In this study, an experimental investigation of separation control using a dielectric barrier discharge plasma actuator was performed on an NACA0015 airfoil over a wide range of Reynolds numbers, angles of attack, and nondimensional burst frequencies. The range of the Reynolds number was based on a chord length ranging from 2.52 × 10${}^5$ to 1.008 × 10${}^6$. A plasma actuator was installed at the leading edge and driven by AC voltage. Burst mode (duty-cycle) actuation was applied, with the nondimensional burst frequency ranging between 0.1–30. The control authority was evaluated using the time-averaged distribution of the pressure coefficient $Cp$ and the calculated value of the lift coefficient $C_l$. The baseline flow fields were classified into three types: (1) leading-edge separation; (2) trailing-edge separation; and (3) the hysteresis between (1) and (2). The results of the actuated cases show that the control trends clearly depend on the differences in the separation conditions. In leading-edge separation, actuation with a burst frequency of approximately $F^{+}=$ 0.5 creates a wide negative pressure region on the suction-side surface, leading to an increase in the lift coefficient. In trailing-edge separation, several actuations alter the position of turbulent separation.

In this study, skin friction around a ½-scale Ahmed body was measured experimentally at a Reynolds number of Re = 2 × 10⁵. The slant angle of the Ahmed body was 25 deg, and the yaw angles ranged from 0 deg to 8 deg. This study focused on the flow structure on the slant surface under different cross-wind conditions. A force balance system was applied to measure the aerodynamic drag of the model. The global skin-friction topology was measured by applying a luminescent oil layer with a subgrid data processing algorithm. The method used to measure the skin friction was conducted for the first time on the Ahmed body. The results indicated that the technique is highly capable of extracting the skin-friction topology. For a yaw angle below 3 deg, the flow on the slant surface was not significantly affected by the cross-wind condition, and the drag of the model was nearly constant. However, at yaw angles above 3 deg, the flow on the slant surface was highly affected by the roof longitudinal vortexes on the windward side, leading to a dramatic increase in the drag of the model. High consistency in the drag and skin-friction fields was observed. The detailed skin-friction structure at different yaw angles will be discussed in this study.

Flow over a flat plate with a 5% thickness ratio is investigated by schlieren visualization in compressible low-Reynolds-number conditions. The results show that flow separates at the leading-edge and laminar separation-bubble forms. The position of the maximum root mean square of the schlieren image which is related to the position of the vortex shedding moves downstream as a Mach number increases. Furthermore, the two-dimensional structure of generated vortices is maintained up to the trailing edge at the Mach number of 0.66. The frequency analysis of the time-series intensity value of the schlieren images also shows that the flow is stabilized with increasing the Mach number. The position of the end of the pressure plateau region matches the position where the root-mean-square value of the intensity image becomes a maximum due to vortex shedding.

We investigated the effect of compressibility on the laminar separation bubbles and aerodynamic properties of a 5%-thick blunt, flat plate in unsteady two-dimensional (2D) laminar and large-eddy simulations at chord-based Reynolds numbers (Re) of 6,100 and 11,000. Instantaneous flow fields and time-averaged pressure distributions indicate that the transition from laminar-to-turbulent flow and reattachment of the separated shear layer is delayed. This indicates that the separated shear layer is stabilized due to the compressibility effect on the laminar separation bubble. However, the magnitude of the impact varies with the Reynolds number. For example, at Re = 6,100, the effect on the transition and reattachment points is limited because the laminar separated shear layer at incompressible flows is already stable. Therefore, the large lift oscillation owing to periodic and two-dimensional spanwise vortex shedding is significantly suppressed with an increasing Mach number. Conversely, at Re = 11,000, the transition and reattachment points move downstream more significantly with the increasing Mach number. Consequently, the negative skin friction shear stress owing to the reverse flow inside the expanded separation bubble becomes relatively large, decreasing the total drag with the increasing Mach number.