Showing papers on "Freestream published in 2015"

Control of vortex on a non-slender delta wing by a nanosecond pulse surface dielectric barrier discharge

Abstract: Wind tunnel experiments are conducted for improving the aerodynamic performance of delta wing using a leading-edge pulsed nanosecond dielectric barrier discharge (NS-DBD). The whole effects of pulsed NS-DBD on the aerodynamic performance of the delta wing are studied by balanced force measurements. Pressure measurements and particle image velocimetry (PIV) measurements are conducted to investigate the formation of leading-edge vortices affected by the pulsed NS-DBD, compared to completely stalled flow without actuation. Various pulsed actuation frequencies of the plasma actuator are examined with the freestream velocity up to 50 m/s. Stall has been delayed substantially and significant shifts in the aerodynamic forces can be achieved at the post-stall regions when the actuator works at the optimum reduced frequency of F + = 2. The upper surface pressure measurements show that the largest change of static pressure occurs at the forward part of the wing at the stall region. The time-averaged flow pattern obtained from the PIV measurement shows that flow reattachment is promoted with excitation, and a vortex flow pattern develops. The time-averaged locations of the secondary separation line and the center of the vortical region both move outboard with excitation.

Effect of High Freestream Turbulence on Flowfields of Shaped Film Cooling Holes

Abstract: Shaped film cooling holes have become a standard geometry for protecting gas turbine components. Few studies, however, have reported flowfield measurements for moderately-expanded shaped holes and even fewer have reported on the effects of high freestream turbulence intensity relevant to gas turbine airfoils. This study presents detailed flowfield and adiabatic effectiveness measurements for a shaped hole at freestream turbulence intensities of 0.5% and 13%. Test conditions included blowing ratios of 1.5 and 3 at a density ratio of 1.5. Measured flowfields revealed a counter-rotating vortex pair and high jet penetration into the mainstream at the blowing ratio of 3. Elevated freestream turbulence had a minimal effect on mean velocities and rather acted by increasing turbulence intensity around the coolant jet, resulting in increased lateral spreading of coolant.Copyright © 2015 by ASME

Hummingbird flight stability and control in freestream turbulent winds

Abstract: Airflow conditions close to the Earth's surface are often complex, posing challenges to flight stability and control for volant taxa. Relatively little is known about how well flying animals can contend with complex, adverse air flows, or about the flight control mechanisms used by animals to mitigate wind disturbances. Several recent studies have examined flight in the unsteady von Karman vortex streets that form behind cylinders, generating flow disturbances that are predictable in space and time; these structures are relatively rare in nature, because they occur only the immediate, downstream vicinity of an object. In contrast, freestream turbulence is characterized by rapid, unpredictable flow disturbances across a wide range of spatial and temporal scales, and is nearly ubiquitous in natural habitats. Hummingbirds are ideal organisms for studying the influence of freestream turbulence on flight, as they forage in a variety of aerial conditions and are powerful flyers. We filmed ruby-throated hummingbirds (Archilochus colubris) maintaining position at a feeder in laminar and strongly turbulent (intensity ∼15%) airflow environments within a wind tunnel and compared their mean kinematics of the head, body, tail and wing, as well as variability in these parameters. Hummingbirds exhibited remarkably stable head position and orientation in both smooth and turbulent flow while maintaining position at the feeder. However, the hummingbird's body was less stable in turbulent flow and appeared to be most sensitive to disturbances along the mediolateral axis, displaying large lateral accelerations, translations and rolling motions during flight. The hummingbirds mitigated these disturbances by increasing mean wing stroke amplitude and stroke plane angle, and by varying these parameters asymmetrically between the wings and from one stroke to the next. They also actively varied the orientation and fan angle of the tail, maintaining a larger mean fan angle when flying in turbulent flow; this may improve their passive stability, but probably incurs an energetic cost as a result of increased drag. Overall, we observed many of the same kinematic changes noted previously for hummingbirds flying in a von Karman vortex street, but we also observed kinematic changes associated with high force production, similar to those seen during load-lifting or high-speed flight. These findings suggest that flight may be particularly costly in fully mixed, freestream turbulence, which is the flow condition that hummingbirds are likely to encounter most frequently in natural habitats.

