Showing papers on "Freestream published in 2001"

Turbine Nozzle Endwall Film Cooling Study Using Pressure-Sensitive Paint

Abstract: Endwall surface film cooling effectiveness was measured on a turbine vane endwall surface using the pressure sensitive paint (PSP) technique. A double staggered row of holes and a single row of discrete slots were used to supply film cooling in front of the nozzle cascade leading edges. Nitrogen gas was used to simulate film cooling flow as well as a tracer gas to indicate oxygen concentration such that film effectiveness by the mass transfer analogy could be obtained. Cooling mass flow was controlled to be 0.5 to 3.0% of the mainstream mass flow. The freestream Reynolds number was about 283000 and Mach number was about 0.11. The freestream turbulence intensity was kept at 6.0% for all the tests, measured by a thermal anemometer. The PSP was calibrated at various temperatures and pressures to obtain better accuracy before being applied to the endwall surface. Film effectiveness distributions were measured on a flat endwall surface for five different mass flow rates. The film effectiveness increased nonlinearly with mass flow rate, indicating a strong interference between the cooling jets and the endwall secondary flows. At lower mass flow ratios, the secondary flow dominated the near wall flow field, resulting in a low film effectiveness. At higher mass flow ratios, the cooling jet momentum dominated the near wall flow field, resulting in a higher film effectiveness. The comparison between hole injection and slot injection was also made.Copyright © 2001 by ASME

Modeling of Laminar-Turbulent Transition for High Freestream Turbulence

Abstract: To simulate transitional skin friction or heat transfer, the conditionally averaged Navier-Stokes equations are used. To describe the diffusion of freestream turbulence into the boundary layer and the intermittent laminar-turbulent flow behavior during transition, a turbulence weighting factor τ is used. A transport equation is presented for this τ-factor including convection, diffusion, production, and sink terms, In combination with the conditioned Navier-Stokes equations, this leads to an accurate calculation of flow characteristics within the transitional layer. The method is validated on transitional skin friction and heat transfer measurements, respectively on a flat plate and in a linear turbine cascade

Characteristics of the compressible shear layer over a cavity

Abstract: The compressible shear layer over a 3:1 rectangular cavity was investigated using schlieren photography and planer laser imaging at freestream Mach numbers of 1.8, 2.1, 2.8, and 3.5. The purpose of this investigation was to study the characteristics of the shear layer structures and their convective velocity over this Mach number range. Schlieren images show leading- and trailing-edge shock waves, as well as shock waves emanating from the shear layer, which became less prevalent as the Mach number increased. Streamwise planar laser sheet lighting images indicate the existence of organized roller-type structures at the lower Mach numbers studied (M =2.1 and 2.8). These structures became less coherent as the Mach number was increased, although the cavity appears to cause the large-scale structures to persist at higher levels of compressibility than found in planar free shear layers. Plan view images indicate that the two dimensionality of the large-scale structures decreased with increasing Mach number. Autocorrelations performed on single-pulse images show that the structure size decreased 63% when the freestream Mach number was increased from 1.8 to 3.5. By double pulsing the laser at delays of 15, 20, and 25 πs, the evolution of the large-scale structures were investigated and quantie ed. The correlations were found to decrease by 23% for the same nondimensional time when increasing the Mach number from 2.1 to 3.5. From the shift of the peak correlation, the convective velocity was calculated for each case and found to vary laterally across the shear layer. At a lateral location most representative of the large-scale structures in the shear layer, the variation in convective velocity with Mach number was best represented by 0.57 times the freestream velocity.

Detailed Boundary Layer Measurements on a Turbine Stator Vane at Elevated Freestream Turbulence Levels

Abstract: High freestream turbulence levels have been shown to greatly augment the heat transfer on a gas turbine airfoil. To better understand these effects, this study has examined the effects elevated freestream turbulence levels have on the boundary layer development along a stator vane airfoil. Low freestream turbulence measurements (0.6 percent) were performed as a baseline for comparison to measurements at combustor simulated turbulence levels (19.5 percent). A two-component LDV system was used for detailed boundary layer measurements of both the mean and fluctuating velocities on the pressure and suction surfaces. Although the mean velocity profiles appeared to be more consistent with laminar profiles, large velocity fluctuations were measured in the boundary layer along the pressure side at the high freestream turbulence conditions. Along the suction side, transition occurred further upstream due to freestream turbulence.

