Film cooling is extensively used in modern gas turbine to protect high heat load components with coolant ejection from discrete holes and slots. In a high pressure gas turbine blade, there are many rows of film cooling hole with various arrangements and configurations. In the numerical simulations, the researchers are often facing a big challenge in mesh generation process because of many complicated geometries and features. From the view of numerical accuracy, it is widely accepted that the structured grid has a higher resolution in flow physics predictions than the unstructured grid especially in boundary layer. However, because many manual splitting, attaching and merging manipulations are involved for the cooling features and curved boundaries, it is very time-consuming for a researcher to generate multi-block structured grids for a full-cooled gas turbine blade. As a result, in the existing studies, almost all researchers preferred to generate unstructured meshes instead of structured meshes for the full-cooled blade. Unlike the previous research, the aim of this study is to show the possibilities of multi-block structured grid generation for full-cooled gas turbine blade based on the proposed background-grids based mapping (BGBM) method. With the strategies described in Part I of this work, the mesh generations for a full-cooled gas turbine blade is also convenient and time-saving. The layout of this paper is: At first, the NASA GE-E3 nozzle guided vane in first turbine stage is selected as a research objective for the mesh generation. This vane has two supply cavities and a slot to feed coolant air to the multi-rows of cooling holes and trailing edge. The front cavity is connected with 11 rows of film cooling holes, while the aft cavity is connected with two rows of film cooling holes, and the slot-cutback in trailing edge has multiple staggered arrays of cylindrical pins and land extensions. Then, with the generated structured grids, numerical simulations are performed to investigate the heat transfer and film cooling performance in gas turbine blade. Results indicate that the temperature distributions on vane surface are much related to the interactions between coolant flow and the main flow. Adiabatic film cooling effectiveness on blade surface near No.7 hole-array is the lower than the others. Compared with the pressure side wall and suction side wall, blade leading edge suffers stronger mainstream impingement and has a low film cooling effectiveness in this region. As the flow develops, the film cooling region expands on the suction side and shrinks on the pressure side. Finally, for blades with and without land extensions in trailing edge cutback, the heat transfer and film cooling effect in cutback region are also analyzed and compared with each other. It is seen that the land extensions have significant influence on the cooling effect on trailing edge cutback.


Background-Grid Based Mapping Approach to Film Cooling Meshing: Part III \u2014 Applications in a Full-Cooled Turbine Vane

Division / Sponsor:
International Gas Turbine Institute (IGTI)

Paper Proceedings: https://asmedigitalcollection.asme.org/GT/GT2020/volume/84058

Turbo Expo - Virtual Conference

Supporting the need among researchers worldwide for additional online and multimedia content, ASME now offers the ASME Conference Video Collection. This platform will collect and provide long-term access to conference presentation videos, encompassing all technology areas within ASME's technical conference and events program. About 30,000-40,000 engineers across all mechanical engineering and related disciplines participate in ASME conferences each year. This includes authors of approximately 6,000 to 8,000 technical papers published by ASME, the majority of whom also submit video presentations for our virtual conferences. The video collection will be available as an add-on to existing product packages or as a separate subscription.