This paper reports the results of a combined experimental and numerical investigation carried out to support the design of a film cooling system for a rotor blade platform. A 7 blade cascade of a high-pressure-rotor stage of a heavy-duty gas turbine has been tested in a low speed wind tunnel. Tests have been carried out at a low Mach number (Ma2is = 0.27) with a relatively high inlet turbulence intensity level of about 7.6% at the leading edge plane. The same cascade model was also numerically tested by means of a 3D RANS approach. Cascade flow and heat transfer behavior was first experimentally assessed without coolant injection and used to validate the numerical approach. These data were also used to design the platform cooling scheme based on shaped holes that was then tested for variable injection rates. The thermal behavior was measured by using the Binary PSP technique, so to obtain film cooling effectiveness distributions over the passage. RANS 3D CFD simulations were also run on the same cooled model and testing conditions, allowing to critically assess the prediction capability of the selected numerical approach and of the design process.