Temperature- and Pressure-Sensitive Paint Measurements in Short-Duration Hypersonic Flow

Abstract: Surface temperatures and pressures were measured on an elliptic cone lifting body in a hypersonic e owe eld using thin-e lm (» 5πm) temperature- and pressure-sensitive paints (TSPs and PSPs ). The tests were conducted in the 48-inch hypersonic shock tunnel (48-inch HST) at Calspan‐University of Buffalo Research Center and were part of a more comprehensive experimental study examining the three-dimensional characteristics of laminar, transitional, and turbulent e ow over the model. Measurement opportunity in the 48-inch HST was limited by the short duration of steady freestream conditions of the driven gas; image acquisition times were » 3 ms. Images of the coatings applied to the broad side of the symmetric elliptic cone were calibrated with in situ static pressure and surface-e lm temperature measurements. The TSP results illustrate the higher heat transfer rates and change in boundary-layer transition over the model surface caused by the nose geometry, and the PSP results show a mild pressure gradient over the interrogated surface region. Submillisecond TSP acquisition using a high-speed imager demonstrated the feasibility of measuring the surface temperature rise.

Experimental investigation of transition to turbulence as affected by passing wakes

Abstract: Experimental results from a study of the effects of passing wakes upon laminar-to-turbulent transition in a low-pressure turbine passage are presented. The test section geometry is designed to simulate the effects of unsteady wakes resulting from rotor-stator interaction upon laminar-to-turbulent transition in turbine blade boundary layers and separated flow regions over suction surfaces. Single-wire, thermal anemometry techniques were used to measure time-resolved and phase-averaged, wall-normal profiles of velocity, turbulence intensity, and intermittency at multiple streamwise locations over the turbine airfoil suction surface. These data are compared to steady state, wake-free data collected in the same geometry to identify the effects of wakes upon laminar-to-turbulent transition. Results are presented for flows with a Reynolds number based on suction surface length and exit velocity of 50,000 and an approach flow turbulence intensity of 2.5 percent. From these data, the effects of passing wakes and associated increased turbulence levels and varying pressure gradients on transition and separation in the near-wall flow are presented. The results show that the wakes affect transition both by virtue of their difference in turbulence level from that of the free-stream but also by virtue of their velocity deficit relative to the freestream velocity, and the concomitant change in angle of attack and temporal pressure gradients. The results of this study seem to support the theory that bypass transition is a response of the near-wall viscous layer to pressure fluctuations imposed upon it from the free-stream flow. The data also show a significant lag between when the wake is present over the surface and when transition begins. The accompanying CD-ROM includes tabulated data, animations, higher resolution plots, and an electronic copy of this report.

Fluorescence velocimetry of the hypersonic, separated flow over a cone

Abstract: Planar laser-induced fluorescence of nitric oxide is used to measure a component of the velocity field for the Mach 7 flow around a 30-deg half-angle, 50-mm-diam cone mounted to a long, 38-mm-diam shaft, or sting. Transverse velocities are measured in the freestream, the shock layer, and the separated region at the junction between the cone and the sting. For most of the flowfield, the uncertainty of the measurements is between ±50 and ±100 m/s for velocities ranging from -300 to 1300 m/s, corresponding to a minimum uncertainty of ±5%. The measurements are compared with the commercial computational fluid dynamics (CFD) code CFD-FASTRAN . The agreement between the theoretical model and the experiment is reasonably good. CFD accurately predicts the size and shape of the shock layer and separated region behind the cone as well as the magnitude of the gas velocity near the reattachment shock. However, the magnitude of the velocity in the shock layer and gas expansion differ somewhat from that predicted by CFD. The discrepancies are attributed to a small systematic error associated with laser-beam attenuation and also to inexact modeling of the flowfield by CFD

