Time-resolved particle image velocimetry was conducted at 40 kHz using a pulse-burst laser in the supersonic wake of a wall-mounted hemisphere. Velocity fields suggest a recirculation region with two lobes, in which flow moves away from the wall near the centerline and recirculates back toward the hemisphere off the centerline, contrary to transonic configurations. Spatio-temporal cross-correlations and conditional ensemble averages relate the characteristic behavior of the unsteady shock motion to the flapping of the shear layer. At Mach 1.5, oblique shocks develop, associated with vortical structures in the shear layer and convect downstream in tandem; a weak periodicity is observed. Shock motion at Mach 2.0 appears somewhat different, wherein multiple weak disturbances propagate from shear-layer turbulent structures to form an oblique shock that ripples as these vortices pass by. Bifurcated shock feet coalesce and break apart without evident periodicity. Power spectra show a preferred frequency of shear-layer flapping and shock motion for Mach 1.5, but at Mach 2.0, a weak preferred frequency at the same Strouhal number of 0.32 is found only for oblique shock motion and not shear-layer unsteadiness.
The mechanism by which aerodynamic effects of jet/fin interaction arise from the flow structure of a jet in crossflow is explored using particle image velocimetry measurements of the crossplane velocity field as it impinges on a downstream fin instrumented with high-frequency pressure sensors. A Mach 3.7 jet issues into a Mach 0.8 crossflow from either a normal or inclined nozzle, and three lateral fin locations are tested. Conditional ensemble-averaged velocity fields are generated based upon the simultaneous pressure condition. Additional analysis relates instantaneous velocity vectors to pressure fluctuations. The pressure differential across the fin is driven by variations in the spanwise velocity component, which substitutes for the induced angle of attack on the fin. Pressure changes at the fin tip are strongly related to fluctuations in the streamwise velocity deficit, wherein lower pressure is associated with higher velocity and vice versa. The normal nozzle produces a counter-rotating vortex pair that passes above the fin, and pressure fluctuations are principally driven by the wall horseshoe vortex and the jet wake deficit. In conclusion, the inclined nozzle produces a vortex pair that impinges the fin and yields stronger pressure fluctuations driven more directly by turbulence originating from the jet mixing.
Fluid-structure interactions were studies on a 7° half-angle cone in the Sandia Hypersonic Wind Tunnel at Mach 5 and 8 and in the Purdue Boeing/AFOSR Mach 6 Quiet Tunnel. A thin composite panel was integrated into the cone and the response to boundary-layer disturbances was characterized by accelerometers on the backside of the panel. Here, under quiet-flow conditions at Mach 6, the cone boundary layer remained laminar. Artificially generated turbulent spots excited a directionally dependent panel response which would last much longer than the spot duration.
The mechanism by which aerodynamic effects of jet/fin interaction arise from the flow structure of a jet in crossflow is explored using particle image velocimetry measurements of the crossplane velocity field as it impinges on a downstream fin instrumented with high-frequency pressure sensors. A Mach 3.7 jet issues into a Mach 0.8 crossflow from either a normal or inclined nozzle, and three lateral fin locations are tested. Conditional ensemble-averaged velocity fields are generated based upon the simultaneous pressure condition. Additional analysis relates instantaneous velocity vectors to pressure fluctuations. The pressure differential across the fin is driven by variations in the spanwise velocity component, which substitutes for the induced angle of attack on the fin. Pressure changes at the fin tip are strongly related to fluctuations in the streamwise velocity deficit, wherein lower pressure is associated with higher velocity and vice versa. The normal nozzle produces a counter-rotating vortex pair that passes above the fin, and pressure fluctuations are principally driven by the wall horseshoe vortex and the jet wake deficit. The inclined nozzle produces a vortex pair that impinges the fin and yields stronger pressure fluctuations driven more directly by turbulence originating from the jet mixing.
The spanwise variation of resonance dynamics in the Mach 0.94 flow over a finite-span cavity was explored using stereoscopic time-resolved particle image velocimetry (TR-PIV) and time-resolved pressure sensitive paint (TR-PSP). The TR-PSP data were obtained along the cavity floor, whereas the TR-PIV measurements were made in a planform plane just above the cavity lip line. The pressure data showed relatively coherent distributions across the span. In contrast, the PIV showed a significant variation in resonance dynamics to occur across the span in the plane above the cavity. A substantial influence of the sidewalls appears to stem from spillage vortices. At the first cavity mode frequency, streamwise velocity fluctuations were several times higher near the sidewalls in comparison to the centerline values. Importantly, PSDs of streamwise velocity in the region of the spillage vortices showed a large peak to occur at mode one, indicating velocity fluctuations in these regions can have a preferred frequency. The resonance fluctuations in the velocity fields at modes two and three demonstrated a complex spatial dependence that varied with spanwise location.
The development of the unsteady pressure field on the floor of a rectangular cavity was studied at Mach 0.9 using high-frequency pressure-sensitive paint. Power spectral amplitudes at each cavity resonance exhibit a spatial distribution with an oscillatory pattern; additional maxima and minima appear as the mode number is increased. This spatial distribution also appears in the propagation velocity of modal pressure disturbances. This behavior was tied to the superposition of a downstream-propagating shear-layer disturbance and an upstream-propagating acoustic wave of different amplitudes and convection velocities, consistent with the classical Rossiter model. The summation of these waves generates an interference pattern in the spatial pressure amplitudes and resulting phase velocity of the resonant pressure fluctuations.
