Kun Du

Effects of the layout of film holes near the vane leading edge on the endwall cooling and phantom cooling of the vane suction side surface

ABSTRACT In the current research, effects of the layout of film holes near the first-stage vane leading edge on the endwall cooling and phantom cooling of the vane suction side surface were numerically studied. The computational results indicate that the case with a positive film-hole angle achieves a higher cooling effectiveness level on the endwall and vane suction side surface compared to the case with a corresponding negative film-hole angle. Furthermore, the location of the film hole has a significant influence on the cooling performance of the endwall and vane suction side surface. In addition, the case with a smaller distance from film holes to the vane stagnation also attains a slightly higher cooling effectiveness (phantom cooling effectiveness) on the vane suction side surface.

Effects of tip cavity geometries on the aerothermal performance of the transonic turbine blade with cavity tip

The inevitable transonic tip leakage flow in unshrouded high pressure gas turbines induces aerodynamic loss and gives rise to heavy thermal loads on the blade tip. In the pursuit of superior gas turbine engine, cavity tips are widely used in high-pressure turbines to reduce the over-tip leakage flow. The present research aims to numerically investigate the influence of the depth of cavity and thickness of squealer rim on the aerothemal performance of the transonic turbine blade. Three-dimensional (3D) Reynolds-Averaged Navier-Stokes (RANS) equations with Spalart-Allmaras turbulence were solved to perform the simulation. The obtained results indicate that the variations of the cavity depth and squealer rim thickness have significant effects on the aerothermal performance of the transonic turbine blade tip regions. The phenomena in this transonic setup are different from the results obtained in subsonic condition. This study quantified that the improvement of aerodynamic efficiency contributed to the enhancement of heat flux on the blade tip in transonic condition by changing the geometry of tip cavity.

Influence of the pressure side injection slot on the cooling performance of endwall surface

ABSTRACT In modern gas turbine engines, the first stage vane endwall endures high thermal load with the increase of the turbine inlet temperature and the uniformity of the temperature distribution at combustor outlet. Moreover, the endwall secondary flow forces the coolant flow toward the suction side, resulting in hot regions along the pressure side endwall. In the worst case, hot regions lead to thermal failure. In order to ensure that the gas turbine operates safely, advanced cooling techniques are urgently needed to be implemented to reduce the hot regions along the pressure side endwall. In the current research, the influences of the pressure side injection slot on the film cooling performance of the endwall surface were numerically investigated. The three-dimensional Reynolds-averaged Navier-Stokes (RANS) equations combined with the shear stress transport (SST) turbulence model were solved to conduct the simulations. Cases with different injection slot configurations have been simulated. The results indicate that the hot region along the pressure side endwall is significantly reduced by introducing the pressure side injection slot. The coolant from the pressure side injection slot is assisted by the pressure side vertical flow toward the adjacent vane suction side. Therefore, the coolant coverage and the cooling effectiveness are increased. In this study, the expanded slot (ES) achieves a larger cooling effectiveness than the normal slot (NS) and convergent slot (CS) at a small blowing ratio M = 0.5. In contrast, the CS obtains a larger cooling effectiveness than the NS and ES at M = 1.0 and M = 1.5. In addition, the introduction of the pressure side injection slot has a small influence on the aerodynamic performance of the vane cascade.