Flowfield Establishment in Hypervelocity Shock-Wave/Boundary-Layer Interactions

Abstract: Shock/boundary-layer interactions generated over double-wedge and double-cone models in high-enthalpy, hypersonic flows are known to be sensitive to the thermochemical state of the gas. In this study, the transient evolution of shock interactions and separated flow is examined in nitrogen and air freestream conditions with stagnation enthalpies ranging from 2 to 8 MJ/kg and Mach numbers from 4 to 7. The time-dependent flowfield and associated time scales required to reach mean values for both viscous and inviscid processes are investigated using fast-response thermocouples and high-speed schlieren and chemiluminescence imaging. In all cases, the oblique/bow-shock triple point is observed to propagate upstream to a mean location as the bow-shock standoff distance increases with time. For all freestream conditions, the triple point reaches a mean position in less time for the conical than the wedge flow. Distinct differences between nitrogen and air both in the evolution and mean flow features are observed...

Film Cooling for Cylindrical and Fan-Shaped Holes Using Pressure-Sensitive Paint Measurement Technique

Abstract: A systematic study was performed to investigate the combined effects of film-hole geometry, blowing ratio, density ratio, and freestream turbulence intensity on a flat-plate film cooling. Detailed film-cooling effectiveness was obtained using a pressure-sensitive-paint technique. Four common geometries were used in this study: simple-angled cylindrical and fan-shaped holes, and compound-angled (β=45 deg) cylindrical and fan-shaped holes. Each plate contained one row with seven holes, and the hole diameter D and hole-length-to-diameter ratio (L/D) are 4 mm and 7.5, respectively. The effects of the blowing ratio M, coolant-to-mainstream-density ratio DR, and freestream turbulence intensity Tu were tested within the ranges of 0.3∼2.0, 1.0∼2.0, and 0.5∼6%, correspondingly. Detailed variations of the laterally averaged effectiveness from low to high blowing ratios were obtained for three density ratios. The results indicated that effectiveness increased as increasing density ratio in general for all geometrie...

High-Speed Boundary-Layer Stability on a Cone with Localized Wall Heating or Cooling

Abstract: A localized heating or cooling effect on stability of the boundary-layer flow on a sharp cone at zero angle of attack and freestream Mach number 6 is analyzed. Experiments were carried out in the Transit-M wind tunnel of the Institute of Theoretical and Applied Mechanics (Novosibirsk, Russia) for different heating/cooling intensities and freestream Reynolds numbers. The mean flows with localized heating/cooling are calculated using axisymmetric Navier–Stokes equations. These solutions are used for the spatial linear stability analysis to estimate the transition onset points using the eN method. Direct numerical simulations of two-dimensional disturbances propagating in the boundary layer through the cooled/heated region are performed. The experiment and computations showed similar qualitative trends. The localized cooling decreases the second-mode amplitude and delays transition. The heating produced an opposite effect, which is less pronounced.

HIFiRE-1 Ascent-Phase Boundary-Layer Transition

Abstract: The Hypersonic International Flight Research Experimentation (HIFiRE) program is a hypersonic flight test program. The primary experiment for flight one, launched in March 2010, was to measure boundary-layer transition in hypersonic flight on a nonablating, 7 deg half-angle axisymmetric cone with a small bluntness of 2.5 mm radius. The flight gathered pressure, temperature, and heat transfer measurements during ascent and reentry. Although the vehicle reentered the atmosphere at a higher-than-intended angle of attack, the ascent portion of the flight provided smooth-body boundary-layer transition data at freestream Mach numbers greater than 5, where transition was presumed to be dominated by second-mode instability. The angle of attack during this portion of the flight was less than 1 deg. The end of turbulent-to-laminar transition occurred at Reynolds numbers between 10.3×106 and 12.2×106, based on x-location and freestream conditions. Transition was correlated with second-mode N-factors of approximately...