Pulse-burst particle image velocimetry has been used to acquire time-resolved data at 37.5 kHz of the flow over a finite-width rectangular cavity at Mach 0.8. Power spectra of the particle image velocimetry data reveal four resonance modes that match the frequencies detected simultaneously using high-frequency wall pressure sensors, but whose magnitudes exhibit spatial dependence throughout the cavity. Spatiotemporal cross correlations of velocity to pressure were calculated after bandpass filtering for specific resonance frequencies. Cross-correlation magnitudes express the distribution of resonance energy, revealing local maxima and minima at the edges of the shear layer attributable to wave interference between downstream-and upstream-propagating disturbances. Turbulence intensities were calculated using a triple decomposition and are greatest in the core of the shear layer for higher modes, where resonant energies ordinarily are lower. Most of the energy for the lowest mode lies in the recirculation region and results principally from turbulence rather than resonance. Together, the velocity-pressure cross correlations and the triple-decomposition turbulence intensities explain the sources of energy identified in the spatial distributions of power spectra amplitudes.
Time-resolved particle image velocimetry recently has been demonstrated in high-speed flows using a pulse-burst laser at repetition rates reaching 50 kHz. Turbulent behavior can be measured at still higher frequencies if the field of view is greatly reduced and lower laser pulse energy is accepted. Current technology allows image acquisition at 400 kHz for sequences exceeding 4,000 frames, but for an array of only 128 × 120 pixels, giving the moniker of “postage-stamp PIV.” The technique has been tested far downstream of a supersonic jet exhausting into a transonic crossflow. Two-component measurements appear valid until 100 kHz at which point a noise floor emerges dependent upon the reduction of peak locking. Stereoscopic measurement offers three-component data for turbulent kinetic energy spectra, but exhibits a reduced signal bandwidth and higher noise in the out-of-plane component due to the oblique camera images. The resulting spectra reveal two regions exhibiting power-law dependence describing the turbulent decay. One is the well-known inertial subrange with a slope of -5/3 at high frequencies. The other displays a -1 power-law dependence for a decade of mid-range frequencies corresponding to the energetic eddies measured by PIV, which appears to have been previously unrecognized for high-speed free shear flows.
Boundary-layer transition was measured on a pitched, 7° half-angle cone in a Mach 8 conventional wind tunnel. On a smooth cone, transition via second-mode waves was ob- served at all angles of attack. In addition, naturally-excited stationary crossow waves were apparent in temperature sensitive paint images, but did not appear to lead to transition. Two patterns of roughness elements were used to generate higher-amplitude stationary crossow waves. Breakdown of the stationary waves was observed. The roughness resulted in instability amplitudes nearly an order of magnitude larger than the smooth cone at the same Reynolds numbers and higher instability growth rates. Transition occurred 30% - 40% sooner using the roughness elements with peak amplitudes near 15 - 20%, for α ≥ 4°. A low-frequency, coherent wave was measured at all angles of attack. The calculated phase velocity shows a strong dependence on angle of attack, but the propagation angle is similar for all non-zero α. The measured wave properties are curiously similar to measurements of a suspected tunnel-noise-driven instability made on an elliptic cone at Mach 6.
Fluid-structure interactions were studied on a 7 * half-angle cone in the Sandia Hypersonic Wind Tunnel at Mach 8 over a range of freestream Reynolds numbers b etween 3 . 3 and 14 . 5 x 10 6 / m . A thin panel with tunable structural natural frequencies was integrated into the cone and exposed to naturally developing boundary layers. An elevated panel re sponse was measured during boundary- layer transition at frequencies corresponding to the turbu lent burst rate, and lower vibrations were measured under a turbulent boundary layer. Controlled pert urbations from an electrical discharge were then introduced into the boundary layer at varying freq uencies corresponding to the struc- tural natural frequencies of the panel. The perturbations w ere not strong enough to drive a panel response exceeding that due to natural transition. Instead at high repetition rates, the perturber modified the turbulent burst rate and intermittency on the co ne and therefore changed the condi- tions for when an elevated transitional panel vibration res ponse occurred. Studies were also conducted in the Boeing/AFOSR Mach 6 Quiet Tunnel at Purdue University. Under quiet flow, natural transition does not occur, and the c ontrolled perturbations are the only disturbance source. A clear panel response to turbulent spo ts created by the controlled pertur- bations was observed at varying frequencies of spot generat ion. The quiet-flow measurements confirm the clear relationship between turbulent spot passa ge and panel vibration.
