Experimental Study of Gas Turbine Blade Film Cooling and Heat Transfer (2010)

Modern gas turbine engines require higher turbine-entry gas temperature to improve their\nthermal efficiency and thereby their performance. A major accompanying concern is the heat-up\nof the turbine components which are already subject to high thermal and mechanical stresses.\nThis heat-up can be reduced by: (i) applying thermal barrier coating (TBC) on the surface, and\n(ii) providing coolant to the surface by injecting secondary air discharged from the compressor.\nHowever, as the bleeding off of compressor discharge air exacts a penalty on engine performance,\nthe cooling functions must be accomplished with the smallest possible secondary air injection.\nThis necessitates a detailed and systematic study of the various flow and geometrical parameters\nthat may have a bearing on the cooling pattern.\nIn the present study, experiments were performed in three regions of a non-rotating gas\nturbine blade cascade: blade platform, blade span, and blade tip. The blade platform and blade\nspan studies were carried out on a high pressure turbine rotor blade cascade in medium flow\nconditions. Film-cooling effectiveness or degree of cooling was assessed in terms of cooling hole\ngeometry, blowing ratio, freestream turbulence, coolant-to-mainstream density ratio, purge flow\nrate, upstream vortex for blade platform cooling and blowing ratio, and upstream vortex for blade\nspan cooling. The blade tip study was performed in a blow-down flow loop in a transonic flow\nenvironment. The degree of cooling was assessed in terms of blowing ratio and tip clearance.\nLimited heat transfer coefficient measurements were also carried out. Mainstream pressure loss\nwas also measured for blade platform and blade tip film-cooling with the help of pitot-static\nprobes. The pressure sensitive paint (PSP) and temperature sensitive paint (TSP) techniques were\nused for measuring film-cooling effectiveness whereas for heat transfer coefficient measurement,\ntemperature sensitive paint (TSP) technique was employed.\nResults indicated that the blade platform cooling requires a combination of upstream purge\nflow and downstream discrete film-cooling holes to cool the entire platform. The shaped cooling\nholes provided wider film coverage and higher film-cooling effectiveness than the cylindrical\nholes while also creating lesser mainstream pressure losses. Higher coolant-to-mainstream density ratio resulted in higher effectiveness levels from the cooling holes. On the blade span, at\nany given blowing ratio, the suction side showed better coolant coverage than the pressure side\neven though the former had two fewer rows of holes. Film-cooling effectiveness increased with\nblowing ratio on both sides of the blade. Whereas the pressure side effectiveness continued to\nincrease with blowing ratio, the increase in suction side effectiveness slowed down at higher\nblowing ratios (M=0.9 and 1.2). Upstream wake had a detrimental effect on film coverage. 0%\nand 25% wake phase positions significantly decreased film-cooling effectiveness magnitude.\nComparison between the compound shaped hole and the compound cylindrical hole design\nshowed higher effectiveness values for shaped holes on the suction side. The cylindrical holes\nperformed marginally better in the curved portion of the pressure side. Finally, the concept tip\nproved to be better than the baseline tip in terms of reducing mainstream flow leakage and\nmainstream pressure loss. The film-cooling effectiveness on the concept blade increased with\nincreasing blowing ratio and tip gap. However, the film-coverage on the leading tip portion was\nalmost negligible.