Shallow wakes generated on smooth and rough surfaces

Abstract: This study deals with the behavior of shallow turbulent wakes generated on smooth and rough surfaces. The wake generator used is a flat plate placed normal to the flow. Experiments were conducted at flow depths of 40 and 80 mm. The boundary layer thickness in the approaching flow occupies 60–75% of the flow depth. The Reynolds number based on the plate width and approaching freestream velocity varies from 13.0 × 103 to 14.5 × 103. Velocity measurements were carried out in the near-wake region (1–10 plate widths) using a laser-Doppler anemometer. The mean velocity distributions at various axial stations collapse onto a single curve by a proper choice of the length and velocity scales. It is important to note that a sense of self-similarity is attained even in the near-wake region. Attempts were made to clarify the relative effects of the transverse shear and bed friction in shallow open channel wakes.

Mole-Fraction-Sensitive Imaging of Hypermixing Shear Layers

Abstract: A theoretical model that determines the optimum excitation frequency for obtaining a e uorescence signal with a strong dependence on fuel mole fraction is presented for supersonic fuel ‐air compressible mixing studies. The challenge associated with this is to maintain a high sensitivity to fuel mole fraction with minimal sensitivity to temperature and pressure in a e ow with large temperature variations and pressure gradients. The results of the modelareappliedtothemixingregion behindvariousscramjetfuelinjectorsinashocktunnelto measurefuelmole fraction. Hydrogen fuel at a Mach numberof 1.7 is injected into a mostly N 2 freestream at Mach 4.8. Experimental e uorescence images are presented in streamwise and spanwise planes.

Pressure Surface Separations in Low Pressure Turbines: Part 1 of 2 — Midspan Behaviour

Abstract: This paper describes an investigation into the behaviour of the pressure surface separation at midspan in a linear cascade. It is f ound that the pressure surface separation can be a significant contributor to the profile loss of a thin, solid, low pressure turbine blade that is typical of current engine designs. Numerical predictions are first used to study the inviscid behaviour of the blade. These show a strong incidence dependence around the leading edge of the profile. Experiments then show clearly that all characteristics of the pressure surface separation are controlled primarily by the incidence. It is also shown that the effects of wake passing, freestream turbulence and Reynolds number are of secondary importance. A simple two-part model of the pressure surface flow is then proposed. This model suggests that the pressure surf ace separation is highly dissipative through the action of its strong turbulent shear. As the incidence is reduced, the increasing blockage of the pressure surface separation then raises the velocity in the separated shear layer to levels at which the separation can create significant loss. NOMENCLATURE Cd dissipation coefficient

Design of viscoelastic coatings to reduce turbulent friction drag

Abstract: A method is provided to select appropriate material properties for turbulent friction drag reduction, given a specific body (1), configuration and freestream velocity (2). The method is based on a mathematical description of the balance of energy at the interface between the viscoelastic surface and the moving fluid, and permits determination of the interaction of turbulent boundary layer fluctuations with a viscoelastic layer by solving two subtasks -- i.e., a hydrodynamic problem and an elasticity problem, which are coupled by absorption and compliancy coefficients. Displacement, velocity, and energy transfer boundary conditions on a viscoelastic surface are determined, and a Reynolds stress type turbulence model is modified to account for redistribution of turbulent energy in the near-wall of the boundary layer. The invention permits drag reduction by a coating with specified density, thickness, and complex shear modulus to be predicted for a given body geometry and freestream velocity. For practical applications, viscoelastic coatings may be combined with additional structure, including underlying wedges to minimize edge effects for coatings of finite length, and surface riblets, for stabilization of longitudinal vortices.

Practical Experience With the Discrete Green’s Function Approach to Convective Heat Transfer

Abstract: The heat transfer from a short uniform heat flux strip beneath a turbulent boundary layer with and without freestream turbulence was measured using a liquid crystal imaging technique. Data were taken for freestream turbulence intensities on the order of 12% and at momentum thickness Reynolds numbers on the order of 1,000 and 2,000 for the turbulent and steady freestreams respectively. Heat transfer enhancement due to the presence of freestream turbulence was measured in terms of the ratio of the average St on the heated strip for the turbulent freestream case to the steady freestream case. Compared to the baseline case of a uniformly heated surface upstream of the strip, the heat transfer enhancement decreased by 20%. The temperature distribution measured on and around the heated strip represented a kernel solution that was used by means of superposition to predict the heat transfer for any arbitrarily specified thermal boundary condition given the same flowfield. Predictions are compared against correlations and numerical predictions as well as data from the literature. The details and practical applications of this approach to handling heat transfer with non-uniform thermal boundary conditions are presented.