Fluid–structure interactions that occur during aircraft internal store carriage were experimentally explored at Mach 0.58–1.47 using a generic, aerodynamic store installed in a rectangular cavity having a length-to-depth ratio of seven. The store vibrated in response to the cavity flow at its natural structural frequencies, and it exhibited a directionally dependent response to cavity resonance frequencies. Cavity tones excited the store in the streamwise and wall-normal directions consistently, whereas the spanwise response to cavity tones was much more limited. Increased surface area associated with tail fins raised vibration levels. The store had interchangeable components to vary its natural frequencies by about 10–300 Hz. By tuning natural frequencies, mode-matched cases were explored where a prominent cavity tone frequency matched a structural natural frequency of the store. Mode matching in the streamwise and wall-normal directions produced substantial increases in peak store vibrations, though the response of the store remained linear with dynamic pressure. Near mode-matched frequencies, changes in cavity tone frequencies of only 1% altered store peak vibrations by as much as a factor of two. In conclusion, mode matching in the spanwise direction did little to increase vibrations.
Pulse-burst Particle Image Velocimetry (PIV) has been employed to acquire time-resolved data at 25 kHz of a supersonic jet exhausting into a subsonic compressible crossflow. Data were acquired along the windward boundary of the jet shear layer and used to identify turbulent eddies as they convect downstream in the far-field of the interaction. Eddies were found to have a tendency to occur in closely spaced counter-rotating pairs and are routinely observed in the PIV movies, but the variable orientation of these pairs makes them difficult to detect statistically. Correlated counter-rotating vortices are more strongly observed to pass by at a larger spacing, both leading and trailing the reference eddy. This indicates the paired nature of the turbulent eddies and the tendency for these pairs to recur at repeatable spacing. Velocity spectra reveal a peak at a frequency consistent with this larger spacing between shear-layer vortices rotating with identical sign. The spatial scale of these vortices appears similar to previous observations of compressible jets in crossflow. Super-sampled velocity spectra to 150 kHz reveal a power-law dependency of -5/3 in the inertial subrange as well as a -1 dependency at lower frequencies attributed to the scales of the dominant shear-layer eddies.
Stereoscopic particle image velocimetry was used to experimentally measure the recirculating flow within finite-span cavities of varying complex geometry at a freestream Mach number of 0.8. Volumetric measurements were made to investigate the side wall influences by scanning a laser sheet across the cavity. Each of the geometries could be classied as an open-cavity, based on L/D. The addition of ramps altered the recirculation zone within the cavity, causing it to move along the streamwise direction. Within the simple rectangular cavity, a system of counter-rotating streamwise vortices formed due to spillage from along the side wall, which caused the mixing layer to develop a steady spanwise waviness. The ramped complex geometry, due to the presence of leading edge and side ramps, appeared to suppress the formation of streamwise vorticity associated with side wall spillage, resulting in a much more two-dimensional mixing layer.
Fluid-structure interactions were studied on a 7° half-angle cone in the Sandia Hypersonic Wind Tunnel at Mach 8 over a range of freestream Reynolds numbers between 3.3 and 14.5 × 106/m. A thin panel with tunable structural natural frequencies was integrated into the cone and exposed to naturally developing boundary layers. An elevated panel response was measured during boundary-layer transition at frequencies corresponding to the turbulent burst rate, and lower vibrations were measured under a turbulent boundary layer. Controlled perturbations from an electrical discharge were then introduced into the boundary layer at varying frequencies corresponding to the structural natural frequencies of the panel. The perturbations were not strong enough to drive a panel response exceeding that due to natural transition. Instead at high repetition rates, the perturber modified the turbulent burst rate and intermittency on the cone and therefore changed the conditions for when an elevated transitional panel vibration response occurred.
Fluid-structure interactions that occur during aircraft internal store carriage were experimentally explored at Mach 0.58-1.47 using a generic, aerodynamic store installed in a rectangular cavity having a length-To-depth ratio of seven. The store vibrated in response to the cavity flow at its natural structural frequencies, and it exhibited a directionally dependent response to cavity resonance frequencies. Cavity tones excited the store in the streamwise and wall-normal directions consistently, whereas the spanwise response to cavity tones was much more limited. Increased surface area associated with tail fins raised vibration levels. The store had interchangeable components to vary its natural frequencies by about 10-300 Hz. By tuning natural frequencies, mode-matched cases were explored where a prominent cavity tone frequency matched a structural natural frequency of the store. Mode matching in the streamwise and wall-normal directions produced substantial increases in peak store vibrations, though the response of the store remained linear with dynamic pressure. Near mode-matched frequencies, changes in cavity tone frequencies of only 1% altered store peak vibrations by as much as a factor of two. Mode matching in the spanwise direction did little to increase vibrations.
Time-resolved PIV has been accomplished in three high-speed flows using a pulse-burst laser: a supersonic jet exhausting into a transonic crossflow, a transonic flow over a rectangular cavity, and a shock-induced transient onset to cylinder vortex shedding. Temporal supersampling converts spatial information into temporal information by employing Taylor’s frozen turbulence hypothesis along local streamlines, providing frequency content until about 150 kHz where the noise floor is reached. The spectra consistently reveal two regions exhibiting power-law dependence describing the turbulent decay. One is the well-known inertial subrange with a slope of-5/3 at high frequencies. The other displays a-1 power-law dependence for as much as a decade of mid-range frequencies lying between the inertial subrange and the integral length scale. The evidence for the-1 power law is most convincing in the jet-in-crossflow experiment, which is dominated by in-plane convection and the vector spatial resolution does not impose an additional frequency constraint. Data from the transonic cavity flow that are least likely to be subject to attenuation due to limited spatial resolution or out-of-plane motion exhibit the strongest agreement with the-1 and-5/3 power laws. The cylinder wake data also appear to show the-1 regime and the inertial subrange in the near-wake, but farther downstream the frozen-turbulence assumption may deteriorate as large-scale vortices interact with one another in the von Kármán vortex street.
