A complementary experimental and computational study was performed to characterize the test section flowfield of a variable-Mach-number supersonic wind tunnel. A planar laser Rayleigh scattering technique using [Formula: see text] nanocrystals was used to capture cross-sectional flow visualizations at various test section streamwise locations. Reynolds-averaged Navier–Stokes simulations of the wind tunnel flow were performed using a [Formula: see text] shear stress transport turbulence model to compare against the flow visualizations. Mach 3.20 and 3.55 flow conditions were investigated in the wind tunnel with an empty test section. The Rayleigh signal profile near the wind tunnel walls was compared against boundary-layer thickness data from shadowgraph visualizations, pitot survey measurements, and numerical simulation predictions to guide the interpretation of the scattering results. A brief discussion regarding the nature and potential source of certain unexpected scattering patterns in the images was provided. The most distinguishable flowfield characteristics observed in the Rayleigh scattering visualizations were the asymmetric thickness of the boundary layers along the test section walls and the appearance of two floor lobes near the flow-path corners. The results of the numerical simulations showed favorable agreement with the experimental data and were used to provide insight into the mechanism that caused the boundary-layer features observed. This study demonstrated the usefulness of [Formula: see text] Rayleigh scattering to obtain flow visualization data of great value to better understand flow quality in supersonic wind tunnels with two-dimensional nozzles.

A complementary experimental and computational investigation of the flow characteristics in a supersonic wind tunnel are presented. Stereoscopic planar laser Rayleigh scattering of CO2 particles, typically several nanometers to tens of nanometers in size, were used in this study allowing cross-sectional visualization of the supersonic flow and turbulent boundary layers throughout the test section of an academic scale facility. The optical diagnostic measurements show asymmetric boundary layer growth between the top and bottom walls of the test section as well as a disturbed bottom boundary layer due to the induced pressure and velocity gradient. The computational fluid dynamics simulations were used to provide additional detail and insight into the mechanisms driving these features.

This paper describes the ship airwake simulated with sinusoidal lateral gusts over the flight deck of a Simple Frigate Shape 2 (SFS2) geometry using Computational Fluid Dynamics (CFD) simulation and experimental measurements using Time-Resolved Particle Image Velocimetry (TRPIV). The airwake with varying gust periods and magnitudes greatly affects the development of the turbulent wake over the flight deck of an aviation ship and can compromise the launch/recovery of helicopters. To derive the high-fidelity simulation of the airwake over the deck for an initial steady-state ship motion, the Large Eddy Simulation (LES) scheme was used and compared with prior experimental results. Furthermore, for the simulation of the gust effects over the ship deck, the moving mesh method with the RANS (Reynolds-Averaged Navier-Stokes) scheme was directly applied in the CFD simulations and also compared with transient wind tunnel flow field measurements. From these analyses, similar trends were investigated between the CFD results and TRPIV data for the initial steady-state ship motion; it was verified that the moving mesh computational results are in good agreement with the experimental results, in certain aspects.

No AccessEngineering NoteMeasurements of Midspan Flow Interactions of a Low-Aspect-Ratio Circulation Control WingDavid Miklosovic, Robin Imber and Michael Britt-CraneDavid MiklosovicUnited States Naval Academy, Annapolis, MD 21402, Robin ImberNaval Air Systems Command, Patuxent River, MD 20670 and Michael Britt-CraneNaval Sea Systems Command, West Bethesda, MD 20376Published Online:5 Jul 2016https://doi.org/10.2514/1.C033852SectionsRead Now ToolsAdd to favoritesDownload citationTrack citations ShareShare onFacebookTwitterLinked InRedditEmail About References [1] Miklosovic D., Imber R., Britt-Crane M. and Colantonio K., "Mid-Span PIV Measurements of a Low Aspect Ratio Circulation Control Wing," AIAA 30th Applied Aerodynamics Conference, AIAA Paper 2012-2893, June 2012. doi:https://doi.org/10.2514/6.2012-2893 LinkGoogle Scholar[2] Riebe J. 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H. and Pope A., Low-Speed Wind Tunnel Testing, 3rd ed., John Wiley and Sons, Hoboken, NJ, 1999, pp. 368–372. Google Scholar[18] Coleman H. W. and Steele W. G., "Engineering Application of Experimental Uncertainty Analysis," AIAA Journal, Vol. 33, No. 10, 1995, pp. 1888–1896. doi:https://doi.org/10.2514/3.12742 AIAJAH 0001-1452 LinkGoogle Scholar[19] Rogers E. and Donnelly M., "Characteristics of a Dual-Slotted Circulation Control Wing of Low Aspect Ratio Intended for Naval Hydrodynamic Applications," AIAA 42nd Aerospace Sciences Meeting, AIAA Paper 2004-1244, Jan. 2004. doi:https://doi.org/10.2514/6.2004-1244 LinkGoogle Scholar[20] Abramson J., "Characteristics of a Cambered Circulation Control Aerofoil Having Both Upper and Lower Surface Trailing Edge Slots," Naval Surface Warfare Center Carderock Division Technical Rept. NSWCCD-50-TR-2004/030, Bethesda, MD, 2004. Google Scholar Previous article Next article

