Leading Edge Separation Control Research Articles

Wavelet analysis of the coherent structures in airfoil leading-edge separation control by bionic coverts

We experimentally investigate leading-edge separation control effect by bionic coverts with various materials and sawteeth shapes in wind tunnel tests. The artificial flexible coverts, bio-inspired by bird covert feathers on upper wings, are hinged at the trailing-edge of a NACA 0018 airfoil at a constant high angle-of-attack of 15°. The chord-based Reynolds number is 1.0 × 105 in the generic range of bird flight in nature and low-speed fixed-wing unmanned aerial vehicles. The velocity profiles in the wake flow are measured by multi-channel hot-wire anemometer. By comparing the mean velocity profiles and root-mean-square velocities, we find the trailing-edge coverts reduced the thickness of the shear layers by 0.05 chord length. The turbulence intensity of the trailing- and leading-edge shear layers are reduced 34% and 5%, respectively. Further wavelet analysis reveals that the large sizes of vortices are considerably suppressed in the time-frequency spectrum. Based on the hot-wire datasets, we develop a novel multi-dimensional genetic algorithm to analyze the featured ordered structures in the shear layers and quantitatively characterize the amplitude modulation between the large- and small-scale flow structures. As a result, we find that the coverts-generated perturbations induce an increase in the high-frequency ( f = 91.2 Hz) coherence between the leading- and trailing-edge shear layers from 40% to 70%, leading to a reduction of the flow separation bubble on the upper wing. The present work reveals that the artificial bionic coverts have leading-edge flow separation control effectiveness and shows the engineering potential for aircrafts and unmanned aerial vehicles.

Control of leading-edge separation on bioinspired airfoil with fluttering coverts.

In this work, the aerodynamic role of the artificial covert feathers (i.e., coverts) on an airfoil is experimentally studied in a wind tunnel to investigate the flow control effect on the leading-edge separation. We apply flexible featherlike devices on a high-angle-of-attack airfoil. We use a hot-wire anemometer to measure the velocity profiles and turbulent fluctuations in the downstream wake flow. As a baseline of flow separation, a two-dimensional NACA 0018 airfoil model is set at the angle of attack of 15 ° at the chord-based Reynolds number of 1.0×10^{5}, causing strong leading-edge and trailing-edge shear layers and a low-speed wake flow area in between as large as 0.35 chord length. When deployed on the upper wing surface, the flexible coverts adaptively flutter under the influence of the local unsteady airflow. Hot-wire measurement results show that the leading-edge coverts effectively suppress the flow separation and reduce the size of the wake flow area. The change of power spectral density shows that the predominant peaks as the fundamental and harmonic frequencies are both attenuated due to the suppression of unsteady motions of the shear layers. On the other hand, the fluttering coverts at the trailing edge modify the trailing-edge shear layer by redistributing the turbulent kinetic energy to the high-frequency components. By simultaneous double-point measurement, we find that the leading-edge and trailing-edge shear layers are drawn closer to each other, and the two shear layers show an increased peak in the coherence spectrum. Further multiscale wavelet analysis shows that the perturbations at the 60% chord length increase the large-scale amplitude modulation of small-scale turbulence and therefore they stabilize the leading-edge and trailing-edge shear layers. Meanwhile, the flow intermittency outside of the wake flow area is attenuated as well. The effective flow control effects in the present work are in good agreement with the previous direct observations of bird flight in literature that the coverts on the upper wing surface play an important role in flow separation control during high-angle-of-attack flight. These findings advance the understanding of aerodynamic contribution of the covers on bird wings and reveal the engineering potential of bioinspired coverts for flow separation control of aircrafts and unmanned air vehicles.

