Jacob S. Wilson

Passive Control of Compressible Dynamic Stall

The passive control of compressible dynamic stall has been experimentally demonstrated in a dynamically oscillating airfoil test at Reynolds numbers near two million and Mach numbers ranging from M = 0.3 to 0.4. A rotorcraft-type airfoil was retrofitted with a leading edge glove, and tested with and without integral molded vortex generators. Even at relatively low free-stream Mach numbers (M = 0.3 to 0.4), the dynamic stall process on the basic airfoil was initiated by a strong shock boundary layer interaction near the airfoil leading edge. When installed on the original airfoil, the vortex generators failed to control leading edge stall as the free-stream Mach number increased from M = 0.3 to 0.4 due to this locally supersonic leading edge flow at high lift. Although not successful in eliminating dynamic stall, the aerodynamic shape of the glove reduced the extent of a supersonic pocket near the leading edge and the subsequent shock strength; as a result, the airfoil retained trailing edge stall behavior during the dynamic stall event. Without the vortex generators, the glove switched the stall from leading edge to trailing edge, but did not alleviate the stall. The combined use of a transonic leading edge glove and vortex generators lead to the alleviation of dynamic stall pitching moments; the two concepts worked together to alleviate the severe pitching moments during stall, even though neither one was successful when used independently. In some light stall cases the adverse pitching moment associated with dynamic stall was completely eliminated up to M = 0.40. For deep stall cases, the adverse pitching moments were alleviated up to a Mach number of M = 0.375, however the method failed for deep stall at M = 0.40 for this configuration.

Drag-Reduction Mechanisms of Suction-and-Oscillatory-Blowing Flow Control

An active-flow-control study, using steady suction-and-oscillatory-blowing actuators, was conducted on an axisymmetric bluff-body model for a range of Reynolds numbers between 2×106 and 5×106. Prev...

Control of Massive Separation on A Thick-Airfoil Wing: A Computational and Experimental Study

The objective of this research is mutual validation of computational analysis and experimental measurements of baseline and controlled flow over a blunt configuration. A rectangular wing with a 36 percent thick airfoil was selected as a test case, representing a rotor pylon fairing. However, the results are relevant to thick airfoils in general, and contribute to the study of turbulent boundary layer separation and control. Despite the apparently simple geometry, previous measurements on this NACA 0036 airfoil proved challenging for computational analysis. Computational analysis, using DES, provided accurate prediction of the the massive separation and the effects of the oscillatory flow control. Both the aerodynamic coefficients variation with angle of attack, and the pressure distribution on the wing were in good agreement with measurements. These results stand in contrast to the previous failure of incompressible RANS solutions to reproduce the massive separation, its effect on aerodynamic coefficients, and the impact of flow control from t different slots.

Interaction of Synthetic Jet with Boundary Layer Using Microscopic Particle Image Velocimetry

The aerodynamic interaction between an unsteady, inclined synthetic jet and a crossflow boundary layer was studied as a precursor toward applying active flow control concepts for rotor applications, such as dynamic stall control and fuselage drag reduction. Because the flowfield offered numerous challenges from a measurement perspective, several experiments were carried out using a phase-locked, two-dimensional microscopic particle image velocity technique in a building block approach, by adding one complexity after another. The procedure began with boundary layer measurements made on a simple flat plate using the microscopic particle image velocity technique. Velocity measurements were made deep in the viscous sublayer, as close as 20μm from the surface. Following this, the synthetic jet actuator was characterized while operating in quiescent air as well as in crossflow. The results showed that the evolution of the synthetic jet in crossflow was substantially different from its evolution in quiescent air, suggesting that any flow physics or performance prediction (for example, the depth of penetration of the jet into the boundary layer) made based on the quiescent flow conditions may not be applicable in crossflow. All the momentum added to the boundary layer had its source from the synthetic jet actuator, and the penetration of the jet was limited to the viscous sublayer and log layer; the outer layer was unaffected, despite using a jet to freestream velocity ratio of four. Significant effort was also made to validate the microscopic particle image velocity technique and evaluate its capability to accurately resolve such a complex flowfield. To this end, microscopic particle image velocity measurements were compared with hot-wire measurements made on a simple steady jet, as well as an unsteady, periodic synthetic jet. Excellent correlation was found between the two techniques, validating microscopic particle image velocity measurements.

