Showing papers on "Freestream published in 1998"

Turbulent shear-layer mixing at high Reynolds numbers: effects of inflow conditions

Abstract: We report on the results from a set of incompressible, shear-layer flow experiments, at high Reynolds number (Re_δ ≡ ρΔUδ_T(x)/μ ≃ 2×10^5), in which the inflow conditions of shear-layer formation were varied (δ_T is the temperature-rise thickness for chemically-reacting shear layers). Both inert and chemically-reacting flows were investigated, the latter employing the (H_2+NO)/F_2 chemical system in the kinetically-fast regime to measure molecular mixing. Inflow conditions were varied by perturbing each, or both, boundary layers on the splitter plate separating the two freestream flows, upstream of shear-layer formation. The results of the chemically-reacting ‘flip experiments’ reveal that seemingly small changes in inflow conditions can have a significant influence not only on the large-scale structure and shear-layer growth rate, as had been documented previously, but also on molecular mixing and chemical-product formation, far downstream of the inflow region.

Experiments on boundary-layer receptivity to freestream turbulence

Abstract: An understanding of the physical mechanisms by which freestream turbulence (FST) affects boundary-layer transition to turbulence is essential for transition prediction and control. Presented here is a brief summary of some experimental studies on the topic of boundarylayer receptivity. All cases mentioned concern layers on either flat-plates or axisymmetric bodies, and only cases of a relatively low turbulence level are included. Those restrictions are necessarily exclusive of the environments and configurations important for turbomachinery. The studies described reveal that a layer can respond to FST in several ways, exhibiting what are called K-mode oscillations, Tollmien-Schlichting (T-S) waves, and other motions not classified. Although the studies have focused upon finding and describing T-S wave growth as the principal cause of the hastening of transition, that link is not firmly established. Much work remains to be done before a description of the interaction based upon 'first principles' can be applied to transition prediction.

