A Numerical Investigation Into Secondary Flows at the Inlet to a Low-Pressure Compressor Transition Duct

To reduce fuel-burn and CO2 emissions from aero gas turbines there is a drive towards very-high bypass ratio and smaller ultra-high-pressure ratio core engine technologies. However, this makes the design of the ducts connecting various compressor spools more challenging as the larger required radius change increases their aerodynamic loading. This is exacerbated for the duct which feeds the engine core as it must accept the relatively low-quality flow produced by the fan root. This is characterised by a hub-low pressure profile and large secondary flow structures which will inevitably increase loss and the likelihood of flow separation. This paper presents a numerical design study aimed at modifying an existing fully annular, low-speed test facility to produce larger, more representative rotor loss cores. The CFD domain comprises a single stage axial compressor and transition duct representative of the low-pressure spool in a high bypass ratio turbofan. The rotating and stationary frames were coupled using a sliding mesh and a uRANS approach with Reynolds Stress closure for the turbulence modelling. The methodology was first validated using existing experimental data before examining a number of parameters such as inlet boundary layer thickness, inlet swirl, tip gap, blade solidity, thickness, lean. It was found that a combination of an increased inlet hub boundary layer and a reduction in solidity sufficiently increased the hub loading and encouraged the development of significantly large loss cores. Preliminary CFD results also showed that the increased rotor loss cores promoted better mixing through the outlet guide which subsequently reduced hub boundary layer thickness and secondary flows. Despite incurring a larger total pressure loss this suggested that the transition duct moved further from stall and could therefore potentially be made shorter.