Type of Document Master's Thesis Author Bubb, James Vernon Author's Email Address firstname.lastname@example.org URN etd-071599-164531 Title The Influence of Pressure Ratio on Film Cooling Performance of a Turbine Blade Degree Master of Science Department Mechanical Engineering Advisory Committee
Advisor Name Title Diller, Thomas E. Committee Co-Chair Ng, Fai Committee Co-Chair Schetz, Joseph A. Committee Member Keywords
- Transonic cascade
- Heat transfer
- Film cooling
Date of Defense 1999-07-26 Availability unrestricted AbstractThe relationship between the plenum to freestream total pressure ratio on film cooling performance is experimentally investigated. Measurements of both the heat transfer coefficient and the adiabatic effectiveness were made on the suction side of the center blade in a linear transonic cascade. Entrance and exit Mach numbers were 0.3 and 1.2 respectively. Reynolds number based on chord and exit conditions is 3 x 10^6. The blade
contour is representative of a typical General Electric first stage, high turning, turbine blade. Tunnel freestream conditions were 10 psig total pressure and approximately 80 C. A chilled air coolant film was supplied to a generic General Electric leading edge showerhead coolant scheme. Pressure ratios were varied from run to run over the ranges of 1.02 to 1.20. The density ratio was near a value of 2. A method to determine both the heat transfer coefficient and film cooling effectiveness from experimental data is outlined.
Results show that the heat transfer coefficient is independent of the pressure ratio over these ranges of blowing parameters. Also, there is shown to be a weak reduction of film cooling
effectiveness with higher pressure ratios. Results are shown for effectiveness and heat transfer coefficient profiles along the
Filename Size Approximate Download Time (Hours:Minutes:Seconds)
28.8 Modem 56K Modem ISDN (64 Kb) ISDN (128 Kb) Higher-speed Access thesismain.pdf 3.10 Mb 00:14:20 00:07:22 00:06:27 00:03:13 00:00:16
If you have questions or technical problems, please Contact DLA.