Direct numerical simulation of transition to turbulence in a supersonic boundary layer

Abstract: Based on full unsteady compressible Navier–Stokes equations a direct numerical simulation of the linear and nonlinear stages of the laminar-turbulent transition in boundary layer of a flate plate at the freestream Mach number M = 2 is carried out.

Control of shock-wave boundary layer interaction using steady micro-jets

Abstract: An experimental investigation was conducted to control the amplitude of shock unsteadiness associated with the interaction induced by a cylindrical protuberance on a flat plate in a Mach 2.18 flow. The control was applied in the form of an array of steady micro air-jets of different configurations with variation in pitch $$(\beta )$$ and skew angle $$(\alpha )$$ of the jets. The effect of air-jet supply pressure on control was also studied. Each of the micro-jet configurations was placed 20 boundary layer thicknesses upstream of the leading edge of the cylinder. The overall interaction is seen to get modified for all control configurations and shows a reduction in both separation- and bow-shock strengths and in triple-point height. A significant reduction in the peak rms value is also observed in the intermittent region of separation for each case. For $$90^{\circ }$$ pitched jets placed in a zig-zag configuration, good control effectiveness is achieved at control pressures similar to the stagnation pressure of the freestream. At higher control pressures, however, their obstruction component increases and if these jets are not spaced sufficiently far apart, the effectiveness of their control begins to drop due to the beginning of spanwise jet-to-jet interaction. On the other hand, pitching or skewing the jets to $$45^{\circ }$$ reduces the obstruction component considerably which at lower control pressures shows lower effectiveness. But at higher control pressure, the effectiveness of these configurations continues to increase unlike the $$90^{\circ }$$ pitched jets.

Effects of nonuniform incident velocity on a dynamic wind turbine airfoil

Abstract: For further insight into the performance of a horizontal axis wind turbine blade section under yaw loads, a 2D numerical simulation of a pitching S809 airfoil under dynamic stall with an unsteady incident velocity is presented. The streamwise incident velocity and pitch angle of incidence oscillated with the same frequency but with a range of phase differences, − π ≤ Φ ≤ π. Changing Φ caused variation of the results, which can be highlighted as significantly augmented and dramatically damped dynamic stall loads, both increasing and decreasing trends for vortex growth time during Φ increase, a shifted location of the maximum loads and a change in the order of the vortex pair circulation in each cycle. The results showed strong dependency on the velocity and acceleration of the freestream during dynamic stall, which categorized the results in four individual subdomains with different behaviors. Copyright © 2014 John Wiley & Sons, Ltd.

Uncertainty Analysis of Mars Entry Flows over a Hypersonic Inflatable Aerodynamic Decelerator

Abstract: A detailed uncertainty analysis for high-fidelity flowfield simulations over a fixed aeroshell of hypersonic inflatable aerodynamic decelerator scale for Mars entry is presented for fully laminar and turbulent flows at peak stagnation-point heating conditions. This study implements a sparse-collocation approach based on stochastic expansions for efficient and accurate uncertainty quantification under a large number of uncertainty sources in the computational model. The convective and radiative heating and shear stress uncertainties are computed over the hypersonic inflatable aerodynamic decelerator surface and are shown to vary due to a small fraction of 65 flowfield and radiation modeling parameters considered in the uncertainty analysis. The main contributors to the convective heating uncertainty near the stagnation point are the CO2–CO2, CO2–O, and CO–O binary collision interactions, freestream density, and freestream velocity for both boundary-layer flows. In laminar flow, exothermic recombination rea...

Width Effects in Transonic Flow over a Rectangular Cavity

Abstract: 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.

