Also, it is identified which of these uncertain parameters have the greatest impact. An iso-thermal test rig and an engine realistic setup with lean burn inflow conditions are investigated. The analysis combines a parameterized model for combustor exit flow fields with uncertainty quantification methods. It is shown that the clocking position of turbine inlet swirl has a large effect on the formation of secondary flows on the vane surface and thus affects the uncertainty of thermal predictions on the hub and vanes.

An algorithm is presented generating a complete set of inlet boundary conditions for Reynolds-averaged Navier–Stokes computational fluid dynamics (RANS CFD) of high-pressure turbines to investigate their interaction with lean and rich burn combustors. The method shall contribute to understanding the sensitivities of turbine aerothermal performance in a systematic approach. The boundary conditions are based on a set of input parameters controlling velocity, temperature, and turbulence fields. All other quantities are derived from operating conditions and additional modelling assumptions. The algorithm is coupled with a CFD solver by applying the generated profiles as inlet boundary conditions. The successive steps to derive consistent flow profiles are described and results are validated against flow fields extracted from combustor CFD.