Problems in high speed flow prediction relevant to control

Abstract: Three flow problems are discussed whose solutions suggest flow control schemes. These are 1) unsteady hypersonic flow over bodies in the Newtonian approximation, 2) a mechanism of hypersonic flow stabilization over acoustically semi-transparent walls and 3) store separation from cavities. Simplified systematic approximations based on asymptotic frameworks lead to compact computational models that elucidate the flow structure and opportunities for control. Besides generalizing the steady model of Cole, the Newtonian approximation in the unsteady context shows that unsteady body perturbations can lead to inflectional velocity profiles that can produce instabilities and boundary layer transition to enhance mixing in combustors and inlets. The absorbing wall illustrates a mechanism that can be exploited to damp 2 mode hypersonic instabilities. Simplified flow modeling based on systematic asymptotics for store separation from cavities shows the influence of the cavity shear layer on apparent mass effects that are important to damping in pitch and clearance from the parent body. Comparisons with free drop experiments are used for initial validations of the analytical models. * Senior Scientist, Fellow, AIAA f Principal Researcher, Member, AIAA * Margaret Damn Distinguished Professor, Mathematical Sciences, Fellow, AIAA § Professor ** Professor, Associate Fellow Copyright© 1998, American Institute of Aeronautics and Astronautics, Inc. 1. Unsteady Newtonian thin shock layers and hypersonic flow stability 1.11ntroduction Although the stability of high speed flows has received much attention in the recent literature, major complicating aspects have not been treated in a unified way. These features include the combined effects of the finite shock displacement on the boundary layer, the nonparallelism of the flow and the vorticity introduced by the shock curvature. The relevant structure of the shock and boundary layers has been treated in [1][9]. In [6] and [7], the aforementioned stability issues were discussed within the Hypersonic Small Disturbance approximation for the inviscid deck strongly interacting with the hypersonic boundary layer. Equations of motion for the mean and fluctuating small amplitude flows were analyzed. Because of nonparallelism in this framework, the spatial part of the waves cannot be treated by the usual Fourier decomposition and an initial value rather than eigenproblem for spatial stability is obtained. The initial value problem leads to partial rather than ordinary differential equations that require a numerical marching method for their solution. Results indicate that the specific heat ratio 7 plays a major role in the stability of flow since it controls the reflection of waves from the shock and the radiation of energy in the shock layer whose thickness scales with 7 -1. Early experiments such as those described in [2] showed that for a practically interesting class of flows, the shock layer becomes very thin compared to the boundary layer near the nose of hypersonic flat plates. This feature and the desire to further understand the shock and boundary layer structure encourage the use of the Newtonian approximation 7 —> 1. The connection with flow stability motivates the study of this approximation in an unsteady context. In this chapter, limit process expansions will be discussed relevant to unsteady viscous interactions as a prelude to the analysis of hypersonic stability and transition. The application of these limits is an unsteady extension of the steady state analysis of [3]. Although the focus here is the treatment of viscous interaction, boundary layer stability, receptivity and transition, the results derived are useful in inviscid hypersonic unsteady aerodynamic methodology and load prediction as well. 1.2 Analysis Figure 1 schematically indicates strong interaction flow near the leading edge of a hypersonic body. The viscous boundary layer which is usually thin, occupies an appreciable fraction of the distance between the shock and body that will be considered without undue loss of generality a flat plate in what follows. Accordingly F(x,f} = Q, in the notation of Fig. 1. The results in this chapter will be expressed in terms of the boundary layer thickness function A(3c,r) = 0, which in the interpretation mentioned in the Introduction could be the body shape in an inviscid context. Copyright© 1998, American Institute of Aeronautics and Astronautics, Inc. The unsteady form of the Hypersonic Small Disturbance Theory (HSDT) equations [9] are applicable and are obtained as in [7] from limit process expansions of hatted variables defined as quantities normalized by their freestream counterparts, with p,T,u,v,fJL the density, temperature, horizontal, vertical components of the velocity vector, and viscosity respectively. If the freestream density, pressure and velocity are denoted as U,p^ and p^ respectively, then a pressure coefficient used in these expansions is defined as p = (P-PJ/P-U. Fig. 1 Schematic of hypersonic strong interaction flow. With these definitions and the coordinate system in Fig. 1 as well the normalization of the Cartesian dimensional coordinates x and y to the unit reference length L and the reference time scale L/U for the time t, unbarred dimensionless normalized counterparts of these independent variables are defined. If M^ and R^ are respectively the freestream Mach and Reynolds numbers, and 5 is a characteristic flow deflection angle, then the expansions are p=a(x,y,t;H,y)+--(1.1) T=T+— p = 8p+M = l+v =• §v+• • (1.2) (1.3) (1.4) (1.5) (1.6) where y = y/(L8}. These expansions are valid in the HSDT limit x, y, t, H = M o are fixed as 8 — > 0 ,

Experimental Investigation of Boundary Layer Behavior in a Simulated Low Pressure Turbine

Abstract: A detailed investigation of the flow physics occurring on the suction side of a simulated Low Pressure Turbine (LPT) blade was performed. A contoured upper wall was designed to simulate the y pressure distribution of an actual LPT blade onto a flat plate. The experiments were carried out at Reynolds numbers of 100,000 and 250,000 with three levels of freestream turbulence. The main emphasis in this paper is placed on flow field surveys performed at a y Reynolds number of 100,000 with levels of freestream turbulence ranging from 0.8% to 3%. Smoke-wire flow visualization data was used to confirm that the boundary layer was separated and formed a bubble. The transition process over the separated flow region is observed to be similar to a laminar free shear layer flow with the formation of a large coherent eddy structure. For each condition, the locations defining the separation bubble were determined by careful examination of pressure and mean velocity profile data. Transition onset location and length determined from intermittency profiles decrease as freestream turbulence levels increase. Additionally, the length and height of the laminar separation bubbles were observed to be inversely proportional to the levels of freestream turbulence.