Shock Wave - Film Cooling Interactions in Transonic Flows

Abstract: Interactions between shock waves and film cooling are described as they affect magnitudes of local and spanwise-averaged adiabatic film cooling effectiveness distributions. A row of three cylindrical holes is employed. Spanwise spacing of holes is 4 diameters, and inclination angle is 30 degrees. Freestream Mach numbers of 0.8 and 1.10–1.12 are used, with coolant to freestream density ratios of 1.5–1.6. Shadowgraph images show different shock structures as the blowing ratio is changed, and as the condition employed for injection of film into the cooling holes is altered. Investigated are film plenum conditions, as well as perpendicular film injection cross-flow Mach numbers of 0.15, 0.3, and 0.6. Dramatic changes to local and spanwise-averaged adiabatic film effectiveness distributions are then observed as different shock wave structures develop in the immediate vicinity of the film-cooling holes. Variations are especially evident as the data obtained with a supersonic Mach number are compared to the data obtained with a freestream Mach number of 0.8. Local and spanwise-averaged effectiveness magnitudes are generally higher when shock waves are present when a film plenum condition (with zero cross-flow Mach number) is utilized. Effectiveness values measured with a supersonic approaching freestream and shock waves then decrease as the injection cross-flow Mach number increases. Such changes are due to altered flow separation regions in film holes, different injection velocity distributions at hole exits, and alterations of static pressures at film hole exits produced by different types of shock wave events.Copyright © 2001 by ASME

Mixing Enhancement in Compressible Base Flows via Generation of Streamwise Vorticity

Abstract: A C-type boundary-layer grid around a NACA-0012 airfoil has been generated up to a normal distance corresponding to 1=Ch D 0:06. The values of N and s are 20 and i4, respectively. This boundary-layermesh is coupled and with unstructuredand algebraicmeshes as shown in Fig. 5. Figure 6a shows the streamlines in laminar  ow with the conditions for freestream Mach number M1 D 0:5, Reynolds number based on chord RechD 5£ 10 and incidence ®D 0 deg, same as those used by Liu.4 The present results, including a small separation zone near the trailing edge, are in good agreementwith Liu.4 Figure 6b shows the isentropicMachnumbercontoursfor the turbulentcompressible owatM1 D 0:754, Rech D 3:76£ 106 , and ®D 3:02 deg. With the help of hyperbolic grid, the Baldwin–Lomax turbulencemodel has been implemented without any complex changes, such as those suggested by Turner and Jennions. The standard k–" model is used in the unstructured and algebraic grid zones. This result is in good agreement with the predictions of Kallinderis.

Fluorescence Visualization of Hypersonic Flow Establishment over a Blunt Fin

Abstract: Fluorescence imaging is used to investigate the separated flow upstream of a blunt fin in a hypersonic freestream with a transitional boundary layer. Images are presented to show the flow development before, during, and after the test time of the free-piston shock tunnel used to generate the flow. These images indicate that the test time in this facility is long enough to achieve a steady flow over the blunt fin. Thermocouple measurements are included to compare the surface heat flux upstream of the fin with that for flow along a flat plate with the same freestream conditions. The heat flux results are consistent with separation in a transitional boundary layer and show that the separated flow is oscillatory

Experimental study of the mixing transition in a gaseous axisymmetric jet

Abstract: Instantaneous, quantitative, planar images of molecularly mixed-jet fluid fraction were obtained for the purpose of studying the mixing transition in a gaseous axisymmetric jet from ReD=16 200–29 200. By using a simultaneous nitric oxide and acetone planar laser-induced fluorescence technique, the mixing transition was detected from sudden changes in the molecularly mixed-jet fluid volume fraction, the growth rate of the shear layer, the preferred mixed-jet fluid fraction, and the character of axial/radial probability density functions. The mixing transition for all Reynolds numbers in this regime was found to begin after the first vortex pairing near Rx/λ=6 and was completed by the second vortex pairing near Rx/λ=12, where R=(1−r)/(1+r), r is the low- to high-speed freestream velocity ratio, and λ is the natural instability wavelength. The statistical quantities at all Reynolds numbers were found to collapse when scaled with Rx/λ, with the exception of the mixing layer width. The latter collapsed for all...

