The application of swirl lean combustion introduces pronounced non-uniform flow features at the turbine inlet. These inlet distortions change the film-cooling behavior of high-pressure turbine stator vanes and affect the downstream transport of coolant and its interaction with the rotor cooling flow

thereby posing additional challenges for aero-thermal design of the first-stage turbine. This work investigates a first-stage cooled turbine vane of an aero-engine under non-uniform inlet boundary conditions that include temperature hot streaks and total pressure distortion. Numerical simulations are conducted to analyze the cooling air discharge characteristics and film cooling efficiency distributions for three different leading-edge hole arrangements under various swirl intensities. The influence of the vane cooling outflow on downstream rotor cooling performance is further examined. Results show that swirl intensity governs the vane cooling distribution

enhancing the leading-edge cooling effectiveness by approximately 2%~16%

while reducing cooling performance on both suction and pressure surfaces. The upstream vane film cooling air lowers the rotor blade surface temperature by about 30~60 K

and in counter-swirl conditions the cooling effectiveness near the rotor pressure-side leading edge increases by more than 20%. The coupled interaction between swirl and coolant flow is identified as a key factor in determining local cooling performance

and the research provides physical insight and design guidance for integrated combustor-turbine cooling optimization.