Internal flow mechanism of an aggressive compressor transition duct

In aero engines, designing shorter aggressive compressor transition ducts contributes to improved performance and weight savings. An experimental and numerical investigation has been carried out to explore the internal flow mechanism of an aggressive compressor transition duct, which laid the foundation for subsequently reducing the axial length of the transition duct. Two different total pressure and flow angle profiles were created through the gauzes to simulate the design condition and the near stall (NS) condition in the aero engine. The results showed that in the downward path, flow separation was more likely to occur in the stator hub corner. Hub corner stall occurred at NS condition as a result of a decrease in total pressure and an increase in flow angle, which significantly increased total pressure loss. The existence of hub corner vortex and the uneven flow field at the stator outlet contribute to reducing the tendency for the flow along hub to separation. However, the mixing of these vortices also increased the duct loss. And the velocity deficit produced by the stator hub corner separation/stall was intensified by the adverse pressure gradient in the hub region after the stator.