Spectroscopic Emission Measurements Within the Blunt-Body Shock Layer in an Arcjet Flow

Abstract: In the present study, radiation emanating from the freestream and shock-layer e ow over a 15.24-cmdiam, e at-faced cylinder model was measured in the NASA Ames Research Center’ s 20-MW Arcjet Facility. The test gas was a mixture of argon and air. Spatially resolved emission spectra were obtained over a 200- to 890-nm wavelength range using a charged-coupled device camera (1024 3 256 array) attached to a spectrograph. The optical system was calibrated using tungsten and deuterium radiation sources. Analytical tools were used to determine the following line-of-sight-averaged thermodynamicproperties from the calibrated spectra: 1 ) rotational temperature of the freestream and 2 ) rotational, vibrational, electronic temperatures, and species number densities within the shock layer. An analysis was performed to estimate the uncertainty bounds of the determined properties.

High-Altitude Capsule Aerodynamics with Real Gas Effects

Abstract: Re-entry capsule aerodynamics within a wide range of angles of attack and flight altitudes are examined by the direct simulation Monte Carlo method. The local bridging method is verified by comparison with results of simulations. Capsule stability is analyzed for flight altitudes from 130 km down to 85 km. Comparison between computed and free flight data shows a good agreement. A qualitative change of heat transfer coefficient behavior for different angles of attack during the descent is revealed. The influence of chemical reactions on aerodynamics and flowfields at 85 km is shown to be significant. For a flow simulation in the near-continuum regime, a parallel version of the direct simulation code, with static and dynamic load balancing techniques, is used. An efficiency of about 80% is obtained for 256 processors using dynamic load balancing. Nomenclature CA = axial force normalized by p^U^S/2 CH - heat transfer normalized by p^U^S/ Cm - pitching moment normalized by p^ CN = normal force coefficient normalized by p^U^S/2 Ck = local coefficient Q cont = continuum coefficient ckjm = free molecular coefficient Fb = bridging function H = altitude, km Kn = Knudsen number KnotOC = Knudsen number based on /z (), TO, and p^ L = characteristic length, m NI = number of molecules in cell / Nm = average number of simulated molecules in the computational domain Wproc = number of processors S = characteristic size, m2 TO = stagnation temperature, K kaic = calculation time rcom = communication time Adic = synchronizatio n time fun = total operation time f/oo = freestream velocity, m/s a = angle of attack, deg £/ = volume of interaction region in cell /, m3 jjio = stagnation viscosity, kg/m-s v"1 = majorant frequency, s"1 Poo = freestream density, kg/m3

Detrimental Effects of Almost Immeasurably Small Freestream Nonuniformities Generated by Wind-Tunnel Screens

Abstract: Measurements are presented that were obtained during the course of a series of flow quality improvements to a small stand-alone wind tunnel that has been modified for boundary-layer transition studies. The objective of establishing a Blasius boundary layer with a high degree of spanwise uniformity has been frustrated by the persistence of Klebanoff modes, i.e., weak streamwise vortices within the layer. The vortices originate at the leading edge and appear to be caused by almost immeasurably small nonuniformities in the freestream introduced by the wind-tunnel screens. The vortices cause a spanwise thickening and thinning of the layer. Contours of the background unsteadiness in a spanwise plane through the layer show locally concentrated regions with elevated levels that are associated with the vortices. These contours are used as a sensitive indicator for the flow quality improvements. Although far from forming a complete parametric study, the observations should act as a valuable guide for others. For example, spanwise variations in the porosity of the screens were discovered by traversing each screen between a laser and a photo detector. Significant improvement in the spanwise uniformity of the layer was obtained by sorting the screens based on these results

Computational/experimental aeroheating predictions for X-33 Phase II vehicle

Abstract: Laminar and turbulent heating-rate calculations from an "engineering" code and laminar calculations from a "benchmark" Navier-Stokes code are compared with experimental wind-tunnel data obtained on several candidate configurations for the X-33 Phase II flight vehicle. The experimental data were obtained at a Mach number of 6 and a freestream Reynolds number ranging from 1 to 8 x 10/ft. Comparisons are presented along the windward symmetry plane and in a circumferential direction around the body at several axial stations at angles of attack from 20 to 40 deg. The experimental results include both laminar and turbulent flow. For the highest angle of attack some of the measured heating data exhibited a "non-laminar" behavior which caused the heating to increase above the laminar level long before "classical" transition to turbulent flow was observed. This trend was not observed at the lower angles of attack. When the flow was laminar, both codes predicted the heating along the windward symmetry plane reasonably well but under-predicted the heating in the chine region. When the flow was turbulent the LATCH code accurately predicted the measured heating rates. Both codes were used to calculate heating rates over the X-33 vehicle at the peak heating point on the design trajectory and they were found to be in very good agreement over most of the vehicle windward surface.