Numerical simulations using computational fluid dynamics are frequently applied to analyze complex flow fields. However, they have to be validated by matching simulation results to those from canonical flows or experimental measurements. The objective of the present research is to compare results from numerical simulations and wind tunnel measurements for air wakes generated behind ships' superstructures to those from direct in situ measurements. Numerical simulations are performed using COBALT on an unstructured grid system, wind tunnel data are collected from a 4%-scale model, and in situ measurement data are sampled using ultrasonic anemometers mounted above an aft flight deck on a 32.9-m (108 ft)-long research vessel. Reynolds numbers are closely matched for all three approaches concurrently. Two different incoming velocity conditions are compared: a head wind condition and wind 15 degrees off the starboard bow (beta = 0 degrees,-15 degrees, respectively, where beta is the wind yaw angle). Differences in velocity and boundary layers between the three approaches are resolved using unique velocity normalization. The flow structures between beta = 0 degrees and beta = -15 degrees are quite different, i.e., there appears to be strong asymmetric vortical structures over the flight deck for beta = -15 degrees. In general, in situ, computational, and wind tunnel data all show large-scale recirculation motion behind the ship's hangar. However, there are nonnegligible differences between the simulations and wind tunnel measurements compared to the in situ measurements. Differences in velocity angles increase with the yaw angle of the incoming flow.

A stereoscopic particle image velocimetry (SPIV) system was used in a low-speed wind tunnel to measure the external flowfield of a circulation control (CC) wing having an aspect ratio of 1.1. The CC wing tested had a 20% elliptic airfoil section with a trailing edge Coanda surface to increase circulation control effectiveness through a segmented, eightplenum system that could create spanwise massflow efflux profiles. Three-component wake velocity data were acquired in six station planes a quarter-chord downstream of the trailing edge. Four spanwise variations in efflux modeled uniform, linear, zero-tip, and elliptic profiles at an overall momentum coefficient of 0.09 and a Reynolds number of 970,000. The SPIV velocity data revealed the nature of the interaction of the jet with the external flowfield through the time-averaged velocity magnitude, the temporal variability, and the vorticity distribution. The results showed that the elliptical efflux profile was the most beneficial profile not only because of a reduced and more localized unsteadiness distribution, but because of the largest lift-to-drag ratio of greater than 4. The zero-tip efflux profile, thought to show promise in lift augmentation for reduced drag and noise, turned out to be the worst of the four efflux profiles tested because of low lift, high drag, and moderately high wake unsteadiness.

A stereoscopic particle image velocimetry (SPIV) system was used in a low-speed wind tunnel to measure the external flowfield of a circulation control (CC) wing having an aspect ratio of 1.1. This ongoing project, sponsored by the Office of Naval Research, sought to further the knowledge of the jet interactions and the wingtip vortex formation through 3-component, planar velocity surveys. The CC wing tested had a 20% elliptic airfoil section with a trailing edge Coanda surface that was intended to increase circulation control effectiveness through a segmented system that could create spanwise massflow efflux profiles. To date, 1.2 TB of raw SPIV data have been acquired in one of two wake station planes at z/c = 1.25 over the vertical region of 0.13 < y/b < 0.87 (i.e, the tip region). The three-component velocity data revealed the nature of the interaction of the jet with the external flowfield and the temporal variability at an overall momentum coefficient of 0.08 with spanwise variations in the trailing edge efflux. The results from these tests will be used to quantify, for the first time, the effects of spanwise massflow distributions on the 3D velocity field near the trailing edge, the stall modes, jet interactions, and the overall performance of a CC wing of this geometry.

