Improvement of Film Cooling Design for Turbine Vane Leading Edge Considering Combustor Outflow

As the interaction between the combustor and the turbine in the aero-engine continues to increase, the film cooling design considering the combustor swirling outflow has become the research focus. The swirling inflow and high-temperature gas first affect the vane leading edge (LE). However, no practical improved solution for the LE cooling design has been proposed considering the combustor swirling outflow. In this paper, the improved scheme of showerhead cooling is carried out around the two ways of adopting the laid-back-fan-shaped hole and reducing the coolant outflow angle. The film cooling effectiveness (& eta;) and the coolant flow state are obtained by PSP (pressure-sensitive-paint) and numerical simulation methods, respectively. The research results show that the swirling inflow increases the film distribution inhomogeneity by imposing the radial pressure gradient on the vane to make the film excessively gather in some positions. The showerhead film cooling adopts the laid-back-fan-shaped hole to reduce the momentum when the coolant flows out. Although this cooling scheme improves the film attachment and increases the surface-averaged film cooling effectiveness (& eta;sur) by as much as 15.4%, the film distribution inhomogeneity increases. After reducing the coolant outlet angle, the wall-tangential velocity of the coolant increases, and the wall-normal velocity decreases. Under the swirl intake condition, both & eta; and the film distribution uniformity are significantly increased, and the growth of & eta;sur is up to 16.5%. This paper investigates two improved schemes to improve the showerhead cooling under the swirl intake condition to provide a reference for the vane cooling design.