Abstract
Panels were tested at different locations around the turbine blade, on both suction and pressure surfaces. Three different surface porosities were also tested. Results demonstrated that the approach can be very successful with high levels of film cooling effectiveness, exceeding 95%, achieved using low coolant mass flow rates. Increasing the surface porosity also proved to be an important parameter in the panel’s performance. Additionally, staggering the film holes lead to significant positive interactions between individual films, resulting in much improved panel performance.
Highlights
The need to design turbine blades able to endure the ever increasing temperatures found in the generation of turbine engines has long motivated continuous improvement in cooling and material technology
Very high cooling effectiveness was achieved around the entire blade at very low mass coolant fractions
Films perform well on the early suction surface, which benefits from the effects of curvature and acceleration, resulting in very constant film coverage downstream of the panel
Summary
The need to design turbine blades able to endure the ever increasing temperatures found in the generation of turbine engines has long motivated continuous improvement in cooling and material technology. Recent advances in porous C/C-SiC and CMC have led to a revival in the field of transpiration in aerospace for uses such as rocket nozzle cooling [2], yet momentum in turbine blade technology has lagged behind. While this may be due to the significant constraints in blade design arising from the high thermal, mechanical and cyclical loading, a fundamental understanding of the true potential of transpiration cooling as applied to turbine blades remained mostly unexplored before the present research
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