Oxygen Atoms' Effect on Vibrational Relaxation of Nitrogen in Blunt-Body Flows

Abstract: Numerical simulations are presented of the steady-state aire ow over a hemisphere cylinder of 1-m radius having hypersonic Mach numbers, where vibrational relaxation is the dominant mechanism and the dissociation of oxygen is small. A Mach 6.5 e ow was analyzed at freestream pressure of 50 Pa with a nonequilibrium freestream translational temperature of 300 K and vibrational temperature of 4000 K; a Mach 1.5 e ow was also studied to delineate effects of vibration ‐translation (V‐T)energy losses due to N 2‐O collisions. The effects on the vibrational population distribution, temperature, and pressure in the e owe eld were studied for various media: pure nitrogen and air mixtures of 0.0001, 0.1, and 1% oxygen atoms. Code validation was performed with previously reported computational results and experimental data for equilibrium e ow in freestream, but nonequilibrium in the shock layer. An upwind difference numerical scheme was used to solve the inviscid Euler equations coupled to a vibrational kineticsmodel ofN 2,assumedasananharmonicoscillatorof40quantum levels. Theshock-standoffdistance comparison with experimental data for a Mach 7.7 and 8.6 aire ow past a blunt body showed good agreement. For the Mach 1.5 e ow at nonequilibrium freestream conditions, the high efe ciency of the V ‐T rates of N2‐O collisions introduces additional heating in the shock layer for 0.1% and higher atomic oxygen, thus increasing the shockstandoff distance; for the Mach 6.5 e ow, a 0.1% atomic oxygen in air decreases the translational temperature in air compared to that of pure nitrogen in the stagnation region.

Effect of Noise on Roughness-Induced Boundary-Layer Transition for Scramjet Inlet

Abstract: Boundary-layer transition was studied on an axisymmetric scramjet-inlet model with compression corners, using thePurdue Mach 4 quiet-e owLudwieg tube. Although theReynoldsnumberwasnotlargeenough to achieve naturaltransitionunderquiet-e owconditionsonthesmoothmodel,transitionwasinvestigatedbyapplyingisolated roughness elements near the nose of the model. Surface hot-e lm measurements were obtained under both quiet and noisy wind-tunnel conditions. Both the noisy tunnel e ow and the roughness elements generated intermittency in the boundary layer. The effect of the tunnel noise was comparable to the effect of the roughness. The change from quiet to noisy e ow also changes the character of the intermittency and the effects of the roughness, which suggests that ground tests for boundary-layer transition should be interpreted with care. These results form the e rst substantial quiet-e ow measurements availablein the open literaturefor supersonic boundary-layer transition over compression corners. Nomenclature I.t/ = smoothed value of i.t/, intermittency is average value i.t/ = trace of provisional turbulence/nonturbulence using " as threshold for the identie cation Re = Reynolds number Res = Reynolds number based on freestream conditions and arclength from the leading edge " = local difference between two points (proportional to the local e rst derivative )

Generation of the karman vortex street at low reynolds number due to cooling a cylinder: cause and fluid type effect by numerical computation

Abstract: The aim is to clarify a main cause and a fluid-type effect of the generation of the Karman vortex street due to cooling a cylinder at a low Reynolds number where the isothermal wake is not a Karman vortex street but the wavy wake. The two-dimensional, laminar, time-dependent continuity equation; Navier–Stokes equations with the buoyancy term; and energy equation are solved numerically by finite difference methods in the wake from a cooled circular cylinder submerged in an upward freestream of mercury, air, or water. The main cause is clarified and is that one is the generation of the wake vorticity, which never occurs in any isothermal wake, and the other the stable arrangement with amplified asymmetry of the vorticity distribution in the wake. When the Prandtl number and / or the cylinder temperature are decreased, the Karman vortex street is generated easily and has a smaller wake frequency, smaller vortex speed, and larger vortex spiral than those in any isothermal wake. The characteristics of the Karm...