Effect of high freestream turbulence with large length scale on blade heat/mass transfer

Abstract: The naphthalene sublimation technique is used to investigate the influence of high free-stream turbutence with large length scale on the heat/mass transfer from a turbine blade in a highly accelerated linear cascade. The experiments are conducted at four exit Reynolds numbers, ranging from 2.4 x 10 5 to 7.8 x 10 5 , with free-stream turbulence of 3, 8.5, and 18 percent and corresponding integral length scales of 0.9 cm, 2.6 cm, and 8 cm, respectively. On the suction surface, the heat/mass transfer rate is significantly enhanced by high free-stream turbulence due to an early boundary layer transition. By contrast, the transition occurs very late, and may not occur at very low Reynolds numbers with low free-stream turbulence. In the turbulent boundary layer, lower heat/mass transfer rates are found for the highest free-stream turbulence level with large length scale than for the moderate turbulence levels with relatively small scales. Similar phenomena also occur at the leading edge. However, the effect of turbulence is not as pronounced in the laminar boundary layer.

Flow Characterization in the ONERA F4 High-Enthalpy Wind Tunnel

Abstract: Experimental results obtained in the ONERA F4 hot-shot wind tunnel have been analyzed to address two important problems for this type of wind tunnel, ie, the determination of reservoir conditions and the thermochemical nature of the nozzle flow Numerical tools (one-dimensional and two-dimensional inviscid or viscous, unsteady or space marching codes) are used to reproduce the results of the high-enthalpy experiments To illustrate this approach, several runs are investigated with different arc-chamber material options, for which diode laser infrared absorption spectrometry (DLAS) has provided freestream velocity, translational temperature, and nitric oxide concentration measurements The reservoir enthalpy can be determined using spherical and sharp-cone probe heat-transfer-rate measurements with adequate correlations Recent direct measurements of freestream velocity with electron beam fluorescence time-of-flight technique are used to cross check DLAS and heat-transfer-rate probe results Concerning the thermochemical nature of F4 nozzle flows, an unexpected conclusion is obtained, as the nozzle wall pressure and translational temperature are observed to be close to equilibrium values at high-enthalpy operating conditions

Numerical Simulation of Combustion and Extinction of a Solid Cylinder in Low-Speed Cross Flow

Abstract: The combustion and extinction behavior of a diffusion flame around a solid fuel cylinder (PMMA) in low-speed forced flow in zero gravity was studied numerically using a quasi-steady gas phase model. This model includes two-dimensional continuity, full Navier Stokes' momentum, energy, and species equations with a one-step overall chemical reaction and second-order finite-rate Arrhenius kinetics. Surface radiation and Arrhenius pyrolysis kinetics are included on the solid fuel surface description and a parameter Phi, representing the percentage of gas-phase conductive heat flux going into the solid, is introduced into the interfacial energy balance boundary condition to complete the description for the quasi-steady gas-phase system. The model was solved numerically using a body-fitted coordinate transformation and the SIMPLE algorithm. The effects of varying freestream velocity and Phi were studied. These parameters have a significant effect on the flame structure and extinction limits. Two flame modes were identified: envelope flame and wake flame. Two kinds of flammability limits were found: quenching at low-flow speeds due to radiative loss and blow-off at high flow speeds due to insufficient gas residence time. A flammability map was constructed showing the existence of maximum Phi above which the solid is not flammable at any freestream velocity.

Diffusion Flame Stabilized on a Porous Plate in a Parallel Airstream

Abstract: Effects of an obstacle on the structure and stability of a laminar diffusion e ame established on a porous plate in a parallel airstream have been investigated experimentally. The obstacle, a backward-facing step or a rectangular cylinder, is located upstream of the porous plate through which gaseous methane is injected uniformly. Structures of the e ame are elucidated by the direct and schlieren photography. Flame shapes are described and stability diagrams are plotted for the freestream velocity and the fuel injection velocity, which are discussed with e ow structures.