Pulse-burst particle image velocimetry (PIV) has been used to acquire time-resolved data at 37.5 kHz of the flow over a finite-width rectangular cavity at Mach 0.6, 0.8, and 0.94. Power spectra of the PIV data reveal four resonance modes that match the frequencies detected simultaneously using high-frequency wall pressure sensors. Velocity resonances exhibit spatial dependence in which the lowest-frequency acoustic mode is active within the recirculation region whereas the three higher modes are concentrated within the shear layer. Spatio-temporal cross-correlations were calculated from velocity data first bandpass filtered for specific resonance frequencies. The low-frequency acoustic mode shows properties of a standing wave without spatial correlation. Higher resonance modes are associated with alternating coherent structures whose size and spacing decrease for higher resonance modes and increase as structures convect downstream. The convection velocity appears identical for the high-frequency resonance modes, but it too increases with downstream distance. This is in contrast to the well-known Rossiter equation, which assumes a convection velocity constant in space.
Time-resolved particle image velocimetry (PIV) using a pulse-burst laser has been acquired of a supersonic jet issuing into a Mach 0.8 crossflow. Simultaneously, the final pulse pair in each burst has been imaged using conventional PIV cameras to produce an independent two-component measurement and two stereoscopic measurements. Each measurement depicts generally similar flowfield features with vorticity contours marking turbulent eddies at corresponding locations. Probability density functions of the velocity fluctuations are essentially indistinguishable but the precision uncertainty estimated using correlation statistics shows that the pulse-burst PIV data have notably greater uncertainty than the three conventional measurements. This occurs due to greater noise in the cameras and a smaller size for the final iteration of the interrogation window. A small degree of peak locking is observed in the aggregate of the pulse-burst PIV data set. However, some of the individual vector fields show peak locking to non-integer pixel values as a result of real physical effects in the flow. Even if peak locking results entirely from measurement bias, the effect occurs at too low a level to anticipate a significant effect on data analysis.
Mach 0.94 flow over a cavity having a length-to-depth ratio of five was explored using time-resolved particle image velocimetry (TR-PIV) with a burst-mode laser. The data were used to probe the resonance dynamics of the first three cavity (Rossiter) tones. Bandpass filtering was employed to reveal the coherent flow structure associated with each tone. The first Rossiter mode was associated with a propagation of large scale structures in the recirculation region, while the second and third modes contained organized structures consistent with convecting vortical disturbances. The wavelengths of the second and third modes were quite similar to those observed in a previous study by the current authors using phase-averaged PIV. Convective velocities computed using cross correlations in the unfiltered data showed the convective velocity increased with streamwise distance in a fashion similar to other studies. Convective velocities during cavity resonance were found to decrease with decreasing mode number, consistent with the modal activity residing in lower portions of the cavity in regions of lower local mean velocities. The convective velocity fields associated with resonance exhibited a streamwise periodicity consistent with wall-normal undulations in the resonant velocity fields; however, additional work is required to confirm this is not an analysis artifact.
The flow over an aircraft bay is often represented using a rectangular cavity; however, this simplification neglects many features of actual flight geometry that 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, internal cavity structure, and doors. A parametric study of these features was conducted based on fluctuating pressure measurements at subsonic and supersonic Mach numbers. Resonance frequencies and amplitudes increased in the complex geometry compared to a simple rectangular cavity that could produce severe loading conditions for store carriage. High-frequency content and dominant frequencies were generated by features that constricted the flow such as leading-edge overhangs, internal cavity variations, and the presence of closed doors. Broadband frequency components measured at the aft wall of the complex cavities were also significantly higher than in the rectangular geometry. Furthermore, these changes highlight the need to consider complex geometric effects when predicting the flight loading of aircraft bays.
A previous experiment by the present authors studied the flow over a finite-width rectangular cavity at freestream Mach numbers 1.5–2.5. In addition, this investigation considered the influence of three-dimensional geometry that is not replicated by simplified cavities that extend across the entire wind-tunnel test section. The latter configurations have the attraction of easy optical access into the depths of the cavity, but they do not reproduce effects upon the turbulent structures and acoustic modes due to the length-to-width ratio, which is becoming recognized as an important parameter describing the nature of the flow within narrower cavities.