The present work was undertaken to measure and analyze the external flowfield around an existing rectangular circulation control wing having a geometric aspect ratio of 0.5 and dual upper/lower surface trailing edge slots adjoining a Coanda surface. A stereoscopic particle image velocimetry system was configured in “dual-mono” mode to image a large interrogation area at the mid-span location; tests were performed in a low-speed wind tunnel at Reynolds numbers of 710,000-1,060,000. The circulation control wing was operated at momentum coefficients between 0 and 0.11 at a zero incidence angle. For pre- and postcirculation stall cases, the effects of a 5% “opposite slot assist” from the lower surface jet were studied. For zero-lift, zero-drag cases, the Reynolds number effect was studied. The results showed how the trailing edge jet augmented the velocity in the wake through upper surface entrainment and shifted the aft separation point onto the lower surface. The measured leeward flow patterns showed velocities from 25-175% of the freestream velocity. At high pre-circulation-stall momentum coefficients, both the wake and the jet were split: one jet was the primary jet effect accelerating around the Coanda surface while the other was entrained momentum from the external upper surface flow. The region of velocity deficit was also split: one between the split jets and one upstream of the stagnation streamline. At momentum coefficients above 0.07, circulation stall shifted the stagnation streamline forward from 95% to 80% of the chord length and resulted in an 8% decrease in lift and a 16% increase in drag. Stall effects were suppressed with the opposite slot assist technique. Even with the non-stalled flow conditions, the flow unsteadiness was high, with variations as high as 50% of the freestream velocity. Operating the wing at a zero-lift, zerodrag condition using the dual slots with equal jet momentum produced a well-defined wake, but the distinction was more due to differences in velocity components than overall velocity magnitude.

P REVIOUS studies on increasing airfoil lift and improving stall characteristics have addressed various passive and active approaches to modifying the leading and trailing edge shapes. The passive approaches have covered such methods as rippling the trailing edge, applying serrated-edge Gurney flaps, or modifying the leading-edge (LE) profile [1,2]. Other efforts have effectively eliminated the dynamic stall of an NACA 0012 airfoil by perturbing the LE contour as little as 0.5–0.9%of the chord [3]. Levshin et al. [4] demonstrated that sinusoidal LE planforms on an NACA 63-021 airfoil section decreased maximum lift, but extended the stall angle by almost 9 deg. The larger amplitude sinusoids created “softer” stall characteristics by maintaining attached flow at the peaks despite separated flow in the troughs. These tests were performed to simulate the effects of LE tubercles on humpback whale (Megaptera novaeangliae) flippers. Prior work by the authors also reported wind tunnel measurements for idealized scale models of humpback whale flippers [5]. One model had a smooth leading edge and a secondmodel had sinusoidal bumps (tubercles) along the leading edge for the outer 2 3 of the span. It was found that the addition of tubercles to a 3-D idealized flipper increased the maximum lift coefficient while reducing the drag coefficient over a portion of the operational envelope. It is thought that the tubercles on the flipper leading-edge enhance the whale’s ability to maneuver to catch prey [6]. Though the work to date regarding sinusoidal or serrated leading-edge planforms is largely motivated by marine mammal locomotion, the effects of extending the stall point for lifting surfaces at similar Reynolds numbers (Re) may have application to small-UAV (unmanned aerial vehicle) design and the inevitable laminar stall problems [7]. However other relevant applications might benefit from the effects of simulated tubercles such as stall alleviation/separation control on sailboat centerboards or wind turbines, where an expanded operating envelope could improve the overall effectiveness of the blade [8,9]. In the present work, a better understanding is sought of the mechanism of the improvements measured in previous experiments, with a greater applicability in mind. The authors seek to determine whether the performance improvements resulted from enhancements to the sectional characteristics of wings with tubercles (i.e., essentially 2-D effects), or from Reynolds number effects on a tapered planform, or from other 3-D effects such as spanwise stall progression.

