elib
DLR-Header
DLR-Logo -> http://www.dlr.de
DLR Portal Home | Impressum | Kontakt | English
Schriftgröße: [-] Text [+]

Reduction of Film Coolant in High Pressure Turbines

Woopen, Michael (2011) Reduction of Film Coolant in High Pressure Turbines. Bachelorarbeit, RWTH Aachen.

Dieses Archiv kann nicht den gesamten Text zur Verfügung stellen.

Kurzfassung

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.

Dokumentart:Hochschulschrift (Bachelorarbeit)
Titel:Reduction of Film Coolant in High Pressure Turbines
Autoren:
AutorenInstitution oder E-Mail-Adresse der Autoren
Woopen, MichaelNICHT SPEZIFIZIERT
Datum:17 März 2011
Seitenanzahl:104
Stichwörter:Turbine, Plane cascade, Sidewall, Secondary Flow, Film-Cooling, Cooling Slot, TRACE
Institution:RWTH Aachen
Abteilung:Faculty of Mechanical Engineering
HGF - Forschungsbereich:Luftfahrt, Raumfahrt und Verkehr
HGF - Programm:Luftfahrt
HGF - Programmthema:Antriebe
DLR - Schwerpunkt:Luftfahrt
DLR - Forschungsgebiet:L ER - Antriebsforschung
DLR - Teilgebiet (Projekt, Vorhaben):L - Turbinentechnologien
Standort: Göttingen
Institute & Einrichtungen:Institut für Antriebstechnik > Turbine
Hinterlegt von: Dr.-Ing. Peter-Anton Giess
Hinterlegt am:13 Feb 2012 15:01
Letzte Änderung:13 Feb 2012 15:01

Nur für Mitarbeiter des Archivs: Kontrollseite des Eintrags

Blättern
Suchen
Hilfe & Kontakt
Informationen
electronic library verwendet EPrints 3.3.12
Copyright © 2008-2013 Deutsches Zentrum für Luft- und Raumfahrt (DLR). Alle Rechte vorbehalten.