Plasma actuation for leading edge separation control on 300-kW rotor blades with chord length around 1 m at a Reynolds number around 1.6 × 106

The second field test of a plasma-assisted 300-kW wind turbine was performed to investigate the flow-control authority of plasma-actuation technology at Reynolds numbers (Re) exceeding 106. The turbine rotational speed was regulated via pitch activation in the power-saving operation mode to limit Re to approximately 106. Additional generator-torque control was adopted to make the angles of attack exceed those allowable via original control. The leading-edge plasma electrode was activated at 10-min intervals and turbine characteristics with and without plasma actuation were compared using 1-min-averaged SCADA data, including wind conditions measured by a nacelle anemometer, obtained during a 6-day test. Results of this study reveal that the tip speed ratio (TSR) reduces under high-wind-speed conditions owing to additional torque control. Further, the power coefficient drops in the low TSR operating range, and improves upon plasma actuation. This has been statistically proven for at least one TSR bin. This result implies the existence of leading-edge separation under the additional torque control, and that plasma-actuation technology can effectively control the flow in the 1.6–1.8 × 106 Re range. Results of this study demonstrate the potential of and need for further investigation of the proposed technology to facilitate its application for multi-megawatt wind-turbine.

A single unsteady DBD plasma actuator excited by applying two high voltages simultaneously for flow control

Dielectric barrier discharge plasma actuators have attracted a lot of attention to use as new technologies for active flow control. In this paper, an experimental investigation of a single unsteady plasma actuator driven by two simultaneous sinusoidal high voltages is carried out. The aerodynamic performance of this plasma actuator is investigated for flow control of a wind turbine blade. Leading edge separation control at Reynolds number of 0.26×106 and in a wide range of angles of attack including linear, stall, and post-stall regions is considered as the test case. The momentum imparted by the plasma actuator to flow is investigated via measuring induced electric wind velocity, which represents that the induced velocities in the steady mode, with vmax≅5 m/s, are higher than that of the unsteady mode. The important aspect of exciting the unsteady dielectric barrier discharge plasma actuator in this new approach is improving its efficiency through increasing the authority of momentum addition to flow and reducing the minimum input power for discharge ignition relative to its typical grounded structure. This can be very important practically in flow control applications where the amount of consumed energy is a substantial factor in determining the actuator's efficiency. The obtained aerodynamic results reveal that the unsteady plasma actuator has the best operation in post-stall angles of attack, which is of great importance for operation characteristics of wind turbine blades. The power spectral density of pressure time series illustrates that the unsteady plasma actuator affects the flow through instabilities of the separation layer and vortex shedding structure.

Leading edge separation control by plasma actuation for 300kW wind turbine operated in partial load.

Leading edge separation control by plasma actuation for 300kW wind turbine operated in partial load.

Investigation of airfoil leading edge separation control with nanosecond plasma actuator

A thorough combined numerical and experimental investigation of nanosecond dielectric barrier discharge actuation provides a description of the dynamics of the flow actuation process and elucidates the associated flow control mechanisms.

Flow Separation Control Using Nanosecond Pulse Driven DBD Plasma Actuators

This work continues an ongoing effort aimed at development and use of dielectric barrier discharge (DBD) plasma actuators driven by repetitive nanosecond pulses for high Reynolds number aerodynamic flow control. These actuators are believed to influence the flow via a thermal mechanism which is fundamentally different from more commonly studied AC-DBD actuators. Leading edge separation control on an 8-inch chord NACA 0015 airfoil is demonstrated at various post-stall angles of attack for Mach numbers up to 0.26 (free stream velocity up to 93 m/s) and Reynolds numbers up to 1.15 X 106. The nanosecond (NS) pulse driven DBD is shown to extend the stall angle at low Reynolds numbers by functioning as an active trip. At post-stall angles of attack, the device is shown to excite shear layer instabilities and generate coherent spanwise vortices that transfer momentum from the freestream to the separated region, thus reattaching the flow. This is observed for all high Reynolds numbers and Mach numbers spanning th...