X-33 Aerodynamic Computations and Comparisons with Wind-Tunnel Data

Abstract: A computational  uid dynamics study conducted at the NASA Langley Research Center to support the phase II development of the X-33 vehicle is detailed. Aerodynamic computations for the X-33 vehicle were performed using two Ž nite volume, Navier–Stokes solvers and an inviscid Euler code. Computations were made for a range of wind-tunnel test conditions from Mach 4.63 to 10.0 with angles of attack from 10 to 48 deg and body  ap de ections of 0,+10, and+20 deg. Additionalcomputationswere performed over a parametric range of freestream conditions with Mach numbers of 4–10 and angles of attack of 10–50 deg. Computational results and comparisons with wind-tunnel aerodynamic data are presented.

Applications of Zero-Net-Mass Jets for Enhanced Rotorcraft Aerodynamic Performance

Abstract: Numerical studies were conducted to investigate the benee cial effects of using arrays of zero-net-mass (ZNM) “ synthetic” jets on the aerodynamic characteristics of the NACA-0012 airfoil. Flowe eld predictions were made using modie ed versions of the NASA Ames ARC2D, U.S. Army 2DBVI unsteady, two-dimensional, compressible thin-layer Navier ‐Stokes e ow solvers. An unsteady surface transpiration boundary condition was enforced over a user-specie ed portion of the airfoil’ s upper, or lower, surface to emulate the time variation of the mass e ux out from and into the airfoil’ s surface. Special emphasis is placed on two-dimensional model problems that are representativeofthemorecomplexthree-dimensionalhelicopterrotore owe eldenvironment.Thenumericalresults have indicated that ZNM jets can be used to enhance the lift characteristics of airfoils (helicopter rotor blades ) and alleviate the impulsive aerodynamic response of a helicopter blade during encounters with the tip vortex wake. The effectiveness of ZNM jets for aerodynamic control is shown to increase with the increase in freestream Mach number and, more importantly, with the decrease in the ratio between the peak jet Mach number to the freestream Mach number. The striking similarities with the aerodynamics of an airfoil having an array of surface protuberances are presented.

Experimental assessment of a water-film-thickness Weber number for scaling of glaze icing

Abstract: The evidence indicates that current practices for scaling of glaze icing tests do not recognize one or more important parameters. A Weber number based on a measure of the thickness of the liquid water film formed on the surface of the accreting ice has been proposed in another paper as the required additional scaling parameter. The present paper reports on experiments specifically designed to assess this proposal. Icing wind tunnel tests were done on 45 mm and 20 mm circular cylinders; the former constituted the reference cases and the latter, the sub-scale cases. In all sub-scale tests the accumulation parameter, the droplet inertia parameter and the calculated freezing fraction were made equal to the corresponding reference values. Freestream velocity for sub-scale test runs was chosen using several scaling parameters, including the newly proposed one. It was found that reasonably good similarity of ice accretion shapes was obtained for all of the sub-scale velocities that were tried, provided that the freestream static temperature was the same as that in the reference case. Possible explanations are suggested. When sub-scale freestream static temperature was the same as the reference value, the freestream velocities determined using Weber numbers based on water-film thickness and on droplet size were approximately equal, and this velocity gave marginally better similarity of ice shapes than velocities chosen on other bases. Solid aluminum and solid Plexiglas models gave essentially the same ice shapes for corresponding conditions. Most of the findings of the work are very preliminary and much more work is required to explore them. Nomenclature Ac accumulation parameter Ca capillary number, |awV/a d, D droplet diameter, cylinder diameter hc convective heat transfer coefficient K droplet inertia parameter L length scale of the overall air flow field (L = D in the present case) LWC liquid water content MVD mean volume diameter of water spray droplets n freezing fraction p., freestream static pressure qe/qc ratio of evaporative to convective heat transfer rates Re Reynolds number, pVD/|-i ^ water-film thickness T, freestream static temperature V freestream velocity Wed, Wer,7 We. Weber number based on droplet OLameter, body diameter and water-film thickness; e.g. Wed = pVM/a P local collection efficiency |i , j^ viscosity of air, viscosity of liquid water * Professor, Dept. of Mechanical and Aerospace Engineering, Assoc. Fellow AIAA. "Senior Research Officer, Aerodynamics Laboratory, Institute for Aerospace Research Copyright © 2000 by the American Institute of Aeronautics and Astronautics, Inc. All rights reserved 1 American Institute of Aeronautics and Astronautics (c)2001 American Institute of Aeronautics & Astronautics or Published with Permission of Author(s) and/or Author(s)' Sponsoring Organization.

