Publications

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Three stereoscopic PIV experiments have been examined to test the effectiveness of self-calibration under varied circumstances. Measurements conducted in a streamwise plane yielded a robust self-calibration that returned common results regardless of the specific calibration procedure, but measurements in the crossplane exhibited substantial velocity bias errors whose nature was sensitive to the particulars of the self-calibration approach. Self-calibration is complicated by thick laser sheets and large stereoscopic camera angles and further exacerbated by small particle image diameters and high particle seeding density. Despite the different answers obtained by varied self-calibrations, each implementation locked onto an apparently valid solution with small residual disparity and converged adjustment of the calibration plane. Therefore, the convergence of self-calibration on a solution with small disparity is not sufficient to indicate negligible velocity error due to the stereo calibration.

The flow over aircraft bays exhibits many characteristics of cavity flows, namely resonant pressures that can create high structural loading. Most studies have represented these bays as rectangular cavities; however, this simplification neglects many features of the actual flight geometry which could affect the unsteady pressure field and resulting loading in the bay. To address this shortcoming, a complex cavity geometry was developed to incorporate more realistic aircraft-bay features including shaped inlets and internal cavity variations. A parametric study of these features at Mach 1.5, 2.0, and 2.5 was conducted to identify key differences from simple rectangular cavity flows. The frequency of the basic rectangular cavity modes could be predicted by theory; however, most complex geometries shifted these frequencies. Geometric changes that constricted the flow tended to enhance cavity modes and create higher pressure fluctuations. Other features, such as a leading edge ramp, lifted the shear layer higher with respect to the aft cavity wall and led to cavity tone suppression. Complex features that introduced spanwise non-uniformity into the shear layer also led to a reduction of cavity tones, especially at the aft end of the cavity.

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Data have been acquired from a spanwise array of fluctuating wall pressure sensors beneath a wind tunnel wall boundary layer at Mach 2, then invoking Taylor's hypothesis allows the temporal signals to be converted into a spatial map of the wall pressure field. Different frequency ranges of pressure fluctuations may be accessed by bandpass filtering the signals. In all frequency ranges, this reveals signatures of coherent structures where negative pressure events are interspersed amongst positive events, with some degree of alternation in the streamwise direction. Within lower frequency ranges, streaks of instantaneously correlated pressure fluctuations elongated in the streamwise direction exhibit a spanwise meander and show apparent merging of pressure events. Coherent length scales based on single-sensor correlations are artificially shortened by neglecting this meander and merging, but are captured correctly using the sensor array. These measurements are consistent with similar observations by other researchers in the velocity field above the wall, and explain the presence of the flat portion of the wall pressure spectrum at frequencies well below those associated with the boundary layer thickness. However, the pressure data lack the common spanwise alternation of positive and negative events found in velocity data, and conversely demonstrate a weak positive correlation in the spanwise direction at low frequencies. © 2013 AIP Publishing LLC.

Particle image velocimetry measurements have been conducted for a Mach 0.8 flow over a wall-mounted hemisphere. The flow is strongly separated, with a mean recirculation length exceeding 5 δ and a mean reverse velocity of -0.2 U∞. The shock foot was found to typically sit just forward of the apex of the hemisphere and move within a range of about ±10 deg. Conditional averages based upon the shock foot location show that the separation shock is positioned upstream along the hemisphere surface when reverse velocities in the recirculation region are strong and is located downstream when they are weaker. The recirculation region appears smaller when the shock is located farther downstream. No correlation was detected of the incoming boundary layer with the shock position, nor with the wake recirculation velocities. These observations are consistent with recent studies concluding that for large strong separation regions, the dominant mechanism is the instability of the separated flow rather than a direct influence of the incoming boundary layer.

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High-frequency pressure sensors were used in conjunction with a high-speed schlieren system to study the growth and breakdown of boundary-layer disturbances into turbulent spots on a 7° cone in the Sandia Hypersonic Wind Tunnel. At Mach 5, intermittent low-frequency disturbances were observed in the schlieren videos. High-frequency secondmode wave packets would develop within these low-frequency disturbances and break down into isolated turbulent spots surrounded by an otherwise smooth, laminar boundary layer. Spanwise pressure measurements showed that these packets have a narrow spanwise extent before they break down. The resulting turbulent fluctuations still had a streaky structure reminiscent of the wave packets. At Mach 8, the boundary layer was dominated by secondmode instabilities that extended much further in the spanwise direction before breaking down into regions of turbulence. The amplitude of the turbulent pressure fluctuations was much lower than those within the second-mode waves. These turbulent patches were surrounded by waves as opposed to the smooth laminar flow seen at Mach 5. At Mach 14, second-mode instability wave packets were also observed. Theses waves had a much lower frequency and larger spanwise extent compared to lower Mach numbers. Only low freestream Reynolds numbers could be obtained, so these waves did not break down into turbulence.

A high-speed schlieren system was developed for the Sandia Hypersonic Wind Tunnel. Schlieren images were captured at 290 kHz and used to study the growth and breakdown of second-mode instabilities into turbulent spots on a 7° cone. At Mach 5, wave packets would intermittently occur and break down into isolated turbulent spots surrounded by an otherwise smooth, laminar boundary layer. At Mach 8, the boundary layer was dominated by second-mode instabilities which would break down into larger regions of turbulence. Second-mode waves surrounded these turbulent patches as opposed to the smooth laminar flow seen at Mach 5. Detailed pressure and thermocouple measurements were also made along the cone at Mach 5, 8 and 14, in a separate tunnel entry. These measurements give an average picture of the transition behavior that complements the intermittent behavior captured by the schlieren system. At Mach 14, the boundary-layer remained laminar so the transition process could not be studied. However, the first measurements of second-mode waves were made in HWT-14.

A reassessment of historical drag coefficient data for spherical particles accelerated in shock-induced flows has motivated new shock tube experiments of particle response to the passage of a normal shock wave. Particle drag coefficients were measured by tracking the trajectories of 1-mm spheres in the flow induced by incident shocks at Mach numbers 1.68, 1.93, and 2.04. The necessary data accuracy is obtained by accounting for the shock tube wall boundary layer growth and avoiding interactions between multiple particles. Similar to past experiments, the current data clearly show that as the Mach number increases, the drag coefficient increases substantially. This increase significantly exceeds the drag predicted by incompressible standard drag models, but a recently developed compressible drag correlation returns values quite close to the current measurements. Recent theoretical work and low particle accelerations indicate that unsteadiness should not be expected to contribute to the drag increase over the relatively long time scales of the experiments. These observations suggest that elevated particle drag coefficients are a quasi-steady phenomenon attributed to increased compressibility rather than true flow unsteadiness. © 2012 American Institute of Physics.

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A reassessment of historical drag coefficient data for spherical particles accelerated in shock-induced flows has motivated new shock tube experiments of particle response to the passage of a normal shock wave. Particle drag coefficients were measured by tracking the trajectories of 1-mm spheres in the wake of incident shocks of Mach numbers 1.68, 1.93, and 2.05. Data clearly show that as the Mach number increases, the drag coefficient increases substantially, consistent with past experiments. This increase significantly exceeds the drag predicted by incompressible standard drag models, but recently developed compressible drag models return values quite close to the current measurements. Low values for the acceleration parameter indicate that unsteadiness should not be expected to contribute to the drag increase. These observations suggest that elevated particle drag coefficients can be attributed to increased compressibility rather than flow unsteadiness.