Woopen, Michael (2011) Reduction of Film Coolant in High Pressure Turbines. Bachelor's, RWTH Aachen.
Full text not available from this repository.
Gas turbine development has been in the direction of increased turbine inlet temperatures to improve efficiency and power output. As a result the allowable metal temperatures of the alloys used for the turbine airfoils and endwalls are exceeded such that the surfaces have to be protected from the hot gases by cooling. Bypassed leakage flows emerging from a slot between the combustor and turbine component endwalls can be utilized for cooling purposes. In order to improve this cooling method the complex flow structures in a turbine have to be considered as they enhance the heat transfer and convect hot core flow gases onto component surfaces. In this thesis four film cooling configurations using this slot which differ from each other in terms of geometrical aspects are examined with the help of the flow solver TRACE (Turbomachinery Research Aerodynamics Computational Environment) developed at the Institute of Propulsion Technology of the German Aerospace Center. Furthermore, in the present study the amount of coolant ejected is varied by adjusting the blowing ratio. Additionally the shape of the leakage slot is tailored to take the pressure distribution in front of the leading edge into account. The results are analysed with respect to film cooling effectiveness and the surface heat transfer. The results are mainly influenced by the three-dimensional flow structures visualized with the streamwise vorticity. The results obtained for the straight slot show for all mass flow ratios an intensified ejection of coolant in the center of the passage due to the wall pressure distribution. Furthermore, the coolant is not only concentrated in the center of the passage but also “trapped” between the two legs of the horseshoe vortex, where the coolant air is scraped away from the endwall in the region of the separation lines. With increasing mass flow rate of the coolant one realizes a broader area of the endwall covered by the coolant. By adapting the contour of the slot concerning the pressure distribution in front of the leading edge the coolant is enabled to leave the slot in a much more uniform manner. This way a better cooling around the leading edge region and the juncture of the blade’s pressure side and the endwall is achieved. Thus, not only the need for complex cooling hole layouts next to the critical areas can be minimized but also the number of cooling holes and thus the amount of coolant can be reduced.
|Document Type:||Thesis (Bachelor's)|
|Title:||Reduction of Film Coolant in High Pressure Turbines|
|Date:||17 March 2011|
|Number of Pages:||104|
|Keywords:||Turbine, Plane cascade, Sidewall, Secondary Flow, Film-Cooling, Cooling Slot, TRACE|
|Department:||Faculty of Mechanical Engineering|
|HGF - Research field:||Aeronautics, Space and Transport|
|HGF - Program:||Aeronautics|
|HGF - Program Themes:||Propulsion Systems|
|DLR - Research area:||Aeronautics|
|DLR - Program:||L ER - Engine Research|
|DLR - Research theme (Project):||L - Turbine Technologies|
|Institutes and Institutions:||Institute of Propulsion Technology > Turbine|
|Deposited By:||Dr.-Ing. Peter-Anton Giess|
|Deposited On:||13 Feb 2012 15:01|
|Last Modified:||13 Feb 2012 15:01|
Repository Staff Only: item control page