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Engine development requires accurate estimates of the heat loads. Estimates of the convective heat fluxes are particularly vital to assess the thermomechanical integrity of the turbomachinery components. This paper reports an experimental heat transfer research in a one and a half turbine stage, composed of a high-pressure turbine and a low-pressure vane. Measurements were performed in a compression tube facility at the von Karman Institute, able to reproduce engine representative Reynolds and Mach numbers. Double-layered thin film gauges were used to monitor the time-dependent temperature distribution around the airfoil. Several tests at different metal temperatures were performed to derive the adiabatic wall temperature. This research allowed quantifying the independent effects on the unsteady heat flux of the gas temperature fluctuations and boundary layer unsteadiness.

This paper reports the external convective heat transfer in an innovative low pressure vane with multisplitter configuration. Three small aerodynamic blades are positioned between each structural vane, providing a novel architecture for ultra-high by-pass ratio aero-engines, with increased LP vane radius and swan-neck diffuser to link the HP turbine. The measurements have been performed in the compression tube test rig of the von Karman Institute, using single layered thin film gauges. Time-averaged and time-resolved heat transfer distributions are presented for the three aerovanes and for the structural blade, at three pressure ratios tested at representative conditions of modern aeroengines, with M 2,is ranging from 0.87 to 1.07 and a Reynolds number of about 10 6 . This facility is specially suited to control the gas-to-wall temperature ratio. Accurate time-averaged heat transfer distributions around the aerovanes are assessed, that allow characterizing the boundary layer status for each position and pressure ratio. The heat transfer distribution around the structural blade is also obtained, depicting clear transition to turbulence, as well as particular flow features on the pressure side, like separation bubbles. Unsteady data analysis reveals the destabilizing effect of the rotor left-running shock on the aerovanes boundary layer, as well as the shift of transition onset for different blade passing events.

During engine development, heat loads in the turbomachinery are analyzed based on theoretical and numerical estimates together with correlations. Accurate models of the convective fluxes are vital to assess the thermo-mechanical integrity. This paper reports an experimental heat transfer research in a 1.5 turbine stage, the researched model is a the structural vane of a multi-splittered low pressure vane located downstream of a high pressure turbine stage. This concept is envisioned for ultra-high bypass-ratio aero-engines with a swan-neck diffuser between the high-pressure turbine and the low-pressure turbine. Measurements were performed in the large compression tube facility of the von Karman Institute, at representative conditions of modern aero-engines. Double-layered thin film gauges were employed for the measurement of the time-dependent temperature distribution around the airfoil. The initial temperature of the structural vane was adjusted using a heating system. The experimental procedure has allowed the determination of the time-mean and unsteady adiabatic wall temperature. Hence this technique allows the determination of the non-dimensional Nusselt number and proper scaling of the surface temperature to engine conditions. Furthermore, the analysis of the unsteady data reveals the contribution of the temperature fluctuations to the unsteady heat fluxes.