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dc.contributorHan, Je-Chin
dc.creatorMhetras, Shantanu
dc.date.accessioned2010-01-15T00:15:15Z
dc.date.accessioned2010-01-16T02:13:57Z
dc.date.accessioned2017-04-07T19:56:46Z
dc.date.available2010-01-15T00:15:15Z
dc.date.available2010-01-16T02:13:57Z
dc.date.available2017-04-07T19:56:46Z
dc.date.created2006-08
dc.date.issued2009-06-02
dc.identifier.urihttp://hdl.handle.net/1969.1/ETD-TAMU-1820
dc.description.abstractFilm cooling effectiveness on a gas turbine blade tip on the near tip pressure side and on the squealer cavity floor is investigated. Optimal arrangement of film cooling holes, effect of a full squealer and a cutback squealer, varying blowing ratios and squealer cavity depth are also examined on film cooling effectiveness. The film-cooling effectiveness distributions are measured on the blade tip, near tip pressure side and the inner pressure and suction side rim walls using a Pressure Sensitive Paint (PSP) technique. A blowing ratio of 1.0 is found to give best results on the pressure side whereas the other tip surfaces give best results for blowing ratios of 2. Film cooling effectiveness tests are also performed on the span of a fully-cooled high pressure turbine blade in a 5 bladed linear cascade using the PSP technique. Film cooling effectiveness over the entire blade region is determined from full coverage film cooling, showerhead cooling and from each individual row with and without an upstream wake. The effect of superposition of film cooling effectiveness from each individual row is then compared with full coverage film cooling. Results show that an upstream wake can result in lower film cooling effectiveness on the blade. Effectiveness magnitudes from superposition of effectiveness data from individual rows are comparable with that from full coverage film cooling. Internal heat transfer measurements are also performed in a high aspect ratio channel and from jet array impingement on a turbulated target wall at large Reynolds numbers. For the channel, three dimple and one discrete rib configurations are tested on one of the wide walls for Reynolds numbers up to 1.3 million. The presence of a turbulated wall and its effect on heat transfer enhancement against a smooth surface is investigated. Heat transfer enhancement is found to decrease at high Re with the discrete rib configurations providing the best enhancement but highest pressure losses. Experiments to investigate heat transfer and pressure loss from jet array impingement are also performed on the target wall at Reynolds numbers up to 450,000. The heat transfer from a turbulated target wall and two jet plates is investigated. A target wall with short pins provides the best heat transfer with the dimpled target wall giving the lowest heat transfer among the three geometries studied.
dc.language.isoen_US
dc.subjectheat transfer
dc.subjectfilm cooling
dc.subjectgas turbines
dc.subjecttip
dc.subjectchannel
dc.subjectimpingement
dc.subjectdimples
dc.subjectribs
dc.titleExperimental study of gas turbine blade film cooling and internal turbulated heat transfer at large Reynolds numbers
dc.typeBook
dc.typeThesis


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