This paper provides an overview of a multi-year research project that involves the systematic investigation of ship air wakes using an instrumented United States Naval Academy (USNA) YP (Patrol Craft, Training). The objective is to validate and improve Computational Fluid Dynamics (CFD) tools that will be useful in determining ship air wake impact on naval rotary wing vehicles. This project is funded by the Office of Naval Research and includes extensive coordination with Naval Air Systems Command. Currently, ship launch and recovery wind limits and envelopes for helicopters are primarily determined through at-sea in situ flight testing that is expensive and frequently difficult to schedule and complete. The time consuming and potentially risky flight testing is required, in part, because computational tools are not mature enough to adequately predict air flow and wake data in the lee of a ship with a complex superstructure. The top-side configuration of USNA YPs is similar to that of a destroyer or cruiser, and their size (length of 108 ft and above waterline height of 24 ft) allows for collection of air wake data with a Reynolds number that is the same order of magnitude as that of modern naval warships, an important consideration in aerodynamic modeling. A dedicated YP has been modified to add a flight deck and hangar structure to produce an air wake similar to that on a modern destroyer. Three axis acoustic anemometers, fog generators and an inertial measurement unit have been installed. Repeated testing on the modified YP is being conducted in the Chesapeake Bay, which allows for the collection of data over a wide range of wind conditions. Additionally, a 4% scale model of the modified YP has been constructed and tested in the 42×60×120 inch USNA wind tunnel. The project involves USNA midshipmen who are participating in test planning, collecting and analyzing data, and in CFD modeling, providing the midshipmen with valuable professional and research experience. Comparison of YP in situ data with similar data from wind tunnel testing and CFD simulations shows reasonable agreement for a headwind condition and for wind 15° off the starboard bow.

The Cobalt V5.0 CFD solver was applied to the prediction of bluff body wake from a model scale hangar configuration at a Reynolds number of 1.610. Results were compared against 144 scale surface pressure coefficient and surface oil flow wind tunnel data. Baseline CFD predictions applied a monotone integrated large eddy simulation (MILES) turbulence modeling approach. Results from delayed detached eddy simulation (DDES) are also discussed.

This paper provides current results of a multi-year research project that involves the systematic investigation of ship air wakes using an instrumented United States Naval Academy (USNA) YP (Patrol Craft, Training). The objective is to validate and improve Computational Fluid Dynamics (CFD) tools that will be useful in determining ship air wake impact on naval rotary wing vehicles. This project is funded by the Office of Naval Research and includes extensive coordination with Naval Air Systems Command. Currently, ship launch and recovery wind limits and envelopes for helicopters are primarily determined through at-sea in situ flight testing that is expensive and frequently difficult to schedule and complete. The time consuming and potentially risky flight testing is required, in part, because computational tools are not mature enough to adequately predict air flow and wake data in the lee of a ship with a complex superstructure. The top-side configuration of USNA YPs is similar to that of a destroyer or cruiser, and their size (length of 108 ft and above waterline height of 24 ft) allows for collection of air wake data with a Reynolds number that is the same order of magnitude as that of modern naval warships, an important consideration in aerodynamic modeling. A dedicated YP has been modified to add a flight deck and hangar-like structure to produce an air wake similar to that on a modern destroyer. Three-axis acoustic anemometers, fog generators and an inertial measurement unit have been installed. Repeated testing on the modified YP is being conducted in the Chesapeake Bay, which allows for the collection of data over a wide range of wind conditions. Additionally, a 4% scale model of the modified YP has been constructed and tested in the 42×60×120 inch USNA wind tunnel. Comparison of YP in situ data with similar data from wind tunnel testing and CFD simulations shows reasonable agreement for a headwind condition and for a relative wind 15° off the starboard bow. Analysis of in situ data and wind tunnel data for a 30° relative wind also show reasonable agreement, though with a greater deviation than in the 15° relative wind condition. Furthermore, analysis indicates that CFD simulations require modeling the velocity profile in the atmospheric boundary layer to improve simulation accuracy.

