Characterization of Boundary Layers on Isothermal and Adiabatic Curved Surfaces of a Supersonic Turbine Cascade

The effects of adiabatic and isothermal boundary conditions are investigated on the shock-boundary layer interactions (SBLIs) in a supersonic turbine cascade. Special attention is given to the characterization of the incoming boundary layers over the convex and concave walls of the blade and their impact in the SBLIs. Large eddy simulations (LES) are performed for an inlet Mach number of $\mathbf{M_\infty = 2.0}$ and Reynolds number based on the axial chord $\mathbf{Re = 200\,000}$. For the isothermal condition, the wall to inlet temperature ratio is $\mathbf{T_w/T_{\infty}=0.75}$, representing a cooled wall. Different incident shock wave topologies occur on the suction and pressure sides of the airfoil. For the former, an oblique shock impinges on the boundary layer leading to a larger separation bubble. On the other hand, a normal shock from a Mach reflection induces a small separation region near the wall for the pressure side. Results are presented in terms of mean velocity and temperature contours, and the incoming boundary layers are characterized by looking at the Clauser parameter, shape factor, dilatation and turbulent kinetic energy (TKE) profiles. Inspection of the shape factors show that the adiabatic wall boundary layers are more prone to separate than the isothermal ones. This is indeed observed in the airfoil suction side, but not on the pressure side, where the flow separates in the same chord position regardless of the thermal boundary condition. This is a topic for investigation in the final version of the paper. An assessment of the dilatation and TKE profiles explains the disparities of the bubble sizes on the pressure and suction sides of the airfoil.