In the boundary layer ingesting propulsion system, the compressor suffers from a non-uniform flow field. The compressor operating with distorted inflow continuously results in the loss of aerodynamic performance and stability margin. In this paper, three non-axisymmetric configurations are described for the stator of a transonic compressor to match the non-uniform flow field. The flow fields with distorted inflow at near stall condition are obtained and analyzed, the effects of the prototype stator and the three non-axisymmetric stators on aerodynamic performance are compared in detail. Results show that the non-axisymmetric stator schemes can effectively improve the stability margin of the transonic compressor and the maximum stability margin is relatively increased by 22.3% in all the three non-axisymmetric stators. The non-axisymmetric stator design is effective on decreasing the aerodynamic losses and improving the performance of the compressor operating with distorted inflow. Overall, the results show that in the design of the non-axisymmetric stator, the adoption of a curved-twisted blade and the increase of cascade solidity have the potential to reduce loss sources caused by distorted inflow.

Based on the structural characteristics of bird-wing bifurcation, the blade tip groove structure of a linear compressor cascade for controlling separation of suction corner was studied with the use of numerical calculation. The obtained results indicated that the effects of the tip groove on the aerodynamic performance and the flow field structure could be improved. The tip groove design caused high-pressure fluid to be injected into the corner position of separation, causing the initial corner position of separation to move backward, so the pitchwise and spanwise corner separation to be controlled. Under the conditions of positive angle of incidence, the blade tip groove had a good loss reduction effect. When the angle of incidence was less than the angle of incidence with the minimum loss, the flow loss in the cascade increased by the adoption of the tip groove. The tip groove needed to be installed near the initial position of the corner separation. As in the groove height increases, the suppression effect of the tip groove on the corner separation initially increases and then decreases. When the groove height is equal to 4% of the blade height (H), the flow loss in the cascade reduces by 14.42%.

The adverse impacts of non-uniform inlet flow have been the focus for several decades with the increase of the operating range of engines. A deep understanding of the flow mechanism of distortion passing through a compressor is needed urgently and the improvement of the compressor performance becomes more and more important. In this paper, a non-axisymmetric stator is presented with significant non-axisymmetric characteristics in a transonic compressor to investigate compressor performance and flow field effects. A time-dependent three-dimensional Reynolds-averaged Navier-Stokes equation composed in ‘Fluent Software Pack’ is validated and used to perform the simulations. The flow fields with distorted inlet are obtained and the effects of original stator and non-axisymmetric stator in a transonic compressor are compared. The results are discussed in terms of the effects of non-axisymmetric stator on compressor performance, blockage of flow passage, rotor and stator. The results show that the non-axisymmetric stator influences not only the stator flow field but also the rotor flow field, so the efficiency and total pressure ratio are improved correspondingly.

In order to study the influence of the suction surface spanwise grooves under different Reynolds numbers ( Re) on the loss characteristics of highly loaded compressor cascade and rotor. In the present study, numerical simulations were conducted to investigate the effects of the number and position of the suction surface spanwise groove on aerodynamic performance of the cascade with the final aim to extend the findings to a highly loaded transonic compressor rotor. Besides, the influence mechanism for the performance change of compressor cascade and rotor by spanwise groove at different Re was clarified. The combination 1-2-3-4-gro was selected as an optimal structure in accordance with the data weighting criterion. The investigation illustrates that, at high Re operating conditions, the spanwise grooves force the transition of boundary layer from laminar to turbulent, which causes a large turbulent boundary layer loss and increases the overall loss. But at low Re operating conditions, the laminar separation bubble is suppressed by spanwise grooves and the aerodynamic performance is improved. At Re = [ ], the total pressure loss can be reduced by 4.63% at the angle of attack of 0°. When this group scheme was applied to the highly loaded axial compressor transonic rotor, the stability margin of the rotor can be increased by 5.62% at Re = [ ] condition. Moreover, at 20 km height condition, with the increase of rotating speed, the influence of 1-2-3-4-gro groove group on rotor stability margin is not great, but its efficiency improvement effect is gradually enhanced, which shows that 1-2-3-4-gro can effectively reduce rotor flow loss and increase its aerodynamic performance within this Re range.

Crosswind and angle of attack are critical factors that influence the safe operation of civil aviation engines, which are also essential criteria in the certification of civil engine airworthiness. Conducting research on the effects and mechanisms of angle of attack and crosswind on the integrated characteristics and flow field of nacelle and fan holds significant engineering value. In this study, a high-bypass-ratio turbofan engine serves as the research subject, and the influence of crosswind and angle of attack coupling on the nacelle intake and fan components is analyzed using numerical simulations. The numerical research findings indicate that both individual angle of attack and crosswind, as well as their combination, lead to a specific range of total pressure distortion in the inlet. At an angle of attack of 25°, the extent and degree of circumferential total pressure distortion formed on the aerodynamic interface exhibit minor variations under different crosswind inflow conditions. Under two operating conditions involving left crosswind and right crosswind, both characterized by an angle of attack of 25° and a wind speed of 20 m/s, the combined effects of crosswind and angle of attack on the fan and outlet guide vane (OGV) components are mainly concentrated at the blade tip. These influencing factors collectively lead to total pressure distortion in the inlet flow field, resulting in the interaction between tip leakage flow and passage shock waves. At the same angle of attack, there are differences in the impact of total pressure distortion caused by crosswinds in two directions on the fan. Under the same operating conditions, the total pressure distortion induced by the right crosswind has a greater impact on OGV.