Aspirated compressor is a promising design concept to enhance the power density of the compression system; however, with regard to the rear stages of multistage aspirated compressor, the blade is fairly thin. Limited by the mechanical constraints, it seems impractical to implement the boundary layer suction on the blade suction surface. So the question arising is can we replace the blade suction surface with other feasible flow control methods without implementing extra device on the blade? To address this issue, a compound flow control method, composed of the endwall boundary layer suction and tandem blade, is proposed. The design philosophy is to utilize the EBLS to suppress the three-dimensional corner stall while to use the tandem blade to control the two-dimensional airfoil flow separation. The endwall boundary layer suction is barely implemented in the forward blade, whereas the corner flow in the rear blade is restrained by the flow through the gap between the forward and rear blades. The preliminary implement strategy of the compound flow control was presented and then applied in the design of a highly loaded aspirated compressor outlet vane. Three-dimensional numerical simulations were carried out to validate its effectiveness with different inlet boundary layer distributions. Both flow fields in the outlet vane and its loss characteristics were analyzed. The results show that, by applying the compound flow control, the outlet vane could not only achieve an aggressive loading without incurring large-scale separation at the design point but also have a considerable available incidence range. Due to the implement of the endwall boundary layer suction, the tandem blade can bring out its full potential in the two-dimensional flow control. Moreover, owing to the flow through the gap of the forward and rear blades, the aspiration flow rate required for the suppression of the three-dimensional corner stall can be reduced.

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Aspirated compressor is a promising design concept to promote the working capacity of compression system and thus may result in an intensive interactions with adjacent rows, however, few extensive unsteady flow analysis have been conducted on this type of highly-loaded compressor. This work presents a three-dimensional numerical simulation of a low speed 1.5 stage low-reaction aspirated compressor (LRAC). The boundary layer suction is only implemented in the outlet vane to control the corner stall and boundary layer separation while no active flow control method is applied in the rotor. The total aspiration flow rate is around 3%. Both aspiration slot and plenum were integrated into the computational domain. Two operating points were selected with the aim to investigate the unsteady effects on the performance of the LRAC and to provide a preliminary unsteady description for this type of aspirated compressor. It is found that compared with the differences in 1D total aerodynamic parameters, evident departures are found in the radial distributions of stage outlet flow parameters between the steady and time-averaged results. For the unsteady case, the radial distributions of pitchwise-averaged parameters become more uniform due to the redistributed aspiration flow rate and the convection behavior of the rotor wake. For the aspiration scheme in the blade, although one-side-aspiration manner is applied, the aspiration flow rate presents a C-type distribution in the radial direction, and this tendency becomes more prominent for the forward slot and time accurate results. Besides, the fluctuation of aspiration flow rate is mainly focused on the upper span due to the intensive rotor outlet secondary flow here. Moreover, the potential effect of aspirated stator on rotor is also examined preliminarily. It is found that for the LRAC investigated in this paper, the rotor suction side is apt to be influenced by the downstream aspirated stator. Finally, some suggestions on the design of the aspirated compressor are provided.

The influence of local surface roughness of rotor blades on the performance of axial compressor stages were investigated through numerical simulation with local surface roughness added on the suction and pressure surfaces of rotor blades of realistic compressor stage NASA Stage35. First of all, the reliability of a commercial computational fluid dynamic code was validated and the computed performance maps showed a good agreement with experimental data from literatures. Numerical results indicated that the increase in surface roughness in most of the local positions may cause the deterioration of compressor stage performance. The amplitude of decrease in compressor performance due to the addition of surface roughness in outer and inner portions of the span and the area near the leading edge of the rotor blades would be much greater than that in the region near the trailing edge. The roughness added to the pressure surface near the leading edge had less impact on the stage characteristics, including the mass flow rate at the choked point. Thus the compressor characteristic got close to that under normal conditions and showed a wider stable operating range. The mass flow in the choked region and the adiabatic efficiency were less affected by the roughness added to the region near the trailing edge of pressure surface from rotor blades. However, this scheme mentioned before would increase the total pressure ratio to some extent, with the adverse effect of adding roughness on the corresponding suction surface.