To explore AC-DBD's ability in controlling dynamic stall, a practical SC-1095 airfoil of a helicopter was selected, and systematic wind tunnel experiments were carried out through direct aerodynamic measurements. The effectiveness of dynamic stall control under steady and unsteady actuation is verified. The influence of parameters such as constant actuation voltage, pulsed actuation voltage, pulsed actuation frequency and duty ratio on dynamic stall control effect is studied under the flow condition of k=0.15 above the airfoil, and the corresponding control mechanism is discussed. Steady actuation can effectively reduce the hysteresis loop area of dynamic lift, and control the peak drag and moment coefficient. For unsteady actuation, there is an optimal duty ratio DC=50%, which has the best effect in improving the lift and drag characteristics, and there is a threshold of pulsed actuation voltage in dynamic stall control. The optimal dimensionless frequency will not be found; different F + have different control advantages in different aerodynamic coefficients of different pitching stages. Unsteady actuation has obvious control advantages in improving the lift-drag characteristics and hysteresis, while steady actuation can better control the large nose-down moment.
The wind tunnel test was conducted with an NACA 0012 airfoil to explore the flow control effects on airfoil dynamic stall by NS-DBD plasma actuation. Firstly, light and deep dynamic stall states were set, based on the static stall characteristics of airfoil at a Reynolds number of 5.8 × 105. Then, the flow control effect of NS-DBD on dynamic stall was studied and the influence law of three typical reduced frequencies (k = 0.05, k = 0.05, and k = 0.15) was examined at various dimensionless actuation frequencies (F + = 1, F + = 2, and F + = 3). For both light and deep dynamic stall states, NS-DBD had almost no effect on upstroke. However, the lift coefficients on downstroke were increased significantly and the flow control effect at F + = 1 is the best. The flow control effect of the light stall state is more obvious than that of deep stall state under the same actuation conditions. For the same stall state, with the reduced frequency increasing, the control effect became worse. Based on the in being principles of flow separation control by NS-DBD, the mechanism of dynamic stall control was discussed and the influence of reduced frequency on the dynamic flow control was analyzed. Different from the static airfoil flow separation control, the separated angle of leading-edge shear layer for the airfoil in dynamic stall state is larger and flow control with dynamic oscillation is more difficult. The separated angle is closely related to the effective angle of attack, so the effect of dynamic stall control is greatly dependent on the history of angles of attack.
At present, the control capability of dielectric barrier discharge (DBD) plasma actuation covers the flow velocity range of helicopter’s retreating blades, so it is necessary to extend it to the dynamic stall control of rotor airfoils. A DBD plasma actuator was adopted to control the dynamic stall of an oscillating CRA309 airfoil in this paper. The effectiveness of alternating current (AC) DBD plasma actuation on reducing the area of lift hysteresis loop of the oscillating airfoil was verified through pressure measurements at a Reynolds number of 5.2 × 105. The influence of actuation parameters on the airfoil’s lift and moment coefficients was studied. Both steady and unsteady actuation could effectively reduce the hysteresis loop area of the lift coefficients. The flow control effect of dynamic stall was strongly dependent on the history of angle of attack. Compared with the steady actuation, unsteady actuation had more obvious advantages in dynamic stall control, with reducing the area of lift hysteresis loop by more than 30%. The effects of plasma actuation on the airfoil’s flow field at both upward and downward stages were discussed at last.
An experimental investigation on the control effects of the high-frequency streamwise pulsed arc discharge array (HS-PADA) on the double compression ramp shock wave/boundary layer interaction (DCR-SWBLI) was carried out at Mach 2.0. Firstly, two types of ramp configurations were designed. The base flow field and actuation flow field were investigated. The actuation frequencies were 10 kHz and 20 kHz. Fast Fourier transform and root-mean-square methods were applied based on schlieren images. The base flow field indicated that the compression effect of ramps decreased with the lengthen of the first ramp. The results of actuation flow field showed that the 20 kHz actuation was superior to the 10 kHz actuation in weakening the shock wave intensity. The HS-PADA exhibited two types of control effects: modifying the shock structure, which was closely related to the separation zone, and modifying the low-frequency unsteadiness of the shock wave, which might not be related to the separation zone. The first separation shock wave, of which the high-frequency motion was more intense under actuation, may be more sensitive to HS-PADA than the second. Finally, the control mechanism of the HS-PADA on DCR-SWBLI was extracted.
In this paper, a pulsed spark discharge plasma actuator array is deployed to control laminar–turbulent transition in a Mach 3.0 flat-plate boundary layer, and the subtle flow structures are visualized by nanoparticle planar laser scattering (NPLS) technique. Results show that the onset location of turbulence can be brought upstream by plasma actuation, corresponding to forced boundary-layer transition. Hairpin vortex packets evolved from the thermal bulbs play a vital role in the breakdown of laminar flow. With the help of a machine learning tool, all the relevant structures induced by plasma actuation are extracted from NPLS images, and a conceptual model of the hairpin vortex generation is proposed, including three stages: production and lift-up of the high-vorticity region, formation of the $\varLambda$ vortex and evolution of the hairpin vortex.
Hypersonic boundary layer transition is a hot yet challenging problem restricting the development and breakthrough of hypersonic aerodynamics. In recent years, despite great progress made by wind tunnel experiment, transition mechanism and transition prediction, only partial knowledge has been gained so far. In this paper, firstly, the specific scenarios of hypersonic boundary layer transition control are clarified. Secondly, the experimental research progress and mechanism of passive control and active control methods under different hypersonic transition control demands are summarized, with their advantages and disadvantages being analyzed separately. Plasma actuation is easy to produce controllable broadband aerodynamic actuation, which has potential in the field of boundary layer transition control. Hence, the following part of the paper focuses on plasma flow control. The feasibility of plasma actuation to control the hypersonic boundary layer transition is demonstrated and the research ideas are presented. Finally, hypersonic boundary layer transition control methods are summarized and the direction of future research is prospected.
In this paper, an experimental study on the stability of hypersonic plate boundary layer is carried out using a spanwise plasma actuation array. The characteristics and evolution of different kinds of unstable waves in the hypersonic plate boundary layer are analyzed based on the results of linear stability theory and high-frequency pulsation sensors. The typical morphological characteristics of the boundary layer and the macro-control effect of the plasma actuation array are explored through a high-speed schlieren method. Finally, based on grayscale mode extraction and proper orthogonal decomposition (POD), the influences of three different actuation frequencies on the instability waves and characteristic structure of the boundary layer are studied, including the dominant frequency of the first mode wave, the half-frequency of the dominant frequency of the first mode wave, and the dominant frequency of the second mode wave, the change of characteristic structures under the regulation of plasma actuation is further discussed. The corresponding regulation rules and mechanisms are summarized. The results show that the plasma actuation array can advance the starting position of laminar discontinuities and that the induced coherent structure can excite instabilities at an earlier flow-direction position. The actuation can be used to control the stability of the boundary layer by acting on the first mode wave to break the original unstable wave spectrum characteristics. This verifies the ability of extensional array plasma actuation to regulate the stability of the hypersonic plate boundary layer and suggests it has great potential in the promotion of hypersonic boundary layer transition.
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