Gas turbines have been widely used in power generation and aircraft propulsion. To improve the gas turbine performance, the turbine inlet temperature is usually elevated higher than the metal melting point. Therefore, cooling of gas turbines becomes very critical for engines' safety and lifetime. One of the effective methods is film cooling, in which the coolant air from the discrete holes blankets the surface from the hot gas flow. The major issues related to film cooling are its poor coverage, aerodynamic loss, and increase of heat transfer coefficient due to strong flow mixing. To improve the cooling performance, this paper examined film cooling with backward injection. It is observed that film cooling with backward injection can produce much more uniform cooling coverage under different conditions, which include cases on flat surface with low or high pressure and temperature. The backward injection also performs better in the presence of blade curvature. The effect of other parameters on the film cooling is also reported. The numerical results are validated by simple experimental test in this study.
The thermal efficiency of a gas turbine is largely dependent on the turbine inlet temperature (TIT). Modern gas turbines may operate at temperatures as high as 2000K, which is higher than the melting point of the material in use. Hence, thermal protection of gas turbine hot components is a big challenge. Film cooling is the most common cooling technique adopted in this application. In film cooling, coolant is injected at discrete locations along the metal surface, which forms a layer of cool air immediately over the hot surface, thus, protecting it from direct contact with hot mainstream air. The cooling is strong along the centerline of the hole in the downstream region and rapidly decreases over the span-wise direction. The distributed cooling can result in large thermal gradients, inducing thermal stresses in the material. The region with least cooling may lead to a cascade failure of the blade. Film cooling with backward injection holes has been proved to reduce this effect. In the current work, backward coolant injection scheme is explored under fan-shaped holes numerically. Fluent, a commercial CFD software, is used in the current work for numerical simulations. Effects of blowing ratio, injection angle, and turbulence are considered. Numerical results show that fan-shaped holes are better than simple cylindrical holes in terms of both cooling effectiveness and its uniformity. Numerical results are validated with experimental results. The image of temperature fields on cooling surface is captured with an Infrared camera.
Combustion chamber or combustor is one of the hottest parts of a gas turbine. Liner is where the actual flame occurs in a combustor and thus, the hottest part of the combustor. The temperature of working fluid inside a liner is about 1200 to 2000K. Because of the hot fluid, the liner is heated up to a temperature almost impossible for the material to withstand. Although the mechanical stresses experienced by the combustor liner are within acceptable limits, high temperatures and large temperature gradients affect the structural integrity of its components, which makes the liner a very critical component of a gas turbine in structural and thermal designs. Film cooling is a traditional method of cooling the inner surface of liner. In film cooling for a combustor, axial holes are drilled along the surface of the liner at discrete locations, through which cold air is injected axially into the liner to provide a film of cool air that prevents direct contact of hot air, and thus, protects the inner wall surface. The film is destroyed in the downstream to the flow because of mixing of cool and hot air. Though this method provides an acceptable cooling, there is a compromise with the increased net benefits of the gas turbine. Therefore, there is a need for new cooling techniques or enhancing the techniques available. The current work is a numerical simulation of film cooling in a model combustor. The effect of coolant injection angles and blowing ratios on film cooling effectiveness is studied. One innovative method, cooling with mist injection, is explored to enhance the performance of film cooling. The effect of droplet size and mist concentration, which can affect the performance of the mist injection, is also analyzed. Fluent, a commercial CFD software, is used in the current work for numerical simulations.
Due to the unique role of gas turbine engines in power generation and aircraft propulsion, significant effort has been made to improve the gas turbine performance. As a result, the turbine inlet temperature is usually elevated to be higher than the metal melting point. Therefore, effective cooling of gas turbines is a critical task for engines’ efficiency as well as safety and lifetime. Film cooling has been used to cool the turbine blades for many years. The main issues related to film cooling are its poor coverage, aerodynamic loss, and increase of heat transfer coefficient due to strong mixing. To overcome these problems, film cooling with backward injection has been found to produce a more uniform cooling coverage under low pressure and temperature conditions and with simple cylindrical holes. Therefore, the focus of this paper is on the performance of film cooling with backward injection at gas turbine operating conditions. By applying numerical simulation, it is observed that along the centerline on both concave and convex surfaces, the film cooling effectiveness decreases with backward injection. However, cooling along the span is improved, resulting in more uniform cooling.
Film cooling has been successfully used in cooling gas turbine components that are exposed to very high temperature environments. One main disadvantage of using film cooling is the aerodynamic losses associated. To address to the needs of obtaining uniform cooling in the downstream regions, backward injection of coolant has proved to be effective. However, there is a need to understand the aerodynamic behaviors of jet and mainstream flows in order to design effective configurations with this scheme of injecting coolant. In this work, the underlying aerodynamic principles of backward injection are studied numerically. All simulations are conducted with Fluent, a commercial CFD software. Results show that the classical counter rotating vortex found in simple cylindrical holes are not seen in the case of backward injections. Backward injection results in reduced coolant requirements and elimination of complex hole designs to avoid jet lift-off.
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