Abstract
An experimental study was performed to investigate the effects of the arrangement of fan-shaped film cooling holes and density ratio (DR) on heat transfer coefficient augmentation. Both single- and multi-row fan-shaped film cooling holes were considered. For the multi-row fan-shaped holes, the heat transfer coefficient was measured at DRs of 1 and 2, and both staggered and inline arrangements of holes were considered. For the single-row fan-shaped holes, DR = 1.0, 1.5, 2.0, and 2.5 and M = 1.0 and 2.0 conditions were tested. The mainstream velocity was 20 m/s, and the turbulence intensity and boundary layer thickness were 3.6% and 6 mm, respectively. The heat transfer coefficient was measured using the one-dimensional transient infrared thermography method. The results show that an increased heat transfer coefficient augmentation is observed between film cooling holes for the case with a smaller hole pitch and higher blowing ratio. For the given fan-shaped hole parameters, the effects of the row-to-row distance and hole arrangement are not significant. In addition, as the velocity difference between the mainstream and coolant increases, the heat transfer coefficient ratio increases.
Highlights
The measured results were expressed as the ratio of the heat transfer coefficient with film cooling (h) to the heat transfer coefficient without film cooling
The mean value of the heat transfer coefficients for both sides of the test section without holes was used as the heat transfer coefficient without film cooling (Figure 6)
A laterally averaged heat transfer coefficient ratio and the overall averaged heat transfer coefficient ratio were used to compare the heat transfer coefficient augmentation quantitatively, and the hlateral was taken within one pitch near the hole, as shown8 in Schematic of of measurement measurement area
Summary
Turbine inlet temperatures have continuously been increased to improve the efficiency of gas turbines. Modern gas turbines operate under conditions that exceed allowable material temperatures. The heat load and thermal stress of gas turbine blades have seen increases, and appropriate cooling techniques are essential to assure the required life of the gas turbine blades. A typical external cooling technique for turbine blades, protects the blade surface from hot gas by injecting coolant through discrete holes or slots installed on the turbine blade surface [1]. The performance of the film cooling technique is generally evaluated through film cooling effectiveness, η, which is defined as follows
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