Aerodynamic Heating of a Hypersonic Projectile with Forward-Facing Ellipsoid Cavity at Nose

Abstract: Numerical experiments are carried out using a commercially available computational-fluid-dynamics code to investigate the aerodynamic heating of a hemisphere–cylinder with a forward-facing ellipsoid cavity at nose in atmosphere. A wide range ellipsoid cavity of depths varying between 3 to 30 mm placed axisymmetrically at the nose of the blunt body with base diameter 40 mm and overall length 70 mm have been investigated. The ratios of semimajor axis to semiminor axis of these cavities (a/b) are 3, 2, or 1.5, where semimajor axis a is equal to the depth of cavity d and the semiminor axis b is close to the lip radius r of the cavity. All computations have been done at a freestream Mach number of 10.1 and static pressure and temperature of 16,066 Pa and 216.65 K respectively. The steady-state solutions of axisymmetric Navier–Stokes equations are obtained using a time marching approach in order to study the aerodynamic heating of a hypersonic projectile with a forward-facing ellipsoid cavity at a nose. The sim...

Integrating Mach–Zehnder interferometry with TPIV to measure the time-resolved deformation of a compliant wall along with the 3D velocity field in a turbulent channel flow

Abstract: A system combining tomographic PIV (TPIV) and Mach–Zehnder interferometry (MZI) simultaneously measures the time-resolved 3D flow field and 2D distribution of wall-normal deformation in a turbulent channel flow over a transparent compliant surface. This paper focuses on the experimental techniques and data analysis procedures, but includes sample results. Standard TPIV analysis resolves the log layer of the mean velocity and the linear decrease in total shear stress with distance from the wall. Single-pixel ensemble correlations reveal the buffer layer and top of the viscous sublayer. Analysis of the MZI data consists of two steps, namely critical spatial filtering of interferograms to remove noise and phase demodulation to calculate the surface shape. A new technique to improve the filtration of noise from interferograms based on spatial correlations of small windows is introduced and optimized. Taking advantage of this enhancement, the phase/deformation distribution is calculated directly from arccosines of the intensity, which avoids edge artifacts affecting spectral calculations. Validations using synthetic noisy interferograms indicate that errors associated with correlation-based enhancement are consistently lower and much less sensitive to fringe shape than spectral band-pass filtering. The experimental wavenumber–frequency spectra show that the deformation consists of patterns that are larger than the field of view, surface waves and small-scale patterns. Some of the latter are advected at the freestream velocity, but mostly at 70 % of the freestream, the mean speed at 10 % of the channel half height. Indeed, spatial correlations of the deformation with velocity components peak at this elevation.

Supersonic Flow over a Shallow Open Rectangular Cavity

Abstract: An experimental investigation was conducted on a supersonic flow at a freestream Mach number of 2 over a shallow open cavity, including the effects of adding streamwise serrated edges. These flows have relevance to weapons bays and airframe gaps on high-speed aircraft. The measurements consisted of single-shot and time-resolved schlieren visualization, as well as unsteady pressure spectra. The length-to-depth ratio of the cavity was 8. The tests conducted at different Reynolds numbers with the baseline cavity (straight leading and trailing edges) showed that increasing the Reynolds number increases the root-mean-square pressure inside the cavity. The addition of serrations to the cavity leading or trailing edge did not show any significant effect on the separating shear layer nor in controlling the oscillations of the shear layer. There was also no noticeable effect on the overall sound pressure levels inside the cavity. A new expression for calculating shallow-cavity resonant frequencies applicable at su...

Experimental Study on Icing and Anti-Icing Characteristics of Engine Inlet Guide Vanes

Abstract: Icing and anti-icing characteristics of engine inlet guide vanes are studied in this paper. Experiments are carried out in an icing wind tunnel under different icing conditions. The tests are performed at a freestream velocity of 85 m/s, an inlet angle of attack of 0 deg, and a freestream static temperature in the range of −8∼−20 °C. The icing environment parameters include a mean volume diameter of 20 μm, liquid water content in the range of 0.5 to 2.0 g/m3. The experiment is performed on full-scale inlet guide vanes. The temperature distributions on the guide vane surface are measured by thermocouples. The leading-edge ice shapes of inlet guide vanes are compared under different icing conditions. The effects of freestream static temperature, liquid water content, and the mass flow rate of anti-icing hot air on the performance of the anti-icing system are presented and analyzed.