Velocimetry and thermometry of supersonic flow around a cylindrical body

Abstract: The supersonic flow around a cylindrical body has been studied using two optical techniques. For both sets of measurements, the cylinder was mounted from the side of the tunnel, allowing investigation of the bow shock region as well as in the wake. A new technique, laser-enhanced ionization flow tagging, was used for streamwise velocity determinations behind the body. From these measurements, it was found that the downstream velocity outside the wake was (1.90 +/- 0.06) km/s, whereas inside the wake the velocity was about 0-500 m/s in the upstream direction. Planar laser induced fluorescence of nitric oxide was employed for temperature determinations. It was established that the freestream temperature was (2120 +/- 100) K, decreasing to around (1550 +/- 400) K in the wake.

Aspects of turbulent-shear-layer dynamics and mixing

Abstract: Experiments have been conducted in the GALCIT Supersonic Shear Layer Facility to investigate some aspects of high-Reynolds-number, turbulent, shearlayer flows in both incompressible- and compressible-flow regimes Experiments designed to address several issues were performed; effects of inflow boundary conditions, freestream conditions (supersonic/subsonic flow), and compressibility, on both large-scale dynamics and small-scale mixing, are described Chemically-reacting and non-reacting flows were investigated, the former relying on the (H2 + NO/F2) chemical system, in the fast-kinetic regime, to infer the structure and amount of molecular-scale mixing through use of "flip" experiments A variety of experimental techniques, including a color-schlieren visualization system developed as part of this work, were used to study the flows Both inflow conditions and compressibility are found to have significant effects on the flow In particular, inflow conditions are "remembered" for long distances downstream, a sensitivity similar to that observed in low-dimensionality, non-linear (chaotic) systems The global flowfields (freestreams coupled by the shear layer) of transonic flows exhibit a sensitivity to imposed boundary conditions, i e, local area ratios A previously-proposed mode-selection rule for turbulent-structure convection speeds, based on the presence of a lab-frame subsonic freestream, was experimentally demonstrated to be incorrect Compressibility, when decoupled from all other parameters, eg, Reynolds number, velocity and density ratios, etc, reduces laxge-scale entrainment and turbulent growth, but slightly enhances smallscale mixing, with an associated change in the structure of the molecularly-mixed fluid This reduction in shear-layer growth rate is examined and a new parameter that interprets compressibility as an energy-exchange mechanism is proposed The parameter reconciles and collapses experimentally-observed growth rates

PLIF thermometry in shock tunnel flows using a Raman-shifted tunable excimer laser

Abstract: Planar laser-induced fluorescence is performed in a free-piston shock tunnel by using a Raman-shifted tunable excimer laser to excite nitric oxide molecules in the flow. Two different flowfields are examined to test the difficulties associated with applying the technique to shock tunnels: the bluff body flow produced by a 25 mm diameter cylinder; and the oblique shock and expansion fan produced by a 35° half-angle wedge. For the cylinder, the maximum flow enthalpy was limited to 4.1 MJ kg $^{-1}$ due to high flow luminosity which is produced by metallic contaminants in the flow. A reflective filter is used to reduce the influence of flow luminosity making these measurements feasible. Freestream temperature measurements are in excellent agreement with those predicted from numerical flow calculations. Large uncertainties were observed for the high-temperature post-shock results. Several higher enthalpy shots (14 MJ kg $^{-1}$ ) were also performed with the wedge and showed an insignificant amount of contaminant emission.

Dns of leading - edge receptivity to sound

Abstract: Numerical simulations of leading-edge acoustic receptivity are performed for a flat plate with a modified-super-elliptic (MSB) leading edge. For small freestream amplitude, the agreement between Branch I receptivity coefficients predicted from the DNS and the experiments of Saric & White (1998) for acoustic waves at zero incidence is excellent. The effect of angle of incidence of the impinging wave is investigated and found to produce higher receptivity coefficients than in the symmetric case. The slope of leading-edge receptivity coefficient versus angle of incidence of the impinging wave is found to be less than 1/4 of the slope predicted by zero-thicknes s flatplate theory. However, there is excellent agreement between the DNS and finite-nose-radius theory of Hammerton & Kerschen (1996). These results clearly demonstrate the importance of including the effects of the finite nose radius in any receptivity study. Finally, downstream of the leading-edge region, linear stability theory is found to accurately reproduce the characteristics of the instability waves. At higher freestream forcing, an instability wave generated by nonlinear interaction is found at double the frequency of the forcing.

