Patrick Bowles

Closed-Loop Dynamic Stall Control Using a Plasma Actuator

A closed-loop plasma-actuated control scheme that relies on the ability to detect incipient flow separation is presented. It is based on the heightened receptivity of the boundary layer to disturbances at the onset of flow separation. A plasma actuator introduces periodic disturbances into the flow at a low-powered “sense state”. A pressure sensor measures the resulting pressure fluctuations close to the leading edge. When the sensed pressure fluctuations exceed a predetermined threshold level, the plasma actuator is switched to a high-powered “control state” that is designed to reattach the flow. The method provides not only a precursor for flow separation but also an indicator of when conditions exist where active flow reattachment is no longer needed. This is demonstrated on a periodically oscillating airfoil that is pitched into “light” and “deep” dynamic stall. The performance of the closed-loop scheme is evaluated by comparing the aerodynamic loading parameters to the baseline and open-loop unsteady...

Leading-Edge Separation Control Using Alternating-Current and Nanosecond-Pulse Plasma Actuators

Wind-tunnel experiments were conducted to quantify the effectiveness of ac and nanosecond-pulse single dielectric barrier discharge plasma actuators to suppress leading-edge stall on a NASA Energy Efficient Transport airfoil at Mach numbers up to 0.4 and chord Reynolds numbers up to 2.3×106. The airfoil model was designed to have a removable leading edge to accommodate two different leading-edge plasma-actuator designs, either with a thick ceramic or a thin Kapton dielectric layer. The exposed electrode for both plasma actuators was located at the leading edge of the airfoil. The covered electrode for both was on the suction side of the leading edge. The model was mounted on stages that measured the lift and drag forces and the pitching moment about the quarter-chord location. Both steady and unsteady ac plasma-actuator operation were examined. By its nature, the nanosecond-pulse plasma actuator only operates in unsteady operation. The optimal unsteady frequencies with regard to lift, lift to drag, and pi...

High Mach Number Leading-Edge Flow Separation Control Using AC DBD Plasma Actuators

Wind tunnel experiments were conducted to quantify the e↵ectiveness of alternating current dielectric barrier discharge flow control actuators to suppress leading-edge stall on a NASA energy e cient transport airfoil at compressible freestream speeds. The objective of this research was to increase lift, reduce drag, and improve the stall characteristics of the supercritical airfoil near stall by flow reattachment at relatively high Mach and Reynolds numbers. In addition, the e↵ect of unsteady (or duty cycle) operation on these aerodynamic quantities was also investigated. The experiments were conducted at the University of Notre Dame Mach 0.6 Wind Tunnel for a range of Mach numbers between 0.1 and 0.4 with an airfoil model of chord 30.48 cm at atmospheric conditions corresponding to a Reynolds number range of 560, 000 through 2, 260, 000. Lift and drag forces, as well as the quarter chord moments were measured directly by a sting which reacted on load cells and torque sensors on the outside of the 0.91⇥0.91 m wind tunnel test section. Two leading-edges of the airfoil were fabricated. The first was covered in a Kapton dielectric film of 0.127 mm and had a 7 μm copper electrode, and the second was a thick-dielectric Macor with a copper tape exposed (76 μm thick) electrode. A high voltage AC signal was applied to electrodes for the flow control case. The results show that the plasma actuators were e↵ective at reattaching the leading-edge separated flow as evidenced by the increase in maximum lift coe cient and stall angle. In the post stalled regime, the lift was dramatically increased, by as much as 90%. Drag in the stalled regime was reduced by as much as 28% and the nose down pitching moment was reduced by as much as 40%. Pressure taps on the suction surface confirmed flow reattachment as evidenced by the return of a pressure peak near the leading-edge and better pressure recovery aft of the leading-edge when the active flow control was enabled. Time-averaged PIV confirmed the airflow following the airfoil surface closely. The experiment also showed that lift was increased the most in deep stall when the plasma actuator was operated unsteady with a reduced frequency of unity, whereas in light stall steady operation was preferred. Overall, both AC DBD plasma actuator designs were able to increase the maximum lift coe cient and stall angle of attack for the full range of Mach numbers, with the thick-dielectric Macor leading-edge performing better at Mach 0.4.

Stall detection on a leading-edge plasma actuated pitching airfoil utilizing onboard measurement

This experimental work develops a new technique to identify between separated and unseparated flow states on a plasma actuated, pitching airfoil undergoing incompressible dynamic stall. The approach exploits the flow’s receptivity to periodic disturbances introduced by the actuator at the airfoil’s leading edge; these disturbances act as a mechanism to prevent separation and, additionally, provide a decisive indication of the flow state through their pressure signature. No, light, and deep dynamic stall regimes are investigated. Short-time Fourier transforms and wavelet transforms map the pressure time-series obtained by an in-situ sensor into the time-frequency domain where the detection scheme is most apparent.

Combustion-Powered Actuation for Dynamic-Stall Suppression: High-Mach Simulations and Low-Mach Experiments

An investigation on dynamic-stall suppression capabilities of combustion-powered actuation (COMPACT) applied to a tabbed VR-12 airfoil is presented. In the first section, results from computational fluid dynamics (CFD) simulations carried out at Mach numbers from 0.3 to 0.5 are presented. Several geometric parameters are varied including the slot chordwise location and angle. Actuation pulse amplitude, frequency, and timing are also varied. The simulations suggest that cycle-averaged lift increases of approximately 4% and 8% with respect to the baseline airfoil are possible at Mach numbers of 0.4 and 0.3 for deep and near-deep dynamic-stall conditions. In the second section, static-stall results from low-speed wind-tunnel experiments are presented. Low-speed experiments and high-speed CFD suggest that slots oriented tangential to the airfoil surface produce stronger benefits than slots oriented normal to the chordline. Low-speed experiments confirm that chordwise slot locations suitable for Mach 0.3-0.4 stall suppression (based on CFD) will also be effective at lower Mach numbers.

Improved Understanding of Aerodynamic Damping Through the Hilbert Transform

The equation of motion used to derive the aerodynamic damping coefficient for a single-degree-of-freedom airfoil oscillating in pitch about its quarter-chord is rewritten in analytic signal form through application of the Hilbert transform. The results yield a mathematical framework that can be used to estimate the aerodynamic damping coefficient throughout the entire pitch cycle. The analysis is then applied to experimental data from attached, light, and deep dynamic stall conditions at freestream Mach numbers ranging from 0.2 to 0.6 and Reynolds numbers up to 3.5×106. The Hilbert-transform-based approach is used to demonstrate that the cycle-integrated aerodynamic damping coefficient masks the physics underlying the stabilizing and destabilizing mechanisms of the dynamic stall process. In particular, conditions that exhibit positive cycle-integrated aerodynamic damping may include time intervals of negative aerodynamic damping during the pitch cycle.