Effects of jet arrangement and impingement spacing on flow and tip-wall heat transfer in a two-pass sharp-turn cooling passage of turbine blade

Abstract

To weaken the extreme thermal loads at the blade tip in gas turbines, the state-of-the-art tip cooling relies on serpentine cooling passages and tip bleed holes to reduce surface temperature. However, there are still issues of cracks and fractures at the blade tip during long-term operation. The present study proposes a novel cooling approach for the blade tip that involves impingement-film cooling near the sharp-turn region. For a two-pass cooling channel, the influence of impingement holes arrangement ($L_x/D$; $L_y/D$) and impingement plate-to-target spacing ($H/D$) on the flow and heat transfer near the sharp turn at different Reynolds numbers are investigated. The results indicate that the new configuration enhances the internal convective heat transfer and improves the effectiveness of film cooling. The temperature at the blade tip is significantly reduced compared with the conventional tip cooling structure. Among studied cases, Case 4 ($L_x/D$=2.25, $L_y/D$=2.4) achieves an optimal balance between internal heat transfer and external film cooling performance, which means that a sparser row of impingement holes helps to the heat transfer increment. As the impingement spacing ($H/D$) increases, the cooling effectiveness of the air impingement diminishes. At $H/D$=1.5, the heat transfer coefficient on the inner wall of the blade tip was enhanced by nearly threefold, while the temperature on the outer wall was reduced by about 16%. Although this configuration is accompanied by an increase in flow resistance, the performance evaluation criterion (PEC) still improved by 27%. The present proposed cooling structure can reduce blade tip temperature, offering novel ideals for enhancing heat transfer at blade tips.