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Foam metal is a foam-like substance with a high porosity; it has been used in flow control, vibration abatement, and acoustic absorption, mainly based on its physical properties. The aim of the current paper is to investigate the effect of foam metal casing treatments on the stability and acoustic level of a low-speed axial flow compressor. The experimental results show that the casing treatment improves the stall margin by 14.9%, without any efficiency loss. In terms of noise, the SPL of the tonal noise at the third order of BPF is reduced by 3.2 dB, while the SPL of the broadband noise is reduced up to 2.4 dB. The comparison in evolutions of the tip structure in a smooth casing condition and with a casing treatment indicates that the casing treatment affects the origination and the development of the tip leakage vortex. The working mechanism is also discussed.

The present work performed a comprehensive investigation to find the effects of a dual-bleeding port recirculation channel on the aerodynamic performance of a single-stage transonic axial compressor, NASA Stage 37, and optimized the channel’s configuration to enhance the operating stability of the compressor. The compressor’s performance was examined using three parameters: The stall margin, adiabatic efficiency, and pressure ratio. Steady-state three-dimensional Reynolds-averaged Navier–Stokes analyses were performed to find the flow field and aerodynamic performance. The results showed that the addition of a bleeding channel increased the recirculation channel’s stabilizing effect compared to the single-bleeding channel. Three design variables were selected for optimization through a parametric study, which was carried out to examine the influences of six geometric parameters on the channel’s effectiveness. Surrogate-based design optimization was performed using the particle swarm optimization algorithm coupled with a surrogate model based on the radial basis neural network. The optimal design was found to increase the stall margin by 51.36% compared to the case without the recirculation channel with only 0.55% and 0.28% reductions in the peak adiabatic efficiency and maximum pressure ratio, respectively.

Numerical simulations have been performed to study the effect of the circumferential single-grooved casing treatment (CT) at multiple locations on the tip-flow stability and the corresponding control mechanism at three tip-clearance-size (TCS) schemes in a transonic axial flow compressor rotor. The results show that the CT is more efficient when its groove is located from 10% to 40% tip axial chord, and G2 (located at near 20% tip axial chord) is the best CT scheme in terms of stall-margin improvement for the three TCS schemes. For effective CTs, the tip-leakage-flow (TLF) intensity, entropy generation and tip-flow blockage are reduced, which makes the interface between TLF and mainstream move downstream. A quantitative analysis of the relative inlet flow angle indicates that the reduction of flow incidence angle is not necessary to improve the flow stability for this transonic rotor. The control mechanism may be different for different TCS schemes due to the distinction of the stall inception process. For a better application of CT, the blade tip profile should be further modified by using an optimization method to adjust the shock position and strength during the design of a more efficient CT.

Axial slot casing treatment is a common method to extend the stall margin of a compressor. Based on the mechanism of unsteady flow control, this paper redesigns axial slots with large circumferential interval and small opening area. To test the effect of this axial slot structure, unsteady numerical simulations were carried out with different slot areas and circumferential intervals. The results show that this novel axial slot casing treatment can significantly improve compressor stall margin. Meanwhile, compared with the traditional axial slot, the efficiency loss is greatly reduced. The flow field analysis shows that the new axial slot structure proposed in this paper can suppress the development of tip leakage vortex and unsteadiness in the tip region at the near stall condition through decreasing the tip loading periodically. Moreover, we find that the slot area is proportional to the improvement of stability margin. Under the same slot area, an excessive number of slots is not conducive to the improvement of the stability margin.

To explain the effect of tip leakage flow on the performance of an axial-flow transonic compressor, the compressors with different rotor tip clearances were studied numerically. The results show that as the rotor tip clearance increases, the leakage flow intensity is increased, the shock wave position is moved backward, and the interaction between the tip leakage vortex and shock wave is intensified, while that between the boundary layer and shock wave is weakened. Most of all, the stall mechanisms of the compressors with varying rotor tip clearances are different. The clearance leakage flow is the main cause of the rotating stall under large rotor tip clearance. However, the stall form for the compressor with half of the designed tip clearance is caused by the joint action of the rotor tip stall caused by the leakage flow spillage at the blade leading edge and the whole blade span stall caused by the separation of the boundary layer of the rotor and the stator passage. Within the investigated varied range, when the rotor tip clearance size is half of the design, the compressor performance is improved best, and the peak efficiency and stall margin are increased by 0.2% and 3.5%, respectively.

The demands to apply transonic centrifugal compressor have increased in the advanced gas turbine engines. Various techniques are used to increase the aerodynamic performance of the centrifugal compressor. The effects of the inclined leading edges in diffuser vanes of a transonic centrifugal compressor on the flow-field unsteadiness and noise generation are investigated by solving the compressible, three-dimensional, transient Navier–Stokes equations. Diffuser vanes with various inclination angles of the leading edge from shroud-to-hub and hub-to-shroud are numerically modeled. The results show that the hub-to-shroud inclined leading edge improves the compressor performance (2.6%), and the proper inclination angle is effective to increase the stall margin (3.88%). In addition, in this study, the transient pressure variations and radiated noise prediction at the design operating point of the compressor are emphasized. The influences of the inclined leading edges on the pressure waves were captured in time/space domain with different convective velocities. The pressure fluctuation spectra are calculated to investigate the tonal blade passing frequency (BPF) noise, and it is shown that the applied inclination angles in the diffuser blades are effective, not only to improve the aerodynamic performance and stall margin, but also to reduce the BPF noise (7.6 dB sound pressure level reduction). Moreover, it is found that the diffuser vanes with inclination angles could suppress the separation regions and eddy structures inside the passages of the diffuser, which results in reduction of the overall sound pressure level and the broadband noise radiated from the compressor.

Downsizing in engine size is pushing the automotive industry to operate compressors at low mass flow rate. However, the operation of turbocharger centrifugal compressor at low mass flow rate leads to fluid flow instabilities such as stall. To reduce flow instability, surface roughness is employed as a passive flow control method. This paper evaluates the effect of surface roughness on a turbocharger centrifugal compressor performance. A realistic validation of SRV2-O compressor stage designed and developed by German Aerospace Center (DLR) is achieved from comparison with the experimental data. In the first part, numerical simulations have been performed from stall to choke to study the overall performance variation at design conditions: 2.55 kg/s mass flow rate and rotational speed of 50,000 rpm. In second part, surface roughness of magnitude range 0–200 μm has been applied on the diffuser shroud to control flow instability. It was found that completely rough regime showed effective quantitative results in controlling stall phenomena, which results in increases of operating range from 16% to 18% and stall margin from 5.62% to 7.98%. Surface roughness as a passive flow control method to reduce flow instability in the diffuser section is the novelty of this research. Keeping in view the effects of surface roughness, it will help the turbocharger manufacturers to reduce the flow instabilities in the compressor with ease and improve the overall performance.

A casing treatment using inclined oblique slots (INOS) is proposed to improve the stability of the single-stage transonic axial compressor, NASA Stage 37, during operation. The slots are installed on the casing of the rotor blades. The aerodynamic performance was estimated using three-dimensional steady Reynolds-Averaged Navier-Stokes analysis. The results showed that the slots effectively increased the stall margin of the compressor with slight reductions in the pressure ratio and adiabatic efficiency. Three geometric parameters were tested in a parametric study. A single-objective optimization to maximize the stall margin was carried out using a Genetic Algorithm coupled with a surrogate model created by a radial basis neural network. The optimized design increased the stall margin by 37.1% compared to that of the smooth casing with little impacts on the efficiency and pressure ratio.