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.
An experiment using fins mounted on a wind tunnel wall has examined the proposition that the interaction between axially-separated aerodynamic control surfaces fundamentally results from an angle of attack superposed upon the downstream fin by the vortex shed from the upstream fin. Particle Image Velocimetry data captured on the surface of a single fin show the formation of the trailing vortex first as a leading-edge vortex, then becoming a tip vortex as it propagates to the fin's spanwise edge. Data acquired on the downstream fin surface in the presence of a trailing vortex shed from an upstream fin may remove this impinging vortex by subtracting its mean velocity field as measured in single-fin experiments, after which the vortex forming on the downstream fin's leeside becomes evident. The properties of the downstream fin's lifting vortex appear to be determined by the total angle of attack imposed upon it, which is a combination of its physical fin cant and the angle of attack induced by the impinging vortex, and are consistent with those of a single fin at equivalent angle of attack.
A sub-scale experiment has been conducted to study the trailing vortex shed from a tapered fin installed on a wind tunnel wall to represent missile configurations. Stereoscopic particle image velocimetry measurements have been acquired in the near-field for several locations downstream of the fin tip and at different fin angles of attack. The vortex's tangential velocity is found to decay with downstream distance while its radius increases, but the vortex core circulation remains constant. Circulation and tangential velocity rise greatly for increased fin angle of attack, but the radius is approximately constant or slightly decreasing. The vortex axial velocity is always a deficit, whose magnitude diminishes with downstream distance and smaller angle of attack. No variation with Mach number can be discerned in the normalized velocity data. Vortex roll-up is observed to be largely complete by about four root chord lengths downstream of the fin trailing edge. Prior to this point, the vortex is asymmetric in the tangential velocity but the core radius stays nearly constant. Vortical rotation draws low-speed turbulent fluid from the wind tunnel wall boundary layer into the vortex core, which appears to hasten vortex decay and produce a larger axial velocity deficit than might be expected. Self-similarity of the vortex is established even while it is still rolling up. Attempts to normalize vortex properties by the fin's lift coefficient proved unsuccessful.
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.
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.
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.
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.