As part of a larger program to develop analytic and computational tools to predict the air wake characteristics of naval vessels, an experimental effort has been undertaken to map the air wake of a scaled patrol craft modified with a representative hangar structure and stern flight deck. A 4% scaled model of a United States Naval Academy YP (Patrol Craft, Training) was fabricated for wind tunnel testing at 0, 15, and 30 degrees of yaw at a Reynolds number of 7.6 million. The topside configuration of the vessel simulates that of larger destroyers and cruisers, which service rotorcraft from stern hangars and flight decks. Flow measurements were made in station planes from 0.45 to 5.14 hangar-heights downstream of the aft hangar face using an 18-hole Omniprobe. The 3D velocity measurements indicated that a recirculation zone occurred over the flight deck and extended from 1.5-3 hangar-heights downstream. The effect of yaw angle on the air wake was to consolidate and shift the rotationality of the flow into a large, leeward vortex with a size on the order of the superstructure height. Furthermore, the unsteadiness of the flow was independent of the survey plane position at zero yaw, but increased by 57% at 30 deg yaw, with a large band of high vorticity where the shear layer separated and rolled up. The velocity distributions, flow angles, and vorticity will be used to compare with computational fluid dynamics (CFD) and full-scale, in-situ measurements to portray a complete, Reynoldsscaled picture of the air wake.

The leading-edge tubercles of humpback whale flippers have been shown to enhance hydrodynamic performance by increasing lift and decreasing drag poststall. To explore this effect, computational simulations of two models based on an idealized humpback whale flipper were conducted, one with a smooth leading edge and one with simulated leading-edge tubercles. Two different commercial computational fluid dynamics packages were used, STAR-CCM+ and Solid Works Flow Simulation, and the results were compared with experiment. Numeric lift predictions in the nonstall region were reasonably accurate (maximum error 6.6% between both codes), while lift predictions in the poststall, region were problematic. Numeric drag predictions in the early nonstall region were within experimental error for STAR-CCM+ using the Spalart-Allmaras turbulence model, while both codes exhibited drag prediction error in the stall region. Flow visualizations showed that the smooth flipper exhibited trailing-edge stall, while the simulated tubercle flipper stalled in the troughs, behind the leading notches, first. At high angles of attack, the simulated tubercle flipper still possessed significant regions of attached flow, which contributes to its ability to maintain increased lift poststall.

An analytic and experimental effort was undertaken to assess the effectiveness and efficiency of three winglets mounted chordwise to the tip of a rectangular wing. The winglets, with an aspect ratio of 4.6, were mounted on a half-span wing having an effective aspect ratio of 6.29. 13 configurations of varying dihedral arrangements were analyzed with a vortex lattice method and tested in a low-speed wind tunnel at a Reynolds number of 600,000. While the analytic method provided fair agreement with the experimental results, the predicted trends in lift, drag, and (to a lesser degree) pitching moment were in good agreement. The analytic distributions of wake velocity, circulation, and downwash angle verified that highly nonplanar configurations tended to reduce and diffuse the regions of highest circulation and to create more moderate downwash angles in the wake. This was manifest as an overall drag reduction. More specifically, the results showed that the winglets could be placed in various optimum orientations to increase the lift coefficient as much as 65% at the same angle of attack, decrease the drag coefficient as much as 54% at the same lift coefficient, or improve the maximum L∕D by up to 57%. The most dramatic findings from this study show that positioning the winglet dihedral angles had the result of adjusting the magnitude and slope of the pitching moment coefficient. These observations suggest that multiple winglet dihedral variations may be feasible for use as actively controlled surfaces to improve the performance of aircraft at various flight conditions and to “tune” the longitudinal stability characteristics of the configuration.

An experimental effort was undertaken to assess the effectiveness and efficiency of three winglets mounted chordwise to the tip of a rectangular wing (NACA 0018 section). The winglets, with an aspect ratio of 3.6, were mounted on a half-span wing having an aspect ratio of 3.1. Twenty configurations of varying dihedral arrangements were analyzed with a vortex lattice method and tested in a low-speed wind tunnel at a Reynolds number of 600,000. In general, the arrangements involving high dihedral angles had lower performance increments, due to lower lift and higher interference drag. More specifically, the results showed that the winglets placed at 60, 45, and 30 degrees, respectively, produced nominal 4% higher lift and 46% lower drag. The most dramatic findings from this study show that positioning the winglet dihedral angles had the result of adjusting the point of maximum L/D and the magnitude of the pitching moment coefficient. These observations suggest that multiple winglet dihedral changes affect the lift, drag, and pitching moment in such a way that they are feasible for use as actively-controlled surfaces to improve the performance of aircraft at various flight conditions and to “tune” the longitudinal stability characteristics of the wing.