The design of large wind turbines requires a comprehensive and accurate analysis of the dynamic loads of airfoils, so it is of great importance to study the dynamic aerodynamic characteristics of a yaw-oscillating airfoil. In this paper, using “electronic cam” technology and synchronous acquisition of dynamic data, a wind tunnel test of yaw oscillation for the airfoil dynamic “sweep effect” is carried out for the first time, providing previously missing lateral dynamic test data. The results show that the aerodynamic curves of the yaw-oscillating airfoil have an obvious hysteresis effect, induced mainly by a periodic pressure fluctuation on the airfoil suction surface, and the aerodynamic hysteresis characteristics are enhanced with increasing oscillation frequency, initial angle of attack, and amplitude. The hysteresis loops of the lift and pressure drag, as a function of yaw angle, follows a “W” shape, the hysteresis loop of the pitching moment follows an “M” shape, and the hysteresis loop of the unsteady lift increment follows an “∞” shape. The aerodynamic force of the airfoil under negative stroke is higher than that under positive stroke, and the aerodynamic coefficients decrease clearly with increasing oscillation frequency under positive stroke. The pressure fluctuation on the airfoil surface is due to a periodic generation, development, movement, breakdown, dissipation, and reconstruction of shear layer vortices, leading edge vortices, trailing edge vortices, and dynamic separation vortices. The dynamic aerodynamic hysteresis of the yaw-oscillating airfoil occurs essentially because of the dynamic interaction between vortex and vortex, or vortex and airfoil surface boundary layer.

The influencing characteristic for the evolution mechanism of a dynamic stall vortex structure and distributed blowing control on rotor airfoils was investigated. Based on the moving-embedded grid method, the finite volume scheme, and Roe’s FDS scheme, a simulation method for the unsteady flow field of a pitch-oscillating airfoil was established. The flow field of the NACA63-218 airfoil was calculated using Reynolds-averaged Navier–Stokes equations. The evolution processes of different vortex structures during dynamic stall and the principal controlled vortex mechanism affecting aerodynamic nonlinearity were analyzed based on the pressure contours C_p and Q of the flow field structure and the spatiotemporal evolution characteristics of the wall pressure distribution. The research indicated that dynamic stall vortices (DSVs) and shear layer vortices (SLVs) were the major sources of the increase in aerodynamic coefficients and the onset of nonlinear hysteresis. Building upon these findings, the concept of distributed blowing control for DSVs and shear layer vortices (SLVs) was introduced. A comparative analysis was conducted to assess the control effectiveness of dynamic stall with different blowing locations and blowing coefficients. The results indicated that distributed blowing control effectively inhibited the formation of DSVs and reduced the intensity of SLVs. This led to a significant decrease in the peak values of the drag and pitch moment coefficients and the disappearance of secondary peaks in the aerodynamic coefficients. Furthermore, an optimal blowing coefficient existed. When the suction coefficient C_μ exceeded 0.03, the effectiveness of the blowing control no longer showed a significant improvement. Finally, with a specific focus on the crucial motion parameters in dynamic stall, the characteristics of dynamic stall controlled by air blowing were investigated. The results showed that distributed air blowing control significantly reduced the peak pitching moment coefficient and drag coefficient. The peak pitching moment coefficient was reduced by 72%, the peak drag coefficient was reduced by 70%, and the lift coefficient hysteresis loop area decreased by 46%. Distributed blowing jet control effectively suppressed the dynamic stall characteristics of the airfoil, making the unsteady load changes gentler.

The environment for the shipboard landing and takeoff of helicopters is extremely complex and significantly affects their safe flight. To address the intricate characteristics of the flow field during these operations, a simulation method suitable for rotor/ship wake vortex interaction is developed. This method couples the Delayed Detached Eddy Simulation (DDES) method and the momentum source method. The simulation of flow field characteristics of the SFS2 ship model under different conditions reveals that, in a rotor/ship coupling scenario, the inflow velocity in the wake zone of the flight deck is distributed in a “W” shape due to the influence of the rotor blade tip vortex. Under wind shear conditions, the rotor’s influence on the wake is reduced, resulting in smaller velocity fluctuations compared to uniform inflow conditions. Moreover, the detached eddy is suppressed to some extent. It can be concluded that shear flow mitigates the unsteady characteristics of the ship’s wake zone to some extent, which is beneficial to helicopter operations during takeoff and landing.