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Pitching airfoil measurements are known to exhibit significant scatter near stall angles of attack that can make meaningful correlation with modeling or simulation difficult. Application of data-driven clustering algorithms to dynamic stall experiments, conducted at two different research facilities, revealed the presence of furcation within the data scatter. Such furcation render the statistical mean and standard deviation as inadequate to represent the observed cycle-to-cycle variations. After ruling out facility effects, an alternative approach to conventional statistical analysis is developed through the use of cluster averages, associated variances, and group probability. Several existing clustering techniques are tested; however, their shortcomings led to the development of two new data-driven algorithms that use proper orthogonal decomposition to cluster data based on flow phenomena that contribute the most energy to flow variations. Several test cases are used to show the physical mechanisms leading to cycle-to-cycle variations, such as differences in separation location, boundary layer reattachment, occurrence of leading-edge/trailing-edge stall, and presence of a dynamic stall vortex (or vortices). In all cases, these physical processes and their effects are obscured by conventional phase-averaging. Further analyses on the effects of the Mach number, reduced frequency, mean angle, and amplitude of oscillation reveal trends in the probability of a given flow behavior. An initial step towards using these results for advanced semiempirical models using Markov analysis is discussed.

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.

The application of statistical Design of Experiments (DOE) to helicopter wind tunnel testing was explored during two powered rotor wind tunnel entries during the summers of 2012 and 2013. These tests were performed jointly by the U.S. Army Aviation Development Directorate Joint Research Program Office and NASA Rotary Wing Project Office, currently the Revolutionary Vertical Lift Project, at NASA Langley Research Center located in Hampton, Virginia. Both entries were conducted in the 14- by 22-Foot Subsonic Tunnel with a small portion of the overall tests devoted to developing case studies of the DOE approach as it applies to powered rotor testing. A 16-47 times reduction in the number of data points required was estimated by comparing the DOE approach to conventional testing methods. The average error for the DOE surface response model for the OH-58F test was 0.95 percent and 4.06 percent for drag and download, respectively. The DOE surface response model of the Active Flow Control test captured the drag within 4.1 percent of measured data. The operational differences between the two testing approaches are identified, but did not prevent the safe operation of the powered rotor model throughout the DOE test matrices.

8. Through the contribution of these authors the development and lifetime of the fully developed wing-tip vortex is relatively well understood. Many of these experimental methods have had difficulties resolving the vortex core. These difficulties stem from the high velocity gradients, and flow fluctuations that typically accompany the vortex core. In particular it has shown to be difficult to resolve the out-of-plane or axial velocity component of the vortex core. Therefore, it is beneficial to instead examine the in-plane velocity components from which much can be learned about the vortex flow field. Additionally, exact solutions to the Navier-Stokes equations exist for vortex flows. The most basic of these solutions involves superimposing a stagnation point flow and a potential vortex, resulting in axial, circumferential, and radial potential solutions. Also, steady and unsteady viscous solutions for a vortex in stretched flow have been developed. An unsteady solution for a viscous vortex in stretched flow was discovered by Rott 9 , who independently arrived at Burgers’ famous solution 10 . Rott’s solution is circumferentially symmetric, and has time and radial dependence. This effort adapted Rott’s solution into a steady solution for circumferential velocity that is dependent on radial and axial coordinates. This experiment sought to detail the evolution of a wing-tip vortex shed into a favorable pressure gradient. Cross sectional velocities were measured at sequential locations in a mild contraction in the Cal Poly 3 x 4 ft lowspeed wind tunnel to examine the circumferential velocity and vorticity evolutions of the vortex in the streamwise direction. Afterwards, these experimental results were compared with previous experiments as well as an approximate analytical solution to a vortex in stretched flow. Test description The experiment was conducted in the California Polytechnic State University 3 ft by 4 ft low-speed open-circuit indraft wind tunnel. The tunnel inlet has an 11:1 contraction ratio, it has a honeycomb of soda straws followed by three screens at the entrance. The variation in mean velocity across an empty test section was less than 1% and the mean free-stream turbulence intensity was less than 0.5% at a centerline speed of 20 m/s. The wind tunnel was powered by a belt driven, 150 hp, 440 volt three-phase motor. A one-dimensional contraction was designed to fit inside the wind tunnel test section. The contraction created a favorable pressure gradient that would cause stretching in a wing-tip vortex passing through the section. It had increasing curvature for the first 30% of the length, and then

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.

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.

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.

A road simulation system has been developed at the Langley Full-Scale Tunnel (LFST) to support the aerodynamic testing of road vehicles, particularly NASCARclass race cars. The leading edge of the existing ground board was recontoured to alleviate a separation bubble and an active suction boundary layer control system, incorporating a bleed slot and axial flow blower, has been implemented. Performance evaluations include boundary layer surveys at various locations in the vicinity of the car balance with the empty tunnel as well as force measurements with a representative vehicle both with and without the boundary layer control system operating.