This paper presents heat transfer characteristics for a Multi-Stage Cooling Scheme (MSCS) design applicable to high temperature gas turbine engines in aerospace and electric power generation. The film cooling and impingement techniques are considered concurrently throughout this study. The proposed design involves passing cooling air from the inside of the turbine blade to the outside through three designed stages. The coolant air is passed through a circular hole into an internal gap creating an impingement of air inside the blade. It then exits through a sequence of two differently shaped holes onto the blades external surface. The film cooling effectiveness is enhanced by increasing the internal gap height and offset distance. This effect is significantly diminished however by changing the inclination angle from 90 to 30 at large gap height. The coolant momentum became more uniform by creating the internal gap consequently the coolant air is spread closer to the external blade surface. This reduces jet liftoff as the air exits its hole and also provides internal cooling for the blade. The hole exit positioned on the outer surface of the blade is designed to give a positive and a wide downstream lateral spreading. The MSCS demonstrates greater film cooling effectiveness performance than traditional schemes.