Experiments were performed to understand the complex fluid-structure interactions that occur during aircraft internal store carriage. A cylindrical store was installed in a rectangular cavity having a length-to-depth ratio of 3.33 and a length-to-width ratio of 1. The Mach number ranged from 0.6 to 2.5 and the incoming boundary layer was turbulent. Fast-response pressure measurements provided aeroacoustic loading in the cavity, while triaxial accelerometers provided simultaneous store response. Despite occupying only 6% of the cavity volume, the store significantly altered the cavity acoustics. The store responded to the cavity flow at its natural structural frequencies, and it exhibited a directionally dependent response to cavity resonance. Specifically, cavity tones excited the store in the streamwise and wall-normal directions consistently, whereas a spanwise response was observed only occasionally. The streamwise and wall-normal responses were attributed to the longitudinal pressure waves and shear layer vortices known to occur during cavity resonance. Although the spanwise response to cavity tones was limited, broadband pressure fluctuations resulted in significant spanwise accelerations at store natural frequencies. The largest vibrations occurred when a cavity tone matched a structural natural frequency, although energy was transferred more efficiently to natural frequencies having predominantly streamwise and wall-normal motions.
The flow over an open aircraft bay is often represented in a wind tunnel with a cavity. In flight, this flow is unconfined, though in experiments, the cavity is surrounded by wind tunnel walls. If untreated, wind tunnel wall effects can lead to significant distortions of cavity acoustics in subsonic flows. To understand and mitigate these cavity–tunnel interactions, a parametric approach was taken for flow over an L/D = 7 cavity at Mach numbers 0.6–0.8. With solid tunnel walls, a dominant cavity tone was observed, likely due to an interaction with a tunnel duct mode. An acoustic liner opposite the cavity decreased the amplitude of the dominant mode and its harmonics, a result observed by previous researchers. Acoustic dampeners were also placed in the tunnel sidewalls, which further decreased the dominant mode amplitudes and peak amplitudes associated with nonlinear interactions between cavity modes. This indicates that cavity resonance can be altered by tunnel sidewalls and that spanwise coupling should be addressed when conducting subsonic cavity experiments. Though mechanisms for dominant modes and nonlinear interactions likely exist in unconfined cavity flows, these effects can be amplified by the wind tunnel walls.
Two-component and stereoscopic particle image velocimetry measurements have been acquired in the streamwise plane for supersonic flow over a rectangular cavity of variable width, peering over the sidewall lip to view the depths of the cavity. The data reveal the turbulent shear layer over the cavity and the recirculation region within it. The mean position of the recirculation region was found to be a function of the length-to-width ratio of the cavity, as was the turbulence intensity within both the shear layer and the recirculation region. Compressibility effects were observed in which turbulence levels dropped, and the shear layer thickness decreased as the Mach number was raised from 1.5 to 2.0 and 2.5. Supplemental measurements in the crossplane and the planform view suggest that zones of high turbulence were affixed to each sidewall centered on the cavity lip, with a strip of turbulence stretched out across the cavity shear layer for which the intensity was a function of the length-to-width ratio. These sidewall features are attributed to spillage, which is greatly reduced for the narrowest cavity. Such effects cannot be found in experiments lacking finite spanwise extent.
Sandia’s Hypersonic Wind Tunnel (HWT) became operational in 1962, providing a test capability for the nation’s nuclear weapons complex. The first modernization program was completed in 1977. A blowdown facility with a 0.46-m diameter test section, the HWT operates at Mach 5, 8, and 14 with stagnation pressures to 21 MPa and temperatures to 1400K. Minimal further alteration to the facility occurred until 2008, but in recent years the HWT has received considerable investment to ensure its viability for at least the next 25 years. This has included reconditioning of the vacuum spheres, replacement of the high-pressure air tanks for Mach 5, new compressors to provide the high-pressure air, upgrades to the cryogenic nitrogen source for Mach 8 and 14, an efficient high-pressure water cooling system for the nozzle throats, and refurbishment of the electric-resistance heaters. The HWT is now returning to operation following the largest of the modernization projects, in which the old variable transformer for the 3-MW electrical system powering the heaters was replaced with a silicon-controlled rectifier power system. The final planned upgrade is a complete redesign of the control console and much of the gas-handling equipment.
Particle image velocimetry (PIV) measurements quantified the coherent structure of acoustic tones in a Mach 0.91 cavity flow. Stereoscopic PIV measurements were performed at 10-Hz and two-component, time-resolved data were obtained using a pulse-burst laser. The cavity had a square planform, a length-to-depth ratio of five, and an incoming turbulent boundary layer. Simultaneous fast-response pressure signals were bandpass filtered about each cavity tone frequency. The 10-Hz PIV data were then phase-averaged according to the bandpassed pressures to reveal the flow structure associated with the resonant tones. The first Rossiter mode was associated with large scale oscillations in the shear layer, while the second and third modes contained organized structures consistent with convecting vortical disturbances. The spatial wavelengths of the cavity tones, based on the vertical coherent velocity fields, were less than those predicted by the Rossiter relation. With increasing streamwise distance the spacing between structures increased and approached the predicted Rossiter value at the aft-end of the cavity. Moreover, the coherent structures appeared to rise vertically with downstream propagation. The time-resolved PIV data were bandpass filtered about the cavity tone frequencies to reveal flow structure. The resulting spacing between disturbances was similar to that in the phase-averaged flowfields.
