Numerical Studies on the Effect of Gurney Flap on Aerodynamic Performance and Stall Margin of a Transonic Axial Compressor Rotor

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.