Abstract
The performance of aircraft gas turbine engines mainly depends on performance of the turbine which expands the combusted air into the atmosphere. The turbine is a critical part which gets affected by the hot gas from combustor exhaust. So in order to enhance the performance of gas turbine engines, a proposed cooling type called film cooling is used for more than six decades. The current work is also an attempt to enhance the performance of the gas turbine engine by enhancing the film cooling performance. The film cooling performance was numerically calculated on a flat plate with micro-hole and compared the cooling performance from the macro-hole. The analysis was carried out for different blowing ratios and found that the coolant from micro-hole performs better in the vicinity region and also spreads well in the lateral direction. The vortex structure is also captured from the proposed turbulence model and discussed. The behaviour of micro flow inside the coolant pipe was also analyzed. The comparison between multiple micro-hole jets and discrete jets was also made and discussed.
Highlights
The advancement in computer technology economically made life easier for researchers in film cooling
The current paper focuses on achieving better film cooling performance by using a lesser amount of coolant with the help of micro-hole
Numerous turbulence models are available in FLUENT but there is no specific turbulence model specified to capture the complex behaviour of film cooling
Summary
Gas turbine engines are an integral part of the aerospace industries. The jet engines works according to the Brayton cycle as shown in Fig., where air enters through the inlet (stage 0), its temperature gets increased due to compression in the compressor (stage 0-2), the temperature of compressed air reaches its peak level due to the addition of fuel in the combustion chamber (stage 2-3), the high temperature gas expands with the help of the turbine and is expelled to the atmosphere through the nozzle (stage 3-4). Over the decades researchers focused on improving the performance of gas turbine engines by increasing TIT. The coolant is ejected through the porous wall of the blade material which thickens the oncoming mainstream flow boundary layer to increase the cooling performance. The disadvantage on transpiration cooling is its difficulty in maintaining the porosity of the holes and aerodynamic loss due to the normal injection of the coolant from the porous wall. Film cooling provides lesser cooling performance than transpiration cooling and is the widely applied cooling scheme by the aircraft turbine engine manufacturers. The current paper focuses on achieving better film cooling performance by using a lesser amount of coolant with the help of micro-hole
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