The high temperature level reached at the exit of combustion chambers of modern aircraft engines and the practical limitations of advanced materials, demand efficient cooling of turbine blades. Optimization of the cooling requires an accurate prediction of aerodynamic losses and heat transfer on turbine blades.
A new two-dimensional compressible, aerothermal boundary layer code has been developed. The formulation includes strong viscous-inviscid interaction, which enhances the stability properties of the code. The boundary layer equations associated with the energy equation are solved with an implicit Keller-box scheme. Viscous-inviscid flow coupling is performed by adding an interaction equation which has an elliptic character. The complete system of equations is solved by a multi-pass procedure. This technique contributes to the stabilization of the method and allows the computation of regions with strong adverse pressure gradients, separation bubbles and injections in case of film cooling.
Comparisons between experimental and theoretical results are provided. Flow characteristics including heat transfer were computed for several cases such as flat plates with strong pressure gradients, and turbine blade boundary layers. Good agreement between computation and experiment is observed, demonstrating the high accuracy and robustness of the code.