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While modern gas turbine engines operate at hot gas path velocities approaching the speed of sound, few facilities have studied the effects that the flow’s compressibility can have on the adiabatic effectiveness. A new facility at the University of Texas at Austin has been developed to investigate these high Mach number effects and how to appropriately scale laboratory film cooling experiments to engine conditions. This study investigates two film cooling hole geometries, a baseline 7-7-7 shaped film cooling hole and a recent design which has been numerically optimized for increased effectiveness. Both holes are tested at mainstream Mach numbers of 0.25 and 0.50 in a flat plate test section. The optimized hole outperforms the effectiveness of the baseline geometry at all blowing ratios tested, matching the trend in the results of previous studies on these geometries. However, there is a marked decrease in film cooling hole performance as the Mach number is increased.
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Transpiration cooling in hypersonic flow and mutual effect on turbulent transition and cooling performance
This work presents recent advancements in the study of film cooling in hypersonic flows, considering experimental and numerical investigations, with the aim to characterize the wall-cooling performance in different coolant injection and baseflow conditions in a Mach number range 2–7.7. The study explores the mutual interaction between the injected coolant film and the boundary-layer flow, with emphasis on the effects of wall blowing on the boundary-layer characteristics, stability, and transition to turbulence, as well as the effect of transition on wall-cooling performance. Considered flow configurations include cases of effusion cooling in both wall-normal or slightly inclined and wall-parallel blowing, different types of coolant, cases of favorable pressure gradient compared to zero pressure gradient, as well as transpiration cooling cases at different blowing ratios and surface geometries. For the transpiration cooling case, experiments in different hypersonic wind tunnel facilities are presented for flat plate and cone geometries, with coolant injected through C/C porous samples, whereas numerical simulations of modeled porous injection are presented for a flat plate and a blunt cone, showing results for the boundary-layer receptivity with coolant injection and the associated effects on transition and cooling performance. A summary of the main findings is provided along with a critical analysis based on a comparative study to evaluate the effect of each configuration, injection strategy, and key parameters on the boundary-layer flow and the feedback on wall-cooling performance. Conclusions are drawn about potential directions of study for the further development and optimization of the film cooling technique for future hypersonic vehicles.
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- Award ID(s):
- 2146100
- PAR ID:
- 10589335
- Publisher / Repository:
- American Institute of Physics
- Date Published:
- Journal Name:
- Physics of Fluids
- Volume:
- 37
- Issue:
- 2
- ISSN:
- 1070-6631
- Format(s):
- Medium: X
- Sponsoring Org:
- National Science Foundation
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