The influence of freestream turbulence representative of the flow downstream of a
modern gas turbine combustor and first stage vane on turbine blade heat transfer has been
measured and analytically modeled in a linear, transonic turbine cascade. Measurements
were performed on a high turning, transonic turbine blade. The facility is capable of
heated flow with inlet total temperature of 120ºC and inlet total pressure of 10 psig. The
Reynolds number based on blade chord and exit conditions (5x106
) and the inlet and exit
Mach numbers (0.4 and 1.2, respectively) are representative of conditions in a modern
gas turbine engine. High intensity, large length-scale freestream turbulence was
generated using a passive turbulence-generating grid to simulate the turbulence generated
in modern combustors after it has passed through the first stage vane row. The grid
produced freestream turbulence with intensity of approximately 10-12% and an integral
length scale of 2 cm near the entrance of the cascade passages, which is believed to be
representative of the core flow entering a first stage gas turbine rotor blade row. Mean
heat transfer results showed an increase in heat transfer coefficient of approximately 8%
on the suction surface of the blade, with increases on the pressure surface on the order of
two times higher than on the suction surface (approximately 17%). This corresponds to
increases in blade surface temperature of 5-10%, which can significantly reduce the life
of a turbine blade. The heat transfer data were compared with correlations from
published literature with good agreement.