Experiments were conducted at freestream Mach numbers of 0.55, 0.80, and 0.90 in open cavity flows having a length-to-depth ratio L/D of 5 and an incoming turbulent boundary having a thickness of about 0.5D. To ascertain aspect ratio effects, the length-to-width ratio L/W was varied between 1.00, 1.67, and 5.00. Two stereoscopic PIV systems were used simultaneously to characterize the flow in the plane at the spanwise center of the cavity. For each aspect ratio, trends in the mean and turbulence fields were identified, regardless of Mach number. The recirculation region had the weakest reverse velocities in the L/W = 1.67 cavity, a trend previously observed at supersonic Mach numbers. Also, like the previous supersonic experiments, the L/W = 1.00 and L/W = 5.00 mean streamwise velocities were similar. The L/W = 1.00 cavity flows had the highest turbulence intensities, whereas the two narrower cavities exhibited lower turbulence intensities of a comparable level. This is in contrast to previous supersonic experiments, which showed the lowest turbulence levels in the L/W = 1.67 cavity.
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 and 8. To relate the intermittent disturbances to the average characteristics of transition on the cone, the statistical distribution of these disturbances must be known. These include the boundarylayer intermittency, burst rate, and average disturbance length. Traditional low-speed methods to characterize intermittency identify only turbulent/nonturbulent regions. However at high M, instability waves become an important part of the transitional region. Algorithms to distinguish instability waves from turbulence in both the pressure and schlieren measurements are being developed and the corresponding intermittency, burst rate, and average burst length of both regions have been provisionally computed for several cases at Mach 5 and 8. Distinguishing instability waves from turbulence gives a better description of the intermittent boundary layer at high M and will allow the fluctuations associated with boundary-layer instabilities to be incorporated into transitional models.
Previous wind tunnel experiments up to Mach 3 have provided fluctuating wall-pressure spectra beneath a supersonic turbulent boundary layer, which essentially are flat at low frequency and do not exhibit the theorized {psi}{sup 2} dependence. The flat portion of the spectrum extends over two orders of magnitude and represents structures reaching at least 100 {delta} in scale, raising questions about their physical origin. The spatial coherence required over these long lengths may arise from very-large-scale structures that have been detected in turbulent boundary layers due to groupings of hairpin vortices. To address this hypothesis, data have been acquired from a dense spanwise array of fluctuating wall pressure sensors, then invoking Taylor's Hypothesis and low-pass filtering the data allows the temporal signals to be converted into a spatial map of the wall pressure field. This reveals streaks of instantaneously correlated pressure fluctuations elongated in the streamwise direction and exhibiting spanwise alternation of positive and negative events that meander somewhat in tandem. As the low-pass filter cutoff is lowered, the fluctuating pressure magnitude of the coherent structures diminishes while their length increases.
Wind tunnel experiments up to Mach 3 have provided fluctuating wall-pressure spectra beneath a supersonic turbulent boundary layer to frequencies reaching 400 kHz by combining signals from piezoresistive silicon pressure transducers effective at low- and mid-range frequencies and piezoelectric quartz sensors to detect high frequency events. Data were corrected for spatial attenuation at high frequencies and for wind-tunnel noise and vibration at low frequencies. The resulting power spectra revealed the {omega}{sup -1} dependence for fluctuations within the logarithmic region of the boundary layer, but are essentially flat at low frequency and do not exhibit the theorized {omega}{sup 2} dependence. Variations in the Reynolds number or streamwise measurement location collapse to a single curve for each Mach number when normalized by outer flow variables. Normalization by inner flow variables is successful for the {omega}{sup -1} region but less so for lower frequencies. A comparison of the pressure fluctuation intensities with fifty years of historical data shows their reported magnitude chiefly is a function of the frequency response of the sensors. The present corrected data yield results in excess of the bulk of the historical data, but uncorrected data are consistent with lower magnitudes. These trends suggest that much of the historical compressible database may be biased low, leading to the failure of several semi-empirical predictive models to accurately represent the power spectra acquired during the present experiments.
The low-frequency meander of a trailing vortex shed from a tapered fin installed on a wind tunnel wall has been studied using stereoscopic particle image velocimetry in the near-wake at Mach 0.8. Distributions of the instantaneous vortex position reveal that the meander amplitude increases with downstream distance and decreases with vortex strength, indicating meander is induced external to the vortex. Trends with downstream distance suggest meander begins on the fin surface, prior to vortex shedding. Mean vortex properties are unaltered when considered in the meandering reference frame, apparently because turbulent fluctuations in the vortex shape and strength dominate positional variations. Conversely, a large peak of artificial turbulent kinetic energy is found centered in the vortex core, which almost entirely disappears when corrected for meander, though some turbulence remains near the core radius. Turbulence originating at the wind tunnel wall was shown to contribute to vortex meander by energizing the incoming boundary layer using low-profile vortex generators and observing a substantial increase in the meander amplitude while greater turbulent kinetic energy penetrates the vortex core. An explanatory mechanism has been hypothesized, in which the vortex initially forms at the apex of the swept leading edge of the fin where it is exposed to turbulent fluctuations within the wind tunnel wall boundary layer, introducing an instability into the incipient vortex core.