Application of FLEET Velocimetry in the NASA Langley 0.3-Meter Transonic Cryogenic Tunnel

Abstract: Femtosecond laser electronic excitation and tagging (FLEET) velocimetry is demonstrated in a large-scale transonic cryogenic wind tunnel. Test conditions include total pressures, total temperatures, and Mach numbers ranging from 15 to 58 psia, 200 to 295 K, and 0.2 to 0.75, respectively. Freestream velocity measurements exhibit accuracies within 1 percent and precisions better than 1 m/s. The measured velocities adhere closely to isentropic flow theory over the domain of temperatures and pressures that were tested. Additional velocity measurements are made within the tunnel boundary layer; virtual trajectories traced out by the FLEET signal are indicative of the characteristic turbulent behavior in this region of the flow, where the unsteadiness increases demonstrably as the wall is approached. Mean velocities taken within the boundary layer are in agreement with theoretical velocity profiles, though the fluctuating velocities exhibit a greater deviation from theoretical predictions.

Experimental investigation of turbulent boundary layers over transversal moving surfaces

Abstract: The influence of a spanwise traveling transversal surface wave on the near-wall flow field of turbulent boundary layers is investigated by particle-image velocimetry (PIV) and micro-particle tracking velocimetry (μ-PTV). The experimental setup consists of a flat plate equipped with an insert to generate a transversal spanwise traveling wave of an aluminum surface. PIV and μ-PTV measurements are conducted for three Reynolds numbers based on the freestream velocity and momentum thickness immediately downstream of the actuated surface Re θ = 1200, 1660, and 2080. The transversal traveling wave is generated by a newly developed electromagnetic actuator system underneath the aluminum surface. Three amplitudes A = 0.25, 0.30, and 0.375 mm at a wave length of $$\lambda \, = \,160\,{\text{mm}}$$ and a frequency of f = 81 Hz are investigated. The detailed analysis of the velocity profile shows the transversal traveling surface motion to redistribute the velocity in the viscous sublayer and in the logarithmic region of the turbulent boundary layer. The streamwise and wall-normal velocity fluctuations in the outer boundary layer are increased and the streamwise momentum in the near-wall regime is lowered. The drag reduction ratio (DR) due to the actuation is determined by the velocity gradient in the viscous sublayer. At the lowest Reynolds number the drag-reducing impact is proportional to the amplitude of the wave. That is, the higher the amplitude, the more pronounced the drag reduction resulting in a friction drag reduction up to 3.4 % compared to the non-actuated configuration.

Cavitation about a jet in crossflow

Abstract: Cavitation occurrence about a jet in crossflow is investigated experimentally in a variable-pressure water tunnel using still and high-speed photography. The 0.012 m diameter jet is injected on the centreplane of a 0.6 m square test section at jet to freestream velocity ratios ranging from 0.2 to 1.6, corresponding to jet-velocity-based Reynolds numbers of 25×103 to 160×103 respectively. Measurements were made at a fixed freestream-based Reynolds number, for which the ratio of the undisturbed boundary layer thickness to jet diameter is 1.18. The cavitation number was varied from inception (up to about 10) down to 0.1. Inception is investigated acoustically for bounding cases of high and low susceptibility to phase change. The influence of velocity ratio and cavitation number on cavity topology and geometry are quantified from the photography. High-speed photographic recordings made at 6 kHz provide insight into cavity dynamics, and derived time series of spatially averaged pixel intensities enable frequency analysis of coherent phenomena. Cavitation inception was found to occur in the high-shear regions either side of the exiting jet and to be of an intermittent nature, increasing in occurrence and duration from 0 to 100 % probability with decreasing cavitation number or increasing jet to freestream velocity ratio. The frequency and duration of individual events strongly depends on the cavitation nuclei supply within the approaching boundary layer. Macroscopic cavitation develops downstream of the jet with reduction of the cavitation number beyond inception, the length of which has a power-law dependence on the cavitation number and a linear dependence on the jet to freestream velocity ratio. The cavity closure develops a re-entrant jet with increase in length forming a standing wave within the cavity. For sufficiently low cavitation numbers the projection of the re-entrant jet fluid no longer reaches the cavity leading edge, analogous to supercavitation forming about solid cavitators. Hairpin-shaped vortices are coherently shed from the cavity closure via mechanisms of shear-layer roll-up similar to those shed from protuberances and jets in crossflow in single-phase flows. These vortices are shed at an apparently constant frequency, independent of the jet to freestream velocity ratio but decreasing in frequency with reducing cavitation number and cavity volume growth. Highly coherent cavitating vortices form along the leading part of the cavity due to instability of the jet upstream shear layer for jet to freestream velocity ratios greater than about 0.8. These vortices are cancelled and condense as they approach the trailing edge in the shear layer of opposing vorticity associated with the cavity closure and the hairpin vortex formation. For lower velocity ratios, where there is decreased jet penetration, the jet upstream shear velocity gradient reverses and vortices of the opposite sense form, randomly modulated by boundary layer turbulence.

