The high-pressure turbine faces inlet distortion caused by swirl at the combustion chamber exit, leading to complex internal flow mechanisms. The rotational momentum associated with the swirl alters the classic internal flow structure of the turbine, resulting in degraded thermodynamic performance. There is an urgent need for extensive fundamental mechanism research to support subsequent design and optimization efforts. Therefore, this paper conducts a detailed numerical simulation study on the aerodynamic performance and internal flow mechanisms of a high-pressure turbine linear blade row, focusing on five different circumferential positions of the swirl center. The results indicate that when the swirl center is aligned with the blade leading edge, the aerodynamic loss of the turbine blades is minimized. When the swirl center is located between two blades, the rotational momentum enhances the secondary flow at the tip region of NGV2, causing the flow on the blade surface to exhibit radial velocity components. This significantly deteriorates the aerodynamic performance of the NGV. Additionally, the swirl leads to a more uneven distribution of the heat transfer coefficient (HTC) on the blade surface, posing challenges for the cooling design of the blades.
Wu et al. (2025) studied this question.