Impulsively Started Flow About a Rigid Parachute Canopy

Abstract: The temporal evolution of the flowfield in the near wake of a parachute canopy is studied computationally with a finite element method. The canopy is assumed to be rigid and impermeable, and the flow is started impulsively. The separated shear layer surrounding the canopy creates a starting vortex ring. As time evolves, flow instabilities cause the vortex ring to become convoluted and eventually lead to the breakup of the ring. This phase of the flow lasts for approximately 16D/U, where D is the mean projected diameter of the canopy and U is the freestream velocity. After the initial phase, the flow goes through a transition phase before settling into its steady state. In the steady-state phase, the drag and base pressure coefficient become nearly constant. The computed drag coefficient matches very well against experimental data. The steady-state phase is reached after a time period of approximately 45D/U. During the steady-state phase, vortex shedding is observed in the near wake despite the nearly constant drag coefficient

Effect of Turbulence With Different Vortical Structures on Stagnation Region Heat Transfer

Abstract: The effect of freestream turbulence with different vortical structures on the stagnation region heat transfer was experimentally studied. Reynolds numbers, based on leading edge diameter of the heat transfer model with a cylindrical leading edge, ranged from 67.750 to 142,250. Turbulence generating grids of parallel rods were placed at several positions upstream of the heat transfer model in orientations where the rods were perpendicular and parallel to the stagnation line. The turbulence intensity and ratio of integral length scale to leading edge diameter were in the range 3.93 to 11.78 percent and 0.07 to 0.70, respectively. The grids with rods perpendicular to the stagnation line, where the primary vortical structures are expected to be perpendicular to the stagnation line, result in higher heat transfer than those with rods parallel to the stagnation line. The measured heat transfer data and turbulence characteristics are compared with existing correlation models

Response of the boundary layer developing over a blunt-nosed flat plate to free-stream non-uniformities

Abstract: The response of the boundary layer on a flat plate with blunt nose to infinitesimally small non-uniformity in the freestream velocity along the span has been studied. The non-uniformity was shown to excite boundary-layer disturbances similar to streaks or Klebanoff modes generally observed in experiments conducted with a high level of free-stream turbulence. The boundary layer disturbances have a predominantly streamwise velocity component and exhibit transient growth. In contrast to streaks generated by streamwise vortices impinging on the sharp nose of a plate, the disturbances produced by free-stream non-uniformity interaction with a blunt nose have a different level of growth. Their maximal amplification scales with the Reynolds number, based on the size of nose bluntness and is almost independent of the spanwise period of disturbances. This difference was shown to be caused by additional amplification of disturbances via vortex lines stretching around the leading edge.

Extent of the reverse flow wake region produced by a body in a uniform flow field

Abstract: A reverse flow is created in front of an object placed in a uniform air stream that originates from behind the object. Gaseous contaminants may then be transported into the breathing zone of a worker from sources located within the reverse flow region. This should be taken into consideration when local ventilation systems are designed. The objective of this study was to characterize the reverse flow zone created in front of a worker in a uniform flow of air,using both experimental data and numerical simulation. Experiments were carried out by moving a point contaminant source on a table placed in front of the worker, and by measuring the contaminant concentration at nose level in front of the worker. The experimentally estimated length of the reverse flow region was smallest (0.5-1.0 m) with a nominal freestream velocity of 0.1 m/s, and similar (1.1-1.4 m) with nominal freestream velocities of 0.3 and 0.5 m/s.