A sub-scale experiment has been constructed using fins mounted on one wall of a transonic wind tunnel to investigate the influence of fin trailing vortices upon downstream control surfaces. Data are collected using a fin balance instrumenting the downstream fin to measure the aerodynamic forces of the interaction, combined with stereoscopic Particle Image Velocimetry to determine vortex properties. The fin balance data show that the response of the downstream fin essentially is shifted from the baseline single-fin data dependent upon the angle of attack of the upstream fin. Freestream Mach number and the spacing between fins have secondary effects. The velocimetry shows that the vortex strength increases markedly with upstream fin angle of attack, though even an uncanted fin generates a noticeable wake. No variation with Mach number can be discerned in the normalized velocity data. Correlations between the force data and the velocimetry suggest that the interaction is fundamentally a result of an angle of attack superposed upon the downstream fin by the vortex shed from the upstream fin tip. The Mach number influence arises from differing vortex lift on the leading edge of the downstream fin even when the impinging vortex is Mach invariant.
The Mach number in the inviscid core of the flow exiting scarfed supersonic nozzles was measured using pitot probes. Nozzle characterization experiments were conducted in a modified section of an obsolete M = 7.3 test section/nozzle assembly on Sandia's Hypersonic Wind Tunnel. By capitalizing on existing hardware, the cost and time required for tunnel modifications were significantly reduced. Repeatability of pitot pressure measurements was excellent, and instrumentation errors were reduced by optimizing the pressure range of the transducers used for each test run. Bias errors in probe position prevented us from performing a successful in situ calibration of probe angle effects using pitot probes placed at an angle to the nozzle centerline. The abrupt throat geometry used in the Baseline and Configuration A and B nozzles modeled the throat geometry of the flight vehicle's spin motor nozzles. Survey data indicates that small (''unmeasurable'') differences in the nozzle throat geometries produced measurable flow asymmetries and differences in the flow fields generated by supposedly identical nozzles. Therefore, data from the Baseline and Configuration A and B nozzles cannot be used for computational fluid dynamics (CFD) code validation. Configuration C and D nozzles replaced the abrupt throat geometry of Baseline and Configuration A and B nozzles with a 0.500-inch streamwise radius of curvature in the throat region. This throat geometry eliminated the flow asymmetries, flow separation in the nozzle throat, and measurable differences between the flow fields from identical nozzles that were observed in Baseline/A/B nozzles. Data from Configuration C and D nozzles can be used for CFD code validation.
Particle image velocimetry (PIV) data have been acquired using three different configurations in the far-field of the interaction of a transverse supersonic jet with a transonic crossflow. The configurations included two-dimensional PIV in the centerline streamwise plane at two overlapping stations, as well as stereoscopic PIV in both the same streamwise plane and the crossplane. The streamwise data show the downstream evolution of the interaction whereas the crossplane data directly reveal its vortex structure. The measurement planes intersect at a common line, allowing a comparison of those mean velocity components and turbulent stresses common to all configurations. All data from the streamwise plane agree to within their estimated uncertainties, but data from the crossplane exhibit reduced velocity and turbulent stress magnitudes by a small but significant degree. Additionally, the vertical positions of the peak velocities are slightly nearer the wall for the crossplane configuration. This comparison suggests that routine methods of uncertainty quantification for data used in the validation of computational models may not fully capture the error sources of an experiment.
The couple on a ball rotating relative to an otherwise quiescent suspension of comparably-sized, neutrally buoyant spheres is studied both experimentally and numerically. Apparent 'slip' relative to the analytical solution for a sphere spinning in a Newtonian fluid (based upon the viscosity of the suspension) is determined in suspensions with volume fractions c ranging from 0.03 to 0.50. This apparent slip results in a decrease of the measured torque on the spinning ball when the radius of the ball becomes comparable with that of the suspended spheres. Over the range of our data, the slip becomes more pronounced as the concentration c increases. At c = 0.25, three-dimensional boundary-element simulations agree well with the experimental data. Moreover, at c = 0.03, good agreement exists between such calculations and theoretical predictions of rotary slip in dilute suspensions.
A stereoscopic particle image velocimetry (PIV) instrument has been constructed for a transonic wind tunnel to study the interaction created by a supersonic axisymmetric jet exhausting from a flat plate into a subsonic compressible crossflow. Data have been acquired in the crossplane of the interaction at a single station in the farfield, in which the bulk particle motion is aligned with the out-of-plane velocity component. The resulting vector fields distinctly show the strength and location of the induced counter-rotating vortex pair as well as the remnant of the horseshoe vortex that wraps around the jet plume as it first exhausts from the nozzle. Data taken for four different values of the jet-to-freestream dynamic pressure ratio reveal the resulting change in vortex strength, size, and position. Vorticity fields were derived from the in-plane velocity data, but limited convergence of the present small data sets prevented any conclusions about the symmetry of the flowfield. Comparison of the present data is made with two-dimensional PIV data previously acquired in the streamwise plane.
A particle image velocimetry instrument has been constructed for a transonic wind tunnel and applied to study the interaction created by a supersonic axisymmetric jet exhausting from a flat plate into a subsonic compressible crossflow. Data have been acquired in two configurations; one is a two-dimensional measurement on the streamwise plane along the wind tunnel centerline, and the other is a stereoscopic measurement in the crossplane of the interaction. The presence of the induced counter-rotating vortex pair is clearly visible in both data sets. The streamwise-plane data determined the strength and location of the vortices using the vertical velocity component while the crossplane data directly provided a measurement of the vortical motion. A comparison of the vertical velocity component measured using each configuration showed reasonable agreement.
