Recently, the number of gas turbine combined cycle plants is rapidly increasing in substitution of nuclear power plants. The turbine inlet temperature (TIT) is being constantly increased in order to achieve higher efficiency. Therefore, the improvement of the cooling technology for high temperature gas turbine blades is one of the most important issue to be solved.
In a gas turbine, the main flow impinging at the leading edge of the turbine blade generates a so called horseshoe vortex by the interaction of its boundary layer and generated pressure gradient at the leading edge. The pressure surface leg of this horseshoe vortex crosses the passage and reaches the blade suction surface, driven by the pressure gradient existing between two consecutive blades. In addition, this pressure gradient generates a crossflow along the endwall. This all results into a very complex flow field in proximity of the endwall. For this reason, burnouts tend to occur at a specific position in the vicinity of the leading edge. In this research, a methodology to cool the endwall of the turbine blade by means of film cooling jets from the blade surface is proposed. The cooling performance and heat transfer coefficient distribution is investigated using the transient thermography method. CFD analysis is also conducted to know the phenomena occurring at the end wall and calculate the heat transfer distribution.