Effects of High Freestream Turbulence Levels and Length Scales on Stator Vane Heat Transfer

Abstract: Gaining a good understanding of how high freestream turbulence augments heat transfer is important for predicting thermal loadings for turbine blades and vanes. This study was aimed at documenting the surface heat transfer and the highly turbulent flowfield around a stator vane. The effects of turbulence levels between 11% and 24% were studied. At the highest turbulence level, two different Reynolds numbers (Reex = 6 × 105 and 1.2 × 106) and two different length scales were also studied. Three-component laser Doppler velocimeter measurements of the velocity fluctuations indicated that downstream of the active grid there was an initial decay of the turbulent kinetic energy which then leveled off at about one leading edge radius upstream of the vane. Inside the vane passage the turbulent kinetic energy increased slightly and then decayed through the passage. The surface heat transfer showed the largest augmentations on the pressure side of the vane with higher augmentations at higher turbulence levels, smaller length scales, and higher Reynolds numbers.Copyright © 1998 by ASME

Heating enhancement caused by a transverse control jet in hypersonic flow

Abstract: Experiments are reported in which the heat flux distribution near a single circular, sonic transverse jet on a flat plate exposed to a hypersonic (Mach 6.7) freestream flow was quantitatively measured using thermochromic liquid crystals. The freestream conditions were such that the boundary layer growth on the plate ahead of the jet was laminar. The results indicate that the interaction of the jet with the freestream flow created a complex flowfield with regions of separation and reattachment which caused localised enhancements to the heat flux upstream and to the side of the jet, the magnitudes of which were sensitive to both jet plenum pressure and jet gas composition.

Heat Flux on Partially Catalytic Surfaces in Hypersonic Flows

Abstract: The combined effects of the chemical nonequilibrium and the surface catalysis on the stagnation-point heat flux of a blunt body are investigated by Navier-Stokes calculations on a spherical model under Scirocco Plasma Wind Tunnel conditions. The freestream properties in the test section are found by means of a preliminary nonequilibrium, full Navier-Stokes computation of the nozzle flow expansion. All available numerical heat flux results are correlated as a function of a dimensionless parameter, including the simultaneous effects of the finite-gas and the surface-phase reactions. The proposed parameter, which takes into account simultaneously the nonequilibrium effects in the bulk and surface phases, correlates well with the numerical results over a blunt-body stagnation region. The final correlation formula for the stagnation heat flux can be used to correlate experimental results on different materials with known catalytic coefficients or, conversely, to determine the surface catalytic efficiency of different thermal protection materials. Nomenclature a = speed of sound, m/s cp =

Hypersonic interference heating induced by a blunt fin

Abstract: Interference heating effects generated by fin-type protuberances on hypersonic vehicles have been in- vestigated. Experiments were carried out at a freestream Mach number of 6.7 under laminar flow conditions with a model comprising a flat plate with a single, blunt fin. The effects of fin leading edge di- ameter, sweep and angle of incidence on the heating distribution on the flat plate surface were explored. The surface heating was measured using liquid crys- tal thermography which provides quantitative data with high spatial resolution. Complementary surface oil flow and schlieren experiments were also carried out to gain a better understanding of the interfer- ence flowfield. The results reveal a highly complex interaction region which extends ~ 7 diameters up- stream of the fin and with heating enhancements up to (and, in some cases, probably exceeding) 7-times the undisturbed value.

