Abstract
Gas turbines are extensively used for aircraft propulsion, land‐based power generation, and industrial applications. Thermal efficiency and power output of gas turbines increase with increasing turbine rotor inlet temperature (RIT). The current RIT level in advanced gas turbines is far above the .melting point of the blade material. Therefore, along with high temperature material development, a sophisticated cooling scheme must be developed for continuous safe operation of gas turbines with high performance. Gas turbine blades are cooled internally and externally. This paper focuses on external blade cooling or so‐called film cooling. In film cooling, relatively cool air is injected from the inside of the blade to the outside surface which forms a protective layer between the blade surface and hot gas streams. Performance of film cooling primarily depends on the coolant to mainstream pressure ratio, temperature ratio, and film hole location and geometry under representative engine flow conditions. In the past number of years there has been considerable progress in turbine film cooling research and this paper is limited to review a few selected publications to reflect recent development in turbine blade film cooling.
Highlights
Advanced gas turbine engines operate at high temperatures (1200-1400C) to improve thermal efficiency and power output
The results indicated that the secondary flow and horseshoe vortex interact with the coolant jet which is converted towards the suction side of the nozzle guide vanes (NGV)
This study presents a comparison of various tip cooling configurations and their effects on film effectiveness and heat transfer coefficient using a transient liquid crystal technique
Summary
Advanced gas turbine engines operate at high temperatures (1200-1400C) to improve thermal efficiency and power output. 1.10, corresponding blowing ratios approximately turbine cooling system designers need to know where heat is transferred from hot mainstream to the airfoil. As mentioned earlier, these film-hole pattern (i.e., film hole location, distribution, angle and shape) would affect film cooling performance. The best film cooling design is to reduce the heat load ratio (i.e., smaller h/ho enhancement with greater r/) for a minimum amount of coolant available for a film cooled airfoil. They performed experiments on a three-vane hot cascade and studied the effects of parameters such as Mach number, Reynolds number, coolant-to-gas temperature ratio, and coolant-to-gas pressure ratio.
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