Session: 12-03 Endwall Film Cooling

Paper Number: 82799

82799 - Investigations on Cooling Hole Patterns Over a Turbine Endwall for Improving Cooling Effectiveness 

Film cooling design over a turbine vane endwall is much challenging because the flows near the endwall are characterized by complicated secondary flow structures, coupled with a strong circumferential pressure gradient. In this study, to improve film cooling performance over this region, the film holes within the endwall passage were redistributed based on the knowledge of the inherent endwall-nearby flow patterns, allowing the film injection to overcome the crossflow and to cover more areas of the endwall passage. The baseline design of the endwall film cooling featured an axial pattern of six rows of discrete holes (12 holes in total) positioned near the passage pressure side and four holes located around the shoulder (the strongest curvature) of the suction side. This original design mainly cooled the pressure side and the shoulder region of the suction side of the passage, leaving the mid-passage region uncooled despite upstream purge flow being included. To protect the mid-passage region, two rows of film holes (13 holes in total) were added at the inlet of the endwall passage and the six rows of holes near the passage pressure side in the baseline design were reduced into two rows of holes (5 holes). The four holes near the suction side shoulder was retained but their positions were shifted upstream to the juncture region of the vane leading edge and endwall in the newly-designed hole distribution. Experimental verification of the newly-design holes and the baseline case were conducted in a four-vane, three-passage linear turbine vane cascade at a low exit Mach number. Adiabatic cooling effectiveness over the surfaces of the two sets of film-cooled endwalls was measured by using the pressure-sensitive paint (PSP) technique. Furthermore, to evaluate the interaction of the film injection with the inherent flow structures within the passage, a five-hole probe was traversed along the pitchwise direction at the downstream of the vane trailing edge, obtaining the change of the total pressure loss and flow vectors at the cascade passage exit relative to the baseline design. Results showed that the newly-designed hole distribution pattern with a blowing ratio of 1.5 for the inlet film holes provided a much better film coverage over the endwall passage than the baseline design with a very slight impact on the aerodynamic performance of the vane cascade.

Presenting Author: Xing Yang Xi'an Jiaotong University

Presenting Author Biography: Xing Yang is a full Associate Professor of Institute of Turbomachinery at the Xi’an Jiaotong University. He obtained his Ph.D. from XJTU in 2018, during which he had a one-year academic stay at the University of Minnesota Twin Cities, US. and in October of that year he joined the XJTU faculty. His major interests include experimental measurements and numerical simulations of flow, heat transfer, and cooling problems within complicated turbine flows, advanced novel turbine cooling designs, and deposition effects on turbine aerothermal performance. Added value of the results has been applied to endwall cooling and internal impingement cooling designs for turbine blades in industrial applications. He has published more than 60 peer-reviewed journal or conference papers and gave a keynote at an international conference. He was awarded the “Annual Top Ten Academic New Stars” of XJTU (2021), Humboldt Research Fellowship for Postdoctoral Researchers (2020), ASME YETEP Award (2020), National Postdoctoral Program for Innovative Talents (2018), ASME IGTI Student Scholarship (2018), and "Aviation Powers Nation with Chinese Heart” Scholarship (2018).

Authors:

Xing Yang Xi'an Jiaotong University
Qiang Zhao Xi'an Jiaotong University
Hang Wu Xi'an Jiaotong University
Zhenping Feng Xi'An Jiaotong University