Particle image velocimetry data have been acquired in the far field of the interaction generated by an overexpanded axisymmetric supersonic jet exhausting transversely from a flat plate into a subsonic compressible crossflow. Mean velocity fields were found in the streamwise plane along the flowfield centerline for different values of the crossflow Mach number M{sub {infinity}} and the jet-to-freestream dynamic pressure ratio J. The magnitude of the streamwise velocity deficit and the vertical velocity component both decay with downstream distance and were observed to be greater for larger J while M{sub {infinity}} remained constant. Jet trajectories derived independently using the maxima of each of these two velocity components are not identical, but show increasing jet penetration for larger J. Similarity in the normalized velocity field was found for constant J at two different transonic M{sub {infinity}}, but at two lower M{sub {infinity}} the jet appeared to interact with the wall boundary layer and data did not collapse. The magnitude and width of the peak in the vertical velocity component both increase with J, suggesting that the strength and size of the counter-rotating vortex pair increase and, thus, may have a stronger influence on aerodynamic surfaces despite further jet penetration from the wall.
Despite many decades of jet-in-crossflow experimentation, a distinct lack of data remains for a supersonic jet exhausting into a subsonic compressible crossflow. The present investigation seeks to address this deficiency by examining the flowfield structure of a Mach 3.73 jet injected transversely from a flat plate into a subsonic compressible freestream. The experimental results described herein include the mean surface pressure field as mapped using static pressure taps on the flat plate and an identification of flow features by employing an oil-based surface flow tracer. The possibility of flow separation within the nozzle itself also is addressed using pressure taps along the nozzle interior wall, as is the asymmetry of the separation line due to the variation of the local backpressure around the perimeter of the nozzle orifice resulting from the jet-in-crossflow interaction. Pressure data both on the flat plate and within the nozzle are presented at numerous angles with respect to the crossflow freestream direction to provide a breadth of measurements throughout the interaction region. Since the data are intended for use in validating computational models, attention is paid to providing details regarding the experimental geometry, boundary conditions, flowfield nonuniformities, and uncertainty analyses. Eight different sets of data are provided, covering a range of values of the jet-to-freestream dynamic pressure ratio from 2.8 to 16.9 and a freestream Mach number range of 0.5 to 0.8.
The present document summarizes the experimental efforts of a three-year study funded under the Laboratory Directed Research and Development program of Sandia National Laboratories. The Innovative Diagnostics LDRD project was designed to develop new measurement capabilities to examine the interaction of a propulsive spin jet in a transonic freestream for a model in a wind tunnel. The project motivation was the type of jet/fin interactions commonly occurring during deployment of weapon systems. In particular, the two phenomena of interest were the interaction of the propulsive spin jet with the freestream in the vicinity of the nozzle and the impact of the spin rocket plume and its vortices on the downstream fins. The main thrust of the technical developments was to incorporate small-size, Lagrangian sensors for pressure and roll-rate on a scale model and include data acquisition, transmission, and power circuitry onboard. FY01 was the final year of the three-year LDRD project and the team accomplished much of the project goals including use of micron-scale pressure sensors, an onboard telemetry system for data acquisition and transfer, onboard jet exhaust, and roll-rate measurements. A new wind tunnel model was designed, fabricated, and tested for the program which incorporated the ability to house multiple MEMS-based pressure sensors, interchangeable vehicle fins with pressure instrumentation, an onboard multiple-channel telemetry data package, and a high-pressure jet exhaust simulating a spin rocket motor plume. Experiments were conducted for a variety of MEMS-based pressure sensors to determine performance and sensitivity in order to select pressure transducers for use. The data acquisition and analysis path was most successful by using multiple, 16-channel data processors with telemetry capability to a receiver outside the wind tunnel. The development of the various instrumentation paths led to the fabrication and installation of a new wind tunnel model for baseline non-rotating experiments to validate the durability of the technologies and techniques. The program successfully investigated a wide variety of instrumentation and experimental techniques and ended with basic experiments for a non-rotating model with jet-on with the onboard jets operating and both rotating and non-rotating model conditions.
An experiment to measure surface pressure data on a series of three stainless steel simulated parachute ribbons was conducted. During the first phase of the test, unsteady pressure measurements were made on the windward and leeward sides of the ribbons to determine the statistical properties of the surface pressures. Particle Image Velocimetry (PIV) measurements were simultaneously made to establish the velocity field in the wake of the ribbons and its correlation with the pressure measurements. In the second phase of the test, steady-state pressure measurements were made to establish the pressure distributions. In the third phase, the stainless steel ribbons were replaced with nylon ribbons and PIV measurements were made in the wake. A detailed error analysis indicates that the accuracy of the pressure measurements was very good. However, an anomaly in the flow field caused the wake behind the stainless steel ribbons to establish itself in a stable manner on one side of the model. This same stability was not present for the nylon ribbon model although an average of the wake velocity data indicated an apparent 2{degree} upwash in the wind tunnel flow field. Since flow angularity upstream of the model was not measured, the use of the data for code validation is not recommended without a second experiment to establish that upstream boundary condition.