Turbulence Measurements of a Two-Dimensional NACA 0036 Airfoil with Synthetic Jet Flow Control

An active ∞ow control experiment was conducted on a 2-ft chord NACA 0036 airfoil in a 3-ft by 4-ft Wind Tunnel at Re = 1:0 £ 10 6 . The model was equipped with oscillatory jet actuators at x/c = 0.30 and 0.65 that provided 120 Hz periodic excitation at a C„ = 0:86% through 0.06-in wide slots. Three difierent slot conflgurations were tested, including a baseline with no slots. Surface pressure data was collected to compare to previous tests and to combine with turbulence data to aid future CFD modeling efiorts. Turbulence data, measured by hotwire anemometry, was compared with and without ∞ow control. Pressure data corroborates previous test data and provides more points for CFD validation. Hotwire results showed ∞ow control reduced the separated wake size and brought the high Reynolds shear stress layer closer to the airfoil surface. The position of this layer to the surface was altered more signiflcantly than the magnitude of the peak shear stresses. Flow control was shown to increase turbulent energy in the attached boundary layer downstream of the slot but to have little efiect upstream. These results provide further justiflcation to continue assessing the potential of active ∞ow control to reduce separation drag of rotorcraft airframe components.

Flow Physics of Drag Reduction Mechanism using Suction and Pulsed Blowing

aft portion of the model. The drag reduction behavior was scaled using multiple AFC parameters associated with the unique features of the SaOB actuators. Results show that the drag reduction mechanisms associated with the SaOB actuation system include boundary layer suction, wall-jet momentum addition, unsteady shear layer excitation, thrust, and streamwise vortices.

Suction and Oscillatory Blowing Interaction with Boundary Layers

This work presents ongoing experiments toward developing fundamental understanding of suction and oscillatory blowing (SaOB) flow control mechanisms along with development of accurate and practical CFD simulation methodologies for complex unsteady active flow control systems. Experimental and computational studies were conducted on SaOB actuators’ internal flow and external interaction with two zero pressure gradient boundary layers. The experimental work incorporates detailed multi component hot-wire measurements of both laminar and turbulent boundary layers with steady suction and pulsed blowing for a single actuator and an array configuration. Large eddy simulation was employed for the computational model. Simulations were carried out on the actuator internal flow, and the resulting oscillatory blowing jet exit velocity profiles are characterized and fit with a functional form, in order to create simplified boundary conditions for future flow control simulations. Both experimental and computational results show that the suction hole geometry and configuration is an important factor in determining the structure and stability of the downstream laminar as well as turbulent boundary layer flow-fields. Measurements of oscillatory blowing jets interacting with a turbulent boundary layer demonstrate this flow control produces unsteady spanwise and streamwise vorticity components that can be interpreted as counter-rotating streamwise vortex patterns.

Flow Response of Active Flow Control Actuators

Numerical and experimental analysis of a synthetic jet actuator in quiescent air is reported. The study focuses on the actuator itself and on the vorticity field as well as on structures that are generated by the actuator. Large eddies simulation was used for numerical analysis and phase-locked, two-dimensional microscopic-particle image velocimetry technique was used in the experiment. In the numerical simulation, the equations were integrated using a nondissipative scheme that enforces discrete conservation of kinetic energy. The internal cavity flow of the actuator was part of the simulation. The actuator under consideration is of a unique design, fabricated to accommodate practical issues associated with helicopter rotor application. However, it has the same operating principles as a conventional synthetic jet actuator. Harmonics of the forcing frequency were traced in the velocity field. Insight into the evolution of the vortical structures, jet flapping, and onset of turbulence in the jet was retrie...

TOWARDS MORE ACCURATE ANALYSIS OF ACTIVE FLOW CONTROL ACTUATORS

Numerical and experimental analysis of synthetic jet actuator is reported. The study focuses on the actuator itself and on the vorticity field and structures that are generated by the actuator. Phase-locked, 2-D microscopic-PIV technique (MPIV) was used in experiment, and large eddies simulation (LES) was used for numerical analysis. The two methods were first validated for a circular steady jet, continuing with synthetic jet, emanated from a practical device design to quiescent air. The development of vortical structures and their interactions was carefully studied. The insight obtained is an important building block for better understanding of the interaction of synthetic jets and boundary layers. Nomenclature