Simulating Turbulence and Mixing in Supersonic Combustors Using Hybrid RANS/LES

Abstract: Simulation results are presented for non-reacting flow within a supersonic cavity flameholder. The freestream is air at Mach 2. A case is simulated with no fuel injection and with ethylene fuel injected through holes located on the back face of the cavity. The simulations correspond to a series of experiments for which particle image velocimetry measurements of two velocity components were made within the cavity. The flow within the cavity is computed using unsteady hybrid Reynolds-averaged Navier-Stokes/large-eddy simulation. A thorough grid resolution study is presented in which the resolution required to resolve the flow within the cavity is determined. The influence of the level of turbulence within the oncoming boundary layer is examined and found to significantly affect the velocity field and mixing within the cavity. The effect of the side walls is investigated by comparing simulations of the full-width duct to simulations of a partial-width duct that uses periodic boundary conditions. Differences in resolved turbulence kinetic energy and mixing are seen between the full-width and partial width simulations. The results of the simulations are also compared to the velocity measurements from the experiments, and the hybrid Reynolds-averaged Navier-Stokes/large-eddy simulation results are found to compare reasonably well with the experiment in most locations. Improved agreement with fluctuating velocity components is found when the simulation results are filtered to a resolution corresponding to the resolution of the velocity measurement technique.

Freestream Effects on Boundary Layer Disturbances for HIFiRE-5

Abstract: A 2:1 elliptic cone model was tested in a variable-Mach-number conventional hypersonic wind tunnel. Freestream Reynolds number, Mach number, model streamwise location, and model wall temperature were all varied to ascertain the effect of each on measured disturbances. A low-frequency disturbance was observed at Mach 5.8. It experienced some growth in excess of the increasing freestream pressure and noise for a narrow Reynolds number range. Disturbance properties were similar to what was measured in another conventional hypersonic wind tunnel. For Mach 6.5 and 7, there was no evidence of traveling crossflow waves. However, higher-frequency disturbances were observed. These disturbances were nearly two-dimensional and had phase speeds near the expected edge velocity. It is possible that these disturbances were second mode waves. None of the measured disturbances corresponded to any feature of the freestream Pitot spectra.

The influence of cornering on the vortical wake structures of an inverted wing

Abstract: The aerodynamic performance of inverted wings on racing-car configurations is most critical when cornering; however, current wind tunnel techniques are generally limited to the straight-line condition. The true cornering condition introduces complexity because of the curvature of the freestream flow. This results in an increase in the tangential velocity with increasing distance from the instantaneous center of rotation and causes the front wing to be placed at a yaw angle. Numerical simulations were used to consider an 80% scale front wing when steady-state cornering with radii ranging from 60m to 7.5m, and yaw angles ranging from 1.25° to 10°. The changes to the pressure distribution near the endplates caused the wake structure to become highly asymmetric. Both the primary longitudinal vortices and the secondary longitudinal vortices differed in strength, and the vortex core positions shifted in the vertical direction and the spanwise direction. The change in the position became more substantial further downstream as the structures tended toward the freestream direction. The effects on the wing surface pressure distribution resulted in the introduction of yawing and rolling moments, as well as a side force and an increase in drag. The results demonstrate the importance of evaluating the cornering condition if that is where a good performance is most sought after.