Characterization of the Effect of Helicopter Isolated Blade Vortex on Dynamic Stall

Document Type : Research Article

Authors

1 Department of Aerospace Engineering, Amirkabir University of Technology, Tehran, Iran

2 Department of Aerospace Engineering, Malek Ashtar University of Technology, Tehran, Iran

Abstract

In this research, dynamic stall at sections near the rotor blade tip at a maximum cruise speed of the helicopter with an advanced ratio of 0.35 and cyclic pitching motion, has been studied using computational fluid dynamics simulation. Unsteady Reynolds-averaged Navier–Stokes equations are solved using  model on a domain discretized into a hybrid mesh using finite volume discretization method. Numerical simulation is validated using experimental results of AH1-G helicopter flight tests. Comparison of results indicates that present numerical results match with experimental data well. Dynamic stall occurs as a result of a shock wave in the advancing side which affects the lift coefficient. Interestingly, the effect of the shock wave on the lift coefficient in the regions closer to the blade tip is weakened due to the tip vortex penetration. As a result, few changes are seen in the lift coefficient in these regions in comparison to those of the inner regions of the blade. In addition, the maximum value of lift coefficient in the section closer to the blade tip reduces by 10.2% in comparison to that of the most inner section. Results show that despite the formation of the leading-edge vortex, especially in the inner most sections of the blade, severe dynamic stall does not occur in the retreating side.  In fact, this is due to the weakening of the leading edge vortex by the effect of the radial flow.

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