The leading edge regions of first stages blades and vanes of heavy-duty gas turbines are subjected to high thermal loads. Efficient cooling allows the reduction of the coolant mass flow required to drive the metal temperatures to a range satisfying mechanical integrity requirements.
This paper investigates the heat transfer and pressure loss behavior for the internal cooling channel of a leading edge of a gas turbine blade. The geometrical profile of the blade leading edge and the operating conditions considered are representative of that normally found in a heavy-duty gas turbine.
The geometries investigated cover angled turbulators of various angles, pitches and heights. Partial and full length rib coverage as well as broken ribs are also considered. In addition, the impact of including fillets in the geometry is assessed.
The experimental and numerical studies are conducted at passage Reynolds numbers ranging from 7.5·104 to 1.3·105.
Experiments are performed using Perspex models at atmospheric conditions. The internal heat transfer coefficients on all internal surfaces are measured via thermochromic liquid crystal method and the pressure drop is measured via pressure taps distributed along the channel.
The predicted and experimental heat transfer enhancements are compared on the leading-edge, pressure, suction and web surfaces. The overall non dimensional cooling performance numbers are also compared for the various geometries.
The results show a large variation of heat transfer enhancement and pressure loss over the various turbulator geometries investigated. Also, the complex flow structures lead to highly differentiated results for leading edge, pressure, suction and web surfaces. Although some configurations with higher ribs lead to increased heat transfer, the associated pressure losses are also shown to increase substantially.