Single-Dielectric Barrier Discharge Plasma Enhanced Aerodynamics: Concepts, Optimization, and Applications

This paper deals with the physics and design of single dielectric barrier discharge plasma actuators for enhanced aerodynamics in a variety of applications. The actuators consist of two electrodes: one exposed to the air and the other covered by a dielectric material. The electrodes are supplied with an alternating current voltage that, at high enough levels, causes the air over the covered electrode to ionize. The ionized air, in the presence of the electric field produced by the electrode geometry, results in a body force vector that acts on the ambient air. The body force is the mechanism for active aerodynamic control. The plasma generation is a dynamic process within the alternating current cycle. The body force per unit volume of plasma has been derived from first principles and implemented in numerical flow simulations. Models for the time and space dependence of the body force on the input voltage amplitude, frequency, electrode geometry, and dielectric properties have been developed and used along with experiments to optimize actuator performance. This paper presents results that highlight the plasma actuator characteristics and modeling approach. This is followed by overviews of some of the applications that include leading-edge separation control on airfoils, dynamic-stall vortex control on oscillating airfoils, and trailing-edge separation control on simulated turbine blades.

Leading-Edge Separation Control by Means of Pulsed Vortex Generator Jets

DOI: 10.2514/1.26176 An airfoil was designed to stall with an abrupt turbulent leading-edge separation and equipped with a pneumatic system for active flow control. The system pulses compressed air through 45 deg skewed slots. Depending on the position of the slots, two very different behaviors were found. With the actuators positioned in the region of the separation line, the control system could not prevent separation. However, a very advantageous influence on the separation resulted with even higher normal forces than in prestall condition. The duty cycle was found to be the major control parameter for this kind of control. When the actuation was positioned well in front of the separation line the influence of the actuation system on the stall behavior changed completely. Here the system was able to prevent the leading-edge separation. The influence of the duty cycle becomes much weaker and maximum normal force scaled mainly with the mean ejected momentum, varied either by the duty cycle or the supply pressure.

Scaling Effects of an Aerodynamic Plasma Actuator

We present experimental results to yield insight into the scalability and control effectiveness of single-dielectricbarrier-discharge plasma actuators for leading-edge separation control on airfoils. The parameters investigated are chord Reynolds number, Mach number, leading-edge radius, actuator amplitude, and unsteady frequency. This includes chord Reynolds numbers up to 1:0 � 106 and a maximum freestream speed of 60 m=s corresponding to a Mach number of 0.176. The main objective of this work is to examine the voltage requirements for the plasma actuators to reattach the flow at the leading edge of airfoils at poststall angles of attack for a range of flow parameters in order to establish scaling between laboratory and full-flight conditions. For the full range of conditions, an optimum unsteady actuator frequency f is found to minimize the actuator voltage needed to reattach the flow, such that F� � fLsep=U1 � 1. At the optimum frequencies, the minimum voltage required to reattach the flow is weakly dependent on chord Reynolds number and strongly dependent on the poststall angle of attack and leading-edge radius. The results indicate that the voltage required to reattach the flow scales as the square of the leading-edge radius.

Active Flow Control on the Stingray Uninhabited Air Vehicle: Transient Behaviour

The application of leading-edge separation control on an uninhabited air vehicle (UAV) with 50-deg leading-edge sweep is investigated experimentally in a full-scale close-return wind tunnel using arrays of synthetic jet actuators. Active flow control is used to enhance vehicle control at moderate and high angles of attack for takeoff and landing activities or gust load alleviation. The surface-mounted synthetic jet actuators are operated in various waveforms where the carrier frequency is at least an order of magnitude higher than the characteristic shedding frequency of the UAV. Actuation yields a suction peak near the leading edge, however, while excitation with a sinusoidal waveform results in a sharp suction peak near the leading edge; pulse modulation yields a larger and wider suction peak. The flow transients associated with controlled reattachment and separation of the flow over the UAV are investigated using amplitude modulation of the actuation waveform by measuring the dynamic surface pressure at different locations on the upper surface of the UAV’s wing. Phase-locked measurements show that the transients, associated with the onset of reattachment and separation, at high angles of attack are accompanied by the shedding of large-scale vortical structures and oscillations of the surface pressure. However, no oscillations are observed when the baseline flow is attached.