The increasing trend of high stage pressure ratio with increased aerodynamic loading has led to reduction in stable operating range of centrifugal compressors with stall and surge initiating at relatively higher mass flow rates. The casing treatment technique of stall control is found to be effective in axial compressors, but very limited research work is published on the application of this technique in centrifugal compressors. Present research was aimed to investigate the effect of casing treatment on the performance and stall margin of a high speed, 4 : 1 pressure ratio centrifugal compressor through numerical simulations using ANSYS CFX software. Three casing treatment configurations were developed and incorporated in the shroud over the inducer of the impeller. The predicted performance of baseline compressor (without casing treatment) was in good agreement with published experimental data. The compressor with different inducer casing treatment geometries showed varying levels of stall margin improvement, up to a maximum of 18%. While the peak efficiency of the compressor with casing treatment dropped by 0.8%–1% compared to the baseline compressor, the choke mass flow rate was improved by 9.5%, thus enhancing the total stable operating range. The inlet configuration of the casing treatment was found to play an important role in stall margin improvement.

Winglets are plane surfaces with certain thickness and different shapes. Winglets are used in aircraft to reduce wing tip vortex which is created due to differential pressure in between pressure surface and suction surface. In transonic axial compressor, rotor tip leakage vortex interaction with shock layer and shroud boundary layer leads to total pressure loss and initiation of stall phenomenon. Effect of tip winglets are investigated in compressor rotor cascade. Cascade investigation shows that rotor tip winglet are able to reduce total pressure loss due to tip leakage flow and blade passage secondary flow. Cascade studies are performed with winglet on blade suction side, pressure side and combination of both. From cascade studies it is revealed that suction side winglet are aerodynamically better than pressure side and combined winglets. Owing to favorable results of tip winglet on compressor cascade performance, it was assumed that tip winglets would enhance overall performance of transonic compressor stage with rotating rotor. Results of present CFD simulations have predicted both positive and negative effects of winglets. Effect of different winglet configurations on pressure side and suction side of rotor blade tip are investigated to analyze the compressor stage overall performance. Rotor tip winglets are able to increase stage total pressure ratio compare to the baseline stage without winglet. Stage with winglets have shown better performance in choke region. Winglets are able to vary rotor blade loading from hub to tip region. Presence of winglet has shown ability to reduce to total pressure loss in trailing edge wake region. Stall margin is decreased in compressor stage with winglets due to more blockage towards trailing edge in tip region.

No abstract

No abstract

Numerical studies have been carried out on the effectiveness of trailing edge Gurney flap on a transonic axial compressor rotor. The baseline geometry of the rotor blade was modified at the trailing edge by introducing Gurney flaps of varying depth and span-wise length, viz. 1 mm, 2 mm and 3 mm depth with 20% span length of Gurney flap from tip (designated as GF1-20, GF2-20 and GF3-20 respectively), and 1 mm depth with 50% and 100% span length (designated as GF1-50 and GF1-100 respectively). Geometric models of the compressor rotor without and with Gurney flaps were generated using CATIA V5 software and CFD simulations at 100% design rotor speed were carried out using ANSYS CFX software. Results have shown that the compressor total pressure ratio increased with increase in both depth and spanwise length of Gurney flap. Peak pressure ratio increased from 1.51 for baseline case to 1.58 for rotor GF1-100. However, the peak isentropic efficiency remained almost constant for various Gurney flap configurations, except for GF1-100 which showed a tendency for improvement in efficiency. The stall margin reduced with the introduction of Gurney flap and was lowest for configuration GF1-100 which gave highest peak pressure ratio. Higher blade loading with Gurney flap was responsible for lowering the stall margin. Analysis of the flow through the blade passages has shown clear formation of trailing end vortex structure in the presence of Gurney flap that resulted in bending of the streamlines towards suction surface of the rotor blade, with consequent reduction in flow deviation and increased flow deflection, and hence increased total pressure ratio.