Study on the flow mechanism of turbine blades considering combustor exit swirling flow

In response to the issue of strong swirl and flow distortion at the inlet of the high-pressure turbine, this paper focuses on the research method of combustor–turbine coupling and conducts relevant research on the aerodynamic flow mechanism of the high-pressure turbine first stage nozzle guide vane (NGV) under different swirl number generated by the swirl generator. The results indicate that the coefficient of total pressure loss (ξ) of the NGV in the non-coupled group increases with the increase in the inlet swirl intensity, while the total pressure loss of the NGV in the coupled group shows a macroscopic decreasing trend with the intensity of swirl generated by the swirl generator. After the swirling flow enters the blade passage, the momentum it carries enables the airflow on the blade surface to obtain a spanwise acceleration, causing significant changes in the intensity and development trend of the secondary flow in the end region. The influence of the swirling flow on the blade surface and the flow field in the end region is remarkable. This study elucidates the development law of swirl generated by the swirl simulator in the combustor and clarifies the migration mode of swirl in the NGV passage and the mechanism of swirl's influence on the flow field of the NGV blade surface and end region.