Emission Spectral Measurements in the Plenum of an Arc Jet Wind Tunnel

Abstract: Arc jet wind tunnel facilities are used to evaluate thermal protection system materials for re-entry vehicles The high speed, high temperature flowfield generated by the arc jet can simulate the extreme aerodynamic heating environment experienced during re-entry so that the survivability of heat shield materials and performance of various designs options can be tested Although the re-entry heating environment can be approximated in the arc jet facility, the flowfield only partially simulates the actual re-entry flight conditions Reynolds numbers are not matched so that surface shear stress distributions and mass transfer rates due to ablation or other mechanisms are not modeled correctly Unlike flight conditions the arc freestream air is in non-equilibrium because of the rapid expansion that occurs in the supersonic nozzle To properly study the actual re-entry flow environment, computational fluid dynamics, computational chemistry and radiation models must be used Arc jet tunnel tests serve to validate these models To perform accurate simulations inlet and boundary-conditions are needed, which come from measurements of the flowfield The present study is concerned with measurements in the plenum region of an arc heater In the past, conditions in the arc heater flowfield have been predicted using simulations since conventional measurement techniques could not be used in the harsh extremely high temperature environment The present study is part of a recent push to utilize optical techniques to help better characterize the arc jet flowfields Emission measurements have been made in the shock layer and the constrictor section of the arc heater to determine temperatures and species number densities LIF measurements have been made in the free stream to determine temperature and velocity

Wake Decay: Effect of Freestream Swirl

Abstract: A study of the effects of freestream swirl on the decay characteristics of wakes shed from a rotating blade row is presented. The freestream swirl behind the rotor causes the wakes to skew tangentially, stretching the wakes as they are convected from the rotor to the stator. The effect of stretching on wake decay is illustrated using a simplified two-dimensional model. The model is described and the results are compared to 1) measurements from a two-dimensional cascade facility where no stretching or skewing of the wakes occurs, 2) solutions obtained using a three-dimensional, Reynolds-averaged Navier-Stokes solver, and 3) experimental wake measurements taken behind a low hub-to-tip ratio fan.For typical fan geometries with hub-to-tip ratios of approximately 0.5 and rotor-stator spacings of one to two rotor chord lengths, the wake can be stretched by over 50 percent. The stretching increases the mixing rate which leads to a reduction in the mean wake velocity deficit of approximately thirty percent and a widening of the wake of about fifteen percent. These effects account for much of the difference seen between cascade wake measurements and those taken behind rotating fan blade rows. It is therefore important to include such effects when using cascade data for prediction of fluid mechanic, acoustic, or structural phenomena associated with fan wakes. Finally, the study also suggests a potential for small (< 3 dB) reductions in wake-stator interaction noise through tailoring the fan loading distribution to produce particular span wise wake decay characteristics.Copyright © 1997 by ASME

Velocity measurement in hypersonic flows using electron-beam-assisted glow discharge

Abstract: A new pseudospark-type electron gun has been utilized in the F4 high-enthalpy, low-density, hot-shot wind tunnel at ONERA. The main goal was to perform accurate measurements of the velocity profile across the boundary layer and the freestream through a time-of-flight principle. In these experiments, an intense pulsed electron beam traces the path of a high-voltage, sustained-glow discharge in some 10 ns. A charge-coupled device camera is opened briefly, 5 μs after the electron gun actuation, to image the position of the luminous column convected by the flow

Measurement and Analysis of the Flowfields Induced by Suction Perforations

Abstract: The flow physics of laminar flow control suction surfaces is revealed by performing a detailed fundamental experimental investigation of an isolated suction perforation. A unique series of nonintrusive, high-resolution measurements are obtained using a three-component laser Doppler velocimetry system, and experiments are conducted in a low-speed, low-turbulence wind tunnel. The suction perforation flowfields are mapped for a range of sub- and supercritical suction rates and are found to be highly three dimensional. A rich variety of flowfield features is observed, including a pair of counter-rotating longitudinal vortices, multiple corotating longitudinal vortices, spanwise variations of the mean flow, and inherently unstable boundary-layer profiles. Critical suction limits, over a range of freestream speeds, are determined, and a new design criterion for critical suction is established. It is also shown that for sufficiently small perforations, irrespective of suction flow, boundary-layer transition does not occur. Further analyses of the measurements explore the possibility of interaction between the crossflow vortices and the suction-induced longitudinal vortices